GB2260580A - Gas turbine engine aerofoil structure - Google Patents

Abstract

A gas turbine engine rotor stage 19 comprises an annular array of aerofoil blades 26 mounted on a ring-shaped support member 24. The ring-shaped support member 24 is supported and driven by a rotor disc 20. Each of the aerofoil blades 26 has a bifurcated root portion 29 which straddles and is bonded by for instance brazing to the ring-shaped support member 24. The ring-shaped support member 24 is preferably formed from ceramic fibre reinforced titanium and the aerofoil blades 26 from ceramic fibre reinforced glass or glass ceramic. <IMAGE>

Description

GAS TURBINE ENGINE AEROFOIL STRUCTURE This invention relates to a gas turbine engine aerofoil structure and particularly, although not exclusively, to an aerofoil bladed rotor of such an engine.

Axial flow gas turbine engines conventionally include compressors and turbines which comprise axially alternate annular arrays of aerofoils. Each array comprises either aerofoil blades or aerofoil vanes: the aerofoil blades being mounted on a suitable disc for rotation about the engine axis while the aerofoil vanes are mounted statically.

The conventional manner of mounting both aerofoil blades and vanes within compressors and turbines is by some form of mechanical fixing. For example, a common way of fixing aerofoil blades on their rotor discs is by the use of the so-called fir-tree root arrangement. In that arrangement, the radially inner or root portion of the aerofoil blade is provided with a fir tree-like cross-sectional configuration. The root is received within a correspondingly shaped cut-out provided on the periphery of the rotor disc.

While this and other forms of mechanical fixing provide adequate blade and vane anchorage, they do tend to be bulky, thereby adding undesirable weight to the engine. Moreover they are not suitable if the aerofoil or the structure upon which it is mounted are formed from composite or ceramic materials.

An alternative form of attaching an aerofoil blade to a rotor disc is described in UK Patent No, 1170592. In that reference, there is described an aerofoil blade having a root made up of a large number of tangs which locate and are bonded within annular grooves in the rotor disc. However such an arrangement is complicated, and therefore expensive to produce. It is, nevertheless suitable for use with composite and ceramic materials. Its major drawback, however, is that it is not amenable to easy repair. If, for instance, one of the aerofoil blades on the rotor disc required repair, it would be very difficult to remove it from the disc.

It is an object of the present invention to provide an aerofoil mounting arrangement in which such difficulties are substantially avoided.

According to the present invention, a gas turbine engine aerofoil structure comprises a ring-shaped support member carrying an annular array of aerofoil members, each of said aerofoil members having a root portion to facilitate its attachment to said ring-shape support member, said root portion being bifurcated to straddle said ring-shaped support member, said bifurcated root portion defining surfaces which abut and are bonded to corresponding surfaces on said ring-shaped support member.

The present invention will now be described, by way of example, with reference to the accompanying drawings in which: Figure 1 is a sectioned side view of the upper half of a ducted fan gas turbine engine having an aerofoil mounting arrangement in accordance with the present invention.

Figure 2 is a partially exploded view of a portion of an aerofoil bladed rotor of the high pressure compressor of the ducted fan gas turbine engine shown in Figure 1.

Figure 3 is a group of four cross-sectional views of alternative aerofoil mounting arrangements in accordance with the present invention.

Figure 4 is a view of a partially manufactured aerofoil blade for use in the aerofoil mounting arrangement in accordance with the present invention.

Figure 5 is a view of the partially manufactured aerofoil blade shown in Figure 4 in the process of being pressed into its final configuration.

Figure 6 is a view of an alternative form of the partially manufactured aerofoil blade shown in Figure 4.

Referring to Figure 1, a ducted fan gas turbine engine generally indicated at 10 comprises, in axial flow series, a ducted fan 11, an intermediate pressure compressor 12, a high pressure compressor 13, combustion equipment 14, high intermediate and low pressure compressors 15,16 and 17 respectively and an exhaust nozzle 18. Drive shafts ila, 12a and 13a respectively interconnect the fan 11 and the low pressure turbine 17, the intermediate pressure compressor 12 and the intermediate pressure turbine 16, and the high pressure compressor 13 and the high pressure turbine 15.

