US3423070A - Sealing means for turbomachinery - Google Patents

Sealing means for turbomachinery Download PDF

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Publication number
US3423070A
US3423070A US596613A US3423070DA US3423070A US 3423070 A US3423070 A US 3423070A US 596613 A US596613 A US 596613A US 3423070D A US3423070D A US 3423070DA US 3423070 A US3423070 A US 3423070A
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United States
Prior art keywords
seal member
layer
shroud
tips
core layer
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Expired - Lifetime
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US596613A
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English (en)
Inventor
Charles E Corrigan
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General Electric Co
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General Electric Co
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Publication date
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Publication of US3423070A publication Critical patent/US3423070A/en
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    • 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
    • F01D11/127—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 a deformable or crushable structure, e.g. honeycomb
    • 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

  • This invention relates to a sealing means for turbomachinery and, more particularly, to improved shroud means for preventing undesired leakage around the tips of rotating airfoils during turbomachine operation.
  • turbomachines such as axial flow compressors and turbines
  • the overall operating eliiciency is adversely affected by leakage of the working fluid around the tips of the rotating airfoils.
  • leakage of the compressed fluid around the blade tips to lowerpressure regions results in -a loss to the system ⁇ of the energy utilized in initially compressing the fluid since it must then be recompressed.
  • leakage of high temperature iiuid around the tips of turbine buckets also results in an energy loss to the system since the fluid performs no useful work on the turbine.
  • the overall operating efficiency is an extremely important operating parameter, and it is therefore very desirable that this tip leakage be prevented or at least maintained at an acceptably low level.
  • abradable shrouds have been proposed and used in the past, these arrangements usually providing smaller clearances since there is no need to anticipate and provide for all situations whic'h might conceivably cause rubbing. While providing smaller clearances, abradable shrouds available heretofore have tended to display performance losses due to surface roughness generating high turbulence on the shroud surface. Efforts have been made to provide abradable shrouds having the more desirable aerodynamic characteristics of non-abradable shrouds, but these configurations have not been altogether satisfactory with respect to strength, durability, and structural integrity.
  • Another object of this invention is to provide for turbomachines an improved shroud means for preventing leakage of working iiuid around the tips of rotating airfoils.
  • Yet another object is to provide for turbomachines an improved shroud means having the small clearances normally associated with abradable shrouds in combination with the aerodynamic char-acteristics of non-abradable shrouds.
  • a still further object of this invention is to provide the foregoing objects with a shroud arrangement having suicient strength, durability, and structural integrity for sustained operation in high performance turbomachines such as aircraft gas turbine engines.
  • a turbomachine shroud or similar sealing device includes a composite seal member of honeycomb sandwich construction. More particularly, the composite seal member includes a core layer of honeycomb material having a pair of substantially parallel faces, the individual cells comprising the core layer being disposed to intersect the pair of intersecting faces and thereby form a large number of open passages extending through the core layer. A layer of imperforate material is secured to each of the parallel faces to close the individual -cell passages to prevent flow into and through the cells, the layer of imperforate material secured to at least one of the faces being relatively smooth and thereby adapted for use as an aerodynamically smooth sealing surface in conjunction with mating seal members, including the tips of rotating airfoils.
  • both the core layer of honeycomb and the smooth layer of imperforate material are of really abradable construction such that any rubbing will result in preferential abrading of the composite seal member. Accordingly, the composite seal member can be mounted with relatively small clearances with respect to the mating seal members.
  • both the core layer of honeycomb material and the smooth layer of imperforate material are comprised of relatively thin and easily abraded sheet metal, the smooth cover layer of sheet metal preferably having a thickness in the range of 0.003 to 0.006 inch.
