Drive Shaft

Abstract

An apparatus comprising a shaft for a gas turbine engine, the shaft comprising a shaft flange; a turbine rotor for the gas turbine engine, the turbine rotor comprising a turbine rotor flange configured to couple to the shaft flange; and a stub shaft comprising a stub shaft flange configured to couple to the shaft flange. The stub shaft is concentric with the shaft and configured to have a first interference fit around a portion of the turbine rotor.

Claims

1. An apparatus comprising: a shaft for a gas turbine engine, the shaft comprising a shaft flange; a turbine rotor for the gas turbine engine, the turbine rotor comprising a turbine rotor flange configured to couple to the shaft flange; and a stub shaft comprising a stub shaft flange configured to couple to the shaft flange, wherein the stub shaft is configured to be concentric with the shaft and to have a first interference fit around a portion of the turbine rotor.

2. The apparatus as claimed in claim 1, wherein the stub shaft flange comprises an axial protrusion with a first surface facing radially inward, and the turbine rotor comprises a shoulder with a second surface facing radially outward, and the first interference fit is between the first surface and the second surface.

3. The apparatus as claimed in claim 1, wherein the stub shaft is configured to have a second interference fit around the shaft.

4. The apparatus as claimed in claim 3, wherein the shaft comprises a radial protrusion having a third surface, facing radially outward, and the stub shaft comprises a fourth surface, facing radially inward, and the second interference fit is between the third surface and the fourth surface.

5. The apparatus as claimed in claim 3, wherein the stub shaft comprises a radial protrusion having a fourth surface facing radially inward, and the shaft comprises a third surface facing radially outward, and the second interference fit is between the third surface and the fourth surface.

6. The apparatus as claimed in claim 3, wherein both the first and second interference fits are provided between surfaces of the stub shaft that face radially inward and a respective surface of the shaft and rotor that faces radially outwards.

7. The apparatus as claimed in claim 1, wherein the first interference fit radially locates the turbine rotor on the stub shaft.

8. The apparatus as claimed in claim 1, wherein each of the shaft flange, turbine rotor flange and stub shaft flange are provided with a plurality of holes on a common bolt circle, each hole configured to receive a bolt that passes through each of the shaft flange, turbine rotor flange and stub shaft flange.

9. The apparatus as claimed in claim 8, wherein each of the holes have parallel walls.

10. The apparatus as claimed in claim 8, further comprising a bolt for each of the holes in the shaft flange, wherein each bolt is configured to have a clearance fit in the corresponding hole of each of the shaft flange, turbine rotor flange and stub shaft flange.

11. The apparatus as claimed in claim 1, wherein the stub shaft comprises a bearing surface configured to be received in a bearing.

12. The apparatus as claimed in claim 1, wherein the stub shaft comprises sealing protrusions for forming a seal between the stub shaft and a stationary further seal element, wherein at least one of the sealing protrusions is provided from an axial protrusion extending from the stub shaft flange.

13. The apparatus as claimed in claim 1, wherein the turbine rotor flange is configured to be received between the shaft flange and the stub shaft flange.

14. The apparatus as claimed in claim 1 in combination with a gas turbine engine for an aircraft comprising: an engine core comprising a turbine, a compressor, and a core shaft connecting the turbine to the compressor; a fan located upstream of the engine core, the fan comprising a plurality of fan blades; and a gearbox that receives an input from the core shaft and outputs drive to the fan so as to drive the fan at a lower rotational speed than the core shaft; wherein the shaft of the apparatus is the core shaft.

15. The apparatus as claimed in claim 1 in combination with a gas turbine engine for an aircraft comprising: an engine core comprising a first turbine, a first compressor, and a first core shaft connecting the first turbine to the first compressor; a fan located upstream of the engine core, the fan comprising a plurality of fan blades; a gearbox that receives an input from the first core shaft and outputs drive to the fan so as to drive the fan at a lower rotational speed than the first core shaft; a second turbine, a second compressor, and a second core shaft connecting the second turbine to the second compressor; the second turbine, second compressor, and second core shaft arranged to rotate at a higher rotational speed than the first core shaft; and an apparatus as claimed in claim 1, wherein the shaft of the apparatus is the first core shaft or the second core shaft.

16. A method of coupling a turbine rotor to a stub shaft, the method comprising the steps of: causing a temperature difference between the stub shaft and the turbine rotor thereby causing the stub shaft to expand relative to the turbine rotor so that there is a clearance fit between the stub shaft and the turbine rotor; fitting the turbine rotor to the stub shaft; and reducing the temperature difference between the stub shaft and the turbine rotor to cause a first interference fit between the stub shaft and the turbine rotor.

