EP1634021B1 - Chambre de combustion annulaire de turbomachine - Google Patents

Chambre de combustion annulaire de turbomachine Download PDF

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Publication number
EP1634021B1
EP1634021B1 EP04767843.8A EP04767843A EP1634021B1 EP 1634021 B1 EP1634021 B1 EP 1634021B1 EP 04767843 A EP04767843 A EP 04767843A EP 1634021 B1 EP1634021 B1 EP 1634021B1
Authority
EP
European Patent Office
Prior art keywords
holes
chamber
combustion chamber
perforations
effectively
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.)
Expired - Lifetime
Application number
EP04767843.8A
Other languages
German (de)
English (en)
French (fr)
Other versions
EP1634021A1 (fr
Inventor
Yves Salan
Denis Sandelis
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.)
Safran Aircraft Engines SAS
Original Assignee
Safran Aircraft Engines SAS
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 Safran Aircraft Engines SAS filed Critical Safran Aircraft Engines SAS
Publication of EP1634021A1 publication Critical patent/EP1634021A1/fr
Application granted granted Critical
Publication of EP1634021B1 publication Critical patent/EP1634021B1/fr
Anticipated expiration legal-status Critical
Expired - Lifetime legal-status Critical Current

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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
    • F23R3/50—Combustion chambers comprising an annular flame tube within an annular casing
    • 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
    • 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
    • F23R3/02—Continuous combustion chambers using liquid or gaseous fuel characterised by the air-flow or gas-flow configuration
    • F23R3/04—Air inlet 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
    • F23R3/00—Continuous combustion chambers using liquid or gaseous fuel
    • F23R3/02—Continuous combustion chambers using liquid or gaseous fuel characterised by the air-flow or gas-flow configuration
    • F23R3/04—Air inlet arrangements
    • F23R3/10—Air inlet arrangements for primary air
    • 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
    • F23R2900/03041—Effusion cooled combustion chamber walls or domes
    • 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
    • F23R2900/03042—Film cooled combustion chamber walls or domes

