WO2022258257A1 - Ensemble rotor pour un moteur à turbine à gaz, procédé d'assemblage d'un ensemble rotor et procédé de fabrication d'un manchon - Google Patents

Ensemble rotor pour un moteur à turbine à gaz, procédé d'assemblage d'un ensemble rotor et procédé de fabrication d'un manchon Download PDF

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Publication number
WO2022258257A1
WO2022258257A1 PCT/EP2022/061515 EP2022061515W WO2022258257A1 WO 2022258257 A1 WO2022258257 A1 WO 2022258257A1 EP 2022061515 W EP2022061515 W EP 2022061515W WO 2022258257 A1 WO2022258257 A1 WO 2022258257A1
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WO
WIPO (PCT)
Prior art keywords
sleeve
rotor assembly
slot
blade
root
Prior art date
Application number
PCT/EP2022/061515
Other languages
English (en)
Inventor
Paul Mathew Walker
Original Assignee
Siemens Energy Global GmbH & Co. KG
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 Siemens Energy Global GmbH & Co. KG filed Critical Siemens Energy Global GmbH & Co. KG
Publication of WO2022258257A1 publication Critical patent/WO2022258257A1/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
    • F01D5/30—Fixing blades to rotors; Blade roots ; Blade spacers
    • F01D5/3084—Fixing blades to rotors; Blade roots ; Blade spacers the blades being made of ceramics
    • 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
    • F01D5/30—Fixing blades to rotors; Blade roots ; Blade spacers
    • F01D5/3092—Protective layers between blade root and rotor disc surfaces, e.g. anti-friction layers
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B33—ADDITIVE MANUFACTURING TECHNOLOGY
    • B33Y—ADDITIVE MANUFACTURING, i.e. MANUFACTURING OF THREE-DIMENSIONAL [3-D] OBJECTS BY ADDITIVE DEPOSITION, ADDITIVE AGGLOMERATION OR ADDITIVE LAYERING, e.g. BY 3-D PRINTING, STEREOLITHOGRAPHY OR SELECTIVE LASER SINTERING
    • B33Y80/00—Products made by additive manufacturing
    • 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
    • F01D25/00—Component parts, details, or accessories, not provided for in, or of interest apart from, other groups
    • F01D25/005—Selecting particular materials
    • 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
    • F01D25/00—Component parts, details, or accessories, not provided for in, or of interest apart from, other groups
    • F01D25/007—Preventing corrosion
    • 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
    • F01D5/02—Blade-carrying members, e.g. rotors
    • F01D5/08—Heating, heat-insulating or cooling means
    • F01D5/081—Cooling fluid being directed on the side of the rotor disc or at the roots of the blades
    • 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
    • F01D5/28—Selecting particular materials; Particular measures relating thereto; Measures against erosion or corrosion
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B22—CASTING; POWDER METALLURGY
    • B22F—WORKING METALLIC POWDER; MANUFACTURE OF ARTICLES FROM METALLIC POWDER; MAKING METALLIC POWDER; APPARATUS OR DEVICES SPECIALLY ADAPTED FOR METALLIC POWDER
    • B22F5/00—Manufacture of workpieces or articles from metallic powder characterised by the special shape of the product
    • B22F5/009—Manufacture of workpieces or articles from metallic powder characterised by the special shape of the product of turbine components other than turbine blades
    • C—CHEMISTRY; METALLURGY
    • C22—METALLURGY; FERROUS OR NON-FERROUS ALLOYS; TREATMENT OF ALLOYS OR NON-FERROUS METALS
    • C22C—ALLOYS
    • C22C19/00—Alloys based on nickel or cobalt
    • C22C19/03—Alloys based on nickel or cobalt based on nickel
    • C22C19/05—Alloys based on nickel or cobalt based on nickel with chromium
    • 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
    • F01D5/28—Selecting particular materials; Particular measures relating thereto; Measures against erosion or corrosion
    • F01D5/282—Selecting composite materials, e.g. blades with reinforcing filaments
    • 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
    • F01D5/28—Selecting particular materials; Particular measures relating thereto; Measures against erosion or corrosion
    • F01D5/284—Selection of ceramic materials
    • 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
    • F05D2230/20—Manufacture essentially without removing material
    • F05D2230/22—Manufacture essentially without removing material by sintering
    • 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
    • F05D2230/20—Manufacture essentially without removing material
    • F05D2230/23—Manufacture essentially without removing material by permanently joining parts together
    • F05D2230/232—Manufacture essentially without removing material by permanently joining parts together by welding
    • F05D2230/234—Laser welding
    • 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
    • F05D2230/30—Manufacture with deposition of material
    • F05D2230/31—Layer deposition
    • 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
    • F05D2230/50—Building or constructing in particular ways
    • F05D2230/54—Building or constructing in particular ways by sheet metal manufacturing
    • 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
    • F05D2300/10—Metals, alloys or intermetallic compounds
    • F05D2300/13—Refractory metals, i.e. Ti, V, Cr, Zr, Nb, Mo, Hf, Ta, W
    • F05D2300/132—Chromium
    • 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
    • F05D2300/10—Metals, alloys or intermetallic compounds
    • F05D2300/17—Alloys
    • F05D2300/175—Superalloys

