Combustor with baffle

Abstract

Turbine engines and associated methods for a combustor as shown and described. The turbine engine includes a combustor having a combustor liner with a dilution opening and a geometry that varies in an axial direction. The combustor further has a damper that surrounds the combustor liner defining a combustion chamber of the combustor. Methods for controlling nitrogen oxides within a combustor include injecting compressed air into an annular combustion chamber through any of the dilution openings described herein.

Description

[0001] Cross-references to related applications

[0002] This application claims priority to U.S. Provisional Application Serial No. 63 / 298,749, filed January 12, 2022, and U.S. Patent Application Serial No. 17 / 694,978, filed March 15, 2022, the contents of which are incorporated herein by reference. Technical Field

[0003] This topic generally relates to combustors for gas turbine engines, and more specifically, to combustors for burning fuels that are lighter than air, either alone or in combination. Background Technology

[0004] A turbine engine is driven by a flow of combustion gases through the engine, which rotates multiple turbine blades. These turbine blades, in turn, rotate a compressor to supply compressed air to a combustor for combustion. The combustor can be located within the turbine engine and fluidly connected to the turbine through which the combustion gases flow.

[0005] The use of hydrocarbon fuels in the combustors of turbine engines is known. Generally, air and fuel are supplied to the combustion chamber, mixed, and then the fuel burns in the presence of air to produce hot gases. These hot gases are then supplied to the turbine, where they are cooled and expand to generate power. Byproducts of fuel combustion typically include environmentally harmful byproducts such as nitrogen oxides and nitrogen dioxide (collectively known as NO). x Carbon monoxide (CO), unburned hydrocarbons (UHC) (e.g., methane and volatile organic compounds that contribute to the formation of atmospheric ozone), and other oxides including sulfur oxides (e.g., SO2 and SO3).

[0006] Various fuels are being explored for use in gas turbine engines. Hydrogen, or hydrogen mixed with another element or compound, can be used for combustion, but hydrogen or hydrogen-blended fuels can result in higher flame temperatures than conventional fuels. That is, hydrogen or hydrogen-blended fuels generally have a wider combustible range and faster combustion rates than conventional fuels such as petroleum-based fuels or mixtures of petroleum and synthetic fuels.

[0007] Standards governing air pollution problems originating worldwide regulate NO generated by turbine engine operation. x Emissions of UHC and CO. In particular, due to the high burner flame temperature during operation, NO is formed within the burner. x The desired effect is to reduce NO by adjusting the profile and / or pattern within the burner. x Emissions can be reduced while maintaining the desired efficiency. Attached Figure Description

[0008] In the attached diagram:

[0009] Figure 1 is a schematic view of a turbine engine.

[0010] Figure 2 depicts a portion of a combustion chamber along a combustion section of a turbine engine. Figure 1

[0011] Figure 3 is a cross-sectional view of line III-III of Figure 2

[0012] Figure 4 is a schematic view of a portion of a combustor taken along line A-A of Figure 3

[0013] Figure 5 is a schematic view of a portion of a combustor taken along line B-B of Figure 3

[0014] Figure 6 is a schematic view of a portion of a combustor taken along line E-E of Figure 3

[0015] Figure 7 is a variant of a portion of a combustor taken along line A-A of Figure 3

[0016] Figure 8 is a portion of a combustor taken along line B-B of Figure 3

[0017] Figure 9 is a portion of a combustor taken along line E-E of Figure 3

[0018] Figure 10 is a close-up view of a portion of a combustor of Figure 3

[0019] Figure 11 is a schematic cross-sectional view of a portion of a combustor of Figure 3

[0020] Figure 12 is a schematic cross-sectional view of a portion of a combustor of Figure 3

[0021] Figure 13 is a schematic cross-sectional view of a portion of a combustor of​​​​​​​​​​​Figure 3 FIG. 6 is a schematic cross-sectional view of a portion of a combustor having a dam defining a shadowed zone of the combustor. DETAILED DESCRIPTION

[0022] The disclosed aspects described herein are directed to combustors, and more particularly to combustors having a dam defining a shadowed zone proximate a fuel injector of the combustor. Further, the combustors as described herein can include a combustion chamber having a primary zone with a transition cross-section from a can to an annular profile. For illustrative purposes, the disclosure will be described with respect to a turbine engine. However, it will be understood that the disclosed aspects described herein are not so limited and the combustors described herein can be implemented in engines including, but not limited to, turbojet engines, turboprop engines, turboshaft engines, and turbofan engines. The disclosed aspects discussed herein can have universal applicability within non-aircraft engines having combustors, such as within other mobile applications and non-mobile industrial, commercial, and residential applications.

[0023] The word "exemplary" is used herein to mean "serving as an example, instance, or illustration." Any implementation described herein as "exemplary" is not necessarily to be construed as preferred or advantageous over other implementations. Furthermore, unless otherwise implicit or the context will dictate otherwise, use of the term "comprises" or variations thereof is not intended to imply that a disclosed aspect comprises only the recited elements, but rather that the disclosed aspect comprises at least the recited elements, and also comprises additional elements that are not expressly listed or enumerated.

[0024] As used herein, the terms "first," "second," and "third" can be used interchangeably to distinguish one component from another and are not intended to signify location or importance of individual components.

