EP2666972A2 - Turbinenummantelungskühlanordnung für ein Gasturbinensystem und zugehöriges Gasturbinensystem - Google Patents

Turbinenummantelungskühlanordnung für ein Gasturbinensystem und zugehöriges Gasturbinensystem Download PDF

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Publication number
EP2666972A2
EP2666972A2 EP13168436.7A EP13168436A EP2666972A2 EP 2666972 A2 EP2666972 A2 EP 2666972A2 EP 13168436 A EP13168436 A EP 13168436A EP 2666972 A2 EP2666972 A2 EP 2666972A2
Authority
EP
European Patent Office
Prior art keywords
passage
cooling
side portion
turbine
cooling assembly
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Withdrawn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
EP13168436.7A
Other languages
English (en)
French (fr)
Inventor
Michelle Jessica Rogers
Gregory Thomas Foster
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.)
General Electric Co
Original Assignee
General Electric Co
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 General Electric Co filed Critical General Electric Co
Publication of EP2666972A2 publication Critical patent/EP2666972A2/de
Withdrawn legal-status Critical Current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
    • F01D9/00—Stators
    • F01D9/02—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles
    • F01D9/04—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles forming ring or sector
    • 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
    • F01D11/08—Preventing or minimising internal leakage of working-fluid, e.g. between stages for sealing space between rotor blade tips and stator
    • F01D11/14—Adjusting or regulating tip-clearance, i.e. distance between rotor-blade tips and stator casing
    • F01D11/20—Actively adjusting tip-clearance
    • F01D11/24—Actively adjusting tip-clearance by selectively cooling-heating stator or rotor components
    • 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
    • F05D2240/00—Components
    • F05D2240/10—Stators
    • F05D2240/11—Shroud seal segments
    • 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
    • F05D2260/00—Function
    • F05D2260/20—Heat transfer, e.g. cooling
    • F05D2260/221—Improvement of heat transfer
    • F05D2260/2214—Improvement of heat transfer by increasing the heat transfer surface
    • F05D2260/22141—Improvement of heat transfer by increasing the heat transfer surface using fins or ribs

