GAS TURBINE COMBUSTION CHAMBER AND GAS TURBINE

Abstract

A gas turbine combustion chamber according to at least one embodiment of the present disclosure comprises: a combustion cylinder containing a combustion chamber and having several through-holes open to the combustion chamber; a housing arranged on an outer circumferential side of the combustion cylinder, defining a sound-absorbing space that communicates with the combustion chamber via the through-hole; an air hole plate having several air holes and positioned on an upstream side of the combustion cylinder; several fuel nozzles corresponding to the several air holes; and an air duct provided between an inner circumferential surface of the combustion cylinder and an outer circumferential surface of the air hole plate, extending in an axial direction of the combustion cylinder.and an air supply flow path for supplying air flowing on an outside of the combustion cylinder to the sound attenuation chamber. At least one of the several through-holes is provided directly below the air hole plate in the axial direction.

Description

Gas turbine combustor and gas turbine

[0001] This application claims priority to 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.

[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).

[0003] Japanese Patent Application Laid-Open No. 2018-150912

[0004] A gas turbine combustor such as that described in Patent Document 1 is also referred to as a multi-cluster combustor. In a multi-cluster combustor such as that 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 arrange 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 area where the acoustic device is arranged 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.

[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 having a plurality of through holes formed therein opening to the combustion chamber; a housing disposed on an outer circumferential side of the combustion liner and defining an acoustic attenuation space communicating with the combustion chamber via the through holes; an air hole plate disposed on an upstream side of the combustion liner and having a plurality of air holes formed therein; a plurality of fuel nozzles corresponding to the plurality of air holes, respectively; an air passage provided between an inner circumferential surface of the combustion liner and an outer circumferential surface of the air hole plate and extending in an axial direction of the combustion liner; and an air supply flow path for supplying air flowing outside the combustion liner to the acoustic attenuation space, wherein at least one of the plurality of through holes is provided immediately 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 that generates compressed air; a gas turbine combustor having the configuration described in (1); and a turbine that is rotationally driven by combustion gas generated by the gas turbine combustor.

[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.

[0009] 2A is a diagram illustrating a schematic configuration of a gas turbine including a gas turbine combustor according to an embodiment of the present disclosure. FIG. 2B is a schematic partial cross-sectional view illustrating an enlarged vicinity of an axially upstream end portion of a combustor liner in a gas turbine combustor according to an embodiment of the present disclosure that is provided in the gas turbine illustrated in FIG. 1. FIG. 2C is an enlarged view of a portion A of FIG. 2A. FIG. 2A is a cross-sectional view taken along arrows III-III of FIG. 2A. FIG. 2B is a cross-sectional view taken along arrows IV-IV of FIG. 2A. FIG. 2C is a cross-sectional view taken along arrows V-V of FIG. 2A. FIG. 2A is a cross-sectional view taken along arrows VI-VI of FIG. 2A. FIG. 2C is a view equivalent to the cross-sectional view taken along arrows VI-VI of FIG. 2A.

[0010] Several embodiments of the present disclosure will be described below 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 not intended to limit the scope of the present disclosure and are merely illustrative examples. For example, expressions expressing relative or absolute arrangements, such as "in a certain direction," "along a certain direction," "parallel," "orthogonal," "center," "concentric," or "coaxial," not only strictly express such arrangements, but also express relative displacements with a tolerance or angle or distance to the extent that the same function is achieved. For example, expressions expressing the equality of things, such as "same," "equal," and "homogeneous," not only express strict equality, but also express tolerance or differences to the extent that the same function is achieved. For example, expressions expressing shapes such as a square or cylindrical shape not only express shapes such as a square or cylindrical shape in the strict geometric sense, but also express shapes including concave and convex portions, chamfered portions, etc., to the extent that the same effect is achieved. 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. 1. Fig. 1 shows a schematic configuration of a gas turbine including 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) 120, and supplies the high-pressure air 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 mixture, 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 an air compressor 110, is introduced into a casing 140 from a diffuser 130, and flows from an air inlet hole 151 provided in a transition piece flow sleeve 150 of the gas turbine combustor 100 into a flow passage formed in a gap between the transition piece flow sleeve 150 and a transition piece 152 disposed 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 a flow path formed in the gap between a 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 that is introduced from a fuel system 22 and injected from a plurality of fuel nozzles 210 that constitute a cluster nozzle, and combusts in a combustion chamber 160 inside the combustor liner 153 to form a flame 156, generating 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 a 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, a direction along a central axis line AXc of the gas turbine combustor 100 will be referred to as an axial direction of the gas turbine combustor 100 or simply as the axial direction. A direction in which the combustion gas 170 flows along the axial direction will be referred to as an axial downstream side or simply as the downstream side, and a direction opposite to the flow of the combustion gas 170 will be referred to as an axial upstream side or simply as the upstream side. Note that in the gas turbine combustor 100 according to an embodiment, the central axis line AXc of the gas turbine combustor 100 is, for example, a central axis line of a combustor liner 153 having a cylindrical shape. That is, the axial direction of the gas turbine combustor 100 is the axial direction of the combustor liner 153. Furthermore, in the following description, a circumferential direction of the combustor liner 153 will be simply referred to as the circumferential direction, and a radial direction of the combustor liner 153 will be simply referred to as the radial direction.

