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WO1998016764A1 - Joint a brosse toroidal pour interface de transition du dispositif combustor et de la turbine a gaz - Google Patents

Joint a brosse toroidal pour interface de transition du dispositif combustor et de la turbine a gaz Download PDF

Info

Publication number
WO1998016764A1
WO1998016764A1 PCT/US1997/013218 US9713218W WO9816764A1 WO 1998016764 A1 WO1998016764 A1 WO 1998016764A1 US 9713218 W US9713218 W US 9713218W WO 9816764 A1 WO9816764 A1 WO 9816764A1
Authority
WO
WIPO (PCT)
Prior art keywords
combustor
gas turbine
brush seal
air
bristles
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.)
Ceased
Application number
PCT/US1997/013218
Other languages
English (en)
Inventor
Graydon L. Whidden
Raymond E. Chupp
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 Energy Inc
Westinghouse Electric Corp
Original Assignee
Siemens Westinghouse Power Corp
Westinghouse Electric Corp
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 Westinghouse Power Corp, Westinghouse Electric Corp filed Critical Siemens Westinghouse Power Corp
Publication of WO1998016764A1 publication Critical patent/WO1998016764A1/fr
Anticipated expiration legal-status Critical
Ceased legal-status Critical Current

Links

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
    • F01D9/00—Stators
    • F01D9/02—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles
    • F01D9/023—Transition ducts between combustor cans and first stage of the turbine in gas-turbine engines; their cooling or sealings
    • 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
    • F01D11/00—Preventing or minimising internal leakage of working-fluid, e.g. between stages
    • F01D11/001—Preventing or minimising internal leakage of working-fluid, e.g. between stages for sealing space between stator blade and rotor
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F16—ENGINEERING ELEMENTS AND UNITS; GENERAL MEASURES FOR PRODUCING AND MAINTAINING EFFECTIVE FUNCTIONING OF MACHINES OR INSTALLATIONS; THERMAL INSULATION IN GENERAL
    • F16J—PISTONS; CYLINDERS; SEALINGS
    • F16J15/00—Sealings
    • F16J15/16—Sealings between relatively-moving surfaces
    • F16J15/32—Sealings between relatively-moving surfaces with elastic sealings, e.g. O-rings
    • F16J15/3284—Sealings between relatively-moving surfaces with elastic sealings, e.g. O-rings characterised by their structure; Selection of materials
    • F16J15/3288—Filamentary structures, e.g. brush seals
    • 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
    • F23R3/60—Support structures; Attaching or mounting means
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2230/00—Manufacture
    • F05D2230/60—Assembly methods
    • F05D2230/64—Assembly methods using positioning or alignment devices for aligning or centring, e.g. pins
    • F05D2230/642—Assembly methods using positioning or alignment devices for aligning or centring, e.g. pins using maintaining alignment while permitting differential dilatation

