JP7149909B2 - gas turbine combustor - Google Patents

Description

The present invention relates to the construction of gas turbine combustors.

As a means of reducing emissions of carbon dioxide (CO ₂ ), which is a cause of global warming, in thermal power plants, improvements in power generation efficiency and active use of fuels other than fossil fuels, such as hydrogen, are being considered. ing. Raising the turbine inlet temperature of gas turbine power generation equipment is effective for improving power generation efficiency.

However, as the temperature of gas turbines rises, the amount of nitrogen oxide (NOx) emissions, which are environmental pollutants, increases. There is a demand for a low NOx combustion system for gas turbines that can handle hydrogen-containing fuels.

Low NOx combustion methods for gas turbine combustors are generally divided into diffusion combustion and premixed combustion. In the diffusion combustion method, fuel is injected directly into the combustion chamber and the fuel and air are mixed in the combustion chamber. A flame is formed from the area where the flame is drawn. Therefore, there is no possibility of backflow of flame to the upstream side of the combustion chamber or self-ignition within the fuel supply system, and combustion stability can be ensured.

However, since a flame is formed in a region in which fuel and air are mixed in a stoichiometric ratio within the combustion chamber, a high-temperature flame is formed locally. In this localized high-temperature region, NOx emissions are high, and it is necessary to inject inert media such as nitrogen, water, and steam to reduce NOx emissions. may decline.

On the other hand, the premixed combustion method is a combustion method in which fuel and air are mixed in advance and supplied to the combustion chamber. However, when the temperature of the combustion air rises as the temperature of the gas turbine rises, the temperature distribution in the combustion chamber tends to become uniform and the combustion state tends to become unstable, causing pressure fluctuations in the combustion chamber called combustion oscillation. Sometimes. The occurrence of combustion oscillation means that the combustor structure is subject to pressure fluctuations and consumes its life, leading to a decrease in reliability of the combustor.

Therefore, when the premixed combustion method is employed, the diffusion combustion method, which is excellent in combustion stability, is often used in combination with the premixed combustion method.

By the way, in order to meet the demand for further reduction of NOx, there are cases where the ratio of premixed combustion is increased when diffusion combustion and premixed combustion are used together, or all premixed combustion is used. In order to suppress the increase in the amplitude of pressure fluctuation due to combustion vibration, it is effective to install an acoustic liner that forms an acoustic space for damping pressure fluctuation due to combustion vibration on the outer peripheral surface of the liner that forms the combustion chamber.

The acoustic liner guides pressure waves into the acoustic space through multiple pressure wave introduction holes provided on the inner wall of the liner. It is composed of purge air holes that introduce purge air into the acoustic space to prevent flames from entering the space.

As an example, Patent Document 1 discloses a structure in which an acoustic space is provided on the outer peripheral surface of a cylindrical member (combustion tube) that forms a combustion chamber.

WO2013/077394

However, in the acoustic liner disclosed in Patent Document 1, a specific amount of purge air is required from the pressure wave introduction hole (through hole group 16) in order to prevent the flame from flowing back into the acoustic space. of air may decrease and NOx emissions may increase.

In addition, in the case of gas turbines using hydrogen-containing fuel, hydrogen has a higher adiabatic flame temperature in the stoichiometric mixture ratio than natural gas, so NOx emissions increase when combined with diffusion combustion.

SUMMARY OF THE INVENTION Accordingly, it is an object of the present invention to provide a gas turbine combustor capable of damping pressure fluctuations due to combustion oscillations and suppressing an increase in NOx emissions due to a reduction in combustion air.

In order to solve the above problems, the present invention provides a combustion chamber for combusting a mixture of fuel and air, a combustor liner forming the combustion chamber, and a combustor liner positioned upstream of the combustion chamber to form a flame. a spring seal disposed between the outer peripheral surface of the burner and the inner peripheral surface of the combustor liner at the fitting portion between the burner and the combustor liner; and the spring seal downstream of the spring seal. an annular chamber installed in the combustor liner and forming an acoustic space, the annular chamber having a purge air hole on a side surface of the spring seal for guiding leaked air that has passed through the spring seal to the acoustic space; It is characterized by having a plurality of pressure wave introduction holes that communicate the combustion chamber and the acoustic space.

