JP2022049136A - Fuel nozzle, and gas turbine combustor - Google Patents

Fuel nozzle, and gas turbine combustor Download PDF

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Publication number
JP2022049136A
JP2022049136A JP2020155193A JP2020155193A JP2022049136A JP 2022049136 A JP2022049136 A JP 2022049136A JP 2020155193 A JP2020155193 A JP 2020155193A JP 2020155193 A JP2020155193 A JP 2020155193A JP 2022049136 A JP2022049136 A JP 2022049136A
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flow path
fuel
nozzle
combustion air
gas turbine
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裕明 長橋
Hiroaki Nagahashi
義隆 寺田
Yoshitaka Terada
祥平 沼田
Shohei Numata
祥太 五十嵐
Shota Igarashi
康弘 和田
Yasuhiro Wada
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Mitsubishi Heavy Industries Ltd
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Mitsubishi Heavy Industries Ltd
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Priority to JP2020155193A priority Critical patent/JP2022049136A/en
Priority to US17/405,372 priority patent/US20220082260A1/en
Priority to CN202111081539.0A priority patent/CN114263930A/en
Priority to DE102021210300.6A priority patent/DE102021210300A1/en
Publication of JP2022049136A publication Critical patent/JP2022049136A/en
Priority to JP2023129450A priority patent/JP2023153214A/en
Pending legal-status Critical Current

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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
    • 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
    • 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
    • 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
    • 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

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  • Engineering & Computer Science (AREA)
  • Chemical & Material Sciences (AREA)
  • Combustion & Propulsion (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)

Abstract

To provide a fuel nozzle comprising a plurality of fuel systems and generating little thermal stress resulting from a temperature difference between fuel and combustion air flowing therethrough, and a gas turbine combustor employing the same.SOLUTION: The fuel nozzle equipped with a plurality of flow channels, comprises: a first flow channel through which fuel or combustion air is conducted; and a second flow channel through which fuel or combustion air is conducted and which is different from the first flow channel, wherein among component members of the fuel nozzle, at least a portion where the first flow channel and the second flow channel are disposed, is constituted of an integral member.SELECTED DRAWING: Figure 3

Description

本発明は、ガスタービン燃焼器に用いられる燃料ノズルの構造に係り、特に、パイロットノズルに適用して有効な技術に関する。 The present invention relates to the structure of a fuel nozzle used in a gas turbine combustor, and more particularly to a technique applicable to a pilot nozzle.

ガスタービンに使用する燃料の種類には様々なものがあり、燃料カロリーと燃焼速度により使用する燃焼器を選定している。カロリーの低い燃料については、拡散燃焼器を使用し、カロリーの高い燃料については予混合燃焼器を使用する。予混合燃焼は拡散燃焼に比べ火炎温度を低減できるため、水や蒸気の噴霧なしでＮＯｘの低減が可能であり、現在ガスタービンに広く適用されている。 There are various types of fuel used for gas turbines, and the combustor to be used is selected according to the fuel calories and combustion speed. For low-calorie fuels, use a diffusion combustor, and for high-calorie fuels, use a premixed combustor. Since premixed combustion can reduce the flame temperature as compared with diffusion combustion, it is possible to reduce NOx without spraying water or steam, and it is currently widely applied to gas turbines.

発電用に使用されているガスタービンでは主に天然ガスを燃料としているが、多くの天然ガス焚き予混合燃焼器はパイロットノズルとメインノズルを備えており、パイロットノズルで形成される火炎によりメイン予混合火炎の安定化を図っている。 Gas turbines used for power generation mainly use natural gas as fuel, but many natural gas-fired premixed combustors are equipped with a pilot nozzle and a main nozzle, and the flame formed by the pilot nozzle is used as the main premixer. We are trying to stabilize the mixed flame.

