• 专利附录
  • 专利说明
  • 权利要求
  • 类似技术
  • 同族专利
  • 引用文献
  • 法律状态
  • 优先权
    • 专利摘要:
    The invention relates to a turbomachine blade, comprising an aerodynamic surface (28) extending in a first direction between a leading edge (8a) and a trailing edge, and in a second direction substantially perpendicular to the first direction between a root and a top (8d) of the blade, and a leading edge reinforcement (32) comprising a fin (32a) partially covering the aerodynamic surface (28) of the blade, characterized in that, the fin (32a) has a radially outer edge (34) arranged in the vicinity of the apex (8d) of the blade and extending between the leading edge (8a) and the trailing edge, this edge (34) radially outer comprising an upstream point (34a) flush with the apex (8d) of the blade at the edge (8a) of attack and a point (34b) said downstream spaced from the vertex (8d) of the blade.
    公开号:FR3058181A1
    申请号:FR1660479
    申请日:2016-10-28
    公开日:2018-05-04
    发明作者:Jean-Louis ROMERO；Jean-Francois Frerot
    申请人:Safran Aircraft Engines SAS；
    IPC主号:
    专利说明:

    Holder (s): SAFRAN AIRCRAFT ENGINES Simplified joint-stock company.
    Extension request (s)
    Agent (s): ERNEST GUTMANN - YVES PLASSERAUD SAS.
    REINFORCEMENT OF EDGE OF ATTACK OF A TURBOMACHINE BLADE.
    FR 3 058 181 - A1
    The invention relates to a turbomachine blade, comprising an aerodynamic surface (28) which extends in a first direction between a leading edge (8a) and a trailing edge, and in a second direction substantially perpendicular to the first direction between a foot and a vertex (8d) of the blade, and a reinforcement (32) of leading edge comprising a fin (32a) partially covering the aerodynamic surface (28) of the blade, characterized in that, the fin (32a) has a radially outer edge (34) arranged in the vicinity of the apex (8d) of the blade and extending between the leading edge (8a) and the trailing edge, this edge (34) Radially outer comprising an upstream point (34a) flush with the apex (8d) of the blade at the leading edge (8a) and a point (34b) said downstream away from the apex (8d) of the blade.
    EDGE REINFORCEMENT OF A TURBOMACHINE BLADE
    The present invention relates to a turbomachine blade, and more particularly a leading edge reinforcement of this blade.
    By blade here is meant both the movable blades and the fixed blades of turbomachinery.
    In order to increase the resistance of the blades to FOD (acronym for English Foreign Object Damages) in the air flow, that is to say to foreign bodies such as birds or hailstones, 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 impact force over a large area of the blade.
    A blade leading edge reinforcement conventionally comprises an upper fin at least partially covering the aerodynamic surface of the upper surface of the blade and a lower fin fin at least partially covering the aerodynamic surface of the lower surface of the blade, these two fins being joined by a leading edge of the reinforcement.
    When the blade is movable relative to the axis of the turbomachine, it turns its lower surface forward, that is to say that the air comes into contact with the lower surface thus creating an overpressure on the lower surface and a depression on its upper surface.
    The impact of a FOD on the leading edge reinforcement tends to cause the upper portion of the lower wing fin to detach. Beyond a certain mass of FOD, the force of the impacts is greater on the reinforcement, which also causes detachment of the upper portion of the upper surface fin. The overpressure generated on the lower surface tends to limit detachment of the lower fin from the lower surface. On the other hand, the combination of centrifugal force, more important at the top of the blade than at the bottom, with the depression generated on the surface of the upper surface, tends to favor the detachment of the upper surface fin.
    When the blade is a fan blade mounted in an external fairing carrying an internal layer of abradable facing the blades, the detachment of the upper surface fin causes damage to the internal layer of abradable. In fact, the upper wing protrudes from the surface of the dawn upper surface and penetrates into the inner layer of abradable, which creates a groove in the inner layer of abradable. It is therefore necessary to immobilize the turbomachine in order to replace both the vane, the leading edge reinforcement of which has come off and the internal layer of abradable. Such immobilization generates a significant cost resulting from the lack of operation 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.
