WO2014096840A1 - Structure de profil aérodynamique ayant des bords de coupe de partie pointe - Google Patents

Structure de profil aérodynamique ayant des bords de coupe de partie pointe Download PDF

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Publication number
WO2014096840A1
WO2014096840A1 PCT/GB2013/053371 GB2013053371W WO2014096840A1 WO 2014096840 A1 WO2014096840 A1 WO 2014096840A1 GB 2013053371 W GB2013053371 W GB 2013053371W WO 2014096840 A1 WO2014096840 A1 WO 2014096840A1
Authority
WO
WIPO (PCT)
Prior art keywords
aerofoil structure
aerofoil
cutting edges
tip
cutting
Prior art date
Application number
PCT/GB2013/053371
Other languages
English (en)
Inventor
Matthew E HANCOCK
Nicholas Michael Merriman
Andrew Hewitt
Glen Pattinson
Original Assignee
Composite Technology And Applications Limited
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 Composite Technology And Applications Limited filed Critical Composite Technology And Applications Limited
Publication of WO2014096840A1 publication Critical patent/WO2014096840A1/fr

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Classifications

    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • 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
    • 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
    • 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
    • F05D2230/00—Manufacture
    • F05D2230/90—Coating; Surface treatment
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2240/00—Components
    • F05D2240/20—Rotors
    • F05D2240/30—Characteristics of rotor blades, i.e. of any element transforming dynamic fluid energy to or from rotational energy and being attached to a rotor
    • F05D2240/307—Characteristics of rotor blades, i.e. of any element transforming dynamic fluid energy to or from rotational energy and being attached to a rotor related to the tip of a rotor blade
    • 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
    • F05D2300/2284—Nitrides of titanium
    • 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

