Combustor air flow path

Abstract

The invention is entitled "Combustor Air Flow Path". A combustor (50) for a turbine is disclosed that includes a head end (72), a liner (66) at least partially defining a hot gas path, and a flow sleeve (68) circumferentially surrounding at least a portion of the liner (66). The flow sleeve (68) is spaced apart from the liner (66) to form a cooling flow annulus (90) therebetween. The cooling flow annulus (90) is in direct fluid communication with a high pressure plenum (58), whereby air from the high pressure plenum flows into the cooling flow annulus (90) and from the cooling flow annulus (90) to the head end (72). The combustor (50) also includes a first combustion zone and a second combustion zone downstream of the first combustion zone along the hot gas path (32, 70). A plurality of fuel injectors (84) are in fluid communication with the second combustion zone and are not in direct fluid communication with the high pressure plenum (58).

Description

TECHNICAL FIELD

[0001] The present disclosure relates generally to combustors for turbomachinery. More particularly, the present disclosure relates to combustors having axially staged fuel injectors and features that define air flow paths for such combustors. BACKGROUND

[0002] Turbomachinery is used in a variety of industries and applications for energy transfer purposes. For example, a gas turbine engine generally includes a compressor section, a combustion section, a turbine section, and an exhaust section. The compressor section progressively increases the pressure of a working fluid entering the gas turbine engine and supplies this compressed working fluid to the combustion section. The compressed working fluid and a fuel, such as natural gas, are mixed within the combustion section and combusted in a combustion chamber to generate high pressure and high temperature combustion gases. The combustion gases flow from the combustion section into the turbine section where the combustion gases expand to do work. For example, the expansion of the combustion gases in the turbine section can cause a rotor shaft connected to, for example, an electrical generator, to rotate to produce electricity. The combustion gases then exit the gas turbine via the exhaust section.

[0003] Gas turbines generally combust a hydrocarbon fuel and produce emissions, such as nitrogen oxides (NOx) and carbon monoxide (CO). It is generally desirable to minimize the production of such emissions. The oxidation of molecular nitrogen in a gas turbine depends on the temperature of the gases located in the combustor, as well as the residence time of the reactants in the highest temperature region within the combustor. Thus, the amount of NOx produced by a gas turbine can be reduced by keeping the combustor temperature below the temperature at which NOx is produced or by limiting the residence time of the reactants in the combustor.

[0004] One method for controlling the temperature of a combustor involves premixing fuel and air to produce a lean fuel-air mixture prior to combustion. The method can include axial staging of fuel injection, where a first fuel-air mixture is injected and ignited at a first or primary combustion zone of the combustor to produce a primary stream of high-energy combustion gases, and where a second fuel-air mixture is injected into and mixed with the primary stream of high-energy combustion gases via a plurality of radially oriented and circumferentially spaced fuel injectors or an axially staged fuel injector assembly positioned downstream of the primary combustion zone, sometimes also referred to as an aft lean injector. The axial staging of injection increases the likelihood of complete combustion of the available fuel, which in turn reduces undesirable emissions.

[0005] During operation of the combustor, it is necessary to cool one or more liners or ducts forming the combustion chamber and / or the hot gas path passing through the combustor. Liner cooling is typically achieved by directing a cooling medium, such as compressed working fluid from a compressor, through cooling flow annuli or flow passages defined between the liner and a flow sleeve and / or an impingement sleeve surrounding the liner. Accordingly, diverting a portion of the compressed working fluid passing through the axial staging injector can reduce the amount of cooling provided to the exterior of the combustion chamber.

[0006] Accordingly, an improved system and method for supplying compressed working fluid to an axial staging injector without reducing the cooling provided to the combustion liner or duct would be useful. SUMMARY

[0007] Aspects and advantages of systems according to the disclosure will be set forth in part in the following description, or can be apparent from the description, or can be learned through practice of the technology.

[0008] According to one embodiment, a combustor for a turbine is provided. The combustor is coupled to an outer casing of the turbine and is in fluid communication with a high pressure plenum within the outer casing. The combustor includes a head end, a liner at least partially defining a hot gas path, and a flow sleeve circumferentially surrounding at least a portion of the liner. The flow sleeve is spaced apart from the liner to form a cooling flow annulus therebetween. The cooling flow annulus is in direct fluid communication with the high pressure plenum, whereby air from the high pressure plenum flows into the cooling flow annulus and from the cooling flow annulus to the head end. The combustor further includes a first combustion zone defined by the liner and a second combustion zone defined by the liner downstream of the first combustion zone along the hot gas path. A plurality of fuel injectors is in fluid communication with the second combustion zone. The plurality of fuel injectors is configured to inject a mixture of fuel and air directly into the second combustion zone. The plurality of fuel injectors is not in direct fluid communication with the high pressure plenum.

