US20080061515A1 - Rim seal for a gas turbine engine - Google Patents

Rim seal for a gas turbine engine Download PDF

Info

Publication number
US20080061515A1
US20080061515A1 US11/530,226 US53022606A US2008061515A1 US 20080061515 A1 US20080061515 A1 US 20080061515A1 US 53022606 A US53022606 A US 53022606A US 2008061515 A1 US2008061515 A1 US 2008061515A1
Authority
US
United States
Prior art keywords
seal
rim seal
gas turbine
turbine engine
rim
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
US11/530,226
Inventor
Eric Durocher
Rene Paquet
Guy Lefebvre
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.)
Pratt and Whitney Canada Corp
Original Assignee
Pratt and Whitney Canada 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 Pratt and Whitney Canada Corp filed Critical Pratt and Whitney Canada Corp
Priority to US11/530,226 priority Critical patent/US20080061515A1/en
Assigned to PRATT & WHITNEY CANADA CORP. reassignment PRATT & WHITNEY CANADA CORP. ASSIGNMENT OF ASSIGNORS INTEREST (SEE DOCUMENT FOR DETAILS). Assignors: DUROCHER, ERIC, LEFEBVRE, GUY, PAQUET, RENE
Priority to CA2598329A priority patent/CA2598329C/en
Publication of US20080061515A1 publication Critical patent/US20080061515A1/en
Priority to US12/426,472 priority patent/US8172514B2/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
    • F01D11/001—Preventing or minimising internal leakage of working-fluid, e.g. between stages for sealing space between stator blade and rotor
    • 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
    • F01D11/12—Preventing or minimising internal leakage of working-fluid, e.g. between stages for sealing space between rotor blade tips and stator using a rubstrip, e.g. erodible. deformable or resiliently-biased part
    • F01D11/122—Preventing or minimising internal leakage of working-fluid, e.g. between stages for sealing space between rotor blade tips and stator using a rubstrip, e.g. erodible. deformable or resiliently-biased part with erodable or abradable material

