EP2154431A2 - Machine thermique - Google Patents

Machine thermique Download PDF

Info

Publication number: EP2154431A2
Authority: EP; European Patent Office
Prior art keywords: cooling; shells; parting plane; kühlhemdsegmente; thermal machine
Prior art date: 2008-08-14
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.): Granted

Application number

EP09167590A

Other languages

German (de)

English (en)

Other versions

EP2154431A3 (fr

EP2154431B1 (fr

Inventor

Remigi Tschuor

Hartmut Hähnle

Uwe Rüdel

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.)

GE Vernova GmbH

Original Assignee

Alstom Technology AG

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.)

2008-08-14

Filing date

2009-08-11

Publication date

2010-02-17

2009-08-11 Application filed by Alstom Technology AG filed Critical Alstom Technology AG

2010-02-17 Publication of EP2154431A2 publication Critical patent/EP2154431A2/fr

2010-08-04 Publication of EP2154431A3 publication Critical patent/EP2154431A3/fr

2017-07-26 Application granted granted Critical

2017-07-26 Publication of EP2154431B1 publication Critical patent/EP2154431B1/fr

Status Active legal-status Critical Current

2029-08-11 Anticipated expiration legal-status Critical

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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
- F23R3/54—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
- 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
- F23R3/42—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
- F23R2900/00017—Assembling combustion chamber liners or subparts
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
- F23R—GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
- F23R2900/00—Special features of, or arrangements for continuous combustion chambers; Combustion processes therefor
- F23R2900/00018—Manufacturing combustion chamber liners or subparts
- 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
- F23R2900/03044—Impingement cooled combustion chamber walls or subassemblies