本技術分野の背景技術として、例えば、特許文献１のような技術がある。特許文献１には「ガスタービンの燃焼器の軸心に配置されてなるガスタービンのパイロット燃焼バーナであって、軸方向に沿って、その内部に複数の予混合燃焼用の燃料流路と、複数の拡散燃焼用の燃料流路とが独立して形成されたパイロット燃焼ノズルと、このパイロット燃焼ノズルに対して同心状で、かつ、その上流側の端部が、前記パイロット燃焼ノズルの下流側の端部を囲繞する状態で配置されるパイロットバーナ筒と、前記パイロット燃焼ノズルの下流側の端部に放射状に配置されて、前記パイロット燃焼ノズルの下流側の端部と、前記バーナ筒の上流側の端部との間に形成されたリング状の空気通路を通過する圧縮空気に旋回力を付与して、この圧縮空気を旋回空気流にする複数枚の旋回翼とを備えているガスタービンのパイロット燃焼バーナ」が開示されている。 As a background technology in this technical field, for example, there is a technology such as Patent Document 1. Patent Document 1 describes, "A pilot combustion burner of a gas turbine arranged at the axis of a gas turbine combustor, and a plurality of fuel flow paths for premixed combustion inside the pilot combustion burner along the axial direction. A pilot combustion nozzle in which a plurality of fuel channels for diffusion combustion are independently formed, and an end concentric with the pilot combustion nozzle and on the upstream side thereof are on the downstream side of the pilot combustion nozzle. A pilot burner cylinder arranged so as to surround the end of the pilot combustion nozzle, and radially arranged at the downstream end of the pilot combustion nozzle, the downstream end of the pilot combustion nozzle and the upstream of the burner cylinder. A gas turbine equipped with a plurality of swivel blades that apply a swirling force to the compressed air passing through a ring-shaped air passage formed between the side ends and make the compressed air a swirling air flow. Pilot Combustion Burner "is disclosed.

特開２０１０－２４９４４９号公報Japanese Unexamined Patent Publication No. 2010-2449449

上述したように、多くの天然ガス焚き予混合燃焼器は１本のパイロットノズルと８本のメインノズルを備えており、燃料系統はメイン系統とパイロット系統の２系統から構成される。パイロット比率（パイロット燃料流量／全体燃料流量）は着火時に最も多く、負荷上昇と共に低下させ、定格負荷ではパイロット比率を最も低くしＮＯｘの排出量を抑えている。 As mentioned above, many natural gas-fired premixed combustors are equipped with one pilot nozzle and eight main nozzles, and the fuel system consists of two systems, a main system and a pilot system. The pilot ratio (pilot fuel flow rate / total fuel flow rate) is the highest at the time of ignition and decreases as the load increases, and the pilot ratio is the lowest at the rated load to suppress NOx emissions.

また、燃料中のメタン濃度が変化した場合、燃焼性が変化するため、空気バイパス弁の調節により燃焼領域の燃空比を調整したり、パイロット比率を変化させて、安定な燃焼状態に調整することが必要となる。 In addition, when the methane concentration in the fuel changes, the combustibility changes, so the fuel-air ratio in the combustion region can be adjusted by adjusting the air bypass valve, or the pilot ratio can be changed to adjust to a stable combustion state. Is required.

ところで、ガスタービン燃焼器の燃料ノズルでは燃焼空気と燃料との温度差に起因した熱応力の発生がしばしば問題となる。過大な熱応力が発生すると低サイクル疲労寿命が不足し、運用上の制限が発生する。特に、上述した天然ガス焚き予混合燃焼器のように複数の燃料系統を備えた燃料ノズルにおいては、運転状態に応じて燃料及び燃焼空気（パージ空気）などの温度が異なる流体が導通し、これに起因して熱応力が増大する場合がある。燃料ノズルに発生する熱応力は、燃料ノズルの信頼性及び耐久性の低下に繋がる。 By the way, in the fuel nozzle of a gas turbine combustor, the generation of thermal stress due to the temperature difference between the combustion air and the fuel is often a problem. If excessive thermal stress is generated, the low cycle fatigue life will be insufficient and operational restrictions will occur. In particular, in a fuel nozzle equipped with a plurality of fuel systems such as the above-mentioned natural gas-fired premixed combustor, fluids having different temperatures such as fuel and combustion air (purge air) are conducted depending on the operating state. The thermal stress may increase due to the above. The thermal stress generated in the fuel nozzle leads to a decrease in the reliability and durability of the fuel nozzle.

上記特許文献１によれば、圧縮空気が流動することによって生じる振動を低減させることができ、かつ、起動時における吹き消えを防止することができるが、上記のような燃料及び燃焼空気（パージ空気）などの温度が異なる流体の導通に起因して発生する燃料ノズルの熱応力については何ら考慮されていない。 According to the above Patent Document 1, it is possible to reduce the vibration caused by the flow of the compressed air and to prevent the blowout at the time of starting, but the fuel and the combustion air (purge air) as described above can be prevented. ), Etc., no consideration is given to the thermal stress of the fuel nozzle generated due to the conduction of fluids having different temperatures.

そこで、本発明の目的は、複数の燃料系統を備える燃料ノズルにおいて、導通する燃料及び燃焼空気の温度差に起因する熱応力の少ない燃料ノズル及びそれを用いたガスタービン燃焼器を提供することにある。 Therefore, an object of the present invention is to provide a fuel nozzle having a small thermal stress due to a temperature difference between conducting fuel and combustion air in a fuel nozzle provided with a plurality of fuel systems, and a gas turbine combustor using the fuel nozzle. be.

