FLUID FLOW GUIDING DEVICE AND A GAS TURBINE ENGINE

Abstract

The invention concerns a flow guide device for a gas turbine engine, in particular an aircraft engine, with a fan and a bypass duct downstream of the fan, wherein the bypass duct is arranged radially between an engine nacelle and a core engine of the gas turbine engine, characterized by a guide vane assembly with load-bearing guide vanes and additional guide vanes in the bypass duct, wherein the load-bearing guide vanes are integrally connected to a structural component of the gas turbine engine, and the additional guide vanes are connected to the structural component via connecting means. The invention furthermore concerns a gas turbine device.

Claims

1. A flow guide device for a gas turbine engine, in particular an aircraft engine, with a fan and a bypass duct downstream of the fan, wherein the bypass duct is arranged radially between an engine nacelle and a core engine of the gas turbine engine, characterized by a guide vane assembly with load-bearing guide vanes and additional guide vanes in the bypass duct, wherein the load-bearing guide vanes are integrally connected to a structural component of the gas turbine engine, and the additional guide vanes are connected to the structural component via connecting means.

2. The flow guide device according to claim 1, wherein the structural component is connected or coupled to a part of the core engine and/or to a part of the engine nacelle.

3. The flow guide device according to claim 1, wherein the connecting means are configured as screw connections, form-fit connections and/or substance-bonded connections, wherein the connecting means are arranged in particular in a region over which air does not flow.

4. The flow guide device according to claim 1, wherein the number of additional guide vanes is equal to or greater than the number of load-bearing guide vanes.

5. The flow guide device according to claim 4, wherein the ratio of the number of additional guide vanes to load-bearing guide vanes is between 1:2 and 10:1, in particular between 1:1 and 10:1.

6. The flow guide device according to claim 1, wherein during operation, the additional guide vanes bear a smaller structural load than the load-bearing guide vanes.

7. The flow guide device according to claim 1, wherein the additional guide vanes have the same aerodynamically active profile and/or the same size as the load-bearing guide vanes.

8. The flow guide device according to claim 1, wherein the additional guide vanes have a different aerodynamically active profile and/or a different size from the load-bearing guide vanes.

9. The flow guide device according to claim 1, wherein the leading-edge and/or the trailing edge of the additional guide vanes and/or the load-bearing guide vanes is provided with a metal coating.

10. The flow guide device according to claim 1, wherein the additional guide vanes are connected to a mechanical damping element, in particular made of elastic plastic.

11. The flow guide device according to claim 1, wherein the additional guide vanes are made of a composite material, in particular a carbon-fiber composite material, or a metallic material.

12. The flow guide device according to claim 1, wherein the connecting means comprise a bolted connection.

13. The flow guide device according to claim 1, wherein the additional guide vanes are connected to an intermediate casing structure.

14. The flow guide device according to claim 13, wherein the intermediate casing structure comprises two concentrically arranged rings.

15. A gas turbine engine for an aircraft, said gas turbine engine comprising the following: a core engine comprising a turbine, a compressor, and a core shaft connecting the turbine to the compressor; a fan, which is positioned upstream of the core engine, wherein the fan comprises a plurality of fan blades; and a gear mechanism, which can be driven by the core shaft, wherein the fan can be driven by means of the gear mechanism at a lower rotational speed than the core shaft, having a flow guide device according to claim 1.

Description

[0052] FIG. 1 shows a sectional side view of a gas turbine engine;

[0053] FIG. 2 shows a close-up sectional side view of an upstream portion of a gas turbine engine;

[0054] FIG. 3 shows a partially cut-away view of a gear mechanism for a gas turbine engine;

[0055] FIG. 4 shows a sectional view through an embodiment of a flow guide device with additional guide vanes;

[0056] FIG. 5 shows a perspective view through a further embodiment of a flow guide device with additional guide vanes;

[0057] FIG. 6 shows a perspective illustration of an additional guide vane;

[0058] FIG. 7 shows a perspective illustration of a further embodiment of an additional guide vane;

[0059] FIG. 8 shows a detail view of a fixing of additional guide vanes;

[0060] FIG. 9 shows an example of a form-fit guide of a distal end of an additional guide vane.

[0061] FIG. 1 illustrates a gas turbine engine 10 having a main axis of rotation 9. The gas turbine engine 10 comprises an air inlet 12 and a 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 engine 11 that receives the core air flow A. When viewed in the order corresponding to the axial direction of flow, the core engine 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 via a shaft 26 and an epicyclic planetary gear mechanism 30. The epicyclic planetary gear mechanism 30 is a reduction gear mechanism.

[0062] During operation, 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 resulting hot combustion products then propagate through the high-pressure and the low-pressure turbines 17, 19 and thereby drive said turbines, before being expelled through the nozzle 20 to provide a certain propulsive thrust. The high-pressure turbine 17 drives the high-pressure compressor 15 by means of a suitable connecting shaft 27. The fan 23 generally provides the major part of the propulsive thrust. For high efficiency of the gas turbine engine, it is important that a rotation in the bypass air flow B is deflected in the direction of the engine axis.