Definitions

the present invention relates to the field of combustion technology. It relates to a thermal machine according to the preamble of claim 1.
IGT Modern industrial gas turbines
IGT are usually designed with annular combustion chambers.
Most smaller IGTs are designed as so-called "Can Annular Combustors".
An IGT with annular combustion chamber of the combustion chamber is limited by the side walls and the inlet and outlet plane of the hot gas.
Such a gas turbine is in the Fig. 1 and 2 shown.
the in the Fig. 1 and 2 Gas turbine 10 shown in the cutout has a turbine housing 11, in which a about an axis 27 rotating rotor 12th is housed.
the compressor 17 compresses air which flows into a plenum 14.
annular combustion chamber 15 is arranged, which is closed on the input side by a front plate cooling air 20 cooled front panel 19 and the output side via a hot gas channel 25 with the input of the turbine 13 in conjunction.
burners 16 are arranged in a ring, which are designed for example as a double cone or EV burner and inject a fuel-air mixture into the combustion chamber 15.
the resulting during the combustion of the mixture hot air stream 26 passes through the hot gas channel 25 in the turbine 13 and is relaxed there under work.
the combustion chamber 15 with the hot gas channel 25 is outside with a distance surrounded by an outer and innerdehemd 21 and 31, which are fastened by means of fasteners 24 to the combustion chamber 15, 25 and between them and the combustion chamber 15, 25 each have a ring-shaped outer and inner Form cooling channel 22 and 32 respectively.
In the cooling channels 22, 32 flows in the opposite direction to the hot gas stream 26 cooling air on the walls of the combustion chamber 15, 25 along in a combustion chamber hood 18 and from there into the burner 16 and as a faceplate cooling air 20 directly into the combustion chamber 15th
the side walls of the combustion chamber 15, 25 are carried out either as shell elements or as solid shells (outer shell 23, inner shell 33).
a parting plane (29 in Fig. 2a ), which allows an upper half of the shell 23, 33 (upper half shell 33a in FIG Fig. 2a ) from the lower half (lower half-shell 33b in FIG Fig. 2a ), for example, to assemble or disassemble the gas turbine rotor 12.
the parting plane 29 has correspondingly two parting plane welding seams 30 (FIG. Fig.
the respective lower and upper half-shells 33a, 33b have to be cooled convectively.
the already mentioned cooling shirts (CoShirts) 21 and 31 are mounted on the half-shell cold side, which redirect ambient air and cause due to the combustion chamber pressure drop or burner pressure drop across the half-shells and thus cause a convective cooling.
the inner and outer shell 33 and 23 of a gas turbine such as the GT13E2 are thermally and mechanically stressed during operation.
the strength properties of the material of the shells 23, 33 are highly temperature dependent.
the shells 23, 33 are convectively cooled.
the shape and the high thermal load near the turbine inlet (hot gas channel 25) require, especially in this area a constant high heat transfer on the cooling air side. This is achieved in the outer shell 23 by impingement cooling.
space and flow conditions and a seal against a cross-flow for such an impingement cooling are not given. Therefore, a conventional convection cooling is resorted to, in which the intensity of the cooling is increased by reducing the channel height of the cooling channel 32.
the previously used configuration of the inner cooling sleeve 31 of 2 axial plates is on the one hand prone to distance tolerances and other irregularities, e.g. in the flow field before the cooling air inlet into the cooling channel, and on the other hand causes an undesirable reduction of the cooling air mass flow in the region of the smaller of the two axial plates.
At least one of the cooling shirts on the Side on which the cooling air enters the cooling channel, to improve the Einström discipline has an outwardly curved, rounded leading edge.
the at least one cooling shirt in the region of the leading edge is flared bell-shaped or trumpet-like.
An embodiment of the invention is characterized in that the innerdehemd on the side on which the cooling air exiting the cooling channel, to reduce the flow losses has an outwardly curved, rounded exit edge.
the cooling shirts are composed of individual, in the circumferential direction adjoiningdehemdsegmenten, wherein thedehemdsegmente are attached by means of distributed arranged fasteners to the associated shells.
a preferred development is characterized in that thedehemdsegmente overlap each other in pairs in the connection areas, and that in each case adehemdsegment a pair in the overlapping region is provided with overlapping elements for a positive connection between the overlappingdehemdsegmenten.
Another embodiment of the invention is characterized in that the fastening elements are arranged in thedehemdsegmenten in the axial direction one behind the other, and that in the axial line with the fasteners additional holes are provided in thedehemdsegmenten, through which to improve the cooling cooling air in the rays from the outside flows into the respective cooling channel.
a further embodiment of the invention is characterized in that the combustion chamber is divided in a parting plane into an upper half with upper half shells and a lower half with lower half shells, that the half shells in the parting plane connected to each other by Trennebenenschweissnähte in that the shells have a shape deviating from the rotational symmetry in the area of the parting plane weld seams, and in that the cooling shrouds in the parting plane are adapted to the deviating shape of the shells.
the entirety of thedehemdsegmente in firstdehemdsegmente, which adjoin the parting plane, and seconddehemdsegmente, which are outside the parting plane, divided, wherein the firstdehemdsegmente to adapt to the different shape of the shells have a raised side edge.
Fig. 3 is shown in a side view of the part of an inner shell with segmented cooling shirt according to an embodiment of the invention.
an annular cooling channel 32 is formed on the outside of the inner shell 33 by a concentric inner cooling jacket 31 arranged concentrically therewith, in which on the in Fig. 3 Cooling air flows in on the left side, flows to the right and leaves the cooling channel 32 on the right side (see flow arrows in Fig. 3 ).
the innerdehemd 31 is composed of individual, extending in the axial directiondehemdsegmenten 34, which overlap each other. In the overlapping area, overlapping elements 36 protruding at the edges on the cooling-shirt segments are welded (see in particular FIG Fig. 7 ), which provide a positive fit between the overlapping segments in the overlap area.
Thedehemdsegmente 34 are distributed by means of fasteners 24 which pass through mounting holes 40 in the segments ( Fig. 5 . 6 and 8th ), attached to the associated inner shell 33.
the fastening elements 24 are arranged one behind the other in the axial direction.
additional holes 35 are provided in the cooling jacket segments 34 in the wake region of the fastening elements 24, flows through the air from the cooling air inlet. Due to its locally high velocity with respect to the incoming cooling air mass flow, the air jet entering the cooling channel 32 leads to an increase in the heat transfer coefficient and thus to a reduction in the wall temperature of the inner shell 33.
the inner cooling jacket 31 is flared bell-shaped or trumpet-like in the region of the leading edge 37.
This rounded "bellmouth-shaped" leading edge 37 of the axially one-piecede Kunststoffbleches allows on the one hand to minimize the pressure loss at the cooling air inlet and on the other an (unintended) variation of the heat transfer coefficient by detachment of the cooling air atdekanaleintritt (leading edge 37), as to Example of sharp-edged entries arise to prevent.
the achieved by the improved inflow conditions decreases the Verwirbelungsproe lead to a reduction of the required cooling air mass flow and thus to a more efficient operation of the combustion chamber.
the flow direction of the cooling air is opposite to the hot gas flow direction.
the inner shell cooling shirt or inner cooling jacket 31 is further designed so that at its outlet side (trailing edge 38) a new transition radius is selected, which causes a much more favorable, ie lower, flow loss than the previous configuration.
the reduction in the flow loss at this point is compensated by a reduction in the cooling channel height, which in turn leads to an increase in the cooling air side heat transfer and thus leads to a reduction of the average material temperature of the inner shell 33.
the cooling jacket 31 recedes outwards in the region of the parting plane welding seams 30 and creates space for a corresponding bulge of the combustion chamber shell 33 in the region of the parting plane weld seam 39.

