Gas turbine combustor and gas turbine

Abstract

A gas turbine combustor according to at least one embodiment of the present disclosure is provided with: a combustion cylinder having therein a combustion chamber and having formed therein a plurality of through-holes that open to the combustion chamber; a housing that is disposed peripherally outwards of the combustion cylinder and defines an acoustic attenuation space that communicates with the combustion chamber via the through-holes; an air hole plate that has formed therein a plurality of air holes and is positioned upstream of the combustion cylinder; a plurality of fuel nozzles respectively corresponding to the plurality of air holes; an air passage that is provided between the inner peripheral surface of the combustion cylinder and the outer peripheral surface of the air hole plate and extends in the axial direction of the combustion cylinder; and an air supply flow path for supplying air flowing outside the combustion cylinder to the acoustic attenuation space. At least one of the plurality of through holes is provided directly below the air hole plate in the axial direction.

Description

[Technical Field]

[0001] The present disclosure relates to gas turbine combustors and gas turbines. This application claims priority based on Japanese Patent Application No. 2022-011689, filed with the Japan Patent Office on January 28, 2022, the contents of which are incorporated herein by reference. [Background technology]

[0002] There is known a gas turbine combustor that has an air hole plate disposed between a fuel nozzle and a combustion chamber, and that is configured to eject a fuel flow and an air flow formed on the outer periphery of the fuel flow into the combustion chamber inside air holes formed in the air hole plate (see, for example, Patent Document 1). [Prior art documents] [Patent documents]

[0003] [Patent Document 1] Japanese Patent Application Publication No. 2018-150912 Summary of the Invention [Problem to be solved by the invention]

[0004] The gas turbine combustor described in Patent Document 1 is also called a multi-cluster combustor. In a multi-cluster combustor described in Patent Document 1, the flame length is shorter than that of a conventional combustor that is not a multi-cluster combustor, and therefore antinodes and nodes of combustion oscillations are formed near the burner outlet. Therefore, in order to suppress combustion oscillations, it is important to place an acoustic device having an acoustic damping space directly below the burner. However, in a gas turbine combustor such as that described in Patent Document 1, the region where the acoustic device is disposed is likely to become relatively hot, and therefore it is necessary to achieve both cooling and acoustic characteristics.

[0005] In view of the above circumstances, at least one embodiment of the present disclosure has an object to achieve both cooling and acoustic characteristics in a gas turbine combustor. [Means for solving the problem]

[0006] (1) A gas turbine combustor according to at least one embodiment of the present disclosure includes: a combustion liner having a combustion chamber therein and a plurality of through holes opening into the combustion chamber; a housing that is disposed on an outer circumferential side of the combustion liner and defines an acoustic attenuation space that communicates with the combustion chamber via the through hole; an air hole plate having a plurality of air holes formed therein and positioned upstream of the combustion liner; a plurality of fuel nozzles corresponding to the plurality of air holes; an air passage provided between an inner peripheral surface of the combustion liner and an outer peripheral surface of the air hole plate, the air passage extending in the axial direction of the combustion liner; an air supply passage for supplying air flowing outside the combustion liner to the acoustic attenuation space; Equipped with At least one of the plurality of through holes is provided directly below the air hole plate in the axial direction.

[0007] (2) A gas turbine according to at least one embodiment of the present disclosure includes: a compressor for generating compressed air; a gas turbine combustor having the configuration of (1) above; a turbine that is rotationally driven by combustion gas generated by the gas turbine combustor; Equipped with. [Effects of the Invention]

[0008] According to at least one embodiment of the present disclosure, it is possible to achieve both cooling and acoustic characteristics in a gas turbine combustor. [Brief explanation of the drawings]

[0009] [Figure 1] 1 is a diagram illustrating a schematic configuration of a gas turbine including a gas turbine combustor according to an embodiment of the present disclosure. [Figure 2A]2 is an enlarged schematic partial cross-sectional view of the vicinity of an axially upstream end of a combustor liner in a gas turbine combustor according to one embodiment of the present invention that is provided in the gas turbine shown in FIG. 1. [Figure 2B] FIG. 2B is an enlarged view of part A in FIG. 2A. [Figure 3] FIG. 3 is a cross-sectional view taken along the line III-III in FIG. 2A. [Figure 4] FIG. 4 is a cross-sectional view taken along the line IV-IV in FIG. 2A. [Figure 5] FIG. 2B is a cross-sectional view taken along the arrows VV in FIG. 2A. [Figure 6A] FIG. 6 is a cross-sectional view taken along the line VI-VI in FIG. 2A. [Figure 6B] 6 is a cross-sectional view taken along the line VI-VI in FIG. 2A. DETAILED DESCRIPTION OF THE INVENTION

[0010] Hereinafter, several embodiments of the present disclosure will be described with reference to the accompanying drawings. However, the dimensions, materials, shapes, relative arrangements, etc. of components described as embodiments or shown in the drawings are merely illustrative examples and are not intended to limit the scope of the present disclosure. For example, expressions expressing relative or absolute arrangement such as "in a certain direction," "along a certain direction," "parallel," "orthogonal," "center," "concentric," or "coaxial" not only express such an arrangement exactly, but also express a state in which there is a relative displacement with a tolerance or an angle or distance to the extent that the same function is obtained. For example, expressions such as "identical," "equal," and "homogeneous" that indicate that something is in an equal state not only indicate a state of strict equality, but also indicate a state in which there is a tolerance or a difference to the extent that the same function is obtained. For example, expressions representing shapes such as a square shape or a cylindrical shape not only represent shapes such as a square shape or a cylindrical shape in the strict geometric sense, but also represent shapes including uneven portions, chamfered portions, etc., to the extent that the same effect can be obtained. On the other hand, the expressions "comprise," "include," "have," "includes," or "have" of one element are not exclusive expressions that exclude the presence of other elements.

