Fuel spray nozzle

Abstract

Fuel spray nozzle for generating a spray of atomised liquid fuel in a combustor of a gas turbine engine. The nozzle includes a flow circuit that has in flow series: a gallery that receives fuel flow, plural circumferentially-spaced restrictor passages arranged in a row around the nozzle, plural conditioning passages configured to impart a circumferential component to their respective portions of the fuel flow, and an annular spin chamber which forms a swirling fuel flow which is discharged at an exit port. The restrictor passages form flow restrictions which in use produce a pressure differential between the gallery and the spin chamber to evenly circumferentially distribute the fuel flow between the restrictor passages. The conditioning passages have increased flow cross-sectional areas relative to the flow cross-sectional areas of the restrictor passages, such that the restrictor passages produce substantially all of the pressure differential between the gallery and the spin chamber.

Claims

1. A fuel spray nozzle for generating a spray of atomised liquid fuel in a combustor of a gas turbine engine, wherein the fuel spray nozzle includes: a flow circuit having an inlet port for receiving a flow of liquid fuel and having an annular exit port for discharging the received flow of liquid fuel as a swirling fuel flow; and an annular prefilming surface downstream of the annular exit port, the annular prefilming surface configured such that the swirling fuel flow received from the annular exit port spreads, as a film of fuel, across the prefilming surface, whereupon one or more swirling air flows generated by the fuel spray nozzle shear the film of fuel towards a trailing edge of the annular prefilming surface and atomise the film of fuel into a spray of fine droplets; wherein the flow circuit has in flow series: a gallery which wraps circumferentially around the fuel spray nozzle and receives the flow of liquid fuel from the inlet port; a plurality of circumferentially-spaced restrictor passages arranged in a row around the fuel spray nozzle, the plurality of restrictor passages receiving respective portions of the flow of liquid fuel from the gallery; a plurality of conditioning passages which respectively receive the respective portions of the flow of liquid fuel from respective restrictor passages of the plurality of restrictor passages, the plurality of conditioning passages configured to impart a circumferential velocity component to their respective portions of the flow of liquid fuel; and an annular spin chamber which receives and recombines the respective portions of the flow of liquid fuel from the plurality of conditioning passages to form the swirling fuel flow which is discharged at the annular exit port; and wherein: each restrictor passage of the plurality of restrictor passages form a flow restrictions which in use produces a pressure differential between the gallery and the annular spin chamber to evenly circumferentially distribute the flow of liquid fuel between the plurality of restrictor passages; and each conditioning passage of the plurality of conditioning passages has increased flow cross-sectional areas relative to the flow cross-sectional areas of respective restrictor passage of the plurality of restrictor passages, such that the plurality of restrictor passages produce substantially all of the pressure differential between the gallery and the annular spin chamber.

2. The fuel spray nozzle of claim 1, wherein each restrictor passage of the plurality of restrictor passages extend substantially parallel to each other in an axial direction of the fuel spray nozzle.

3. The fuel spray nozzle of claim 1, wherein each conditioning passage of the plurality of conditioning passages smoothly increase in flow cross-sectional area with downstream distance from the plurality of restrictor passages.

4. The fuel spray nozzle of claim 1, wherein each conditioning passage of the plurality of conditioning passages is further configured to impart a radial velocity component to its respective portion of the flow of liquid fuel.

5. The fuel spray nozzle of claim 1, wherein the flow cross-sectional areas of each conditioning passage of the plurality of conditioning passages is at least two times the flow cross-sectional areas of its respective restrictor passage of the plurality of restrictor passages.

6. The fuel spray nozzle of claim 1, wherein the flow circuit is a mains flow circuit, the flow of liquid fuel received at the inlet port being a mains fuel flow, and wherein the fuel spray nozzle further includes a pilot flow circuit for receiving and discharging a separate pilot fuel flow, whereby the fuel spray nozzle is able to implement staged combustion of the mains and pilot fuel flows.

7. The fuel spray nozzle of claim 1, wherein an atomiser subassembly of the fuel spray nozzle defines the flow circuit, the atomiser subassembly being formed by additive layer manufacture.

