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

FIELD OF INVENTION

[0001] The invention relates to the stator parts of a turbomachine comprising a blade such as flow straighteners located downstream of a compressor and in particular fixed pitch straighteners. STATE OF THE ART

[0002] In an aircraft turbomachine, particularly in passenger aircraft, it is the air propelled by a fan and the combustion gases exiting the turbomachine through an exhaust nozzle that exert reaction thrust on the turbomachine and, through it, on the aircraft. The flow of gases through the turbomachine is influenced by rotating and stationary blades. The stationary or stator blades include outlet guide vanes (OGVs), inlet guide vanes (IGVs), and variable stator vanes (VSVs). The stator blades of an aircraft gas turbine engine may each have two platforms (inner and outer) attached to the blade.There are also unfaired designs featuring stator blades with only a single inner platform. In all cases, these stator blades form fixed rows that guide the gas flow through the engine at the appropriate speed and angle. Document WO 2021 / 148751 A1 concerns a turbomachine component comprising blades and a platform with a non-axisymmetric surface.

[0003] Within a flow straightener comprising a plurality of fixed blades, the gas flow generally occurs between the blades in an upstream-downstream direction. However, it is known that the blade root area can be the site of secondary aerodynamic flows.

[0004] For each pair of blades facing each other, a pressure gradient between the pressure side (intrados) of the first blade and the depression side (extrados) of the second blade generates a crossflow (also known as "crossflow" in English) which transports the gases towards the extrados.

[0005] At the blade tip, that is, at the junction between the blade and the hub or between the blade and the casing, corner separation and corner vortexing can occur. This separation generates pressure losses and aerodynamic stall. The latter is problematic in terms of operability. For high angles of attack of the flow arriving at the stator, that is, when the gas flow direction upstream of the stator makes a significant angle with the direction of the blade's leading edge, this corner separation intensifies until it causes boundary layer separation on the blade, which can no longer deflect the flow.

[0006] Therefore, there is a need for a new geometry to correct these problems and improve performance in terms of equipment efficiency, particularly with a high impact of the flux entering the rectifier. DESCRIPTION OF THE INVENTION

[0007] One aim of the invention is to provide a stator part of a turbomachine whose geometry improves fluid flow compared to the prior art.

[0008] The objective is achieved within the framework of the present invention by means of a stator component of a turbomachine comprising: a platform defining a wall of a gas flow duct, a blade extending radially from the platform relative to a central axis of the turbomachine, and a blade extending radially in the duct from a blade root located on the platform to a blade tip, the blade comprising an intrados and an extrados situated radially between the blade root and the blade tip, each point of the intrados or respectively of the extrados defining a radial axis passing through the point, each plane including the radial axis defining a section of the intrados or respectively of the extrados, an angle defined in the plane between the foot profile and a tangent to the section at an intersection of the section and the foot profile being less than or equal to 45 degrees, the section being located between the foot profile and the tangent.

[0009] Such a stator component is advantageously and optionally complemented by the following various features, taken alone or in combination: the platform is a first platform, the part comprising a second platform so as to define the flow channel between the first platform and the second platform, the channel extending radially over a channel height, the fin extending radially over a fin height, a ratio of the fin height to the channel height being greater than or equal to 0.01 and less than or equal to 0.25; the blade comprises a leading edge and a trailing edge separated by a blade chord, the fin comprising a plurality of profiles stacked radially between the fin root and the fin tip, each profile defining a chord between the leading edge of the fin and the trailing edge of the fin, 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 root profile chord to the blade chord being greater than or equal to 0.3 and less than or equal to 1.1; each fin profile defines a maximum profile thickness between the lower and upper surfaces in a direction perpendicular to a chord line, a ratio of the maximum thickness of the root profile to the root profile chord 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 tip profile to the tip profile chord being greater than or equal to 0.05 and less than or equal to 0.25; the blade includes a lower surface opposite the upper surface of the fin, the fin including a trailing edge, the trailing edge including a trailing point located on the platform, a tangent to the trailing edge at the trailing point extending from the platform away from the axis between the blade and a radial trailing plane passing through the axis and the trailing point;the trailing edge includes a point of asymmetry such that for any current point of the trailing edge located between the point of asymmetry and the trailing point, a tangent to the trailing edge at the current point extends from the trailing edge away from the axis between the blade and a radial trailing plane passing through the axis and the current point;