Definitions

  • ROTOR ASSEMBLY FOR A GAS TURBINE ENGINE METHOD OF ASSEMBLING A ROTOR ASSEMBLY AND METHOD OF MANUFACTURING A SLEEVE
  • the present invention relates to a blade assembly for a rotor assembly for a gas turbine engine.
  • the turbine In a gas turbine engine, the turbine is subject to very high temperatures of working gas from the combustor.
  • the turbine, particularly the first two rotor blade stages are subject to environments that can degrade the components by means of corrosion attack.
  • the corrosion attack preferentially attacks the gamma prime, the primary strengthening phase in nickel-based superalloys, in the substrate resulting in a reduced service-life because of cracking and eventually resulting in a blade failure in the worst cases.
  • the corrosion is driven by gas-born salt deposits and Sulphur, which are well known to accelerate corrosion attack of nickel-based alloys such as those used in turbine blades and vanes.
  • WO9641068 discloses a cold-worked low-friction metallic shim or a low- friction non-metallic composite shim that is bonded onto the surface of a component with a layer of adhesive to form an anti-fretting barrier.
  • This barrier is incorporated between the tightly fit components which are subject to high cyclic stresses, e.g., a blade and a disc slot in a gas turbine engine.
  • the barrier is bonded to the surface of only one component. Bonding the shim significantly delays fretting damage on the surface of the component.
  • US 4207029 A discloses a turbine rotor assembly wherein ceramic blades are secured to slots provided in the periphery of a metallic rotor disc, a metallic jacket is interposed between each blade root and disc slot, and ceramic potting material is, in turn, interposed between the metallic jacket and the ceramic blade root.
  • the metallic jacket also includes locking means to prevent upstream or downstream axial movement of the blade in the disc slot, with the resulting assembly effectively isolating the ceramic root from high localized loads and permitting a more uniform stress distribution on the critical root load bearing surfaces.
  • FR 3099520 A1 discloses a sheet comprising a first part and a second part.
  • the first part is placed between a root of the blade and a bottom of the disc pocket in which said blade root is engaged.
  • the second part is arranged radially opposite the gap between the platforms of two adjacent vanes.
  • the first part of the sheet provides an anti-wear function, a function generally provided by a foil, to limit the wear of the blade roots when the blade wheel is rotated.
  • the second part of the sheet ensures a sealing function in combination with the sealing devices. During operation of the impeller, the centrifugal force forces the sealing device to press against the radially inner surface of this second part. Sealing devices can also dampen the vibrations to which the blades are subjected
  • US 2002/044870 A1 discloses compliant shim for use between the root of a gas turbine fan blade and a dovetail groove in a gas turbine rotor disk to reduce fretting therebetween.
  • the compliant shim has first and second slots for engaging tabs extending from the fan blade root. The slots and tabs cooperate to hold the shim during engine operation.
  • An oxidation layer covers the compliant shim.
  • the shim is augmented with an upstanding wall and a seal element to seal the gap that exists between platform edges of adjacent fan blades. This simple combination solves two complex problems, fatigue of fan assembly parts and loss of operating efficiency caused by fluid flow leakage.
  • US 2014/286781 A1 discloses a fan section for a gas turbine engine that includes a fan hub having a slot. A platform is supported by the fan hub. A fan blade has a root that is received in the slot. A wear pad is provided between the fan hub and the root and includes a flap integral with the wear pad to provide a seal relative to the platform.
  • a fan blade for a gas turbine engine includes an airfoil extending from a root. A wear pad is secured to the root and has a free end providing a flap canted toward the root.
  • a rotor assembly for a gas turbine engine, the rotor assembly comprises a rotational axis, a disc and at least one blade.
  • the disc comprises at least one slot.
  • the slot has at least one pair of bearing surfaces.
  • the blade has a pressure side and a suction side and comprises a root portion having a root fixture and radially outwardly a platform.
  • the platform extends circumferentially on both the pressure side and the suction side and has a radially inner surface.
  • the root fixture comprises at least one pair of bearing surfaces.
  • a sleeve is located at least partly between the slot and the root fixture. The sleeve is in contact with the slot’s bearing surfaces and the root fixture’s bearing surfaces.