[0009] According to another embodiment, a turbine is provided. The turbine includes a compressor extending from an inlet to a discharge. The discharge of the compressor provides a high pressure air flow directly into a high pressure plenum defined within an outer casing of the turbine. The turbine further includes a combustor. The combustor includes a head end, a liner at least partially defining a hot gas path, and a flow sleeve circumferentially surrounding at least a portion of the liner. The flow sleeve is spaced apart from the liner to form a cooling flow annulus therebetween. The cooling flow annulus is in direct fluid communication with the high pressure plenum, whereby air from the high pressure plenum flows into the cooling flow annulus and from the cooling flow annulus to the head end. The combustor further includes a first combustion zone defined by the liner and a second combustion zone defined by the liner downstream of the first combustion zone along the hot gas path. A plurality of fuel injectors are in fluid communication with the second combustion zone. The plurality of fuel injectors are configured to inject a mixture of fuel and air directly into the second combustion zone. The plurality of fuel injectors are not in direct fluid communication with the high pressure plenum. The turbine further includes a turbine downstream of the combustor and an exhaust outlet downstream of the turbine.

[0010] These and other features, aspects, and advantages of the present components will become better understood with reference to the following description and appended claims. The accompanying drawings, which are incorporated in and constitute a part of this specification, illustrate embodiments of the technology and serve to explain the principles of the technology. BRIEF DESCRIPTION OF DRAWINGS

[0011] A complete and enabling disclosure of the inventive system, including the best mode thereof, directed to those of ordinary skill in the art in this field of technology is set forth in the specification which makes reference to the appended drawings in which:

[0012] Figure 1 is a schematic illustration of a turbine according to an embodiment of the present disclosure;

[0013] Figure 2 shows a cross-sectional side view of a portion of an exemplary turbine including an exemplary combustor that can encompass various embodiments of the present disclosure;

[0014] Figure 3 shows a simplified side cross-sectional view of a portion of a combustor according to one or more embodiments of the present disclosure;

[0015] Figure 4 shows a perspective view of a portion of a combustor for a turbine according to one or more embodiments of the present disclosure;

[0016] Figure 5 shows a cross-sectional view of a flange of a combustor for a turbine according to one or more embodiments of the present disclosure;

[0017] Figure 6a cross-sectional view of a flange of a combustor for a turbomachine is shown in accordance with one or more additional embodiments of the present disclosure; and

[0018] Figure 7 a schematic cross-sectional view of a portion of certain components of a combustor for a turbomachine is shown in accordance with one or more embodiments of the present disclosure. DETAILED DESCRIPTION

[0019] Reference will now be made in detail to implementations of the present system, one or more examples of which are illustrated in the drawings. Each example is provided by way of explanation of the present technology and not as a limitation thereto. In fact, it will be apparent to those skilled in the art that modifications and variations to the implementations can be made without departing from the scope or spirit of the technology as disclosed in the specification. For instance, features illustrated or described as part of one implementation can be used with another implementation to yield a still further implementation. Thus, it is intended that this disclosure encompass such modifications and variations as come within the scope of the appended claims and their equivalents.

[0020] The DETAILED DESCRIPTION uses numerical and letter designations to refer to features in the drawings. Like or similar designations in the drawings and description have been used to identify like or similar elements. As used herein, the terms "first," "second," and "third" can be used interchangeably to distinguish one component from another and are not intended to signify location or importance of the individual components.

[0021] As used herein, the terms "upstream" (or "up") and "downstream" (or "down") refer to the relative direction with respect to the flow of fluid in a fluid pathway. For example, "upstream" refers to the direction from which the fluid flows, and "downstream" refers to the direction to which the fluid flows. The term "radially" refers to a relative direction substantially perpendicular to an axial centerline of a particular component, the term "axially" refers to a relative direction substantially parallel and / or coaxially aligned with an axial centerline of a particular component, and the term "circumferentially" refers to a relative direction extending around an axial centerline of a particular component. Approximating language, such as "generally," or "approximately," is used herein to include values that are within ten percent of a stated magnitude. Such terminology includes amounts that are the same as and amounts that are closer to the stated magnitude when compared to the amounts that are outside of ten percent of the stated magnitude. When used in the context of an angle or a direction, such terminology includes angles or directions that are within ten degrees of the stated angle or direction. For example, "generally vertical" includes directions that are within ten degrees of vertical in either direction (e.g., clockwise or counterclockwise).

