EP1124039A1 - Dispositif de refroidissement par impact pour une bande de protection de turbine à gaz - Google Patents

Dispositif de refroidissement par impact pour une bande de protection de turbine à gaz Download PDF

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Publication number
EP1124039A1
EP1124039A1 EP00301015A EP00301015A EP1124039A1 EP 1124039 A1 EP1124039 A1 EP 1124039A1 EP 00301015 A EP00301015 A EP 00301015A EP 00301015 A EP00301015 A EP 00301015A EP 1124039 A1 EP1124039 A1 EP 1124039A1
Authority
EP
European Patent Office
Prior art keywords
shroud
inner shroud
wall
segment
cooling air
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Withdrawn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
EP00301015A
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German (de)
English (en)
Inventor
Steven Sebastian Burdgick
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 EP00301015A priority Critical patent/EP1124039A1/fr
Publication of EP1124039A1 publication Critical patent/EP1124039A1/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Withdrawn legal-status Critical Current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
    • 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
    • F01D11/10—Preventing or minimising internal leakage of working-fluid, e.g. between stages for sealing space between rotor blade tips and stator using sealing fluid, e.g. steam
    • 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
    • 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
    • F05D2260/00—Function
    • F05D2260/20—Heat transfer, e.g. cooling
    • F05D2260/201—Heat transfer, e.g. cooling by impingement of a fluid

