An improved combustor apparatus

Abstract

A fuel swirl nozzle 100 for a gas turbine engine (10 Fig. 1) has a first array 116 of circumferential vanes 118, a second array 144 of circumferential vanes 146 and a third array 164 of circumferential vanes 166 in an axial flow sequence. The vanes are arranged as an outer air swirler array 164 and an inner air swirler array 144, with the inner air swirler array positioned concentrically within the outer air swirler array. The outer air swirler is an annular array and in use imposes a first rotational flow component on an air flow entering the outer air swirler. The inner air swirler in an annular array within the inner air swirler body and in use imposes a second rotational flow component on an air flow entering the inner air swirler. Upstream of the inner and outer array is a further array 116 that in use imposes a third rotational flow component on the air flow. The second and / or third circumferential array of vanes can include a single oversized vane that has a greater thickness and / or a greater chord than the remaining second and / or third vanes. By including an oversized vane, it becomes possible to accommodate a fuel supply line within the oversized vane and therefore eliminate the pressure losses associated with a fuel line extending radially through the air flow.

Description

Field of the Disclosure The present disclosure relates to an improved combustor apparatus and particularly, but not exclusively, to an improved combustor apparatus for a gas turbine engine. Background to the Disclosure It is known to use fuel swirl nozzles as part of a combustion apparatus in a gas turbine engine. The combustion apparatus instigates and facilitates combustion of fuel with relatively high-pressure air received from a compressor stage of the gas turbine engine and thereby adds thermal energy to the relatively high-pressure air prior to the air being expanded in a turbine stage of the gas turbine engine and subsequently being exhausted at an exhaust nozzle to provide propulsive thrust. In the turbine stage, energy is extracted from the expanding air in the form of mechanical work. The turbine stage may drive, for example, a propulsive fan which provides additional propulsive thrust to the gas turbine engine. The performance of the fuel swirl nozzle and the combustor is important for effective and efficient operation of the gas turbine engine. In particular, the performance of both the fuel swirl nozzle and the combustor is directly linked to the thermal efficiency of the gas turbine engine and / or emissions produced by the gas turbine engine. It is therefore desirable to provide an improved fuel swirl nozzle and combustor for a gas turbine engine. Statements of Disclosure According to a first aspect of the present disclosure, there is provided a fuel swirl nozzle for a gas turbine engine, the fuel swirl nozzle comprising an inner air portion, an outer air portion, a dome air portion, and a fuel delivery portion, the inner air portion comprising an inner air passage, the inner air passage comprising a first annular upstream portion fluidly connected to an opposite second cylindrical downstream portion, a downstream end of the second cylindrical downstream portion defining a pre-filmer edge, the outer air portion comprising an outer air passage, the outer air passage being positioned concentrically with, and radially outwardly of, the second downstream portion of the inner air passage, and comprising a first intake end, and a second opposite exhaust end, the dome air portion comprising a dome air passage, the dome air passage being positioned concentrically with, and radially outwardly of, the outer air passage, and comprising a first intake end, and a second opposite exhaust end, and the fuel delivery portion comprising a fuel delivery passage, the fuel delivery passage comprising an inlet portion and an outlet portion, the inlet portion being configured to receive a flow of fuel, and the outlet portion being arranged concentrically between a radially outwardly facing surface of the second downstream portion of the inner air passage, and a radially inwardly facing surface of the second exhaust end of the outer air passage, and wherein the first annular upstream portion of the inner air passage comprises a first circumferential array of first vanes configured to impart a first rotational swirl to a first air flow entering the first annular upstream portion, the first rotational swirl having a first swirl angle relative to the longitudinal axis of the engine, the outer air passage comprises a second circumferential array of second vanes configured to impart a second rotational swirl to a second air flow entering the first intake end of the outer air passage, the second rotational swirl having a second swirl angle relative to the longitudinal axis of the engine, the dome air passage comprises a third circumferential array of third vanes configured to impart a third rotational swirl to a third air flow entering the first intake end of the dome air passage, the third rotational swirl having a third swirl angle relative to the longitudinal axis of the engine, and the flow of fuel is delivered into an annular volume defined between the radially outwardly facing surface of the second downstream portion of the inner air passage, and the radially inwardly facing surface of the second exhaust end of the outer air passage, and the flow of fuel is exhausted from the annular volume into, in axial flow sequence, the first air flow exhausted from the inner air passage, the second air flow exhausted from the outer air passage, and the third air flow exhausted from the dome air passage. The fuel stem portion passes radially through the outer air swirler to deliver fuel to the region between the outer air swirler and the inner air swirler. This enables an axial length of the fuel swirl nozzle to be reduced, which in turn allows the size and mass of the fuel / air swirler portion to be reduced. A reduced mass fuel / air swirler portion then enables the fuel stem, that supports the fuel / air swirler assembly, to also be reduced in weight. By moving the outer air swirler axially upstream (relative to prior art arrangements of fuels spray nozzle) it becomes possible to increase the pre-mixing length for the air and fuel upstream of the fuel swirl nozzle throat. This increases the fuel-air mixing efficiency of the fuel swirl nozzle. By forming the fuel / air swirler assembly from only two components, namely the outer air swirler and the inner air swirler, there is only a single component interface whose dimensions must be carefully controlled. This makes the fuel swirl nozzle simpler to assemble and easier for a user to maintain the necessary component tolerances during assembly, than for conventional fuel swirl nozzles whose fuel / air swirlers comprise multiple sub-components all requiring close tolerances in order to achieve efficient operation. In one arrangement, the plurality of first vanes induce a rotational flow component having a first swirl angle relative to the longitudinal axis of the engine, and the plurality of second vanes induce a rotational flow component having a second swirl angle relative to the longitudinal axis of the engine, with the first swirl angle being different to the second swirl angle. In one arrangement, the first swirl angle is approximately 10°, while the second swirl angle is approximately 20°. In this way, the flow of the air / fuel mixture through the inner air swirler