US20150013345A1 - Gas turbine shroud cooling - Google Patents

Gas turbine shroud cooling Download PDF

Info

Publication number
US20150013345A1
US20150013345A1 US13/939,727 US201313939727A US2015013345A1 US 20150013345 A1 US20150013345 A1 US 20150013345A1 US 201313939727 A US201313939727 A US 201313939727A US 2015013345 A1 US2015013345 A1 US 2015013345A1
Authority
US
United States
Prior art keywords
portion along
edge
gas turbine
critical process
casing
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.)
Abandoned
Application number
US13/939,727
Other languages
English (en)
Inventor
Christopher Donald Porter
Gregory Thomas Foster
Aaron Ezekiel Smith
David Wayne Weber
Michelle J. Rogers
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
Priority to US13/939,727 priority Critical patent/US20150013345A1/en
Assigned to GENERAL ELECTRIC COMPANY reassignment GENERAL ELECTRIC COMPANY ASSIGNMENT OF ASSIGNORS INTEREST (SEE DOCUMENT FOR DETAILS). Assignors: ROGERS, MICHELLE J., FOSTER, GREGORY THOMAS, Porter, Christopher Donald, Smith, Aaron Ezekiel, Weber, David Wayne
Priority to DE102014109288.0A priority patent/DE102014109288A1/de
Priority to JP2014138206A priority patent/JP2015017608A/ja
Priority to CH01042/14A priority patent/CH708326A2/de
Priority to CN201420385965.2U priority patent/CN204253116U/zh
Publication of US20150013345A1 publication Critical patent/US20150013345A1/en
Abandoned legal-status Critical Current

Links

Images

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
    • 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
    • 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
    • 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
    • F05D2260/201—Heat transfer, e.g. cooling by impingement of a fluid
    • 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
    • F05D2260/202—Heat transfer, e.g. cooling by film cooling
    • Y—GENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
    • Y02—TECHNOLOGIES OR APPLICATIONS FOR MITIGATION OR ADAPTATION AGAINST CLIMATE CHANGE
    • Y02T—CLIMATE CHANGE MITIGATION TECHNOLOGIES RELATED TO TRANSPORTATION
    • Y02T50/00—Aeronautics or air transport
    • Y02T50/60—Efficient propulsion technologies, e.g. for aircraft

