GB2589885A - Combustor - Google Patents

Abstract

A combustor 41’ for a gas turbine engine has an inner liner 80, an outer liner 81, an annular metering panel 51’, and between which is a combustion chamber 42. The metering panel has an inner circumferential sealing surface 82 sealed to an upstream end of the inner liner, and an outer circumferential sealing surface 83 sealed to an upstream end of the outer liner. The metering panel includes a plurality of apertures 84 each for receipt of a respective fuel spray nozzle 58. The metering panel can have a convex gradient between the inner and outer circumferential sealing surfaces, and the metering panel can be segmented into a plurality of annulus sectors (96 Fig. 9). The metering panel may include cooling channels and have a downstream face that is coated with a thermal barrier coating, and additionally may have an upstream face that has a surface texture to increase the upstream surface area. The metering panel may be formed my additive layer manufacturing, and the arrangement provides a metering panel with improved temperature capabilities.

Description

COMBUSTOR

Field of the Disclosure

The present disclosure relates to the field of gas turbine engines. In particular, the present invention relates to a combustor and combustion equipment for a gas turbine engine.

Background of the Disclosure

The core of a gas turbine engine has, in axial flow series, a compressor section, combustion equipment and a turbine section. Figure 1 shows a longitudinal cross-section through conventional combustion equipment 40 for a gas turbine engine, the equipment being coaxial with a principle rotational axis of the engine. Block arrows show general directions of compressed air flow through the equipment.

The combustion equipment 40 has a combustor 41 in the form of a canister which defines a combustion chamber 42. Surrounding the combustor 41 is a casing enclosure formed from: at the front (upstream end), a high pressure outlet guide vane assembly (HP OGV assembly) 43; on the outside, an annular combustion chamber inner casing (CCIC) 44; on the inside, an annular combustion chamber outer casing (CCOC) 45; and at the rear (downstream end), a high pressure nozzle guide vane assembly (HP NGV assembly) 46.

The HP OGV assembly 43 has a circumferential row of high pressure outlet guide vanes (HP OGVs) 50 at an outlet from the engine compressor section. Similarly, the HP NGV assembly 46 has a circumferential row of high pressure nozzle guide vanes (HP NGV) 52 at an inlet to the engine turbine section. A circumferential row of fuel injectors 54 having spray nozzles 58 which spray fuel into the combustion chamber 42 penetrate through the CCOC 45 and a head wall 53 at the front end of the combustor 41.

Figure 2 shows a conventional combustor head wall 53 in more detail. The head wall 53 comprises a metering panel 51 and a heat shield 55. The metering panel 51 supports the fuel spray nozzles 58 and reacts loads arising from differential movements between the CCOC 45 and CCIC 47/48. The heat shield 55 is provided to protect the metering panel 51 from radiative and convector heat from the combustion chamber 42. It is typically cast from high temperature nickel alloy and has a thermal barrier coating on its downstream (hot) surface to increase its heat shielding effect. It has pedestals on its upstream surface which are fed with cooling air through laser drilled cooling holes provided in the metering panel 51 (which is typically a machined ring rolled high temperature nickel alloy forging). The heat shield 55 is typically secured to the metering panel 51 via threaded studs and nuts.

In use, compressed air passes through the HP OGVs 50 and enters the combustion chamber 42 to mix with fuel sprayed into the combustion chamber 42 by the fuel spray nozzles 58 of the injectors 54. The resulting air-fuel mixture combusts, and hot gas products of the combustion reaction exit the combustion chamber 42 to the turbine section via the HP NGVs 52.

Thermal responses cause radial movement of the fuel spray nozzles 58 and the head wall 53. Head seals 59 are provided between the fuel spray nozzles 58 and the head wall 53 to prevent uncontrolled flow of air into the combustion chamber 42. Each head seal 59 is typically cast in two parts -an upstream and a downstream part which are brazed together on either side of the heat shield 55. They are typically cast of the same high temperature nickel alloy as the heat shield 55.

Other flows of compressed air leaving the HP OGVs 50 are directed, as indicated, by a cowl 49 extending from the head wall 53 between the combustor 41 and the CCOC 45, between the combustor and the CCIC 44, and inside the CCIC. The spacing between the HP OGVs 50 and the cowl 49 is typically called the "dump gap" 60 and this is sized so as to reduce the dynamic pressure of the air expelled from the HP OGVs 50 to an acceptable Mach number to ensure a smooth transition into the annuli between the combustor 41 and the CCOC 45/CCIC 44. This reduction in the Mach number of the flow exiting the HP OGVs 50 results in an undesirable reduction in pressure energy available for the fuel spray nozzle 58 mixing.

The dump gap 60 has a direct impact on engine length. The spacing between the cowl 49 and the metering panel 51 also adds to the engine length and is functionally redundant.

There are a number of drawbacks of the current head wall design: 1) the temperature capability of the high temperature nickel alloy used to produce the heat shield and the pedestal cooling system will be unlikely to meet temperature requirements for future combustors; 2) the pedestal cooling system and the studs/nuts used to secure the metering panel and heat shield add significant weight and complexity to the current design; 3) the nuts have been found to seize and shear the studs from the heat shield during servicing thus reducing the life of the heat shield; 4) the current design is expensive due to the large forging required to produce the metering panel and complexity of the cast design for the heat shield- 5) assembly of the heat shield is difficult due to the restricted access to the downstream side of the metering panel; and 6) the head wall is restricted to a planar or conical shape as a result of the current manufacturing processes meaning that it cannot be easily shaped to match the shape to the combustion gases flow field to reduce unwanted recirculation zones and thus reduce emissions.

