COMBUSTION LINER AND GAS TURBINE ENGINE COMPRISING A COMBUSTION LINER

Abstract

A combustion liner for a gas turbine engine. The combustion liner defines a bypass direction which extends between an upstream portion and a downstream portion of the combustion liner in use, the combustion liner. The combustion liner comprises a liner wall for defining at least a portion of a combustion chamber, the liner wall having an outer surface and an inner surface and a depth between the outer and inner surfaces of the combustion liner in use. The combustion liner comprises a chute formed through the liner wall for conveying fluid from the outer surface through the liner wall and ejecting fluid from an exhaust hole of the chute on the inner surface, the exhaust hole having a length L along the bypass direction. The combustion liner also comprises a fluid-guiding surface at an upstream side of the chute which defines an arc terminating at the exhaust hole, the fluid guiding surface configured to guide fluid from a direction generally parallel to the outer surface into the chute to be ejected from the chute exhaust hole at a direction generally perpendicular to the inner surface.

Claims

1. A combustion liner for a gas turbine engine, the combustion liner defining a bypass direction which extends between an upstream portion and a downstream portion of the combustion liner, the combustion liner comprising: a liner wall for defining at least a portion of a combustion chamber, the liner wall having an outer surface and an inner surface and a depth D between the outer and inner surfaces of the combustion liner; a chute formed through the liner wall for conveying fluid from the outer surface through the liner wall and ejecting fluid from an exhaust hole of the chute on the inner surface, the exhaust hole having a length L parallel to the bypass direction; and a fluid-guiding surface on the outer surface of the chute, the fluid guiding surface defines an arc terminating at the exhaust hole, the fluid guiding surface configured to guide fluid from a direction generally parallel to the outer surface into the chute to be ejected from the chute exhaust hole at a direction generally perpendicular or upstream, relative to the inner surface, the arc comprising a minimum radius to prevent boundary layer separation of the fluid.

2. The combustion liner as claimed in claim 1, wherein the fluid guiding surface defines an arc having a minimum radius R in a plane aligned with the bypass direction and extending vertically between the outer and inner surfaces of the liner wall, and wherein the radius R of the arc of the fluid-guiding surface is at least 50% of the length L of the exhaust hole of the chute.

3. The combustion liner as claimed in claim 2, wherein the radius R is between 50% and 150% of the length L of the exhaust hole of the chute.

4. The combustion liner as claimed in claim 2, wherein the radius R is between 75% and 125% of the length L of the exhaust hole of the chute.

5. The combustion liner as claimed in claim 2, wherein the radius R is between 90% and 110% of the length L of the exhaust hole of the chute.

6. The combustion liner as claimed in claim 2, wherein the radius R is substantially equal to the length L of the exhaust hole of the chute.

7. The combustion liner as claimed in claim 1, wherein the arc extends through at least 90 degrees.

8. The combustion liner as claimed in claim 1, wherein the arc terminates at the exhaust hole at an angle of 90 degrees to the inner surface or a plane parallel thereto.

9. The combustion liner as claimed in claim 1, wherein the radius R is substantially equal to the depth D of the liner wall, such that the outer surface upstream of the chute forms a tangent to the arc of the fluid-guiding surface.

10. The combustion liner as claimed in claim 1, wherein the radius R is greater than the depth D of the liner wall, such that the fluid-guiding surface forms a protrusion from the outer surface of the liner wall upstream of the chute.

11. The combustion liner as claimed in claim 1, further comprising a fluid deflector arranged on a downstream side of the chute on the outer surface of the liner wall, the fluid deflector being configured to deflect fluid into the chute.

12. The combustion liner as claimed in claim 11, wherein the fluid deflector is a scoop-like element configured to deflect fluid along an arcuate path into the chute.

13. The combustion liner as claimed in claim 1, further comprising an exhaust deflector surface arranged in the chute proximate the exhaust hole for deflecting fluid in the chute at least partially upstream relative to the bypass direction during ejection from the chute exhaust hole.

