EP1295014A1 - Transition piece for non-annular gas turbine combustion chambers - Google Patents
Transition piece for non-annular gas turbine combustion chambersInfo
- Publication number
- EP1295014A1 EP1295014A1 EP01956465A EP01956465A EP1295014A1 EP 1295014 A1 EP1295014 A1 EP 1295014A1 EP 01956465 A EP01956465 A EP 01956465A EP 01956465 A EP01956465 A EP 01956465A EP 1295014 A1 EP1295014 A1 EP 1295014A1
- Authority
- EP
- European Patent Office
- Prior art keywords
- transition piece
- gas turbine
- combustion chambers
- terminal end
- axis
- 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.)
- Granted
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
Definitions
- the present invention relates to a transition piece for non-annular gas turbine combustion chambers.
- gas turbines comprise a compressor, to which air from the external environment is supplied in order to pressurise the compressor.
- the compressed air passes into a series of combustion chambers which end in a nozzle, into each of which an injector supplies fuel, which is mixed with the air in order to form a mixture of combustible air to be burnt .
- the burnt gases are conveyed towards the turbine, which transforms the enthalpy of the burnt gases in the said combustion chamber into mechanical energy which is available to a user.
- transition pieces in a gas turbine are substantially tubular elements, each of which is used to connect a combustion chamber, which belongs to the combustion system of the gas turbine, to the first stage of the said gas turbine.
- the combustion chambers comprise a plurality of elements with a generally cylindrical shape, which are provided with nozzles disposed around an annular development.
- the transition pieces are disposed with an annular development, in order to direct the flow of hot burnt gases from the combustion chambers to the first stage of the gas turbine.
- the transition pieces have an upstream aperture for the flow of gas, which is generally cylindrical, and is used to receive the flow of gas directly from the corresponding combustion chamber, and are configured in a longitudinal direction such that their downstream ends comprise arcuate segments which open towards the first stage of the gas turbine.
- transition pieces can direct the flow of gas with a high level of enthalpy obtained from spaced, generally cylindrical flow configurations, towards arcuate segments which form a ring-type configuration relative to the first stage of the gas turbine.
- the combustion system since the combustion system has an energising function in relation to the thermal carrier fluid, it is the true heart of the turbine engine, and thus defines its level of emission, and, according to the service life of its own components, also defines the intervals of functioning between machine stoppages, which are necessary in order to carry out inspections of the combustion chambers.
- the object of the present invention is therefore to provide a transition piece for non-annular gas turbine combustion chambers, which permits optimisation of the performance in operation.
- a further object of the invention is to provide a transition piece for non-annular gas turbine combustion chambers, which permits improved efficiency in operation.
- Another object of the invention is to provide a transition piece for non-annular gas turbine combustion chambers, which permits a longer service life of the machine.
- a further object of the invention is to provide a transition piece for a non-annular gas turbine combustion chamber system, which generates fewer pollutant emissions.
- the transition piece has a support arm, with a circular or elliptical base, which is present at the said second terminal end, and surrounds the flow aperture of the said transition piece, wherein the said support arm with a circular or elliptical base has an upper side and a lower side, for its own connection to the first stage of the said gas turbine.
- the cylindrical section relative to the first terminal end, has an anti-wear deposit made of Stellite 6 or another, similar material.
- a deposit of the TBC type is provided on the entire inner surface of the said transition piece, in order to reduce the temperature of the metal.
- the distance between the centre of the lead-in of the said transition piece and the support plane of the flange for anchorage to the stator of the said gas turbine is between 350 mm and 380 mm.
- the body of the transition piece can be made of Nimonic 263 or of Hastelloy-X. Further characteristics of the invention are defined in the claims attached to the present patent application.
- - figure 1 represents a view, partially in cross- section, of a gas turbine to which there is fitted the transition piece according to the present invention, for non-annular gas turbine combustion chambers
- - figure 2 represents a plan view of a transition piece according to the present invention, for non- annular gas turbine combustion chambers
- figure 3 represents a view in cross-section according to the plane III-III, of the transition piece in figure 2
- figure 4 represents a front view of the transition piece in figures 2-3
- figure 5 represents a view in cross-section of a detail belonging to the transition piece in figures 2-4.