Definitions

  • the present invention relates to impingement cooling apparatus for a shroud system surrounding the rotating components in the hot gas path of a gas turbine and particularly relates to inner and outer shroud segments employing a feed of cooling air directly into the inner shroud body for impingement cooling of the inner shroud wall surface opposite the wall surface surrounding the hot gas path.
  • Shrouds employed in gas turbines surround and in part define the hot gas path through the turbines.
  • Systems for cooling the shrouds, particularly those directly surrounding the rotating parts, i.e., the gas turbine buckets or blades, in the hot gas path of the gas turbine are oftentimes necessary in gas turbines to reduce the temperature of the surrounding shrouds.
  • Shrouds are typically characterized by a plurality of circumferentially extending shroud segments arranged about the hot gas path with each segment including discrete inner and outer shroud bodies.
  • the inner shroud body includes a wall which in part defines the hot gas path and which must be cooled, for example, with cooling air from the compressor discharge of the turbine.
  • an impingement plate has been provided in the outer shroud body for receiving the cooling air and directing the cooling air through apertures in the plate for impingement cooling of the inner shroud body wall. This arrangement is not optimum from the standpoint of efficient cooling and requires substantial cooling flow. More particularly, the impingement plate mounted on the outer shroud body in this conventional design is spaced a substantial distance from the wall being cooled by the impingement air flow through the apertures of the plate.
  • the inner shroud body has axially extending reinforcing or structural ribs projecting radially outwardly from the wall being cooled, previously believed to necessitate the location of the impingement plate mounted to the outer shroud a substantial distance from that wall.
  • cooling efficiency is lost as the impingement cooling air flows over this very substantial distance before impacting and cooling the inner shroud wall.
  • the impingement cooling air sees secondary leakage paths prior to passing through the impingement plate apertures, which causes further inefficiencies in cooling and requires additional cooling flow.
  • an impingement cooling system which will substantially reduce these cooling inefficiencies, eliminate leakage paths and substantially reduce the impingement flow distance between the impingement plate and the inner shroud body wall being cooled by the impingement cooling air flow.
  • an impingement cooling apparatus for a shroud system surrounding rotating components in the hot gas path of a turbine and which system employs a plurality of shroud segments each comprising an outer shroud segment and one or more inner shroud segments secured to the outer shroud segment.
  • the inner shroud segment mounts an impingement plate in a manner which eliminates leakage paths between the outer shroud segment and impingement plate and locates the impingement plate directly adjacent the inner shroud segment wall being cooled by the impingement air flow, thereby affording efficient impingement cooling.
  • the inner shroud segment includes an inner shroud segment body having a bottom wall, the radially innermost surface of which in part defines the hot gas path through the turbine.
  • One or more cavities are provided in the inner shroud body on a side thereof remote from the wall surface defining the hot gas path.
  • the inner shroud segment also includes a cover for overlying the inner shroud body.
  • the cover has one or more depending closed compartments for reception in the respective cavities of the plate.
  • the cover is secured to the inner shroud body by welding, brazing or the like, with the one or more compartments lying in respective cavities.
  • An air inlet opens through the cover in communication with an air inlet passageway through the body of the outer shroud segment for supplying cooling air to the compartments.
  • each compartment has a plurality of apertures for flowing cooling air received in the compartment directly onto and hence impingement cooling the bottom walls or floors of the cavities defining in part the hot gas path.
  • Passages through the inner shroud body lie in communication with the space between the compartments and the cavities for exhausting the spent cooling flow into the hot gas stream.
  • the cavities are defined by radially outwardly projecting structural ribs which extend between the compartments of the cover, thereby maintaining the structural integrity of the inner shroud body. At least one or more compartments with corresponding registering cavities are preferred and preferably two or four compartments with corresponding cavities are most preferred. Four cavities are used if a circumferential rib is needed for stiffening.
  • the ribs of the inner shroud body in the latter preferred embodiment extend axially, radially and circumferentially, thereby maintaining the structural integrity of the plate.
  • the air inlet passages to the compartments of the cover of the inner shroud segment are provided with a spoolie which can be disposed in a passageway formed through the outer shroud body.
  • the spoolie is coupled at its inner end to a nipple forming an air inlet for the inner shroud segment cover.