A need exists for combustion equipment which addresses one or more of these drawbacks. Summary of the Disclosure In a first aspect there is provided a combustor for a gas turbine engine, the combustor comprising: an inner liner and an outer liner defining a combustion chamber therebetween; and an annular metering panel having an inner circumferential sealing surface sealed to an upstream end of the inner liner and an outer circumferential sealing surface sealed to an upstream end of the outer liner, the metering panel defining a plurality of apertures each for receipt of a respective fuel spray nozzle, wherein the metering panel has a convex gradient between the inner and outer circumferential sealing surfaces.

By providing a metering panel having a convex gradient between the inner and outer circumferential sealing surfaces, it is possible to dispense with the prior art cowl as the curved outer (upstream) surface of the metering panel can direct air exiting the HP OGVs into the annuli between the combustion chamber and the CCIC/CCOC. The absence of the cowl also allows the engine length to be reduced (e.g. by around 15mm) as the functionally redundant space between the cowl and metering panel is eliminated. The reduction in engine length leads directly to a weight reduction through material removal. The part count during assembly is also advantageously reduced. Furthermore, the elimination of the cowl allows the HPOGVs to be positioned closer to the fuel spray nozzles allowing an increase in dynamic pressure to the fuel spray nozzle for improved atomisation and mixing of fuel and emission reductions.

In some embodiments, the metering panel has a convex gradient such that it forms an arc between the inner and outer circumferential sealing surfaces. The arc may subtend an angle of between 100 and 140 degrees e.g. around 1200 degrees.

The plurality of apertures may be provided midway between the sealing surfaces i.e. at the axial centre of the arc. Thus, the metering panel may comprise a first convex portion extending from each aperture to the inner sealing surface and a second convex portion extending from each aperture to the outer sealing surface. The first and second convex portions may have equal and opposite convex gradients.

In some embodiments, the inner and outer liners also comprise a respective convex portion proximal the join with the inner/outer circumferentially sealing surfaces of the metering panel respectively. The combined arc created by the convex metering panel and the convex liner portions may subtend an angle of between 250 and 290 degrees e.g. around 270 degrees.

This provides a combustor chamber volume proximal each aperture that is substantially spherical. This combustion chamber shape reduces corner recirculation of the combustion gases in each sector which leads to a narrower population distribution of residence times which improves emissions. The substantially spherical volumes around each aperture intersect around the circumference of the combustion chamber. In this manner, the circumferential sealing surfaces and abutting end surfaces of the inner/outer liners may be undulating.

In some embodiments, the inner and/or outer circumferential sealing surface may each comprise a plurality of fixing elements (e.g. fixing lugs or fixing nuts) for alignment/cooperation with a respective one of a plurality of corresponding fixing elements (e.g. fixing lugs/nuts) on the inner and/or outer liners respectively. The aligned pairs of fixing elements/lugs may each receive a fastener such as an axial fastener (e.g. a threaded bolt) to seal the metering panel to the liners. One or both of the fixing elements/lugs/nuts in each pair may have a threaded bore for cooperation with the axial fastener/threaded bolt. The fixing elements are provided externally to the combustion chamber and thus reducing the risk of seizing and damage on service disassembly.

In some embodiments, the inner and/or outer circumferential sealing surface may each comprise a tongue or groove (e.g. proximal the fixing elements) for cooperation with a respective groove/tongue on the inner and/or outer liners respectively.

In some embodiments, the metering panel may comprise cooling channels e.g. between the upstream and downstream faces. The cooling channels may be integrally formed within the wall. By providing cooling channels within the metering panel, the heat shield and associated cooling pedestals can be eliminated. These cooling channels will provide greater heat loss to the annulus air between the inner/outer liners and the CCIC/CCOC as the metering panel is directly exposed to these annular air flows and greater heat pick up of the cooling flow entering into the combustor.

In some embodiments, the downstream (hot) face of the metering panel may be coated with a thermal barrier coating such as a ceramic thermal barrier coating to provide additional heat protection form the combustion gases.

In some embodiments, the upstream (cold) face of the metering panel may be provided with a surface texture e.g. comprising protrusions/ribs and/or recesses to increase its surface area to thus increase heat loss by interaction with the HP OGV air.

In some embodiments, the metering panel is formed by additive layer manufacturing e.g. using a high temperature nickel alloy. It is envisaged that the metering panel could be manufactured by ALM with the upstream face facing the building plate, oriented such that the groove/tongues on the inner/outer sealing surfaces are vertical. This is to ensure the optimum surface finish to the radial interfaces. ALM facilitates the formation of integral cooling channels within the metering panel.

In other embodiments, the metering panel is formed of ceramic metal composite (CMC). Cooling channels can be drilled into a CMC metering panel. In yet further embodiments, the metering panel can be formed by metal injection moulding or casting.

In some embodiments, the metering panel further comprises a plurality of fuel spray nozzle (FSN) seals, each FSN seal lining a respective one of the plurality of apertures. The plurality of FSN seals may be manufactured integrally e.g. by ALM (using, for example, a high temperature nickel alloy). In other embodiments, they may be manufactured in two parts e.g. in an upstream and a downstream part, and secured (e.g. brazed) together, the upstream part on the upstream face of the metering panel and the downstream part on the downstream face of the metering panel. The apertures may each be circumscribed by a respective seal land (e.g. having an increased thickness) onto which the respective FSN seal is mounted.

In some embodiments, a secondary wall having a convex gradient is provided upstream of the metering panel to further improve the aerodynamic profile to the annulus air without increasing the dump-gap. This may be attached as the seal land and extend to the fasting elements on the inner and outer circumferential sealing surfaces.

