EP1818612B1 - Chambre de combustion annulaire d'une turbomachine - Google Patents

Chambre de combustion annulaire d'une turbomachine Download PDF

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Publication number
EP1818612B1
EP1818612B1 EP07102014A EP07102014A EP1818612B1 EP 1818612 B1 EP1818612 B1 EP 1818612B1 EP 07102014 A EP07102014 A EP 07102014A EP 07102014 A EP07102014 A EP 07102014A EP 1818612 B1 EP1818612 B1 EP 1818612B1
Authority
EP
European Patent Office
Prior art keywords
combustion chamber
chamber
sectors
sector
turbomachine
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.)
Active
Application number
EP07102014A
Other languages
German (de)
English (en)
French (fr)
Other versions
EP1818612A1 (fr
Inventor
Mario De Sousa
Didier Hernandez
Thomas Noel
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
SNECMA SAS
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Publication date
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Publication of EP1818612A1 publication Critical patent/EP1818612A1/fr
Application granted granted Critical
Publication of EP1818612B1 publication Critical patent/EP1818612B1/fr
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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
    • F23R3/002—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
    • F23R3/007—Continuous combustion chambers using liquid or gaseous fuel constructed mainly of ceramic 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
    • 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
    • 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

Definitions

  • the invention relates to an annular combustion chamber of a turbomachine, of the type comprising an inner wall, an outer wall, a chamber bottom disposed between said walls in the upstream region of said chamber, and two fastening flanges arranged downstream. of the chamber bottom and for attaching respectively said walls to other parts of the turbomachine, usually inner and outer casings surrounding the combustion chamber.
  • said inner and outer walls of the chamber were made of metal or metal alloy and it was necessary to cool these walls so that they can withstand the temperatures reached during operation of the turbomachine.
  • Ceramic materials are more resistant to high temperatures and have a lower density than commonly used metals.
  • the gains made in cooling air and in mass make it possible to improve the efficiency of the turbomachine.
  • the ceramic materials used are preferably ceramic matrix composite materials chosen for their good mechanical properties.
  • the state of the art leads to making these pieces of metal or metal alloy, rather than ceramic material, in order to be able to use the known fastening methods and tested to date, to fix the attachment flanges to the metal casings of the combustion chamber and the injection systems at the bottom of the chamber. It may be, for example, fasteners by welding or bolting.
  • the ceramics used to make the walls often have a coefficient of expansion about three times less than that of the metal materials used to make the chamber bottom and said flanges.
  • Such a gap generates stresses in the assembled parts during their assembly, as well as during the temperature rise thereof during operation. These stresses can be the cause of cracking in the fastening flanges or in the walls, if these flanges are not sufficiently flexible, the ceramic materials being by nature quite fragile.
  • a solution described in the document FR 2,855,249 consists in providing a plurality of flexible fastening tabs connecting the chamber bottom audites walls, these tabs being able to deform elastically depending on the expansion gap between these parts.
  • FR 2,825,781 and FR 2,825,784 consisting of connecting the walls to the casings of the combustion chamber by a plurality of flexible fasteners, elastically deformable, replacing the annular attachment flanges.
  • the inner and outer walls of the combustion chamber are made in one piece of generally frustoconical shape.
  • FR 2,855,249 there are spaces between the fastening tabs at the bottom of the chamber in which the fresh air rushes, which can degrade the efficiency of the combustion chamber by promoting the formation of pollutant emissions, such as for example, unburnt and / or carbon monoxide.
  • the annular combustion chamber of the aforementioned type is such that each wall of the chamber is divided into several adjacent sectors, each sector being attached to the bottom of the chamber and to one of the snap flanges.
  • the walls Thanks to the partitioning of the walls, they can deform depending on the expansion of the chamber bottom and the attachment flanges (this expansion being greater than that of the walls). For example, during a rise in temperature, during which the chamber bottom and / or The attachment flanges expand (ie their diameters increase), the adjacent sectors of the walls deviate circumferentially so that the diameters of these walls increase. This avoids the creation of thermomechanical stresses in these parts.
  • the sectors of the walls are provided with side edges and the side edges of two adjacent sectors overlap, so as to limit the passage of fresh air between the sectors, from the outside to the inside of the combustion chamber. Indeed, such an air passage, if it is not controlled, causes the introduction of too much air into the chamber, which causes the formation of pollutant emissions such as, for example, unburned and carbon monoxide, and thus reduces the efficiency of the chamber. On the other hand, this passage of air, if it is controlled, can serve for the cooling of the walls, as explained hereafter.
