US20220074595A1 - Turbomachine combustion chamber - Google Patents

Turbomachine combustion chamber Download PDF

Info

Publication number
US20220074595A1
US20220074595A1 US17/416,956 US201917416956A US2022074595A1 US 20220074595 A1 US20220074595 A1 US 20220074595A1 US 201917416956 A US201917416956 A US 201917416956A US 2022074595 A1 US2022074595 A1 US 2022074595A1
Authority
US
United States
Prior art keywords
connection
peripheral tube
turbine engine
wall
tube
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.)
Abandoned
Application number
US17/416,956
Other languages
English (en)
Inventor
Damien Fauvet
Marc Nguyen
Baptiste Guerin
Jean-Michel Guimbard
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.)
Turbotech
Original Assignee
Turbotech
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 Turbotech filed Critical Turbotech
Assigned to TURBOTECH reassignment TURBOTECH ASSIGNMENT OF ASSIGNORS INTEREST (SEE DOCUMENT FOR DETAILS). Assignors: FAUVET, Damien, GUERIN, BAPTISTE, GUIMBARD, JEAN-MICHEL, NGUYEN, Marc
Publication of US20220074595A1 publication Critical patent/US20220074595A1/en
Abandoned legal-status Critical Current

Links

Images

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
    • F23R3/42—Continuous combustion chambers using liquid or gaseous fuel characterised by the arrangement or form of the flame tubes or combustion chambers
    • F23R3/46—Combustion chambers comprising an annular arrangement of several essentially tubular flame tubes within a common annular casing or within individual casings
    • F23R3/48—Flame tube interconnectors, e.g. cross-over tubes
    • 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
    • F23R3/002—Wall structures
    • 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
    • F23R3/005—Combined with pressure or heat exchangers
    • 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
    • F23R3/42—Continuous combustion chambers using liquid or gaseous fuel characterised by the arrangement or form of the flame tubes or combustion chambers
    • F23R3/54—Reverse-flow combustion chambers
    • 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
    • F23R2900/00001—Arrangements using bellows, e.g. to adjust volumes or reduce thermal stresses