Definitions

  • the present invention relates generally to the field of turbomachine annular combustion chambers, and more particularly to that of means for thermally protecting these combustion chambers.
  • annular turbomachine combustion chamber comprises an outer axial wall and an inner axial wall, these walls being arranged coaxially and interconnected via a chamber bottom.
  • the combustion chamber is provided with angularly spaced injection ports, each of which is intended to receive a fuel injector in order to allow the combustion reactions to take place. inside this combustion chamber. It is furthermore noted that these injectors can also make it possible to introduce at least part of the air intended for combustion, which occurs in a primary zone of the combustion chamber situated upstream of a secondary zone. said dilution zone.
  • deflectors are arranged on the chamber bottom, in order to protect it from thermal radiation.
  • Each deflector also called cup or heat shield, then has at least one injection port for receiving a fuel injector, and a plurality of perforations for passing cooling air inside the nozzle. combustion chamber.
  • the invention therefore aims to provide a turbomachine annular combustion chamber, at least partially overcoming the disadvantages mentioned above relating to the embodiments of the prior art.
  • the object of the invention is to present an annular combustion chamber of turbomachine, whose means used to cool the chamber bottom generate neither significant disturbance of the combustion reactions inside the combustion chamber nor thermal discontinuities at the junctions between the chamber bottom and the external axial walls and internal.
  • the subject of the invention is an annular turbomachine combustion chamber, comprising an outer axial wall, an inner axial wall and a chamber bottom connecting the axial walls, the chamber bottom having a plurality of orifices. injection ports and a plurality of perforations, the injection ports being intended to allow at least the injection of fuel inside the combustion chamber and the perforations being intended to allow the passage of a flow cooling air suitable for cooling the chamber bottom.
  • the bottom chamber is provided on the one hand with an outer portion on which the perforations are made so as to direct a portion of the cooling air flow towards the outer axial wall, and on the other hand, an inner portion on which the perforations are made to direct another portion of the cooling air flow to the inner axial wall, and the chamber is designed such that in axial half-section , taken in any manner between two directly consecutive injection orifices, the value of the acute angles formed between a substantially median line of the half-section situated between the external axial wall and the axial wall internal, and main directions, in this half-section, perforations of the outer portion, evolves decreasingly depending on the distance between the perforations and this substantially middle line, and the value of acute angles formed between the line substantially median and principal directions, in this half-section, perforations of the inner portion, evolves decreasingly depending on the distance between the perforations and this substantially median line.
  • the combustion chamber according to the invention is such that the perforations located near a junction between the outer portion and the inner portion of the chamber bottom, that is to say substantially opposite a central annular ring of the combustion chamber, are more inclined towards the axial walls than can be the perforations located near these same axial walls, that is to say substantially facing the annular rings d end of this same combustion chamber.
  • the perforations located near the junction between the outer portion and the inner portion of the chamber bottom can therefore be strongly inclined towards the axial walls, and therefore allow the cooling air from these perforations s' flow easily and directly along the inner surface of the chamber bottom, substantially radially to the outer and inner axial walls.
  • this strong possible inclination indicates that the air of The cooling is only very slightly directed towards the center of the primary zone of the combustion chamber, so that it does not cause a significant disturbance of the combustion reactions.
  • the perforations located near the axial walls may be inclined only slightly towards these axial walls, so that the cooling air from these perforations can easily and directly flow along the surfaces. inside these same axial walls. It is specified that at these levels of the chamber bottom where the cooling air can be ejected inside the combustion chamber in a substantially axial direction of the latter, that is to say substantially parallel to the walls axial, the primary zone is sufficiently distant that the introduced cooling air does not cause significant disturbance of combustion reactions.
  • the combustion chamber according to the invention is therefore perfectly adapted not to cause significant disturbance of the combustion reactions inside the primary zone, which is essential for the stability and ignition of the combustion chamber.
  • the specific design of this chamber simultaneously ensures a satisfactory thermal continuity at the junctions between the chamber bottom and the outer and inner axial walls.
  • the two acute angles formed between the main directions of these perforations and the substantially median line have different values
  • the two acute angles formed between the main directions of these perforations and the substantially median line have different values.
  • This particular configuration makes it possible to obtain a very gradual inclination of the perforations of the chamber bottom.
  • the chamber base is provided with primary sectors of perforations as well as secondary sectors of perforations, the primary sectors being situated substantially between two directly consecutive injection orifices, and the secondary sectors lying on either side of each injection orifice, in a substantially radial direction of the combustion chamber.
  • annular combustion chamber 1 of a turbomachine With reference jointly to figures 1 and 2 , there is shown an annular combustion chamber 1 of a turbomachine, according to a preferred embodiment of the present invention.
  • the combustion chamber 1 comprises an external axial wall 2, as well as an internal axial wall 4, these two walls 2 and 4 being disposed coaxially along a principal longitudinal axis 6 of the chamber 1, this axis 6 also corresponding to the axis longitudinal main of the turbomachine.
  • the axial walls 2 and 4 are connected to each other via a chamber bottom 8, the latter being assembled for example by welding to an upstream portion of each of the axial walls 2 and 4.
  • the chamber bottom 8 preferably takes the form of a substantially flat annular ring, of axis identical to the longitudinal main axis 6 of the chamber 1.
  • this chamber bottom 8 could also have any other suitable shapes, such as a frustoconical shape of the same axis, without departing from the scope of the invention.
  • Each of these injection orifices 10 is designed so as to be able to cooperate with a fuel injector 12, in order to allow the combustion reactions inside this combustion chamber 1.
  • these injectors 12 are also designed to so as to allow the introduction of at least a portion of the air for combustion, the latter occurring in a primary zone 14 located in an upstream portion of the combustion chamber 1.
  • the air for combustion can also be introduced inside the chamber 1 via primary orifices 16, located all around the external axial walls 2 and internal 4.
  • the primary orifices 16 are arranged upstream of a plurality of dilution orifices 18, the latter also being placed all around the external axial walls 2 and internal walls 4, and whose main function is to allow the supply of air to a dilution zone 20 located downstream of the primary zone 14.
  • a cooling air flow D serving mainly to cool the inner surface 21 of the chamber floor 8.
  • an additional cooling air flow rate (not shown) is generally allocated to cool all of these interior hot surfaces 22 and 24.
  • the chamber bottom 8 is of the multi-perforated type, namely that it has a plurality of perforations 26, preferably cylindrical of circular sections, and intended to allow the passage of the cooling air flow D inside the combustion chamber 1.
  • the bottom chamber 8 is divided into an outer portion 28 connected to the outer axial wall 2, and an inner portion 30 connected to the inner axial wall 4.
  • these annular portions 28 and 30 are usually formed in one piece, and their virtual separation may then consist of a center circle C located on the longitudinal main axis 6, and radius R corresponding to a mean radius between an outer radius and an inner radius from the bedroom floor 8.
  • the perforations 26 located on the outer portion 28 are then made to direct a portion D1 of the cooling air flow D towards the outer axial wall 2, in order to cool the entire this outer portion 28, and an upstream portion of the outer axial wall 2.
  • the perforations 26 on the inner portion 30 are made to direct another portion D2 of the cooling air flow D in the direction of the inner axial wall 4, in order to cool all of this inner portion 30, as well as an upstream portion of the inner axial wall 4.
  • the perforations 26 of the outer portion 28 are such that the value of the acute angles A formed between a line substantially median 32 of the half-section and principal directions 34 of the perforations 26 in this half-section, evolves decreasingly as a function of the distance between these perforations 26 and this substantially median line 32.
  • a substantially median line 32 of the half-section it is naturally to understand that it is the virtual line located approximately equal distance from the upstream portions of the external axial walls 2 and internal 4 considered in half-section, this line 32 can also be noted in that in addition to forming an axis of symmetry of the half-section shown, it virtually separates the outer portions 28 and inner 30 of the chamber bottom 8.
  • this substantially median line 32, passing through the circle C, is also substantially perpendicular to the chamber bottom 8, insofar as it itself is substantially perpendicular to the axial walls 2 and 4 .
  • each main direction 34 of the perforations 26 correspond respectively to their main axes, in the sense that these perforations 26 are all traversed diametrically by the section plane.
  • each main direction 34 can then be considered as being a line substantially parallel to the two line segments symbolizing the perforation 26 concerned.
  • the perforations 26 located near the substantially median line 32 can therefore be strongly inclined, for example so that the acute angle A reaches a value of about 60 °.
  • the cooling air coming from these perforations 26 can therefore flow easily and directly along the inner surface 21 of the outer portion 28 of the chamber bottom 8, substantially radially to the outer axial wall 2, without disturbing combustion reactions in the primary zone 14.
  • the perforations 26 located near the outer axial wall 2 may be inclined only slightly towards the wall 2, for example so that the acute angle A reaches a value of about 5 °.
  • the cooling air coming from these perforations 26 can then easily and directly flow along the hot inner surface 22 of the external axial wall 2, without stagnating at the junction between the chamber bottom 8 and this same wall. axial 2.
  • the perforations 26 of the inner portion 30 are such that the value of the acute angles B formed between the substantially median line 32 and principal directions 36 of the perforations 26 in this half-section, changes in a decreasing manner as a function of the distance between these perforations 26 and this substantially median line 32.
  • the value of the acute angles B formed between on the one hand the main directions 36 of the perforations 26 of the inner portion 30, and on the other hand the line substantially median 32 can progressively evolve from about 60 ° to about 5 °, approaching the inner axial wall 4.
  • the chamber bottom 8 is provided with primary sectors 38 of perforations 26, these primary sectors 38 being situated substantially between two directly consecutive injection orifices 10.
  • the perforations 26 of each primary sector 38 are arranged to define rows in the form of curved lines centered on the center of the injection port 10 near which these perforations 26 are located.
  • the chamber bottom 8 is also provided with secondary sectors 40 of perforations 26, these secondary sectors 40 being each between two consecutive primary sectors 38, on either side of an injection orifice 10 in one direction. substantially radial of the combustion chamber 1.
  • a secondary sector 40 is both above and below the injection port 10 concerned.
  • the perforations 26 of the outer portion 28 are such that the value of acute angles C formed between a substantially median line 42 of the half-section and principal directions 44 of the perforations 26 in this half-section, evolves in a decreasing manner as a function of the distance between these perforations 26 and this substantially median line 42.
  • the perforations 26 of the inner portion 28 are such that the value of the acute angles D formed between the substantially median line 42 of the half-section and directions main 46 perforations 26 in this half-section, evolves decreasingly depending on the distance between these perforations 26 and this substantially median line 42.
  • the perforations 26 of the secondary sectors 38 are preferably of greater dimensions than those of the perforations 26 of the primary sectors 40, because of their presence. in a lower number.