Definitions

  • the present invention relates to combustors capable of burning fuel in compressed air for a power generation system, such as a gas turbine power system. More particularly, the present invention relates to a self- accommodating brush seal at the interface of the combustor and the transition area of the gas turbine.
  • the combustor- transition seal significantly reduces the amount of leakage of air into the transition section that would otherwise flow into the combustor, thus reducing the amount of harmful NOx produced by combustion and increasing the efficiency of the turbine system.
  • fuel is burned in compressed air, produced by a compressor, in one or more combustors
  • a chamber is formed by an outer casing and a plurality of combustors are circumferentially arrayed within the chamber.
  • a transition section one for each combustor which distributes the hot combustion gases to the turbine for expansion.
  • Air from the compressor section is distributed into the combustion section chamber such that the air is expose-i to the outer surfaces of the combustor liner and the transition section before the air flows through the inlet mcc the combustors.
  • an adequate supply of air intake into the combustor is necessary for, among other things, reducing to acceptable levels the emissions of environmentally harmful by-products of the c ⁇ moustion process, primarily NOx.
  • the availability of more air input to the combustor provides for burning of the fuel at a lean fuel-to-air mixture ratio.
  • m known turbine designs there is a substantial amount of leakage of air through the mechanical coupling where the combustor interfaces with the transition piece that would otherwise flow across the outer surfaces of the combustor and transition section and into the combustor inlet.
  • Sealing devices are known to minimize leakage at the interface of the aft end of the transition section and the first stage nozzle support elements.
  • Such known seal mechanisms are disclosed in U.S Patent Nos. 5,265,412 and 5,400586 -- both to Bagepalli et al .
  • these patents disclose a brush seal device for sealing the space between the transition piece and the first stage nozzle to minimize leakage of the combustion products into the transition-nozzle interface.
  • These patents do not disclose means for sealing the interface of the combustor and the transition area to minimize leakage of compressed air into the transition area, wherein that air would otherwise flow into the combustor for the generation of the combustion products. It is therefore desirable to provide a gas turbine combustion section having a seal at the interface of the aft end of the combustor and the front end of the transition section, so as to reduce the formation of NOx and to increase the overall efficiency of the turbine system.
  • curren t invention provides a gas turbine combustion section wherein the amount of leakage of compressed air that would otherwise be available to supply the combustors is reduced, with the attendant benefits of operation of the combustors with very lean mixtures of fuel and air and increased turbine efficiency.
  • a gas turbine comprising a compressor for compressing air, a combustor for producing a hot gas by burning a fuel in the compressed air, a transition piece disposed proximate the combustor with a gap between the transition piece and the combustor, a brush seal disposed in the gap to minimize the compressed air from leaking through the gap, and a turbine for expanding the hot gas produced by the combustor.
  • the brush seal is a densely packed bed of directionally compliant bristles clamped between an upstream and downstream retainer . The upstream and downstream retainer provide mechanical support for the sealed pressure load.
  • a retaining ring that provides support for the brush seal can be attached to either the transition piece or the combustor, such that the brush seal can be mounted to either the transition, with the bristles slightly deflected by the combustor, or to the combustor, with the bristles contacting the transition piece. Axial and radial movement between the combustor and the transition piece is accommodated by the bristles in the seal.
  • Figure 1 is a longitudinal cross-section through a portion of a gas turbine in the vicinity of the combustion section.
  • Figure 2A is an exploded view of Section A-A of FIG. 1, showing a brush seal in accordance with a preferred embodiment of the invention.
  • Figure 2B is an exploded view of Section B-B of FIG. 2A.
  • Figure 2C is a cross-sectional view of a brush seal taken along lines C-C of FIG. 2A.
  • Figure 1 a portion of a gas turbine 1 in the vicinity of the combustion section 6.
  • the combustion section 6 In addition to the combustion section
  • the gas turbine comprises a compressor section 2 and a turbine section 3.
  • the current invention concerns a combustor
  • the combustion section 6 comprises a chamber 7 formed by an outer casing 18 of the gas turbine and in which a plurality of combustors 4 are circumferentially arrayed.
