GAS TURBINE

Abstract

The invention relates to a gas turbine engine, in particular an aircraft engine, comprising: a turbine connected via an input shaft device to a gearbox device having a sun gear, a planet carrier having a plurality of planet gears attached thereto, and a ring gear, the sun gear is connected to the input shaft device, the planet carrier or the ring gear is connected to a propulsive fan via an output shaft device of the gearbox device, with an inter-shaft bearing system being positioned radially between the input shaft device and the planet carrier of the gearbox device.

Claims

1. A gas turbine engine, in particular an aircraft engine, comprising: a turbine connected via an input shaft device to a gearbox device having a sun gear, a planet carrier having a plurality of planet gears attached thereto, and a ring gear, the sun gear is connected to the input shaft device, the planet carrier or the ring gear is connected to a propulsive fan via an output shaft device of the gearbox device, with an inter-shaft bearing system being positioned radially between the input shaft device and the planet carrier of the gearbox device, with a carrier bearing system being located radially between the input shaft device and a static structure supporting the carrier bearing system, the support connection being axially in front of the input side of the gearbox device.

2. The gas turbine of claim 1, wherein the inter-shaft bearing system is located axially within or in front of a low-pressure compressor or an intermediate compressor.

3. The gas turbine of claim 1, wherein the inter-shaft bearing system is axially adjacent to the gearbox device on the input and/or the output side, in particular with an axial distance measured from the centreline of the gearbox between 0.001 and 4 times the inner radius of the inter-shaft bearing system.

4. The gas turbine of claim 1, wherein the inter-shaft bearing device comprises at least one ball bearing.

5. The gas turbine of claim 1, wherein a fan shaft bearing system is radially located between a fan shaft as part of the output shaft device and a static structure, in particular a static front cone structure, in particular the fan shaft bearing system being axially positioned within the width of the propulsive fan or axially behind it.

6. The gas turbine of claim 5, wherein the fan shaft bearing system has an outer diameter between 0.05 to 0.35 the diameter of the propulsive fan, in particular between 0.1 and 0.3 times the diameter of the propulsive fan.

7. The gas turbine of claim 1, wherein the carrier bearing system comprises at least one roller bearing.

8. The gas turbine of claim 7, wherein the carrier bearing system is axially adjacent to the gearbox device on the input side or output side, in particular with an axial distance measured from the centreline of the gearbox between 0.1 and 4 times the inner radius of the inter-shaft bearing system.

9. The gas turbine of claim 1, wherein in the planet carrier comprises a seat element extending axially to the front and/or the rear of the gearbox device providing a radial seat for the inter-shaft bearing system and/or the carrier bearing system.

10. The gas turbine of claim 1, wherein the inter-shaft bearing system and the carrier bearing system are essentially located in one vertical plane or have an axial offset between 0.1 and 4 times the inner radius of the inter-shaft bearing system.

11. The gas turbine of claim 1, wherein an input shaft bearing system is radially located between the input shaft device and a static structure, in particular a static rear structure, the input shaft bearing system in particular comprising at least one roller bearing.

12. The gas turbine of claim 1, wherein the output shaft device comprises at least one axial cross-section with a conical, sigmoidal or logarithmical shape.

13. The gas turbine of claim 1, wherein the output shaft device comprises a curvic or a spline coupling.

14. The gas turbine of claim 1, wherein the load path for force and/or torque from the driving turbine to the propulsive fan exclusively extends via the input shaft device, the gearbox device and the output shaft device, in particular without a through shaft extending through the gearbox device.

15. The gas turbine of claim 1, wherein the ring gear is rigidly connected to the static front cone structure.

16. The gas turbine of claim 1, wherein the gearbox device comprises an epicyclic gearbox with the ring gear being fixed relative to the other parts of the gearbox device and the output shaft device being connected to the planet carrier.

17. The gas turbine of claim 1, wherein the gearbox device comprises a planetary gearbox in star arrangement with the planet carrier fixed relative to the other parts of the gearbox device and the output shaft device being connected to the ring gear.

18. The gas turbine of claim 1, wherein the input shaft device comprises a bent section with folded back part of the shaft and/or a corrugated section with corrugations around the circumference of the shaft.

