EP3315721B1 - Renfort de bord d'attaque d'une aube de turbomachine - Google Patents

Renfort de bord d'attaque d'une aube de turbomachine Download PDF

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Publication number
EP3315721B1
EP3315721B1 EP17197595.6A EP17197595A EP3315721B1 EP 3315721 B1 EP3315721 B1 EP 3315721B1 EP 17197595 A EP17197595 A EP 17197595A EP 3315721 B1 EP3315721 B1 EP 3315721B1
Authority
EP
European Patent Office
Prior art keywords
blade
point
fin
radially outer
edge
Prior art date
Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
Active
Application number
EP17197595.6A
Other languages
German (de)
English (en)
French (fr)
Other versions
EP3315721A1 (fr
Inventor
Jean-Louis Romero
Jean-François FREROT
Current Assignee (The listed assignees may be inaccurate. Google has not performed a legal analysis and makes no representation or warranty as to the accuracy of the list.)
Safran Aircraft Engines SAS
Original Assignee
Safran Aircraft Engines SAS
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 Safran Aircraft Engines SAS filed Critical Safran Aircraft Engines SAS
Publication of EP3315721A1 publication Critical patent/EP3315721A1/fr
Application granted granted Critical
Publication of EP3315721B1 publication Critical patent/EP3315721B1/fr
Active legal-status Critical Current
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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
    • F01D5/147—Construction, i.e. structural features, e.g. of weight-saving hollow 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
    • F01D5/288—Protective coatings for 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
    • F01D21/00—Shutting-down of machines or engines, e.g. in emergency; Regulating, controlling, or safety means not otherwise provided for
    • F01D21/04—Shutting-down of machines or engines, e.g. in emergency; Regulating, controlling, or safety means not otherwise provided for responsive to undesired position of rotor relative to stator or to breaking-off of a part of the rotor, e.g. indicating such position
    • F01D21/045—Shutting-down of machines or engines, e.g. in emergency; Regulating, controlling, or safety means not otherwise provided for responsive to undesired position of rotor relative to stator or to breaking-off of a part of the rotor, e.g. indicating such position special arrangements in stators or in rotors dealing with breaking-off of part of rotor
    • 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
    • F04—POSITIVE - DISPLACEMENT MACHINES FOR LIQUIDS; PUMPS FOR LIQUIDS OR ELASTIC FLUIDS
    • F04D—NON-POSITIVE-DISPLACEMENT PUMPS
    • F04D29/00—Details, component parts, or accessories
    • F04D29/26—Rotors specially for elastic fluids
    • F04D29/32—Rotors specially for elastic fluids for axial flow pumps
    • F04D29/321—Rotors specially for elastic fluids for axial flow pumps for axial flow compressors
    • F04D29/324—Blades
    • 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
    • F04D29/26—Rotors specially for elastic fluids
    • F04D29/32—Rotors specially for elastic fluids for axial flow pumps
    • F04D29/38—Blades
    • F04D29/388—Blades characterised by construction
    • 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
    • F05D2220/00—Application
    • F05D2220/30—Application in turbines
    • F05D2220/36—Application in turbines specially adapted for the fan of turbofan engines
    • 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
    • F05D2240/303—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 leading edge 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
    • F05D2250/00—Geometry
    • F05D2250/70—Shape
    • F05D2250/71—Shape curved
    • F05D2250/711—Shape curved convex
    • 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
    • F05D2300/133—Titanium
    • 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
    • F05D2300/174—Titanium alloys, e.g. TiAl