- the transition piece 10 consists of a body 11, which receives upstream the flow of burnt gases, directly from the corresponding combustion chamber 25, and is configured longitudinally such as to have a downstream end which opens towards the first stage 17 of the gas turbine 18.
- the body 11 of the transition piece 10 has a cylindrical section 13, relative to a first terminal end, which ends with an aperture 14 for connection to a combustion chamber 25.
- the body 11 of the transition piece 10 also has a second terminal end 16, which is connected to the first stage 17 of a gas turbine 18, and thus has a flow aperture 15.
- the transition piece 10 thus consists of a body 11, which is provided with a support arm, with a circular or elliptical base, which is present at the second terminal end 16, and surrounds the flow aperture 15.
- the support arm with a circular or elliptical base has an upper side 20 and a lower side 21.
- the body 11 is provided with characteristics of resilience, such as to minimise the stresses associated with its anchorage onto the stator of the said gas turbine 18.
- the body 11 is preferably made of Nimonic 263, or, alternatively, of Hastelloy-X, and is produced in a single component part, indicated by the reference number 11, which reduces the parts necessary and the assembly time, and thus permits considerable savings.
- the cylindrical section 13, relative to the first terminal end, has an anti-wear deposit made of Stellite 6, or of another material homologous with that used on the Hula Seals of the liners.
- a deposit of the TBC type is provided on the entire inner surface of the body 11 of the transition piece 10, in order to reduce the temperature of the metal.
- the transition piece 11 also has a projecting connection arm, indicated as a whole by the reference number 22, for support on the turbine stator.
- the projecting connection arm 22 also has a hole 30, for its own connection to the first stage of the gas turbine 18, and a centring pin 31.
- the geometry of the transition piece 11 is thus completely re-shaped, compared with the known art. In fact, it has a lead-in axis with an angle of between 5° and 7°, relative to an axis which is parallel to the axis of the machine, and is perpendicular to the plane on which there lies the frame for interfacing of the transition piece 10 with the ring of nozzles which belong to the first stage 17 of the gas turbine 18.
- the distance between the centre of the lead-in of the transition piece 10 and the support plane of the flange for anchorage to the stator of the gas turbine is between
- the most significant functioning parameters are: Tmax of the gas ⁇ 1300 °C, and Pmax of the gas ⁇ 10 Ata.
- the transition piece 11 thus formed has an optimal geometric profile, which, inter alia, makes it possible to keep the thermal stresses within acceptable limits.
- the particular design of the transition piece 11 described makes it possible to obtain increased structural stability, with consequent reduction of the vibratory motion.
Landscapes
- Engineering & Computer Science (AREA)
- Mechanical Engineering (AREA)
- General Engineering & Computer Science (AREA)
- Turbine Rotor Nozzle Sealing (AREA)
- Gas Burners (AREA)
Abstract
Description
Claims
Applications Claiming Priority (3)
Application Number | Priority Date | Filing Date | Title |
---|---|---|---|
IT2000MI001362A IT1317978B1 (en) | 2000-06-16 | 2000-06-16 | TRANSITION PIECE FOR COMBUSTION CHAMBERS OF NONANULAR GAS TURBINES. |
ITMI001362 | 2000-06-16 | ||