8. Combustion equipment for a gas turbine engine, the combustion equipment including a combustor and a plurality of the fuel spray nozzles of claim 1 for generating respective sprays of atomised liquid fuel in the combustor.

9. A gas turbine engine for an aircraft having an engine core, the gas turbine engine comprising in axial flow series a compressor, the combustion equipment of claim 8, and a turbine, a core shaft connecting the turbine to the compressor.

10. The gas turbine engine of claim 9, further comprising: a fan located upstream of the engine core, the fan comprising a plurality of fan blades; and a gearbox that receives an input from the core shaft and outputs drive to the fan so as to drive the fan at a lower rotational speed than the core shaft.

11. The gas turbine engine of claim 9, wherein: the compressor is a first compressor, the turbine is a first turbine, and the core shaft is a first core shaft; the engine core further comprises a second compressor between the first compressor and the combustion equipment, a second turbine between the combustion equipment and the first turbine, and a second core shaft connecting the second turbine to the second compressor; and the second core shaft is arranged to rotate at a higher rotational speed than the first core shaft.

Description

DESCRIPTION OF THE DRAWINGS

(1) Embodiments will now be described by way of example only, with reference to the Figures, in which:

(2) FIG. 1 is a sectional side view of a gas turbine engine;

(3) FIG. 2 is a close up sectional side view of an upstream portion of a gas turbine engine;

(4) FIG. 3 is a partially cut-away view of a gearbox for a gas turbine engine;

(5) FIG. 4 shows a fuel spray nozzle with a cut away portion revealing part of an atomiser subassembly that defines a mains flow circuit;

(6) FIG. 5 shows the mains flow circuit in perspective view as a “negative” body

(7) FIG. 6 shows the mains flow circuit in top view as a “negative” body;

(8) FIG. 7 shows schematically restrictor passages and the conditioning passage of the mains flow circuit;

(9) FIG. 8 shows the result of a comparative CFD analysis in which a mains flow circuit is provided with J-shaped passages; and

(10) FIG. 9 shows the result of CFD analysis in which a mains flow circuit has restrictor passages and conditioning passages of the type shown in FIGS. 5 and 6.

DETAILED DESCRIPTION

(11) 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.

(12) FIG. 1 illustrates a gas turbine engine 10 having a principal rotational axis 9. The engine 10 comprises an air intake 12 and a propulsive fan 23 that generates two airflows: a core airflow A and a bypass airflow B. The gas turbine engine 10 comprises a core 11 that receives the core airflow A. The engine core 11 comprises, in axial flow series, a low pressure compressor 14, a high-pressure compressor 15, combustion equipment 16, a high-pressure turbine 17, a low pressure turbine 19 and a core exhaust nozzle 20. A nacelle 21 surrounds the gas turbine engine 10 and defines a bypass duct 22 and a bypass exhaust nozzle 18. The bypass airflow B flows through the bypass duct 22. The fan 23 is attached to and driven by the low pressure turbine 19 via a shaft 26 and an epicyclic gearbox 30.

(13) 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.

(14) An exemplary arrangement for a geared fan gas turbine engine 10 is shown in FIG. 2. The low pressure turbine 19 (see FIG. 1) drives the shaft 26, which is coupled to a sun wheel, or sun gear, 28 of the epicyclic gear arrangement 30. Radially outwardly of the sun gear 28 and intermeshing therewith is a plurality of planet gears 32 that are coupled together by a planet carrier 34. The planet carrier 34 constrains the planet gears 32 to precess around the sun gear 28 in synchronicity whilst enabling each planet gear 32 to rotate about its own axis. The planet carrier 34 is coupled via linkages 36 to the fan 23 in order to drive its rotation about the engine axis 9. Radially outwardly of the planet gears 32 and intermeshing therewith is an annulus or ring gear 38 that is coupled, via linkages 40, to a stationary supporting structure 24.

(15) 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.

(16) The epicyclic gearbox 30 is shown by way of example in greater detail in FIG. 3. Each of the sun gear 28, planet gears 32 and ring gear 38 comprise teeth about their periphery to intermesh with the other gears. However, for clarity only exemplary portions of the teeth are illustrated in FIG. 3. There are four planet gears 32 illustrated, although it will be apparent to the skilled reader that more or fewer planet gears 32 may be provided within the scope of the claimed invention. Practical applications of a planetary epicyclic gearbox 30 generally comprise at least three planet gears 32.