Advantageous Effects of Invention According to the present invention, it is possible to realize a gas turbine combustor capable of attenuating pressure fluctuations due to combustion oscillation and suppressing an increase in NOx emissions due to a reduction in combustion air.

This makes it possible to improve the reliability of the gas turbine combustor and reduce NOx, contributing to improved power generation efficiency and reduced NOx emissions in thermal power plants.

Problems, configurations, and effects other than those described above will be clarified by the following description of the embodiments.

It is a sectional view of a combustor burner part concerning Example 1 of the present invention. 1 is a schematic configuration diagram of a gas turbine plant provided with a gas turbine combustor according to Embodiment 1 of the present invention; FIG. It is a sectional view of a combustor burner part in a conventional example. FIG. 5 is a cross-sectional view of a combustor burner section according to Embodiment 2 of the present invention; FIG. 5 is a front view of an air hole plate according to Example 2 of the present invention; It is a sectional view of a combustor burner part concerning Example 3 of the present invention. FIG. 5 is a front view of an air hole plate according to Example 3 of the present invention; FIG. 10 is a front view of an air hole plate and a liner according to Example 3 of the present invention;

Hereinafter, embodiments of the present invention will be described with reference to the drawings. In addition, in each drawing, the same configurations are denoted by the same reference numerals, and detailed descriptions of overlapping portions are omitted.

A gas turbine combustor according to a first embodiment of the present invention will be described with reference to FIGS. 1 to 3. FIG. FIG. 3 is a cross-sectional view of a combustor burner portion of a conventional gas turbine combustor shown as a comparative example to facilitate understanding of the present invention.

FIG. 2 is a schematic configuration diagram of a gas turbine plant 1 equipped with a gas turbine combustor 3 of this embodiment. In the gas turbine plant 1 , compressed air 102 compressed by the compressor 2 flows into the casing 13 after passing through the diffuser 9 . Compressed air 102 flowing into the casing 13 passes through the flow sleeve 14 and flows between the outer cylinder 10 and the liner 12 .

A portion of the compressed air 102 enters the combustion chamber 5 as cooling air 103 for the liner 12 . The compressed air 102 that has passed between the outer cylinder 10 and the liner 12 is divided into combustion air 104 that passes through the diffusion burner 15 and the premixed burner 16, and leak air 105 that passes through the spring seal 26 (reference numeral 105 in FIG. 1). distributed. Combustion air 104 is mixed with fuel 100 (reference numeral 100 in FIG. 1) ejected from combustion nozzle 17 and fuel nozzle 22 and combusted to become high-temperature and high-pressure combustion gas 110 in combustion chamber 5, and turbine 4 is driven by fluid force. and take out the rotational power of the turbine 4 as electric power.

In the combustor 3 of this embodiment, the diffusion burner 15 is provided on the central axis of the combustion chamber 5, and the combustion air 104 swirled by the swirler 18 and the fuel 100 ejected from the fuel ejection holes 17 are mixed. , forming a diffusion flame downstream of the diffusion burner 15 .

A plurality of premixed burners 16 are provided on the outer periphery of the diffusion burner 15 , and the fuel 100 ejected from the fuel nozzle 22 is mixed with the combustion air 104 to form a sufficiently mixed air-fuel mixture, which is used as a flame stabilizer 19 . A premixed flame is formed downstream.

Since the premixed flame receives heat energy from the diffusion flame and can stably burn in the combustion chamber 5, low NOx combustion is possible.

A fuel supply system 201 and a fuel supply system 202 for supplying the fuel 100 to the fuel ejection hole 17 and the fuel nozzle 22 are branched from the fuel supply system 200 having the fuel cutoff valve 60 . The fuel supply systems 201 and 202 are provided with fuel pressure control valves 61a and 62a, respectively, so that the supply pressure of the supplied fuel 100 can be individually controlled. In addition, downstream thereof, fuel flow control valves 61b and 62b are provided, respectively.