Definitions

  • the present disclosure relates to the field of regenerative-cycle turbines intended, in particular, for the on-board production of electrical or mechanical energy from fuels for aeronautical, terrestrial, and maritime vehicles, and lightweight mobile units.
  • a turbine is made up of three elements:
  • the exhaust gases emerge from the turbine at high temperatures (greater than 500° C.), while the temperature of the air emerging from the compressor is lower (typically between 200° C. and 400° C.), having a circulation through a heat exchanger inserted between the exhaust gases and the compressed air, which makes it possible to partially reheat the air before it enters the combustion chamber, making it possible to reduce the fuel consumption.
  • the present disclosure relates more particularly to the combustion chamber and the injection of fuel into the combustion chamber.
  • German patent DE1254911 proposes injection nozzles in the form of hooks, mounted in the body of the injection nozzle, which is fixed in the flame holder by the nozzle tip thereof, so as to be displaceable with respect to the walls of the combustion chamber.
  • the external part of the body of the injection nozzle in the form of a hook, is slidably mounted in a guide, optionally associated with the exterior wall of the combustion chamber, and a clamping device, directed in parallel with the mouthpiece of the nozzle is provided for keeping the front end of the mouthpiece against a stop fixed to a hub of the flame holder.
  • British patent GB2097112 describes a fuel burner for a gas turbine engine, comprising a fuel feed arm and a fuel injector, the fuel feed arm and the fuel injector being joined together, the fuel feed arm having at least one fuel passage, the fuel injector 65 comprising a body having a passage in communication with the at least one fuel passage in the fuel feed arm, the body having an air duct, the axis of the air duct being coaxial with the axis of the fuel injector, the fuel injector having one or more 70 fuel passages to inject fuel into the air duct, the fuel burner having locating means at the end adjacent to the fuel injector, the locating means being arranged to engage with corresponding locating means on an engine component.
  • the problem posed by the solutions of the prior art relates to the turbines having a combustion chamber that is insulated from the outside by a double wall, two walls defining an annular channel for the circulation of a flow of compressed air originating from the compressor, and the third wall being the exterior wall of the combustion chamber, allowing for the circulation of the same air flow, previously reheated upon crossing a heat exchanger.
  • the injector or injectors must cross the three walls in a sealed manner, or at least in a manner having controlled leakage. This results in hyperstatic mounting, which does not allow for absorption of the longitudinal thermal expansions of the injector, or the radial and longitudinal thermal expansions of the metal walls exposed to extremely different temperatures.
  • the injector passes through the walls of the combustion chamber via simple holes, references 38 , 48 and 52 .
  • This document proposes positioning the injector coaxially inside each of the coaxial holes 52 , 48 and 38 provided in the housing 50 .
  • This solution thus results in several disadvantages: the radial expansion of the injector is different from the surface expansion of the walls, resulting either in leaks between the periphery of the injector and the edge of the through-holes in the wall, or in clamping of the edges of the holes around the wall of the injector, which limits the radial displacement possibilities and may lead to deformations and to fatigue of the walls.
  • the present disclosure relates to a combustion chamber of a turbine engine, surrounded by two coaxial axisymmetric walls, extending one inside the other and delimiting therebetween an annular air-circulation space, and a second air-circulation space delimited by the axisymmetric wall of smaller diameter and the exterior wall of the combustion chamber, and at least one injector that crosses the walls via ports, wherein the injector comprises a peripheral tube that is connected to the walls by three connections, at least two connections being flexible sealed connections allowing for multidirectional clearance, for example, of the slide type and/or of the ball joint type, or of the bellows type.
  • bellows means a sealed casing that can be deformed at least axially and radially, and optionally in a torsional or tilting manner.
  • the present disclosure also relates to a turbine comprising a combustion chamber of this kind.
  • FIG. 1 is a cross section of a turbine engine according to the present disclosure.
  • FIGS. 2 to 8 are schematic views of different variants.
  • FIG. 1 is a perspective view of the turbine engine, comprising a heat exchanger ( 1 ), a compressor ( 2 ), a combustion chamber ( 3 ), and a turbine ( 4 ).
  • a conical deflector ( 11 ) which is coaxial with the heat exchanger ( 1 ), causes the hot gases, originating from the turbine ( 4 ), to circulate towards a discharge outlet ( 12 ) after having passed through the heat exchanger ( 1 ), passing through two cassettes ( 5 , 6 ) between the tubes.
  • the parts formed by the compressor ( 2 ), the combustion chamber ( 3 ) and the turbine ( 4 ) are known to a person skilled in the art, and are in accordance with the state of knowledge in the field of turbine engines.
  • the heat exchanger ( 1 ) is formed by a tube heat exchanger, comprising two coaxial annular cassettes ( 5 , 6 ).
  • the external cassette ( 5 ) is formed by an assembly of parallel tubes, made of a metal alloy that is resistant to high temperatures, for example, refractory stainless steel 347.