[0025] The terms "forward" and "aft" refer to relative locations within a gas turbine engine or vehicle and refer to the normal operating attitude of the gas turbine engine or vehicle. For example, for a gas turbine engine, forward refers to a location closer to the engine inlet, and aft refers to a location closer to the engine nozzle or exhaust.

[0026] As used herein, the term "upstream" refers to a direction opposite to the direction of fluid flow, while the term "downstream" refers to a direction the same as the direction of fluid flow. The terms "forward" or "front" mean in front of something, and "aft" or "back" mean behind something. For example, when used in relation to fluid flow, forward / aft can mean upstream / downstream.

[0027] The term "fluid" can be a gas or a liquid. The term "fluid communication" means that a fluid is able to establish a connection between designated areas.

[0028] Further, as used herein, the terms "radial" or "radially" refer to a direction away from a common center. For example, in the overall context of a turbine engine, radial refers to a direction along a ray extending between a central longitudinal axis of the engine and an outer periphery of the engine.

[0029] All directional references (e.g., radial, axial, proximal, distal, upper, lower, upward, downward, left, right, lateral, front, back, top, bottom, above, below, vertical, horizontal, clockwise, counterclockwise, upstream, downstream, forward, aft, etc.) are used only for identification purposes to aid the reader's understanding of the present disclosure, and do not create limitations, particularly as to the position, orientation, or use of the disclosure described herein. Connection references (e.g., attached, coupled, connected, and joined) are to be construed broadly and will be given their ordinary and accustomed meaning to an artisan of ordinary skill in the art to which the disclosure pertains, and will include intermediate member(s) for connection and will not exclude the intermediate member(s) unless explicitly indicated otherwise, such as, for example, when an explicit reference to an intermediate member(s) is provided. As such, connection references will typically include an intermediate member(s) unless specifically stated otherwise. The exemplary drawings are for purposes of illustrative only and the dimensions, positions, order and relative sizes reflected in the attached drawings attached hereto can vary.

[0030] The singular forms "a," "an," and "the" include plural references unless the context clearly dictates otherwise. Further, as used herein, the term "set" or "a set" of elements can be any number of elements, including only one.

[0031] Approximating language as used herein throughout the description and claims for example, "about," "approximately," "substantially," and "essentially" are terms concerning the imprecision of measurement often due to real world circumstances or unintended experimental error, and do not mean to be a precise or perfect measurement or representation in any way. A value modified by these terms is not limited to the precise value specified. In at least some instances, the approximating language can correspond to the precision of an instrument for measuring the value, or the precision of the methods or machines for constructing or manufacturing the components and / or systems. In at least some instances, the approximating language can correspond to the precision of an instrument for measuring the value, or the precision of the methods or machines for constructing or manufacturing the components and / or systems. For example, the approximating language can refer to within 1%, 2%, 4%, 5%, 10%, 15%, or 20% of a single value, a range of values, and / or an end point of a range of values. Ranges are combined and interchanged, such ranges are identified and include all subranges therein, unless context or language indicates otherwise. For example, all ranges disclosed herein are inclusive of the endpoints, and endpoints are combinable with one another to form new ranges, independently of the other ranges disclosed herein.

[0032] “Proximate” as used herein is not limiting, but is a descriptor for positioning the parts described herein. Further, the term “proximate” means closer or more adjacent to the recited part than the following part. For example, a first hole proximate a wall, with the first hole upstream of a second hole, means that the first hole is closer to the wall than the first hole is to the second hole.

[0033] Figure 1 is a schematic illustration of a turbine engine 10. As a non-limiting example, the turbine engine 10 can be used within an aircraft. The turbine engine 10 can include at least a compressor section 12, a combustion section 14, and a turbine section 16. A drive shaft 18 rotationally couples the compressor section 12 and the turbine section 16 such that rotation of one affects rotation of the other and defines an axis or centerline 20 of rotation of the turbine engine 10.

[0034] The compressor section 12 can include a low pressure (LP) compressor 22 and a high pressure (HP) compressor 24 fluidly coupled in series with one another. The turbine section 16 can include a LP turbine 26 and a HP turbine 28 fluidly coupled in series with one another. The drive shaft 18 can operably couple the LP compressor 22, the HP compressor 24, the LP turbine 26, and the HP turbine 28 together. Alternatively, the drive shaft 18 can include a LP drive shaft (not shown) and a HP drive shaft (not shown). The LP drive shaft can couple the LP compressor 22 to the LP turbine 26, and the HP drive shaft can couple the HP compressor 24 to the HP turbine 28. A LP train can be defined as the combination of the LP compressor 22, the LP turbine 26, and the LP drive shaft such that rotation of the LP turbine 26 can impart a driving force to the LP drive shaft, which in turn can cause the LP compressor 22 to rotate. A HP train can be defined as the combination of the HP compressor 24, the HP turbine 28, and the HP drive shaft such that rotation of the HP turbine 28 can impart a driving force to the HP drive shaft, which in turn can cause the HP compressor 24 to rotate.

