GAS TURBINE ENGINE AND AIRCRAFT WITH A GAS TURBINE ENGINE

Abstract

A gas turbine engine including an engine core and including a bypass channel of an aircraft is described. The bypass channel radially surrounds the engine core at least in part. At least one core shaft extending in the axial direction is provided, which shaft is operatively connected, by means of a drive train, to a core accessory gearbox arranged between the engine core and the bypass channel, and to an aircraft accessory gearbox. The aircraft accessory gearbox is arranged radially outside the bypass channel. The drive train extends substantially in the radial direction between the core shaft and the aircraft accessory gearbox. The drive train has an angle drive between a radial shaft and the core shaft. Furthermore, the drive train comprises another angle drive between the radial shaft and the core accessory gearbox and an additional angle drive between the radial shaft and the aircraft accessory gearbox.

Claims

1. A gas turbine engine, comprising an engine core and comprising a bypass channel which radially surrounds the engine core at least in part, wherein at least one core shaft extending in the axial direction is provided, wherein the core shaft is operatively connected to a core accessory gearbox arranged between the engine core and the bypass channel and to an aircraft accessory gearbox by means of a drive train, wherein the aircraft accessory gearbox is arranged radially outside the bypass channel, wherein the drive train extends substantially in the radial direction between the core shaft and the aircraft accessory gearbox, wherein the drive train has an angle drive between a radial shaft and the core shaft, and wherein the drive train comprises another angle drive between the radial shaft and the core accessory gearbox and an additional angle drive between the radial shaft and the aircraft accessory gearbox.

2. The gas turbine engine according to claim 1, wherein the radial shaft is formed as a single piece.

3. The gas turbine engine according to claim 1, wherein the radial shaft has at least two radial shaft portions which are arranged coaxially with one another and one behind the other in the radial direction at least in some regions and are interconnected for conjoint rotation.

4. The gas turbine engine according to claim 1, wherein engine accessories are substantially operatively connected to the drive train by means of the core accessory gearbox, which engine accessories are provided to carry out functions of the gas turbine engine and are arranged between the engine core and the bypass channel.

5. The gas turbine engine according to claim 1, wherein aircraft accessories are substantially operatively connected to the drive train by means of the aircraft accessory gearbox, which aircraft accessories are provided to carry out functions of an aircraft which is configured with the gas turbine engine, the aircraft accessories being arranged radially outside the bypass channel.

6. The gas turbine engine according to claim 1, wherein the core accessory gearbox and the aircraft accessory gearbox overlap at least in regions of connection to the drive train in the circumferential direction of the gas turbine engine.

7. The gas turbine engine according to claim 1, wherein the aircraft accessory gearbox is arranged in the circumferential direction at least in part in a region of the gas turbine engine in which the gas turbine engine comprises means which are designed to connect the gas turbine engine to an aircraft, the region of the gas turbine engine overlapping a fuselage of the aircraft when the engine is installed on the aircraft.

8. The gas turbine engine according to claim 1, wherein the core accessory gearbox is designed to transmit a torque and to drive an oil pump, a fuel pump, an air/oil separator and/or a permanent magnet alternator.

9. The gas turbine engine according to claim 1, wherein the aircraft accessory gearbox is designed to transmit a torque and to drive a pneumatic air turbine starter, a hydraulic pump and/or an alternator.

10. The gas turbine engine according to claim 1, wherein the aircraft accessory gearbox is designed to transmit a torque and to drive an electric machine, the electric machine being arranged radially outside the bypass channel and designed, during motor operation, to start the gas turbine engine and to generate electrical energy during alternator operation.

11. The gas turbine engine according to claim 1, wherein an electrically operable hydraulic pump is connected to the electric machine radially outside the bypass channel and can receive electrical energy from this machine.

12. The gas turbine engine according to claim 1, wherein the core accessory gearbox is arranged in the axial direction in the region of a rear side of a front engine frame between the engine core and the bypass channel.

13. The gas turbine engine according to claim 1, wherein the aircraft accessory gearbox comprises a separate oil system which is designed to supply the aircraft accessory gearbox with oil.

14. An aircraft, comprising at least one gas turbine engine according to claim 1, wherein the gas turbine engine is arranged on the fuselage or in the fuselage of the aircraft, and wherein the aircraft accessory gearbox and the aircraft accessories which are operatively connected thereto are arranged radially outside the bypass channel in a region of overlap between the gas turbine engine and the fuselage.

15. The aircraft according to claim 14, wherein the aircraft accessory gearbox and the aircraft accessories which are operatively connected thereto are arranged radially in an engine nacelle, in part radially in the engine nacelle and in part radially outside the engine nacelle in an engine pylon and/or the fuselage or radially outside the engine nacelle in the engine pylon and/or in the fuselage.

