GB2496887A - Gas turbine engine abradable liner - Google Patents

Gas turbine engine abradable liner Download PDF

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Publication number
GB2496887A
GB2496887A GB1120338.7A GB201120338A GB2496887A GB 2496887 A GB2496887 A GB 2496887A GB 201120338 A GB201120338 A GB 201120338A GB 2496887 A GB2496887 A GB 2496887A
Authority
GB
United Kingdom
Prior art keywords
text
abradable
liner
gas turbine
matrix
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Withdrawn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
GB1120338.7A
Other versions
GB201120338D0 (en
Inventor
Nicholas Michael Merriman
Allan Frederick Collard
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.)
Rolls Royce PLC
Original Assignee
Rolls Royce PLC
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 Rolls Royce PLC filed Critical Rolls Royce PLC
Priority to GB1120338.7A priority Critical patent/GB2496887A/en
Publication of GB201120338D0 publication Critical patent/GB201120338D0/en
Priority to PCT/EP2012/072228 priority patent/WO2013075954A1/en
Publication of GB2496887A publication Critical patent/GB2496887A/en
Withdrawn legal-status Critical Current

Links

Classifications

    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F04—POSITIVE - DISPLACEMENT MACHINES FOR LIQUIDS; PUMPS FOR LIQUIDS OR ELASTIC FLUIDS
    • F04D—NON-POSITIVE-DISPLACEMENT PUMPS
    • F04D29/00—Details, component parts, or accessories
    • F04D29/40—Casings; Connections of working fluid
    • F04D29/52—Casings; Connections of working fluid for axial pumps
    • F04D29/522—Casings; Connections of working fluid for axial pumps especially adapted for elastic fluid pumps
    • F04D29/526—Details of the casing section radially opposing blade tips
    • 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
    • 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
    • F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
    • F01D5/12—Blades
    • F01D5/14—Form or construction
    • F01D5/20—Specially-shaped blade tips to seal space between tips and stator
    • 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
    • F02C7/04—Air intakes for gas-turbine plants or jet-propulsion plants
    • F02C7/05—Air intakes for gas-turbine plants or jet-propulsion plants having provisions for obviating the penetration of damaging objects or particles
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02K—JET-PROPULSION PLANTS
    • F02K9/00—Rocket-engine plants, i.e. plants carrying both fuel and oxidant therefor; Control thereof
    • F02K9/08—Rocket-engine plants, i.e. plants carrying both fuel and oxidant therefor; Control thereof using solid propellants
    • F02K9/32—Constructional parts; Details not otherwise provided for
    • F02K9/34—Casings; Combustion chambers; Liners thereof
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02K—JET-PROPULSION PLANTS
    • F02K9/00—Rocket-engine plants, i.e. plants carrying both fuel and oxidant therefor; Control thereof
    • F02K9/08—Rocket-engine plants, i.e. plants carrying both fuel and oxidant therefor; Control thereof using solid propellants
    • F02K9/32—Constructional parts; Details not otherwise provided for
    • F02K9/34—Casings; Combustion chambers; Liners thereof
    • F02K9/346—Liners, e.g. inhibitors
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F04—POSITIVE - DISPLACEMENT MACHINES FOR LIQUIDS; PUMPS FOR LIQUIDS OR ELASTIC FLUIDS
    • F04D—NON-POSITIVE-DISPLACEMENT PUMPS
    • F04D29/00—Details, component parts, or accessories
    • F04D29/02—Selection of particular materials
    • F04D29/023—Selection of particular materials especially adapted for elastic fluid pumps
    • 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
    • F05D2300/00—Materials; Properties thereof
    • F05D2300/20—Oxide or non-oxide ceramics
    • F05D2300/22—Non-oxide ceramics
    • F05D2300/224—Carbon, e.g. graphite
    • 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
    • F05D2300/00—Materials; Properties thereof
    • F05D2300/20—Oxide or non-oxide ceramics
    • F05D2300/22—Non-oxide ceramics
    • F05D2300/228—Nitrides
    • F05D2300/2282—Nitrides of boron
    • 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
    • F05D2300/00—Materials; Properties thereof
    • F05D2300/20—Oxide or non-oxide ceramics
    • F05D2300/22—Non-oxide ceramics
    • F05D2300/229—Sulfides
    • F05D2300/2291—Sulfides of molybdenum
    • 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
    • F05D2300/00—Materials; Properties thereof
    • F05D2300/40—Organic materials
    • F05D2300/43—Synthetic polymers, e.g. plastics; Rubber
    • 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
    • F05D2300/00—Materials; Properties thereof
    • F05D2300/40—Organic materials
    • F05D2300/43—Synthetic polymers, e.g. plastics; Rubber
    • F05D2300/432—PTFE [PolyTetraFluorEthylene]
    • 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
    • F05D2300/00—Materials; Properties thereof
    • F05D2300/40—Organic materials
    • F05D2300/44—Resins
    • 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
    • F05D2300/00—Materials; Properties thereof
    • F05D2300/50—Intrinsic material properties or characteristics
    • F05D2300/509—Self lubricating materials; Solid lubricants

