COMBUSTOR TILE FOR A COMBUSTOR OF A GAS TURBINE ENGINE

Abstract

An annular combustor for a gas turbine engine including inner and outer combustor walls, wherein each wall defines an annulus and the inner wall is radially inward of the outer wall. The combustor includes a primary zone where the inner and outer combustor walls converge in a downstream direction, and a secondary zone downstream of the primary zone. In the secondary zone, the inner and outer walls are arranged to converge at a different rate to the primary zone, are non-convergent or are divergent in a downstream direction, a rate of change of radial width of the combustor is different in the zones. A transition is provided from the primary zone to the secondary zone. A plurality of combustor cooling tiles lines the inner and outer walls. One or more of the tiles are arranged to extend from the primary to secondary zone and across the transition from the zones.

Claims

1. An annular combustor for a gas turbine engine, the combustor comprising: an inner combustor wall and an outer combustor wall, the inner and outer combustor walls each define an annulus and the inner combustor wall is radially inward of the outer combustor wall; a primary zone, wherein within the primary zone the inner and outer combustor walls converge in a downstream direction; a secondary zone downstream of the primary zone, wherein in the secondary zone the inner and outer combustor walls are arranged to converge at a different rate to the primary zone, are non-convergent or are divergent in a downstream direction, such that a rate of change of radial width of the combustor is different in the primary zone to the secondary zone; a transition from the primary zone to the secondary zone; and a plurality of combustor cooling tiles lining the inner and outer combustor walls, wherein one or more of the tiles are arranged to extend from the primary zone to the secondary zone and across the transition from the primary zone to the secondary zone.

2. The annular combustor according to claim 1, wherein the tiles have a first portion provided in the primary zone, a second portion provided in the secondary zone and a transition region, the first portion, second portion and transition region being contiguous, and wherein the transition region of the tile has a greater radius of curvature than the inner and/or outer walls of the combustor at the transition between the primary and secondary zone.

3. The annular combustor according to claim 1, wherein a plurality of combustor cooling tiles are adjacently arranged in a circumferential direction to define an annulus that extends from the primary zone to the secondary zone across the transition from the primary zone to the secondary zone.

4. The annular combustor according to claim 1, wherein four tiles are adjacently arranged in a circumferential direction and together define a full annulus that extends from the primary zone to the secondary zone across the transition from the primary zone to the secondary zone of the inner and/or outer combustor wall.

5. The annular combustor according to claim 1, wherein the tile extending across the primary zone, secondary zone, and transition, extends across the entire primary zone and/or secondary zone.

6. The annular combustor according to claim 1, wherein the combustor cooling tiles are effusion tiles.

7. The annular combustor according to claim 1, wherein the inner and outer combustor walls are angled so as to reduce the radial width of the primary zone in a downstream direction.

8. The annular combustor according to claim 1, wherein the inner and outer combustor walls are angled so as to increase the radial width of the secondary zone in a downstream direction.

9. A combustor for a gas turbine engine, the combustor comprising: a primary zone; a secondary zone downstream of the primary zone; and a plurality of tiles lining the primary and secondary zone of the combustor, wherein a series of tiles are arranged to extend axially such that a portion of each of the tiles in said series of tiles is in the primary zone and a portion of said same tiles is in the secondary zone.

10. The combustor according to claim 9, wherein the combustor comprises an inner and outer combustor wall that converge at a first rate in the primary zone and are divergent, non-convergent or converge at a second rate in the secondary zone such that there is a change in axial direction of the walls at a transition between the primary and secondary zones, and wherein each of the tiles of the series of tiles are curved in a region coincident with the change in axial direction of the walls, the curve of the tiles having a greater radius than the change in direction of the walls.

