BLADE MOUNTING

Abstract

An aerofoil blade having a root portion provided with a low friction coating layer. The low friction coating layer is affixed to the root portion by an adhesive layer. The adhesive layer has a service temperature of 125 C. or more, for example 140 C. or more.

Claims

1. An aerofoil blade for a gas turbine engine, the aerofoil blade having a root portion provided with a low friction coating layer, the low friction coating layer being affixed to the root portion by an adhesive layer, wherein the adhesive layer has a service temperature of 125 C. or more.

2. The aerofoil blade of claim 1, wherein the aerofoil blade is a composite aerofoil blade.

3. The aerofoil blade of claim 1, wherein the adhesive layer has a service temperature of 130 C. or more,

4. The aerofoil blade of claim 1, wherein the adhesive layer has a service temperature of 140 C. or more.

5. The aerofoil blade of claim 1, wherein the low friction coating layer comprises a PTFE impregnated fabric layer.

6. The aerofoil blade of claim 1, wherein the root portion further comprises a metallic foil interposed between the adhesive layer and the low friction coating layer.

7. A method of affixing a low friction coating layer to a root portion of an aerofoil blade for a gas turbine engine, said method comprising the steps of: applying an adhesive to the root portion, affixing the low friction coating layer to the adhesive, and curing the adhesive to form an adhesive layer having a service temperature of 125 C. or more.

8. The method of claim 7, comprising curing the adhesive to form an adhesive layer having a service temperature of 130 C. or more.

9. The method of claim 7, comprising curing the adhesive to form an adhesive layer having a service temperature of 140 C. or more.

10. The method of claim 7, further comprising providing a metallic foil layer between the adhesive layer and the low friction coating layer.

11. The method of claim 7, further comprising applying a lubricator to the root portion with the low friction coating layer.

12. A gas turbine engine comprising at least one aerofoil blade of claim 1.

13. The gas turbine of claim 12, comprising a fan section formed of a plurality of aerofoil blades of claim 1, circumferentially-arranged around a rotor disc.

Description

DESCRIPTION OF THE DRAWINGS

[0054] Embodiments will now be described by way of example only, with reference to the Figures, in which:

[0055] FIG. 1 is a sectional side view of a gas turbine engine;

[0056] FIG. 2 is a close up sectional side view of an upstream portion of a gas turbine engine;

[0057] FIG. 3 is a partially cut-away view of a gearbox for a gas turbine engine; and

[0058] FIG. 4 is view of a first example of a root portion of a composite fan blade; and

[0059] FIG. 5 is a view of a second example of a root portion of a composite fan blade.

DETAILED DESCRIPTION

[0060] Embodiments will now be described by way of example only, with reference to the Figures.

[0061] 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.

[0062] 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 system 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.

[0063] Arrangements of the present disclosure may be particularly, although not exclusively, beneficial for fans 23 that are driven via a gearbox 30. Accordingly, the gas turbine engine may comprise a gearbox 30 that receives an input from the core shaft 26 and outputs drive to the fan 23 so as to drive the fan 23 at a lower rotational speed than the core shaft 26. The input to the gearbox 30 may be directly from the core shaft 26, or indirectly from the core shaft 26, for example via a spur shaft and/or gear.

[0064] 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.

[0065] 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.

[0066] 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.

[0067] 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.

[0068] 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.

[0069] 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.

[0070] Optionally, the gearbox may drive additional and/or alternative components (e.g. the intermediate pressure compressor and/or a booster compressor).

[0071] 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 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. 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.

[0072] 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.

[0073] The fan 23 comprises a plurality of circumferentially-arranged, radially-extending composite fan blades formed of laminated layers of carbon epoxy composite. As shown in FIG. 4, each fan blade has dovetail root portion 100 which is received in a slot in a rotor disc 101. The root portion 100 has a low friction coating layer comprising a PTFE impregnated fabric layer 102 such as Dupont's Vespel CP-0664. The PTFE fabric layer 102 faces the contact surface 103 of the slot in the rotor disc 101.

[0074] The PTFE fabric layer 102 is affixed to the root portion by an adhesive layer 104 having a service temperature higher than 125 C. For example, the PTFE fabric layer may be affixed to the root portion using Henkel's EA-9695 which has a service temperature greater than 149 C.

[0075] FIG. 5 shows another example where the PTFE fabric layer 102 is backed by a metallic foil layer 105 which is interposed between the PTFE fabric layer 102 and the adhesive layer 104.

[0076] By using an adhesive layer 104 having a service temperature greater than 125 C., 130 C., 140 C. or 145 C. etc, localised heating arising from the relative slip movement between the root portion 100 and rotor disc 101 during blade vibration does not result in thermal degradation of the adhesive layer 104 and thus the PTFE fabric layer 102 remains mechanically sound.

[0077] It will be understood that the disclosure 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.