CN114263930A - Fuel nozzle and gas turbine combustor - Google Patents
Fuel nozzle and gas turbine combustor Download PDFInfo
- Publication number
- CN114263930A CN114263930A CN202111081539.0A CN202111081539A CN114263930A CN 114263930 A CN114263930 A CN 114263930A CN 202111081539 A CN202111081539 A CN 202111081539A CN 114263930 A CN114263930 A CN 114263930A
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- Prior art keywords
- flow path
- fuel
- nozzle
- combustion air
- fuel nozzle
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Classifications
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
- F23D—BURNERS
- F23D14/00—Burners for combustion of a gas, e.g. of a gas stored under pressure as a liquid
- F23D14/20—Non-premix gas burners, i.e. in which gaseous fuel is mixed with combustion air on arrival at the combustion zone
- F23D14/22—Non-premix gas burners, i.e. in which gaseous fuel is mixed with combustion air on arrival at the combustion zone with separate air and gas feed ducts, e.g. with ducts running parallel or crossing each other
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
- F23R—GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
- F23R3/00—Continuous combustion chambers using liquid or gaseous fuel
- F23R3/28—Continuous combustion chambers using liquid or gaseous fuel characterised by the fuel supply
- F23R3/283—Attaching or cooling of fuel injecting means including supports for fuel injectors, stems, or lances
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
- F23R—GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
- F23R3/00—Continuous combustion chambers using liquid or gaseous fuel
- F23R3/28—Continuous combustion chambers using liquid or gaseous fuel characterised by the fuel supply
- F23R3/286—Continuous combustion chambers using liquid or gaseous fuel characterised by the fuel supply having fuel-air premixing devices
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
- F23R—GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
- F23R3/00—Continuous combustion chambers using liquid or gaseous fuel
- F23R3/28—Continuous combustion chambers using liquid or gaseous fuel characterised by the fuel supply
- F23R3/34—Feeding into different combustion zones
- F23R3/343—Pilot flames, i.e. fuel nozzles or injectors using only a very small proportion of the total fuel to insure continuous combustion
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
- F23R—GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
- F23R3/00—Continuous combustion chambers using liquid or gaseous fuel
- F23R3/28—Continuous combustion chambers using liquid or gaseous fuel characterised by the fuel supply
- F23R3/36—Supply of different fuels
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
- F23D—BURNERS
- F23D2209/00—Safety arrangements
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
- F23R—GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
- F23R2900/00—Special features of, or arrangements for continuous combustion chambers; Combustion processes therefor
- F23R2900/00005—Preventing fatigue failures or reducing mechanical stress in gas turbine components
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
- F23R—GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
- F23R2900/00—Special features of, or arrangements for continuous combustion chambers; Combustion processes therefor
- F23R2900/00012—Details of sealing devices
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
- F23R—GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
- F23R2900/00—Special features of, or arrangements for continuous combustion chambers; Combustion processes therefor
- F23R2900/00017—Assembling combustion chamber liners or subparts
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
- F23R—GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
- F23R2900/00—Special features of, or arrangements for continuous combustion chambers; Combustion processes therefor
- F23R2900/00018—Manufacturing combustion chamber liners or subparts
Landscapes
- Engineering & Computer Science (AREA)
- Chemical & Material Sciences (AREA)
- Combustion & Propulsion (AREA)
- Mechanical Engineering (AREA)
- General Engineering & Computer Science (AREA)
Abstract
The invention provides a fuel nozzle with a plurality of fuel systems, which has small thermal stress generated by temperature difference of conducted fuel and combustion air, and a gas turbine combustor using the fuel nozzle. The fuel nozzle is provided with a plurality of flow paths, and is characterized by comprising: a first flow path that conducts fuel or combustion air; and a second flow path which conducts fuel or combustion air and is different from the first flow path, wherein at least a portion where the first flow path and the second flow path are arranged among the components of the fuel nozzle is formed of an integral member.
Description
Technical Field
The present invention relates to the construction of fuel nozzles for use in gas turbine combustors, and more particularly to techniques that are effectively applied to pilot nozzles.
