Multi-fuel stage gas turbine combustor and method of operation

Abstract

A combustor for a gas turbine engine includes a head end section defining a head end plenum and housing a fuel nozzle assembly, and a liner extending from downstream of the head end to an aft frame and defining a combustion chamber therein. A first injector disposed at a first axial position directs a first fuel / air mixture through the liner. A second injector disposed at a different second axial position directs a second fuel / air mixture through the liner. The head end section, the first injector, and the second injector each receive a respective air supply from a compressor discharge plenum at least partially surrounding the combustor. Each air supply is directed to only one of the fuel nozzle assembly, the first injector, and the second injector. The first injector and the second injector receive more than 50% of their air supply from the compressor discharge plenum.

Description

[Technical Field]

[0001] The present disclosure relates generally to gas turbine combustors, and more particularly to combustors having multiple axial stages of fuel injection and methods of operating such combustors. [Background technology]

[0002] Gas turbine engines (such as those used to generate power) typically include a compressor section, a combustion section with one or more combustors, and a turbine section. The compressor section progressively increases the pressure of a working fluid and supplies the compressed working fluid to the combustion section. The compressed working fluid is channeled through one or more axially extending fuel nozzles within the forward or head end of the combustor. Fuel combines with the compressed working fluid flow to form a combustible mixture. The combustible mixture is combusted within the combustion chamber, generating high-temperature, high-pressure, and high-velocity combustion gases. The combustion chamber is bounded by one or more liners or ducts that define a hot gas path through which the combustion gases are conveyed to the turbine section. In a can-annular combustion system, multiple combustion cans (each with its own fuel nozzle and liner) generate the combustion gases that drive the turbine section.

[0003] The combustion gases expand as they flow through the turbine section and produce work. For example, the expansion of the combustion gases in the turbine section can rotate a shaft connected to a generator to produce electricity. The turbine can also drive a compressor or another mechanical load (e.g., a generator or a propeller) through a common shaft or rotor.

[0004] In recent years, manufacturers of large gas turbine engines have invested considerable effort in developing combustion systems that produce low emissions (e.g., NOx emissions), often using fuel nozzles commonly referred to as "micromixers," "advanced premixers," or "bundle-tube fuel nozzles." Each such fuel nozzle, located at the head end of a combustor, includes a group of premixer tubes disposed within a housing and extending through a common fuel plenum defined by the housing. Each premixer tube has one or more fuel injection holes in fluid communication with the fuel plenum. Air enters the upstream ends of the tubes and mixes with fuel through the fuel injection holes within the tubes, resulting in a fuel / air mixture exiting the tubes through the exit ends into the combustion chamber. The flames produced by bundle-tube fuel nozzles are characteristically short.

[0005] Designing a robust bundle-tube fuel nozzle with dozens of premixer tubes presents challenges. Some designs include multiple components of different materials, which increases product cost and complexity. Various approaches have been taken to secure the tubes and associated housing components within the bundle-tube fuel nozzle, join the bundle-tube fuel nozzle to the fuel delivery conduit, ensure sufficient mixing time, adequately cool hot surfaces, and minimize dynamics. Thus, for a single combustor with many (e.g., five or six) bundle-tube fuel nozzles, hundreds of braze joints and multiple seal locations are employed. These numbers are multiplied for gas turbine engines with six to eighteen combustors.

[0006] Further efforts to reduce emissions and improve turndown of gas turbine engines have led to the development of axial fuel staging ("AFS") systems (sometimes referred to as "distributed combustion systems") that include injectors located downstream of the head end that introduce a fuel / air mixture as a crossflow into the combustion gases generated by the head-end fuel nozzles. The region to which the axial fuel staging injectors deliver the fuel / air mixture is often referred to as the "secondary combustion zone," downstream of the "primary combustion zone" supplied by the head-end fuel nozzles. The ability to control multiple fuel delivery locations (e.g., head-end and downstream injectors) provides gas turbine operators with greater flexibility to turndown (i.e., reduce) the power output of the gas turbine engine and distribute heat release, which can reduce engine dynamics.

[0007] In the H-Class combustion system, which has both bundle-tube fuel nozzles and axial-staging fuel injectors, the air flowing into the bundle-tube fuel nozzles is primarily used to cool the combustion liners. That is, air from the compressor discharge case is directed through impingement holes in the flow sleeve, which surrounds the liner and travels through the annulus between the liner and the flow sleeve, thereby convectively cooling the liner. As a result, there is a significant drop in air pressure between the compressor discharge case and the bundle-tube fuel nozzles. A portion of the air is also directed into the axial-staging fuel nozzles.

[0008] Applying the above-described techniques to smaller combustors and gas turbine engine frames (e.g., F-class engines) represents an advancement in the technology. In particular, a retrofittable F-class combustor offering improved hydrogen combustion capabilities and significantly increased turndown, while achieving the same pressure ratio (dP / P), better NOx / T3.90 capability, and control of the exit temperature profile, and offered at a reasonable cost, represents a significant advancement in gas turbine combustion technology. [Prior art documents] [Patent documents]

[0009] [Patent Document 1] US Patent Application Publication No. 2014 / 0174090 Summary of the Invention

[0010] a first plurality of injectors disposed at a first axial position to direct a first fuel / air mixture through the liner; and a second plurality of injectors disposed at a second axial position downstream of the first plurality of injectors to direct a second fuel / air mixture through the liner, wherein the head end section, the first plurality of injectors, and the second plurality of injectors each receive a respective air supply from a compressor discharge plenum defined by a compressor discharge casing at least partially surrounding the combustor, and each air supply is directed to only one of the fuel nozzle assembly, the first plurality of injectors, and the second plurality of injectors, and wherein the first plurality of injectors and the second plurality of injectors receive more than 50% of the air supply from the compressor discharge plenum.

[0011] According to another aspect, a bundle-tube fuel nozzle assembly for a gas turbine combustor includes: a forward plate facing a head-end air plenum; an aft plate facing a combustion chamber; a first plurality of premixer tubes extending from the forward plate to the aft plate; an inner sidewall extending circumferentially around the plurality of premixer tubes and extending axially from the forward plate to the aft plate; and an outer sidewall extending circumferentially around the inner sidewall and extending from the forward plate to the aft plate, wherein the forward plate, the aft plate, the inner sidewall, and the outer sidewall define an external fuel plenum; the forward plate, the aft plate, and the inner sidewall define an internal fuel plenum in fluid communication with the external fuel plenum; each premixer tube of the plurality of premixer tubes includes at least one fuel injection hole therethrough in fluid communication with the internal fuel plenum; and the head-end air plenum is in fluid communication with the combustion chamber through inlet ends of the first plurality of premixer tubes.

[0012] In another aspect, a combustor for a gas turbine engine comprises: a head end section including a fuel nozzle assembly; a liner extending from downstream of the head end section to an aft frame and defining a combustion chamber therein; a first plurality of injectors disposed at a first axial location spaced from the head end section to direct a first fuel / air mixture through the liner; and a second plurality of injectors disposed at a second axial location downstream of the first plurality of injectors to direct a second fuel / air mixture through the liner, wherein each of the head end section, the first plurality of injectors, and the second plurality of injectors receives a respective air supply from a compressor discharge plenum defined by a compressor discharge casing at least partially enclosing the combustor, wherein each air supply is directed to only one of the fuel nozzle assembly, the first plurality of injectors, and the second plurality of injectors, and wherein the second plurality of injectors receives a respective second air supply that is greater than each of a respective first air supply to the first plurality of injectors and a respective third air supply to the head end section.

[0013] According to another aspect, a method of operating a gas turbine combustor having a plurality of axially spaced fuel stages includes selectively directing a fuel and a first air supply through a fuel nozzle assembly in a head end section of the gas turbine combustor to generate a first fuel / air mixture and igniting it within a liner defining a combustion chamber to generate combustion gases, the liner extending from downstream of the head end section to an aft frame; selectively directing a second fuel / air mixture through the liner from at least one of a first plurality of injectors disposed at a first axial location spaced from the head end section; and selectively directing a second fuel / air mixture through the liner from at least one of a first plurality of injectors disposed at a second axial location downstream of the first plurality of injectors. and selectively directing a third fuel / air mixture through the liner from at least one of a second plurality of injectors disposed in a position adjacent the first plurality of injectors, wherein the fuel nozzle assembly, the first plurality of injectors, and the second plurality of injectors each receive a respective air supply from a compressor discharge plenum defined by a compressor discharge casing that at least partially surrounds the combustor, the respective air supply being directed to only one of the fuel nozzle assembly, the first plurality of injectors, and the second plurality of injectors, the second plurality of injectors receiving a respective air supply that is greater than each of the respective air supplies of the fuel nozzle assembly and the first plurality of injectors.

[0014] A further aspect of the present disclosure includes a combustor head end section with an integrated cooling system, the combustor head end section comprising a bundle-tube fuel nozzle assembly for a gas turbine combustor, the bundle-tube fuel nozzle assembly comprising: a forward plate facing a head-end air plenum, an aft plate facing a combustion chamber, a first plurality of premixer tubes extending from the forward plate to the aft plate, and a sidewall extending circumferentially around the first plurality of premixer tubes and axially from the forward plate to the aft plate, wherein the forward plate, the aft plate, and the sidewall define a fuel plenum, each premixer tube of the first plurality of premixer tubes including at least one fuel injection hole therethrough in fluid communication with the fuel plenum, the head-end air plenum being in fluid communication with the combustion chamber through the inlet ends of the first plurality of premixer tubes, and the integrated cooling system being integrated with an outer surface of each premixer tube and comprising at least one air flow passage extending from the forward plate to the aft plate.

[0015] A still further aspect of the present disclosure provides a combustor head end section with an air supply system, the combustor head end section comprising a bundle-tube fuel nozzle assembly for a gas turbine combustor, the bundle-tube fuel nozzle assembly comprising: a forward plate facing a head-end air plenum; an aft plate facing a combustion chamber; a first plurality of premixer tubes extending from the forward plate to the aft plate; and a sidewall extending circumferentially around the first plurality of premixer tubes and axially from the forward plate to the aft plate, The plate and sidewall define a fuel plenum, each premixer tube of the first plurality of premixer tubes including at least one fuel injection hole therethrough in fluid communication with the fuel plenum, the air supply system including a head-end air plenum, a first inlet flow conditioner partially defining the head-end air plenum, and a second inlet flow conditioner, the head-end air plenum in fluid communication with the combustion chamber through the first plurality of premixer tubes, the first inlet flow conditioner surrounding the bundle-tube fuel nozzle assembly, and the second inlet flow conditioner surrounding the first inlet flow conditioner.