[0010] The invention also relates to a turbomachine comprising a stator piece as just described and to an aircraft comprising such a turbomachine. DESCRIPTION OF THE FIGURES

[0011] Other features and advantages of the invention will become apparent from the following description, which is purely illustrative and not limiting, and should be read in conjunction with the accompanying drawings on which: there figure 1 is a schematic representation of a turbomachine; the figure 2is a schematic representation of a stator component according to a first embodiment; the figure 3 is a schematic cross-sectional view in a plane perpendicular to a radial axis of the turbomachine of a stator component according to the first embodiment; the figure 4 is a schematic cross-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 Turbomachine

[0012] With reference to the figure 1A turbomachine is schematically represented, more specifically an axial turbojet 1 with a double-flow design. The turbojet 1 shown extends along an axis Δ and comprises successively, in the direction of gas flow in the turbomachine, a fan 2, a compression section which may include a low-pressure compressor 3 and a high-pressure compressor 4, a combustion chamber 5, and a turbine section which may include a high-pressure turbine 6, a low-pressure turbine 7 and an exhaust nozzle.

[0013] The blower 2 and the low pressure compressor 3 are driven in rotation by the low pressure turbine 7 via a first drive shaft 9, while the high pressure compressor 4 is driven in rotation by the high pressure turbine 6 via a second drive shaft 10.

[0014] In operation, a flow of compressed air from the low and high pressure compressors 3 and 4 feeds combustion in the combustion chamber 5, the expansion 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 1 through an exhaust nozzle downstream of the turbines 6, 7 exert a reaction thrust on the turbojet 1 and, through it, on a vehicle or machine such as an aircraft (not shown). Stator component

[0015] Downstream of the fan or a compression stage, the turbomachine may include a straightening blade stage. Such a straightening blade stage may include a stator piece 20 as shown with reference to the figure 2 .

[0016] The stator piece 20, or the set 20 of stator pieces 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 piece may for example include two adjacent blades 24, 26 which extend from the platform 22.

[0017] The term "platform" here refers to any element of the turbomachine from which blades 24, 26 can be mounted. The platform may be, in particular, a hub or a housing surrounding the turbomachine's axis. The platform may have a cylindrical surface at a constant radial distance from the turbomachine's axis Δ. The platform 22 has either an internal or external wall against which the air flows; that is, the platform 22 defines a wall of a gas flow path. The blades 24, 26 extend from the platform 22 into the flow path either radially outwards, away from the turbomachine's axis Δ, or radially inwards, towards the turbomachine's axis Δ.

[0018] There figure 2is a schematic perspective representation of stator part 20. The axis Δ of the turbomachine is shown oriented positively in the direction of the gas flow within the turbomachine. figure 2 It also represents a radial axis r and a circumferential axis θ passing through a point 34 on the platform 22. At each point in space, and for example at a point 34 on the platform 22, we can define a radial axis r that is perpendicular to the axis Δ of the turbomachine and that passes through the point and the axis Δ of the turbomachine. The radial axis is oriented positively in the direction away from the axis Δ of the turbomachine. We can also define a circumferential axis θ that passes through the point and is perpendicular to the radial axis r and the axis Δ of the turbomachine. The circumferential axis is oriented positively in the direction away from the axis Δ of the turbomachine.

[0019] In the example of the 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 single situation.

[0020] There 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 Δ of the turbomachine. Such a circumferential plane parallel to the axis Δ of the turbomachine allows us to define a cross-section of the blades 24, 26

[0021] The direction of the axis Δ is given on the figure 3 by the x-axis, whose orientation is the direction of gas flow. The radial axis r is perpendicular to the plane of the figure 3 and directed towards the reader of the figure 3 The θ axis corresponds to the circumferential direction perpendicular simultaneously to the Δ axis and the radial axis.

[0022] Blades 24 and 26 each have an intrados 624, 126 and an extrados 124, 626.

[0023] 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 is from upstream to downstream in the direction and sense of the turbomachine's axis Δ.

[0024] 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 at a constant radius or at a constant distance from the axis Δ, a circumferential plane which can be described as a cutting plane.

[0025] Similarly, in a circumferential section plane, each blade exhibits a camber line 43, 41, which is the curve equal to the average of the upper surface (extrados) and lower surface (intrados) curves. More precisely, the camber line is formed by all points equidistant from the upper and lower surfaces. The distance from a particular point to the upper (or lower) surface is defined here as the minimum distance between that particular point and a point on the upper (or lower) surface. Fin

[0026] The stator piece 20 also includes 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 duct radially with respect to the axis Δ of the turbomachine from the platform 22.

