FAN CASE ASSEMBLY FOR A GAS TURBINE ENGINE

Abstract

Aspects of the disclosure regard a fan case assembly for a gas turbine engine, the fan case assembly comprising a fan case having an inner surface, a fan track liner comprising abradable material layer, and a rear acoustic panel arranged aft of the fan track liner. The fan track liner and the rear acoustic panel are integrated into a single panel structure attached to the fan case inner surface.

Claims

1. A fan case assembly for a gas turbine engine, the fan case assembly comprising: a fan case having an inner surface; a fan track liner comprising an abradable material layer; and a rear acoustic panel arranged aft of the fan track liner; wherein the fan track liner and the rear acoustic panel are integrated into a single panel structure attached to the fan case inner surface.

2. The fan case assembly of claim 1, wherein the single panel structure comprises an outer tray extending between a front end and an aft end of the single panel structure, wherein the outer tray forms an outer structure of the single panel structure holding the fan track liner and the rear acoustic panel, wherein the outer tray is connected to the fan case.

3. The fan case assembly of claim 2, wherein the outer tray is connected at three axial positions only to the fan case.

4. The fan case assembly of claim 3, wherein the outer tray is connected to the fan case at a front end, at an aft end, and at a middle position.

5. The fan case assembly of claim 2, wherein the connection between the outer tray and the fan case is by means of fasteners, the fasteners comprising at least one of screws, bolts and flange connections.

6. The fan case assembly of claim 2, wherein the fan track liner and the rear acoustic panel are formed by two different compartments of the single panel structure.

7. The fan case assembly of claim 6, wherein the fan track liner compartment and the rear acoustic panel compartment each comprise a honeycomb core structure, wherein the honeycomb core structures of the fan track liner compartment and the rear acoustic panel compartment differ in at least one of cell configuration, cell size and density.

8. The fan case assembly of claim 7, wherein the honeycomb core structure of the fan track liner compartment comprises a Flex-Core cell configuration.

9. The fan case assembly of claim 7, wherein the honeycomb core structure of the rear acoustic panel compartment comprises a hexagonal cell configuration.

10. The fan case assembly of claim 7, wherein the honeycomb core structure of the fan track liner compartment has a higher density than the honeycomb core structure of the rear acoustic panel compartment.

11. The fan case assembly of claim 7, wherein the honeycomb core structure of the fan track liner compartment is covered by a septum inner sheet to which a layer of abradable material is attached.

12. The fan case assembly of claim 7, wherein the honeycomb core structure of the rear acoustic panel is covered by a perforated inner face sheet.

13. The fan case assembly of claim 6, wherein an inner skin of the rear acoustic panel compartment and the aft end of the outer tray together form an aft flange which is connected to the fan case.

14. The fan case assembly of claim 6, wherein an inner skin of the rear acoustic panel compartment and the aft end of the outer tray are connected to the fan case by means of a through fastener, wherein the structure of the rear acoustic panel is reinforced in the area of the through fastener.

15. The fan case assembly of claim 6, wherein an inner skin of the rear acoustic panel compartment forms a radial offset that is radially trapped with a retention ring that is bolted to the structure of the fan case.

16. The fan case assembly of claim 2, wherein the outer tray is a tray formed by carbon fiber composite material.

17. The fan case assembly of claim 7, wherein the single panel further comprises an ice impact liner compartment arranged between the fan track liner compartment and the rear acoustic panel compartment, wherein the ice impact liner compartment comprises a honeycomb core structure different from the honeycomb core structures of the fan track liner compartment and the rear acoustic panel compartment.

18. The fan case assembly of claim 17, wherein the ice impact liner compartment comprises a non-perforated inner face sheet formed by a laminate material comprised of a plurality of plies.

19. The fan case assembly of claim 1, wherein the single panel structure comprises an inner skin delimiting the outer flow path through the fan, wherein the inner skin comprises at least in sections a layer made of high modulus polypropylene fibers or hybrid fibers containing high modulus polypropylene.

20. The fan case assembly of claim 1, wherein at least the rear acoustic panel compartment of the single panel structure in cross section converges along its full length in the aft direction.