[0016] According to another aspect of the present disclosure, a combustor head end section with an air supply system includes a bundle-tube fuel nozzle assembly for a gas turbine combustor, the bundle-tube fuel nozzle assembly including: a forward plate facing a head-end air plenum; an aft plate facing a combustion chamber; a first plurality of premixer tubes extending from the forward plate to the aft plate; a sidewall extending circumferentially around the first plurality of premixer tubes and axially from the forward plate to the aft plate; and an outer sidewall extending from the forward plate to the aft plate and circumferentially around the sidewall; The front plate, the aft plate, and the sidewalls define a fuel plenum, each premixer tube of the first plurality of premixer tubes including at least one fuel injection hole therethrough in fluid communication with the fuel plenum, and the air supply system includes a head-end air plenum, an annular air plenum defined between the outer sidewalls, and a circumferential array of openings defined through the forward plate around the bundle-tube fuel nozzle assembly, the head-end air plenum in fluid communication with the annular air plenum via the circumferential array of openings and in fluid communication with the combustion chamber via the first plurality of premixer tubes.

[0017] According to another aspect, a combustor for a gas turbine engine comprises: a head end section including a fuel nozzle assembly and defining a head end air plenum; a liner extending from downstream of the head end section toward an aft frame and defining a combustion chamber therein; a first plurality of injectors positioned at a first axial location spaced from the head end section to direct a first fuel / air mixture through the liner; and a dynamics mitigation system comprising: a plurality of cold side resonators positioned entirely within the head end air plenum, each resonator of the plurality of cold side resonators having a resonator body with a closed end and an open neck extending from the resonator body opposite the closed end, the resonator bodies defining a respective volume, the open neck in fluid communication with the head end air plenum.

[0018] Two or more aspects described in this disclosure, including those described in this Summary section, may be combined to form implementations not specifically described herein, i.e., all embodiments described herein can be combined with each other.

[0019] The details of one or more exemplary implementations are set forth in the accompanying drawings and the following detailed description. Other features, objects, and advantages will become apparent from the drawings and description, and from the claims.

[0020] A full and enabling disclosure of the present products and methods, including the best mode thereof, directed to one of ordinary skill in the art, is set forth in this specification, which makes reference to the accompanying figures, in which: [Brief explanation of the drawings]

[0021] [Figure 1] 1 is a schematic diagram of a gas turbine assembly including a combustion system as described herein. [Figure 2] 2 is a perspective view of a combustor with multiple axial fuel stages for use in the gas turbine assembly of FIG. 1 in accordance with the present disclosure. [Figure 3A] FIG. 3 is a detailed cross-sectional view of the combustor of FIG. 2. [Figure 3B] FIG. 3 is a schematic cross-sectional view of the combustor of FIG. 2. [Figure 4] FIG. 2 is a side perspective view of a combustor head end section according to aspects of the present disclosure. [Figure 5] FIG. 2 is a perspective view of the forward side of a bundle-tube fuel nozzle assembly and associated damper according to aspects of the present disclosure. [Figure 6] FIG. 6 is a perspective view of the aft side of the bundle-tube fuel nozzle assembly of FIG. 5, including an expanded view of cooling features associated with the bundle-tube fuel nozzle assembly, according to an embodiment of the present disclosure. [Figure 7] FIG. 2 is a partial cross-sectional view of a bundle-tube fuel nozzle assembly of the present disclosure. [Figure 8] FIG. 2 is an interior plan view of a portion of a bundle-tube fuel nozzle assembly of the present disclosure. [Figure 9] FIG. 2 is an enlarged cross-sectional view of a head end section of a combustor according to the present disclosure. [Figure 10] FIG. 2 is a perspective view of a bundle-tube fuel nozzle assembly and associated fuel conduits according to aspects of the present disclosure. [Figure 11] FIG. 2 is a schematic view of the aft face of the bundle-tube fuel nozzle assembly of the present disclosure; [Figure 12] FIG. 2 is an enlarged cross-sectional view of a plurality of premixer tubes of a bundle-tube fuel nozzle assembly according to aspects of the present disclosure. [Figure 13] FIG. 2 is an enlarged cross-sectional view of a plurality of premixer tubes and cooling mechanisms of a bundle-tube fuel nozzle assembly according to aspects of the present disclosure. [Figure 14] FIG. 2 is a side perspective view of an aft portion of the combustor showing the relative positions of a first plurality of fuel injectors and a second plurality of fuel injectors. [Figure 15] FIG. 2 is a schematic diagram illustrating a fuel circuit associated with the combustor, in accordance with various aspects provided herein. [Figure 16] 1 is a series of images illustrating various modes of operation of the present combustor from ignition to full load, in accordance with aspects of the present disclosure. [Figure 17] 1 is a series of images illustrating various operating modes of the present combustor from full load to flameout, in accordance with aspects of the present disclosure. DETAILED DESCRIPTION OF THE INVENTION

[0022] The following detailed description illustrates, by way of example and not limitation, a gas turbine combustor having multiple axially spaced fuel stages, and a method of operating such a gas turbine combustor. This description will enable one skilled in the art to manufacture and use the combustion system. This description provides several embodiments of a combustor assembly, including what is currently believed to be the best mode for making and using the combustion system. An exemplary combustion system is described herein as part of a heavy-duty gas turbine assembly used for power generation. However, it is contemplated that the combustion system described herein may have general application to a wide range of systems in various fields other than power generation.

[0023] As used herein, the terms "first," "second," and "third" may be used interchangeably to distinguish one component from another, but do not denote the location or importance of the individual components. The terms "upstream" and "downstream" refer to a direction relative to fluid flow in a fluid pathway. For example, "upstream" refers to the direction from which the fluid is flowing, and "downstream" refers to the direction from which the fluid is flowing. The terms "forward" and "aft" refer to directions, unless otherwise specified, with "forward" referring to the forward or compressor end of a gas turbine engine and "aft" referring to the aft or turbine end of a gas turbine engine.

[0024] The term "radially" refers to a relative direction substantially perpendicular to the axial centerline of a particular component, and the term "axially" refers to a relative direction substantially parallel to the axial centerline of a particular component. As used herein, the term "radius" (or any variation thereof) refers to a dimension extending outward from the center of any suitable shape (e.g., square, rectangle, triangle, etc.), including but not limited to a dimension extending outward from the center of a circular shape. Similarly, as used herein, the term "circumference" (or any variation thereof) refers to a dimension extending around the center of any suitable shape (e.g., square, rectangle, triangle, etc.), including but not limited to a dimension extending around the center of a circular shape. When a first component is located closer to an axis (i.e., an axial centerline) than a second component, the first component may be referred to herein as being "radially inward" or "inward" of the second component. On the other hand, if a first component is farther from the axis than a second component, the first component may be referred to herein as being "radially outward" or "outward" of the second component. As noted above, and depending on the context, such terms may be applied with respect to the combustor axis or the gas turbine engine axis.

[0025] When an element or layer is referred to as being "on," "engaged," "connected," "coupled," or "attached" to another element or layer, it may be directly on, engaged, connected, coupled, or attached to the other element or layer, or intervening elements or layers may be present. Conversely, when an element is referred to as being "directly on," "directly engaged," "directly connected," or "directly coupled" to another element or layer, there are no intervening elements or layers present. Other terms used to describe relationships between elements should be interpreted similarly (e.g., "between" versus "directly between," "adjacent to" versus "directly adjacent to," etc.). As used herein, the term "and / or" includes any and all combinations of one or more of the associated listed items. The verb forms of "couple" and "attach" may be used interchangeably herein.

[0026] Each example is provided by way of illustration, not limitation, of the invention. Indeed, it will be apparent to those skilled in the art that modifications and variations can be made in the present combustion system and its components without departing from the scope or spirit of the present disclosure. For example, features illustrated or described as part of one embodiment may be used on another embodiment to yield yet a further embodiment. Accordingly, the present disclosure is intended to cover such modifications and variations as fall within the scope of the appended claims and their equivalents. While the exemplary embodiments of the present combustion system and method are generally described with reference to a heavy-duty power-generating gas turbine engine for purposes of illustration, those skilled in the art will readily appreciate that embodiments of the present disclosure are applicable to any combustor incorporated in any turbomachine and are not limited to gas turbine combustors, unless otherwise recited in the claims.

[0027] Reference will now be made in detail to various embodiments of the present combustion systems and methods, one or more examples of which are illustrated in the accompanying drawings. Numerical and letter designations are used in the Detailed Description to refer to features in the drawings. The same or similar designations are used in the drawings and the description to refer to the same or similar parts.

[0028] 1 provides a functional block diagram of an exemplary gas turbine engine 10 that may incorporate various embodiments of the present disclosure. As shown, the gas turbine engine 10 generally includes an inlet section 12 that may include a series of filters, cooling coils, water separators, and / or other devices for cleaning and otherwise conditioning a working fluid (e.g., air) 14 entering the gas turbine engine 10. The working fluid 14 flows to a compressor section where a compressor 16 progressively imparts kinetic energy to the working fluid 14 to produce a compressed working fluid 18.

[0029] The compressed working fluid 18 is mixed with fuel 20 (gaseous or liquid) to form a combustible mixture in one or more combustors 100 of the combustion section or system 24. The combustible mixture, whether gaseous and / or liquid fuel, is combusted to produce high-temperature, high-pressure, and high-velocity combustion gases 26. The combustion gases 26 flow through an expansion turbine 28 of the turbine section to produce work. For example, the expansion turbine 28 may be connected to a shaft 30 such that rotation of the expansion turbine 28 drives the compressor 16 to produce the compressed working fluid 18. Alternatively or additionally, the shaft 30 may connect the expansion turbine 28 to a load, such as a generator 32, for generating electricity.

[0030] Exhaust gases 34 from the expansion turbine 28 flow through an exhaust section (not shown) that connects the expansion turbine 28 to a downstream exhaust stack of the gas turbine engine 10. The exhaust section may include, for example, a heat recovery steam generator (not shown) to clean the exhaust gases 34 and extract additional heat before being released to the environment.

[0031] In an F-class engine (e.g., a nine-frame F-class engine manufactured by GE Vernova, Greenville, South Carolina), the compressor discharge casing partially defines a compressor discharge plenum that surrounds a portion of each combustor (sometimes referred to as a "combustor can"). The expansion turbine casing or shell forms the aft outer boundary of the plenum, and the inner barrel defines the innermost boundary. Each combustor is installed through one of the circumferentially arranged openings in the compressor discharge casing and is positioned between struts that support the compressor discharge casing. For an F-class engine, the compressor discharge casing openings have a diameter of approximately 20 inches. The combustors also join hot gas path hardware (e.g., expansion turbine first stage hardware) at their aft ends. Thus, the subject combustor 100 is configured to be suitable for retrofit into the existing environment of an F-class engine due to its defined opening size, strut location, and expansion turbine hardware location. A conventional F-class combustor has a head end section with swozzle-type (swirl nozzle) fuel nozzles, a large head end diameter, and a length-to-diameter ratio of approximately 3.4 when measured from the head end to the aft frame.

