GB2588136A

GB2588136A - Fuel spray nozzle for a gas turbine engine

Info

Publication number: GB2588136A
Application number: GB1914569.7A
Authority: GB
Inventors: Masters Jonathan; Stiles Andrew
Original assignee: Rolls Royce PLC
Current assignee: Rolls Royce PLC
Priority date: 2019-10-09
Filing date: 2019-10-09
Publication date: 2021-04-21
Also published as: GB201914569D0

Abstract

A fuel spray nozzle 100 for a combustion system of a gas turbine engine (10 Fig 1) that has a principal rotational axis (9 Fig 1). The nozzle has a mounting flange 110, a nozzle head 150 which in use is located in the engine radially inwardly of the mounting flange and a feed arm 135 extending between the mounting flange and the nozzle head. The arm has a central axis (165 Fig. 5), and a cross-section through the arm (‘A’ Fig 5) that is perpendicular to the axis at a defined radial position between the mounting flange and the nozzle head, is generally elliptical in shape. The cross-section may have a major axis whose length is at least 1.4 times greater than that of its minor axis. The cross-section may also have its major and minor axis vary along the length of the feed arm, and the length of the major axis may decrease further away from the mounting flange. The nozzle may be additively manufactured and the elliptical design increases frequency separation between vibration modes and improves aerodynamic and strength characteristics. Bird strike tolerance is also improved.

Description

FUEL SPRAY NOZZLE FOR A GAS TURBINE ENGINE

Field of the disclosure

The present disclosure relates to a fuel spray nozzle for a combustion system of a gas turbine engine.

Background

Gas turbine engines for aircraft may be provided with fuel spray nozzles that distribute fuel into an airflow for subsequent combustion in a combustor. Conventional fuel spray nozzles may be provided with a nozzle head on a feed arm projecting radially inwardly from a suitable surface such as a combustor outer casing. When the fuel spray nozzle is subjected to vibration from the combustor casing it will respond at modal frequencies determined by the stiffness of the structure and the modulus of the material at particular temperatures. The stiffnesses of the feed arm and of the flange/bolted joint which fixes the nozzle to the casing, along with the dynamic mass are the dominant parameters that determine the frequencies of the first two modes of the fuel spray nozzle. The first two dynamic modes are cantilever modes in the axial and tangential directions, and existing feed arms suffer from a small frequency separation between these modes providing a complex mode shape with orbiting momentum.

It is desirable to provide a feed arm configuration that not only improves the complex mode shape but, where possible, also improves other considerations such as the ability to handle bird strikes, weight considerations, manufacturability and drag coefficients.

Summary

The present disclosure provides a fuel spray nozzle, a combustion system, a gas turbine engine, and the use of additive layer manufacture to form a fuel spray nozzle, as set out in the appended claims.

According to a first aspect there is provided a fuel spray nozzle for a combustion system of a gas turbine engine having a principal rotational axis. The fuel spray nozzle comprises a mounting flange, a nozzle head which, in use, is located in the engine radially inwardly of the mounting flange, and a feed arm extending between the mounting flange and the nozzle head. The feed arm has a central axis and is configured such that a cross-section through the feed arm perpendicular to the central axis at a given radial position between the mounting flange and the nozzle head is generally elliptical in shape.

By providing a stem having a generally elliptical shape the frequency separation between the first two modes is improved and the fuel spray nozzle may benefit from improved aerodynamic and strength characteristics.

According to a second aspect there is provided a combustion system of a gas turbine engine, the combustion system having a circumferential row of the fuel spray nozzles according to the first aspect.

According to a third aspect there is provided a gas turbine engine for an aircraft, the gas turbine engine comprising a combustion system according to the second aspect.

According to a fourth aspect there is provided the use of additive layer manufacturing to form the fuel spray nozzle according to the first aspect.

Optional features of the present disclosure will now be set out. These are applicable singly or in any combination with any aspect of the present disclosure.

The generally elliptical cross-section through the feed arm at a radial position adjacent the mounting flange may have a major axis whose length is at least 1.4 times, and preferably at least 1.6 times, greater than that of its minor axis.

The feed arm is typically for positioning the nozzle head in an airflow and may have a leading edge, a trailing edge, and first and second air-washed surfaces which extend from the leading edge to the trailing edge to form respective sides of the feed arm, wherein the major axis of the generally elliptical cross-section extends between the leading and trailing edges. In this way, although the feed arm is formed with a generally elliptical cross-section rather than with a more conventional circular cross-section, the projected area of the feed arm presented to the airflow is not necessarily increased.