has a greater amount of swirl rotation than the flow of the air / fuel mixture through the outer air swirler. In an alternative arrangement, the first swirl angle is approximately 20°, while the second swirl angle is approximately 10°. This difference in swirl rotation angle between the flows through the outer air swirler and the inner air swirler increases the efficiency of the atomisation of the fuel into the air flow entering the fuel swirl nozzle, which in turn increases the combustion efficiency of the combustor. In an alternative arrangement, the plurality of first vanes in the outer air swirler direct an airflow entering the outer air swirler in a first rotational direction, while the plurality of second vanes in the inner air swirler direct an airflow entering the inner air swirler in a second rotational direction. The first rotational direction is in an opposite sense to the second rotational direction. For example, the first rotational direction may be clockwise, when viewed towards an inlet face of the fuel / air swirler assembly, while the second rotational direction may be anti-clockwise when viewed in the same direction. Directing the airflow through the first vanes in a first rotational direction assists in the atomisation of the fuel into the air flow through the first vanes. Subsequently directing the atomised fuel / air flow through the second vanes in a second opposite rotational direction will further atomise the fuel / air mixture prior to entry of the fuel / air mixture into the combustor volume, Optionally, the second exhaust end of the outer air passage directs the second air flow exhausted from the outer air passage in a radially inwardly direction. By directing the airflow that is exhausted from the outer air passage in a radially inwardly direction, it is possible to improve the mixing of the air flow with the air flow exhausted from the inner air passage. This in turn improves the mixing of te fuel with the air flow. Optionally, each of the first vanes has radial height H, and an axial spacing S is defined between a downstream-most edge of the first vane array and an upstream-most edge of the second vane array, and the axial spacing S is in a range of between H and 5*H. The choice of the axial spacing S between the downstream-most edge of the first vane array and the upstream-most edge of the second vane array has been found to be optimally selected in a range of between H and 5*H in order to ensure that the air flow entering the second vane array remains in a laminar flow regime. This in turn promotes the efficient mixing of the fuel with the air flow. Optionally, at least one of the first circumferential array of first vanes, the second circumferential array of second vanes, and the third circumferential array of third vanes are circumferentially equally spaced. In one arrangement, the first circumferential array of first vanes is arranged in a circumferentially equi-spaced array. This arrangement generates a circumferentially evenly distributed flow pattern, which serves to more efficiently entrain the fuel that is supplied to the nozzle into the air flow. Likewise, for the circumferentially equi-spaced second array of second vanes, and circumferentially equi-spaced third array of third vanes. Optionally, at least one of the first circumferential array of first vanes, the second circumferential array of second vanes, and the third circumferential array of third vanes are circumferentially unequally spaced. In an alternative arrangement, one or more of the first circumferential array of first vanes, the second circumferential array of second vanes, and the third circumferential array of third vanes, may be arranged in a circumferentially unequispaced array. By arranging these vanes in an unequally spaced configuration, it is possible to generate an asymmetric fuel / air distribution exiting the fuel swirl nozzle. This may be helpful for flow of the air / fuel mixture into and through the downstream combustor assembly. Optionally, the first circumferential array of first vanes comprises a circumferential array of between three (3) vanes and six (6) vanes. In one arrangement, the first circumferential array of first vanes comprises a circumferential array of three (3) vanes, which provides a balance between the pressure loss resulting from a higher vane count and providing the ability to impart sufficient rotational swirl to the air flow passing through the vane array. In another arrangement, in which a higher degree of rotational swirl is required of the air flow exiting the first circumferential array, the first array may comprise a circumferential array of six (6) vanes Optionally, the second circumferential array of second vanes comprises a circumferential array of between eight (8) vanes and fourteen (14) vanes. In one arrangement, the second circumferential array of second vanes comprises a circumferential array of eight (8) vanes, which provides a balance between the pressure loss resulting from a higher vane count and providing the ability to impart sufficient rotational swirl to the air flow passing through the vane array. In another arrangement, in which a higher degree of rotational swirl is required of the airflow exiting the second circumferential array, the second array may comprise a circumferential array of fourteen (14) vanes Optionally, the third circumferential array of third vanes comprises a circumferential array of between eight (8) vanes and fourteen (14) vanes. In one arrangement, the third circumferential array of third vanes comprises a circumferential array of eight (8) vanes, which provides a balance between the pressure loss resulting from a higher vane count and providing the ability to impart sufficient rotational swirl to the air flow passing through the vane array. In another arrangement, in which a higher degree of rotational swirl is required of the airflow exiting the third circumferential array, the second array may comprise a circumferential array of fourteen (14) vanes Optionally, the second circumferential array of second vanes comprises a single oversized second vane, the single oversized vane having a greater thickness and / or a greater chord than the remaining second vanes. By increasing the thickness of one of the second vanes (the “oversized” vane) in the second circumferential array, together with a corresponding consequential increase in chord of the same vane, it becomes possible to accommodate a fuel supply line within the oversized vane. This means that it is not necessary to locate the fuel supply line that delivers a fuel flow to the circumferential region between the first circumferential vane array and the second circumferential vane array within the flow region between second vanes. Consequently, this eliminate the pressure losses associated with a fuel line extending radially through the airflow passing through the second circumferential vane array flow region. Optionally, the third circumferential array of third vanes comprises a single oversized third vane, the single oversized third vane having a greater thickness and / or a greater chord than the remaining third vanes. Likewise, by increasing the thickness of one of the third vanes (the “oversized” vane) in the third circumferential array, together with a corresponding consequential increase in chord of the same vane, it becomes possible to accommodate a fuel supply line within the oversized vane. This means that it is not necessary to locate the fuel supply line that delivers a fuel flow