14. The combustion liner as claimed in claim 13, wherein the exhaust deflector surface is arranged on a downstream side of the chute and extends at least partially upstream relative to the bypass direction such that an acute angle is formed between the deflector surface and the inner surface at the exhaust hole.

15. The combustion liner as claimed in claim 14, wherein the exhaust deflector surface is formed on a projection which extends laterally into the chute.

16. The combustor liner as claimed in claim 13, wherein the exhaust deflector surface is arranged on an upstream side of the chute and extends at least partially upstream relative to the bypass direction such that an obtuse angle is formed between the exhaust deflector surface and the inner surface.

17. The combustor liner as claimed in claim 1, wherein the chute and/or the exhaust hole is elliptical, having a major axis of the elliptical shape extending in the bypass direction.

18. A combustion assembly for a gas turbine engine comprising the combustor liner as claimed in claim 1.

19. 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 the combustion assembly as claimed in claim 18; and optionally further comprising a gearbox that receives an input from the core shaft and outputs drive to the fan to drive the fan at a lower rotational speed than the core shaft.

Description

BRIEF DESCRIPTION OF THE DRAWINGS

[0063] Embodiments will now be described by way of example only, with reference to the Figures, in which:

[0064] FIG. 1 is a sectional side view of a gas turbine engine;

[0065] FIG. 2 is a close up sectional side view of an upstream portion of a gas turbine engine;

[0066] FIG. 3 is a partially cut-away view of a gearbox for a gas turbine engine;

[0067] FIG. 4 is a detailed cross-sectional view of a combustion apparatus of the gas turbine engine;

[0068] FIG. 5 is a detailed cross-sectional view of a combustion liner comprising a chute;

[0069] FIG. 6 is a plan view of the combustion liner of FIG. 5;

[0070] FIG. 7 is a bottom view of the combustion liner of FIG. 5;

[0071] FIG. 8 is a cross-sectional view of the combustion liner of FIG. 5;

[0072] FIG. 9 is a detailed cross-sectional view of an alternative combustion liner comprising a chute; and

[0073] FIG. 10 is an image from a slightly upstream position of an alternative combustion liner showing a fluid guiding surface and the outer surface.

DETAILED DESCRIPTION OF THE DISCLOSURE

[0074] 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.

[0075] 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.

[0076] 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 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.

[0077] 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.

[0078] 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 disclosure. Practical applications of a planetary epicyclic gearbox 30 generally comprise at least three planet gears 32.

[0079] 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.

[0080] 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.

[0081] 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.

[0082] Optionally, the gearbox may drive additional and/or alternative components (e.g. the intermediate pressure compressor and/or a booster compressor).

[0083] 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.

[0084] FIG. 4 shows a cross-sectional view of combustion apparatus 100 of the combustion equipment 16 of the gas turbine engine 10.

[0085] The combustion apparatus 100 comprises a fuel injector 102 which is configured to inject fuel into a combustion chamber 104. The combustion chamber 104 is formed by a combustion liner 106 which defines the size and shape of the combustion chamber 104. The combustion liner is substantially annular in cross section and defines a generally tube- shaped combustion chamber 104 therein. The combustion chamber 104 defines an axis, having an axial direction which runs along the centre of the combustion chamber 104. The term axial direction may also apply to a combustion liner in isolation from a combustion chamber, in which case it may refer to the axis the liner is intended to be arranged around when used in a combustion chamber. The combustion liner, when incorporated into a combustion chamber or combustion apparatus may be spaced from the combustion chamber axis in the radial direction. The term radial direction as used herein may refer to the direction perpendicular to the inner or outer surfaces of the liner. This may be equivalent to the radial direction from the combustion chamber axis were the liner to be incorporated into a combustion chamber. Similarly, the terms circumferential or lateral direction may refer to the direction perpendicular to the radial direction and in the plane defined by the inner or outer layers. The combustion liner 106 may be arranged in an annular arrangement with the bypass channel 108 formed around it, such that air flowing into the combustion equipment 16 flows along the combustion outer surface of the combustion liner 106. The flow of air through the combustion apparatus 100 is illustrated by the solid arrows shown on FIG. 4. In use, the outer surface bounds part of the bypass channel and directs bypass air flow roughly parallel to the outer surface. The outer surface 116 of the combustion liner defines a bypass direction parallel to the outer surface, extending from the upstream end of the liner to the downstream end. When the combustion liner is assembled into a combustion apparatus, the bypass direction may be parallel with the axial direction or may radially diverge from the axial direction in the downstream direction.

