Fuel nozzle with an improved swirler vane structure

Abstract

To provide a fuel nozzle with an improved swirler vane structure.SOLUTION: A fuel nozzle 60 for a turbomachine includes a centerbody 102 that extends axially with respect to a centerline 200 of the fuel nozzle. A confining tube 104 is positioned radially outward of the centerbody. Multiple swirler vanes 106 are disposed between the centerbody and the confining tube. Each of the multiple swirler vanes includes a radially inner base and a radially outer tip. Each of the swirler vanes further includes an upstream portion that extends generally axially from a leading edge. A downstream portion extends from the upstream portion to a trailing edge. The downstream portion defines a bend length between the upstream portion and the trailing edge. The bend length at the radially outer tip is greater than the bend length at the radially inner base.SELECTED DRAWING: Figure 3

Description

Technical Field

[0001] The present disclosure generally relates to turbomachinery fuel nozzles. In particular, the present disclosure relates to a swirler vane structure for use in a turbomachinery fuel nozzle.

Background Art

[0002] Turbomachines are utilized in various industries and applications for energy transfer purposes. For example, a gas turbine engine generally includes a compressor section, a combustion section, a turbine section, and an exhaust section. The compressor section gradually increases the pressure of the working fluid entering the gas turbine engine and supplies this compressed working fluid to the combustion section. The compressed working fluid and fuel (e.g., natural gas) are mixed within the combustion section and combust within the combustion chamber to produce high-pressure and high-temperature combustion gases. The combustion gases flow from the combustion section to the turbine section, where they expand to generate work. For example, the expansion of the combustion gases in the turbine section can rotate a rotor shaft connected to a generator to generate electricity. The combustion gases are then discharged from the gas turbine via the exhaust section.

[0003] Turbomachines typically include fuel nozzles in the combustor section. Each fuel nozzle is a component having one or more passages for delivering a mixture of fuel and air to the combustion chamber for ignition. The fuel nozzle often includes a swirler to improve the mixing of fuel and air into a consistent and uniform mixture prior to ignition. The swirler portion of the fuel nozzle includes a plurality of aerodynamic vanes that extend radially from the central body of the nozzle and circumferentially around the central body of the nozzle. The swirler vanes often include internal passages that supply fuel through fuel holes defined in the surface of the swirler vanes. When the fuel exits the fuel holes, the fuel mixes with the fluid passing between the swirler vanes, typically air. The fuel / air mixture is then ignited within the combustion chamber to produce combustion gases that power the turbine section.

[0004] Often, to reduce emissions and / or improve turndown capabilities, old turbomachinery models are retrofitted to include a secondary combustion stage, which generally includes one or more axial fuel injectors disposed downstream of a primary combustion stage, such as a fuel nozzle. Typically, the axial fuel injectors require most of the compressed air that was previously sent only through the fuel nozzle to operate. As a result of the reduction in the compressed air flow to the primary fuel nozzle, conventional swirler vanes can cause flow separation at or downstream of the swirler, which can adversely affect the performance of the fuel nozzle, such as flameholding. Generally, the reduction in the compressed air flow is often accompanied by a reduction in the bulk velocity of the air flow across the swirler, increasing the risk of flameholding at the swirler surface. SUMMARY OF THE INVENTION

[0005] Accordingly, there is a need in the art for a fuel nozzle having an improved swirler vane structure that can operate with a reduced air flow while maintaining an appropriate flameholding margin.

[0006] Aspects and advantages of the swirler assembly and turbomachinery according to the present disclosure are described in part in the following description, or will be apparent from the description, or can be learned by practice of the technology.

[0007] According to one embodiment, a fuel nozzle is provided. The fuel nozzle includes a central body that extends axially with respect to a centerline of the fuel nozzle. A confinement tube is disposed radially outward of the central body. A plurality of swirler vanes are disposed between the central body and the confinement tube. Each of the plurality of swirler vanes includes a radially inner base and a radially outer tip. Each of the swirler vanes further includes an upstream portion that extends generally axially from a leading edge. A downstream portion extends from the upstream portion to a trailing edge. The downstream portion defines a bend length between the upstream portion and the trailing edge. The bend length at the radially outer tip is greater than the bend length at the radially inner base.