Definitions

  • the present invention generally involves cooling a turbine shroud element that may be located in a hot gas path of the turbine.
  • Turbines are widely used in a variety of aviation, industrial, and power generation applications to perform work.
  • Each turbine generally includes alternating stages of peripherally mounted stator vanes and rotating blades.
  • the stator vanes may be attached to a stationary component such as a casing that surrounds the turbine, and the rotating blades may be attached to a rotor located along an axial centerline of the turbine.
  • a compressed working fluid such as steam, combustion gases, or air, flows along a gas path through the turbine to produce work.
  • the stator vanes accelerate and direct the compressed working fluid onto the subsequent stage of rotating blades to impart motion to the rotating blades, thus turning the rotor and performing work. If any compressed working fluid moves radially outside of the desired flow path, the efficiency of the turbine may be reduces.
  • the casing surrounding the turbine often includes a radially inner shell of shrouds, often formed in segments.
  • the shrouds surround and define the outer perimeter of the hot gas path and may be located around both stator vanes and rotating blades.
  • the turbine shrouds are typically cooled in some fashion to remove heat transferred by the hot gas path.
  • U.S. Pat. No. 7,284,954 describes a turbine shroud segment that includes many small cooling fluid passages machined throughout the turbine shroud. A fluid such as compressed air from an upstream compressor may be supplied through the fluid passages to cool the turbine shroud.
  • Other shroud segments utilize a single larger “core” flow path cast in place rather than multiple small machined passages as above. The core extends along an entire side of the shroud segment from an axially upstream end to an axially downstream end.
  • a shroud segment for a casing of gas turbine may include a body configured for attachment to the casing proximate a localized critical process location within the casing.
  • the body has a leading edge, a trailing edge, and two side edges, as well as a first surface for facing the casing and a second surface opposite the first surface for facing a hot gas path.
  • the critical process location is located between the leading edge and the trailing edge when the body is attached to the casing.
  • At least two cooling passages are defined in the body along one of the side edges.
  • a first of the cooling passages has an inlet and extends to an outlet, one of the inlet or outlet being adjacent the critical process location.
  • a second of the cooling passages has an inlet and extends to an outlet, one of the inlet or the outlet being adjacent the critical process location.
  • the first and second cooling passages are configured large enough to cool the one side edge to a desired level during operation of the gas turbine.
  • a gas turbine may include a compressor section, a combustion section downstream from the compressor section, and a turbine section downstream from the combustion section.
  • the turbine section includes a casing defining a localized critical process location and a plurality of shroud segments circumferentially attached to the casing.
  • Each shroud segment includes a body configured for attachment to the casing.
  • At least one of the bodies has a leading edge, a trailing edge, and two side edges, as well as a first surface facing the casing and a second surface opposite the first surface facing a hot gas path.
  • the critical process location is located between the leading edge and the trailing edge when the body is attached to the casing.
  • At least two cooling passages are defined in the body along one of the side edges.
  • a first of the cooling passages has an inlet and extends to an outlet, one of the inlet or outlet being adjacent the critical process location.
  • a second of the cooling passages has an inlet and extends to an outlet, one of the inlet or the outlet being adjacent the critical process location.
  • the first and second cooling passages are configured large enough to cool the one side edge to a desired level during operation of the gas turbine. As above, various options and modifications are possible.
  • FIG. 1 is a schematic view of an exemplary gas turbine incorporating aspects of the present disclosure
  • FIG. 2 is a simplified cross-section view of a portion of the gas turbine of FIG. 1 showing a shroud segment;
  • FIG. 3 is a top view of a shroud segment as in FIG. 2 ;
  • FIG. 4 is a side view of the shroud segment of FIG. 3 ;
  • FIG. 5 is a sectional view of the shroud segment taken along line 5 - 5 in FIG. 3 ;
  • FIG. 6 is an isometric view of the shroud segment of FIG. 3 ;
  • FIG. 7 is a top view of a first alternate shroud segment
  • FIG. 8 is a top view of a second alternate shroud segment.
  • FIG. 9 is a top view of a third alternate shroud segment.
  • FIG. 1 is a schematic view of an exemplary gas turbine that can incorporate a shroud element according to the present disclosure.
  • gas turbine 110 includes an inlet section 111 , a compressor section 112 , a combustion section 114 , a turbine section 116 , and an exhaust section 117 .
  • a shaft (rotor) 122 may be common to compressor section 112 and turbine section 116 , and may further connect to a generator 105 for generating electricity.