The length of the major axis of the generally elliptical cross-section may vary for cross-sections at different given radial positions along the length of the feed arm.

The length of the major axis may decrease for generally elliptical cross-sections at given radial positions which are progressively more distant from the mounting flange.

The cross-section through the feed arm at a radial position adjacent the nozzle head may be generally circular in shape.

The length of the minor axis of the generally elliptical cross-section may vary for cross-sections at different given radial distances along the length of the feed arm.

The length of the minor axis may decrease for generally elliptical cross-sections at given radial positions which are progressively more distant from the mounting flange.

The generally elliptical cross-section may be elliptical, oval, racetrack, teardrop or aerofoil in shape.

The fuel nozzle may be manufactured by additive layer manufacturing.

The gas turbine engine may further comprise an engine core comprising a turbine, a compressor, a core shaft connecting the turbine to the compressor, a fan located upstream of the engine core, the fan comprising a plurality of fan blades, and a gearbox that receives an input from the core shaft and outputs drive to the fan so as to drive the fan at a lower rotational speed than the core shaft.

As noted elsewhere 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.

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 (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 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.

Brief description of the drawings

Embodiments will now be described by way of example only, with reference to the Figures, in which: Figure 1 is a sectional side view of a gas turbine engine; Figure 2 is a close up sectional side view of an upstream portion of a gas turbine engine; Figure 3 is a partially cut-away view of a gearbox for a gas turbine engine; Figure 4 is an isometric view of a partial fuel spray nozzle; Figure 5 is a trailing edge view of the fuel spray nozzle of Figure 4; and Figure 6 is a side view of the fuel spray nozzle of Figure 4.

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 1 illustrates a gas turbine engine 10 having a principal rotational axis 9. The engine 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 2.

The low pressure turbine 19 (see Figure 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.

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 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 Figure 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.

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

It will be appreciated that the arrangement shown in Figures 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 Figure 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 Figure 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 Figure 2.

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 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 On 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 On 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.

The combustion equipment 16 includes at least one fuel spray nozzle 100, an exemplary arrangement of which is partially shown in isometric view in Figure 4, trailing edge view in Figure 5 and side view in Figure 6. Some detail has been omitted from Figures 4 to 6 for clarity. The combustion system 16 typically has a circumferential row of such nozzles 100.

The fuel spray nozzle 100 is provided with a mounting flange 110 for mounting the fuel spray nozzle 100 onto the gas turbine engine 10.A feed arm 135 projects from the mounting flange 110 and extends to a fuel spray nozzle head 150 which, in use, is located in the engine 10 radially inwardly of the mounting flange 110. A transition zone 160 may be provided between the feed arm 135 and the fuel spray nozzle head 150. The transition zone 160 may be shaped and formed in any suitable manner to have the required strength and other desirable characteristics for that area where the feed arm 135 transitions into the fuel spray nozzle head 150.

The feed arm 135 is configured to be positioned in an airflow and has a central axis 165, a leading edge 170, a trailing edge 180, and a first airwashed surface 190 and a second airwashed surface 200 which extend from the leading edge 170 to the trailing edge 180 to form respective sides of the feed arm 135. A cross-section A through the feed arm 135 perpendicular to the central axis 165 at a given radial position between the mounting flange 110 and the nozzle head 150 is generally elliptical in shape. This may provide several advantages such as a greater frequency separation between the feed arm's first two dynamic modes, thereby improving high cycle fatigue performance since the generally elliptical shape will lead to a more defined cantilever mode rather than the orbiting momentum typically associated with conventional circular cylindrical feed arms. The generally elliptical cross section (A) may also help reduce overall mass since the second moment of area of the feed arm 135 will be higher in the axial load direction and hence the wall thickness of the feed arm may be reduced. Bird strike impact loads act in the axial direction 9 of the engine, and feed arms 135 are sized to withstand bird strike impact loads. A generally elliptical shape may improve the capability of the feed arm 135 to withstand bird strike impact in general, whilst still possibly allowing a reduced wall thickness of the feed arm 135.

A first distance 210 extending between the leading and trailing edges 170, 180 may be the major axis of the generally elliptical shape and a second distance 220 extending between the first and second airwashed surfaces 190, 200 may be the minor axis. The length of the major axis at a radial position adjacent the mounting flange 110 may be at least 1.4 times, and preferably at least 1.6 times, greater than that of the minor axis.

The generally elliptical cross-section may be elliptical, oval, racetrack or teardrop in shape.