to the circumferential region between the first circumferential vane array and the second circumferential vane array within the flow region between third vanes. Consequently, this eliminate the pressure losses associated with a fuel line extending radially through the airflow passing through the third circumferential vane array flow region. Optionally, the second exhaust end of the outer air passage extends axially past the second downstream end of the inner air passage. Extending the second exhaust end of the outer air passage axially downstream past the second downstream end of the inner air passage, increases the efficiency of the atomisation of the fuel into the airflow entering the fuel swirl nozzle, which in turn increases the combustion efficiency of the combustor. Optionally, the second exhaust end of the dome air passage extends axially past the second exhaust end of the outer air passage. Likewise, extending the second exhaust end of the dome air passage axially downstream past the second exhaust end of the outer air passage, increases the efficiency of the atomisation of the fuel into the air flow entering the fuel swirl nozzle, which in turn increases the combustion efficiency of the combustor. Optionally, each of the first rotational swirl, the second rotational swirl, and the third rotational swirl is in the same rotational direction around a longitudinal axis of the fuel swirl nozzle. The selection of the first rotational swirl, the second rotational swirl, and the third rotational swirl being in the same rotational direction provides for an atomised fuel / air mixture flow that itself has a rotational swirl component. This in turn increases the combustion efficiency of the combustor. Optionally, at least one of the inner air portion, the outer air portion, the dome air portion, and the fuel delivery portion is formed by an additive layer manufacturing process. The use of an additive layer manufacturing process to form either or both of the outer air swirler and the inner air swirler enables the plurality of corresponding first and second vanes to be formed integrally with the respective one of the outer air swirler body and inner air swirler body. This makes the air swirler more compact and lighter than the prior art equivalent. Forming either or both of the outer air swirler and the inner air swirler by an additive layer manufacturing process further enables the forming of a geometrical arrangement of corresponding first and second vanes that cannot be produced by conventional (i.e. casting, forging, machining, etc.,) forming processes. For example, the first and second vanes may have a re-entrant geometry. Optionally, the inner air portion and the outer air portion are joined to one another by a welding process or a brazing process. In one arrangement, the outer air swirler and the inner air swirler are each formed by an additive layer manufacturing process and are joined to one another by a brazing process. This enables the outer air swirler to be machined in the region where the fuel stem passes radially inwardly through the outer air swirler, prior to assembly of the inner air swirler, outer air swirler and fuel stem. Optionally, the fuel delivery portion is formed as a monolithic additive layer manufactured component. The use of an additive layer manufacturing process to form the fuel inlet portion allows the internal flow passage(s) of the fuel inlet portion to be formed with a streamlined axial profile. This minimises flow losses in the fuel flow and makes the flow of the fuel more laminar in profile. Optionally, a heat shielding portion is provided between the outer air portion and the inner air portion, the heat shielding portion being configured to reduce a transfer of thermal energy between the outer air portion and the inner air portion. A reduced transfer of thermal energy between the outer air swirler and the inner air swirler is advantageous because it minimises a transfer of thermal energy into the fuel prior to the fuel being atomised with the incoming airflow in the fuel / air swirler assembly. The transfer of thermal energy to the fuel prior to its atomisation is undesirable because it can reduce the effectiveness of the atomisation process, which in turn reduces the efficiency of the engine in operation. In one arrangement, the heat shielding portion is integrally formed with the outer air swirler either on or at a radially inwardly facing surface of the outer air swirler that is oriented towards a radially outwardly facing surface of the inner air swirler. This could be a surface treatment that is applied to the radially inwardly facing surface of the outer air swirler, or alternatively could be a discrete material layer that is moulded or cast into the radially inwardly facing surface of the outer air swirler. In another arrangement, the heat shielding portion is integrally formed with the inner air swirler either on or at a radially outwardly facing surface of the inner air swirler that is oriented towards a radially inwardly facing surface of the outer air swirler. This could be a surface treatment that is applied to the radially outwardly facing surface of the inner air swirler, or alternatively could be a discrete material layer that is moulded or cast into the radially outwardly facing surface of the inner air swirler. In a further alternative arrangement, the heat shielding portion may be an annular air gap that is maintained between a radially inwardly facing surface of the outer air swirler and a radially outwardly facing surface of the inner air swirler. Optionally, the gas turbine engine comprises, in axial flow sequence, a fan assembly, a compressor assembly, a combustor assembly, a turbine assembly, and an exhaust assembly, wherein the combustor assembly comprises a plurality of fuel swirl nozzles. A typical turbofan engine assembly will comprise inter alia a circumferential array of fuel swirl nozzles. For efficient operation of the engine, it is desirable that each of the fuel swirl nozzles has a common fuel / air atomisation characteristic. Optionally, the fan assembly comprises a plurality of fan blades extending radially from a hub, the plurality of fan blades defining a fan diameter (Dfan), and wherein the fan diameter Dfan is within the range of 0.3m to 2.0m, preferably within the range 0.4m to 1,5m, and more preferably in the range of 0.7m to 1,0m. In one embodiment of the disclosure, the fan diameter is 0.9m. Consequently, for the same heat energy loading rejected to the air flow through the heat exchanger, the loss in propulsive efficiency of the turbofan engine is proportionately smaller for a large diameter (for example, approximately 1.5 to 2.0m in diameter) turbofan engine than for a small diameter turbofan engine. The fan tip diameter, measured across a centreline of the engine and between an outermost tip of opposing fan blades at their leading edge, may be in the range from 95 cm to 200 cm, for example in the range from 110 cm to 150 cm, or alternatively in the range from 155 cm to 200 cm. The fan tip diameter may be greater than any of: 110 cm, 115 cm, 120 cm, 125 cm, 130 cm, 135cm, 140 cm, 145 cm, 150 cm, 155 cm, 160 cm, 165 cm, 170 cm, 175 cm, 180 cm, 185 cm, 190 cm or 195 cm. The fan tip diameter may be around 110 cm, 115 cm, 120 cm, 125 cm, 130 cm, 135cm, 140 cm, 145 cm, 150 cm, 155 cm, 160 cm, 165 cm, 170 cm, 175 cm, 180 cm, 185 cm, 190 cm or 195 cm. The fan tip diameter may be greater than 160 cm. The fan tip diameter may be in the range from 95 cm to 