Landscapes

  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Chemical & Material Sciences (AREA)
  • Combustion & Propulsion (AREA)
  • Structures Of Non-Positive Displacement Pumps (AREA)

Abstract

A gas turbine fan section or compressor section abradable liner 50 comprises an abradable polymeric matrix 52 having a plurality of dispersed dry lubricant particulates or metallic fibres 54. The dry lubricant particulates or metallic fibres reduce heat caused by friction and which may affect either the material of the blade or the casing. The dry lubricant particles may be selected from the group molybdenum disulphide, polytetrafluoroethylene (PTFE), graphite, boron nitride, talc, calcium fluoride, cerium fluoride and tungsten disulfide. The particulates may be dispersed in the matrix at a volume of between 5% and 15% of the matrix volume. The abradable polymeric matrix may be a low density epoxy resin. The dry lubricant particles may have an average size of between 10 microns and 500 microns. The fibres may be less than 5mm in length. The liner may be provided in a fan casing 32. The liner may be provided on the tip of an aerofoil 12

Description

Adbradable Liner This invention relates to abradable liners for gas turbine fan sections and / or gas turbine compressor sections.
Gas turbine engines comprise, in flow series, a fan, one or more higher pressure compressors, a combustion chamber for burning fuel in the compressed air from the compressor and one or more turbines driven by the exhaust gas from the turbines.
The fan is surrounded by a casing which clears the tips of the fan blades at a minimum acceptable spacing to ensure a high efficienoy by minimising flow leakage over the tips of the blades.
However, during operation of the engine, the rotor blades can thermally expand, extend due to centrifugal force or any vibration or movement of the rotating shaft on which the blades are mounted may cause radial movement which can rub against the inside of the casing and abrade the tip of the rotor.
To mitigate this damage the blade or the casing is provided with an abradable layer which sacrificially wears when rubbed.
It is an object of the present invention to seek to provide an improved abradable liner for gas turbine fan sections and / or gas turbine compressor sections.
According to a first aspect of the invention there is provided a gas turbine fan or compressor abradable liner comprising an abradable polymeric matrix having a plurality of dispersed dry lubricant particulates.
Preferably the dry lubricant particulates are selected from the group consisting of: molybdenum disulfide, polytetrafluoroethylene, graphite, boron nitride, talc, calcium fluoride, cerium fluoride and tungsten disulfide.
Preferably the particulates are dispersed in the matrix at a volume of between 5 and 25% of the matrix volume.
The abradable polymeric matrix may be a low density epoxy resin.
Preferably the dry lubricant particulates have an average size of between 10 microns and 500 microns.
The abradable liner may further comprise a plurality of dispersed metallic fibres. Preferably the fibres are less than 5mm in length. The diameter of the metallic fibres may be between 10 microns and 500 microns.
The liner may be provided on a fan casing. The casing may be composite or metallic. The liner may be provided on the tip of an aerofoil. The aerofoil may be composite or metallic. The casing may be composite and the aerofoil metallic.
The liner may be a gas turbine fan abradable liner. The liner may be a gas turbine compressor abradable liner.
Embodiments of the invention will now be described by way of example only, with reference to the accompanying drawings, in which: Fig. 1 is a cross-sectional view of half of a gas turbine engine; Fig. 2 is a diagrammatic view of a fan casing assembly according to a first embodiment of the present invention.
Referring to Fig. 1, a gas turbine engine generally indicated at 10 comprises, in axiai flow series, an air intake 11, fan section 9 which has a propulsive fan 12 and a fan casing 32, a compressor section which includes, in this embodiment, an intermediate pressure compressor 13, a high pressure compressor 14, a combustor 15, a turbine arrangement comprising a high pressure turbine 16, an intermediate pressure turbine 17 and a low pressure turbine 18, and an exhaust nozzle 19.
The gas turbine engine 10 operates in a conventional manner so that air entering the intake 11 is accelerated by the fan 12 which produces two air flows, a first air flow into the intermediate pressure compressor 13 and a second air flow which provides propulsive thrust. The intermediate pressure compressor compresses the air flow directed into it before delivering that air to the high pressure compressor 14 where further compression take place.
The compressed air exhausted from the high pressure compressor 14 is directed into the combustor 15 where it is mixed with fuel and the mixture combusted. The resultant hot combustion products then expand through and thereby drive the high, intermediate and low pressure turbines 16, 17 and 18, before being exhausted through the nozzle 19 to provide additional propulsive thrust. The high, intermediate and low pressure turbines 16,17 and 18 respectively drive the high and intermediate pressure compressor 14 and 13 and the fan 12 by suitable interconnecting shafts.
The intermediate pressure compressor 13 is connected to the intermediate pressure turbine 17 via an interconnecting shaft 26. Similarly the high pressure compressor 14 is connected to the high pressure turbine 16 by an interconnecting shaft 28 and the fan 12 is connected to the low pressure turbine 18 via an interconnecting shaft 30.
These drive shafts 26,28 and 30 are concentrically mounted around a central axis X-X this also being the central axis of the generally annular gas turbine engine.
The high and intermediate pressure compressors 14 and 13 comprise alternate axially spaced arrays of rotor blades 20 and stator vanes 22 (the respective arrays being referred to as stages of the compressor) . The compressors 13 and 14 (together with the combusto: 15 and the turbines 16, 17 and 18) are encased within an annular casing structure 24.