The gas turbine engine 10 functions in the conventional manner whereby air accelerated by the fan 11 is divided into two flows: one to provide propulsive thrust and the other directed into the intermediate pressure compressor 12. The intermediate and high pressure compressors 12 and 13 compress the air before directing it into the combustion equipment 14. There fuel is mixed with the compressed air and the mixture combusted. The resultant combustion products then expand through, and thereby drive, the high, intermediate and low pressure turbines 15,16 and 17 before being exhausted through the exhaust nozzle 18 to provide propulsive thrust.

The high pressure compressor 13 comprises a plurality of axially alternate rotor and stator aerofoil stages. Each rotor stage comprises a rotor disc carrying an annular array of radially extending aerofoil blades. Similarly each stator stage comprises an annular array of statically mounted radially extending aerofoil vanes.

The manner of construction of one of the high pressure compressor 13 rotor stages can be seen if reference is now made to Figure 2. The rotor stage generally indicated at 19 comprises a rotor disc 20 which rotates with the other rotor stages of the high pressure compressor 13. The radially outer edge of the disc 20 is defined by an axially extending flange 21. The flange 21 carries a plurality of radially extending pegs 22, one of which can be seen in the drawing.

Each of the pegs 22 locates in a corresponding hole provided in the radially inner surface of a ring 23. The ring 23 is formed from a metal matrix composite material. A suitable material would, for instance, comprise a titanium alloy matrix reinforced by continuous circumferentially extending inorganic fibres of silicon carbide. Other suitable inorganic fibres could be used if so desired.

The ring 23 is radially spaced apart from the disc 20 and is of generally triangular cross-sectional shape.

However its two surfaces having radially extending components 24 and 25 are, as can be seen in the drawing, slightly convex.

The ring 23 carries a plurality of similar radially extending circumferentially adjacent aerofoil blades, one of which 26 can be seen in Figure 2. Each aerofoil blade 26 comprises an aerofoil portion 27 and a root portion 28. The aerofoil portion 27 is of conventional configuration.

However, the root portion 28 is bifurcated to define two similar sheets 29 (only one of which can be seen in full in Figure 2). The bifurcated root 28 straddles the ring 25 and the sheets 29 are so configured that their surfaces adjacent the ring 25 correspond in shape with and abut the ring 25.

The sheets 29 are bonded to the ring 23 in such a manner that the bond can be readily broken if it is desired to replace the blade 26. Such a replacement may, for instance, be necessary if the blade 26 has been damaged in service. A variety of different bonds may be used if desired. Thus, for instance a suitable adhesive could be employed. Alternatively brazing could be used. The important characteristic of the type of bonding employed being that bond remains sound during normal operation but can be readily broken in the event of blade 26 replacement.

In order to provide the gas passage defining platforms which are a necessary feature of gas turbine engine aerofoil blades, separate platform pieces 30 are bonded to the bifurcated blade root 28: one platform piece 30 to each root sheet 29. Again the manner of bonding used is one which facilitates easy removal of the platform pieces 30.

The platform pieces 30 may be formed from any suitable material. However we prefer to use a glass ceramic foam material. This ensures that the ring 25 is shielded from the hot air flowing through the high pressure compressor 13.

Moreover it is of low density, thereby reducing engine weight. A further advantage is that in the event of one of the platform pieces 30 becoming detached, it will rapidly disintegrate into a powder which will cause little damage to the remainder of the engine.

Although in the embodiment shown in Figure 2, the ring 25 is shown as being of generally triangular cross-sectional shape, other configurations could be utilised if so desired.

A selection of such alternative configurations is shown in Figure 3. In each case, the sheets 29 of the aerofoil bifurcated roots 28 are so configured as to correspond in shape with that of the ring 25.

The aerofoil blade 26 is also of composite structure consisting of fibres of silicon carbide reinforcing a matrix of glass or glass ceramic. Such materials, and in particular those which include a glass ceramic matrix, are capable of withstanding high temperatures (up to 10000C) and are therefore fully capable of withstanding the high temperatures encountered in the latest high pressure compressors. In addition, the use of titanium in the aerofoil blades 26 is avoided, thereby removing the risk, at least in this particular part of the engine, of titanium fires.

The aerofoil rotor blades 26 are constructed by continuous filament winding silicon carbide fibres on to a generally y-shaped tool until a structure similar to that shown at 31 in Figure 4 is formed. A glass matrix is then incorporated into the resultant fibre structure by, for instance a sol-gel infiltration process. In that process a glass sol is incorporated into the fibre structure whereupon the sol is gelled to produce a self-supporting structure.