[0022] Referring now to the drawings, Figure 1A schematic diagram of one embodiment of a turbomachine, which in the illustrated embodiment is a gas turbine 10, is shown. Although an industrial or land-based gas turbine is shown and described herein, the present disclosure is not limited to industrial and / or land-based gas turbines unless otherwise specified in the claims. For example, the systems as described herein can be used in any type of turbomachine, including but not limited to a steam turbine, an aircraft gas turbine, or a marine gas turbine.

[0023] As shown, the gas turbine 10 generally includes an inlet section 12, a compressor section 14 disposed downstream of the inlet section 12, a plurality of combustors 50 (one example of which is shown) disposed within a combustor section 16 downstream of the compressor section 14, a turbine section 18 disposed downstream of the combustor section 16, and an exhaust section 20 disposed downstream of the turbine section 18. Additionally, the gas turbine 10 can include one or more shafts 22 coupled between the compressor section 14 and the turbine section 18. Figure 2 One example of which is shown in the center) disposed within a combustor section 16 downstream of the compressor section 14, a turbine section 18 disposed downstream of the combustor section 16, and an exhaust section 20 disposed downstream of the turbine section 18. Additionally, the gas turbine 10 can include one or more shafts 22 coupled between the compressor section 14 and the turbine section 18.

[0024] The compressor section 14 can generally include a plurality of rotor disks 24 (one of which is shown) and a plurality of rotor blades 26 extending radially outward from and connected to each rotor disk 24. Each rotor disk 24 can in turn be coupled to or form a portion of the shaft 22 extending through the compressor section 14.

[0025] The turbine section 18 can generally include a plurality of rotor disks 28 (one of which is shown) and a plurality of rotor blades 30 extending radially outward from and interconnected to each rotor disk 28. Each rotor disk 28 can in turn be coupled to or form a portion of the shaft 22 extending through the turbine section 18. The turbine section 18 also includes an outer casing 31 that circumferentially surrounds portions of the shaft 22 and the rotor blades 30, thereby at least partially defining a hot gas path 32 through the turbine section 18.

[0026] During operation, a working fluid such as air flows through the inlet section 12 and into the compressor section 14, where the air is progressively compressed, thereby providing pressurized air to the combustors of the combustor section 16. The pressurized air is mixed with fuel and combusted within each combustor to produce combustion gases 34. The combustion gases 34 flow from the combustor section 16 through the hot gas path 32, into the turbine section 18, where energy (kinetic and / or thermal) is transferred from the combustion gases 34 to the rotor blades 30, thereby causing the shaft 22 to rotate. The mechanical rotational energy can then be used to power the compressor section 14 and / or generate electricity. The combustion gases 34 exiting the turbine section 18 can then be exhausted from the gas turbine 10 via the exhaust section 20.

[0027] Figure 2A cross-sectional side view of a portion of an exemplary gas turbine 10 including an exemplary combustor 50, which can be one of several combustors provided in a combustor section 16 as shown in FIG. 1 and described above, is provided. The exemplary combustor 50 shown can encompass various embodiments of the present disclosure. As shown, the combustor 50 is at least partially surrounded by an outer casing 52, such as a compressor discharge casing 54 provided downstream of the compressor 14, and / or an outer turbine casing 56. The outer casing 52 is in fluid communication with the compressor 14 and at least partially defines a high pressure plenum 58 surrounding at least a portion of the combustor 50. An end cover 60 is coupled to the outer casing 52 at one end of the combustor 50. Figure 1

[0028] As shown, the combustor 50 generally includes at least one axially extending fuel nozzle 62 extending downstream from the end cover 60, an annular end cap assembly 64 extending radially and axially within the outer casing 52 downstream of the end cover 60, an annular hot gas path conduit or combustion liner 66 extending downstream from the end cap assembly 64, and an annular flow sleeve 68 surrounding at least a portion of the combustion liner 66. The combustion liner 66 defines a hot gas path 70 for directing the combustion gases 34 through the combustor 50. The end cover 60 and the end cap assembly 64 at least partially define a head end 72 of the combustor 50. Figure 2