In some embodiments, the metering panel is segmented i.e. comprises a plurality of annulus sectors that are joined at their radii (at radial interfaces) to form the annular metering panel.

This is described further below in relation to the second aspect and all features below described may be incorporated into the first aspect.

In a second aspect, there is provided a combustor for a gas turbine engine, the combustor comprising: an inner liner and an outer liner defining a combustion chamber therebetween; and an annular metering panel having an inner circumferential sealing surface sealed to an upstream end of the inner liner and an outer circumferential sealing surface sealed to an upstream end of the outer liner, the metering panel defining a plurality of apertures each for receipt of a respective fuel spray nozzle, wherein the metering panel is segmented into a plurality of annulus sectors.

Providing a segmented metering panel, facilitates manufacture and improves serviceability (as individual sectors can be replaced). The metering panel annulus sectors can be secured together (e.g. temporally secured together) and assembled (e.g. vertically) to the inner liner to create a 'mushroom" structure. This structure can then be installed (e.g. vertically) into the outer liner.

There could be any number of annulus sectors but, in some embodiments, each annulus sector subtends an angle of substantially 90 degrees such that there are four annulus sectors joined to form the annular metering panel. In some embodiments, each annulus sector of the metering panel comprises a plurality e.g. four apertures.

In some embodiments, the segmented metering panel comprises radial seals between adjacent annulus sectors. For example, each annulus sector may comprise a radial interface for interfacing with a radial interface on the adjacent annulus sector. A mating structure may be provided between the adjacent radial interfaces for securing the adjacent annulus sectors together.

In some embodiments, each radial interface may comprise a groove/slot facing an opposing radially-extending groove/slot on the radial interface of the adjacent annulus sector and the mating structure may further comprise a radially-extending seal strip housed within the opposing grooves to secure the adjacent annulus sectors.

In other embodiments, the mating structure may be integrally formed (e.g. by ALM) with the annulus sector. For example, one radial interface may comprise a projection (e.g. a ridge or tongue) for mating with a recess (e.g. a channel or groove) on the radial interface of the adjacent annular sector. The or each mating structure may be coated with a low friction coating. In some embodiments, the inner and/or outer liners defining the combustion chamber may be radially segmented and the radial joins between the metering panel annulus sectors may be circumferentially staggered from any axial joins between inner/outer liner segments.

The metering panel may further comprise any of the features described above in relation to the first aspect either alone or in combination.

In a third aspect there is provided combustion equipment for a gas turbine engine a comprising: a combustor according to the first or second aspect; an annular combustion chamber inner casing (CCIC) located radially inwardly of the combustor; an annular combustion chamber outer casing (CCOC) located radially outwardly of the combustor; a high pressure nozzle guide vane assembly downstream of the combustor and having a circumferential row of high pressure nozzle guide vanes at an inlet to a turbine section of the engine, the high pressure nozzle guide vane assembly joining at a radially inner side to a rear end of the combustion chamber inner casing and joining at a radially outer side to a rear end of the combustion chamber outer casing; and a high pressure outlet guide vane assembly upstream of the combustor and having a circumferential row of annular high pressure outlet guide vanes at an outlet from a compressor section of the engine, the high pressure outlet guide vane assembly joining at a radially inner side to a front end the combustion chamber inner casing and joining at a radially outer side to a front end the combustion chamber outer casing.

According to a fourth aspect there is provided a gas turbine engine for an aircraft including the combustor or combustion equipment of any one of the first to third aspects.

As noted herein, the present disclosure may relate to a gas turbine engine. Such a gas turbine engine may comprise an engine core comprising a turbine, a combustor, a compressor, and a core shaft connecting the turbine to the compressor. Such a gas turbine engine may comprise a fan (having fan blades) located upstream of the engine core.

Arrangements of the present disclosure may be particularly, although not exclusively, beneficial for fans that are driven via a gearbox. Accordingly, the gas turbine engine may comprise 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. The input to the gearbox may be directly from the core shaft, or indirectly from the core shaft, for example via a spur shaft and/or gear. The core shaft may rigidly connect the turbine and the compressor, such that the turbine and compressor rotate at the same speed (with the fan rotating at a lower speed).

The gas turbine engine as described and/or claimed herein may have any suitable general architecture. For example, the gas turbine engine may have any desired number of shafts that connect turbines and compressors, for example one, two or three shafts. Purely by way of example, the turbine connected to the core shaft may be a first turbine, the compressor connected to the core shaft may be a first compressor, and the core shaft may be a first core shaft. The engine core may further comprise a second turbine, a second compressor, and a second core shaft connecting the second turbine to the second compressor. The second turbine, second compressor, and second core shaft may be arranged to rotate at a higher rotational speed than the first core shaft.

In such an arrangement, the second compressor may be positioned axially downstream of the first compressor. The second compressor may be arranged to receive (for example directly receive, for example via a generally annular duct) flow from the first compressor.

The gearbox may be arranged to be driven by the core shaft that is configured to rotate (for example in use) at the lowest rotational speed (for example the first core shaft in the example above). For example, the gearbox may be arranged to be driven only by the core shaft that is configured to rotate (for example in use) at the lowest rotational speed (for example only be the first core shaft, and not the second core shaft, in the example above). Alternatively, the gearbox may be arranged to be driven by any one or more shafts, for example the first and/or second shafts in the example above.

