System and method for fuel and inert gas injection in the turbine section of a gas turbine engine

Abstract

The present invention provides a system and method for fuel and inert gas injection in the turbine section of a gas turbine engine. [Solution] At least one component includes a fuel port 56 formed in the wall 111 of at least one component, the fuel port 56 being configured to inject fuel 52 into the chamber of the turbine section. At least one component also includes an inert gas port 57 formed in the wall upstream of the fuel port. The inert gas port is configured to inject an inert gas 54 so as to flow at least partially around the fuel injected in a buffer region downstream of the fuel port. The presence of the inert gas in the buffer region can suppress ignition of the fuel near the surface of the wall downstream of the fuel port (e.g., flame anchoring) and / or help lift the flame away from the ignited fuel into the hot gas flow path.

Description

Technical Field

[0001] The subject matter disclosed herein relates to a gas turbine engine having fuel injection in a turbine section. The invention claimed herein relates to the subject matter recited in the claims.

Background Art

[0002] A gas turbine engine generally includes a compressor section, a combustor section, and a turbine section (e.g., an expansion turbine). The turbine section can include one or more turbine stages each having a plurality of turbine blades. The combustion gas flow expands through the turbine section, driving the rotation of the turbine blades in each of the one or more stages. In some configurations, the turbine section can include a plurality of fuel injection ports for injecting fuel for combustion in the turbine section, thereby adding additional heat to the turbine section and resulting in an isothermal expansion of the gas through the one or more turbine stages. Such isothermal expansion can increase the efficiency and performance of the gas turbine.

[0003] Unfortunately, certain conditions in the turbine section can create a vulnerability to flameholding near the wall of the turbine section at the plurality of fuel injection ports. Therefore, it is necessary to reduce the possibility of flameholding near the wall of the turbine section.

Summary of the Invention

[0004] The invention claimed herein relates to the subject matter recited in the appended claims. Some embodiments within the scope equivalent to the invention recited in the original claims at the time of filing are summarized below. These embodiments are not intended to limit the scope of the claimed invention; rather, these embodiments are only intended to provide a brief overview of possible forms of the invention. In fact, the invention can encompass various forms, which may be the same as or different from the embodiments described below.

[0005] In one embodiment, the system includes at least one component of a turbine section. The at least one component includes a fuel port formed in the wall of the at least one component. The fuel port is configured to inject fuel into the chamber of the turbine section. The at least one component also includes an inert gas port formed in the wall upstream of the fuel port. The inert gas port is configured to inject an inert gas so that it flows at least partially around the fuel injected from the fuel port into a buffer region downstream.

[0006] In another embodiment, the system includes a gas turbine engine having a turbine section having circumferentially spaced airfoils. At least one component of the turbine section includes a fuel injection port formed in the wall of at least one component. The fuel injection port is configured to guide fuel between the successive circumferentially spaced airfoils of the turbine section. At least one component also includes an inert gas injection port formed in the wall upstream of the fuel port. The inert gas port is configured to guide an inert gas so as to flow at least partially around the fuel from the fuel port to a downstream buffer region.

[0007] In another embodiment, the system comprises at least one turbine section, the turbine section having a flow path through the turbine section, the flow path defining an upstream direction and a downstream direction. The turbine section further comprises a fuel port formed in the wall of the turbine section, the fuel port configured to inject fuel into the turbine section. Furthermore, an inert gas port is formed in the wall upstream of the fuel port, the inert gas port configured to inject an inert gas so as to flow at least partially around the fuel injected from the fuel port into a buffer region downstream.

[0008] In another embodiment, the method includes the step of injecting fuel into the turbine section of a gas turbine engine from a fuel port formed in the wall of at least one component of the turbine section. The method also includes the step of injecting an inert gas from an inert gas port located upstream of the fuel port toward the injected fuel. The method also includes the step of causing the injected inert gas to occupy at least a portion of a buffer area downstream of the fuel port. In another view, a method for injecting fuel into the turbine section of a gas turbine engine includes the steps of injecting fuel into the turbine section of a gas turbine engine from a fuel port formed in the wall of the turbine section, injecting an inert gas from an inert gas port located upstream of the fuel port toward the injected fuel along the flow direction through the flow path of the turbine section, and causing the injected inert gas to flow at least partially around the injected fuel in a buffer area downstream of the fuel port.

[0009] These and other features, aspects, and advantages of this disclosure will be better understood by reading the following detailed description with reference to the accompanying drawings. In the accompanying drawings, similar reference numerals represent similar parts throughout the drawings. [Brief explanation of the drawing]

[0010] [Figure 1] This is a schematic flowchart of one embodiment of a gas turbine engine having a turbine section in which a fuel injection system is installed, the fuel injection system being configured to suppress fuel ignition near the surface of the wall of the turbine section. [Figure 2] This is a cross-sectional side view of one embodiment of the gas turbine engine shown in Figure 1, cut through the longitudinal axis. [Figure 3] Figure 2 is a schematic top view of one embodiment of the turbine section of a gas turbine engine, further illustrating the locations of the fuel ports and inert gas ports of the fuel injection system. [Figure 4]Figure 2 is a schematic perspective view of one embodiment of the fuel injection system, further showing the fuel port and inert gas port of the fuel injection system. [Figure 5] Figure 2 is a schematic top view of one embodiment of the fuel injection system, further showing the fuel port and inert gas port. [Figure 6] Figure 2 is a cross-sectional side view of one embodiment of the fuel injection system, further showing the flow path of fuel injected through the fuel port and the inert gas flow path of inert gas injected through the inert gas port. [Figure 7] Figure 6 is a schematic perspective view of one embodiment of the flow path. [Figure 8] This is a cross-sectional plot showing one embodiment of the velocity profile of the flow path in Figure 6. [Figure 9] This plot is orthogonal to the plot in Figure 8, which shows one embodiment of the flow path velocity profile. [Figure 10] Figure 2 is a flowchart illustrating an exemplary process for operating the fuel injection system. [Modes for carrying out the invention]

[0011] One or more specific embodiments of this disclosure are described below. Not all features of the actual embodiments are described herein in order to provide a concise description of these embodiments. It should be understood that, as with any engineering or design project, the development of such actual embodiments requires numerous decisions specific to the embodiment, such as compliance with system-related and business-related constraints, in order to achieve the developer's specific goals, although these constraints may differ from embodiment to embodiment. Furthermore, it should be understood that such development efforts, while complex and time-consuming, are still routine design, fabrication, and manufacturing tasks for those skilled in the art who are interested in this disclosure.