11. A combustor for a gas turbine engine, the combustor comprising: inner and outer combustor walls defining a primary zone and a secondary zone, the inner and outer combustor walls in the primary zone being arranged at a first angle relative to each other, and the inner and outer combustor walls in the secondary zone being arranged at a second angle relative to each other, the second angle being different to the first angle; a fuel injector provided in the primary zone; an ignitor provided in the primary zone; a plurality of air inlets provided in the secondary zone for injecting air into the fit, combustor; and a plurality of combustor cooling tiles lining the combustor walls and connected thereto; wherein one or more of the combustor cooling tiles extend from the primary zone to the secondary zone such that a portion of the tile is in the primary zone, a portion of the tile is in the secondary zone and the tile extends across a transition from the primary zone to the secondary zone.

12. A gas turbine engine comprising the combustor according to claim 1.

13. A gas turbine engine for an aircraft comprising: an engine core comprising a turbine, a compressor, and a core shaft connecting the turbine to the compressor; a fan located upstream of the engine core, the fan comprising a plurality of fan blades; and a gearbox that receives an input from the core shaft and outputs drive to the fan so as to drive the fan at a lower rotational speed than the core shaft; a combustor provided downstream of the compressor and upstream of the turbine, the combustor being the combustor according to claim 1.

14. The gas turbine engine according to claim 13, wherein: the turbine is a first turbine, the compressor is a first compressor, and the core shaft is a first core shaft; the engine core further comprises a second turbine, a second compressor, and a second core shaft connecting the second turbine to the second compressor; and the second turbine, second compressor, and second core shaft are arranged to rotate at a higher rotational speed than the first core shaft.

15. A combustor tile for a gas turbine engine, the combustor tile comprising: a body having a first portion and a second portion, the first portion being contiguous with the second portion; and wherein the first portion is angled to the second portion by an angle between 185 and 210 degrees.

16. The combustor tile according to claim 15, wherein the first portion is angled to the second portion by an angle between 185 and 195 degrees.

17. The combustor tile according to claim 15, wherein a transition from the first portion to the second portion is curved such that there is a gradual change in angle of the tile from the first portion to the second portion.

Description

BRIEF DESCRIPTION OF THE DRAWINGS

[0057] Embodiments will now be described by way of example only, with reference to the Figures, in which:

[0058] FIG. 1 is a sectional side view of a gas turbine engine;

[0059] FIG. 2 is a close up sectional side view of an upstream portion of a gas turbine engine;

[0060] FIG. 3 is a partially cut-away view of a gearbox for a gas turbine engine;

[0061] FIG. 4 is a sectional side view of combustor equipment of a gas turbine engine;

[0062] FIG. 5 is a perspective view of a combustor cooling tile;

[0063] FIG. 6 is a sectional side view of alternative combustor equipment for a gas turbine engine;

[0064] FIG. 7 is a sectional side view of further alternative combustor equipment for a gas turbine engine;

[0065] FIG. 8 is a perspective view of a combustor cooling tile;

[0066] FIG. 9 is a sectional side view of further alternative combustor equipment for a gas turbine engine; and

[0067] FIG. 10 is a sectional side view of further alternative combustor equipment for a gas turbine engine.

DETAILED DESCRIPTION OF THE DISCLOSURE

[0068] 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.

[0069] 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.

[0070] 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.

[0071] 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.

[0072] 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.

[0073] 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.

[0074] 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.

[0075] 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.

[0076] Optionally, the gearbox may drive additional and/or alternative components (e.g. the intermediate pressure compressor and/or a booster compressor).

[0077] 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 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.

[0078] 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.

[0079] Referring to FIG. 4, exemplary combustion equipment 16 is illustrated. The combustion equipment is an annular combustor. The annular combustor has a radially inner wall 48 and a radially outer wall 49. The walls 48, 49 of the combustor define a primary zone 42 and a diffusion zone 44. The diffusion zone is downstream of the primary zone. In the present example, the radially inner and the radially outer walls of the combustor are arranged such that the radial width r of the combustor diverges and then converges in downstream direction.

