Stator part having a fin, in a turbine engine

Abstract

The invention relates to a stator part (20) of a turbine engine, comprising a platform (22), a blade (24, 26) extending radially relative to a central axis (A), and a fin (28) extending radially from a fin root (44) to a fin tip (46), the fin comprising a lower side (48) and an upper side (50), each point (100) of the lower side or of the upper side defining a radial axis (Ar) passing through the point, each plane (Pr) that includes the radial axis defining a section (S) of the lower side or of the upper side, an angle defined in the plane between the root profile and a tangent to the section at an intersection (104) of the section and of the root profile being less than or equal to 45 degrees, the section being located between the root profile and the tangent.

Description

[0001] Stator part with fin in a turbomachine

[0002] FIELD OF THE INVENTION

[0003] The invention relates to the stator parts of a turbomachine comprising a blade such as the flow straighteners located downstream of a compressor and in particular the fixed-pitch straighteners.

[0004] STATE OF THE ART

[0005] In an aircraft turbomachine, and in particular aircraft intended for passenger transport, it is the air propelled by a fan and combustion gases exiting the turbomachine through an exhaust nozzle that exerts a reaction thrust on the turbomachine and, through it, on the aircraft. The circulation of gases through the turbomachine is influenced by rotating blades and fixed blades. The fixed or stator blades include in particular outlet guide vanes (also known as "Outlet Guide Vane" or "OGV"), inlet guide vanes (also known as "Inlet Guide Vane" or "IGV"), and variable stator vanes (also known as "Variable Stator Vane" or "VSV"). The stator vanes of a gas turbine aircraft engine may each have two platforms (inner and outer) which are attached to the blades.There are also unducted designs with stator blades that have only one inner platform. In all cases, these stator blades form rows of fixed vanes that guide the gas flow through the engine at an appropriate speed and angle.

[0006] Within a flow straightener comprising a plurality of fixed blades, the flow of gases generally takes place between the blades in an upstream-downstream direction. It is known, however, that the blade root area can be the site of secondary aerodynamic flows.

[0007] For each pair of blades facing each other, a pressure gradient between the pressure face (intrados) of the first blade and the depression face (extrados) of the second blade generates a passage flow (also known as "crossflow" in English) which transports the gases towards the extrados.

[0008] At the blade tip, i.e. at the junction between the blade and the hub or between the blade and the casing, a corner separation (also known as "corner separation" in English) and a vortex (also known as "corner vortex" in English) can occur. This separation generates pressure losses as well as aerodynamic blocking. The latter is problematic in terms of operability. For high incidences of the flow arriving at the rectifier, i.e. when the direction of gas flow upstream of the rectifier makes a significant angle with a direction of the leading edge of the blade, this corner separation is amplified to the point of causing a detachment of the boundary layer on the blade which can no longer ensure the deflection of the flow.

[0009] There is therefore a need for a new geometry to correct these problems and improve performance in terms of equipment efficiency, particularly with high incidence of flow entering the rectifier.

[0010] STATEMENT OF THE INVENTION

[0011] An aim of the invention is to propose a stator part of a turbomachine whose geometry improves the flow of fluids compared to the prior art.

[0012] The aim is achieved within the framework of the present invention thanks to a stator part of a turbomachine comprising:

[0013] - a platform defining a wall of a gas flow vein,

[0014] - a blade extending radially relative to a central axis of the turbomachine from the platform, and

[0015] - a fin extending in the vein radially from a fin root located on the platform to a fin head, the fin comprising a lower surface and an upper surface located radially between the fin root and the fin head, each point of the lower surface or respectively of the upper surface defining a radial axis passing through the point, each plane comprising the radial axis defining a section of the lower surface or respectively of the upper surface, an angle defined in the plane between the root profile and a tangent to the section at an intersection of the section and the root profile being less than or equal to 45 degrees, the section being located between the root profile and the tangent.

