WO2025103669A1 - Brûleur à turbine à gaz - Google Patents

Brûleur à turbine à gaz Download PDF

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Publication number
WO2025103669A1
WO2025103669A1 PCT/EP2024/078461 EP2024078461W WO2025103669A1 WO 2025103669 A1 WO2025103669 A1 WO 2025103669A1 EP 2024078461 W EP2024078461 W EP 2024078461W WO 2025103669 A1 WO2025103669 A1 WO 2025103669A1
Authority
WO
WIPO (PCT)
Prior art keywords
burner
combustion air
fuel
gas turbine
pilot
Prior art date
Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
Pending
Application number
PCT/EP2024/078461
Other languages
German (de)
English (en)
Inventor
Stefan Clauss
Stefan Hoffmann
Christian Steinbach
Bernhard Cosic
David Kluß
Dominik Wassmer
Marc BLÜMER
Current Assignee (The listed assignees may be inaccurate. Google has not performed a legal analysis and makes no representation or warranty as to the accuracy of the list.)
MAN Energy Solutions SE
Original Assignee
MAN Energy Solutions SE
Priority date (The priority date is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the date listed.)
Filing date
Publication date
Application filed by MAN Energy Solutions SE filed Critical MAN Energy Solutions SE
Publication of WO2025103669A1 publication Critical patent/WO2025103669A1/fr
Pending legal-status Critical Current
Anticipated expiration legal-status Critical

Links

Classifications

    • 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
    • F23R2900/00—Special features of, or arrangements for continuous combustion chambers; Combustion processes therefor
    • F23R2900/00002—Gas turbine combustors adapted for fuels having low heating value [LHV]
    • 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
    • F23R2900/00014—Reducing thermo-acoustic vibrations by passive means, e.g. by Helmholtz resonators
    • 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
    • F23R2900/03343—Pilot burners operating in premixed mode

