ROTOR ASSEMBLY

Abstract

The present disclosure relates to a rotor assembly for a gas turbine engine, the rotor assembly comprising a first rotor stage having a first disc portion with a peripheral first rim portion and a second rotor stage, the second rotor stage having a second disc portion with a peripheral second rim portion. The second rotor stage is axially adjacent and downstream of the first rotor stage and the second rim portion has an axial extension extending towards the first rim portion such that the axial extension of the second rim portion defines a rotor drum cavity between the first and second disc portions. The second rotor stage further comprises a drive arm extending within the drum cavity to the first disc portion, the drive arm being connected to the first disc portion by at least one connector. The drive arm divides the drum cavity into radially outer rim cavity portion and a radially inner main cavity portion. The rotor assembly further comprises a rim seal located between the axial extension of the second rim portion and the first rim portion, and a pressure equalisation path extending from the rim cavity portion to the main cavity portion.

Claims

1. A rotor assembly for a gas turbine engine, the rotor assembly comprising: a first rotor stage having a first disc portion with a peripheral first rim portion; a second rotor stage, the second rotor stage having a second disc portion with a peripheral second rim portion, wherein the second rotor stage is axially adjacent and downstream of the first rotor stage and the second rim portion has an axial extension extending towards the first rim portion such that the axial extension of the second rim portion defines a rotor drum cavity between the first and second disc portions, wherein the second rotor stage further comprises a drive arm extending within the drum cavity to the first disc portion, the drive arm being connected to the first disc portion by at least one connector, wherein the drive arm divides the drum cavity into radially outer rim cavity portion and a radially inner main cavity portion, wherein the rotor assembly further comprises a rim seal located between the axial extension of the second rim portion and the first rim portion, wherein the rotor assembly comprises a pressure equalisation path extending from the rim cavity portion to the main cavity portion, and wherein the drive arm comprises an oblique portion extending from a radially inner surface of the axial extension of the second rim portion.

2. The assembly according to claim 1, wherein the drive arm comprises a radial portion extending from the oblique portion and wherein the connector(s) extend through the first disc portion and the radial portion of the drive arm.

3. The assembly according to claim 1, wherein the pressure equalisation path is provided through the drive arm.

4. The assembly according to claim 2, wherein the oblique portion of the drive arm comprises at least one aperture extending from a radially outer surface in the rim cavity portion to a radially inner surface in the main cavity portion.

5. The assembly according to claim 1, wherein the rim seal is a folded sheet seal having a first leaf in abutment with the first rim portion and a second leaf in abutment with the axial extension of the second rim portion.

6. The assembly according to claim 1, wherein the rim seal is an interference seal.

7. The assembly according to claim 1, wherein the rotor assembly is a compressor assembly and the first rotor stage is a first compressor stage and the second rotor stage is a second, adjacent compressor stage.

8. A gas turbine engine comprising a rotor assembly or a compressor assembly according to claim 1.

Description

DESCRIPTION OF THE DRAWINGS

[0059] Embodiments will now be described by way of example only, with reference to the Figures, in which:

[0060] FIG. 1 is a sectional side view of a gas turbine engine;

[0061] FIG. 2 is a close-up sectional side view of an upstream portion of a gas turbine engine;

[0062] FIG. 3 is a partially cut-away view of a gearbox for a gas turbine engine;

[0063] FIG. 4 is a sectional view of a prior art compressor assembly;

[0064] FIG. 5 is a sectional view of a first embodiment; and

[0065] FIG. 6 is a sectional view of a second embodiment.

DETAILED DESCRIPTION

[0066] FIG. 5 shows a first embodiment of a high-pressure compressor assembly comprising a first compressor stage having a first disc portion 100a with a peripheral first rim portion 101a and, axially adjacent and downstream of the first compressor stage, a second compressor stage, the second compressor stage having a second disc portion with a peripheral second rim portion (not shown).

[0067] The first rim portion (and the second rim portion) carries a series of circumferentially arranged compressor blades 102a.

[0068] The second rim portion has an axial extension 104b extending axially upstream towards the first rim portion 101a. The first compressor stage comprises a series of circumferentially arrange stators 103 which are radially aligned and radially outwards of the axial extension 104 of the second rim portion.

[0069] The axial extension 104b defines a compressor drum cavity between the first and second disc portions. The second compressor stage further comprises a drive arm 106 extending from a radially inner surface of the axial extension 104b of the second rim portion within the drum cavity to the first disc portion 100a. The drive arm 106 comprises a radial portion 106a extending from an oblique portion 106b.

[0070] The radial portion 106b lies in abutment with the first disc portion 100a and a series of circumferentially arranged bolts 107 extend through the first disc portion 100a and the radial portion 106a of the drive arm 106.

[0071] The oblique portion 106b of the drive arm 106 divides the drum cavity into radially outer rim cavity portion 108 and a radially inner main cavity portion 109.

[0072] In order to provide a pressure equalisation path from the rim cavity portion 108 to the main cavity portion 109, the oblique portion 106b of the drive arm 106 comprises a plurality of circumferentially spaced apertures 111 extending from a radially outer surface in the rim cavity portion 108 to a radially inner surface in the main cavity portion 109.

[0073] The rim cavity 108 houses a rim seal 110 located between the axial extension 104b of the second rim portion and the first rim portion 101a. The rim seal 110 comprises a folded metal sheet seal having a first leaf 110a in abutment with the first rim portion 101a and a second leaf 110b in abutment a chamfered edge of with the axial extension 104b of the second rim portion.

[0074] The rim seal 110 and the apertures 111 which provide a pressure equalisation path between the rim cavity portion 108 and the main cavity portion 109, allow the pressure within the rim cavity portion 108 to substantially match the pressure in the main cavity portion 109 rather than the pressure of the gas path (which is radially outwards of the rim portions). Accordingly, the load applied to the bolts 107 is significantly reduced because the lower pressure in the rim cavity 108 no longer imparts a high axial separation pressure load on the bolts 107. The bolts 107 only have to provide the structural connection between the two compressor stages and no longer have to provide the seal against the pressure in the gas path.

[0075] FIG. 6 shows a second embodiment which is the same as the first embodiment except that the rim seal comprises an interference seal 110 and the pressure equalisation path is provided via the aperture 111′.

[0076] Other embodiments (not shown) may combine the rim seal of the first embodiment with the pressure equalisation path of the second and vice versa.

[0077] The compressor assemblies described above are for use in a gas turbine engine such as that shown in FIG. 1 and discussed above.

[0078] Such a gas turbine engine 10 may comprise an engine core 11 comprising at least one turbine 17, 19, a combustor 16, at least one compressor 14, 15 which each comprise a compressor assembly as described above, and a core shaft 26. Such a gas turbine engine may comprise a fan 23 (having fan blades) located upstream of the engine core 11.

[0079] 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.

[0080] 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.

[0081] 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.

[0082] 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 disclosure. Practical applications of a planetary epicyclic gearbox 30 generally comprise at least three planet gears 32.

[0083] 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.

[0084] 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.

[0085] 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.

[0086] Optionally, the gearbox may drive additional and/or alternative components (e.g. the intermediate pressure compressor and/or a booster compressor).

[0087] 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 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.

[0088] 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.

[0089] 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.