[0063] A flow guide device 100, the function of which is described in more detail in connection with FIGS. 4 to 9, is arranged axially at the exit from the engine nacelle 21. In other embodiments, the engine nacelle 21 may extend axially over the region of the low-pressure compressor 14 up to the region of the high-pressure turbine 17.

[0064] 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 planetary gear mechanism 30. Multiple planet gears 32, which are coupled to one another by a planet carrier 34, are situated radially to the outside of the sun gear 28 and mesh therewith. The planet carrier 34 guides the planet gears 32 in such a way that they circulate synchronously around the sun gear 28, 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 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 support structure 24.

[0065] It is noted that the terms “low-pressure turbine” and “low-pressure compressor” as used herein may be taken to mean the lowest-pressure turbine stage and lowest-pressure compressor stage (i.e. not including the fan 23) respectively, and/or the turbine and compressor stages that are connected together by the connecting 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 documents, the “low-pressure turbine” and the “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 can be referred to as a first, or lowest-pressure, compression stage.

[0066] The epicyclic planetary gear mechanism 30 is shown by way of example in greater detail in FIG. 3. The sun gear 28, planet gears 32 and ring gear 38 in each case comprise teeth on their periphery to allow intermeshing with the other gearwheels. 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 may be provided within the scope of protection of the claimed invention. Practical applications of an epicyclic planetary gear mechanism 30 generally comprise at least three planet gears 32.

[0067] The epicyclic planetary gear mechanism 30 illustrated by way of example in FIGS. 2 and 3 is a planetary gear mechanism in which the planet carrier 34 is coupled to an output shaft via linkages 36, with the ring gear 38 being fixed. However, any other suitable type of planetary gear mechanism 30 may be used. As a further example, the planetary gear mechanism 30 may be a star arrangement, in which the planet carrier 34 is held fixed, with the ring gear (or external gear) 38 allowed to rotate. In such an arrangement, the fan 23 is driven by the ring gear 38. As a further alternative example, the gear mechanism 30 can be a differential gear mechanism in which both the ring gear 38 and the planet carrier 34 are allowed to rotate.

[0068] 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 protection of the present disclosure. Purely by way of example, any suitable arrangement may be used for locating the gear mechanism 30 in the gas turbine engine 10 and/or for connecting the gear mechanism 30 to the gas turbine engine 10. As a further example, the connections (for example the linkages 36, 40 in the example of FIG. 2) between the gear mechanism 30 and other parts of the gas turbine engine 10 (such as the input shaft 26, the output shaft and the fixed structure 24) may have a certain degree of stiffness or flexibility. As a further example, any suitable arrangement of the bearings between rotating and stationary parts of the gas turbine engine 10 (for example between the input and output shafts of the gear mechanism and the fixed structures, such as the gear casing) may be used, and the disclosure is not limited to the exemplary arrangement of FIG. 2. For example, where the gear mechanism 30 has a star arrangement (described above), a person skilled in the art would readily understand that the arrangement of output and supporting linkages and bearing positions would usually be different from that shown by way of example in FIG. 2.

[0069] Accordingly, the present disclosure extends to a gas turbine engine having any arrangement of gear mechanism types (for example star or epicyclic-planetary), supporting structures, input and output shaft arrangement, and bearing positions.

[0070] Optionally, the gear mechanism may drive additional and/or alternative components (for example the intermediate-pressure compressor and/or a booster compressor).

[0071] Other gas turbine engines in which the present disclosure can be used may have alternative configurations. For example, such engines may have an alternative number of compressors and/or turbines and/or an alternative number of connecting shafts. By way of a 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 from and radially outside the engine core nozzle 20. However, this is not restrictive, 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 may be referred to as a mixed flow nozzle. One or both nozzles (whether mixed or split flow) can have a fixed or variable area. In some arrangements, the gas turbine engine 10 may not comprise a gear mechanism 30.

[0072] The geometry of the gas turbine engine 10, and components thereof, is/are defined by a conventional axis system, comprising an axial direction (which is aligned with the axis of rotation 9), a radial direction (in the bottom-to-top direction 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.

[0073] On the basis of FIGS. 4 to 9, embodiments of the flow guide device 100 are now described which serve to reduce the rotational part of the bypass air flow B.

[0074] FIG. 4 shows an extract from the illustration in FIG. 2 which shows the flow guide device 100 behind the fan 23 (not shown here). The rotating bypass air flow B flows from left to right through the bypass duct 22.

[0075] The purpose of a guide vane assembly 50 as part of the flow guide device 100 in the bypass duct 22 is to align the bypass air flow B, flowing in from the left here, such that the rotation in the flow is reduced, ideally even reduced to zero. For this, the guide vane assembly 50 comprises load-bearing guide vanes 51 and additional guide vanes 52, which are illustrated as a whole in FIG. 5.