  • the sleeve extends and is in contact with at least a part of the radially inner surface on both the pressure side and the suction side.
  • the sleeve may extend and be in contact with all of either one or both the radially inner surface on the pressure side and the suction side.
  • the sleeve prevents deposits of harmful substances on surfaces of the blade which might otherwise cause corrosion of the blade below the platform.
  • the sleeve enables the blade to be formed having a single crystal structure and with a relatively low chromium content. Such a single crystal blade with low Cr content is particularly robust and particularly resistant to creep. Low Cr content superalloys are susceptible to corrosion. Parts of the blade exposed to hot working gases are less susceptible to corrosive particle deposits.
  • the root portion may comprise a neck, the neck is between the root fixture and the platform.
  • the sleeve may be in contact with the neck.
  • the sleeve may be formed of a superalloy comprising a first composition having a Cr content.
  • the blade may be formed of a superalloy comprising a second composition having a Cr content.
  • the second composition may have a Cr content greater than the first composition.
  • the second composition may have a Cr content % by weight greater than 3% Cr the first composition having a Cr content % by weight.
  • the sleeve may be formed from the first composition having a Cr content at least 12% by weight.
  • the sleeve may be formed from the first composition having a Cr content less than 25% by weight.
  • the sleeve may be formed from the first composition having a Cr content approximately 22% by weight.
  • the blade may be formed from the second composition having a Cr content between 3% and 18% by weight.
  • the blade may be formed from the second composition having a Cr content less than 12%.
  • a line may have a constant radius from the rotational axis and may pass through the radially outward edge of the bearing surfaces of the slot.
  • the sleeve may be in contact with substantially all the surface of the slot radially inwardly of the line.
  • a line may have a constant radius from the rotational axis and may pass through the radially outward edge of the bearing surfaces of the root portion.
  • the sleeve may be in contact with substantially all the surface of the root portion radially inwardly of the line.
  • the root portion may comprise a radially inner surface and a cooling passage through the radially inner surface.
  • the sleeve may comprise a cooling air inlet that is aligned with cooling passage to allow cooling air to pass into the root portion.
  • the sleeve may comprise an end plate. Preferably the end plate is at a leading end of the sleeve. The end plate is useful to locate the sleeve to the blade and to protect the end of the blade’s root fixture against corrosive deposits.
  • the method may comprise inserting the sleeve into the slot and inserting the root fixture into the sleeve.
  • the method may comprise peening the sleeve against the bearing surfaces of the slot prior to inserting the root fixture into the sleeve.
  • the method may comprise inserting the root fixture into the sleeve and inserting the root fixture and the sleeve into the slot.
  • the method may comprise peening the sleeve against the bearing surfaces of the root fixture prior to inserting the root fixture and the sleeve into the slot.
  • the sleeve may be formed monolithically by a casting process or an additive manufacturing process.
  • FIG. 1 shows part of a turbine engine in a sectional view and in which a rotor assembly in accordance with the present invention is incorporated
  • FIG. 2 shows a first embodiment of a radial section through a rotor assembly in accordance with the present invention
  • FIG. 3 shows a second embodiment of a radial section through a rotor assembly in accordance with the present invention
  • FIG. 4 is a perspective view of a sleeve which is part of the rotor assembly and in accordance with the present invention
  • FIG. 5 is an enlarged view on area A shown in FIG. 3.
  • FIG. 1 shows an example of a gas turbine engine 10 in a sectional view.
  • the gas turbine engine 10 comprises, in flow series, an inlet 12, a compressor section 14, a combustor section 16 and a turbine section 18 which are generally arranged in flow series and generally about and in the direction of a longitudinal or rotational axis 20.
  • the gas turbine engine 10 further comprises a shaft 22 which is rotatable about the rotational axis 20 and which extends longitudinally through the gas turbine engine 10.
  • the shaft 22 drivingly connects the turbine section 18 to the compressor section 14.
  • air 24 which is taken in through the air inlet 12 is compressed by the compressor section 14 and delivered to the combustion section or burner section 16.
