EP1508680A1 - Diffuseur situé entre le compresseur et la chambre de combustion d'une turbine à gaz - Google Patents

Diffuseur situé entre le compresseur et la chambre de combustion d'une turbine à gaz Download PDF

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Publication number
EP1508680A1
EP1508680A1 EP03018565A EP03018565A EP1508680A1 EP 1508680 A1 EP1508680 A1 EP 1508680A1 EP 03018565 A EP03018565 A EP 03018565A EP 03018565 A EP03018565 A EP 03018565A EP 1508680 A1 EP1508680 A1 EP 1508680A1
Authority
EP
European Patent Office
Prior art keywords
combustion chamber
turbine
diffuser
longitudinal axis
wall
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.)
Withdrawn
Application number
EP03018565A
Other languages
German (de)
English (en)
Inventor
Christian Dr. Cornelius
Reinhard Dr. Mönig
Peter Tiemann
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.)
Siemens AG
Original Assignee
Siemens 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.)
Filing date
Publication date
Application filed by Siemens AG filed Critical Siemens AG
Priority to EP03018565A priority Critical patent/EP1508680A1/fr
Priority to PCT/EP2004/007946 priority patent/WO2005019621A1/fr
Priority to PL04741084T priority patent/PL1656497T3/pl
Priority to EP04741084A priority patent/EP1656497B1/fr
Priority to ES04741084T priority patent/ES2275226T3/es
Priority to DE502004001924T priority patent/DE502004001924D1/de
Priority to US10/568,736 priority patent/US8082738B2/en
Priority to CNB2004800235393A priority patent/CN100390387C/zh
Publication of EP1508680A1 publication Critical patent/EP1508680A1/fr
Withdrawn legal-status Critical Current

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Classifications

    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D25/00—Component parts, details, or accessories, not provided for in, or of interest apart from, other groups
    • F01D25/08—Cooling; Heating; Heat-insulation
    • F01D25/12—Cooling
    • 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