[0027] The winglet 28 includes an extrados 50 which is opposite the intrados 126 of the blade 26.

[0028] 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.

[0029] The winglet 28 includes an intrados 48 which is opposite the extrados 124 of the first blade and an extrados 50 which is opposite the intrados 126 of the second blade 26.

[0030] 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.

[0031] The leading edge 30 includes an attack point 34 located on the platform 22. The attack point 34 corresponds to the intersection of the leading edge 30 and the platform 22. A radial attack plane Pa is defined which passes through the axis Δ of the turbomachine and the attack point 34.

[0032] Every radial plane includes the Δ axis of the turbomachine.

[0033] The trailing edge 32 includes a trailing point 36 located on the platform 22. The trailing point 36 corresponds to the intersection of the trailing edge 32 and the platform 22. A radial trailing plane Pf is defined which passes through the axis Δ of the turbomachine and the trailing point 36.

[0034] There figure 4 is a schematic representation in a Pr plane of certain parameters of the fin profile.

[0035] The most general embodiment of the invention corresponds to the following two characteristics in relation to the figures 2 And 4 .

[0036] 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 Δ of the turbomachine and passes through the central axis Δ of the turbomachine.

[0037] Each plane Pr including the radial axis Ar defines a section S of the intrados 48 or respectively of the extrados 50.

[0038] The Pr plane is defined by the direction of the radial axis Ar and any other direction in the plane, such as the circumferential direction, the direction of the central axis Δ of the turbomachine, or another direction. The Pr plane may or may not be a radial plane.

[0039] By defining this plane Pr, we define a cutting plane of the intrados 48 if point 100 belongs to the intrados 48, or a cutting plane of the extrados 50 if point 100 belongs to the extrados 50. The cutting plane then defines a section of the intrados 48 or the extrados 50. The section S passes through point 100 and through a point at the intersection 104 of the base profile 44 and the intrados 48 or the extrados 50. This point is also at the intersection 104 of the section S and the base profile 44.

[0040] With reference to the figure 4, we define in the Pr plane an angle 106 between the foot profile 44 and a tangent T to the section S, the tangent being constructed at the point at the intersection 104.

[0041] This angle 106 is less than or equal to 45 degrees.

[0042] In the Pr plane, 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.

[0043] Any section S as defined above is located between platform 22 and a tangent T as shown above. Since the tangent is relatively close to the platform, in relation to the value of angle 106, the fin has a compact, pyramidal shape. This shape allows it to act on: a first part of the flow passage located upstream of the fin by guiding this first part without separation along the intrados of the fin, and on a second part of the flow passage which is closest to the blade whose intrados is opposite the fin by confining this second part between the intrados of the blade and the extrados of the fin.

[0044] Corner separation is then greatly reduced.

[0045] Furthermore, the upstream section of the fin has a relatively gentle slope compared to platform 22. This reduces the risk of aerodynamic stall, even when this upstream section is located in the region of smallest cross-section of the airflow most prone to stall. This is particularly advantageous at high Mach numbers, where excessive stall can induce shock waves.

[0046] The slow change in height prevents the initial flow from separating: the gases follow the underside of the blade and are thus guided to the trailing edge. Guiding this initial portion of the flow can locally cancel the transverse pressure gradient, thereby limiting stator blade stall.

[0047] It should be noted that if point 100 is on the leading edge 30 or on the trailing edge 32 of the fin 28, 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.

[0048] There figure 4 represents the situation where section S includes a common point with head profile 46, but this is not necessarily the case. Fin height

[0049] The platform 22 as described so far defines an inner radial wall or respectively outer radial wall of the gas flow vein.

[0050] When the stator component 20 corresponds to a shrouded architecture, it includes a second platform located radially opposite the first platform 22. This second platform defines the outer radial wall or, respectively, the inner radial wall of the gas flow path. The gas flow path therefore passes radially between the first platform 22 and the second platform, the path extending radially over a certain height designated by the reference Hv on the figure 4 .

[0051] When the stator piece 20 corresponds to an unshod architecture, it comprises only a single platform 22 defining the inner radial wall of the gas flow duct. The duct extends radially over a certain duct height defined by the height of the stator piece blades 20, blades which project radially outwards from the platform 20.