[0016] Such a stator part is advantageously and optionally completed by the following different characteristics taken alone or in combination:

[0017] - the platform is a first platform, the part comprising a second platform so as to define the flow vein between the first platform and the second platform, the vein extending radially over a vein height, the fin extending radially over a fin height, a ratio of the fin height to the vein height being greater than or equal to 0.01 and less than or equal to 0.25;

[0018] - the blade comprises a leading edge and a trailing edge separated by a blade chord, the winglet comprising a plurality of profiles stacked radially between the winglet root and the winglet tip, each profile defining a chord between the leading edge of the winglet and the trailing edge of the winglet, the chord of the tip profile being less than the chord of the root profile; a ratio of the chord of the tip profile to the blade chord being greater than or equal to 0.1 and less than or equal to 0.6; a ratio of the chord of the root profile to the blade chord being greater than or equal to 0.3 and less than or equal to 1.1;

[0019] - each fin profile defines a maximum thickness of the profile between the intrados and the extrados in a direction perpendicular to a chord line, a ratio of the maximum thickness of the root profile to the chord of the root profile being greater than or equal to 0.05 and less than or equal to 0.25; and a ratio of the maximum thickness of the head profile to the chord of the head profile being greater than or equal to 0.05 and less than or equal to 0.25;

[0020] - the blade comprises a lower surface facing the upper surface of the fin, the fin comprising a trailing edge, the trailing edge comprising a vanishing point located on the platform, a tangent to the trailing edge at the vanishing point extending from the platform away from the axis between the blade and a radial vanishing plane passing through the axis and the vanishing point;

[0021] - the trailing edge includes a point of asymmetry such that for any running point of the trailing edge between the point of asymmetry and the trailing point, a tangent to the trailing edge at the running point extends from the trailing edge away from the axis between the blade and a radial trailing plane passing through the axis and the running point;

[0022] The invention also relates to a turbomachine comprising a stator part such as has just been presented and to an aircraft comprising such a turbomachine.

[0023] DESCRIPTION OF FIGURES

[0024] Other characteristics and advantages of the invention will emerge from the following description, which is purely illustrative and non-limiting, and must be read in conjunction with the appended drawings in which:

[0025] - figure 1 is a schematic representation of a turbomachine;

[0026] - figure 2 is a schematic representation of a stator part according to a first embodiment;

[0027] - figure 3 is a schematic sectional view in a plane perpendicular to a radial axis of the turbomachine of a stator part according to the first embodiment;

[0028] - Figure 4 is a schematic sectional view in a plane perpendicular to the axis of the turbomachine of a stator part according to the first embodiment. DETAILED DESCRIPTION OF THE INVENTION

[0029] Turbomachine

[0030] With reference to Figure 1, a turbomachine is shown schematically, more specifically a double-flow axial turbojet 1. The turbojet 1 illustrated extends along an axis A and successively comprises, in the direction of flow of the gases in the turbomachine, a fan 2, a compression section which may comprise a low-pressure compressor 3 and a high-pressure compressor 4, a combustion chamber 5, and a turbine section which may comprise a high-pressure turbine 6, a low-pressure turbine 7 and an exhaust nozzle.

[0031] The blower 2 and the low pressure compressor 3 are driven in rotation by the low pressure turbine 7 via a first transmission shaft 9, while the high pressure compressor 4 is driven in rotation by the high pressure turbine 6 via a second transmission shaft 10.

[0032] In operation, a flow of air compressed by the low and high pressure compressors 3 and 4 feeds combustion in the combustion chamber 5, the expansion of the combustion gases of which drives the high and low pressure turbines 6, 7. The air propelled by the fan 2 and the combustion gases exiting the turbojet engine 1 through an exhaust nozzle downstream of the turbines 6, 7 exert a reaction thrust on the turbojet engine 1 and, through it, on a vehicle or machine such as an aircraft (not shown).

[0033] Stator part

[0034] Downstream of the fan or a compression stage, the turbomachine may comprise a stage of straightening vanes. Such a stage of straightening vanes may comprise a stator part 20 as shown with reference to FIG. 2.