上記課題を解決するために、本発明は、複数の流路を備える燃料ノズルであって、燃料または燃焼空気が導通する第１の流路と、燃料または燃焼空気が導通し、前記第１の流路とは異なる第２の流路と、を有し、前記燃料ノズルの構成部材の内、少なくとも前記第１の流路と前記第２の流路が配置される部位は一体の部材で構成されていることを特徴とする。 In order to solve the above problems, the present invention is a fuel nozzle provided with a plurality of flow paths, wherein the fuel or combustion air conducts with the first flow path through which the fuel or combustion air conducts, and the first said. It has a second flow path different from the flow path, and among the constituent members of the fuel nozzle, at least the portion where the first flow path and the second flow path are arranged is composed of an integral member. It is characterized by being done.

また、本発明は、燃料と燃焼空気の混合気を燃焼させる燃焼室を構成する燃焼器ライナと、前記燃焼室からタービンに燃焼ガスを導く尾筒と、前記燃焼室に燃料と燃焼空気を供給するパイロットノズルと、前記パイロットノズルの周囲に複数配置され、前記燃焼室に燃料と燃焼空気を供給するメインノズルと、を備え、前記パイロットノズルは、燃料または燃焼空気が導通する第１の流路と、燃料または燃焼空気が導通し、前記第１の流路とは異なる第２の流路と、を有し、前記パイロットノズルの構成部材の内、少なくとも前記第１の流路と前記第２の流路が配置される部位は一体の部材で構成されていることを特徴とする。 Further, the present invention supplies fuel and combustion air to a combustor liner constituting a combustion chamber for burning a mixture of fuel and combustion air, a tail tube for guiding combustion gas from the combustion chamber to a turbine, and the combustion chamber. The pilot nozzle is provided with a plurality of pilot nozzles arranged around the pilot nozzle and a main nozzle for supplying fuel and combustion air to the combustion chamber, and the pilot nozzle is a first flow path through which the fuel or the combustion air is conducted. And a second flow path in which fuel or combustion air is conducted and is different from the first flow path, and at least the first flow path and the second flow path among the constituent members of the pilot nozzle are provided. The portion where the flow path is arranged is characterized in that it is composed of an integral member.

本発明によれば、複数の燃料系統を備える燃料ノズルにおいて、導通する燃料及び燃焼空気の温度差に起因する熱応力の少ない燃料ノズル及びそれを用いたガスタービン燃焼器を実現することができる。 According to the present invention, in a fuel nozzle provided with a plurality of fuel systems, it is possible to realize a fuel nozzle having less thermal stress due to a temperature difference between conducting fuel and combustion air, and a gas turbine combustor using the fuel nozzle.

これにより、信頼性及び耐久性に優れた高性能なガスタービン燃焼器を提供することができる。 This makes it possible to provide a high-performance gas turbine combustor having excellent reliability and durability.

上記した以外の課題、構成及び効果は、以下の実施形態の説明により明らかにされる。 Issues, configurations and effects other than those described above will be clarified by the following description of the embodiments.

一般的なガスタービンの構成例を示す図である。It is a figure which shows the structural example of a general gas turbine. 一般的な燃焼器の構成例を示す図である。It is a figure which shows the structural example of a general combustor. 本発明の実施例１に係る燃料ノズルの構造を示す断面図である。It is sectional drawing which shows the structure of the fuel nozzle which concerns on Example 1 of this invention. 図３のＡ－Ａ’断面図である。FIG. 3 is a cross-sectional view taken along the line AA'in FIG. 図３のＢ－Ｂ’断面図である。FIG. 3 is a cross-sectional view taken along the line BB'in FIG. 従来の燃料ノズルの構造を示す断面図である。It is sectional drawing which shows the structure of the conventional fuel nozzle. 図５のＣ－Ｃ’断面図である。FIG. 5 is a cross-sectional view taken along the line CC'of FIG. 図５のＤ－Ｄ’断面図である。FIG. 5 is a cross-sectional view taken along the line DD'in FIG.

以下、図面を用いて本発明の実施例を説明する。なお、各図面において同一の構成については同一の符号を付し、重複する部分についてはその詳細な説明は省略する。 Hereinafter, embodiments of the present invention will be described with reference to the drawings. In each drawing, the same components are designated by the same reference numerals, and the detailed description of the overlapping portions will be omitted.