[0032] Figure 2 is a perspective view of combustor 100 of combustion section 24 shown schematically in Figure 1. Figures 3A, 3B, and 4 are cross-sectional views of combustor 100.

[0033] The combustor 100 includes an end cover 102 to which a bundle-tube fuel nozzle assembly 200 is coupled. The end cover 102 defines a forward boundary of a combustor head end section 104. The bundle-tube fuel nozzle assembly 200 defines an aft boundary of the combustor head end section 104 and a forward boundary of the combustion chamber. The head end section 104 further includes a flanged forward casing 103 extending circumferentially around an inner support barrel 105 and extending axially between the end cover 102 and a combustor mounting flange 107. The combustor mounting flange 107 surrounds a portion of all combustors 100 of the gas turbine engine 10 and mounts the combustors 100 to a compressor discharge casing 17 that defines a compressor discharge plenum 19 that receives compressed working fluid (e.g., air) 18 from a compressor 16.

[0034] An inner or first inlet flow conditioner (IFC) 109 extends between the mounting flange 107 and a bundle-tube fuel nozzle assembly 200 that defines a downstream boundary of the head end section 104 and separates the head end section 104 from the primary combustion zone 120. The end cover 102, inner support barrel 105, and inner inlet flow conditioner 109 collectively define a head-end air plenum 111 that supplies air to the bundle-tube fuel nozzle assembly 200. The bundle-tube fuel nozzle assembly 200 has a diameter sized appropriately for retrofitting onto existing F-class engines (e.g., a diameter of 12 to 14 inches, or about 30.5 to about 35.5 centimeters, compared to a diameter of 18 to 20 inches, or about 45.7 to about 50.8 centimeters, for an H-class engine). That is, in contrast to conventional combustion systems in which multiple fuel nozzles are mounted on a cap plate assembly that spans a diameter and defines the upstream boundary of the combustion chamber, the bundle-tube fuel nozzle assembly 200 spans the entire diameter of the forward end of the liner 106.

[0035] The inner inlet flow conditioner 109 is surrounded by an outer or second inlet flow conditioner 113 such that an annular gap is provided between the inner IFC 109 and the outer IFC 113. Both the inner IFC 109 and the outer IFC 113 are cylindrical ducts perforated with a plurality of holes to allow airflow from the compressor discharge casing 17 surrounding the combustor 100 to the head-end air plenum 111. The two sets of perforations in the outer inlet flow conditioner 113 and the inner inlet flow conditioner 109 are configured (in size, number, and alignment) to achieve a desired pressure drop of the compressed air 18 entering the bundle-tube fuel nozzle assembly 200. In particular, the perforations in the outer IFC 113 are larger than the perforations in the inner IFC 109, and the perforations in the outer IFC 113 are aligned with the perforations in the inner IFC 109. Such sizing and alignment is important because the perforations in the inner inlet flow conditioner 109 have been found to have the greatest impact on the pressure drop of the airflow.

[0036] 4 provides a perspective view of the internal components of the head end section 104 of the combustor 100, including, in series from the forward end, a plurality of fuel conduits 402 (specifically, fuel connection conduits 412), the end cover 102, the inner support barrel 105, the inner support barrel flange 121, the inner IFC mounting flange 115, the inner IFC 109, the bundle-tube fuel nozzle assembly 200, and the inner hula seal 219 surrounding the bundle-tube fuel nozzle assembly 200. In the exemplary embodiment, the inner IFC 109 is welded at its aft end to the bundle-tube fuel nozzle assembly 200 and at its forward end to the inner IFC mounting flange 115. The inner support barrel flange 121 may be made of two circumferential segments that collectively define an annular ring. The inner support barrel 105 is formed from two circumferential halves that are bolted or otherwise fastened together to collectively define a complete cylindrical shape. The inner support barrel 105 may include one or more windows 119 to facilitate joining of the inner support barrel 105 to the inner IFC 109. Two circumferential segments or halves of the inner support barrel 105 are welded to circumferential segments of the inner support barrel flange 121, and segments of the inner support mounting flange 121 are bolted to the inner IFC mounting flange 115. The forward end (i.e., each circumferential half) of the inner support barrel 105 is welded to a corresponding semicircular panel of the end cover 102. The two semi-circumferential panels of the end cover 102 are bolted or otherwise removably fastened to one another to collectively define the forward boundary of the head end air plenum 111.

[0037] It should be noted that the above combination of internal components of the head end section 104 is exemplary and should not be construed as limiting the subject technology. The head end section 104, including the above-described internal components, is configured to achieve various technical and practical objectives, such as providing airflow at a desired pressure to the bundle-tube fuel nozzle assembly 200, facilitating assembly and maintenance of the head end section 104, and meeting mechanical and aerodynamic requirements for operation and durability. The above-described combination of components of the head end section 104 may allow the head end section 104 (e.g., including the bundle-tube fuel nozzle assembly 200 and inner IFC 109) to be collectively removable from the liner 106 of the combustor 100. Additionally, an igniter, a flame detector, a dynamic pressure sensor, and / or other sensors may be located within the head end air plenum 111 or in proximity to the head end section 104 (e.g., slightly downstream of the bundle-tube fuel nozzle 200).

[0038] 2, 3A, and 3B, the combustor includes a liner 106 and an outer sleeve 116. The liner 106 defines a combustion chamber having a length “L” from the head end section 104 (e.g., from the aft face of the bundle-tube fuel nozzle assembly 200) to the aft frame 112, with a length-to-head-end-diameter ratio of approximately 2.7. The liner 106 has a generally cylindrical upstream portion 108 and a tapered downstream portion 110 extending between the upstream portion 108 and the aft frame 112. The upstream and downstream portions 108, 110 of the liner 106 may be continuous along the length of the combustor 100 or may be formed as first (e.g., upper or radially outer) and second (e.g., lower or radially inner) clamshell components joined by welding. Unlike many conventional F-series combustors that have separate cylindrical and tapered sections joined by a sealed joint, the present liner 106 can be considered a "unibody" in that the cylindrical section 108 and tapered section 110 are integral with one another. As a result, the elimination of seals prevents loss of airflow between sections 108 and 110.

[0039] Unlike the liner 106, which extends axially between the head end section 104 and the aft frame 112, the outer sleeve 116 includes a forward (first) impingement sleeve 122 and an aft (second) impingement sleeve 124 that span the length L of the combustor 100. The forward impingement sleeve 122 circumferentially surrounds the first plurality of injectors 510 and extends axially from the head end section 104 to the aft impingement sleeve 124. The aft impingement sleeve 124 circumferentially surrounds the second plurality of injectors 550 and extends axially from the forward impingement sleeve 122 to the aft frame 112. The forward impingement sleeve 122 includes a plurality of openings configured (sized and shaped) to deliver airflow from the compressor discharge plenum 19 at a predetermined first pressure drop for the first plurality of injectors.

[0040] The aft impingement sleeve 124 includes a plurality of openings configured (e.g., sized and shaped) to deliver airflow from the compressor discharge plenum 19 to the second plurality of injectors at a predetermined second pressure drop, the second pressure drop being different from the first pressure drop. In one embodiment, the first pressure drop is less than the second pressure drop. A first annulus is defined between the first impingement sleeve 122 and the liner 106, and a second annulus is defined between the second impingement sleeve 124 and the liner 106. The first annulus is fluidly isolated from the second annulus such that the airflow through each impingement sleeve 122, 124 is dedicated to a respective one of the plurality of injectors. As described herein, the airflows to the head end section 104 (e.g., bundle-tube fuel nozzle assembly 200), the first plurality of injectors 510, and the second plurality of injectors 550 are separate from one another.

[0041] To help direct air from the compressor discharge plenum to the respective first or second annulus, the impingement sleeve(s) 122, 124 may be provided with one or more aerodynamic scoops 126 ( FIG. 2 ) that protrude from the outer surface of the respective impingement sleeve 122, 124 into the high-velocity airflow passing through the respective impingement sleeve 122, 124. The aerodynamic scoops 126 function as flow capture devices that, through a combination of stagnation and redirection, capture and redirect air that would previously have passed through the impingement opening but cannot due to a lack of static pressure differential, driving the airflow through the impingement opening. The scoops 126 direct the airflow inward over the hot surface of the liner 106, thus reducing metal temperatures to acceptable levels and increasing the cooling capacity of the impingement sleeve(s) 122, 124. Additionally, the scoops 126 help ensure that sufficient airflow is directed to each of the multiple AFS injectors 510, 550.

[0042] The scoops 126 are attached to or formed in the surface of each impingement sleeve 122, 124 along the side panels of the impingement sleeves 122, 124 and adjacent the impingement sleeve openings. The scoops 126 may at least partially surround one of the impingement sleeve openings and may include a first portion attached to or formed in the surface and an edge defining the open side of the scoop 126. The edge may be oriented in a plane substantially perpendicular to the surface of the impingement sleeves 122, 124 or at other angles toward the direction of airflow. If the impingement sleeves 122, 124 are manufactured by additive manufacturing, the scoops 126 may be integrally formed with the respective impingement sleeves 122, 124. Alternatively, the scoops 126 may be welded to the respective impingement sleeves 122, 124. The number and location of the scoops 126 are defined by the geometry of the impingement sleeves 122, 124, the flow within the compressor discharge casing 17, and the heat load on the liner 106 by the combustor 100.

[0043] Each of the fuel injection systems (bundle-tube fuel nozzle assembly and axial fuel staging injector) will be described in detail below with reference to Figures 5-13, beginning with the bundle-tube fuel nozzle assembly 200. While the head end section 104 is shown as including the bundle-tube fuel nozzle assembly 200 described herein, it should be understood that the head end section 104 may include other types of fuel nozzle assemblies attached to, for example, a cap plate that defines the forward boundary of the combustion chamber. In other words, the aft plate 204 of the bundle-tube fuel nozzle assembly 200 may replace a conventional cap plate, which may itself benefit from the cooling system disclosed herein.