[0012] When introducing elements of the various embodiments of this disclosure, the articles “a, an,” “the,” and “said” are intended to mean that there is one or more of those elements. The terms “equip,” “include,” and “have” are intended to be inclusive and mean that there may be further elements other than those listed.

[0013] As will be described in more detail below, the disclosed embodiment includes a fluid injection system in the turbine section of a gas turbine engine, the fluid injection system including a fuel injection port for injecting fuel for combustion in the turbine section and an inert gas injection port that helps reduce the possibility of flame holding near the walls of the turbine section. For example, the fuel injection port is formed in the wall of the turbine section. In addition, the inert gas injection port is located in the wall upstream of the fuel injection port. The inert gas injection port injects an inert gas at a sharp angle (relative to the wall) toward the injected fuel. The injected inert gas remains near the surface of the wall as it flows toward the injected fuel (e.g., the fuel jet region). Once the injected inert gas reaches the injected fuel (e.g., the fuel jet), the injected inert gas wraps around the side of the fuel jet region and flows into a buffer region (e.g., a low-pressure region) located adjacent to the downstream of the fuel injection port. In response to the inert gas flowing into this region, the mass fraction of the inert gas in the buffer region increases, thereby suppressing the possibility of ignition of the injected fuel near the wall surface (this is known as flame anchoring or flame holding).

[0014] With the above in mind, Figure 1 is a schematic flowchart of one embodiment of a gas turbine system 10 having a gas turbine engine 12 with a fuel injection system 11. As will be described in detail below, the fuel injection system 11 includes a fluid supply source 14 (see Figure 2) configured to supply fuel and inert gas to the fuel injection system 11. In certain embodiments, the gas turbine system 10 may include an aircraft, a locomotive, a power generation system, or a combination thereof. In series flow order, the illustrated gas turbine engine 12 includes an intake section 16, a compressor or compressor section 18, a combustor or combustor section 20, a turbine or turbine section 22 (i.e., an expansion turbine), and an exhaust section 24. The turbine section 22 is coupled to the compressor 18 via a shaft 26.

[0015] As indicated by the arrows, air enters the gas turbine engine 12 through the intake section 16 and can flow into the compressor 18, which compresses the air before it enters the combustor section 20. The illustrated combustor section 20 includes a combustor housing 28 arranged concentrically or annularly around a shaft 26 between the compressor 18 and the turbine section 22. Compressed air from the compressor 18 enters the combustor section 20, where the compressed air and fuel are mixed and burned in the combustor to drive the turbine section 22. From the combustor section 20, the hot combustion gases flow through the turbine section 22 and drive the compressor 18 via the shaft 26. For example, the combustion gases can add prime mover power to the turbine blades in the turbine section 22 and rotate the shaft 26.

[0016] In the illustrated embodiment, the fuel injection system 11 is located in the turbine section 22. In the illustrated embodiment, the fuel injection system 11 may include one or more combined or separate fuel injection systems configured to inject fuel into the combustor section 20 and the turbine section 22. For example, as shown in Figure 1, the fuel injection system 11 injects fuel into the turbine section 22 to provide additional heat to the turbine section 22, while simultaneously injecting an inert gas to reduce the possibility of flame holding along the surface in the turbine section 22. After flowing through the turbine section 22, the hot combustion gases can exit the gas turbine engine 12 through the exhaust section 24. The gas turbine engine 12 can be described with respect to the longitudinal or axis 32 (e.g., axial), radial or axis 34, and circumferential or axis 36.

[0017] Figure 2 is a cross-sectional side view of one embodiment of the gas turbine engine 12 of Figure 1, cut through the longitudinal axis 32, showing one embodiment of a fuel injection system 11 (e.g., separate or integrated fuel injection systems) coupled to a combustor section 20 and a turbine section 22. As described above with respect to Figure 1, air can enter the gas turbine engine 12 through the intake section 16 and be compressed by the compressor 18. The compressed air from the compressor 18 may then be led to the combustor section 20, where it can be mixed with fuel from the fuel injection system 11. The combustor section 20 includes one or more combustors 38. In certain embodiments, the gas turbine engine 12 may include a plurality of combustors 38 arranged in an annular arrangement (e.g., a so-called can-annular combustor arrangement). Furthermore, each combustor 38 may include a plurality of fuel nozzles 40 mounted at or near the head end of each combustor 38 in an annular or other arrangement. During operation, the fuel nozzles 40 can inject a fuel-air mixture into the combustor 38 in appropriate ratios for optimal combustion, emissions, fuel consumption, and power output.

[0018] Within the combustor section 20, the fuel-air mixture can burn to produce high-temperature, pressurized combustion gases. After combustion, the high-temperature, pressurized combustion gases can flow through the transition piece 42 and exit from the combustor section 20 to the turbine section 22. Within the turbine section 22, the pressurized combustion gases extend radially within the turbine section 22, causing the blades 44 (e.g., rotating turbine blades) positioned between the vanes 46 (e.g., stationary turbine vanes) to rotate the shaft 26 before exiting as exhaust gas through the exhaust section 24.

[0019] In certain embodiments, the turbine section 22 may include one or more turbine stages (e.g., one, two, three, four, or more) arranged at different axial positions along the longitudinal axis 32 of the turbine section 22. Each turbine stage may have a plurality of blades 44 (e.g., rotary turbine blades) spaced apart from each other in a circumferential arrangement (e.g., spaced apart circumferentially) around the longitudinal axis 32 of the turbine section 22 at a common axial position, and the blades 44 may be coupled to the central turbine rotor or shaft of the turbine section 22. Similarly, each turbine stage may have a plurality of vanes 46 (e.g., stationary turbine vanes) spaced apart from each other in a circumferential arrangement (e.g., spaced apart circumferentially) around the longitudinal axis 32 of the turbine section 22 at a common axial position offset from the blades 44, and the vanes 46 may be coupled to the outer casing or wall of the turbine section 22. In certain embodiments, each of the blades 44 and vanes 46 may include a structure with an airfoil shape that is oriented or extends radially (e.g., radial airfoil).