先ず、図１，図２及び図５から図６Ｂを参照して、本発明の対象となるガスタービン燃焼器と従来の問題点について説明する。図１は、一般的なガスタービンの構成例を示す図である。図２は、一般的な燃焼器の構成例を示す図であり、燃焼室１５を構成する燃焼器ライナ４及び尾筒（トランジションピース）５を含む燃焼器として示している。図５は、従来のパイロットノズル７の構造を示す断面図であり、図６Ａ，図６Ｂは、それぞれ図５のＣ－Ｃ’断面、Ｄ－Ｄ’断面を示している。 First, a gas turbine combustor, which is a subject of the present invention, and conventional problems will be described with reference to FIGS. 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 is shown as a combustor including a combustor liner 4 and a tail tube (transition piece) 5 constituting a combustion chamber 15. 5 is a cross-sectional view showing the structure of the conventional pilot nozzle 7, and FIGS. 6A and 6B show a CC'cross section and a DD' cross section of FIG. 5, respectively.

図１に示すように、ガスタービンは大きく分けて圧縮機１、燃焼器２及びタービン３から構成されている。圧縮機１は大気から吸い込んだ空気を作動流体として断熱圧縮し、燃焼器２は圧縮機１から供給された圧縮空気に燃料を混合し燃焼させることで高温高圧の燃焼ガスを生成し、タービン３では燃焼器２から導入された燃焼ガスが膨張する際に回転動力を発生する。タービン３からの排気は大気中に放出される。 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 the air sucked from the atmosphere as a working fluid, and the combustor 2 produces high-temperature and high-pressure combustion gas by mixing fuel with the compressed air supplied from the compressor 1 and burning it, and the turbine 3 Then, when the combustion gas introduced from the compressor 2 expands, rotational power is generated. The exhaust gas from the turbine 3 is released into the atmosphere.

図２に示すように、燃焼器２は、燃料と燃焼空気の混合気を燃焼させる燃焼室１５を構成する燃焼器ライ４と、燃焼室１５からタービン３に燃焼ガスを導く（燃焼ガスの流れ方向８）尾筒（トランジションピース）５と、燃焼室１５に燃料と燃焼空気を供給するメインノズル６及びパイロットノズル７を備えている。メインノズル６は、上述したように、１本のパイロットノズル７の周囲に複数（例えば８本）配置されている。 As shown in FIG. 2, the combustor 2 guides combustion gas from the combustion chamber 15 to the turbine 3 and the combustor lie 4 constituting the combustion chamber 15 that burns the mixture of fuel and combustion air (combustion gas flow). Direction 8) A tail tube (transition piece) 5 and a main nozzle 6 and a pilot nozzle 7 for supplying fuel and combustion air to the combustion chamber 15 are provided. As described above, a plurality of (for example, eight) main nozzles 6 are arranged around one pilot nozzle 7.

図５に示すように、従来のパイロットノズル７は、ドリル等による穴加工により流路Ａ１３や流路Ｂ１４が予め形成されたノズル構成部材９，１０，１１を接合部１２で互いに接合することで構成されている。ノズル構成部材９，１０，１１の接合には、例えば、ロウ付けによる溶接が用いられる。 As shown in FIG. 5, in the conventional pilot nozzle 7, the nozzle constituent members 9, 10 and 11 having the flow paths A13 and the flow path B14 formed in advance by drilling holes or the like are joined to each other at the joint portion 12. It is configured. For joining the nozzle components 9, 10 and 11, for example, welding by brazing is used.

一般的に、ガスタービンの定格負荷時には、流路Ａ１３に相対的に温度の高いスイープ空気（燃焼空気）が導通し、流路Ｂ１４に相対的に温度の低い天然ガス等の燃料が導通する。このため、パイロットノズル７の主として径方向の温度差、及びそれに起因する径方向、軸方向の熱伸び差により、熱応力が発生する。一般に溶接部においては、未溶着部等による形状の不連続性に起因して熱応力が助長されやすく、また溶接部では母材に比べて疲労強度が低下する。 Generally, at the rated load of the gas turbine, sweep air (combustion air) having a relatively high temperature conducts to the flow path A13, and fuel such as natural gas having a relatively low temperature conducts to the flow path B14. Therefore, thermal stress is generated mainly due to the temperature difference mainly in the radial direction of the pilot nozzle 7 and the thermal expansion difference in the radial direction and the axial direction caused by the temperature difference. Generally, in the welded portion, thermal stress is likely to be promoted due to the discontinuity of the shape due to the unwelded portion or the like, and the fatigue strength of the welded portion is lower than that of the base metal.