Description

[0065] The invention will be explained in more detail on the basis of exemplary embodiments with reference to the accompanying drawings in which:

[0066] FIG. 1 is a sectional side view of a gas turbine engine;

[0067] FIG. 2 is a close up sectional side view of an upstream portion of a gas turbine engine;

[0068] FIG. 3 is a partially cut-away view of a gearbox for a gas turbine engine;

[0069] FIG. 4 is an embodiment of a fan case assembly comprising a fan track liner and a rear acoustic panel integrated into a single panel structure, wherein the aft end of the single panel structure forms an aft flange;

[0070] FIG. 4A an enlarged view of the single panel structure of the fan case assembly of FIG. 4;

[0071] FIG. 5 is a further embodiment of a fan case assembly comprising a fan track liner and a rear acoustic panel integrated into a single panel structure, wherein the aft end of the single panel structure comprises a through connection; and

[0072] FIG. 6 is a still further embodiment of a fan case assembly comprising a fan track liner and a rear acoustic panel integrated into a single panel structure, wherein the aft end of the single panel structure forms a radial offset.

[0073] 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 epicyclical gearbox 30.

[0074] 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 epicyclical gearbox 30 is a reduction gearbox.

[0075] 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 epicyclical 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 process 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.

[0076] 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.

[0077] The epicyclical 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 epicyclical gearbox 30 generally comprise at least three planet gears 32.

[0078] The epicyclical 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. However, any other suitable type of epicyclical gearbox 30 may be used. By way of further example, the epicyclical 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.

[0079] 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 30 in the engine 10 and/or for connecting the gearbox 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 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 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.

[0080] 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.

[0081] Optionally, the gearbox may drive additional and/or alternative components (e.g. the intermediate pressure compressor and/or a booster compressor).

[0082] 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 30.

[0083] 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.

[0084] In the context of this invention, the design of a fan case assembly enclosing the fan 23 is of relevance. It is pointed out that the fan case assembly that will be discussed in the following may be implemented in a geared turbofan engine as discussed with respect to FIGS. 1 to 3 but may generally be implemented in any gas turbine engine. The principles of the present invention are not dependent on a particular kind of gas turbine engine.

[0085] More particularly, a particularly useful application lies with Civil Small and Medium Engines. which may have a fan diameter in the range between 35 to 55″. The rotational speed of the fan of such Civil Small and Medium Engines may be in the range between 5000 and 9000 rpm at Maximum Takeoff Thrust.

[0086] FIG. 4 depicts a first embodiment of a fan case assembly for a gas turbine engine. The fan case assembly comprises a fan case 4 circumferentially surrounding a fan 23. The fan case 4 comprises a front end at which it is connected to an engine inlet (not shown), wherein the connection may be realized by means of a flange connection 401. The flange connection 401 is also referred to as an A1 connection. The fan case 4 further comprises an aft end at which it is connected to further structural elements of the gas turbine engine nacelle. The connection may be realized by means of a flange connection 402 which is also referred to as A3 connection.

[0087] The fan case 4 comprises an outer surface 41 and an inner surface of 42, wherein the inner surface 42 faces the flow path through the fan 3. Several panels are arranged along the inner surface 42. Upstream of the fan 3 is located a front acoustic panel 6 absorbing sound which, however, is of no particular concern in the present context. There is further provided a single panel structure 5 attached to the fan case 4 which includes a plurality of different compartments providing different structures and associated functions.

[0088] More particularly, the single panel structure 5 forms radially outward to the fan 23 and adjacent to the fan blade tips 230 a fan track liner compartment 51 which comprises a layer 512 of abradable material. The layer 512 of abradable material minimizes air leakage around the blade tips 230. The single panel structure 5 further forms an ice impact liner compartment 52 aft of the fan track liner compartment 51 and a rear acoustic panel compartment 53 aft of the ice impact liner compartment 52. The ice impact liner compartment 52 is configured to withstand ice shed from the fan rotor blades. The rear acoustic panel compartment 53 is configured to attenuate noise and improve flutter margin.

[0089] In addition, the fan track liner compartment 51 may be structurally embodied in such a manner that it is suited for receiving fan fragments in the event that a fan blade breaks and for avoiding that they penetrate the engine nacelle in an outward direction.