Background
The types of fuel used in the gas turbine are various, and the combustor to be used is selected according to the fuel calorie and the combustion speed. For low calorie fuels, diffusion burners are used, and for high calorie fuels, premix burners are used. Compared with diffusion combustion, premixed combustion can reduce flame temperature, so that there is no spray of water or steam, and can reduce NOx, and is widely used in current gas turbines.
In many gas turbines for power generation, natural gas is mainly used as a fuel, and a natural gas burning premix burner includes a pilot nozzle and a main nozzle, and the main premix flame is stabilized by a flame formed by the pilot nozzle.
As a background art in this field, for example, there is a technology as in patent document 1. Patent document 1 discloses "a pilot burner of a gas turbine, which is disposed at an axial center of a combustor of the gas turbine, and includes: a pilot combustion nozzle in which a plurality of fuel flow paths for premixed combustion and a plurality of fuel flow paths for diffusion combustion are independently formed in an axial direction; a pilot burner cylinder which is concentric with respect to the pilot combustion nozzle and whose upstream end portion is disposed so as to surround the downstream end portion of the pilot combustion nozzle; and a plurality of swirl vanes which are disposed radially at the downstream end of the pilot combustion nozzle, and which impart a swirl force to the compressed air passing through an annular air passage formed between the downstream end of the pilot combustion nozzle and the upstream end of the burner tube, thereby turning the compressed air into a swirl flow ".
Documents of the prior art
Patent document
Patent document 1: japanese patent laid-open publication No. 2010-249449
Disclosure of Invention
Problems to be solved by the invention
As described above, many natural gas burning premix combustors include one pilot nozzle and eight main nozzles, and the fuel system includes two systems, i.e., a main system and a pilot system. The ignition ratio (ignition fuel flow rate/total fuel flow rate) is the largest at the time of ignition, decreases as the load increases, and is the lowest at the rated load, thereby suppressing the amount of NOx discharged.
Further, since combustibility changes when the methane concentration in the fuel changes, it is necessary to adjust the fuel-air ratio in the combustion region by adjusting the air bypass valve, or to adjust the ignition ratio to a stable combustion state by changing the ignition ratio.
However, in a fuel nozzle of a gas turbine combustor, the generation of thermal stress caused by a temperature difference between combustion air and fuel often becomes a problem. When excessive thermal stress is generated, the low cycle fatigue life is insufficient, and the operation is limited. In particular, in a fuel nozzle provided with a plurality of fuel systems, such as the natural gas burning premix burner described above, there are cases where: the thermal stress increases by conducting a fluid having a different temperature, such as fuel and combustion air (purge air), depending on the operating state. Thermal stresses generated at the fuel nozzle result in a reduction in the reliability and durability of the fuel nozzle.
According to patent document 1, although it is possible to reduce vibration caused by the flow of compressed air and to prevent blow-out at the time of start-up, no consideration is given to the thermal stress of the fuel nozzle caused by conduction of fluids having different temperatures, such as fuel and combustion air (purge air), as described above.
Accordingly, an object of the present invention is to provide a fuel nozzle having a plurality of fuel systems, in which thermal stress generated by a temperature difference between conducted fuel and combustion air is small, and a gas turbine combustor using the fuel nozzle.
Means for solving the problems
In order to solve the above problem, the present invention provides a fuel nozzle including a plurality of flow paths, the fuel nozzle including: a first flow path that conducts fuel or combustion air; and a second flow path which conducts fuel or combustion air and is different from the first flow path, wherein at least a portion where the first flow path and the second flow path are arranged among the components of the fuel nozzle is formed of an integral member.
Further, the present invention is a gas turbine combustor including: a combustor liner constituting a combustion chamber in which a mixture of fuel and combustion air is combusted; a transition piece that guides the combustion gas from the combustion chamber to a turbine; a pilot nozzle that supplies fuel and combustion air to the combustion chamber; and a main nozzle which is disposed in a plurality of numbers around the pilot nozzle and supplies fuel and combustion air to the combustion chamber, the pilot nozzle including: a first flow path that conducts fuel or combustion air; and a second flow path that conducts fuel or combustion air, and that is different from the first flow path, wherein at least a portion where the first flow path and the second flow path are arranged among constituent members of the pilot nozzle is formed of an integral member.