150 cm, optionally in the range from 110 cm to 150 cm, optionally in the range of from 110 cm to 145 cm, and further optionally in the range from 120 cm to 140 cm. The fan tip diameter may be in the range from 155 cm to 200 cm, optionally in the range from 160 cm to 200 cm, and further optionally in the range from 165 cm to 190 cm. Optionally, the turbofan gas turbine engine further comprises an outer casing, the outer casing enclosing the sequential arrangement of fan assembly, compressor module, and turbine module, an annular bypass duct being defined between the outer casing and the sequential arrangement of compressor module, and turbine module, a bypass ratio being defined as a ratio of a mass airflow rate through the bypass duct to a mass air flow rate through the sequential arrangement of modules, and wherein the bypass ratio is less than 4.0. A turbofan engine having a bypass ratio (BPR) of less than approximately 4.0 will have a generally smaller bypass duct (the annular duct surrounding the core gas turbine engine) than a turbofan engine having a BPR greater than approximately 4.0. For a turbofan engine with a BPR greater than, say, 4.0, the correspondingly larger bypass duct volume provides more scope for positioning a heat exchanger within the bypass duct than would be the case for a low BPR turbofan engine. A reduced transfer of thermal energy between the outer air swirler and the inner air swirler is advantageous because it minimises a transfer of thermal energy into the fuel prior to the fuel being atomised with the incoming airflow in the fuel / air swirler assembly. The transfer of thermal energy to the fuel prior to its atomisation is undesirable because it can reduce the effectiveness of the atomisation process, which in turn reduces the efficiency of the engine in operation. In one arrangement, the heat shielding portion is integrally formed with the outer air swirler either on or at a radially inwardly facing surface of the outer air swirler that is oriented towards a radially outwardly facing surface of the inner air swirler. This could be a surface treatment that is applied to the radially inwardly facing surface of the outer air swirler, or alternatively could be a discrete material layer that is moulded or cast into the radially inwardly facing surface of the outer air swirler. In another arrangement, the heat shielding portion is integrally formed with the inner air swirler either on or at a radially outwardly facing surface of the inner air swirler that is oriented towards a radially inwardly facing surface of the outer air swirler. This could be a surface treatment that is applied to the radially outwardly facing surface of the inner air swirler, or alternatively could be a discrete material layer that is moulded or cast into the radially outwardly facing surface of the inner air swirler. In a further alternative arrangement, the heat shielding portion may be an annular air gap that is maintained between a radially inwardly facing surface of the outer air swirler and a radially outwardly facing surface of the inner air swirler. As noted elsewhere herein, the present disclosure may relate to a turbofan 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. The fan may comprise any number of stages, for example multiple stages. Each fan stage may comprise a row of fan blades and a row of stator vanes. The stator vanes may be variable stator vanes (in that their angle of incidence may be variable). 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. 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 compressor stages, for example multiple stages. Each compressor stage may comprise a row of rotor blades and a row of stator vanes. The stator vanes may be variable stator vanes (in 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 turbine stages, for example multiple stages. Each turbine 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 diameter of the fan may be measured across the engine centreline and between the tips of opposing fan blades at their leading edge. The fan diameter may be greater than (or on the order of) any of: 50 cm, 60 cm, 70 cm (around 27.5 inches), 80 cm (around 31.5 inches), 90 cm, 100 cm (around 39 inches), 110 cm (around 43 inches), 120 cm (around 47 inches), 130 cm (around 51 inches), 140 cm (around 55 inches), 150 cm (around 59 inches), or 160 cm (around 63 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 50 cm to 70 cm or 90 cm to 130 cm. The fan face area may be calculated as it multiplied by the square of the fan tip radius. 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 10000 rpm, for example less than 9000 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 50 cm to 90 cm (for example 60 cm to 80 cm or 65 cm to 75 cm) may be in the range of from 7000 rpm to 22000 rpm, for example in the range of from 7000 rpm to 16000 rpm, for example in the range of from 7500 rpm to 14000 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 90 cm to 150 cm may be in the range of from 4500 rpm to 12500 rpm, for example in the range of from 4500 rpm to 10000 rpm, for example in the range of from 6000 rpm to 10000 rpm. In use of the turbofan 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 lltip. 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: 0.22, 0.23, 0.24, 0.25, 0.26, 0.27, 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 0.40 (all values being dimensionless). 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 0.30. Turbofan 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: 0.4, 0.5, 0.6, 0.7, 0.8, 0.9, 1.0, 1.1, 1.2, 1.3, 1.4, 1.5, 1.6, 1.7, 1.8, 1.9, 2.0, 2.4, 2.8, 3.2, 3.6, or 4.0. 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 of 0.4 to 1.0, 0.5 to 0.9, or 0.6 to 0.9. Alternatively, the bypass ratio may be in a bounded range in the form of 1.0 to 4.0, 1.8 to 3.6, or 2.4 to 3.6. The bypass duct may be substantially annular. The bypass duct may be radially outside the core engine. 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 turbofan 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-limitative 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: 10, 15, 20, 25, 30, 35 or 40. 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 20 to 35. Specific thrust of a turbofan gas turbine 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 as described and / or claimed herein may be less than (or on the order of) any of the following: 800 Nkg1s, 850 Nkg~1s, 900 Nkg1s, 950 Nkg-1s, 1000 Nkg-1s, 1050 Nkg'1s, 1100 Nkg'1s, 1150 Nkg'1s, 1200 Nkg'1s, 1250 Nkg-1s, 1300 Nkg-1s, 1350 Nkg-1s, or 1400 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 800 Nkg~ 1s to 950 Nkg~1s, or 900 Nkg~1s to 1350 Nkg~1s. Such engines may be particularly efficient in comparison with conventional gas turbine engines. A turbofan 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: 20kN, 40kN, 60kN, 80kN, 100kN, 120kN, 140kN, 160kN, 180kN, or200kN. 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 60kN to 160kN, for example 70kN to 120kN. The thrust referred to above may be the maximum net thrust at standard atmospheric conditions at sea level plus 15 degrees C (ambient pressure 101.3kPa, 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: 1500K, 1550K, WOOK, 1650K, 1700K, 1750K, WOOK, 1850K, WOOK, 1950K, or2000K. 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: 1700K, 1750K, WOOK, 1850K, WOOK, 1950K, 2000K, 2050K, 2100K, 2150K, 2200K, 2250K or 2300K. 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 WOOK to 2200K. 