[0035] The compressor section 12 can include a plurality of axially spaced stages. Each stage includes a set of circumferentially spaced rotating blades and a set of circumferentially spaced stationary vanes. Compressor blades for a stage of the compressor section 12 can be mounted to a disk, which is mounted to the drive shaft 18. Each set of blades for a given stage can have its own disk. The vanes of the compressor section 12 can be mounted to a casing, which can extend circumferentially around the turbine engine 10. It should be understood that the representation of the compressor section 12 is merely schematic and there can be any number of stages. Further, it is contemplated that there can be any other number of components within the compressor section 12.

[0036] Similar to the compressor section 12, the turbine section 16 can include a plurality of axially spaced stages, with each stage having a set of circumferentially spaced rotating blades and a set of circumferentially spaced stationary vanes. The turbine blades for a stage of the turbine section 16 can be mounted to a disk, which is mounted to the drive shaft 18. Each set of blades for a given stage can have its own disk. The vanes of the turbine section can be mounted to the casing in a circumferential manner. It is noted that there can be any number of blades, vanes, and turbine stages, as the illustrated turbine section is merely a schematic representation. Further, it is contemplated that there can be any other number of components within the turbine section 16.

[0037] The combustion section 14 can be serially disposed between the compressor section 12 and the turbine section 16. The combustion section 14 can be fluidly coupled to at least a portion of the compressor section 12 and the turbine section 16, such that the combustion section 14 at least partially fluidly couples the compressor section 12 to the turbine section 16. As a non-limiting example, the combustion section 14 can be fluidly coupled to the HP compressor 24 at an upstream end of the combustion section 14, and fluidly coupled to the HP turbine 28 at a downstream end of the combustion section 14.

[0038] During operation of the turbine engine 10, ambient or atmospheric air is drawn into the compressor section 12 via a fan (not shown) upstream of the compressor section 12, the air is compressed at the compressor section 12, defining pressurized air. The pressurized air can then flow into the combustion section 14, where the pressurized air is mixed with fuel and ignited, generating combustion gases. The HP turbine 28 extracts some work from these combustion gases, the HP turbine 28 driving the HP compressor 24. The combustion gases are exhausted into the LP turbine 26, which extracts additional work to drive the LP compressor 22, and the exhaust is ultimately exhausted from the turbine engine 10 via an exhaust section (not shown) downstream of the turbine section 16. The driving of the LP turbine 26 drives the LP spool to rotate the fan (not shown) and the LP compressor 22. The pressurized air stream and the combustion gases can together define a working airflow that flows through the fan, the compressor section 12, the combustion section 14, and the turbine section 16 of the turbine engine 10.

[0039] Figure 2 depicted along Figure 1A cross-sectional view of combustion section 14 along line II-II. Combustion section 14 may include a set of fuel injectors 30 arranged annularly around the centerline 20 of the turbine engine 10. Combustor 34 is fluidly connected to the set of fuel injectors 30 to define at least a portion of a set of fuel cups 32. The set of fuel cups 32 may include a rich cup, a lean cup, or a combination of both, arranged annularly around the engine centerline 20. It should be understood that the annularly arranged fuel injectors 30 may be one or more fuel injectors 30, and the one or more fuel injectors 30 may have different characteristics.

[0040] A burner liner 38 defines a burner 34 and includes an outer burner liner 40 and an inner burner liner 42 that are concentric with each other and annularly surrounding the engine centerline 20. The burner liner 38 further defines a set of fuel cups 32. A dome wall 44 and the burner liner 38 together define a combustion chamber 46 of the burner 34 annularly surrounding the engine centerline 20. The set of fuel cups 32 is fluidly connected to the combustion chamber 46.

[0041] The burner 34 may have a canister-shaped, canister-annular, or annular arrangement depending on the type of engine in which it is located. The burner liner 38 may have a varied geometry as further described herein. The burner 34 may be completely enclosed by the housing 36. The compressed air passage 48 may be defined at least partially by both the burner liner 38 and the housing 36.

[0042] Figure 3 Depicting along the combustion zone 14 Figure 2 The cross-sectional view is taken from line III-III. The dome assembly 50 may house the fuel injector 30. The fuel injector 30 may be fluidly coupled to the fuel inlet 52 via a fuel passage 54 adapted to receive the fuel flow (F). Compressed air (C) may be supplied from the compressor section 12 to the combustion section 14 via a compressed air passage 48. The fuel injector 30 may terminate at a dome inlet 56 to define a fuel cup 32. The dome inlet 56, together with the burner liner 38, may further define at least a portion of the primary zone 58 in the combustion chamber 46. A swirl generator 60 may be fluidly coupled to the fuel injector 30. A first set of dilution openings 62 may be provided in the burner liner 38 for connecting the compressed air passage 48 and the combustion chamber 46. At least one igniter 64 may be coupled to the burner liner 38.

[0043] A baffle 70 can be mounted to an outer surface 66 of the combustor liner 38 that surrounds the primary zone 58. The baffle 70 can be spaced apart from the combustor liner 38 to define a shadowed zone 72 having a shadowed zone inlet 74. The baffle 70 can direct a specific amount of compressed air (C) for use along the combustor liner 38. A second set of dilution openings 76 can be provided in the combustor liner 38 for fluidly coupling the shadowed zone 72 to the primary zone 58 of the combustion chamber 46. A third set of dilution openings 78 can be provided in the combustor liner 38 downstream of the second set of dilution openings 76 for connecting different portions of the shadowed zone 72 to the primary zone 58 of the combustion chamber 46. The first, second, and third sets of dilution openings can be circular or shaped holes, slots, or annular gaps and are designed to form radial or angled air jets.

