DE-ICING SYSTEM

Abstract

A de-icing system for one or more vanes of a gas turbine engine. The system comprises a coating of electrically conductive material on at least a portion of a vane and a plurality of magnets located circumferentially about a rotor shaft of the gas turbine engine and configured to be driven by the rotor shaft in the circumferential direction to generate a rotating magnetic field. A coiled wire is configured to remain stationary within the rotating magnetic field so as to induce a current therein, the coiled wire being connected to the coating for supplying the induced current to the electrically conductive material. In this way, the coating of electrically conductive material is heated by the induced current, for inhibiting ice accretion on the vane.

Claims

1. A de-icing system for one or more vanes of a gas turbine engine, the de-icing system comprising: a coating of electrically conductive material on at least a portion of a vane; a plurality of magnets located circumferentially about a rotor shaft) of the gas turbine engine and configured to be driven by the rotor shaft in the circumferential direction to generate a rotating magnetic field in a space within the gas turbine engine; and a coiled wire located within the space and configured to be relatively stationary within the rotating magnetic field so as to induce a current therein; wherein the coiled wire is electrically connected to the coating to supply the induced current to the electrically conductive material, thereby heating the electrically conductive material for inhibiting the accretion of ice on the vane.

2. The de-icing system of claim 1, wherein the vane is a variable inlet guide vane or a variable stator vane of the gas turbine engine.

3. The de-icing system of claim 1, wherein the coating of electrically conductive material is formed on the leading edge and/or the trailing edge of the vane.

4. The de-icing system of claim 1, wherein the coating of electrically conductive material covers the entire surface area of the vane.

5. The de-icing system of claim 1, wherein the coating of electrically conductive material is an activated graphite ink.

6. The de-icing system of claim 1, wherein the coating of electrically conductive material is load-bearing.

7. The de-icing system of claim 1, wherein the coating of electrically conductive material is not load-bearing.

8. The de-icing system of claim 1, wherein the plurality of magnets are permanent magnets.

9. The de-icing system of claim 1, wherein the plurality of magnets are electromagnets powered by at least one power source.

10. The de-icing system of claim 9, wherein the plurality of electromagnets are connected to the at least one power source via a rotary electrical interface.

11. The de-icing system of claim 1, wherein the plurality of magnets is arranged in alternating polarity in the circumferential direction.

12. The de-icing system of claim 1, wherein the rotor shaft is a high-pressure shaft or an intermediate-pressure shaft of the gas turbine engine.

13. A gas turbine engine that includes a de-icing system of claim 1.

14. A method of de-icing one or more vanes of a gas turbine engine, the method comprising the steps of: providing a coating of electrically conductive material on at least a portion of a vane; generating a rotating magnetic field in a space within the gas turbine engine by the rotation of a plurality of magnets located circumferentially about and driven by a rotor shaft of the gas turbine engine; using the rotating magnetic field to induce a current in a coiled wire that is relatively stationary within the space; and supplying the induced current from the coiled wire to the coating of electrically conductive material to electrically heat the coating, thereby inhibiting the accretion of ice on the vane.

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 close up sectional side view of a compressor section of a gas turbine engine in accordance with a first embodiment of the technology described herein;

[0061] FIG. 5 schematically illustrates a close up sectional front view of the extension and the plurality of permanent magnets described with respect to FIG. 4;

[0062] FIG. 6 is a close up sectional side view of a compressor section of a gas turbine engine in accordance with a second embodiment of the technology described herein; and

[0063] FIG. 7 is a flow chart schematically illustrating a method of de-icing a guide vane of a gas turbine engine.

DETAILED DESCRIPTION OF THE DISCLOSURE

[0064] Aspects and embodiments of the present disclosure will now be discussed with reference to the accompanying figures. Further aspects and embodiments will be apparent to those skilled in the art.

[0065] FIG. 1 illustrates a gas turbine engine 10 having a principal rotational axis 9 (aka engine axis). 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.

[0066] 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 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. The epicyclic gearbox 30 is a reduction gearbox.

[0067] 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.

[0068] 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.

[0069] 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.

[0070] 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.

[0071] 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.

[0072] 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.

[0073] Optionally, the gearbox may drive additional and/or alternative components (e.g. the intermediate pressure compressor and/or a booster compressor).

[0074] 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 18, 20 meaning that the flow through the bypass duct 22 has its own nozzle 18 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.

[0075] 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.