[0029] The end cap assembly 64 generally includes a forward end 74 positioned downstream of the end cover 60, an aft end 76 provided downstream of the forward end 74, and one or more annular shrouds 78 extending at least partially therebetween. In particular embodiments, the axially extending fuel nozzle 62 extends at least partially through the end cap assembly 64 to provide a first combustible mixture 80, which consists primarily of fuel and a portion of the compressed working fluid 19, e.g., air, from the compressor 14 to a main combustion zone 82 defined within the combustion liner 66 downstream of the aft end 76 of the end cap assembly 64.

[0030] In particular embodiments, the combustor 50 further includes one or more radially extending fuel injectors 84, also referred to as axial staging fuel injectors or post-lean fuel injectors, extending through the flow sleeve 68 and the combustion liner 66 at a point downstream of the at least one axially extending fuel nozzle 62. The combustion liner 66 defines a combustion chamber 86 within the combustor 50. In particular embodiments, the combustion liner 66 further defines a secondary combustion zone 88 proximate the fuel injectors 84 and downstream of the main combustion zone 82. In particular embodiments, the combustion liner 66, the flow sleeve 68, and the fuel injectors 84 are provided as part of a combustion module 100 that extends axially through the outer casing 52 and circumferentially around at least a portion of the end cap assembly 64. ​​

[0031] Combustion module 100 includes a forward or upstream end 102 axially separated from an aft or downstream end 104 relative to an axial centerline 106 of combustion module 100. Figure 4 As shown, combustion liner 66 extends downstream and terminates at an aft frame 130. A mounting bracket 131 can be coupled to aft frame 130. In some embodiments, aft frame 130 and / or mounting bracket 131 can be coupled to an outer turbine casing 56, and mounting flange 112 can be connected to a compressor discharge casing 54 so as to constrain combustion module 100 at forward end 102 and aft end 104 of combustion module 100. Figure 2

[0032] A simplified side cross-sectional view of a portion of combustor 50 according to various embodiments of the present disclosure is provided. As can be seen in Figure 3 flow sleeve 68 can be spaced apart from liner 66 to form a cooling flow annulus 90 therebetween. Compressed working fluid 19 from compressor discharge plenum 58 can flow along an exterior of liner 66 through cooling flow annulus 90 to provide convective cooling to liner 66, then reverse direction to flow through head end 72 and axially extending fuel nozzles 62. Figure 3 Figure 2 A plurality of fuel injectors 84 can be arranged circumferentially around liner 66 and flow sleeve 68 downstream of one or more of fuel nozzles 62. Fuel injectors 84 provide fluid communication through liner 66 and flow sleeve 68 and into combustion chamber 86. Fuel injectors 84 can receive the same or different fuel as that supplied to fuel nozzles 62, and mix the fuel with a portion of compressed working fluid 19 and then inject the mixture into combustion chamber 86 or simultaneously inject the mixture into the combustion chamber. In this way, fuel injectors 84 can supply a mixture of fuel and compressed working fluid 19 directly to a secondary combustion zone 88 for additional combustion, thereby raising the temperature of combustor 50 and thus improving efficiency. In exemplary embodiments, fuel is delivered through passages defined in flow sleeve 68, but fuel conduits disposed radially outward of flow sleeve 68 (as shown in

[0033] In some embodiments, as shown in Figure 4 combustor 50 can include at least one air shroud 92 around some or all of plurality of fuel injectors 84. For example, in some embodiments, a single air shroud 92 (such as

[0034] In some embodiments, as shown in Figure 3 combustor 50 can include at least one air shroud 92 around some or all of plurality of fuel injectors 84. For example, in some embodiments, a single air shroud 92 (such as Figure 3 ​The air shrouds 92 can circumferentially surround the fuel injectors 84 to prevent the fuel injectors 84 from being directly impinged by the compressed working fluid 19 flowing from the compressor 14. Thus, the plurality of fuel injectors 84 are not in direct fluid communication with the high pressure plenum 58. The air shrouds 92 can be pressure fit or otherwise connected to the mounting flange 112 and / or around the circumference of the flow sleeve 68 to provide a substantially enclosed volume or second annular passage 94 between the air shroud 92 and the flow sleeve 68. The air shroud 92 can extend axially along a portion or the entire length of the flow sleeve 50, terminating at or slightly aft of the fuel injectors 84. In Figure 3 In the particular embodiment shown, for example, the air shroud 92 extends axially along the entire length of the flow sleeve 68 such that the air shroud 92 is substantially coextensive with the flow sleeve 68.