[0008] According to another embodiment, a turbomachine is provided. The turbomachine includes a compressor section, a turbine section, and a combustion section including a plurality of fuel nozzles. Each fuel nozzle of the plurality of fuel nozzles includes a central body extending axially with respect to the centerline of the fuel nozzle. A confinement tube is disposed radially outside the central body. A plurality of swirler vanes are disposed between the central body and the confinement tube. Each of the plurality of swirler vanes includes a radially inner base and a radially outer tip. Each of the swirler vanes further includes an upstream portion extending substantially axially from a leading edge. A downstream portion extends from the upstream portion to a trailing edge. The downstream portion defines a bend length between the upstream portion and the trailing edge. The bend length at the radially outer tip is greater than the bend length at the radially inner base.

[0009] These and other features, aspects, and advantages of the present fuel nozzle and turbomachine will be better understood with reference to the following description and the appended claims. The accompanying drawings, which are incorporated herein and form a part of this specification, illustrate embodiments of the technology and, together with the description, serve to explain the principles of the technology.

[0010] A complete and enabling disclosure of the present swirler assembly, and turbomachine, including the best mode of making and using the system and method for those skilled in the art, is set forth herein with reference to the accompanying drawings.

Brief Description of the Drawings

[0011]

Figure 1

Figure 2

Figure 3

Figure 4

Figure 5

Figure 6

Figure 7

Figure 8

Figure 9

Figure 10

[0012] Here, embodiments of the present fuel nozzle having an improved swirler vane structure and a turbomachine having such a fuel nozzle will be referred to in detail, one or more examples of which are shown in the drawings. Each example is provided for the purpose of explaining the present technology and is not intended to limit the present technology. Indeed, it will be apparent to those skilled in the art that modifications and changes are possible in the present technology without departing from the scope or spirit of the claimed technology. For example, features illustrated or described as part of one embodiment can be used in another embodiment to bring about further embodiments. Accordingly, the present disclosure is intended to cover such modifications and changes within the scope of the appended claims and their equivalents.

[0013] For the detailed description, numerical and alphabetical reference signs are used to refer to the features of the drawings. Similar or identical reference signs in the drawings and the description are used to refer to similar or identical parts of the present invention. As used herein, the terms "first", "second", and "third" can be used interchangeably to distinguish one component from another, and are not intended to indicate the position or importance of individual components.

[0014] As used herein, the terms "upstream" (or "frontward") and "downstream" (or "rearward") refer to the relative directions with respect to the flow of fluid in a fluid path. For example, "upstream" refers to the direction from which the fluid flows, and "downstream" refers to the direction in which the fluid flows.

[0015] The term "radially" refers to a relative direction substantially perpendicular to the axial centerline of a particular component, the term "axially" refers to a relative direction substantially parallel and / or coaxial with the axial centerline of a particular component, and the term "circumferentially" refers to a relative direction extending around the axial centerline of a particular component.

[0016] Approximating terms such as "substantially" or "about" include values within a range of plus or minus 10 percent of the recited value. When used in the context of an angle or direction, such terms include a range of plus or minus 10 degrees of the recited angle or direction. For example, "substantially perpendicular" includes any direction within 10 degrees of perpendicular, either clockwise or counterclockwise.

[0017] Referring now to the drawings, FIG. 1 shows a schematic view of one embodiment of a turbomachine, which in the illustrated embodiment is a gas turbine 10. Although industrial or land-based gas turbines are shown and described herein, the present disclosure is not limited to land-based and / or industrial gas turbines unless specifically recited in the claims. For example, the swirler assembly described herein can be used in any type of turbomachine including, but not limited to, steam turbines, aircraft gas turbines, or marine gas turbines.

[0018] As shown, the gas turbine 10 generally includes an inlet section 12, a compressor section 14 disposed downstream of the inlet section 12, a plurality of combustors 17 (shown in FIG. 2) within a combustor section 16 disposed downstream of the compressor section 14, a turbine section 18 disposed downstream of the combustor section 16, and an exhaust section 20 disposed downstream of the turbine section 18. Additionally, the gas turbine 10 can include one or more shafts 22 coupled between the compressor section 14 and the turbine section 18.