  • the compressor section 112 may include an axial flow compressor in which a working fluid 100 , such as ambient air, enters the compressor from the inlet section 111 and passes through alternating stages 113 of stationary vanes and rotating blades (shown schematically in FIG. 1 ).
  • Compressor casing 118 contains the working fluid 100 as the stationary vanes and rotating blades accelerate and redirect the working fluid to produce a continuous flow of compressed working fluid. The majority of the compressed working fluid flows downstream through the combustion section 114 and then the turbine section 116 .
  • the combustion section 114 may include any type of combustor known in the art.
  • a combustor casing 115 may circumferentially surround some or all of the combustion section 114 to direct the compressed working fluid 100 from the compressor section 112 to a combustion chamber 119 .
  • Fuel 101 is also supplied to the combustion chamber 119 .
  • Possible fuels include, for example, one or more of blast furnace gas, coke oven gas, natural gas, vaporized liquefied natural gas (LNG), hydrogen, and propane.
  • the compressed working fluid 100 mixes with fuel 101 in the combustion chamber 119 where it ignites to generate combustion gases having a high temperature and pressure. The combustion gases then enter the turbine section 116 .
  • alternating stages of rotating blades (buckets) 124 and stationary blades (nozzles) 126 are attached to rotor 122 and turbine casing 120 , respectively.
  • Working fluid 100 such as steam, combustion gases, or air, flows along a hot gas path through gas turbine 110 from left to right as shown in FIG. 2 .
  • the first stage of stationary nozzles 126 accelerates and directs the working fluid 100 onto the first stage of rotating blades 124 , causing the first stage of rotating blades 124 and rotor 122 to rotate.
  • Working fluid 100 then flows across the second stage of stationary nozzles 126 which accelerates and redirects the working fluid to the next stage of rotating blades (see FIG. 1 ), and the process repeats for each subsequent stage.
  • the radially inward portion of turbine casing 120 may include a series of shrouds 128 .
  • Shrouds 128 in FIG. 1 are formed around blades 124 .
  • FIG. 2 shows shrouds 128 formed around both blades 124 and nozzles 126 .
  • Shrouds 128 may be formed in segments, such as segment 130 of FIGS. 2-6 . It should be understood that, although an example of a shroud segment related to a blade 124 is shown, the present disclosure incorporates shroud segments formed around nozzles 124 as well. Therefore, no limitation as to location of shrouds within casing 120 should be made.
  • each shroud segment 130 may generally comprise a body having a plurality of sides. Specifically, each segment 130 has a leading edge 132 , a trailing edge 134 , and two side edges 136 and 138 . A first surface 140 faces (radially outwardly) toward casing 120 and a second surface 142 opposite the first surface faces (radially inwardly) toward the hot gas path where the working fluid 100 flows.
  • a critical process location (defined below) 144 is located between leading edge 132 and the trailing edge 134 , generally in alignment with rotating blades 124 .
  • the critical process location 144 could be, for example, a maximum or other critical temperature location along the segment during gas turbine use, a maximum or other critical pressure location along the segment during gas turbine use, a maximum or other critical gas side heat transfer coefficient location, or a maximum or other critical stress location.
  • the critical process location 144 could be a location where cooling gases can enter or exit the segment after travelling through a passageway allowing for sufficient cooling of the segment, while still respecting back flow margin limitations.
  • the critical process location need not be an absolute maximum, it could be any desired value that can be used to determine optimal flow and heat transfer characteristics within the gas turbine or within the segment itself. Much depends on the desired characteristics of the gas turbine, flow at the location of segment 130 , etc.
  • the critical process location along segment 130 could vary at different stages within a gas turbine. Further, two or more of such critical process locations could exist along a single segment 130 .
  • At least two cooling passages 146 , 148 are defined in segment 130 along one of side edges 136 .
  • First cooling passage 146 has an inlet 150 that may be (as shown) on first surface 140 near leading edge 132 .
  • First cooling passage 146 also has an outlet 152 adjacent critical process location 144 .
  • Second cooling passage 148 has an inlet 154 that may be (as shown) on first surface 140 adjacent critical process location 144 .
  • Second cooling passage 148 has at least one outlet 156 that may be (as shown) near trailing edge 134 .
  • passages 146 , 148 could be reverse of that which is shown.
  • flow in first passage 146 could be counter (directed upstream) to that through second passage 148 .
  • flow could run from opening 152 to opening 150 (reversing inlet/outlet functions), if desired.
  • Flow through second passage 148 could also be similarly reversed.
  • First and second cooling passages 146 , 148 may be formed by casting rather than machining.