[0011] A gas turbine combustor according to an embodiment of the present disclosure will be described with reference to FIG. FIG. 1 shows a schematic configuration of a gas turbine equipped with a gas turbine combustor according to an embodiment of the present disclosure. The gas turbine 1 shown in FIG. 1 includes an air compressor 110, a gas turbine combustor 100, and a turbine 180.

[0012] A gas turbine combustor 100 according to one embodiment includes a combustor liner (inner cylinder) 153, a liner flow sleeve (outer cylinder) 154, a transition piece 152, a transition piece flow sleeve 150, a burner 200, and a fuel system 22. Note that in Fig. 1, the burner 200 and the fuel system 22 are illustrated in a simplified manner.

[0013] 1 , an air compressor 110 is rotationally driven by a turbine 180, compresses air (intake air) drawn from the atmosphere via an intake section (not shown), generates high-pressure air (combustion air), and supplies the high-pressure air 120 to a gas turbine combustor 100. The gas turbine combustor 100 mixes the high-pressure air 120 supplied from the air compressor 110 with fuel supplied from a fuel system 22, combusts the air, generates high-temperature combustion gas 170, and supplies the high-temperature combustion gas 170 to the turbine 180.

[0014] That is, in the gas turbine 1 shown in FIG. 1 , high-pressure air 120, which is combustion air discharged from the air compressor 110, is introduced into the casing 140 from the diffuser 130, and flows from an air inlet hole 151 provided in the transition piece flow sleeve 150 of the gas turbine combustor 100 into a flow passage formed in the gap between the transition piece flow sleeve 150 and a transition piece 152 arranged inside the transition piece flow sleeve 150.

[0015] The high-pressure air 120 that has flowed into the flow path formed in this gap then flows through the flow path formed in the gap between the combustor liner 153 of the gas turbine combustor 100 and a liner flow sleeve 154 that is arranged concentrically with the combustor liner 153 on the outer periphery of the combustor liner 153, then reverses its flow, mixes with fuel introduced from the fuel system 22 and injected from multiple fuel nozzles 210 that constitute the cluster nozzle, and combusts in the combustion chamber 160 inside the combustor liner 153 to form a flame 156 and generate high-temperature, high-pressure combustion gas 170.

[0016] The high-temperature, high-pressure combustion gas 170 thus generated in the gas turbine combustor 100 flows down the transition piece 152 and is introduced into the turbine 180 .

[0017] In the turbine 180 that constitutes the gas turbine 1, the amount of work generated when the high-temperature, high-pressure combustion gas 170 introduced into the turbine 180 undergoes adiabatic expansion is converted into shaft rotational force by the turbine 180, thereby driving the generator 190 that is connected to the turbine 180 by a turbine shaft, and output is obtained from the generator 190.

[0018] The air compressor 110 and the generator 190 that constitute the gas turbine 1 are connected to the turbine 180 by a turbine shaft. However, the air compressor 110, the turbine 180, and the generator 190 do not have to be configured with a single turbine shaft, and may be configured with two or more turbine shafts.

[0019] Generally, gas turbines widely used in thermal power plants and the like have a configuration in which a plurality of gas turbine combustors are arranged radially around a turbine shaft.

[0020] In the following description, the direction along the central axis AXc of the gas turbine combustor 100 will be referred to as the axial direction of the gas turbine combustor 100, or simply as the axial direction. The direction in which the combustion gas 170 flows along the axial direction will be referred to as the axial downstream side, or simply as the downstream side. Direction The opposite direction is referred to as the axial upstream side, or simply as the upstream side. Note that in the gas turbine combustor 100 according to the one embodiment, a central axis line AXc of the gas turbine combustor 100 is a central axis line of the combustor liner 153 having, for example, a cylindrical shape. That is, the axial direction of the gas turbine combustor 100 is the axial direction of the combustor liner 153. In the following description, the circumferential direction of the combustor liner 153 will also be simply referred to as the circumferential direction, and the radial direction of the combustor liner 153 will also be simply referred to as the radial direction.

[0021] (Overview of the gas turbine combustor 100) The gas turbine combustor 100 according to one embodiment is a hydrogen-fired combustor. This makes it possible to prevent carbon dioxide from being discharged from the gas turbine combustor 100. The gas turbine combustor 100 according to an embodiment may be a gas turbine combustor that is capable of combusting hydrogen fuel and another fuel other than hydrogen fuel, and may combust, for example, natural gas fuel as the other fuel. The gas turbine combustor 100 according to an embodiment may be a gas turbine combustor that is capable of exclusively burning hydrogen fuel, exclusively burning natural gas fuel, or co-combusting hydrogen fuel and natural gas fuel.

[0022] 1 , in a gas turbine combustor 100 according to one embodiment, a burner 200 is arranged so as to be perpendicular to a central axis AXc of the gas turbine combustor 100, and is provided at an end portion on an upstream side in the axial direction of the combustor liner 153. In the gas turbine combustor 100 according to one embodiment, the burner 200 includes a fuel header 230, a plurality of fuel nozzles 210, and an air hole plate 25.