[0035] The stator part 20, or the set 20 of stator parts if it is not a single piece, has at least one blade 24, 26 and a platform 22 from which the blade 24, 26 extends. The stator part may for example comprise two adjacent blades 24, 26 which extend from the platform 22.

[0036] The term "platform" here designates any element of the turbomachine from which blades 24, 26 are capable of being mounted. The platform may in particular be a hub or a casing which surrounds the axis of the turbomachine. The platform may have a cylindrical surface at a constant radial distance from the axis A of the turbomachine. The platform 22 has an inner wall or an outer wall against which the air circulates, that is to say that the platform 22 defines a wall of a gas flow vein. The blades 24, 26 extend from the platform 22 in the vein either radially outwards away from the axis of the turbomachine A or radially inwards approaching the axis of the turbomachine A.

[0037] Figure 2 is a schematic representation of the stator part 20 in perspective. The axis A of the turbomachine is shown therein oriented positively in the direction of the gas flows in the turbomachine. Figure 2 also shows a radial axis r and a circumferential axis 0 passing through a point 34 of the platform 22. At each point in space and for example at a point 34 of the platform 22, a radial axis r can be defined which is perpendicular to the axis A of the turbomachine and which passes through the point and the axis A of the turbomachine. The radial axis is oriented positively in the direction which moves away from the axis A of the turbomachine. It is also possible to define a circumferential axis 0 which passes through the point and which is perpendicular to the radial axis r and to the axis A of the turbomachine. The circumferential axis is oriented positively in the direction which moves away from the axis A of the turbomachine.

[0038] In the example of Figure 2, the blades 24, 26 extend radially from the platform 22 away from the axis of the turbomachine, but the invention is not limited to this situation alone.

[0039] Figure 3 is a schematic representation of the stator part 20 in a circumferential plane passing through the platform 22, a circumferential plane which is at a constant distance from the axis A of the turbomachine. Such a circumferential plane parallel to the axis A of the turbomachine makes it possible to define a section of the blades 24, 26.

[0040] The direction of axis A is given in Figure 3 by the x axis whose orientation is the direction of gas flow. The radial axis r is perpendicular to the plane of Figure 3 and directed towards the reader of Figure 3. The 0 axis corresponds to the circumferential direction perpendicular simultaneously to axis A and the radial axis.

[0041] The blades 24 and 26 each have a lower surface 624, 126 and an upper surface 124, 626.

[0042] The blades 24 and 26 each comprise a leading edge 224, 226 on the upstream side and a trailing edge 324, 326 on the downstream side. The terms upstream and downstream are defined in relation to the general flow of gases through the turbomachine which takes place from upstream to downstream in the direction and sense of the axis A of the turbomachine.

[0043] The blades define a blade chord 424 which is the length of the segment connecting the leading edge and the trailing edge in a circumferential plane with a constant radius or a constant distance from the axis A, a circumferential plane which can be called a cutting plane. Similarly in a circumferential cutting plane, each blade has a camber line 43, 41 which is the curve equal to the average between the curve of the extrados and the curve of the intrados. More precisely, the camber line is formed of all the points located at an equal distance from the extrados and the intrados. The distance from a particular point to the extrados (or the intrados) is defined here as the minimum distance between the particular point and a point on the extrados (or the intrados).

[0044] Fin

[0045] The stator part 20 also comprises a fin 28 which extends from the platform 22 in the same direction and the same sense of extension as the blade(s) 24, 26. The fin 28 extends in the vein radially relative to the axis A of the turbomachine from the platform 22.

[0046] The fin 28 comprises an extrados 50 which is opposite the intrados 126 of the blade 26.

[0047] When the part comprises two blades 24, 26, the fin 28 is located between the blades 24 and 26. More precisely, the fin 28 is located opposite the extrados 124 of the first blade 24 and the intrados 126 of the second blade 26.

[0048] The fin 28 comprises a lower surface 48 which is opposite the upper surface 124 of the first blade and an upper surface 50 which is opposite the lower surface 126 of the second blade 26.

[0049] The fin 28 comprises a leading edge 30 and a trailing edge 32, the leading edge 30 being located upstream of the trailing edge 32.