[0052] The fin 28 extends radially over a fin height Ha indicated in figure 4 .

[0053] According to a first optional variant of the most general embodiment, a ratio of fin height to vein height being greater than or equal to 0.01 and less than or equal to 0.25.

[0054] 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. Circumferential distance from fin to first blade

[0055] The first blade 24 and the second blade 26 are separated circumferentially by a pitch 42. The pitch 42 separating the blades is an angle between a radial direction of the first blade 24 and a radial direction of the second blade 26. The pitch is determined by the total number of blades performing the same function and having the same axial position as blades 24 and 26, and which are located all around the axis Δ of the turbomachine. The distance between the first blade 24 and the second blade is given by this angle and the radius to the axis Δ at which this distance is to be measured.

[0056] We can also define an angle separating the first blade 24 and the fin 28 in the circumferential direction, or an angle separating a radial direction of the first blade 24 and a radial direction of the fin 28.

[0057] The angular separation of the first blade 24 and the fin 28 in the circumferential direction can be freely chosen 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. Fin profiles

[0058] As seen previously, the first blade 24 defines a blade chord 424 between its leading edge 224 and its trailing edge 324.

[0059] The blade 28 can be modeled or represented as a stack of profiles arranged radially between a blade root 44 and a blade tip 46. The blade root 44 is located on the platform 22 and corresponds to the intersection of the blade 28 and the platform 22. The blade tip is located at a distance from the platform 22 within the gas flow channel. Each blade profile extends in a circumferential plane parallel to the axis Δ of the turbomachine, like a cross-section of the blade made in this circumferential plane at a constant radius or constant distance from the axis Δ, a circumferential plane which can be described as a cutting plane.

[0060] With reference to the figure 3Each fin profile defines a fin chord 54 between the leading edge 30 and the trailing edge 32 of the fin. More precisely, 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 and second points. The chord line designates the segment connecting the first and second points.

[0061] 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.

[0062] Each fin profile defines a camber line, which is the curve equal to the average of the curve on the upper surface (extrados 50) of the fin and the curve on the lower surface (intrados 48) of the fin. More precisely, in a given fin profile, the camber line is formed by all points equidistant between the intersection of the upper surface (extrados 50) and the cutting plane, and the intersection of the lower surface (intrados 48) and the cutting plane. The fin's camber line can be chosen to be close to the camber line of the first blade (24).

[0063] The fin exhibits a continuous variation of profiles in the radial direction, that is to say that all parameters of camber, chord, thickness vary continuously from the foot profile to the tip profile.

[0064] A second optional variant of the most general embodiment and / or its variant includes the following three characteristics: the chord of the head profile is less than the chord of the foot profile; a ratio of the chord of the head profile to the blade chord is greater than or equal to 0.1 and less than or equal to 0.6; a ratio of the chord of the foot profile to the blade chord is greater than or equal to 0.3 and less than or equal to 1.1.

[0065] The root profile chord is then relatively large, which allows the flow passage to be blocked over a large part of the chord length of the blade.

[0066] A third optional variant of the most general embodiment and / or its variants includes the following two characteristics: 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 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.

[0067] The fin is therefore relatively thick, which facilitates its manufacture and mechanical strength, particularly in the event of ingestion of particles or foreign bodies by the engine. "Metal" angles of the fin and the first blade

[0068] For the first blade 24 and the second blade 26, we can define: a blade leading metal angle as the angle between the tangent to the camber line at the leading edge of the blade and the Δ axis of the turbomachine, the angle being oriented from the Δ axis 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 Δ axis, the angle being oriented from the Δ axis towards the tangent.

[0069] For each fin profile, we can define: a leading metal angle of the fin as the angle between the tangent 61 to the line of camber at the leading edge of the fin and the axis Δ of the turbomachine, the angle being oriented from the axis Δ towards the tangent 61, and a trailing metal angle of the fin as the angle between the tangent 63 to the line of camber at the trailing edge of the fin and the axis Δ, the angle being oriented from the axis Δ towards the tangent 63.

[0070] In the particular case of the fins described in this invention, one can choose to estimate the metal angles as a function of the local camber angle of the blade 24 or 26.

[0071] To achieve this, an axial position and a radial position can be associated with each point on each camber line of each fin profile. The fin point is associated with a reference point on a camber line 43 of a profile of the first blade 24, the reference point presenting the axial and radial positions of the fin point.