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)

EP09167590.0A 2008-08-14 2009-08-11 Machine thermique Active EP2154431B1 (fr)

Applications Claiming Priority (1)

Application Number	Priority Date	Filing Date	Title
CH01277/08A CH699309A1 (de)	2008-08-14	2008-08-14	Thermische maschine mit luftgekühlter, ringförmiger brennkammer.

Publications (3)

Publication Number	Publication Date
EP2154431A2 true EP2154431A2 (fr)	2010-02-17
EP2154431A3 EP2154431A3 (fr)	2010-08-04
EP2154431B1 EP2154431B1 (fr)	2017-07-26

Family

ID=40342516

Family Applications (1)

Application Number	Title	Priority Date	Filing Date
EP09167590.0A Active EP2154431B1 (fr)	2008-08-14	2009-08-11	Machine thermique

Country Status (4)

Country	Link
US (1)	US8434313B2 (fr)
EP (1)	EP2154431B1 (fr)
AU (1)	AU2009208110B2 (fr)
CH (1)	CH699309A1 (fr)

Cited By (6)

* Cited by examiner, † Cited by third party
Publication number	Priority date	Publication date	Assignee	Title
EP2589873A1 (fr) *	2011-11-04	2013-05-08	General Electric Company	Chambre de combustion à flux inversé avec une injection d'air de sillage
WO2014099091A3 (fr) *	2012-10-01	2014-08-28	Alstom Technology Ltd.	Mécanisme de rétention de chemise thermiquement libre
US9267687B2 (en)	2011-11-04	2016-02-23	General Electric Company	Combustion system having a venturi for reducing wakes in an airflow
US9322553B2 (en)	2013-05-08	2016-04-26	General Electric Company	Wake manipulating structure for a turbine system
US9435221B2 (en)	2013-08-09	2016-09-06	General Electric Company	Turbomachine airfoil positioning
US9739201B2 (en)	2013-05-08	2017-08-22	General Electric Company	Wake reducing structure for a turbine system and method of reducing wake

Families Citing this family (13)

* Cited by examiner, † Cited by third party
Publication number	Priority date	Publication date	Assignee	Title
MY161317A (en) *	2008-02-20	2017-04-14	General Electric Technology Gmbh	Gas turbine
US20140208771A1 (en) *	2012-12-28	2014-07-31	United Technologies Corporation	Gas turbine engine component cooling arrangement
EP2868898A1 (fr) *	2013-10-30	2015-05-06	Siemens Aktiengesellschaft	Mode de fonctionnement en charge partielle amélioré d'une turbine à gaz comprenant un canal d'écoulement à dérivation réglable
GB201501817D0 (en) *	2015-02-04	2015-03-18	Rolls Royce Plc	A combustion chamber and a combustion chamber segment
EP3236155B1 (fr) *	2016-04-22	2020-05-06	Rolls-Royce plc	Chambre de combustion à paroi segmentée
US10641174B2 (en)	2017-01-18	2020-05-05	General Electric Company	Rotor shaft cooling
US10801730B2 (en)	2017-04-12	2020-10-13	Raytheon Technologies Corporation	Combustor panel mounting systems and methods
US10697634B2 (en) *	2018-03-07	2020-06-30	General Electric Company	Inner cooling shroud for transition zone of annular combustor liner
US11377970B2 (en)	2018-11-02	2022-07-05	Chromalloy Gas Turbine Llc	System and method for providing compressed air to a gas turbine combustor
EP3874129A4 (fr) *	2018-11-02	2022-10-05	Chromalloy Gas Turbine LLC	Système et procédé pour fournir de l'air comprimé à une chambre de combustion de turbine à gaz
US11248797B2 (en)	2018-11-02	2022-02-15	Chromalloy Gas Turbine Llc	Axial stop configuration for a combustion liner
US11047575B2 (en) *	2019-04-15	2021-06-29	Raytheon Technologies Corporation	Combustor heat shield panel
US11215367B2 (en)	2019-10-03	2022-01-04	Raytheon Technologies Corporation	Mounting a ceramic component to a non-ceramic component in a gas turbine engine