Effects of the invention
According to the present invention, a fuel nozzle having a small thermal stress due to a temperature difference between conducted fuel and combustion air in a fuel nozzle including a plurality of fuel systems, and a gas turbine combustor using the fuel nozzle can be realized.
This makes it possible to provide a high-performance gas turbine combustor having excellent reliability and durability.
Problems, structures, and effects other than those described above will become apparent from the following description of the embodiments.
Drawings
Fig. 1 is a diagram showing a configuration example of a general gas turbine.
Fig. 2 is a diagram showing a configuration example of a general combustor.
Fig. 3 is a sectional view showing the structure of a fuel nozzle according to embodiment 1 of the present invention.
Fig. 4A is a sectional view a-a' of fig. 3.
Fig. 4B is a sectional view B-B' of fig. 3.
Fig. 5 is a sectional view showing the structure of a conventional fuel nozzle.
Fig. 6A is a cross-sectional view C-C' of fig. 5.
Fig. 6B is a cross-sectional view D-D' of fig. 5.
In the figure:
1-compressor, 2-combustor, 3-turbine, 4-combustor liner, 5-transition piece, 6-main nozzle, 7-pilot nozzle, 8-flow direction of combustion gas, 9, 10, 11-nozzle component, 12-junction, 13-flow path a, 14-flow path B, 15-combustion chamber.
Detailed Description
Hereinafter, embodiments of the present invention will be described with reference to the drawings. In the drawings, the same components are denoted by the same reference numerals, and detailed description of overlapping portions will be omitted.
Example 1
First, a gas turbine combustor and a conventional problem to be solved by the present invention will be described with reference to fig. 1, 2, and 5 to 6B. Fig. 1 is a diagram showing a configuration example of a general gas turbine. Fig. 2 is a diagram showing a configuration example of a general combustor, and shows a combustor including a combustor liner 4 and a transition piece 5 constituting a combustion chamber 15. Fig. 5 is a sectional view showing the structure of a conventional pilot burner 7, and fig. 6A and 6B show a section C-C 'and a section D-D' of fig. 5, respectively.
As shown in fig. 1, the gas turbine is roughly divided into a compressor 1, a combustor 2, and a turbine 3. The compressor 1 adiabatically compresses air taken in from the atmosphere as a working fluid, and the combustor 2 mixes and combusts fuel with the compressed air supplied from the compressor 1 to generate high-temperature and high-pressure combustion gas, and the turbine 3 generates rotational power when the combustion gas introduced from the combustor 2 expands. The exhaust gas from the turbine 3 is discharged to the atmosphere.
As shown in fig. 2, the combustor 2 includes: a combustor liner 4 constituting a combustion chamber 15 for combusting a mixture of fuel and combustion air; a transition piece 5 that guides the combustion gases from the combustion chamber 15 to the turbine 3 (flow direction 8 of the combustion gases); and main nozzles 6 and pilot nozzles 7 that supply fuel and combustion air to the combustion chamber 15. As described above, the main nozzle 6 is arranged in plural (for example, eight) around one pilot nozzle 7.
As shown in fig. 5, the conventional pilot burner 7 is configured by joining together the nozzle component members 9, 10, and 11, in which the flow paths a13 and B14 have been formed in advance by drilling with a drill or the like, at the joint 12. For example, brazing welding is used for joining the nozzle component members 9, 10, and 11.
Generally, at the rated load of the gas turbine, purge air (combustion air) having a relatively high temperature is conducted to the flow path a13, and fuel such as natural gas having a relatively low temperature is conducted to the flow path B14. Thermal stresses are thus generated due to the mainly radial temperature differences of the pilot burner 7 and the resulting radial and axial thermal expansion differences. In general, thermal stress is easily promoted in the welded portion due to the shape discontinuity caused by the unwelded portion or the like, and the fatigue strength is lower in the welded portion than in the base material.
As described above, in the conventional pilot nozzle 7, the joint 12, in particular, where both the flow path a13 and the flow path B14 are disposed, becomes a bottleneck of strength due to the temperature difference between the fuel and the combustion air that are conducted through the flow path a13 and the flow path B14, and is restricted in operation due to low cycle fatigue.