  • impingement cooling apparatus for a shroud system surrounding components rotatable about an axis in the hot gas path of a turbine, comprising a shroud segment forming part of a shroud for surrounding the rotating components of the turbine, the shroud segment including a shroud segment body having a circumferentially extending wall, in part, defining the hot gas path, a plurality of cavities on a side of the segment body remote from the hot gas path and a cover for the shroud segment body having a cooling air inlet and a plurality of radially inwardly projecting compartments in communication with the air inlet and received in the cavities, respectively, each compartment having a bottom wall in spaced registration with the wall of the segment body and having a plurality of impingement apertures opening therethrough for flowing impingement cooling air from the compartments through the apertures and against the segment body wall for cooling the segment body wall and at least one passage through the segment body in communication with the space between the segment body wall
  • impingement cooling apparatus for a shroud system surrounding components rotatable about an axis in the hot gas path of a turbine, comprising an inner shroud segment forming part of the shroud system for surrounding the rotating components of the turbine, the inner shroud segment including an inner shroud body having a circumferentially and axially extending wall defining in part the hot gas path, at least four cavities formed in the inner shroud body on a side thereof remote from the hot gas path with radial innermost portions of the cavities formed by portions of the inner shroud body wall and a cover having a cooling air inlet and a plurality of radially inwardly projecting closed compartments in communication with the inlet for receiving cooling air, the compartments being received in the cavities, respectively, the compartments having bottom walls in spaced registration with the inner shroud body wall portions and a plurality of impingement apertures through each of the bottom walls for flowing impingement cooling air from the compartments against the inner sh
  • impingement cooling apparatus for a shroud system surrounding components rotatable about an axis in the hot gas path of a turbine, comprising an inner shroud segment forming part of the shroud system for surrounding the rotating components of the turbine, the inner shroud segment including an inner shroud body having a circumferentially and axially extending wall defining in part the hot gas path, at least one cavity formed in the inner shroud body on a side thereof remote from the hot gas path and opening radially outwardly, radial innermost portions of one cavity being formed by portions of the inner shroud body wall, and a cover having a cooling air inlet and at least one radially inwardly projecting closed compartment in communication with the inlet for receiving cooling air, one compartment being received in one cavity, one compartment having a bottom wall in spaced registration with the inner shroud body wall portions and a plurality of impingement apertures through the bottom wall for flowing impingement cooling air from one compartment against
  • the present invention seeks to provide a novel and improved impingement cooling apparatus for the shroud of a gas turbine wherein impingement cooling efficiencies are maximized by eliminating leakage paths for the cooling inlet flow to the inner shroud segment and minimizing the distance of impingement flow between the impingement plate apertures and the wall surface being cooled.
  • shroud system 10 for surrounding the rotating components in the hot gas path of a turbine and which shroud system is generally designated 10.
  • Shroud system 10 is secured to a stationary frame 12 of a turbine housing and surrounds the rotating buckets or vanes 14 disposed in the hot gas path 16 of the turbine, shroud system 10 for the first stage of the turbine being illustrated.
  • the direction of flow of the hot gas is indicated by the arrow 18.
  • the shroud system 10 includes outer and inner shroud segments, generally designated 20 and 22, respectively. It will be appreciated that the shroud system includes a plurality of such segments arranged circumferentially relative to one another with two or three inner shroud segments 22 connected to each of the outer shroud segments 20. For example, there may be on the order of forty-two outer shroud segments circumferentially adjacent one another and eighty-four inner shroud segments circumferentially adjacent one another, with each pair of inner shroud segments being secured to an outer shroud segment.
  • Each outer shroud segment 20 preferably has a pair of axially extending flanges 24 and an axially reduced neck portion 26 forming a dovetail connection with locating flanges or hooks 28 formed on the stationary frame 12.
  • the outer shroud segments 20 can be fitted to the frame 12 in a circumferential direction for securement thereto.
  • Radially inner portions of the outer shroud segment 20 define locating hooks 30 extending axially toward one another.
  • Inner shroud segment 22 has axially projecting flanges 32 which cooperate with the hooks 30 to secure the inner shroud segments 22 to the outer shroud segments 20.
  • the outer shroud segment 20 also includes a passageway 34 for receiving cooling air, for example, compressor discharge air.
  • a spoolie 36 is disposed in passage 34 for transmitting the cooling air into compartments of the inner shroud segment as described below.
