GB2530553A

GB2530553A - Fan blade and method of manufacture

Info

Publication number: GB2530553A
Application number: GB1416984.1A
Authority: GB
Inventors: Simon Mark Barlow
Original assignee: Rolls Royce PLC
Current assignee: Rolls Royce PLC
Priority date: 2014-09-26
Filing date: 2014-09-26
Publication date: 2016-03-30
Also published as: GB201416984D0

Abstract

An aerofoil blade 400 for a turbo machine has a root portion 408 shaped for attachment to the turbo machine, where the blade is formed to define an internal cavity 316 in at least the root portion of the blade. The cavity can extend from the root portion into the blade portion. The blade may include one or more membranes 404 arranged to subdivide the internal cavity in at least the root portion. The blade is manufactured by providing a stacked assembly including a pair of opposing panels each having a root member projecting outward. The assembly is hot formed to form an internal cavity at least in the root portion. The hot forming may use a cavity die and the assembly can be inflated to form the internal cavity. The hot forming may include diffusion bonding and super plastic forming. The assembly may include a membrane sandwiched between the opposing panels an may include a stop off layer for preventing boding in predetermined regions. The fan blade is lighter than prior art fan blades.

Description

FAN BLADE AND METHOD OF MANUFACTURE

Field of the Invention

The present invention relates to a fan (aerofoil) blade for a turbo machine, and its manufacture; in particular but not exclusively to a fan (aerofoil) blade for a gas turbine engine, and its manufacture.

Backaround of the Invention

Diffusion bonded and superplastically formed line core hollow fan blades are known. A typical design presently in service includes a wide chord fan blade (WCFB), a method for the manufacture of which is represented in Figs. 2A-C for example.

Such fan blades are typically formed from a fan blade assembly comprised of two opposing panels 202, sandwiching a membrane 204. An internal core pattern 206 is typically printed on to an internal surface of one or both of the opposing panels 202, as shown in Fig. 2A for example. The sandwich assembly is then typically diffusion bonded into a flat pack and placed into a cavity die 208 as shown in Fig. 2B. Typically, the core pattern 206 printed on to the internal surface of the or each respective panel 202 is formed of a stop-off material. A stop-off material located between the panels 202 prevents bonding of the panels in the region of the stop-off material during the diffusion bonding process. Thus, during the diffusion bonding, only the region(s) of the interface(s) of the assembly which were not printed with the pattern 206 bond.

Subsequently, by hot forming processes, the flat pack is typically inflated by superplastic forming to conform to the cavity die 208 thereby forming an internal cavity 210, typically referred to as the internal line core, as shown in Fig. 20. The extent to which the internal cavity extends within the fan blade is governed by the printed pattern.

In the prior art, the panels are typically each provided with a root portion 203, projecting laterally outward from the assembly. The root portions 203 are diffusion bonded together to form a solid root 203' to the fan blade. The panels are also provided with a blade portion 205 extending from the root portion 203 to the tip of the panel.

The cavity die 208 is provided with shoulders 209 to prevent at least any movement or deformation of the root portions 203 forming the solid root 203' during at least the superplastic forming step, which is performed at temperatures and pressures which could otherwise undesirably deform the solid root 203'. Thus, during the superplastic forming step, only the blade portions 205 of panels 202 are caused to inflate in order to conform to the cavity die 208. In other words, only the central region of the interface between the panels, e.g. generally between the solid root and a tip region of the assembly is caused to inflate to form the internal cavity 210. Thus, the internal cavity 210 is present only between the blade portions 205.

After the superplastic forming process the solid root 203' is typically machined and shaped for complementary engagement with e.g. a slot formed in the hub of a fan. In use, the solid root 203' retains the fan blade in place in the fan and prevents undesirable movement (e.g. radial movement) of the fan blade during use.

By suitable patterning of the stop-off material, during the superplastic forming step the membrane 204 can be shaped as a corrugated, or concertina-type, membrane alternatingly bonded to the respective panels 202 in the direction from leading edge to trailing edge of the fan blade 200. Thus, in some fan blades, the membrane 204 is formed to subdivide the cavity 210 into a series of sub-cavities each formed between the blade portions 205 to extend in the general direction between the root and tip of the blade.

Whilst fan blades produced in accordance with the above method have proved successful, the present inventors have realised that the prior art fan blades can be improved yet further.

Summary of the Invention

Therefore, in an aspect, the present invention proposes a fan blade for a turbo machine as set forth in claim 1.

Accordingly, the present invention aims to provide an aerofoil blade with suitable rigidity and integrity for use in a turbomachine, e.g. a gas turbine engine, but which is lighter than the

prior art aerofoil fan blades.