The combustor 3 of this embodiment includes a plurality of fuel nozzles 22 each connected to a fuel header 23 (reference numeral 23 in FIG. 1) that distributes the fuel 100 . A fuel header 23 is provided inside the end cover 7 , and fuel 100 is supplied to the fuel header 23 from fuel supply systems 201 and 202 . Although the fuel 100 is distributed to two systems in this embodiment, it may be distributed to more systems.

By distributing the fuel supply system to a plurality of systems in this way, the degree of freedom of operation (combustion control) can be expanded by increasing the number of systems. The combustor 3 of this embodiment can be applied to many mixed gas fuels including liquefied natural gas (LNG) by controlling the fuel distribution.

FIG. 1 shows a sectional view of the combustor burner portion of this embodiment. A feature of this embodiment is that an annular chamber 27 having a rectangular cross section is arranged in the liner 12 on the downstream side of the spring seal 26 , and the radial height of the annular chamber 27 covers the spring seal 26 and is located on the outer peripheral side of the liner 12 . , a purge air hole 91 is provided on the side surface of the annular chamber 27 on the spring seal 26 side, and a plurality of pressure wave introduction holes 90 are provided on the combustion chamber 5 side. .

FIG. 3 shows a sectional view of the combustor upstream side of the conventional example. The effects of this embodiment will be described in comparison with the cross-sectional structure on the upstream side of the combustor of the conventional example. As shown in FIG. 3, the cross-sectional structure of the upstream side of the combustor of the conventional example has a rectangular annular cover 29 with a substantially U-shaped cross section so as to form an acoustic space 92 on the outer peripheral surface 12b of the liner 12. A plurality of pressure wave introduction holes 90 are provided in the liner 12 formed with 92 , and a plurality of purge air holes 91 are provided in the outer peripheral wall of the rectangular annular lid 29 . The combustor structure and fuel system of the conventional example are the same as those of the present embodiment except for the features described above.

In both the combustor of the present embodiment and the conventional example, when the amplitude of pressure fluctuation due to the occurrence of combustion vibration increases, the pressure wave generated by the combustion vibration is guided from the pressure wave introduction hole 90 to the acoustic space 92, Pressure fluctuations in a specific frequency band according to the radial height L can be attenuated.

In the conventional example shown in FIG. 3, a purge air hole 91 is provided in the rectangular annular cover 29, part of the combustion air 104 is introduced into the acoustic space 92 as the purge air 106, and the purge air 106 is introduced into the combustion chamber 5 through the pressure wave introduction hole 90. The flame is prevented from entering the acoustic space 92 by flowing the .

In order to prevent flames from entering the acoustic space 92, the purge air 106 must always flow at a constant flow rate or higher. As a result, the amount of combustion air 104 flowing to the premixed burner decreases, and the fuel concentration relatively increases, so local high-temperature regions are likely to be formed, and NOx emissions may increase.

Furthermore, since the effect of damping pressure fluctuations in the acoustic space 92 increases as the number of pressure wave introduction holes 90 increases, it is conceivable that it becomes difficult to reduce NOx emissions when dealing with larger pressure fluctuations. .

On the other hand, in this embodiment of FIG. 1, the acoustic space 92 is formed by the annular chamber 27 installed in the liner 12 downstream of the spring seal 26, and the purge air hole 91 is provided on the side surface of the annular chamber 27 on the spring seal 26 side. , the leakage air 105 passing through the spring seal 26 can be guided to the acoustic space 92 as the purge air 106, and the purge air 106 can be flowed from the pressure wave introduction hole 90 to the combustion chamber 5, so that the flame can be prevented from entering the acoustic space 92. can be prevented.