[0020] In the illustrated embodiment, the fuel injection system 11 includes a fluid source 14 having a fuel source 48 and an inert gas source 50. As shown, the fluid source 14 (e.g., the fuel source 48 and the inert gas source 50) is fluidly coupled to the combustor section 20 and the turbine section 22 of the gas turbine engine 22. However, in some embodiments, separate fuel injection systems 11 can be coupled to the combustor section 20 and the turbine section 22. The fuel source 48 can be configured to store and supply one or more fuels 52 such as natural gas, syngas, methane [CH4], hydrogen [H2], ammonia [NH3], biofuel, or combinations thereof. The inert gas source 50 can be configured to store and supply one or more inert gases such as nitrogen [N2], argon [Ar], carbon dioxide [CO2], recirculated exhaust gas (e.g., from an exhaust gas recirculation [EGR] system and containing carbon dioxide), or combinations thereof.

[0021] The fuel injection system 11 can be configured to supply the same or different fuels 52 (e.g., natural gas, syngas, methane [CH4], hydrogen [H2], ammonia [NH3], biofuel, or combinations thereof) to the combustor section 20 and the turbine section 22. For example, the fuel injection system 11 can be configured to use a relatively low-carbon or carbon-free fuel (e.g., hydrogen, ammonia, biofuel, or combinations thereof) for both the combustor section 20 and the turbine section 22, or the fuel injection system 11 can be configured to use a carbon-containing fuel (e.g., natural gas, syngas, methane, or combinations thereof) for the combustor section 20 and a relatively low-carbon or carbon-free fuel (e.g., hydrogen, ammonia, biofuel, or combinations thereof) for the turbine section 22. In certain embodiments, the fuel injection system 11 can be configured to use different fuels at different operating stages of the gas turbine system 10 such as startup, steady state, shutdown, partial load, or transient states.

[0022] The fuel supply source 48 is configured to supply fuel 52 (e.g., natural gas, syngas, methane, hydrogen, ammonia, biofuel, or a combination thereof) to the turbine section 22, and the inert gas supply source 50 is configured to supply inert gas 54 (e.g., carbon dioxide, nitrogen, argon, recirculated exhaust gas, or a combination thereof) to the turbine section 22. As will be described in more detail herein, the fuel 52 is injected into the turbine section 22 via a plurality of fuel injection ports 56 (e.g., fuel ports), and the inert gas 54 is injected into the turbine section 22 via a plurality of inert gas injection ports 57 (e.g., inert gas ports). The fuel injection ports 56 and the inert gas injection ports 57 are disposed directly above and / or between the airfoils (e.g., blades 44 and / or vanes 46) in one or more turbine stages of the turbine section 22. The fuel injection ports 56 and the inert gas injection ports 57 may extend from channels formed in the structure of the turbine section 22, and the channels may be cylindrical, rectangular, or other shapes. Similarly, depending on the interface between the channels and the surface of the turbine section 22, the fuel injection ports 56 and the inert gas injection ports 57 may be circular, rectangular, oval, triangular, polygonal, or other (not necessarily symmetric) shapes.

[0023] In the illustrated embodiment, the fuel injection system 11 includes a plurality of valves 58 that regulate the flow rate of fuel 52 and / or inert gas 54 from a fluid source 14 to a fuel injection port 56 and / or an inert gas injection port 57. As shown, the fuel injection system 11 further includes a controller 60 having a processor 62, a memory 64, and a communication circuit 68 coupled to various sensors and actuators of the gas turbine system 10. The processor 62 includes instructions 66 stored in the memory 64 and executable by the processor 62 to control the gas turbine system 10, in particular the fuel injection system 11. The controller 60 is communicably coupled (e.g., via the communication circuit 68) to the fluid source 14, the plurality of valves 58, and, in a particular embodiment, to sensors 70 located in the turbine section 22. In a particular embodiment, the sensors 70 may provide signals indicating the flow rate of fuel 52 and / or inert gas 54. The controller 60 can be configured to determine an estimated flow rate of fuel 52 and / or inert gas 54 based on the received signal, and to control a plurality of valves 58 to control the flow rate of fuel 52 and / or inert gas 54 based on the estimated flow rate.

[0024] Figure 3 is a top view of the turbine section 22 of the gas turbine engine 12 of Figure 2, further showing the locations of one or more fuel injection ports 56 and one or more inert gas injection ports 57 of the fuel injection system 11 along the flow path 23 (e.g., high-temperature combustion gas flow) through the turbine section 22. The fuel injection ports 56 are configured to direct (e.g., inject) fuel 52 (e.g., a jet, flow, or spray of fuel) into a chamber 94 between continuous, circumferentially spaced airfoil sections 96 (e.g., airfoil sections 98, 100) of the turbine section 22, and the inert gas injection ports 57 are configured to direct inert gas 54 (e.g., a film of inert gas, a jet of inert gas) into the chamber 94. In certain embodiments, the airfoil sections 96 may include turbine blades 44, turbine vanes 46, or a combination thereof. As shown, the inert gas injection ports 57 are located upstream of the fuel injection ports 56 (e.g., upstream 102 opposite to the flow direction of the flow path 23). As will be described in more detail herein, the inert gas injection port 57 is configured to cross (e.g., contact) the inert gas 54 so that it fills and pressurizes a buffer region (e.g., a low-pressure region) downstream of the fuel injection port 56, and to guide it to flow around the fuel 52 injected through the fuel injection port 56.

[0025] The size and shape of the fuel injection port 56 and the inert gas injection port 57 may be the same or different from each other. For example, the width of the inert gas injection port 57 may be greater than the width of the fuel injection port 56, thereby helping to spread the inert gas 54 more widely than the fuel 52 and to facilitate the flow of the inert gas 54 around the fuel 52 into the buffer area. Where used herein, the width or width dimension of the fuel injection port or fuel injection port opening, or the inert gas injection port or inert gas port opening, may be defined in particular in a direction traversing the upstream-downstream direction of the flow path or turbine section, or more specifically in a direction perpendicular thereto. As a further example, the inert gas injection port 57 may be elongated laterally with respect to the flow path 23, and more specifically, may be molded as a rectangular port or an oval port. The fuel injection port 56 may be molded as a circular port. However, the fuel injection port 56 and the inert gas injection port 57 may include circular, square, rectangular, oval, triangular, polygonal, or any combination thereof, and the inert gas injection port 57 may have larger dimensions than the fuel injection port 56.