  • a known solution is to make a multitude of small perforations in said walls, through which calibrated volumes of fresh air pass. We usually talk about multiperforations. This solution nevertheless has the disadvantage of significantly increasing the cost price of said walls and cause a significant decrease in the characteristics of behavior and mechanical damage.
  • This objective is achieved thanks to the fact that there is a radial clearance (ie in a direction perpendicular to the axis at the axis of rotation of the turbomachine) between two overlapping adjacent sectors, this clearance allowing the passage of air cool from the outside to the inside of said chamber to cool the inner face of at least one of the sectors.
  • the fresh air coming from the outside of the chamber does not penetrate radially inside the chamber since the sectors overlap: it penetrates circumferentially while skirting, at least in part, the inner face of the chambers. inner and outer walls, so as to cool them.
  • it controls the amount of cooling air entering the chamber.
  • the invention aims to increase the cooling efficiency of the inner faces of the inner and outer walls.
  • the lateral edges of the sectors are inclined circumferentially relative to the main axis of the combustion chamber, this main axis corresponding to the axis of rotation of the turbomachine.
  • the circumferential direction at a point on the surface of a wall of the chamber is defined as the direction of the tangent to the wall, at this point, in a plane perpendicular to the axis of rotation. of the turbomachine.
  • a lateral sectoral edge is inclined circumferentially with respect to the axis of rotation of the turbomachine, when this edge is inclined with respect to a generator of the wall concerned.
  • the wall sectors are not attached to the chamber bottom and to the attachment flanges by means of flexible fasteners but, on the contrary, they are rigidly attached to these elements, for example by bolting.
  • the structure has a better dynamic behavior in operation than a structure with flexible fastening tabs.
  • the chamber bottom 30 and the attachment flanges 27 and 29 are made of metal alloy, while the walls 26 and 28 of the chamber 24 are made of ceramic matrix composite material.
  • the walls 26 and 28 are respectively divided into several adjacent sectors 126 and 128.
  • Each sector 126 (128) is attached to the chamber bottom 30, on the one hand, and to one of the attachment flanges 27 (29), on the other hand. At least one of these sectors may have multiperforations.
  • each wall sector 126 (128) is attached to the bottom of the chamber. 30 or one of the attachment flanges 27 (29) at two points of attachment, at least.
  • each sector 126 (128) is prevented from pivoting relative to the chamber bottom and / or said flange, thereby preventing the angular displacement of the chamber floor 30.
  • each sector 126 (128) is attached at the bottom of the chamber 30 and at an attachment flange 27 (29) at two attachment points 36 and 36 '.
  • At least one of these two attachment points 36 ' is made by bolting, by passing a bolt 52, through at least one oblong hole 50.
  • This oblong hole 50 can be formed in the flap 32 (34). of the chamber bottom 30, in the sector 126 (128) or in these two pieces at a time.
  • This oblong hole 50 is circumferentially oriented and the bolt 52 can therefore move circumferentially, inside the hole 50 as indicated by the double arrow B on the figure 4 .
  • Each sector 128 (126) includes a lip 60 extending along one of its side edges 128a (126a), preferably substantially the entire length thereof.
  • the other side edge of the sector is devoid of lip and will be hereinafter referred to as simple edge 128b (126b).
  • the lip 60 projects from one of the inner or outer faces of the sector 128 (126) so as to cover the single edge 128b (126b) of the adjacent sector.
  • the lip 60 is offset radially inwards or outwards with respect to the sector 128.
  • the lip 60 is projecting (outwardly) relative to the outer face of the sector 128.
  • it could be projecting (inwards) relative to the inner face of the sector.
  • the outer and inner faces 126, 128 being turned respectively outwardly and inwardly of the combustion chamber 24.
  • the lip 60 may be made directly during the manufacture of sector 128 (126), or during a subsequent machining step in its manufacture.
  • the lip 60 may also consist of an added band, for example by gluing, on the lateral edge 128a (126a) of the sector.
  • the fresh air circulates outside the chamber 24 according to the arrows F represented on the figure 1 , that is to say in a direction more axial than radial.
  • the clearance J and the slot 66 form a passage which deviates little enough the flow of fresh air F 'entering the combustion chamber 24.
  • this air flow F' remains sufficiently inclined relative to the radial direction as represented on the figures 1 and 4 to, on the one hand, disturb as little as possible the combustion inside the chamber 24 and, on the other hand, create a protective film of fresh air along the inner face of the wall segments 126, 128, which makes it possible to limit the heating of these segments.
  • the side edges 126a, 126b, 128a, 128b of the sectors 126, 128 are inclined circumferentially with respect to the main axis of the combustion chamber. As indicated above, this circumferential inclination corresponds to an angle inclination y of the lateral edges relative to the generatrices G of the walls 126, 128.
  • Tilting the side edges 126a, 126b, 128a, 128b and therefore the slots 66 for fresh air inlet allows to distribute the flow of fresh air F 'entering the chamber 24 in a cooling zone Z plus important that if said lateral edges were oriented along a generatrix G.
  • This cooling zone Z is hatched on the figure 2 . The more the lateral edges 126, 128 are inclined, the more the zone Z is extended, and the better is the cooling of the wall sectors 126, 128.