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JPH0752014B2 (ja) *	1986-03-20	1995-06-05	株式会社日立製作所	ガスタ−ビン燃焼器
US4896510A (en) *	1987-02-06	1990-01-30	General Electric Company	Combustor liner cooling arrangement
EP0489193B1 (fr) *	1990-12-05	1997-07-23	Asea Brown Boveri Ag	Chambre de combustion pour turbine à gaz
DE4232442A1 (de) *	1992-09-28	1994-03-31	Asea Brown Boveri	Gasturbinenbrennkammer
DE4239856A1 (de) *	1992-11-27	1994-06-01	Asea Brown Boveri	Gasturbinenbrennkammer
DE4242721A1 (de) *	1992-12-17	1994-06-23	Asea Brown Boveri	Gasturbinenbrennkammer
DE19751299C2 (de) *	1997-11-19	1999-09-09	Siemens Ag	Brennkammer sowie Verfahren zur Dampfkühlung einer Brennkammer
DE59810347D1 (de) *	1998-09-10	2004-01-15	Alstom Switzerland Ltd	Schwingungsdämpfung in Brennkammern
EP1001224B1 (fr) *	1998-11-12	2006-03-22	Mitsubishi Heavy Industries, Ltd.	Chambre de combustion pour une turbine à gaz
JP2002195565A (ja) *	2000-12-26	2002-07-10	Mitsubishi Heavy Ind Ltd	ガスタービン燃焼器
US7082770B2 (en) *	2003-12-24	2006-08-01	Martling Vincent C	Flow sleeve for a low NOx combustor
GB2420614B (en) *	2004-11-30	2009-06-03	Alstom Technology Ltd	Tile and exo-skeleton tile structure
GB2434199B (en) *	2006-01-14	2011-01-05	Alstom Technology Ltd	Combustor liner with heat shield

2008
- 2008-08-14 CH CH01277/08A patent/CH699309A1/de not_active Application Discontinuation
2009
- 2009-08-11 AU AU2009208110A patent/AU2009208110B2/en not_active Ceased
- 2009-08-11 EP EP09167590.0A patent/EP2154431B1/fr active Active
- 2009-08-13 US US12/540,453 patent/US8434313B2/en active Active

Cited By (8)

* Cited by examiner, † Cited by third party
Publication number	Priority date	Publication date	Assignee	Title
EP2589873A1 (fr) *	2011-11-04	2013-05-08	General Electric Company	Chambre de combustion à flux inversé avec une injection d'air de sillage
US8899975B2 (en)	2011-11-04	2014-12-02	General Electric Company	Combustor having wake air injection
US9267687B2 (en)	2011-11-04	2016-02-23	General Electric Company	Combustion system having a venturi for reducing wakes in an airflow
WO2014099091A3 (fr) *	2012-10-01	2014-08-28	Alstom Technology Ltd.	Mécanisme de rétention de chemise thermiquement libre
CN104685296A (zh) *	2012-10-01	2015-06-03	阿尔斯通技术有限公司	热自由衬套固持机构
US9322553B2 (en)	2013-05-08	2016-04-26	General Electric Company	Wake manipulating structure for a turbine system
US9739201B2 (en)	2013-05-08	2017-08-22	General Electric Company	Wake reducing structure for a turbine system and method of reducing wake
US9435221B2 (en)	2013-08-09	2016-09-06	General Electric Company	Turbomachine airfoil positioning

Also Published As

Publication number	Publication date
US20100037621A1 (en)	2010-02-18
CH699309A1 (de)	2010-02-15
US8434313B2 (en)	2013-05-07
EP2154431A3 (fr)	2010-08-04
EP2154431B1 (fr)	2017-07-26
AU2009208110A1 (en)	2010-03-04
AU2009208110B2 (en)	2014-07-10

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