17. The method as claimed in claim 16, further comprising coupling the stub shaft to a shaft, by: causing a temperature difference between the stub shaft and shaft thereby causing the stub shaft to expand relative to the shaft so that there is a clearance fit between the stub shaft and the shaft; fitting the stub shaft to the shaft; and reducing the temperature difference between the stub shaft and the shaft to cause a second interference fit between the stub shaft and the shaft.

18. The method as claimed in claim 17, comprising causing a temperature difference between the stub shaft and both of the shaft and the turbine rotor at the same time, so as to cause a clearance fit between the stub shaft and each of the shaft and the turbine rotor at the same time.

19. The method as claimed in claim 17, wherein the temperature difference is caused by applying heat to the stub shaft, to cause the temperature of the stub shaft to increase relative to the turbine rotor and/or the shaft.

20. The method as claimed in claim 17, wherein the method is performed using the apparatus as claimed in claim 1, or the gas turbine as claimed in claim 16.

Description

BRIEF DESCRIPTION OF THE DRAWINGS

[0074] Embodiments will now be described by way of example only, with reference to the Figures, in which:

[0075] FIG. 1 is a sectional side view of a gas turbine engine;

[0076] FIG. 2 is a close up sectional side view of an upstream portion of a gas turbine engine;

[0077] FIG. 3 is a partially cut-away view of a gearbox for a gas turbine engine;

[0078] FIG. 4 is a sectional view of a coupling between a core shaft and a turbine rotor, not in accordance with an embodiment; and,

[0079] FIG. 5 is a sectional view of an example embodiment.

DETAILED DESCRIPTION OF THE DISCLOSURE

[0080] 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 and an epicyclic gearbox 30.

[0081] 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.

[0082] 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.

[0083] 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.

[0084] 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.

[0085] 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.

[0086] 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.

[0087] 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.

[0088] Optionally, the gearbox may drive additional and/or alternative components (e.g. the intermediate pressure compressor and/or a booster compressor).

[0089] 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 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.

[0090] 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.

[0091] Referring to FIG. 4, a sectional view of a coupling not in accordance with an embodiment is shown, comprising a turbine rotor 130, a shaft 110, and a stub shaft 120.

[0092] The turbine rotor 130 comprises blades that derive power from the core fluid flow, which comprises combustion products and core air B. The blades are radially outward from the coupling area that is shown in FIG. 4, and therefore not visible in FIG. 4.

[0093] The turbine rotor 130 comprises a rotor flange 136 that includes a plurality of holes on a common bolt circle about the rotational axis 9 (the sectional view being of a plane including axis of one of these holes and the rotational axis 9).

[0094] The shaft 110 includes a shaft flange 116 that includes a plurality of holes on the common bolt circle, each of which lines up with a corresponding hole on the rotor flange 136.

[0095] The stub shaft 120 comprises a stub shaft flange 126 that includes a plurality of holes on the common bolt circle, each of which lines up with a corresponding hole on the rotor flange 136.

[0096] The stub shaft 120 further comprises a bearing surface 121, supported by a bearing 105 in contact therewith.

[0097] A bolt 140 is provided for each of the holes in the shaft flange 116. The bolt 140 is a taper bolt, and each of the holes in in the shaft flange 116 and the turbine rotor flange 136 have a corresponding taper, so that each bolt 140 forms a precision taper fit with each of the tapering holes in the shaft flange and turbine rotor flange 136. The bolt 140 has a clearance fit with the hole in the stub shaft flange 126. A nut 141 is provided adjacent to the stub shaft flange 126.

[0098] A load path 150 illustrates the path of radial loads from the rotor 130 to the bearing 105. The tolerance of this load path 150 is dependent on the fit of the taper bolts 140 with the rotor flange 136 and shaft flange 116, and on the fit between the shaft 110 and the stub shaft (between surfaces 125 and 115).

[0099] FIG. 5 shows a sectional view of an apparatus according to an embodiment of the present disclosure. The apparatus comprises a turbine rotor 130, stub shaft 120 and shaft 110.

[0100] The turbine rotor 130 comprises blades that derive power from the core fluid flow, which comprises combustion products and core air B. The blades are radially outward from the coupling area that is shown in FIG. 5, and therefore not visible in FIG. 5.

[0101] The turbine rotor 130 comprises a rotor flange 136 that includes a plurality of holes on a common bolt circle about the rotational axis 9 (the sectional view being of a plane including axis of one of these holes and the rotational axis 9).

[0102] The shaft 110 includes a shaft flange 116 that includes a plurality of holes on the common bolt circle, each of which lines up with a corresponding hole on the rotor flange 136.