  • the present disclosure relates to blades for gas turbine engines and the like and relates particularly but not exclusively to improvements relating to the tips of such blades and to modifications thereto for enabling said blades to abrade the liners of casings associated therewith.
  • the blades of gas turbine engines are arranged with a minimum clearance between the tips of the blades and the casings associated therewith, as any gap therebetween will contribute to a reduction in efficiency.
  • the tips of such blades are often provided with an abrasive coating and the corresponding portion of the casing is provided with an abradable layer which is removed by the tip if the tip comes into contact with the casing.
  • the abrasive coating is applied to a fiat surface, machined or otherwise formed on the tip and, on occasion, the flat surface rather than an edge thereof can come into contact with the casing abradable layer.
  • the present disclosure seeks to provide a novel blade, which at least reduces the above problem whilst providing an efficient cutting surface to the blade tip.
  • an aerofoil structure for a gas turbine engine, the aerofoil structure comprising an aerofoil portion and a tip portion, wherein the tip portion comprises a tip surface configured to face a corresponding casing structure, the tip portion further comprising a plurality of cutting edges provided on at least a portion of the tip surface, the cutting edges being configured to cut into the casing structure.
  • the plurality of cutting edges may be arranged in a regular fashion on the tip surface. Alternatively, the plurality of cutting edges may be arranged in an irregular fashion on the tip surface.
  • the irregular arrangement of the cutting edges may be random.
  • the irregular arrangement of the cutting edges may be pseudorandom.
  • the irregular arrangement may be repeatabie across a plurality of aerofoil structures.
  • the cutting edges may be formed of a cutting line or a cutting point.
  • the cutting edges may have a length L and the length L may extend at an angle 8 s to a direction of rotation R of the aerofoil structure.
  • the length L may extend substantially perpendicular to the direction of rotation R of the aerofoil structure.
  • the tip portion may comprise one or more first cutting edges having a length L1 angled at a first angle with respect to a direction of rotation R.
  • the tip portion may comprise one or more second cutting edges having a length L2 angled at a second angle with respect to the direction of rotation.
  • the first and second cutting edges may define rhombus shaped cutting tips on the tip surface.
  • the first and second angles may be equal in value.
  • the first and second angles may be opposite in direction.
  • the orientation of the cutting edges may be configured for when the aerofoil structure is an untwisted condition (e.g. in a normal mode of operation) or a twisted condition (e.g. following a bird strike or blade-off event).
  • the cutting edge angles may vary along the chord of the tip surface.
  • the cutting edges may be spaced apart with a regular or irregular distribution across the tip surface.
  • the cutting edges may define tips or peaks and troughs or valleys therebetween.
  • the troughs may be provided between neighbouring cutting edge tips.
  • the cutting edge tips' cross-sectional shape may be substantially square, rectangular, triangular or any other shape.
  • the cutting edge tips may at least partially extend over neighbouring troughs.
  • a height of the cutting tips and/or depth of the troughs may be regular or irregular across the tip surface. The height of the cutting tips and/or depth of the troughs may be measured in a substantially radial direction, e.g. relative to a root of the aerofoil structure.
  • the cutting edges may be elongate.
  • the cutting edges may be formed by a material removal process, such as Electrical Discharge machining (EDM), Electro Chemical Machining (ECM), machining, milling, mechanical blasting, chemical etching and laser etching.
  • EDM Electrical Discharge machining
  • ECM Electro Chemical Machining
  • the tip portion may comprise one or more further cutting edges.
  • the further cutting edges may be provided on a suction and/or pressure surface of the aerofoil structure, e.g. in a region adjacent to the tip surface.
  • the cutting edges may extend beyond the tip surface.
  • the cutting edges may extend from the tip surface down at least a portion of the aerofoil portion.
  • the aerofoil structure may further comprise a coating provided over the tip surface.
  • the coating may extend beyond the tip surface, e.g. over the rest of the tip portion.
  • the coating may comprise a material having a higher hardness than the tip portion.
  • the coating may comprise a material having a hardness in excess of 1000 Vickers.
  • the coating may comprise a thin-film coating having a thickness of between 2 and 10 microns.
  • the coating may comprise a deposited coating.