Definitions

  • the invention relates generally to a rim seal for a gas turbine engine, and in particular to a rim seal for use within an annular space between rotating blades and a non-rotating adjacent structure in a gas turbine engine.
  • rotating elements such as compressors and turbine rotors
  • Their blades are also subjected to intense pressure and heat.
  • Compressors and turbine rotors are mounted between non-rotating structures within the engine. These structures are designed to be as close as possible to the rotating blade platforms. This mitigates pressurized air ingestion inside the gas turbine engine.
  • the present concept provides a rim seal for an annular space between blade platforms and a non-rotating adjacent structure in a gas turbine engine, the rim seal being connectable to the non-rotating structure and made of an abradable material.
  • the present concept provides an annular abradable rim seal for mitigating combustion gas ingestion on a side of blades in a gas turbine engine, the seal having an outer peripheral portion configured and disposed to be at least partially in friction engagement with blade platforms during operation of the engine.
  • the present concept provides a method of sealing an annular space between blade platforms and a non-rotating structure immediately adjacent to the blade platforms in a gas turbine engine, the method comprising securing to the non-rotating structure an abradable annular seal provided in the annular space; and operating the gas turbine engine to carve a notch in the seal with the side of the blades.
  • FIG. 1 is a schematic cross-sectional view of an example of a gas turbine engine
  • FIG. 2 is a schematic longitudinal cross-sectional view of an example of an improved rim seal.
  • FIG. 1 illustrates a gas turbine engine 10 of a type preferably provided for use in subsonic flight, generally comprising in serial flow communication a fan 12 through which ambient air is propelled, a multistage compressor 14 for pressurizing the air, a combustor 16 in which the compressed air is mixed with fuel and ignited for generating an annular stream of hot combustion gases, and a turbine section 18 for extracting energy from the combustion gases.
  • the turbine section 18 includes a high pressure turbine stage 20 and a low pressure turbine stage 22 .
  • FIG. 2 schematically shows the downstream side of a turbine wheel disc 24 which can be the rotor of either one of the high pressure turbine stage 20 or the low pressure turbine stage 22 .
  • the wheel disc 24 has a plurality of radially interspaced blades 26 .
  • a blade 26 can be seen having an airfoil section 28 extending radially outwardly from a blade platform 30 .
  • a non-rotating structure 32 is present adjacent to the blades 26 .
  • the non-rotating structure 32 can be the inner wall of an interturbine duct in the case of the high pressure turbine stage 20 or the inner wall of an exhaust duct in the case of the low pressure turbine stage 22 , for example.
  • the improved rim seal is not limited for use with turbine blades or at the outlet of a turbine stage.
  • the rim seal can also be used on either sides of a compressor rotor or on the inlet of the turbine rotor.
  • An annular space 34 is defined immediately adjacent to the blades of the wheel disc 24 , between the side of the blade platforms 30 and an end 36 of the non-rotating structure 32 .
  • a rim seal 38 connected to the end 36 of the non-rotating structure 32 , substantially fills the inner side of the annular space 34 .
  • the rim seal 38 is made of an abradable material such as honeycomb-shaped light material, for example.
  • each blade platform 26 has a protruding portion 40 on the side thereof. Together, the protruding portion 40 defines an annular recess 42 .
  • the rim seal 38 is set within the annular recess 42 along an overlap distance with respect to the edge of the protruding portions 40 .
  • a gap 44 referred to as a cold gap 44 is provided between the protruding portions 40 and the rim seal 38 along the overlap distance at ambient conditions.
  • the temperature rises and causes thermal expansion to close the cold gap 44 .
  • a light rub then occurs between the protruding portions 40 and the rim seal 38 . This increases the sealing effect. Interference between the rim seal and the protruding portions results in abrasion of the rim seal abradable material and the creation of a notch 46 .
  • the relative radial position of the flat portion 48 adjacent the notch 46 can be selected to arrive as flush as possible with the outer surface 50 of the blade platforms 26 and the outer surface 52 of the adjacent non-rotating structure 32 during operation of the engine, to minimize aerodynamic disruptions in the gas flow.
  • a carefully selected flat portion 48 configuration can thus contribute to more closely obtain a smooth surface transition between the outer surface 50 of the blade platform 26 and the outer surface 52 of the non-rotating structure 32 .
  • the notch 46 can be machined prior to installation of the rim seal 38 . Alternately, it can be carved in the rim seal 38 by abrasion with the protruding portions 40 during engine operation, or can be made by a combination of pre-machining and abrasion during operation.
  • a flanged support bracket 54 is connected to the end 36 and provides a support flange 56 on which the rim seal 38 can be brazed.
  • the abradable rim seal 38 can be secured both to the flange 56 and to the end 36 of the non-rotating structure 32 .
  • annular rim seal can be used with other types of non-rotating structures than the one described and illustrated herein.
  • abradable materials exist and the exact choice thereof is left to those skilled in the art.
  • the seal-holding bracket is optional, many different configurations can be used to connect the abradable rim seal to the edge of the non-rotating structure. Still other modifications which fall within the scope of the present invention will be apparent to those skilled in the art, in light of a review of this disclosure, and such modifications are intended to fall within the appended claims.

Abstract

The rim seal is positioned in an annular space between blades and a non-rotating adjacent structure in a gas turbine engine. The rim seal is connectable to the non-rotating structure and is made of an abradable material.