[0044] During operation, compressed air (C) can be fed into the fuel injectors 30 and mixed with fuel (F) to define a fuel / air mixture. The mixture can be ignited within the combustion chamber 46 by the at least one igniter 64 to generate combustion gases (G). The swirler 60 can swirl the incoming compressed air (C) with the fuel (F) entering the fuel cup 32 to provide a uniform mixture of air and fuel that enters the combustion chamber 46 via the dome inlet 56.

[0045] Further, compressed air (C) can be fed into the shadowed zone 72 via the shadowed zone inlet 74. The compressed air (C) can act as a dilution jet provided through both the second set of dilution openings 76 and the third set of dilution openings 78. The primary zone 58 between the second set of dilution openings 76 and the third set of dilution openings 78 can define an extension of the at least one fuel cup 32. The baffle 70 provides the shadowed zone 72 for the fuel cup 32 such that the primary zone 58 operates similarly to a single can combustor arrangement. The baffle 70 allows for stability of the individual fuel cup 32 by minimizing cup-to-cup interaction.

[0046] Compressed air (C) can additionally enter the combustion chamber 46 via the first set of dilution openings 62 to provide a dilution flow (D) within the combustion chamber 46. The combustion gases (G) can be mixed using the dilution flow (D) or simply controlled by the dilution flow (D) to move through the combustor outlet 80 and exit into the turbine section 16.

[0047] Figures 4-6 A varying cross-section of the combustor 34 is shown at varying positions moving axially from a first end 43 at the dome wall 44 to a second end 45 at the combustor outlet 80. As shown, the combustor liner 38 has a transition geometry. The combustor liner 38 defines a cup-shaped geometry 100 proximate the set of fuel injectors 30 and a cylindrical geometry 102 downstream of the cup-shaped geometry 100. Figures 4-6 The cup-shaped geometry 100 is defined by a first end 104 proximate the set of fuel injectors 30 and a second end 106 distal from the set of fuel injectors 30. The cup-shaped geometry 100 can be defined by a first diameter 108 proximate the set of fuel injectors 30 and a second diameter 110 distal from the set of fuel injectors 30. The first diameter 108 can be greater than the second diameter 110. The cup-shaped geometry 100 can be defined by a first height 112 proximate the set of fuel injectors 30 and a second height 114 distal from the set of fuel injectors 30. The first height 112 can be greater than the second height 114. Figure 4) transitions to an annular geometry 102 downstream of the set of fuel injectors 30 Figure 6 ) defining a cup geometry means rounding the combustor liner 38 into a cylindrical shape to define distinguishable cups. More specifically, the geometry of the combustor liner 38 transitions from the modified can shape 104 Figure 4 ) to a can-annular shape 106 Figure 5 ) to an annular shape 108 Figure 6

[0048] Figure 4 is a cross-sectional view along line A-A of Figure 3 looking toward the first end 43 and showing the modified can shape 104. The second set of dilution openings 76 can include a plurality of dilution openings annularly surrounding the fuel cups 32. The combustor liner 38 can be shaped to define different fuel cups, a first fuel cup 32a, a second fuel cup 32b, and a third fuel cup 32c fluidly connected to one another by the spacing of the inner and outer combustor liners 40, 42 defining circumferentially spaced apart openings 82. The modified can shape 104 is defined to have openings 82 small enough to define a plurality of discrete circumferentially spaced apart fuel cups 32a, 32b, 32c while still enabling fluid connection between the different fuel cups 32a, 32b, 32c. The baffles 70 Figure 3 described herein can properly feed compressed air (C) to the second set of dilution openings 76 and the third set of dilution openings 78 in the shadowed zones 72 Figure 3 to produce similar operation as a single can combustor while being geometrically shaped to have the modified can shape 104.

[0049] Figure 5 is a cross-sectional view along line B-B of Figure 3 looking toward the plurality of fuel injectors 30 just downstream of the third set of dilution openings 78. The circumferentially spaced apart openings 82 can gradually widen as the combustor liner 38 extends downstream toward the second end 45 between line A-A and line B-B of Figure 3 It can be appreciated that the combustor liner 38 at line B-B is shaped such that the different fuel-rich cups 32a, 32b, 32c are still distinguishable but more fluidly dynamic relative to one another. In this manner, the geometry of the combustor liner 38 transitions to define the combustor 34 having the modified can shape 104 at line A-A to the can-annular shape 106 at line B-B.

[0050] Figure 6 is a cross-sectional view along line E-E of Figure 3 looking toward the plurality of fuel injectors 30 just downstream of line B-B. The circumferentially spaced apart openings 82​Figure 4 and 5 ) are no longer distinguishable. It will be appreciated that the combustor liner 38 at line E-E is shaped such that the different fuel cups 32a, 32b, 32c ( Figure 4 and 5 ) are no longer distinguishable, and the combustion chamber 46 is annular about the engine centerline 20 ( Figure 2 ). The outer combustor liner 40 and the inner combustor liner 42 at line E-E are circular in shape and annular about the engine centerline 20 ( Figure 2 ). In this manner, the geometry of the combustor liner 38 transitions to define the combustor 34 having a can annular shape 106 at line B-B to an annular shape 108 at E-E.