  • the burner section 16 comprises a burner plenum 26, one or more combustion chambers 28 and at least one burner 30 fixed to each combustion chamber 28.
  • the combustion chambers 28 and the burners 30 are located inside the burner plenum 26.
  • the compressed air passing through the compressor 14 enters a diffuser 32 and is discharged from the diffuser 32 into the burner plenum 26 from where a portion of the air enters the burner 30 and is mixed with a gaseous or liquid fuel.
  • the air/fuel mixture is then burned and the combustion gas 34 or working gas from the combustion is channelled through the combustion chamber 28 to the turbine section 18 via a transition duct 17.
  • This exemplary gas turbine engine 10 has a cannular combustor section arrangement 16, which is constituted by an annular array of combustor cans 19 each having the burner 30 and the combustion chamber 28, the transition duct 17 has a generally circular inlet that interfaces with the combustor chamber 28 and an outlet in the form of an annular segment.
  • An annular array of transition duct outlets form an annulus for channelling the combustion gases to the turbine 18.
  • the turbine section 18 comprises a number of blade carrying discs 36 attached to the shaft 22.
  • two discs 36 each carry an annular array of turbine blades 38.
  • the number of blade carrying discs could be different, i.e. only one disc or more than two discs.
  • guiding vanes 40 which are fixed to a stator 42 of the gas turbine engine 10, are disposed between the stages of annular arrays of turbine blades 38. Between the exit of the combustion chamber 28 and the leading turbine blades 38 inlet guiding vanes 44 are provided and turn the flow of working gas onto the turbine blades 38.
  • the turbine section 18 comprises a turbine casing 51 that surrounds the turbine blades 38 and defines a radially outer surface 55 of a working gas path 57.
  • a radially inner surface 59 further defines the working gas path 57. Parts of the radially inner surface 59 and the radially outer surface 55 are defined by platforms of the blades and stator vanes as is well known. A tip gap is defined between the tip of the blades 38 and the casing 51 or radially outer surface 55.
  • the combustion gas from the combustion chamber 28 enters the turbine section 18 and drives the turbine blades 38 which in turn rotate the shaft 22.
  • the guiding vanes 40, 44 serve to optimise the angle of the combustion or working gas on the turbine blades 38.
  • the turbine section 18 drives the compressor section 14.
  • the compressor section 14 comprises an axial series of vane stages 46 and rotor blade stages 48.
  • the rotor blade stages 48 comprise a rotor disc supporting an annular array of blades.
  • the compressor section 14 also comprises a compressor casing 50 that surrounds the rotor stages and supports the vane stages 46.
  • the guide vane stages include an annular array of radially extending vanes that are mounted to the compressor casing 50. The vanes are provided to present gas flow at an optimal angle for the blades at a given engine operational point.
  • Some of the guide vane stages have variable vanes, where the angle of the vanes, about their own longitudinal axis, can be adjusted for angle according to airflow characteristics that can occur at different engine operations conditions.
  • the compressor casing 50 defines a radially outer surface 52 of the passage 56 of the compressor 14.
  • a radially inner surface 54 of the passage 56 is at least partly defined by a rotor drum 53 of the rotor and which is partly defined by the annular array of blades 48.
  • the present invention is described with reference to the above exemplary turbine engine having a single shaft or spool connecting a single, multi-stage compressor and a single, one or more stage turbine.
  • the present invention is equally applicable to two or three shaft engines and which can be used for industrial, aero or marine applications.
  • upstream and downstream refer to the flow direction of the airflow and/or working gas flow through the engine unless otherwise stated.
  • forward and rearward refer to the general flow of gas through the engine.
  • axial, radial and circumferential are made with reference to the rotational axis 20 of the engine.
  • Figures 2 and 3 are sections through part of a rotor assembly 70 and are viewed in the axial direction. Figures 2 and 3 show first and second embodiments of the rotor assembly 70 respectively.
  • the rotor assembly 70 comprises a disc 36 and when fully assembled an annular array of radially extending blades 38, only one of which is shown.
  • the disc 36 comprises an annular array of disc slots 72, only one of which is shown, evenly spaced around the radially outer periphery 74 of the disc 36.