Description

    TECHNICAL FIELD
  • The invention relates generally to a rim seal for a gas turbine engine, and in particular to a rim seal for use within an annular space between rotating blades and a non-rotating adjacent structure in a gas turbine engine.
    • BACKGROUND
  • In a gas turbine engine, rotating elements, such as compressors and turbine rotors, operate at a very high rotation speed. Their blades are also subjected to intense pressure and heat.
  • Compressors and turbine rotors are mounted between non-rotating structures within the engine. These structures are designed to be as close as possible to the rotating blade platforms. This mitigates pressurized air ingestion inside the gas turbine engine.
  • Although various rim seal arrangements have been suggested in the past, there is always a need to provide an improved rim seal yielding better results than previous seals.
    • SUMMARY
  • In one aspect, the present concept provides a rim seal for an annular space between blade platforms and a non-rotating adjacent structure in a gas turbine engine, the rim seal being connectable to the non-rotating structure and made of an abradable material.
  • In a second aspect, the present concept provides an annular abradable rim seal for mitigating combustion gas ingestion on a side of blades in a gas turbine engine, the seal having an outer peripheral portion configured and disposed to be at least partially in friction engagement with blade platforms during operation of the engine.
  • In a third aspect, the present concept provides a method of sealing an annular space between blade platforms and a non-rotating structure immediately adjacent to the blade platforms in a gas turbine engine, the method comprising securing to the non-rotating structure an abradable annular seal provided in the annular space; and operating the gas turbine engine to carve a notch in the seal with the side of the blades.
  • Further details of these and other aspects of the present invention will be apparent from the detailed description and figures included below.
    • DESCRIPTION OF THE DRAWINGS
  • Reference is now made to the accompanying figures depicting aspects of the present invention, in which:
  • FIG. 1 is a schematic cross-sectional view of an example of a gas turbine engine; and
  • FIG. 2 is a schematic longitudinal cross-sectional view of an example of an improved rim seal.
    •  DETAILED DESCRIPTION
  • FIG. 1 illustrates a gas turbine engine 10 of a type preferably provided for use in subsonic flight, generally comprising in serial flow communication a fan 12 through which ambient air is propelled, a multistage compressor 14 for pressurizing the air, a combustor 16 in which the compressed air is mixed with fuel and ignited for generating an annular stream of hot combustion gases, and a turbine section 18 for extracting energy from the combustion gases. In this example, the turbine section 18 includes a high pressure turbine stage 20 and a low pressure turbine stage 22.
  • FIG. 2 schematically shows the downstream side of a turbine wheel disc 24 which can be the rotor of either one of the high pressure turbine stage 20 or the low pressure turbine stage 22. The wheel disc 24 has a plurality of radially interspaced blades 26. In the figure, a blade 26 can be seen having an airfoil section 28 extending radially outwardly from a blade platform 30. A non-rotating structure 32 is present adjacent to the blades 26. The non-rotating structure 32 can be the inner wall of an interturbine duct in the case of the high pressure turbine stage 20 or the inner wall of an exhaust duct in the case of the low pressure turbine stage 22, for example.
  • It should be noted that the improved rim seal is not limited for use with turbine blades or at the outlet of a turbine stage. The rim seal can also be used on either sides of a compressor rotor or on the inlet of the turbine rotor.
  • An annular space 34 is defined immediately adjacent to the blades of the wheel disc 24, between the side of the blade platforms 30 and an end 36 of the non-rotating structure 32. A rim seal 38, connected to the end 36 of the non-rotating structure 32, substantially fills the inner side of the annular space 34. The rim seal 38 is made of an abradable material such as honeycomb-shaped light material, for example.
  • In the illustrated example, each blade platform 26 has a protruding portion 40 on the side thereof. Together, the protruding portion 40 defines an annular recess 42. The rim seal 38 is set within the annular recess 42 along an overlap distance with respect to the edge of the protruding portions 40. A gap 44 referred to as a cold gap 44 is provided between the protruding portions 40 and the rim seal 38 along the overlap distance at ambient conditions. During operation of the gas turbine engine, the temperature rises and causes thermal expansion to close the cold gap 44. A light rub then occurs between the protruding portions 40 and the rim seal 38. This increases the sealing effect. Interference between the rim seal and the protruding portions results in abrasion of the rim seal abradable material and the creation of a notch 46.
  • The relative radial position of the flat portion 48 adjacent the notch 46 can be selected to arrive as flush as possible with the outer surface 50 of the blade platforms 26 and the outer surface 52 of the adjacent non-rotating structure 32 during operation of the engine, to minimize aerodynamic disruptions in the gas flow. A carefully selected flat portion 48 configuration can thus contribute to more closely obtain a smooth surface transition between the outer surface 50 of the blade platform 26 and the outer surface 52 of the non-rotating structure 32. The notch 46 can be machined prior to installation of the rim seal 38. Alternately, it can be carved in the rim seal 38 by abrasion with the protruding portions 40 during engine operation, or can be made by a combination of pre-machining and abrasion during operation.
  • In the illustrated example, a flanged support bracket 54, also made of sheet material, is connected to the end 36 and provides a support flange 56 on which the rim seal 38 can be brazed. The abradable rim seal 38 can be secured both to the flange 56 and to the end 36 of the non-rotating structure 32.
  • The above description is meant to be exemplary only, and one skilled in the art will recognize that changes may be made to the embodiments described without departing from the scope of the invention disclosed. For example, the annular rim seal can be used with other types of non-rotating structures than the one described and illustrated herein. Many different types of abradable materials exist and the exact choice thereof is left to those skilled in the art. The seal-holding bracket is optional, many different configurations can be used to connect the abradable rim seal to the edge of the non-rotating structure. Still other modifications which fall within the scope of the present invention will be apparent to those skilled in the art, in light of a review of this disclosure, and such modifications are intended to fall within the appended claims.