Definitions

  • the invention relates to a gas turbine with an annular combustion chamber and one of these upstream, essentially parallel to a turbine longitudinal axis vorströmbaren and of this less than the annular combustion chamber spaced diffuser, in which a compressed gas at a branch point can be divided into sub-streams.
  • Gas turbines are used in many areas to drive generators or used by work machines. It is the Energy content of a fuel for generating a rotational movement used a turbine shaft.
  • the fuel will burned in a combustion chamber, being used by an air compressor compressed air is supplied. That in the combustion chamber produced by the combustion of the fuel, under high Pressure and high temperature working medium is doing via a turbine downstream of the combustion unit led, where it relaxes work.
  • a gas turbine which one upstream of a combustion chamber and into a diffuser having opening air compressor.
  • a partial stream of compacted Air can diverted in the diffuser from this and the Cooling of structural parts, such as turbine blades the gas turbine, are used.
  • the cooling air branch from the diffuser is only for a branch of a relatively low partial flow from the air compressor leaving airflow suitable.
  • the led by the diffuser Main air flow however, in the diffuser in the direction deflected towards the combustion chamber and this as combustion air fed. Cooling downstream of the diffuser, that is, with respect to the flow direction of the turbine flowing working medium, arranged downstream Components is therefore limited possible.
  • the invention is based on the object, one with a Indicate ring combustor equipped compact gas turbine, which a flow favorable leadership of the Compressor air for a particularly even and effective Coolability of thermally loaded components allows.
  • the Gas turbine an annular combustion chamber and one of these upstream annular diffuser, which at least partially between the turbine longitudinal axis and the Ring combustion chamber is arranged.
  • the diffuser which in the Essentially can be flowed parallel to the turbine longitudinal axis, is a compressed gas in several cooling gas streams divisible.
  • the diffuser has a Main deflection area, which at an acute angle of the turbine longitudinal axis pointing the way on the inner wall of the Ring combustion chamber is directed.
  • the main deflection area is in Direction of the gas flowing through the diffuser, in particular Air, a branch point downstream, at which the the Diffuser gas flowing through partial flows by means of a Flow dividing element is divisible.
  • Two the walls of the Diffuser opposite Ablenkflanken run in one acute angles towards each other and meet at the Branching point. There they enclose an angle bisector, the turbine longitudinal axis at an acute pitch angle greater than 20 ° cuts.
  • the main deflection area is behind in the axial direction the compressor and before the annular combustion chamber, whereas the Flow dividing element between ring combustion chamber and turbine longitudinal axis is arranged.
  • This geometry allows for the gas turbine a compact and in particular a in Axial direction shortened design. Furthermore, the flow losses in the compressed coolant sub-streams reduced.
  • the flow direction is a particularly good cooling of radially spaced from the turbine longitudinal axis components, in particular the annular combustion chamber reached.
  • the two divided in the diffuser cooling gas partial flows in Connection also used for combustion.
  • the compressed gas which at this point the diffuser leaves, directed directly into a flow transfer space, which the fluidic connection to the Wandungskühlraum the annular combustion chamber manufactures.
  • the flow transfer space adjoins the outside of the combustion chamber wall, so that thereby an additional cooling of the Brenncrowandung is achieved.
  • the annular combustion chamber is preferably closed cooled educated.
  • the cooling medium is preferably Combustion air in countercurrent to the flue gas through a Wandungsraum the annular combustion chamber out.
  • the by the Brennschdung flowing combustion air is here preferably at least with a partial flow of the compressed air identical, which previously flowed through the diffuser.
  • the air flowing through the diffuser completely the wall of the annular combustion chamber as cooling air and further fed to the annular combustion chamber as combustion air.
  • the division of the air flow at the branch point of the Diffuser serves to several parts of the annular combustion chamber, for example, an inner shell and an outer shell, to provide evenly with cooling air.
  • annular combustion chamber at least in one Subarea essentially flat combustion chamber rear wall has, is below the wall angle of the annular combustion chamber understood the angle that the combustion chamber rear wall with the turbine longitudinal axis includes.
  • a special uniform all-round cooling of the combustion chamber wall is preferably achieved in that the pitch angle of the Flow dividing element of the wall angle of the Combustion chamber rear wall by no more than 20 °, in particular around not more than 15 °, deviates.
  • the advantage of the invention is in particular that in a gas turbine compressed air, as cooling and then serves as combustion air, low pressure loss of an air compressor through a compact diffuser Ring combustion chamber is supplied, wherein a flow dividing element at the outlet of the diffuser a uniform Cooling air is applied to the annular combustion chamber.
  • the gas turbine 1 has a compressor 2 for Combustion air, an annular combustion chamber 4 and a turbine. 6 for driving the compressor 2 and a not shown Generator or a working machine.
  • a turbine. 6 for driving the compressor 2 and a not shown Generator or a working machine.
  • the annular combustion chamber 4 is provided with a number of burners 10 for combustion of a liquid or gaseous fuel stocked. She is also on her combustion chamber wall 23 with a wall lining 24 provided.
  • the turbine 6 has a number of with the turbine shaft. 8 connected, rotatable blades 12.
  • the blades 12 are arranged in a ring on the turbine shaft 8 and thus form a number of blade rows.
  • the turbine 6 includes a number of stationary vanes 14, which is also coronal under the formation of Guide vane rows attached to an inner housing 16 of the turbine 6 are.