[0026] The fuel injection port 56 and the inert gas injection port 57 may be oriented at the same or different angles with respect to the surface of the wall 111. For example, downstream of the flow path 23, the fuel injection port 56 may be oriented at a 90-degree angle and the inert gas injection port 57 may be oriented at an acute angle (e.g., 5-70, 10-60, or 15-50 degrees). However, the fuel injection port 56 and the inert gas injection port 57 may be oriented at the same or different angles of 10 degrees or less, 20 degrees or less, 30 degrees or less, 40 degrees or less, 50 degrees or less, 60 degrees or less, 70 degrees or less, 80 degrees or less, or 90 degrees or less.

[0027] In the illustrated embodiment, the turbine section includes sets of ports 104 (e.g., a first set 106, a second set 108, a third set 110) formed in a plurality of walls 111 (e.g., walls 107, 109, and 112) of the turbine section 22. As shown, each set of ports 104 includes a fuel injection port 56 and an inert gas injection port 57 located upstream 102 relative to the fuel injection port 56. In the illustrated embodiment, the turbine section 22 includes a first set 106 formed in the negative pressure side wall 107 of the airfoil section 98, a second set 108 formed in the wall 109 between continuous circumferentially spaced airfoil sections 96, and a third set 110 formed in the positive pressure side wall 112 of the airfoil section 100. In certain embodiments, the walls 111 may include the outer diameter walls of the turbine section 22 (e.g., if the airfoil section 96 is a stationary vane). Additionally or alternatively, the wall 111 may include an inner diameter wall of the turbine section (for example, if the airfoils 96 are rotating blades). In certain embodiments, the turbine section 22 may include a set of ports 104 formed on both the blades 44 and vanes 46 of the turbine section 22. In certain embodiments, the turbine section 22 may include any combination of the sets of ports 104. The illustrated embodiment shows the turbine section 22 including three sets of ports 104 between consecutive circumferentially spaced airfoils 96 (two sets for each airfoil 96 and one set between adjacent airfoils 98, 100), but it can be recognized that the turbine section 22 may include more or fewer sets of ports 104 for each pair of airfoils 96 and / or adjacent airfoils (e.g., 98, 100). For example, the turbine section may include one, two, four, five, six, or more sets of ports 104 for each pair of airfoil sections 96 and / or adjacent airfoil sections (e.g., 98, 100).

[0028] Figure 4 is a perspective view of one embodiment of the fuel injection system 11 of Figure 2, further schematicly showing the fuel injection port 56 and the inert gas injection port 57 of the fuel injection system 11. As shown, the fuel injection port 56 includes a fuel port channel 130 fluidly coupled to a fuel port opening 132 (e.g., outlet) on the surface 134 of the wall 111. In addition, the inert gas injection port 57 includes an inert gas port channel 136 fluidly coupled to an inert gas port opening 138 (e.g., outlet) on the surface 134.

[0029] In the illustrated embodiment, the fuel port channel 130 is circular in shape (e.g., cylindrical channel) and substantially perpendicular to the surface 134 (e.g., vertical), so the fuel injection port 56 is configured to inject fuel 52 substantially perpendicular to the surface 134. In addition, as described herein, the inert gas port channel 136 is rectangular in shape (e.g., rectangular channel) and acute to the surface 134, so the inert gas injection port 57 is configured to inject inert gas 54 toward the injected fuel 52 at an acute angle to the surface 134. For example, the acute angle of the inert gas injection port 57 may be about 5 to 70, 10 to 60, 15 to 50, or 3 to 15 degrees, for example, about 10 degrees or less, 15 degrees or less, 20 degrees or less, 25 degrees or less, 30 degrees or less, 40 degrees or less, 50 degrees or less, or 60 degrees or less.

[0030] In the illustrated embodiment, the inert gas port channel 136 includes a first channel portion 139 and a second channel portion 140. As shown, the height dimension 142 of the first channel portion 139 is smaller than the width dimension 144 of the first channel portion 139. In certain embodiments, the smaller height dimension 142 of the first channel portion 139 may allow the inert gas injection port 57 to inject the inert gas 54 as a membrane (e.g., a membrane layer) toward the injected fuel 52. For example, the ratio of the width dimension 144 to the height dimension 142 may be 2:1 to 20:1, 2:1 to 10:1, or 2:1 to 5:1. In some embodiments, the ratio of the width dimension 144 of the inert gas port channel 136 to the width dimension 133 (e.g., diameter) of the fuel port channel 130 may be 1.5:1 to 10:1, 2:1 to 5:1, or 2:1 to 4:1. In addition, as shown in the figure, the upper surface 146 of the second channel portion 140 is inclined upward at an acute angle (e.g., obliquely) with respect to the bottom surface 148 of the second channel portion 140 (e.g., and the first channel portion 139), so that the height dimension 150 of the second channel portion 140 gradually increases along the downstream direction 151, and as a result the inert gas port opening 138 has a length dimension 152 that is substantially equal to the width dimension 144.

[0031] Figure 5 is a top view of one embodiment of the fuel injection system 11, further showing the fuel injection port 56 and the inert gas injection port 57. As shown, the inert gas port opening 138 includes an inert gas opening width dimension 170 (e.g., a first width dimension) that is slightly larger (e.g., less than 10% larger) than the width dimension 144 of the first channel portion 139. As shown, the fuel injection port 56 is located downstream 151 of the flow path 23 than the inert gas injection port 57. In the illustrated embodiment, the fuel port opening 132 is circular in shape and has a diameter 172 (e.g., a second width dimension equal to the width dimension 133 of the fuel port channel 130 shown in Figure 4). In certain embodiments, the fuel port opening 132 may have a different shape (e.g., rectangular, triangular, polygonal, etc.). In the illustrated embodiment, the inert gas port opening 138 is shown as rectangular. In certain embodiments, the inert gas port opening 138 may have a different shape (e.g., triangular, elliptical, etc.).

