DEVICE WITH TWO STRUCTURAL COMPONENTS AND GAS TURBINE ENGINE

Abstract

A device has two components with at least one of the components being rotatable relative to the other component. The components each have oil-conducting regions. The components are operatively connected to one another via an overlap region to transfer oil. The overlap region is delimited by a sealing unit which has at least one slide ring seal. The slide ring seal includes at least one recess which is operatively connected to the oil-conducting regions and which is formed to run in the direction of a sealing side, averted from said oil-conducting regions, of the slide ring seal. Via the sealing side, the slide ring seal bears sealingly against at least one of the components. An end, which faces toward the sealing side, of the recess is spaced apart from the sealing side in an axial direction of the slide ring seal.

Claims

1. A device having two components, wherein at least one of the components is designed to be rotatable relative to the other component, and the components each have oil-conducting regions, which, for the purposes of transferring oil, are operatively connected to one another via an overlap region between the components wherein the overlap region between the components is delimited by means of a sealing unit which has at least one slide ring seal, wherein the slide ring seal is designed with at least one recess which is connected to the oil-conducting regions and which is formed so as to run in the direction of a sealing side, averted from the oil-conducting regions, of the slide ring seal, by means of which sealing side the slide ring seal bears sealingly against at least one of the components, wherein an end, which faces toward the sealing side, of the recess is spaced apart from the sealing side in an axial direction of the slide ring seal.

2. The device according to claim 1, wherein, proceeding from a defined degree of wear of the slide ring seal in the region of the sealing side, proceeding from which the recess opens out in the region of the sealing side, an oil volume flow can be conducted through the recess from the oil-conducting regions in the direction of the sealing side.

3. The device according to claim 2, wherein the oil volume flow that can be conducted through the recess is greater than or equal to a predefined leakage oil volume flow.

4. The device according to claim 3, wherein the predefined leakage oil volume flow is smaller than a degree of leakage proceeding from which an oil volume flow conducted from the component into the rotatable component is smaller than a threshold value.

5. The device according to claim 1, wherein the recess is designed as a blind bore running in an axial direction in the slide ring seal.

6. The device according to claim 1, wherein the recess is formed as an axial groove which is arranged in an outer side of the slide ring seal.

7. The device according to claim 1, wherein the spacing between the end of the recess and the sealing side is configured such that an expected degree of operational wear of the slide ring seal in the region of the sealing surface over a defined operating duration, which is associated with a defined abrasive removal of material, is smaller than the wall thickness, corresponding to the spacing, of the slide ring seal between the sealing surface and the end of the recess.

8. The device according to claim 1, wherein the rotatable component radially surrounds the other component, and the slide ring seal is arranged in a radial groove of the other component, wherein the slide ring seal bears sealingly with its radial outer side against an inner side of the rotating component and bears sealingly with the axial sealing side against a wall of the radial groove.

9. The device according to claim 1, wherein the recess is connected, in the region of a sealing side situated opposite the sealing side, to the oil-conducting regions.

10. The device according to claim 1, wherein the slide ring seal is formed with multiple recesses arranged so as to be distributed over the circumference, the flow cross sections of which recesses are, in sum total, configured such that, proceeding from a defined degree of wear of the slide ring seal in the region of the sealing side, proceeding from which the recesses open out in the region of the sealing side, an oil volume flow greater than or equal to the predefined leakage oil volume flow can be conducted through the recesses from the oil-conducting regions in the direction of the sealing side.

11. A gas turbine engine for an aircraft, comprising the following: an engine core, which comprises a turbine, a compressor, and a core shaft, which connects the turbine to the compressor; a fan, which is positioned upstream of the engine core, wherein the fan comprises a plurality of fan blades; a gear box, which receives an input from the core shaft and outputs drive for the fan in order to drive the fan at a lower speed than the core shaft; and a device according to claim 1, wherein one component of the device is coupled to a rotatable shaft of the gear box and the other component of the device is operatively connected rotationally conjointly to a housing of the gas turbine engine.

12. The gas turbine engine according to claim 11, wherein the turbine is a first turbine, the compressor is a first compressor, and the core shaft is a first core shaft; the engine core further comprises a second turbine, a second compressor, and a second core shaft, which connects the second turbine to the second compressor; and the second turbine, the second compressor and the second core shaft are arranged so as to rotate at a higher speed than the first core shaft.

Description

[0058] Embodiments will now be described by way of example with reference to the figures.