[0044] The bundle-tube fuel nozzle assembly 200 includes a forward plate 202 facing the head-end air plenum 111, an aft plate 204 facing the combustion chamber (e.g., primary combustion zone 120), a plurality of premixer tubes 210 extending from the forward plate 202 to the aft plate 204, and a sidewall 206 extending circumferentially around the plurality of premixer tubes 210 and axially from the forward plate 202 to the aft plate 204. The forward plate 202, the aft plate 204, and the sidewall 206 collectively define a fuel plenum 220 (e.g., an internal fuel plenum) that surrounds the first plurality of premixer tubes 210. The head-end air plenum 111 is in fluid communication with the combustion chamber 120 via inlets in the premixer tubes 210. Each premixer tube 210 includes at least one through-going fuel injection hole 216 (seen most clearly in FIG. 13 ) in fluid communication with the fuel plenum 220. The bundle-tube fuel nozzle assembly 200 spans the entire diameter of the head end section 104 .

[0045] More specifically, each premixer tube 210 defines a premix passage and extends from an inlet 212 defined in the forward plate 202, through a fuel plenum 220, to an outlet 214 defined in the aft plate 204. At least one fuel injection hole 216 may be defined therethrough to provide fluid communication between the fuel plenum 220 and the internal premix passage. During operation, each premixer passage 210 may receive air at the inlet 212 and fuel at the fuel injection holes 216, which are mixed together and discharged as a fuel / air mixture at the outlet 214 for combustion in the primary combustion zone 120. It should be understood that the bundle-tube fuel nozzle assembly 200 may include any number of premixer tubes 210, and the present disclosure should not be limited to any particular number of tubes 210 unless specifically recited in the claims.

[0046] The sidewall 206 of the bundle-tube fuel nozzle assembly 200 defines an inner sidewall, and the fuel plenum 220 is an inner fuel plenum. The inner sidewall (i.e., the sidewall 206) is surrounded by an outer sidewall 226 having an axial length 208 suitable for facilitating additive manufacturing. By way of example, the axial length 208 may be approximately 4 inches (approximately 10 centimeters). An outer fuel plenum 230 is defined between the inner sidewall 206 and the outer sidewall 226 (specifically, between the inner sidewall 206 and the plenum wall 224) and extends circumferentially around the first plurality of premixer tubes 210. At least one fuel conduit 402 ( FIG. 4 ) is coupled to the forward plate 202 in flow communication with the outer fuel plenum 230. Notably, the external fuel plenum 230 is devoid of premixer tubes 210 (i.e., the premixer tubes 210 do not extend through or are in direct fluid communication with the external fuel plenum 230). As shown in FIG. 6, the periphery of the bundle-tube fuel nozzle assembly 200 includes only cooling mechanisms (described below) and is devoid of premixer tubes 210.

[0047] More specifically, as shown in FIG. 7 , a plurality of slot-shaped fuel inlets 232 extend outwardly (i.e., upstream) from the forward plate 202 to fluidly couple the external fuel plenum 230 to respective fuel conduits 402. The slot-shaped fuel inlets 232, integrally constructed with the tube-bundle fuel nozzle assembly 200, are circumferentially spaced about the circumference of the tube-bundle fuel nozzle assembly 200. Any number of slot-shaped fuel inlets 232 may be used, two of which are shown in FIG. 7 , which is a partial cross-sectional view of an embodiment having three slot-shaped fuel inlets 232. Thus, as shown, each of the plurality of circumferentially spaced coupling conduits 242 is in fluid communication with the external fuel plenum 230.

[0048] Fuel is delivered by a fuel coupling conduit 242 ( FIGS. 3A , 9 , 10 ) welded to the slot-shaped fuel inlet 232 and enters the annular external fuel plenum 230. In one embodiment, from the external fuel plenum 230, the fuel enters the internal fuel plenum 220 by flowing through a group of fuel delivery chutes 260 ( FIG. 8 ) located in groups around the periphery of the bundle-tube fuel nozzle assembly 200. The chutes 260 are positioned tangentially to the radius of the bundle-tube fuel nozzle assembly 200 and are oriented to direct fuel between adjacent rows of premixer tubes 210, as represented by the arrows. Such an orientation facilitates fuel distribution from the periphery toward the center of the bundle-tube fuel nozzle assembly 200. Because the air 18 flowing through the premixer tubes 210 is significantly hotter than the fuel, the fuel removes heat from the premixer tubes 210 as it traverses the internal fuel plenum 220. The fuel passes through one or more fuel injection holes 216 (FIG. 13) and is mixed within the premixer tubes 210 before being delivered to the combustion chamber (eg, primary combustion zone 120).

[0049] 9 provides an alternative view of the chute 260 located around the periphery of the tube-bundle fuel nozzle assembly 200. The openings 223 extend inward from the outer fuel plenum 230 relative to the radius of the tube-bundle fuel nozzle assembly 200, fluidly connecting the outer fuel plenum 230 to the inner fuel plenum 220. Because the air 18 flowing through the premixer tubes 210 is significantly hotter than the fuel, the fuel removes heat from the premixer tubes 210 as it traverses the inner fuel plenum 220. The fuel passes through one or more fuel injection holes 216 and is mixed within the premixer tubes 210 before being delivered to the combustion chamber (e.g., the primary combustion zone 120).

[0050] 9 and described in co-pending U.S. patent application Ser. Nos. 18 / 081,924 and 18 / 081,949, both filed December 15, 2022, the plenum wall 224 defining the outer fuel plenum 230 has a corrugated bellows shape, and the inner sidewall 206 extending axially between the forward plate 202 and the aft plate 204 may include a self-breaking portion 264. The bellows-shaped plenum wall 224 intersects with the inner sidewall 206 aft of the self-breaking portion 264. The bellows-shaped plenum wall 224 and the self-breaking portion 264 of the inner sidewall 206 are spaced apart from the axial centerline C of the bundle-tube fuel nozzle assembly 200. L to help manage thermal stresses resulting from temperature differences between the fuel and compressed air.

[0051] The inner sidewall 206 is radially spaced from the outer sidewall 226 such that an annular plenum 225 is defined therebetween, and more specifically, between the bellows-shaped plenum wall 224 and the outer sidewall 226. The annular plenum 225 is an air plenum that is fed by air entering through a plurality of openings 223 ( FIG. 7 ) defined around the periphery of the bundle-tube fuel nozzle assembly 200. The air entering the openings 223 is supplied from the head-end air plenum 111. As shown in FIG. 6 , the air from the annular plenum 225 is channeled through peripheral cooling channels 294 defined in the aft plate 204 and exits the combustion chamber through swirl funnels 298, as described below.

[0052] A fuel delivery system 270 for the bundle-tube fuel nozzle assembly 200 is shown in FIG. 10 . In the exemplary embodiment, three fuel conduits 402 and a central fuel conduit 404 are shown. As described above, multiple (e.g., three) fuel coupling conduits 242 ( FIG. 3A ) are welded to the slot-shaped fuel inlets 232 of the bundle-tube fuel nozzle assembly 200. Each fuel coupling conduit 242 has a forward end with a circular cross-section and an aft end with a slot-shaped cross-section. The forward end of each fuel coupling conduit 242 is welded to a straight pipe section 252 that extends between the fuel coupling conduit 242 and a fuel line bellows 262 (e.g., a single or dual fuel bellows assembly). The fuel line bellows 262 reduces the possibility of fuel leakage outside the head end section 104 of the combustor 100. The fuel line bellows 262 is attached to a mounting flange 272 that attaches each fuel conduit 402 to the end cover 102. A fuel connection conduit 412 extends forward of the end cover 102 for connection to a fuel supply line (not shown) external to the combustor 100. Because the fuel conduit 402 extends from the fuel conduit 242 through the end cover 102, the fuel conduit 402 is straight (no bends) and there are no fuel seals or fittings (e.g., compression fittings) within the end cover 102 that could potentially deteriorate or fail, resulting in a leak. The absence of fuel seals and fittings is especially important when operating with highly reactive fuels such as hydrogen.

[0053] As shown in Figures 5, 7, 9, and 10, the bundle-tube fuel nozzle assembly 200 includes a central fuel inlet 234 that is integrally formed with the bundle-tube fuel nozzle assembly 200 and extends axially outward from the forward plate 202. A connecting conduit 244 is attached to the central fuel inlet 234, and a fuel conduit 254 is joined to the connecting conduit 244. In this case, the connecting conduit 244 is straight, and the fuel conduit 254 may have a curved portion. As shown, the central fuel inlet 234, the connecting conduit 244, and the fuel conduit 254 may each have a circular cross-section, although other cross-sectional shapes may alternatively be used. As with the fuel conduit 402, a fuel line bellows 262 is used to reduce the possibility of fuel leakage from the central fuel conduit 404. The central fuel conduit 404 also includes a mounting flange 272 for coupling to the end cover 102 and a fuel connection conduit 412 for coupling to an external fuel supply line (not shown).

[0054] The central fuel conduit 404 supplies fuel to the central fuel plenum 222, which is shown schematically in FIG. 11 . For simplicity, FIG. 11 depicts only a portion of the premixer tubes 210. The central fuel plenum 222 is defined by a bulkhead 280 that includes a plurality of boundary premixer tubes 282 and a boundary air passage 284 located between each pair of adjacent boundary premixer tubes 282. The bulkhead 280 extends axially from the forward plate 202 to the aft plate 204, and the boundary air passages 284 are in fluid communication with the head-end air plenum 111. Each boundary air passage 284 of the plurality of boundary air passages has a uniform cross-sectional shape from an inlet in the forward plate 202 to an outlet defined in the aft plate 204. As described further below, the outlet is fluidly coupled to one or more cooling channels 294 defined in the aft plate 204.

[0055] As a result of the partition 280, the premixer tubes 210 define a radially outer group of radially outer premixer tubes 210, 278 in fluid communication with the internal fuel plenum 220, a radially inner group of premixer tubes 210, 286 in fluid communication with the central fuel plenum 222, and the aforementioned boundary premixer tubes 210, 282 positioned along and forming part of the partition 280. As best shown in FIG. 13 , the boundary premixer tubes 210, 282 have a plurality of fuel injection holes 216, with at least one fuel injection hole 290 in fluid communication with the internal fuel plenum 220 and at least one fuel injection hole 292 in fluid communication with the central fuel plenum 222. The presence of the boundary premixer tubes 282 between the inner fuel plenum 220 and the central fuel plenum 222 relieves thermal stresses and ensures the absence of flame front voids that would otherwise be caused by a solid boundary wall between the fuel plenums 220 and 222, even when the fuel plenums 220 and 222 are supplied separately or at different flow rates.