  • the blades 12 serve to drive the turbine shaft 8 by momentum transfer from the turbine. 6 flowing flue gas or working medium M.
  • the vanes 14, however, serve to guide the flow of the working medium M between each two in the flow direction of the working medium M seen consecutive blade rows or blade wreaths.
  • a successive pair out a ring of vanes 14 or a row of vanes and a ring of blades 12 or a blade row is also referred to as a turbine stage.
  • Each vane 14 has one also referred to as blade root 19 Platform 18 on which is to fix the respective Guide blade 14 is determined in the gas turbine 1. Every blade 12 is analogously via a platform as well 18 designated blade root 19 on the turbine shaft. 8 fastened, wherein the blade root 19 each one along a Blade axis extended profiled airfoil 20 wearing.
  • each guide ring 21 on the inner housing 16 of Turbine 6 is arranged between the spaced apart platforms 18 of the vanes 14 of two adjacent rows of vanes.
  • the outer surface of each guide ring 21 is also the hot, the turbine 6 flowing through Working medium M exposed and in the radial direction from the outer end 22 of the blade opposite it 12 spaced by a gap.
  • the between adjacent Guide blade rows arranged guide rings 21st serve in particular as cover elements that the inner wall 16 or other housing-mounted components before a thermal Overuse by the turbine 6 flowing through hot working medium M protects.
  • the combustion chamber wall 23 is compressed in the compressor 2 Cooling air as coolant K coolable. Between the combustion chamber wall 23 and the wall lining 24 flows cooling air K in one Wandungsraum or wall lining room 26 in countercurrent to Working medium M on the burner 10 to.
  • the cooling air K which also serves as combustion air is discharged from the compressor 2 through a diffuser 27 in the direction of the annular combustion chamber 4th directed. Through the diffuser 27, the cooling and Combustion air K defines a split one hand outer combustion chamber shell 28 and on the other hand an inner Combustion chamber 29 fed.
  • the diffuser 27 has a Hauptablenk Scheme 30, which is connected to the compressor 2 connects.
  • the compressed cooling air K flows parallel to Central axis or turbine longitudinal axis 9 from the compressor. 2 from and into the main deflection region 30 of the diffuser 27 a.
  • the seen in the axial direction between the compressor 2 and the annular combustion chamber 4 arranged Mannablenk Scheme 30 of the Diffuser 27 extends radially under cross-sectional expansion to the outside, i. away from the turbine longitudinal axis 9. hereby In the main deflection region 30, the flow velocity is reduced of the used as coolant K compressed Gas. If there is a flow separation on the inner wall and outer wall of the diffuser 27 comes, such occurs Replacement only at low flow rate and accordingly low pressure loss.
  • a flow dividing element 32 is disposed adjacent to the outer combustion chamber shell 29.
  • the arranged between the annular combustion chamber 4 and the turbine longitudinal axis 9 flow dividing element 32 has an approximately triangular in cross-section, also referred to as a dividing fork 33 shape with an outer Ablenkflanke 34 and an inner Ablenkflanke 35.
  • the deflection flanks 34, 35 converge toward a division tip 36 directed toward the main deflection region 30 and enclose an acute angle of less than 90 °, in particular an angle of 60 °, in the division tip 36.
  • the dividing point or edge 36 forming a branching point divides the cooling air K flowing through the main deflecting region 30 of the diffuser 27 approximately equally into an outer cooling air flow K a and an inner cooling air flow K i .
  • the outer cooling air flow K a is fed through an outer flow transfer chamber 37 of an outer combustion chamber shell 28, while the inner cooling air flow K i is fed via an inner flow transfer chamber 38 of the inner combustion chamber shell 29.
  • the diffuser 27 dividing the cooling air K at the flow dividing element 32 is also referred to as a split diffuser.
  • the cooling air K flowing through the main deflecting region 30 is directed approximately C-shaped radially, relative to the turbine longitudinal axis 9, outwardly to the dividing point 36 of the flow dividing element 32.
  • a line extending as an angle bisector 39 between the curved Ablenkflanken 34,35 through the divisional peak 36 includes with the turbine longitudinal axis 9 a pitch angle ⁇ of about 45 °.
  • the bisector 39 includes an approximately right angle.
  • the inner cooling air flow K i is, starting from the division tip 36, forced by the inner Ablenkflanke 35 first in a horizontal flow direction, ie parallel to the turbine longitudinal axis 9 and further through the outside of the combustion chamber wall 23 radially inward, ie towards the turbine longitudinal axis 9, directed.
  • the inner cooling air flow K i is thus, initially within the undiluted in Hauptablenk Scheme 30 cooling air K, guided radially outwardly in an approximately C-shaped path and thereby delayed and then in a reverse direction approximately C-shaped curved path radially inwardly guided.
  • the flow through the diffuser 27 and further into the internal flow transfer space 38 describes approximately a double S-shaped path. The radii of curvature within this path are large enough to cause only small energy losses in the flow.
  • the outer cooling air flow K a is guided by the dividing fork 33 radially, perpendicular to the turbine longitudinal axis 9, to the outside.
  • the outer cooling air flow K a is guided past the outer combustion chamber shell 28 and introduced into the wall lining room or wall cooling space 26.
  • Similar to the inner cooling air flow Ki results in a flow guide with large deflection radii, with no sudden cross-sectional enlargements occur.
  • Due to the cooling air streams or partial streams K a , K i , the combustion chamber shells 28, 29 are also cooled from the outside.
  • the burner 10 is approximately centered in a combustion chamber rear wall 42 arranged.
  • the wall angle ⁇ corresponds thus about the pitch angle ⁇ . That around the pitch angle ⁇ arranged obliquely to the turbine longitudinal axis 9
  • Flow divider 32 splits main deflection region 30 in an upper sub-channel 43 and a lower sub-channel 44th on, which both have approximately the same cross-section.