  • the coating may comprise Titanium Nitride and/or Chromium Nitride.
  • the aerofoil structure may comprise a metallic or carbon composite aerofoil portion.
  • the aerofoil structure may comprise a metallic tip portion.
  • a method of manufacturing an aerofoil structure for a gas turbine engine comprising an aerofoil portion and a tip portion, wherein the tip portion comprises a tip surface configured to face a corresponding casing structure, the method comprising: forming a plurality of cutting edges on at least a portion of the tip surface, the cutting edges being configured to cut into the casing structure.
  • Forming the plurality of cutting edges may comprise forming by one or more of: Electrical Discharge machining (EDM), Electro Chemical Machining (EC ), machining, milling, stamping, mechanical blasting, chemical etching, laser etching or any other forming process.
  • the tip surface may comprise an impact or pressure generated surface.
  • the aerofoil structure may comprise a compressor blade, e.g. a compressor fan blade, of a gas turbine engine.
  • the aerofoil structure may comprise a turbine blade of a gas turbine engine.
  • the present disclosure may also provide a gas turbine engine having an aerofoil structure as described above.
  • Figure 1 is a diagrammatic representation of a gas turbine engine incorporating the present invention
  • Figure 2 is a view of an exemplar blade for a gas turbine engine incorporating the present invention at a tip thereof;
  • Figure 3 is a simplified view of the front elevation of part of a typical compressor fan blade assembly from a gas turbine engine of figure 1 and illustrates the direction of rotation about central axis CL and the radial axis of the blades RA;
  • Figure 4 is an enlarged view of a typical blade tip of the blades shown in figures 2 & 3;
  • Figure 5 is a schematic plan view of the tip shown in figure 4 and illustrates the normal direction of rotation R relative to the centre line CL and the Leading Surface LS and Trailing Surface TS;
  • Figures 6 and 7 are cross-sectional views of a metal (or metallic alloy) blade and a composite blade incorporating the present invention shown in dose proximity to a casing; and Figures 8 to 13 are cross-sectional views of alternative forms of cutting profiles that may be applied to the tip in accordance with the present disclosure.
  • a turbofan gas turbine engine comprises in flow series an inlet 12, a fan section 14, a compressor section 16, a combustion section 18, a turbine section 20 and an exhaust 22.
  • the fan section 14 comprises a fan rotor 24 carrying a plurality of circumferentially spaced radially outwardly extending fan blades 26.
  • the fan blades 26 are arranged in a bypass duct 28 defined by a fan casing 30 having an abradable lining 30a, which surrounds the fan rotor 24 and fan blades 26.
  • the fan casing 30 is secured to a core engine casing 34 by a plurality of circumferentially spaced radially extending fan outlet guide vanes 32.
  • the fan rotor 24 and fan blades 26 are arranged to be driven by a turbine (not shown) in the turbine section 20 via a shaft (not shown).
  • the compressor section 16 comprises one or more compressors (not shown) arranged to be driven by one or more turbines (not shown) in the turbine section 20 via respective shafts (not shown).
  • the engine includes a longitudinally extending centre line CL around which the blades rotate in a direction R.
  • the fan blade 26 comprises a root portion 36 at a radially inner end and an aerofoil portion 38.
  • the root portion 36 is arranged to locate in a slot 40 in the rim of the disc 42 of the fan rotor 24, and for example the root portion 36 may be dovetail shape, fir-tree shape, or other conventional shape, in cross-section and hence the corresponding slot 40 in the rim of the disc 42 of the fan rotor 24 is a similar shape.
  • the aerofoil portion 38 has a leading edge 44, a trailing edge 46 and a tip 48 at a radially outer end remote from the root portion 36 and the fan rotor 24.
  • a concave pressure surface 50 extends from the leading edge 44 to the trailing edge 46 and a convex suction surface 52 also extends from the leading edge 44 to the trailing edge 46.
  • Figure 3 illustrates a plurality of blades 26 assembled in the rotor disc 42 and also illustrates the relationship between the centre line of the engine CL, a Radial Axis RA of the blades 26 and the tip 48. The radial axis RA is shown from the centre line CL in a direction radially outwards towards the casing. Arrow T illustrates the direction of a tangent referred to later herein. In particular, figure 3 illustrates the position of the blade tip 48 of a blade 26 within part of a blade set. As the blade set rotates in direction R the leading edge 44 of the blade is incident with the incoming air flow.
  • a nose cone 43 covers the shaft and provides an air-washed surface from the engine centreline CL into the root of the blade.