The gearbox may be a reduction gearbox (in that the output to the fan is a lower rotational rate than the input from the core shaft). Any type of gearbox may be used. For example, the gearbox may be a "planetary" or "star' gearbox, as described in more detail elsewhere herein. The gearbox may have any desired reduction ratio (defined as the rotational speed of the input shaft divided by the rotational speed of the output shaft), for example greater than 2.5, for example in the range of from 3 to 4.2, or 3.2 to 3.8, for example on the order of or at least 3, 3.1, 3.2, 3.3, 3.4, 3.5, 3.6, 3.7, 3.8, 3.9, 4, 4.1 or 4.2. The gear ratio may be, for example, between any two of the values in the previous sentence. Purely by way of example, the gearbox may be a "star" gearbox having a ratio in the range of from 3.1 or 3.2 to 3.8. In some arrangements, the gear ratio may be outside these ranges.

In any gas turbine engine as described and/or claimed herein, a combustor may be provided axially downstream of the fan and compressor(s). For example, the combustor may be directly downstream of (for example at the exit of) the second compressor, where a second compressor is provided. By way of further example, the flow at the exit to the combustor may be provided to the inlet of the second turbine, where a second turbine is provided. The combustor may be provided upstream of the turbine(s).

The or each compressor (for example the first compressor and second compressor as described above) may comprise any number of stages, for example multiple stages. Each stage may comprise a row of rotor blades and a row of stator vanes, which may be variable stator vanes On that their angle of incidence may be variable). The row of rotor blades and the row of stator vanes may be axially offset from each other.

The or each turbine (for example the first turbine and second turbine as described above) may comprise any number of stages, for example multiple stages Each stage may comprise a row of rotor blades and a row of stator vanes. The row of rotor blades and the row of stator vanes may be axially offset from each other.

Each fan blade may be defined as having a radial span extending from a root (or hub) at a radially inner gas-washed location, or 0% span position, to a tip at a 100% span position.

The ratio of the radius of the fan blade at the hub to the radius of the fan blade at the tip may be less than (or on the order of) any of: 0.4, 0.39, 0.38 0.37, 0.36, 0.35, 0.34, 0.33, 0.32, 0.31, 0.3, 0.29, 0.28, 0.27, 0.26, or 0.25. The ratio of the radius of the fan blade at the hub to the radius of the fan blade at the tip may be in an inclusive range bounded by any two of the values in the previous sentence (i.e. the values may form upper or lower bounds), for example in the range of from 0.28 to 0.32. These ratios may commonly be referred to as the hub-to-tip ratio. The radius at the hub and the radius at the tip may both be measured at the leading edge (or axially forwardmost) part of the blade. The hub-to-tip ratio refers, of course, to the gas-washed portion of the fan blade, i.e. the portion radially outside any platform.

The radius of the fan may be measured between the engine centreline and the tip of a fan blade at its leading edge. The fan diameter (which may simply be twice the radius of the fan) may be greater than (or on the order of) any of: 220 cm, 230 cm, 240 cm, 250 cm (around 100 inches), 260 cm, 270 cm (around 105 inches), 280 cm (around 110 inches), 290 cm (around 115 inches), 300 cm (around 120 inches), 310 cm, 320 cm (around 125 inches), 330 cm (around 130 inches), 340 cm (around 135 inches), 350cm, 360cm (around 140 inches), 370 cm (around 145 inches), 380 (around 150 inches) cm, 390 cm (around 155 inches), 400 cm, 410 cm (around 160 inches) or 420 cm (around 165 inches). The fan diameter may be in an inclusive range bounded by any two of the values in the previous sentence (i.e. the values may form upper or lower bounds), for example in the range of from 240 cm to 280 cm or 330 cm to 380 cm.

The rotational speed of the fan may vary in use. Generally, the rotational speed is lower for fans with a higher diameter. Purely by way of non-limitative example, the rotational speed of the fan at cruise conditions may be less than 2500 rpm, for example less than 2300 rpm. Purely by way of further non-limitative example, the rotational speed of the fan at cruise conditions for an engine having a fan diameter in the range of from 220 cm to 300 cm (for example 240 cm to 280 cm or 250 cm to 270cm) may be in the range of from 1700 rpm to 2500 rpm, for example in the range of from 1800 rpm to 2300 rpm, for example in the range of from 1900 rpm to 2100 rpm. Purely by way of further non-limitative example, the rotational speed of the fan at cruise conditions for an engine having a fan diameter in the range of from 330 cm to 380 cm may be in the range of from 1200 rpm to 2000 rpm, for example in the range of from 1300 rpm to 1800 rpm, for example in the range of from 1400 rpm to 1800 rpm.

In use of the gas turbine engine, the fan (with associated fan blades) rotates about a rotational axis. This rotation results in the tip of the fan blade moving with a velocity Ufip. The work done by the fan blades 13 on the flow results in an enthalpy rise dH of the flow. A fan tip loading may be defined as dH/Utip2, where dH is the enthalpy rise (for example the 1-D average enthalpy rise) across the fan and Utip is the (translational) velocity of the fan tip, for example at the leading edge of the tip (which may be defined as fan tip radius at leading edge multiplied by angular speed). The fan tip loading at cruise conditions may be greater than (or on the order of) any of: 0.28, 0.29, 0.30, 0.31, 0.32, 0.33, 0.34, 0.35, 0.36, 0.37, 0.38, 0.39 or 0.4 (all units in this paragraph being Jkg-1K-1/(ms-1)2). The fan tip loading may be in an inclusive range bounded by any two of the values in the previous sentence (i.e. the values may form upper or lower bounds), for example in the range of from 0.28 to 0.31, or 0.29 to 0.3.