  • Each combustor comprises a pre-mixing zone 14 and a combustion zone 10 downstream of the pre-mixing zone. Fuel 11, 12 and compressed air 8 from the compressor 2 are mixed in the pre- mixing zone 14 and burned in the combustion zone 10.
  • a transition section (or transition piece) 5 comprises a duct that is mechanically coupled to a vessel 16, which vessel 16 encloses the combustion zone 10 and generally forms the aft end of the combustor 4.
  • the transition section 5 directs the hot gas 9 produced by the combustor 4 to the turbine 3 for expansion.
  • the forward portion of the transition section 5 can be mechanically coupled to the aft end of the combustor vessel
  • one currently known mechanical coupling is a spring clip 20.
  • the spring clip 20 is an annular flexible member that comprises a plurality of leaves spaced closely together in the circumferential direction.
  • the spring clip 20 is mounted to the combustor vessel 16 and the transition section 5 rests on the upper curved surface 21 of the spring clip. Due to the presence of the small gap (not shown) between each of the leaves of the spring clip, the leaves deflect in the radial direction under the influence of the transition section such that the transition section generally remains in contact with the spring clip.
  • the spring clip 20 ensures that the combustor is centered with respect to the inlet of the transition section and the contact between the transition section 5 and the upper curved surface 21 of the spring clip minimizes the amount air leaking into the transition area over the top of the spring clip.
  • air flowing from the compressor section to the combustor inlet is known to leak through the gaps and into the transition section. Accordingly, and as described above, spring clips allow for substantial leakage of air into the transition section that would otherwise be available for input into the combustors, resulting in decreased turbine efficiency and undesirable emissions of NOx.
  • FIGS. 2A-2C a seal mechanism in accordance with the present invention is shown in FIGS. 2A-2C.
  • Brush seal 22 comprises a densely packed bed of directionally compliant bristles 24 clamped between an upstream retainer 26 and a downstream retainer 28.
  • the upstream and downstream retainers 26, 28 provide mechanical support for the sealed pressure load on the brush seal.
  • Retaining r ng 30 provides further mechanical support for the brush seal onto the inner surfac of the transition section 5.
  • the means for mechanically affixing the brush seal within the combustor-transitier. section interface is not intended to be limited to the coupling described herein and other suitable means for mechanically securing the brush seal into position will be known to those skilled in the art .
  • the ends of the bristles 24 of the brush seal 22 are formed into a circumferential surface 32 such that the bristles conform to the outer circumferential surface of the combustor vessel 16.
  • the densely packed bed of directionally compliant bristles 24 can be optimally designed to minimize leakage of air through the seal in accordance with the description set forth in "Generalizing Circular Brush Seal Leakage Through a Randomly Distributed Bristle Bed," R.E. Chupp and G.F. Holle, Journal of Turbomachinery, Volume 118, January 1996, which is herein incorporated by reference as if set forth in its entirety.
  • a brush seal in accordance with the present invention is shown mounted to the transition section 5 with the bristles 24 slightly deflected by the combustor vessel 16.
  • the brush seal can also be mounted to the combustor vessel, using the same type of retention mechanism described herein, such that the bristles 24 contact the transition piece 5.
  • the operation of the combustor 4 is as follows.
  • the compressor 2 is spun-up to ignition speed and as the compressor rotor accelerates, compressed air 8 from the compressor 2 flows into the chamber 7 of the combustion section.
  • a portion of the compressed air 8 passes across the outer surfaces of the combustor vessel 16 and the transition piece 5.
  • a portion of the compressed air 8 has a tendency to flow through the spring clip 20 at the interface of the combustor and the transition section.
  • the brush seal 22 minimizes leakage of the air that tends to flow through the spring clip and intc the transition section such that this air is available fcr input into the combustor to provide a more lean mixture cf fuel and air in the combustor.
  • leakage through known spring clips would be reduced to less than one percen .
  • a brush seal in accordance with the present invention also accommodates thermal growth and other movement between the combustor and transition section in the axial and radial direction.
  • Sealing mechanisms such as the brush seal 22, in accordance with the present invention can be used for sealing the interface between the combustor and transition section in conjunction with any other known mechanical coupling devices for mechanically securing the combustor to the transition, such as known stand-off pins.
  • the sealing mechanism must only be positioned downstream of the combustor-transition mechanical coupling to minimize the amount of air leaking through the coupling and into the transition section.