[0076] FIG. 4 shows the load-bearing guide vane 51 behind the additional guide vane 52. The load-bearing guide vane absorbs forces and moments from the bypass air flow B and conducts them in the known fashion to a structural component 60 (see FIG. 5), which is here connected to the core engine 11 and engine nacelle 21. The structural component 60 is a cast component which is integrally connected to the load-bearing guide vanes 51 (see FIG. 5).

[0077] The additional guide vanes 52 are not integrally connected to the structural component 60, but via a connecting means 55; in the embodiment illustrated, via a screw connection. The connecting means 55 is here arranged such that it is situated in a cavity and not exposed to the air flow.

[0078] Thus the additional guide vanes 52 transfer significantly lower forces and moments to the structural component 60 than the load-bearing guide vanes 51.

[0079] To damp vibrations, the additional guide vanes 52 are coupled to a damping means 53, e.g. a rubber part, which is illustrated schematically in FIG. 4.

[0080] Since the number of load-bearing guide vanes 51 is limited by the design in the casting, the additional guide vanes 52 serve to increase the number of guide vanes as a whole, whereby the efficiency of the gas turbine engine 10 is increased.

[0081] In FIG. 5, an additional guide vane 52 is arranged between every two load-bearing guide vanes 51, so that the ratio of the number of additional guide vanes 52 to the number of load-bearing guide vanes 51 is 1:1. In alternative embodiments, the ratio may be higher, e.g. 5:1. It is however in principle also possible that the number of additional guide vanes 52 is smaller than the number of load-bearing guide vanes 51.

[0082] FIG. 5 also shows that the additional guide vanes 52 have the same aerodynamically active profile and the same size as the load-bearing guide vanes 51. In principle however, it is also possible that the aerodynamically active profile and/or the size of the additional guide vanes 52 differs from those of the load-bearing guide vanes 51.

[0083] In the embodiment illustrated here, the additional guide vanes 52 are made of a composite material, in particular a carbon-fiber composite material or a metallic material. In an embodiment not shown here, the additional guide vanes 52 have a metallic coating on the leading edge and/or trailing edge.

[0084] FIG. 5 furthermore shows that the additional guide vanes 52, and also the load-bearing guide vanes 51, are connected to an intermediate casing structure 62 which comprises two concentrically arranged rings 62, 63.

[0085] FIG. 6 shows a single additional guide vane 52 which, at its radially inner end, has a bolt opening 56 for a bolt connection as part of the connecting means 55.

[0086] FIG. 7 shows a derivative of the embodiment of the additional guide vane 52 in FIG. 6. The additional guide vane 52 has guide elements 57 at the radially distal end, which may be inserted in a corresponding receiver 58, e.g. in the engine nacelle 21 (see also FIG. 9).

[0087] FIG. 8 shows that the additional guide elements 52 are connected at the radially proximal end to the structural component 60 via an elastic casting compound as a connecting means 55. The casting compound may here also serve as a damping element.

[0088] The concept of the additional guide vanes 52 described here allows good access to the fixing points. During maintenance of the gas turbine engine 10, for example individual additional guide vanes 52 may be replaced.

[0089] It will be understood that the invention is not limited to the embodiments described above, and various modifications and improvements can be made without departing from the concepts described herein. Any of the features may be used separately or in combination with any other features, unless they are mutually exclusive, and the disclosure extends to and includes all combinations and subcombinations of one or more features which are described here.

LIST OF REFERENCE SIGNS

[0090] 9 Main axis of rotation [0091] 10 Gas turbine engine, aircraft engine [0092] 11 Core engine [0093] 12 Air inlet [0094] 14 Low-pressure compressor [0095] 15 High-pressure compressor [0096] 16 Combustion device [0097] 17 High-pressure turbine [0098] 18 Bypass thrust nozzle [0099] 19 Low-pressure turbine [0100] 20 Core thrust nozzle [0101] 21 Engine nacelle [0102] 22 Bypass duct [0103] 23 Fan [0104] 24 Stationary supporting structure [0105] 26 Shaft [0106] 27 Connecting shaft [0107] 28 Sun gear [0108] 30 Gear mechanism [0109] 32 Planet gears [0110] 34 Planet carrier [0111] 36 Linkage [0112] 38 Ring gear [0113] 40 Linkage [0114] 50 Guide vane assembly [0115] 51 Load-bearing guide vane [0116] 52 Additional guide vane [0117] 53 Damping element [0118] 55 Connecting means, bolted connection [0119] 56 Bolt opening [0120] 57 Guide element [0121] 58 Receiver for guide element [0122] 60 Structural component [0123] 61 Intermediate casing structure [0124] 62 First ring of intermediate casing structure [0125] 62 Second ring of intermediate casing structure [0126] 100 Flow guide device [0127] A Core air flow [0128] B Bypass air flow