[0050] The leading edge 30 comprises a point of attack 34 located on the platform 22. The point of attack 34 corresponds to the intersection of the leading edge 30 and the platform 22. A radial plane of attack Pa is defined which passes through the axis A of the turbomachine and the point of attack 34.

[0051] Any radial plane includes the A axis of the turbomachine.

[0052] The trailing edge 32 comprises a vanishing point 36 located on the platform 22. The vanishing point 36 corresponds to the intersection of the trailing edge 32 and the platform 22. A radial vanishing plane Pf is defined which passes through the axis A of the turbomachine and the vanishing point 36.

[0053] Figure 4 is a schematic representation in a Pr plane of some parameters of the fin profile.

[0054] The most general embodiment of the invention corresponds to the following two characteristics in relation to figures 2 and 4.

[0055] Each point 100 of the intrados 48 of the fin 28 or respectively of the extrados 50 of the fin 28 defines a radial axis Ar passing through the point 100. The radial axis is orthogonal to the central axis A of the turbomachine and passes through the central axis A of the turbomachine.

[0056] Each plane Pr comprising the radial axis Ar defines a section S of the intrados 48 or respectively of the extrados 50.

[0057] The plane Pr is defined by the direction of the radial axis Ar and any other direction of the plane such as the circumferential direction, the direction of the central axis A of the turbomachine, or another direction. The plane Pr can be a radial plane or not.

[0058] By defining this plane Pr, a cutting plane of the intrados 48 is defined if the point 100 belongs to the intrados 48, or a cutting plane of the extrados 50 if the point 100 belongs to the extrados 50. The cutting plane then defines a section of the intrados 48 or of the extrados 50. The section S passes through the point 100 and through a point at the intersection 104 of the root profile 44 and the intrados 48 or of the extrados 50. This point is also at the intersection 104 of the section S and the root profile 44.

[0059] With reference to Figure 4, an angle 106 is defined in the plane Pr between the foot profile 44 and a tangent T to the section S, the tangent being constructed at the point at the intersection 104.

[0060] This angle 106 is less than or equal to 45 degrees.

[0061] In the plane Pr, the section S is located between the foot profile 44 and the tangent T, that is to say that each point of the section S defines a radial axis and on this radial axis this point is located between a point of the foot profile 44 and a point of the tangent T.

[0062] Any section S as defined above is located between the platform 22 and a tangent T as presented above. As the tangent is relatively close to the platform, in relation to the value of the angle 106, the fin has a compact and pyramidal shape. This shape makes it possible to act on:

[0063] - a first part of the flow passage located upstream of the fin by guiding this first part without detachment along the intrados of the fin, and on

[0064] - a second part of the passage flow which is closest to the blade whose intrados is opposite the fin, confining this second part between the intrados of the blade and the extrados of the fin.

[0065] Corner detachment is then greatly reduced.

[0066] Furthermore, the upstream part of the fin has a relatively gentle slope compared to platform 22. This reduces the risk of aerodynamic blockage even when this upstream part is located in the region of the smallest section of the flow path most subject to blockage. This is particularly advantageous at high Mach where excessive solid blockage can induce shock waves. The slow evolution of the height prevents the first part of the flow from separating: the gases follow the lower surface of the fin and are thus guided to the trailing edge of the fin. Guiding this first part of the flow passage can locally cancel the transverse pressure gradient, which limits the stalling of the stator blades.

[0067] It should be noted that if the point 100 is on the leading edge 30 or on the trailing edge 32 of the fin 28, the point 100 is part of both the intrados 48 and the extrados 50. The characteristics presented above are then verified for the intrados and for the extrados.

[0068] Figure 4 represents the situation where section S includes a point common to the head profile 46, but this is not necessarily the case.

[0069] Fin height

[0070] The platform 22 as described so far defines a radial inner or respectively radial outer wall of the gas flow vein.