[0032] In the illustrated embodiment, the diameter 172 of the fuel port opening 132 is smaller than the inert gas opening width dimension 170 of the inert gas port opening 138. In certain embodiments, the ratio between the diameter 172 of the fuel port opening 132 and the inert gas opening width dimension 170 is 1:1.5 to 1:10, 1:2 to 1:5, or 1:2 to 1:4. As shown, the fuel port central axis 174 of the fuel port opening 132 is offset from the downstream edge 176 of the inert gas port opening 138 by a length 178. In certain embodiments, the ratio between the diameter 172 of the fuel port opening 132 and the length 178 is 0.8:1 to 1.2:1. In certain embodiments, the length 178 may be about 1, 1.5, 2, 2.5, 3, 3.5, or 4 times or less the diameter 172 of the fuel port opening 132.

[0033] In the illustrated embodiment, the fuel port central axis 174 of the fuel port opening 132 is aligned with the inert gas port central axis 180 of the inert gas port opening 138 in the downstream direction 151 such that axes 174 and 180 intersect each other. That is, the fuel port central axis 174 of the fuel port opening 132 is equidistant from the distal end or lateral margin 182 (e.g., distal end or lateral margin 184, distal end or lateral margin 186) of the downstream edge 176 of the inert gas port opening 138, so that the fuel port opening 132 is positioned between the distal ends 182, respectively, or centered relative to the inert gas port opening when viewed along the upstream-downstream direction of the flow path. When viewed along the upstream-downstream direction of the flow path, the inert gas port opening 138 has a width extending between the two lateral margins 184 and 186, and the fuel port opening 132 is positioned between the two lateral margins of the width of the inert gas port opening. This configuration of the fuel port opening 132 and the inert gas port opening 138 will be understood to allow the inert gas 54 injected from the inert gas port opening 138 to flow around the sides 188 (e.g., sides 190, sides 192) of the fuel port opening 132 (e.g., and the injected fuel 52). Although the illustrated embodiment shows one fuel port opening 132, it can be recognized that the fuel injection system 11 may include multiple fuel port openings 132 for each inert gas port opening 138. For example, the fuel injection system 11 may include two, three, four, or more fuel port openings 132 for each inert gas port opening 138. In a particular embodiment, different sets of ports may each include a different number of fuel port openings 132.

[0034] Figure 6 is a cross-sectional side view of the fuel injection system 11, further showing the flow path 210 for fuel 52 injected through the fuel injection port 56, and the inert gas flow path 212 for inert gas 54 injected through the inert gas injection port 57. As shown, the fuel port central axis 174 is substantially perpendicular (e.g., vertical) to the surface 134. In certain embodiments, the angle 214 extending from the surface 134 of the wall 111 to the fuel port central axis 174 is 45–95 degrees, 65–95 degrees, or 85–95 degrees. In addition, as shown, the inert gas port central axis 180 is acute (e.g., less than 30 degrees) to the surface 134. In certain embodiments, the angle 216 extending from the surface 134 to the inert gas port central axis 180 is approximately 5 to 70 degrees, 10 to 60 degrees, 15 to 50 degrees, or 3 to 15 degrees, for example, approximately 10 degrees or less, 15 degrees or less, 20 degrees or less, 25 degrees or less, 30 degrees or less, 40 degrees or less, 50 degrees or less, or 60 degrees or less.

[0035] In the illustrated embodiment, the crossflow of the core gas 218 (e.g., combustion gas) flows downstream 151 along the flow path 23 above the injected fuel 52 and inert gas 54. The crossflow of the core gas 218 causes the injected fuel 52 to bend downstream 151 along the flow path 23. As shown, the injected fuel 52 is ignited after entering the chamber 94 from the fuel injection port 56, forming a fuel jet region 220 (e.g., flame region). In the illustrated embodiment, the mass fraction (e.g., density) of the fuel 52 is greater near the interior portion 222 of the jet region 220 than near the edge portion 224 of the fuel jet region 220. The fuel jet region 220 transitions to a rear region 225 (e.g., wake region) containing ignited fuel 52 with a lower mass fraction as the ignited fuel 52 moves downstream from the fuel injection port 56. As shown in the figure, when the inert gas 54 is injected into the chamber 94, it flows through the inert gas jet region 226 toward the fuel jet region 220, remaining near the surface 111 due to the angle 216 at which the inert gas 54 enters the crossflow of the core gas 218. In certain embodiments, the pressure ratio of the fuel jet region 220 is in the range of 1.5 to 1.7, 1.55 to 1.65, 1.575 to 1.625, or a combination thereof. In certain embodiments, the pressure ratio of the inert gas jet region 226 is in the range of 1.0 to 1.2, 1.05 to 1.15, 1.075 to 1.125, or a combination thereof.

[0036] As shown in the figure, upon reaching the fuel jet region 220, the inert gas 54 flows around the side surface 188 of the fuel jet region 220, at least partially filling and pressurizing the buffer region 228 located between the downstream and rear region 225 of the fuel jet region 220 and the surface 111. The inert gas 54 fills this region with a nearly uniform mass fraction. It will be understood that filling the buffer region 228 with the inert gas 54 suppresses the ignition of the injected fuel 52 near the surface 111 downstream of the fuel injection port 56.

[0037] Figure 7 is a schematic diagram of one embodiment of a flow path 250 for injected fuel 52 and inert gas 54. In the illustrated embodiment, the flow path 250 includes a fuel jet region 220 that collides with the core gas 218 in the flow path 23 by crossflow. In the illustrated embodiment, the fuel jet region 220 transitions from a single jet portion 251 to a pair of counter-rotating vortices 252 (e.g., a first vortex 254, a second vortex 256). As will be further described herein, the velocity of the injected fuel 52 is lower in the vortex interior portions 258 of the counter-rotating vortices 252 (e.g., vortex interior portions 260, vortex interior portions 262) than in the vortex edge portions 264 of the counter-rotating vortices 252 (e.g., vortex edge portions 266, vortex edge portions 268).

[0038] In the illustrated embodiment, a single jet portion 251 of the fuel jet region 220 includes a shear wave vortex 270 formed in the outer portion 272 of the single jet portion 251. As shown, the flow path 250 further includes a plurality of wake vortices 274 (e.g., wake vortices 276, 278, 280, and 282), so that each wake vortex 274 extends from the single jet portion 251 and / or the counter-rotating vortex 252 to the wall 111. Although the illustrated embodiment shows four wake vortices 274, the flow path 250 may include fewer or more wake vortices 274. As shown, the plurality of wake vortices 274 extend through a buffer region 228 located above the wall 111 and below the single jet portion 251 and / or the counter-rotating vortex 252.