[0035] In some embodiments, for example, as Figure 4 shown, a plurality of air shrouds 92 can be provided, such as one air shroud 92 for each fuel injector 84, for example, a one-to-one correspondence between the air shrouds 92 and the fuel injectors 84, such that there is one air shroud 92 for each fuel injector 84 and each air shroud 92 surrounds one fuel injector 84. As Figure 4 shown, each fuel injector 84 can be fluidly coupled to a fuel source by a fluid conduit 126 extending between the fuel injector 84 and the mounting flange 112. Additionally as Figure 4 shown, the aft frame 130 can be positioned at and extend around the aft or downstream end 128 of the combustion liner 66. For example, the aft frame 130 can circumferentially surround the aft end 128, for example, as Figure 4 shown.

[0036] Figure 5 A cross-sectional view of the flange 112 is shown for use with embodiments including a single air shroud 92, for example, as Figure 3 shown. In such embodiments, the flange 112 can include a single passage 96 that is continuous around the flange 112, for example, the single passage 96 can extend circumferentially around the entire flange 112, as Figure 5 shown. Figure 6 A cross-sectional view of the flange 112 is shown for use with embodiments including a plurality of air shrouds 92, for example, as Figure 4 shown. In such embodiments, the flange 112 can include a plurality of passages 96 therethrough and the plurality of passages 96 can be arranged in a circumferential array on the flange 112, for example, the plurality of passages 96 can be spaced around the circumference of the flange 112, as Figure 6 shown, such that each passage 96 is circumferentially aligned with a respective air shroud 92 and fuel injector 84.

[0037] The enclosed volume 94 defined by each air shroud 92 can be in direct fluid communication with the slots or passages 96 in the flanges 112 to receive a direct flow of compressed working fluid (e.g., air) 19 from the flanges 112 and deliver the air 19 to the plurality of fuel injectors 84, such as by one or more of the passages 96 in the mounting flanges 112, for example, as shown in Figure 3 and Figure 7 For example, in various embodiments, a portion of the air 19 flowing through the cooling flow annulus 90 can be directed radially outward into one or more of the passages 96, for example, as shown in Figure 7 Furthermore, the flow of air 19 from the flanges 112 to the plurality of fuel injectors 84 can be the only flow of air to the plurality of fuel injectors 84. Thus, the plurality of fuel injectors 84 can be located completely downstream of the flanges 112 and be in indirect fluid communication with the high-pressure plenum 58 and the cooling flow annulus 90, for example, via the flanges 112 only, with the compressed working fluid (e.g., air) 19 reaching the plurality of fuel injectors 84 only after traveling completely through the cooling flow annulus 90 and then to at least the flanges 112. Thus, the compressed working fluid can flow to the plurality of fuel injectors 84 only after flowing through the entire cooling flow annulus 90, for example, along a continuous and uninterrupted flow path from the high-pressure plenum 58 to the flanges 112 (e.g., via the cooling flow annulus 90). For example, the flow path from the high-pressure plenum 58 to the flanges 112 can be uninterrupted at least because the compressed working fluid 19 from the high-pressure plenum 58 is not diverted to the plurality of fuel injectors 84 before reaching the flanges 112. For example, such a flow path can advantageously provide improved or increased cooling of the liner 66 as compared to a design that allows some of the compressed working fluid 19 to flow directly from the high-pressure plenum 58 to the plurality of fuel injectors 84 before reaching the cooling flow annulus 90, such as before the compressed working fluid 19 flows through the entire cooling flow annulus 90.

[0038] This written description uses examples to disclose the application, including the best mode, and also to enable any person skilled in the art to practice the application, including making and using any devices or systems and performing any incorporated methods. The patentable scope of the application is defined by the claims, and can include other examples that occur to those skilled in the art. Such other examples are intended to be within the scope of the claims if they have structural elements that do not differ from the literal language of the claims, or if they include equivalent structural elements with insubstantial differences from the literal languages of the claims.