  • Each blade 38 comprises a root portion 84 having a root fixture 86 and a neck 88. The neck 88 may be optional on other blades.
  • the root fixture 86 blends into a platform 60 which has a radially inner surface 64 and a radially outer surface 62 or gas washed surface 62.
  • Extending, in a radially outward direction, from the platform 60 is an aerofoil 66 comprising a leading edge and a tailing edge.
  • the blade 38 has a pressure side and a suction side and the platform 60 extends circumferentially from the blade 38 on both the pressure side and the suction side.
  • a generally concave pressure surface and a generally convex suction surface each extend between the leading and trailing edges in conventional manner.
  • the aerofoil’s 66 pressure and suction surfaces also extend in a radial direction from the platform 60 to a tip of the aerofoil 66 as is well known.
  • the blade 80 shown here is an unshrouded blade, but in other examples the blade 80 could be a shrouded blade or a blade having one or more winglets as known in the art.
  • a tip gap is formed between the tip of the blade and the casing 51.
  • the root fixture 86 comprises a fir- tree root configuration and the slot 72 comprises a generally corresponding shape as is well known.
  • each disc slot 72 has two (or more) pairs of bearing surfaces 76, 78 and 80, 82 where one bearing surface 76, 80 opposes the other bearing surface 78, 82 of its pair respectively.
  • the root fixture 86 comprises two pairs of lobes 90, 92 and 94, 96 respectively. Each lobe 90, 92, 94,
  • the root fixture 86 is a single lobe dovetail design and the slot 72 has a generally corresponding shape as is well known.
  • the slot 72 has a pair of bearing surfaces 76, 78 where one bearing surface 76 opposes the other bearing surface 78.
  • the root fixture 86 comprises a pair of lobes 90, 92 respectively.
  • Each lobe 90, 92 defines a bearing surface 98, 100 respectively.
  • loads from the blade 38 are transmitted through corresponding bearing surfaces, that is the blade’s bearing surfaces 98, 100 and the slot’s bearing surfaces 76, 78 respectively.
  • the rotor assembly 70 further comprises a sleeve 110 disposed at least partly between the root fixture 86 and the slot 72.
  • the sleeve 110 is located between the blade’s bearing surfaces 98, 100, 102, 104 and the slot’s bearing surfaces 76, 78, 80, 82 and as such transmits loads, in use, from the blade 38 to the disc 72 via the respective bearing surfaces.
  • the sleeve 110 extends and is in contact with at least a part of the radially inner surface 64 on both the pressure side and the suction side of the platform 60.
  • the sleeve 110 covers the whole of the inner surface 64 of the platform 60.
  • the sleeve 110 is assembled to the slot 72 first.
  • the sleeve 110 has an external surface 114 that is in contact with the surface 73 of the slot 72 radially inwardly of the radially outermost part of the bearing surfaces 76, 78 as shown by line 112.
  • Line 112 is defined by a constant radius from the rotational axis 20 and intersects the radially outermost part of the bearing surfaces 76, 78 of the slot 72.
  • Dashed line R denotes a radial line from the engine rotational axis 20.
  • an internal surface 116 of the sleeve 110 is in contact with the surface of the neck 88 and the radially inner surface 64 of the platform 64 and, radially inwardly of the line 112, the sleeve 110 only contacts the bearing surfaces 98, 100, 102, 104 of the blade 38.
  • a gap is defined between the sleeve 110 and the non-bearing surface of the lobes 90, 92 and another gap 113 is defined between the non-bearing surface of the lobes 94, 96.
  • the sleeve 110 is assembled to the blade 38 first. Then the blade 38 and sleeve 110 are inserted into the slot 72.
  • the internal surface 116 of the sleeve 110 is in contact with the surfaces of the blade 38 comprising the radially inner surface 64 of the platform 64, the surface of the neck 88 and the surface of the root portion 84.
  • a line 113 has a constant radius from the rotational axis 20 and passes through the radially outward edge of the bearing surfaces 98, 100 of the root portion 68.
  • the inner surface 116 of the sleeve 110 is in contact with substantially all the surface 69 of the root portion 68 radially inwardly of the line 112. When the rotor assembly 70 is assembled the sleeve 110 is in contact with slot’s bearing surfaces 76, 78, 80, 82 only.
  • the sleeve 110 may be assembled to the blade 38 (first) whether the root fixture 86 comprises a fir-tree configuration (multiple lobe design) or a single lobe dovetail design as described with reference to the first or the second embodiments respectively. Similarly, the sleeve 110 may be assembled to the slot 72 whether the slot 72 comprises a fir-tree configuration or a single lobe dovetail design as described with reference to the first or the second embodiments respectively.