[0059] In the figures:

[0060] FIG. 1 shows a longitudinal section through a gas turbine engine;

[0061] FIG. 2 shows an enlarged partial longitudinal sectional view of an upstream portion of a gas turbine engine;

[0062] FIG. 3 shows an isolated illustration of a gear box for a gas turbine engine;

[0063] FIG. 4 shows a highly simplified detail sectional view of a region IV indicated in more detail in FIG. 2;

[0064] FIG. 5 shows an illustration corresponding to FIG. 4, wherein a slide ring seal shown in FIG. 5 exhibits a higher degree of wear in the region of a sealing surface than the state shown in FIG. 4; and

[0065] FIG. 6 shows a schematic three-dimensional illustration of the slide ring seal as per FIG. 4 and FIG. 5 on its own.

[0066] FIG. 1 illustrates a gas turbine engine 10 with a primary axis of rotation 9. The engine 10 comprises an air intake 12 and a thrust fan 23 that generates two air flows: a core air flow A and a bypass air flow B. The gas turbine engine 10 comprises a core 11 which receives the core air flow A. In the sequence of axial flow, the engine core 11 comprises a low-pressure compressor 14, a high-pressure compressor 15, a combustion device 16, a high-pressure turbine 17, a low-pressure turbine 19, and a core thrust nozzle 20. An engine nacelle 21 surrounds the gas turbine engine 10 and defines a bypass duct 22 and a bypass thrust nozzle 18. The bypass air flow B flows through the bypass duct 22. The fan 23 is attached to and driven by the low pressure turbine 19 by way of a shaft 26 and an epicyclic gear box 30. In this context, the shaft 26 is also referred to as a core shaft.

[0067] During use, the core air flow 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 expelled from the high-pressure compressor 15 is directed into the combustion device 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 expelled through the nozzle 20 to provide some propulsive thrust. The high-pressure turbine 17 drives the high-pressure compressor 15 by way of a suitable connecting shaft 27, which is also referred to as the core shaft. The fan 23 generally provides the majority of the propulsion force. The epicyclic gear box 30 is a reduction gear box.

[0068] 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 gear 28 of the epicyclic gear box assembly 30. A plurality of planet gears 32, which are coupled to one another by way of a planet carrier 34, are situated radially outside the sun gear 28 and mesh with the latter, and are in each case disposed so as to be rotatable on carrier elements 29 that are connected to the planet carrier 34 for conjoint rotation therewith. The planet carrier 34 limits the planet gears 32 to orbiting around the sun gear 28 in a synchronous manner while enabling each planet gear 32 to rotate about its own axis on the carrier elements 29. The planet carrier 34 is coupled by way of linkages 36 to the fan 23 so as to drive the rotation of the latter about the engine axis 9. Radially to the outside of the planet gears 32 and meshing therewith is an annulus or ring gear 38 that is coupled, via linkages 40, to a stationary supporting structure 24.

[0069] It is noted that the terms low-pressure turbine and low-pressure compressor as used herein can be taken to mean the lowest-pressure turbine stage and the lowest-pressure compressor stage (that is to say not including the fan 23) respectively and/or the turbine and compressor stages that are connected to one another by the connecting shaft 26 with the lowest rotational speed in the engine (that is to say not including the gear box output shaft that drives the fan 23). In some documents, the low-pressure turbine and the low-pressure compressor referred to herein can alternatively be known as the intermediate-pressure turbine and intermediate-pressure compressor. Where such alternative nomenclature is used, the fan 23 can be referred to as a first compression stage or lowest-pressure compression stage.

[0070] The epicyclic gear box 30 is shown in greater detail by way of example in FIG. 3. Each of the sun gear 28, the planet gears 32 and the ring gear 38 comprise teeth about their periphery to mesh with the other gears. However, for clarity, only exemplary portions of the teeth are illustrated in FIG. 3. Although four planet gears 32 are illustrated, it will be apparent to the person skilled in the art that more or fewer planet gears 32 can be provided within the scope of protection of the claimed invention. Practical applications of an epicyclic gear box 30 generally comprise at least three planet gears 32.

[0071] The epicyclic gear box 30 illustrated by way of example in FIGS. 2 and 3 is of the planetary type, in which the planet carrier 34 is coupled to an output shaft via linkages 36, wherein the ring gear 38 is fixed. However, any other suitable type of epicyclic gear box 30 can be used. By way of a further example, the epicyclic gear box 30 can be a star arrangement, in which the planet carrier 34 is held so as to be fixed, with the ring gear (or annulus) 38 allowed to rotate. In the case of such an arrangement, the fan 23 is driven by the ring gear 38. By way of a further alternative example, the gear box 30 can be a differential gear box in which the ring gear 38 and the planet carrier 34 are both allowed to rotate.