  • the external cassette ( 5 ) is formed of 2000 tubes having a length of 300 millimeters, an internal cross-section of 2.8 millimeters, and an external cross-section of 3 millimeters.
  • the tubes are held in a known manner by means of inserts for defining the passages of hot gases originating from the turbine.
  • the tubes form a sleeve having an external radius of 158 millimeters and an internal radius of 128 millimeters.
  • the internal cassette ( 6 ) is formed of 2000 tubes having a length of 300 millimeters, an internal cross-section of 2.8 millimeters, and an external cross-section of 3 millimeters.
  • the tubes form a sleeve having an external radius of 123 millimeters and an internal radius of 67 millimeters.
  • the two cassettes ( 5 , 6 ) are coaxial and are fitted into one another.
  • the two cassettes ( 5 , 6 ) are united, at the end opposite the compressor ( 2 ), by an annular closure structure ( 8 ).
  • Each of the cassettes ( 5 , 6 ) comprises, at each end, a front sealing plate that is pierced for the tubes to pass through, and ensures the constant center distance of the tubes.
  • the tubes are brazed or soldered in order to ensure sealing in the region of the connection thereof to the front plates.
  • the closure structure ( 8 ) is formed of two coaxial parts that are fitted together and have the general shape of a rum baba mold, which parts are made of refractory stainless steel 347 of a thickness of 2 millimeters.
  • the outer part ( 9 ) has an external cross section that corresponds to the external cross section of the external cassette ( 5 ), and an internal cross section that corresponds to the internal cross section of the internal cassette ( 6 ).
  • the inner part ( 10 ) has an external cross section that corresponds to the internal cross section of the external cassette ( 5 ), and an internal cross section that corresponds to the external cross section of the internal cassette ( 6 ).
  • Each of the parts ( 9 , 10 ) is rotationally symmetric according to the axis of the turbine engine, having a constant longitudinal cross section.
  • the closure structure ( 8 ) ensures the deflection of the gases, originating from the external cassette ( 5 ), towards the tubes that make up the internal cassette ( 6 ).
  • This solution ensures a double passage of the gases in the heat exchanger ( 1 ), which significantly increases the thermal efficiency thereof for a given bulk, and, in particular, length.
  • the combustion chamber ( 3 ) of the annular type has a double interior casing formed by a sheath ( 30 ) (“liner”) and an intermediate wall ( 31 ).
  • the liner ( 30 ) and the intermediate wall ( 31 ) define a tubular volume for circulation of the air flow originating from the heat exchanger ( 1 ).
  • An exterior wall ( 32 ) and the intermediate wall ( 31 ) define a tubular volume for circulation of the air flow originating from the compressor ( 2 ) and travelling towards the heat exchanger ( 1 ).
  • the tube ( 35 ) of the injector passes through the three walls ( 30 to 32 ) via three ports.
  • the walls ( 30 to 32 ) as well as the tube ( 35 ) of the injector are subjected to longitudinal and radial expansions. The fixing is ensured by a combination of connections, avoiding the hyperstatic situations.
  • connection between the tube ( 35 ) of the injector and the interior wall ( 30 ) is ensured by a sliding connection formed by a calibrated port defining, together with the outside surface of the tube ( 35 ), a calibrated annular clearance.
  • connection between the tube ( 35 ) of the injector and the intermediate wall ( 31 ) is ensured by a fixed connection.
  • the first variant is illustrated schematically by FIG. 2 .
  • the tube ( 35 ) of the injector passes through the three walls ( 30 to 32 ) having the respective connections:
  • the second variant is illustrated schematically by FIG. 3 .
  • the tube ( 35 ) of the injector passes through the three walls ( 30 to 32 ) having the respective connections:
  • the third variant is illustrated schematically by FIG. 4 .
  • the tube ( 35 ) of the injector passes through the three walls ( 30 to 32 ) having the respective connections:
  • the fourth variant is illustrated schematically by FIG. 5 .
  • the tube ( 35 ) of the injector passes through the three walls ( 30 to 32 ) having the respective connections:
  • the fifth variant is illustrated schematically by FIGS. 6 to 8 .
  • the tube ( 35 ) of the injector passes through the three walls ( 30 to 32 ) having the respective connections:
  • connection between the peripheral tube ( 35 ) and the intermediate wall ( 31 ) of the liner is formed by a connection ( 80 ) having several degrees of freedom for allowing axial displacement and tangential displacement of the tube, and a tolerance for a ball joint.
  • connection between the peripheral tube ( 35 ) and the exterior wall ( 32 ) of the liner is formed by a sealed rigid assembly.
  • the head ( 38 ) comprises a discal flange ( 38 ) that is engaged between the two parts of the gland ( 37 ), which ensures clamping and sealing of the discal flange ( 38 ).
  • the inside end ( 40 ) passes through the interior wall ( 30 ), via passage in a simple hole formed in the interior wall ( 30 ).
  • the hole is oblong in this case, in order to take into account the inclination of the axis of the nozzle ( 35 ) with respect to the radial axis.
  • connection between the nozzle ( 35 ) and the intermediate wall ( 31 ) is achieved by a part having a conical upper portion ( 41 ) that is flared towards the outside and is extended at the base thereof by a discal flange ( 42 ) that is movable in radial translation in a slit ( 42 ) formed in the head ( 44 ) of a tubular extension ( 43 ) soldered to the surface of the interior wall ( 30 ).
  • the discal flange ( 42 ) is flexible, which furthermore allows for a lightweight ball joint with respect to the tubular extension ( 43 ).

