Gas turbine engine and combustion section for a gas turbine engine

Abstract

A gas turbine engine has a compressor section, a turbine section, a combustion section downstream of the compressor section and upstream of the turbine section, the combustion section including a dome inlet, a combustor outlet fluidly coupled to the turbine section, a liner, and a dome assembly that together at least partially define a combustion chamber extending between the dome inlet and the combustor outlet, a primary fuel injector fluidly coupled to the dome inlet, and a second fuel injector fluidly coupled to the combustion chamber.

Description

[0001] Cross-references to related applications

[0002] This application is a continuation of and claims priority to U.S. Application No. 17 / 587,634, filed January 28, 2022, entitled “Gas Turbine Engine Combustor with Primary and Secondary Fuel Injectors,” the entire contents of which are expressly incorporated herein by reference. Technical Field

[0003] This topic generally relates to a burner with a fuel injector, and more specifically, to a burner with a primary fuel injector and a secondary fuel injector. Background Technology

[0004] A gas turbine engine is driven by a flow of combustion gases through the engine to rotate multiple turbine blades. A combustor may be located within the gas turbine engine and fluidly connected to the turbine through which the combustion gases flow.

[0005] The use of hydrocarbon fuels in the combustors of gas turbine engines is known. Typically, air and fuel are fed into the combustion chamber, where they mix, and the fuel is then burned in the presence of air to produce hot gases. These hot gases are then fed to the turbine, where they are cooled and expanded to generate energy. Byproducts of fuel combustion often include toxins harmful to the environment, such as nitrogen oxides and nitrogen dioxide (collectively known as NO). x ), CO, UHC (such as methane and volatile organic compounds that contribute to the formation of atmospheric ozone), and other oxides, including oxides of sulfur (such as SO2 and SO3).

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

[0007] Standards originating from global air pollution problems regulate nitrogen oxides (NOx) produced by the operation of gas turbine engines. x Emissions include unburned hydrocarbons (UHC) and carbon monoxide (CO). In particular, due to the high burner flame temperature during operation, nitrogen oxides (NOx) are formed within the burner. x The aim is to reduce NO by adjusting the distribution and / or mode within the burner. x Emissions are maintained while still achieving the desired efficiency. Attached Figure Description

[0008] The complete and practical disclosure of this disclosure, including its best mode, is set forth in the specification with reference to the accompanying drawings, for those skilled in the art, wherein:

[0009] Figure 1 This is a schematic diagram of a gas turbine engine.

[0010] Figure 2 The combustion zone of the gas turbine engine is depicted along... Figure 1 Cross-sectional view of line II-II.

[0011] Figure 3 It is in accordance with one aspect of this disclosure along Figure 2 A cross-sectional view of the burner in the combustion zone with primary and secondary fuel injectors, taken from line III-III.

[0012] Figure 4 yes Figure 3 A cross-sectional view, in which some figures have been removed to show the burner during operation according to one aspect of this disclosure.

[0013] Figure 5 It is along Figure 4 The cross-sectional view of line VV shows the streamline extending from the secondary fuel injector to the first set of impeller blades according to one aspect of this disclosure.

[0014] Figure 6 According to one aspect of this disclosure, a fuel system is available from Figure 1 A schematic diagram.

[0015] Figure 7 It is based on one aspect of this disclosure. Figure 6 A flowchart of a method for supplying reforming fuel to an engine. Detailed Implementation

[0016] The aspects of this disclosure described herein relate to a combustor, and more specifically to a combustor having a primary fuel injector and a second fuel injector. For illustrative purposes, this disclosure will be described with respect to a gas turbine engine. However, it should be understood that the aspects of this disclosure described herein are not limited thereto, and the combustor described herein can be implemented in engines, including but not limited to turbojet engines, turboprop engines, turboshaft engines, and turbofan engines. The aspects of this disclosure discussed herein have general applicability in non-aircraft engines with combustors, such as other mobile applications and non-mobile industrial, commercial, and residential applications.

[0017] The term "exemplary" as used herein means "serving as an example, instance, or illustration." Any implementation described herein as "exemplary" is not necessarily to be construed as superior or better than other implementations. Furthermore, unless specifically stated otherwise, all embodiments described herein should be considered exemplary.

[0018] As used herein, the terms “first,” “second,” and “third” are used interchangeably to distinguish one component from another and are not intended to indicate the location or importance of the individual components.