Therefore, pressure fluctuations due to combustion vibration can be damped in the acoustic space 92 without reducing the combustion air 104 by installing the purge air holes 91 as in the conventional example of FIG. 3, and NOx emissions can be reduced. It becomes possible.

Furthermore, the purge air 106 jetted from the pressure wave introduction hole 90 can lower the local flame temperature of the premixed flame formed by the premixed burner 16 (reference numeral 16 in FIG. liner 12 can be reduced.

In this embodiment, it is assumed that the purge air holes 91 for guiding the leak air 105 to the acoustic space 92 are provided in a plurality of circular holes in the circumferential direction of the liner 12. You can expect results.

Furthermore, in this embodiment, since the annular chamber 27 has a shape protruding toward the outer circumference of the liner 12, the setting range of the radial height L of the acoustic space 92 can be expanded, and the frequency band that can be attenuated in the acoustic space 92 can be increased. can be selected more widely.

Moreover, the annular chamber 27 is preferably provided so that the outermost peripheral surface of the burner 8 and the inner peripheral surface of the annular chamber 27 (side surface on the side of the combustion chamber 5) are substantially flush with each other. If the inner peripheral surface of the annular chamber 27 (the side surface on the side of the combustion chamber 5) protrudes excessively toward the combustion chamber 5, there is a risk of thermal deformation and shortening of the service life.

As described above, the gas turbine combustor of this embodiment includes the combustion chamber 5 for burning the mixture of the fuel 100 and air (combustion air 104), the combustor liner 12 forming the combustion chamber 5, A burner 8 which is located upstream of the combustion chamber 5 and forms a flame, and is arranged between the outer peripheral surface of the burner 8 and the inner peripheral surface of the combustor liner 12 at the fitting portion between the burner 8 and the combustor liner 12 . and an annular chamber 27 located in the combustor liner 12 downstream of the spring seal 26 and defining an acoustic space 92, the annular chamber 27 having the spring seal 26 on the spring seal side. It has a purge air hole 91 that guides the passed leak air 105 to the acoustic space 92 and has a plurality of pressure wave introduction holes 90 that communicate the combustion chamber 5 and the acoustic space 92 .

Also, the annular chamber 27 is arranged downstream of the spring seal 26 so as to cover the spring seal 26 .

Further, the annular chamber 27 is arranged so as to protrude from the combustor liner 12 to the outer peripheral side.

The purge air holes 91 are circular holes provided in the circumferential direction of the combustor liner 12 or annular slits provided in the circumferential direction of the combustor liner 12 .

The annular chamber 27 is provided so that the outermost peripheral surface of the burner 8 and the inner peripheral surface of the annular chamber 27 (side surface on the side of the combustion chamber 5) are substantially flush with each other.

The burner 8 is composed of a diffusion burner 15 arranged on the central axis of the combustion chamber 5 and a plurality of premixed burners 16 arranged around the diffusion burner 15 .

As in the present embodiment, pressure fluctuations caused by combustion oscillations can be damped in the acoustic space 92, and flames can be prevented from entering the acoustic space 92 without reducing the combustion air 104. Therefore, NOx emissions can be reduced at the same time. can be reduced.

Although the effect of reducing NOx emissions is reduced, an embodiment in which purge air holes 91 are provided on the outer peripheral surface of the annular chamber 27 as in the conventional example of FIG. 3 is also within the scope of the present invention. Needless to say.

A gas turbine combustor according to a second embodiment of the present invention will be described with reference to FIGS. 4 and 5. FIG. FIG. 4 is a sectional view showing the combustor burner section of this embodiment.

A feature of the present embodiment is that a plurality of fuel nozzles 22 and a plurality of air holes 21 are arranged coaxially on the upstream side of the combustion chamber 5, and a multi-hole nozzle for supplying fuel 100 and combustion air 104 to the combustion chamber 5 as coaxial flows. This is because the burner 8 is of the coaxial jet type.

The point that the annular chamber 27 is installed downstream of the spring seal 26 of the liner 12, and the configuration and operation method of the gas turbine plant 1 having the gas turbine combustor of this embodiment are basically the same as those of the first embodiment. is.