[0071] When the stator part 20 corresponds to a streamlined architecture, it comprises a second platform located radially opposite the first platform 22, this second platform defining the outer radial wall or respectively the inner radial wall of the gas flow vein. The gas flow vein therefore passes radially between the first platform 22 and the second platform, the vein extending radially over a certain vein height designated by the reference Hv in FIG. 4.

[0072] When the stator part 20 corresponds to an unducted architecture, it comprises only a single platform 22 defining the inner radial wall of the gas flow vein. The vein extends radially over a certain vein height defined by the height of the blades of the stator part 20, blades which project radially from the platform 20 towards the outside.

[0073] The fin 28 extends radially over a fin height Ha indicated in Figure 4.

[0074] According to a first optional variant of the most general embodiment, a ratio of the fin height to the vein height being greater than or equal to 0.01 and less than or equal to 0.25.

[0075] It may be advantageous for the fin described above to remain relatively small in height so as not to block the flow in the vein.

[0076] Circumferential distance fin - first blade The first blade 24 and the second blade 26 are separated in a circumferential direction by a pitch 42. The pitch 42 which separates the blades is an angle separating a radial direction of the first blade 24 and a radial direction of the second blade 26. The pitch is fixed by the total number of blades playing the same role and having the same axial position as the blades 24 and 26 and which are located all around the axis A of the turbomachine. The distance separating the first blade 24 and the second blade is given by this angle and the radius at the axis A at which this distance is to be evaluated.

[0077] It is also possible to define an angle separating the first blade 24 and the fin 28 in the circumferential direction, i.e. an angle separating a radial direction of the first blade 24 and a radial direction of the fin 28.

[0078] The angular separation of the first blade 24 and the fin 28 in the circumferential direction can be freely chosen to be less than or equal to the angular pitch 42. In other words, the fin can be located between the first blade and the second blade at any distance from the first blade.

[0079] Fin profiles

[0080] As seen previously, the first blade 24 defines a blade chord 424 between its leading edge 224 and its trailing edge 324.

[0081] The fin 28 can be modeled or represented as a stack of profiles in a radial direction between a fin root 44 and a fin head 46. The fin root 44 is located on the platform 22 and corresponds to the intersection of the fin 28 and the platform 22. The fin head is located at a distance from the platform 22 in the gas flow path. Each fin profile extends in a circumferential plane parallel to the axis A of the turbomachine, like a section of the fin made in this circumferential plane at constant radius or constant distance from the axis A, a circumferential plane which can be described as a section plane.

[0082] Referring to Figure 3, each fin profile defines a fin chord 54 between the leading edge 30 of the fin and the trailing edge 32 of the fin. More specifically, the fin chord 54 is defined between, on the one hand, a first point at the intersection of the leading edge 30 and the cutting plane and, on the other hand, a second point at the intersection of the trailing edge 32 and the cutting plane. The fin chord designates the length of the segment connecting the first point and the second point. The chord line designates the segment connecting the first point and the second point.

[0083] Each fin profile also defines a maximum thickness 52 between the lower surface 48 of the fin and the upper surface 50 of the fin in a direction perpendicular to the chord line. Each fin profile makes it possible to define a camber line which is the curve equal to the average between the curve of the upper surface 50 of the fin and the curve of the lower surface 48 of the fin. More precisely, in a given profile of the fin the camber line is formed of all the points located at an equal distance between on the one hand the intersection of the upper surface 50 and the cutting plane and on the other hand the intersection of the lower surface 48 and the cutting plane. The camber line of the fin can be chosen close to the camber line of the first blade 24.

[0084] The fin exhibits a continuous variation of profiles in the radial direction, i.e. all parameters of camber, chord, thickness vary continuously from the root profile to the head profile.

[0085] A second optional variation of the most general embodiment and / or its variant includes the following three features:

[0086] - the chord of the head profile is less than the chord of the foot profile;

[0087] - a ratio of the head profile chord to the blade chord is greater than or equal to 0.1 and less than or equal to 0.6;

[0088] - a ratio of the root profile chord to the blade chord is greater than or equal to 0.3 and less than or equal to 1.1.