[0019] The terms "front" and "rear" refer to relative positions within a gas turbine engine or vehicle, and specifically to the normal operating posture of the gas turbine engine or vehicle. For example, in the case of a gas turbine engine, "front" refers to the position closer to the engine inlet, while "rear" refers to the position closer to the engine nozzle or exhaust port.

[0020] As used herein, the term "upstream" refers to the direction opposite to the direction of fluid flow, and the term "downstream" refers to the direction in the same direction as the fluid flow. The terms "in front" or "ahead" indicate something in front of it, and "behind" or "behind" indicate something behind it. For example, when used in relation to fluid flow, "in front" or "ahead" can indicate upstream, while "behind" or "behind" can indicate downstream.

[0021] The term "fluid" can refer to either a gas or a liquid. The term "fluid connectivity" refers to the ability of fluids to establish connections between specified areas.

[0022] Furthermore, as used herein, the term "radial" or "radially" refers to a direction away from a common center. For example, in the overall context of a gas turbine engine, radial refers to the direction along a ray extending between the engine's central longitudinal axis and the engine's outer perimeter.

[0023] All directional references (e.g., radial, axial, proximal, distal, up, down, upward, downward, left, right, lateral, front, rear, top, bottom, above, below, vertical, horizontal, clockwise, counterclockwise, upstream, downstream, front, rear, etc.) are used for identification purposes only to aid the reader in understanding this disclosure and do not impose limitations, particularly regarding the location, orientation, or use of the aspects of the disclosure described herein. Unless otherwise stated, connective references (e.g., attachment, connection, link, and linkage) are to be interpreted broadly and may include intermediate structural elements between sets of elements and relative movement between elements. Therefore, a connective reference does not necessarily mean that two elements are directly connected and fixed to each other. Exemplary figures are for illustrative purposes only, and the dimensions, positions, order, and relative sizes reflected in the accompanying figures may vary.

[0024] The singular forms “a,” “one,” and “the” include plural references unless the context clearly specifies otherwise. Furthermore, as used herein, the term “group” or “set” of elements can refer to any number of elements, including only one.

[0025] As used herein and throughout the specification and claims, approximate language is applied to modify any quantitative expression that may allow variation without altering its underlying function. Therefore, values ​​modified by terms such as “approximately,” “about,” “usually,” and “substantially” are not limited to specified precise values. In at least some cases, approximate language may correspond to the precision of the instrument used to measure the value, or the precision of the method or machine used to construct or manufacture the part and / or system. For example, approximate language may refer to a margin of 1%, 2%, 4%, 5%, 10%, 15%, or 20% of the endpoints of a single value, a range of values, and / or a defined range of values. Scope limitations are combined and interchanged herein and throughout the specification and claims; such scopes are identified and include all subscopes contained herein, unless the context or language otherwise indicates. For example, all scopes disclosed herein include endpoints, and endpoints may be combined independently of each other.

[0026] Figure 1 This is a schematic diagram of a gas turbine engine 10. As a non-limiting example, the gas turbine engine 10 may be used within an aircraft. The gas turbine engine 10 may include at least a compressor section 12, a combustion section 14, and a turbine section 16. A drive shaft 18 rotatably connects the compressor section 12 and the turbine section 16 such that rotation of one affects rotation of the other, and defines the axis of rotation or centerline 21 of the gas turbine engine 10.

[0027] Compressor section 12 may include a low-pressure (LP) compressor 22 and a high-pressure (HP) compressor 24 fluidly connected in series. Turbine section 16 may include an LP turbine 26 and an HP turbine 28 fluidly connected in series. Drive shaft 18 may operatively connect the LP compressor 22, HP compressor 24, LP turbine 26, and HP turbine 28 together. Alternatively, drive shaft 18 may include an LP drive shaft (not shown) and an HP drive shaft (not shown). The LP drive shaft may connect the LP compressor 22 to the LP turbine 26, and the HP drive shaft may connect the HP compressor 24 to the HP turbine 28. The LP spool may be defined as a combination of the LP compressor 22, LP turbine 26, and LP drive shaft, such that rotation of the LP turbine 26 can apply a driving force to the LP drive shaft, which in turn can rotate the LP compressor 22. The HP spool may be defined as a combination of the HP compressor 24, HP turbine 28, and HP drive shaft, such that rotation of the HP turbine 28 can apply a driving force to the HP drive shaft, which in turn can rotate the HP compressor 24.