Definitions

  • the invention relates to a gas turbine burner.
  • DE 10 2018 125 848 A1 discloses a gas turbine burner of a gas turbine configured to burn a mixture of combustion air and a gaseous fuel in a gas fuel operating mode and liquid fuel in the presence of combustion air in a liquid fuel operating mode.
  • the gas turbine burner disclosed therein has a combustion chamber having a flame tube and a prechamber upstream of the flame tube.
  • a mixture of the gaseous fuel and the combustion air is supplied to the combustion chamber via a main swirl body.
  • liquid fuel operating mode liquid fuel is supplied to the combustion chamber via an atomization device, and the combustion air is supplied to the combustion chamber via the main swirl body.
  • DE 10 2020 116 245 A1 discloses another gas turbine burner with a combustion chamber comprising a flame tube and a prechamber. Combustion air is guided past the outside of the flame tube and supplied to the combustion chamber via a swirl body. The mixing of the combustion air with a gaseous fuel takes place in the area of the swirl body. Part of the combustion air can penetrate past the swirl body directly into the flame tube.
  • DE 10 2021 123 513 A1 covers a gas turbine burner designed as a swirl burner.
  • the swirl burner has several swirl tubes wound in a spiral or helix-like manner around a central axis, allowing gaseous fuel and combustion air to be mixed in the inlet area of the swirl tubes.
  • the outlet areas of the swirl tubes open into a combustion chamber.
  • DE 197 37 997 A1 and DE 195 10 743 A1 disclose gas turbine burners designed as conical burners. In the area of the conical burner, also referred to as a conical burner, gas can be mixed with combustion air to provide the actual combustion chamber with a mixture of gas and combustion air.
  • the object of the invention is therefore to provide a gas turbine burner that can stably combust both natural gas and hydrogen with low emissions. This object is achieved by a gas turbine burner according to claim 1.
  • the gas turbine burner according to the invention has a combustion chamber which has a flame tube, wherein the flame tube defines a combustion zone of the gas turbine burner for burning a fuel in the presence of combustion air.
  • the gas turbine burner according to the invention has a main burner and a pilot burner, wherein the main burner encloses the pilot burner radially outwardly.
  • the main burner of the gas turbine burner according to the invention is designed as a jet burner with preferably between 40 and 120 straight jet tubes and is designed to introduce a mixture of fuel and combustion air into a radially outer section of the combustion zone in a straight flow.
  • the pilot burner of the gas turbine burner according to the invention is designed either as a swirl burner with a defined number of spiral or helical swirl tubes or as a conical burner or cone burner and is configured to introduce a mixture of fuel and combustion air into a radially inner section of the combustion zone in a wired flow.
  • a ratio FH/FP between the combustion air flow area FH of the main burner and the combustion air flow area FP of the pilot burner is between 3 and 6.
  • the combustion air flow area FH of the main burner is calculated from the sum of the smallest cross-sectional areas of all jet pipes of the main burner along their length or axial extent. Assuming that all jet pipes are identical, the combustion air flow area FH of the main burner is calculated from the smallest cross-sectional area of one of the jet pipes multiplied by the number of jet pipes.
  • the combustion air flow area FP of the pilot burner if the burner is designed as a swirl burner, is calculated from the sum of the smallest cross-sectional areas of all swirl tubes along their length or axial extent. Assuming that all swirl tubes are of the same design, the combustion air flow area FP of the pilot burner designed as a swirl burner is calculated from the smallest cross-sectional area of one of the swirl tubes multiplied by the number of swirl tubes.
  • the combustion air flow area FP of the pilot burner results, if the same is designed as a conical burner or conical burner, from the smallest cross-sectional area of a cavity of the conical burner or conical burner, the body of the conical burner or conical burner, the segments of a cone or cone of the conical burner or conical burner, limit radially inward.
  • the gas turbine burner according to the invention comprises a main burner and a pilot burner.
  • the main burner is designed as a jet burner and radially surrounds the pilot burner.
  • the pilot burner is either a swirl burner with swirl tubes or a conical burner. Both hydrogen and natural gas can be burned stably with low emissions.
  • the main burner is configured to mix fuel and combustion air in the region of the straight jet pipes in the inlet region of the jet pipes, wherein a ratio l/d between a mixing section l of the jet pipes and a diameter d of the jet pipes is greater than 5, preferably greater than 10.
  • This diameter d of the jet pipes is the smallest diameter of the same along their length or axial extent.
  • the swirl burner is designed to mix fuel and combustion air in the region of the swirl tubes in the inlet region thereof, wherein a ratio l/d between a mixing section l of the swirl tubes and a diameter d of the swirl tubes is greater than 5, preferably greater than 10.
  • This diameter d of the swirl tubes is the smallest diameter of the same along their length or axial extent.
  • the pilot burner is designed as a conical burner, it is configured to mix a first portion of the gas with combustion air immediately upstream of the flame tube and a second, smaller portion of the gas with combustion air further upstream of the flame tube. This is preferred for stably combusting both natural gas and hydrogen with low emissions.
  • the ratio FH/FP between the combustion air flow area FH of the main burner and the combustion air flow area FP of the pilot burner is between 3 and 5, preferably between 3.5 and 4.5, or between 4 and 6, preferably between 4.5 and 5.5. These ratios between the combustion air flow area of the main burner and the combustion air flow area of the pilot burner are particularly suitable for stably combusting both natural gas and hydrogen with low emissions.
  • the main burner preferably has between 40 and 100, preferably between 60 and 80, straight jet tubes. This is also preferred for stable combustion of both natural gas and hydrogen with low emissions.