[0019] The compressor section 14 can generally include a plurality of rotor disks 24 (one of which is shown) and a plurality of rotor blades 26 extending radially outward from each rotor disk 24 and connected to each rotor disk 24. Next, each rotor disk 24 may be coupled to or form a part of a portion of the shaft 22 extending through the compressor section 14.

[0020] The turbine section 18 can generally include a plurality of rotor disks 28 (one of which is shown) and a plurality of rotor blades 30 that extend radially outward from each rotor disk 28 and are connected to each rotor disk 28. Next, each rotor disk 28 may be coupled to or form a part of a portion of the shaft 22 that extends through the turbine section 18. The turbine section 18 further includes an outer casing 31 that circumferentially surrounds a portion of the shaft 22 and the rotor blades 30, thereby at least partially defining a hot gas path 32 through the turbine section 18.

[0021] During operation, a working fluid such as air flows through the inlet section 12 into the compressor section 14, where the air is gradually compressed, thereby providing pressurized air 27 to the combustors of the combustor section 16. The pressurized air 27 is mixed with fuel and burned in each combustor to generate combustion gases 34. The combustion gases 34 flow from the combustor section 16 into the turbine section 18 through the hot gas path 32, where energy (kinetic energy and / or thermal energy) is transferred from the combustion gases 34 to the rotor blades 30, causing the shaft 22 to rotate. The mechanical rotational energy can then be used to power the compressor section 14 and / or generate electricity. The combustion gases 34 discharged from the turbine section 18 can then be exhausted from the gas turbine 10 via the exhaust section 20.

[0022] As shown in FIG. 2, the combustor 17 may be at least partially surrounded by an outer casing 31, which can be referred to as a compressor discharge casing. The outer casing 31 can at least partially define a high-pressure plenum 35 that at least partially surrounds various components of the combustor 17. The high-pressure plenum 35 is in fluid communication with the compressor 14 (FIG. 1) and can receive compressed air 27 therefrom. An end cover 36 may be coupled to the outer casing 31. In certain embodiments, the outer casing 31 and the end cover 36 can at least partially define a head-end volume or head-end portion 38 of the combustor 17.

[0023] In certain embodiments, the head-end portion 38 is in fluid communication with the high-pressure plenum 35 and / or the compressor 14. One or more liners or ducts 40 can at least partially define a combustion chamber or zone 42 for combusting the fuel-air mixture and / or at least partially define a hot gas path through the combustor, as indicated by arrow 43, for directing the combustion gas 34 towards the inlet of the turbine 18.

[0024] In various embodiments, the combustor 17 includes at least one fuel nozzle 60 in the head-end portion 38. As shown in FIG. 2, the fuel nozzle 60 can be disposed within the outer casing 31 downstream of and / or spaced from the end cover 36 of the combustor 17 and upstream of the combustion chamber 42. In certain embodiments, the fuel nozzle assembly 60 may be in fluid communication with a fuel source 48 via one or more fluid conduits 50. In many embodiments, the fluid conduit 50 may be fluidly coupled to and / or connected at one end to the end cover 36.

[0025] Figure 3 shows an example of the fuel nozzle 60 described herein. The fuel nozzle 60 can be used with a combustor 17 or the like. As shown, the fuel nozzle 60 can include a swirler portion 100. The fuel nozzle 60 can include a hub or central body 102 that is radially spaced from the confinement tube 104. As shown, the central body 102 can be connected to the confinement tube by one or more swirler vanes 106. The swirler vanes 106 can have a generally aerodynamic profile and can be configured to impart a swirl to the air passing through the fuel nozzle 60. Each swirler vane 106 can include one or more fuel supply passages 58 therethrough. These fuel supply passages 58 can distribute gaseous fuel to gaseous fuel injection holes (not shown). The gaseous fuel can enter the swirler assembly 100 through one or more annular passages 61 that supply the gaseous fuel to the fuel supply passages 58. The gaseous fuel can be mixed with the compressed air 27 as the fuel and air move through the swirler portion 100, and after mixing within the confinement tube 104 of the fuel nozzle 60, the fuel / air mixture can enter the combustion zone 42 (Figure 2) where combustion occurs.

