• 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 turbine engine blade (10) comprising a foot (12) and a blade (14) comprising a main median main plane (14) longitudinal and radial, which is carried by the foot, the blade (14). ) having a leading edge (24) at an upstream longitudinal end, a trailing edge (26) at a longitudinal end downstream of the leading edge (24) from a flow of gas flowing around the blade (14), a lower surface wall (28) and an extrados wall laterally spaced from each other and each connecting the leading edge (24) to the trailing edge (26), and a top (32) located at the free outer radial end of the blade (14), the blade (14) further comprising a fin (34) of principal orientation (14) longitudinal which is carried by the face of the lower face (28), which is located at the top (32) of the blade (14), characterized in that the fin (34) is shifted radially inwardly by port at the top (32) of the blade (14).
    公开号:FR3022295A1
    申请号:FR1455527
    申请日:2014-06-17
    公开日:2015-12-18
    发明作者:Erwan Daniel Botrel；Laurent Patrick Robert Coudert
    申请人:SNECMA SAS；
    IPC主号:
    专利说明:

    [0001] TECHNICAL FIELD The invention relates to a turbomachine blade made to limit the disturbances that may form at the head of the blade. The invention relates more particularly to a blade for a turbomachine high pressure turbine having a lateral fin located on the underside face of the blade. STATE OF THE PRIOR ART In a turbomachine high pressure turbine, the gases from the combustion chamber exert on the blades of the turbine significant temperature and pressure constraints. Thus, the combustion gases can bring the blades of the blades to temperatures beyond the limit of temperature allowed by the blades, which leads to incorporate the blades solutions for their cooling continuously. Also, the heating and cooling of the blades cause dimensional variations of the blades. These dimensional variations must be taken into account to prevent the top of each blade from coming into contact with the fixed casing of the turbine, which surrounds the blades.
    [0002] Thus, a radial operating clearance is present in the turbine between the top of the blades and the fixed casing. Because of the difference in pressure between the lower and upper surfaces of each blade, a vortex is formed at the radial clearance, which consequently produces aerodynamic losses which reduce turbine efficiency and produce additional localized heating. at the top of the blade.
    [0003] To limit this formation of turbulence, US-A-1,955,929 discloses a blade having a fin laterally extending the crown face of the blade, projecting from the underside face of the blade. Although this embodiment makes it possible to limit the turbulence that forms at the top of the blade, the presence of the blade increases the general moving mass of the blade, which induces additional constraints on the blade during the rotation of the blade. the turbomachine. Also, the fin can come into contact with the face opposite the housing when the blade expands too much, under the action of heat or under the action of centrifugal forces. The object of the invention is to propose a blade comprising a fin which is designed in such a way as to overcome the drawbacks mentioned above. DISCLOSURE OF THE INVENTION The invention proposes a turbomachine blade comprising a foot and a blade having a median main plane of longitudinal and radial main orientation, which is carried by the foot, the blade having a leading edge located at a longitudinal end upstream, a trailing edge located at a longitudinal end downstream of the leading edge with respect to a flow of gas flowing around the blade, a lower pressure wall and an extrados wall located laterally at a distance; one of the other and each connecting the leading edge to the trailing edge, and a vertex located at the free outer radial end of the blade, the blade further comprising a longitudinally oriented longitudinal fin which is carried by the intrados face, which is located at the top of the blade, characterized in that the fin is offset radially inward with respect to the top of the blade. The radially inward shift of the vane prevents the formation of vortices at the top of the blade and also limits the impact of the mass of the vane on the inertia of the blade.
    [0004] It will be understood that the main longitudinal orientation of the fin means that the fin may be longitudinal as well as inclined with respect to this longitudinal orientation. Preferably, the top of the blade has a hollow recess which is open radially outwardly and has a bottom wall located radially away from the top of the blade, and the fin is offset radially inwards relative to at the bottom wall of the recess. Preferably, the fin has an upstream longitudinal end which is offset longitudinally downstream relative to the leading edge of the blade. Preferably, the fin has a downstream longitudinal end which is offset longitudinally upstream relative to the trailing edge of the blade. Preferably, the blade comprises at least one bore which opens into a lateral end face of the fin and which opens into a cavity formed in the blade. Preferably, a median longitudinal plane of the fin is perpendicular to the main plane of the blade. Preferably, a median longitudinal plane of the fin is inclined relative to the main plane of the blade. Preferably, the connection between the underside face and the radially inner face of the fin is convex.
    [0005] Preferably, the upstream longitudinal end of the fin is offset radially relative to the downstream longitudinal end of the fin. The invention also relates to a turbomachine high-pressure turbine rotor comprising a plurality of blades according to the invention which are regularly distributed around the main axis of rotation of the rotor.