Gas turbine engines in accordance with the present disclosure may have any desired bypass ratio, where the bypass ratio is defined as the ratio of the mass flow rate of the flow through the bypass duct to the mass flow rate of the flow through the core at cruise conditions. In some arrangements the bypass ratio may be greater than (or on the order of) any of the following: 10, 10.5, 11, 11.5, 12, 12.5, 13, 13.5, 14, 14.5, 15, 15.5, 16, 16.5, 17, 17.5, 18, 18.5, 19, 19.5 or 20. The bypass ratio may be in an inclusive range bounded by any two of the values in the previous sentence (i.e. the values may form upper or lower bounds), for example in the range of form 12 to 16, 13 to 15, or 13 to 14. The bypass duct may be substantially annular. The bypass duct may be radially outside the engine core. The radially outer surface of the bypass duct may be defined by a nacelle and/or a fan case.

The overall pressure ratio of a gas turbine engine as described and/or claimed herein may be defined as the ratio of the stagnation pressure upstream of the fan to the stagnation pressure at the exit of the highest pressure compressor (before entry into the combustor).

By way of non-limitafive example, the overall pressure ratio of a gas turbine engine as described and/or claimed herein at cruise may be greater than (or on the order of) any of the following: 35, 40, 45, 50, 55, 60, 65, 70, 75. The overall pressure ratio may be in an inclusive range bounded by any two of the values in the previous sentence (i.e. the values may form upper or lower bounds), for example in the range of from 50 to 70.

Specific thrust of an engine may be defined as the net thrust of the engine divided by the total mass flow through the engine. At cruise conditions, the specific thrust of an engine described and/or claimed herein may be less than (or on the order of) any of the following: Nkg-1s, 105 Nkg-1s, 100 Nkg-1s, 95 Nkg-1s, 90 Nkg-1s, 85 Nkg-1s or 80 Nkg-1s. The specific thrust may be in an inclusive range bounded by any two of the values in the previous sentence (i.e. the values may form upper or lower bounds), for example in the range of from 80 Nkg-1s to 100 Nkg-1s, or 85 Nkg-1s to 95 Nkg-1s. Such engines may be particularly efficient in comparison with conventional gas turbine engines.

A gas turbine engine as described and/or claimed herein may have any desired maximum thrust. Purely by way of non-limitative example, a gas turbine as described and/or claimed herein may be capable of producing a maximum thrust of at least (or on the order of) any of the following: 160 kN, 170 kN, 180 kN, 190 kN, 200 kN, 250 kN, 300 kN, 350 kN, 400 kN, 450 kN, 500 kN, or 550 kN. The maximum thrust may be in an inclusive range bounded by any two of the values in the previous sentence (i.e. the values may form upper or lower bounds). Purely by way of example, a gas turbine as described and/or claimed herein may be capable of producing a maximum thrust in the range of from 330 kN to 420 kN, for example 350 kN to 400 kN. The thrust referred to above may be the maximum net thrust at standard atmospheric conditions at sea level plus 15 degrees C (ambient pressure 101.3 kPa, temperature 30 degrees C), with the engine static.

In use, the temperature of the flow at the entry to the high pressure turbine may be particularly high. This temperature, which may be referred to as TET, may be measured at the exit to the combustor, for example immediately upstream of the first turbine vane, which itself may be referred to as a nozzle guide vane. At cruise, the TET may be at least (or on the order of) any of the following: 1400 K, 1450 K, 1500 K, 1550 K, 1600 K or 1650 K. The TET at cruise may be in an inclusive range bounded by any two of the values in the previous sentence (i.e. the values may form upper or lower bounds). The maximum TET in use of the engine may be, for example, at least (or on the order of) any of the following: 1700 K, 1750 K, 1800 K, 1850 K, 1900 K, 1950 K or 2000 K. The maximum TET may be in an inclusive range bounded by any two of the values in the previous sentence (i.e. the values may form upper or lower bounds), for example in the range of from 1800 K to 1950 K. The maximum TET may occur, for example, at a high thrust condition, for example at a maximum take-off (MTO) condition.

A fan blade and/or aerofoil portion of a fan blade described and/or claimed herein may be manufactured from any suitable material or combination of materials. For example, at least a part of the fan blade and/or aerofoil may be manufactured at least in part from a composite, for example a metal matrix composite and/or an organic matrix composite, such as carbon fibre. By way of further example at least a part of the fan blade and/or aerofoil may be manufactured at least in part from a metal, such as a titanium based metal or an aluminium based material (such as an aluminium-lithium alloy) or a steel based material. The fan blade may comprise at least two regions manufactured using different materials.

For example, the fan blade may have a protective leading edge, which may be manufactured using a material that is better able to resist impact (for example from birds, ice or other material) than the rest of the blade. Such a leading edge may, for example, be manufactured using titanium or a titanium-based alloy. Thus, purely by way of example, the fan blade may have a carbon-fibre or aluminium based body (such as an aluminium lithium alloy) with a titanium leading edge.

A fan as described and/or claimed herein may comprise a central portion, from which the fan blades may extend, for example in a radial direction. The fan blades may be attached to the central portion in any desired manner. For example, each fan blade may comprise a fixture which may engage a corresponding slot in the hub (or disc). Purely by way of example, such a fixture may be in the form of a dovetail that may slot into and/or engage a corresponding slot in the hub/disc in order to fix the fan blade to the hub/disc. By way of further example, the fan blades maybe formed integrally with a central portion. Such an arrangement may be referred to as a bladed disc or a bladed ring. Any suitable method may be used to manufacture such a bladed disc or bladed ring. For example, at least a part of the fan blades may be machined from a block and/or at least part of the fan blades may be attached to the hub/disc by welding, such as linear friction welding.

