US20070212214A1 - Segmented component seal - Google Patents

Segmented component seal Download PDF

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Publication number
US20070212214A1
US20070212214A1 US11/372,404 US37240406A US2007212214A1 US 20070212214 A1 US20070212214 A1 US 20070212214A1 US 37240406 A US37240406 A US 37240406A US 2007212214 A1 US2007212214 A1 US 2007212214A1
Authority
US
United States
Prior art keywords
gap
bridging element
seal
end portions
slots
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
US11/372,404
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English (en)
Inventor
Corneil Paauwe
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.)
RTX Corp
Original Assignee
United Technologies Corp
Priority date (The priority date is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the date listed.)
Filing date
Publication date
Application filed by United Technologies Corp filed Critical United Technologies Corp
Priority to US11/372,404 priority Critical patent/US20070212214A1/en
Assigned to UNITED TECHNOLOGIES CORPORATION reassignment UNITED TECHNOLOGIES CORPORATION ASSIGNMENT OF ASSIGNORS INTEREST (SEE DOCUMENT FOR DETAILS). Assignors: PAAUWE, CORNEIL
Priority to EP07250994A priority patent/EP1832715B1/de
Priority to DE602007008001T priority patent/DE602007008001D1/de
Publication of US20070212214A1 publication Critical patent/US20070212214A1/en
Abandoned legal-status Critical Current

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Classifications

    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F16—ENGINEERING ELEMENTS AND UNITS; GENERAL MEASURES FOR PRODUCING AND MAINTAINING EFFECTIVE FUNCTIONING OF MACHINES OR INSTALLATIONS; THERMAL INSULATION IN GENERAL
    • F16J—PISTONS; CYLINDERS; SEALINGS
    • F16J15/00—Sealings
    • F16J15/02—Sealings between relatively-stationary surfaces
    • F16J15/06—Sealings between relatively-stationary surfaces with solid packing compressed between sealing surfaces
    • F16J15/08—Sealings between relatively-stationary surfaces with solid packing compressed between sealing surfaces with exclusively metal packing
    • F16J15/0887—Sealings between relatively-stationary surfaces with solid packing compressed between sealing surfaces with exclusively metal packing the sealing effect being obtained by elastic deformation of the packing
    • 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
    • F01D11/005—Sealing means between non relatively rotating elements
    • 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
    • F01D11/005—Sealing means between non relatively rotating elements
    • F01D11/006—Sealing the gap between rotor blades or blades and rotor
    • F01D11/008—Sealing the gap between rotor blades or blades and rotor by spacer elements between the blades, e.g. independent interblade platforms
    • 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
    • F01D25/24—Casings; Casing parts, e.g. diaphragms, casing fastenings
    • F01D25/246—Fastening of diaphragms or stator-rings
    • 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
    • F01D9/041—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles forming ring or sector using blades
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02C—GAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS
    • F02C7/00—Features, components parts, details or accessories, not provided for in, or of interest apart form groups F02C1/00 - F02C6/00; Air intakes for jet-propulsion plants
    • F02C7/28—Arrangement of seals
    • 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
    • F05D2240/00—Components
    • F05D2240/55—Seals