Landscapes

  • Engineering & Computer Science (AREA)
  • General Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • Chemical & Material Sciences (AREA)
  • Combustion & Propulsion (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)

Abstract

L'invention porte sur un mécanisme d'étanchéité permettant de réduire les fuites d'air dans l'accouplement mécanique du dispositif combustor et de la section de transition d'une turbine à gaz. Selon une réalisation préférée de l'invention, le mécanisme d'étanchéité comprend un joint à brosse toroïdal qui peut être monté sur la section de transition ou sur le dispositif combustor. Ce joint à brosse toroïdal réduit les fuites d'air dans la section de transition étant donné que l'air a tendance à s'écouler devant l'accouplement mécanique du dispositif combustor et de la pièce de transition. En conséquence, l'air qui pourrait fuir dans la section de transition s'écoule dans le dispositif combustor de façon à créer des mélanges plus pauvres de combustible et d'air dans le dispositif combustor, les turbines à gaz de cette invention ayant ainsi de faibles taux d'émissions de NOx et un meilleur rendement global.
PCT/US1997/013218 1996-10-16 1997-07-29 Joint a brosse toroidal pour interface de transition du dispositif combustor et de la turbine a gaz Ceased WO1998016764A1 (fr)

Applications Claiming Priority (2)

Application Number Priority Date Filing Date Title
US73303996A 1996-10-16 1996-10-16
US08/733,039 1996-10-16

Publications (1)

Publication Number Publication Date
WO1998016764A1 true WO1998016764A1 (fr) 1998-04-23

Family

ID=24945974

Family Applications (1)

Application Number Title Priority Date Filing Date
PCT/US1997/013218 Ceased WO1998016764A1 (fr) 1996-10-16 1997-07-29 Joint a brosse toroidal pour interface de transition du dispositif combustor et de la turbine a gaz

Country Status (2)

Country Link
JP (1) JPH10153129A (fr)
WO (1) WO1998016764A1 (fr)

Cited By (8)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
GB2361304A (en) * 2000-04-14 2001-10-17 Rolls Royce Plc Combustor wall tile
WO2005108869A1 (fr) * 2004-05-05 2005-11-17 Alstom Technology Ltd Chambre de combustion pour une turbine a gaz
US7096668B2 (en) * 2003-12-22 2006-08-29 Martling Vincent C Cooling and sealing design for a gas turbine combustion system
EP1767835A1 (fr) * 2005-09-22 2007-03-28 Siemens Aktiengesellschaft Dispositif d'étanchéité résistant aux hautes températures, en particulier pour turbines à gaz
EP2085575A1 (fr) * 2008-02-01 2009-08-05 Siemens Aktiengesellschaft Combinaison de joints de brosses dotée d'une bague de piston pour grands espaces d'étanchéité
CN104373965A (zh) * 2014-10-28 2015-02-25 北京华清燃气轮机与煤气化联合循环工程技术有限公司 过渡段后密封结构
EP3136000A1 (fr) * 2015-08-24 2017-03-01 General Electric Company Système de patin d'usure pour systèmes de combustion à turbine et procédé de couplage d'un patin d'usure dans un système de combustion à turbine
US9840924B2 (en) 2014-08-15 2017-12-12 Siemens Aktiengesellschaft Gas turbine system with a transition duct having axially extending cooling channels

Families Citing this family (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP1118806A1 (fr) * 2000-01-20 2001-07-25 Siemens Aktiengesellschaft Structure de paroi sous charge thermique et méthode pour fermer des fentes dans une tel structure
JP3795036B2 (ja) * 2003-03-14 2006-07-12 三菱重工業株式会社 タービン尾筒のシール構造およびシール装置
EP2128524A1 (fr) * 2008-05-26 2009-12-02 Siemens Aktiengesellschaft Agencement de composant, agencement de chambre de combustion et turbine à gaz
WO2025229806A1 (fr) * 2024-04-30 2025-11-06 三菱パワー株式会社 Chambre de combustion et turbine à gaz pourvue de celle-ci