Definitions

  • the subject matter disclosed herein relates to gas turbine systems, and more particularly to a turbine shroud cooling assembly for cooling turbine shrouds of such gas turbine systems.
  • a combustor converts the chemical energy of a fuel or an air-fuel mixture into thermal energy.
  • the thermal energy is conveyed by a fluid, often compressed air from a compressor, to a turbine where the thermal energy is converted to mechanical energy.
  • hot gas is flowed over and through portions of the turbine as a hot gas path. High temperatures along the hot gas path can heat turbine components, causing degradation of components.
  • a turbine shroud assembly is an example of a component that is subjected to the hot gas path and often comprises two separate pieces, such as an inner shroud and an outer shroud.
  • various cooling schemes have been employed to maintain the structural integrity, as well as the intended functionality, of the inner shroud. Such cooling schemes typically result in excessive cooling flow from a cooling source, thereby sacrificing overall efficiency of the gas turbine system.
  • a turbine shroud cooling assembly for a gas turbine system includes an inner shroud component disposed within a turbine section of the gas turbine system and proximate a hot gas path therein, wherein the inner shroud component includes a base portion in direct contact with the hot gas path. Also includes is a rib protruding radially away from the base portion and disposed proximate at least one cavity configured to receive a cooling flow from a cooling source, wherein the cooling flow passes through the main passage of the rib for cooling the inner shroud component.
  • a turbine shroud cooling assembly for a gas turbine system includes an inner shroud component disposed within a turbine section of the gas turbine system and proximate a hot gas path therein, wherein the inner shroud component includes a leading edge and a trailing edge disposed at an aft location of the inner shroud component relative to the leading edge. Also included is a base portion extending from the leading edge to the trailing edge, wherein the base portion is in direct contact with the hot gas path.
  • a rib extending from a first side portion to a second side portion and radially outward from the base portion, wherein the rib includes a main passage extending between the first side portion and the second side portion and configured to receive a cooling flow from a cooling source.
  • a gas turbine system includes a compressor for distributing a cooling flow at a high pressure. Also included is a turbine casing operably supporting a turbine shroud assembly for receiving the cooling flow for cooling therein. Further included is an inner shroud component comprising a leading edge, a trailing edge spaced axially rearward of the leading edge, and a base portion connecting the leading edge to the trailing edge. Yet further included is a rib disposed between the leading edge and the trailing edge, and extending between a first side portion and a second side portion, wherein the rib includes a main passage configured to receive the cooling flow for cooling the inner shroud component.
  • the gas turbine system 10 includes a compressor 12, a combustor 14, a turbine 16, a shaft 18 and a fuel nozzle 20. It is to be appreciated that one embodiment of the gas turbine system 10 may include a plurality of compressors 12, combustors 14, turbines 16, shafts 18 and fuel nozzles 20. The compressor 12 and the turbine 16 are coupled by the shaft 18.
  • the shaft 18 may be a single shaft or a plurality of shaft segments coupled together to form the shaft 18.
  • the combustor 14 uses a combustible liquid and/or gas fuel, such as natural gas or a hydrogen rich synthetic gas, to run the gas turbine system 10.
  • fuel nozzles 20 are in fluid communication with an air supply and a fuel supply 22.
  • the fuel nozzles 20 create an air-fuel mixture, and discharge the air-fuel mixture into the combustor 14, thereby causing a combustion that creates a hot pressurized exhaust gas.
  • the combustor 14 directs the hot pressurized gas through a transition piece into a turbine nozzle (or "stage one nozzle"), and other stages of buckets and nozzles causing rotation of the turbine 16 within a turbine casing 24.
  • hot gas path components are located in the turbine 16, where hot gas flow across the components causes creep, oxidation, wear and thermal fatigue of turbine components. Controlling the temperature of the hot gas path components can reduce distress modes in the components and the efficiency of the gas turbine system 10 increases with an increase in firing temperature. As the firing temperature increases, the hot gas path components need to be properly cooled to meet service life and to effectively perform intended functionality.
  • a shroud assembly is an example of a component disposed in the turbine 16 proximate the turbine casing 24 and subjected to the hot gas path described in detail above, the hot gas path referred to with numeral 32.
  • the turbine shroud cooling assembly 30 includes an inner shroud component 34 with an inner surface 36 proximate the hot gas path 32 within the turbine 16.
  • the turbine shroud cooling assembly 30 also includes an outer shroud component (not illustrated) that is generally proximate to a relatively cool fluid and/or air in the turbine 16, with the inner shroud component 34 being operably coupled to the outer shroud component.
  • a cooling flow 38 supplied by a cooling source is introduced into the outer shroud component and directed toward the inner shroud component 34.
  • a plenum within the outer shroud component may be present to ingest and direct the cooling flow 38 toward the inner shroud component 34.
  • the inner shroud component 34 includes a base portion 40 having an outer surface 42, as well as the inner surface 36 that is directly exposed to the hot gas path 32, as described above.
  • the base portion 40 typically arcuately extends between a leading edge 44 and a trailing edge 46 of the inner shroud component 34.
  • Both the leading edge 44 and the trailing edge 46 include at least one fastening device 48, such as a rail or clip for example, that operably couples the inner shroud component 34 with the outer shroud component.
  • the inner shroud component 34 also includes a first side portion 50 and a second side portion 52 extending along the base portion 40 between, and connected to, the leading edge 44 and the trailing edge 46.
  • the outer surface 42 of the base portion 40 combines with the outer shroud component to form at least one cavity 54, such as an impingement cavity, into which the cooling flow 38 is directed toward and into.
  • a first side portion passage 60 is disposed proximate the first side portion 50 and a second side portion passage 62 is disposed proximate the second side portion 52. Additionally, a fore passage 64 and an aft passage 68 may be included at locations proximate the leading edge 44 and the trailing edge 46, respectively. Numerous other internal passages may be provided in addition to, or alternatively to, the internal passages described above. In the illustrated embodiment, the first side portion passage 60, the second side portion passage 62, the fore passage 64 and the aft passage 68 are disposed proximate the perimeter of the inner shroud component 34.
  • a rib 70 integrally formed with the base portion 40 protrudes radially away from the remainder of the outer surface 42 of the base portion 40 and extends between the first side portion 50 and the second side portion 52. It is to be appreciated that in other embodiments, the rib 70 may extend at various angles across the base portion 40, including relatively perpendicular to that illustrated, where the rib 70 extends from proximate the leading edge 44 to the trailing edge 46. Irrespective of the precise location and orientation of the rib 70, in order to effectively and efficiently cool portions of the inner shroud component 34 other than those proximate the perimeter, a main passage 72 is formed within the rib 70.
  • the main passage 72 extends between, and connects with, the first side portion passage 60 and the second side portion passage 62, thereby allowing the cooling flow 38 to be transferred through the main passage 72, the first side portion passage 60 and the second side portion passage 62, in any direction.
  • the fore passage 64 and the aft passage 68 extend between, and connect to, the first side portion passage 60 and the second side portion passage 62, thereby forming a continuous, interconnected cooling flow circuit 74. It is to be appreciated that a discontinuous circuit may be formed by including one or more breaks in any of the passages, including the main passage 72, the first side portion passage 60, the second side portion passage 62, the fore passage 64 and/or the aft passage 68.
  • Cooling of the inner shroud component 34 is achieved by ingesting an airstream of the cooling flow 38 from a cooling source (not illustrated) that provides the cooling flow 38, which may include air, a water solution and/or a gas.
  • the cooling flow 38 is any suitable fluid that cools the inner shroud component 34.
  • the cooling source is a supply of compressed air from the compressor 12, where the compressed air is diverted from the air supply that is routed to the combustor 14.
  • the supply of compressed air bypasses the combustor 14 and is used to cool the turbine shroud cooling assembly 30.
  • the inner shroud component 34 receives the cooling flow 38 at the at least one cavity 54 and introduces the cooling flow 38 into at least one of the first side portion passage 60, the second side portion passage 62, the fore passage 64 and the aft passage 68.
  • Such an arrangement allows the cooling flow 38 to be transferred to the main passage 72 for cooling therein.
  • the main passage 72 may be the sole, or an additional, ingestion point for the cooling flow 38 into the internal passages.
  • the main passage 72 may include at least one, but typically a plurality of channels 76 formed in the rib 70 to fluidly connect the at least one cavity 54 and the main passage 72.
  • the plurality of channels 76 may be drilled or formed in any suitable manner.
  • One or more exit paths for the cooling flow 38 may be formed throughout one or more portions of the inner shroud component 34 to allow dumping of the cooling flow 38 to external regions, such as the hot gas path 32.
  • One contemplated location of the exit paths is through the inner surface 36 of the inner shroud component 34.
  • the main passage 72 within the rib 70 allows the cooling flow 38 to flow through the rib 70 that is disposed away from the perimeter of the inner shroud component 34, thereby leading to improved cooling of the overall inner shroud component 34.
  • Such a feature ultimately decreases the high temperatures of various regions of the inner shroud component 34, including an aft edge of the rib 70.
  • Overall gas turbine system 10 efficiency is improved based on the reduction of the cooling flow 38 that is required to effectively cool the inner shroud component 34. Additionally, service life of the inner shroud component 34 is increased due to the lower temperature experienced during exposure to the hot gas path 32.