[0039] In addition, as shown in the figure, the flow path 250 includes a horseshoe vortex 284 flowing around the side 188 of a single jet portion 251. As shown in the figure, the horseshoe vortex 284 flows from the upstream area 286 upstream of the fuel jet region 220 to the buffer region 228 downstream of the fuel jet region 220. In certain embodiments, the horseshoe vortex 284 transports inert gas 54 from the inert gas port to the buffer region 228. As described herein, a higher mass fraction of inert gas 54 in the buffer region 228 suppresses the ignition of injected fuel 52 in the buffer region 228, thereby suppressing the ignition of injected fuel 52 near the surface 111 downstream of the fuel injection port 56. The inert gas 54 in the buffer region 228 lifts the fuel 52 away from the surface 111, promoting ignition and flame formation in the hot gas flow path 23.

[0040] Figure 8 is a cross-sectional plot 298 showing the velocity profile 300 of the flow path 250 for fuel 52 and inert gas 54. As shown, the cross-sectional plot 298 is a perspective view with the downstream direction 151 in the page. As shown, the cross-sectional plot 298 shows the flow path 250 as including a pair of counter-rotating vortices 252 having a first vortex 254 and a second vortex 256. As shown, the velocity of the injected fuel 52 is lower in the vortex interior portion 258 of the counter-rotating vortex 252 (e.g., vortex interior portion 260, vortex interior portion 262) than in the vortex edge portion 264 of the counter-rotating vortex 252 (e.g., vortex edge portion 266, vortex edge portion 268). In the illustrated embodiment, the vortex interior portion 258 forms a kidney bean-shaped region where the velocity of the injected fuel 52 is lower than that of the surrounding vortex edge portion 264.

[0041] In addition, a cross-sectional plot 298 shows a horseshoe vortex 284 flowing around the side 188 of a single jet section. As shown in the figure, the horseshoe vortex 284 is shown as a region with a higher velocity, sandwiched between the counter-rotating vortex 252 and the wall 111. The horseshoe vortex 284 helps to transport the inert gas 54 from the inert gas port 57, around the side 188 of the fuel port 56, and to the buffer region 228 downstream of the fuel port 56.

[0042] Figure 9 is a longitudinal plot 320 perpendicular to the transverse plot 298 of Figure 8, showing the velocity profile 300 of the flow path 250. In the illustrated embodiment, line 322 illustrates the cross section used for the transverse plot 298 of Figure 8. As shown, the velocity of the flow path 250 is lower in the buffer region 228 compared to the cross-flow region 324 in which the core gas 218 flows freely. As shown, the buffer region 228 includes a low-velocity region 326 immediately downstream of the fuel injection port 56, in which the velocity of the flow path 250 is lower than that of the rest of the buffer region 228. It will be understood that the inert gas 54 flows into the buffer region 228 from the inert gas injection port 57 and, due to the low-velocity profile and low pressure of the buffer region 228, can diffuse at least partially throughout the buffer region 228. In other words, the inert gas 54 is at least partially encompassed into the buffer region 228 by being carried by the horseshoe vortex 284, due to the low speed and low pressure of the buffer region 228, which draws the inert gas 54 into the buffer region 228. As will be understood herein, the encompassion of the inert gas 54 into the buffer region 228 can suppress the ignition of the fuel 52 downstream of the fuel injection port 56 along the wall (e.g., flame anchoring).

[0043] Figure 10 is a flowchart illustrating an exemplary process 350 for operating the fuel injection system 11. Process 350 can be performed by the controller 60 in Figure 2 or any other suitable computing device or controller. Furthermore, the operation of process 350 may be performed in the order disclosed herein or in any other suitable order. For example, certain operations of process 350 can be performed simultaneously.

[0044] In block 352 of process 350, fuel is injected from fuel ports formed in the wall of the turbine section between continuous, circumferentially spaced airfoil sections of the turbine section. In certain embodiments, the airfoil sections may include turbine vanes, turbine blades, or a combination thereof. Additionally or alternatively, the wall may include the wall of the turbine vanes, the wall of the turbine blades, the inner diameter wall of the turbine section, the outer diameter wall of the turbine section, or a combination thereof. The fuel may be injected substantially perpendicular to the surface of the wall. In certain embodiments, the fuel may include natural gas, synthesis gas, methane, hydrogen, ammonia, biofuels, or a combination thereof.

[0045] In block 354 of process 350, an inert gas is injected toward the injected fuel from an inert gas port located upstream of the fuel port. The inert gas may be injected toward the injected fuel (e.g., fuel jet) at a sharp angle to the wall surface. In certain embodiments, the injected inert gas may include an injection layer of the inert gas (e.g., a membrane). In certain embodiments, the inert gas may include carbon dioxide, nitrogen, argon, recirculated exhaust gas, or a combination thereof. In certain embodiments, the injected fuel may generate a set of counter-rotating vortices, and the velocities of the fuel, inert gas, or a combination thereof are lower near the center of each vortex of the set of counter-rotating vortices.

[0046] In block 356 of process 350, the injected inert gas is made to occupy at least a portion of the buffer region downstream of the fuel port. As described herein, the mass fraction of the inert gas may be relatively uniform across the entire portion of the buffer region it occupies. It will be understood that the presence of the inert gas in the buffer region can suppress fuel ignition near the surface of the wall downstream of the fuel port (e.g., flame anchoring) and / or help lift the flame away from the ignited fuel into the hot gas flow path.

[0047] The technical effect of the disclosed embodiment is to cause an inert gas injected upstream of a fuel port formed in the wall of the turbine section to flow into a buffer region downstream of the fuel port, thereby at least partially occupying the buffer region. The increase in the mass fraction of the inert gas in the buffer region suppresses flame anchoring (e.g., sustaining ignition of the injected fuel) near the wall surface downstream of the fuel port.

[0048] The subject matter described in detail above may be defined by one or more of the following clauses.