The fan blades 12 of the fan section 9 are arranged in a circumferential array with each blade extending radially outwardly in relation to the axis X-X. The tips of the fan blades 12 are positioned close to an inner surface of the casing 32, such that there is a minimum clearance therebetween.
Referring to Fig. 2, in one embodiment of the invention, a casing structure 32 surrounds the fan 12. The casing structure may be formed of a plurality of layers and has, on its radially inner surface, an abracIable layer 50 whiob is formed of a resin or polymeric matrix 52 within which is dispersed a plurality of lubricant particulates 54.
The resin or polymeric matrix may be located within cells defined by a honeycomb made of aramid paper dipped in phenolic resin (an example of such a honeycomb material is the NOF4EX (a registered trademark of DuPont) range of products produced by the DuPont company. It will be appreciated that the abradable liner may be provided with the honeycomb. It will be further appreciated that the abradable liner may be provided without the honeycomb.
7 preferred material for the polymeric matrix is an organic adhesive of closed-cell expanding syntactic epoxy with a Shore D hardness of around 70 or less. In alternative embodiments the polymeric matrix may be a polyamide. In further alternative embodiments the polymeric matrix may be provided by a combination of appropriate materials. The polymeric matrix used where the abradable liner is in the fan section may be different to the polymeric matrix used where the abradable liner is in the compressor section. For the Intermediate Pressure compressor a polyester polymer may be used. Cf course, where conditions permit and which may be determined by appropriate experimentation the same or other polymers may be used in both the fan and / or compressor sections.
The rotating fan blade 12 can cut a fan track 56 which is shown in exaggerated form in Figure 2. It will be appreciated that any rub between blade and casing which generates wear will also generate heat in the blade and / or casing through friction. Excess heat generation within the casing or blade can cause mechanical or chemical changes. For example, where the blade or casing is a composite component formed by multiple layers or infused uni or multi directional plies of material the temperature should be kept below the temperature at which the infused adhesive becomes disassociated and which may result in delamination of the layers or a crack initiation point which may allow a crack or other failure to transfer to the rest of the aerofoil or casing. Where the blade or casing is metallic the temperature must not chemically affect the material which could result in softening or embrittlement.
The friction caused by cutting or rubbing is reduced by the dispersed lubricant particles and there is a ccncomitant reduction in heat generation. ecause the particles are dispersed within the matrix any removal of material from the abradable liner inherently exposes new lubricant material which will reduce heating during subseguent rubs that may occur during steady engine running.
Preferred particulates for the lubricant are selected from the group comprising molybdenum disulfide (t4o52), polytetrafluoroethylene, graphite, boron nitrde, talc, calcium fluoride, cerium fluoride, or tugsten disulfide.
The choice of particles may be made such that there is a single particle type dispersed within the matrix.
7&lternatively there may be multiple particle types dispersed within the matrix.
The selected particles are dispersed in the matrix in a ratio between 5% and 15% of the total material making up the abradable layer 50. Particle sizes of between 10 and 500 microns are preferred but the size of the particle and the percentage fill within the matrix should be selected such that the strength of the matrix does not fall below a minimum acceptable value.
In the embodiment of Figure 2 the matrix contains a plurality of metallic fibres 58 which serve to conduct heat away from the vicinity of any rubbing or cutting by the blade on the abradable liner.
The metallic fibres may be chopped fibres or otherwise formed and are preferably less than 5mm in length with a diameter of between 10 and 500 microns. Preferred materials are copper and aluminium due to availability and price but it will be understood that other metals may be used.
In further embodiments the lubricant particles or metallic fibres may not be provided together in the matrix but may instead be used independently of each other. i.e. the matrix comprises lubricant particles or the matrix comprises the metallic fibers.
In the embodiments described above the abradable liner is provided on the radially inner surface of the casing.
Although not shown the casing may include multiple layers which together provide the reguired noise, strength and containment abilities. In some circumstances it may be desirable to locate the abradable liner on the aerofoil tip that is presented to the casing. rhen the abradable liner is provided on the aerofoil tip a further abradable or abrading liner may be provided on the casing, or this further liner may be omitted.
The abradable liner is manufactured by combining the lubricant particles and / or the metallic fibres with the matrix material in an uncured form and mixing till the particles and / or the metallic fibres are dispersed through the matrix. The combination is then formed into the reguired shape and the matrix cured. Curing may be achieved by heating, applying UV or other radiation, or by chemical reaction or any other appropriate process.
Whilst endeavouring in the foregoing specification to draw attention to those features of the invention believed to be of particular importance it should be understood that the Applicant claims protection in respect of any patentable feature or combination of features hereinbefore referred to and/or shown in the drawings whether or not particular emphasis has been placed thereon.

