STABILIZATION BEARING SYSTEM FOR GEARED TURBOFAN ENGINES

Abstract

The invention relates to a stabilization bearing system for a geared turbofan engine, in particular an aircraft engine, with a stabilization bearing for an input shaft device to an epicyclic gearbox device of the geared turbofan engine, the stabilization bearing being located axially in front of an input shaft device location bearing system.

Claims

1. A stabilization bearing system for a geared turbofan engine, in particular an aircraft engine, with a stabilization bearing for an input shaft device to an epicyclic gearbox device of the geared turbofan engine, the stabilization bearing being located axially in front of an input shaft device location bearing system.

2. A stabilization bearing system according to claim 1, with location bearing systems and the stabilization bearing being statically indeterminate.

3. A stabilization bearing system according to claim 1, wherein the axial distance between the input shaft device location bearing system and the stabilization bearing is at least greater than the smallest diameter of input shaft device.

4. A stabilization bearing system according to claim 1, wherein the axial distance between the input shaft device location bearing system and the stabilization bearing is 1 to 8 times, in particular 2 to 6 times the mean diameter of the input shaft device.

5. A stabilization bearing system according to claim 1, wherein the load path of the stabilization bearing is different from the load paths of the locations bearing systems.

6. A stabilization bearing system according to claim 1, wherein the axial distance of the stabilization bearing from the centerline of the epicyclic gearbox device is 2 to 4 times the mean diameter of the input shaft device at the epicyclic gearbox device.

7. A stabilization bearing system according to claim 1, wherein the stabilization bearing comprises a cylindrical roller bearing.

8. A stabilization bearing system according to claim 1, wherein the stabilization bearing is taking only a radial load.

9. A stabilization bearing system according to claim 1, wherein the ratio of the radial load applied by the shaft to stabilization bearing and the radial load applied input shaft location bearing system of the input shaft is in the range between 0.3 and 4.5, in particular 0.5 to 3.

10. 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, with a stabilization bearing system according to claim 1.

Description

[0041] Embodiments will now be described by way of example only, with reference to the Figures, in which:

[0042] FIG. 1 is a sectional side view of a gas turbine engine;

[0043] FIG. 2 is a close up sectional side view of an upstream portion of a gas turbine engine;

[0044] FIG. 3 is a partially cut-away view of a gearbox for a gas turbine engine;

[0045] FIG. 4 shows schematically the upper half of a front section of an embodiment of a gas turbine with a drive train with an input shaft device of an epicyclic gearbox device, a gearbox device and an output shaft device extending to a propulsive fan and an embodiment of a stabilization bearing.

[0046] 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 device 30.

[0047] 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 device 30 is a reduction gearbox.

[0048] 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.

[0049] 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.

[0050] The epicyclic gearbox device 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 epicyclic gearbox device 30 generally comprise at least three planet gears 32.

[0051] The epicyclic gearbox device 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 device 30 may be used. By way of further example, the epicyclic gearbox device 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 device 30 may be a differential gearbox in which the ring gear 38 and the planet carrier 34 are both allowed to rotate.

[0052] 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 device 30 in the engine 10 and/or for connecting the gearbox device 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 device 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 device 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.

[0053] 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.

[0054] Optionally, the gearbox may drive additional and/or alternative components (e.g. the intermediate pressure compressor and/or a booster compressor).

[0055] 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 device 30.

[0056] 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.

[0057] In FIG. 4 a schematic view of the front section of geared turbofan engine 10 is shown. The view axially extends from the propulsive fan 23 in the front to the low-pressure compressor 14 towards the rear. The low-pressure compressor 14 is only shown symbolically to indicate the relative position of the drive train and its units.

[0058] The drive train comprises an input shaft device 50 (e.g. comprising the shaft 26 shown in FIG. 1), here driven by the not shown low-pressure turbine 19. The input shaft device 50 is driveably connected to the sun gear 28 of the epicyclical gearbox device 30.

[0059] The output of the gearbox device 30 in the embodiment shown takes place via the planet carrier 34 which is connected with an output shaft device 60 which has a portion acting as a fan shaft 61. That portion is rigidly connected with the propulsive fan 23.

[0060] Therefore, the input torque is transmitted from the input shaft device 50 to the sun gear 28 of the gearbox device 30, and to some extent to the ring gear mount. The planet carrier 34 transmits the output torque (at a reduced rotational speed) to the output gear device 60 and eventually to the propulsive fan 23.