[0028] Compressor section 12 may include multiple axially spaced stages. Each stage includes a set of circumferentially spaced rotating blades and a set of circumferentially spaced stationary blades. Compressor blades for a stage of compressor section 12 may be mounted to a disc that is mounted to drive shaft 18. Each set of blades for a given stage may have its own disc. The blades of compressor section 12 may be mounted to a housing that extends circumferentially around gas turbine engine 10. It should be understood that the representation of compressor section 12 is merely illustrative and any number of stages may be possible. Furthermore, it is conceivable that any other number of components may be present within compressor section 12.

[0029] Similar to compressor section 12, turbine section 16 may include multiple axially spaced stages, each stage having a set of circumferentially spaced rotating blades and a set of circumferentially spaced stationary blades. Turbine blades for one stage of turbine section 16 may be mounted to a disc, which is mounted to drive shaft 18. Each set of blades for a given stage may have its own disc. The blades of the turbine section may be circumferentially mounted to the housing. It should be noted that any number of blades, blades, and turbine stages are possible, as the turbine section shown is merely schematic. Furthermore, it is conceivable that any other number of components may be present within turbine section 16.

[0030] Combustion section 14 may be arranged in series between compressor section 12 and turbine section 16. Combustion section 14 may be fluidly connected to at least a portion of compressor section 12 and turbine section 16, such that combustion section 14 at least partially fluidly connects compressor section 12 to turbine section 16. As a non-limiting example, combustion section 14 may be fluidly connected to HP compressor 24 at its upstream end and to HP turbine at its downstream end 28.

[0031] During operation of the gas turbine engine 10, ambient or atmospheric air is drawn into the compressor section 12 via a fan (not shown) upstream of the compressor section 12, where the air is compressed and defined as pressurized air. This pressurized air can then flow into the combustion section 14, where it mixes with fuel and is ignited to produce combustion gases. The HP turbine 28 extracts some work from these combustion gases, which drives the high-pressure compressor 24. The combustion gases are discharged into the LP turbine 26, which extracts additional work to drive the LP compressor 22, and the exhaust gases are finally discharged from the gas turbine engine 10 via an exhaust section (not shown) downstream of the turbine section 16. The drive of the LP turbine 26 drives the LP spool to rotate the fan (not shown) and the LP compressor 22. The pressurized airflow and combustion gases together define the working airflow flowing through the fan, compressor section 12, combustion section 14, and turbine section 16 of the gas turbine engine 10.

[0032] Figure 2 A cross-sectional view along line II-II is depicted for combustion section 14. Combustion section 14 may include an annular arrangement of primary fuel injectors 30 arranged around the centerline 21 of the gas turbine engine 10. Each of the primary fuel injectors 30 may be connected to a combustor 32. It should be understood that the annular arrangement of fuel injectors may be one or more fuel injectors, and the one or more primary fuel injectors 30 may have different characteristics. Depending on the type of engine in which the combustor 32 is located, the combustor 32 may have a canister, canister ring, or annular arrangement. In a non-limiting example, an annular arrangement is shown and arranged within a housing 34. Combustor 32 is defined by a combustor liner 36, which includes an outer combustor liner 38 and an inner combustor liner 40 that are concentric with each other and annular around the engine centerline 21. A dome assembly 42 including a dome wall 44, together with the inner combustor liner 36, may define an annular combustion chamber 46 around the engine centerline 21. A first set of dilution openings 48 may be located in the dome wall 44. At least one primary fuel injector 30 (shown as a plurality of primary fuel injectors arranged in a ring around the engine centerline 21) is fluidly coupled to the combustion chamber 46. The compressed air passage 50 may be defined at least partially by both the burner liner 36 and the housing 34.

[0033] Figure 3 Depicting along Figure 1 The cross-sectional view taken by line III-III shows combustion section 14. A second set of dilution openings 52 may be located in the burner liner 36, connecting the compressed air passage 50 and the burner 32. It should be understood that any number of dilution openings may be part of the first set of dilution openings 48 and / or the second set of dilution openings 52.