As shown in fig. 6A, in the root of the conventional pilot burner 7, since both the flow path a13 and the flow path B14 are arranged in a ring shape (Annular shape) in the circumferential direction of the pilot burner 7, the pilot burner 7 has a structure thermally divided in the radial direction by the flow path a13 and the flow path B14. Therefore, thermal stress on the pilot nozzle 7 due to a difference in temperature between the fuel and the combustion air that are conducted through the flow paths a13 and B14, respectively, is further promoted.
As shown in fig. 6B, in the vicinity of the tip of the conventional pilot nozzle 7, the flow passage B14 is disposed in a plurality of divided portions in the circumferential direction of the pilot nozzle 7, but the flow passage a13 is disposed in an annular (ring) shape in the circumferential direction of the pilot nozzle 7 similarly to the root portion, and the pilot nozzle 7 is thermally divided by the flow passage a13 in the radial direction.
Next, a fuel nozzle according to example 1 of the present invention will be described with reference to fig. 3 to 4B. Fig. 3 is a sectional view showing the structure of the pilot nozzle 7 of the present embodiment, and fig. 4A and 4B show a section a-a 'and a section B-B' of fig. 3, respectively.
As shown in fig. 3, the pilot nozzle 7 of the present embodiment has a flow path a13 (first flow path) for conducting fuel or combustion air and a flow path B14 (second flow path) different from the flow path a13 (first flow path) for conducting fuel or combustion air, and at least a portion of the nozzle constituting members 9 and 10 of the pilot nozzle 7 where both the flow path a13 (first flow path) and the flow path B14 (second flow path) are arranged is constituted by an integral nozzle constituting member 10 without a joint portion 12.
As shown in fig. 3, by forming the portion where both the flow path a13 (first flow path) and the flow path B14 (second flow path) are arranged by the integrated nozzle constituent member 10 without the joint portion 12, it is possible to prevent the joint portion 12 from becoming a bottleneck in strength due to the temperature difference of the fuel or the combustion air conducted through the flow paths a13 and B14, respectively, as described above, and to improve the reliability and durability of the pilot nozzle 7.
As shown in fig. 4A and 4B, in the pilot nozzle 7 of the present embodiment, the flow path a13 (first flow path) and the flow path B14 (second flow path) are arranged in a plurality of segments in the circumferential direction of the pilot nozzle 7.
As shown in fig. 4A and 4B, by arranging both the flow path a13 (first flow path) and the flow path B14 (second flow path) in a plurality of divisions in the circumferential direction of the pilot nozzle 7, the pilot nozzle 7 can be prevented from being thermally completely divided by the flow path a13 (first flow path) and the flow path B14 (second flow path) in the radial direction. This can alleviate thermal stress on the pilot nozzle 7 due to a temperature difference between the fuel and the combustion air that are conducted through the flow paths a13 and B14, respectively.
For example, even when the combustion air is conducted through the flow path a13 (first flow path) and the fuel having a temperature lower than that of the combustion air is conducted through the flow path B14 (second flow path), thermal stress to the pilot nozzle 7 due to a temperature difference between the fuel and the combustion air can be alleviated, and therefore, the reliability and durability of the pilot nozzle 7 can be further improved in addition to the effect of the integral nozzle component 10 without the joint 12.
As shown in fig. 3, the nozzle component 9 near the tip of the pilot nozzle 7 is provided with only the flow path a13 (first flow path) of the flow path a13 (first flow path) and the flow path B14 (second flow path), and the nozzle component 9 provided with only the flow path a13 (first flow path) and the nozzle component 10 provided with both the flow path a13 (first flow path) and the flow path B14 (second flow path) are joined by, for example, brazing welding or Hot Isostatic Pressing (HIP).
As shown in fig. 3, by configuring the joint 12 to be disposed limited to the region in which only one of the flow path a13 (first flow path) of the flow path a13 (first flow path) and the flow path B14 (second flow path) is formed, it is possible to prevent the occurrence of thermal stress on the pilot nozzle 7 due to a temperature difference between the fuel and the combustion air that are conducted through the flow paths, and it is possible to ensure the joint reliability of the joint 12.
The joint 12 of the nozzle component 9 and the nozzle component 10 is preferably joined by the Hot Isostatic Pressing (HIP) method described above. By using the Hot Isostatic Pressing (HIP) method, the unwelded portion can be eliminated as much as possible, and therefore, thermal stress due to shape discontinuity in the joint 12 can be suppressed.