[0086] The fuel injector 102 introduces fuel (as shown by the dotted arrow) into the combustion chamber 104 at the upstream end of the chamber 104. The combustion liner 104 and the fuel injector 102 may comprise apertures and mixing equipment at the upstream end of the liner 106 for air flowing into the combustion equipment 16 to enter the combustion chamber 104 and mix with the fuel. The mixture of fuel and air is ignited in the combustion chamber. The combusting fuel-air mix travels downstream and exits the combustion chamber at high speed, temperature, and pressure (as illustrated by the dashed arrows) to drive the turbines. It will be understood that the combustion liner 106, when assembled into combustion apparatus, generally extends in an axial direction in the engine, from upstream to downstream.

[0087] One or a plurality of chutes 110 may be provided through the combustion liner 106. The chutes 110 allow air from the bypass channels 108 to enter the combustion chamber 104. The chutes inject a jet of air into the combustion chamber 104, which contains a hot, fast-flowing fuel-gas mix. The further the air from the chutes is injected into the centre of the combustion chamber, the better the mixing of the fuel-gas mix.

[0088] An example of a chute 110 of the combustion liner 106 is shown in FIGS. 5-8. FIG. 5 is a cross-sectional view of the combustion liner 106 and chute 110 as marked by box X in FIG. 4. FIG. 5 is a cross-sectional view of the chute 110 and liner 106 in the plane X-X as illustrated in FIG. 6. The plane on which the view of FIG. 5 extends parallel to the bypass direction A and bisects the chute 110 in the bypass direction. The plane extends vertically (or radially with respect to the combustion chamber) and generally perpendicularly between the outer and inner surfaces 116, 118 of the liner 106.

[0089] As can be seen clearly in FIG. 5, the combustion liner 106 comprises a combustion liner wall 111 of two-layer construction, having an outer layer 112 which faces the bypass channel 108 and an inner layer 114 which faces the combustion chamber 104. Each layer may be constructed differently to optimise its characteristics; for example, the inner layer 114 may have better heat-resistant properties than the outer layer 112 because it is exposed to the high temperatures in the combustion chamber 104. In this and other examples, cooling air flow may be passed between the layers 112, 114. In some other examples, the combustion liner may instead be of single layer construction or may comprise more than two layers. The outer layer 112 defines an outer surface 116 of the combustion liner wall 111, while the inner layer 114 defines an inner surface 118 of the combustion liner wall 111. The radial or vertical distance between the upper and lower surfaces 116, 118 defines the depth D of the combustion liner wall 111.

[0090] The chute 110 extends through the combustion liner wall 111 from the upper surface 116 to the lower surface 118. The chute 110 is configured to convey air from the bypass channel 108 at the outer surface 116 through the liner wall 111 and eject it from an exhaust hole 120 of the chute 110 formed on the inner surface 118. In this example, a protruding rim 122 is formed around the exhaust hole 120 on the inner surface 118 to improve flow characteristics in the combustion chamber 104 around the exhaust hole 120, but this rim may not be provided in other embodiments (see FIG. 9 for example). The low profile and angled edges of the protruding rim 122 may be less susceptible to erosion in the combustion chamber 104 and may reduce disruption to the internal film cooling of the inner surface 118 of the liner 106, thereby reducing the overall cooling requirement for the combustion apparatus 16.