  • a mold may be used in which a fill substance is provided matching the path of first and second cooling passages 146 , 148 , the fill substance being burned off and/or chemically removed afterward leaving the passages.
  • Such manufacture using casting of at least some portion of the passages may be more cost effective than machining the passages or multiple smaller passages. Even if the passages are formed substantially by casting, inlets and outlets to the passages or other features may be machined as part of the manufacture.
  • First and second cooling passages 146 , 148 may be configured large enough to cool side edge 136 and/or a related area to a desired level during operation of the gas turbine.
  • the passage sizes are configured to allow sufficient flow that back flow margins are respected, and heat transfer is sufficient to cool segment 130 to a desired temperature.
  • a segment 130 with a length of about 6.5 inch, a width of about 3.0 inch, and general thickness of about 0.25 inch, passages 146 , 148 may be of a cross-section of about 0.025 square inch. Accordingly, numerous small passages spread along the locations of cooling passages 146 , 148 are not required to cool segment 130 .
  • an additional set of cooling passages 158 , 160 can be provided on other side edge 138 .
  • Passages 158 , 160 may if desired but not necessarily be substantially symmetrical to passages 146 , 148 along a central axis running between leading edge 132 and trailing edge 134 .
  • first cooling passage 158 has an inlet 162 which may be (as shown) on first surface 140 near leading edge 132 .
  • First cooling passage 158 also has an outlet 164 adjacent critical process location 144 .
  • Second cooling passage 160 has an inlet 166 which may be (as shown) on first surface 140 adjacent critical process location 144 .
  • Second cooling passage 160 has at least one outlet 168 which may be (as shown) near trailing edge 134 .
  • inlet 150 and inlet 162 are a common, single inlet. However, as discussed below, the inlets 150 , 162 may be separate.
  • both second passages 148 , 160 may have multiple outlets 156 , 168 , which may be along trailing edge 134 .
  • Such multiple exits may be machined or cast, and may be employed to cool trailing edge 134 if spaced sufficiently from second passages 148 , 160 to require additional cooling.
  • Some or all of such multiple outlets could instead or also exit segment 130 at locations other than trailing edge 134 if desired.
  • modified segment 130 ′ has first passages 146 ′, 158 ′ each with their own individual inlets 150 , 162 with first portions 170 , 172 and second portions 174 , 176 leading to outlets 152 , 164 .
  • first portions 170 , 172 are in communication with each other; as shown. If desired, some or all of the inlets in segment 130 or 130 ′ could also be located elsewhere other than first surface 140 .
  • modified segment 130 ′′ has second passages 148 ′, 160 ′ each with individual inlets 154 , 166 leading to first portions 178 , 180 and second portions 182 , 184 and then outlet(s) 156 , 168 , as above.
  • second portions 182 , 184 in FIG. 8 are in communication with each other. Therefore, instead of the construction shown in FIG. 3 , having one first passage upstream of location 144 and two second passages downstream of location 144 , a shroud segment could be made as shown in FIG. 8 with one upstream passage and one downstream passage split along the side edges 136 , 138 at location 144 .
  • splits could be provided at two or more critical process locations along the shroud segment.
  • a first passageway 146 extends from inlet 150 to outlet 152
  • a second passageway 148 ′ extends from inlet 154 to outlet 153
  • a third passageway 179 extends from inlet 155 to outlets 156 .
  • a first passageway 158 extends from inlet 162 to outlet 164
  • a second passageway 160 ′ extends from inlet 166 to outlet 165
  • a third passageway 181 extends from inlet 167 to outlets 168 .
  • FIG. 9 illustrates that more than one split can be made between leading edge 132 and trailing edge 134 at critical process locations, as desired. It should also be understood that splits need not be symmetrical or even along a given side of the segments or between sides of the segments.
  • the segments above can be mounted to turbine casings in various known ways, via hooks, impingement plates, clips, etc.
  • the present invention is not limited to any such mounting arrangement, cooling mode, or any particular fluid used to cool the shroud segment.
  • such mounting may or may not provide that the cooling fluid first impacts the segments to provide impingement cooling to the bulk of the segment before some fluid flows through the disclosed passageways.
  • the segments may include mounting structures, cooling passage openings, etc., for receiving, contacting or cooling nozzles 126 if the segments are located along a row of nozzles as opposed to a row of blades 124 .
  • the various embodiments of the shroud segments shown above may be manufactured at lower costs than previous designs.
  • the segments may be cast or forged, with reduced machining required for inlets and outlets and the larger passages being formed by casting.
  • the shroud may be readily manufactured to include the desired fluid passages that provide cooling to the sides of the segments.
  • cooling can be beneficially located at a desired point while providing a more efficient flow, with less leakage.
  • the segments can thus be tuned in various ways to improve thermal and flow performance.