[0090] The single panel structure 5 further comprises an outer tray 55 which extends between a front end and an aft end of the single panel structure 5. The outer tray 55 represents an outer structure which serves to receive and hold the different compartments 51, 52, 53. The outer tray 55 is formed in a straight manner along most of its length. The outer tray 55 is connected at three axial positions to the fan case 4. The first axial position is a front position at which the outer tray 55 is connected to the fan case 4 by a row of fasteners 501 such as a row of screws. The second position is an aft position at which the outer tray 55 is connected to the fan case 4 by a second row of fasteners 502. The third position is a middle position at which the outer tray 55 is connected to the fan case 4 by a third row of fasteners 503.

[0091] The outer tray 55 may be formed as one piece. It may be formed by a glass fiber composite or from a carbon fiber composite material.

[0092] In the depicted embodiment, there are provided non-structural outlet guide vanes 25 at the beginning of the bypass channel of the gas turbine engine. Such outlet guide vanes 25 limit the length of the rear acoustic panel compartment 53.

[0093] The single panel structure 5 is depicted in FIG. 4 in a cross-sectional view. In the circumferential direction, the single panel structure 5 may consist of a plurality of panels each extending in the circumferential direction over less than 360°. In embodiments, the single panel structure 5 may consist of five to thirteen panel structures each extending 360°/(number of panel structures) in the circumferential direction.

[0094] The single panel structure 5 of FIG. 4 is depicted in enlarged view in FIG. 4A. With respect to attachment of the outer tray 55 to the fan case 4, FIG. 4A shows in more detail on that the inner skin of the single panel structure 5 and the outer tray 55 form at the front end of the single panel structure 5 a front flange 551 which is fastened by the row of screws 501 to the fan case 4. In a similar manner, the inner skin of the single panel structure 5 and the outer tray 55 form at the aft end of the single panel structure 5 an aft flange 552 which is fastened by the row of screws 502 to the fan case 4. By the outer tray 55 extending to the aft flange 552, particular stiffness for the unsupported span is provided for. Also, by avoiding through fasteners, there is a reduced fire and vent concern. A third row of fasteners such as bolts 503 provides for attachment of the outer tray 52 to the fan case 4 in a middle region.

[0095] The first, second and third compartments 51 to 53 of the single panel structure 5 which form the fan track liner, the ice impact panel and the rear acoustic panel each comprise a honeycomb core structure, wherein the honeycomb core structures of the different compartments 51 to 53 differ in at least one of cell configuration, cell size and density. Generally, the honeycomb core structure is part of a honeycomb sandwich structure additionally comprising an inner face sheet and an outer face sheet, wherein the outer face sheet may be formed by respective sections of the outer tray 55 or is a separate outer face sheet (not shown) attached to the inside of outer tray 55. For example, there may be a glass layer between the outer tray 55 carbon fiber and the honeycomb structure.

[0096] More particularly, the first compartment 51 which forms the fan track liner comprises a honeycomb core structure 510, an inner septum sheet 511 and the layer 512 of abradable material. The layer 512 of abradable material is attached to the inner septum sheet 511. The honeycomb core structure 510 may comprise a Flex-core cell configuration available from the company Hexcel Corporation.

[0097] The second compartment 52 which forms the ice impact panel comprises a honeycomb core structure 520 and an inner face sheet 521. The face sheet, in embodiments, is a non-perforated face sheet formed by a laminated material comprised of a plurality of plies. For example, the face sheet of 521 may comprise several plies of glass fiber composites and several plies of high modulus polypropylene composites. The total thickness of the face sheet 521 may be in the range between 1.5 and 4 millimeter. The honeycomb core structure 520 may comprise the typical hexagonal cell configuration but may be higher density and/or wall thickness relative to honeycomb core structure 510. Alternatively, the honeycomb core structure 520 and the aft portion of honeycomb core structure 510 may be of the same honeycomb core structure and density. Generally, the aft section of honeycomb core structure 510 may be denser than the front section of honeycomb cores structure 510, i.e., honeycomb core structure 510 could be divided into two sections. In such case, in embodiments, the front section has a density in the range between 4 and 6 pcf and the aft section has a density in the range between 7.5 and 12.5 pcf.