[0089] The root profile chord is then relatively large, which allows the flow to be blocked over a large part of the blade chord length.

[0090] A third optional variation of the most general embodiment and / or its variations includes the following two features:

[0091] - a ratio of the maximum thickness of the foot profile to the chord of the foot profile is greater than or equal to 0.05 and less than or equal to 0.25; and

[0092] - a ratio of the maximum thickness of the head profile to the chord of the head profile being greater than or equal to 0.05 and less than or equal to 0.25.

[0093] The fin is then relatively thick, which facilitates its manufacture and mechanical resistance, particularly in the event of ingestion of particles or foreign bodies by the engine.

[0094] “Metal” angles of the fin and the first blade

[0095] For the first blade 24 and the second blade 26, we can define:

[0096] - a blade leading metal angle as the angle between the tangent to the camber line at the leading edge of the blade and the axis A of the turbomachine, the angle being oriented from the axis A towards the tangent, and - a blade trailing metal angle as the angle between the tangent to the camber line at the trailing edge of the blade and the axis A, the angle being oriented from the axis A towards the tangent.

[0097] For each fin profile we can define:

[0098] - a fin leading metal angle as the angle between the tangent 61 to the camber line at the leading edge of the fin and the axis A of the turbomachine, the angle being oriented from the axis A towards the tangent 61, and

[0099] - a fin trailing metal angle as the angle between the tangent 63 to the camber line at the trailing edge of the fin and the axis A, the angle being oriented from the axis A towards the tangent 63.

[0100] In the particular case of the fins described in this invention, it is possible to choose to estimate the metal angles as a function of the local camber angle of the blade 24 or 26.

[0101] For this, it is possible to associate with each point of each camber line of each fin profile an axial position and a radial position. The fin point is associated with a reference point of a camber line 43 of a profile of the first blade 24, the reference point having the axial position and the radial position of the fin point.

[0102] The blade attack metal angle can be associated with the local camber angle which is the angle oriented from the axis A to the tangent to the camber line of the profile of the first blade 24 at the associated reference point. The blade attack metal angle can be chosen to be described with reference to this local camber angle as the difference between this metal angle and this local camber angle.

[0103] The blade trailing metal angle can be associated with the local camber angle which is the angle oriented from the axis A to the tangent to the camber line of the profile of the first blade 24 at the associated reference point. The blade trailing metal angle can be chosen to be described with reference to this local camber angle as the difference between this metal angle and this local camber angle.

[0104] Generally speaking, the fins described in this invention have at the leading edge and trailing edge of the head profile and the root profile metal angles close to the local camber angle. More precisely:

[0105] - the angle formed by a first tangent to the camber line of the fin head profile at the leading edge 30 and by a second tangent to the camber line of the profile of the first blade at the reference point associated with the leading edge of the fin head profile is less than or equal in absolute value to 10 degrees;

[0106] - the angle formed by a third tangent to the camber line of the fin head profile at the trailing edge 32 and a fourth tangent to the camber line of the profile of the first blade at the reference point associated with the trailing edge of the fin head profile is less than or equal in absolute value to 10 degrees;

[0107] - the angle formed by a fifth tangent to the camber line of the winglet root profile at the point of attack 34 and a sixth tangent to the camber line of the profile of the first blade at the reference point associated with the point of attack 34 is less than or equal in absolute value to 10 degrees; and

[0108] - the angle formed by a seventh tangent to the camber line of the winglet root profile at the vanishing point 36 and an eighth tangent to the camber line of the profile of the first blade at the reference point associated with the vanishing point 36 is less than or equal in absolute value to 10 degrees.

[0109] The head profile and the root profile then have a metallic angle at the trailing edge whose value is close to the local camber of the blades. In this way the fin guides the flow at the trailing edge like the blades.

[0110] Axial position of the fin relative to the first blade

[0111] To characterize the position of the fin in the direction of the A axis of the turbomachine, an axial coordinate of a point at the mid-chord of a fin profile is used.

[0112] The mid-chord point is located equidistant from the leading edge 30 of the fin and the trailing edge 32 of the fin.