The gas turbine engines described and/or claimed herein may or may not be provided with a variable area nozzle (VAN). Such a variable area nozzle may allow the exit area of the bypass duct to be varied in use. The general principles of the present disclosure may apply to engines with or without a VAN The fan of a gas turbine as described and/or claimed herein may have any desired number of fan blades, for example 14, 16, 18, 20, 22, 24 or 26 fan blades.

As used herein, cruise conditions have the conventional meaning and would be readily understood by the skilled person. Thus, for a given gas turbine engine for an aircraft, the skilled person would immediately recognise cruise conditions to mean the operating point of the engine at mid-cruise of a given mission (which may be referred to in the industry as the "economic mission") of an aircraft to which the gas turbine engine is designed to be attached. In this regard, mid-cruise is the point in an aircraft flight cycle at which 50% of the total fuel that is burned between top of climb and start of descent has been burned (which may be approximated by the midpoint -in terms of time and/or distance-between top of climb and start of descent. Cruise conditions thus define an operating point of, the gas turbine engine that provides a thrust that would ensure steady state operation (i.e. maintaining a constant altitude and constant Mach Number) at mid-cruise of an aircraft to which it is designed to be attached, taking into account the number of engines provided to that aircraft. For example, where an engine is designed to be attached to an aircraft that has two engines of the same type, at cruise conditions the engine provides half of the total thrust that would be required for steady state operation of that aircraft at mid-cruise.

In other words, for a given gas turbine engine for an aircraft, cruise conditions are defined as the operating point of the engine that provides a specified thrust (required to provide -in combination with any other engines on the aircraft -steady state operation of the aircraft to which it is designed to be attached at a given mid-cruise Mach Number) at the mid-cruise atmospheric conditions (defined by the International Standard Atmosphere according to ISO 2533 at the mid-cruise altitude). For any given gas turbine engine for an aircraft, the mid-cruise thrust, atmospheric conditions and Mach Number are known, and thus the operating point of the engine at cruise conditions is clearly defined.

Purely by way of example, the forward speed at the cruise condition may be any point in the range of from Mach 0.7 to 0.9, for example 0.75 to 0.85, for example 0.76 to 0.84, for example 0.77 to 0.83, for example 0.78 to 0.82, for example 0.79 to 0.81, for example on the order of Mach 0.8, on the order of Mach 0.85 or in the range of from 0.8 to 0.85. Any single speed within these ranges may be part of the cruise condition. For some aircraft, the cruise conditions may be outside these ranges, for example below Mach 0.7 or above Mach 0.9.

Purely by way of example, the cruise conditions may correspond to standard atmospheric conditions (according to the International Standard Atmosphere, ISA) at an altitude that is in the range of from 10000 m to 15000 m, for example in the range of from 10000 m to 12000 m, for example in the range of from 10400 m to 11600 m (around 38000 ft), for example in the range of from 10500 m to 11500 m, for example in the range of from 10600 m to 11400 m, for example in the range of from 10700 m (around 35000 ft) to 11300 m, for example in the range of from 10800 m to 11200 m, for example in the range of from 10900 m to 11100 m, for example on the order of 11000 m. The cruise conditions may correspond to standard atmospheric conditions at any given altitude in these ranges.

Purely by way of example, the cruise conditions may correspond to an operating point of the engine that provides a known required thrust level (for example a value in the range of from 30 kN to 35 kN) at a forward Mach number of 0.8 and standard atmospheric conditions (according to the International Standard Atmosphere) at an altitude of 38000 ft (11582 m).

Purely by way of further example, the cruise conditions may correspond to an operating point of the engine that provides a known required thrust level (for example a value in the range of from 50 kN to 65 kN) at a forward Mach number of 0.85 and standard atmospheric conditions (according to the International Standard Atmosphere) at an altitude of 35000 ft (10668 m).

In use, a gas turbine engine described and/or claimed herein may operate at the cruise conditions defined elsewhere herein. Such cruise conditions may be determined by the cruise conditions (for example the mid-cruise conditions) of an aircraft to which at least one (for example 2 or 4) gas turbine engine may be mounted in order to provide propulsive thrust.

According to an aspect, there is provided an aircraft comprising a gas turbine engine as described and/or claimed herein. The aircraft according to this aspect is the aircraft for which the gas turbine engine has been designed to be attached. Accordingly, the cruise conditions according to this aspect correspond to the mid-cruise of the aircraft, as defined elsewhere herein.

According to an aspect, there is provided a method of operating a gas turbine engine as described and/or claimed herein. The operation may be at the cruise conditions as defined elsewhere herein (for example in terms of the thrust, atmospheric conditions and Mach Number).

According to an aspect, there is provided a method of operating an aircraft comprising a gas turbine engine as described and/or claimed herein. The operation according to this aspect may include (or may be) operation at the mid-cruise of the aircraft, as defined elsewhere herein.

The skilled person will appreciate that except where mutually exclusive, a feature or parameter described in relation to any one of the above aspects may be applied to any other aspect. Furthermore, except where mutually exclusive, any feature or parameter described herein may be applied to any aspect and/or combined with any other feature or parameter described herein.