[0023] A gas turbine combustor 100 according to one embodiment is a combustor of a type known as a multi-cluster combustor. In the gas turbine combustor 100 according to the one embodiment, a plurality of air holes 250 are formed in an air hole plate 25. The plurality of fuel nozzles 210 correspond one-to-one to the plurality of air holes 250 formed in the air hole plate 25 arranged adjacently and downstream of the fuel nozzles 210 in the axial direction, and are arranged coaxially with the plurality of air holes 250.

[0024] The hydrogen fuel injected from the plurality of fuel nozzles 210 toward the plurality of air holes 250 formed in the air hole plate 25 is ejected into the combustion chamber 160 together with the combustion air supplied from the air compressor 110, where it is rapidly mixed and burned, forming the flame 156 as described above and generating high-temperature, high-pressure combustion gas 170.

[0025] (Details of the gas turbine combustor 100) FIG. 2A is a schematic partial cross-sectional view illustrating an enlarged vicinity of an axially upstream end portion of a combustor liner 153 in a gas turbine combustor 100 according to one embodiment of the present invention that is provided in the gas turbine 1 shown in FIG. 1. FIG. 2B is an enlarged view of part A in FIG. 2A. FIG. 3 is a cross-sectional view taken along the line III-III in FIG. 2A. FIG. 4 is a cross-sectional view taken along the line IV-IV in FIG. 2A. FIG. 5 is a cross-sectional view taken along the arrows VV in FIG. 2A. FIG. 6A is a cross-sectional view taken along the line VI-VI in FIG. 2A. FIG. 6B is a cross-sectional view taken along the line VI-VI in FIG. 2A, and shows another example of the cooling passage 312, which will be described later. Hereinafter, with reference to FIGS. 2A to 6B, the gas turbine combustor 100 according to one embodiment will be described in detail.

[0026] In the gas turbine combustor 100 according to one embodiment, the burner 200 includes an air hole plate outer periphery support 26 that supports the outer periphery of the air hole plate 25 . In the gas turbine combustor 100 according to one embodiment, the combustor liner 153 includes a combustion liner body 31 forming a combustion chamber 160, and a combustion liner fixing adapter 33 connected to an axially upstream end portion 31 u of the combustion liner body 31 by a welded portion 35. In the gas turbine combustor 100 according to one embodiment, the downstream portion of the air hole plate outer periphery support body 26 in the axial direction and the upstream end of the combustion liner fixing adapter 33 in the axial direction are connected by a welded portion 27 .

[0027] The gas turbine combustor 100 according to one embodiment includes the acoustic device 40 attached to the combustion liner body 31 of the combustor liner 153 . The acoustic device 40 has a housing 41 and an acoustic hole 43. The combustion liner body 31 of the combustor liner 153 has an area 311 covered by the housing 41, and at least one acoustic hole 43 is formed in the area 311. For example, a plurality of acoustic holes 43 are formed in the area 311, and each acoustic hole 43 has a circular cross-sectional shape. The acoustic hole 43 is a through hole that opens into the combustion chamber 160.

[0028] The housing 41 is disposed on the outer circumferential side of the combustion liner body 31 and defines an acoustic attenuation space 45 that communicates with the combustion chamber 160 via the acoustic hole 43. The housing 41 extends along the circumferential direction of the combustion liner body 31. The housing 41 is secured to the combustor liner 153 by, for example, welding.

[0029] In the gas turbine combustor 100 according to one embodiment, the acoustic device 40 attached to the combustor liner 153 detects combustion vibrations due to combustion vibrations. room Pressure fluctuations within 160 are damped. The acoustic device 40 according to one embodiment is an acoustic device known as an acoustic liner, and is capable of absorbing relatively high-frequency sounds caused by combustion vibrations. However, the acoustic device 40 may also be an acoustic damper that is capable of absorbing relatively low-frequency sounds caused by combustion vibrations. Furthermore, an acoustic damper (not shown) may be provided together with the acoustic device 40 according to one embodiment. In the gas turbine combustor 100 according to an embodiment, the plurality of acoustic devices 40, the acoustic devices having different frequencies of combustion oscillations to be attenuated by varying the heights of the housing 41 in the radial direction, may be disposed in different regions in the circumferential direction.

[0030] In the gas turbine combustor 100 according to one embodiment, at least one of the plurality of acoustic holes 43 may be provided immediately below in the axial direction of the air hole plate 25. In the example shown in Fig. 2A and Fig. 2B, all of the plurality of acoustic holes 43 are provided immediately below the air hole plate 25 in the axial direction. The area directly below the air hole plate 25 in the axial direction refers to, for example, the area upstream in the axial direction of the reaction zone of the flame 156. After combustion in the burner 200, a trace of the flame can be seen on the inner circumferential surface 153b of the combustor liner 153 (the inner circumferential surface 31b of the combustion liner body 31). The area directly below the air hole plate 25 in the axial direction refers to, for example, the area upstream in the axial direction of the flame trace.

[0031] In a gas turbine combustor called a multi-cluster combustor, such as the gas turbine combustor 100 according to one embodiment, the length of the flame is shorter than that of a conventional combustor that is not a multi-cluster combustor. of Therefore, in order to suppress combustion oscillation, it is important to place the acoustic device 40 having the acoustic damping space 45 directly below the burner 200 (air hole plate 25). However, in a gas turbine combustor such as the gas turbine combustor 100 according to an embodiment, the region 311 where the acoustic device 40 is disposed is likely to reach a relatively high temperature, and therefore it is necessary to achieve both cooling and acoustic characteristics. In particular, when hydrogen is used as the fuel as in the gas turbine combustor 100 according to an embodiment, which is a hydrogen-only combustor, or when a fuel containing a relatively large amount of hydrogen is used, the flame moves upstream in the axial direction as the combustion velocity increases. Therefore, a flame is formed directly below the burner 200, and the most upstream portion of the combustor liner 153 is exposed to the flame, making cooling difficult.