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 turbofan 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 stage of a turbofan gas turbine engine as described and / or claimed herein may have any desired number of fan blades, for example 12, 14, 16, 18, 20, 22, 24, 26, 28, 30, 32, or 34 fan blades. As used herein, cruise conditions have the conventional meaning and would be readily understood by the skilled person. In this regard, cruise conditions encompass both subsonic cruise conditions and supersonic cruise conditions. Thus, for a given turbofan 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 turbofan 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 subsonic cruise condition may be any point in the range of from Mach 0.80 to 0.99, for example 0.80 to 0.85, for example 0.85 to 0.90, for example 0.90 to 0.95, for example 0.95 to 0.99, for example in the region of Mach 0.80, in the region of Mach 0.85 or in the range of from 0.80 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 0.80. Purely by way of example, the forward speed at the supersonic cruise condition may be any point in the range of from Mach 1.20 to 2.20, for example 1.35 to 2.10, for example 1.50 to 2.05, for example in the region of Mach 2.00 or in the range of from 1.80 to 2.00. 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 between Mach 1.0 and 1.20, or above Mach 2.20. Purely by way of example, the supersonic 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 11000m to 19000m, for example in the range of from 12500m to 17000m, for example in the range of from 15000m to 17000m (around 56000 ft), for example in the range of from 16000m to 17000m, for example in the region of 17000m. The cruise conditions may correspond to standard atmospheric conditions at any given altitude in these ranges. Purely by way of example, the subsonic 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 40kN to 65kN) at a forward Mach number of 0.8 and standard atmospheric conditions (according to the International Standard Atmosphere) at an altitude of 38000ft (11582m). Purely by way of further example, the supersonic 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 70kN to 120kN) at a forward Mach number of 1.50 and standard atmospheric conditions (according to the International Standard Atmosphere) at an altitude of 56000ft (17000m). In use, a turbofan 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 of the disclosure, there is provided an aircraft comprising a turbofan 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 of the disclosure, there is provided a method of operating a turbofan 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 of the disclosure, there is provided a method of operating an aircraft comprising a turbofan 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. Other aspects of the disclosure provide devices, methods and systems which include and / or implement some or all of the actions described herein. The illustrative aspects of the disclosure are designed to solve one or more of the problems herein described and / or one or more other problems not discussed. Brief Description of the Drawings There now follows a description of an embodiment of the disclosure, by way of nonlimiting example, with reference being made to the accompanying drawings in which: Figure 1 shows a schematic sectional view of a turbofan gas turbine engine according to the prior art; Figure 2 shows a schematic sectional view of a combustion chamber with a fuel swirl nozzle according to an embodiment of the disclosure; and Figure 3 shows a sectional view of a fuel swirl nozzle according to an embodiment of the disclosure; Figure 4 shows a perspective view of the fuel swirl nozzle shown in Figure 2; and Figure 5 shows a perspective, cutaway view of the fuel swirl nozzle of Figure 4 highlighting the oversized first vane, oversized second vane, and oversized third vane. It is noted that the drawings may not be to scale. The drawings are intended to depict only typical aspects of the disclosure, and therefore should not be considered as limiting the scope of the disclosure. In the drawings, like numbering represents like elements between the drawings. Detailed Description Figure 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. 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. 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. 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. 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 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. 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 1), and a circumferential direction (perpendicular to the page in the Figure 1 view). The axial, radial and circumferential directions are mutually perpendicular. Figure 2 shows a schematic sectional view of a combustion equipment 16 of a turbofan gas turbine engine 10 having a fuel swirl nozzle 100 according to an embodiment of the disclosure. The combustion equipment 16 comprises a fuel swirl nozzle 100 at an upstream end of a combustion chamber. The fuel swirl nozzle 100 is downstream of a high-pressure compressor 15. The fuel swirl nozzle 100 is positioned at a proximal end of a fuel stem portion (not indicated) that itself is positioned facing radially inwardly from a casing of the turbofan gas turbine engine 10. Figure 3, 4 and 5 show an exemplary arrangement for the fuel swirl nozzle 100 for a gas turbine engine 10 according to an embodiment of the disclosure. The fuel swirl nozzle 100 comprises an inner air portion 110, an outer air portion 140, a dome air portion 160, and a fuel delivery portion 180. The inner air portion 110 comprises an inner air passage 112. The inner air passage 112 comprises a first annular upstream portion 114 that is fluidly connected to an opposite second cylindrical downstream portion 126. The downstream end 130 of the second cylindrical downstream portion 126 defines a pre-filmer edge 132. The outer air portion 140 comprises an outer air passage 142. The outer air passage 142 is positioned concentrically with, and radially outwardly of, the second downstream portion 126 of the inner air passage 112. The outer air passage 142 comprises a first intake end 152, and a second opposite exhaust end 154. The dome air portion 160 comprises a dome air passage 162. The dome air passage 162 is positioned concentrically with, and radially outwardly of, the outer air passage 142. The dome air passage 162 comprises a first intake end 172, and a second opposite exhaust end 174. The fuel delivery portion 180 comprises a fuel delivery passage 182. The fuel delivery passage 182 comprises an inlet portion 184 and an outlet portion 186. The inlet portion 184 is configured to receive a flow of fuel, and the outlet portion 186 is arranged concentrically between a radially outwardly facing surface 128 of the second downstream portion 126 of the inner air passage 112, and a radially inwardly facing surface 156 of the second exhaust end 154 of the outer air passage 142. The first annular upstream portion 114 of the inner air passage 112 comprises a first circumferential array 116 of first vanes 118. The first circumferential array 116 is configured to impart a first rotational swirl to a first air flow 120 entering the first annular upstream portion 114. The first rotational swirl has a first swirl angle relative to a longitudinal axis 9 of the turbofan gas turbine engine 10. The outer air passage 142 comprises a second circumferential array 144 of second vanes 146 configured to impart