  • a resonator chamber is formed between the main burner and the pilot burner, which opens into the combustion zone of the gas turbine burner defined by the flame tube. This prevents self-excited acoustic instabilities during operation.
  • the resonator chamber which is formed between the main burner and the pilot burner, acoustically stabilizes the gas turbine burner and reduces the risk of damage to it.
  • the gas turbine burner comprises a fuel gallery configured to supply fuel to the main burner and the pilot burner.
  • the fuel gallery is plate-shaped at a portion facing the main burner and the pilot burner, and roof-shaped or curved at a portion facing away from the main burner and the pilot burner. This makes it possible to reduce thermal stresses in the fuel gallery and thus reduce the risk of damage to the fuel gallery due to material embrittlement and/or plastic deformation.
  • the features relating to the fuel gallery and the features relating to the resonator chamber can also be used advantageously in a gas turbine combustor independently of the features relating to the FH/FP ratio and the l/d ratios.
  • Fig. 1 shows a partial cross-section through a first gas turbine burner according to the invention
  • Fig. 2 shows a partial cross-section through a second gas turbine burner according to the invention.
  • Fig. 1 shows a partial cross-section of a first gas turbine burner 10 according to the invention.
  • the gas turbine burner 10 has a combustion chamber 11 comprising a flame tube 12.
  • the flame tube 12 defines a combustion zone 13 of the gas turbine burner 10, within which fuel is combusted in the presence of combustion air.
  • the gas turbine burner 10 has a main burner 14 and a pilot burner 15, with the main burner 14 radially surrounding the pilot burner 15.
  • a mixture of fuel and combustion air can be introduced into a radially outer portion of the combustion zone 13, which is defined by the flame tube 12, via the main burner 14.
  • a mixture of fuel and combustion air can be introduced into a radially inner portion of the combustion zone 13 via the pilot burner 15.
  • the main burner 14 shown in Fig. 1 is designed as a jet burner with rectilinear jet tubes 16, wherein the main burner 14 or the rectilinear jet tubes 16 are designed and configured to introduce the mixture of fuel and combustion air into the radially outer portion of the combustion zone 13 defined by the flame tube 12 in a rectilinear flow. All jet tubes 16 of the main burner 14 can be of the same and thus identical design.
  • the pilot burner 15, positioned radially inside the main burner 14, is designed in the embodiment of Fig. 1 as a swirl burner with spiral or helical swirl tubes 17, wherein the swirl burner is configured to introduce a mixture of fuel and combustion air into the radially inner section of the combustion zone 13 in a continuous flow.
  • All swirl tubes 17 of the pilot burner 15 can be of the same and thus identical design.
  • Fig. 1 shows, viewed in the direction of flow through the jet pipes 16 and the swirl pipes 17, upstream of the same fuel nozzles 18, wherein a fuel nozzle 18 cooperates with each jet pipe 16 and each swirl pipe 17 in order to introduce fuel into the respective jet pipe 16 or respective swirl pipe 17 on an inlet side thereof.
  • Combustion air can also be introduced into the jet tubes 16 and swirl tubes 17 at this inlet side.
  • This combustion air initially flows past the outside of the flame tube 12 according to the arrows 19 shown in Fig. 1 and then flows through a flow conditioner 20 toward the fuel nozzles 18 to flow around the fuel nozzles 18 and, after flowing around the fuel nozzles 18, enters the jet tubes 16 and swirl tubes 17 in the inlet area.
  • the fuel nozzles 18 can be supplied with fuel from a fuel gallery 21.
  • the combustion air flow can be uniformed via the flow conditioner 20.
  • a ratio FH/FP between the combustion air flow area FH of the main burner 14 and the combustion air flow area FP of the pilot burner 15 is between 3 and 6, preferably between 3 and 5, particularly preferably between 3.5 and 4.5, or between 4 and 6, preferably between 4.5 and 5.5.
  • the area FH of the main burner 14 through which combustion air flows is calculated from the sum of the smallest cross-sectional areas of all jet tubes 16 along their length or axial extent. Assuming that all jet tubes 16 are of identical design, the area FH of the main burner 14 through which combustion air flows is calculated from the smallest cross-sectional area of one of the jet tubes 16 multiplied by the number of jet tubes 16.
  • the area FP of the pilot burner 15 designed as a swirl burner through which combustion air flows is calculated from the sum of the smallest cross-sectional areas of all swirl tubes 17 along their length or axial extent. Assuming that all swirl tubes 17 are of identical design, the area FP of the pilot burner 15 designed as a swirl burner through which combustion air flows is calculated from the smallest cross-sectional area of one of the swirl tubes 17 multiplied by the number of swirl tubes 17.
  • Such a gas turbine burner is particularly preferred for burning both natural gas and hydrogen with full flexibility, i.e. both 100% natural gas and 100% hydrogen as well as any mixtures of natural gas and hydrogen, stably with low emissions.
  • the main burner 14, designed as a jet burner preferably has between 40 and 120, in particular between 60 and 100, particularly preferably between 60 and 80, straight jet pipes 16. These extend straight from the inlet side to the outlet side thereof, wherein in the region of the inlet side, gas and combustion air enters the jet pipes 16 and in the region of the outlet side thereof, the mixture of fuel and combustion air exits from the same and enters the combustion zone 13.
  • the pilot burner 15, which is designed as a swirl burner in Fig. 1, has in particular between 10 and 35, preferably between 15 and 30, particularly preferably between 15 and 25, spiral or helical swirl tubes 17.
  • the winding direction of the spiral or helical swirl tubes 17 extends around a longitudinal center axis of the pilot burner 15, which preferably coincides with a longitudinal center axis of the flame tube 12.
  • Fig. 2 shows a second embodiment of a gas turbine burner 10, which differs from the embodiment of Fig. 1 only in that the pilot burner 15, which is surrounded radially on the outside by the main burner 14, is not designed as a swirl burner with a plurality of swirl tubes 17, but rather as a conical burner.
