WO2013075954A1 - Abradable liner for a gas turbine - Google Patents

Abradable liner for a gas turbine Download PDF

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Publication number
WO2013075954A1
WO2013075954A1 PCT/EP2012/072228 EP2012072228W WO2013075954A1 WO 2013075954 A1 WO2013075954 A1 WO 2013075954A1 EP 2012072228 W EP2012072228 W EP 2012072228W WO 2013075954 A1 WO2013075954 A1 WO 2013075954A1
Authority
WO
WIPO (PCT)
Prior art keywords
abradable
liner
gas turbine
abradable liner
liner according
Prior art date
Application number
PCT/EP2012/072228
Other languages
French (fr)
Inventor
Nicholas MERRIMAN
Allan COLLARD
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
Publication of WO2013075954A1 publication Critical patent/WO2013075954A1/en

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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

Definitions

  • 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 efficiency by minimising flow leakage over the tips of the blades.
  • 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.
  • the blade or the casing is provided with an abradable layer which sacrificially wears when rubbed.
  • a gas turbine fan or compressor abradable liner comprising an abradable polymeric matrix having a plurality of dispersed dry lubricant particulates and metallic fibres.
  • 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 .
  • 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.
  • the dry lubricant particulates have an average size of between 10 microns and 500 microns.
  • the dispersed metallic fibres are preferably less than 5mm in length.
  • the diameter of the metallic fibres may be between 10 microns and 500 microns.
  • Advantageously the fibres permit conduction of thermal energy within the polymer matrix.
  • the elongate form permits a more uniform spread of temperature as well as helping to reinforce the polymeric matrix.
  • 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 ai the aerofoil metallic.
  • the liner may be a gas turbine fan abradable liner.
  • the liner may be a gas turbine compressor abradable liner.
  • 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.
  • a gas turbine engine generally
  • indicated at 10 comprises, in axial 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 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, 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 intermediate pressure compressor 13 is connected to the intermediate pressure turbine 17 via an interconnecting shaft 26.
  • 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 combustor 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
  • 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 abradable layer 50 which 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 NOMEX (a registered trademark of DuPont) range of products produced by the DuPont company. It will be described in detail below.
  • the abradable liner may be provided with the honeycomb. It will be further appreciated that the abradable liner may be provided without the honeycomb.
  • a 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.
  • the polymeric matrix may be a polyamide.
  • 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.
  • a polyester polymer may be used for the Intermediate Pressure compressor.
  • 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
  • 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.
  • 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.
  • the blade or casing is metallic the temperature must not chemically affect the material which could result in softening or embrittlement .
  • Preferred particulates for the lubricant are selected from the group comprising molybdenum disulfide (MoS2),
  • 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.
  • 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.
  • the matrix also 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.
  • the fibres permit conduction of thermal energy within the polymer matrix and therefore produce a more uniform spread of temperature.
  • Advantageously the fibres help to reinforce the polymeric matrix which is desirable in the case of a rub between the abradable liner and the blade .
  • the abradable liner is provided on the radially inner surface of the casing.
  • the casing may include multiple layers which together provide the required noise, strength and containment abilities.
  • 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 the metallic fibres with the matrix material in an uncured form and mixing till the particles and metallic fibres are dispersed through the matrix. The combination is then formed into the required 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.

Abstract

A gas turbine fan or compressor abradable liner comprises an abradable polymeric matrix having a plurality of dispersed dry lubricant particulates. The dry lubricant particulates reduce heat caused by friction and which may affect either the material of the blade or the casing.

Description

Abradable Liner for a Gas Turbine
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 efficiency 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 an 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 and metallic fibres.
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 dispersed metallic fibres are preferably less than 5mm in length. The diameter of the metallic fibres may be between 10 microns and 500 microns. Advantageously the fibres permit conduction of thermal energy within the polymer matrix. The elongate form permits a more uniform spread of temperature as well as helping to reinforce the polymeric matrix.
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 ai 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 axial 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 combustor 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 abradable layer 50 which 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 NOMEX (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.
A 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. Of 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 concomitant reduction in heat generation. Because 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 subsequent rubs that may occur during steady engine running.
Preferred particulates for the lubricant are selected from the group comprising molybdenum disulfide (MoS2),
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.
Alternatively 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.
The matrix also 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. The fibres permit conduction of thermal energy within the polymer matrix and therefore produce a more uniform spread of temperature. Advantageously the fibres help to reinforce the polymeric matrix which is desirable in the case of a rub between the abradable liner and the blade . 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 required 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. When 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 the metallic fibres with the matrix material in an uncured form and mixing till the particles and metallic fibres are dispersed through the matrix. The combination is then formed into the required 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. A gas turbine fan section or compressor section abradable liner comprising an abradable polymeric matrix having a plurality of dispersed dry lubricant particulates and metallic fibres.
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 .
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 .
4. An abradable liner according to any preceding claim, wherein the abradable polymeric matrix is a low density epoxy resin.
5. A gas turbine fan liner according to any preceding claim, wherein the dry lubricant particulates have an average size of between 10 microns and 500 microns .
6. An abradable liner according to any preceding claim, wherein the fibres are less than 5mm in length.
7. An abradable liner according to any preceding claims, wherein the diameter of the metallic fibres are between 10 microns and 500 microns.
8. An abradable liner according to any preceding claim, wherein the liner is provided on a fan casing.
9. An abradable liner according to any one of claims 1 to 8, wherein the liner is provided on the tip of an aerofoil .
10. An abradable liner according to any preceding claim, wherein the liner is a gas turbine fan abradable liner.
11. An abradable liner according to any of claims 1 to 10, wherein the liner is a gas turbine compressor abradable liner.
PCT/EP2012/072228 2011-11-25 2012-11-09 Abradable liner for a gas turbine WO2013075954A1 (en)