[0091] Referring to FIG. 7, in the exemplary embodiment shown, the exhaust hole 120 is substantially elliptical. The exhaust hole 120 has a length L parallel with the bypass direction, and a shorter width W along the transverse direction, perpendicular to the bypass direction. In other examples, the exhaust hole may have other shapes, (e.g. circular or polygonal). Irrespective of the shape of the exhaust hole, the exhaust hole will still have a length L which is the maximum dimension of the hole measured parallel to the bypass direction. In the case of a circular hole, the length L is the diameter of the hole. The elliptical shape of the exhaust hole 120 and chute 110 may reduce the overall footprint of the chute 110 on the liner 106, may improve jet penetration into the chamber 104 and may also reduce disruption to the cooling flow on the hot side of the combustor wall. An elliptical hole may also further improve ease of manufacture by additive layer manufacture.

[0092] Referring to FIG. 5, a fluid-guiding surface 124 is shown upstream of the chute 110. The fluid-guiding surface 124 is configured to guide fluid from the outer surface 116 into the chute 110. The arrows in FIG. 5 illustrate the direction of airflow around and through the chute 110. As will be appreciated, the prevailing direction of air flow on the outer surface is in the bypass direction A as shown. However, on exiting the chute 110, the prevailing direction of the injecting air jet is substantially perpendicular to the inner surface 118, or partially upstream. Therefore, the fluid-guiding surface 124 is configured to gradually redirect air from the outer surface through a rotation to be ejected from the chute exhaust hole 120. Fluid is preferably ejected close to perpendicular to the inner surface (or bypass direction) or angled partially upstream. The fluid-guiding surface 124 gradually redirects air to prevent separation of the airflow from the fluid guiding surface.

[0093] The embodied fluid-guiding surface 124 defines an arc 126 that extends from the inner surface to the exit hole in the plane X-X (as shown in FIG. 6) and has a radius R. Computational fluid dynamics simulations may be used to determine an appropriate minimum radius for particular operating conditions and hole size, such that flow separation does not occur. In embodiments, the minimum radius R of the arc of the fluid-guiding surface 124 may be at least 50% of the length L of the exhaust hole 120 of the chute 110. If the radius of the arc varies along its length, then the minimum radius of the arc may be at least 50% of the length L of the exhaust hole 120 of the chute 110. By providing a fluid-guiding surface having at least this radius, it has been found that the airflow over the fluid-guiding surface may be less likely to separate from the surface 124 during travel into the chute 110 and through the necessary rotation, and may have an increased velocity entering the combustion chamber 104 and an improved direction (i.e. closer to perpendicular or partially upstream). This may improve the penetration depth of the air jet into the combustion chamber 104 from the exhaust hole 120 and may also improve mixing of the air jet from the chute 100 into the combustion chamber flow, thereby improving combustion conditions and combustion efficiency. The improved depth of air jet penetration and mixing may also enable the radius of the combustion chamber to be increased and the length of the combustor reduced. This may enable the axial length of the entire engine to be decreased. The improved penetration of airflow may also obviate the need for guide chutes that extend a substantial distance inwardly into the combustor from the inner surface. Such guide chutes have been used to increase penetration depth of the air jet, but they are exposed to harsher conditions in the centre of the combustor and may require more frequent maintenance.

[0094] In an embodiment the radius R of the arc 126 of the fluid-guiding surface 124 may be substantially equal to the length L of the exhaust hole 120 of the chute, but more generally the radius R may be optionally between 50-150%, 60-140%, 75-125% 90-110% of the length L of the exhaust hole 120.

[0095] Referring to FIG. 5, the arc 124 terminates at the exhaust hole 120 at an angle of perpendicular to a plane parallel to the inner surface. The angle between the plane and the arc may be close to perpendicular, i.e. within 15 degrees, 10 degrees, 5 degrees or 2 degrees. Or, the angle may be greater than 90 degrees, i.e. the angle may direct the jet partially upstream. The term terminates refers to the end of the arc, starting from the upstream end to the downstream end. The angle is taken as a tangent from the arc where it terminates and is measured in the axial-radial plane through the bisection of the chute as shown in FIG. 5.