In this combustor, the fuel 100 and the combustion air 104 are dispersed and supplied as coaxial flows, thereby rapidly promoting mixing and reducing NOx emissions.

In addition, by shortening the mixing distance, it is possible to reduce the possibility that the fuel self-ignites during mixing of the fuel 100 and the combustion air 104, and it is possible to prevent the flame from flowing back to the mixing area. It can improve the reliability of things. In particular, in the case of a gas turbine using a hydrogen-containing fuel, hydrogen has a low ignition energy and a high combustion speed, so it is highly likely that flame reverse flow and self-ignition will occur, so it is effective.

In addition, hydrogen has a higher adiabatic flame temperature at a stoichiometric mixture ratio than natural gas, and the amount of NOx emissions in diffusion combustion tends to increase. is mixed, low NOx combustion is possible.

Therefore, in this embodiment, hydrogen-containing fuels such as coke oven gas, refinery off-gas, and coal gasification gas can be used as the fuel 100, and many gas fuels such as liquefied natural gas (LNG) can also be used.

FIG. 5 shows a front view of the air hole plate 20 arranged on the upstream side of the combustion chamber 5 of FIG. A plurality of air holes 21 are arranged concentrically around the center axis of the air hole plate 20, and in FIG. ing. The central axis of the air hole 21 is inclined in the pitch circumferential direction of each row, and a swirling angle is imparted so that the jetted combustion air 104 swirls around the central axis of the air hole plate 20 .

Further, as shown in FIG. 4, the air hole plate 20 is coaxial with the liner 12, so that the combustion air 104 is swirled and introduced into the combustion chamber 5, thereby swirling around the central axis of the combustion chamber 5. acts to form a circulation flow (circulation vortex) and stabilize the flame.

As described above, in the gas turbine combustor of this embodiment, the burner 8 is positioned upstream of the combustion chamber 5 and includes an air hole plate 20 having a plurality of air holes 21 concentrically arranged in a plurality of rows. , has a plurality of fuel nozzles 22 arranged coaxially with the air holes 21 .

By installing the annular chamber 27 in the liner 12 of the combustor equipped with the multi-hole coaxial jet burner as in this embodiment, the following effects can be obtained in addition to the same effects as in the first embodiment.

In this embodiment, leaked air 105 that has passed through the spring seal 26 is ejected as purge air 106 from the pressure wave introduction hole 90 into the combustion chamber 5. The air-fuel mixture with has a lower fuel concentration than the air-fuel mixture ejected from the other air holes 21, and can form a relatively low temperature region when burning in the combustion chamber 5, so NOx emissions can be reduced. can be reduced.

A gas turbine combustor according to a third embodiment of the present invention will be described with reference to FIGS. 6 to 8. FIG. FIG. 6 is a sectional view showing the combustor burner section of this embodiment.

The configuration and operating method of the gas turbine plant 1 having a gas turbine combustor of this embodiment and the installation of the annular chamber 27 downstream of the spring seal 26 of the liner 12 are the same as those of the second embodiment.

However, in order to cope with a larger power generation output and various operation modes, a multi-hole coaxial jet type burner 8 consisting of three rows of air holes 21 and fuel nozzles 22 arranged substantially coaxially shown in Embodiment 2 is installed in the center. It has a structure in which one is arranged and six are arranged around it.

In addition, in the combustor shown in this embodiment, the starting fuel nozzle 24 is installed in the central part of the burner 8 provided in the center, and the fuel system is divided for each burner 8, so that the operating state of the gas turbine can be adjusted. This differs from the first embodiment in that the fuel distribution for each burner 8 can be individually controlled accordingly.

FIG. 7 is a front view of the air hole plate 20 of FIG. 6 as seen from the combustion chamber 5 side. As shown in FIG. 7, in this embodiment, one air hole plate 20 is constructed by arranging a plurality of air hole plates 20 of the second embodiment (FIG. 5). That is, the combustor of this embodiment includes one central burner 32 positioned at the center of the combustor 3 and six peripheral burners 33 positioned outside the central burner 32 .