  • the blades 26 are not subjected to excessive loads and do not materially deflect.
  • the blade may be caused to deflect or (un)twist into an abnormal position.
  • portions of the tip 48 may come into contact with the casing 30 or any abradable lining 30a provided on the inner surface thereof.
  • the tip may also come into contact with the casing in other situations, such as during engine run in, aircraft take off or other aircraft manoeuvres. Whilst a certain degree of contact can normally be accommodated by abrading the abradable lining 30a, excessive contact may cause the tip of a normal blade or the casing itself to be damaged.
  • the present disclosure relates to an aerofoil structure, e.g. blade 26, comprising an aerofoil portion 38 and a tip portion 48, wherein the tip portion comprises a tip surface 60 configured to face a corresponding casing structure 30, the tip portion further comprising a plurality of cutting edges provided on at least a portion of the tip surface.
  • the cutting edges are configured to cut into the casing structure, for example to avoid damage to the blade.
  • Figure 4 illustrates the tip of the blade 26 in more detail than in figure 2 and from which it will be appreciated that the tip 48 is shown in exaggerated form and comprises a tip surface 60 which may be substantially provided in a plane P.
  • the plane P may be curved, e.g. due to the stagger angle of the blade and/or curvature of the casing.
  • the plane P may extend at an angle ⁇ to a plane containing the radial axis RA and the centreline CL.
  • the specific angle may be altered and may be arranged such that plane P is generally tangential to the radial axis RA of the blade such as to allow a generally tangential surface to be presented to an abradable lining 30a provided on an inner surface 58 of casing 30.
  • angle ⁇ 2 may be more than 90 degrees.
  • the angle ⁇ may be varied to account for the blade twisting either in normal operation or abnormal operation, e.g. following a bird-strike or blade-off event.
  • the angle ⁇ may be varied along the chord of the blade, e.g. to account for twist of the blade about the radial axis RA.
  • the tip surface 60 may not necessarily lie entirely in plane P, for example the tip surface may comprise a ridge (e.g. in the shape of a roof top) extending along the tip surface in a chordwise direction.
  • the plane P indicates the general orientation of the tip surface 60 with respect to the radial axis RA.
  • the tip 48 of the blade 26 is formed so as to create a cutting tip surface 60 directly at the tip 48 at an end of the blade 26 which, in operation, is used to abrade the abradable lining 30a in the event that the tip 48 does come into contact therewith.
  • the specific form of the cutting surface is discussed in more detail below but comprises a foundation surface 62, which may be formed by mechanically modifying the surface of the tip such as to produce a regular or irregular roughened surface suitable for cutting the abradable lining.
  • the foundation surface 62 may be supplemented with a hardened coating 64 or an outer layer 66.
  • a metal tip portion made from, for example, titanium is provided with a base portion 68 for bonding or otherwise joining to the main portion of the blade 26 which, as shown, may comprise an inner portion of carbon composite material 70 and an outer protective layer 72, which may be metallic, and an adhesive 74 to bond the base portion 68 in position.
  • a base portion 68 for bonding or otherwise joining to the main portion of the blade 26 which, as shown, may comprise an inner portion of carbon composite material 70 and an outer protective layer 72, which may be metallic, and an adhesive 74 to bond the base portion 68 in position.
  • Other forms of securing the base portion 68 to the remainder of the blade 26 will present themselves to those skilled in the art.
  • Figure 4 illustrates in diagrammatic form a regular cut tip surface 60 in which the surface may be cut or mechanically modified to provide a particular and regular surface having a regular orientation that is used to define a cutting surface which can be used to abrade the lining 30a of the casing 30.
  • the regular pattern provides the surface with a plurality of individual cutting edges 80 (best seen in figures 8 to 13), which, in operation, will act as a cutting edge.
  • the cutting edges 80 are provided on the tip surface 60 in a region between the tip surface edges at the leading side LS and trailing side TS of the blade.
  • the tip portion comprises one or more first cutting edges 80 having a length L1 angled at a first angle with respect to the direction of rotation R and one or more second cutting edges 80 having a length L2 angled at a second angle with respect to the direction of rotation.
  • the cutting edges 80 are arranged to extend in two directions A and B with direction A being angled at an angle ⁇ 5 relative to the direction of rotation R whilst direction B is angled at an angle ⁇ 6 relative to direction A.