Definitions

  • the present invention relates to a turbomachine blade, and more particularly to a leading edge reinforcement of this blade.
  • blade is meant here both the moving blades and the stationary blades of the turbomachines.
  • these include a leading edge reinforcement whose role is to protect the leading edge from deterioration during an impact with a FOD and to distribute the force of the impact over a large surface of the blade.
  • a reinforcement of the blade leading edge conventionally comprises an upper surface fin covering at least partially the aerodynamic surface of the upper surface of the blade and a lower surface fin covering at least partially the aerodynamic surface of the lower surface of the blade. blade, these two fins being joined by a leading edge of the reinforcement.
  • Document EP2540974 A discloses a reinforcement of the leading edge of a fan blade.
  • the detachment of the upper surface fin causes damage to the internal abradable layer.
  • the extrados fin protrudes from the extrados surface of the blade and penetrates into the internal abradable layer, which creates a groove in the internal abradable layer. It is then necessary to immobilize the turbine engine to replace both the blade, the leading edge reinforcement of which has come off, and the internal abradable layer. Such an immobilization generates a significant cost resulting from the lack of exploitation of the turbomachine which it is important to reduce, or even eliminate.
  • the object of the invention is in particular to provide a simple, effective and economical solution to this problem.
  • the invention proposes a blade according to claim 1.
  • the spacing of the downstream point from the upper edge of the extrados fin makes it possible to limit the penetration of the fin into the internal abradable layer of the turbomachine, in the event of separation of the downstream point of the fin since this it then finds itself far from the abradable due to its remoteness when mounting the blade tip.
  • the upstream point is located at the level of the upstream end of the upper edge, that is to say at the level of the leading edge of the blade and the downstream point is located at the downstream end of the radially outer edge of the fin.
  • downstream point is spaced radially inward from the tip of the blade.
  • the separation into two portions offers a good compromise between limiting penetration of the fin into the internal layer of abradable in the event of separation of the fin, and good distribution of the forces in the event of impact of a FOD on the backstop leading edge. downstream of the blade since the latter then finds itself far from the abradable due to its distance from the assembly of the blade tip.
  • the second portion of the radially outer edge of the extrados fin is curved convex. This particular shape makes it possible to facilitate the manufacture of the reinforcement and also to limit the creation of disturbances in the flow of the air flow.
  • the distance indicated in claim 1 offers a good compromise between limitation of penetration of the fin into the internal layer of abradable in the event of separation of the fin, and good distribution of the forces in the event of impact of a FOD on the leading edge sleeps.
  • the reinforcement comprises a lower surface fin partly covering an aerodynamic surface of the lower surface of the blade.
  • This lower surface fin also protects the aerodynamic surface of the lower surface of the blade against FOD.
  • the leading edge reinforcement is made of a metallic material.
  • the invention proposes, secondly, an assembly comprising a central disc on which are mounted a plurality of vanes as previously described, said vanes being regularly distributed around the periphery of the central disc, and extending substantially radially to the disc central.
  • the invention proposes, thirdly, a turbomachine comprising an assembly as previously described.
  • a turbine engine 2 having an assembly 4 comprising a central disc 6 rotating around a longitudinal axis A of the turbine engine 2, and on which is mounted a plurality of blades 8.
  • the blades 8 are regularly distributed around the periphery 6a of the central disc 6, and extending substantially radially to the central disc 6.
  • assembly 4 is the fan of turbomachine 2, and blades 8 are fan blades.
  • the turbomachine 2 also comprises, from upstream to downstream, and downstream of the fan, a low-pressure compressor 10, a high-pressure compressor 12, a combustion chamber 14, a high-pressure turbine 16, a low-pressure turbine 18 , and an exhaust casing.
  • the turbine engine 2 comprises attachment means 22, in this case two, each carried by an intermediate fan casing 24 carrying an internal abradable layer 24a (visible on the figure 4 ), and a casing 26 of the turbine.
  • radial means any direction substantially perpendicular to the axis A of the turbomachine 2, upstream the side by which the air reaches a part of the turbomachine 2, and downstream the side by which the air moves away from said part of the turbomachine 2.
  • the direction of air flow is represented by figure 2 by arrow F.
  • blade 8 is meant here both the mobile blades 8 (for example the rotor blades) and the stationary blades (for example the stator blades) of the turbomachines 2.