  • the inner shroud segment 22 includes an inner shroud segment body 38 and a cover 40.
  • Inner shroud segment body 38 extends axially and circumferentially and includes a radially inner circumferentially and axially extending wall 42 defining in part the hot gas path 16 flowing past the rotating components, i.e., buckets 14.
  • Body 38 also includes a plurality of cavities 44 formed in the radially outermost wall surface of body 38. Cavities 44 are defined by radially outwardly projecting structural ribs 46 and 48, the ribs 46 extending axially, while the ribs 48 extend circumferentially.
  • the cavities 44 have a plurality of exit openings along side wall portions thereof for flowing spent cooling air through passages 45 opening through the outer walls of the body 38 for egress into the hot gas path 16.
  • the openings 50 through the side walls of the cavity thus communicate with openings in the circumferentially and axially extending faces of the inner shroud body 38 radially inwardly of seals, not shown, between the inner shroud bodies and between the inner shroud bodies and outer shroud bodies.
  • the inner shroud body cover 40 carries a plurality of depending compartments 52.
  • the compartments lie in communication with a plenum 54 located along the radially outermost surface of cover 40 and which plenum lies in communication with the inner end of the spoolie 36 via plenum inlet 55 for receiving cooling air.
  • Plenum 54 also lies in communication through openings in the cover with each of the compartments 52.
  • Each of the compartments 52 has a plurality of apertures 56 through bottom walls 60 of compartments 52, the compartments 52 being otherwise closed except for plenum inlet 55 and apertures 56.
  • Compartments 52 are spaced from one another to define recesses 57 therebetween for receiving the ribs 46 and 48 when the cover 40 overlies the inner shroud body 38.
  • Additional apertures 58 are provided through corner portions of the compartments 52.
  • the compartments 52 reside in cavities 44 with the ribs 46 and 48 extending in the recesses 57 between the respective compartments.
  • the depth of the compartments is such that the bottom walls 60 and hence the apertures 56 therethrough lie in close spaced relation to the wall portions or floors 64 of the cavities 44.
  • cooling air is supplied to the spoolie 36, which in turn supplies the air to plenum 54 via inlet 55 and compartments 52 via openings through the cover into compartments 52.
  • the cooling air flows through the impingement apertures 56 of compartments 52 for impingement cooling against the floors 64 of the cavities lying on the opposite side of the inner shroud body from the hot gas path 16, thus cooling the radially innermost wall 42 of the inner shroud segments.
  • Additional impingement cooling air flow flows through the corner apertures 58 of compartments 52 and against the side walls (corners) of the cavities 44.
  • the spent cooling air flows out of the cavities 44 through the apertures 50 and passages 45 and into the hot gas stream 16 by way of openings on the axial sides, circumferential sides , or floor of the inner shroud body.
  • the impingement openings 56 in the compartments 52 lie closely spaced to the wall 42 of the inner shroud bodies for efficient impingement air cooling. That is, the distance between the bottom walls 60 of the compartments 52 and the floors 64 is minimal to maximize the cooling effect of the impingement air flow.
  • the inner shroud body is also structurally maintained by the arrangement of the ribs 46 and 48.
  • the inner shroud body 38a includes two compartments 52a circumferentially spaced one from the other, with an axially extending rib 46a between the compartments.
  • the impingement cooling on the inner shroud wall is accomplished similarly as previously described.
  • the impingement cooling flow is supplied to the spoolie 36a, which in turn supplies the air to plenum 54a via inlet 55a and compartments 52a.
  • the cooling air flows through the impingement apertures 56a of compartments 52a for impingement cooling against the floors of the cavities 44a, thus cooling the radially innermost wall of the inner shroud body 38a.
  • the spent cooling air flows out of the cavities 44a through forward and aft passages 45a.
  • the inner shroud body may include only one cavity 44 formed in the radially outermost wall thereof, with exit openings along forward and aft walls and/or side walls for flowing spent cooling air through the exit openings for egress into the hot gas path.
  • the inner shroud body cover carries a single depending compartment which lies in communication with the plenum at the inner end of the spoolie.
  • the compartment has a plurality of apertures through bottom walls spaced closely adjacent the radially outer wall of the inner shroud body for flowing impingement cooling air against the latter wall.
  • the spent cooling air then flows through the forward and aft and/or side openings for egress into the hot gas stream.
  • the spent cooling impingement air may flow into the hot gas stream through openings, for example, openings 64 illustrated in Figure 4, through the radially innermost floor of the cavities 44, i.e., the wall defining the hot gas path.
  • the flow of cooling air to the shrouds can be altered, for example, during an engine retrofit.
  • the size of the spoolie can be changed to admit additional cooling air if the engine is running too hot or to limit the flow of cooling air if the cooling effect is too substantial.