    To this end, the invention proposes, firstly, a turbomachine blade extending along a longitudinal axis, comprising an aerodynamic surface which extends in a first direction between a leading edge and a trailing edge, and in a second direction substantially perpendicular to the first direction between a foot and a top of the blade, and a leading edge reinforcement comprising a fin partially covering the aerodynamic surface of the blade, characterized in that, the fin has a radially outer edge arranged in the vicinity of the apex of the blade and extending between the leading edge and the trailing edge, this radially outer edge comprising an upstream point flush with the apex of the blade at the edge of attack and a downstream point away from the summit of dawn.
    The spacing of the downstream point of the upper edge of the upper surface fin limits the penetration of the fin into the internal abradable layer of the turbomachine, in the event of detachment of the downstream point of the blade since it it is then found far from the abradable due to its distance when mounting the blade tip.
    In a particular embodiment of the invention, the upstream point is located at the upstream end of the upper edge, that is to say at the leading edge of the blade and the downstream point is located at the downstream end of the radially outer edge of the fin.
    In the reference frame of the turbomachine, it can thus be considered that the downstream point is moved radially towards the inside of the blade tip.
    Advantageously, the aerodynamic surface is a surface of the upper surface, and the fin is a fin of the upper surface, the upper surface part of the reinforcement being more particularly subject to detachment, detachment increased in particular by centrifugal force for a movable blade.
    Advantageously, the radially outer edge of the fin comprises an intermediate point situated between the upstream point and the downstream point and defining with the upstream point a first portion of the radially outer edge, flush with the top of the blade and, with the downstream point , a second portion of the radially outer edge gradually deviating from the top of the blade towards the vanishing point.
    The separation into two portions offers a good compromise between limiting the penetration of the fin into the internal abradable layer in the event of separation of the fin, and good distribution of the forces in the event of an impact of a FOD on the joint. leading edge.
    The intermediate point can be arranged longitudinally equidistant from the upstream point and the downstream point.
    This protects the dawn over the entire height since the first portion is flush with the top of the dawn.
    Preferably, the second portion of the radially outer edge of the upper wing is curved convex. This particular shape facilitates the manufacture of the reinforcement and, also, limits the creation of disturbances in the flow of the air flow.
    Advantageously, the intermediate point and the vanishing point are separated from each other by a distance, measured along a longitudinal median axis of the fin, between 0 and sina x L 4 where:
    - L is the length of the fin before optimization, that is to say between the upstream point and a fictitious end point corresponding to the symmetry of the upstream point with respect to the median axis substantially perpendicular to the longitudinal axis of the turbomachine , and passing at least through the center of the top of the fin, and
    - a is the angle measured between a line passing through the upstream point and the intermediate point of the radially outer edge and a tangent to the radially outer edge, parallel to the longitudinal axis and passing through the intermediate point.
    This distance also offers a good compromise between limiting the penetration of the fin into the internal abradable layer in the event of separation of the fin, and good distribution of forces in the event of an impact of a FOD on the edge joint. of attack.
    Preferably, the reinforcement comprises a lower surface fin partially covering an aerodynamic surface of the lower surface of the blade.
    This lower wing also protects the aerodynamic surface of the lower surface from dawn against FOD.
    To ensure good dawn protection, the leading edge reinforcement is made of a metallic material.
    The invention proposes, secondly, an assembly comprising a central disk on which are mounted a plurality of blades as previously described, said blades being regularly distributed around the periphery of the central disk, and extending substantially radially to the disk. central.
    The invention proposes, thirdly, a turbomachine comprising an assembly as previously described.
    The invention will be better understood and other details, characteristics and advantages of the invention will appear on reading the following description given by way of nonlimiting example with reference to the appended drawings in which:
    - Figure 1 is a schematic view of a turbomachine comprising an assembly having a plurality of blades;
    - Figure 2 is a perspective view of a blade according to the invention, in particular a fan blade, this blade carrying a leading edge reinforcement limiting the degradation of the internal abradable layer of the turbomachine;
    - Figure 3 is a sectional view of the blade along the section plane III - III of Figure 2;
    FIG. 4 is a detailed view of an upper portion of the blade according to insert IV of FIG. 2, and
    FIG. 5 is a detail view on an enlarged scale of detail V of FIG. 4.