    [0006] The invention also relates to an aircraft turbine engine comprising a high pressure turbine in which a plurality of blades according to the invention are mounted.
    [0007] BRIEF DESCRIPTION OF THE DRAWINGS Other features and advantages of the invention will appear on reading the detailed description which follows for the understanding of which reference will be made to the appended figures among which: FIG. 1 is a diagrammatic representation in perspective of a turbine blade for a turbomachine comprising a fin according to the invention; FIG. 2 is a diagrammatic representation in perspective of the blade of the blade shown in FIG. 1; - Figures 3 to 6 are sections of the blade shown in Figure 2, showing different embodiments of the fin according to the invention. DETAILED DESCRIPTION OF PARTICULAR EMBODIMENTS FIG. 1 shows a turbine engine blade 10, and more particularly a blade 10 for a turbomachine high-pressure turbine which comprises a base 12 for mounting the blade 10 on a rotor disk. rotor of the turbomachine (not shown) and a blade 14 of radial main orientation which is adapted to cooperate with the gases from the combustion chamber of the turbomachine. The blade 10 also has an intermediate portion 16 whose radially outer face is profiled, to partially delimit the flow of flue gas flow.
    [0008] The blade 10 is intended to be subjected to hot gases from the combustion chamber (not shown). A plurality of orifices 18 and grooves 20 are distributed on the blade 14 to allow the circulation of a flow of cooling air which is injected through inlet orifices 22 formed in the root 12 of the blade. The blade 14 is defined by a leading edge 24 which is located at an upstream longitudinal end of the blade, a trailing edge 26 situated at a longitudinal end downstream of the blade, and therefore downstream of the leading edge, a concave intrados wall 28 on which the pressure of the combustion gases is exerted, a convex extrados wall located laterally away from the intrados wall 28, and a top 32 located at the outer radial end of the pale, farthest from foot 12 of dawn 10.
    [0009] The blade 14 also has a lateral fin 34 which is carried by the intrados face 28 and which opposes the formation of vortices at the top of the blade 14. Indeed, as can be seen in FIGS. 6, the combustion gases flow partly radially outwards along the underside face 28. They are then blocked by the fin 34 before reaching the top 32 of the blade 14. It is formed then a vortex 36 along the intrados wall 28 which is less energetic than vortices that form at the top of a blade 14 according to the prior art.
    [0010] According to the embodiment shown in the figures, this fin 34 extends in a plane of longitudinal orientation, that is to say that it is generally parallel to the end edge of the intrados face 28 of the 14 and the two longitudinal ends 38, 40 of the fin 34 are located at the same radial position. According to an alternative embodiment not shown, the fin is inclined relative to the longitudinal direction, that is to say that its upstream longitudinal end 38 is offset radially relative to the downstream longitudinal end 40 of the fin 34 Thus, either the upstream longitudinal end 38 of the fin 34 is located radially outwardly with respect to the downstream longitudinal end 40 of the fin 34, or vice versa, the longitudinal upstream end 38 of the fin 34 is located radially inward with respect to the downstream longitudinal end 40 of the fin 34. According to the invention, and as can be seen in the figures, the fin 34 is offset radially inwards relative to at the top 32 of the blade 14. This radial positioning of the fin 34 makes it possible to move it away from the wall of the stationary housing near which the top 32 of the blade 14 evolves, in order to avoid any contact during the operation of the turbomachine and therefore li miter any risk of damaging the fin 34. The interaction of the fin 34 with the flow of hot gases causes a heating of the fin 34 which must be mastered. For this purpose, the fin 34 comprises cooling means similar to the orifices 18 and grooves 20 formed in the blade 14.