Landscapes

  • Engineering & Computer Science (AREA)
  • Chemical & Material Sciences (AREA)
  • Combustion & Propulsion (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Combustion Methods Of Internal-Combustion Engines (AREA)
  • Fuel-Injection Apparatus (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
  • Cylinder Crankcases Of Internal Combustion Engines (AREA)
EP04767843.8A 2003-06-18 2004-06-18 Chambre de combustion annulaire de turbomachine Expired - Lifetime EP1634021B1 (fr)

Applications Claiming Priority (2)

Application Number Priority Date Filing Date Title
FR0350232A FR2856467B1 (fr) 2003-06-18 2003-06-18 Chambre de combustion annulaire de turbomachine
PCT/FR2004/050281 WO2004113794A1 (fr) 2003-06-18 2004-06-18 Chambre de combustion annulaire de turbomachine

Publications (2)

Publication Number Publication Date
EP1634021A1 EP1634021A1 (fr) 2006-03-15
EP1634021B1 true EP1634021B1 (fr) 2018-08-29

Family

ID=33484726

Family Applications (1)

Application Number Title Priority Date Filing Date
EP04767843.8A Expired - Lifetime EP1634021B1 (fr) 2003-06-18 2004-06-18 Chambre de combustion annulaire de turbomachine

Country Status (8)