Description

[0067] Embodiments will now be described by way of example only, with reference to the Figures, in which:

[0068] FIG. 1 is a sectional side view of a gas turbine engine;

[0069] FIG. 2 is a close up sectional side view of an upstream portion of a gas turbine engine;

[0070] FIG. 3 is a partially cut-away view of a gearbox for a gas turbine engine;

[0071] 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;

[0072] FIG. 4a shows different arrangement of an inter-shaft bearing and a rear carrier bearing;

[0073] FIG. 5 shows a variation of the embodiment shown in FIG. 4;

[0074] FIG. 6 shows a variation of the embodiment shown in FIG. 5;

[0075] FIG. 7 shows a variation of the embodiment shown in FIG. 6;

[0076] FIG. 8 shows a schematic view of an embodiment indicating certain dimensions.

[0077] 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.

[0078] 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.

[0079] 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.

[0080] 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.

[0081] 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.

[0082] 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. In another embodiment the carrier and the output shaft can be manufactured as one part. 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.

[0083] 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.

[0084] 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.

[0085] Optionally, the gearbox may drive additional and/or alternative components (e.g. the intermediate pressure compressor and/or a booster compressor).

[0086] 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.

[0087] 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.

[0088] 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.

[0089] 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 connected to the sun gear 28 of the epicyclical gearbox device 30. The output of the gearbox device 30 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. In an alternative embodiment, the output shaft 60 can be replaced by a direct connection of the fan disk 61 to the carrier 34.

[0090] 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.

[0091] 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.

[0092] The shafting arrangement of the embodiment shown in FIG. 4 also comprises several bearing systems, e.g. for taking the mechanical loads or for locating the propulsive fan 23 and the gearbox 30.

[0093] The first bearing to be described is an inter-shaft bearing system 70 being positioned radially between the input shaft device 50 and the planet carrier 34. This inter-shaft bearing system 70 here comprises one ball bearing. In alternative embodiments, more than one ball bearing (e.g. double bearings, two bearings of different design) or other bearing designs can be used. It is also possible that different bearings of the inter-shaft bearing system 70 are positioned at different locations.

[0094] The inter-shaft bearing system 70 is, in the embodiment shown in FIG. 4, located axially in front of the low-pressure compressor 14. In other embodiments, the inter-shaft bearing system 70 can be located axially within the low-pressure compressor 14 to support the compressor tip clearance and to ensure appropriate engine performance.

[0095] The inter-shaft bearing system 70 is, in this embodiment, axially adjacent to the gearbox device 30 on the input side. The axial distance between the inter-shaft bearing system 70 to the gearbox device 30 can e.g. be between 0.001 and 4 times the inner radius of the inter-shaft bearing system 70. This could be in the range of 1 to 100 mm measured from the axial front side of the inter-shaft bearing system 70 to a centreline 31 of the gearbox device 30.

[0096] The fan axial load is transferred via the fan-shaft bearing system 80 (roller bearing), via the gearbox device 30 and into the input-shaft bearing 95 towards the rear. With this arrangement the support structures of the bearings can be reduced.

[0097] The similar load path would apply when the inter-shaft bearing system 70 would comprise a roller bearing and the carrier bearing system 90 would comprise a ball bearing (i.e. inverse situation to the embodiment of FIG. 4). On the input side of the gearbox device 30 the load would be transferred radially outwards towards the ESS and the input shaft bearing system 95. The inter-shaft bearing system 70 would only be supported radially.

[0098] If both, the inter-shaft bearing system 70 and the carrier bearing system 90, would comprise roller bearings, the fan axial load would be transferred via the static front cone section 81 and the ESS. In this case the gearbox device would not carry an axial load.

[0099] This inter-shaft bearing system 70 locates the propulsive fan 23 and transmits axial loads towards a further bearing system towards the rear of the gearbox device 30, the input shaft bearing system 95. This bearing system is radially located between the input shaft device 50 and a static rear structure 96. In the embodiment shown here, the input shaft bearing system 95 comprises at least one ball bearing. In alternative embodiments, more than one roller bearing (e.g. double bearings, two bearings of different design) or other bearing designs can be used. In a further alternative, it would be possible to transfer the fan axial load via the bearing 90 and the rear cone structure. The inter-shaft bearing 70 would then transfer only radial load to control the gears relative displacements.

[0100] At the input side of the gearbox device 30 a further bearing system, the carrier bearing system 90, is located; the carrier bearing system 90 in this case could also be rear carrier bearing system 90. The carrier bearing system 90 is positioned axially to the rear of the inter-shaft bearing 70.