[0076] FIG. 4 is a close up sectional side view of a compressor section of a gas turbine engine, such as the low pressure compressor 14 described with respect to FIGS. 1 and 2.

[0077] The low pressure compressor 14 comprises a radially inner casing 41 and a radially outer casing 42 that are stationary and held apart in the radial direction by a stationary supporting structure 43. The radially inner and outer casings 41, 42 define a flowpath 44 there between for receiving the core airflow A to be accelerated and compressed by the low pressure compressor 14, as described with respect to FIG. 1.

[0078] The compressor 14 also comprises a rotor structure comprising a rotor shaft 45 which is centred about the principal rotational axis 9, and a disc 46 mounted thereon (by conventional means) for rotation about the principal rotational axis 9. Attached along the radially outer circumference of the disc 46 is a plurality of compressor blades 47 extending in the radial direction, although only one such compressor blade 47 is shown here for clarity. During operation, the rotor structure is driven at high speed, e.g. by the low pressure turbine (not shown), such that the compressor blades 47 rotate and impart energy to the airflow passing through the low pressure compressor 14.

[0079] As can be seen in FIG. 4, the low pressure compressor 14 comprises a variable inlet guide vane 48 disposed between the inner and outer casing 41, 42 such that an aerofoil portion thereof is disposed within the flowpath 44. Although only a single guide vane 48 is shown in FIG. 4, it will be appreciated that the low pressure compressor 14 will include a plurality of such guide vanes 48 arranged circumferentially about the inner casing 41.

[0080] The guide vane 48 is rotatable about a rotational axis for controlling the airflow within the flowpath 44 to achieve efficient engine and compressor operation. For example, at low engine speed, the variable inlet guide vane 48 may be rotated to reduce the incidence of airflow onto the compressor blades 47 to tolerable angles. In order to achieve such rotation, each of the vanes 48 is mounted in a bush 49 in the inner casing 41 and a corresponding bush 49 in the outer casing 42. The vane 48 also has a lever 417 (or other linkage) fitted to its outer end on the outer side of the outer casing 42. The levers 417 of the plurality of variable inlet guide vanes 48 may be connected to a common unison ring (as known in the art) so that, when the unison ring is rotated, so do the vanes 48.

[0081] The variable inlet guide vane 48 often experiences icing conditions and is therefore susceptible to ice accretion. Accordingly, there is provided a de-icing system for inhibiting the formation of ice and reducing the amount of ice after its formation on the guide vane 48. However, in contrast to conventional systems in which the guide vanes are heated by hot air bled from the compressor, the system of the technology described herein heats the guide vanes by electrical heating of a coating of electrically conductive material on the guide vane.

[0082] The electrically conductive material 410 coats one or more or all portions of the guide vane 48 for this purpose. However, in the example shown in FIG. 4, a coating of electrically conductive material 410 is provided on the surface of the guide vane 48 in regions that have been found (by appropriate testing) as being particularly susceptible to the formation of ice, in this case portions encompassing a leading edge 411 and a trailing edge 412 of the guide vane 48. The electrically conductive material 410 serves as a heating element to convert an electric current passing there through into heat energy (through resistive heating) for directly heating the guide vane 48 by conduction. To serve as a heating element, the electrically conductive material 410 is an activated graphite ink, for example.

[0083] The coating may be applied as a thin film layer, which is non-load bearing. In this way, the coating may be used to heat the guide vane 48 without substantially increasing the weight of the guide vane 48. Alternatively, however, the coating may be applied as a thick layer, which is load-bearing. The coating is thick in that it is of sufficient thickness to carry a portion of the loads acting upon the guide vane (in addition to the guide vane itself), e.g at least 5% of the total centrifugal load across the cross section on the vane. In particular, the thickness of the coating may be selected to serve a dual purpose of not only heating the guide vane 48, but also to provide edge retention. Furthermore, the material properties of the electrically conductive material 410 that is used as the coating may be chosen to provide sufficient edge retention.

[0084] The coating of electrically conductive material 410 on the leading edge 411 is electrically connected at the radially inner side of the guide vane 48 to a first end of a coiled wire 413 (or solenoid) via electrical cabling 414. The coating of electrically conductive material 410 on the trailing edge 412, meanwhile, is electrically connected at the radially inner side of the guide vane 48 to a second end of the coiled wire 413 (via electrical cabling 414) opposite the first end. As can be seen in FIG. 4, the coating of electrically conductive material 410 extends at the radially outer side of the guide vane 48 between the portions of guide vane 48 that include the leading edge 411 and the trailing edge 412 of the guide vane 48. In this way, for example, the coating of electrically conductive material 410, the electrical cabling 414 and the coiled wire 413 form a closed circuit about which current can flow.