(17) The epicyclic gearbox 30 illustrated by way of example in FIGS. 2 and 3 is of the planetary type, in that the planet carrier 34 is coupled to an output shaft via linkages 36, with the ring gear 38 fixed. However, any other suitable type of epicyclic gearbox 30 may be used. By way of further example, the epicyclic gearbox 30 may be a star arrangement, in which the planet carrier 34 is held fixed, with the ring (or annulus) gear 38 allowed to rotate. In such an arrangement the fan 23 is driven by the ring gear 38. By way of further alternative example, the gearbox 30 may be a differential gearbox in which the ring gear 38 and the planet carrier 34 are both allowed to rotate.

(18) It will be appreciated that the arrangement shown in FIGS. 2 and 3 is by way of example only, and various alternatives are within the scope of the present disclosure. Purely by way of example, any suitable arrangement may be used for locating the gearbox 30 in the engine 10 and/or for connecting the gearbox 30 to the engine 10. By way of further example, the connections (such as the linkages 36, 40 in the FIG. 2 example) between the gearbox 30 and other parts of the engine 10 (such as the input shaft 26, the output shaft and the fixed structure 24) may have any desired degree of stiffness or flexibility. By way of further example, any suitable arrangement of the bearings between rotating and stationary parts of the engine (for example between the input and output shafts from the gearbox and the fixed structures, such as the gearbox casing) may be used, and the disclosure is not limited to the exemplary arrangement of FIG. 2. For example, where the gearbox 30 has a star arrangement (described above), the skilled person would readily understand that the arrangement of output and support linkages and bearing locations would typically be different to that shown by way of example in FIG. 2.

(19) 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.

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

(21) Other gas turbine engines to which the present disclosure may be applied may have alternative configurations. For example, such engines may have an alternative number of compressors and/or turbines and/or an alternative number of interconnecting shafts. By way of further example, the gas turbine engine shown in FIG. 1 has a split flow nozzle 18, 20 meaning that the flow through the bypass duct 22 has its own nozzle 18 that is separate to and radially outside the core exhaust nozzle 20. However, this is not limiting, and any aspect of the present disclosure may also apply to engines in which the flow through the bypass duct 22 and the flow through the core 11 are mixed, or combined, before (or upstream of) a single nozzle, which may be referred to as a mixed flow nozzle. One or both nozzles (whether mixed or split flow) may have a fixed or variable area. Whilst the described example relates to a turbofan engine, the disclosure may apply, for example, to any type of gas turbine engine, such as an open rotor (in which the fan stage is not surrounded by a nacelle) or turboprop engine, for example. In some arrangements, the gas turbine engine 10 may not comprise a gearbox 30.

(22) The geometry of the gas turbine engine 10, and components thereof, is defined by a conventional axis system, comprising an axial direction (which is aligned with the rotational axis 9), a radial direction (in the bottom-to-top direction in FIG. 1), and a circumferential direction (perpendicular to the page in the FIG. 1 view). The axial, radial and circumferential directions are mutually perpendicular.

(23) The combustion equipment 16 of the engine 10 includes a plurality of fuel injectors having fuel spray nozzles which combine pilot and mains fuel flows and air flows to generate sprays of atomised liquid fuel into a combustion chamber. FIG. 4 shows one these nozzles 40 with a cut away portion revealing part of an atomiser subassembly 41 that defines the mains flow circuit 50 for the mains fuel flow within the nozzle. The nozzle has a connector 42 for joining the nozzle to a feed arm (not shown) of the injector, the feed arm providing passages through which the pilot and mains fuel flows are transported to the nozzle. The pilot fuel flow has its own pilot flow circuit (not shown) in the atomiser subassembly. This circuit has upstream passages which are in close proximity to the mains flow circuit enabling the pilot fuel flow to reduce the wetted wall temperature of the mains flow circuit. Thereafter the pilot flow circuit carries the pilot fuel flow into a central portion of nozzle where a spray is generated from the pilot fuel flow separately from the spray formed from the mains fuel flow. The air flows for forming these fuel sprays (derived from the high pressure air produced by high-pressure compressor 15) enter the right hand side of the nozzle as drawn in FIG. 4, and the fuel sprays exit from the left hand side of the nozzle.