Claims (9)

1. A rim seal for an annular space between blade platforms and a non-rotating adjacent structure in a gas turbine engine, the rim seal being connectable to the non-rotating structure and made of an abradable material.
2. The rim seal as defined in claim 1 wherein each blade platform has a side protruding portion in engagement with the abradable material.
3. The rim seal as defined in claim 2 wherein a cold gap is provided between the protruding portion and the rim seal at ambient conditions, the cold gap being designed to close due to thermal expansion during operation of the gas turbine engine.
4. The rim seal as defined in claim 1 wherein the non-rotating structure is made of sheet metal the structure having an end folded radially inwards, the abradable seal being connected to the end.
5. The rim seal as defined in claim 4 wherein the non-rotating structure comprises a seal-holding bracket secured to the end.
6. An annular abradable rim seal for mitigating combustion gas ingestion on a side of blades in a gas turbine engine, the seal having an outer peripheral portion configured and disposed to be at least partially in friction engagement with blade platforms during operation of the engine.
7. The rim seal as defined in claim 6 wherein a cold gap is provided between the blades and the rim seal at ambient conditions, the cold gap being designed to close due to thermal expansion during operation of the gas turbine engine.
8. The rim seal as defined in claim 6 wherein the non-rotating structure comprises a seal-holding bracket secured to a side thereof.
9. A method of sealing an annular space between blade platforms and a non-rotating structure immediately adjacent to the blade platforms in a gas turbine engine, the method comprising:
securing to the non-rotating structure an abradable annular seal provided in the annular space; and
operating the gas turbine engine to carve a notch in the seal with the side of the blades.
US11/530,226 2006-09-08 2006-09-08 Rim seal for a gas turbine engine Abandoned US20080061515A1 (en)

Priority Applications (3)

Application Number Priority Date Filing Date Title
US11/530,226 US20080061515A1 (en) 2006-09-08 2006-09-08 Rim seal for a gas turbine engine
CA2598329A CA2598329C (en) 2006-09-08 2007-08-22 Rim seal for a gas turbine engine
US12/426,472 US8172514B2 (en) 2006-09-08 2009-04-20 Rim seal for a gas turbine engine

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
US11/530,226 US20080061515A1 (en) 2006-09-08 2006-09-08 Rim seal for a gas turbine engine

Related Child Applications (1)