  • Dawn 8 illustrated in perspective on the figure 2 and in section on picture 3 , comprises an aerodynamic upper surface 28 and an aerodynamic lower surface 30 which extend in a first direction between a leading edge 8a and a trailing edge 8b of the blade 8.
  • the blade 8 of a fan being twisted, the first direction evolves in an XY plane along the section taken in a radial direction along the Z axis which forms with the X and Y axes an orthonormal reference mark on the figure 2 .
  • the aerodynamic upper surface 28 and the aerodynamic lower surface 30 extend between a root 8c and a tip 8d of the blade 8.
  • the blade 8 also comprises a leading edge reinforcement 32 comprising an extrados fin 32a partly covering the aerodynamic extrados surface 28 of the substantially radial blade 8, and an intrados fin 32b partly covering the aerodynamic surface 30 of the lower surface of the blade 8.
  • a leading edge reinforcement 32 comprising an extrados fin 32a partly covering the aerodynamic extrados surface 28 of the substantially radial blade 8, and an intrados fin 32b partly covering the aerodynamic surface 30 of the lower surface of the blade 8.
  • the two fins 32a, 32b are joined by a leading edge 32c which covers the leading edge 8a of the blade 8 and has, in section, a thickness greater than the maximum thickness of the fins 32a, 32b.
  • the reinforcement 32 of the leading edge 8a of the blade 8 extends substantially from the root 8c of the blade 8, up to its top 8d.
  • the leading edge reinforcement 32 is preferably made of a highly resistant metallic material, such as for example a titanium alloy.
  • the detail view of the figure 4 highlights a feature of the extrados fin 32a of the leading edge reinforcement 32.
  • the fin 32a of extrados has a radially outer (also called upper) edge 34 arranged in the vicinity of the blade tip 8d and which extends from the leading edge 8a towards the edge 8b ( figure 2 ) leak.
  • This radially outer edge 34 comprises an upstream point 34a which is flush with the top 8d of the blade 8 at the level of the leading edge 8a and a downstream point 34b which is spaced from the top 8d of the blade 8.
  • the term "upper” is understood according to the orientation of the figure 4 .
  • the radially outer edge 34 is arranged radially outward with respect to the axis A of the turbomachine 2.
  • upstream point 34a is arranged on the side of the leading edge 8a of the blade 8 and the downstream point 34b is arranged on the side of the trailing edge 8b of the blade 8 according to the direction of air flow F ( figure 2 ) on the blade 8 from the leading edge 8a to the trailing edge 8b.
  • the upper radially outer edge 34 of the extrados fin 32a comprises an intermediate point 34c located between the upstream point 34a and the downstream point 34b and defining with the upstream point 34a a first portion 36 of the radially outer edge, flush the top 8d of the blade 8 and, with the downstream point 34b, a second portion 38 of the upper edge gradually diverging from the top 8d of the blade 8.
  • the connection of the first portion 36 of the radially outer edge 34 with the second portion 38 of the upper edge is substantially tangential.
  • the intermediate point 34c is arranged at an equal distance from the upstream point 34a and from the downstream point 34b, in an axial direction parallel to the longitudinal axis A.
  • a fictitious extreme point 34th corresponding to the symmetry of the upstream point 34a with respect to a median axis M substantially perpendicular to the axis A of the turbomachine 2, and passing at least through the center of the top of the upper surface fin 32a .
  • This fictitious extreme point 34e corresponds to an extreme point of the extrados fin 32a before optimization of the latter.
  • this extreme point 34e makes it possible to define the progressive spacing of the downstream point 34b with respect to the top 8d of the blade 8.
  • the spacing of the second portion 38 from the radially outer edge 34 of the extrados fin 32a is preferably curved convex.
  • the second portion 38 has substantially a curved shape which deviates continuously from the top 8d of the blade 8 in the direction of the foot 8c ( figure 2 ) of the latter, and this from upstream to downstream.
  • the second portion 38 of the radially outer edge 34 of the extrados fin 32a could be rectilinear or, on the contrary, comprise an alternation of bumps and hollows.
  • the distance L, the tangent T and the angle ⁇ are illustrated on the figure 5 .
  • the lower surface fin 32b also comprises an upper edge having an upstream point flush with the top 8d of the blade 8 and a downstream point distant from the upstream point and separated from the top 8d of the blade 8, that is to say distant radially internally.
  • the upper edge of the intrados fin 32b may also comprise an intermediate point located between the point of attack and the vanishing point and defining with the point of attack a first portion of the upper edge, flush with the top 8d of the blade 8 and, with the vanishing point, a second portion of the upper edge gradually deviating from the top 8d of the blade 8 in the direction of the foot 8c.
  • the shapes and dimensions of the portions of the intrados fin 32b are reduced compared to the shapes and dimensions of the portions 36, 38 of the upper edge 34 of the extrados fin 32a.