Landscapes

  • Engineering & Computer Science (AREA)
  • Chemical & Material Sciences (AREA)
  • Combustion & Propulsion (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Ceramic Engineering (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
  • Structures Of Non-Positive Displacement Pumps (AREA)
EP07102014A 2006-02-10 2007-02-09 Chambre de combustion annulaire d'une turbomachine Active EP1818612B1 (fr)

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
FR0650475A FR2897418B1 (fr) 2006-02-10 2006-02-10 Chambre de combustion annulaire d'une turbomachine

Publications (2)

Publication Number Publication Date
EP1818612A1 EP1818612A1 (fr) 2007-08-15
EP1818612B1 true EP1818612B1 (fr) 2010-09-29

Family

ID=37102414

Family Applications (1)

Application Number Title Priority Date Filing Date
EP07102014A Active EP1818612B1 (fr) 2006-02-10 2007-02-09 Chambre de combustion annulaire d'une turbomachine

Country Status (7)

Country Link
US (1) US7788928B2 (ru)
EP (1) EP1818612B1 (ru)
JP (1) JP2007212129A (ru)
CA (1) CA2577520C (ru)
DE (1) DE602007009436D1 (ru)
FR (1) FR2897418B1 (ru)
RU (1) RU2429418C2 (ru)

Families Citing this family (24)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
FR2920525B1 (fr) * 2007-08-31 2014-06-13 Snecma Separateur pour alimentation de l'air de refroidissement d'une turbine
US8266914B2 (en) * 2008-10-22 2012-09-18 Pratt & Whitney Canada Corp. Heat shield sealing for gas turbine engine combustor
US10240790B2 (en) 2013-11-04 2019-03-26 United Technologies Corporation Turbine engine combustor heat shield with multi-height rails
EP3066390B1 (en) 2013-11-04 2020-10-21 United Technologies Corporation Gas turbine engine wall assembly with offset rail
EP3084310A4 (en) 2013-12-19 2017-01-04 United Technologies Corporation Gas turbine engine wall assembly with circumferential rail stud architecture
US10234140B2 (en) 2013-12-31 2019-03-19 United Technologies Corporation Gas turbine engine wall assembly with enhanced flow architecture
DE102014204482A1 (de) * 2014-03-11 2015-09-17 Rolls-Royce Deutschland Ltd & Co Kg Brennkammer einer Gasturbine
US9752447B2 (en) * 2014-04-04 2017-09-05 United Technologies Corporation Angled rail holes
US10648669B2 (en) 2015-08-21 2020-05-12 Rolls-Royce Corporation Case and liner arrangement for a combustor
US20170059159A1 (en) 2015-08-25 2017-03-02 Rolls-Royce Corporation Cmc combustor shell with integral chutes
FR3045137B1 (fr) * 2015-12-11 2018-05-04 Safran Aircraft Engines Chambre de combustion de turbomachine
US10473332B2 (en) 2016-02-25 2019-11-12 General Electric Company Combustor assembly
US10393380B2 (en) * 2016-07-12 2019-08-27 Rolls-Royce North American Technologies Inc. Combustor cassette liner mounting assembly
GB201613299D0 (en) 2016-08-02 2016-09-14 Rolls Royce Plc A method of assembling an annular combustion chamber assembly