PCT/EP2001/006484 WO2001096712A1 (en) | 2000-06-16 | 2001-06-08 | Transition piece for non-annular gas turbine combustion chambers |
Publications (2)
Publication Number | Publication Date |
---|---|
EP1295014A1 true EP1295014A1 (en) | 2003-03-26 |
EP1295014B1 EP1295014B1 (en) | 2005-11-23 |
Family
ID=11445283
Family Applications (1)
Application Number | Title | Priority Date | Filing Date |
---|---|---|---|
EP01956465A Expired - Lifetime EP1295014B1 (en) | 2000-06-16 | 2001-06-08 | Transition piece for non-annular gas turbine combustion chambers |
Country Status (12)
Country | Link |
---|---|
US (1) | US20030167776A1 (en) |
EP (1) | EP1295014B1 (en) |
AR (1) | AR028133A1 (en) |
AT (1) | ATE310896T1 (en) |
AU (1) | AU7844301A (en) |
BR (1) | BR0111577A (en) |
DE (1) | DE60115236T2 (en) |
IT (1) | IT1317978B1 (en) |
MX (1) | MXPA02012451A (en) |
NO (1) | NO330416B1 (en) |
RU (1) | RU2275554C2 (en) |
WO (1) | WO2001096712A1 (en) |
Families Citing this family (31)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
US7178341B2 (en) * | 2004-06-17 | 2007-02-20 | Siemens Power Generation, Inc. | Multi-zone tubing assembly for a transition piece of a gas turbine |
US7278254B2 (en) * | 2005-01-27 | 2007-10-09 | Siemens Power Generation, Inc. | Cooling system for a transition bracket of a transition in a turbine engine |
US8015818B2 (en) * | 2005-02-22 | 2011-09-13 | Siemens Energy, Inc. | Cooled transition duct for a gas turbine engine |
US7827801B2 (en) * | 2006-02-09 | 2010-11-09 | Siemens Energy, Inc. | Gas turbine engine transitions comprising closed cooled transition cooling channels |
US8151570B2 (en) * | 2007-12-06 | 2012-04-10 | Alstom Technology Ltd | Transition duct cooling feed tubes |
US8113003B2 (en) * | 2008-08-12 | 2012-02-14 | Siemens Energy, Inc. | Transition with a linear flow path for use in a gas turbine engine |
US8091365B2 (en) * | 2008-08-12 | 2012-01-10 | Siemens Energy, Inc. | Canted outlet for transition in a gas turbine engine |
US20120186269A1 (en) * | 2011-01-25 | 2012-07-26 | General Electric Company | Support between transition piece and impingement sleeve in combustor |
KR101613096B1 (en) * | 2011-10-24 | 2016-04-20 | 제네럴 일렉트릭 테크놀러지 게엠베하 | Gas turbine |
US9593853B2 (en) * | 2014-02-20 | 2017-03-14 | Siemens Energy, Inc. | Gas flow path for a gas turbine engine |
WO2017039567A1 (en) * | 2015-08-28 | 2017-03-09 | Siemens Aktiengesellschaft | Non-axially symmetric transition ducts for combustors |
US10830442B2 (en) | 2016-03-25 | 2020-11-10 | General Electric Company | Segmented annular combustion system with dual fuel capability |
US10563869B2 (en) | 2016-03-25 | 2020-02-18 | General Electric Company | Operation and turndown of a segmented annular combustion system |
US10584880B2 (en) | 2016-03-25 | 2020-03-10 | General Electric Company | Mounting of integrated combustor nozzles in a segmented annular combustion system |
US11428413B2 (en) | 2016-03-25 | 2022-08-30 | General Electric Company | Fuel injection module for segmented annular combustion system |
US10520194B2 (en) | 2016-03-25 | 2019-12-31 | General Electric Company | Radially stacked fuel injection module for a segmented annular combustion system |
US10605459B2 (en) | 2016-03-25 | 2020-03-31 | General Electric Company | Integrated combustor nozzle for a segmented annular combustion system |
US10641175B2 (en) | 2016-03-25 | 2020-05-05 | General Electric Company | Panel fuel injector |
US10641491B2 (en) | 2016-03-25 | 2020-05-05 | General Electric Company | Cooling of integrated combustor nozzle of segmented annular combustion system |