Application Number Title Priority Date Filing Date
US12/426,472 Division US8172514B2 (en) 2006-09-08 2009-04-20 Rim seal for a gas turbine engine

Publications (1)

Publication Number Publication Date
US20080061515A1 true US20080061515A1 (en) 2008-03-13

Family

ID=39153702

Family Applications (2)

Application Number Title Priority Date Filing Date
US11/530,226 Abandoned US20080061515A1 (en) 2006-09-08 2006-09-08 Rim seal for a gas turbine engine
US12/426,472 Active 2026-10-21 US8172514B2 (en) 2006-09-08 2009-04-20 Rim seal for a gas turbine engine

Family Applications After (1)

Application Number Title Priority Date Filing Date
US12/426,472 Active 2026-10-21 US8172514B2 (en) 2006-09-08 2009-04-20 Rim seal for a gas turbine engine

Country Status (2)

Country Link
US (2) US20080061515A1 (en)
CA (1) CA2598329C (en)

Cited By (11)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US20090110548A1 (en) * 2007-10-30 2009-04-30 Pratt & Whitney Canada Corp. Abradable rim seal for low pressure turbine stage
US20090208326A1 (en) * 2006-09-08 2009-08-20 Eric Durocher Rim seal for a gas turbine engine
CN101858230A (en) * 2009-04-06 2010-10-13 通用电气公司 Method, system and/or device about the sealed department that is used for turbogenerator
US20110023499A1 (en) * 2006-09-15 2011-02-03 Nicolas Grivas Gas turbine combustor exit duct and hp vane interface
US9068469B2 (en) 2011-09-01 2015-06-30 Honeywell International Inc. Gas turbine engines with abradable turbine seal assemblies
WO2015181489A1 (en) * 2014-05-27 2015-12-03 Snecma Sealing plate with fuse function
US20160076454A1 (en) * 2014-09-16 2016-03-17 Alstom Technology Ltd Sealing arrangement at the interface between a combustor and a turbine of a gas turbine and gas turbine with such a sealing arrangement
EP3112602A1 (en) * 2015-07-01 2017-01-04 United Technologies Corporation Break-in system for gapping and leakage control
FR3039225A1 (en) * 2015-07-20 2017-01-27 Snecma TURBOMACHINE, SUCH AS A TURBO AIRCRAFT
FR3080646A1 (en) * 2018-04-26 2019-11-01 Safran Aircraft Engines SEALING BETWEEN A FIXED WHEEL AND A MOBILE WHEEL OF A TURBOMACHINE
US10633992B2 (en) 2017-03-08 2020-04-28 Pratt & Whitney Canada Corp. Rim seal

Families Citing this family (10)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US8888446B2 (en) * 2011-10-07 2014-11-18 General Electric Company Turbomachine rotor having patterned coating
US8905716B2 (en) 2012-05-31 2014-12-09 United Technologies Corporation Ladder seal system for gas turbine engines
US10378451B2 (en) 2013-09-13 2019-08-13 United Technologies Corporation Large displacement high temperature seal
EP2886801B1 (en) * 2013-12-20 2019-04-24 Ansaldo Energia IP UK Limited Seal system for a gas turbine and corresponding gas turbine
US10774666B2 (en) 2014-01-24 2020-09-15 Raytheon Technologies Corporation Toggle seal for a rim seal
US9957826B2 (en) 2014-06-09 2018-05-01 United Technologies Corporation Stiffness controlled abradeable seal system with max phase materials and methods of making same
EP3085900B1 (en) * 2015-04-21 2020-08-05 Ansaldo Energia Switzerland AG Abradable lip for a gas turbine
CN105134306B (en) * 2015-09-18 2017-01-18 西安交通大学 Radial rim sealing structure with damping holes and flow guide blades
US11193389B2 (en) 2019-10-18 2021-12-07 Raytheon Technologies Corporation Fluid cooled seal land for rotational equipment seal assembly
CN114909188B (en) * 2022-05-13 2023-02-24 北京航空航天大学 Sealing structure for rim of turbine disc of gas turbine