Applications Claiming Priority (2)

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

Publications (1)

Publication Number Publication Date
WO2013075954A1 true WO2013075954A1 (en) 2013-05-30

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Cited By (8)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP2946870A1 (en) * 2014-04-24 2015-11-25 Rolls-Royce plc A boroscope and a method of processing a component within an assembled apparatus using a boroscope
JP2016102499A (en) * 2014-11-17 2016-06-02 一夫 有▲吉▼ Jet engine easily rejecting sucked foreign items
US20170015413A1 (en) * 2015-07-14 2017-01-19 Northrop Grumman Systems Corporation Magrail, bleed air driven lift fan
EP3239466A1 (en) * 2016-04-29 2017-11-01 United Technologies Corporation Organic matrix abradable coating
US10422242B2 (en) 2016-04-29 2019-09-24 United Technologies Corporation Abrasive blade tips with additive resistant to clogging by organic matrix abradable
US10655492B2 (en) 2016-04-29 2020-05-19 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
EP4361406A1 (en) * 2022-10-28 2024-05-01 RTX Corporation Abradable material and design for jet engine applications

Families Citing this family (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
WO2014120117A1 (en) * 2013-01-29 2014-08-07 United Technologies Corporation Blade rub material
DE102013223585A1 (en) * 2013-11-19 2015-06-03 MTU Aero Engines AG Inlet lining based on metal fibers
US9957819B2 (en) 2014-03-28 2018-05-01 United Technologies Corporation Abrasive tip blade manufacture methods
GB201405704D0 (en) * 2014-03-31 2014-05-14 Rolls Royce Plc Gas turbine engine

Citations (7)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
GB791568A (en) * 1956-03-26 1958-03-05 Gen Motors Corp Improvements in axial flow compressors
US4139376A (en) * 1974-02-28 1979-02-13 Brunswick Corporation Abradable seal material and composition thereof
US5304032A (en) * 1991-07-22 1994-04-19 Bosna Alexander A Abradable non-metallic seal for rotating turbine engines
US20030228483A1 (en) * 2002-06-07 2003-12-11 Petr Fiala Thermal spray compositions for abradable seals
US20040126225A1 (en) * 2002-12-31 2004-07-01 General Electric Grc Rotary machine sealing assembly
US20050220612A1 (en) * 2003-11-17 2005-10-06 Ingo Jahns Inner shroud for the stator blades of the compressor of a gas turbine
US20070134411A1 (en) * 2005-12-14 2007-06-14 General Electric Company Method for making compositions containing microcapsules and compositions made thereof

Family Cites Families (8)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
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
US5196471A (en) * 1990-11-19 1993-03-23 Sulzer Plasma Technik, Inc. Thermal spray powders for abradable coatings, abradable coatings containing solid lubricants and methods of fabricating abradable coatings
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
SG88799A1 (en) * 1999-12-17 2002-05-21 United Technologies Corp Abradable seal having improved properties
US20010055652A1 (en) * 1999-12-17 2001-12-27 William John Dalzell Method of making abradable seal having improved properties
AU2001288459A1 (en) * 2000-08-29 2002-03-13 Andrew W. Suman Abradable dry powder coatings, methods for making and coating, and coated articles therefrom
US6660405B2 (en) * 2001-05-24 2003-12-09 General Electric Co. High temperature abradable coating for turbine shrouds without bucket tipping

Patent Citations (7)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
GB791568A (en) * 1956-03-26 1958-03-05 Gen Motors Corp Improvements in axial flow compressors
US4139376A (en) * 1974-02-28 1979-02-13 Brunswick Corporation Abradable seal material and composition thereof
US5304032A (en) * 1991-07-22 1994-04-19 Bosna Alexander A Abradable non-metallic seal for rotating turbine engines
US20030228483A1 (en) * 2002-06-07 2003-12-11 Petr Fiala Thermal spray compositions for abradable seals
US20040126225A1 (en) * 2002-12-31 2004-07-01 General Electric Grc Rotary machine sealing assembly
US20050220612A1 (en) * 2003-11-17 2005-10-06 Ingo Jahns Inner shroud for the stator blades of the compressor of a gas turbine
US20070134411A1 (en) * 2005-12-14 2007-06-14 General Electric Company Method for making compositions containing microcapsules and compositions made thereof

Cited By (11)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP2946870A1 (en) * 2014-04-24 2015-11-25 Rolls-Royce plc A boroscope and a method of processing a component within an assembled apparatus using a boroscope
US9703090B2 (en) 2014-04-24 2017-07-11 Rolls-Royce Plc Boroscope and a method of processing a component within an assembled apparatus using a boroscope
JP2016102499A (en) * 2014-11-17 2016-06-02 一夫 有▲吉▼ Jet engine easily rejecting sucked foreign items
US20170015413A1 (en) * 2015-07-14 2017-01-19 Northrop Grumman Systems Corporation Magrail, bleed air driven lift fan
US9950788B2 (en) * 2015-07-14 2018-04-24 Northrop Grumman Systems Corporation Magrail, bleed air driven lift fan
EP3239466A1 (en) * 2016-04-29 2017-11-01 United Technologies Corporation Organic matrix abradable coating
US10233938B2 (en) 2016-04-29 2019-03-19 United Technologies Corporation Organic matrix abradable coating resistant to clogging of abrasive blade tips
US10422242B2 (en) 2016-04-29 2019-09-24 United Technologies Corporation Abrasive blade tips with additive resistant to clogging by organic matrix abradable
US10655492B2 (en) 2016-04-29 2020-05-19 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
EP4361406A1 (en) * 2022-10-28 2024-05-01 RTX Corporation Abradable material and design for jet engine applications

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