US10584876B2 (en) | 2016-03-25 | 2020-03-10 | General Electric Company | Micro-channel cooling of integrated combustor nozzle of a segmented annular combustion system |
JP6650849B2 (en) * | 2016-08-25 | 2020-02-19 | 三菱日立パワーシステムズ株式会社 | gas turbine |
US11156362B2 (en) | 2016-11-28 | 2021-10-26 | General Electric Company | Combustor with axially staged fuel injection |
US10690350B2 (en) | 2016-11-28 | 2020-06-23 | General Electric Company | Combustor with axially staged fuel injection |
US10837299B2 (en) * | 2017-03-07 | 2020-11-17 | General Electric Company | System and method for transition piece seal |
US11994292B2 (en) | 2020-08-31 | 2024-05-28 | General Electric Company | Impingement cooling apparatus for turbomachine |
US11460191B2 (en) | 2020-08-31 | 2022-10-04 | General Electric Company | Cooling insert for a turbomachine |
US11994293B2 (en) | 2020-08-31 | 2024-05-28 | General Electric Company | Impingement cooling apparatus support structure and method of manufacture |
US11371702B2 (en) | 2020-08-31 | 2022-06-28 | General Electric Company | Impingement panel for a turbomachine |
US11614233B2 (en) | 2020-08-31 | 2023-03-28 | General Electric Company | Impingement panel support structure and method of manufacture |
US11255545B1 (en) | 2020-10-26 | 2022-02-22 | General Electric Company | Integrated combustion nozzle having a unified head end |
US11767766B1 (en) | 2022-07-29 | 2023-09-26 | General Electric Company | Turbomachine airfoil having impingement cooling passages |
Family Cites Families (3)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
US5964091A (en) * | 1995-07-11 | 1999-10-12 | Hitachi, Ltd. | Gas turbine combustor and gas turbine |
US6006523A (en) * | 1997-04-30 | 1999-12-28 | Mitsubishi Heavy Industries, Ltd. | Gas turbine combustor with angled tube section |
US6393828B1 (en) * | 1997-07-21 | 2002-05-28 | General Electric Company | Protective coatings for turbine combustion components |
-
2000
- 2000-06-16 IT IT2000MI001362A patent/IT1317978B1/en active
-
2001
- 2001-06-08 AT AT01956465T patent/ATE310896T1/en active
- 2001-06-08 MX MXPA02012451A patent/MXPA02012451A/en unknown
- 2001-06-08 BR BR0111577-4A patent/BR0111577A/en not_active Application Discontinuation
- 2001-06-08 RU RU2003101058/06A patent/RU2275554C2/en not_active IP Right Cessation
- 2001-06-08 US US10/311,251 patent/US20030167776A1/en not_active Abandoned
- 2001-06-08 EP EP01956465A patent/EP1295014B1/en not_active Expired - Lifetime
- 2001-06-08 DE DE60115236T patent/DE60115236T2/en not_active Expired - Lifetime
- 2001-06-08 WO PCT/EP2001/006484 patent/WO2001096712A1/en active IP Right Grant
- 2001-06-08 AU AU78443/01A patent/AU7844301A/en not_active Abandoned
- 2001-06-15 AR ARP010102872A patent/AR028133A1/en unknown
-
2002
- 2002-12-13 NO NO20025995A patent/NO330416B1/en not_active IP Right Cessation
Non-Patent Citations (1)
Title |
---|
See references of WO0196712A1 * |
Also Published As
Publication number | Publication date |
---|---|
RU2275554C2 (en) | 2006-04-27 |
WO2001096712A1 (en) | 2001-12-20 |
MXPA02012451A (en) | 2003-04-25 |
ITMI20001362A1 (en) | 2001-12-16 |
ITMI20001362A0 (en) | 2000-06-16 |
AU7844301A (en) | 2001-12-24 |
EP1295014B1 (en) | 2005-11-23 |
ATE310896T1 (en) | 2005-12-15 |
DE60115236D1 (en) | 2005-12-29 |
AR028133A1 (en) | 2003-04-23 |
DE60115236T2 (en) | 2006-07-27 |
BR0111577A (en) | 2003-03-18 |
NO20025995D0 (en) | 2002-12-13 |
US20030167776A1 (en) | 2003-09-11 |
NO330416B1 (en) | 2011-04-11 |
IT1317978B1 (en) | 2003-07-21 |
NO20025995L (en) | 2003-02-13 |
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