[0021] (Overview of Gas Turbine Combustor 100) The gas turbine combustor 100 according to an embodiment is a hydrogen-fired combustor. This makes it possible to prevent carbon dioxide from being emitted from the gas turbine combustor 100. Note that 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 performing the exclusive combustion of hydrogen fuel, the exclusive combustion of natural gas fuel, and the mixed combustion of 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 line 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 the 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 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 the gas turbine combustor 100 according to an embodiment that is provided in the gas turbine 1 illustrated in FIG. 1. FIG. 2B is an enlarged view of portion A in FIG. 2A. FIG. 3 is a cross-sectional view taken along arrows III-III in FIG. 2A. FIG. 4 is a cross-sectional view taken along arrows IV-IV in FIG. 2A. FIG. 5 is a cross-sectional view taken along arrows V-V in FIG. 2A. FIG. 6A is a cross-sectional view taken along arrows VI-VI in FIG. 2A. FIG. 6B is a view corresponding to the cross-sectional view taken along arrows VI-VI in FIG. 2A, and illustrates another example of a cooling passage 312, which will be described later. Hereinafter, details of the gas turbine combustor 100 according to an embodiment will be described with reference to FIGS. 2A to 6B.

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

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

[0029] In the gas turbine combustor 100 according to an embodiment, pressure fluctuations in the combustor 160 due to combustion oscillations are damped by the acoustic device 40 attached to the combustor liner 153. The acoustic device 40 according to the embodiment is a so-called acoustic liner and is capable of absorbing relatively high-frequency sounds caused by combustion oscillations. However, the acoustic device 40 may be an acoustic damper that can absorb relatively low-frequency sounds caused by combustion oscillations. Moreover, an acoustic damper (not shown) may be provided together with the acoustic device 40 according to the embodiment. Moreover, in the gas turbine combustor 100 according to an embodiment, a plurality of acoustic devices 40 whose radial heights of the housing 41 are different to damp combustion oscillations of different frequencies may be disposed in different circumferential regions.

[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 the air hole plate 25 in the axial direction. In the example shown in FIGS. 2A and 2B , all of the plurality of acoustic holes 43 are provided immediately below the air hole plate 25 in the axial direction. Note that, "immediately below the axial direction of the air hole plate 25" refers to, for example, a region axially upstream of a reaction zone of a flame 156. Note that, after combustion in the burner 200, a flame trace is observed on an inner circumferential surface 153b of the combustor liner 153 (inner circumferential surface 31b of the combustion liner body 31). "Immediately below the axial direction of the air hole plate 25" refers to, for example, a region axially upstream of this flame trace.

[0031] In a gas turbine combustor called a multi-cluster combustor, such as the gas turbine combustor 100 according to an embodiment, the flame length is shorter than in 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 arrange 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 arranged is likely to become relatively hot, and therefore it is necessary to achieve both cooling and acoustic characteristics. In particular, when using hydrogen as fuel, such as in the gas turbine combustor 100 according to an embodiment, which is a hydrogen-only combustor, or when using a fuel containing a relatively large amount of hydrogen, the flame moves upstream in the axial direction as the combustion velocity increases. Therefore, the 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 an embodiment, the 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-most portion of the combustor liner 153 which is likely to become relatively hot. Furthermore, 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 153 b of the combustor liner 153 and the outer circumferential surface 25 a of the air hole plate 25 as will be described later, thereby film-cooling the upstream-most portion of the combustor liner 153 which is likely to become relatively hot.