Alternatively glass or glass ceramic fibres could be incorporated into the fibre structure during the filament winding process.

The resultant fibre/matrix structure is then removed from the Y-shaped former and hot pressed at its final shape with suitably shaped dies 32 as shown in Figure 5.

Instead of being formed by filament winding, the fibre structure 31 could be formed by weaving. This would involve the weaving by conventional means of a tube-like structure as shown in Figure 6. Slits 33 would then be cut in the tubes to define the fibre reinforcement component of the aerofoil sheets 29.

It will be seen therefore that the present invention facilitates the use of two different composite materials in a high pressure compressor, taking advantage of the best properties of those materials i.e. the high temperature resistance of the aerofoil blades 26 and the strength properties of the ring 25. This is achieved with a low weight structure which may be readily repaired in the event of the failure of one or more of the aerofoil blades 26.

Although the present invention has been described with reference to a bladed rotor for a gas turbine engine high speed compressor, it is applicable to other gas turbine engine bladed rotors, for instance those in the engine's turbine. Appropriate materials would of course have to be used in such applications. Additionally, the present invention is also applicable to aerofoil stator structures.

Typically in such a structure, the ring 25 would be mounted radially outwardly of the stator aerofoils.

Claims (18)

Claims:

1. A gas turbine engine aerofoil structure comprising a ring-shaped support member carrying an annular array of aerofoil members, each of said aerofoil members having a root portion to facilitate its attachment to said ringshaped support member, said root portion being bifurcated to straddle said ring-shaped support member, said bifurcated root portions defining surfaces which abut and are bonded to correspondingly surfaces on said ring-shaped support member.

2. A gas turbine engine aerofoil structure as claimed in claim 1 wherein each of said aerofoil members is provided with means to define at least one platform, said at least one platform in turn defining a portion of the boundary of the gas passage in which its associated aerofoil member is operationally located.

3. A gas turbine engine aerofoil structure as claimed in claim 2 wherein each of said platform defining means is bonded to the bifurcated root portion of its associated aerofoil member.

4. A gas turbine engine aerofoil structure as claimed in claim 3 wherein each of said platform defining means is formed from a foamed ceramic material.

5. A gas turbine engine aerofoil structure as claimed in claim 4 wherein said foamed ceramic material is foamed glass ceramic.

6. A gas turbine engine aerofoil structure as claimed in any one preceding claim wherein said structure is so constructed as to function as a rotor stage of a gas turbine engine.

7. A gas turbine engine aerofoil structure as claimed in claim 6 wherein said ring-shaped support member is both supported and driven by a rotor disc.

8. A gas turbine engine aerofoil structure as claimed in claim 7 wherein said ring-shaped support member is located radially outwardly of the rotor disc which supports it.

9. A gas turbine engine aerofoil structure as claimed in any one preceding claim wherein said ring-shaped support member is formed from a fibre reinforced metal matrix composite material said fibres being substantially continuous and circumferentially extending.

10. A gas turbine engine aerofoil structure as claimed in claim 9 wherein said fibre is ceramic.

11. A gas turbine engine aerofoil structure as claimed in claim 9 or claim 10 wherein said metal matrix comprises titanium.

12. A gas turbine engine aerofoil structure as claimed in any one preceding claim wherein said aerofoil members are formed from a fibre reinforced non-metal matrix composite material.

13. A gas turbine engine aerofoil-structure as claimed in claim 12 wherein the fibres in said aerofoil members are ceramic.

14. A gas turbine engine aerofoil structure as claimed in claim 12 or claim 13 wherein said non-metal matrix of said aerofoil members is either glass or glass ceramic.

15. A gas turbine engine aerofoil structure as claimed in any one preceding claim wherein ring-shaped support member is of generally triangular cross-sectional configuration.

16. A gas turbine engine aerofoil structure as claimed in claim 15 wherein the surfaces of said ring-shaped support member to which are bonded said bifurcated root portions are slightly convex.

17. A gas turbine engine aerofoil structure as claimed in any one preceding claim wherein said aerofoil members are bonded to said ring-shaped support member by brazing.

18. A gas turbine engine aerofoil structure substantially as hereinbefore described with reference to and as shown in the accompanying drawings.