US10670269B2 (en) 2016-10-26 2020-06-02 Raytheon Technologies Corporation Cast combustor liner panel gating feature for a gas turbine engine combustor
US10669939B2 (en) 2016-10-26 2020-06-02 Raytheon Technologies Corporation Combustor seal for a gas turbine engine combustor
US10830448B2 (en) 2016-10-26 2020-11-10 Raytheon Technologies Corporation Combustor liner panel with a multiple of heat transfer augmentors for a gas turbine engine combustor
US10823410B2 (en) 2016-10-26 2020-11-03 Raytheon Technologies Corporation Cast combustor liner panel radius for gas turbine engine combustor
US10935243B2 (en) 2016-11-30 2021-03-02 Raytheon Technologies Corporation Regulated combustor liner panel for a gas turbine engine combustor
CN106812556B (zh) * 2017-03-16 2018-05-25 中国科学院工程热物理研究所 一种燃气轮机热端冷却结构及具有其的燃气轮机
US10598380B2 (en) * 2017-09-21 2020-03-24 General Electric Company Canted combustor for gas turbine engine
US11073285B2 (en) * 2019-06-21 2021-07-27 Raytheon Technologies Corporation Combustor panel configuration with skewed side walls
CN112902230A (zh) * 2021-03-11 2021-06-04 西北工业大学 一种倾斜式入口双头部的双级旋流器燃烧室
US11747019B1 (en) * 2022-09-02 2023-09-05 General Electric Company Aerodynamic combustor liner design for emissions reductions

Family Cites Families (9)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US2544538A (en) * 1948-12-01 1951-03-06 Wright Aeronautical Corp Liner for hot gas chambers
US3854503A (en) * 1971-08-05 1974-12-17 Lucas Industries Ltd Flame tubes
US4543781A (en) * 1981-06-17 1985-10-01 Rice Ivan G Annular combustor for gas turbine
US5025622A (en) * 1988-08-26 1991-06-25 Sol-3- Resources, Inc. Annular vortex combustor
US5636508A (en) * 1994-10-07 1997-06-10 Solar Turbines Incorporated Wedge edge ceramic combustor tile
FR2825779B1 (fr) * 2001-06-06 2003-08-29 Snecma Moteurs Chambre de combustion munie d'un systeme de fixation de fond de chambre
FR2825781B1 (fr) 2001-06-06 2004-02-06 Snecma Moteurs Montage elastique de chambre ce combustion cmc de turbomachine dans un carter metallique
FR2825784B1 (fr) 2001-06-06 2003-08-29 Snecma Moteurs Accrochage de chambre de combustion cmc de turbomachine utilisant les trous de dilution
FR2855249B1 (fr) 2003-05-20 2005-07-08 Snecma Moteurs Chambre de combustion ayant une liaison souple entre un fond de chambre et une paroi de chambre

Also Published As

Publication number Publication date
CA2577520A1 (fr) 2007-08-10
CA2577520C (fr) 2015-03-31
US7788928B2 (en) 2010-09-07
RU2429418C2 (ru) 2011-09-20
US20070186559A1 (en) 2007-08-16
RU2007105075A (ru) 2008-08-20
DE602007009436D1 (de) 2010-11-11
JP2007212129A (ja) 2007-08-23
FR2897418A1 (fr) 2007-08-17
EP1818612A1 (fr) 2007-08-15
FR2897418B1 (fr) 2013-03-01

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