  • Angle ⁇ 5 may be between zero and 89 degrees whilst angle ⁇ 6 may be between zero and 179 degrees.
  • angle ⁇ 5 may be 45 degrees and angle ⁇ 6 may be 90 degrees.
  • the directions A, B may be orientated in equal and opposite directions with respect to the direction of rotation R.
  • Figure 5 illustrates the simpler arrangement in which the cutting edges extend in just one direction A.
  • the tip portion comprises one or more first cutting edges 80 having a length L angled at a first angle with respect to a direction of rotation R.
  • the direction A is angled at angle ⁇ 7 relative to the direction of rotation R.
  • Angle ⁇ 7 may be between zero and 90 degrees.
  • ⁇ 5 may be 90 degrees such that the cutting edges are perpendicular to the direction of rotation.
  • the orientation of directions A and/or B relative to the direction of rotation R referred to above may be either when the blade is an untwisted condition (e.g. in a normal mode of operation) or in a twisted condition (e.g. following a bird strike or b!ade-off event).
  • the cutting edge directions A and/or B may be selected to provide efficient cutting when the blade is in a twisted condition as this may be the condition which results in a stronger interaction between the tip and the casing.
  • the cutting edges 80 will be employed to cut into the liner 30a such as to create a clearance (not shown) and, thus, avoid the blade impacting or rubbing with the lining 30a which may cause the "blueing" or excessive oxidation referred to above.
  • the arrangements of figures 6 and 7 are illustrative of the other form of cutting surface 60 in which the cutting edges on the tip surface are arranged in an irregular fashion.
  • the size, height, shape, distribution, cutting edge direction and/or orientation of cutting tips 90, which form the cutting edges may be arranged in a generally random or pseudo-random manner.
  • the irregular pattern on a particular blade may be, or at least appear to be, random, but the same pattern may be applied to other blades, e.g. by virtue of a repeatable manufacturing process.
  • the tips 90 perform the same function in the event that they come into contact with the liner 30a of the casing.
  • An additional benefit may reside in arrangements in which the height H is not the same for each tip 90 as, a more progressive cutting action may be created as taller tips 90 cut first and the lower tips are not used unless a higher degree of impact with the liner 30a has been created.
  • the irregular arrangement of the cutting tips, and the troughs therebetween, may also assist in clearing debris resulting from the cutting of the liner 30a, thereby preventing the cutting surface from clogging up.
  • the pattern of the cutting tips may be arranged so as to minimise heat flow into the blade tip during cutting. This may be achieved by virtue of having a more efficient cutting action thanks to the cutting edges provided on the tip surface.
  • the cutting edges may be formed by randomly or pseudo-randomly impacting the surface 60 such as to create a cratered or peaked imprint thereon.
  • a surface could be made by grit blasting the surface 60 or may be formed by machining, chemical etching, electro discharge machining or the like.
  • FIG. 8 With reference to figures 8 to 13, three forms of cutting edge are shown at 80a, 80b and 80c in which 80a are each representative of bluff edges, 80b is representative of a peaked edge and the 80c references illustrate cutting edges with an undercut 82 at an angle ⁇ 3 and optionally a rake angle ⁇ 4 applied to a trailing edge 84.
  • the forms shown may be milled, machined, stamped, ground or produced by electron discharge machining.
  • the cutting edges 80 may be cut as multiple single rows each extending in the same direction or as cross-cut rows in which the cutting edges each extend in two or more directions A, B.
  • Each of the above arrangements may have a heat treatment applied to the cutting surfaces such as to make them more resistant to wear.
  • a wear resistant coating may be applied to the cutting surface as shown at 92 in figures 6 and 8 to 13.
  • Such a coating may include titanium nitride and/or chromium nitride which may be applied as a coating.
  • the coating may have a thickness of between 2 and 10 microns.
  • the blade 26 as described above may be a compressor blade, such as a fan blade 26, or may, in certain circumstances, be a turbine blade (not shown).
  • the blade 26 may be part of an engine 10 as discussed with reference to figure 1 above and the present disclosure extends to such an engine with such a blade 26.
  • the blade 26 as described above may also be applied to ducted fans, e.g., future aircraft engine architectures, hovercraft propelling fans, ducted helicopter rear rotors, air-conditioning fans, wind tunnel propulsors, marine propulsors and marine power generators or any other ducted fan.
  • the blade 26 of the present disclosure may also be applied to disc or drum seals and to labyrinth seal fins.