[0032] Thus, in the gas turbine combustor 100 according to one embodiment, cooling air is blown into the combustion chamber 160 from the acoustic holes 43 as will be described later, thereby suppressing the flame from entering the acoustic attenuation space 45 from the acoustic holes 43 and cooling the upstream portion of the combustor liner 153 which is likely to become relatively hot. In addition, in the gas turbine combustor 100 according to one embodiment, as will be described later, Burning Cooling air is blown into the combustion chamber 160 from the air passage 50 between the inner peripheral surface 153b of the combustor liner 153 and the outer peripheral surface 25a of the air hole plate 25, so that the most upstream portion of the combustor liner 153, which tends to become relatively hot, is film-cooled by the cooling air blown out from the air passage 50.

[0033] In the gas turbine combustor 100 according to one embodiment, at least one of the plurality of acoustic holes 43 is provided directly below the air hole plate 25 in the axial direction, whereby combustion oscillation can be effectively suppressed.

[0034] (Cooling by cooling air from acoustic hole 43) The gas turbine combustor 100 according to one embodiment includes an air supply passage 49 for supplying high-pressure air 120 , which is air flowing outside the combustor liner 153 , to the acoustic attenuation space 45 . 2A, 3, 6A, and 6B, the combustion liner body 31 of the combustor liner 153 has a plurality of cooling passages 312 formed at intervals in the circumferential direction and extending axially inside a wall 31W constituting the combustion liner body 31. The plurality of cooling passages 312 each have an inlet opening 312a that opens into the outer peripheral surface 31a of the combustion liner body 31 downstream in the axial direction from the housing 41, and an outlet opening 312b that opens into the outer peripheral surface 31a of the combustion liner body 31 facing the acoustic attenuation space 45, and form an air supply flow path 49.

[0035] In the example shown in FIG. 6A, the axially most upstream portion of the plurality of cooling passages 312 is the outlet opening 312b, which is located axially downstream of the axially upstream wall portion 41UW of the housing 41. In another example shown in FIG. 6B, among the multiple cooling passages 312, for one cooling passage (first cooling passage) 312A, the most upstream portion in the axial direction is the outlet opening 312b, which is located axially downstream of the axially upstream wall portion 41UW of the housing 41. 6B , the axially most upstream portions of some cooling passages (second cooling passages) 312B among the plurality of cooling passages 312 are located axially upstream of the axially upstream wall portion 41UW of the housing 41. That is, in the other example shown in FIG. 6B , the plurality of second cooling passages 312B extend axially upstream from inlet openings 312a located axially downstream of the axially downstream wall portion 41DW of the housing 41 to a position axially upstream of the axially upstream wall portion 41UW of the housing 41, and reach a region axially upstream of the region 311 in which the acoustic device 40 is disposed. The plurality of second cooling passages 312B then turn back axially downstream at turning portions 312c located in a region axially upstream of the region 311 in which the acoustic device 40 is disposed, and extend to an outlet opening 312b that opens into the outer circumferential surface 31a of the combustion liner body 31 and faces the acoustic attenuation space 45. The folded portion 312c is located downstream of the welded portion 35 in the axial direction.

[0036] 2A and 6A, the high-pressure air 120 flowing outside the combustor liner 153 flows into the multiple cooling passages 312 from the inlet opening 312a as shown by arrow a, and flows axially upstream. The high-pressure air 120 flowing through the multiple cooling passages 312 flows into the sound attenuation space 45 from the outlet opening 312b as shown by arrow b, while cooling the combustor liner 153 as cooling air. 6B , the high-pressure air 120 flowing outside the combustor liner 153 flows into the multiple cooling passages 312 from the inlet opening 312a as indicated by arrow a, and flows axially upstream. The high-pressure air 120 flowing through the multiple first cooling passages 312A flows into the acoustic attenuation space 45 from the outlet opening 312b as indicated by arrow b, while acting as cooling air to cool the combustor liner 153. The high-pressure air 120 flowing through the multiple second cooling passages 312B flows into the acoustic attenuation space 45 from the outlet opening 312b as indicated by arrow b, while acting as cooling air to cool the combustor liner 153 in the region 311 where the acoustic device 40 is disposed, and also flows into the acoustic attenuation space 45 from the outlet opening 312b as indicated by arrow b, while cooling the region axially upstream of the region 311. Thus, in the example shown in FIG. 6B, compared to the example shown in FIG. 6A, the area axially upstream of the area 311 in which the acoustic device 40 is arranged can be further cooled without increasing the flow rate of the cooling air circulating through the multiple cooling passages 312.

[0037] In a gas turbine combustor 100 according to one embodiment, a housing 41 has at least one housing through-hole 47 penetrating the housing 41, which constitutes an air supply passage 49. In the example shown in FIG. 2A , the housing 41 has a plurality of housing through-holes 47. The plurality of housing through-holes 47 are purge holes for directing the high-pressure air 120 into the sound-attenuating space 45 . As a result, the high-pressure air 120 flowing outside the combustor liner 153 flows into the sound-attenuating space 45 via the plurality of housing through-holes 47 as shown by arrows c in FIG. 2A.