Citations (13)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US2564503A (en) * 1949-01-29 1951-08-14 Gen Electric Baffle means for rotating bodies
US3606349A (en) * 1969-08-07 1971-09-20 Rolls Royce Seal with pressure responsive tolerance control
US3755870A (en) * 1971-12-20 1973-09-04 Entoleter Pressure retaining face seal and method
US3908361A (en) * 1972-12-16 1975-09-30 Rolls Royce 1971 Ltd Seal between relatively moving components of a fluid flow machine
US3918925A (en) * 1974-05-13 1975-11-11 United Technologies Corp Abradable seal
US4309145A (en) * 1978-10-30 1982-01-05 General Electric Company Cooling air seal
US4320903A (en) * 1978-09-27 1982-03-23 Societe Nationale D'etude Et De Construction De Moteurs D'aviation Labyrinth seals
US5575486A (en) * 1992-10-01 1996-11-19 Atlas Copco Tools Ab Grooved run-in face seal
US5785492A (en) * 1997-03-24 1998-07-28 United Technologies Corporation Method and apparatus for sealing a gas turbine stator vane assembly
US5962076A (en) * 1995-06-29 1999-10-05 Rolls-Royce Plc Abradable composition, a method of manufacturing an abradable composition and a gas turbine engine having an abradable seal
US6776573B2 (en) * 2000-11-30 2004-08-17 Snecma Moteurs Bladed rotor disc side-plate and corresponding arrangement
US6899339B2 (en) * 2001-08-30 2005-05-31 United Technologies Corporation Abradable seal having improved durability
US7040857B2 (en) * 2004-04-14 2006-05-09 General Electric Company Flexible seal assembly between gas turbine components and methods of installation

Family Cites Families (26)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3023998A (en) * 1959-03-13 1962-03-06 Jr Walter H Sanderson Rotor blade retaining device
US3146938A (en) * 1962-12-28 1964-09-01 Gen Electric Shrouding for compressor stator vanes
US3262635A (en) * 1964-11-06 1966-07-26 Gen Electric Turbomachine sealing means
GB1318654A (en) * 1970-12-05 1973-05-31 Secr Defence Bladed rotors
US3841792A (en) * 1973-03-09 1974-10-15 Westinghouse Electric Corp Turbomachine blade lock and seal device
US4084919A (en) * 1976-06-29 1978-04-18 United Technologies Corporation Means of attaching a seal to a disk
GB2070700B (en) * 1980-03-01 1983-10-05 Rolls Royce Gas turbine seals
US4747750A (en) * 1986-01-17 1988-05-31 United Technologies Corporation Transition duct seal
US4867639A (en) * 1987-09-22 1989-09-19 Allied-Signal Inc. Abradable shroud coating
US4813848A (en) * 1987-10-14 1989-03-21 United Technologies Corporation Turbine rotor disk and blade assembly
US5228195A (en) * 1990-09-25 1993-07-20 United Technologies Corporation Apparatus and method for a stator assembly of a rotary machine
US5217348A (en) * 1992-09-24 1993-06-08 United Technologies Corporation Turbine vane assembly with integrally cast cooling fluid nozzle
DE4232385A1 (en) * 1992-09-26 1994-03-31 Asea Brown Boveri Gas turbine with flanged exhaust housing
US5522698A (en) * 1994-04-29 1996-06-04 United Technologies Corporation Brush seal support and vane assembly windage cover
GB2294732A (en) * 1994-11-05 1996-05-08 Rolls Royce Plc Integral disc seal for turbomachine
JP3416447B2 (en) * 1997-03-11 2003-06-16 三菱重工業株式会社 Gas turbine blade cooling air supply system
JPH10252412A (en) * 1997-03-12 1998-09-22 Mitsubishi Heavy Ind Ltd Gas turbine sealing device
JPH10259703A (en) * 1997-03-18 1998-09-29 Mitsubishi Heavy Ind Ltd Shroud for gas turbine and platform seal system
JP3327814B2 (en) * 1997-06-18 2002-09-24 三菱重工業株式会社 Gas turbine sealing device
US6220815B1 (en) * 1999-12-17 2001-04-24 General Electric Company Inter-stage seal retainer and assembly
JP4494658B2 (en) * 2001-02-06 2010-06-30 三菱重工業株式会社 Gas turbine stationary blade shroud
US6887039B2 (en) * 2002-07-10 2005-05-03 Mitsubishi Heavy Industries, Ltd. Stationary blade in gas turbine and gas turbine comprising the same
US6837676B2 (en) * 2002-09-11 2005-01-04 Mitsubishi Heavy Industries, Ltd. Gas turbine
US7229247B2 (en) * 2004-08-27 2007-06-12 Pratt & Whitney Canada Corp. Duct with integrated baffle
US7909569B2 (en) * 2005-06-09 2011-03-22 Pratt & Whitney Canada Corp. Turbine support case and method of manufacturing
US20080061515A1 (en) * 2006-09-08 2008-03-13 Eric Durocher Rim seal for a gas turbine engine