Landscapes

  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
US13/939,727 2013-07-11 2013-07-11 Gas turbine shroud cooling Abandoned US20150013345A1 (en)

Priority Applications (5)

Application Number Priority Date Filing Date Title
US13/939,727 US20150013345A1 (en) 2013-07-11 2013-07-11 Gas turbine shroud cooling
DE102014109288.0A DE102014109288A1 (de) 2013-07-11 2014-07-02 Gasturbinen-Deckbandkühlung
JP2014138206A JP2015017608A (ja) 2013-07-11 2014-07-04 ガスタービン・シュラウド冷却
CH01042/14A CH708326A2 (de) 2013-07-11 2014-07-09 Gasturbinen-Deckbandkühlung.
CN201420385965.2U CN204253116U (zh) 2013-07-11 2014-07-11 用于燃气涡轮机外壳的防护罩节段

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
US13/939,727 US20150013345A1 (en) 2013-07-11 2013-07-11 Gas turbine shroud cooling

Publications (1)

Publication Number Publication Date
US20150013345A1 true US20150013345A1 (en) 2015-01-15

Family

ID=52107479

Family Applications (1)

Application Number Title Priority Date Filing Date
US13/939,727 Abandoned US20150013345A1 (en) 2013-07-11 2013-07-11 Gas turbine shroud cooling

Country Status (5)

Country Link
US (1) US20150013345A1 (de)
JP (1) JP2015017608A (de)
CN (1) CN204253116U (de)
CH (1) CH708326A2 (de)
DE (1) DE102014109288A1 (de)

Cited By (12)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CN107178397A (zh) * 2016-03-10 2017-09-19 通用电气公司 用于冷却热气体流路构件的后缘和/或前缘的系统和方法
US10335900B2 (en) 2016-03-03 2019-07-02 General Electric Company Protective shield for liquid guided laser cutting tools
US10337411B2 (en) 2015-12-30 2019-07-02 General Electric Company Auto thermal valve (ATV) for dual mode passive cooling flow modulation
US10337739B2 (en) 2016-08-16 2019-07-02 General Electric Company Combustion bypass passive valve system for a gas turbine
US20190368377A1 (en) * 2018-05-31 2019-12-05 General Electric Company Shroud for gas turbine engine
US10677084B2 (en) 2017-06-16 2020-06-09 Honeywell International Inc. Turbine tip shroud assembly with plural shroud segments having inter-segment seal arrangement
US10712007B2 (en) 2017-01-27 2020-07-14 General Electric Company Pneumatically-actuated fuel nozzle air flow modulator
US10738712B2 (en) 2017-01-27 2020-08-11 General Electric Company Pneumatically-actuated bypass valve
US10900378B2 (en) 2017-06-16 2021-01-26 Honeywell International Inc. Turbine tip shroud assembly with plural shroud segments having internal cooling passages
US10961864B2 (en) 2015-12-30 2021-03-30 General Electric Company Passive flow modulation of cooling flow into a cavity
US20220275734A1 (en) * 2021-01-22 2022-09-01 Mitsubishi Heavy Industries, Ltd. Flow channel forming plate, blade and gas turbine including this, and method of manufacturing flow channel forming plate
US11746663B2 (en) 2019-03-29 2023-09-05 Mitsubishi Power, Ltd. High-temperature component and method of producing the high-temperature component

Citations (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US6264426B1 (en) * 1997-02-20 2001-07-24 Mitsubishi Heavy Industries, Ltd. Gas turbine stationary blade
US6761529B2 (en) * 2002-07-25 2004-07-13 Mitshubishi Heavy Industries, Ltd. Cooling structure of stationary blade, and gas turbine
US20090304520A1 (en) * 2006-06-07 2009-12-10 General Electric Company Serpentine cooling circuit and method for cooling tip shroud
US8596962B1 (en) * 2011-03-21 2013-12-03 Florida Turbine Technologies, Inc. BOAS segment for a turbine

Family Cites Families (1)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US7284954B2 (en) 2005-02-17 2007-10-23 Parker David G Shroud block with enhanced cooling

Patent Citations (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US6264426B1 (en) * 1997-02-20 2001-07-24 Mitsubishi Heavy Industries, Ltd. Gas turbine stationary blade
US6761529B2 (en) * 2002-07-25 2004-07-13 Mitshubishi Heavy Industries, Ltd. Cooling structure of stationary blade, and gas turbine
US20090304520A1 (en) * 2006-06-07 2009-12-10 General Electric Company Serpentine cooling circuit and method for cooling tip shroud
US8596962B1 (en) * 2011-03-21 2013-12-03 Florida Turbine Technologies, Inc. BOAS segment for a turbine