16. The aircraft according to claim 14, wherein the electric machine is arranged so as to be radially aligned with the radial shaft in the engine nacelle, in the engine pylon and/or in the fuselage.

17. The aircraft according to claim 14, wherein in each case one gas turbine engine is provided at least on both sides of the fuselage, the aircraft accessory gearbox and the core accessory gearbox each being arranged in the circumferential direction of the gas turbine engines in such a way that the core accessory gearbox and engine accessories operatively connected thereto is shielded against components which, in the event of damage, escape from a gas turbine engine which is arranged on the opposite side of the fuselage with correspondingly high kinetic energy towards the other gas turbine engine by the aircraft accessory gearbox and aircraft accessories connected thereto.

Description

[0058] It is self-evident to a person skilled in the art that a feature or parameter described in relation to one of the above aspects can be applied to any other aspect, unless they are mutually exclusive. Furthermore, any feature or any parameter described here may be applied to any aspect and/or combined with any other feature or parameter described, unless they are mutually exclusive. Further advantages and advantageous developments of the invention can be found in the claims and the exemplary embodiments described based on the concept with reference to the drawings, in which:

[0059] FIG. 1 is a simplified three-dimensional view of an aircraft with gas turbine engines arranged in the rear region on an aircraft fuselage;

[0060] FIG. 2 is a simplified longitudinal sectional view of a gas turbine engine of the aircraft according to FIG. 1;

[0061] FIG. 3 is a simplified cross-sectional view of the gas turbine engine according to FIG. 2; and

[0062] FIG. 4 is an illustration corresponding to that of FIG. 3 of another embodiment of the gas turbine engine according to FIG. 2.

[0063] FIG. 1 shows an aircraft or a passenger aircraft 1 which has three gas turbine engines 2, 3, 4. The first gas turbine engine 2 is arranged on a left-hand side of the aircraft in the rear region of the aircraft 1, in the region of a vertical stabilizer 6, and is attached in the region of an engine pylon 7 to a fuselage 8 of the aircraft 1. The second gas turbine engine 3 is connected to the fuselage 8 substantially mirror-symmetrically on a right-hand side of the aircraft.

[0064] The third gas turbine engine 4 is positioned at the rear end of the fuselage 8 and is attached to an inner fuselage strut, which is arranged below the vertical stabilizer 6 of the aircraft 1. An air inlet 10 is provided to supply air to the third gas turbine engine 4. The air inlet 10 is arranged, in front of the vertical stabilizer 6 in a direction of flight, on a top side of the fuselage 8 and is connected, within the aircraft fuselage 8, to the third gas turbine engine 4.

[0065] FIG. 2 shows the gas turbine engine 2 of the aircraft 1 according to FIG. 1 in a simplified longitudinal sectional view. The gas turbine engine 2 comprises a subsidiary flow channel or bypass channel 11 and an inlet region 12. Downstream of the inlet region 12, a blower 13 is connected in a manner which is known per se.

[0066] After the blower 13, the fluid flow in the gas turbine engine 2 is divided into a bypass flow and a core flow. The bypass flow flows through the bypass channel 11, whereas the core flow flows into an engine core 14. The engine core 14 is configured with a compressor device 15, with a burner 16, with a low-pressure turbine 17 which is provided to drive the blower 13, and with a high-pressure turbine 18 provided to drive the compressor device 15.

[0067] In addition, FIG. 2 is a schematic view of a core accessory gearbox 19 which is arranged substantially in the region of an intermediate casing 20 of the gas turbine engine 2. The intermediate casing 20 is located in the radial direction R of the gas turbine engine 2 in a region between the engine core 14 and the bypass channel 11. Furthermore, an aircraft accessory gearbox 21 is arranged radially outside the bypass channel 11. The core accessory gearbox 19 and the aircraft accessory gearbox 21 are driven by a radial shaft 22 of a drive train 9 which is operatively connected to a core shaft 24 of the gas turbine engine 2 which core shaft extends in the axial direction A of the gas turbine engine 2. The radial shaft 22 is connected to the core shaft 24 by means of an angle drive 5. In the present case, the core shaft 24 is a high-pressure shaft of the gas turbine engine 2 which, in the operation of the gas turbine engine 2, rotates at a higher speed than another core shaft 23 arranged coaxially therewith which is what is known as a low-pressure shaft.

[0068] Starting from the core shaft 24, the radial shaft 22 extends substantially in the radial direction R of the gas turbine engine 2 through what is known as an inner strut 25, that is to say a strut formed with a hollow profile or an aerodynamic profile formed with a hollow profile, outwards through the engine core 14 to the intermediate casing 20. In the region of the intermediate casing 20, the radial shaft 22 interacts with a drive shaft 27 by means of another angle drive 26 in the form of a bevel gear set.