  • the sleeve 110 is formed of a superalloy comprising a first composition having a Cr content % by weight and the blade 38 is formed of a superalloy comprising a second composition having a Cr content % by weight.
  • the second composition has a Cr % by weight content that is greater than the first composition of the sleeve 110.
  • the Cr content of the second composition is greater than 3% by weight Cr of the first composition.
  • the sleeve 110 is formed from the first composition having a Cr content at least 12% by weight.
  • the maximum Cr content may be less than 25% by weight.
  • the first composition has a Cr content approximately 22% by weight.
  • the blade 38 is formed from the second composition having a Cr content between 3% and 18% by weight. Although preferably the second composition may have a Cr content less than 12% by weight.
  • its superalloy composition may be dependant on commercially available alloys.
  • one example of a superalloy composition has nominal constituents, Cr 22%, Co 19%, W2%, Ta 1.1%, Al 2.3%, Ti 3.5%, Nb 0.8%, Hf 0.8%, Ni balance by weight.
  • the longevity of the blade 38 is thereby not compromised by exposure to chemicals, particularly, sulphates which can corrode the blade 38 reducing it life. This corrosion is particularly severe radially inwardly of the upper surface 62 of the platform 64 where such corrosive chemicals can collect.
  • the sleeve 110 has a higher Cr content than the blade 38.
  • This higher Cr content means that the sleeve 110 has a greater corrosion resistance than the material of the blade.
  • the lower Cr content of the blade means it has a greater strength and resistance to creep in service.
  • the Cr content is at least 3% by weight greater in the composition of the sleeve 110 than in the composition of the blade 38.
  • the sleeve 110 may be formed by pressing, an additive manufacturing technique or casting. It is possible to fabricate the sleeve 110 in two halves, split approximately at the intersection with the radial line 112 as shown in Figures 2 and 3, assemble the two halves to the blade or slot and then using a known joining technique join the two halves together.
  • the sleeve 110 may be provided with cooling air inlets 118 that align with cooling passage 119 in the radially inner surface 117 of the blade 38.
  • the sleeve 110 may be peened against the bearing surfaces 76, 78, 80, 82 of the slot 72 prior to inserting the root fixture 86 into the sleeve 110.
  • the method may comprise peening the sleeve 110 against the bearing surfaces 98, 100, 102, 104 of the root fixture 86 prior to inserting the root fixture 86 and the sleeve 110 into the slot 72.
  • the sleeve 100 may be peened against all of the slot 72 in the first embodiment or against the surfaces of the blade the sleeve is in contact with in the second embodiment.
  • the sleeve 110 is shown in a perspective view.
  • the internal surface 116 of the sleeve 110 can be seen with the blade 38 removed.
  • the internal surface 116 would otherwise be in contact with the underside surface 64 of the platform 60.
  • the sleeve 110 comprises a slot bearing surface 124 and a blade bearing surface 126.
  • the slot bearing surface 124 and the blade bearing surface 126 are on opposite sides of the thickness 120 of the sleeve 110.
  • an end plate 127 is provided at a leading end 128 of the sleeve 110.
  • the leading end 128 is the axially forward end of the blade and slot.
  • the sleeve 110 is monolithically formed with the end plate 127, particularly where the sleeve 110 is manufactured by an additive manufacturing process or casting, or the end plate is integrally formed by welding or other joining technique if the sleeve is formed by pressing.
  • the end plate 127 may be peened against the root portion of the blade if applied to the blade first as in the second embodiment.
  • the slot 72 may require modifying such that a depth 120 of material is removed from its bearing surfaces 76, 78, 80, 82.
  • the depth of material removed may be equivalent to the thickness of the sleeve 110 such that the tip of the blade 38 is in the same position as in the existing rotor assembly 70 without the sleeve 110.
  • the tip of the blade 38 remains in the same nominal position prior to application of the sleeve 110 and the tip gap is maintained.
  • the sleeve 110 has a thickness of 3mm, but its thickness may be between 1mm and 4mm.
  • the superalloys referred to herein are any commercially available compositions such as Hastelloy ® , Inconel ® , Waspaloy ® , Rene ® alloys, Incoloy, MP98T, TMS alloys, and CMSX ® single crystal alloys. Other superalloy compositions are intended to be included herein.
  • the superalloys are nickel-based superalloys, but other alloys are possible.
  • the constituents of these superalloys may be adapted to comprising the prescribed Cr content % by weight.