Cited By (15)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US10337411B2 (en) 2015-12-30 2019-07-02 General Electric Company Auto thermal valve (ATV) for dual mode passive cooling flow modulation
US10961864B2 (en) 2015-12-30 2021-03-30 General Electric Company Passive flow modulation of cooling flow into a cavity
US10335900B2 (en) 2016-03-03 2019-07-02 General Electric Company Protective shield for liquid guided laser cutting tools
EP3228821A1 (de) * 2016-03-10 2017-10-11 General Electric Company System und verfahren zur kühlung der abströmkante und/oder der vorderkante einer heissgaspfadkomponente
CN107178397A (zh) * 2016-03-10 2017-09-19 通用电气公司 用于冷却热气体流路构件的后缘和/或前缘的系统和方法
US10337739B2 (en) 2016-08-16 2019-07-02 General Electric Company Combustion bypass passive valve system for a gas turbine
US10738712B2 (en) 2017-01-27 2020-08-11 General Electric Company Pneumatically-actuated bypass valve
US10712007B2 (en) 2017-01-27 2020-07-14 General Electric Company Pneumatically-actuated fuel nozzle air flow modulator
US10677084B2 (en) 2017-06-16 2020-06-09 Honeywell International Inc. Turbine tip shroud assembly with plural shroud segments having inter-segment seal arrangement
US10900378B2 (en) 2017-06-16 2021-01-26 Honeywell International Inc. Turbine tip shroud assembly with plural shroud segments having internal cooling passages
US11181006B2 (en) 2017-06-16 2021-11-23 Honeywell International Inc. Turbine tip shroud assembly with plural shroud segments having inter-segment seal arrangement
US20190368377A1 (en) * 2018-05-31 2019-12-05 General Electric Company Shroud for gas turbine engine
US10989070B2 (en) * 2018-05-31 2021-04-27 General Electric Company Shroud for gas turbine engine
US11746663B2 (en) 2019-03-29 2023-09-05 Mitsubishi Power, Ltd. High-temperature component and method of producing the high-temperature component
US20220275734A1 (en) * 2021-01-22 2022-09-01 Mitsubishi Heavy Industries, Ltd. Flow channel forming plate, blade and gas turbine including this, and method of manufacturing flow channel forming plate

Also Published As

Publication number Publication date
CH708326A2 (de) 2015-01-15
CN204253116U (zh) 2015-04-08
DE102014109288A1 (de) 2015-01-15
JP2015017608A (ja) 2015-01-29

Similar Documents

Publication Publication Date Title
US20150013345A1 (en) Gas turbine shroud cooling
US7284954B2 (en) Shroud block with enhanced cooling
US9464538B2 (en) Shroud block segment for a gas turbine
EP2388435B1 (de) Turbinenlaufschaufel
US10184342B2 (en) System for cooling seal rails of tip shroud of turbine blade
US9080451B2 (en) Airfoil
US9382810B2 (en) Closed loop cooling system for a gas turbine
JP2015092076A (ja) タービンアセンブリに冷却を提供するための方法およびシステム
EP3203024B1 (de) Laufschaufel und zugehörige gasturbine
US7011492B2 (en) Turbine vane cooled by a reduced cooling air leak
EP3156609B1 (de) Turbinenleitschaufel mit auslassplenum für kühlluft.
US9528380B2 (en) Turbine bucket and method for cooling a turbine bucket of a gas turbine engine
US20160186574A1 (en) Interior cooling channels in turbine blades
US10385727B2 (en) Turbine nozzle with cooling channel coolant distribution plenum
JPWO2017158637A1 (ja) タービン及びタービン静翼
EP3412869B1 (de) Turbomaschinenrotorschaufel
EP3252270B1 (de) Düsenkühlsystem für einen gasturbinenmotor
US10590777B2 (en) Turbomachine rotor blade
US10472974B2 (en) Turbomachine rotor blade
US10494932B2 (en) Turbomachine rotor blade cooling passage
US9039370B2 (en) Turbine nozzle
GB2584299A (en) Heatshield for gas turbine engine
US20140050558A1 (en) Temperature gradient management arrangement for a turbine system and method of managing a temperature gradient of a turbine system

Legal Events

Date Code Title Description
AS Assignment

Owner name: GENERAL ELECTRIC COMPANY, NEW YORK

Free format text: ASSIGNMENT OF ASSIGNORS INTEREST;ASSIGNORS:PORTER, CHRISTOPHER DONALD;FOSTER, GREGORY THOMAS;SMITH, AARON EZEKIEL;AND OTHERS;SIGNING DATES FROM 20130702 TO 20130709;REEL/FRAME:030779/0836

STCB Information on status: application discontinuation

Free format text: ABANDONED -- FAILURE TO RESPOND TO AN OFFICE ACTION