  • a shroud or other sealing device formed in accordance with this invention may be comprised of a single composite seal member or a plurality of cooperating seal members, such as arcuate shroud segments forming a complete annular shroud.
  • FIG. l is a view partially in section of a gas turbine engine having a turbine shroud constructed in accordance with the present invention
  • FIG. 2 is an enlarged view of the turbine shroud arrangement illustrated by FIG. 1;
  • FIG. 3 is a view taken along line 3 3 of FIG. 2 illustrating one of the arcuate shroud segments comprising the turbine shroud;
  • FIG. 4 is a pictorial view, partially cut-away, of the shroud segment of FIG. 3;
  • FIG. 5 is an end view, partially cut-away of Ia modified shroud segment in which the honeycomb core is encased in sheet metal.
  • an axial iiow gas turbine engine 10 of the turbojet type is illustrated, the engine having an outer cylindrical casing 11 circumferentially surrounding an axial flow compressor 12, an annular combustor 13, and a turbine 14 axially disposed in serial ow relationship between a compressor inlet 15 and an exhaust nozzle 16. More particularly, the engine components just described cooperate to form an annular passageway extending axially between the inlet and the nozzle 16 for the flow of motive fluid, which is initially air and later combustion products.
  • the air ows through a number of axially space-d-apart .and alternating rows of rotor blades 20 and stator vanes 21, each adjacent pair of rotor blades 20 and stator vanes 21 comprising a compression stage for increasing the pressure of the air.
  • annular shroud assemblies 23 are provided in close running clearance with the blade tips; similarly, annular seal assemblies 24 are provided for preventing undesired leakage around the inner tips of the stator vanes 21.
  • an annular nozzle diaphragm 26 is located at the downstream end of the combustor 13 for supplying the combustion gases to .a first stage row of turbine buckets 27 at the proper velocity and angle. From the turbine buckets 27, the hot gases flow through a second stage turbine nozzle ⁇ diaphragm 28 from which they are redirected to a second stage row of turbine buckets 29.
  • the turbine buckets 27 are peripherally mounted on a turbine wheel 32 which, along with its associated shaft 33 and a second turbine wheel 34 upon which the turbine buckets 29 are mounted, is rotatably mounted on the engine axis 35 by suitable mounting means including bearing arrangements 36.
  • the turbine unit compris-ing the wheels 32 and 34 and the shaft 33 drives the compressor rotor 38 upon which the rotor blades 20 .are mounted.
  • annular shroud assemblies 40 and 41 are provided at the outer tips of the turbine buckets 27 and 29, respectively, andan annular seal assembly 42 is provided at the inner end of the nozzle diaphragm 28.
  • the annular shroud assembly 40 surrounding the tips 27 of the turbine buckets 27 is formed in accordance with the present invention and is illustrated in detail by FIGS. 2-4.
  • the shroud assembly 40 includes a number of arcuate seal members 44 which .abut to form a complete annular shroud ring surrounding the tips 27 with a relatively small clearance C.
  • the size of this clearance C is selected such that rubbing will not occur between the tips 27 and the arcuate seal members 44 under ordinary operating conditions; the clearance C need not, however, be large enough to assure that rubbing never occurs since the shroud assembly of this invention can accommodate rubs.
  • the arcuate seal members 44 are composite structures including a core layer 45 of open-celled honeycomb material, the individual cells 46 comprising the core layer 45 being radially disposed with respect to the engine axis 35 to intersect inner and outer substantially cylindrical faces 47 and 48, respectively, of the layer 45. In other words, the individual cells 46 thus form open passages between the faces 47 and 48.
  • This core layer 45 of honeycomb material is of abradable construction and is preferably formed of expanded sheet metal as described in United States Patent 2,963,307 to Bobo, issued Dec. 6, 1960, and assigned to the assignee of the present invention.
  • a backing layer 50 of imperforate material such as metal is secured to the outer face 48 to close the outer ends of the cells 46, and support means such as the flanges 52 and 53 are provided for cooperating with casing flanges 54 and 55 to support the seal members 44 within the casing 11 with the proper clearance C.
  • each of the composite sealing members 44 has an inner layer 58 of imperforate material secured to the inner face 47 of the core layer 45.
  • the inner layer 58 is both smooth and abradable.
  • this inner layer 58 may be comprised of relatively thin sheet metal, preferably having a thickness in the range of 0.003 to 0.006 inch.
  • the material from which the inner layer 58 is formed should, of course, possess any other necessary characteristics for satisfactory use in its -operating environment, including suitable resistance to normal operating temperatures.
  • the clearance C between the arcuate seal members 44 and the bucket tips 27 will be quite small, and the leakage around the tips Will be correspondingly small.