[0034] The primary fuel injector 30 can be connected upstream of the primary fuel / air mixture outlet 54 defined by the flared cone 56 to and disposed within the dome assembly 42. The primary fuel / air mixture outlet 54 can define a dome inlet such that the primary fuel / air mixture outlet and the dome inlet are the same, referred to herein as the dome inlet 54. The primary fuel injector 30 includes a fuel inlet 60 and a fuel passage 62, the fuel inlet 60 being adapted to receive a fuel (F) flow, and the fuel passage 62 extending between the fuel inlet 60 and the primary fuel / air mixture outlet 54. A first set of dilution openings 48 can define a swirler disposed at the dome inlet 54 to swirl incoming compressed air (C) to the vicinity of the fuel / air mixture exiting the primary fuel injector 30. The swirler provides both a homogeneous mixture of air and fuel entering the combustor 32 and aids in the cooling of the combustor liner 36.

[0035] Both the outer burner liner 38 and the inner burner liner 40 can have an outer surface 64 and an inner surface 66 that at least partially define the combustion chamber 46. The burner liner 36 can be made from a single continuous integral portion or from multiple integral portions assembled together to define the outer burner liner 38 and the inner burner liner 40. As a non-limiting example, the outer surface 64 can define a first part of the burner liner 36, while the inner surface 66 can define a second part of the burner liner 36 when assembled together to form the burner liner 36. Further contemplation suggests that the burner liner 36 can be any type of burner liner 36, including but not limited to double-walled liners or tile liners. The igniter 68 can be disposed at any suitable location, including at the burner liner 36, and is fluidly connected to the combustion chamber 46 at any location downstream of the dome inlet 54 by a non-limiting example.

[0036] Structural element 70 may extend into combustion chamber 46 downstream of primary fuel injector 30. Structural element 70 may be any structural element; as a non-limiting example, it may be a strut. Structural element 70 may include internal cooling channels and mechanisms and / or high-temperature materials as needed to allow the structural element to be positioned in the path of combustion gases (G1) generated in combustion chamber 46. Structural element 70 may further extend between housing 34 and outer burner liner 38. Structural element 70 may define a housing for fuel passage 72 for second fuel injector 74. Fuel passage 72 may include fuel inlet 76, which may be adapted to receive another fuel (F) flow. A set of openings 78 may define a second fuel / air mixture outlet 79 for air and fuel exiting second fuel injector 74. This set of openings 78 may be a single opening or a plurality of openings radially spaced along structural element 70 as shown. It is further envisioned that this set of openings 78 is arranged along the leading edge of structural element 70 (referred to herein as structural element leading edge 92). It can be further envisioned that the set of openings 78 is located near the leading edge 92 of the structural element. By being located near the leading edge 92 of the structural element, the set of openings 78 is within 5% of the total axial length of the structural element.

[0037] Combustor outlet 80 fluidly connects combustor 32 to turbine section 16. Around engine centerline 21 ( Figure 2 A circumferentially spaced set of blades 82 may define a nozzle 84 for the turbine section 16. A set of blades 86 is circumferentially spaced around the engine centerline 21 and located downstream of the set of blades 82. The set of blades 82 and the set of blades 86 together define the first stage 58 of the turbine section 16. In one aspect disclosed herein, the set of blades 82 is uncooled. By defining the "uncooled" blades of the set 82, there are no cooling chambers, cooling holes, or other cooling features fluidly connected to the cooling air. It is further conceivable that either the set of blades 82 and / or the set of blades 86 is cooled and / or made of ceramic material.

[0038] Structural element 70 divides combustion chamber 46 into a primary region 88 having a first volume (V1) and a secondary region 90 having a second volume (V2). The primary region 88 may have a first length (L1), axially measured from the dome wall 44 to the trailing edge of the structural element (referred to herein as the trailing edge 94 of the structural element). The secondary region 90 may have a second length (L2), axially measured from the trailing edge 94 of the structural element to the leading edge 96 of a set of blades 82. In one aspect disclosed herein, the total length (L) of burner 32 is less than a typical combustion length. The first length (L1) is less than or equal to twice the second length (L2), such that the primary region 88 is smaller than a typical primary region of a burner.