[0080] A fuel injector 52 is provided at an upstream end of the combustor and an ignitor 54 is provided to create a spark in the primary zone 42 of the combustor. A diffuser 50 is provided upstream of the combustor.

[0081] In use, flow A from the compressor is directed through the diffuser 50 to the combustor and around the combustor. The diffuser reduces the axial velocity of the air received from the compressor. The flow of air around the combustor contributes to cooling the combustor and to control of the combustion process, as will be described later. Air enters the combustor at an upstream end of the primary zone 42. Fuel is injected into the primary zone from the fuel injector 52. Flow within the primary zone is turbulent to encourage mixing of the fuel and air mixture and to reduce the axial velocity of the fuel air mixture. The ignitor 56 ignites the fuel air mixture in the primary zone. Air inlets 56 (only one labelled) are provided in the inner and outer radial walls of the combustor in the region of the diffusion zone. In the present example the air inlets are holes but in alternative examples chutes may be provided. Airflow from the compressor enters the combustor through the air inlets 58 into the combustion chamber to cool the combustion gases G and to control emissions.

[0082] The temperature of the combustion gases G is higher than the melting point of the walls 48, 49 of the combustor. As such, cooling of the walls of the combustor is required. One method of achieving wall cooling is to line the walls of the combustor with tiles 60 (only one labelled). In the present example, a conventional gas turbine engine is illustrated, and there are a plurality of tiles provided in the primary zone and the diffusion zone, and the tiles are axially adjacent one another in a transition between the primary zone and the diffusion zone, such that there is no single tile extending from the primary zone to the diffusion zone. The tiles 60 may be mounted to the walls using for example mechanical fasteners. The tiles 60 may be metallic and coated in a thermal barrier coating, for example a ceramic coating, or may be made from ceramic. Cooling holes are provided in the combustor walls and cooling air flows through these holes and impinges on the tiles. An example tile is shown in more detail in FIG. 5.

[0083] With reference to FIG. 5, the tile 60 includes a body 62 and a series of pedestals 64 protruding radially from the body. When mounted on the combustor wall, the pedestals are arranged to protrude towards the wall 48, 49 to which the tile is mounted. In use, cooling flow from the compressor flows through the wall 48, 49 of the combustor and impinges on the body 62 of the tile. The flow then flows through the pedestals before exiting the region of the tile. The pedestals increase convective cooling. As will be described later, an alternative cooling tile may not have pedestals and may be what is referred to in the art as an effusion tile.

[0084] Referring now to FIG. 6, the walls 48, 49 of the combustor may be arranged such that the radial width of the combustor decreases from an upstream end to a downstream end of the primary zone 42. That is the walls 48, 49 converge in a downstream direction. In the present example, the walls of the combustor in the primary zone are angled so as to reduce the radial width of the primary zone of the combustor in a downstream direction, but in alternative examples the walls may be curved to reduce the radial width of the primary zone. A secondary zone 44 is provided downstream of the primary zone. In the present example the secondary zone may be considered the diffusion zone. In the present example, the walls of the combustor in the secondary zone 44 are substantially parallel such that the radial width of the secondary zone is substantially constant. For example, the walls can be considered as non-convergent in the secondary zone. In the present example, the angle of the inner and outer walls changes by a step change at the transition 66. However, in alternative embodiments, the change in angle may be more gradual, e.g. the transition 66 may be curved to provide a gradual change in angle instead of being a step change.

[0085] A tile 60a is provided at the transition 66. The tile 60a extends from the primary zone 42, across the transition 66, to the secondary zone 44. A series of tiles are provided circumferentially adjacent to each other so as to define an annulus that extends from the primary zone to the secondary zone across the transition. The junctions between the tiles in the region of the transitions are all axially extending, i.e. there are no circumferentially extending junctions between tiles at the transition.