[0085] The electrical cabling 414 and the coiled wire 413 are both fixed to the internal wall of the inner casing 41 such that they remain stationary compared to the rotor structure of the low pressure compressor 14. In this way, as an example, the coiled wire 413 is configured to remain stationary in a space 419 which is to be subjected to a rotating magnetic field to generate a current directly in the coiled wire 413 by electromagnetic induction. As mentioned above, the rotating magnetic field is generated by a plurality of permanent magnets 415, only a first of which is shown in FIG. 4, which are located circumferentially about the rotor shaft 45 and configured to be driven by the rotor shaft 45 in the circumferential direction. The plurality of permanent magnets 415 are attached to a disc 46 of the rotor structure via an extension 416 bolted to the disc 46.

[0086] Each one of the plurality of permanent magnets 415 comprises a ferromagnetic core having two poles, a South pole at a first end of the magnet 415 and a North pole at a second end of the magnet 415 opposite the first end. As such, each permanent magnet 41 can be said to have a magnetic polarity in a direction from the South pole to the North pole.

[0087] As best shown in FIG. 5, the plurality of permanent magnets 415 are fixed on a radially outer surface of the extension 416 in an arrangement where each magnet 415 is at a position that is equidistant between its two closest neighbouring magnets 415. Although in FIG. 5 a common, continuous extension 416 is bolted to the disc 46 along the entire circumference of the rotor shaft 45, there may instead be a plurality of discrete extensions 416 bolted to the disc 46, one for each permanent magnet 415. It will also be appreciated that although FIG. 5 shows only four permanent magnets 415 fixed to the extension 416, in practice this number may be much greater.

[0088] The extension 416 has a radial extend from the rotor shaft 45 such that the plurality of permanent magnets 415 are disposed radially inwards of the coiled wire 413 and such that the coiled wire 413 is located within a rotational plane of the plurality of permanent magnets 415. Alternatively, the plurality of permanent magnets may be disposed at, e.g., the same radial position as the coiled wire, but offset therefrom (i.e. offset from the rotational plane) in the axial direction 9, for more efficient use of space. The plurality of permanent magnets 415 create a magnetic field of a magnitude suitable for extending across a gap (radial or axial gap) between the permanent magnets 415 and the coiled wire 413.

[0089] The plurality of permanent magnets 415 are arranged in a facing arrangement with the coiled wire 413, such that that when each permanent magnet 415 is at the same angular position about the rotor shaft 45 as the coiled wire 413, e.g. only, one pole of the magnet will face and confront the coiled wire 413. This facing arrangement may increase the number of magnetic field lines of each magnet 415 that are cut by the coiled wire 413 during rotation of the permanent magnets 415, e.g. as compared to hypothetical arrangements in which the magnets are arranged in a non-facing arrangement where a magnetic pole of the magnet does not confront the coiled wire 413.

[0090] A first permanent magnet 415a of the plurality of magnets 415 is oriented such that its North pole faces and confronts the coiled wire 413 in the radial direction perpendicular to the principal rotational axis 9. However, a second permanent magnet 415b, which is adjacent to the first permanent magnet 415a in the circumferential direction 51, is oriented such that its South pole will face and confront the coiled wire 413 in the radial direction. This pattern of permanent magnets having alternating North and South poles facing the coiled wire 413 is repeated for all of the plurality of permanent magnets located circumferentially about the rotor shaft 45. Alternating the polarity of the plurality of permanent magnets 415 may maximise the variation in magnetic field strength about the circumference of the rotor shaft 45, to increase the extent of electromagnetic induction.

[0091] Although the arrangement of FIGS. 4 and 5 has been described with respect to using permanent magnets 415 located circumferentially about the rotor shaft 45, a plurality of electromagnets 61 may be used instead of the permanent magnets, as will now be described with respect to FIG. 6.

[0092] FIG. 6 is a close up sectional side view of a compressor section of a gas turbine engine in accordance with a second embodiment of the technology described herein.

[0093] The compressor section of FIG. 6 corresponds to that described with respect to FIGS. 4 and 5 in that it comprises the same casing 41, 42 and rotor assembly, at least. However, the de-icing system of FIG. 6 differs from that of FIGS. 4 and 5 in that the plurality of permanent magnets 415 have been replaced by a plurality of electromagnets 61.