The central axis of each air hole 21 is inclined in the pitch circumferential direction of each row, and the flow of combustion air 104 that has passed through the air holes 21 spirally swirls downstream of the air holes 21 to form a swirling flow. . Also, as shown in FIG. 6, the central burner 32 and the peripheral burner 33 are connected to three fuel systems 203, 204 and 205 so that the fuel flow rate can be controlled independently. Although the number of peripheral burners 33 is six in this embodiment, it is desirable that there are three or more peripheral burners 33 concentrically with respect to the central burner 32 .

A central burner fuel system 203 and a gas turbine start-up fuel system 206 are connected to the central burner 32, and are mainly used for starting the gas turbine and stabilizing the combustion of the entire combustor during load operation. operation to ensure reliability. In this embodiment, the fuel supplied to the gas turbine start-up fuel system 206 is liquid fuel such as light oil and heavy oil.

On the other hand, an outer burner inner peripheral fuel system 204 and an outer peripheral burner outer peripheral fuel system 205 are connected to the outer peripheral burner 33 . The coaxial jet group arranged on the first row concentric circle of the peripheral burner 33 forms the starting point of the flame, so it is particularly related to combustion stability. Therefore, as in the present embodiment, by independently controlling the flow rate of fuel supplied to the first row (inner circumference) of the outer circumference burner 33, the starting point of the flame is strengthened and stable combustion is achieved over a wider load range. can be maintained.

FIG. 8 is a front view of the air hole plate 20 and the liner 12 of FIG. 6 as seen from the combustion chamber 5 side. A feature of this embodiment is that, as shown in FIG. 8, an annular chamber 27 is divided into a plurality of parts in the circumferential direction of the liner 12 in an air hole plate 20 having one central burner 32 and a plurality of peripheral burners 33. It is what I did.

In the gas turbine combustor of this embodiment, the burner has air holes 21 arranged in a plurality of rows concentrically and a plurality of fuel nozzles 22 arranged coaxially with the air holes 21 on the central axis of the air hole plate 20. , and a plurality of peripheral burners 33 are arranged concentrically around the central burner 32 to form one combustor.

By dividing the annular chamber 27 in the circumferential direction in this way, in addition to the same effects as in the first embodiment, the following effects can be obtained.

In this embodiment, the ring-shaped chamber 27 is divided into a plurality of sections in the circumferential direction of the liner 12, so that the ring-shaped chamber 27 is provided only in the region between the peripheral burners 33 adjacent in the circumferential direction in the liner 12. be able to.

In the region between two adjacent peripheral burners 33, the leakage air 105 passing through the spring seal 26 flows downstream of the liner 12 without affecting the flame 83 formed in the combustion chamber 5, but in this region the annular chamber 27, leak air 105 can be jetted out from pressure wave introduction hole 90 into combustion chamber 5 as purge air 106, flame can be prevented from entering acoustic space 92, and a decrease in flow rate of combustion air 104 can be suppressed. Therefore, NOx emissions can be reduced.

Further, by arranging the ring-shaped chambers 27 separately, the radial height L of the acoustic space 92 can be set individually for each ring-shaped chamber 27, so that the frequency band of pressure fluctuations that can be attenuated in each of the sound spaces 92 can be determined. It can be varied and can accommodate a wide band of combustion oscillations.

In addition, the present invention is not limited to the above-described embodiments, and includes various modifications. For example, the above-described embodiments have been described in detail in order to explain the present invention in an easy-to-understand manner, and are not necessarily limited to those having all the described configurations. In addition, it is possible to replace part of the configuration of one embodiment with the configuration of another embodiment, and it is also possible to add the configuration of another embodiment to the configuration of one embodiment. Moreover, it is possible to add, delete, or replace a part of the configuration of each embodiment with another configuration.