[0038] 2A, when the cooling air (high-pressure air 120) that has flowed into the acoustic attenuation space 45 is ejected into the combustion chamber 160 through the multiple acoustic holes 43, it flows through the multiple acoustic holes 43 and cools the inner circumferential surfaces of the acoustic holes 43. In addition, the cooling air that has flowed into the combustion chamber 160 from the acoustic holes 43 film-cools the inner circumferential surface 153b of the combustor liner 153 (the inner circumferential surface 31b of the combustion liner main body 31).

[0039] In the gas turbine combustor 100 according to an embodiment, the high-pressure air 120 flowing along the outside of the combustor liner 153 can be circulated through the plurality of cooling passages 312 as cooling air, thereby enabling the combustor liner 153 to be effectively cooled. Furthermore, in the gas turbine combustor 100 according to an embodiment, the cooling air after flowing through the plurality of cooling passages 312 can be supplied to the acoustic attenuation space 45, and therefore can be blown into the combustion chamber 160 from the plurality of acoustic holes 43. This enables efficient use of the cooling air and suppresses flames from entering the acoustic attenuation space 45 from the plurality of acoustic holes 43.

[0040] In the gas turbine combustor 100 according to one embodiment, the high-pressure air 120 flowing outside the combustor liner 153 can be introduced into the acoustic attenuation space 45 from the housing through hole 47 and blown out into the combustion chamber 160 from the plurality of acoustic holes 43. This makes it possible to cool a region that is likely to become relatively hot by the cooling air blown out from the plurality of acoustic holes 43, and to suppress the intrusion of flames into the acoustic attenuation space 45 from the plurality of acoustic holes 43.

[0041] In the gas turbine combustor 100 according to an embodiment, the high-pressure air 120 (cooling air) supplied into the acoustic attenuation space 45 via the air supply passage 49 can be blown into the combustion chamber 160 from the at least one acoustic hole 43 provided immediately below the air hole plate 25 in the axial direction. This makes it possible to cool the region immediately below the air hole plate 25 in the axial direction, which is likely to become relatively hot, by the cooling air blown out from the at least one acoustic hole 43. Furthermore, in the gas turbine combustor 100 according to an embodiment, blowing out the cooling air from the plurality of acoustic holes 43 can suppress flames from entering the acoustic attenuation space 45 from the plurality of acoustic holes 43.

[0042] (Cooling by cooling air from air passage 50) In the gas turbine combustor 100 according to one embodiment, as illustrated in FIGS. 2A , 2B , 4 , and 5 , an inner circumferential surface 153 b of the combustor liner 153 and an outer circumferential surface 25 a of the air hole plate 25 are spaced apart in the radial direction, forming an annular space. An inner peripheral surface 153b of the combustor liner 153 and an outer peripheral surface 25a of the air hole plate 25 define, in order from the upstream side, an annular cavity 53 and an air passage 50. That is, the gas turbine combustor 100 according to the one embodiment includes the annular cavity 53 which is formed between an inner circumferential surface 153b of the combustor liner 153 (the inner circumferential surface 33b of the combustion liner fixing adapter 33) and the outer circumferential surface 25a of the air hole plate 25 and extends in the circumferential direction. A gas turbine combustor 100 according to one embodiment includes an air passage 50 that is provided between an inner peripheral surface 153b of a combustor liner 153 and an outer peripheral surface 25a of an air hole plate 25 and extends in the axial direction.

[0043] The annular cavity 53 according to one embodiment is formed so that the cross-sectional area thereof when viewed in the axial direction is larger than the cross-sectional area of ​​the air passage 50 .

[0044] As shown in Fig. 2B, the combustion cylinder fixing adapter 33 of the combustor liner 153 is formed with a plurality of adapter through holes 331. The adapter through holes 331 have inlet openings 331a that open to the outer peripheral surface 33a of the combustion cylinder fixing adapter 33 and outlet openings 331b that open to the inner peripheral surface 33b of the combustion cylinder fixing adapter 33. As shown in Fig. 4, the plurality of adapter through holes 331 are arranged circumferentially at intervals. Each of the outlet openings 331b faces the annular cavity 53.

[0045] The air passage 50 according to one embodiment is an annular air passage extending in the circumferential direction, and includes a first region 51 and a second region 52 located axially downstream of the first region 51. A radial height h2 of the second region 52 is greater than a radial height h1 of the first region 51. In the air passage 50 according to one embodiment, the radial height changes in a stepwise manner from the first region 51 to the second region 52. Therefore, in the air passage 50 according to one embodiment, the flow path area increases in a stepwise manner from the first region 51 to the second region 52. The air passage 50 is connected to an annular cavity 53 at an axially upstream end 50u of the air passage 50.

[0046] In the gas turbine combustor 100 according to one embodiment, high-pressure air 120 flowing outside the combustor liner 153 flows into the multiple adapter through-holes 331 from inlet openings 331a of the adapter through-holes 331, as indicated by arrows e in Fig. 2A , and then flows into the annular cavity 53 from outlet openings 331b. The high-pressure air 120 that has flowed into the annular cavity 53 flows into the air passage 50 while the circumferential deviation of the flow velocity within the annular cavity 53 is suppressed.

[0047] When the high-pressure air 120 that flows into the air passage 50 flows from the first region 51 to the second region 52, the flow path area increases in a step-like manner, causing the flow to become turbulent and pressure loss to occur, thereby suppressing the circumferential deviation of the flow velocity toward the downstream axial direction. The high-pressure air 120 flowing from the air passage 50 into the combustion chamber 160 serves as cooling air to perform film cooling on the region directly below the air hole plate 25 in the axial direction.