[0039] Turning Figure 4 , from Figure 3The same schematic diagram, with some figures omitted for clarity, shows the burner 32 during operation. Compressed air (C) can flow from the compressor section 12 to the burner 32 through the dome assembly 42. Compressed air (C) can also flow from the compressor section 12 to the burner 32 through the compressed air passage 50. A second set of dilution openings 52 in the burner liner 36 allows some compressed air (C) to flow from the compressed air passage 50 to the combustion chamber 46. Another portion of the compressed air (C) can flow through the first set of dilution openings 48. The compressed air (C) received in the burner 32 via the dilution openings 48, 52 can create swirling currents to mix the fuel / air mixture, to cool the burner liner 36, and / or to provide dilution within the combustion chamber 46.

[0040] The first volume (V1) of combustion chamber 46 is sized and derived by the high re-ignition Darmco number (Da) necessary to provide sufficient temperature rise to restart gas turbine engine 10. The Darmco number (Da) is the ratio of reaction rate to transport rate. In other words, a Darmco number (Da) greater than one is associated with a combustion rate higher than the rate of movement of combustion gases through combustor 32. For combustion chamber 46, the Darmco number is greater than one (Da>1). At least a portion of compressed air (C) may flow through primary fuel injector 30. A first amount of fuel (F1) may be injected at primary fuel injector 30 and mixed with compressed air (C). The first amount of fuel (F1) may be hydrogen fuel, or any hydrocarbon fuel, including jet fuel. Upon entering combustor 32, the mixture is ignited within combustion chamber 46 by one or more igniters 68 to produce combustion gases (G1). Combustion gases (G1) may be mixed as previously described to define a first portion of combustion gases (G1) with a temperature rise to a first amount (T1). The first volume (V1) is sized to allow the first portion of the combustion gas (G1) to reach the maximum temperature. When a set of blades 82 defines an uncooled nozzle 84, a maximum temperature of 1800°F (982°C) applies. The first volume (V1) mixes all the air in the combustion gas (G1) and produces the temperature limit of 1800°F (982°C). When a set of blades 82 defines a cooled nozzle 84, or when the set of blades 82 is made of a non-metallic material, this maximum temperature can be increased. For example, ceramic materials can provide a limit of 2700°F (1482°C). Therefore, for an uncooled nozzle 84, the maximum temperature can be 1800°F (982°C), while when advanced non-metallic or cooling technologies are used on the nozzle 84, the maximum temperature can be 2700°F (1482°C). When the turbine nozzle 84 is actively cooled, then the maximum temperature may approach 3400°F (1871°C). Since bleed air is used for cooling, increasing the maximum temperature would reduce the performance of the gas turbine engine 10. Therefore, providing an uncooled nozzle 84 and combining this with dimensionalizing the first volume (V1) to reduce the maximum temperature can increase engine performance.

[0041] A second quantity of fuel (F2), such as hydrogen fuel as a non-limiting example, can be injected at a second fuel injector 74 and mixed with another portion of compressed air (C). This mixture can exit as a fuel / air mixture from a set of openings 78, which can be ignited to produce a second portion of combustion gases (G2), which can flow through burner outlet 80 and enter turbine section 16 via nozzle 84. The temperature of the second portion of combustion gases (G2) can be increased to a second quantity (T2) greater than the first quantity (T1). The second fuel injector 74 can be used to provide the temperature rise required for takeoff and cruise.

[0042] Turning Figure 5 It shows along Figure 4 A schematic diagram of the secondary region 90 intercepted by line VV. The structural element 70 may have an airfoil shape 98 and extend in a downstream direction between the leading edge 92 and the trailing edge 94 of the structural element. Further, the airfoil shape 98 may have a first side 100 extending between the leading edge 92 and the trailing edge 94 of the structural element and a second side 102 opposite to the first side 100. Each blade in a set of blades 82 may also have an airfoil shape 104 and extend downstream between the leading edge 96 and the trailing edge 106.

[0043] A fuel / air mixture (F / A) exits through a set of openings 78 to define a second portion of combustion gas (G2). The fuel / air mixture (F / A) may exit from one or both sides 100, 102 of the structural element 70. The second portion of combustion gas (G2) may be elevated to a second quantity (T2) exceeding 1800°F (982°C) to 3400°F (1870°C). The second portion of combustion gas (G2) passes through nozzle 84 at a temperature equal to or between 1800°F (982°C) and 3400°F (1870°C). The location of the set of openings 78, along with the angle and positioning of the set of blades 82, is oriented to form a streamline 108 located at the center of nozzle 84 along which combustion gas (G2) may flow. This prevents the high-temperature combustion gas (G2) from contacting the set of blades 82, which allows the use of uncooled blades as described herein. All combustion occurs between the turbine blades 82, thus providing sufficient temperature rise to the turbine blades 86.