  • the smoot-h surface provided by the inner layer 58 will maintain turbulence and accompanying areodynamic losses at reasonably low levels.
  • the arcuate seal members 44 provide extremely satisfactory leakage and aerodynamic char- ⁇ acteristics under ordinary operating conditions. Under unusual conditions where the relative growth between the bucket tips 27 and the shroud assembly 40 is greater than anticipated, rubbing can occur with preferential abrading of the inner layer 58 of the arcuate seal member 44 and possibly the core layer 45 since these elements have substantially less mass than the bucket tips 27.
  • the backing and inner layers 50 ⁇ and 58, respectively, of imperforate material are separate and distinct elements. It may be desirable, however, to provide additional strength and durability by encasing the entire core layer 45 in imperforate material as illustrated by FIG. 5, the inner layer 58 being wrapped around the sides and ends of the core layer 45' and secured by brazing or other means at 60 to the backing layer 50. Similarly, it may be desirable to wrap the entire core layer with a single sheet of sheet metal to provide both the inner rubbing layer and the backing layer.
  • shroud assemblies of FIGS. 2-5 are preferably formed of arcuate seal memybers, it would be quite possible to provide an arcuate seal member that is a complete, unsegmented ring having the composite structure of this invention.
  • seal construction of this invention has been described only in conjunction with the shroud assembly 40 for the turbine buckets 27, it will be appreciated that the invention could also be utilized with respect to the shroud assemblies 41 and 23 and the stator seal assemblies 42 and 24.
  • seal members could be made in accordance with this invention in which the faces are substantially flat, parallel surfaces or inclined surfaces. Accordingly, the term parallel faces as used in the claims appended hereto is hereby defined to means surfaces that are spaced apart a substantially uniform distance; as such, the surfaces may be flat or curved.
  • this invention provides improved sealing means capable of maintaining small seal clearances in combination with a high degree of aerodynamic eiciency. More specifically, as applied to shrouds for rows of airfoils, the invention provides in combination the small clearances normally associated with abradable shrouds and the aerodynamic characteristics of nonabradable shrouds.
  • a composite seal member comprising:
  • the individual cells comprising said core layer of honeycom'b material being disposed to intersect said pair of parallel faces
  • the layer of imperforate material secured to at least one of said parallel faces being relatively smooth and thereby adapted for use as a sealing surface in conjunction with a mating seal member, said smooth layer of imperforate material and said core layer of honeycomb material being readily abradable, whereby any rubbing between said composite seal member and a mating seal ymember will result in preferential abrading of said composite seal member.
  • a composite seal member as defined by claim 1 in which said core layer of honeycomb material is comprised of sheet metal and in which the layers of imperforate material secured to said parallel faces are comprised of sheet metal substantially encasing the entire core layer of honeycomb material, the sheet metal being relatively thin so as to be readily abraded in the event of rubbing between said composite seal member and a mating seal member.
  • a composite seal member as defined by claim 1 in which said core layer of honeycomb material is comprised of sheet metal and in which the layers of imperforate material secured to said Parallel faces cornprise separate sheets of sheet metal secured to said respective faces, the sheet metal being relatively thin so as to be readily abraded in the event of rubbing between said composite seal member and a mating seal member.
  • a composite seal member as defined by claim 4 having a generally arcuate configuration for use in a turbomachine, said parallel faces being curved surfaces spacedapart radially a substantiallyruniform distance.
  • annular shroud comprising at least one composite seal member as defined by claim 5, the parallel faces of said composite seal member being of substantially cylindrical configuration and said smooth layer of sheet metal being secured to the radially inner one of said faces.
  • a turbomachine including a rotor mounted for rotation about an axis and a row of radial airfoils peripherally mounted on said rotor, an annular shroud assembly peripherally surrounding said airfoils, said annular shroud assembly comprising:
  • At least one arcuate composite seal member At least one arcuate composite seal member
  • said composite seal member comprising:
  • a core layer of honeycomb material having inner and outer radially spaced, substantially cylindrical faces
  • the individual cells comprising said core layer of honeycomb material being radially disposed to intersect said inner and outer faces
  • the layer of imperforate material secured to said inner face being relatively smooth