[0048] In the gas turbine combustor 100 according to an embodiment, the cooling air is blown into the combustion chamber 160 from the air passage 50 between the inner circumferential surface 153b of the combustor liner 153 and the outer circumferential surface 25a of the air hole plate 25, whereby a region directly below the air hole plate 25 in the axial direction, which is likely to become relatively hot, can be film-cooled by the cooling air blown out of the air passage 50. Furthermore, in the gas turbine combustor 100 according to an embodiment, the cooling air is blown out from the air passage 50 into the combustion chamber 160, whereby intrusion of the flame into the acoustic attenuation space through the plurality of acoustic holes 43 can be suppressed.

[0049] The gas turbine combustor 100 according to one embodiment includes the annular cavity 53 that is formed between the inner circumferential surface 153b of the combustor liner 153 (the inner circumferential surface 33b of the combustion liner fixing adapter 33) and the outer circumferential surface 25a of the air hole plate 25 and extends in the circumferential direction. This makes it possible to suppress the occurrence of a deviation in the circumferential direction in the flow velocity of the cooling air that is blown out of the air passage 50 and performs film cooling on the combustor liner 153. This enables the temperature difference in the circumferential direction of the combustor liner 153 to be suppressed.

[0050] In the gas turbine combustor 100 according to one embodiment, as described above, the radial height of the air passage 50 changes in a stepwise manner from the first region 51 to the second region 52. Therefore, even if there is a circumferential deviation in the flow velocity of the high-pressure air 120 (cooling air) flowing through the first region 51, the high-pressure air 120 (cooling air) is turbulent when it flows from the first region 51 into the second region 52, thereby making it possible to suppress the circumferential deviation in the flow velocity of the cooling air blown out of the air passage 50 and performing film cooling on the combustor liner 153.

[0051] In the gas turbine combustor 100 according to one embodiment, the welded portion 35 connecting the combustion liner fixing adapter 33 and the combustion liner body 31 is located upstream of a downstream end 50d of the air passage 50 in the axial direction. That is, the welded portion 35 is located upstream of the downstream end 25d of the air hole plate 25. This allows the cooling air flowing through the air passage 50 to effectively cool the welded portion 35, which is relatively susceptible to the effects of heat.

[0052] In the gas turbine combustor 100 according to one embodiment, the weld 35 is located within the axial extension range of the second region 52 . This allows the welded portion 35 to be cooled with high-pressure air 120 (cooling air) whose circumferential deviation in flow rate is suppressed as it flows from the first region 51 to the second region 52, thereby suppressing the circumferential deviation in temperature of the welded portion 35.

[0053] As described above, in the gas turbine combustor 100 according to one embodiment, both cooling and acoustic characteristics in the gas turbine combustor 100 can be achieved. Moreover, in the gas turbine 1 including the gas turbine combustor 100 according to one embodiment, it is possible to realize the gas turbine 1 including the gas turbine combustor 100 that achieves both cooling and acoustic characteristics.

[0054] The present disclosure is not limited to the above-described embodiments, but also includes modifications to the above-described embodiments and appropriate combinations of these modifications.

[0055] The contents described in each of the above embodiments can be understood, for example, as follows. (1) A gas turbine combustor 100 according to at least one embodiment of the present disclosure includes: a combustion liner (combustor liner 153) having a combustion chamber 160 therein and having a plurality of through holes (acoustic holes 43) opening to the combustion chamber 160 formed therein; a housing 41 disposed on an outer circumferential side of the combustion liner (combustor liner 153) and defining an acoustic attenuation space 45 communicating with the combustion chamber 160 via the through holes (acoustic holes 43); and an air hole plate 25 formed with a plurality of air holes 250 and positioned upstream of the combustion liner (combustor liner 153). A gas turbine combustor 100 according to at least one embodiment of the present disclosure includes a plurality of fuel nozzles 210 corresponding to the plurality of air holes 250, an air passage 50 provided between an inner circumferential surface 153b of the combustion liner (combustor liner 153) and an outer circumferential surface 25a of the air hole plate 25 and extending in the axial direction of the combustion liner (combustor liner 153), and an air supply flow path 49 for supplying air (high-pressure air 120) flowing outside the combustion liner (combustor liner 153) to an acoustic attenuation space 45. At least one of the plurality of through-holes (acoustic hole 43) is provided directly below the air hole plate 25 in the axial direction.

[0056] According to the above configuration (1), sound is blown through the air supply passage 49. Reduced sound The air (cooling air) supplied into the acoustic attenuation space 45 can be blown into the combustion chamber 160 from at least one through-hole (acoustic hole 43) provided directly below the air hole plate 25 in the axial direction. This allows areas that tend to become relatively hot to be cooled by the cooling air blown out from at least one through-hole (acoustic hole 43). Furthermore, according to the configuration (1) above, blowing out cooling air from multiple through-holes (acoustic holes 43) can prevent flames from entering the acoustic attenuation space 45 from multiple through-holes (acoustic holes 43). According to the above-mentioned configuration (1), by blowing cooling air into the combustion chamber 160 from the air passage 50 between the inner peripheral surface 153b of the combustion liner (combustor liner 153) and the outer peripheral surface 25a of the air hole plate 25, it is possible to perform film cooling on the above-mentioned region that is likely to become relatively hot with the cooling air blown out from the air passage 50. Furthermore, according to the above-mentioned configuration (1), by blowing cooling air into the combustion chamber 160 from the air passage 50, it is possible to suppress the flame from entering the acoustic attenuation space 45 through the multiple through-holes. According to the above configuration (1), at least one of the plurality of through holes (acoustic hole 43) is provided directly below the air hole plate 25 in the axial direction, so that combustion vibration can be effectively suppressed. That is, according to the above configuration (1), it is possible to achieve both cooling and acoustic characteristics in the gas turbine combustor 100.