Citations (8)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
FR943933A (fr) * 1946-03-25 1949-03-22 Rolls Royce Perfectionnements aux groupes moto-propulseurs à réaction et à turbine à gaz
GB622384A (en) * 1947-03-14 1949-05-02 Arthur Alexander Rubbra Improvements relating to gas turbine engines
US2494659A (en) * 1944-12-22 1950-01-17 Lucas Ltd Joseph Pipe joint
US3759038A (en) * 1971-12-09 1973-09-18 Westinghouse Electric Corp Self aligning combustor and transition structure for a gas turbine
US5265412A (en) 1992-07-28 1993-11-30 General Electric Company Self-accommodating brush seal for gas turbine combustor
US5400586A (en) 1992-07-28 1995-03-28 General Electric Co. Self-accommodating brush seal for gas turbine combustor
GB2306594A (en) * 1995-10-25 1997-05-07 Europ Gas Turbines Ltd Interconnecting and sealing two tubular structures
WO1997026485A1 (fr) * 1996-01-17 1997-07-24 Mitsubishi Jukogyo Kabushiki Kaisha Unite a joints elastiques de chambre de combustion

Patent Citations (8)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US2494659A (en) * 1944-12-22 1950-01-17 Lucas Ltd Joseph Pipe joint
FR943933A (fr) * 1946-03-25 1949-03-22 Rolls Royce Perfectionnements aux groupes moto-propulseurs à réaction et à turbine à gaz
GB622384A (en) * 1947-03-14 1949-05-02 Arthur Alexander Rubbra Improvements relating to gas turbine engines
US3759038A (en) * 1971-12-09 1973-09-18 Westinghouse Electric Corp Self aligning combustor and transition structure for a gas turbine
US5265412A (en) 1992-07-28 1993-11-30 General Electric Company Self-accommodating brush seal for gas turbine combustor
US5400586A (en) 1992-07-28 1995-03-28 General Electric Co. Self-accommodating brush seal for gas turbine combustor
GB2306594A (en) * 1995-10-25 1997-05-07 Europ Gas Turbines Ltd Interconnecting and sealing two tubular structures
WO1997026485A1 (fr) * 1996-01-17 1997-07-24 Mitsubishi Jukogyo Kabushiki Kaisha Unite a joints elastiques de chambre de combustion

Cited By (13)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
GB2361304A (en) * 2000-04-14 2001-10-17 Rolls Royce Plc Combustor wall tile
US7096668B2 (en) * 2003-12-22 2006-08-29 Martling Vincent C Cooling and sealing design for a gas turbine combustion system
US7752846B2 (en) 2004-05-05 2010-07-13 Alstom Technology Ltd Combustion chamber for a gas turbine
WO2005108869A1 (fr) * 2004-05-05 2005-11-17 Alstom Technology Ltd Chambre de combustion pour une turbine a gaz
EP1767835A1 (fr) * 2005-09-22 2007-03-28 Siemens Aktiengesellschaft Dispositif d'étanchéité résistant aux hautes températures, en particulier pour turbines à gaz
WO2007033974A1 (fr) * 2005-09-22 2007-03-29 Siemens Aktiengesellschaft Garniture d'etancheite resistant aux temperatures elevees, conçue en particulier pour des turbines a gaz
US8365538B2 (en) 2005-09-22 2013-02-05 Siemens Aktiengesellschaft High temperature-resistant sealing assembly, especially for gas turbines
WO2009095402A1 (fr) * 2008-02-01 2009-08-06 Siemens Aktiengesellschaft Combinaison d’un joint d’étanchéité à broche et d’un segment de piston pour de grandes fentes d’étanchéité
EP2085575A1 (fr) * 2008-02-01 2009-08-05 Siemens Aktiengesellschaft Combinaison de joints de brosses dotée d'une bague de piston pour grands espaces d'étanchéité
US9840924B2 (en) 2014-08-15 2017-12-12 Siemens Aktiengesellschaft Gas turbine system with a transition duct having axially extending cooling channels
CN104373965A (zh) * 2014-10-28 2015-02-25 北京华清燃气轮机与煤气化联合循环工程技术有限公司 过渡段后密封结构
EP3136000A1 (fr) * 2015-08-24 2017-03-01 General Electric Company Système de patin d'usure pour systèmes de combustion à turbine et procédé de couplage d'un patin d'usure dans un système de combustion à turbine
US10634349B2 (en) 2015-08-24 2020-04-28 General Electric Company Wear pad system for turbine combustion systems and method for coupling wear pad into turbine combustion system

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