Landscapes

  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
EP13168436.7A 2012-05-25 2013-05-20 Turbinenummantelungskühlanordnung für ein Gasturbinensystem und zugehöriges Gasturbinensystem Withdrawn EP2666972A2 (de)

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
US13/480,906 US20130315719A1 (en) 2012-05-25 2012-05-25 Turbine Shroud Cooling Assembly for a Gas Turbine System

Publications (1)

Publication Number Publication Date
EP2666972A2 true EP2666972A2 (de) 2013-11-27

Family

ID=48428400

Family Applications (1)

Application Number Title Priority Date Filing Date
EP13168436.7A Withdrawn EP2666972A2 (de) 2012-05-25 2013-05-20 Turbinenummantelungskühlanordnung für ein Gasturbinensystem und zugehöriges Gasturbinensystem

Country Status (5)

Country Link
US (1) US20130315719A1 (de)
EP (1) EP2666972A2 (de)
JP (1) JP2013245676A (de)
CN (1) CN103422917A (de)
RU (1) RU2013123451A (de)

Cited By (1)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP3246523A1 (de) * 2016-05-19 2017-11-22 United Technologies Corporation Gekühlte schaufel-aussenluftdichtung

Families Citing this family (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US20150198063A1 (en) * 2014-01-14 2015-07-16 Alstom Technology Ltd Cooled stator heat shield
US20180223681A1 (en) * 2017-02-09 2018-08-09 General Electric Company Turbine engine shroud with near wall cooling
US10989068B2 (en) 2018-07-19 2021-04-27 General Electric Company Turbine shroud including plurality of cooling passages
US10837315B2 (en) * 2018-10-25 2020-11-17 General Electric Company Turbine shroud including cooling passages in communication with collection plenums

Family Cites Families (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US7306424B2 (en) * 2004-12-29 2007-12-11 United Technologies Corporation Blade outer seal with micro axial flow cooling system
US8123466B2 (en) * 2007-03-01 2012-02-28 United Technologies Corporation Blade outer air seal
CH700686A1 (de) * 2009-03-30 2010-09-30 Alstom Technology Ltd Schaufel für eine gasturbine.
US8371800B2 (en) * 2010-03-03 2013-02-12 General Electric Company Cooling gas turbine components with seal slot channels

Non-Patent Citations (1)

* Cited by examiner, † Cited by third party
Title
None

Cited By (2)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP3246523A1 (de) * 2016-05-19 2017-11-22 United Technologies Corporation Gekühlte schaufel-aussenluftdichtung
US10344611B2 (en) 2016-05-19 2019-07-09 United Technologies Corporation Cooled hot section components for a gas turbine engine

Also Published As

Publication number Publication date
JP2013245676A (ja) 2013-12-09
CN103422917A (zh) 2013-12-04
RU2013123451A (ru) 2014-11-27
US20130315719A1 (en) 2013-11-28

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