[0096] As will be appreciated from FIG. 6, for example, the fluid-guiding surface 124 may have a different profile to the arc 126 at the extremities of its lateral width, but generally the central area or the fluid-guiding surface 124 in the bypass direction A may conform closely to the arc 126 and may have a radius substantially similar to R.

[0097] In the embodiment of FIG. 5, to provide a fluid-guiding surface having a sufficiently large radius R, the fluid-guiding surface 124 forms a raised aerodynamic bump-like protrusion 128 upstream of the chute 110. That is the fluid guiding surface starts by deviating upwards away from the inner surface before arcing downwards towards the inner surface. This is because the depth D of the liner wall 111 is less than the radius R. However, as the radius R is sufficiently large, the airflow from the outer surface 116 may stay attached to the fluid-guiding surface 124 as it rises over the protrusion 128 and into the chute 110. In other embodiments, e.g. as shown in FIG. 9, the depth D of the liner wall may be sufficient that it is equal to or larger than the radius R, such that the fluid guiding surface begins with the outer surface of the liner as a tangent, then arcs towards the inner surface.

[0098] In addition to the fluid-guiding surface 124, the liner 106 and chute 110 may have further features which promote improved air jet penetration from the exhaust hole 120 of the chute 110. In some aspects, these further features may be provided in addition to the fluid-guiding surface 124 or, in some aspects, may be provided without the fluid guiding surface.

[0099] Referring to FIG. 5, the embodiment shown comprises an exhaust deflector surface 130 arranged in the chute 110 proximate the exhaust hole 120 for deflecting a portion of fluid in the chute 110 at least partially upstream during ejection from the inner surface 118. The exhaust deflector surface 130 is formed as a projection or surface which extends from the downstream wall of the chute. The exhaust deflector surface may be positioned immediately adjacent the exhaust hole 120 or may be positioned further radially outwards towards the outer surface. The exhaust deflector surface may be angled obliquely or acutely relative to the downstream side of the inner surface 118 and angled partially upstream. Alternatively, the exhaust deflector may comprise a gradual curve extending from the downstream wall of the chute, also arranged to direct a portion of the airflow upstream. The exhaust surface 130 may extend at least partially upstream relative to the axial direction/bypass direction from the downstream side of the chute. Air travelling over or proximate to the downstream side of the chute 110 will be redirected by an exhaust deflection surface partially in the upstream direction as it is ejected from the exhaust hole 120 (as shown by the airflow arrow in FIG. 5). This direction may further delay the jet being consumed by the prevailing airflow in the combustion chamber and may increase the penetration depth of the jet into the chamber 104, improving the mixing of the air jet in the chamber 104. It should be appreciated that the exhaust deflector surface may also improve jet penetration without being combined with a fluid-guiding surface 124 or a fluid deflector 132.

[0100] In other embodiments, additionally or alternatively, an exhaust deflector surface may be arranged in a circumferential position between the inner and outer surfaces.

[0101] A further feature which may improve the jet penetration and mixing of the chute 110 optionally in isolation from or in combination with the fluid-guiding surface 124 and/or exhaust deflection surface 130, is the fluid deflector 132. The fluid deflector 132 is arranged on a downstream side of the chute 110 on the outer surface 116 of the liner wall 111 and is shaped to deflect airflow into the chute 110.

[0102] The fluid deflector 132 may be shaped to redirect airflow which is more distant from the outer surface 116 and is therefore not guided by the fluid guiding-surface 124 into the chute 110 through boundary layer attachment. The fluid deflector may be planar, angled, curved or a scoop-like structure that projects upwards from the outer surface and optionally extends at least partially over the chute. The fluid deflector 132, as shown in FIG. 5, is scoop-shaped and substantially arcuate in cross-section, the arcuate scoop extending upstream in the bypass direction A. Accordingly, air which travels towards the deflector 132 will be directed into the chute 110, where in the absence of the deflector, the airflow would have bypassed the chute. More generally, the deflector 132 may be configured to redirect fluid along an arcuate path into the chute 110.