a second rotational swirl to a second airflow 148 entering the first intake end 152 of the outer air passage 142. The second rotational swirl has a second swirl angle relative to the longitudinal axis 9 of the turbofan gas turbine engine 10. The dome air passage 162 comprises a third circumferential array 164 of third vanes 166 configured to impart a third rotational swirl to a third airflow 168 entering the first intake end 172 of the dome air passage 162. The third rotational swirl has a third swirl angle relative to the longitudinal axis 9 of the turbofan gas turbine engine 10. The flow of fuel is delivered into an annular volume 188 that is defined between the radially outwardly facing surface 128 of the second downstream portion 126 of the inner air passage 112, and the radially inwardly facing surface 156 of the second exhaust end 154 of the outer air passagel 42. The flow of fuel is exhausted from the annular volume 188 into, in axial flow sequence, the first airflow 120 exhausted from the inner air passage 112, the second airflow 148 exhausted from the outer air passagel 42, and the third air flow 168 exhausted from the dome air passagel 62. The second exhaust end 154 of the outer air passage 142 directs the second air flow 148 that is exhausted from the outer air passage 142, in a radially inwardly direction. Each of the first vanes 118 has radial height (H) 122. An axial spacing (S) 124 is defined between a downstream-most edge of the first circumferential vane array 116 and an upstream-most edge of the second circumferential vane array 144, and the axial spacing (S) 124 is in a range of between H and 5*H. In the present arrangement, each of the first circumferential array 116 of first vanes 118, the second circumferential array 144 of second vanes 146, and the third circumferential array 164 of third vanes 166 are circumferentially equally spaced. In other arrangements of the disclosure, one or two of the first circumferential array 116, the second circumferential array 144, and the third circumferential array 164 may be circumferentially unequally spaced. In the present embodiment, the first circumferential array 116 of first vanes 118 comprises a circumferential array of between four (4) vanes 118. In other arrangements, the first circumferential array 116 may comprise three (3) vanes, five (5) vanes, or six (6) vanes. In the present arrangement, the second circumferential array 144 of second vanes 146 comprises a circumferential array 144 of twelve (12) vanes 146. In other arrangements of the disclosure, the second circumferential array 144 may comprise between eight (8) vanes and fourteen (14) vanes. In the present arrangement, the third circumferential array 164 of third vanes 166 comprises a circumferential array 164 of twelve (12) vanes 166. In other arrangements of the disclosure, the third circumferential array 164 may comprise between eight (8) vanes and fourteen (14) vanes. In the present arrangement, the second circumferential array 144 of second vanes 146 comprises a single oversized second vane 150. The single oversized vane 150 has a greater thickness and a greater chord than the remaining second vanes 148. In the present arrangement, the third circumferential array 164 of third vanes 166 comprises a single oversized third vane 170. The single oversized vane 170 has a greater thickness and a greater chord than the remaining second vanes 166. The purpose of the oversized second vane 150, and the oversized third vane 170 is to allow for the passage of fluids in a radially inwardly direction without requiring a separate fluid pipe or route that would obstruct the corresponding air flow passing through the second circumferential array 144 and third circumferential array 164. The second exhaust end 154 of the outer air passage 142 extends axially past the second downstream end 130 of the inner air passage 112. The second exhaust end 174 of the dome air passage 162 extends axially past the second exhaust end 154 of the outer air passage 142. Each of the first rotational swirl, the second rotational swirl, and the third rotational swirl is in the same rotational direction around the longitudinal axis of the fuel swirl nozzle 100. In the present arrangement, each of the inner air portion 110, the outer air portion 140, the dome air portion 160, and the fuel delivery portion 180 is formed by an additive layer manufacturing process. This enables the geometrical features of the first, second, and third circumferential arrays 116,144,164 to be generated without the need for additional manufacturing finishing operations. In the present arrangement, the inner air portion 110 and the outer air portion 140 are joined to one another by a welding or brazing process. In the present embodiment, the inner air portion 110 and the outer air portion 140 are manufactured separately by an additive layer manufacturing process. The inner air portion 110 and the outer air portion 140 are subsequently joined to one another by a brazing process. In this way the fuel wetted surfaces of the inner air portion 110 and the outer air portion 140 can be machined after they are manufactured. This enables a very smooth surface finish on these surfaces to be obtained, which in turn reduces the propensity for coke to form and accumulate on the surfaces. This reduces the need for a user to maintain the fuel swirl nozzle 100, which is advantageous. In the present arrangement, the fuel delivery portion 180 is formed as a monolithic additive layer manufactured component. In the present arrangement, a heat shielding portion 134 is provided between the outer air portion 140 and the inner air portion 110. The heat shielding portion 134 is configured to reduce a transfer of thermal energy between the outer air portion 140 and the inner air portion 110. The gas turbine engine 100 comprises, in axial flow sequence, a fan assembly 23, a compressor assembly 14,15, a combustor assembly 16, a turbine assembly 17,19, and an exhaust assembly 20, wherein the combustor assembly 16 comprises a plurality of fuel swirl nozzles 100. The fan assembly 23 comprises a plurality of fan blades extending radially from a hub, the plurality of fan blades defining a fan diameter (Dfan), and wherein the fan diameter Dfan is within the range of 0.3m to 2.0m, preferably within the range 0.4m to 1.5m, and more preferably in the range of 0.7m to 1.0m. The turbofan gas turbine engine 10 comprises an outer casing , the outer casing enclosing the sequential arrangement of fan assembly 23, compressor module 14,15, and turbine module 17,19, an annular bypass duct 22 being defined between the outer casing and the sequential arrangement of compressor module, and turbine module, a bypass ratio being defined as a ratio of a mass airflow rate through the bypass duct 22 to a mass air flow rate through the sequential arrangement of modules, and wherein the bypass ratio is less than 4.0 While the invention has been described in conjunction with the exemplary embodiments described above, many equivalent modifications and variations will be apparent to those skilled in the art when given this disclosure. Accordingly, the exemplary embodiments of the invention set forth above are considered to be illustrative and not limiting. Moreover, in determining extent of protection, due account shall be taken of any element which is equivalent to an element specified in the claims. Various changes to the described embodiments may be made without departing from the spirit and scope of the invention.

