FAN CONTAINMENT

Abstract

A fan containment arrangement for a gas turbine engine comprises: a fan containment casing; and a fan track liner radially within the fan containment casing and extending radially inwardly from the fan containment casing to a gas-washed surface delimiting a gas path for a fan. The fan track liner comprises an impact resistant layer formed from titanium or titanium alloy. No greater than about 20% of the radial thickness of the fan track liner is formed from titanium or titanium alloy. The fan track liner comprises first and second cellular impact structures separated from one another by a first septum layer, the first cellular impact structure being the impact resistant layer formed from titanium or titanium alloy, or the impact resistant layer formed from titanium or titanium alloy is a septum layer, separating the first and second cellular impact structures. (FIG. 5)

Claims

1. A fan containment arrangement for a gas turbine engine comprising: a fan containment casing; and a fan track liner radially within the fan containment casing and extending radially inwardly from the fan containment casing to a gas-washed surface delimiting a gas path for a fan; wherein the fan track liner comprises an impact resistant layer formed from titanium or titanium alloy, wherein no greater than about 20% of the radial thickness of the fan track liner is formed from titanium or titanium alloy, and (i) the fan track liner comprises first and second cellular impact structures separated from one another by a first septum layer, the first cellular impact structure being the impact resistant layer formed from titanium or titanium alloy, or (ii) the impact resistant layer formed from titanium or titanium alloy is a septum layer, separating the first and second cellular impact structures.

2. The fan containment arrangement of claim 1, wherein the impact resistant layer formed from titanium or titanium alloy is a cellular impact structure formed from titanium or titanium alloy.

3. The fan containment arrangement of claim 2, wherein the cellular impact structure formed from titanium or titanium alloy constitutes from about 10% to about 20%, of the radial thickness of the fan track liner.

4. The fan containment arrangement of claim 1, wherein the fan track liner comprises first and second cellular impact structures separated from one another by a first septum layer, the first cellular impact structure being the impact resistant layer formed from titanium or titanium alloy.

5. The fan containment arrangement of claim 4, wherein the first cellular impact structure is a radially outboard cellular impact structure and the second cellular impact structure is a radially inboard cellular impact structure.

6. The fan containment arrangement of claim 4, wherein one or each of the first cellular impact structure and the second cellular impact structure is a honeycomb structure.

7. The fan containment arrangement of claim 4, wherein the second cellular impact structure is formed from aluminium, aluminium alloy or a polymeric material such as an aramid polymeric material.

8. The fan containment arrangement of claim 4, wherein the second cellular impact structure constitutes from about 70% to about 80% of the radial thickness of the fan track liner.

9. The fan containment arrangement of claim 4, further comprising an abradable structure separated from the second cellular impact structure by a second septum layer.

10. The fan containment arrangement of claim 4, wherein the first and/or second septum layers, where present, are formed from fibre-reinforced polymer material.

11. The fan containment arrangement of claim 1, wherein the impact resistant layer formed from titanium or titanium alloy is a septum layer, formed from titanium or titanium alloy, separating the first and second cellular impact structures.

12. The fan containment arrangement of claim 11, wherein the septum layer formed from titanium or titanium alloy is a sheet of titanium or titanium alloy.

13. The fan containment arrangement of claim 1, wherein the impact resistant layer is formed from a titanium alloy comprising titanium, aluminium and vanadium, such as Ti-3Al-2.5V or Ti-6Al-4V.

14. The fan containment arrangement of claim 1, wherein titanium or titanium alloys account for no greater than about 30% of the total weight of the fan track liner.

15. A gas turbine engine comprising a fan containment arrangement of claim 1.

Description

BRIEF DESCRIPTION OF THE DRAWINGS

[0091] Embodiments will now be described by way of example only, with reference to the Figures, in which:

[0092] FIG. 1 is a sectional side view of a gas turbine engine;

[0093] FIG. 2 is a close up sectional side view of an upstream portion of a gas turbine engine;

[0094] FIG. 3 is a partially cut-away view of a gearbox for a gas turbine engine;

[0095] FIG. 4 is a schematic sectional side view of a fan containment arrangement;

[0096] FIG. 5 is a schematic sectional view through a portion of a first example fan track liner; and

[0097] FIG. 6 is a schematic sectional view through a portion of a second example fan track liner.