以上より、従来のパイロットノズル７では、特に流路Ａ１３と流路Ｂ１４の両方が配置される部位の接合部１２が、流路Ａ１３と流路Ｂ１４のそれぞれを導通する燃料または燃焼空気の温度差に起因して強度のボトルネックとなり、低サイクル疲労により運用が制限される。 From the above, in the conventional pilot nozzle 7, in particular, the temperature difference between the fuel or the combustion air in which the joint portion 12 at the portion where both the flow path A13 and the flow path B14 are arranged conducts each of the flow path A13 and the flow path B14. This causes a strong bottleneck, and low cycle fatigue limits operation.

また、図６Ａに示すように、従来のパイロットノズル７の根元部においては、流路Ａ１３と流路Ｂ１４の両方がパイロットノズル７の周方向に環状（アニュラ状：Annular）に配置されているため、パイロットノズル７は径方向において流路Ａ１３及び流路Ｂ１４により熱的に分断された構造となる。このため、流路Ａ１３と流路Ｂ１４のそれぞれを導通する燃料または燃焼空気の温度差によるパイロットノズル７への熱応力がさらに助長される。 Further, as shown in FIG. 6A, in the root portion of the conventional pilot nozzle 7, both the flow path A13 and the flow path B14 are arranged in an annular shape in the circumferential direction of the pilot nozzle 7. The pilot nozzle 7 has a structure that is thermally separated by the flow path A13 and the flow path B14 in the radial direction. Therefore, the thermal stress on the pilot nozzle 7 due to the temperature difference between the fuel or the combustion air conducting each of the flow paths A13 and the flow path B14 is further promoted.

また、図６Ｂに示すように、従来のパイロットノズル７の先端近傍においては、流路Ｂ１４はパイロットノズル７の周方向に複数に分割して配置されているが、流路Ａ１３は根元部と同様に、パイロットノズル７の周方向に環状（アニュラ状）に配置されており、パイロットノズル７は径方向において流路Ａ１３により熱的に分断された構造となる。 Further, as shown in FIG. 6B, in the vicinity of the tip of the conventional pilot nozzle 7, the flow path B14 is divided into a plurality of parts in the circumferential direction of the pilot nozzle 7, but the flow path A13 is the same as the root portion. In addition, the pilot nozzle 7 is arranged in an annular shape (annular shape) in the circumferential direction, and the pilot nozzle 7 has a structure that is thermally divided by the flow path A13 in the radial direction.

次に、図３から図４Ｂを参照して、本発明の実施例１に係る燃料ノズルについて説明する。図３は、本実施例のパイロットノズル７の構造を示す断面図であり、図４Ａ，図４Ｂは、それぞれ図３のＡ－Ａ’断面、Ｂ－Ｂ’断面を示している。 Next, the fuel nozzle according to the first embodiment of the present invention will be described with reference to FIGS. 3 to 4B. FIG. 3 is a cross-sectional view showing the structure of the pilot nozzle 7 of this embodiment, and FIGS. 4A and 4B show a cross section taken along the line AA'and a cross section taken along the line BB'of FIG. 3, respectively.

図３に示すように、本実施例のパイロットノズル７は、燃料または燃焼空気が導通する流路Ａ１３（第１の流路）と、燃料または燃焼空気が導通し、流路Ａ１３（第１の流路）とは異なる流路Ｂ１４（第２の流路）を有しており、パイロットノズル７のノズル構成部材９，１０の内、少なくとも流路Ａ１３（第１の流路）と流路Ｂ１４（第２の流路）の両方が配置されている部位は、接合部１２の無い一体のノズル構成部材１０で構成されている。 As shown in FIG. 3, the pilot nozzle 7 of this embodiment has a flow path A13 (first flow path) through which fuel or combustion air conducts, and a flow path A13 (first flow path) through which fuel or combustion air conducts. It has a flow path B14 (second flow path) different from the flow path), and among the nozzle constituent members 9 and 10 of the pilot nozzle 7, at least the flow path A13 (first flow path) and the flow path B14. The portion where both (second flow paths) are arranged is composed of an integral nozzle component 10 without a joint portion 12.

図３のように、流路Ａ１３（第１の流路）と流路Ｂ１４（第２の流路）の両方が配置される部位を接合部１２の無い一体のノズル構成部材１０で構成することにより、上述したような、流路Ａ１３と流路Ｂ１４のそれぞれを導通する燃料または燃焼空気の温度差に起因して接合部１２が強度のボトルネックとなるのを防ぐことができ、パイロットノズル７の信頼性及び耐久性を向上することができる。 As shown in FIG. 3, the portion where both the flow path A13 (first flow path) and the flow path B14 (second flow path) are arranged is configured by the integrated nozzle constituent member 10 without the joint portion 12. Thereby, as described above, it is possible to prevent the joint portion 12 from becoming a strength bottleneck due to the temperature difference between the fuel or the combustion air conducting each of the flow paths A13 and the flow path B14, and the pilot nozzle 7 It is possible to improve the reliability and durability of the.