  • a pilot burner 15 designed as a conical burner provides a wired flow of a mixture of fuel and combustion air, which it introduces into the radially inner section of the combustion zone 13 defined or delimited by the flame tube 12.
  • Fig. 2 shows bodies 22 that form segments of a cone or cone. These bodies 22 are designed in particular as conical or cone-like shells with an extension of approximately 90°, which are offset from one another parallel to their central axis with the formation of slots 23 between them. Combustion air reaches a cavity defined radially inward by the conical or cone-like bodies 22 via the slots 23, wherein fuel can be introduced into this cavity via a nozzle 24 and via fuel holes 25.
  • gas is mixed with combustion air directly upstream of the combustion zone 13 and thus directly upstream of the flame tube 12 by fuel flowing radially inward via the fuel openings 25 into the space defined by the bodies 22, where combustion air entering via the slots 23 is mixed with the same.
  • fuel can be introduced upstream via the nozzle 24 into the cavity delimited by the conical or cone-like bodies 22.
  • the area FP of the pilot burner 15 through which combustion air flows corresponds to the smallest cross-sectional area of the cavity of the conical burner or conical burner, which is delimited radially on the inside by the bodies 22 of the conical burner or conical burner, which form the segments of the cone or cone of the conical burner or conical burner.
  • a premixing of fuel and combustion air takes place both in the area of the main burner 14 and in the area of the respective pilot burner 15. In the area of the main burner 14, this mixing takes place on the inlet side in the area of the jet pipes 16, with the jet pipes 16 then providing a mixing section for the fuel and the combustion air.
  • a ratio l/d between the mixing section l of the jet pipes 16 and a diameter d thereof is preferably greater than 5, particularly preferably greater than 10. This diameter d of the jet pipes 16 of the ratio l/d is the smallest diameter of the jet pipes along their length or axial extent.
  • a premixing of fuel and combustion air also takes place in the swirl tubes 17 in the area of the pilot burner 15 designed as a swirl burner, so that the swirl tubes 17 then provide a mixing section for fuel and combustion air, the ratio l/d of which between the mixing section l of the swirl tubes and the diameter d of the swirl tubes is again greater than 5, preferably greater than 10.
  • This diameter d of the swirl tubes 17 of the ratio l/d is the smallest diameter of the same along their length or axial extent.
  • a premixing of fuel and combustion air also takes place upstream of the combustion zone 13, namely on the one hand directly upstream of the flame tube 12 in the area of the fuel openings 25 and on the other hand upstream thereof in the area of the nozzle 24.
  • a resonator chamber 26 is formed between the main burner 14 and the pilot burner 15.
  • This resonator chamber 26 is delimited by a wall 27 and encloses an air volume, which opens into the combustion zone 13 delimited by the flame tube 12 via at least one resonator neck 29 with an opening 28.
  • the resonator chamber 26 acts as a Helmholtz resonator and provides acoustic damping for the gas turbine burner 10.
  • the resonator frequency depends in particular on the length of the resonator neck 29 and the volume of the resonator chamber 26.
  • the gas turbine burner 10 can thus be acoustically stabilized, thereby reducing the risk of damage to the same due to thermoacoustic instabilities.
  • the at least one resonator neck 29 also provides cooling air to cool the burner surface.
  • the features relating to the resonator chamber 26 can also be used advantageously in a gas turbine burner independently of the features relating to the FH/FP ratio and the l/d ratios.
  • fuel can be supplied to both the main burner 14 and the pilot burner 15 from the fuel gallery 21.
  • the fuel gallery 21 has a plate-like contour on a section 21a facing the main burner 14 and pilot burner 15, while the fuel gallery 21 has a roof-like or curved shape on a section 21b facing away from the main burner 14 and pilot burner 15. This allows for the compensation of deformations in the area of the fuel gallery caused by thermal temperature cycles.
  • the risk of excessively high thermal stresses in the fuel gallery 21 can be reduced, thereby counteracting material embrittlement or plastic deformation of the fuel gallery 21.
  • the features relating to the fuel gallery 21 can also be used advantageously in a gas turbine burner independently of the features relating to the FH/FP ratio and the l/d ratios.
  • Both the main burner 14 with its jet tubes 16 and the pilot burner 15 and the fuel gallery 21 can be manufactured by 3D printing. This also applies to the pilot burner 15 in Fig. 2, which is designed as a conical burner, and the pilot burner 15 in Fig. 1, which is designed as a swirl burner.
  • thermocouples to the jet pipes 16 of the jet burner 14 as well as to the swirl pipes 17 of the pilot burner 15 in Fig. 1, which is designed as a swirl burner, in order to detect a possible flashback into the pipes with the aid of these thermocouples.
  • the gas turbine burner 10 accordingly combines a main burner 14 designed as a jet burner with a pilot burner 15 which, in contrast to the jet burner 14, introduces the mixture of fuel and combustion air into the combustion zone 13 of the flame tube 12 not via a straight-line flow, but rather via a wired flow, in such a way that the wired flow of fuel and combustion air enters a radially inner region of the combustion zone 13, whereas the straight-line flow of combustion air and fuel, which is available to the main burner 14, enters a radially outer section of the combustion zone 13.
  • Both natural gas and hydrogen can be burned with high stability and low emissions, with a flexible natural gas and hydrogen blend ratio between 100% hydrogen and 100% natural gas.
  • the fuel distribution between the main burner 14 and the pilot burner 15 can be flexibly adjusted, in particular between partial load operation and full load operation.
  • a resonator chamber 27 can be integrated into the gas turbine burner 10 in order to acoustically stabilize the gas turbine burner 10.
  • the fuel gallery 21 of the gas turbine burner 10 is preferably designed to avoid thermal stresses, as well as material embrittlement and plastic deformation thereof.
  • a geometry integrated on the main burner 14 and/or pilot burner 15 can provide a holder for thermocouples in order to provide flashback detection.