DETAILED DESCRIPTION OF THE DISCLOSURE

[0098] Aspects and embodiments of the present disclosure will now be discussed with reference to the accompanying figures. Further aspects and embodiments will be apparent to those skilled in the art.

[0099] 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. A fan containment arrangement 41 extends around the fan 23 inboard the nacelle 21.

[0100] 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 core exhaust 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.

[0101] 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.

[0102] 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.

[0103] 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.

[0104] 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. 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.

[0105] 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.

[0106] 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.

[0107] Optionally, the gearbox may drive additional and/or alternative components (e.g. the intermediate pressure compressor and/or a booster compressor).

[0108] 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 18, 20 meaning that the flow through the bypass duct 22 has its own nozzle 18 that is separate to and radially outside the core exhaust 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.

[0109] 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.

[0110] The structure of the fan containment arrangement 41 is illustrated schematically in FIG. 4 which shows a sectional view of one portion of the fan containment arrangement in a radial plane intersecting the engine centreline. The fan containment arrangement 41 comprises a fan containment casing 42 which extends between a forward end 43 and an aft end 44. The fan containment casing 42 is formed predominantly from fibre-reinforced composite material, for example carbon-fibre reinforced polymer (CFRP). A fan impact liner 45 is adhered or otherwise secured to an inboard surface of the fan containment casing 42 intermediate the forward and aft ends. The fan impact liner 45 is formed from layers of fibre-reinforced composite material and honeycomb material, as explained in more detail below, and is configured to absorb a substantial amount of energy on impact of a blade during a fan blade-off (FBO) event. Forward and aft acoustic liners 46 and 47 are adhered to the fan containment casing 42 proximate the forward 43 and aft 44 ends respectively. The fan containment casing 42 acts as a rigid structural support for the fan impact liner 45 and the forward and aft acoustic liners 46 and 47.

[0111] The structure of the fan impact liner 45, in a first example embodiment, is illustrated schematically in FIG. 5 which shows a sectional view of one portion of the fan impact liner 45 adhered to the fan containment casing 42. The fan impact liner 45 includes three different layers (48, 49 and 50) of hexagonal honeycomb material: a radially outer honeycomb layer, a radially intermediate honeycomb layer 49, and a radially inner honeycomb layer.

[0112] Radially outer honeycomb layer 48 is formed from titanium or a titanium alloy such as Ti-3Al-2.5V or Ti-6Al-4V and, in this example, accounts for about 10% to about 15% of the total radial thickness of the fan impact liner 45. For example, the honeycomb layer 48 may be about 6.6 mm thick.

[0113] Radially intermediate honeycomb layer 49 is formed from aluminium or aluminium alloy and, in this example, accounts for about 70% to about 80% of the total radial thickness of the fan impact liner 45. For example, the honeycomb layer 49 may be about 38 mm thick. A suitable example aluminium honeycomb material has a cell size of inch (i.e. about 3.175 mm) and a foil thickness of 0.003 inches (i.e. about 0.0762 mm), but cell sizes between about 0.004 inches (i.e. about 0.1 mm) and about 0.8 inches (i.e. about 20.3 mm), and foil thicknesses between about 0.0005 inches (i.e. about 0.013 mm) and about 0.01 inches (i.e. about 0.25 mm), are possible.

[0114] Radially inner honeycomb layer 50 is formed from an aramid polymeric material such as NOMEX flame-resistant meta-aramid material, the cells of the honeycomb being filled with a cured epoxy resin (such as those available in the Scotch-Weld line available from 3M). The honeycomb layer 50 may be about 6 mm thick. Radially inner honeycomb layer 50 is an abradable layer having a gas-washed exterior surface S which faces the fan blades 23. In use, the fan blade is run-in to cut a track in the abradable layer for good clearance.

[0115] In this example, the honeycomb layer 48 is adhered directly to the fan containment casing 42 by a layer of cured epoxy adhesive (not shown). The radially outer and intermediate honeycomb layers 48 and 49 are separated from one another by an intervening composite outer septum layer 51 formed from two carbon-fibre reinforced plies suspended in an epoxy resin matrix, the outer septum layer having a thickness of about 0.5 mm. Both the radially outer and intermediate honeycomb layers 48 and 49 are adhered to the septum layer 51 by a cured epoxy adhesive. The intermediate and inner honeycomb layers 49 and 50 are also separated from one another by an intervening composite inner septum layer 52 formed from twelve carbon-fibre reinforced plies suspended in an epoxy resin matrix (having a total thickness of about 3 mm), and again both the intermediate and inner honeycomb layers 49 and 50 are adhered to the septum layer 52 by a cured epoxy adhesive.