[0056] 12 and 13 , at least one premixer tube of the plurality of premixer tubes 210, 278, 286 includes at least one air passage 296 integral with an outer surface of the respective premixer tube 210, 278, 286 and extending from the forward plate 202 to the aft plate 204 of the bundle-tube fuel nozzle assembly 200. Each respective air passage 296, which is part of the integrated cooling system of the head end section 104, is in fluid communication with the head-end air plenum 211 and the combustion chamber (e.g., the primary combustion zone 120). More specifically, each respective air passage 296 has an inlet defined in the forward plate 202 and an outlet defined in the aft plate 204, the outlet being fluidly coupled to at least one cooling channel 294 defined in the aft plate 204 and extending between the respective outlets of adjacent premixer tubes 210, 278, 286 to provide convection cooling for the aft plate 204. An inlet flow restriction (not shown) may be provided to optimize the amount of cooling flow provided by the cooling channels 294. In various embodiments, the integrated cooling system includes one or two air passages 296 integrated with the exterior surface of each premixer tube 210, 278, 286.

[0057] Each of the cooling channels 294 terminates in a swirl funnel 298, which is a conical structure that is in fluid communication with the combustion zone to deliver a swirling flow of air to the combustion zone. Specifically, each of the cooling channels 294 intersects the swirl funnel 298 tangentially, thereby promoting a swirling flow as the air exits the swirl funnel 298. In some embodiments, the swirl funnel 298 can include a lip that reduces the diameter of the opening of the swirl funnel 298, which accelerates the flow from the swirl funnel 298. Introducing a small swirl flow between the non-swirling flows generated by the premixer tubes 210, 278, 282, and 286 promotes combustion in the primary combustion zone 120. The swirl funnel 298 is included in the head end section 104's integrated cooling system.

[0058] The bundle-tube fuel nozzle assembly 200 is formed from an additively manufactured body in which the premixer tubes 210, 278, 282, 286 are integrally connected to the forward plate 202 and the aft plate 204, as described above. The premixer tubes 210, 278, 282, 286 each have a relatively thin tube wall thickness when compared to the thickness of the forward plate 202. Other features, including the inner sidewall 206, the outer sidewall 226, the plenum wall 224 (e.g., a corrugated plenum wall), the fuel inlets 232, 234, the bulkhead 280, and the cooling features of the head-end section integrated cooling system (e.g., the boundary air passage 284, the air passage 296, the convection cooling channels 294, and the swirl funnel 298), are incorporated into the additively manufactured body.

[0059] It has been found that additively manufactured parts having elements of different thicknesses can experience thermal stresses. To distribute the transition between the thickness of the forward plate 202 and the wall thickness of each premixer tube 210, the forward plate 202 includes ridges 213 (FIGS. 5 and 7) that protrude radially outward from the forward plate 202 in a honeycomb pattern between the premixer tubes 210 (e.g., between the premixer tubes 278 of the first plurality of premixer tubes). The inlet ends of the premixer tubes 210, 278, 282, 286 are concentric with and recessed relative to the recesses defined by the ridges 213. Such a configuration (i.e., ridge, recess, concentric arrangement) also helps prevent the formation of steps on the interior surface of the premixer tubes 210, which could pose flame retention concerns.

[0060] Additively manufacturing all components together as a single body eliminates the need for multiple welded or brazed joints between multiple small components, reducing the possibility of leaks from the fuel plenums 220, 222, 230. Additionally, such an assembly eliminates seals between the fuel nozzles in the head end section, providing greater engineering flexibility for premixer tube placement and avoiding additional hardware to complete the assembly.

[0061] Another feature additively manufactured with the bundle-tube fuel nozzle assembly is a mount 302 ( FIG. 7 ) for one or more quarter-wave tube dampers 300 (visible in FIGS. 5 , 9 , and 10 ). The mount 302 is integrally coupled to the aft plate 204, extends through the fuel plenums 220, 222, and terminates at the upstream end of the forward plate 202, where the term “upstream” refers to the flow of air through the premixer tubes 210 and / or fluid through the combustion chamber. As seen most clearly in FIG. 5 , an extended damper body 306 is welded to the damper mount 302 to form a quarter-wave tube damper of a desired length to mitigate specific frequency dynamics. The damper end 308 (of the extended damper body 306) has one or more openings 310 to allow airflow through the damper 300. The damper end 308 is distal to the forward plate 202.

[0062] The damper mount 302 of each of the one or more quarter-wave tube dampers 300 is thermally decoupled from the interior sidewall 206 to prevent distortion, such as may result from temperature differences between the cold fuel and hot air flowing through the damper mount 302. The damper mount 302 is surrounded by a bellows mount 304 that protrudes from the forward plate 202 into the head-end air plenum 111, as shown in FIG. 7 . The bellows mount 304 is closer to the forward plate 202 than the damper mount 302. The annular gap between the damper mount 302 and the bellows mount 304 can be used for powder removal after additive manufacturing of the bundle-tube fuel nozzle assembly 200.

[0063] A bellows assembly 312 (shown in FIG. 5 ) is coupled to the bellows mount 304 of each of the one or more quarter-wave tube dampers 300 to prevent fuel from leaking from the fuel plenums 220, 222 into the head-end plenum 111. The bellows assembly 312 includes a bellows with multiple folds and a thin shield that protects the bellows. The bellows assembly 312 may be coupled to the bellows mount 304 by welding. The segmented structure of the quarter-wave tube damper 300 (e.g., including the bellows assembly 312) allows for a significant reduction in stresses due to thermal growth of the quarter-wave tube damper 300 distributed at different locations within the bundle-tube fuel nozzle assembly 200. As a result, the likelihood of cracking at these locations is minimized, improving the life and durability of the component.

[0064] 5 , for example, one or more quarter-wave tube dampers 300 may comprise a plurality of quarter-wave tube dampers 300 having extended damper bodies 306 of at least two different lengths to mitigate combustion dynamics of different frequencies. In the illustrated embodiment, seven quarter-wave tube dampers 300 are provided with damper bodies 306 of four different lengths. More or fewer dampers 300 may be used as part of the combustion dynamics mitigation system of the combustor 100. Alternatively, the damper bodies 306 may have a uniform length, with inserts (not shown) of different lengths installed therein to achieve desired damping characteristics.

[0065] It is envisioned that the damper mount 302, bellows mount 304, and bellows assembly 312 can be used with one or more liquid fuel cartridges (not shown). In such a case, the liquid fuel cartridges would be replaced by one or more of the extended damper bodies 306. For example, every other one of the three extended damper bodies 306 around the periphery of the bundle-tube fuel nozzle assembly 200 can be replaced with liquid fuel cartridges, which can be designed with a spray pattern that extends across approximately one-third of the aft plate 206 of the head end section 104. Additionally or alternatively, the central damper mount 302 can be coupled to a pilot fuel nozzle that can provide pilot fuel for starting with a highly reactive fuel, such as hydrogen.

[0066] 3A and 3B, the dynamics mitigation system may also include a resonator 320 and / or a damper 340. The resonator 320 may be provided within the head-end plenum 111 (i.e., as a “cold-side” damper), and the damper 340 may be provided through the liner 106 and the outer sleeve 122 and / or 124 (i.e., as a “hot-side” damper). Specifically, multiple resonators 320 may be disposed singly or in groups within the head-end plenum 111. Each resonator 320 has a resonator body defining a volume. The resonator body is closed at one end and has an open neck extending from the resonator body opposite the closed end. The neck, having a smaller diameter than the closed end, is in fluid communication with the head-end air plenum 111. In an exemplary embodiment, the multiple resonators 320 include one or more sets of at least one resonator 320 (e.g., three sets of three resonators, each having a different volume). Each set is coupled to an upstream inner support barrel 105 that partially surrounds a fuel conduit 402 coupled to the bundle-tube fuel nozzle assembly 200 .

[0067] One or more (“hot side”) dampers 340 may extend through the respective outer sleeves 122, 124 and liners 106 such that the damper volumes are in fluid communication with the combustion chamber. Such dampers 340 are useful for mitigating mode shape dynamics experienced during different operating conditions of the combustor 100. In the exemplary embodiment, a first damper 340 is positioned forward of the first plurality of injectors 510, a second damper 340 is positioned between the first plurality of injectors 510 and the second plurality of injectors 550, and a third damper 340 is positioned aft of the second plurality of injectors 550. More or fewer dampers 340 may be used and may be positioned at various locations in fluid communication with the combustion chamber. The first damper 340, second damper 340, and third damper 340 may have two or more different damper volumes to damp different combustion dynamics frequencies. Therefore, the illustrated number and placement of dampers 340 should not be considered limiting of the present subject matter.

[0068] The combustor 100, its head end section 104, and its axial fuel staging system 500 are sized for retrofit into an existing opening in a conventional 9F-class compressor discharge casing 17. The head end section 104 with its bundle-tube fuel nozzle assembly 200 is described above. As shown in FIGS. 2, 3A, and 3B, the axial fuel staging system 500 includes a first plurality of injectors 510 disposed at a first axial location to direct a first fuel / air mixture through the liner 106 and into a secondary combustion zone 512, and a second plurality of injectors 550 disposed at a second axial location downstream of the first plurality of injectors to direct a second fuel / air mixture through the liner 106 and into a tertiary combustion zone 514. Each of the head end section 104 (e.g., bundle-tube fuel nozzle assembly 200), the first plurality of injectors 510, and the second plurality of injectors 550 receives a respective, separate air supply from a compressor discharge plenum 19 defined by a compressor discharge casing 17 that at least partially surrounds the combustor 100. Each air supply is directed to only one of the head end section 104 (e.g., bundle-tube fuel nozzle assembly 200), the first plurality of injectors 510, and the second plurality of injectors 550.

[0069] The first and second injectors 510 and 550 together receive more than 50% of their air supply from the compressor discharge casing 17. The first air supply to the first plurality of injectors 510 differs in volume from the second air supply to the second plurality of injectors 550 and the third air supply to the head end section 104. The second plurality of injectors 550 receive their respective (second) air supplies that are greater than the respective (third) air supplies of the head end section 104 (e.g., bundle-tube fuel nozzle assembly 200) and greater than the respective (first) air supplies of the first plurality of injectors 510. Additionally, in various preferred embodiments, the second air supply to the second plurality of injectors 550 is less than the sum of the first air supply to the first plurality of injectors 550 and the third air supply to the head end section 104.