Landscapes

  • Engineering & Computer Science (AREA)
  • Chemical & Material Sciences (AREA)
  • Combustion & Propulsion (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Pre-Mixing And Non-Premixing Gas Burner (AREA)

Abstract

L'invention se rapporte à un brûleur à turbine à gaz (10), comprenant : une chambre de combustion (11) qui comporte un tube à flamme (12), le tube à flamme (12) définissant une zone de combustion (13) du brûleur à turbine à gaz pour brûler un combustible en présence d'air de combustion ; un brûleur principal (14) ; et un brûleur pilote (15), le brûleur principal (14) entourant le brûleur pilote (15) radialement à l'extérieur de celui-ci, le brûleur principal (14) étant conçu sous la forme d'un brûleur à jet avec des tubes à jet (16) s'étendant en ligne droite et étant conçu pour introduire, dans un écoulement droit, un mélange de combustible et d'air de combustion dans une partie radialement externe de la zone de combustion (13), le brûleur pilote (15) étant conçu soit en tant que brûleur à tourbillon avec un nombre défini de tubes de tourbillonnement en forme de spirale ou hélicoïdaux (17) soit en tant que brûleur à cône ou brûleur conique et étant conçu pour introduire, dans un écoulement tourbillonnant, un mélange de combustible et d'air de combustion dans une partie radialement interne de la zone de combustion (13), le rapport FH/FP entre la surface (FH) du brûleur principal (14) à travers laquelle s'écoule l'air de combustion et la zone (FP) du brûleur pilote (15) à travers laquelle s'écoule l'air de combustion (FP) étant compris entre 3 et 6.
PCT/EP2024/078461 2023-11-17 2024-10-10 Brûleur à turbine à gaz Pending WO2025103669A1 (fr)

Applications Claiming Priority (2)

Application Number Priority Date Filing Date Title
DE102023132050.5 2023-11-17
DE102023132050.5A DE102023132050A1 (de) 2023-11-17 2023-11-17 Gasturbinenbrenner

Publications (1)

Publication Number Publication Date
WO2025103669A1 true WO2025103669A1 (fr) 2025-05-22

Family

ID=93119726

Family Applications (1)

Application Number Title Priority Date Filing Date
PCT/EP2024/078461 Pending WO2025103669A1 (fr) 2023-11-17 2024-10-10 Brûleur à turbine à gaz

Country Status (2)