Definitions

  • the invention relates to gas turbine engine components in general, and specifically to a seal for preventing leakage of high pressure air or other fluids between segmented components found in such engines.
  • a gas turbine engine 10 comprises one or more forward compressor sections 12 , a central combustor section 14 and one or more rearward turbine sections 16 .
  • the engine 10 operates by compressing ambient air 18 with the compressors 12 , adding fuel upstream of the combustor 14 and burning a fuel-air mixture 20 in the combustor 14 .
  • High temperature combustion gases 22 are directed axially rearward from the combustor 14 , through an annular duct 24 disposed in the turbines 16 .
  • the combustion gases 22 interact with one or more turbine rotors 26 disposed in the duct 24 .
  • the turbine rotors 26 are coupled to compressor rotors 28 via concentric shafts 30 rotating about a central longitudinal axis 32 of the engine 10 .
  • Gas turbine engines are known to power aircraft, ships and electrical generators.
  • Extending into the annular gas duct 24 are alternating circumferential stages of rotating blades 34 and stationary vanes 36 .
  • the stationary vanes 36 extend radially inwardly from a casing structure 38 surrounding the turbines 16 .
  • low temperature compressor air 40 is directed radially inboard and outboard of the duct 24 to the components.
  • the compressor air 40 is maintained at a higher pressure than the combustion gas 22 pressure, to ensure a continuous supply of compressor air 40 reaches the components.
  • TMF thermal mechanical fatigue
  • axial and radial gaps must be included between adjacent components to allow for thermal expansion. These gaps require sealing to ensure an adequate pressure differential exists between the compressor air and the combustion gas. Maintaining a compressor air pressure that is greater than the combustion gas pressure ensures a continuous flow of compressor air and prevents backflow of the combustion gas. Excessive leakage of the compressor air may cause premature oxidation of the components and can increase the engine's fuel burn. With jet fuel accounting for up to sixty five percent of the operating expense of a commercial airliner, any reduction in fuel burn is beneficial.
  • Feather seals are the type most commonly used between segmented components in gas turbine engines. Feather seals comprise a slot in the adjacent components that are open to the gap, and a bridging element disposed in the slots, spanning across the gap.
  • Flat bridging elements such as those disclosed in U.S. Pat. No. 5,154,577 to Kellock, et al, are fit into the adjoining slots. They depend on the higher-pressure compressor air to seat the bridging elements against the slots to form the seal. Assembly damage, misaligned slots, slot surface finish and low compressor air pressure may negatively affect the performance of flat bridging elements.
  • Resilient bridging elements such as those disclosed in U.S. Pat. No. 4,537,024 to Grosjean, are press fit into the adjoining slots. They rely on the contact pressure between the bridging element and the slot being greater than the compressor air pressure to form the seal. However, the single loop ends of the Grosjean bridging element offer limited contact pressure with the slot and are subject to compression about their minor axis.
  • a seal for restricting leakage of a high-pressure fluid from a first chamber, through a gap between two adjoining components, to a second chamber.
  • a slot is formed in each of the two components.
  • the slots face one another and are open to the gap.
  • Each slot contains a longitudinal axis, an upstream surface proximate the first chamber and a downstream surface proximate the second chamber.
  • Disposed in the slots and spanning the gap is a bridging element.
  • the bridging element contains a sectional profile, transverse to the longitudinal slot axis, that includes a flat central portion disposed between two approximately wave shaped end portions. The bridging element spans the gap between the two components with the end portions disposed in the slots, and the central portion disposed against a slot surface.
  • a primary feature of the seal is the approximately wave shaped profile of the end portions.
  • the approximate wave shape increases the contact force between the ends of the bridging element and the slot surfaces. Also, the approximately wave shaped ends force the flat center portion against a slot surface.
  • a primary advantage of the seal is an increased leakage restriction over conventional seals with minimum increase in weight and cost.
  • FIG. 1 is a simplified sectional view of an axial flow gas turbine engine.
  • FIG. 2 is a partial sectional view of a high pressure turbine of the type used in the gas turbine engine of FIG. 1 .
  • FIG. 3 is a partial isometric view of a segmented vane assembly of the type used in the high pressure turbine of FIG. 2 .
  • FIG. 4 a is a simplified sectional view, taken perpendicular to the longitudinal axis of the slots, of a seal in accordance with an embodiment of the invention disposed between segmented components with aligned slots.
  • FIG. 4 b is a simplified sectional view, taken perpendicular to the longitudinal axis of the slots, of a seal in accordance with an embodiment of the invention disposed between segmented components with misaligned slots.