[0103] The stub shaft 120 comprises a stub shaft flange 126 that includes a plurality of holes on the common bolt circle, each of which lines up with a corresponding hole on the rotor flange 136. The stub shaft 120 further comprises a bearing surface 121 supported by a bearing 105 in contact therewith.

[0104] A bolt 140 is provided for each of the holes in the shaft flange 116. In contrast to the arrangement of FIG. 4, the bolt 140 is not required to be a taper bolt, and may be a conventional cylindrical bolt configured for a clearance fit with each of the holes in the shaft flange 116, rotor flange 136 and stub shaft flange 126. A nut 141 is provided adjacent to the stub shaft flange 126. The stub shaft flange 126 may be configured to prevent rotation of the nut 141, so that the bolt 140 may be tightened without access to the nut 141.

[0105] In embodiments, the stub shaft 120 is configured to have an interference fit with the rotor 130. This simplifies the load path for reacting lateral forces from the rotor 130. In some embodiments and may obviate the need for taper bolts. The tolerance stack-up of the load path 150 (from the rotor 130 to the bearing 105) may be simplified because fewer tolerances are involved, and the through life costs may be reduced by the elimination of taper bolts. The interference fit between the rotor 130 and the stub shaft 120 enables loads to be transmitted directly from the rotor 130 to the stub shaft 120, and thence to the bearing that supports the stub shaft 120 The interference fit between the rotor 130 and the stub shaft 120 may also carry torque. The torque is carried by the friction between the rotor flange 136 and shaft flange 116 caused by the pre-load arising from tension in the bolts 140.

[0106] The use of an interference fit between the stub shaft 120 and the rotor 130 may be particularly useful in the context of a gas turbine engine comprising a gearbox that receives an input from a core shaft and outputs drive to a fan, so as to drive the fan at a lower speed than the core shaft. In such engines it is advantageous for the core shaft to be relatively long, which causes problems in achieving taper reaming operations because very long shafts are typically difficult to accommodate in existing machine tools for taper reaming operations. The interference fit between the stub shaft 120 and the rotor 130 enables a long core shaft to be used.

[0107] In order to provide the first interference fit, between the stub shaft 120 and the rotor 130, in the example embodiment the stub shaft flange 126 is provided with an axial protrusion 124 (or first spigot) extending from the stub flange 126 in the direction of the rotor flange 136. The axial protrusion 124 defines a first surface 127, which is substantially cylindrical and faces radially inward. The rotor flange 136 comprises a shoulder with a second, corresponding second surface 137, facing radially outwards. The first surface 127 is configured to be an interference fit with the second surface 137, so that the axial protrusion 124 of the stub shaft 120 grips the shoulder of the rotor 130.

[0108] In this embodiment, a second interference fit is provided between the stub shaft 120 and the shaft 110. In this embodiment this second fit is at an outward radial protrusion from the shaft 110 (with the stub shaft 120 defining a second spigot over this protrusion). The radial protrusion ends in a third surface 115, facing radially outward. The stub shaft 120 comprises a corresponding fourth surface 125, facing radially inward, which is configured to be an interference fit with the third surface 115. It will be understood that the second interference fit may be at a protrusion facing radially inward from the stub shaft instead of, or in addition to any radial protrusion from the shaft 110. In some embodiments the second interference fit is not associated with any protrusions (from either the shaft 110 or stub shaft 120).

[0109] Both the first and second interference fits are provided between surfaces 125, 127 of the stub shaft 120 that face radially inward and a respective surface 115, 137 of the shaft 110 and rotor 130 that faces radially outwards. This enables the stub shaft 120 to be shrink fitted to both the shaft 110 and rotor 130, by heating the stub shaft 120 relative to the shaft 110 and rotor 130 (and/or cooling the shaft 110 and rotor 130 relative to the stub shaft 120).

[0110] A load path 150 illustrates the path of transverse loads from the rotor 130 to the bearing 105. The tolerance of this load path 150 is dependent only on the first interference fit (between the rotor 130 and stub shaft 110). This is a simpler and more controllable tolerance stack than in the arrangement of FIG. 4.

[0111] The stub shaft 120 may be provided with sealing protrusions 122. In the example embodiment, the stub shaft 120 comprises a further axial protrusion 123 from the stub shaft flange 126, extending in the opposite direction to the protrusion 124. At least one sealing protrusion may be provided on the further axial protrusion (the sealing protrusions facing radially outward). The sealing protrusions may be configured to form a labyrinth seal with a corresponding stationary part of the engine. Further sealing protrusions 122 may be provided on other portions of the stub shaft, for example between the bearing surface 121 and the stub shaft flange 126.

[0112] 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.