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  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
  • Sealing Using Fluids, Sealing Without Contact, And Removal Of Oil (AREA)
  • Gasket Seals (AREA)
US596613A 1966-11-23 1966-11-23 Sealing means for turbomachinery Expired - Lifetime US3423070A (en)

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
US59661366A 1966-11-23 1966-11-23

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US3423070A true US3423070A (en) 1969-01-21

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US (1) US3423070A (de)
BE (1) BE701461A (de)
CH (1) CH473307A (de)
DE (1) DE1551183A1 (de)
GB (1) GB1175287A (de)
NL (1) NL6710449A (de)

Cited By (33)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3529905A (en) * 1966-12-12 1970-09-22 Gen Motors Corp Cellular metal and seal
US3603599A (en) * 1970-05-06 1971-09-07 Gen Motors Corp Cooled seal
US3846899A (en) * 1972-07-28 1974-11-12 Gen Electric A method of constructing a labyrinth seal
US4135851A (en) * 1977-05-27 1979-01-23 The United States Of America As Represented By The Administrator Of The National Aeronautics And Space Administration Composite seal for turbomachinery
US4207024A (en) * 1977-05-27 1980-06-10 The United States Of America As Represented By The Administrator Of The National Aeronautics And Space Administration Composite seal for turbomachinery
US4318666A (en) * 1979-07-12 1982-03-09 Rolls-Royce Limited Cooled shroud for a gas turbine engine
FR2502696A1 (fr) * 1981-03-25 1982-10-01 Rolls Royce Moteur a turbine a gaz presentant une immunite accrue aux dommages causes par l'ingestion de corps etrangers
EP0081010A2 (de) * 1981-08-10 1983-06-15 George T. Straza Unterbrochene Wabenstruktur
US4478552A (en) * 1982-11-08 1984-10-23 Thompson Stanley E Method and apparatus for fan blade tip clearance
US4497610A (en) * 1982-03-23 1985-02-05 Rolls-Royce Limited Shroud assembly for a gas turbine engine
US4669955A (en) * 1980-08-08 1987-06-02 Rolls-Royce Plc Axial flow turbines
US4767267A (en) * 1986-12-03 1988-08-30 General Electric Company Seal assembly
DE4110244A1 (de) * 1990-04-03 1991-10-10 Gen Electric Zwischenstufendichtungsanordnung fuer schaufelstufen von gegenlaeufigen turbinentriebwerksrotoren
US5165848A (en) * 1991-07-09 1992-11-24 General Electric Company Vane liner with axially positioned heat shields
US5169159A (en) * 1991-09-30 1992-12-08 General Electric Company Effective sealing device for engine flowpath
US5174714A (en) * 1991-07-09 1992-12-29 General Electric Company Heat shield mechanism for turbine engines
US5195868A (en) * 1991-07-09 1993-03-23 General Electric Company Heat shield for a compressor/stator structure
US5292138A (en) * 1992-09-21 1994-03-08 General Elecric Company Rotor to rotor split ring seal
US5314304A (en) * 1991-08-15 1994-05-24 The United States Of America As Represented By The Secretary Of The Air Force Abradeable labyrinth stator seal
US5524846A (en) * 1993-12-21 1996-06-11 The Boeing Company Fire protection system for airplanes
US6382905B1 (en) * 2000-04-28 2002-05-07 General Electric Company Fan casing liner support
GB2382380A (en) * 2001-11-24 2003-05-28 Rolls Royce Plc A removable abradable lining for the casing assembly of a gas turbine engine
DE102004018585A1 (de) * 2004-04-16 2005-12-01 Rolls-Royce Deutschland Ltd & Co Kg Triebwerkeinlaufkonus für ein Gasturbinentriebwerk
US20090053045A1 (en) * 2007-08-22 2009-02-26 General Electric Company Turbine Shroud for Gas Turbine Assemblies and Processes for Forming the Shroud
FR2922950A1 (fr) * 2007-10-31 2009-05-01 Snecma Sa Cartouche de materiau abradable
US20090263239A1 (en) * 2004-03-03 2009-10-22 Mtu Aero Engines Gmbh Ring structure with a metal design having a run-in lining
US20110206502A1 (en) * 2010-02-25 2011-08-25 Samuel Ross Rulli Turbine shroud support thermal shield
FR2979664A1 (fr) * 2011-09-01 2013-03-08 Snecma Paroi stator de turbomachine recouverte d'un revetement abradable a faible rugosite aerodynamique
EP2872763A4 (de) * 2012-07-16 2015-07-15 United Technologies Corp Äussere luftdichtung für eine turbinenschaufel mit kühlkonfiguration
US20170089213A1 (en) * 2015-09-28 2017-03-30 United Technologies Corporation Duct with additive manufactured seal
US10472980B2 (en) * 2017-02-14 2019-11-12 General Electric Company Gas turbine seals
US11674405B2 (en) 2021-08-30 2023-06-13 General Electric Company Abradable insert with lattice structure
US11674396B2 (en) 2021-07-30 2023-06-13 General Electric Company Cooling air delivery assembly