また、図４Ａ及び図４Ｂに示すように、本実施例のパイロットノズル７は、流路Ａ１３（第１の流路）及び流路Ｂ１４（第２の流路）が、パイロットノズル７の周方向において、いずれも複数に分割して配置されている。 Further, as shown in FIGS. 4A and 4B, in the pilot nozzle 7 of this embodiment, the flow path A13 (first flow path) and the flow path B14 (second flow path) are in the circumferential direction of the pilot nozzle 7. In, all of them are divided into a plurality of arrangements.

図４Ａ及び図４Ｂのように、流路Ａ１３（第１の流路）と流路Ｂ１４（第２の流路）の両方を、パイロットノズル７の周方向において、複数に分割して配置することで、パイロットノズル７が径方向において、流路Ａ１３（第１の流路）及び流路Ｂ１４（第２の流路）により熱的に完全に分断されるのを防いでいる。これにより、流路Ａ１３と流路Ｂ１４のそれぞれを導通する燃料または燃焼空気の温度差によるパイロットノズル７への熱応力を緩和することができる。 As shown in FIGS. 4A and 4B, both the flow path A13 (first flow path) and the flow path B14 (second flow path) are divided into a plurality of parts in the circumferential direction of the pilot nozzle 7 and arranged. This prevents the pilot nozzle 7 from being completely thermally separated by the flow path A13 (first flow path) and the flow path B14 (second flow path) in the radial direction. This makes it possible to relieve the thermal stress on the pilot nozzle 7 due to the temperature difference between the fuel or the combustion air conducting each of the flow paths A13 and the flow path B14.

例えば、流路Ａ１３（第１の流路）に燃焼空気を導通させ、流路Ｂ１４（第２の流路）に燃焼空気より温度の低い燃料を導通させるような場合であっても、燃料及び燃焼空気の温度差によるパイロットノズル７への熱応力が緩和されるため、接合部１２の無い一体のノズル構成部材１０で構成することによる効果に加えて、パイロットノズル７の信頼性及び耐久性をさらに向上することができる。 For example, even in the case where 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), the fuel and the fuel and the flow path B14 (second flow path) are conducted. Since the thermal stress on the pilot nozzle 7 due to the temperature difference of the combustion air is relaxed, in addition to the effect of being composed of the integrated nozzle constituent member 10 without the joint portion 12, the reliability and durability of the pilot nozzle 7 are improved. It can be further improved.

また、図３に示すように、パイロットノズル７の先端近傍のノズル構成部材９は、流路Ａ１３（第１の流路）及び流路Ｂ１４（第２の流路）の内、流路Ａ１３（第１の流路）のみが配置されており、流路Ａ１３（第１の流路）のみが配置されているノズル構成部材９は、流路Ａ１３（第１の流路）と流路Ｂ１４（第２の流路）の両方が配置されているノズル構成部材１０と、例えば、ロウ付けによる溶接または熱間等方圧加圧（ＨＩＰ：Hot Isostatic Pressing）法により接合されている。 Further, as shown in FIG. 3, the nozzle constituent member 9 in the vicinity of the tip of the pilot nozzle 7 is the flow path A13 (the flow path A13 (second flow path) in the flow path A13 (first flow path) and the flow path B14 (second flow path). The nozzle component member 9 in which only the flow path (first flow path) is arranged and only the flow path A13 (first flow path) is arranged includes the flow path A13 (first flow path) and the flow path B14 (the first flow path). It is joined to the nozzle component 10 in which both of the second flow paths) are arranged, for example, by welding by brazing or by a hot isostatic pressing (HIP) method.

図３のように、流路Ａ１３（第１の流路）及び流路Ｂ１４（第２の流路）の内、一方の流路Ａ１３（第１の流路）のみが形成されている領域に限定して接合部１２が配置されるような構造とすることで、流路を導通する燃料または燃焼空気の温度差によるパイロットノズル７への熱応力の発生を防止することができ、接合部１２の接合信頼性を担保することができる。 As shown in FIG. 3, in the region where only one of the flow paths A13 (first flow path) and the flow path B14 (second flow path) is formed, as shown in FIG. By making the structure such that the joint portion 12 is arranged in a limited manner, it is possible to prevent the generation of thermal stress on the pilot nozzle 7 due to the temperature difference of the fuel or the combustion air conducting through the flow path, and the joint portion 12 can be prevented. Bonding reliability can be guaranteed.