[0094] Each one of the plurality of electromagnets 61 has two poles, a South pole at a first end of the electromagnet 61 and a North pole at a second end of the electromagnet 61 opposite the first end. Furthermore, the plurality of electromagnets 61 are provided in the same facing arrangement described with respect to the permanent magnets 415 in FIGS. 4 and 5. The electromagnets 61 are also arranged in alternating polarity about the circumference of the rotor shaft 45, substantially as described with respect to FIG. 5.

[0095] To energise the plurality of electromagnets 61, there is provided a power source 62, e.g. one or more batteries, fixed to a stationary support platform 63 located radially inwards of the extension 416. The power source 62 is connected to the electromagnets 61 via a rotary electrical interface, in particular a slip ring 64, between the stationary support platform 63 and the extension 416. The slip ring 64 comprises a stationary electrical contact 65 (e.g. a brush) on the stationary support platform 63 which rubs on and electrically connects to a rotating electrical contact 66 on the extension 416, to allow the transmission of power from the power source 62 connected to the stationary electrical contact 65 and the electromagnet 61 connected to the rotating electrical contact 66. The provision of a slip ring 64 allows the power source to be positioned on a stator structure of the engine instead of the rotor structure. This reduces the amount of weight on the rotor and in turn the drag on the rotor structure.

[0096] The provision of electromagnets instead of permanent magnets allows the magnetic field to be selectively turned on and off by respectively energising and de-energising the power source. This may provide a more sophisticated and efficient system that heats the guide vanes only when required, and reduces drag caused by the magnetic field when heating is not required.

[0097] Turning to FIG. 7, the de-icing system described above with respect to FIGS. 4 to 6 is used to de-ice one or more guide vanes 48 of a gas turbine engine.

[0098] The method begins at step 71 by providing or assembling the de-icing system in a gas turbine engine. This includes providing the coating of electrically conductive material 410 on at least a portion of the guide vane 48 and electrically connecting the coating to the coiled wire 413. The de-icing method then proceeds to step 72, at which point the rotor shaft 45 is driven to rotate about the axis 9. This rotation in turn drives the plurality of permanent magnets or suitably energised electromagnets to rotate around the axis 9, thereby causing a magnetic field to rotate in the space 419 that is proximate the coiled wire 413. At step 73, the coiled wire 413 is maintained at a position within space 419 and therefore the rotating magnetic field during rotation of the rotor shaft 45, such that the coiled wire 413 remains relatively stationary within the rotating magnetic field. In this way, the relatively stationary coiled wire 413 continuously cuts the magnetic field lines so that the coiled wire 413 is subjected to a continual fluctuation of positive and negative magnetic field intensity, thereby inducing an alternating current in the coiled wire 413. The induced current is then supplied to the coating of electrically conductive material 410, at step 74, to heat the coating and in turn the guide vane, thereby preventing or reducing ice accretion thereon.

[0099] Although the de-icing method of FIG. 7 has been described above as having four distinct steps, it will be appreciated that the steps may be carried out substantially simultaneously.

[0100] It will be appreciated that although the de-icing system has been described with respect to FIGS. 4 to 7 as having a plurality of magnets arranged in alternating polarity across the circumferential direction, this is not essential. For example, it would be possible to generate a magnetic field of varying intensity with a plurality of magnets having a common polarity with respect to the coiled wire, to induce a current in the coiled wire. Furthermore, the plurality of magnets may be arranged circumferentially about the rotator shaft according to a Halbach array.

[0101] It will also be appreciated that although the plurality of magnets have been described as being in a facing requirement with the coiled wire in FIGS. 4 to 7, the plurality of magnets and the coiled wire may be arranged at any desired position or orientation relative to each other, as long as the coiled wire is configured to be within the rotating magnetic field generated by the plurality of magnets to induce a current therein.

[0102] Furthermore, although FIGS. 4 to 6 show a de-icing system for a single guide vane, it will be appreciated that a plurality of such guide vanes may be coated with the electrically conductive material and be in connection with respective coiled wires for heating the guide vanes in the manner described above. For example, there may be a plurality of coiled wires arranged circumferentially about the rotor shaft 45 perpendicular to the principal rotational axis 9 and configured to remain stationary within the rotating magnetic field created by the plurality of magnets rotated by the rotor shaft 45. Furthermore, although the de-icing system has been described with respect to a variable inlet guide vane, the technology described herein extends to any static vane, such a variable stator vane of the gas turbine engine.

[0103] 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.