In particular, but not exclusively, the present invention aims to provide a turbomachine fan blade which has suitable rigidity and integrity, but which is lighter than the prior art turbomachine fan blades.

The blade may preferably has a leading edge, a trailing edge and a tip region distal to the root portion mutually arranged to define a blade portion of the blade, wherein the internal cavity extends from the root portion into the blade portion.

The root portion may include one or more root members, projecting laterally outward from the blade relative to the leading and trailing edges.

The root portion may include a pair of root members projecting laterally outward from the blade in mutually opposite directions; and wherein the blade may be formed to define the internal cavity in at least a region of the root portion between the pair of root members.

The or each root member may be configured for interlocking engagement with a complementary element of the turbo machine for maintaining (retaining) attachment of the blade to the turbo machine during use.

The blade may include one or more membranes arranged to sub-divide the internal cavity into a plurality of internal sub-cavities in at least the root portion of the blade.

At least one of the internal sub-cavities may extend from the blade portion into the root portion.

The aerofoil blade is preferably a fan blade for a gas turbine engine.

In an aspect the present invention provides a turbomachine including an aerofoil blade as described herein, for example a gas turbine engine including an aerofoil blade as described herein.

In an aspect, the present invention provides a method of manufacture of an aerofoil blade for a turbo machine, the method including the steps of providing a stacked assembly, including a pair of opposing panels, each panel including a respective root member projecting outward from the assembly and mutually cooperating to define a root portion of the assembly; performing a hot forming process to form an internal cavity within the assembly at least in the root portion.

The method may further include the steps of: arranging the stacked assembly in a cavity die; and the hot forming process including the step of inflating the assembly to form the internal cavity by conforming the assembly to an internal region of the cavity die.

The step of conforming the assembly to the internal region of the cavity die may include deforming the root portions of the assembly to conform to a root defining sub-region of the internal region of the cavity die.

The method may further include the steps of performing a bonding process to bond only predetermined regions of the interface(s) within the assembly; performing the hot forming process to form the internal cavity within the assembly in one or more unbonded regions of the interface(s) within the assembly.

The assembly may include a membrane sandwiched between the opposing panels.

The bonding process may include the step of providing at least one interface with a stop-off layer, for preventing bonding of the interface(s) in the region of the layer, to define said predetermined regions.

At least a portion of the at least one interface provided with the stop-off layer may be located between the root portions.

The bonding process may include the step of diffusion bonding the assembly to bond the predetermined regions.

The hot forming process may include the steps of applying suitable temperature and pressure to the assembly to inflate the assembly to conform to the internal region of the cavity die.

The hot forming process includes the step of superplastically forming the assembly to thereby define the internal cavity.

In an aspect, the present invention may provide a method of manufacture and/or assembly of a gas turbine engine including the step of incorporating into the engine an aerofoil (fan) blade as described herein.

The present invention may provide a method of manufacture and/or assembly of a turbomachine, e.g. a gas turbine engine, including the step of incorporating into the turbomachine an aerofoil (fan) blade according to the present invention.

Brief Description of the Drawings

Embodiments of the invention will now be described by way of example with reference to the accompanying drawings in which: Figure 1 shows a cross section of a gas turbine engine incorporating a fan blade according to the present invention; Figures 2A-2C exemplify a prior art process by which prior art fan blades are formed; Figure 3A-3C exemplify an example of a process by which aerofoil (fan) blades according to the present invention can be formed; and Figure 4 shows a fan blade according to an aspect of the present invention, including a cross-section of the fan blade to show the inclusion of an optional corrugated membrane located in the internal cavity of the fan blade.

Detailed Description and Further Optional Features of the Invention With reference to Figure 1, a ducted fan gas turbine engine incorporating the invention is generally indicated at 10 and has a principal and rotational axis X-X. The engine comprises, in axial flow series, an air intake 11, a propulsive fan 12, an intermediate pressure compressor 13, a high-pressure compressor 14, combustion equipment 15, a high-pressure turbine 16, an intermediate pressure turbine 17, a low-pressure turbine 18 and a core engine exhaust nozzle 19. A nacelle 21 generally surrounds the engine 10 and defines the intake 11, a bypass dud 22 and a bypass exhaust nozzle 23.

During operation, air entering the intake 11 is accelerated by the fan 12 to produce two air flows: a first air flow A into the intermediate pressure compressor 13 and a second air flow B which passes through the bypass duct 22 to provide propulsive thrust. The intermediate pressure compressor 13 compresses the air flow A directed into it before delivering that air to the high pressure compressor 14 where further compression takes place.