[0070] Axial fuel staging system 500 includes a first plurality of injectors 510 and a second plurality of injectors 550. First plurality of injectors 510 are disposed in a forward (upstream) portion of liner 106, and second plurality of injectors 550 are disposed in an aft (downstream) portion of liner 106 extending between the upstream portion and aft frame 112. More specifically, as shown in FIG. 3A , the upstream portion of liner 106 is defined between the aft face of bundle-tube fuel nozzle assembly 200 (or a cap plate, if a different fuel nozzle is used) and first plurality of injectors 510 (i.e., immediately downstream of first plurality of injectors 510), and the downstream portion of liner 106 is defined between the upstream portion and aft frame 112. In the exemplary embodiment, first plurality of injectors 510 includes at least two injectors, and second plurality of injectors 550 includes three or more injectors. In the illustrated embodiment, the first plurality of injectors 510 includes three injectors and the second plurality of injectors 550 includes six injectors. In the exemplary embodiment, the injectors 510, 550 are identical to one another. In other embodiments, different injectors may be used between the first and second injection stages or within each injection stage.

[0071] In one embodiment, as shown in FIG. 14 , one injector of the first plurality of injectors 510 is disposed in a forward lower section (e.g., half) 106A of the liner 106, and two injectors of the first plurality of injectors 510 are disposed in a forward upper section (e.g., half) 106B of the liner 106, the forward lower section 106A of the liner 106 being aligned with the axial centerline GT C of the gas turbine engine 10. L (FIG. 1), and the forward upper section 106B of the liner 106 is located proximal to the axial centerline GT C of the gas turbine engine 10. LIn the illustrated embodiment, four injectors of the second plurality of injectors 550 are disposed in an aft upper section (e.g., half) 106C of the liner 106, and two injectors of the second plurality of injectors 550 are disposed in an aft lower section (e.g., half) 106D of the liner 106, which is located distal to the axial centerline GT C of the gas turbine engine 10. L (FIG. 1), and the aft lower section 106D of the liner 106 is located distal to the axial centerline GT C of the gas turbine engine 10. L (FIG. 1). In other words, the injectors 510, 550 in the upper section of the liners 106A, 106B are disposed in the radially outer section of the liners 106A, 106B, and the injectors 510, 550 in the lower section of the liners 106A, 106B are disposed in the radially inner section of the liners 106A, 106B.

[0072] The injectors 510 may be collectively supplied with fuel via an "AFS-1" circuit. Alternatively, a first subset of the injectors 510 may be coupled to a first fuel circuit (e.g., "AFS-1A"), and the remaining injectors 510 (a second subset) may be coupled to a second fuel circuit (e.g., "AFS-1B"). Each of the first and second subsets of injectors 510 may include one or more injectors.

[0073] Similarly, injectors 550 may be collectively fueled via an "AFS-2" fuel circuit. Alternatively, a third subset of injectors 550 may be coupled to a third fuel circuit (e.g., "AFS-2A"), and the remainder of injectors 550 (a fourth subset) may be coupled to a fourth fuel circuit (e.g., "AFS-2B"). Each of the third and fourth subsets of injectors 550 may include one or more injectors.

[0074] 14, the injector 510 on the forward lower section 106A of the liner 106 is considered to be an "AFS-1A" circuit, and the injector 510 on the forward upper section 106B of the liner 106 is considered to be an "AFS-1B" circuit. The injector 550 on the aft lower section 106C of the liner 106 is considered to be an "AFS-2A" circuit, and the injector 550 on the aft upper section 106D of the liner 106 is considered to be an "AFS-2B" circuit.

[0075] During operation, the first and second injectors 510 and 550 receive more than 50% of the airflow from the compressor discharge casing 17. The first and second injectors 510 are surrounded by a first impingement sleeve 122 having a first plurality of impingement openings. A first air supply from the compressor discharge plenum 17 is in fluid communication with the first injectors 510 through the first plurality of impingement openings, such that each first air supply experiences a first pressure drop from flowing through the first plurality of impingement openings. Similarly, the second injectors 550 are surrounded by a second impingement sleeve 124 having a second plurality of impingement openings. A second air supply from the compressor discharge plenum 17 is in fluid communication with the second injectors 550 through the second plurality of impingement openings, such that each second air supply experiences a second pressure drop from flowing through the second plurality of impingement openings. The second pressure drop is different from the first pressure drop.

[0076] 2, 3A, and 3B, each injector of the first and second pluralities of injectors 510 and 550 includes an injector body 515 having an elongated shape that is longer axially than laterally (e.g., a geometric stadium shape) and an injector fuel conduit 520 that extends from the injector body 515, through the mounting flange 107, and along the outer surface of the outer sleeve 116. Each injector fuel conduit 520 includes a straight portion 522 adjacent the mounting flange 107 and may include a curved portion 524 between the straight portion and the respective injector body 515. An arched conduit shield or cover 526 is removably mounted over the straight portion 522 by bolting to the outer inlet flow conditioner 113 that is part of the head end section 104 of the combustor 100.

[0077] In various embodiments, at least one injector of the first plurality of injectors 510 is oriented at a first angle relative to the combustor centerline, and at least one injector of the second plurality of injectors 550 is oriented at a second angle relative to the combustor centerline, the first angle and the second angle being different from one another. For example, the first angle may be approximately 10 degrees (±5 degrees) and the second angle may be approximately 30 degrees (±5 degrees). In some embodiments, at least two injectors of the first plurality of injectors 510 are oriented at one or more first angles relative to the combustor centerline that are different from one or more second angles of at least two injectors of the second plurality of injectors 550. In some embodiments, at least two injectors of the second plurality of injectors 550 are oriented at different angles relative to the combustor centerline. For example, the aft ends of adjacent pairs of injectors of the second plurality of injectors may be oriented toward one another as shown in Figures 3A and 14. The one or more first angles of at least two injectors of the first plurality of injectors 510 and the one or more second angles of at least two injectors of the second plurality of injectors 550 are configured to promote complete combustion and optimize the exit temperature profile, the emissions profile, or both.

[0078] 15 is a schematic diagram of the control of fuel delivery for the combustor 100 described herein. Fuel from a fuel supply is directed through a series of gas control valves (GCV1-GCV4) and optional separate valves 602, 604 associated with the fuel circuit of a given combustor component or region. The control valves GCV1-GCV4 and separate valves 602, 604 are in electronic communication with a controller 600, as represented by dashed lines.

[0079] More specifically, the controller 600 controls GCV1 to regulate fuel flow to the central plenum 222 of the bundled tube fuel nozzle assembly (“BTFN”) 200 (or, for example, to the central fuel nozzle in the head end section 104). The controller 600 controls GCV2 to regulate fuel flow to the inner plenum 220 of the bundled tube fuel nozzle assembly 200 (or, for example, to outer fuel nozzles positioned around the central nozzle in the head end section 104). The controller 600 controls GCV3 to regulate fuel flow to a first plurality of injectors 510 that are a first axial distance from the head end section 104. In an embodiment in which the first plurality of injectors 510 includes a first subset of injectors and a second subset of injectors, the first subset of injectors may be fueled via an “AFS-1A” circuit, and the second subset of injectors may be fueled via an “AFS-1B” circuit. A separate (on / off) valve 602 can be used to direct flow to (or stop flow from) the "AFS-1B" circuit. Similarly, in embodiments where the second plurality of injectors 550 includes a first subset of injectors and a second subset of injectors, the first subset of injectors can be fueled through the "AFS-2A" circuit and the second subset of injectors can be fueled through the "AFS-2B" circuit. A separate (on / off) valve 604 can be used to direct flow to (or stop flow from) the "AFS-2B" circuit.

[0080] Figure 16 provides a pictorial representation of the ramp-up of gas turbine load from light-off to full-speed full-load. Figure 17 provides a pictorial representation of the turn-down of a gas turbine engine from full-speed full-load to flameout (shutdown of the combustion system). In Figures 16 and 17, fueled subsets are designated with the letter "H" for "hot" and unfueled subsets are designated with the letter "C" for "cold." Load percentages referenced in the descriptions of Figures 16 and 17 are expressed based on the operating temperature of the gas turbine combustor 100. Therefore, such load percentages should not be considered limiting.

[0081] As described herein, the present disclosure provides a method of operating a gas turbine combustor 100 having multiple axially spaced fuel stages. The method includes selectively directing a fuel and a first air supply through a fuel nozzle assembly (e.g., bundle-tube fuel nozzle assembly 200) in a head end section 104 of the gas turbine combustor 100 to generate a first fuel / air mixture and igniting the mixture in a liner 106 defining a combustion chamber to generate combustion gases, the liner 106 extending from downstream of the head end section 104 to an aft frame 112. The method further includes selectively directing a second fuel / air mixture through the liner 106 from at least one of a first plurality of injectors 210 disposed at a first axial location spaced from the head end section 104, and selectively directing a third fuel / air mixture through the liner 104 from at least one of a second plurality of injectors 550 disposed at a second axial location downstream of the first plurality of injectors 510. Each of the bundle-tube fuel nozzle assembly 200, the first plurality of injectors 510, and the second plurality of injectors 550 receives a respective, separate air supply from a compressor discharge plenum 19 defined by a compressor discharge casing 17 that at least partially surrounds the combustor 100. Each air supply is directed to only one of the bundle-tube fuel nozzle assembly 200, the first plurality of injectors 510, and the second plurality of injectors 550. The second plurality of injectors 550 receives a respective air supply that is greater than the respective air supply of the bundle-tube fuel nozzle assembly 200 and greater than the respective air supply of the first plurality of injectors 510.

[0082] As represented by the image in FIG. 16, the method steps of selectively directing fuel and a first air supply through the head end section 104 including the bundle-tube fuel nozzle assembly 200 are performed in all combustion modes from ignition to full speed, full load operation.

[0083] As described herein, the bundle-tube fuel nozzle assembly 200 defines a central fuel plenum 222 and an internal fuel plenum 220 surrounding the central fuel plenum 222. The method step of selectively directing fuel and a first air supply through the head end section 104 including the bundle-tube fuel nozzle assembly 200 includes directing fuel through both the central fuel plenum 222 and the internal fuel plenum 220. In some embodiments, fuel is supplied to the central fuel plenum 222 and the internal fuel plenum 220 to generate different fuel / air ratios between the central fuel plenum 222 and the internal fuel plenum 220.

[0084] As described above, the first plurality of injectors 510 comprises one or more injectors in a first subset (e.g., AFS-1A) and the remaining injectors of the first plurality of injectors in a second subset (e.g., AFS-1B). As shown in image 2 of FIG. 16 , the method step of selectively directing a second fuel / air mixture from at least one of the first plurality of injectors 510 through the liner 106 from about 93% of full speed up to about 7% load includes supplying fuel to the first subset (e.g., AFS-1A) of one or more injectors of the first plurality of injectors 510 and not supplying fuel to the second subset (e.g., AFS-1B) of the remaining injectors of the first plurality of injectors 510.