[0044] Figure 6 A fuel system 118 for the engine 10 described herein is shown. The fuel system 118 includes a main fuel line 120 fluidly connected to a primary fuel line 122 and a secondary fuel line 126. The secondary fuel line 126 fluidly connects the main fuel line 120 to a second fuel injector 74. A liquid fuel supply unit 128 is fluidly connected to the main fuel line 120. The liquid fuel supply unit 128 can be filled with hydrogen, ethylene, methane, or any other suitable hydrocarbon liquid fuel.

[0045] Turning Figure 7 In one aspect disclosed herein, a method for controlling nitrogen oxides or NO present in combustion gases (G) within burner 32. xThe method includes injecting a fuel / air mixture into a primary region 88 at block 202 as described herein. Further, the method may include injecting the fuel / air mixture into a secondary region 90 via a second fuel injector 74 at block 204. At block 206, the method may include burning a hydrocarbon fuel in the secondary region 90 to produce a second portion of combustion gas (G2). At 208, the second portion of combustion gas (G2) is allowed to flow into at least a portion of the first stage 58 of the turbine section 16. Further, the method may include controlling the flow of the second portion of combustion gas (G2) via nozzle 84 to prevent the second portion of combustion gas (G2) from contacting a set of turbine blades 82.

[0046] Benefits related to the disclosure in this article include, but are not limited to, reducing NO x Emissions and improved engine efficiency. Delayed injection of hydrocarbon fuel via the second fuel injector 74 reduces NO. x Emissions. Furthermore, the use of uncooled impellers 82 in the first stage 58 reduces weight while improving engine efficiency and meeting all engine operability requirements. Moreover, the smaller temperature rise in the primary region 88 relative to the typical temperature rise of the combustor enables lower NO emissions. x This reduces emissions, improves engine operability, and enhances high-altitude restart capability. Furthermore, the overall length (L) of the combustor 32 is shorter than that of a typical combustor. This reduction in the length (L) of the primary region results in a reduction in the overall engine length, thereby enabling weight reduction.

[0047] Typically, using hydrogen as fuel can significantly reduce burner volume because the fuel is already in a gaseous state. Evaporation time is short, mixing is rapid, and the reaction rate is high. Therefore, engine length and weight can be reduced by decreasing burner volume and length.

[0048] It should be understood that the dilution openings described herein, as shown in the figures, are exemplary. Dilution openings can be organized in numerous different ways and, as non-limiting examples, can include ribs, pin arrays, circuits, sub-circuits, membrane openings, gas chambers, meshes, and turbulence generators of any shape or size. Dilution openings can include other flow enhancement devices, as non-limiting examples, including small openings located behind the dilution openings. It is further conceivable that dilution openings can be part of an assembly of dilution openings. It is also conceivable that dilution openings can be separate from and outside of an assembly of cooling openings positioned along the burner liner.

[0049] Although a gas turbine engine has been described, it should be understood that the combustor described herein can be used in applications with NO emissions. xAny engine with a combustor. It should be understood that the application of the aspects of this disclosure discussed herein is applicable to engines having a propeller section or a fan and supercharger section, as well as turbojet engines and turbine engines.

[0050] Within the scope not yet described, various features and structures of the various embodiments may be combined or substituted for each other as needed. The fact that a feature is not shown in all embodiments does not mean it cannot be shown, but is done for the sake of brevity. Therefore, various features of different embodiments may be mixed and matched as needed to form new embodiments, regardless of whether the new embodiments are explicitly described. All combinations or permutations of the features described herein are covered by this disclosure.

[0051] This written description uses examples to illustrate aspects of the disclosure described herein, including best practices, and also enables any person skilled in the art to practice aspects of this disclosure, including making and using any apparatus or system and performing any incorporated methods. The patent scope of each aspect of this disclosure is defined by the claims and may include other examples that would occur to a person skilled in the art. Such other examples are intended to be within the scope of the claims if they have structural elements that are not indistinguishable from the literal language of the claims, or if they include equivalent structural elements that are not substantially different from the literal language of the claims.