Brief Description of the Drawings

Embodiments of the invention will now be described by way of example with reference to the accompanying drawings in which: Figure 1 is a sectional longitudinal view of prior art combustion equipment of a gas turbine engine; Figure 2 is an enlarged view of the prior art combustor; Figure 3 is a sectional longitudinal view of a gas turbine engine; Figure 4 is a close up sectional side view of an upstream portion of a gas turbine engine; Figure 5 is a partially cut-away view of a gearbox for a gas turbine engine; Figure 6 is a sectional longitudinal view of combustion equipment of a gas turbine engine; Figure 7 shows an enlarged portion of the combustion chamber; Figure 8 shows details of the fixing elements; Figure 9 shows details of an annulus sector of a segmented metering panel; Figure 10 shows a radial seal between two adjacent annulus sectors; Figure 11 shows an exploded view of components of a combustor and Figure 12 shows an alternative combustor with a secondary wall. Detailed Description 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 Figure 3 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.

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.

An exemplary arrangement for a geared fan gas turbine engine 10 is shown in Figure 4. The low pressure turbine 19 (see Figure 3) 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.

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.

The epicyclic gearbox 30 is shown by way of example in greater detail in Figure 5. 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 Figure 4. 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.

The epicyclic gearbox 30 illustrated by way of example in Figures 4 and 5 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.

It will be appreciated that the arrangement shown in Figures 4 and 5 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 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 Figure 4 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 Figure 4. 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 Figure 4.

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.

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

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 Figure 3 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.

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 Figure 3), and a circumferential direction (perpendicular to the page in the Figure 3 view). The axial, radial and circumferential directions are mutually perpendicular.

Figure 6 shows a section view of the combustion equipment 40' for the gas turbine engine 10.

The combustion equipment 40' has a combustor 41' comprising an inner liner 80 and an outer liner 81 defining a combustion chamber 42 therebetween. An annular metering panel 51' having an inner circumferential sealing surface 82 sealed to an upstream end of the inner liner 80 and an outer circumferential sealing surface 83 sealed to an upstream end of the outer liner 81 is provided. The metering panel 51' defines a plurality of apertures 84 each for receipt of a respective fuel spray nozzle 58.

The metering panel 51' has a convex gradient such that it forms an arc subtending an angle of around 120 degrees between the inner and outer circumferential sealing surfaces 82, 83.

The plurality of apertures 84 are provided mid-way between the sealing surfaces 82, 83 i.e. substantially along the axis of symmetry of the arc.

The inner and outer liners 80, 81 also comprise a respective convex portion 80a, 81a proximal the join with the inner/outer circumferentially sealing surfaces 82, 83 respectively.

The arc created by the convex metering panel 51' and the convex liner portions 80a, 81a subtends an angle of between around 270 degrees. This provides a combustor chamber 42 volume proximal each aperture 84 that is substantially spherical as shown in Figure 7. This combustion chamber shape reduces corner recirculation of the combustion gases which leads to a narrower population distribution of residence times which improves emissions.

The substantially spherical volumes around each aperture 84 intersect around the circumference of the combustion chamber 42. This leads to undulating inner and outer sealing surfaces 82, 83 as can be seen most clearly in Figures 8 and 9.

The inner and/or outer circumferential sealing surfaces 82, 83 each comprise a plurality of fixing lugs 87, 88 as can be seen in Figure 8. These fixing lugs 87, 88 align with corresponding fixing lugs 89, 90 on the inner and outer liners 80, 81 respectively to receive a respective threaded axial fastener 85 which mates with a threaded bore within one or both of the pair of fixing lugs 87/89, 88/90. The fixing lugs 87-90 and threaded fasteners 85 are provided externally to the combustion chamber 42.

The inner and outer circumferential sealing surfaces 82, 83 and upstream ends of the inner and outer liners 80, 81 also comprise a mating tongue and groove arrangement 86.

The metering panel 51' is formed of additive layer manufacturing and comprises integrally formed cooling channels between the upstream and downstream faces. These cooling channels will provide for increased heat loss to the annulus air between the inner/outer liners 80, 81 and the CCIC 44/CCOC 45 as the metering panel 51' is directly exposed to these annular air flows.

The downstream (hot) face of the metering panel 51' is coated with a thermal barrier coating such as a ceramic thermal barrier coating to provide additional heat protection form the combustion gases. The upstream (cold) face of the metering panel 51' is provided with a surface texture to increase its surface area to thus increase heat loss by interaction with the air exiting the HP OGV 50.

The metering panel 51' further comprises a plurality of fuel spray nozzle (FSN) seals 91 which can clearly be seen On part) in Figure 8. Each FSN seal 91 lines a respective one of the plurality of apertures 84. They are integrally formed by ALM and fitted around a seal land 92 which circumscribes the aperture 84. The FSN seal 91 seals around the fuel spray nozzle 58 as seen in Figure 6.

The metering panel 51' is segmented i.e. comprises four annulus sectors 96, one of which is shown in plan view and side view in Figure 9. The convex gradient formed between the inner and outer sealing surfaces 82, 83 is clearly shown in the side view in Figure 9.

Each annulus sector 96 subtends an angle of 90 degrees and comprises four apertures 84.

As can be seen in Figure 10, radial seals 95 are provided between adjacent annulus sectors 96. Each annulus sector 96, 96a comprises a radial interface 94 for interfacing with a radial interface on the adjacent annulus sector 96, 96a. Each radial interface 94 comprises a groove/slot 97 facing an opposing radially-extending groove/slot on the radial interface of the adjacent annulus sector 96a and a radially-extending seal strip 95 is housed within the opposing grooves 97 to secure the adjacent annulus sectors 96, 96a.

Figure 11 shows exploded components of the combustor 41' comprising the inner liner 80, outer liner 81 and segmented metering ring 51'.