[0057] (2) In some embodiments, in the configuration of (1) above, the combustion liner (combustor liner 153) may include a combustion liner body 31 that forms the combustion chamber 160, and a combustion liner fixing adapter 33 that is connected to an upstream end 31u of the combustion liner body 31 in the axial direction by a welded portion 35. The welded portion 35 may be located upstream of the downstream end of the air passage 50 in the axial direction.

[0058] According to the above configuration (2), the cooling air flowing through the air passage 50 can effectively cool the welded portion 35, which is relatively susceptible to the effects of heat.

[0059] (3) In some embodiments, in the configuration of (1) or (2) above, an annular cavity 53 may be formed between the inner peripheral surface 153b of the combustion liner (combustor liner 153) and the outer peripheral surface 25a of the air hole plate 25 and extend in the circumferential direction of the combustion liner (combustor liner 153). The air passage 50 may be connected to the annular cavity 53 at its upstream end.

[0060] According to the configuration (3) above, by providing the annular cavity 53, it is possible to suppress deviation in the circumferential direction of the combustion liner (combustor liner 153) in the flow velocity of the cooling air that is blown out from the air passage 50 and film-cools the combustion liner (combustor liner 153). Therefore, it is possible to suppress temperature differences in the circumferential direction of the combustion liner (combustor liner 153).

[0061] (4) In some embodiments, in any of the configurations (1) to (3) above, the air passage 50 may be an annular air passage 50 extending circumferentially of the combustion liner (combustor liner 153) and may include a first region 51 and a second region 52 located axially downstream of the first region 51, and the radial height h2 of the combustion liner (combustor liner 153) of the second region 52 may be greater than the radial height h1 of the first region 51.

[0062] According to the configuration (4) above, the radial height changes from the first region 51 to the second region 52, and therefore the flow of cooling air flowing through the air passage 50 is turbulent when it flows from the first region 51 into the second region 52. Therefore, even if there is a circumferential deviation in the flow velocity of the cooling air flowing through the first region 51, the circumferential deviation can be suppressed by the cooling air being turbulent when it flows from the first region 51 into the second region 52. This makes it possible to suppress the occurrence of a circumferential deviation in the flow velocity of the cooling air that is blown out of the air passage 50 and performs film cooling on the combustion liner (combustor liner 153).

[0063] (5) In some embodiments, in the configuration of (4) above, the combustion liner (combustor liner 153) may include a combustion liner body 31 that forms the combustion chamber 160, and a combustion liner fixing adapter 33 that is connected to an upstream end 31u of the combustion liner body 31 in the axial direction by a welded portion 35. The welded portion 35 may be located within an axial extension range of the second region 52.

[0064] According to the above configuration (5), the welded portion 35 can be cooled with cooling air whose flow velocity deviation in the circumferential direction is suppressed, so that the temperature deviation in the circumferential direction of the welded portion 35 can be suppressed.

[0065] (6) In some embodiments, in any of the configurations (1) to (5) above, the combustion liner (combustor liner 153) may have a plurality of cooling passages 312 extending axially inside the wall 31W constituting the combustion liner (combustor liner 153) and spaced apart along the circumferential direction of the combustion liner (combustor liner 153). The plurality of cooling passages 312 may have inlet openings 312a that open to the outer peripheral surface 153a (outer peripheral surface 31a of the combustion liner main body 31) of the combustion liner (combustor liner 153) downstream in the axial direction from the housing 41, and outlet openings 312b that open to the outer peripheral surface 153a (outer peripheral surface 31a of the combustion liner main body 31) of the combustion liner (combustor liner 153) facing the acoustic attenuation space 45, thereby forming the air supply flow path 49.

[0066] According to the configuration (6) above, the air (high-pressure air 120) flowing outside the combustion liner (combustor liner 153) can be circulated as cooling air through the multiple cooling passages 312, thereby effectively cooling the combustion liner (combustor liner 153). Furthermore, the cooling air that has flowed through the multiple cooling passages 312 can be supplied to the acoustic damping space 45, and can be blown out into the combustion chamber 160 from the multiple through-holes (acoustic holes 43). This allows for efficient use of the cooling air and also prevents flames from entering the acoustic damping space 45 through the multiple through-holes (acoustic holes 43).

[0067] (7) In some embodiments, in any of the configurations (1) to (6) above, the housing 41 may have a housing through-hole 47 that penetrates the housing 41 and forms the air supply passage 49 .

[0068] According to the configuration (7) above, the air (high-pressure air 120) flowing outside the combustion liner (combustor liner 153) can be introduced into the acoustic attenuation space 45 from the housing through-hole 47 and then blown out from the plurality of through-holes (acoustic holes 43) into the combustion chamber 160. This allows the cooling air blown out from the plurality of through-holes (acoustic holes 43) to cool the area that is likely to become relatively hot, and also prevents the flame from spreading through the plurality of through-holes. Hole (acoustic hole 43) Therefore, it is possible to prevent the sound from entering the sound attenuation space 45 from the outside.

[0069] (8) In some embodiments, in any of the configurations (1) to (7) above, the gas turbine combustor 100 may be a hydrogen-fired combustor.