[0051] Figures 7-9 A variant cross-sectional view of a combustor 34 according to another aspect disclosed herein is shown. Moving axially from a first end 43 at the dome wall 44 toward a second end 45 at the combustor outlet 80, the combustor liner 138 has a transitional geometry. The combustor liner 138 is substantially similar to the combustor liner 38, and thus, similar parts will be identified with like numerals increased by 100. It will be appreciated that the description of similar parts of the combustor liner 38 apply to the combustor liner 138, unless otherwise noted.

[0052] The geometry of the combustor liner 138 transitions from a can shape 110 ( Figure 7 ) to a can annular shape 106 ( Figure 8 ) to an annular shape 108 ( Figure 9 ). The can shape 110 ( Figure 7 ) is defined as having no openings such that the combustor liner 138 defines a plurality of discrete circumferentially spaced apart fuel cups 132a, 132b, 132c, while at the first end 43, there is no fluid connection between the different fuel cups 132a, 132b, 132c. More specifically, at line A-A ( Figure 3 ), the inner combustor liner 140 and the outer combustor liner 142 of the combustor liner 138 meet to physically separate the fuel cups 132a, 132b, 132c. At the same time at line B-B ( Figure 3 ), the inner combustor liner 140 and the outer combustor liner 142 are separated to define a set of spaced apart openings 182, which in turn define the can annular shape 106 ( Figure 8 ). It will be appreciated that the combustor liner 138 at line E-E ( Figure 3 ) is shaped such that the different fuel cups 132a, 132b, 132c ( Figure 7 and 8 ) are no longer distinguishable, which in turn define the combustion chamber 146 of the annular shape 108 ( Figure 9 ).

[0053] Figure 10 is a schematic view of an example combustor 34 showing a variation of the baffle 70 in accordance with an aspect disclosed herein. The baffle 270 is substantially similar to the baffle 70, and thus, similar parts will be identified with like numerals increased by 200. It should be understood that the description of similar parts of the baffle 70 apply to the baffle 170 unless otherwise noted.

[0054] The baffle 170 can be spaced apart from the outer surface 66 of the combustor liner 38 to define a shadowed zone 172. The baffle 170 can include a body 184 extending axially from a mounting leg 186. The baffle 170 can be mounted to the outer surface 66 of the combustor liner 38 at the mounting leg 186. The mounting leg 186 can be mounted to the outer surface 66 between the second set of dilution openings 76 and the third set of dilution openings 78. The body 184 can be spaced apart from the outer surface 66 proximate the dome wall 44 to define a shadowed zone inlet 174 fluidly coupled to the shadowed zone 172.

[0055] The baffle 170 can divide an area surrounding the combustor 34 into a total pressure zone 188 and a static pressure zone 190. During operation, the total pressure zone 188 feeds the second set of dilution openings 76, while the static pressure zone 190 feeds the third set of dilution openings 78. It is contemplated that any number of dilution openings are part of the first set of dilution openings 62, which can also be fed by the static pressure zone 190.

[0056] Figure 11 is a schematic view of an example combustor 34 showing a variation of the baffle 70 in accordance with an aspect disclosed herein. The baffle 270 is substantially similar to the baffle 70, and thus, similar parts will be identified with like numerals increased by 200. It should be understood that the description of similar parts of the baffle 70 apply to the baffle 170 unless otherwise noted.

[0057] The baffle 270 can be spaced apart from the outer surface 66 of the combustor liner 38 to define a shadowed zone 272. The baffle 270 can include a body 284 extending axially from a mounting leg 286. The baffle 270 can be mounted to the outer surface 66 of the combustor liner 38 at the mounting leg 286. The body 284 can be spaced apart from the outer surface 66 proximate the dome wall 44 to define a shadowed zone inlet 274 fluidly coupled to the shadowed zone 272. The mounting leg 286 can be mounted to the outer surface 66 downstream of the first set of dilution openings 62.

[0058] The baffle 270 can divide the area around the combustor 34 into a total pressure zone 288 and a static pressure zone 290. During operation, the total pressure zone 288 feeds the first set of dilution openings 62, the second set of dilution openings 76, and the third set of dilution openings 78 described herein. Feeding the dilution openings described herein with total pressure or static pressure enables adjustment of the amount of pressure drop across the dilution openings. This amount of pressure drop is directly related to the amount of penetration of the compressed air (C) as a dilution jet into the combustion chamber 46. This adjustment has been shown to affect emissions.

[0059] Figure 12 is a schematic view of an exemplary combustor 34 showing a variation of the baffle 70 according to another aspect disclosed herein. The baffle 370 is substantially similar to the baffle 70, and therefore, similar parts will be identified with like numerals increased by 300. It should be understood that the description of similar parts of the baffle 70 apply to the baffle 370 unless otherwise noted.

[0060] The baffle 370 can be spaced apart from the outer surface 66 of the combustor liner 38 to define a shadowed zone 372. The baffle 370 can include a body 384 extending axially between a mounting leg 386 and an inlet leg 392. The baffle 370 can be mounted to the outer surface 66 of the combustor liner 38 at the mounting leg 386. The inlet leg 392 can be spaced apart from the dome wall 44 to define a shadowed zone inlet 374 fluidly coupled to the shadowed zone 372. The baffle 370 can have a substantially trapezoidal shape in cross-section with the shadowed zone inlet 374 positioned along the base of the trapezoidal shape.