Families Citing this family (3)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
WO1995021319A1 (en) * 1994-02-01 1995-08-10 United Technologies Corporation Honeycomb abradable seals
DE10020673C2 (de) * 2000-04-27 2002-06-27 Mtu Aero Engines Gmbh Ringstruktur in Metallbauweise
US20160215646A1 (en) * 2013-09-06 2016-07-28 General Electric Company Gas turbine laminate seal assembly comprising first and second honeycomb layer and a perforated intermediate seal plate in-between

Citations (12)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
GB450524A (en) * 1934-10-15 1935-04-23 Andre Abel Auguste Brugier Improvements in or relating to heat-insulating panels
US2477852A (en) * 1945-07-04 1949-08-02 Owens Corning Fiberglass Corp Structural panel construction
US2952442A (en) * 1957-05-28 1960-09-13 Studebaker Packard Corp Rotating shroud
US2963307A (en) * 1954-12-28 1960-12-06 Gen Electric Honeycomb seal
US3053694A (en) * 1961-02-20 1962-09-11 Gen Electric Abradable material
US3068016A (en) * 1958-03-31 1962-12-11 Gen Motors Corp High temperature seal
US3126149A (en) * 1964-03-24 Foamed aluminum honeycomb motor
US3146992A (en) * 1962-12-10 1964-09-01 Gen Electric Turbine shroud support structure
US3241813A (en) * 1964-01-21 1966-03-22 Garrett Corp Turbine wheel burst containment means
US3291382A (en) * 1964-05-08 1966-12-13 Rolls Royce Bladed structure, for example, for a gas turbine engine compressor
US3339933A (en) * 1965-02-24 1967-09-05 Gen Electric Rotary seal
US3365172A (en) * 1966-11-02 1968-01-23 Gen Electric Air cooled shroud seal

Patent Citations (12)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3126149A (en) * 1964-03-24 Foamed aluminum honeycomb motor
GB450524A (en) * 1934-10-15 1935-04-23 Andre Abel Auguste Brugier Improvements in or relating to heat-insulating panels
US2477852A (en) * 1945-07-04 1949-08-02 Owens Corning Fiberglass Corp Structural panel construction
US2963307A (en) * 1954-12-28 1960-12-06 Gen Electric Honeycomb seal
US2952442A (en) * 1957-05-28 1960-09-13 Studebaker Packard Corp Rotating shroud
US3068016A (en) * 1958-03-31 1962-12-11 Gen Motors Corp High temperature seal
US3053694A (en) * 1961-02-20 1962-09-11 Gen Electric Abradable material
US3146992A (en) * 1962-12-10 1964-09-01 Gen Electric Turbine shroud support structure
US3241813A (en) * 1964-01-21 1966-03-22 Garrett Corp Turbine wheel burst containment means
US3291382A (en) * 1964-05-08 1966-12-13 Rolls Royce Bladed structure, for example, for a gas turbine engine compressor
US3339933A (en) * 1965-02-24 1967-09-05 Gen Electric Rotary seal
US3365172A (en) * 1966-11-02 1968-01-23 Gen Electric Air cooled shroud seal