The compressed air exhausted from the high-pressure compressor 14 is directed into the combustion equipment 15 where it is mixed with fuel and the mixture combusted. The resultant hot combustion products then expand through, and thereby drive the high, intermediate and low-pressure turbines 16, 17, 18 before being exhausted through the nozzle 19 to provide additional propulsive thrust. The high, intermediate and low-pressure turbines respectively drive the high and intermediate pressure compressors 14, 13 and the fan 12 by suitable interconnecting shafts.

The present invention provides an advantageously lightweight fan blade for such a gas turbine engine by providing a fan blade having an internal cavity which preferably extends from the blade legion and into the root of the fan blade; the root being the portion of the fan blade which is used to attach the fan blade to the gas turbine engine. For example, the root provides the means by which the fan is attached to the fan hub, which may be rotatable by the aforementioned interconnecting shaft.

In a fan blade according to the present invention the external dimensions of the fan blade (the blade portion and root portion) are preferably identical to the external dimensions of equivalent prior art fan blades. However, the internal cavity of a fan blade according to the present invention is larger than the internal cavity of equivalent prior art fan blades, by virtue of extending into the root portion of the fan blade. Therefore, advantageously, the overall mass of the fan blade is reduced relative to prior art fan blades. Thus, a lightweight fan blade according to the present also contributes to a reduction in fuel usage in a gas turbine engine, for example. Furthermore, the amount of material incorporated in a fan blade according to the present invention is less than the prior art fan blades, thereby reducing material usage and the associated cost.

Figs. 3A-C shows an example of a process by which a fan blade according to an embodiment of the present invention is formed.

A fan blade assembly 300 is provided comprising a pair of opposing panels 302. The panels are typically metallic, preferably an alloy of titanium, for example 6/4 titanium.

Each panel 302 provides a root portion 304 projecting outward from the panel 302. The root portions 304 cooperate in the subsequently processed (e.g. bonded and hot formed) assembly 300 to define the aforementioned root of the final fan blade. Preferably, the assembly is arranged so that the root portions 304 of the respective panels project laterally outward from the assembly in mutually opposing directions to cooperatively form a root of the assembly (and subsequently, ultimately of the fan blade).

Each panel 302 also provides a blade portion 306 (or aerofoil portion 306) extending from the root portion 304 to the tip 308 of the panel 302. In essence, the blade portion 306 contributes to defining the aerofoil region of the final fan blade.

Typically, a portion of at least one, but preferably both, of the respective opposing faces of the panels 302 is treated, e.g. patterned, with a stop-off material 310. The stop-off material may be yttrium oxide for example. The skilled person knows of other suitable materials for acting as a stop-off material to achieve the desirable results described herein.

Optionally, a membrane 312 may be provided between the opposing panels 302, to be sandwiched therebetween. The membrane is typically metallic, for example an alloy of titanium. Preferably, the membrane is formed of the same material as the respective panels 302. The membrane may also be printed with stop-off material if desired.

A bonding process is used to bond the assembly 300 into a single integral body. A diffusion bonding process, as known in the art, is preferred. The skilled person knows of diffusion bonding processes to achieve the desirable results described herein. During the bonding process the interface regions patterned with the stop-off material do not form a bond, whereas the other interface regions do bond. By the bonding process, a single body blade is formed from the assembly 300.

The bonded assembly 300 is arranged in a cavity die mould 314. The internal dimensions of the cavity die mould are such that gaps 315 are provided between regions of the internal surfaces of the cavity die mould and the assembly 300.

A hot forming process is performed during which the assembly is re-shaped to conform to the cavity die mould. Typically, this is achieved by inflating (expanding) the assembly such that regions of the respective panels 302 are caused to move apart and conform to the cavity die mould 314 thereby removing the gaps 315. This is typically achieved by heating the assembly to a temperature sufficient to allow superplastic forming of the assembly 300, and by providing a pressure differential between the internal region of the assembly 300 and the gaps between the assembly 300 and the cavity die mould 314.

In conforming to the internal dimensions of the cavity die mould 314, the bonded and hot formed assembly 300 (hereafter referred to as the fan blade) defines an internal cavity 316 between the panels 302 of the fan blade. The extent to which the internal cavity extends within the fan blade, i.e. between the panels 302, is governed by the pattern of the stop-off material with which the or each panel 302 is treated. This is because the stop-off material pattern defines the regions of the interface(s) which do not bond, and thus which are capable of being moved apart by the hot forming process to define one or more internal cavities.

According to an aspect of the present invention, the stop-off material is provided in a region of the assembly 300 between the root portions 304. This can be seen in Fig. 3A. Therefore, at least a region of the internal cavity is formed between the root portions 304. However, in the prior art, stop-off material 206 is not provided in this region, as can be seen from Fig. 2A.