[0072] It will be appreciated that the arrangement shown in FIGS. 2 and 3 is only exemplary, and various alternatives are within the scope of protection of the present disclosure. Purely by way of example, any suitable arrangement can be used for positioning the gear box 30 in the engine 10 and/or for connecting the gear box 30 to the engine 10. By way of a further example, the connections (such as the linkages 36, 40 in the example of FIG. 2) between the gear box 30 and other parts of the engine 10 (such as the input shaft 26, the output shaft and the fixed structure 24) can have a certain degree of stiffness or flexibility. By way of a further example, any suitable arrangement of the bearings between rotating and stationary parts of the engine (for example between the input and output shafts of the gear box and the fixed structures, such as the gear box casing) can be used, and the disclosure is not limited to the exemplary arrangement of FIG. 2. For example, where the gear box 30 has a star arrangement (described above), the person skilled in the art would readily understand that the arrangement of output and support linkages and bearing positions would usually be different than that shown by way of example in FIG. 2.

[0073] Accordingly, the present disclosure extends to a gas turbine engine having an arbitrary arrangement of gear box types (for example star-shaped or planetary), support structures, input and output shaft arrangement, and bearing positions.

[0074] Optionally, the gear box may drive additional and/or alternative components (e.g. the intermediate-pressure compressor and/or a booster compressor).

[0075] Other gas turbine engines to which the present disclosure can be applied may have alternative configurations. For example, engines of this type may have an alternative number of compressors and/or turbines and/or an alternative number of connecting 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 a dedicated nozzle that is separate from and radially outside the engine core nozzle 20. However, this is not limiting, and any aspect of the present disclosure can 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 can be referred to as a mixed-flow nozzle. One or both nozzles (whether mixed or split flow) can have a fixed or variable region. While the example described relates to a turbofan engine, the disclosure can be applied, for example, to any type of gas turbine engine, such as an open-rotor engine (in which the fan stage is not surrounded by an engine nacelle) or a turboprop engine.

[0076] The geometry of the gas turbine engine 10, and components thereof, is or are defined using a conventional axis system which comprise an axial direction (which is aligned with the axis of rotation 9), a radial direction (in the direction from bottom to top in FIG. 1), and a circumferential direction (perpendicular to the view in FIG. 1). The axial, radial and circumferential directions run so as to be mutually perpendicular.

[0077] FIG. 4 shows a highly schematic sectional view of a region IV, indicated in more detail in FIG. 2, of the gas turbine engine 10. Here, the illustration in FIG. 4 shows a device 60 having two components, wherein the first component corresponds in the present case to the planet carrier 34 and the second component corresponds to the support structure 24. The first component 34 is thus designed to be rotatable, whereas the second component 24 is fixed to a housing and is thus formed so as not to be rotatable. Components 24 and 34 each comprise oil-conducting regions 40, 41.

[0078] For the purposes of transferring oil, the oil-conducting regions 40 and 41 are operatively connected to one another via an overlap region 42 between the components 24 and 34. In the present case, the overlap region 42 is delimited by means of a sealing unit 43 which comprises two slide ring seals 44, wherein only one of the two slide ring seals 44 is shown in FIG. 4.

[0079] The second slide ring seal of the sealing unit 43 is arranged on the opposite side, in the axial direction x, of the two oil-conducting regions 40 and 41. The two slide ring seals 44 of the sealing unit 43 are each arranged in a radial groove 45 of the component 24, wherein, again, only one of the two radial grooves 45 is illustrated in FIG. 4. The rotatable component 34 surrounds the static component 24 in a radial direction. For the sealing of the oil-conducting regions 40 and 41, the slide ring seal 44 bears with its radial outer side 46 sealingly against an inner side 47 of the rotatable component 34. Additionally, the slide ring seal 44 bears with an axial sealing surface 48 sealingly against a side wall 49 of the radial groove 45.

[0080] During the operation of the gas turbine engine 10, the planet carrier 34 rotates at a high rotational speed, whereas the support structure 24 is static. Additionally, the slide ring seal 44 rotates together with the planet carrier 34, whereby a large rotational speed difference prevails in the region between the sealing surface 48 and the side wall 49 of the radial groove 45. This rotational speed difference, together with the radial offset movements between the components 24 and 34 and in a manner dependent on a friction coefficient between the sealing surface 48 and the side wall 49, gives rise to increasing wear over the course of the operating duration. In the long term, this wear leads to a reduction in the sealing performance of the sealing unit 43, as a result of which the sealing unit 43, or the slide ring seals 44 thereof, must be replaced with new slide ring seals after predefined maintenance intervals have elapsed.