[0061] The baffle 370 can divide the area around the combustor 34 into a total pressure zone 388 and a static pressure zone 390. During operation, the total pressure zone 388 feeds the second set of dilution openings 76 and the third set of dilution openings 78, while the static pressure zone 390 feeds the first set of dilution openings 62. As previously described, feeding the dilution openings with total pressure or static pressure enables adjustment of the amount of pressure drop across the dilution openings. Although illustrated with respect to the baffle 370, it should be understood that the inlet leg 392 described herein can be associated with any of the baffles 70, 170, 270 described herein.

[0062] Turning to Figure 13 is an enlarged view of the baffle 370 according to another aspect disclosed herein. At least one acoustic damper 394 can be disposed in the shadowed zone 372 and extend radially from the outer surface 66 to the body 384. The at least one acoustic damper 394 can be a plurality of dampers, a first acoustic damper 394a placed between the second set of dilution openings 76 and the third set of dilution openings 78, and a second acoustic damper 394b placed downstream of the third set of dilution openings 78 and upstream of the mounting leg 386.

[0063] The baffle 370 provides coupling between the acoustic damper 394 and the acoustic pressure source caused by the acoustic dynamics associated with the chamber volume, heat release variation during combustion, and swirler flow vorticity. The acoustic damper 394 can be provided to reduce pressure fluctuations at specific frequencies. The acoustic damper 394 can dampen the associated amplitude at specific frequencies. The acoustic damper 394 can also be referred to as a Helmholtz resonator. The size of the damper is directly related to the acoustic frequency. The acoustic damper 394 is particularly beneficial in combustors with relatively high amplitude associated with specific frequencies. While illustrated with respect to the baffle 370, it should be understood that the acoustic dampers described herein can be located in any of the baffles 70, 170, 270 described herein.

[0064] Any combination of the arrangements described herein with respect to the baffles is contemplated. The rich and lean cup arrangements can be in any of the forms described herein.

[0065] A method for controlling nitrogen oxides or NOx present in combustion gases (G) within a combustor includes injecting compressed air (C) into a combustion chamber through any of the dilution openings described herein. The method can further include swirling the compressed air (C) with a fuel (F) prior to injection into the combustion chamber. The mixing of the compressed air (C) with the fuel (F) can occur in the primary zone. The method can further include transitioning the combustion gas flow (G) through the primary zone 58 to the combustor outlet 80 from the distinguishable cups 32 to a single annular geometry 102. The transition of the combustor liner geometry can also be used to control the flame of the fuel / air mixture in the primary zone 58.

[0066] The benefits associated with the baffle and transition geometries of the combustor liner and the methods described herein are the reduction and / or elimination of CO emissions. Further, the transition geometry of the combustor liner aids in flame control. The baffles described herein and the positioning of the baffles also aid in controlling the flame produced by H2 fuel to achieve lower NO x , lower kinetics, and better component life.

[0067] While described with respect to a turbine engine, it should be understood that the combustors described herein can be used in any engine having a combustor that emits NO x x. It should be understood that the application of the aspects of the disclosure discussed herein also apply to engines having a propeller section or fan and a booster section as well as turbojet and turbine engines.

[0068] The various features and structures of the various embodiments can be used separately or in any combination depending on the needs of the situation at hand. The fact that a feature is not included in every embodiment is not intended to exclude it from the scope of the claims. Rather, it is intended to cover all combinations and permutations of features described herein.

[0069] This written description uses examples to describe the aspects of the disclosure described herein, including the best mode, and also to enable any person skilled in the art to practice the aspects of the disclosure, including making and using any devices or systems and performing any incorporated methods. The patentable scope of the aspects of the disclosure is defined by the claims, and can include other examples that occur to those skilled in the art. Such other examples are intended to fall within the scope of the claims if they have structural elements that do not differ from the literal language of the claims, or if they include equivalent

[0070] Further aspects are provided by the subject matter of the following clauses:

[0071] A gas turbine engine comprising: a compressor section, a combustion section, and a turbine section in a serial flow arrangement, the combustion section comprising: an annular combustion chamber having a first end and a second end, wherein the second end is spaced apart from the first end; a plurality of fuel cups located at the first end and annularly arranged within the annular combustion chamber, wherein at least one fuel cup of the plurality of fuel cups has a defined cup geometry; and a combustor liner at least partially defining the annular combustion chamber and having a varying cross section extending in an axial direction from the first end toward the second end, the varying cross section defining at least a portion of the defined cup geometry at the first end and a single annular geometry at the second end.

[0072] The gas turbine engine of any of the preceding clauses, wherein the defined cup geometry is a modified pot shape defining a plurality of discrete circumferentially spaced fuel cups fluidly connected to each other proximate the first end.

[0073] The gas turbine engine of any of the preceding clauses, wherein the defined cup geometry is a pot shape defining a plurality of discrete circumferentially spaced fuel cups fluidly separated from each other proximate the first end.