[0101] It is possible that there is an axial offset between 0.1 and 4 times the inner radius of the inter-shaft bearing system 70.

[0102] In other embodiments, the inter-shaft bearing system 70 and the carrier bearing system 90 can be positioned in the same plane.

[0103] The radial seat on the inner diameter is a structure coupled to the planet carrier 34 or the planet carrier 34 itself, such as the seat element 39 axially extending into the rear part of the engine 10. The radial seat of the carrier bearing system 90 is connected to a static cone structure 91 at the support connection 92.

[0104] The axial distance between the carrier bearing system 90 measured from the centreline 41 of the gearbox device 30, can be between 0.1 and 4 times the inner radius of the inter-shaft bearing system 70. That can be between 1 mm and 400 mm.

[0105] The support connection 92 is positioned axially in front of the input side of the gearbox 30 device. This results a compact overall structure.

[0106] On the output side of the gearbox device 30, the output shaft device 60 only has one bearing system, a fan shaft bearing system 80. The radial inner seat of that bearing system 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.

[0107] In the embodiment described herein, the fan shaft bearing system 80 can have an outer diameter between 0.05 to 0.35 times the diameter of the propulsive fan 13. This range can be between 175 and 1250 mm.

[0108] In an alternative embodiment (e.g. FIG. 5), the fan shaft bearing system 80 is not directly located underneath the propulsive fan 23. The fan shaft 61 then extends to the front from the fan shaft bearing system 80. The fan shaft 61 connects directly to the carrier 34.

[0109] 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.

[0110] In the embodiment shown in FIG. 4 the static front cone structure 81 and the static structure 91 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.

[0111] 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.

[0112] The shafting arrangement described in connection with FIG. 4 allows a control of the dynamic modes of the power gearbox device 30, a control of relative deflections between the planet carrier 34 and the rotor of the low-pressure compressor 14 and deflections of the gearbox device 30 itself.

[0113] In the embodiment shown in FIG. 4 there is no shaft extending from the rear part of the engine 10 axially beyond the gearbox device 30 to the front. A through shaft, as it is known from the prior art, is not present. The load path for force and/or torque from the driving turbine 19, i.e. the low-pressure turbine 19 to the propulsive fan 23, exclusively extends via the input shaft device 50, the gearbox device 30 and the output shaft device 60. The omission of the through shaft saves weight, costs and complexity. Furthermore, the number of bearings is reduced and overconstraining the low pressure rotor is avoided.

[0114] In the embodiment of FIG. 4, the input shaft device 50 is deliberately made flexible. It comprises a bent section 51, i.e. a section where the tubular input shaft is folded towards itself by 180. And it comprises a corrugated section 52 with corrugations around the circumference of the shaft. The corrugations can be in a plane perpendicular to the axis 9 of the input shaft device 50.

[0115] The input shaft device 50 can also be designed to be so stiff (see FIGS. 6 and 7), that there is only a small tilt of a compressor 14, 15. The limitation of the tilting behaviour improves the tip clearance in the compressor 14, 15 and thereby improving the overall efficiency of the engine 10.

[0116] In FIG. 4A the embodiment shown in FIG. 4 is modified in that the inter-shaft bearing system 70 and the carrier bearing system 90 are positioned essentially in one plane on the input side of the gearbox device 30. Furthermore, the static structure 91 is connected to the inner wall of the bypass duct 22.

[0117] In FIG. 5, a variation of the embodiment shown in FIG. 4 or 4a is shown. Reference can be made to the respective description.

[0118] Like in the other embodiment, the input shaft device 50 is a flexible structure and the fan shaft 61 is a rigid structure. The support connection 92 of the carrier bearing system is positioned axially in front of the input side of the gearbox device 30.

[0119] This embodiment differs e.g. in the connection of the output shaft device 60 to the propulsive fan 23. The fan shaft bearing system 80 is located in a plane which is axially behind the propulsive fan 23. Compared to the embodiment in FIG. 5, the output shaft device 60 is shorter and essentially parallel to the rotational axis 9. The fan shaft bearing system 80 comprises a roller bearing and is made part of the static front cone section 81.

[0120] In FIG. 6, a variation of the embodiment shown in FIG. 5 is shown, i.e. the fan shaft bearing system 80 is arranged in the same way. The input shaft device 50 is made more rigid by neither comprising a bent section 51 nor a corrugated section 52.