Country Link
DE (1) DE102023132050A1 (fr)
WO (1) WO2025103669A1 (fr)

Citations (12)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
DE19510743A1 (de) 1995-02-20 1996-09-26 Abb Management Ag Brennkammer mit Zweistufenverbrennung
DE19737997A1 (de) 1997-08-30 1999-03-04 Asea Brown Boveri Plenum
US20050164138A1 (en) * 2002-08-12 2005-07-28 Thomas Ruck Premixed exit ring pilot burner
US20110179795A1 (en) * 2009-07-08 2011-07-28 General Electric Company Injector with integrated resonator
US20180066847A1 (en) * 2016-09-06 2018-03-08 General Electric Company Fuel nozzle assembly with resonator
DE102018125848A1 (de) 2018-10-18 2020-04-23 Man Energy Solutions Se Brennkammer einer Gasturbine, Gasturbine und Verfahren zum Betreiben derselben
EP3770502A1 (fr) * 2019-07-22 2021-01-27 Delavan, Inc. Collecteurs de carburant segmentés
DE102020116245A1 (de) 2020-06-19 2021-12-23 Man Energy Solutions Se Baugruppe einer Gasturbine mit Brennkammerluftbypass
EP3974725A1 (fr) * 2020-09-29 2022-03-30 General Electric Company Ensemble d'injection de carburant pour une turbomachine
WO2022214384A1 (fr) * 2021-04-06 2022-10-13 Siemens Energy Global GmbH & Co. KG Chambre de combustion pour turbine à gaz
DE102021123513A1 (de) 2021-09-10 2023-03-16 Man Energy Solutions Se Brenner und Verfahren zu dessen Herstellung
US20230135396A1 (en) * 2021-11-03 2023-05-04 Power Systems Mfg., Llc Multitube pilot injector having a split airflow for a gas turbine engine

Family Cites Families (5)

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Publication number Priority date Publication date Assignee Title
JP5438727B2 (ja) * 2011-07-27 2014-03-12 株式会社日立製作所 燃焼器、バーナ及びガスタービン
US9188342B2 (en) * 2012-03-21 2015-11-17 General Electric Company Systems and methods for dampening combustor dynamics in a micromixer
DE102013204309A1 (de) * 2013-03-13 2014-09-18 Siemens Aktiengesellschaft Strahlbrenner
WO2019020350A1 (fr) * 2017-07-27 2019-01-31 Siemens Aktiengesellschaft Brûleur de turbine à gaz à jets de flamme prémélangés
US11371706B2 (en) * 2017-12-18 2022-06-28 General Electric Company Premixed pilot nozzle for gas turbine combustor

Patent Citations (12)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
DE19510743A1 (de) 1995-02-20 1996-09-26 Abb Management Ag Brennkammer mit Zweistufenverbrennung
DE19737997A1 (de) 1997-08-30 1999-03-04 Asea Brown Boveri Plenum
US20050164138A1 (en) * 2002-08-12 2005-07-28 Thomas Ruck Premixed exit ring pilot burner
US20110179795A1 (en) * 2009-07-08 2011-07-28 General Electric Company Injector with integrated resonator
US20180066847A1 (en) * 2016-09-06 2018-03-08 General Electric Company Fuel nozzle assembly with resonator
DE102018125848A1 (de) 2018-10-18 2020-04-23 Man Energy Solutions Se Brennkammer einer Gasturbine, Gasturbine und Verfahren zum Betreiben derselben
EP3770502A1 (fr) * 2019-07-22 2021-01-27 Delavan, Inc. Collecteurs de carburant segmentés
DE102020116245A1 (de) 2020-06-19 2021-12-23 Man Energy Solutions Se Baugruppe einer Gasturbine mit Brennkammerluftbypass
EP3974725A1 (fr) * 2020-09-29 2022-03-30 General Electric Company Ensemble d'injection de carburant pour une turbomachine
WO2022214384A1 (fr) * 2021-04-06 2022-10-13 Siemens Energy Global GmbH & Co. KG Chambre de combustion pour turbine à gaz
DE102021123513A1 (de) 2021-09-10 2023-03-16 Man Energy Solutions Se Brenner und Verfahren zu dessen Herstellung
US20230135396A1 (en) * 2021-11-03 2023-05-04 Power Systems Mfg., Llc Multitube pilot injector having a split airflow for a gas turbine engine

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Publication number Publication date
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