[0116] It will be apparent to the skilled person that many of the materials used in the construction of the fan impact liner shown in FIG. 5 may be varied without departing from the general structure of the liner. For example, the honeycomb layers 48, 49 and 50 may take any suitable honeycomb structures other than hexagonal structures (for example, they may have square honeycomb structures). The intermediate honeycomb layer 49 formed from aluminium or aluminium alloy may be replaced by a honeycomb layer formed from any suitable material having a density less than that of titanium or titanium alloy, such as a polymeric material like NOMEX flame-resistant meta-aramid material. The septum layers 51 and 52 may be formed from any suitable composite materials, for example fibreglass. The thickness of the septum layers may be adjusted to take into account any changes in honeycomb materials. For example, in embodiments in which the honeycomb layer 49 is formed from a polymeric material like NOMEX flame-resistant meta-aramid material, septum layer 52 may be formed from two carbon fibre plies instead of two plies, reducing the septum layer thickness from about 3 mm to about 0.5 mm. However, in all such variations of this first example embodiment, the honeycomb layer 48 is formed from titanium or titanium alloy.

[0117] In use, the fan impact liner 45 as shown in FIG. 5 absorbs a substantial proportion of the kinetic energy of an impacting projectile, such as a detached fan blade. As a projectile begins to impact the fan impact liner 45, contact with the inner septum layer 52 distributes the impact force over a broad area, such that a large portion of the honeycomb layer 49 undergoes deformation. As the honeycomb layer 49 deforms, energy is absorbed. Accordingly, the speed at which the projectile moves through the fan impact liner 45 is reduced. However, an impacting projectile such as a fan blade is not typically arrested fully in the honeycomb layer 49. Instead, as the projectile reaches the septum layer 51, the impact force is again spread over a large area of the honeycomb layer 48. Because the honeycomb layer 48 is formed from titanium or titanium alloy, which is significantly stronger than the aluminium or polymeric materials used to form the honeycomb layer 49, the cell walls of honeycomb layer 48 are more resistant to deformation and the radially outer honeycomb layer 48 is able to absorb the remaining kinetic energy of the projectile to fully arrest the projectile. The honeycomb layer 48 therefore acts as a barrier which hinders or prevents the projectile from penetrating the fan containment casing 42.

[0118] The honeycomb layer 48 formed from titanium or titanium alloy is thin relative to the non-titanium honeycomb layer 49. In particular, the honeycomb layer 48 generally constitutes less than 20%, and preferably between 10% and 15%, of the total radial thickness of the fan track liner. Accordingly, despite titanium having a higher density than aluminium or polymeric materials, incorporation of the titanium honeycomb layer 48 does not lead to significant weight gains compared to previously-considered fan impact liners. In addition, because of its higher strength compared to aluminium or polymeric materials, the titanium honeycomb layer 48 is able to absorb a substantial amount of kinetic energy from an impacting projectile over a short distance. The honeycomb layer 48 is therefore able to function effectively as a ballistic barrier to impacting projectiles, despite its relative thinness. Indeed, it is the use an outboard titanium structure, which has a high crush strength and a high compressive strength compared to other commonly-used materials, which permits incorporation of a large volume of relatively lower strength inboard layers. The lower-strength inboard layers are provided primarily for their energy absorption and deformation capabilities, while the titanium structure acts as a final barrier to stop an impacting projectile. Use of a titanium honeycomb layer is estimated to reduce stress levels in the fan containment casing, on impact, by up to 40%.

[0119] The inventors have found that the fan impact liner 45 is able to arrest impacting projectiles more effectively than known fan impact liners which do not incorporate titanium or titanium alloy layers. The fan impact liner 45 is therefore suitable for use in larger engines which make use of larger, heavier fan blades. In particular, the fan impact liner is able to reduce and spread the impact energy due to an impacting projectile, such as a fan blade, protecting the casing against ice and blade impacts and damping vibrations, thereby also reducing noise.