Landscapes

  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Chemical & Material Sciences (AREA)
  • Combustion & Propulsion (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
  • Structures Of Non-Positive Displacement Pumps (AREA)
EP03018565A 2003-08-18 2003-08-18 Diffuseur situé entre le compresseur et la chambre de combustion d'une turbine à gaz Withdrawn EP1508680A1 (fr)

Priority Applications (8)

Application Number Priority Date Filing Date Title
EP03018565A EP1508680A1 (fr) 2003-08-18 2003-08-18 Diffuseur situé entre le compresseur et la chambre de combustion d'une turbine à gaz
PCT/EP2004/007946 WO2005019621A1 (fr) 2003-08-18 2004-07-16 Diffuseur place entre le compresseur et la chambre de combustion d'une turbine a gaz
PL04741084T PL1656497T3 (pl) 2003-08-18 2004-07-16 Dyfuzor umieszczony pomiędzy sprężarką i komorą spalania turbiny gazowej
EP04741084A EP1656497B1 (fr) 2003-08-18 2004-07-16 Diffuseur situe entre le compresseur et la chambre de combustion d'une turbine a gaz
ES04741084T ES2275226T3 (es) 2003-08-18 2004-07-16 Difusor localizado entre un compresor y una camara de combustion de una turbina de gas.
DE502004001924T DE502004001924D1 (de) 2003-08-18 2004-07-16 Diffusor zwischen verdichter und brennkammer einer gasturbine angeordnet
US10/568,736 US8082738B2 (en) 2003-08-18 2004-07-16 Diffuser arranged between the compressor and the combustion chamber of a gas turbine
CNB2004800235393A CN100390387C (zh) 2003-08-18 2004-07-16 布置在一种燃气轮机的压缩机和燃烧室之间的扩散器

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
EP03018565A EP1508680A1 (fr) 2003-08-18 2003-08-18 Diffuseur situé entre le compresseur et la chambre de combustion d'une turbine à gaz

Publications (1)

Publication Number Publication Date
EP1508680A1 true EP1508680A1 (fr) 2005-02-23

Family

ID=34042857

Family Applications (2)

Application Number Title Priority Date Filing Date
EP03018565A Withdrawn EP1508680A1 (fr) 2003-08-18 2003-08-18 Diffuseur situé entre le compresseur et la chambre de combustion d'une turbine à gaz
EP04741084A Expired - Lifetime EP1656497B1 (fr) 2003-08-18 2004-07-16 Diffuseur situe entre le compresseur et la chambre de combustion d'une turbine a gaz

Family Applications After (1)

Application Number Title Priority Date Filing Date
EP04741084A Expired - Lifetime EP1656497B1 (fr) 2003-08-18 2004-07-16 Diffuseur situe entre le compresseur et la chambre de combustion d'une turbine a gaz

Country Status (7)

Country Link
US (1) US8082738B2 (fr)
EP (2) EP1508680A1 (fr)
CN (1) CN100390387C (fr)
DE (1) DE502004001924D1 (fr)
ES (1) ES2275226T3 (fr)
PL (1) PL1656497T3 (fr)
WO (1) WO2005019621A1 (fr)

Cited By (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
JP2011027403A (ja) * 2009-07-24 2011-02-10 General Electric Co <Ge> ガスタービン燃焼器のためのシステム及び方法
JP2011153815A (ja) * 2010-01-27 2011-08-11 General Electric Co <Ge> ガスタービンの二次燃焼システムに送給するブリードディフューザ
EP2921779A1 (fr) * 2014-03-18 2015-09-23 Alstom Technology Ltd Chambre de combustion avec manchon de refroidissement
EP3023695A1 (fr) * 2014-11-20 2016-05-25 Siemens Aktiengesellschaft Machine à énergie thermique

Families Citing this family (15)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP1508747A1 (fr) * 2003-08-18 2005-02-23 Siemens Aktiengesellschaft Diffuseur de turbine à gaz et turbine à gaz pour la production d'énergie
US20110303390A1 (en) * 2010-06-14 2011-12-15 Vykson Limited Combustion Chamber Cooling Method and System
US9476429B2 (en) 2012-12-19 2016-10-25 United Technologies Corporation Flow feed diffuser
US10704468B2 (en) 2013-02-28 2020-07-07 Raytheon Technologies Corporation Method and apparatus for handling pre-diffuser airflow for cooling high pressure turbine components
US10267229B2 (en) 2013-03-14 2019-04-23 United Technologies Corporation Gas turbine engine architecture with nested concentric combustor
WO2015009449A1 (fr) * 2013-07-17 2015-01-22 United Technologies Corporation Conduit d'alimentation en air de refroidissement
US20150047358A1 (en) * 2013-08-14 2015-02-19 General Electric Company Inner barrel member with integrated diffuser for a gas turbomachine
US11732892B2 (en) 2013-08-14 2023-08-22 General Electric Company Gas turbomachine diffuser assembly with radial flow splitters
US10060631B2 (en) 2013-08-29 2018-08-28 United Technologies Corporation Hybrid diffuser case for a gas turbine engine combustor
US10465907B2 (en) 2015-09-09 2019-11-05 General Electric Company System and method having annular flow path architecture
JP6625427B2 (ja) * 2015-12-25 2019-12-25 川崎重工業株式会社 ガスタービンエンジン
JP6586389B2 (ja) * 2016-04-25 2019-10-02 三菱重工業株式会社 圧縮機ディフューザおよびガスタービン
US10598380B2 (en) 2017-09-21 2020-03-24 General Electric Company Canted combustor for gas turbine engine
US11808178B2 (en) * 2019-08-05 2023-11-07 Rtx Corporation Tangential onboard injector inlet extender
EP4033073A1 (fr) * 2021-01-25 2022-07-27 Siemens Energy Global GmbH & Co. KG Section de combustion dotée d'un blindage de boîtier