  • FIG. 4 c is a simplified sectional view of a gap bridging element of FIGS. 4 a and 4 b prior to installation.
  • FIG. 2 An exemplary turbine 16 of a gas turbine engine 10 is illustrated in FIG. 2 .
  • the high temperature combustion gases 22 discharge rearward from the combustor 14 , at a pressure (P 1 ), to an annular duct 24 defined by an inner periphery 42 and an outer periphery 44 .
  • Stationary vanes 36 guide the combustion gases 22 to rotating blades 34 , extending radially outwardly from rotor disks 46 .
  • the vanes 36 span radially between inner 48 and outer 50 shrouds, which are suspended from an inner support 52 and/or outer casing 38 structures.
  • Inner seals 54 restrict leakage of the combustion gases 22 from beneath the vanes 36 at the inner periphery 42 .
  • Outer seals 56 restrict leakage of the combustion gases 22 from above tips 58 of the blades 34 at the outer periphery 44 .
  • each of the above-described turbine 16 components must be actively cooled, because the combustion gas 22 temperature typically exceeds the melting temperatures of the components' base alloy.
  • relatively low temperature compressor air 40 is distributed from the compressor 12 ( FIG. 1 ), at a pressure (P 2 ), to the inner 42 and outer 44 duct peripheries and away from the annular duct 24 .
  • the compressor air 40 pressure (P 2 ) is maintained at a higher level than the combustion gas 22 pressure (P 1 ) in order to allow compressor air to flow through turbine 16 components for cooling and thus preventing overheating and premature oxidation of the components.
  • Seals ensure a typical pressure ratio (P 2 :P 1 ) of approximately 1.03, but certainly greater than 1.0, exists during all engine operating conditions.
  • Circumferentially segmented components such as the vanes 36 , inner seals 54 , outer seals 56 and the like, include a seal 60 between adjacent segments.
  • a seal 60 in accordance with an embodiment of the invention contains a bridging element 62 that fits into axial and/or radial slots 64 machined into a mate face 66 of the vanes 36 .
  • a gap 38 between the vanes 36 typically between 0:010 inch (0.254 mm) and 0.030 inch (0.762 mm) depending on the size of the components accounts for thermal growth and reduces TMF.
  • the slots 64 face one another and are open to the gap 68 .
  • the bridging element 62 fits into the slots 64 , while spanning across the gap 68 .
  • FIGS. 4 a - 4 c Further details of a segmented component seal 60 according to an embodiment of the invention are generally illustrated in FIGS. 4 a - 4 c .
  • Opposed slots 64 are open to a gap 68 and each contain a longitudinal axis 70 , an upstream surface 72 proximate a first chamber 74 and a downstream surface 76 proximate a second chamber 78 . Although the upstream 72 and downstream 76 surfaces are shown parallel in the illustration, they could also converge or diverge away from the gap 68 .
  • the slots 64 have an opening width (W) of between about 0.030 inch (0.762 mm) and 0.060 inch (1.524 mm).
  • the slots 64 are preferably aligned as illustrated in FIG.
  • the slot axis 70 is linear, but a curvilinear slot axis 70 may also be used.
  • the slots are produced by casting, abrasive machining, electrodischarge machining, or other suitable means.
  • the bridging element 62 contains a sectional profile, transverse to the longitudinal slot axis 64 that includes a flat shaped central portion 84 disposed between two, approximately wave shaped, end portions 86 .
  • the central portion 84 spans across the gap 68 and the end portions 86 seat against surfaces 72 , 76 of each slot 64 .
  • the bridging element 62 may be installed with the central portion 84 positioned adjacent the upstream surface 72 .
  • the bridging element 62 is installed with the central portion 84 positioned adjacent the downstream surface 76 .
  • the end portions 86 are resiliently sprung into the slots 64 , and are in direct contact with each of the upstream 72 and downstream 76 surfaces.
  • the end portions 86 alternate in direction, first away from and then back toward the central portion 84 and the gap 68 , thus approximating a waveform.
  • W slot width
  • the number of wave cycles depends on the slot width (W) and the amount of resilient spring force necessary to positively seat the end portions 86 against the upstream 72 and downstream 76 surfaces.
  • a free height (H) of the uninstalled bridging element 62 is slightly larger than the opening width (W) of the slot 64 .
  • the free height (H) is between 0.005 inch (0.127 mm) and 0.020 inch (0.508 mm) larger than the opening width (W).
  • W opening width
  • the bridging element 62 is made of a material with suitable low temperature ductility and high temperature strength.
  • a Nickel or Cobalt based alloy strip approximately 0.010 inch (0.254 mm) thick is used. Stamping, progressive rolling or other suitable forming process may be used to form the profile of the strip.
  • the end portions 86 are compressed together and resiliently sprung into the slots 64 during assembly.
  • the interference fit between the end portions 86 and the upstream 80 and downstream 82 surfaces creates four independent leakage restrictions 90 .
  • the multiple restrictions 90 significantly reduce leakage from the first chamber 74 to the second chamber 78 .
  • the multiple restrictions 90 remain intact, even if the slots 64 are slightly misaligned.