Landscapes

  • Engineering & Computer Science (AREA)
  • Chemical & Material Sciences (AREA)
  • Combustion & Propulsion (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
  • Fuel-Injection Apparatus (AREA)
  • Structures Of Non-Positive Displacement Pumps (AREA)
  • Joints Allowing Movement (AREA)
  • Combustion Methods Of Internal-Combustion Engines (AREA)
US17/416,956 2018-12-21 2019-12-17 Turbomachine combustion chamber Abandoned US20220074595A1 (en)

Applications Claiming Priority (3)

Application Number Priority Date Filing Date Title
FR1874016A FR3090747B1 (fr) 2018-12-21 2018-12-21 Chambre de combustion d'une turbomachine
FR1874016 2018-12-21
PCT/FR2019/053108 WO2020128292A1 (fr) 2018-12-21 2019-12-17 Chambre de combustion d'une turbomachine

Publications (1)

Publication Number Publication Date
US20220074595A1 true US20220074595A1 (en) 2022-03-10

Family

ID=67441198

Family Applications (1)

Application Number Title Priority Date Filing Date
US17/416,956 Abandoned US20220074595A1 (en) 2018-12-21 2019-12-17 Turbomachine combustion chamber

Country Status (6)

Country Link
US (1) US20220074595A1 (de)
EP (1) EP3899371B1 (de)
CN (1) CN113454390B (de)
CA (1) CA3124209A1 (de)
FR (1) FR3090747B1 (de)
WO (1) WO2020128292A1 (de)

Citations (1)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US4441323A (en) * 1981-04-16 1984-04-10 Rolls-Royce Limited Combustion equipment for a gas turbine engine including a fuel burner capable of accurate positioning and installation as a unit in a flame tube

Family Cites Families (20)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
GB859805A (en) * 1958-07-21 1961-01-25 Gen Motors Corp Improvements relating to gas turbine engine combustion systems
DE1254911B (de) * 1965-09-23 1967-11-23 Daimler Benz Ag Anordnung des Einspritzduesenkoerpers an bzw. in der Brennkammer von Gasturbinentriebwerken
GB2093584B (en) * 1981-02-21 1984-12-19 Rolls Royce Improvements in or relating to fuel burners and combustion equipment for use in gas turbine engines
US4903476A (en) * 1988-12-27 1990-02-27 General Electric Company Gas turbine igniter with ball-joint support
US5966926A (en) 1997-05-28 1999-10-19 Capstone Turbine Corporation Liquid fuel injector purge system
US6438940B1 (en) * 1999-12-21 2002-08-27 General Electric Company Methods and apparatus for providing uniform ignition in an augmenter
US6438936B1 (en) 2000-05-16 2002-08-27 Elliott Energy Systems, Inc. Recuperator for use with turbine/turbo-alternator
US6442929B1 (en) * 2001-06-04 2002-09-03 Power Systems Mfg., Llc Igniter assembly having spring biasing of a semi-hemispherical mount
US7024863B2 (en) * 2003-07-08 2006-04-11 Pratt & Whitney Canada Corp. Combustor attachment with rotational joint
GB2433984B (en) * 2006-01-04 2007-11-21 Rolls Royce Plc A combustor assembly
FR2921463B1 (fr) * 2007-09-26 2013-12-06 Snecma Chambre de combustion d'une turbomachine
EP2726787B1 (de) 2011-06-30 2019-10-30 General Electric Company Brennkammer und verfahren zur brennstoffversorgung der brennkammer
US9032735B2 (en) * 2012-04-26 2015-05-19 General Electric Company Combustor and a method for assembling the combustor
FR3000522B1 (fr) * 2012-12-27 2018-11-02 Safran Aircraft Engines Dispositif de liaison a double tube
US9803555B2 (en) 2014-04-23 2017-10-31 General Electric Company Fuel delivery system with moveably attached fuel tube
US9803863B2 (en) * 2015-05-13 2017-10-31 Solar Turbines Incorporated Controlled-leak combustor grommet
FR3038699B1 (fr) * 2015-07-08 2022-06-24 Snecma Chambre de combustion coudee d'une turbomachine
GB2543803B (en) * 2015-10-29 2019-10-30 Rolls Royce Plc A combustion chamber assembly
GB2548585B (en) * 2016-03-22 2020-05-27 Rolls Royce Plc A combustion chamber assembly
FR3059363B1 (fr) 2016-11-25 2019-04-05 Turbotech Turbomachine, notamment turbogenerateur et echangeur pour une telle turbomachine