    [0011] The fin 34 thus comprises at least one bore 42 which extends from a free lateral end face 44 of the fin to a cavity 46 formed inside the blade 14. This cavity 46 allows the communication between the orifices 18 and the grooves 20 with the inlet orifices 22. It is thus supplied with cooling air, which consequently also supplies each piercing 42 of the vane 34. As can be seen in FIGS. 3 to 6, the top 32 of the blade 14 comprises a recess 48 recessed commonly called "bathtub", which is open radially outwardly. This recess 48 has a bottom wall 50 which separates the recess 48 from the cavity 46 and the lateral walls 52, 54 formed by the outer radial ends of the intrados and extrados walls 30. Preferably, the fin 34 is offset radially inward with respect to the bottom wall 50 of the recess. This further protects the fin 34 against contact wear with the stationary casing of the turbine since it is firstly the side walls of the recess 48 which will wear by contact before the fin 34. radial offset of the vane 34 relative to the bottom wall 50 allows the (x) drilling (s) 42 to open into the cavity 46, that is to say radially inward relative to the wall of bottom 50. According to a first embodiment shown in Figure 3, the fin 34 extends in a plane perpendicular to the radial main axis of the blade 10, that is to say in a lateral longitudinal plane . According to alternative embodiments shown in FIGS. 4 and 5, the fin 34 is inclined with respect to the radial main axis of the blade 10. According to the variant shown in FIG. 4, the fin 34 is inclined radially towards the radial axis. in the variant shown in Figure 5, the fin 34 is inclined radially inwardly. As can be seen in Figure 4, the orientation of each bore 42 is similar or identical to the inclination of the fin 34. This allows in particular to keep enough material around the holes 42, thus strengthening the fin 34 This embodiment can also be applied when the fin 34 extends in a plane perpendicular to the radial main axis of the blade 10 or when the fin 34 is inclined radially outwards or towards the inside. According to another embodiment shown for example in Figures 3 and 5, the orientation of each bore 42 differs from the inclination of the fin 34. For example, as shown in Figure 3, each bore 42 is inclined radially to the outside, to allow a connection of the bore 42 with the cavity 46 when the fin 34 is offset radially outwardly relative to the bottom wall 50. In another example, as shown in Figure 5, each drilling 42 is perpendicular to the radial main axis of the blade 10, regardless of the orientation of the fin 34. According to yet another aspect of the fin 34, the connection 56 between the intrados wall 28 and the face radially internal 58 of the fin 34 is shaped to prevent the so-called phenomenon of "corner vortex" which consists of the formation of a vortex located on the connection 56.
    [0012] This phenomenon is controlled either by adapting the radius of curvature of this connection 56, or, as can be seen in FIG. 6, by forming a convex connection which is curved laterally and radially inwards. The fin 34 represents a certain mass which is rotatable at a distance from the axis of rotation of the rotor.
    [0013] To limit the impacts of the blade on the blade, by the efforts it generates, including centrifugal forces, the general mass of the blade 34 is defined so that it is at least allowable to perform its function. For this purpose, the upstream longitudinal end 38 of the fin 34 is offset longitudinally downstream with respect to the leading edge 24 of the blade 14 and the downstream longitudinal end 40 of the fin 34 is shifted longitudinally towards upstream with respect to the trailing edge 26 of the blade 14. Consequently, the longitudinal dimension of the fin is less than the longitudinal dimension of the blade 14. Also, the optimization of the mass of the fin 34 is performed by defining a radial thickness of the fin to achieve the holes 42, and the lateral dimension of the fin allowing a blockage of gas currents that could generate vortices at the top of the blade. The number of holes 42 also affects the mass of the fin 34. Thus, by multiplying the number of holes 42, it is possible to reduce its mass, while improving the cooling of the fin 34. The invention also relates to a rotor of high pressure turbine on which vanes 10 as described above are mounted, this rotor comprises a disk on the periphery of which grooves distributed over its entire periphery are formed to receive the feet 12 of the blades 10. The rotor thus carries a plurality of blades 10 according to the invention, which are evenly distributed around the main axis of rotation of the rotor.
    权利要求:
    Claims (10)
    [0001]
    REVENDICATIONS1. A turbomachine blade (10) comprising a foot (12) and a blade (14) having a median main plane (14) of longitudinal and radial main direction, which is carried by the foot, the blade (14) having an edge etching (24) located at an upstream longitudinal end, a trailing edge (26) located at a longitudinal end downstream of the leading edge (24) with respect to a flow of gas circulating around the blade (14), an intrados wall (28) and an extrados wall (30) laterally spaced from each other and each connecting the leading edge (24) to the trailing edge (26), and a top (32) located at the free outer radial end of the blade (14), the blade (14) further comprising a fin (34) longitudinal principal orientation (14) which is carried by the underside face (28). ), which is located at the apex (32) of the blade (14), characterized in that the fin (34) is offset radially inwards with respect to the apex (32) of the blade (14).
    [0002]
    A turbomachine blade (10) according to claim 1, wherein the top (32) of the blade (14) has a hollow recess (48) which is open radially outward and has a bottom wall (50). ) located radially away from the top (32) of the blade (14), characterized in that the fin (34) is offset radially inwardly relative to the bottom wall (50) of the recess (48).
    [0003]
    3. blade (10) turbomachine according to any one of the preceding claims, characterized in that the fin (34) has an upstream longitudinal end (38) which is offset longitudinally downstream relative to the leading edge (24) of the blade (14).
    [0004]
    4. blade (10) according to any one of the preceding claims, turbomachine characterized in that the fin (34) has a downstream longitudinal end (40) which is offset longitudinally upstream relative to the trailing edge ( 26) of the blade (14).