[0121] The embodiment shown in FIG. 7 is based on the one depicted in FIG. 6, i.e. the input shaft device 50 is relatively rigid. It is different in that the static structure 91 is connected to a wall of the bypass duct 22. In the direction of the static structure 91 a first strut 93 is extending across the bypass duct 22, the direction of the static structure 91 and the first strut 93 being essentially collinear (e.g. less than 10 deviation from a straight line). From the first strut 93 the load can be further distributed into other internal structures 94a which are here shown only symbolically.

[0122] The static front cone structure 81 is connected to the wall of the bypass duct 22 where a second strut 94 extending across the bypass duct 22 has a connection. Here, the front cone structure 82 and the second strut 94 are essentially collinear (e.g. less than 10 deviation from a straight line). The load from the static front cone structure 81 can be directed towards other parts of the gas turbine engine 10.

[0123] The positions of the front cone structure 81, the static structure 91 and/or the struts 93, 94 relative to each other can be used in connection with the other embodiments as well.

[0124] The struts 93, 94 and the vanes can be coupled with a de-icing device 97 which can e.g. be fed by hot scavenge oil from the gearbox device 30. In FIG. 7, the oil feed lines are indicated by dashed lines.

[0125] In FIG. 8 the geometric relationships between the fan, the fan shaft and the some of the bearing systems in an embodiment are shown.

[0126] As a difference to the embodiment shown in FIGS. 4 to 7, there is a front carrier bearing system 98 present at the output side of the gearbox device. It should be noted that the schematic embodiment of FIG. 8 is otherwise a further embodiment in addition to the ones discussed above.

[0127] The axial distance between the centreline of the propulsive fan 23 and the fan shaft bearing system 81 (also measured from its centreline) is denoted as A.

[0128] The axial distance between the centrelines of the fan shaft bearing system 81 and the inter-shaft bearing system 70 is denoted as B.

[0129] The axial distance between the centreline of the propulsive fan 23 and the front carrier bearing system 98 (also measured at the centreline) is denoted as C.

[0130] The axial distance between the centreline of the front carrier bearing system 98 to the rear carrier system 90 is denoted as D.

[0131] Certain geometric relationships are found to be particularly effective.

[0132] Ratio of A/Diameter of propulsive fan: 0.02 to 0.15, in particular 0.04 to 0.1, more in particular 0.05 to 0.08.

[0133] Ratio of B/Diameter of propulsive fan: 0.5 to 0.9, in particular 0.6 to 0.8, more in particular 0.6 to 0.64.

[0134] Ratio of C/Diameter of propulsive fan: 0.15 to 0.4, in particular 0.2 to 0.35.

[0135] Ratio of D/Diameter of propulsive fan: 0.2 to 0.4, in particular 0.2 to 0.5.

[0136] It is understood that all these ratios can be applied to all embodiments described herein, as far as applicable (e.g. C and D not present in all embodiments).

[0137] 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

[0138] A core airflow [0139] B bypass airflow [0140] 9 principal rotational axis [0141] 10 gas turbine engine [0142] 11 engine core [0143] 12 air intake [0144] 14 low-pressure compressor [0145] 15 high-pressure compressor [0146] 16 combustion equipment [0147] 17 high-pressure turbine [0148] 18 bypass exhaust nozzle [0149] 19 low-pressure turbine [0150] 20 core exhaust nozzle [0151] 21 nacelle [0152] 22 bypass duct [0153] 23 propulsive fan [0154] 24 stationary support structure [0155] 26 shaft [0156] 27 interconnecting shaft [0157] 28 sun gear [0158] 30 gearbox (gearbox device) [0159] 32 planet gears [0160] 34 planet carrier [0161] 36 linkages [0162] 38 ring gear [0163] 39 seat element [0164] 40 linkages [0165] 41 centreline gearbox [0166] 50 input shaft device (sun shaft) [0167] 51 bent section in input shaft [0168] 52 corrugated section in input shaft [0169] 60 output shaft device [0170] 61 fan shaft [0171] 62 conical section [0172] 63 gearbox device oil feed [0173] 70 inter-shaft bearing system [0174] 70 second part of the inter-shaft bearing system [0175] 80 fan shaft bearing system [0176] 81 static front cone structure [0177] 90 carrier bearing system [0178] 91 static structure [0179] 92 support connection [0180] 95 input shaft bearing system [0181] 96 static rear structure [0182] 97 de-icing device