Patent Citations (13)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US2564503A (en) * 1949-01-29 1951-08-14 Gen Electric Baffle means for rotating bodies
US3606349A (en) * 1969-08-07 1971-09-20 Rolls Royce Seal with pressure responsive tolerance control
US3755870A (en) * 1971-12-20 1973-09-04 Entoleter Pressure retaining face seal and method
US3908361A (en) * 1972-12-16 1975-09-30 Rolls Royce 1971 Ltd Seal between relatively moving components of a fluid flow machine
US3918925A (en) * 1974-05-13 1975-11-11 United Technologies Corp Abradable seal
US4320903A (en) * 1978-09-27 1982-03-23 Societe Nationale D'etude Et De Construction De Moteurs D'aviation Labyrinth seals
US4309145A (en) * 1978-10-30 1982-01-05 General Electric Company Cooling air seal
US5575486A (en) * 1992-10-01 1996-11-19 Atlas Copco Tools Ab Grooved run-in face seal
US5962076A (en) * 1995-06-29 1999-10-05 Rolls-Royce Plc Abradable composition, a method of manufacturing an abradable composition and a gas turbine engine having an abradable seal
US5785492A (en) * 1997-03-24 1998-07-28 United Technologies Corporation Method and apparatus for sealing a gas turbine stator vane assembly
US6776573B2 (en) * 2000-11-30 2004-08-17 Snecma Moteurs Bladed rotor disc side-plate and corresponding arrangement
US6899339B2 (en) * 2001-08-30 2005-05-31 United Technologies Corporation Abradable seal having improved durability
US7040857B2 (en) * 2004-04-14 2006-05-09 General Electric Company Flexible seal assembly between gas turbine components and methods of installation