[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 with Cooling Air from Acoustic Hole 43) The gas turbine combustor 100 according to one embodiment includes an air supply passage 49 for supplying the high-pressure air 120, which is air flowing outside the combustor liner 153, to the acoustic attenuation space 45. Specifically, as illustrated in FIGS. 2A , 3 , 6A and 6B , the combustion liner body 31 of the combustor liner 153 is provided with a plurality of cooling passages 312 extending in the axial direction inside a wall 31W constituting the combustion liner body 31, the plurality of cooling passages 312 being formed at intervals in the circumferential direction. The plurality of cooling passages 312 each have an inlet opening 312a that opens in the outer circumferential surface 31a of the combustion liner body 31 downstream in the axial direction from the housing 41, and an outlet opening 312b that opens in the outer circumferential surface 31a of the combustion liner body 31 facing the acoustic attenuation space 45, thereby configuring the air supply passage 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 , for some cooling passages 312A (first cooling passages) among the plurality of cooling passages 312, the axially most upstream portion 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 , for other cooling passages 312 (second cooling passages) among the plurality of cooling passages 312, the axially most upstream portion is located axially upstream of the axially upstream wall portion 41UW of the housing 41. 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 turn-back portions 312c located in a region axially upstream of the region 311 in which the acoustic device 40 is disposed, and extend to outlet openings 312b that open into the acoustic attenuation space 45 on the outer circumferential surface 31a of the combustion liner body 31. The turn-back portions 312c are located axially downstream of the welded portion 35.

[0036] 2A and 6A, the high-pressure air 120 flowing outside the combustor liner 153 flows from the inlet opening 312a into the plurality of cooling passages 312 as indicated by arrow a, and then flows axially upstream. The high-pressure air 120 flowing through the plurality of cooling passages 312 flows into the acoustic attenuation space 45 as indicated by arrow b, while cooling the combustor liner 153. In the example shown in FIG. 6B, the high-pressure air 120 flowing outside the combustor liner 153 flows from the inlet opening 312a into the plurality of cooling passages 312 as indicated by arrow a, and then flows axially upstream. The high-pressure air 120 flowing through the plurality of first cooling passages 312A flows into the acoustic attenuation space 45 as indicated by arrow b, while cooling the combustor liner 153. The high-pressure air 120 flowing through the plurality of second cooling passages 312B acts as cooling air to cool the combustor liner 153 in the region 311 where the acoustic device 40 is disposed, and flows into the acoustic attenuation space 45 from the outlet opening 312b as shown by arrow b while cooling the region axially upstream of the region 311. In this way, in the example shown in Figure 6B, compared to the example shown in Figure 6A, it is possible to further cool the region axially upstream of the region 311 where the acoustic device 40 is disposed, without increasing the flow rate of the cooling air flowing through the plurality of cooling passages 312.

[0037] In the gas turbine combustor 100 according to one embodiment, the housing 41 has at least one housing through-hole 47 penetrating the housing 41 and constituting the 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 guiding the high-pressure air 120 to the acoustic attenuation space 45. As a result, the high-pressure air 120 flowing outside the combustor liner 153 flows into the acoustic attenuation space 45 through the plurality of housing through-holes 47 as indicated by arrow 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 acoustic holes 43, it flows through the 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 performs film cooling on the inner circumferential surface 153b of the combustor liner 153 (the inner circumferential surface 31b of the combustion liner body 31).

[0039] In the gas turbine combustor 100 according to the embodiment, the high-pressure air 120 flowing outside the combustor liner 153 can be circulated through the plurality of cooling passages 312 as cooling air, thereby enabling effective cooling of the combustor liner 153. Furthermore, in the gas turbine combustor 100 according to the 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 flames from entering 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 damping 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 a 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, the cooling air is blown out from the plurality of acoustic holes 43, thereby suppressing flame intrusion into the acoustic damping space 45 from the plurality of acoustic holes 43.