なお、ノズル構成部材９とノズル構成部材１０との接合部１２での接合には、上述した熱間等方圧加圧（ＨＩＰ：Hot Isostatic Pressing）法を用いるのが望ましい。熱間等方圧加圧（ＨＩＰ）法を用いることにより、未溶着部を極力無くすことができるため、接合部１２における形状の不連続性に起因する熱応力を抑制することができる。 It is desirable to use the above-mentioned hot isostatic pressing (HIP) method for joining the nozzle constituent member 9 and the nozzle constituent member 10 at the joint portion 12. By using the hot isotropic pressure pressurization (HIP) method, the unwelded portion can be eliminated as much as possible, so that the thermal stress caused by the discontinuity of the shape at the joint portion 12 can be suppressed.

以上説明したように、本発明によれば、導通する燃料及び燃焼空気の温度差に起因する熱応力の少ない燃料ノズル及びそれを用いたガスタービン燃焼器を実現することができ、ガスタービン燃焼器の信頼性及び耐久性を向上することができる。 As described above, according to the present invention, it is possible to realize a fuel nozzle having less thermal stress due to the temperature difference between the conductive fuel and the combustion air and a gas turbine combustor using the fuel nozzle, and the gas turbine combustor. The reliability and durability of the fuel can be improved.

なお、本発明は上記した実施例に限定されるものではなく、様々な変形例が含まれる。例えば、上記した実施例は本発明を分かりやすく説明するために詳細に説明したものであり、必ずしも説明した全ての構成を備えるものに限定されるものではない。また、ある実施例の構成の一部を他の実施例の構成に置き換えることが可能であり、また、ある実施例の構成に他の実施例の構成を加えることも可能である。また、各実施例の構成の一部について、他の構成の追加・削除・置換をすることが可能である。 The present invention is not limited to the above-described embodiment, and includes various modifications. For example, the above-described embodiment has been described in detail in order to explain the present invention in an easy-to-understand manner, and is not necessarily limited to the one including all the described configurations. Further, it is possible to replace a 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. Further, it is possible to add / delete / replace a part of the configuration of each embodiment with another configuration.

１…圧縮機
２…燃焼器
３…タービン
４…燃焼器ライナ
５…尾筒（トランジションピース）
６…メインノズル
７…パイロットノズル
８…燃焼ガスの流れ方向
９，１０，１１…ノズル構成部材
１２…接合部
１３…流路Ａ
１４…流路Ｂ
１５…燃焼室 1 ... Compressor 2 ... Combustor 3 ... Turbine 4 ... Combustor liner 5 ... Tail tube (transition piece)
6 ... Main nozzle 7 ... Pilot nozzle 8 ... Combustion gas flow direction 9, 10, 11 ... Nozzle component 12 ... Joint 13 ... Flow path A
14 ... Flow path B
15 ... Combustion chamber

Claims (10)