Claims

1. A fuel swirl nozzle for a gas turbine engine, the fuel swirl nozzle comprising an inner air portion, an outer air portion, a dome air portion, and a fuel delivery portion, the inner air portion comprising an inner air passage, the inner air passage comprising a first annular upstream portion fluidly connected to an opposite second cylindrical downstream portion, a downstream end of the second cylindrical downstream portion defining a pre-filmer edge, the outer air portion comprising an outer air passage, the outer air passage being positioned concentrically with, and radially outwardly of, the second downstream portion of the inner air passage, and comprising a first intake end, and a second opposite exhaust end, the dome air portion comprising a dome air passage, the dome air passage being positioned concentrically with, and radially outwardly of, the outer air passage, and comprising a first intake end, and a second opposite exhaust end, andthe fuel delivery portion comprising a fuel delivery passage, the fuel delivery passage comprising an inlet portion and an outlet portion, the inlet portion being configured to receive a flow of fuel, and the outlet portion being arranged concentrically between a radially outwardly facing surface of the second downstream portion of the inner air passage, and a radially inwardly facing surface of the second exhaust end of the outer air passage, and wherein the first annular upstream portion of the inner air passage comprises a first circumferential array of first vanes configured to impart a first rotational swirl to a first air flow entering the first annular upstream portion, the first rotational swirl having a first swirl angle relative to the longitudinal axis of the engine, the outer air passage comprises a second circumferential array of second vanes configured to impart a second rotational swirl to a second airflow entering the first intake end of the outer air passage, the second rotationalswirl having a second swirl angle relative to the longitudinal axis of the engine,the dome air passage comprises a third circumferential array of third vanes configured to impart a third rotational swirl to a third air flow entering the first intake end of the dome air passage, the third rotational swirl having a third swirl angle relative to the longitudinal axis of the engine, andthe flow of fuel is delivered into an annular volume defined between the radially outwardly facing surface of the second downstream portion of the inner air passage, and the radially inwardly facing surface of the second exhaust end of the outer air passage, and the flow of fuel is exhausted from the annular volume into, in axial flow sequence, the first airflow exhausted from the inner air passage, the second air flow exhausted from the outer air passage, and the third air flow exhausted from the dome air passage.