Claims (1)

  1. <claim-text>CLAIMS1. A gas turbine fan section or compressor section abradable liner comprising an abradable polymeric matrix having a plurality of dispersed dry lubricant particulates and / or metallic fibres.</claim-text> <claim-text>2. An abradable liner according to claim 1, wherein the dry lubricant particulates are selected from the group comprising: molybdenum disulfide, polytetrafluoroethylene, graphite, boron nitride, talc, calcium fluoride, cerium fluoride and tungsten disulfide.</claim-text> <claim-text>3. An abradable liner according to claim 1 or claim 2, wherein the particulates are dispersed in the matrix at a volume of between 5 and 15% of the matrix volume.</claim-text> <claim-text>4. An abradable liner according to any preceding claim, wherein the abradable polymeric matrix is a low density epoxy resin.</claim-text> <claim-text>5. A gas turbine fan liner according to any preceding claim, wherein the dry lubricant oarticulates have an average size of between 10 microns and 500 microns.</claim-text> <claim-text>6. An abradable liner according to any preceding claim, wherein the fibres are less than 5mm in length.</claim-text> <claim-text>7. An abradable liner according to any preceding claims, wherein the diameter of the metallic fibres are between 10 microns and 500 microns.</claim-text> <claim-text>8. An abradable liner according to any preceding claim, wherein the liner is provided on a fan casing.</claim-text> <claim-text>9. An abradable liner according tc any one of claims 1 to 8, wherein the liner is provided on the tip of an aerofoil.</claim-text> <claim-text>10. An abradable liner according to any preceding claim, wherein the liner a gas turbine fan abradable liner.</claim-text> <claim-text>11. An abradabie liner according to any of claims 1 to 10, wherein the liner a gas turbine compressor abradable liner.13. An abradable lining arrangement substantially as described herein with reference to any of Figs. 2 to 7 of the accompany drawings.</claim-text>
GB1120338.7A 2011-11-25 2011-11-25 Gas turbine engine abradable liner Withdrawn GB2496887A (en)