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  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Chemical & Material Sciences (AREA)
  • Materials Engineering (AREA)
  • Manufacturing & Machinery (AREA)
  • Ceramic Engineering (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
  • Structures Of Non-Positive Displacement Pumps (AREA)

Abstract

Ensemble rotor (70) pour un moteur à turbine à gaz (10), l'ensemble rotor (70) comprenant un axe de rotation (20), un disque (36) et au moins une pale (38). Le disque (36) comprend au moins une fente (72), la fente (72) présente au moins une paire de surfaces d'appui (76, 78, 80, 82). La pale (38) comprend une partie de base (84) comportant un dispositif de fixation de base (86) et radialement vers l'extérieur une plateforme (60), la plateforme (60) présente une surface radialement interne (64), le dispositif de fixation de base (86) comprend au moins une paire de surfaces d'appui (98, 100, 102, 104). Un manchon (110) est situé au moins partiellement entre la fente (72) et le dispositif de fixation de base (86), le manchon (110) est en contact avec les surfaces d'appui (76, 78, 80, 82) de la fente et les surfaces d'appui (98, 100, 102, 104) du dispositif de fixation de base, le manchon (110) s'étend et est en contact avec au moins une partie de la surface radialement interne (64).
PCT/EP2022/061515 2021-06-11 2022-04-29 Ensemble rotor pour un moteur à turbine à gaz, procédé d'assemblage d'un ensemble rotor et procédé de fabrication d'un manchon WO2022258257A1 (fr)

Applications Claiming Priority (2)

Application Number Priority Date Filing Date Title
GB2108382.9A GB2607886A (en) 2021-06-11 2021-06-11 Rotor assembly and method of assembling a rotor assembly for a gas turbine engine
GB2108382.9 2021-06-11

Publications (1)

Publication Number Publication Date
WO2022258257A1 true WO2022258257A1 (fr) 2022-12-15

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Citations (8)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US4207029A (en) 1978-06-12 1980-06-10 Avco Corporation Turbine rotor assembly of ceramic blades to metallic disc
WO1996041068A1 (fr) 1995-06-07 1996-12-19 National Research Council Of Canada Barriere anti-usure par contact
US20020044870A1 (en) 2000-10-17 2002-04-18 Honeywell International, Inc. Fan blade compliant layer and seal
US20110268564A1 (en) * 2008-11-05 2011-11-03 Martin Nicholas F Gas Turbine with Securing Plate Between Blade Base and Disk
US20140286781A1 (en) 2013-01-11 2014-09-25 United Technologies Corporation Integral fan blade wear pad and platform seal
US20160090841A1 (en) * 2014-09-29 2016-03-31 United Technologies Corporation Gas turbine engine blade slot heat shield
EP3075880A1 (fr) 2015-04-01 2016-10-05 Siemens Aktiengesellschaft Lame à double alliage
FR3099520A1 (fr) 2019-07-31 2021-02-05 Safran Aircraft Engines Roue de turbomachine

Patent Citations (8)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US4207029A (en) 1978-06-12 1980-06-10 Avco Corporation Turbine rotor assembly of ceramic blades to metallic disc
WO1996041068A1 (fr) 1995-06-07 1996-12-19 National Research Council Of Canada Barriere anti-usure par contact
US20020044870A1 (en) 2000-10-17 2002-04-18 Honeywell International, Inc. Fan blade compliant layer and seal
US20110268564A1 (en) * 2008-11-05 2011-11-03 Martin Nicholas F Gas Turbine with Securing Plate Between Blade Base and Disk
US20140286781A1 (en) 2013-01-11 2014-09-25 United Technologies Corporation Integral fan blade wear pad and platform seal
US20160090841A1 (en) * 2014-09-29 2016-03-31 United Technologies Corporation Gas turbine engine blade slot heat shield
EP3075880A1 (fr) 2015-04-01 2016-10-05 Siemens Aktiengesellschaft Lame à double alliage
FR3099520A1 (fr) 2019-07-31 2021-02-05 Safran Aircraft Engines Roue de turbomachine

Also Published As

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GB2607886A (en) 2022-12-21
GB202108382D0 (en) 2021-07-28

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