The inner and outer liners 80, 81 defining the combustion chamber 42 are radially segmented with axially-extending joins 98, 99. The radial joins between the metering panel annulus sectors 96, 96a-c are circumferentially staggered from the axial joins 98/99 between inner/outer liner segments.

Finally, Figure 12 shows a further embodiment of the combustion equipment previously described in relation to Figures 6 to 11 where a secondary wall 100 having a convex gradient is provided upstream of the metering panel 51' to further improve the aerodynamic profile to the annulus air without increasing the dump-gap 60. This attached as the seal land 92 and extends to the fasting lugs 87, 88.

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.

Claims (17)

1. CLAIMS1. A combustor (41') for a gas turbine engine (10), the combustor (41') comprising: an inner liner (80) and an outer liner (81) defining a combustion chamber (42) therebetween; and an annular metering panel (51') having an inner circumferential sealing surface (82) sealed to an upstream end of the inner liner (80) and an outer circumferential sealing surface (83) sealed to an upstream end of the outer liner (81), the metering panel (51') defining a plurality of apertures (84) each for receipt of a respective fuel spray nozzle (58), wherein the metering panel (51') has a convex gradient between the inner and outer circumferential sealing surfaces (82, 83).
2. 2. The combustor (41') according to claim 1 wherein the metering panel (51') has a convex gradient such that it forms an arc between the inner and outer circumferential sealing surfaces (82, 83).
3. 3. The combustor (41') according to claim 2 wherein the inner and outer liners (80, 81) comprise a respective convex portion (80a, 81a) proximal the seal with the inner/outer circumferentially sealing surfaces (82, 83) respectively.
4. 4. The combustor (41') according to any one of the preceding claims wherein the inner and/or outer circumferential sealing surface (82, 83) comprise a plurality of fixing elements (87, 88) for alignment/cooperation with a respective one of a plurality of corresponding fixing elements (89, 90) on the inner and/or outer liners (80, 81) respectively, the aligned pairs of fixing elements each configured to receive an axial fastener (85).
5. 5. The combustor (41') according to any one of the preceding claims wherein the metering panel (51') comprises cooling channels.
6. 6. The combustor (41') according to any one of the preceding claims wherein a downstream face of the metering panel (51') is coated with a thermal barrier coating.
7. 7. The combustor (41') according to any one of the preceding claims wherein an upstream face of the metering panel (51') is provided with surface texture to increase its surface area.
8. 8. The combustor (41') according to any one of the preceding claims wherein the metering panel (51') is formed by additive layer manufacturing.
9. 9. The combustor (41') according to any one of the preceding claims wherein the metering panel (51') is segmented and comprises a plurality of annulus sectors (96) that are joined at their radii to form the annular metering panel (51').
10. 10. A combustor (41') for a gas turbine engine, the combustor comprising: an inner liner (80) and an outer liner (81) defining a combustion chamber (42) therebetween; and an annular metering panel (51') having an inner circumferential sealing surface (82) sealed to an upstream end of the inner liner (80) and an outer circumferential sealing surface (83) sealed to an upstream end of the outer liner (81), the metering panel (51') defining a plurality of apertures 84 each for receipt of a respective fuel spray nozzle 58, wherein the metering panel (51') is segmented into a plurality of annulus sectors 96.
11. 11. The combustor (41') according to claim 9 or 10 wherein each annulus sector 96 subtends an angle of substantially 90 degrees.
12. 12. The combustor (41') according to any one of claims 9 to 11 wherein each annulus sector (96) comprises a radial interface (94) for interfacing with a radial interface on the adjacent annulus sector, the combustor (41') further comprising a mating structure (95) between each pair of radial interfaces.
13. 13. The combustor (41') according to any one of claims 9 to 12 wherein the inner and/or outer liners (80, 81) defining the combustion chamber (42) are radially segmented and radial joins between the metering panel annulus sectors (96) are circumferentially staggered from axial joins (98, 99) between adjacent inner/outer liner segments.
14. 14. A combustion equipment for a gas turbine engine a comprising: a combustor (41') according to any one of the preceding claims; an annular combustion chamber inner casing (CCIC) (44) located radially inwardly of the combustor (41'); an annular combustion chamber outer casing (CCOC) (45) located radially outwardly of the combustor (41'); a high pressure nozzle guide vane assembly (43) downstream of the combustor and having a circumferential row of high pressure nozzle guide vanes (50) at an inlet to a turbine section of the engine, the high pressure nozzle guide vane assembly (43) joining at a radially inner side to a rear end of the CCIC (44) and joining at a radially outer side to a rear end of the CCOC (45); and a high pressure outlet guide vane assembly (46) upstream of the combustor (41') and having a circumferential row of annular high pressure outlet guide vanes (52) at an outlet from a compressor section of the engine, the high pressure outlet guide vane assembly (46) joining at a radially inner side to a front end the CCIC (44) and joining at a radially outer side to a front end the CCOC (45).
15. 15. A gas turbine engine for an aircraft including the combustion equipment (16) according claim 14.
16. 16. The gas turbine engine according to claim 15 further including: an engine core (11) comprising a turbine (19), a compressor (14), and a core shaft (26) connecting the turbine to the compressor; a fan (23) located upstream of the engine core, the fan comprising a plurality of fan blades; and a gearbox (30) that receives an input from the core shaft (26) and outputs drive to the fan so as to drive the fan at a lower rotational speed than the core shaft.
17. 17. The gas turbine engine according to claim 16, wherein: the turbine is a first turbine (19), the compressor is a first compressor (14), and the core shaft is a first core shaft (26); the engine core further comprises a second turbine (17), a second compressor (15), and a second core shaft (27) 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.