[0042] 2A , 2B , 4 , and 5 , in the gas turbine combustor 100 according to an embodiment, an inner circumferential surface 153 b of the combustor liner 153 and the outer circumferential surface 25 a of the air hole plate 25 are spaced apart in the radial direction to form an annular space. The inner circumferential surface 153 b of the combustor liner 153 and the outer circumferential surface 25 a of the air hole plate 25 define, in this order from the upstream side, the annular cavity 53 and the air passage 50. That is, the gas turbine combustor 100 according to an embodiment includes the annular cavity 53 that is formed between the inner circumferential surface 153 b of the combustor liner 153 (the inner circumferential surface 33 b of the combustion liner fixing adapter 33) and the outer circumferential surface 25 a 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 153 b of a combustor liner 153 and an outer peripheral surface 25 a 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 spaced apart from one another around the entire circumference. 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 downstream in the axial direction of the first region 51. The radial height h2 of the second region 52 is greater than the 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 the annular cavity 53 at an upstream end 50u of the air passage 50 in the axial direction.

[0046] In the gas turbine combustor 100 according to one embodiment, high-pressure air 120 flowing outside the combustor liner 153 flows into the plurality of adapter through-holes 331 from inlet openings 331a of the adapter through-holes 331, as indicated by arrows e in Figure 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 has flowed into the air passage 50 flows from the first region 51 to the second region 52, the flow path area increases stepwise, causing turbulence and pressure loss, thereby suppressing the circumferential deviation of the flow velocity toward the axial downstream side. The high-pressure air 120 that flows from the air passage 50 into the combustion chamber 160 serves as cooling air to film-cool the region immediately 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 153 b of the combustor liner 153 and the outer circumferential surface 25 a of the air hole plate 25, thereby enabling film cooling of a region directly below the air hole plate 25 in the axial direction, which region is likely to become relatively hot, by the cooling air blown out from 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, thereby enabling to suppress the flame from entering the acoustic attenuation space through the plurality of acoustic holes 43.

[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 flow velocity of the cooling air that is blown from 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 to perform 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 main body 31 is located upstream of a downstream end 50 d of the air passage 50 in the axial direction. That is, the welded portion 35 is located upstream of a downstream end 25 d of the air hole plate 25. This allows the welded portion 35, which is relatively susceptible to the effect of heat, to be effectively cooled by the cooling air flowing through the air passage 50.

[0052] In the gas turbine combustor 100 according to the one embodiment, the welded portion 35 is located within an axial extending range of the second region 52. Thus, the welded portion 35 can be cooled by the high-pressure air 120 (cooling air) in which the circumferential deviation of the flow velocity is suppressed while flowing from the first region 51 into the second region 52, thereby suppressing the circumferential deviation of the temperature of the welded portion 35.

[0053] As described above, it is possible to achieve both cooling and good acoustic characteristics in the gas turbine combustor 100 according to an embodiment. Moreover, it is possible to realize the gas turbine 1 including the gas turbine combustor 100 according to an embodiment, whereby the gas turbine 1 including the gas turbine combustor 100 achieving both cooling and good acoustic characteristics can be realized.

[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 the above embodiments may be understood as follows, for example. (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 that is provided between an inner circumferential surface 153 b of the combustion liner (combustor liner 153) and an outer circumferential surface 25 a of the air hole plate 25 and extends 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 configuration (1) above, the air (cooling air) supplied into the acoustic damping space 45 via the air supply passage 49 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 the cooling air blown out from at least one through-hole (acoustic hole 43) to cool the area that is likely to become relatively hot. 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 damping space 45 through the multiple through-holes (acoustic holes 43). According to the configuration (1) above, blowing out cooling air into the combustion chamber 160 from the air passage 50 between the inner circumferential surface 153b of the combustion liner (combustor liner 153) and the outer circumferential surface 25a of the air hole plate 25 allows film cooling of the area that is likely to become relatively hot with the cooling air blown out from the air passage 50. Furthermore, according to the above-described configuration (1), by blowing cooling air from the air passage 50 into the combustion chamber 160, it is possible to suppress the flame from entering the acoustic attenuation space 45 through the plurality of through holes. According to the above-described 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 oscillation can be effectively suppressed. That is, according to the above-described 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 above configuration (3), the provision of the annular cavity 53 can 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), thereby suppressing 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 to 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 turbulence of the cooling air when it flows from the first region 51 to the second region 52. This makes it possible to suppress the circumferential deviation in the flow velocity of the cooling air blown out of the air passage 50 and used for film cooling of 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 circumferential deviation in flow velocity is suppressed, so that the circumferential deviation in temperature 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 a wall 31W constituting the combustion liner (combustor liner 153) and spaced apart in a circumferential direction of the combustion liner (combustor liner 153). The plurality of cooling passages 312 may have inlet openings 312a that open to an outer circumferential surface 153a (outer circumferential 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 circumferential surface 153a (outer circumferential surface 31a of the combustion liner main body 31) of the combustion liner (combustor liner 153) facing the acoustic attenuation space 45, thereby constituting 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 into the combustion chamber 160 through 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 flow path 49 .