[0113] The axial coordinate of the mid-chord point may be chosen to be greater than or equal to an axial position of the leading edge 224 of the first blade and less than or equal to a sum of the axial position of the leading edge of the first blade and the blade chord 424.

[0114] In particular, the mid-chord point of the fin head profile can be used as the mid-chord point.

[0115] The fin root profile is positioned axially relative to the head profile thanks to the fin stacking law which gives the relative positioning of the fin head profile relative to the fin root profile.

[0116] For this relative positioning, the mid-chord point is chosen as the positioning reference for each profile of the stack. Two axes are defined:

[0117] - a first axis “t” directed along the direction of the chord of the root profile and oriented in the same direction as the axis A of the turbomachine from upstream to downstream; - a second axis “n” perpendicular to the first axis “t” and oriented in the same direction as the circumferential axis 0 of the first blade 24 towards the second blade 26.

[0118] Two angles are defined:

[0119] - a sweep angle (also known as "sweep") which is the angle between the "t" axis and a direction defined by the mid-chord point of the root profile and the mid-chord point of the head profile;

[0120] - a dihedral angle (also known as "lean") which is the angle between the "n" axis and the direction defined by the mid-chord point of the root profile and the mid-chord point of the head profile.

[0121] In the particular case of the fins described here, the sweep and dihedral angles as just defined take values ​​greater than or equal to - 10 degrees and less than or equal to + 10 degrees. These angles are sufficiently small so that the axial position of the fin can be considered well described by an axial coordinate 78 of a mid-chord point 76 of a fin profile.

[0122] Positive dihedral effect at the trailing edge

[0123] A second embodiment depending on the most general embodiment or its first variant comprises the following feature.

[0124] The fin 28 has a tangent to the trailing edge 32 at the vanishing point 36 which extends in the flow path from the platform 22 between the second blade 26 and the radial vanishing plane Pf. In other words, in a radial plane orthogonal to the axis A of the turbomachine and passing through the vanishing point, the tangent to the trailing edge, moving away from the platform 22 on the flow path side of the platform, approaches the second blade 26. Formulated in a different way again, the trailing edge 32 is inclined at the vanishing point 36 towards the second blade 26. It should be noted that, moreover, this tangent to the trailing edge may have a non-zero projection along the axis A of the turbomachine.

[0125] The inclination towards the second blade on the trailing edge side allows a dihedral effect at the trailing edge which more effectively blocks the passage flow. This makes it possible to prevent the passage flow coming from the second blade 26 from going up the extrados of the fin 28 and passing beyond the fin to reach the intrados of the first blade 24. The fin is therefore “lying down” on the trailing edge side so as to have a wall inclined towards the intrados of the second blade 26. This is the positive dihedral effect. The passage flow is then strongly pushed back towards the intrados of the second blade 26. It should be noted that the fin may have a trailing edge 32 which is straight and therefore coincides with the tangent to the trailing edge 32 at the vanishing point 36.

[0126] More generally, the fin 28 may comprise a point of asymmetry and the following characteristic: at any current point of the trailing edge located between this point of asymmetry and the trailing point 36, the tangent to the trailing edge which extends in the vein from the platform 22 is located between the second blade 26 and a radial plane passing the current point. In other words, in a radial plane passing through the current point, the trailing edge, moving away from the current point towards the point of asymmetry, approaches the second blade 26. Formulated in a different manner again, the trailing edge 32 is inclined between the trailing point 36 and the point of asymmetry towards the second blade 26.

[0127] The trailing edge upstream of the asymmetry point can be symmetrical, that is to say that each tangent to the trailing edge is contained in a radial plane which includes the axis of the turbomachine.

[0128] The leading edge can also be symmetrical, that is, each tangent to the leading edge is contained in a radial plane that includes the axis of the turbomachine.

[0129] In this situation, the fin can be described as symmetrical upstream of the point of asymmetry and asymmetrical downstream.

[0130] When the point of asymmetry is located on the head profile, this point of asymmetry can be a first point of asymmetry and the winglet can include a second point of asymmetry this time included on the leading edge.