[0098] However, in other embodiments, the face sheet 521 may be a perforated face sheet which allows to the second compartment 52 to participate in noise reduction and flutter mitigation provided by the third compartment.

[0099] The third compartment 53 which forms the rear acoustic panel comprises a honeycomb core structure 530 and an inner face sheet 531. The face sheet 531 is a perforated face sheet which may be formed by a glass fiber composite. Alternatively, carbon fiber composites may be included to form the face sheet 531. The honeycomb core structure 530 may comprise the typical hexagonal cell configuration.

[0100] The cell configurations 510, 520, 530 of the three compartments 51, 52, 53 may differ to provide for the respective desired capability needs. For example, the honeycomb core structure 510 may have a higher density than the honeycomb core structure 530. The honeycomb core structure 520 of compartment 52 may have a greater cell depth than the honeycomb core structure 530 of compartment 53. This reflects that in the sections of the single panel structure 5 which form the ice impact panel compartment 52 and the rear acoustic panel department 53, the single panel structure 5 converges in cross-section in the aft direction.

[0101] FIG. 5 shows a further embodiment of a fan case assembly which differs from the embodiment of FIG. 4 only in the manner that the aft end of the single panel structure 5 is connected to the fan case 4. Accordingly, with respect to all other issues, reference is made to FIGS. 4, 4A.

[0102] In the embodiment of FIGS. 4, 4A the outer tray 55 is bent inward at the aft end to form an aft flange together with the inner skin of the rear acoustic panel compartment. The embodiment of FIG. 5 in comparison is easier to produce, wherein a through fastener 504 is used to connect the fan case 4, the outer tray 55 and the compartment 53. To this end, the structure of the rear acoustic panel compartment 53 is reinforced in the area of the through fastener. In an embodiment, this may include potting in the panel to provide compression strength to the honeycomb so a screw or bolt would not crush it. Also the outer tray 55 can be seen offset from the fan case inner surface 42 with a laid-up pad spacing it off at the bolt through location 504. This potting reduces the effective acoustic area, but is implemented at the panel's very end which typically comprises foaming film adhesive or similar to close out the aft edge such that there is already a reduction in effective area required for structural reasons.

[0103] FIG. 6 shows a still further embodiment of a fan case assembly which also differs from the embodiment of FIG. 4 only in the manner that the aft end of the single panel structure 5 is connected to the fan case 4. In the embodiment of FIG. 6, the rear acoustic panel compartment 53 forms a radial offset 506 which is captured by a retention block or segmented ring 505 which is fastened to the fan case. Axial fixation of the single panel structure 5 is provided for by front and middle fasteners 501 and 503. This embodiment is associated with a particularly simple manner of installation, wherein the retention block 505 only needs to be fastened into the case and not through the panel. Another benefit is to avoid a perforated firewall as well as enabling the thinnest case possible there.

[0104] The retention block or retention ring 505 may be a removable flange that is bolted to the aft end of compartment 53.

[0105] Another potential option which yields the same benefits is to have a screw through a small strap and then have a potted insert in the aft end of the compartment 53.

[0106] A general preference to enable the concept described with respect to FIGS. 4 to 6 is a generally converging profile of the fan case 4 between the front fasteners 501 and the outlet guide vanes 25. This supports the possibility that the outer tray 55 is one piece (even if a step for a radial offset from the case is needed over perforated areas).

[0107] It should be understood that the above description is intended for illustrative purposes only and is not intended to limit the scope of the present disclosure in any way. Also, those skilled in the art will appreciate that other aspects of the disclosure can be obtained from a study of the drawings, the disclosure and the appended claims. All methods described herein can be performed in any suitable order unless otherwise indicated herein or otherwise clearly contradicted by context. Various features of the various embodiments disclosed herein can be combined in different combinations to create new embodiments within the scope of the present disclosure. In particular, the disclosure extends to and includes all combinations and sub-combinations of one or more features described herein. Any ranges given herein include any and all specific values within the range and any and all sub-ranges within the given range.