[0070] According to the above configuration (8), it is possible to prevent carbon dioxide from being discharged from the gas turbine combustor 100.

[0071] (9) A gas turbine 1 according to at least one embodiment of the present disclosure includes a compressor (air compressor 110) that generates compressed air, a gas turbine combustor 100 having any of the configurations (1) to (8) described above, and a turbine 180 that is rotationally driven by combustion gas 170 generated by the gas turbine combustor 100.

[0072] According to the above configuration (9), it is possible to realize the gas turbine 1 including the gas turbine combustor 100 that achieves both cooling and acoustic characteristics. [Explanation of symbols]

[0073] 1. Gas turbine 25 Air hole plate 25a Outer surface 31 Combustion tube body 31a Outer surface 31u end 31W wall 33 Combustion tube fixing adapter 33b Inner surface 35 Welded section 40 Sound equipment 41 Housing 43 Acoustic hole 45 Acoustic attenuation space 47 Housing through hole 49 Air supply passage 50 Air passage 50u upstream end 50d downstream end 51 First area 52 Second area 53 Annular Cavity 100 Gas turbine combustor 110 Air compressor (compressor) 120 compressed air 153 Combustor liner (inner tube, combustion tube) 153b Inner surface 180 Turbine 250 air vents 312 Cooling passage 312a Entrance opening 312b Exit opening

Claims

1. a combustion liner having a combustion chamber therein and a plurality of through holes opening into the combustion chamber; a housing that is disposed on an outer circumferential side of the combustion liner and defines an acoustic attenuation space that communicates with the combustion chamber via the through hole; an air hole plate having a plurality of air holes formed therein and positioned upstream of the combustion liner; a plurality of fuel nozzles corresponding to the plurality of air holes; an air passage provided between an inner peripheral surface of the combustion liner and an outer peripheral surface of the air hole plate, the air passage extending in the axial direction of the combustion liner; an air supply passage for supplying air flowing outside the combustion liner to the acoustic attenuation space; Equipped with At least one of the plurality of through holes is provided directly below the air hole plate in the axial direction, the combustion liner includes a combustion liner body that forms the combustion chamber, and a combustion liner fixing adapter that is connected to an upstream end of the combustion liner body in the axial direction by a welded portion, the welded portion is located upstream of a downstream end of the air passage in the axial direction. Gas turbine combustor.

2. a combustion liner having a combustion chamber therein and a plurality of through holes opening into the combustion chamber; a housing that is disposed on an outer circumferential side of the combustion liner and defines an acoustic attenuation space that communicates with the combustion chamber via the through hole; an air hole plate having a plurality of air holes formed therein and positioned upstream of the combustion liner; a plurality of fuel nozzles corresponding to the plurality of air holes; an air passage provided between an inner peripheral surface of the combustion liner and an outer peripheral surface of the air hole plate, the air passage extending in the axial direction of the combustion liner; an air supply passage for supplying air flowing outside the combustion liner to the acoustic attenuation space; Equipped with At least one of the plurality of through holes is provided directly below the air hole plate in the axial direction, an annular cavity formed between an inner peripheral surface of the combustion liner and an outer peripheral surface of the air hole plate and extending in a circumferential direction of the combustion liner; Equipped with the air passage is connected to the annular cavity at an upstream end of the air passage; Gas turbine combustor.

3. a combustion liner having a combustion chamber therein and a plurality of through holes opening into the combustion chamber; a housing that is disposed on an outer circumferential side of the combustion liner and defines an acoustic attenuation space that communicates with the combustion chamber via the through hole; an air hole plate having a plurality of air holes formed therein and positioned upstream of the combustion liner; a plurality of fuel nozzles corresponding to the plurality of air holes; an air passage provided between an inner peripheral surface of the combustion liner and an outer peripheral surface of the air hole plate, the air passage extending in the axial direction of the combustion liner; an air supply passage for supplying air flowing outside the combustion liner to the acoustic attenuation space; Equipped with At least one of the plurality of through holes is provided directly below the air hole plate in the axial direction, the air passage is an annular air passage extending in a circumferential direction of the combustion liner and includes a first region and a second region located downstream of the first region in the axial direction, The height of the second region in the radial direction of the combustion liner is greater than the height of the first region in the radial direction. Gas turbine combustor.

4. the combustion liner includes a combustion liner body that forms the combustion chamber, and a combustion liner fixing adapter that is connected to an upstream end of the combustion liner body in the axial direction by a welded portion, The weld is located within the axial extension range of the second region. The gas turbine combustor according to claim 3 .

5. The combustion liner has a plurality of cooling passages formed at intervals along a circumferential direction of the combustion liner inside a wall that constitutes the combustion liner and extending along the axial direction, The plurality of cooling passages have inlet openings that open onto an outer peripheral surface of the combustion liner downstream of the housing in the axial direction, and outlet openings that open onto the outer peripheral surface of the combustion liner facing the acoustic attenuation space, and constitute the air supply flow path. The gas turbine combustor according to any one of claims 1 to 4.

6. The housing has a housing through-hole that penetrates the housing and forms the air supply passage. The gas turbine combustor according to any one of claims 1 to 4.

7. The gas turbine combustor is a hydrogen-fired combustor. The gas turbine combustor according to any one of claims 1 to 4.

8. a compressor for generating compressed air; The gas turbine combustor according to any one of claims 1 to 4; a turbine that is rotationally driven by combustion gas generated by the gas turbine combustor; A gas turbine comprising:

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