[0085] In some embodiments, as shown in FIG. 14 , a first subset (e.g., AFS-1A) of the first plurality of injectors 510 includes one or more injectors located in the forward lower section 106A of the liner 106, and the remaining number of injectors of the first plurality of injectors 510 in a second subset (e.g., AFS-1B) includes one or more injectors located in the forward upper section 106B of the liner 106.

[0086] When increasing the load from approximately 8% to approximately 17%, as shown in image 3 of Figure 16, the method step of selectively directing a second fuel / air mixture from at least one of the first plurality of injectors 510 through the liner 106 includes supplying fuel to one or all of the injectors of the first plurality of injectors 510.

[0087] As described above, the second plurality of injectors 550 includes one or more injectors in a third subset (e.g., AFS-2A) and the remaining number of injectors of the second plurality of injectors 550 in a fourth subset (e.g., AFS-2B). From ignition to approximately 18% load, as shown in images 1-3 of FIG. 16, the method step of selectively directing a third fuel / air mixture from at least one of the second plurality of injectors 550 through the liner 106 includes not supplying fuel to the second plurality of injectors 550. Thus, as shown in image 3, the first plurality of injectors 510 are supplied with fuel (represented by the letter "H") and the second plurality of injectors 550 are not supplied with fuel (represented by the letter "C").

[0088] As shown in Image 4 of FIG. 16 , the method step of selectively directing a third fuel / air mixture from at least one of the second plurality of injectors 550 through the liner 106 from approximately 18% load to approximately 35% load includes supplying fuel to a third subset of one or more injectors (e.g., AFS-2A) and not supplying fuel to a fourth subset of one or more remaining injectors of the second plurality of injectors 550 (e.g., AFS-2B).

[0089] In some embodiments, as shown in FIG. 14 , a third subset (e.g., AFS-2A) of the second plurality of injectors 550 includes one or more injectors located in the rear lower section 106C of the liner 106, and the remaining number of injectors of a fourth subset (e.g., AFS-2B) of the second plurality of injectors 550 includes one or more injectors located in the rear upper section 106D of the liner 106.

[0090] From approximately 36% load to full load, as shown in image 5 of FIG. 16 , the method of selectively directing a second fuel / air mixture from at least one of the first plurality of injectors 510 through the liner 106 includes supplying fuel to each injector of the first plurality of injectors 510, and the method of selectively directing a third fuel / air mixture from at least one of the second plurality of injectors 550 through the liner 106 includes supplying fuel to each injector of the second plurality of injectors 550.

[0091] 17, the method of operating the combustor 100 preferably further includes turning down the gas turbine combustor to maintain compliance with emissions regulations. As shown in image 1 of FIG. 17, all circuits of the combustor 100 are fueled from 100% load to approximately 36% load. The gas turbine can be unloaded by reducing fuel to each circuit until the gas turbine reaches approximately 36% load.

[0092] The mode switch begins at approximately 35% load, and the method step of selectively directing a third fuel / air mixture from at least one of the second plurality of injectors 550 through the liner 106 includes ceasing to supply fuel to a fourth subset of one or more injectors (e.g., AFS-2B). As described above, the second plurality of injectors 550 includes the third subset of one or more injectors and the remaining number of injectors of the second plurality of injectors in the fourth subset. As shown in image 2 of FIG. 17 , the gas turbine combustor 100 may continue to unload from approximately 35% load to approximately 18% load, while fuel is redistributed among the remaining active circuits (i.e., circuits supplying the bundle-tube fuel nozzle assembly, AFS-1A injectors, AFS-1B injectors, and AFS-2A injectors).

[0093] As described, a third subset (e.g., AFS-2A) of the second plurality of injectors 550 includes one or more injectors located in the rear lower section 106C of the liner 106, and the remaining number of injectors of a fourth subset (e.g., AFS-2B) of the second plurality of injectors 550 includes one or more injectors located in the rear upper section 106D of the liner 106.

[0094] Another mode switch begins at approximately 17% load, where the method step of selectively directing a third fuel / air mixture from at least one of the second plurality of injectors 550 through the liner 106 includes ceasing fuel supply to each injector of the second plurality of injectors 550 by ceasing fuel supply to a third subset of the second plurality of injectors 550 (e.g., AFS-2A), as shown in Image 3 of Figure 17. As shown in Image 3 of Figure 17, the gas turbine combustor 100 can continue to unload from approximately 17% load to approximately 8% load, while fuel is redistributed among the remaining active circuits (i.e., circuits supplying the bundle-tube fuel nozzle assembly, the AFS-1A injector, and the AFS-1B injector).

[0095] As described above, the first plurality of injectors 510 includes a first subset (e.g., AFS-1A) of one or more injectors and a second subset (e.g., AFS-1B) of the remaining injectors of the first plurality of injectors 510. In some embodiments, the first subset (e.g., AFS-1A) of the first plurality of injectors 510 includes one or more injectors located in the forward lower section 106A of the liner 106, and the remaining injectors of the first plurality of injectors 510 in the second subset (e.g., AFS-1B) include one or more injectors located in the forward upper section 106B of the liner 106.

[0096] Another mode switch begins at approximately 7% load, where the method step of selectively directing a second fuel / air mixture from at least one of the first plurality of injectors 510 through the liner 106 includes ceasing to supply fuel to a second subset of one or more injectors (e.g., AFS-1B). As shown in image 4 of Figure 17, the gas turbine combustor 100 can continue unloading from approximately 7% load to approximately 93% speed while fuel is redistributed among the remaining active circuits (i.e., the circuits supplying the bundle-tube fuel nozzle assemblies and the AFS-1A injectors).

[0097] Another mode switch begins at approximately 92% speed, and the method step of selectively directing a second fuel / air mixture from at least one of the first plurality of injectors 510 through the liner 106 includes ceasing to supply fuel to each injector of the first plurality of injectors 510. Thus, as shown in Image 5, only the bundle-tube fuel nozzle assembly 200 remains fueled, and two subsets of each plurality of injectors 510, 550 are not fueled.

[0098] The method further includes turning off the gas turbine combustor. When turning off the gas turbine combustor, the method step of selectively directing fuel and a first air supply through a bundle-tube fuel nozzle assembly 200 in the head end section 104 of the gas turbine combustor 100 to generate a first fuel / air mixture includes reducing the supply of fuel to the bundle-tube fuel nozzle assembly 200 until a flameout occurs.

[0099] To facilitate implementation of the methods described herein, the gas turbine combustor 100 further includes a controller 600 (FIG. 15) configured to selectively direct fuel to the bundle-tube fuel nozzle assembly 200, the first plurality of injectors 510, and the second plurality of injectors 550.

[0100] Exemplary embodiments of a combustion system, its various components, and methods of operation thereof have been described in detail above. The methods, systems, and components described herein are not limited to the specific embodiments described herein; rather, components of the methods and systems can be utilized independently and separately from other components described herein. For example, the methods and systems described herein may have other applications that are not limited to practice in the turbine assemblies described herein. Rather, the methods and systems described herein can be implemented and utilized in connection with various other industries.

[0101] While the claimed subject matter has been described in terms of various specific embodiments, those skilled in the art will recognize that the techniques can be practiced with modification within the spirit and scope of the claims.

[0102] Exemplary clauses that can be used to describe the present combustor and its bundle-tube fuel nozzle assembly are as follows: Any and all features of all clauses can be combined in any practical manner to provide further embodiments.

[0103] According to a first aspect, a combustor for a gas turbine engine includes: a head end section defining a head end plenum and housing a fuel nozzle assembly; a liner extending from downstream of the head end section to an aft frame and defining a combustion chamber therein; a first plurality of injectors disposed at a first axial position spaced from the head end section to direct a first fuel / air mixture through the liner; and a second plurality of injectors disposed at a second axial position different from the first plurality of injectors to direct a second fuel / air mixture through the liner, wherein the head end section, the first plurality of injectors, and the second plurality of injectors each receive a respective air supply from a compressor discharge plenum defined by a compressor discharge casing at least partially surrounding the combustor, wherein each air supply is directed to only one of the fuel nozzle assembly, the first plurality of injectors, and the second plurality of injectors, and wherein the first plurality of injectors and the second plurality of injectors receive more than 50% of the air supply from the compressor discharge plenum.

[0104] Another aspect of the present disclosure includes any of the preceding aspects, wherein the first air supply to the first plurality of injectors is volumetrically different from both the second air supply to the second plurality of injectors and the third air supply to the head end section.

[0105] Another aspect of the present disclosure includes any of the preceding aspects, wherein an upstream portion of the liner is defined between the head end section and the first plurality of injectors, a downstream portion of the liner is defined between the first plurality of injectors and the aft frame, and a second plurality of injectors is disposed in the downstream portion of the liner.

[0106] Another aspect of the present disclosure includes any of the preceding aspects, wherein the second air supply to the second plurality of injectors is greater than each of the first air supply to the first plurality of injectors and the third air supply to the head end section.

[0107] Another aspect of the present disclosure includes any of the preceding aspects, wherein the second air supply to the second plurality of injectors is less than the sum of the first air supply to the first plurality of injectors and the third air supply to the head end section.

[0108] Another aspect of the present disclosure includes any of the preceding aspects, wherein the first plurality of injectors are surrounded by a first impingement sleeve having a first plurality of impingement openings, and a first air supply from the compressor discharge plenum is in fluid communication with the first plurality of injectors through the first plurality of impingement openings, whereby each first air supply experiences a first pressure drop from flowing through the first plurality of impingement openings.

[0109] Another aspect of the present disclosure includes any of the preceding aspects, wherein the second plurality of injectors is surrounded by a second impingement sleeve having a second plurality of impingement openings, and a second air supply from the compressor discharge plenum is in fluid communication with the second plurality of injectors through the second plurality of impingement openings, whereby the second air supply experiences a second pressure drop from flowing through the second plurality of impingement openings, the second pressure drop being different from the first pressure drop.

[0110] Another aspect of the present disclosure includes any of the preceding aspects, wherein the first plurality of injectors includes at least two injectors and the second plurality of injectors includes three or more injectors.

[0111] Another aspect of the present disclosure includes any of the preceding aspects, wherein at least one injector of the first plurality of injectors is disposed in a radially outer section of the liner and at least one injector of the first plurality of injectors is disposed in a radially inner section of the liner, the radially outer section of the liner being distal to an axial centerline of the gas turbine engine, and the radially inner section of the liner being proximal to the axial centerline of the gas turbine engine.

[0112] Another aspect of the present disclosure includes any of the preceding aspects, wherein at least one injector of the second plurality of injectors is disposed in a radially outer section of the liner and at least one injector of the second plurality of injectors is disposed in a radially inner section of the liner, the radially outer section of the liner being distal to an axial centerline of the gas turbine engine, and the radially inner section of the liner being proximal to the axial centerline of the gas turbine engine.