    [0005]
    5. A turbomachine blade (10) according to any one of the preceding claims, characterized in that the blade (14) comprises at least one bore (42) which opens into a lateral end face (44) of the fin (34) and which opens into a cavity (46) formed in the blade (14).
    [0006]
    6. A turbomachine blade (10) according to any one of the preceding claims, characterized in that a median longitudinal plane of the fin (34) is perpendicular to the main plane of the blade (14). 15
    [0007]
    7. A turbomachine blade (10) according to any one of claims 1 to 5, characterized in that a median longitudinal plane of the fin (34) is inclined relative to the main plane of the blade (14).
    [0008]
    8. Turbomachine blade (10) according to any one of the preceding claims, characterized in that the connection (56) between the lower face (28) and the radially inner face (58) of the fin (34) ) is convex in shape.
    [0009]
    9. A turbomachine blade (10) according to any one of the preceding claims, characterized in that the upstream longitudinal end (38) of the fin (34) is offset radially with respect to the downstream longitudinal end (40). ) of the fin (34).
    [0010]
    10. A turbomachine high pressure turbine rotor having a plurality of vanes according to any one of the preceding claims, which are evenly distributed about the main axis of rotation of the rotor. 10
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    同族专利:
    公开号 | 公开日
    FR3022295B1|2019-07-05|
    US20150361808A1|2015-12-17|
    US10260361B2|2019-04-16|
    引用文献:
    公开号 | 申请日 | 公开日 | 申请人 | 专利标题
    US1955929A|1932-03-18|1934-04-24|Voith Gmbh J M|Impeller|
    DE19913269A1|1999-03-24|2000-09-28|Asea Brown Boveri|Turbine blade|
    EP1312754A2|2001-11-16|2003-05-21|FIATAVIO S.p.A.|Bladed member, in particular for an axial turbine of an aircraft engine|
    US20100135813A1|2008-11-28|2010-06-03|Remo Marini|Turbine blade for a gas turbine engine|
    WO2013072610A1|2011-11-17|2013-05-23|Snecma|Gas turbine vane offset towards the lower surface of the head sections and with cooling channels|CN110662884A|2017-05-30|2020-01-07|西门子公司|Turbine blade with recessed tip and dense oxide dispersion strengthened layer|US3706512A|1970-11-16|1972-12-19|United Aircraft Canada|Compressor blades|
    US6905309B2|2003-08-28|2005-06-14|General Electric Company|Methods and apparatus for reducing vibrations induced to compressor airfoils|
    FR2885645A1|2005-05-13|2006-11-17|Snecma Moteurs Sa|Hollow rotor blade for high pressure turbine, has pressure side wall presenting projecting end portion with tip that lies in outside face of end wall such that cooling channels open out into pressure side wall in front of cavity|
    DE102009036406A1|2009-08-06|2011-02-10|Mtu Aero Engines Gmbh|airfoil|
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    GB201015006D0|2010-09-09|2010-10-20|Rolls Royce Plc|Fan blade with winglet|
    US9188017B2|2012-12-18|2015-11-17|United Technologies Corporation|Airfoil assembly with paired endwall contouring|CN107407290B|2015-04-08|2019-07-26|雷顿股份公司|Fan blade and correlation technique|
    DE102017216620A1|2017-09-20|2019-03-21|MTU Aero Engines AG|Shovel for a turbomachine|
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    FR3081497B1|2018-05-23|2020-12-25|Safran Aircraft Engines|GROSS FOUNDRY BLADE WITH MODIFIED LEAKING EDGE GEOMETRY|
    法律状态:
    2015-06-18| PLFP| Fee payment|Year of fee payment: 2 |
    2015-12-18| PLSC| Search report ready|Effective date: 20151218 |
    2016-06-06| PLFP| Fee payment|Year of fee payment: 3 |
    2017-04-27| PLFP| Fee payment|Year of fee payment: 4 |
    2018-02-09| CD| Change of name or company name|Owner name: SAFRAN AIRCRAFT ENGINES, FR Effective date: 20170717 |
    2018-06-05| PLFP| Fee payment|Year of fee payment: 5 |
    2020-05-20| PLFP| Fee payment|Year of fee payment: 7 |
    2021-05-19| PLFP| Fee payment|Year of fee payment: 8 |
    优先权:
    申请号 | 申请日 | 专利标题
    FR1455527|2014-06-17|
    FR1455527A|FR3022295B1|2014-06-17|2014-06-17|TURBOMACHINE DAWN COMPRISING AN ANTIWINDER FIN|FR1455527A| FR3022295B1|2014-06-17|2014-06-17|TURBOMACHINE DAWN COMPRISING AN ANTIWINDER FIN|
    US14/740,405| US10260361B2|2014-06-17|2015-06-16|Turbomachine vane including an antivortex fin|
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