[0049] According to a first embodiment, the system includes at least one component of a turbine section, the turbine section comprising: a fuel port formed in the wall of the at least one component, the fuel port being configured to inject fuel into a chamber of the turbine section; and an inert gas port formed in the wall upstream of the fuel port, the inert gas port being configured to inject an inert gas so as to flow at least partially around the injected fuel into a buffer region downstream from the fuel port. In another aspect of the first embodiment, the system comprises at least one turbine section, the turbine section having a flow path through the turbine section, the flow path defining an upstream direction and a downstream direction. The turbine section further comprises a fuel port formed in the wall of the turbine section, the fuel port being configured to inject fuel into the turbine section. Furthermore, an inert gas port is formed in the wall upstream of the fuel port, the inert gas port being configured to inject an inert gas so as to flow at least partially around the fuel injected into a buffer region downstream from the fuel port.

[0050] The system according to any one of the preceding clauses and / or appended claims, wherein the chamber is located between a series of circumferentially spaced airfoil sections of the turbine section.

[0051] The system according to any of the preceding clauses and / or any one of the appended claims, wherein the continuous, circumferentially spaced airfoil sections include blades or vanes in each turbine stage.

[0052] The system according to any of the preceding clauses and / or any one of the appended claims, wherein the wall includes at least one of the blade wall of the blade, the vane wall of the vane, the inner diameter wall of the turbine section, or the outer diameter wall of the turbine section.

[0053] The system according to any of the preceding clauses and / or any one of the appended claims, wherein the first ratio of the first width dimension of the fuel port opening of the fuel port to the second width dimension of the inert gas port opening of the inert gas port is 1:2 to 1:4.

[0054] The system according to any of the preceding clauses and / or any one of the appended claims, wherein the length is defined by the distance between the central axis of the fuel port opening and the downstream edge of the inert gas port opening, and the second ratio of the first width dimension of the fuel port opening to the length is 1:2 to 3:1.

[0055] The system according to any of the preceding clauses and / or any one of the appended claims, wherein the first angle between the central axis of the inert gas port and the surface of the wall is less than 30 degrees, and the second angle between the central axis of the fuel port and the surface is between 45 and 95 degrees.

[0056] The system according to any of the preceding clauses and / or any one of the appended claims, wherein the fuel port opening is located between the first and second distal ends of the second width dimension of the inert gas port opening.

[0057] The system according to any of the preceding clauses and / or any one of the appended claims, wherein the central axis of the fuel port is equidistant from the first and second distal ends.

[0058] The system according to any of the preceding clauses and / or any one of the appended claims, wherein the first width dimension of the fuel port opening is defined by the diameter of the fuel port opening, and the shape of the inert gas port opening is rectangular.

[0059] The system according to any of the preceding clauses and / or any one of the appended claims, wherein the first pressure ratio of the fuel injected from the fuel port is 1.5 to 1.7, and the second pressure ratio of the inert gas injected from the inert gas port is 1.0 to 1.2.

[0060] A system according to any of the preceding clauses and / or any one of the appended claims, comprising the at least one component of the turbine section, the turbine section, a gas turbine engine, or a combination thereof.

[0061] According to a second embodiment, the system includes a gas turbine engine having a turbine section having circumferentially spaced airfoils, at least one component of the turbine section being a fuel injection port formed in the wall of the at least one component, the fuel injection port being configured to guide fuel between a series of circumferentially spaced airfoils of the turbine section, and an inert gas injection port formed in the wall upstream of the fuel port, the inert gas injection port being configured to guide inert gas so as to flow at least partially around the fuel from the fuel port to a buffer region downstream.

[0062] The system according to any one of the preceding clauses and / or appended claims, wherein the continuous circumferentially spaced airfoil sections include blades or vanes of each turbine stage of the turbine section.

[0063] The system according to any of the preceding clauses and / or any one of the appended claims, wherein the first ratio of the first width dimension of the fuel port opening of the fuel injection port to the second width dimension of the inert gas port opening of the inert gas injection port is 1:2 to 1:4.

[0064] The system according to any of the preceding clauses and / or any one of the appended claims, wherein the length is defined by the distance between the fuel port central axis of the fuel port opening and the downstream edge of the inert gas port opening, and the second ratio of the second width dimension of the fuel injection port to the length is 1:2 to 3:1.

[0065] The system according to any of the preceding clauses and / or any one of the appended claims, wherein the inert gas port is a rectangular port and is oriented with respect to the wall at an acute angle of 30 degrees or less.

[0066] According to a third aspect, the method includes the steps of: injecting fuel into the turbine section of a gas turbine engine from a fuel port formed in the wall of at least one component of the turbine section of the gas turbine engine; injecting an inert gas from an inert gas port located upstream of the fuel port toward the injected fuel; and occupying at least a portion of a buffer area downstream of the fuel port with the injected inert gas. In another aspect of the third aspect, a method for injecting fuel into the turbine section of a gas turbine engine includes the steps of: injecting the fuel into the turbine section of a gas turbine engine from a fuel port formed in the wall of the turbine section; injecting an inert gas from an inert gas port located upstream of the fuel port toward the injected fuel along the flow direction through the flow path of the turbine section; and flowing the injected inert gas at least partially around the injected fuel in a buffer area downstream of the fuel port.

[0067] The method according to any one of the preceding clauses and / or appended claims, wherein the step of injecting the fuel includes injecting the fuel perpendicular to the surface of the wall into a chamber located between a series of circumferentially spaced airfoils of the turbine section, and the step of injecting the inert gas includes injecting the inert gas toward the injected fuel into the chamber at an acute angle to the surface.

[0068] The method according to any of the preceding clauses and / or any one of the appended claims, wherein the injected fuel generates a set of counter-rotating vortices, and the velocity of the fuel, the inert gas, or a combination thereof is lower near the center of each vortex of the set of counter-rotating vortices.