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  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Chemical & Material Sciences (AREA)
  • Materials Engineering (AREA)
  • Architecture (AREA)
  • Composite Materials (AREA)
  • Structures Of Non-Positive Displacement Pumps (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
EP17197595.6A 2016-10-28 2017-10-20 Renfort de bord d'attaque d'une aube de turbomachine Active EP3315721B1 (fr)

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
FR1660479A FR3058181B1 (fr) 2016-10-28 2016-10-28 Renfort de bord d'attaque d'une aube de turbomachine

Publications (2)

Publication Number Publication Date
EP3315721A1 EP3315721A1 (fr) 2018-05-02
EP3315721B1 true EP3315721B1 (fr) 2022-03-02

Family

ID=58314355

Family Applications (1)

Application Number Title Priority Date Filing Date
EP17197595.6A Active EP3315721B1 (fr) 2016-10-28 2017-10-20 Renfort de bord d'attaque d'une aube de turbomachine

Country Status (4)

Country Link
US (1) US10316669B2 (zh)
EP (1) EP3315721B1 (zh)
CN (1) CN108005730B (zh)
FR (1) FR3058181B1 (zh)

Families Citing this family (10)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CN108454829A (zh) * 2018-05-30 2018-08-28 安徽卓尔航空科技有限公司 一种螺旋桨叶片
FR3085300B1 (fr) * 2018-08-31 2022-01-21 Safran Aircraft Engines Aube en materiau composite a film anti-erosion renforce et procede de protection associe
US10760428B2 (en) 2018-10-16 2020-09-01 General Electric Company Frangible gas turbine engine airfoil
US10746045B2 (en) 2018-10-16 2020-08-18 General Electric Company Frangible gas turbine engine airfoil including a retaining member
US11149558B2 (en) 2018-10-16 2021-10-19 General Electric Company Frangible gas turbine engine airfoil with layup change
US10837286B2 (en) 2018-10-16 2020-11-17 General Electric Company Frangible gas turbine engine airfoil with chord reduction
US11434781B2 (en) 2018-10-16 2022-09-06 General Electric Company Frangible gas turbine engine airfoil including an internal cavity
US11111815B2 (en) 2018-10-16 2021-09-07 General Electric Company Frangible gas turbine engine airfoil with fusion cavities
FR3103215B1 (fr) 2019-11-20 2021-10-15 Safran Aircraft Engines Aube de soufflante rotative de turbomachine, soufflante et turbomachine munies de celle-ci
FR3115079B1 (fr) * 2020-10-12 2022-10-14 Safran Aircraft Engines Aube en materiau composite comprenant un bouclier de bord d’attaque, turbomachine comprenant l’aube

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GB1304678A (zh) * 1971-06-30 1973-01-24
US5908285A (en) * 1995-03-10 1999-06-01 United Technologies Corporation Electroformed sheath
JP4390026B2 (ja) * 1999-07-27 2009-12-24 株式会社Ｉｈｉ 複合材翼
US7736130B2 (en) * 2007-07-23 2010-06-15 General Electric Company Airfoil and method for protecting airfoil leading edge
US8858182B2 (en) * 2011-06-28 2014-10-14 United Technologies Corporation Fan blade with sheath
FR2987867B1 (fr) * 2012-03-09 2016-05-06 Snecma Aube de turbomachine comportant un insert de protection de la tete de l'aube

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Also Published As

Publication number Publication date
FR3058181B1 (fr) 2018-11-09
EP3315721A1 (fr) 2018-05-02
CN108005730A (zh) 2018-05-08
US10316669B2 (en) 2019-06-11
FR3058181A1 (fr) 2018-05-04
US20180119551A1 (en) 2018-05-03
CN108005730B (zh) 2022-07-08

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