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  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
EP00301015A 2000-02-09 2000-02-09 Dispositif de refroidissement par impact pour une bande de protection de turbine à gaz Withdrawn EP1124039A1 (fr)

Priority Applications (1)

Application Number Priority Date Filing Date Title
EP00301015A EP1124039A1 (fr) 2000-02-09 2000-02-09 Dispositif de refroidissement par impact pour une bande de protection de turbine à gaz

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
EP00301015A EP1124039A1 (fr) 2000-02-09 2000-02-09 Dispositif de refroidissement par impact pour une bande de protection de turbine à gaz

Publications (1)

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EP1124039A1 true EP1124039A1 (fr) 2001-08-16

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Cited By (18)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP1154126A3 (fr) * 2000-05-08 2003-02-26 General Electric Company Virole de turbine refroidie par vapeur dans un circuit fermé
WO2003054360A1 (fr) * 2001-12-13 2003-07-03 Alstom Technology Ltd Sous-groupe de parcours de gaz chauds de turbine a gaz
WO2003054359A1 (fr) * 2001-12-13 2003-07-03 Alstom Technology Ltd Unite d'etancheification de composants d'une turbomachine
EP1990507A1 (fr) * 2006-03-02 2008-11-12 IHI Corporation Structure de refroidissement par contact
EP1676981A3 (fr) * 2004-12-29 2009-09-16 United Technologies Corporation Refroidissement d'une virole de turbine
WO2010009997A1 (fr) * 2008-07-22 2010-01-28 Alstom Technology Ltd. Joint annulaire d'enveloppe pour turbine à gaz
EP2505787A1 (fr) * 2011-03-28 2012-10-03 Rolls-Royce plc Composant de moteur à turbine à gaz et moteur à turbine à gaz associé
US8814507B1 (en) 2013-05-28 2014-08-26 Siemens Energy, Inc. Cooling system for three hook ring segment
RU2538985C1 (ru) * 2013-12-30 2015-01-10 Открытое акционерное общество "Авиадвигатель" Статор высокотемпературной турбины
US9416671B2 (en) 2012-10-04 2016-08-16 General Electric Company Bimetallic turbine shroud and method of fabricating
EP2527599A3 (fr) * 2011-04-18 2017-03-15 General Electric Company Appareil pour former un étanchéité avec un étage d'aube de turbine dans une turbine à gaz
EP3173583A1 (fr) * 2015-11-24 2017-05-31 Rolls-Royce North American Technologies, Inc. Tubes d'impact pour refroidissement de segment d'étanchéité en cmc
RU2624691C1 (ru) * 2016-05-10 2017-07-05 Акционерное общество "Научно-производственный центр газотурбостроения "Салют" (АО "НПЦ газотурбостроения "Салют") Устройство охлаждения уплотнительных гребней бандажных полок рабочих лопаток турбины
EP3092373A4 (fr) * 2013-12-17 2017-09-27 United Technologies Corporation Plaque de dosge pour élément d'étanchéité à l'air externe d'aube
US10436041B2 (en) 2017-04-07 2019-10-08 General Electric Company Shroud assembly for turbine systems
US10619514B2 (en) 2017-10-18 2020-04-14 Rolls-Royce Corporation Ceramic matrix composite assembly with compliant pin attachment features
US10633996B2 (en) 2016-11-17 2020-04-28 Rolls-Royce Corporation Turbine cooling system
US10801350B2 (en) 2018-02-23 2020-10-13 Rolls-Royce Corporation Actively cooled engine assembly with ceramic matrix composite components

Citations (9)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3583824A (en) * 1969-10-02 1971-06-08 Gen Electric Temperature controlled shroud and shroud support
US4303371A (en) * 1978-06-05 1981-12-01 General Electric Company Shroud support with impingement baffle
US4329113A (en) * 1978-10-06 1982-05-11 Societe Nationale D'etude Et De Construction De Moteurs D'aviation Temperature control device for gas turbines
US5092735A (en) * 1990-07-02 1992-03-03 The United States Of America As Represented By The Secretary Of The Air Force Blade outer air seal cooling system
WO1994012775A1 (fr) * 1992-11-24 1994-06-09 United Technologies Corporation Ensemble d'etancheite d'air externe a refroidissement pour turbine
EP0690205A2 (fr) * 1994-06-30 1996-01-03 General Electric Company Dispositif de refroidissement d'une virole de turbine
EP0940562A2 (fr) * 1998-03-03 1999-09-08 Mitsubishi Heavy Industries, Ltd. Turbine à gaz
EP0959230A2 (fr) 1998-03-23 1999-11-24 General Electric Company Refroidissement pour une virole d'une turbine à gaz
US5993150A (en) * 1998-01-16 1999-11-30 General Electric Company Dual cooled shroud

Patent Citations (9)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3583824A (en) * 1969-10-02 1971-06-08 Gen Electric Temperature controlled shroud and shroud support
US4303371A (en) * 1978-06-05 1981-12-01 General Electric Company Shroud support with impingement baffle
US4329113A (en) * 1978-10-06 1982-05-11 Societe Nationale D'etude Et De Construction De Moteurs D'aviation Temperature control device for gas turbines
US5092735A (en) * 1990-07-02 1992-03-03 The United States Of America As Represented By The Secretary Of The Air Force Blade outer air seal cooling system
WO1994012775A1 (fr) * 1992-11-24 1994-06-09 United Technologies Corporation Ensemble d'etancheite d'air externe a refroidissement pour turbine
EP0690205A2 (fr) * 1994-06-30 1996-01-03 General Electric Company Dispositif de refroidissement d'une virole de turbine
US5993150A (en) * 1998-01-16 1999-11-30 General Electric Company Dual cooled shroud
EP0940562A2 (fr) * 1998-03-03 1999-09-08 Mitsubishi Heavy Industries, Ltd. Turbine à gaz
EP0959230A2 (fr) 1998-03-23 1999-11-24 General Electric Company Refroidissement pour une virole d'une turbine à gaz