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  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Structures Of Non-Positive Displacement Pumps (AREA)

Abstract

L'invention concerne une structure de profil aérodynamique (26) pour un moteur de turbine à gaz (10), la structure de profil aérodynamique comprenant une partie profil aérodynamique (38) et une partie pointe (48), la partie pointe comprenant une surface de pointe (60) configurée pour faire face à une structure d'enveloppe correspondante (30), la partie pointe comprenant en outre une pluralité de bords de coupe (80) disposés sur au moins une partie de la surface de pointe, les bords de coupe étant configurés pour être coupés dans la structure d'enveloppe. L'invention concerne un procédé correspondant de fabrication d'un profil aérodynamique d'un moteur de turbine à gaz.
PCT/GB2013/053371 2012-12-19 2013-12-19 Structure de profil aérodynamique ayant des bords de coupe de partie pointe WO2014096840A1 (fr)

Applications Claiming Priority (2)

Application Number Priority Date Filing Date Title
GB1222975.3 2012-12-19
GBGB1222975.3A GB201222975D0 (en) 2012-12-19 2012-12-19 An aerofoil structure

Publications (1)

Publication Number Publication Date
WO2014096840A1 true WO2014096840A1 (fr) 2014-06-26

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PCT/GB2013/053371 WO2014096840A1 (fr) 2012-12-19 2013-12-19 Structure de profil aérodynamique ayant des bords de coupe de partie pointe

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GB (1) GB201222975D0 (fr)
WO (1) WO2014096840A1 (fr)

Cited By (3)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
BE1022882B1 (fr) * 2015-03-30 2016-10-05 Safran Aero Boosters S.A. Tournage a choc d'extremites d'aubes de blum de compresseur de turbomachine axiale
ITUB20155442A1 (it) * 2015-11-11 2017-05-11 Ge Avio Srl Stadio di un motore a turbina a gas provvisto di una tenuta a labirinto
US10724535B2 (en) 2017-11-14 2020-07-28 Raytheon Technologies Corporation Fan assembly of a gas turbine engine with a tip shroud

Citations (7)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
GB2052644A (en) * 1979-06-18 1981-01-28 Gen Electric Staircase blade tip
US4738586A (en) * 1985-03-11 1988-04-19 United Technologies Corporation Compressor blade tip seal
JPH0913904A (ja) * 1995-06-27 1997-01-14 Ishikawajima Harima Heavy Ind Co Ltd セラミック製タービン動翼
US6171351B1 (en) * 1994-09-16 2001-01-09 MTU Motoren-und Turbinen Union M{umlaut over (u)}nchen GmbH Strip coatings for metal components of drive units and their process of manufacture
EP2196631A2 (fr) * 2008-12-15 2010-06-16 Rolls-Royce plc Composant doté d'une couche abrasive et procédé d'application d'une couche abrasive sur un composant
EP2253803A2 (fr) * 2009-05-21 2010-11-24 Rolls-Royce plc Aube composite avec sommet résistant à l'usure
FR2961846A1 (fr) * 2010-06-28 2011-12-30 Snecma Propulsion Solide Aube de turbomachine a geometrie asymetrique complementaire

Patent Citations (7)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
GB2052644A (en) * 1979-06-18 1981-01-28 Gen Electric Staircase blade tip
US4738586A (en) * 1985-03-11 1988-04-19 United Technologies Corporation Compressor blade tip seal
US6171351B1 (en) * 1994-09-16 2001-01-09 MTU Motoren-und Turbinen Union M{umlaut over (u)}nchen GmbH Strip coatings for metal components of drive units and their process of manufacture
JPH0913904A (ja) * 1995-06-27 1997-01-14 Ishikawajima Harima Heavy Ind Co Ltd セラミック製タービン動翼
EP2196631A2 (fr) * 2008-12-15 2010-06-16 Rolls-Royce plc Composant doté d'une couche abrasive et procédé d'application d'une couche abrasive sur un composant
EP2253803A2 (fr) * 2009-05-21 2010-11-24 Rolls-Royce plc Aube composite avec sommet résistant à l'usure
FR2961846A1 (fr) * 2010-06-28 2011-12-30 Snecma Propulsion Solide Aube de turbomachine a geometrie asymetrique complementaire

Cited By (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
BE1022882B1 (fr) * 2015-03-30 2016-10-05 Safran Aero Boosters S.A. Tournage a choc d'extremites d'aubes de blum de compresseur de turbomachine axiale
ITUB20155442A1 (it) * 2015-11-11 2017-05-11 Ge Avio Srl Stadio di un motore a turbina a gas provvisto di una tenuta a labirinto
EP3168427A1 (fr) * 2015-11-11 2017-05-17 Ge Avio S.r.l. Étage de moteur de turbine à gaz muni d'un joint à labyrinthe
US10724535B2 (en) 2017-11-14 2020-07-28 Raytheon Technologies Corporation Fan assembly of a gas turbine engine with a tip shroud

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