[0072] The metal leading edge angle of the fin can be associated with the local camber angle, which is the angle oriented from the axis Δ to the tangent to the camber line of the first blade's airfoil 24 at the associated reference point. The metal leading edge angle of the fin can be described with reference to this local camber angle as the difference between this metal angle and this local camber angle.

[0073] The local camber angle, which is the angle oriented from the axis Δ to the tangent to the camber line of the first blade's airfoil 24 at the associated reference point, can be associated with the metal-to-fin trailing angle. The metal-to-fin trailing angle with reference to this local camber angle can be described as the difference between this metal-to-fin trailing angle and this local camber angle.

[0074] In general, the fins described in this invention have leading and trailing edges of the tip and foot profiles with metal angles close to the local camber angle. More precisely: the angle formed by a first tangent to the camber line of the fin tip profile at the leading edge 30 and by a second tangent to the camber line of the first blade profile at the reference point associated with the leading edge of the fin tip profile is less than or equal to 10 degrees in absolute value; the angle formed by a third tangent to the camber line of the fin tip profile at the trailing edge 32 and a fourth tangent to the camber line of the first blade profile at the reference point associated with the trailing edge of the fin tip profile is less than or equal to 10 degrees in absolute value; the angle formed by a fifth tangent to the camber line of the fin root profile at the leading edge 34 and a sixth tangent to the camber line of the first blade profile at the reference point associated with the leading edge 34 is less than or equal to 10 degrees in absolute value;and the angle formed by a seventh tangent to the camber line of the fin root profile at vanishing point 36 and an eighth tangent to the camber line of the first blade profile at the reference point associated with vanishing point 36 is less than or equal in absolute value to 10 degrees. ;

[0075] The tip and root profiles then present 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 in the same way as the blades. Axial position of the fin relative to the first blade

[0076] To characterize the position of the fin in the direction of the axis Δ of the turbomachine, an axial coordinate of a point at mid-chord of a fin profile is used.

[0077] The mid-chord point is located equidistant from the leading edge 30 of the fin and the trailing edge 32 of the fin.

[0078] The axial coordinate of the mid-chord point can be chosen 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.

[0079] In particular, the mid-chord point of the fin head profile can be used as the mid-chord point.

[0080] The fin's foot profile is positioned axially relative to the tip profile thanks to the fin stacking law which gives the relative positioning of the fin's tip profile relative to the fin's foot profile.

[0081] For this relative positioning, the mid-chord point is chosen as the positioning reference for each stacking profile. Two axes are defined: a first axis “t” directed along the direction of the chord of the foot profile and oriented in the same direction as the axis Δ 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 θ from the first blade 24 to the second blade 26.

[0082] Two angles are defined: a sweep angle (also known by the English term "sweep") which is the angle between the axis "t", and a direction defined by the mid-chord point of the foot profile and the mid-chord point of the head profile; a dihedral angle (also known by the English term "lean") which is the angle between the axis "n", and the direction defined by the mid-chord point of the foot profile and the mid-chord point of the head profile.

[0083] In the specific case of the fins described here, the sweep and dihedral angles, as defined above, take values ​​greater than or equal to -10 degrees and less than or equal to +10 degrees. These angles are sufficiently small that the axial position of the fin can be considered well described by an axial coordinate 78 of a point at mid-chord 76 of a fin profile. Positive dihedral effect at the trailing edge

[0084] A second embodiment dependent on the most general embodiment or its first variant includes the following characteristic.

[0085] The blade 28 has a tangent to the trailing edge 32 at the trailing point 36, extending along the flow path from the platform 22 between the second blade 26 and the radial trailing plane Pf. In other words, in a radial plane orthogonal to the axis Δ of the turbomachine and passing through the trailing point, the tangent to the trailing edge, as it moves away from the platform 22 on the flow path side of the platform, approaches the second blade 26. Put another way, the trailing edge 32 is inclined at the trailing point 36 towards the second blade 26. It should be noted that this tangent to the trailing edge may also have a non-zero projection along the axis Δ of the turbomachine.

[0086] The inclination towards the second blade on the trailing edge side creates a trailing-edge dihedral effect that more effectively blocks the throughflow. This prevents the throughflow from the second blade 26 from flowing up the upper surface of the winglet 28 and beyond the winglet to reach the lower surface of the first blade 24. The winglet is therefore "flattened" on its trailing edge so that its surface slopes towards the lower surface of the second blade 26. This is the positive dihedral effect. The throughflow is then strongly deflected towards the lower surface of the second blade 26.