[0068] According to the configuration of (7) above, the air (high-pressure air 120) flowing outside the combustion liner (combustor liner 153) can be introduced into the acoustic damping 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 makes it possible to cool areas that tend to become relatively hot with the cooling air blown out from the plurality of through-holes (acoustic holes 43), and also to prevent flames from entering the acoustic damping space 45 from the plurality of through-holes (acoustic holes 43).

[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 described above in (1) to (8); 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.

[0073] REFERENCE SIGNS LIST 1 Gas turbine 25 Air hole plate 25a Outer peripheral surface 31 Combustion liner body 31a Outer peripheral surface 31u End 31W Wall 33 Combustion liner fixing adapter 33b Inner peripheral surface 35 Welded portion 40 Acoustic device 41 Housing 43 Acoustic hole 45 Acoustic attenuation space 47 Housing through hole 49 Air supply flow path 50 Air passage 50u Upstream end 50d Downstream end 51 First region 52 Second region 53 Annular cavity 100 Gas turbine combustor 110 Air compressor (compressor) 120 High-pressure air 153 Combustor liner (inner liner, combustion liner) 153b Inner peripheral surface 180 Turbine 250 Air hole 312 Cooling passage 312a Inlet opening 312b Outlet opening

Claims

1. A gas turbine combustor comprising: a combustion liner having a combustion chamber therein and having a plurality of through holes formed therein that open to the combustion chamber; a housing arranged on the outer periphery of the combustion liner and defining an acoustic attenuation space that communicates with the combustion chamber via the through holes; an air hole plate arranged on the upstream side of the combustion liner and having a plurality of air holes formed therein; a plurality of fuel nozzles corresponding to each of the plurality of air holes; an air passage arranged between an inner circumferential surface of the combustion liner and an outer circumferential surface of the air hole plate, extending in an axial direction of the combustion liner; and an air supply flow path for supplying air flowing outside the combustion liner to the acoustic attenuation space, wherein at least one of the plurality of through holes is arranged directly below the air hole plate in the axial direction.

2. The gas turbine combustor according to claim 1, wherein the combustion liner includes a combustion liner main body that forms the combustion chamber, and a combustion liner fixing adapter that is connected to an upstream end of the combustion liner main body in the axial direction by a welded portion, and the welded portion is located upstream of the downstream end of the air passage in the axial direction.

3. The gas turbine combustor according to claim 1 or 2, further comprising: 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; and the air passage is connected to the annular cavity at an upstream end of the air passage.

4. The gas turbine combustor according to claim 1 or 2, wherein 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, and a height of the second region in the radial direction of the combustion liner is greater than a height of the first region in the radial direction.

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

6. The gas turbine combustor according to claim 1 or 2, wherein the combustion liner has a plurality of cooling passages formed at intervals along the circumferential direction of the combustion liner, the cooling passages extending along the axial direction inside a wall constituting the combustion liner, and the plurality of cooling passages have inlet openings that open into the outer circumferential surface of the combustion liner downstream in the axial direction from the housing, and outlet openings that open into the outer circumferential surface of the combustion liner facing the acoustic attenuation space, and form the air supply flow path.

7. The gas turbine combustor according to claim 1 or 2, wherein the housing has a housing through-hole that penetrates the housing and forms the air supply passage.

8. The gas turbine combustor according to claim 1 or 2, wherein the gas turbine combustor is a hydrogen-fired combustor.

9. A gas turbine comprising: a compressor that generates compressed air; the gas turbine combustor according to claim 1 or 2; and a turbine that is rotationally driven by combustion gas generated by the gas turbine combustor.

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