[0069] By means of gear pairs 30 of the core accessory gearbox 19, which in the present case are in the form of spur-gear stages, the drive shaft 27 is connected to what are known as engine accessories 28. In the present case, the engine accessories 28 are an air/oil separator, an oil pump, a fuel pump, a permanent-magnet alternator and other accessories which are provided for the operation of the gas turbine engine 2. In addition, in the intermediate casing 20, an oil tank and an oil cooler which can be temperature-controlled by fuel are also arranged radially inside the bypass channel 11 in the gas turbine engine 2.

[0070] The aircraft accessory gearbox 21 is arranged radially in an engine nacelle 29 which is delimited radially outwardly by an outer face of the engine nacelle 29 and radially inwardly by an outer face 31 of the bypass channel 11. An additional angle drive 40 is provided between the radial shaft 22 and the aircraft accessory gearbox 21. The radial shaft 22 extends through an outer strut 45, that is to say a strut formed with a hollow profile, or an aerodynamic profile formed with a hollow profile, through the bypass channel 11.

[0071] FIG. 3 is a simplified cross-sectional view of a first embodiment of the gas turbine engine 2 according to FIG. 2, in which the aircraft accessory gearbox 21 is arranged radially outside the engine nacelle 29 in the engine pylon 7. In the present case, a mechanically drivable hydraulic pump 32, a gas turbine engine starter 33 (air turbine starter, ATS) and an alternator 34 are what are known as aircraft accessories.

[0072] In the exemplary embodiment of the gas turbine engine 2 shown in FIG. 3, the core accessory gearbox 19 is arranged together with the engine accessories radially between the bypass channel 11 and the engine core 14. In the present case, the engine accessories are inter alia the previously mentioned fuel pump 35, the air/oil separator or a breather 36 and the oil pump 37. Furthermore, a fuel metering unit 38 (FMU), a measuring nozzle for controlling the amount of fuel which arrives at the burner 16, a fuel filter and an oil filter, an oil tank 41 and an oil cooler 42 which can be temperature controlled by means of fuel are also provided on the engine core 14 so as to be distributed in the circumferential direction U.

[0073] FIG. 4 shows a view corresponding to FIG. 3 of a second exemplary embodiment of the gas turbine engine 2 which differs from the design of the gas turbine engine 2 according to FIG. 3 only in some regions. The gas turbine engine 2 according to FIG. 4 comprises, instead of the gas turbine engine starter 33, what is known as an electric starter alternator 39, which, in the present case, is operatively connected to the radial shaft 22 by means of the aircraft accessory gearbox 21. The electric starter alternator 39 can be operated both as a motor and as an alternator so as to be able to start the gas turbine engine 2 and to generate electrical energy in the operation of the gas turbine engine 2. By means of the electrical energy of the starter alternator 39, for example an electric hydraulic pump and an on-board network of the aircraft 1 can be operated.

[0074] Both in the case of the gas turbine engine 2 according to FIG. 3 and in the case of the gas turbine engine 2 according to FIG. 4, the aircraft accessory gearbox 21, the aircraft accessories 32, 33 and 34, and the electric starter alternator 39 are arranged radially in part in the engine pylon 7 and in part in the fuselage 8.

LIST OF REFERENCE SIGNS

[0075] 1 Aircraft

[0076] 2 to 4 Gas turbine engine

[0077] 5 Angle drive

[0078] 6 Vertical stabilizer

[0079] 7 Engine pylon

[0080] 8 Fuselage

[0081] 9 Drive train

[0082] 10 Air inlet

[0083] 11 Bypass channel

[0084] 12 Inlet region

[0085] 13 Blower

[0086] 14 Engine core

[0087] 15 Compressor device

[0088] 16 Burner

[0089] 17 Low-pressure turbine

[0090] 18 High-pressure turbine

[0091] 19 Core accessory gearbox

[0092] 20 Intermediate casing

[0093] 21 Aircraft accessory gearbox

[0094] 22 Radial shaft

[0095] 23 Core shaft, low-pressure shaft

[0096] 24 Core shaft, high-pressure shaft

[0097] 25 Inner strut

[0098] 26 Additional angle drive

[0099] 27 Drive shaft

[0100] 28 Engine accessory

[0101] 29 Engine nacelle

[0102] 30 Gear pair

[0103] 31 Outer side of the bypass channel

[0104] 32 Hydraulic pump

[0105] 33 Gas turbine engine starter

[0106] 34 Integrated alternator

[0107] 35 Fuel pump

[0108] 36 Air/oil separator

[0109] 37 Oil pump

[0110] 38 Fuel metering unit

[0111] 39 Electric starter alternator, electric machine

[0112] 40 Additional angle drive

[0113] 41 Oil tank

[0114] 42 Oil cooler

[0115] 45 Outer strut

[0116] A Axial direction

[0117] R Radial direction of the gas turbine engine

[0118] U Circumferential direction