As described above, according to the present invention, a fuel nozzle having a small thermal stress due to a temperature difference between conducted fuel and combustion air and a gas turbine combustor using the fuel nozzle can be realized, and the reliability and durability of the gas turbine combustor can be improved.
The present invention is not limited to the above-described embodiments, and includes various modifications. For example, the above-described embodiments are examples described in detail to explain the present invention easily and understandably, and are not limited to having all the configurations described. In addition, a part of the structure of one embodiment may be replaced with the structure of another embodiment, and the structure of another embodiment may be added to the structure of one embodiment. In addition, a part of the configuration of each embodiment can be added, deleted, or replaced with another configuration.
Claims (10)
1. A fuel nozzle provided with a plurality of flow paths, characterized by comprising:
a first flow path that conducts fuel or combustion air; and
a second flow path which conducts fuel or combustion air and is different from the first flow path,
at least a portion where the first flow path and the second flow path are arranged among the components of the fuel nozzle is formed of an integral member.
2. The fuel nozzle of claim 1,
the first flow path and the second flow path are arranged so as to be divided into a plurality of sections in the circumferential direction of the fuel nozzle.
3. The fuel nozzle of claim 1 or 2,
the combustion air is conducted through the first flow path,
the second flow path is configured to conduct fuel having a temperature lower than that of the combustion air.
4. The fuel nozzle of claim 1 or 2,
only the first flow passage of the first flow passage and the second flow passage is disposed in the vicinity of the tip of the fuel nozzle,
the portion where only the first flow path is arranged is joined to the portion where the first flow path and the second flow path are arranged.
5. The fuel nozzle of claim 4,
only the portion where the first flow path is arranged is joined to the portion where the first flow path and the second flow path are arranged by welding or a Hot Isostatic Pressing (HIP) method.
6. A gas turbine combustor, comprising:
a combustor liner constituting a combustion chamber in which a mixture of fuel and combustion air is combusted;
a transition piece that guides the combustion gas from the combustion chamber to a turbine;
a pilot nozzle that supplies fuel and combustion air to the combustion chamber; and
a plurality of main nozzles arranged around the pilot nozzle and configured to supply fuel and combustion air to the combustion chamber,
the pilot nozzle includes:
a first flow path that conducts fuel or combustion air; and
a second flow path which conducts fuel or combustion air and is different from the first flow path,
at least a portion where the first flow path and the second flow path are arranged among the constituent members of the pilot nozzle is formed of an integral member.
7. The gas turbine combustor of claim 6,
the first flow path and the second flow path are arranged so as to be divided into a plurality of sections in the circumferential direction of the fuel nozzle.
8. The gas turbine combustor of claim 6 or 7,
the combustion air is conducted through the first flow path,
the second flow path is configured to conduct fuel having a temperature lower than that of the combustion air.
9. The gas turbine combustor of claim 6 or 7,
only the first flow passage of the first flow passage and the second flow passage is disposed in the vicinity of the tip of the fuel nozzle,
the portion where only the first flow path is arranged is joined to the portion where the first flow path and the second flow path are arranged.
10. The gas turbine combustor of claim 9,
only the portion where the first flow path is arranged is joined to the portion where the first flow path and the second flow path are arranged by welding or a Hot Isostatic Pressing (HIP) method.