[0074] The gas turbine engine of any of the preceding clauses, wherein the combustor liner is shaped to define a pot annular shape downstream of the first end and upstream of the second end.

[0075] The gas turbine engine of any of the preceding clauses, further comprising at least one set of dilution openings in the combustor liner.

[0076] The gas turbine engine of any of the preceding clauses, further comprising a damper mounted to an outer surface of the combustor liner to define a shadow zone.

[0077] The gas turbine engine of any of the preceding clauses, wherein the at least one set of dilution openings fluidly connects the shadow zone to the annular combustion chamber.

[0078] The gas turbine engine of any of the preceding clauses, wherein the at least one set of dilution openings is a plurality of sets of dilution openings axially spaced apart from one another along the combustor liner.

[0079] The gas turbine engine of any of the preceding clauses, wherein the at least one set of dilution openings is at least two sets of dilution openings axially spaced apart from one another along the combustor liner.

[0080] The gas turbine engine of any of the preceding clauses, wherein the damper is mounted to the outer surface between the at least two sets of dilution openings.

[0081] The gas turbine engine of any of the preceding clauses, wherein at least one set of dilution openings fluidly connects the shadow zone to the annular combustion chamber.

[0082] The gas turbine engine of any of the preceding clauses, wherein the damper comprises a body extending axially between a mounting leg and an inlet leg, and the damper is mounted to the outer surface of the combustor liner at the mounting leg and the inlet leg is spaced apart from the first end to define a shadow zone inlet fluidly coupled to the shadow zone.

[0083] The gas turbine engine of any of the preceding clauses, wherein the damper comprises an axially extending body.

[0084] The gas turbine engine of any of the preceding clauses, wherein a mounting leg extends from the body.

[0085] The gas turbine engine of any of the preceding clauses, wherein the mounting leg is mounted to the combustor liner and the body is spaced apart from the combustor liner to define a shadow zone.

[0086] The gas turbine engine of any of the preceding clauses, wherein the dam divides an area surrounding the combustor into a total pressure zone and a static pressure zone, wherein the total pressure zone and the shielded zone are in the same zone.

[0087] The gas turbine engine of any of the preceding clauses, wherein at least one set of dilution openings are fed by the total pressure zone and another set of dilution openings are fed by the static pressure zone.

[0088] The gas turbine engine of any of the preceding clauses, wherein all dilution openings are fed by the total pressure zone.

[0089] The gas turbine engine of any of the preceding clauses, wherein an inlet leg extends from the body to further define the shielded zone and a shielded zone inlet.

[0090] A combustor comprising: an annular combustion chamber having a first end and a second end, wherein the second end is spaced apart from the first end; a plurality of fuel cups located at the first end and annularly arranged within the annular combustion chamber, wherein at least some of the plurality of fuel cups have a defined cup geometry; and a combustor liner at least partially defining the annular combustion chamber and having a varying cross section extending in a downstream direction from the first end toward the second end, the varying cross section defining at least a portion of the defined cup geometry for the fuel cups at the first end and defining a single annular geometry at the second end.

[0091] The combustor of any of the preceding clauses, wherein the defined cup geometry is a modified can shape defining a plurality of discrete circumferentially spaced fuel cups fluidly connected to one another proximate the first end.

[0092] The combustor of any of the preceding clauses, wherein the defined cup geometry is a can shape defining a plurality of discrete circumferentially spaced fuel cups fluidly separated from one another proximate the first end.

[0093] The combustor of any of the preceding clauses, wherein the combustor liner is shaped to define a can annulus shape downstream of the defined cup geometry and upstream of the single annular geometry.

[0094] The combustor of any of the preceding clauses, further comprising at least one set of dilution openings in the combustor liner and a dam mounted to an outer surface of the combustor liner to define a shielded zone, at least one of the at least one set of dilution openings fluidly connecting the shielded zone to the annular combustion chamber.

[0095] A method for controlling nitrogen oxides present within a combustor of a turbine engine, the method comprising: injecting a fuel and compressed air mixture through a fuel injector into a combustion chamber of the combustor; mixing the compressed air and the fuel to define a fuel / air mixture; igniting the fuel / air mixture in a primary zone of the combustion chamber to define a flame and generate combustion gases; flowing the combustion gases from a plurality of circumferentially spaced apart discrete cups through the combustion chamber to a single annular geometry; and exhausting the combustion gases at a combustor exit.

[0096] The method of any of the preceding clauses, further comprising controlling the flame of the fuel / air mixture in the primary zone by feeding compressed air into a shadow zone defined by a baffle mounted to an outer surface of the combustor.

[0097] The method of any of the preceding clauses, further comprising reducing pressure fluctuations at a particular frequency with an acoustic dam located within the shadow zone.

[0098] The method of any of the preceding clauses, further comprising feeding a set of dilution openings from the shadow zone.

[0099] The method of any of the preceding clauses, further comprising feeding a set of dilution openings from a static pressure zone.

[0100] The method of any of the preceding clauses, further comprising feeding a set of dilution openings from a total pressure zone.