The feed arm 135 causes a resistance to the airflow through the engine 10, but providing the generally elliptical cross-section A with the major axis extending between the leading and trailing edges 170, 180) may reduce the wake in the airflow downstream of the feed arm 135. In addition, the generally elliptical cross-section may provide a larger effective flow area for fuel flow through the fuel spray nozzle 100. To further enhance the aerodynamic characteristics the generally elliptical cross-section may be aerofoil shaped with the major axis forming the chord and the minor axis forming the thickness.

The cross-sectional shape of the feed arm 135 at different radial positions along the length of the feed arm 135 along the may be generally similar. Generally similar cross-sectional shapes means in this context that geometrically the shapes are generally similar but that the areas of different cross-sections A may differ and/or the relative lengths of major axis 210 to minor axis 220 may differ.

For example, the feed arm 135 may taper in one or more directions from the mounting flange 110 to the fuel spray nozzle head 150, hence the length of at least one of the major and minor axes may vary for different cross-sections A along the length of the feed arm 135. In particular, the length of the major axis may decrease for cross-sections at different given radial positions which are progressively more distant from the mounting flange 110, and/or the length of the minor axis may decrease for cross-sections at different given radial positions which are progressively more distant from the mounting flange 110. As shown in Figures 5 and 6, the feed arm 135 may thus taper from a more highly elliptical cross-sectional shape adjacent the mounting flange 110 to a generally circular cross-sectional shape adjacent the transition zone 160.

Conveniently, the fuel spray nozzle may be manufactured using additive layer manufacturing.

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

CLAIMS1. A fuel spray nozzle (100) for a combustion system of a gas turbine engine (10) having a principal rotational axis (9), the fuel spray nozzle (100) comprising: a mounting flange (110); a nozzle head (150) which, in use, is located in the engine (10) radially inwardly of the mounting flange (110); and a feed arm (135) extending between the mounting flange (110) and the nozzle head (150), the feed arm (135) having a central axis (165) and being configured such that a cross-section (A) through the feed arm (135) perpendicular to the central axis (180) at a given radial position between the mounting flange and the nozzle head is generally elliptical in shape.
2. The fuel spray nozzle (100) according to claim 1, wherein the generally elliptical cross-section (A) through the feed arm (135) at a radial position adjacent the mounting flange (110) has a major axis whose length is at least 1.4 times greater than that of its minor axis.
3. The fuel spray nozzle (100) according to claim 1 or 2, wherein the feed arm (135) is for positioning in an airflow, and has a leading edge (170), a trailing edge (180), and first and second air-washed surfaces (190, 200) which extend from the leading edge to the trailing edge to form respective sides of the feed arm (135), and wherein the major axis of the generally elliptical cross-section (A) extends between the leading and trailing edges (170, 180).
4. The fuel spray nozzle (100) according to any of the preceding claims, wherein the length of the major axis of the generally elliptical cross-section (A) varies for cross-sections at different given radial positions along the length of the feed arm (135).
5. The fuel spray nozzle (100) according to claim 4, wherein the length of the major axis decreases for generally elliptical cross-sections (A) at given radial positions which are progressively more distant from the mounting flange (110).
6. The fuel spray nozzle (100) according to claim 5, wherein the cross-section through the feed arm at a radial position adjacent the nozzle head (150) is generally circular in shape.
7. The fuel spray nozzle (100) according to any of the preceding claims, wherein the length of the minor axis of the generally elliptical cross-section (A) varies for cross-sections at different given radial distances along the length of the feed arm (135).
8. The fuel spray nozzle (100) according to claim 7, wherein the length of the minor axis decreases for generally elliptical cross-sections (A) at given radial positions which are progressively more distant from the mounting flange (110).
9. The fuel spray nozzle (100) according to any of the preceding claims, wherein the generally elliptical cross-section (A) is elliptical, oval, racetrack, teardrop or aerofoil in shape.
10. The fuel spray nozzle (100) according to any of the preceding claims, wherein the fuel spray nozzle is manufactured by additive layer manufacturing.
11. A combustion system of a gas turbine engine (10), the combustion system having a circumferential row of the fuel spray nozzles according to any of the preceding claims.
12. A gas turbine engine (10) for an aircraft, the gas turbine engine comprising a combustion system according to claim 11.
13. The gas turbine engine (10) according to claim 12, further comprising: an engine core (11) comprising a turbine (19), a compressor (14), 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.
14. Use of additive layer manufacture to form a fuel spray nozzle (100) according to any one of claims 1-10.