[0113] Another aspect of the present disclosure includes any of the preceding aspects, wherein the first plurality of injectors is grouped into a first subset and a second subset, the first subset being coupled to a first fuel circuit and the second subset being coupled to a second fuel circuit.

[0114] Another aspect of the present disclosure includes any of the preceding aspects, wherein the second plurality of injectors is grouped into a third subset and a fourth subset, the third subset being coupled to a third fuel circuit, and the fourth subset being coupled to a fourth fuel circuit.

[0115] Another aspect of the present disclosure includes any of the preceding aspects, wherein the fuel nozzle assembly is a bundle-tube fuel nozzle assembly including a forward plate facing a head-end air plenum, an aft plate facing a combustion chamber, a plurality of premixer tubes extending from the forward plate to the aft plate, and a sidewall extending circumferentially around the plurality of premixer tubes and extending axially from the forward plate to the aft plate, whereby the forward plate, the aft plate, and the sidewall define a fuel plenum, the head-end air plenum in fluid communication with the combustion chamber through the plurality of premixer tubes, each premixer tube of the plurality of premixer tubes including at least one fuel injection hole therethrough in fluid communication with the fuel plenum, and wherein the bundle-tube fuel nozzle assembly spans the entire diameter of the head-end section.

[0116] Another aspect of the present disclosure includes any of the preceding aspects, wherein the at least one resonator is disposed through a fuel plenum of the bundle-tube fuel nozzle assembly and extends into a head-end plenum.

[0117] Another aspect of the present disclosure includes any of the preceding aspects, wherein at least one resonator is disposed entirely within the head end section, away from the bundle-tube nozzle assembly, the at least one resonator having a body and a neck extending therefrom, the neck in fluid communication with a head end air plenum defined in part by an inlet flow conditioner surrounding the head end section.

[0118] Another aspect of the present disclosure includes any of the preceding aspects, wherein the head end section including the bundle-tube fuel nozzle assembly and the inlet flow conditioner surrounding the bundle-tube fuel nozzle assembly are collectively removable from the liner. [Explanation of symbols]

[0119] 10. Gas Turbine Engine 12 Entrance Section 14 Working fluid 16 Compressor 17 Compressor discharge casing, compressor discharge plenum 18 Working fluid (air) 19 Compressor discharge plenum 20 fuel 24 Combustion section or combustion system 26 Combustion Gas 28 Expansion turbine 30 shaft 32 Generator 34 Exhaust gas 100 Combustor 102 End cover 103 Flanged forward casing 104 Head End Section 105 Inner support barrel 106 Liner 106A Radial inner section, liner 106B Radially outer section, liner 106C Radial outer section 106D Radial Inner Section 107 Combustor mounting flange 108 Upstream part 109 Inner Inlet Flow Regulator 110 Downstream part 111 Head end air plenum 112 Rear frame 113 Outer or Second Inlet Flow Regulator 115 Inner IFC mounting flange 116 Outer sleeve 119 Window 120 Combustion chamber (primary combustion zone) 121 Inner support barrel flange 122 Front (first) collision sleeve 124 Rear (second) collision sleeve 126 Aerodynamic Scoop 200 Bundle-tube fuel nozzle assembly, bundle-tube fuel nozzle 202 Front Plate 204 Rear Plate 206 Inner side wall 208 Axial length 210 Premix tube, premix passage 212 Entrance 213 Ridge 214 Exit 216 Fuel injection hole 219 Inner Hula Seal 220 Internal Fuel Plenum 222 Central Fuel Plenum 223 Opening 224 Plenum Wall 225 Annular Plenum 226 Exterior side wall 230 External Fuel Plenum 232 Fuel inlet 234 Central fuel inlet 242 Combined Conduit 244 Combined Conduit 252 Straight Pipe Section 254 Fuel conduit 260 Fuel Delivery Chute 262 Fuel line bellows 264 Self-breaking part 270 Fuel Delivery System 272 Mounting flange 278 Premix tube 280 Bulkhead 282 Boundary premixing tube 284 Boundary Air Flow Path 286 Premixing tube 290 Fuel injection hole 292 Fuel injection hole 294 Cooling Channel 298 Vortex Funnel 300 1 / 4 wavelength tube damper 302 Mounting body 304 Bellows mounting body 306 Damper body 308 Damper end 310 Aperture 312 Bellows Assembly 320 resonator 340 Damper 402 Fuel conduit 404 Central fuel conduit 412 Fuel connection pipe 500 Axial Fuel Staging System 510 first plurality of injectors 512 Secondary Combustion Zone 514 Tertiary Combustion Zone 515 injector body 520 Injector fuel conduit 522 Straight section 524 curved part 526 Arched conduit shield or cover 550 second plurality of injectors 600 Controller 602 Valve 604 Valve GCV1 Control Valve GCV2 control valve GCV3 Control Valve GCV4 control valve GTC L Gas turbine engine axial centerline AFS-1A First Fuel Circuit AFS-1B second fuel circuit AFS-2A third fuel circuit AFS-2B 4th Fuel Circuit

Claims

1. A combustor (100) for a gas turbine engine (10), the combustor (100) comprising: a head end section (104) defining a head end plenum (111) and housing a fuel nozzle assembly (200); a liner (106) extending downstream from the head end section (104) to an aft frame (112) and defining a combustion chamber (120, 512, 514) therein; a first plurality of injectors (510) disposed at a first axial location spaced from the head end section (104) to direct a first fuel / air mixture through the liner (106); a second plurality of injectors (550) disposed at a second axial position different from the first plurality of injectors (510) to direct a second fuel / air mixture through the liner (106); Equipped with each of the head end section, the first plurality of injectors, and the second plurality of injectors receives a respective air supply from a compressor discharge plenum defined by a compressor discharge casing that at least partially surrounds the combustor, the respective air supplies being directed to only one of the fuel nozzle assembly, the first plurality of injectors, and the second plurality of injectors; the first plurality of injectors (510) and the second plurality of injectors (550) receive more than 50% of the air supply from the compressor discharge plenum (17); A combustor (100).

2. 2. The combustor of claim 1, wherein a first air supply to the first plurality of injectors is volumetrically distinct from both a second air supply to the second plurality of injectors and a third air supply to the head end section.

3. 2. The combustor of claim 1, wherein an upstream portion of the liner is defined between the head end section and the first plurality of injectors, a downstream portion of the liner is defined between the first plurality of injectors and the aft frame, and the second plurality of injectors are disposed in the downstream portion of the liner.

4. 4. The combustor of claim 3, wherein a second air supply to the second plurality of injectors is greater than each of a first air supply to the first plurality of injectors and a third air supply to the head end section.

5. 4. The combustor of claim 3, wherein a second air supply to the second plurality of injectors is less than a sum of a first air supply to the first plurality of injectors and a third air supply to the head end section.

6. 2. The combustor of claim 1, wherein the first plurality of injectors are surrounded by a first impingement sleeve having a first plurality of impingement openings, and wherein a first air supply from the compressor discharge plenum is in fluid communication with the first plurality of injectors through the first plurality of impingement openings, whereby each first air supply experiences a first pressure drop from flowing through the first plurality of impingement openings.

7. 7. The combustor of claim 6, wherein the second plurality of injectors are surrounded by a second impingement sleeve having a second plurality of impingement openings, and wherein a second air supply from the compressor discharge plenum is in fluid communication with the second plurality of injectors through the second plurality of impingement openings, whereby the second air supply experiences a second pressure drop from flowing through the second plurality of impingement openings, the second pressure drop being different from the first pressure drop.

8. The combustor (100) of any preceding claim, wherein the first plurality of injectors (510) comprises at least two injectors and the second plurality of injectors (550) comprises three or more injectors.

9. At least one injector of the first plurality of injectors is disposed in a radially outer section of the liner, and at least one injector of the first plurality of injectors is disposed in a radially inner section of the liner, and the radially outer section of the liner is aligned with an axial centerline (GT C) of the gas turbine engine. Ｌ ) and the radially inner section (106A) of the liner (106) is distal to the axial centerline (GT C Ｌ The combustor (100) of claim 8, wherein the combustor (100) is proximate to the

10. At least one injector of the second plurality of injectors is disposed in a radially outer section of the liner, and at least one injector of the second plurality of injectors is disposed in a radially inner section of the liner, and the radially outer section of the liner is aligned with an axial centerline (GT C) of the gas turbine engine. Ｌ ) and the radially inner section (106D) of the liner (106) is distal to the axial centerline (GT C Ｌ The combustor (100) of claim 8, wherein the combustor (100) is proximate to the

11. 9. The combustor of claim 8, wherein the first plurality of injectors are grouped into a first subset and a second subset, the first subset being coupled to a first fuel circuit (AFS-1A) and the second subset being coupled to a second fuel circuit (AFS-1B).

12. 9. The combustor of claim 8, wherein the second plurality of injectors is grouped into a third subset and a fourth subset, the third subset being coupled to a third fuel circuit (AFS-2A), and the fourth subset being coupled to a fourth fuel circuit (AFS-2B).

13. The fuel nozzle assembly (200) is a bundle-tube fuel nozzle assembly (200) including a forward plate (202) facing a head-end air plenum (111), an aft plate (204) facing the combustion chamber (120), a plurality of premixer tubes (210) extending from the forward plate (202) to the aft plate (204), and a sidewall (206) extending circumferentially around the plurality of premixer tubes (210) and axially from the forward plate (202) to the aft plate (204), whereby the forward plate (20) 2), the aft plate (204), and the sidewall (206) define a fuel plenum (220), the head-end air plenum (111) is in fluid communication with the combustion chamber (120) through the plurality of premixer tubes (210), each premixer tube (210) of the plurality of premixer tubes including at least one fuel injection hole (216) therethrough in fluid communication with the fuel plenum (220), and the bundle-tube fuel nozzle assembly (220) spans a diameter of the head-end section (104).

14. The combustor of claim 13, wherein at least one resonator is disposed through the fuel plenum of the bundle-tube fuel nozzle assembly and extends into the head end plenum.

15. 14. The combustor of claim 13, wherein at least one resonator is disposed completely within the head end section away from the bundle-tube nozzle assembly, the at least one resonator having a body and a neck extending therefrom, the neck in fluid communication with a head end air plenum defined in part by an inlet flow conditioner surrounding the head end section.

16. 14. The combustor of claim 13, wherein the head end section including the bundle-tube fuel nozzle assembly and an inlet flow conditioner surrounding the bundle-tube fuel nozzle assembly are collectively removable from the liner.

Images