複数の流路を備える燃料ノズルであって、
燃料または燃焼空気が導通する第１の流路と、
燃料または燃焼空気が導通し、前記第１の流路とは異なる第２の流路と、を有し、
前記燃料ノズルの構成部材の内、少なくとも前記第１の流路と前記第２の流路が配置される部位は一体の部材で構成されていることを特徴とする燃料ノズル。 A fuel nozzle with multiple channels
A first flow path through which fuel or combustion air conducts,
The fuel or combustion air conducts and has a second flow path different from the first flow path.
The fuel nozzle is characterized in that at least a portion of the constituent members of the fuel nozzle in which the first flow path and the second flow path are arranged is composed of an integral member. 請求項１に記載の燃料ノズルであって、
前記第１の流路および前記第２の流路は、前記燃料ノズルの周方向において、いずれも複数に分割して配置されていることを特徴とする燃料ノズル。 The fuel nozzle according to claim 1.
The fuel nozzle is characterized in that the first flow path and the second flow path are each divided into a plurality of parts in the circumferential direction of the fuel nozzle. 請求項１または２に記載の燃料ノズルであって、
前記第１の流路を燃焼空気が導通し、
前記第２の流路を前記燃焼空気より温度の低い燃料が導通することを特徴とする燃料ノズル。 The fuel nozzle according to claim 1 or 2.
Combustion air conducts through the first flow path,
A fuel nozzle characterized in that a fuel having a temperature lower than that of the combustion air conducts through the second flow path. 請求項１から３のいずれか１項に記載の燃料ノズルであって、
前記燃料ノズルの先端近傍は、前記第１の流路および前記第２の流路の内、前記第１の流路のみが配置されており、
前記第１の流路のみが配置されている部位は、前記第１の流路と前記第２の流路が配置されている部位と接合されていることを特徴とする燃料ノズル。 The fuel nozzle according to any one of claims 1 to 3.
In the vicinity of the tip of the fuel nozzle, only the first flow path of the first flow path and the second flow path is arranged.
A fuel nozzle characterized in that 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. 請求項４に記載の燃料ノズルであって、
前記第１の流路のみが配置されている部位は、前記第１の流路と前記第２の流路が配置されている部位と、溶接または熱間等方圧加圧（ＨＩＰ）法により接合されていることを特徴とする燃料ノズル。 The fuel nozzle according to claim 4.
The portion where only the first flow path is arranged is formed by welding or hot isotropic pressure pressurization (HIP) method with the portion where the first flow path and the second flow path are arranged. A fuel nozzle characterized by being joined. 燃料と燃焼空気の混合気を燃焼させる燃焼室を構成する燃焼器ライナと、
前記燃焼室からタービンに燃焼ガスを導く尾筒と、
前記燃焼室に燃料と燃焼空気を供給するパイロットノズルと、
前記パイロットノズルの周囲に複数配置され、前記燃焼室に燃料と燃焼空気を供給するメインノズルと、を備え、
前記パイロットノズルは、燃料または燃焼空気が導通する第１の流路と、
燃料または燃焼空気が導通し、前記第１の流路とは異なる第２の流路と、を有し、
前記パイロットノズルの構成部材の内、少なくとも前記第１の流路と前記第２の流路が配置される部位は一体の部材で構成されていることを特徴とするガスタービン燃焼器。 A combustor liner that constitutes a combustion chamber that burns a mixture of fuel and combustion air,
The tail tube that guides the combustion gas from the combustion chamber to the turbine,
A pilot nozzle that supplies fuel and combustion air to the combustion chamber,
A plurality of main nozzles arranged around the pilot nozzle and supplying fuel and combustion air to the combustion chamber are provided.
The pilot nozzle has a first flow path through which fuel or combustion air conducts, and
The fuel or combustion air conducts and has a second flow path different from the first flow path.
A gas turbine combustor, characterized in that at least a portion of the components of the pilot nozzle in which the first flow path and the second flow path are arranged is composed of an integral member. 請求項６に記載のガスタービン燃焼器であって、
前記第１の流路および前記第２の流路は、前記パイロットノズルの周方向において、いずれも複数に分割して配置されていることを特徴とするガスタービン燃焼器。 The gas turbine combustor according to claim 6.
A gas turbine combustor, wherein the first flow path and the second flow path are each divided into a plurality of parts in the circumferential direction of the pilot nozzle. 請求項６または７に記載のガスタービン燃焼器であって、
前記第１の流路を燃焼空気が導通し、
前記第２の流路を前記燃焼空気より温度の低い燃料が導通することを特徴とするガスタービン燃焼器。 The gas turbine combustor according to claim 6 or 7.
Combustion air conducts through the first flow path,
A gas turbine combustor characterized in that a fuel having a temperature lower than that of the combustion air conducts through the second flow path. 請求項６から８のいずれか１項に記載のガスタービン燃焼器であって、
前記パイロットノズルの先端近傍は、前記第１の流路および前記第２の流路の内、前記第１の流路のみが配置されており、
前記第１の流路のみが配置されている部位は、前記第１の流路と前記第２の流路が配置されている部位と接合されていることを特徴とするガスタービン燃焼器。 The gas turbine combustor according to any one of claims 6 to 8.
In the vicinity of the tip of the pilot nozzle, only the first flow path is arranged among the first flow path and the second flow path.
A gas turbine combustor, wherein 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. 請求項９に記載のガスタービン燃焼器であって、
前記第１の流路のみが配置されている部位は、前記第１の流路と前記第２の流路が配置されている部位と、溶接または熱間等方圧加圧（ＨＩＰ）法により接合されていることを特徴とするガスタービン燃焼器。 The gas turbine combustor according to claim 9.
The portion where only the first flow path is arranged is formed by welding or hot isotropic pressure pressurization (HIP) method with the portion where the first flow path and the second flow path are arranged. A gas turbine combustor characterized by being joined.
JP2020155193A 2020-09-16 2020-09-16 Fuel nozzle, and gas turbine combustor Pending JP2022049136A (en)

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US17/405,372 US20220082260A1 (en) 2020-09-16 2021-08-18 Combustor Fuel Nozzle Structure
CN202111081539.0A CN114263930A (en) 2020-09-16 2021-09-15 Fuel nozzle and gas turbine combustor
DE102021210300.6A DE102021210300A1 (en) 2020-09-16 2021-09-16 Fuel nozzle and gas turbine combustor
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