[0026] Figure 4 shows a swirler portion 100 with a portion of the confinement tube 104 of the fuel nozzle 60 cut away, showing the swirler vanes 106, according to an embodiment of the present disclosure. The swirler vanes 106 may be disposed radially between the central body 102 and the confinement tube 104. As shown, the swirler portion 100 can include a plurality of swirler vanes 106 that function to promote fuel / air mixing and improve flame stabilization. In the embodiment shown in Figures 4 and 5, the swirler portion 100 includes ten circumferentially spaced swirler vanes 106. In other embodiments, the number of swirler vanes 106 may vary.

[0027] Compressed air 27 from the compressor section 14 can flow through the annular space 105 between the central body 102 and the confinement tube 104, and the air 27 encounters the swirler vanes 106. The swirler vanes 106 can induce a swirling motion in the air in the clockwise or counterclockwise direction in the circumferential direction C. The swirler portion 100 can also include a plurality of fuel injection ports (not shown) defined through the swirler vanes 106. The fuel injection ports can direct fuel into the annular space 105 of the swirler portion 100 (i.e., between adjacent swirler vanes 106) where the fuel contacts and mixes with the air. The swirler vanes 106 can induce a swirling motion in the fuel / air mixture as the fuel / air mixture moves through the confinement tube 104 to the combustion zone 42.

[0028] As shown in FIG. 4, the swirler portion 100 can define an axial direction A and a circumferential direction C extending around the axial direction A. The swirler portion 100 can also define a radial direction R perpendicular to the axial direction A.

[0029] As shown in FIG. 4, the swirler portion 100 can further include a maximum radial distance or R max value. As shown, the R max value can be measured in the radial direction R from the axial centerline 200 of the swirler portion 100 to the confinement tube 104. Specifically, the R max value can be measured from the axial centerline 200 to the inner surface 107 of the confinement tube 104. Further, as used herein, the R / R max value can be used to indicate a radial position and can be a percentage and / or fraction of the R max value. For example, as shown in FIG. 4, when R / R max is equal to 0.5 (or 50% of the R max value), the position along the radial direction R is the outer surface 103 of the central body 102 and / or the radially inner base 114 (shown in FIG. 6) of the swirler vanes 106.

[0030] FIG. 5 shows a swirler vane 106 separated from the central body 102 and the confinement tube 104, according to an embodiment of the present disclosure. As shown in FIGS. 4 and 5, each of the swirler vanes 106 can include a radius 108 that extends between the central body 102 and the confinement tube 104. Each of the swirler vanes 106 includes a leading edge 122 defined at the upstream end 110 and a trailing edge 124 defined at the downstream end 112. Air and / or fuel generally flow from the upstream end 110 to the downstream end 112.

[0031] In many embodiments, the swirler vane 106 includes a radially inner base 114 coupled to the central body 102. The swirler vane 106 can extend radially between the radially inner base 114 and the radially outer tip 116. Each of the swirler vanes 106 can include a positive pressure side 118 and a negative pressure side 120. The positive pressure side 118 extends from the leading edge 122 to the trailing edge 124 and can form a positive pressure side surface 126. The positive pressure side surface 126 may have a generally aerodynamic profile and, in many embodiments, may be substantially arcuate. Air and / or fuel can generally flow relative to the positive pressure side 118 and take a path corresponding to the positive pressure side surface 126. Similarly, the negative pressure side 120 also extends from the leading edge 122 to the trailing edge 124 and forms a negative pressure side surface 128. The positive pressure side surface 126 may be different from the negative pressure side surface 128, i.e., may have a different aerodynamic profile. Thus, the surfaces 126, 128 can vary along the radius 108 of the swirler vane 106 to form a varied air swirl angle downstream of the swirler vane 106 and / or downstream of the swirler portion 100.

[0032] As shown in FIGS. 4 and 5, the positive pressure side 118 and the negative pressure side 120 can converge toward each other at the upstream end 110 to at least partially form the leading edge 122. Similarly, the positive pressure side 118 and the negative pressure side 120 also converge toward each other at the downstream end 112 to at least partially form the trailing edge 124. The surface shapes of the positive pressure side 118 and the negative pressure side 120 vary along the swirler vane 106 to ensure a smooth transition from the leading edge 122 to the trailing edge 124 at any radial position.