(24) The mains flow circuit 50 is shown in perspective view in FIG. 5 and top view in FIG. 6 as a “negative” body (i.e. a solid body representing what is in reality a cavity defined by the atomiser subassembly 41). The mains fuel flow enters the flow circuit at an inlet port 51 from the feed arm, and then flows into a gallery 52 which wraps circumferentially around the nozzle. Plural circumferentially-spaced restrictor passages 53 arranged in a row around the nozzle, receive respective portions of the mains fuel flow from the gallery. Each restrictor passage delivers its portion into a respective a conditioning passage 54, which in turn delivers the portion into an annular spin chamber 55 where the portions from all the conditioning passages are recombined. From an annular exit port 56 at the downstream end of the spin chamber, the mains fuel flow is discharged as a swirling flow onto an annular prefilming surface (not shown) of the nozzle for atomisation at a trailing edge of the surface into a spray of fine droplets.

(25) The restrictor passages 53 (in use) produce a pressure differential between the gallery 52 and the spin chamber 55 to evenly circumferentially distribute the fuel flow between the restrictor passages for the entire range of flow conditions of the mains fuel flow. The conditioning passages 54 then impart a circumferential component to their respective portions of the mains fuel flow, but without producing any significant further pressure differential between the gallery and the spin chamber. This is achieved, at least in part, by the conditioning passages having increased flow cross-sectional areas relative to the flow cross-sectional areas of the restrictor passages.

(26) Advantageously, the atomiser subassembly 41 can be formed by additive layer manufacturing. This enables configurations for the restrictor passages 53 and the conditioning passage 54 to be achieved which would be impossible by conventional subtractive machining. For example, the conditioning passages can be provided with precisely shaped deflector walls, in close proximity to the exits from the restrictor passages, which assist with the flow turning performance of the passages but without impinging on the measuring functionality of the restrictor passages. More generally, computational fluid dynamics (CFD) modelling can be used to inform the shaping of the restrictor 53 and conditioning 54 passages, with the confidence that additive layer manufacturing allows optimised configurations resulting from such modelling to be implemented. Such optimised configurations can minimise losses and gently guide the flow into a swirling motion within the spin chamber 55.

(27) FIG. 7 shows schematically the restrictor passages 53 and the conditioning passage 54. The restrictor passages extend substantially parallel to each other in the axial direction of the nozzle. The conditioning passages have deflector walls 54′ which intercept and align the flow ejected from the axially oriented restrictor passages. Additive layer manufacturing allows the circumferential distance p between the exit from each restrictor passage and the respective deflector wall to be reduced or eliminated entirely. The resultant swirl condition from the conditioning passages is a function of the geometry of the deflector walls, such as exit slope angle β, and the length s of the deflector walls measured circumferentially from the exits from the restrictor passages. Furthermore, to avoid creating further flow constrictions downstream of the restrictor passages, the flow cross-sectional area x of the conditioning passages may be at least two times larger, and preferably at least three of four times larger than the flow cross-sectional area d of the restrictor passages (FIG. 7 shows x at least four times greater than d). The flow cross-sectional area of the conditioning passages also preferably smoothly increases with downstream distance from the restrictor passages. Optionally, the conditioning passages may be further configured to impart a radial component to the fuel flow.

(28) FIG. 8 shows the result of comparative CFD analysis in which a mains flow circuit is provided with J-shaped passages 60 of constant flow cross-sectional area between the gallery and the spin chamber which provide both flow measuring and swirl-producing functions, i.e. the circuit does not have restrictor passages and separate conditioning passages as described above. The modelled velocity field of the flow in the spin chamber shows high variation and local disturbances 61.

(29) FIG. 9 shows by contrast the result of CFD analysis in which the mains flow circuit has the restrictor passages 53 and conditioning passages 54 of the type shown in FIGS. 5 and 6. Advantageously, the velocity field of the flow in the spin chamber is much more even and the local disturbances are eliminated.

(30) 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.