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US5269133A (en) * 1991-06-18 1993-12-14 General Electric Company Heat exchanger for cooling a gas turbine
US5557921A (en) * 1994-05-02 1996-09-24 Abb Management Ag Power plant
US5592821A (en) * 1993-06-10 1997-01-14 Societe Nationale D'etude Et De Construction De Moteurs F'aviation S.N.E.C.M.A. Gas turbine engine having an integral guide vane and separator diffuser
DE19544927A1 (de) * 1995-12-01 1997-04-17 Siemens Ag Gasturbine
DE19639623A1 (de) * 1996-09-26 1998-04-09 Siemens Ag Mischung von zwei Fluidströmen an einem Verdichter
US20030010014A1 (en) * 2001-06-18 2003-01-16 Robert Bland Gas turbine with a compressor for air

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GB8928378D0 (en) * 1989-12-15 1990-02-21 Rolls Royce Plc A diffuser
US5077967A (en) * 1990-11-09 1992-01-07 General Electric Company Profile matched diffuser
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US5555721A (en) * 1994-09-28 1996-09-17 General Electric Company Gas turbine engine cooling supply circuit
GB9917957D0 (en) * 1999-07-31 1999-09-29 Rolls Royce Plc A combustor arrangement
DE50109870D1 (de) * 2001-03-26 2006-06-29 Siemens Ag Gasturbine
EP1400751A1 (fr) * 2002-09-17 2004-03-24 Siemens Aktiengesellschaft Chambre de combustion pour turbine à gaz
GB0229307D0 (en) * 2002-12-17 2003-01-22 Rolls Royce Plc A diffuser arrangement

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Publication number Priority date Publication date Assignee Title
US5269133A (en) * 1991-06-18 1993-12-14 General Electric Company Heat exchanger for cooling a gas turbine
US5592821A (en) * 1993-06-10 1997-01-14 Societe Nationale D'etude Et De Construction De Moteurs F'aviation S.N.E.C.M.A. Gas turbine engine having an integral guide vane and separator diffuser
US5557921A (en) * 1994-05-02 1996-09-24 Abb Management Ag Power plant
DE19544927A1 (de) * 1995-12-01 1997-04-17 Siemens Ag Gasturbine
DE19639623A1 (de) * 1996-09-26 1998-04-09 Siemens Ag Mischung von zwei Fluidströmen an einem Verdichter
US20030010014A1 (en) * 2001-06-18 2003-01-16 Robert Bland Gas turbine with a compressor for air

Cited By (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
JP2011027403A (ja) * 2009-07-24 2011-02-10 General Electric Co <Ge> ガスタービン燃焼器のためのシステム及び方法
JP2011153815A (ja) * 2010-01-27 2011-08-11 General Electric Co <Ge> ガスタービンの二次燃焼システムに送給するブリードディフューザ
EP2921779A1 (fr) * 2014-03-18 2015-09-23 Alstom Technology Ltd Chambre de combustion avec manchon de refroidissement
EP3023695A1 (fr) * 2014-11-20 2016-05-25 Siemens Aktiengesellschaft Machine à énergie thermique

Also Published As

Publication number Publication date
EP1656497A1 (fr) 2006-05-17
US8082738B2 (en) 2011-12-27
CN1836097A (zh) 2006-09-20
DE502004001924D1 (de) 2006-12-14
ES2275226T3 (es) 2007-06-01
US20100257869A1 (en) 2010-10-14
PL1656497T3 (pl) 2007-03-30
CN100390387C (zh) 2008-05-28
EP1656497B1 (fr) 2006-11-02
WO2005019621A1 (fr) 2005-03-03

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