[0033] According to an embodiment of the present disclosure, FIG. 6 shows a side view of a single swirler vane 106, FIG. 7 shows a side contour view of the radially inner base 114 of the swirler vane 106, and FIG. 8 shows a side contour view of the radially outer tip 116 of the swirler vane 106. As shown in FIGS. 6-8, the swirler vane 106 can include a camber line 131. The camber line 131 may be defined midway between the positive pressure side surface 126 and the negative pressure side surface 128.

[0034] As shown, the positive pressure side surface 126, the negative pressure side surface 128, and the camber line 131 can each further include an upstream portion 130 and a downstream portion 132. In many embodiments, the upstream portion 130 of the surfaces 126, 128 may extend from the leading edge 122 to the downstream portion 132. Similarly, the downstream portion 132 may extend from the upstream portion 130 to the trailing edge 124. As shown, the upstream portions 130 of the surfaces 126, 128 and the camber line 131 may be substantially flat and substantially axially aligned. The downstream portion 132 can include an aerodynamic contour and / or curvature in the circumferential direction C that functions to induce swirl in the air and / or fuel moving within the swirler portion 100.

[0035] As shown, the upstream portion 130 can extend axially from the leading edge 122 and terminate where the surfaces 126, 128 begin to have curvature and / or contour, i.e., where the downstream portion 132 begins. The curvature of the surfaces 126, 128 can begin at different positions along the swirler vane 106 depending on the radial position. Thus, the lengths of the upstream portion 130 and the downstream portion 132 of the positive pressure side surfaces 126, the negative pressure side surfaces 128, and the camber line 131 may vary along the radius 108 of the swirler vane 106.

[0036] As best shown in FIG. 5, in some embodiments, the downstream portion 132 of the swirler vane 106' in the plurality of swirler vanes 106 can extend circumferentially beyond the leading edge 122 of the adjacent swirler vane 106'' in the plurality of swirler vanes 106. For example, the trailing edge 124 of the swirler vane 106' and at least a portion of the downstream portion 132 may axially overlap the leading edge 122 and the upstream portion 130 of the adjacent swirler vane 106'' in the plurality of swirler vanes 106. Further, in many embodiments, the trailing edge 124 may be circumferentially offset from the leading edge 122.

[0037] As collectively shown in FIGS. 6 - 8, the swirler vane 106 can further include a bend length 134 or L (FIGS. 9 and 10). The bend length 134 may be the length of the downstream portion 132, i.e., the length of the swirler vane 106 that is substantially arcuate, curved, and / or aerodynamically contoured. Specifically, the bend length 134 may be the length of the downstream portion 132 of the positive pressure side surface 126, the length of the downstream portion 132 of the negative pressure side surface 128, or the length of the downstream portion 132 of the camber line 131. As used herein, "bend length 134" generally refers to the bend length 134 of the camber line 131 unless otherwise specified. In various embodiments, the bend length 134 of the positive pressure side surface 126, the bend length 134 of the negative pressure side surface 128, and the bend length 134 of the camber line 131 may be the same or different.

[0038] As shown in FIG. 6 and as described above, the bend lengths 134 of each of the positive pressure side surface 126, the negative pressure side surface 128, and the camber line 131 may vary along the radius 108 of the swirler vane 106. For example, in many embodiments, the bend length 134 of each of the positive pressure side surface 126, the negative pressure side surface 128, and the camber line 131 may be substantially longer at the radially outer tip 116 (FIG. 8) compared to the bend length 134 at the radially inner base 114 (FIG. 7). However, in other embodiments, the bend length 134 of each of the positive pressure side surface 126, the negative pressure side surface 128, and the camber line 131 may be the same at the radially outer tip 116 and the radially inner base 114.

[0039] In many embodiments, the bend length 134 of the camber line 131 at the radially inner base 114 may be from about 40% to about 90% of the bend length 134 of the camber line 131 at the radially outer tip 116. In other embodiments, the bend length 134 of the camber line 131 at the radially inner base 114 may be from about 45% to about 85% of the bend length 134 of the camber line 131 at the radially outer tip 116. In some embodiments, the bend length 134 of the camber line 131 at the radially inner base 114 may be from about 50% to about 80% of the bend length 134 of the camber line 131 at the radially outer tip 116. In various embodiments, the bend length 134 of the camber line 131 at the radially inner base 114 may be from about 55% to about 75% of the bend length 134 of the camber line 131 at the radially outer tip 116.