Country Link
US (1) US7328582B2 (ru)
EP (1) EP1634021B1 (ru)
JP (1) JP2006527834A (ru)
KR (1) KR20060029203A (ru)
CN (1) CN1701203A (ru)
FR (1) FR2856467B1 (ru)
RU (1) RU2351849C2 (ru)
WO (1) WO2004113794A1 (ru)

Families Citing this family (17)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
FR2881813B1 (fr) * 2005-02-09 2011-04-08 Snecma Moteurs Carenage de chambre de combustion de turbomachine
US7540152B2 (en) * 2006-02-27 2009-06-02 Mitsubishi Heavy Industries, Ltd. Combustor
US7654091B2 (en) * 2006-08-30 2010-02-02 General Electric Company Method and apparatus for cooling gas turbine engine combustors
US8438853B2 (en) * 2008-01-29 2013-05-14 Alstom Technology Ltd. Combustor end cap assembly
US8763399B2 (en) * 2009-04-03 2014-07-01 Hitachi, Ltd. Combustor having modified spacing of air blowholes in an air blowhole plate
FR2948988B1 (fr) 2009-08-04 2011-12-09 Snecma Chambre de combustion de turbomachine comprenant des orifices d'entree d'air ameliores
FR2958013B1 (fr) * 2010-03-26 2014-06-20 Snecma Chambre de combustion de turbomachine a compresseur centrifuge sans deflecteur
FR2964725B1 (fr) * 2010-09-14 2012-10-12 Snecma Carenage aerodynamique pour fond de chambre de combustion
FR2980554B1 (fr) * 2011-09-27 2013-09-27 Snecma Chambre annulaire de combustion d'une turbomachine
US9377198B2 (en) * 2012-01-31 2016-06-28 United Technologies Corporation Heat shield for a combustor
FR3011317B1 (fr) * 2013-10-01 2018-02-23 Safran Aircraft Engines Chambre de combustion pour turbomachine a admission d'air homogene au travers de systemes d'injection
US10267521B2 (en) 2015-04-13 2019-04-23 Pratt & Whitney Canada Corp. Combustor heat shield
FR3042023B1 (fr) * 2015-10-06 2020-06-05 Safran Helicopter Engines Chambre de combustion annulaire pour turbomachine
US10808929B2 (en) * 2016-07-27 2020-10-20 Honda Motor Co., Ltd. Structure for cooling gas turbine engine
FR3070751B1 (fr) * 2017-09-01 2022-05-27 Safran Aircraft Engines Chambre de combustion comportant une repartition amelioree de trous de refroidissement
US11313560B2 (en) 2018-07-18 2022-04-26 General Electric Company Combustor assembly for a heat engine
US20240200778A1 (en) * 2022-12-20 2024-06-20 General Electric Company Gas turbine engine combustor with a set of dilution passages

Family Cites Families (9)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US5307637A (en) * 1992-07-09 1994-05-03 General Electric Company Angled multi-hole film cooled single wall combustor dome plate
DE19502328A1 (de) * 1995-01-26 1996-08-01 Bmw Rolls Royce Gmbh Hitzeschild für eine Gasturbinen-Brennkammer
FR2733582B1 (fr) * 1995-04-26 1997-06-06 Snecma Chambre de combustion comportant une multiperforation d'inclinaison axiale et tangentielle variable
FR2751731B1 (fr) * 1996-07-25 1998-09-04 Snecma Ensemble bol-deflecteur pour chambre de combustion de turbomachine
US6155056A (en) * 1998-06-04 2000-12-05 Pratt & Whitney Canada Corp. Cooling louver for annular gas turbine engine combustion chamber
US6145319A (en) * 1998-07-16 2000-11-14 General Electric Company Transitional multihole combustion liner
US6546733B2 (en) * 2001-06-28 2003-04-15 General Electric Company Methods and systems for cooling gas turbine engine combustors
DE10158548A1 (de) * 2001-11-29 2003-06-12 Rolls Royce Deutschland Brennkammerschindel für eine Gasturbine mit mehreren Kühllöchern mit unterschiedlicher Winkelausrichtung
US6751961B2 (en) * 2002-05-14 2004-06-22 United Technologies Corporation Bulkhead panel for use in a combustion chamber of a gas turbine engine

Also Published As

Publication number Publication date
RU2351849C2 (ru) 2009-04-10
FR2856467B1 (fr) 2005-09-02
FR2856467A1 (fr) 2004-12-24
WO2004113794A1 (fr) 2004-12-29
US7328582B2 (en) 2008-02-12
CN1701203A (zh) 2005-11-23
EP1634021A1 (fr) 2006-03-15
KR20060029203A (ko) 2006-04-05
JP2006527834A (ja) 2006-12-07
US20070056289A1 (en) 2007-03-15
RU2005107793A (ru) 2005-11-20

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