[0086] Referring to FIGS. 6 and 8, the tile 60a includes a first portion 70 provided in the primary zone and a second portion 72 provided in the secondary zone. A transition region 68 is provided between the first portion and the second portion. In the present example the tile is an effusion tile, but in alternative examples the tile may include pedestals. By way of example only, in the present example, the angle between the first portion and the second portion is between (and including) 185 and 210 degrees (e.g. 185 to 195). The angle is measured from the combustion gas washed side of the tile (i.e. the side of the tile exposed to the combustion gases).

[0087] The transition 66 in the walls 48, 49 of the combustor creates a kink which can be a location of high stress concentration. The tile 60a may be arranged such that the transition 68 between the first portion (i.e. from the primary zone) to the second portion (i.e. to the secondary zone) is a smoother transition than on the combustor walls. For example, instead of a step change there may be a curved transition as illustrated in FIG. 8. Such a transition may be easier to manufacture on the tiles than on the combustor walls themselves.

[0088] The provision of a smooth transition 68 from the primary zone to the secondary zone reduces the risk of a high stress concentration at the transition. Further the provision of a single tile that extends from the primary zone to the secondary zone removes any possible cooling film disruption at the transition 66 of the combustor walls 48,49 which could otherwise be caused by a circumferentially extending joint between tiles at the transition. Removing the risk of cooling film disruption improves cooling of the combustor walls and reduces the risk of failure of the combustor walls.

[0089] Additional tiles 60b, 60c may be placed axially adjacent the tiles 60a. For example, a plurality of tiles forming an annulus may be provided entirely in the primary zone, a plurality of tiles forming an annulus may be provided entirely in the secondary zone, and the annulus of tiles that extend across the transition 66 of the combustor walls 48, 49 may be axially adjacent (and optionally overlapping) the tiles that are entirely in the primary zone and the tiles that are entirely within the secondary zone.

[0090] Referring to FIG. 7, the tiles 60 that extend across the transition 66 of the combustor walls 48, 49 may have a first portion 70 that extends substantially the entire axial length of the primary zone and a second portion 72 that extends substantially the entire axial length of the secondary zone. The transition 68 between the first portion 70 and the second portion 72 may be more curved than the transition 66 of the combustor walls. In such an embodiment a plurality of tiles 60 may be arranged circumferentially adjacent to each other to define an annulus. In some embodiments only two tiles may be provided, the two tiles defining the entire annulus. Alternatively only three tiles may be provided, the three tiles defining the entire annulus, or further alternatively only four tiles may be provided, the four tiles defining the entire annulus.

[0091] Conventionally, the upstream end of the diffusion zone is defined as the position where bulk air enters the combustor for the control of combustion gas temperature. In the present examples, the bulk air enters the combustor in the region of the transition between the primary and secondary zone. Referring to FIG. 9, in some examples the bulk air may enter the combustor through air inlets provided in the primary zone. In this example the air inlets are holes 74. These air inlets are provided upstream of the change in convergence/divergence of the combustor walls, i.e. upstream of the secondary zone. However, in other examples the bulk air inlets may alternatively or additionally be provided downstream of the change in convergence/divergence of the combustor walls, or in some examples may be provided in the transition region. In FIG. 9, chutes 76 are example air inlets provided downstream of the change in convergence/divergence (or in other words change in axial direction) of the combustor walls. The bulk air inlets differ from the impingement holes 78, because the air from the impingement holes impinges on the tiles 60a, 60b, 60c, rather than entering the combustor to cool the combustor gases. In the present example, the impingement holes are smaller in diameter than the bulk air inlets.

[0092] In the described embodiments, the walls of the combustor have been angled to change the radial width of the combustor. However, referring to FIG. 10, the walls may be curved to achieve the desired change in radial width of the combustor. In the example shown in FIG. 10, the walls are convergent in the primary zone and divergent in the secondary zone. The walls 48, 49 of the combustor are curved at the transition 66 between the primary zone and the secondary zone so that there is a more gradual change in radial width of the combustor than with a step change.

[0093] 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.