DESCRIPTION OF SYMBOLS 1... Gas turbine plant, 2... Compressor, 3... (gas turbine) combustor, 4... Turbine, 5... Combustion chamber, 6... Transition piece, 7... End cover, 8... Burner, 9... Diffuser, 10... Outside Cylinder 12 Liner (combustor) 12b Outer peripheral surface of liner 13 Casing 14 Flow sleeve 15 Diffusion burner 16 Premixed burner 17 Fuel ejection hole 18 Swirler , 19... Flame stabilizer, 20... Air hole plate, 21... Air hole, 22... Fuel nozzle, 23... Fuel header, 24... Starting fuel nozzle, 25... Plate lip, 26... Spring seal, 27... Annular chamber, 29... Rectangular annular cover 32... Central burner 33, 112... Peripheral burner 51... 1st row air hole 52... 2nd row air hole 53... 3rd row air hole 60... Fuel cutoff valve 61a, 62a... fuel pressure control valve 61b, 62b... fuel flow rate control valve 80... circulation flow (circulation vortex) 83... flame 90... pressure wave introduction hole 91... purge air hole 92... acoustic space 100... fuel , 102 Compressed air 103 Cooling air 104 Combustion air 105 Leak air 106 Purge air 110 Combustion gas 200, 201, 202 Fuel supply system 203 Central burner fuel system 204 .

Claims (11)

a combustion chamber for burning a mixture of fuel and air;
a combustor liner that defines the combustion chamber;
a burner positioned upstream of the combustion chamber and forming a flame;
a spring seal disposed between an outer peripheral surface of the burner and an inner peripheral surface of the combustor liner at a fitting portion between the burner and the combustor liner;
an annular chamber located in the combustor liner downstream of the spring seal and defining an acoustic space;
The annular chamber has a purge air hole for guiding leaked air that has passed through the spring seal to the acoustic space on a side surface on the spring seal side, and has a plurality of pressure wave introduction holes that communicate the combustion chamber and the acoustic space. Gas turbine combustor. A gas turbine combustor according to claim 1, comprising:
The gas turbine combustor, wherein the annular chamber is positioned downstream of and overlies the spring seal. A gas turbine combustor according to claim 1, comprising:
A gas turbine combustor, wherein the annular chamber is arranged to project outward from the combustor liner. A gas turbine combustor according to claim 1, comprising:
The gas turbine combustor, wherein the purge air holes are circular holes provided in a circumferential direction of the combustor liner or annular slits provided in a circumferential direction of the combustor liner. A gas turbine combustor according to claim 1, comprising:
A gas turbine combustor, wherein the outermost peripheral surface of the burner and the inner peripheral surface of the annular chamber are substantially flush with each other. A gas turbine combustor according to claim 1, comprising:
A gas turbine combustor, wherein the burners are composed of a diffusion burner arranged on the central axis of the combustion chamber and a plurality of premixed burners arranged around the diffusion burner. A gas turbine combustor according to claim 1, comprising:
The burner is positioned upstream of the combustion chamber and has an air hole plate having a plurality of air holes arranged concentrically in a plurality of rows, and a gas turbine having a plurality of fuel nozzles arranged coaxially with the air holes. combustor. A gas turbine combustor according to claim 7,
The burner includes air holes arranged in a plurality of concentric rows and a burner having a plurality of fuel nozzles arranged coaxially with the air holes, and one burner is arranged as a central burner on the central axis of the air hole plate,
A gas turbine combustor that is a multi-cluster burner in which a plurality of peripheral burners are concentrically arranged around the central burner to form one combustor. A gas turbine combustor according to claim 8,
A gas turbine combustor in which the annular chamber is divided into a plurality of parts so as to be arranged only in regions between adjacent outer peripheral burners in the circumferential direction of the combustor liner. A gas turbine combustor according to claim 8,
A gas turbine combustor having fuel nozzles for supplying liquid fuel to the central burner. A gas turbine combustor according to claim 1, comprising:
A gas turbine combustor that burns a fuel containing hydrogen in its fuel composition.

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