Applications Claiming Priority (2)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| JP2020155193A JP2022049136A (en) | 2020-09-16 | 2020-09-16 | Fuel nozzle, and gas turbine combustor |
| JP2020-155193 | 2020-09-16 |
Publications (1)
| Publication Number | Publication Date |
|---|---|
| CN114263930A true CN114263930A (en) | 2022-04-01 |
Family
ID=80351630
Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| CN202111081539.0A Pending CN114263930A (en) | 2020-09-16 | 2021-09-15 | Fuel nozzle and gas turbine combustor |
Country Status (4)
| Country | Link |
|---|---|
| US (1) | US20220082260A1 (en) |
| JP (2) | JP2022049136A (en) |
| CN (1) | CN114263930A (en) |
| DE (1) | DE102021210300A1 (en) |
Citations (7)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US20120000203A1 (en) * | 2009-03-17 | 2012-01-05 | Matthias Hase | Method for operating a burner and burner, in particular for a gas turbine |
| CN106133294A (en) * | 2014-04-30 | 2016-11-16 | 三菱日立电力系统株式会社 | gas turbine combustor, gas turbine, control device and control method |
| CN205746972U (en) * | 2014-12-23 | 2016-11-30 | 通用电气公司 | The system of the cooling air in utilizing burner |
| CN106461223A (en) * | 2014-09-19 | 2017-02-22 | 三菱日立电力系统株式会社 | Combustion burner, combustor, and gas turbine |
| CN106705121A (en) * | 2015-11-13 | 2017-05-24 | 三菱日立电力系统株式会社 | Gas turbine combustor |
| CN107735617A (en) * | 2015-07-03 | 2018-02-23 | 三菱日立电力系统株式会社 | Burner nozzle, gas turbine combustor and gas turbine and cover ring, the manufacture method of burner nozzle |
| CN109296460A (en) * | 2014-07-08 | 2019-02-01 | 八河流资产有限责任公司 | For heating the method and electricity-generating method of recirculated air |
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| JPH07286730A (en) * | 1994-04-18 | 1995-10-31 | Hitachi Ltd | Flame stabilizer for gas turbine burner |
| JP2011099654A (en) * | 2009-11-09 | 2011-05-19 | Mitsubishi Heavy Ind Ltd | Combustion burner for gas turbine |
| JP5631223B2 (en) * | 2011-01-14 | 2014-11-26 | 三菱重工業株式会社 | Fuel nozzle, gas turbine combustor including the same, and gas turbine including the same |
| JP6206648B2 (en) * | 2013-07-08 | 2017-10-04 | 三菱日立パワーシステムズ株式会社 | Chip holder, combustor nozzle including the same, combustor including the combustor nozzle, and method for manufacturing the combustor nozzle |
| JP6452298B2 (en) * | 2014-03-25 | 2019-01-16 | 三菱日立パワーシステムズ株式会社 | Injection nozzle, gas turbine combustor and gas turbine |
| JP6839571B2 (en) * | 2017-03-13 | 2021-03-10 | 三菱パワー株式会社 | Combustor nozzles, combustors, and gas turbines |
-
2020
- 2020-09-16 JP JP2020155193A patent/JP2022049136A/en active Pending
-
2021
- 2021-08-18 US US17/405,372 patent/US20220082260A1/en active Pending
- 2021-09-15 CN CN202111081539.0A patent/CN114263930A/en active Pending
- 2021-09-16 DE DE102021210300.6A patent/DE102021210300A1/en active Pending
-
2023
- 2023-08-08 JP JP2023129450A patent/JP2023153214A/en active Pending
Patent Citations (7)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US20120000203A1 (en) * | 2009-03-17 | 2012-01-05 | Matthias Hase | Method for operating a burner and burner, in particular for a gas turbine |
| CN106133294A (en) * | 2014-04-30 | 2016-11-16 | 三菱日立电力系统株式会社 | gas turbine combustor, gas turbine, control device and control method |
| CN109296460A (en) * | 2014-07-08 | 2019-02-01 | 八河流资产有限责任公司 | For heating the method and electricity-generating method of recirculated air |
| CN106461223A (en) * | 2014-09-19 | 2017-02-22 | 三菱日立电力系统株式会社 | Combustion burner, combustor, and gas turbine |
| CN205746972U (en) * | 2014-12-23 | 2016-11-30 | 通用电气公司 | The system of the cooling air in utilizing burner |
| CN107735617A (en) * | 2015-07-03 | 2018-02-23 | 三菱日立电力系统株式会社 | Burner nozzle, gas turbine combustor and gas turbine and cover ring, the manufacture method of burner nozzle |
| CN106705121A (en) * | 2015-11-13 | 2017-05-24 | 三菱日立电力系统株式会社 | Gas turbine combustor |
Also Published As
| Publication number | Publication date |
|---|---|
| DE102021210300A1 (en) | 2022-03-17 |
| JP2022049136A (en) | 2022-03-29 |
| JP2023153214A (en) | 2023-10-17 |
| US20220082260A1 (en) | 2022-03-17 |
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