Therefore, the internal cavity 210 in the prior art is not formed between the root portions 203.

Furthermore, according to an aspect of the present invention, the root portions are subject to the hot forming process, and are thus preferably superplastically formed (deformed), to conform to the cavity die mould, i.e. to the internal dimensions of the cavity die mould.

For example, the cavity die mould 314 may provide gaps 315 around the assembly 302 arranged to allow the root portions 304 to be deformed or re-shaped to conform to the internal dimensions of the cavity die mould 314 during the hot forming process.

So, according to aspects of the present invention, the root portions may be re-shaped, e.g. by being moved apart, during the hot forming process so as to define (provide) at least a region of an internal cavity therebetween. The region of the internal cavity formed by the hot forming process is thus defined (provided) within the root of the final fan blade.

However, in the prior art, no gaps are provided between the root portions 203 and the cavity die mould 208 to allow the root portions to deform, and certainly not to move apart.

Therefore, in the prior art, the root portions 203 are not subject to a hot forming process whereby they are deformed to conform to the internal dimensions of the cavity die mould 208. Instead, in the prior art, the cavity die mould is typically shaped to prevent deformation of the root portions 203 during the hot forming process.

Additional processing time (relative to the prior art process in which a cavity is not formed in the root portion) may be required in order to suitably conform the root portion of the fan blade to the cavity die mould by superplastic forming.

As mentioned previously, the present invention advantageously provides a fan blade (aerofoil blade) which is lighter than prior art fan blades, because the internal cavity in a fan blade (aerofoil blade) according to the present invention is relatively larger than the internal

cavity in prior art fan blades.

Where the optional membrane 312 is provided in the assembly 300, the stop-off material may be patterned such that the bonding process and the hot forming process results in corrugation of the membrane into a concertina-type (or corrugated) membrane 404, alternatingly bonded across its width with the respective panels 406 of the fan blade (aerofoil blade) 400, as shown in Fig. 4. View X-X shows a cross-section of the fan blade 400 across the indicated region.

The corrugated membrane 404 thus sub-divides the internal cavity 316 into respective sub-cavities 402. Therefore, at least a region of one or more of the sub-cavities 402 is provided between root portions 304 in the fan blade 400, as described above.

In other words, the root 408 of the fan blade 400 includes at least a portion of the internal cavity defined within the fan blade 400. The root 408 being the portion of the fan blade at least partially defined by the root portions 304, and being the portion of the fan blade which is used to attach the fan blade 400 to a machine in which the fan blade is to be used. For example, the machine may be a turbomachine, such as a gas turbine engine. The root 408 may be used to attach the fan blade 400 to the rotating hub (shaft) of the fan of the turbomachine for example. The hub may include a disk. The disk may include a slot with which the blade is engaged. A plurality of blades may be engaged with respective slots of the disk. The slots, and therefore the blades, may be arranged circumferentially around the disk, although the slots may extend generally parallel to the rotational axis of the disk. When the blade is engaged with the slot, the root portion may be under compressive stress. The slot may be a general T-shape when viewed along the long axis of the slot. Each of the root members may be received by a respective lateral arm of the cross bar of the general T-shape slot to prevent the blade from moving radially relative to the disk during use.

In a fan blade according to the present invention the external dimensions of the fan blade (the blade portion and root portion) are preferably identical to the external dimensions of equivalent prior art fan blades. Therefore, fan blades according to the present invention can replace the existing prior art fan blades, i.e. they can be retrofitted, without any need for modification of the machines in which they are to be installed.

In essence, aspects of the present invention provide for extension of the hollow volume of an aerofoil blade into an area of the blade in which it is predicted that there will be compressive stress under normal running conditions, thereby decreasing the aerofoil blade weight without any loss of integrity of the blade or increasing risk of failure of the blade.

Aerofoil blades according to the present invention include turbomachine (e.g. gas turbine engine) fan blades, outlet guide vanes e.g. in a gas turbine engine, and aerofoil blades incorporated into a blisk, e.g. a hollow blisk.

Reference numbers used in the description:

Ducted fan gas turbine engine 10, air intake 11, propulsive fan 12, intermediate compressor 13, a high-pressure compressor 14, combustion equipment 15, a high-pressure turbine 16, an intermediate pressure turbine 17, a low-pressure turbine 18, a core engine exhaust nozzle 19, a nacelle 21, a bypass duct 22, a bypass exhaust nozzle 23; fan blade assembly 200, panel 202, root portion 203, solid root 203', membrane 204, blade portion 205, internal core pattern 206, cavity die 208, cavity die shoulder 209, internal cavity 210; fan blade assembly 300, panel 302, root portion 304, blade (or aerofoil) portion 306, blade tip 308, stop-off material 310, membrane 312, cavity die (mould) 314, gaps 315, internal cavity 316; fan or aerofoil blade 400, sub-cavity 402, corrugated or concertinaed membrane 404, panel 406, root 408.