Claims

1. A gas turbine engine characterized by, Comprising: a compressor section, a combustion section, and a turbine section in a serial flow arrangement, the combustion section comprising: an annular combustion chamber having a first end and a second end, wherein the second end is spaced apart from the first end; a plurality of fuel cups located at the first end and annularly arranged within the annular combustion chamber, wherein at least one fuel cup of the plurality of fuel cups has a defined cup geometry; and a combustor liner at least partially defining the annular combustion chamber and having a varying cross-section extending in an axial direction from the first end toward the second end, the varying cross-section defining at least a portion of the defined cup geometry at the first end and a single annular geometry at the second end, the combustor liner comprising an inner combustor liner and an outer combustor liner, the inner combustor liner spaced apart from the outer combustor liner at the first end to define a circumferentially spaced apart opening between and fluidly coupling adjacent ones of the plurality of fuel cups; wherein the circumferentially spaced apart opening gradually widens as the combustor liner extends from the first end to the second end.

2. The gas turbine engine of claim 1, wherein, wherein the defined cup geometry is a modified can geometry defining a plurality of discrete circumferentially spaced apart fuel cups fluidly connected to one another proximate the first end.

3. The gas turbine engine of claim 1, wherein, wherein the defined cup geometry is a can geometry defining a plurality of discrete circumferentially spaced apart fuel cups fluidly separated from one another proximate the first end.

4. The gas turbine engine of claim 1, wherein, wherein the combustor liner is shaped to define a can annulus geometry downstream of the first end and upstream of the second end.

5. The gas turbine engine of any one of claims 1-4, wherein, further comprising at least one set of dilution openings in the combustor liner.

6. The gas turbine engine of claim 5, wherein, further comprising a baffle mounted to an outer surface of the combustor liner to define a shadowed zone.

7. The gas turbine engine of claim 6, wherein, wherein the at least one set of dilution openings fluidly connects the shadowed zone to the annular combustion chamber.

8. The gas turbine engine of claim 7, wherein, wherein the at least one set of dilution openings is a plurality of sets of dilution openings axially spaced apart from one another along the combustor liner.

9. The gas turbine engine of claim 6, wherein, wherein the at least one set of dilution openings is at least two sets of dilution openings axially spaced apart from one another along the combustor liner.

10. The gas turbine engine of claim 9, wherein, wherein the baffle is mounted to the outer surface between the at least two sets of dilution openings.

11. The gas turbine engine of claim 10, wherein, wherein at least one set of dilution openings fluidly connects the shadowed zone to the annular combustion chamber.

12. The gas turbine engine of claim 6, wherein, wherein the baffle comprises a body extending axially between a mounting leg and an inlet leg, and the baffle is mounted to the outer surface of the combustor liner at the mounting leg and the inlet leg is spaced apart from the first end to define a shadowed zone inlet fluidly coupled to the shadowed zone.

13. A burner characterized by, Comprising: an annular combustion chamber having a first end and a second end, wherein the second end is spaced apart from the first end; a plurality of fuel cups located at the first end and annularly arranged within the annular combustion chamber, wherein at least some of the plurality of fuel cups have a defined cup geometry; and A combustor liner at least partially defining the annular combustion chamber and having a varying cross section extending in a downstream direction from the first end toward the second end, the varying cross section defining at least a portion of the defined cup geometry for the fuel cups at the first end and defining a single annular geometry at the second end, the combustor liner including an inner combustor liner and an outer combustor liner, the inner combustor liner spaced apart from the outer combustor liner at the first end to define a circumferentially spaced apart opening between and fluidly coupling adjacent ones of the plurality of fuel cups; wherein the circumferentially spaced apart opening gradually widens as the combustor liner extends from the first end to the second end.

14. The burner of claim 13, wherein wherein the defined cup geometry is a modified can geometry defining a plurality of discrete circumferentially spaced apart fuel cups fluidly connected to one another proximate the first end.

15. The burner of claim 13, wherein wherein the defined cup geometry is a can geometry defining a plurality of discrete circumferentially spaced apart fuel cups fluidly separated from one another proximate the first end.

16. The burner of claim 13, wherein wherein the combustor liner is shaped to define a can annulus geometry downstream of the defined cup geometry and upstream of the single annular geometry.

17. Burner according to any of claims 13-16, characterized in that further comprising at least one set of dilution openings in the combustor liner and a damper mounted to an outer surface of the combustor liner to define a shadowed zone, at least one of the at least one set of dilution openings fluidly connecting the shadowed zone to the annular combustion chamber.

18. A method for controlling nitrogen oxides present in a combustor of a turbine engine, characterized by, the combustor of claim 13, the method comprising: injecting a fuel and compressed air mixture into a combustion chamber of the combustor through a fuel injector; mixing the compressed air and the fuel to define a fuel / air mixture; igniting the fuel / air mixture in a primary zone of the combustion chamber to define a flame and generate a combustion gas; flowing the combustion gas from a plurality of circumferentially spaced apart discrete fuel cups through the combustion chamber to a single annular geometry; and discharging the combustion gas at a combustor outlet.

19. The method of claim 18, wherein, further comprising controlling the flame of the fuel / air mixture in the primary zone by feeding compressed air into a shadowed zone defined by a damper mounted to an outer surface of the combustor.

20. The method of claim 19, wherein, further comprising utilizing an acoustic damper located within the shadowed zone to reduce pressure fluctuations at a particular frequency. further comprising utilizing an acoustic damper located within the shadowed zone to reduce pressure fluctuations at a particular frequency.

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