    FIG. 1 shows a turbomachine 2 having an assembly 4 comprising a central disc 6 rotating around a longitudinal axis A of the turbomachine 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. In this case, the assembly 4 is the fan of the turbomachine 2, and the blades 8 are the fan blades.
    Conventionally, 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. In addition, for its attachment to the aircraft, the turbomachine 2 comprises attachment means 22, in this case two, each carried by an intermediate fan casing 24 carrying an internal abradable layer 24a (visible in the figure 4), and a turbine casing 26.
    In the remainder of this description, the term radial (e) 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 air flow direction is represented in FIG. 2 by the arrow F.
    By blade 8 is meant here both the movable blades 8 (for example the rotor blades) and the fixed blades (for example the stator blades) of the turbomachines 2.
    The blade 8, illustrated in perspective in FIG. 2 and in section in FIG. 3, comprises an aerodynamic surface 28 of the upper surface and an aerodynamic surface 30 of the lower surface which extend in a first direction between an edge 8a of attack 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 axes X and Y an orthonormal reference in FIG. 2. In a second direction substantially perpendicular to the first direction, the aerodynamic surface 28 of the upper surface and the aerodynamic surface 30 of the lower surface extend between a foot 8c and a vertex 8d of l 'dawn 8.
    The blade 8 also comprises a reinforcement 32 of the leading edge comprising a fin 32a of upper surface partially covering the aerodynamic surface 28 of upper surface of the substantially radial blade 8, and a fin 32b of lower surface partially covering the Aerodynamic surface 30 of the lower surface of the blade 8. These two fins 32a, 32b have, as visible in FIG. 3, a section which is thinner from upstream to downstream.
    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.
    As seen in Figure 2, the reinforcement 32a 8a leading edge of the blade 8 extends substantially from the foot 8c of the blade 8, 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 FIG. 4 highlights a particularity of the fin 32a of the upper surface of the leading edge reinforcement 32. Indeed, the upper wing 32a has a radially outer edge 34 (also called upper) arranged in the vicinity of the tip 8d of the blade and which extends from the leading edge 8a towards the edge 8b (FIG. 2) leak. This radially outer edge 34 comprises an upstream point 34a which is flush with the apex 8d of the blade 8 at the leading edge 8a and a downstream point 34b which is spaced from the apex 8d of the blade 8. The term "upper" is understood according to the orientation of FIG. 4. In other words the radially outer edge 34 is disposed radially outward relative to the axis A of the turbomachine 2.
    It will be understood that the 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 in the direction of air flow F (Figure 2) on the blade 8 from the leading edge 8a to the trailing edge 8b.
    In addition, the radially outer upper edge 34 of the upper wing 32a includes 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 apex 8d of the blade 8 and, with the downstream point 34b, a second portion 38 of the upper edge progressively deviating from the apex 8d of the blade 8. The connection of the first portion 36 of the edge 34 radially outside with the second portion 38 of the upper edge is substantially tangential.
    According to one aspect, the intermediate point 34c is arranged at equal distance from the upstream point 34a and the downstream point 34b, in an axial direction parallel to the longitudinal axis A. However, the intermediate point 34c could be closer to the upstream point 34a or the downstream point 34b.
    FIG. 5 shows a fictitious extreme point 34e corresponding to the symmetry of the point 34a upstream 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 apex of the upper wing 32a. This fictitious extreme point 34e corresponds to an end point of the upper wing 32a before optimization of the latter.
    Advantageously, this extreme point 34e makes it possible to define the progressive spacing of the downstream point 34b relative to the apex 8d of the blade 8.
    The spacing of the second portion 38 of the radially outer edge 34 of the upper wing 32a is preferably curved convex. In other words, the second portion 38 has a substantially curved shape which deviates continuously from the top 8d of the blade 8 in the direction of the foot 8c (FIG. 2) of the latter, and from upstream to downstream.
    However, according to variant embodiments not shown in the figures, the second portion 38 of the radially outer edge 34 of the upper wing 32a of the upper surface could be rectilinear or, on the contrary, include an alternation of bumps and depressions.