Patent Citations (1)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US4441323A (en) * 1981-04-16 1984-04-10 Rolls-Royce Limited Combustion equipment for a gas turbine engine including a fuel burner capable of accurate positioning and installation as a unit in a flame tube

Also Published As

Publication number Publication date
CA3124209A1 (fr) 2020-06-25
FR3090747A1 (fr) 2020-06-26
WO2020128292A1 (fr) 2020-06-25
CN113454390B (zh) 2023-02-24
FR3090747B1 (fr) 2021-01-22
EP3899371C0 (de) 2024-02-07
EP3899371B1 (de) 2024-02-07
EP3899371A1 (de) 2021-10-27
CN113454390A (zh) 2021-09-28

Similar Documents

Publication Publication Date Title
EP3431878B1 (de) Kraftstoffdüsen
US8181440B2 (en) Arrangement of a semiconductor-type igniter plug in a gas turbine engine combustion chamber
US10197285B2 (en) Gas turbine engine wall assembly interface
US2958183A (en) Rocket combustion chamber
EP3220047B1 (de) Gasturbinenstromhülsenhalterung
EP3039347B1 (de) Wandanordnung für gasturbinenmotor mit stützschalenkonturregionen
US10088161B2 (en) Gas turbine engine wall assembly with circumferential rail stud architecture
RU2696158C2 (ru) Теплоизолированная топливная форсунка для газотурбинного двигателя
CN102607064A (zh) 燃烧室喷嘴及制造燃烧室喷嘴的方法
US8752389B2 (en) Fuel nozzle assembly for use with a gas turbine engine and method of assembling same
US10808937B2 (en) Gas turbine engine wall assembly with offset rail
JP2020514657A (ja) マイクロチャネル冷却を備えた燃料ノズル組立体
US20150176841A1 (en) Apparatus for mixing fuel in a gas turbine nozzle
US10174946B2 (en) Nozzle guide for a combustor of a gas turbine engine
US7287383B2 (en) Afterburner arrangement
US20220074595A1 (en) Turbomachine combustion chamber
CN214247529U (zh) 航空发动机核心机尾喷管以及航空发动机核心机
CN115949968A (zh) 燃烧器旋流器到伪圆顶附接以及与cmc圆顶接口
CN108870444A (zh) 燃气涡轮发动机的燃烧器组件
US10982856B2 (en) Fuel nozzle with sleeves for thermal protection
JPH01131821A (ja) ガスタービン燃焼器の支承構造
JP2009192195A (ja) ガスタービンエンジンの燃焼装置
EP4339513A2 (de) Brennkammeranordnung
US20240093869A1 (en) Fuel nozzle
CN116379474B (zh) 一种航空发动机燃油喷嘴热防护结构

Legal Events

Date Code Title Description
AS Assignment

Owner name: TURBOTECH, FRANCE

Free format text: ASSIGNMENT OF ASSIGNORS INTEREST;ASSIGNORS:FAUVET, DAMIEN;NGUYEN, MARC;GUERIN, BAPTISTE;AND OTHERS;REEL/FRAME:057549/0675

Effective date: 20210901

STPP Information on status: patent application and granting procedure in general

Free format text: DOCKETED NEW CASE - READY FOR EXAMINATION

STPP Information on status: patent application and granting procedure in general

Free format text: NON FINAL ACTION MAILED

STPP Information on status: patent application and granting procedure in general

Free format text: RESPONSE TO NON-FINAL OFFICE ACTION ENTERED AND FORWARDED TO EXAMINER

STPP Information on status: patent application and granting procedure in general

Free format text: FINAL REJECTION MAILED

STCB Information on status: application discontinuation

Free format text: ABANDONED -- FAILURE TO RESPOND TO AN OFFICE ACTION