[0040] In many embodiments, the bend length 134 may increase substantially linearly from the radially inner base 114 to the radially outer tip 116. Thus, the bend length 134 may increase at a constant rate of change from the radially inner base 114 to the radially outer tip 116.

[0041] As shown in FIGS. 7 and 8, the swirler vane 106 can further include an exit flow angle 136. The exit flow angle 136 can be defined between the axial centerline 200 of the fuel nozzle 60 and a line 202 that contacts the camber line 131 at the trailing edge 124. Air and / or fuel may be deflected from a generally axial flow path defined by the upstream portion 130 of the surfaces 126, 128 toward the exit flow angle 136 by the downstream portion 132 of the surfaces 126, 128. Further, the exit flow angle 136 may be constant along the radius 108, i.e., the exit flow angle 136 does not vary in the radial direction R. Thus, the distance required to shift air and / or fuel from a generally axial flow path to a flow direction along line 202, i.e., offset the axial direction by the amount of the exit flow angle 136, can vary depending on the radial position of the air / fuel on the swirler vane 106. For example, the closer the fuel and / or air is to the central body 102, i.e., the more radially inward, the shorter the bend length 134 utilized to shift the air / fuel toward the exit flow angle 136.

[0042] In many embodiments, the exit flow angle may be from about 30° to about 60°. In other embodiments, the exit flow angle may be from about 35° to about 55°. In some embodiments, the exit flow angle 136 may be from about 40° to about 50°. In a particular embodiment, the exit flow angle 136 may be about 45°.

[0043] FIG. 9 is a perspective view of a swirler vane 106 according to an embodiment of the present disclosure. In FIG. 9, the R / R max value, i.e., the radial position along the swirler vane 106, is shown in increments of 0.1, and the radially outer tip 116 is transparent in FIG. 9 for purposes of showing the perspective view. As shown, the bend length L (134 in FIGS. 6 - 8) increases generally linearly along the radial direction R.

[0044] FIG. 10 is a graph 300 plotting the relationship between the radial position of the swirler vane 106 and the bending length L. As shown in FIG. 10, the bending length L refers to the length of the downstream portion 132 of the camber line 131. Specifically, FIG. 10 shows the bending length L normalized with respect to the max value, i.e., the maximum radial distance R max and the graph of a line 302 showing the relationship between the normalized bending length L and the R / R max value. As shown by the plot, the bending length L increases linearly as the radius R increases. Similarly, as shown by the plot, the bending length L can increase at a constant (positive) rate of change due to the substantially uniform gradient of the line 302.

[0045] As summarized in FIGS. 9 and 10, the ratio between the bending length L and the maximum radial distance R max is approximately equal to 0.65 at the radially inner base 114 represented by the point 304 in FIG. 10. Similarly, the ratio between the bending length L and the maximum radial distance R max is approximately equal to 1.45 at the radially outer tip 116 represented by the point 306 in FIG. 10. In other embodiments, the L / R max value can be as low as about 0.4 at the radially inner base 114. The L / R max value at the radially inner base 114 should not be less than 0.4. Otherwise, there may be flow separation on the swirler vane 106.

[0046] During operation, linearly increasing the bend length L of the swirler vane 106 functions to increase the overall flameholding margin, thereby enabling the use of larger quantities of more reactive fuels (fuels rich in hydrogen and dicarbon). The improved structure of the swirler vane 106 described herein can advantageously enable the use of an axial fuel staging system (or secondary combustion system) without adversely affecting the flameholding margin of the fuel nozzle (or primary combustion system). Specifically, the structure of the swirler vane 106 prevents flow separation in the primary fuel nozzle 60 that might otherwise occur when a majority of the total airflow volume to the head end portion 38 of the combustor 17 is diverted to a downstream axial fuel staging injector (not shown) for secondary combustion.

[0047] This specification uses examples to disclose the invention, including the best mode, and to enable any person skilled in the art to practice the invention, including making and using any device or system and performing any incorporated method. The patentable scope of the invention is defined by the claims and may include other examples that occur to those skilled in the art. Such other examples are intended to be within the scope of the claims if they include structural elements that do not differ from the literal language of the claims, or if they include equivalent structural elements with insubstantial differences from the literal language of the claims.