[0131] At any current point of the leading edge located between this point of asymmetry and the tip profile, the tangent to the leading edge which extends in the vein from the leading edge is located between the second blade 26 and the radial plane passing through the current point. In other words, in a radial plane passing through the current point, the trailing edge, moving away from the current point towards the tip profile, approaches the second blade 26. Formulated in a different way again, the leading edge 30 is inclined between the second point of asymmetry and the tip profile towards the second blade 26. The fin can then be described as symmetrical upstream of the second point of asymmetry and asymmetrical downstream. Optionally, the second point of asymmetry can be the leading point 34 at which the entire fin can be described as asymmetrical.

Claims

DEMANDS 1. Stator component (20) of a turbomachine comprising: - a platform (22) defining a wall of a gas flow vein, - a blade (24, 26) extending radially from a central axis (A) of the turbomachine from the platform (22), and - a fin (28) extending radially in the vein from a fin root (44) located on the platform (22) to a fin tip (46), the fin comprising an intrados (48) and an extrados (50) situated radially between the fin root (44) and the fin tip (46), each point (100) of the intrados (48) or respectively of the extrados (46) defining a radial axis (Ar) passing through the point (100), each plane (Pr) comprising the radial axis (Ar) defining a section (S) of the intrados (48) or respectively of the extrados (46), an angle (106) defined in the plane (Pr) between the root profile (44) and a tangent (T) to the section (S) at an intersection (104) of the section (S) and the root profile (44) being less or equal to 45 degrees, the section (S) being located between the foot profile (44) and the tangent (T).

2. Stator piece according to claim 1, wherein the platform (22) is a first platform, the piece comprising a second platform so as to define the flow channel between the first platform (22) and the second platform, the channel extending radially over a channel height (Hv), the fin (28) extending radially over a fin height (Ha), a ratio of the fin height (Ha) to the channel height (Hv) being greater than or equal to 0.01 and less than or equal to 0.

25.

3. Stator component according to any one of the preceding claims, wherein the blade (24,26) comprises a leading edge (224, 226) and a trailing edge (324, 326) separated by a blade chord (424, 426), the vane (28) comprising a plurality of profiles stacked radially between the vane root (44) and the vane tip (46), each profile defining a chord (54) between the leading edge (30) of the vane (28) and the trailing edge (32) of the vane (28), the chord of the tip profile being less than the chord of the root profile; a ratio of the chord of the root profile to the blade chord being greater than or equal to 0.1 and less than or equal to 0.6; a ratio of the chord of the root profile to the blade chord being greater than or equal to 0.3 and less than or equal to 1.

1.

4. Stator component according to claim 3, wherein each fin profile defines a maximum thickness (52) of the profile between the intrados (48) and the extrados (50) in a direction perpendicular to a chord line, a ratio of the maximum thickness of the foot profile to the chord of the foot profile being greater than or equal to 0.05 and less than or equal to 0.25; and a ratio of the maximum thickness of the head profile to the chord of the head profile being greater than or equal to 0.05 and less than or equal to 0.

25.

5. Stator part according to any one of the preceding claims, wherein the blade comprises an intrados (126) opposite the extrados (50) of the fin (28), the fin comprising a trailing edge (32), the trailing edge comprising a trailing point (36) located on the platform (22), a tangent to the trailing edge (32) at the trailing point (36) extending from the platform (22) away from the axis (A) between the blade (26) and a radial trailing plane (Pf) passing through the axis (A) and the trailing point (36).

6. Stator part according to claim 5, wherein the trailing edge (32) includes an asymmetry point such that for any current point of the trailing edge (32) located between the asymmetry point and the trailing point (36), a tangent to the trailing edge (32) at the current point extends from the trailing edge (32) away from the axis (A) between the blade (26) and a radial trailing plane passing through the axis (A) and the current point.

7. Turbomachine comprising a stator part (20) according to any one of claims 1 to 6.

8. Aircraft comprising a turbomachine according to claim 7.

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