    According to a preferred embodiment represented in FIG. 5, the intermediate point 34c and the downstream point 34b are separated from each other by a distance H1 measured along the median longitudinal axis, ie ie in the radial direction Z, H1 being between 0 and sina x L -h 4 where:
    - L is the length of the fin 32a before optimization, that is to say between the point 34a upstream and the fictitious point 34e, and
    - A is the angle measured between a line passing through the upstream point 34a and the intermediate point 34c of the radially outer edge 34 and a tangent T to said radially outer edge 34, parallel to the longitudinal axis A of the turbomachine 2 and passing through point 34 c intermediate.
    The distance L, the tangent T and the angle a are illustrated in the figure
    5.
    Thus, in the event of a FOD impacting on the leading edge reinforcement 32, if the upper wing 32a comes to come off, it will not come into contact with the internal abradable layer 24a carried by the intermediate fan casing 24. Consequently, it will only be necessary to repair the blade 8 which has been impacted (or the blades 8 impacted), which is simpler, faster and less expensive than the complete immobilization of the turbomachine 2 for the replacement of vane 8 impacted (or vanes 8 impacted) and the intermediate casing 24 of the fan and its internal layer 24a of abradable.
    For reasons of simplicity of manufacture of the leading edge reinforcement 32, the underside fin 32b also includes an upper edge having an upstream point flush with the apex 8d of the blade 8 and a downstream point distant from the upstream point and spaced from the apex 8d of the blade 8, that is to say radially distant internally.
    The upper edge of the underside fin 32b may also include 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 deviating gradually from the top 8d of the blade 8 towards the foot 8c.
    However, the shapes and dimensions of the fin portions
    32b of the lower surface are reduced in relation to the shapes and dimensions of the portions 36, 38 of the upper edge 34 of the fin 32a of the upper surface.
    Thus, an asymmetrical reinforcement 32 will be obtained.
    权利要求:
    Claims (10)
    [1" id="c-fr-0001]
    1. Turbine engine blade (8) extending along a longitudinal axis (A), comprising an aerodynamic surface (28, 30) which
    5 extends in a first direction between a leading edge (8a) and a trailing edge (8b), and in a second direction substantially perpendicular to the first direction between a foot (8c) and a top (8d) of the blade (8), and a leading edge reinforcement (32) comprising a fin (32a, 32b) partially covering the aerodynamic surface (28, 30) of
    10 the blade (8), characterized in that, the fin (32a, 32b) has a radially outer edge (34) arranged in the vicinity of the apex (8d) of the blade (8) and extending between the leading edge (8a) and trailing edge (8b), this radially outer edge (34) comprising an upstream point (34a) flush with the apex (8d) of the blade (8) at the leading edge and a point (34b)
    15 downstream radially distant from the top (8d) of the blade (8).
    [2" id="c-fr-0002]
    2. A vane (8) according to claim 1, in which the aerodynamic surface is a surface (28) of upper surfaces, and the fin is a fin (32a) of upper surfaces.
    [3" id="c-fr-0003]
    3. Dawn (8) according to claim 1 or 2, wherein the edge
    20 (34) radially outside of the fin (32a, 32b) 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) from the radially outer edge (34), flush with the top (8d) of the blade (8) and, with the downstream point (34b), a second portion (38) of the edge (34) radially
    25 outside progressively deviating from the top (8d) of the blade (8) in the direction of the downstream point (34b).
    [4" id="c-fr-0004]
    4. Dawn (8) according to claim 3, wherein the intermediate point (34c) is arranged longitudinally equidistant from the upstream point (34a) and the downstream point (34b).
    [5" id="c-fr-0005]
    5. Dawn (8) according to any one of claims 3 to 4, wherein the second (38) portion of the edge (34) radially outside of the fin (32a, 32b) is curved convex.
    [6" id="c-fr-0006]
    6. Dawn (8) according to one of claims 3 to 5, in which the intermediate point (34c) and the downstream point (34b) are separated from each other by a distance (H1), measured on along a median longitudinal axis (M) of the fin, between 0 and sina x L 4- 4 where:
    - L is the length of the fin (32a) before optimization, that is to say between the upstream point (34a) and a fictitious extreme point (34e) corresponding to the symmetry of the upstream point (34a) with respect to the median axis (M) substantially perpendicular to the longitudinal axis (A) of the turbomachine (2), and passing at least through the center of the top of said fin, and
    - a is the angle measured between a line passing through the upstream point (34a) and the intermediate point (34c) of the radially outer edge (34) and a tangent (T) at the radially outer edge (34), parallel to the longitudinal axis (A) and passing through the intermediate point (34c).