[0120] In an alternative second example embodiment, the fan impact liner 45 takes the structure illustrated schematically in FIG. 6 which again shows a sectional view of one portion of the fan impact liner 45 adhered to the fan containment casing 42. In this embodiment, the fan impact liner 45 includes three different layers (53, 54 and 55) of hexagonal honeycomb material: a radially outer honeycomb layer 53, an intermediate honeycomb layer 54, and a radially inner honeycomb layer. The radially outer honeycomb layer 53 is formed from aluminium or aluminium alloy and accounts for about 40% of the total radial thickness of the fan impact liner 45. Intermediate honeycomb layer 54 is formed from an aramid polymeric material such as NOMEX flame-resistant meta-aramid material and accounts for about 40% of the total radial thickness of the fan impact liner 45. Radially inner honeycomb layer 55 is also formed from an aramid polymeric material such as NOMEX flame-resistant meta-aramid material, the cells of the honeycomb being filled with a cured epoxy resin (such as those available in the Scotch-Weld line available from 3M). Radially inner honeycomb layer 55 is an abradable layer having a gas-washed exterior surface S which faces the fan blades 23. In use, the fan blade is run-in to cut a track in the abradable layer for good clearance.

[0121] The honeycomb layer 53 is adhered directly to the fan containment casing 42 by a layer of cured epoxy adhesive (not shown). The radially outer and intermediate honeycomb layers 53 and 54 are separated from one another by an intervening septum layer 56 formed from a sheet of titanium or a titanium alloy such as Ti-3Al-2.5V or Ti-6Al-4V. Both the radially outer and intermediate honeycomb layers 53 and 54 are adhered to the septum layer 56 by a cured epoxy adhesive. The intermediate and radially inner honeycomb layers 54 and 55 are separated from one another by an intervening composite septum layer 57 formed from carbon-fibre reinforced plies suspended in an epoxy resin matrix. Both the intermediate and radially inner honeycomb layers 54 and 55 are adhered to the septum layer 57 by a cured epoxy adhesive.

[0122] Again, it will be apparent to the skilled person that many of the materials used in the construction of the fan impact liner shown in FIG. 6 may be varied without departing from the general structure of the liner. For example, the honeycomb layers 53, 54 and 55 may take any suitable honeycomb structures other than hexagonal structures (for example, they may have square honeycomb structures). Either or both of honeycomb layers 53 and 54 may be formed from aluminium, aluminium alloy, or a polymeric material like NOMEX flame-resistant meta-aramid material. The septum layer 57 may be formed from any suitable composite material, for example fibreglass. However, in all such variations of this second example embodiment, the septum layer 56 is formed from titanium or titanium alloy.

[0123] Again, in use, the fan impact liner 45 as shown in FIG. 6 absorbs a substantial proportion of the kinetic energy of an impacting projectile, such as a detached fan blade. As a projectile begins to impact the fan impact liner 45, contact with the septum layer 57 distributes the impact force over a broad area, such that a large portion of the intermediate honeycomb layer 54 is deformed. As the intermediate honeycomb layer 54 deforms, energy is absorbed. Accordingly, the speed at which the projectile moves through the fan impact liner 45 is reduced. However, an impacting projectile such as a fan blade is not typically arrested fully in the honeycomb layer 54. Instead, as the projectile reaches the septum layer 56, the impact force is again spread over a large area of the radially outer honeycomb layer 53. In addition, the septum layer 56, being formed of titanium or titanium alloy, is itself highly resistant to deformation and contributes substantially to the arresting effect of the fan impact liner. The septum layer 56, being thin, does not contribute significantly to the overall weight of the fan impact liner, despite the higher density of titanium relative to the composite materials commonly used to form septum layers.

[0124] Both of the examples embodiments of fan impact liners shown in FIGS. 5 and 6 may be manufactured using standard techniques for forming composite structures known in the field. For example, the various layers of honeycomb structures and septa may be assembled in the appropriate order and then cured, for example, in an oven or an autoclave. Curable adhesives, such as epoxy resins, may be used to adhere adjacent layers to one another. Composite septa may be formed by laying up reinforcing fibre plies impregnated with adhesive, followed by curing. The fan impact liner 45 and the fan containment casing 42 may be manufactured separately, followed by adhering the fan impact liner to an inboard surface of the fan containment casing. Alternatively, the various layers of the fan impact liner 45 and the fan containment casing 42 may assembled and cured together.

[0125] 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.