Priority Applications (2)

Application Number Priority Date Filing Date Title
GB1120338.7A GB2496887A (en) 2011-11-25 2011-11-25 Gas turbine engine abradable liner
PCT/EP2012/072228 WO2013075954A1 (en) 2011-11-25 2012-11-09 Abradable liner for a gas turbine

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
GB1120338.7A GB2496887A (en) 2011-11-25 2011-11-25 Gas turbine engine abradable liner

Publications (2)

Publication Number Publication Date
GB201120338D0 GB201120338D0 (en) 2012-01-04
GB2496887A true GB2496887A (en) 2013-05-29

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WO (1) WO2013075954A1 (en)

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EP2873811A1 (en) * 2013-11-19 2015-05-20 MTU Aero Engines GmbH Abradable material based on metal fibres
EP2927432A1 (en) * 2014-03-31 2015-10-07 Rolls-Royce plc Gas turbine engine, method of manufacturing a gas turbine engines and fan casing
EP3118113A1 (en) * 2015-07-14 2017-01-18 Northrop Grumman Systems Corporation Bleed air driven lift fan
US9957819B2 (en) 2014-03-28 2018-05-01 United Technologies Corporation Abrasive tip blade manufacture methods
US10480330B2 (en) * 2013-01-29 2019-11-19 United Technologies Corporation Blade rub material

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GB2525410B (en) 2014-04-24 2018-01-17 Rolls Royce Plc A boroscope and a method of processing a component within an assembled apparatus using a boroscope
JP2016102493A (en) * 2014-11-17 2016-06-02 一夫 有▲吉▼ Anti-bird strike net
US10655492B2 (en) 2016-04-29 2020-05-19 United Technologies Corporation Abrasive blade tips with additive resistant to clogging by organic matrix abradable
US10422242B2 (en) 2016-04-29 2019-09-24 United Technologies Corporation Abrasive blade tips with additive resistant to clogging by organic matrix abradable
US10670045B2 (en) 2016-04-29 2020-06-02 Raytheon Technologies Corporation Abrasive blade tips with additive layer resistant to clogging
US10233938B2 (en) * 2016-04-29 2019-03-19 United Technologies Corporation Organic matrix abradable coating resistant to clogging of abrasive blade tips
US11939871B1 (en) * 2022-10-28 2024-03-26 Rtx Corporation Abradable material and design for jet engine applications

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US4336276A (en) * 1980-03-30 1982-06-22 The United States Of America As Represented By The Administrator Of The National Aeronautics And Space Administration Fully plasma-sprayed compliant backed ceramic turbine seal
EP0487273A1 (en) * 1990-11-19 1992-05-27 Sulzer Plasma Technik, Inc. Thermal spray powder
US5364543A (en) * 1991-07-22 1994-11-15 Bosna Alexander A Abradable non-metallic seal for rotating turbine engine
US5472315A (en) * 1993-11-09 1995-12-05 Sundstrand Corporation Abradable coating in a gas turbine engine
US5704759A (en) * 1996-10-21 1998-01-06 Alliedsignal Inc. Abrasive tip/abradable shroud system and method for gas turbine compressor clearance control
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