[0061] The input shaft device 50 and the output shaft device 60 are here shown in a simplified manner. It is possible that the shape of the shaft devices 50, 60 can be more complex and comprises more than one piece.

[0062] The shafting arrangement of the embodiment shown in FIG. 4 also comprises several bearing systems 70, 72, 75, e.g. for taking the mechanical loads or for locating the propulsive fan 23 and the gearbox device 30.

[0063] The input shaft device 50 is supported by input shaft location bearing system 70 and an additional radial support bearing (not shown in FIG. 4) located in the turbine section of the engine 10 which are required to locate the input shaft device 50, i.e. to obtain a statically determined system.

[0064] In addition, an input shaft device stabilization bearing 75 is axially located in front of the input shaft location bearing system 70. With three bearing systems 70, 72, 75 the supporting of the input shaft device 50 becomes statically indetermined.

[0065] The bearing span, i.e. the axial distance D between the input shaft device stabilization bearing 75 and the input shaft location bearing system 70 is larger than the smallest diameter of the input shaft device 50 or 1 to 8 times the mean diameter of the input shaft device 50. The mean diameter of the input shaft device 50 is used since the diameter of the input shaft device 50 varies along its axis.

[0066] The axial distance D is measured between the centrelines of the bearing systems 70, 75.

[0067] The stabilization bearing 75 is mounted to a static structure of a front compressor case, e.g. the stationary support structure 24. The input shaft location bearing system 70 is mounted to the case of the low-pressure compressor 14.

[0068] On the output side of the gearbox device 30, the output shaft device 60 comprises a fan shaft bearing system 80 and the axial location bearing system 72 (ball bearing).

[0069] The radial inner seat of the fan shaft bearing system 80 is on the fan shaft 61, being a part of the output shaft device 60. The radial outer seat of the fan shaft bearing system 80 is connected to a static front cone structure 81. In the embodiment shown a roller bearing is used in the fan shaft bearing system 80. In alternative embodiments, more than one roller bearing (e.g. double bearings, two bearings of different design) or other bearing designs can be used. It would be possible to install a ball bearing and transfer the axial load to the fan 13 via the static front cone structure 81. The output shaft device 60 in the embodiment shown in FIG. 4 comprises essentially a cylindrical section adjacent to the output side of the gearbox device 30 and under the propulsive fan 23 (i.e. the fan shaft section 61). In-between there is a conical section 62 linking the two cylindrical sections. Conical in this context means that the axial cross-section in this part of the output shaft device 60 is a straight line inclined radially inwards. In other embodiments this linking section can have different shapes than the conic shape in FIG. 4.

[0070] In the embodiment shown in FIG. 4 the static front cone structure 81 and a static structure 90 form together one cavity around the gearbox device 30. The already mentioned sealing device secures the oil within that cavity. In an alternative design, the static structure 91 and the sealing device are combined in one device.

[0071] The ring gear 38 is rigidly connected to the static front cone structure 81 but alternatively, it can be connected to a different static part within the engine 10.

[0072] 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.

LIST OF REFERENCE NUMBERS

[0073] 9 principal rotational axis [0074] 10 gas turbine engine [0075] 11 engine core [0076] 12 air intake [0077] 14 low-pressure compressor [0078] 15 high-pressure compressor [0079] 16 combustion equipment [0080] 17 high-pressure turbine [0081] 18 bypass exhaust nozzle [0082] 19 low-pressure turbine [0083] 20 core exhaust nozzle [0084] 21 nacelle [0085] 22 bypass duct [0086] 23 propulsive fan [0087] 24 stationary support structure [0088] 26 shaft [0089] 27 interconnecting shaft [0090] 28 sun gear [0091] 30 epicyclic gearbox device [0092] 32 planet gears [0093] 34 planet carrier [0094] 36 linkages [0095] 38 ring gear [0096] 40 linkages [0097] 41 centreline gearbox [0098] 50 input shaft device (sun shaft) [0099] 60 output shaft device [0100] 61 fan shaft [0101] 62 conical section [0102] 70 input shaft rear bearing system [0103] 71 static rear structure [0104] 72 axial location bearing system [0105] 73 radial support bearing [0106] 75 input shaft stabilization bearing [0107] 80 fan shaft bearing system [0108] 81 static front cone structure [0109] 90 static structure [0110] A core airflow [0111] B bypass airflow [0112] D bearing span