Explanation of Reference Numerals

[0048] 10 Gas turbine 12 Inlet section 14 Compressor section 16 Combustor section 17 Combustor 18 Turbine section 20 Exhaust section 22 Shaft 24 Rotor disk 26 Rotor blade 27 Compressed air, pressurized air 28 Rotor disk 30 Rotor Blade 31 Outer Casing 32 High-Temperature Gas Path 34 Combustion Gas 35 High-Pressure Plenum 36 End Cover 38 Head End Portion 40 Liner, Duct 42 Combustion Zone, Combustion Chamber 43 Arrow 48 Fuel Supply Source 50 Fluid Conduit 58 Fuel Supply Passage 60 Fuel Nozzle, Fuel Nozzle Assembly 61 Annular Passage 100 Swirler Portion, Swirler Assembly 102 Central Body 103 Outer Surface 104 Confining Tube 105 Annular Space 106 Swirler Vane 106’ Swirler Vane 106’’ Swirler Vane 107 Inner Surface 108 Radius 110 Upstream End 112 Downstream End 114 Radial Inner Base 116 Radial Outer Tip 118 Positive Pressure Side 120 Negative Pressure Side 122 Leading Edge 124 Trailing Edge 126 Positive Pressure Side Surface 128 Negative Pressure Side Surface 130 Upstream Portion 131 Camber Line 132 Downstream Portion 134 Buckling Length 136 Exit Flow Angle 200 Axial Center Line 202 Line 300 Graph 302 Line 304 Point 306 Point

Claims

1. A fuel nozzle (60), comprising: a central body (102) extending axially with respect to a center line (200) of a swirler portion (100) of the fuel nozzle (60); a confinement tube (104) radially outside the central body (102); a plurality of swirler vanes (106) disposed between the central body (102) and the confinement tube (104); wherein each of the plurality of swirler vanes (106) comprises: a radially inner base (114) and a radially outer tip (116); an upstream portion (130) extending from a leading edge (122); a downstream portion (132) extending from the upstream portion (130) to a trailing edge (124), the downstream portion (132) defining a bend length (134) between the upstream portion (130) and the trailing edge (124), and the bend length (134) at the radially outer tip (116) is greater than the bend length (134) at the radially inner base (114); and each swirler vane (106) of the plurality of swirler vanes (106) includes a positive pressure side (118) and a negative pressure side (120); the trailing edge (124) of each swirler vane (106) of the plurality of swirler vanes (106) further includes an exit flow angle (136) defined between the center line (200) of the fuel nozzle (60) and a camber line (131); the exit flow angle (136) is constant along a radius (108) of each swirler vane (106) of the plurality of swirler vanes (106) and does not vary in the direction of the radius (108), the fuel nozzle (60).

2. The fuel nozzle (60) according to claim 1, wherein the upstream portion (130) of each swirler vane (106) is substantially flat and axially oriented with respect to the center line (200) of the swirler portion (100).

3. The fuel nozzle (60) according to claim 1, wherein a maximum radial distance is defined between the center line (200) of the swirler portion (100) and the confinement tube (104), and a ratio between the bend length (134) at the radially inner base (114) and the maximum radial distance is greater than 0.

4.

4. The fuel nozzle (60) according to claim 1, wherein the bend length (134) at the radially inner base (114) is about 40% to about 90% of the bend length (134) at the radially outer tip (116).

5. The fuel nozzle (60) according to claim 1, wherein the bending length (134) of each swirler vane (106) of the plurality of swirler vanes (106) increases substantially linearly from the radially inner base (114) to the radially outer tip (116).

6. The fuel nozzle (60) according to claim 1, wherein the downstream portion (132) of the swirler vane (106) in the plurality of swirler vanes (106) extends circumferentially beyond the leading edge (122) of the adjacent swirler vane (106) in the plurality of swirler vanes (106).

7. The fuel nozzle (60) according to claim 1, wherein the outlet flow angle (136) is from about 30° to about 60°.

8. A turbomachine (10), comprising: a compressor section (14); a turbine section (18); a combustion section including a plurality of fuel nozzles (60) according to any one of claims 1 to 7; The turbomachine (10) comprising the same.

9. The turbomachine (10) according to claim 8, wherein the downstream portion (132) of each swirler vane (106) is substantially arcuate.

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