Claims

1. A combustor for a turbine, the combustor coupled to an outer casing of the turbine and in fluid communication with a high pressure plenum within the outer casing, the combustor comprising: a flange coupled to the outer casing; a head end; a liner at least partially defining a hot gas path, the hot gas path including a first combustion zone and a second combustion zone downstream of the first combustion zone; a fuel nozzle at least partially extending through the head end, the fuel nozzle configured to provide a first combustible mixture to the first combustion zone; a flow sleeve circumferentially surrounding at least a portion of the liner, wherein the flow sleeve is spaced apart from the liner to form a cooling flow annulus therebetween, the cooling flow annulus in direct fluid communication with the high pressure plenum, whereby air from the high pressure plenum flows into the cooling flow annulus and from the cooling flow annulus to the head end; a plurality of fuel injectors in fluid communication with the second combustion zone, the plurality of fuel injectors configured to inject a mixture of fuel and air directly into the second combustion zone; and at least one air shroud extending from the flange and surrounding at least one of the plurality of fuel injectors, wherein the flange is disposed proximate the head end and the flange defines at least one passage in fluid communication with the cooling flow annulus for directing a portion of an air flow from the cooling flow annulus to the plurality of fuel injectors, and wherein the portion of the air flow from the cooling flow annulus reaches only the plurality of fuel injectors and the flange after traveling entirely through the cooling flow annulus and not the fuel nozzle, wherein the plurality of fuel injectors are not in direct fluid communication with the high pressure plenum.

2. The combustor of claim 1, wherein the combustor defines a continuous and uninterrupted flow path from the high pressure plenum to the head end.

3. The combustor of claim 2, wherein the continuous and uninterrupted flow path extends from the high pressure plenum to the head end via the cooling flow annulus.

4. The combustor of claim 1, wherein the at least one shroud surrounds the plurality of fuel injectors.

5. The combustor of claim 4, wherein the at least one air shroud comprises a single air shroud surrounding each of the plurality of fuel injectors.

6. The combustor of claim 4, wherein the at least one air shroud comprises a plurality of air shrouds.

7. The combustor of claim 6, wherein each of the plurality of air shrouds surrounds a corresponding one of the plurality of fuel injectors and each of the plurality of fuel injectors is surrounded by a corresponding one of the plurality of air shrouds.

8. A turbine, the turbine comprising: a compressor section extending from an inlet to a discharge, the discharge of the compressor section providing a high pressure air flow directly into a high pressure plenum defined within an outer casing of the turbine; a combustor section, the combustor section comprising: ​ a flange coupled to the outer casing; a head end; a liner at least partially defining a hot gas path, the hot gas path including a first combustion zone and a second combustion zone downstream of the first combustion zone; a fuel nozzle at least partially extending through the head end, the fuel nozzle configured to provide a first combustible mixture to the first combustion zone; a flow sleeve circumferentially surrounding at least a portion of the liner, wherein the flow sleeve is spaced apart from the liner to form a cooling flow annulus therebetween, the cooling flow annulus being in direct fluid communication with the high pressure plenum, whereby air from the high pressure plenum flows into the cooling flow annulus and from the cooling flow annulus to the head end; a plurality of fuel injectors in fluid communication with the second combustion zone, the plurality of fuel injectors configured to inject a mixture of fuel and air directly into the second combustion zone; and at least one air shroud extending from the flange and surrounding at least one of the plurality of fuel injectors, wherein the flange is disposed proximate the head end and the flange defines at least one passage in fluid communication with the cooling flow annulus for directing a portion of an air flow from the cooling flow annulus to the plurality of fuel injectors, and wherein the portion of the air flow from the cooling flow annulus reaches only the plurality of fuel injectors and the flange after traveling entirely through the cooling flow annulus and not the fuel nozzle, wherein the plurality of fuel injectors are not in direct fluid communication with the high pressure plenum; and a turbine section downstream of the combustor section.

9. The turbomachine of claim 8, wherein the combustor section defines a continuous and uninterrupted flow path from the high pressure plenum to the head end.

10. The turbomachine of claim 9, wherein the continuous and uninterrupted flow path extends from the high pressure plenum to the head end via the cooling flow annulus.

11. The turbomachine of claim 8, wherein the at least one air shroud surrounds the plurality of fuel injectors.

12. The turbomachine of claim 11, wherein the at least one air shroud comprises a single air shroud surrounding each of the plurality of fuel injectors.

13. The turbomachine of claim 11, wherein the at least one air shroud comprises a plurality of air shrouds.

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