Claims

CLAIMS1. An aerofoil blade for a turbo machine, the blade (400) having a root portion (408) shaped for attachment to the turbo machine, wherein the blade (400) is formed to define an internal cavity (316) in at least the root portion (408) of the blade.
2. An aerofoil blade for a turbo machine according to claim 1, wherein the blade 400 has a leading edge, a trailing edge and a tip region distal to the root portion mutually arranged to define a blade portion of the blade, wherein the internal cavity (316) extends from the root portion (408) into the blade portion.
3. An aerofoil blade for a turbo machine according to claim 1 or 2, wherein the root portion (408) includes one or more root members (304), projecting laterally outward from the blade relative to the leading and trailing edges.
4. An aerofoil blade for a turbo machine according to any one of claims 1 to 3, wherein the root portion (408) includes a pair of root members (304) projecting laterally outward from the blade in mutually opposite directions; and wherein the blade (400) is formed to define the internal cavity (316) in at least a region of the root portion (408) between the pair of root members (304).
5. An aerofoil blade for a turbo machine according to any one of claims 3 and 4 wherein the or each root member (304) is configured for interlocking engagement with a complementary element of the turbo machine for maintaining attachment of the blade to the turbo machine during use.
6. An aerofoil blade for a turbo machine according to claim 5 wherein the root portion is under compressive stress during use.
7. An aerofoil blade for a turbo machine according to any one of the preceding claims, wherein the blade includes one or more membranes (404) arranged to sub-divide the internal cavity into a plurality of internal sub-cavities (402) in at least the root portion (408) of the blade.
8. An aerofoil blade for a turbo machine according to claim 7 as dependent on claim 2, wherein at least one of the internal sub-cavities (402) extends from the blade portion into the root portion.
9. An aerofoil blade for a turbo machine according to any one of the preceding claims wherein the aerofoil blade is a fan blade for a gas turbine engine.
10. A turbomachine including an aerofoil blade according to any one of the preceding claims.
11. A gas turbine engine including an aerofoil blade according to any one of claims 1 to 9.
12. A method of manufacture of an aerofoil blade for a turbo machine, the method including the steps of: providing a stacked assembly (300), including a pair of opposing panels (302), each panel including a respective root member (304) projecting outward from the assembly and mutually cooperating to define a root portion of the assembly; performing a hot forming process to form an internal cavity (316) within the assembly (300) at least in the root portion.
13. A method according to claim 12, the method further including the steps of: arranging the stacked assembly (300) in a cavity die (314); and the hot forming process including the step of inflating the assembly (300) to form the internal cavity (316) by conforming the assembly (300) to an internal region of the cavity die (316).
14. A method according to claim 13, wherein the step of conforming the assembly (300) to the internal region of the cavity die includes deforming the root portions (304) of the assembly to conform to a root defining sub-region of the internal region of the cavity die (314).
15. A method according to any one of claims 12 to 14 further including the steps: performing a bonding process to bond only predetermined regions of the interface(s) within the assembly (300); performing the hot forming process to form the internal cavity (316) within the assembly (300) in one or more unbonded regions of the interface(s) within the assembly.
16. A method of manufacture according to claim 15, wherein the assembly (300) includes a membrane (312) sandwiched between the opposing panels.
17. A method of manufacture according to claim 15 or 16 wherein the bonding process includes the step of providing at least one interface with a stop-off layer (312), for preventing bonding of the interface(s) in the region of the layer, to define said predetermined regions.
18. A method according to claim 17 wherein at least a portion of the at least one interface provided with the stop-off layer (312) is located between the root portions (304).
19. A method of manufacture according to any one of claims 15 to 18 wherein the bonding process includes the step of diffusion bonding the assembly (300) to bond the predetermined regions.
20. A method of manufacture according to any one of claims 12 to 19 wherein the hot forming process includes the steps of applying suitable temperature and pressure to the assembly to inflate the assembly (300) to conform to the internal region of the cavity die (314).
21. A method of manufacture according to any one of claims 12 to 20 wherein the hot forming process includes the step of superplastically forming the assembly (300) to thereby define the internal cavity (316).
22. A method of manufacture and/or assembly of a gas turbine engine including the step of incorporating into the engine an aerofoil (fan) blade according to any one of claims 1 to 11.