Landscapes

  • Engineering & Computer Science (AREA)
  • General Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • Chemical & Material Sciences (AREA)
  • Combustion & Propulsion (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
  • Gasket Seals (AREA)
US11/372,404 2006-03-09 2006-03-09 Segmented component seal Abandoned US20070212214A1 (en)

Priority Applications (3)

Application Number Priority Date Filing Date Title
US11/372,404 US20070212214A1 (en) 2006-03-09 2006-03-09 Segmented component seal
EP07250994A EP1832715B1 (de) 2006-03-09 2007-03-09 Dichtung für segmentierte Gasturbinenkomponenten
DE602007008001T DE602007008001D1 (de) 2006-03-09 2007-03-09 Dichtung für segmentierte Gasturbinenkomponenten

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
US11/372,404 US20070212214A1 (en) 2006-03-09 2006-03-09 Segmented component seal

Publications (1)

Publication Number Publication Date
US20070212214A1 true US20070212214A1 (en) 2007-09-13

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Application Number Title Priority Date Filing Date
US11/372,404 Abandoned US20070212214A1 (en) 2006-03-09 2006-03-09 Segmented component seal

Country Status (3)

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US (1) US20070212214A1 (de)
EP (1) EP1832715B1 (de)
DE (1) DE602007008001D1 (de)

Cited By (13)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US20090033036A1 (en) * 2006-03-06 2009-02-05 Peter Marx Gas turbine with annular heat shield
US20130209250A1 (en) * 2012-02-13 2013-08-15 General Electric Company Transition piece seal assembly for a turbomachine
US20130266435A1 (en) * 2012-04-10 2013-10-10 General Electric Company Turbine shroud assembly and method of forming
US20140225334A1 (en) * 2013-02-13 2014-08-14 Mitsubishi Heavy Industries, Ltd. Combustor seal structure and a combustor seal
JP2014532831A (ja) * 2011-11-06 2014-12-08 ゼネラル・エレクトリック・カンパニイ ガスタービンエンジン用の非対称半径方向スプラインシール
WO2015013503A1 (en) 2013-07-24 2015-01-29 United Technologies Corporation Trough seal for gas turbine engine
CN104696023A (zh) * 2013-10-08 2015-06-10 通用电气公司 便于在燃气涡轮中密封的方法和系统
US20160003079A1 (en) * 2013-03-08 2016-01-07 United Technologies Corporation Gas turbine engine component having variable width feather seal slot
JP2016512865A (ja) * 2013-03-21 2016-05-09 シーメンス アクティエンゲゼルシャフト 間隙を密封するためのシール要素
US20180320539A1 (en) * 2017-05-02 2018-11-08 Safran Aircraft Engines Assembly for gas turbine, associated gas turbine
US20200173295A1 (en) * 2018-12-04 2020-06-04 United Technologies Corporation Gas turbine engine arc segments with arced walls
US11168574B2 (en) * 2015-06-29 2021-11-09 Raytheon Technologies Corporation Segmented non-contact seal assembly for rotational equipment
US11649732B2 (en) * 2021-03-11 2023-05-16 Raytheon Technologies Corporation Vane assembly with spring device for biasing mate face seal