    [7" id="c-fr-0007]
    7. Dawn (8) according to any one of the preceding claims, in which the reinforcement (32) of the leading edge comprises a fin (32b) of pressure surface partially covering an aerodynamic surface (30) of pressure surface 'dawn (8).
    [8" id="c-fr-0008]
    8. Dawn (8) according to any one of the preceding claims, in which the leading edge reinforcement (32) is made of a metallic material.
    [9" id="c-fr-0009]
    9. assembly (4) comprising a central disc (6) on which are mounted a plurality of blades (8) according to any one of the preceding claims, said blades (8) being regularly distributed around the periphery (6a) of the central disc (6), and extending substantially radially with respect to the central disc (6).
    [10" id="c-fr-0010]
    10. Turbomachine (2) comprising an assembly (4) according to claim 9.
    1/3
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    同族专利:
    公开号 | 公开日
    CN108005730A|2018-05-08|
    EP3315721A1|2018-05-02|
    EP3315721B1|2022-03-02|
    US10316669B2|2019-06-11|
    FR3058181B1|2018-11-09|
    US20180119551A1|2018-05-03|
    引用文献:
    公开号 | 申请日 | 公开日 | 申请人 | 专利标题
    GB2298653A|1995-03-10|1996-09-11|United Technologies Corp|Electroformed sheath|
    EP2540974A2|2011-06-28|2013-01-02|United Technologies Corporation|Fan blade with sheath|
    GB1304678A|1971-06-30|1973-01-24|
    JP4390026B2|1999-07-27|2009-12-24|株式会社Ｉｈｉ|Composite wing|
    US7736130B2|2007-07-23|2010-06-15|General Electric Company|Airfoil and method for protecting airfoil leading edge|
    FR2987867B1|2012-03-09|2016-05-06|Snecma|TURBOMACHINE DAWN COMPRISING A PROTECTIVE INSERT FOR THE HEAD OF THE DAWN|US11149558B2|2018-10-16|2021-10-19|General Electric Company|Frangible gas turbine engine airfoil with layup change|
    US10760428B2|2018-10-16|2020-09-01|General Electric Company|Frangible gas turbine engine airfoil|
    US11111815B2|2018-10-16|2021-09-07|General Electric Company|Frangible gas turbine engine airfoil with fusion cavities|
    US10837286B2|2018-10-16|2020-11-17|General Electric Company|Frangible gas turbine engine airfoil with chord reduction|
    US10746045B2|2018-10-16|2020-08-18|General Electric Company|Frangible gas turbine engine airfoil including a retaining member|
    法律状态:
    2017-09-20| PLFP| Fee payment|Year of fee payment: 2 |
    2018-05-04| PLSC| Publication of the preliminary search report|Effective date: 20180504 |
    2018-09-19| PLFP| Fee payment|Year of fee payment: 3 |
    2019-09-19| PLFP| Fee payment|Year of fee payment: 4 |
    2020-09-17| PLFP| Fee payment|Year of fee payment: 5 |
    2021-09-22| PLFP| Fee payment|Year of fee payment: 6 |
    优先权:
    申请号 | 申请日 | 专利标题
    FR1660479|2016-10-28|
    FR1660479A|FR3058181B1|2016-10-28|2016-10-28|REINFORCEMENT OF THE EDGE OF ATTACK OF A TURBOMACHINE BLADE|FR1660479A| FR3058181B1|2016-10-28|2016-10-28|REINFORCEMENT OF THE EDGE OF ATTACK OF A TURBOMACHINE BLADE|
    EP17197595.6A| EP3315721B1|2016-10-28|2017-10-20|Leading-edge reinforcement of a turbine engine blade|
    US15/794,765| US10316669B2|2016-10-28|2017-10-26|Reinforcement for the leading edge of a turbine engine blade|
    CN201711019798.4A| CN108005730A|2016-10-28|2017-10-27|Turbine engine blade leading edge reinforcer|
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