[0087] 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. Positive dihedral effect downstream of a first point of asymmetry of the edge of leak

[0088] More generally, the fin 28 may include a point of asymmetry and the following characteristic: at any current point on the trailing edge located between this point of asymmetry and the trailing point 36, the tangent to the trailing edge extending in the groove from the platform 22 lies between the second blade 26 and a radial plane passing through the current point. In other words, in a radial plane passing through the current point, the trailing edge, as it moves away from the current point towards the point of asymmetry, approaches the second blade 26. Put another way, the trailing edge 32 is inclined between the trailing point 36 and the point of asymmetry towards the second blade 26.

[0089] The trailing edge upstream of the point of asymmetry 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.

[0090] The leading edge can also be symmetrical, that is, each tangent to the leading edge is contained in a radial plane which includes the axis of the turbomachine.

[0091] In this situation, the fin can be described as symmetrical upstream of the point of asymmetry and asymmetrical downstream. Positive dihedral effect downstream of a second leading edge asymmetry point

[0092] When the point of asymmetry is located on the head profile, this point of asymmetry can be a first point of asymmetry and the fin can include a second point of asymmetry located this time on the leading edge.

[0093] At any current point on the leading edge located between this point of asymmetry and the leading edge profile, the tangent to the leading edge extending along the leading edge's path lies 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, as it moves away from the current point towards the leading edge profile, approaches the second blade 26. Put another way, the leading edge 30 is inclined between the second point of asymmetry and the leading edge profile towards the second blade 26.

[0094] The fin can then be described as symmetrical upstream of the second point of asymmetry and asymmetrical downstream. Optionally, the second point of asymmetry could be the leading edge 34, in which case the entire fin could be described as asymmetrical.

Claims

1. A stator part (20) of a turbine engine comprising: - a platform (22) defining a wall of a gas flow stream, - a blade (24, 26) extending radially relative to a central axis (Δ) of the turbine engine from the platform (22), and - a fin (28) extending radially into the stream from a fin root (44) located on the platform (22) to a fin tip (46), the fin comprising a lower side (48) and an upper side (50) located radially between the fin root (44) and the fin tip (46), each point (100) of the lower side (48) or respectively of the upper side (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 lower side (48) or respectively of the upper side (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 of the root profile (44) being less than or equal to 45 degrees, the section (S) being located between the root profile (44) and the tangent (T).

2. The stator part according to claim 1, wherein the platform (22) is a first platform, the part comprising a second platform so as to define the flow stream between the first platform (22) and the second platform, the stream extending radially over a stream height (Hv), the fin (28) extending radially over a fin height (Ha), a ratio of the fin height (Ha) to the stream height (Hv) being greater than or equal to 0.01 and less than or equal to 0.25.

3. The stator part according to 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 fin (28) comprising a plurality of radially stacked profiles between the fin root (44) and the fin tip (46), each profile defining a chord (54) between the leading edge (30) of the fin (28) and the trailing edge (32) of the fin (28), the tip profile chord being less than the root profile chord; a ratio of the tip profile chord to the blade chord being greater than or equal to 0.1 and less than or equal to 0.6; a ratio of the root profile chord to the blade chord being greater than or equal to 0.3 and less than or equal to 1.1.

4. The stator part according to claim 3, wherein each fin profile defines a maximum thickness (52) of the profile between the lower side (48) and the upper side (50) in a direction perpendicular to a chord line, a ratio of the maximum thickness of the root profile to the root profile chord 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 tip profile to the tip profile chord being greater than or equal to 0.05 and less than or equal to 0.25.

5. The stator part according to one of the preceding claims, wherein the blade comprises a lower side (126) facing the upper side (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) while separating itself from the axis (Δ) between the blade (26) and a radial trailing plane (Pf) passing through the axis (Δ) and the trailing point (36).

6. The stator part according to claim 5, wherein the trailing edge (32) comprises an asymmetry point, so 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) while separating itself from the axis (Δ) between the blade (26) and a radial trailing plane passing through the axis (Δ) and the current point.

7. A turbine engine comprising a stator part (20) according to one of claims 1 to 6.

8. An aircraft comprising a turbine engine according to claim 7.

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