Cited By (27)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP1154126A3 (fr) * 2000-05-08 2003-02-26 General Electric Company Virole de turbine refroidie par vapeur dans un circuit fermé
WO2003054360A1 (fr) * 2001-12-13 2003-07-03 Alstom Technology Ltd Sous-groupe de parcours de gaz chauds de turbine a gaz
WO2003054359A1 (fr) * 2001-12-13 2003-07-03 Alstom Technology Ltd Unite d'etancheification de composants d'une turbomachine
US7104751B2 (en) 2001-12-13 2006-09-12 Alstom Technology Ltd Hot gas path assembly
EP1676981A3 (fr) * 2004-12-29 2009-09-16 United Technologies Corporation Refroidissement d'une virole de turbine
EP1990507A4 (fr) * 2006-03-02 2014-04-23 Ihi Corp Structure de refroidissement par contact
EP1990507A1 (fr) * 2006-03-02 2008-11-12 IHI Corporation Structure de refroidissement par contact
WO2010009997A1 (fr) * 2008-07-22 2010-01-28 Alstom Technology Ltd. Joint annulaire d'enveloppe pour turbine à gaz
US8353663B2 (en) 2008-07-22 2013-01-15 Alstom Technology Ltd Shroud seal segments arrangement in a gas turbine
CH699232A1 (de) * 2008-07-22 2010-01-29 Alstom Technology Ltd Gasturbine.
KR101584974B1 (ko) 2008-07-22 2016-01-13 알스톰 테크놀러지 리미티드 가스 터빈의 시라우드 시일 부분 구성
EP2505787A1 (fr) * 2011-03-28 2012-10-03 Rolls-Royce plc Composant de moteur à turbine à gaz et moteur à turbine à gaz associé
EP2527599A3 (fr) * 2011-04-18 2017-03-15 General Electric Company Appareil pour former un étanchéité avec un étage d'aube de turbine dans une turbine à gaz
US9416671B2 (en) 2012-10-04 2016-08-16 General Electric Company Bimetallic turbine shroud and method of fabricating
US8814507B1 (en) 2013-05-28 2014-08-26 Siemens Energy, Inc. Cooling system for three hook ring segment
EP3092373A4 (fr) * 2013-12-17 2017-09-27 United Technologies Corporation Plaque de dosge pour élément d'étanchéité à l'air externe d'aube
US10364706B2 (en) 2013-12-17 2019-07-30 United Technologies Corporation Meter plate for blade outer air seal
RU2538985C1 (ru) * 2013-12-30 2015-01-10 Открытое акционерное общество "Авиадвигатель" Статор высокотемпературной турбины
EP3173583A1 (fr) * 2015-11-24 2017-05-31 Rolls-Royce North American Technologies, Inc. Tubes d'impact pour refroidissement de segment d'étanchéité en cmc
US10100654B2 (en) 2015-11-24 2018-10-16 Rolls-Royce North American Technologies Inc. Impingement tubes for CMC seal segment cooling
US11002143B2 (en) 2015-11-24 2021-05-11 Rolls-Royce North American Technologies Inc. Impingement tubes for gas turbine engine assemblies with ceramic matrix composite components
RU2624691C1 (ru) * 2016-05-10 2017-07-05 Акционерное общество "Научно-производственный центр газотурбостроения "Салют" (АО "НПЦ газотурбостроения "Салют") Устройство охлаждения уплотнительных гребней бандажных полок рабочих лопаток турбины
US10633996B2 (en) 2016-11-17 2020-04-28 Rolls-Royce Corporation Turbine cooling system
US10436041B2 (en) 2017-04-07 2019-10-08 General Electric Company Shroud assembly for turbine systems
US10619514B2 (en) 2017-10-18 2020-04-14 Rolls-Royce Corporation Ceramic matrix composite assembly with compliant pin attachment features
US11215082B2 (en) 2017-10-18 2022-01-04 Rolls-Royce Corporation Ceramic matrix composite assembly with compliant pin attachment features
US10801350B2 (en) 2018-02-23 2020-10-13 Rolls-Royce Corporation Actively cooled engine assembly with ceramic matrix composite components

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