[0069] This specification discloses the invention in best mode and uses examples to enable those skilled in the art to practice the invention, including the fabrication and use of any device or system, and the implementation of any incorporated method. The scope of the invention as claimed herein is defined by the appended claims and may include other examples beyond those explicitly exemplified, which would be conceivable to those skilled in the art. [Explanation of symbols]

[0070] 10 Gas Turbine Systems 11. Fuel Injection System 12 Gas turbine engines 14 Fluid supply source 16. Intake Section 18 Compressor Section 20 Combustor Section 22 Turbine Section 23 Flow channels 24 Exhaust Section 26 shafts 28 Combustor Housing 32 Longitudinal axis 34 Radial axis 36 Circumferential axis 38 Combustor 40 Fuel Nozzles 42 transition pieces 44 Turbine Blades 46 Turbine vanes 48 Fuel supply source 50 Inert gas supply source 52 Fuel 54 Inert gas 56 Fuel Injection Ports 57 Inert gas injection port 58 valves 60 Controllers 62 processors 64 memory 66 command 68 Communication Circuit 94 Chambers 96 Airfoil 98 Airfoil 100 Airfoil 102 Upstream direction 104 pairs of ports 106 Group 1 107 Negative pressure sidewall 108 Group 2 109 Wall 110 Third group 111 Wall 112 Positive pressure sidewall 130 Fuel port channels 132 Fuel port opening 133 width dimension 134 Surface 136 Inert gas port channels 138 Inert gas port opening 139 First channel portion 140 Second channel portion 142 Height dimensions 144 width dimensions 146 Top surface 148 Bottom 150 Height dimension 151 Downstream direction 152 Length dimensions 170 Inert gas opening width dimension 172 diameter 174 Fuel port central axis 176 Downstream edge 178 Length 180 Inert gas port central axis 182 Lateral margin / distal end 184 Lateral margin / distal end 186 Lateral margin / distal end 188 Side view 190 Side view 192 Side view 210 flow path 212 Inert gas flow path 214 angle 216 angle 218 Core gas 220 Fuel jet area 222 Internal part 224 Edge 225 Posterior area 226 Inert gas jet area 228 buffer area 250 flow paths 251 Single jet section 252 Reverse low vortex 254 The First Vortex 256 The Second Vortex 258 Internal part of vortex 260 Internal part of vortex 262 Internal part of vortex 264 Vortex edge part 266 Vortex edge part 268 Vortex edge part 270 Shear vortex 272 Outer part 274 Wake Vortex 276 Wake Vortex 278 Wake Vortex 280 Wake Vortex 282 Wake Vortex 284 Horseshoe Vortex 286 Upstream area 298 Cross-sectional plot 300 Speed ​​Profiles 320 Longitudinal Plot 322 line 324 Cross-flow area 326 Low speed area 350 Methods / Processes

Claims

1. At least one turbine section (22), the turbine section (22) having a flow path (23) through the turbine section (22), the flow path (23) defining an upstream direction (102) and a downstream direction (151) of the at least one turbine section (22). Equipped with, The turbine section (22) is a fuel port (56) formed in the wall (111) of the turbine section (22), and the fuel port (56) is configured to inject fuel (52) into the turbine section (22), An inert gas port (57) formed in the wall (111) upstream of the fuel port (56), wherein the inert gas port (57) is configured to inject an inert gas (54) so ​​as to flow at least partially around the fuel (52) injected from the fuel port (56) into the buffer region (228) downstream of the fuel port (56), and Furthermore, system.

2. The system according to claim 1, wherein the inert gas port (57) is configured to inject the inert gas (54) into the chamber (94) of the turbine section (22).

3. The system according to claim 2, wherein the turbine section (22) includes airfoil sections (96, 98, 100) spaced apart in the circumferential direction.

4. The system according to claim 3, wherein the chamber (94) is positioned between continuous, circumferentially spaced airfoil sections (98, 100) of the turbine section (22).

5. The system according to claim 3, wherein the fuel port (56) is configured to guide fuel (52) between the continuous circumferentially spaced airfoil sections (96, 98, 100) of the turbine section (22).

6. The system according to claim 3, wherein the wall (111) comprises at least one of the airfoil wall of the airfoil section (96, 98, 100), the inner diameter wall of the turbine section (22), or the outer diameter wall of the turbine section (22).

7. The system according to claim 1, wherein the ratio of the width dimension of the fuel port opening (132) of the fuel port (56) to the width dimension (144) of the inert gas port opening (138) of the inert gas port (57) is 1:2 to 1:

4.

8. The system according to claim 1, wherein the length (178) is defined by the distance between the fuel port central axis (174) of the fuel port opening (132) and the downstream edge (176) of the inert gas port opening (138), and the ratio of the width dimension of the fuel port opening (132) to the length (178) is 1:2 to 3:

1.

9. The system according to claim 1, wherein the angle between the central axis (180) of the inert gas port (57) and the surface (134) of the wall (111) is 30 degrees or less.

10. The system according to claim 1, wherein the angle (214) between the central axis (174) of the fuel port (56) and the surface (134) of the wall (111) is 45 to 95 degrees.

11. The system according to any one of claims 1 to 10, wherein, when viewed along the upstream-downstream direction (102, 151) of the flow path (23), the inert gas port opening (138) has a width extending between two lateral margins (184, 186), and the fuel port opening (132) is positioned between the two lateral margins (184, 186) of the width of the inert gas port opening (138), and in particular the fuel port opening (132) is centered with respect to the lateral margins (184, 186) of the inert gas port opening (138) when viewed along the upstream-downstream direction (102, 151) of the flow path (23).

12. The system according to any one of claims 1 to 11, wherein the shape of the inert gas port (57) opening (138) is rectangular.

13. A method (350) of injecting fuel (52) into the turbine section (22) of a gas turbine engine (12), Step (352) of injecting the fuel (52) from a fuel port (56) formed in the wall (111) of the turbine section (22) into the turbine section (22) of the gas turbine engine (12), Step (354) involves injecting an inert gas (54) from an inert gas port (57) located upstream of the fuel port (56) toward the injected fuel (52) along the flow direction through the flow path (23) of the turbine section (22), Step (356) of flowing the injected inert gas (54) at least partially around the fuel (52) injected into the buffer region (228) downstream of the fuel port (56) and Method (350), including the method (350).

14. The method according to claim 13 (350), wherein the step (352) of injecting the fuel (52) includes injecting the fuel (52) perpendicular to the surface (134) of the wall (111) between continuous circumferentially spaced airfoil sections (98, 100) of the turbine section (22), and the step (354) of injecting the inert gas (54) includes injecting the inert gas (54) toward the injected fuel (52) at an acute angle to the surface (134) of the wall (111), the method according to claim 13 (350).

15. The method according to any one of claims 13 or 14 (350), comprising the step of injecting (352) the fuel (52) to generate a set of counter-rotating vortices (252), wherein the velocities of the fuel (52), the inert gas (54), or a combination thereof are lower near the center of each vortex of the set of counter-rotating vortices (252).

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