[0052] Further aspects are provided by the following topics:

[0053] A gas turbine engine includes: a compressor section; a turbine section; and a combustion section located downstream of the compressor section and upstream of the turbine section, the combustion section including: a dome inlet; a combustor outlet fluidly coupled to the turbine section; a liner and a dome assembly together defining at least partially a combustion chamber extending between the dome inlet and the combustor outlet; a primary fuel injector fluidly coupled to the dome inlet; and a second fuel injector fluidly coupled to the combustion chamber at a location downstream of the dome inlet and upstream of the combustor outlet.

[0054] The gas turbine engine according to any of the preceding clauses further includes a structural element that defines a fuel passage for the second fuel injector and extends through the liner.

[0055] In any of the preceding clauses of the gas turbine engine, the structural element extends at least partially into the combustion chamber.

[0056] The gas turbine engine according to any of the preceding clauses further includes a set of openings that are radially spaced along the structural element and define the outlet of the second fuel injector.

[0057] The gas turbine engine according to any of the preceding clauses, wherein the structural element extends axially between the leading edge and the trailing edge.

[0058] In any of the preceding clauses of the gas turbine engine, wherein the set of openings is located near the leading edge.

[0059] The gas turbine engine according to any of the preceding clauses further includes a structural element that defines the second fuel injector and extends at least partially into the combustion chamber, the structural element dividing the combustion chamber into a primary region downstream of the primary fuel injector and a secondary region downstream of the second fuel injector.

[0060] The gas turbine engine according to any of the preceding clauses further includes a set of blades located downstream of the combustor outlet, wherein a first length is less than or equal to twice the second length, the first length defining the primary region and axially measured between the dome inlet and the leading edge of the set of blades, and the second length defining the secondary region and axially measured between the trailing edge of the structural element and the leading edge of the set of blades.

[0061] According to any of the preceding clauses, the first length is minimized to allow for a high re-ignition Darmcol number to generate sufficient temperature rise.

[0062] According to any of the preceding clauses, the gas turbine engine wherein the sufficient temperature rises to a first amount between 1800°F and 2700°F.

[0063] According to any of the preceding clauses of the gas turbine engine, the second amount of temperature rise in the secondary region is equal to or between 1800°F and 3400°F.

[0064] According to any of the preceding clauses, in a gas turbine engine, the structural elements and the set of blades are oriented to form streamlines that prevent combustion gases formed in the secondary region from contacting the set of blades.

[0065] In any of the preceding clauses, the gas turbine engine wherein the set of blades is uncooled.

[0066] In any of the preceding clauses of the gas turbine engine, the size of the primary region depends on a Darmquerel number greater than or equal to 1.

[0067] According to any of the preceding clauses, the gas turbine engine wherein the second fuel injector is supplied with hydrogen fuel.

[0068] According to any of the preceding clauses, the gas turbine engine wherein the primary fuel injector is supplied with hydrogen fuel.

[0069] A combustion section for a gas turbine engine, the combustion section comprising: a dome inlet; a combustor outlet fluidly coupled to a turbine section of the gas turbine engine; a liner and a dome assembly together defining at least partially a combustion chamber extending between the dome inlet and the combustor outlet; a primary fuel injector fluidly coupled to the dome inlet; and a second fuel injector fluidly coupled to the combustion chamber at a location downstream of the dome inlet and upstream of the combustor outlet.

[0070] The combustion zone according to any of the preceding clauses further includes a structural element that defines the second fuel injector and extends at least partially into the combustion chamber, the structural element dividing the combustion chamber into a primary region downstream of the primary fuel injector and a secondary region downstream of the second fuel injector.

[0071] The combustion section according to any of the preceding clauses further includes a set of blades located downstream of the burner outlet, wherein a first length is less than or equal to twice the second length, the first length defining the primary region and measured axially between the dome inlet and the leading edge of the set of blades, and the second length defining the secondary region and measured axially between the trailing edge of the structural element and the leading edge of the set of blades.

[0072] According to any of the preceding clauses, in the combustion section, the structural element and the set of blades are oriented to form streamlines that prevent combustion gases formed in the secondary region from contacting the set of blades.