Cited By (40)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3529905A (en) * 1966-12-12 1970-09-22 Gen Motors Corp Cellular metal and seal
US3603599A (en) * 1970-05-06 1971-09-07 Gen Motors Corp Cooled seal
US3846899A (en) * 1972-07-28 1974-11-12 Gen Electric A method of constructing a labyrinth seal
US4135851A (en) * 1977-05-27 1979-01-23 The United States Of America As Represented By The Administrator Of The National Aeronautics And Space Administration Composite seal for turbomachinery
US4207024A (en) * 1977-05-27 1980-06-10 The United States Of America As Represented By The Administrator Of The National Aeronautics And Space Administration Composite seal for turbomachinery
US4318666A (en) * 1979-07-12 1982-03-09 Rolls-Royce Limited Cooled shroud for a gas turbine engine
US4669955A (en) * 1980-08-08 1987-06-02 Rolls-Royce Plc Axial flow turbines
DE3209960A1 (de) * 1981-03-25 1982-10-14 Rolls-Royce Ltd., London Gasturbinentriebwerk
FR2502696A1 (fr) * 1981-03-25 1982-10-01 Rolls Royce Moteur a turbine a gaz presentant une immunite accrue aux dommages causes par l'ingestion de corps etrangers
EP0081010A2 (de) * 1981-08-10 1983-06-15 George T. Straza Unterbrochene Wabenstruktur
EP0081010A3 (de) * 1981-08-10 1983-08-03 George T. Straza Unterbrochene Wabenstruktur
US4497610A (en) * 1982-03-23 1985-02-05 Rolls-Royce Limited Shroud assembly for a gas turbine engine
US4478552A (en) * 1982-11-08 1984-10-23 Thompson Stanley E Method and apparatus for fan blade tip clearance
US4767267A (en) * 1986-12-03 1988-08-30 General Electric Company Seal assembly
DE4110244A1 (de) * 1990-04-03 1991-10-10 Gen Electric Zwischenstufendichtungsanordnung fuer schaufelstufen von gegenlaeufigen turbinentriebwerksrotoren
US5195868A (en) * 1991-07-09 1993-03-23 General Electric Company Heat shield for a compressor/stator structure
US5165848A (en) * 1991-07-09 1992-11-24 General Electric Company Vane liner with axially positioned heat shields
US5174714A (en) * 1991-07-09 1992-12-29 General Electric Company Heat shield mechanism for turbine engines
US5314304A (en) * 1991-08-15 1994-05-24 The United States Of America As Represented By The Secretary Of The Air Force Abradeable labyrinth stator seal
US5169159A (en) * 1991-09-30 1992-12-08 General Electric Company Effective sealing device for engine flowpath
US5292138A (en) * 1992-09-21 1994-03-08 General Elecric Company Rotor to rotor split ring seal
US5524846A (en) * 1993-12-21 1996-06-11 The Boeing Company Fire protection system for airplanes
US6382905B1 (en) * 2000-04-28 2002-05-07 General Electric Company Fan casing liner support
GB2382380A (en) * 2001-11-24 2003-05-28 Rolls Royce Plc A removable abradable lining for the casing assembly of a gas turbine engine
US20090263239A1 (en) * 2004-03-03 2009-10-22 Mtu Aero Engines Gmbh Ring structure with a metal design having a run-in lining
US8061965B2 (en) * 2004-03-03 2011-11-22 Mtu Aero Engines Gmbh Ring structure of metal construction having a run-in lining
DE102004018585A1 (de) * 2004-04-16 2005-12-01 Rolls-Royce Deutschland Ltd & Co Kg Triebwerkeinlaufkonus für ein Gasturbinentriebwerk
DE102004018585B4 (de) * 2004-04-16 2013-12-05 Rolls-Royce Deutschland Ltd & Co Kg Triebwerkeinlaufkonus für ein Gasturbinentriebwerk
US20090053045A1 (en) * 2007-08-22 2009-02-26 General Electric Company Turbine Shroud for Gas Turbine Assemblies and Processes for Forming the Shroud
FR2922950A1 (fr) * 2007-10-31 2009-05-01 Snecma Sa Cartouche de materiau abradable
US20110206502A1 (en) * 2010-02-25 2011-08-25 Samuel Ross Rulli Turbine shroud support thermal shield
FR2979664A1 (fr) * 2011-09-01 2013-03-08 Snecma Paroi stator de turbomachine recouverte d'un revetement abradable a faible rugosite aerodynamique
US9574455B2 (en) 2012-07-16 2017-02-21 United Technologies Corporation Blade outer air seal with cooling features
EP2872763A4 (de) * 2012-07-16 2015-07-15 United Technologies Corp Äussere luftdichtung für eine turbinenschaufel mit kühlkonfiguration
US10323534B2 (en) 2012-07-16 2019-06-18 United Technologies Corporation Blade outer air seal with cooling features
US20170089213A1 (en) * 2015-09-28 2017-03-30 United Technologies Corporation Duct with additive manufactured seal
US11459905B2 (en) 2015-09-28 2022-10-04 Raytheon Technologies Corporation Duct with additive manufactured seal
US10472980B2 (en) * 2017-02-14 2019-11-12 General Electric Company Gas turbine seals
US11674396B2 (en) 2021-07-30 2023-06-13 General Electric Company Cooling air delivery assembly
US11674405B2 (en) 2021-08-30 2023-06-13 General Electric Company Abradable insert with lattice structure

Also Published As

Publication number Publication date
GB1175287A (en) 1969-12-23
DE1551183A1 (de) 1970-04-16
NL6710449A (de) 1968-05-24
BE701461A (de) 1968-01-02
CH473307A (de) 1969-05-31

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