Families Citing this family (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
FR2957969B1 (fr) * 2010-03-26 2013-03-29 Snecma Dispositif d'etancheite entre les talons d'aubes adjacentes en materiau compositie d'une roue mobile de turbomachine
US8834109B2 (en) * 2011-08-03 2014-09-16 United Technologies Corporation Vane assembly for a gas turbine engine
EP2657455A1 (de) * 2012-04-27 2013-10-30 Siemens Aktiengesellschaft Hitzeschild und Herstellungsverfahren dafür
FR3081188B1 (fr) * 2018-05-15 2021-03-19 Safran Aircraft Engines Aubage de stator pour une turbomachine

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US4311432A (en) * 1979-11-20 1982-01-19 United Technologies Corporation Radial seal
US4537024A (en) * 1979-04-23 1985-08-27 Solar Turbines, Incorporated Turbine engines
US5125796A (en) * 1991-05-14 1992-06-30 General Electric Company Transition piece seal spring for a gas turbine
US5154577A (en) * 1991-01-17 1992-10-13 General Electric Company Flexible three-piece seal assembly
US5435576A (en) * 1992-12-28 1995-07-25 Rode; John E. Spring gasket
US5709530A (en) * 1996-09-04 1998-01-20 United Technologies Corporation Gas turbine vane seal
US6431825B1 (en) * 2000-07-28 2002-08-13 Alstom (Switzerland) Ltd Seal between static turbine parts
US20040179937A1 (en) * 2001-09-25 2004-09-16 Erhard Kreis Seal arrangement for reducing the seal gaps within a rotary flow machine
US20060082074A1 (en) * 2004-10-18 2006-04-20 Pratt & Whitney Canada Corp. Circumferential feather seal
US7101147B2 (en) * 2003-05-16 2006-09-05 Rolls-Royce Plc Sealing arrangement
US7128323B2 (en) * 2003-08-20 2006-10-31 Eagle Engineering Aerospace Co., Ltd. Seal device
US7316402B2 (en) * 2006-03-09 2008-01-08 United Technologies Corporation Segmented component seal

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GB2335470B (en) * 1998-03-18 2002-02-13 Rolls Royce Plc A seal
US7152864B2 (en) * 2003-10-02 2006-12-26 Alstom Technology Ltd. Seal assembly

Patent Citations (12)

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Publication number Priority date Publication date Assignee Title
US4537024A (en) * 1979-04-23 1985-08-27 Solar Turbines, Incorporated Turbine engines
US4311432A (en) * 1979-11-20 1982-01-19 United Technologies Corporation Radial seal
US5154577A (en) * 1991-01-17 1992-10-13 General Electric Company Flexible three-piece seal assembly
US5125796A (en) * 1991-05-14 1992-06-30 General Electric Company Transition piece seal spring for a gas turbine
US5435576A (en) * 1992-12-28 1995-07-25 Rode; John E. Spring gasket
US5709530A (en) * 1996-09-04 1998-01-20 United Technologies Corporation Gas turbine vane seal
US6431825B1 (en) * 2000-07-28 2002-08-13 Alstom (Switzerland) Ltd Seal between static turbine parts
US20040179937A1 (en) * 2001-09-25 2004-09-16 Erhard Kreis Seal arrangement for reducing the seal gaps within a rotary flow machine
US7101147B2 (en) * 2003-05-16 2006-09-05 Rolls-Royce Plc Sealing arrangement
US7128323B2 (en) * 2003-08-20 2006-10-31 Eagle Engineering Aerospace Co., Ltd. Seal device
US20060082074A1 (en) * 2004-10-18 2006-04-20 Pratt & Whitney Canada Corp. Circumferential feather seal
US7316402B2 (en) * 2006-03-09 2008-01-08 United Technologies Corporation Segmented component seal