[0103] FIG. 8 shows the liner 106 as viewed into the plane Y-Y shown in FIG. 6. This view is therefore along the bypass direction A, i.e. the prevailing direction of airflow over the outer surface 116.

[0104] Referring to FIG. 8, the scoop-like fluid deflector 132 protrudes a significant distance from the outer surface 116 of the liner wall 111 and from the bump-like protrusion 128 of the fluid-guiding surface. The fluid deflector 132 therefore defines a scoop 134 which extends into the airflow above the outer surface 116 to inhibit flow of air in the bypass direction A past the chute. Accordingly, the fluid deflector 132 captures and directs airflow which would otherwise have simply travelled past the chute 110 in the absence of the fluid deflector 132. The fluid deflector 132 therefore increases the capture of air flow, increasing flow rate of air into the chute 110, and increasing jet penetration into the chamber 104.

[0105] Furthermore, the fluid deflector 132 alters the pressure field of air around the chute 110 which can particularly assist the operation of other features of the present disclosure. For example, the pressure field created by the fluid deflector 132 may augment the attachment of the airflow to the fluid-guiding surface 124 and the upstream side of the chute 110, thereby improving jet penetration and fuel to air mixing.

[0106] FIG. 9 shows an alternative example of a combustion liner 206. Like features of the liner 206 and the liner 106 are indicated with references differing by 100.

[0107] The liner wall 211 of the liner 206 has a depth D which is greater than the depth D of the liner 106. Accordingly, the liner 206 can accommodate a fluid-guiding surface 224 which has a greater radius R without the need for a bump-like protrusion upstream of the chute 110. As a larger radius R can be achieved relative to the length L of the exhaust hole 220, this may provide yet further improved flow attachment of air entering the chute 210. Furthermore, as no bump-like protrusion is needed, in some examples, air may travel more smoothly into the chute 210 from the outer surface 218 as the outer surface 218 forms a tangent to the arc of the fluid-guiding surface 224, which may further increase jet penetration and mixing.

[0108] The example liner 206 of FIG. 9 entirely omits any protrusion from the inner surface 218 of the liner wall 211. Accordingly, the liner 206 may have further improved resistance to erosion from the combustion flow in the combustion chamber 204 and may have further improved internal film cooling on the inner surface 218 proximate the exhaust hole 220.

[0109] Referring to FIG. 10 a fluid guiding surface 324 is shown on the outer surface 316 of a combustion liner. The fluid guiding surface is shown extending around the chute 310. Fluid guiding surface transitions from a smooth gradual curve with a minimum radius to an angled wall 332 proximate the downstream-most point of the fluid guiding surface 301. The wall 332 forms the fluid deflecting surface. The fluid guiding surface may extend around a circumference of the chute as shown in FIG. 9. The fluid guiding surface may alternatively extend partway around the circumference of the chute. In some embodiments, this may be through an angle no less than 45 degrees, or an angle no less than 90 degrees, or an angle no less 135 degrees, or an angle no less than 180 degrees or an angle no less than 270 degrees around the chute. This may improve the liners sensitivity to directionality of airflow. The angle through which the fluid guiding surface extends around the circumference is understood to be measured from the centre point of the chute in a plane incident with the outer surface, and the midpoint of the angle is aligned with the apex of the upstream side of the chute 302.

[0110] The combustion liner may be formed from metallic alloys, carbon matrix composites or ceramics. The combustion liner may be made by additive layer manufacture. In embodiments, the combustion liner may be formed using additive layer manufacture building from the downstream end to the upstream end, or from the upstream end to the downstream end.

[0111] It will be understood that the invention is not limited to the examples 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.