Cited By (17)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US20090208326A1 (en) * 2006-09-08 2009-08-20 Eric Durocher Rim seal for a gas turbine engine
US8172514B2 (en) 2006-09-08 2012-05-08 Pratt & Whitney Canada Corp. Rim seal for a gas turbine engine
US20110023499A1 (en) * 2006-09-15 2011-02-03 Nicolas Grivas Gas turbine combustor exit duct and hp vane interface
US8166767B2 (en) * 2006-09-15 2012-05-01 Pratt & Whitney Canada Corp. Gas turbine combustor exit duct and hp vane interface
US20090110548A1 (en) * 2007-10-30 2009-04-30 Pratt & Whitney Canada Corp. Abradable rim seal for low pressure turbine stage
EP2239422A3 (en) * 2009-04-06 2017-05-24 General Electric Company Seal for a gas turbine engine
CN101858230A (en) * 2009-04-06 2010-10-13 通用电气公司 Method, system and/or device about the sealed department that is used for turbogenerator
US9068469B2 (en) 2011-09-01 2015-06-30 Honeywell International Inc. Gas turbine engines with abradable turbine seal assemblies
WO2015181489A1 (en) * 2014-05-27 2015-12-03 Snecma Sealing plate with fuse function
US10047621B2 (en) 2014-05-27 2018-08-14 Safran Aircraft Engines Sealing plate with fuse function
US20160076454A1 (en) * 2014-09-16 2016-03-17 Alstom Technology Ltd Sealing arrangement at the interface between a combustor and a turbine of a gas turbine and gas turbine with such a sealing arrangement
US10393025B2 (en) * 2014-09-16 2019-08-27 Ansaldo Energia Switzerland AG Sealing arrangement at the interface between a combustor and a turbine of a gas turbine and gas turbine with such a sealing arrangement
US9803496B2 (en) 2015-07-01 2017-10-31 United Technologies Corporation Break-in system for gapping and leakage control
EP3112602A1 (en) * 2015-07-01 2017-01-04 United Technologies Corporation Break-in system for gapping and leakage control
FR3039225A1 (en) * 2015-07-20 2017-01-27 Snecma TURBOMACHINE, SUCH AS A TURBO AIRCRAFT
US10633992B2 (en) 2017-03-08 2020-04-28 Pratt & Whitney Canada Corp. Rim seal
FR3080646A1 (en) * 2018-04-26 2019-11-01 Safran Aircraft Engines SEALING BETWEEN A FIXED WHEEL AND A MOBILE WHEEL OF A TURBOMACHINE

Also Published As

Publication number Publication date
US8172514B2 (en) 2012-05-08
US20090208326A1 (en) 2009-08-20
CA2598329C (en) 2015-03-24
CA2598329A1 (en) 2008-03-08

Similar Documents

Publication Publication Date Title
US8172514B2 (en) Rim seal for a gas turbine engine
EP2369138B1 (en) Gas turbine engine with non-axisymmetric surface contoured vane platform
EP3594452B1 (en) Seal segment for a gas turbine engine
CA2552214C (en) Blades for a gas turbine engine with integrated sealing plate and method
US8684680B2 (en) Sealing and cooling at the joint between shroud segments
US9631557B2 (en) Sealing device for a turbomachine turbine nozzle
US20060182633A1 (en) Turbine blade
US20030206799A1 (en) Casing section
CN109519224B (en) Gas turbine engine including turbine rotor assembly
US20180230839A1 (en) Turbine engine shroud assembly
JP6208922B2 (en) Blade used with a rotating machine and method for assembling such a rotating machine
US20130119617A1 (en) Turbomachinery seal
US11585230B2 (en) Assembly for a turbomachine
US9945240B2 (en) Power turbine heat shield architecture
EP2636852B1 (en) Hybrid inner air seal for gas turbine engines
EP3450685B1 (en) Gas turbine engine component
US7534085B2 (en) Gas turbine engine with contoured air supply slot in turbine rotor
EP3553279B1 (en) Blade outer air seal cooling fin
EP3000966B1 (en) Method and assembly for reducing secondary heat in a gas turbine engine
US20200217214A1 (en) Rim seal
EP2995778B1 (en) Method and assembly for reducing secondary heat in a gas turbine engine
US11939880B1 (en) Airfoil assembly with flow surface
US20090060736A1 (en) Compressor
CN112302732A (en) Inner band for a turbine engine

Legal Events

Date Code Title Description
AS Assignment

Owner name: PRATT & WHITNEY CANADA CORP., CANADA

Free format text: ASSIGNMENT OF ASSIGNORS INTEREST;ASSIGNORS:DUROCHER, ERIC;PAQUET, RENE;LEFEBVRE, GUY;REEL/FRAME:018221/0730

Effective date: 20060908

STCB Information on status: application discontinuation

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