Cited By (23)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US20090033036A1 (en) * 2006-03-06 2009-02-05 Peter Marx Gas turbine with annular heat shield
JP2014532831A (ja) * 2011-11-06 2014-12-08 ゼネラル・エレクトリック・カンパニイ ガスタービンエンジン用の非対称半径方向スプラインシール
US9810086B2 (en) 2011-11-06 2017-11-07 General Electric Company Asymmetric radial spline seal for a gas turbine engine
US20130209250A1 (en) * 2012-02-13 2013-08-15 General Electric Company Transition piece seal assembly for a turbomachine
US9115808B2 (en) * 2012-02-13 2015-08-25 General Electric Company Transition piece seal assembly for a turbomachine
US20130266435A1 (en) * 2012-04-10 2013-10-10 General Electric Company Turbine shroud assembly and method of forming
US9316109B2 (en) * 2012-04-10 2016-04-19 General Electric Company Turbine shroud assembly and method of forming
US9500132B2 (en) * 2013-02-13 2016-11-22 Mitsubishi Heavy Industries, Ltd. Combustor seal structure and a combustor seal
US20140225334A1 (en) * 2013-02-13 2014-08-14 Mitsubishi Heavy Industries, Ltd. Combustor seal structure and a combustor seal
US10072517B2 (en) * 2013-03-08 2018-09-11 United Technologies Corporation Gas turbine engine component having variable width feather seal slot
US20160003079A1 (en) * 2013-03-08 2016-01-07 United Technologies Corporation Gas turbine engine component having variable width feather seal slot
JP2016512865A (ja) * 2013-03-21 2016-05-09 シーメンス アクティエンゲゼルシャフト 間隙を密封するためのシール要素
US20160177767A1 (en) * 2013-07-24 2016-06-23 United Technologies Corporation Trough seal for gas turbine engine
EP3025030A4 (de) * 2013-07-24 2017-03-15 United Technologies Corporation Rinnendichtung für gasturbinenmotor
US9714580B2 (en) * 2013-07-24 2017-07-25 United Technologies Corporation Trough seal for gas turbine engine
WO2015013503A1 (en) 2013-07-24 2015-01-29 United Technologies Corporation Trough seal for gas turbine engine
CN104696023A (zh) * 2013-10-08 2015-06-10 通用电气公司 便于在燃气涡轮中密封的方法和系统
US11168574B2 (en) * 2015-06-29 2021-11-09 Raytheon Technologies Corporation Segmented non-contact seal assembly for rotational equipment
US20180320539A1 (en) * 2017-05-02 2018-11-08 Safran Aircraft Engines Assembly for gas turbine, associated gas turbine
US10760440B2 (en) * 2017-05-02 2020-09-01 Safran Aircraft Engines Assembly for gas turbine, associated gas turbine
US20200173295A1 (en) * 2018-12-04 2020-06-04 United Technologies Corporation Gas turbine engine arc segments with arced walls
US10890079B2 (en) * 2018-12-04 2021-01-12 Raytheon Technologies Corporation Gas turbine engine arc segments with arced walls
US11649732B2 (en) * 2021-03-11 2023-05-16 Raytheon Technologies Corporation Vane assembly with spring device for biasing mate face seal

Also Published As

Publication number Publication date
DE602007008001D1 (de) 2010-09-09
EP1832715A2 (de) 2007-09-12
EP1832715A3 (de) 2008-07-09
EP1832715B1 (de) 2010-07-28

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