2. The fuel swirl nozzle as claimed in Claim 1, wherein the second exhaust end of the outer air passage directs the second air flow exhausted from the outer air passage in a radially inwardly direction.

3. The fuel swirl nozzle as claimed in Claim 1 or Claim 2, wherein each of the first vanes has radial height H, and an axial spacing S is defined between a downstream-most edge of the first vane array and an upstream-most edge of the second vane array, and the axial spacing S is in a range of between H and 5*H.

4. The fuel swirl nozzle as claimed in any one of Claims 1 to 3, wherein at least one of the first circumferential array of first vanes, the second circumferential array of second vanes, and the third circumferential array of third vanes are circumferentially equally spaced.

5. The fuel swirl nozzle as claimed in any one of Claims 1 to 4, wherein at least one of the first circumferential array of first vanes, the second circumferential array of second vanes, and the third circumferential array of third vanes are circumferentially unequally spaced.

6. The fuel swirl nozzle as claimed in any one of Claims 1 to 5, wherein the first circumferential array of first vanes comprises a circumferential array of between three (3) vanes and six (6) vanes.

7. The fuel swirl nozzle as claimed in any one of Claims 1 to 6, wherein the second circumferential array of second vanes comprises a circumferential array of between eight (8) vanes and fourteen (14) vanes.

8. The fuel swirl nozzle as claimed in any one of Claims 1 to 7, wherein the third circumferential array of third vanes comprises a circumferential array of between eight (8) vanes and fourteen (14) vanes.

9. The fuel swirl nozzle as claimed in any one of Claims 1 to 8, wherein the second circumferential array of second vanes comprises a single oversized second vane, the single oversized vane having a greater thickness and / or a greater chord than the remaining second vanes.

10. The fuel swirl nozzle as claimed in any one of Claims 1 to 9, wherein the third circumferential array of third vanes comprises a single oversized third vane, the single oversized third vane having a greater thickness and / or a greater chord than the remaining third vanes.

11. The fuel swirl nozzle as claimed in any one of Claims 1 to 10, wherein the second exhaust end of the outer air passage extends axially past the second downstream end of the inner air passage.

12. The fuel swirl nozzle as claimed in any one of Claims 1 to 10, wherein the second exhaust end of the dome air passage extends axially past the second exhaust end of the outer air passage.

13. The fuel swirl nozzle as claimed in any one of Claims 1 to 12, wherein each of the first rotational swirl, the second rotational swirl, and the third rotational swirl is in the same rotational direction around a longitudinal axis of the fuel swirl nozzle.

14. The fuel swirl nozzle as claimed in any one of Claims 1 to 13, wherein at least one of the inner air portion, the outer air portion, the dome air portion, and the fuel delivery portion is formed by an additive layer manufacturing process.

15. The fuel swirl nozzle as claimed in any one of Claims 1 to 14, wherein the inner air portion and the outer air portion are joined to one another by a welding or brazing process.

16. The fuel swirl nozzle as claimed in any one of Claims 1 to 15, wherein the fuel delivery portion is formed as a monolithic additive layer manufactured component.

17. The fuel swirl nozzle as claimed in any one of Claims 1 to 16, wherein a heat shielding portion is provided between the outer air portion and the inner air portion, the heat shielding portion being configured to reduce a transfer of thermal energy between the outer air portion and the inner air portion.

18. The fuel swirl nozzle as claimed in any one of Claims 1 to 17, wherein the gas turbine engine comprises, in axial flow sequence, a fan assembly, a compressor assembly, a combustor assembly, a turbine assembly, and an exhaust assembly, wherein the combustor assembly comprises a plurality of fuel swirl nozzles.

19. The fuel swirl nozzle as claimed in Claim 18, wherein the fan assembly comprises a plurality of fan blades extending radially from a hub, the plurality of fan blades defining a fan diameter (Dfan), and wherein the fan diameter Dfan is within the range of 0.3m to 2.0m, preferably within the range 0.4m to 1,5m, and more preferably in the range of 0.7m to 1,0m.

20. The fuel swirl nozzle as claimed in Claim 18 or Claim 19, wherein the turbofan gas turbine engine further comprises an outer casing, the outer casing enclosing the sequential arrangement of fan assembly, compressormodule, and turbine module, an annular bypass duct being defined between the outer casing and the sequential arrangement of compressor module, and turbine module, a bypass ratio being defined as a ratio of a mass air flow rate through the bypass duct to a mass air flow rate through the sequential5 arrangement of modules, and wherein the bypass ratio is less than 4.0.

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