Claims

1. A gas turbine engine characterized by, Comprising: a compressor section; a turbine section; and a combustion section downstream of the compressor section and upstream of the turbine section, the combustion section comprising: a dome inlet; a combustor outlet fluidly coupled to the turbine section; a liner and a dome assembly together at least partially defining a combustion chamber extending between the dome inlet and the combustor outlet, a first set of dilution openings in a dome wall of the dome assembly, a second set of dilution openings in the liner, and compressed air received via the first and second sets of dilution openings creating a swirl in the combustor; a primary fuel injector fluidly coupled to the dome inlet; and a secondary fuel injector fluidly coupled to the combustion chamber at a location downstream of the dome inlet and upstream of the combustor outlet. Further comprising a structural element defining a fuel passage for the secondary fuel injector and extending through the liner.

2. The gas turbine engine of claim 1, wherein, Wherein the structural element extends at least partially into the combustion chamber.

3. The gas turbine engine of claim 2, wherein, Further comprising a set of openings radially spaced along the structural element and defining an outlet of the secondary fuel injector.

4. The gas turbine engine of claim 3, wherein, Wherein the structural element extends axially between a leading edge and a trailing edge.

5. The gas turbine engine of claim 4, wherein, Wherein the set of openings is located proximate the leading edge.

6. The gas turbine engine of claim 5, wherein, Further comprising a structural element defining the secondary fuel injector and extending at least partially into the combustion chamber, the structural element separating the combustion chamber into a primary region downstream of the primary fuel injector and a secondary region downstream of the secondary fuel injector.

7. The gas turbine engine of claim 1, wherein, Further comprising a set of vanes downstream of the combustor outlet, wherein a first length defining the primary region and measured axially between the dome inlet and a leading edge of the set of vanes is less than or equal to twice a second length defining the secondary region and measured axially between a trailing edge of the structural element and the leading edge of the set of vanes.

8. The gas turbine engine of claim 7, wherein, Wherein the first length is minimized to allow for a high primary Damkohler number for generating a sufficient temperature rise.

9. The gas turbine engine of claim 8, wherein, Wherein the sufficient temperature rise is to a first amount between 1800°F and 2700°F.

10. The gas turbine engine of claim 9, wherein, Wherein a second amount of temperature rise in the secondary region is greater than or equal to 1800°F and less than or equal to 3400°F.

11. The gas turbine engine of claim 10, wherein, Wherein the structural element and the set of vanes are oriented to form a flow line that prevents combustion gases formed in the secondary region from contacting the set of vanes.

12. The gas turbine engine of claim 11, wherein, Wherein the set of vanes is uncooled.

13. The gas turbine engine of claim 12, wherein, Wherein a size of the primary region is dependent on a Damkohler number greater than or equal to 1.

14. The gas turbine engine of any one of claims 7-13, wherein, Wherein the secondary fuel injector is supplied with hydrogen fuel.

15. The gas turbine engine of claim 1 or 7, wherein, Wherein the primary fuel injector is supplied with hydrogen fuel.

16. The gas turbine engine of claim 15, wherein, The combustion section comprising:

17. A combustion section for a gas turbine engine characterized by, a dome inlet; a combustor outlet fluidly coupled to a turbine section of the gas turbine engine; ​ a liner and a dome assembly together at least partially defining a combustion chamber extending between the dome inlet and the combustor outlet, a first set of dilution openings located in a dome wall of the dome assembly, a second set of dilution openings located in the liner, and compressed air received via the first and second sets of dilution openings creating a swirl in the combustor; a primary fuel injector fluidly coupled to the dome inlet; and a second fuel injector fluidly coupled to the combustion chamber at a location downstream of the dome inlet and upstream of the combustor outlet.

18. The combustion section of claim 17, wherein, further comprising a structural element defining the second fuel injector and extending at least partially into the combustion chamber, the structural element separating the combustion chamber into a primary zone downstream of the primary fuel injector and a secondary zone downstream of the second fuel injector.

19. The combustion section of claim 18, wherein, further comprising a set of vanes downstream of the combustor outlet, wherein a first length defining the primary zone and measured axially between the dome inlet and a leading edge of the set of vanes is less than or equal to twice a second length defining the secondary zone and measured axially between a trailing edge of the structural element and the leading edge of the set of vanes.

20. The combustion section of claim 19, wherein, wherein the structural element and the set of vanes are oriented to form a flow line that prevents combustion gases formed in the secondary zone from contacting the set of vanes.

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