[0081] Here, these maintenance intervals are configured such that the slide ring seals 44 are exchanged before the loss of sealing performance of the sealing unit 43. Since unfavorable operating state profiles of the gas turbine engine 10 can cause undesirably high levels of wear in the region of the slide ring seals 44, it is possible that the sealing performance of the sealing unit 43 decreases already before such a maintenance interval has elapsed. A supply of oil to the gear box 30 from the support structure 24 and via the planet carrier 34 is then no longer ensured to the required degree.

[0082] For this reason, the slide ring seals 44 are formed with multiple recesses 50, to the extent illustrated in more detail in FIG. 4 to FIG. 6. The recesses 50 are arranged so as to be distributed over the circumference of the slide ring seal 44 and are formed as blind bores. Furthermore, the recesses 50 extend from a sealing side 51 of the slide ring seal 44 in an axial direction x in the direction of the opposite axial sealing surface or of the sealing side 48. Here, an axial depth t50 of the recesses 50 is in each case smaller than an axial width B44 of the slide ring seal 44. The difference between the axial width B44 and the axial depth t50 corresponds to a spacing between the axial sealing side 48 and an end 52 of the recesses 50.

[0083] The spacing between the end 52 of the recesses 50 and the axial sealing side 48 is configured such that an expected degree of operational wear of the slide ring seal 44 in the region of the sealing surface 48 over a defined operating duration, which is associated with a defined abrasive removal of material in the region of the sealing surface 48, is smaller than the wall thickness, corresponding to the spacing, of the slide ring seal between the sealing surface 48 and the end 52 of the recesses 50.

[0084] In the event that the degree of wear in the region of the sealing surface 48 over the course of the operating duration is greater than the spacing between the sealing surface 48 and the end 52 of the recesses 50, the recesses 50 open out in the region of the sealing surface 48. In such a state of the slide ring seal 44, the oil-conducting regions 40 and 41 are connected via the recesses 50 to the region, facing toward the sealing surface 48, between the components 24 and 34. Thus, oil can be conducted through the recesses 50 from the oil-conducting regions 40 and 41 through the sealing unit 43. Here, the oil volume flow that can be conducted out of the oil-conducting regions 40 and 41 via the recesses 50 is so small that a supply of oil to the gear box 30 via the oil-conducting regions 40 and 41 is not impaired.

[0085] This outflowing leakage oil volume flow has the effect that the pressure in the region of the oil-conducting regions 40 and 41 decreases abruptly. This pressure drop is detected by measurement in the region of a pressure sensor 53, and is fed as a sensor signal or input signal to a control unit of the gas turbine engine 10. In the presence of such a sensor signal, the control unit outputs a corresponding warning signal to the effect that, in the region of the sealing unit 43, a defined degree of wear is present which necessitates an exchange of the slide ring seals 44.

LIST OF REFERENCE SIGNS

[0086] 9 Primary axis of rotation [0087] 10 Gas turbine engine [0088] 11 Core [0089] 12 Air intake [0090] 14 Low-pressure compressor [0091] 15 High-pressure compressor [0092] 16 Combustion installation [0093] 17 High-pressure turbine [0094] 18 Bypass thrust nozzle [0095] 19 Low-pressure turbine [0096] 20 Core thrust nozzle [0097] 21 Engine nacelle [0098] 22 Bypass duct [0099] 23 Thrust fan [0100] 24 Support structure [0101] 26 Shaft, connecting shaft [0102] 27 Connecting shaft [0103] 28 Sun gear [0104] 30 Gear box, planetary gear box [0105] 32 Planet gear [0106] 34 Planet carrier [0107] 36 Linkage [0108] 38 Ring gear [0109] 40, 41 Oil-conducting region [0110] 42 Overlap region [0111] 43 Sealing unit [0112] 44 Slide ring seal [0113] 45 Radial groove [0114] 46 Radial outer side of the slide ring seal [0115] 47 Inner side of the planet carrier 34 [0116] 48 Axial sealing surface of the slide ring seal [0117] 49 Side wall [0118] 50 Recess [0119] 51 Sealing side [0120] 52 End of the recess [0121] 53 Pressure sensor [0122] 60 Device [0123] A Core air flow [0124] B Bypass air flow [0125] B44 Axial width of the slide ring seal [0126] t50 Axial depth of the recess [0127] x Axial direction of the sealing unit