Stator component with fins in a turbomachine

Abstract

A stator component (20) of a turbomachine comprising a platform (22), a blade (24, 26) extending radially about a central axis (∆), and a blade (28) extending radially from a blade root (44) to a blade tip (46), the blade comprising an intrados (48) and an extrados (50), 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 defined in the plane (Pr) between the root profile (44) and a tangent to the section (S) at an intersection (104) of the section (S) and the root profile (44) being less than or equal to 45 degrees, the section (S) being located between the toe profile (44) and the tangent. Figure for the abbreviation: Fig. 2

Description

Title of the invention: Stator component with fins in a turbine 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, and 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 a 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 (also known as OGVs), inlet guide vanes (also known as IGVs), and variable stator vanes (also known as VSVs). The stator blades of an aircraft gas turbine engine may each have two platforms (inner and outer) that are attached to the blade.There are also unfaired designs featuring stator blades with only a single internal platform. In all cases, these stator blades form fixed rows that guide the gas flow through the engine at the appropriate speed and angle.

[0003] Within a flow straightener comprising a plurality of fixed blades, the gas flow generally occurs 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.

[0004] 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 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 a vortex can occur. This separation generates losses of pressure as well as aerodynamic blockage. The latter is problematic in terms of operability. For high angles of incidence of the flow arriving at the stator, that is to say when the direction of gas flow upstream of the stator makes a large angle with a direction of the leading edge of the blade, this wedge separation is amplified until it causes a separation of the boundary layer on the blade which can no longer ensure the deflection of the flow.

[0006] There is therefore a need for a new geometry to correct these problems and improve performance in terms of equipment efficiency, particularly with a high incidence of the flux entering the rectifier. Description of the invention

[0007] An object 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:

[0009] - a platform defining a wall of a gas flow vein,

[0010] - a blade extending radially with respect to a central axis of the turbomachine from the platform, and

[0011] - a fin extending radially in the vein from a fin base located on the platform up to a fin tip, the fin comprising an intrados and an extrados situated radially between the fin root and the fin tip,

[0012] 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.

[0013] Such a stator component is advantageously and optionally complemented by the following various features, taken alone or in combination:

[0014] - 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;

[0015] - 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 head profile being less than the chord of the foot profile; a ratio of the chord of the head 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 foot profile to the blade chord being greater than or equal to 0.3 and less than or equal to 1.1;

[0016] - each fin profile defines a maximum profile thickness between the intrados and the extrados 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;

[0017] - the blade comprises an intrados opposite the extrados of the fin, the fin including a trailing edge, the trailing edge including 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 trailing plane passing through the axis and the vanishing point;

[0018] - the trailing edge includes a point of asymmetry such that for any current point from the trailing edge located between the point of asymmetry and the vanishing 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;

[0019] The invention also relates to a turbomachine comprising a stator component as described above and to an aircraft comprising such a turbomachine. DESCRIPTION OF FIGURES

[0020] 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:

[0021] - the [Fig.1] is a schematic representation of a turbomachine; - [Fig. 2] is a schematic representation of a stator component according to a first embodiment; - the [Fig.3] is a schematic cross-sectional view in a plane perpendicular to a radial axis of the turbomachine of a stator part according to the first embodiment; - [Fig.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

[0022] With reference to [Fig. 1], a turbomachine is schematically represented, more specifically an axial turbojet 1 with double flow. The turbojet 1 illustrated extends along an axis A 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.

[0023] 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.

[0024] In operation, a flow of air compressed by the low and high pressure compressors 3 and 4 feeds a 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 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 illustrated). Stator component

[0025] Downstream of the blower or a compression stage, the turbomachine may include a stage of straightening blades. Such a stage of straightening blades may include a stator piece 20 as shown with reference to [Fig.2].

[0026] The stator piece 20, or the assembly 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, comprise two adjacent blades 24, 26 which extend from the platform 22.

[0027] The term "platform" here refers to any element of the turbomachine from which blades 24, 26 are suitable for mounting. The platform may, in particular, be a hub or a housing surrounding 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 internal or external wall against which the air flows; that is to say, 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 axis of the turbomachine A, or radially inwards, towards the axis of the turbomachine A.

[0028] Figure 2 is a schematic perspective view of the stator part 20. The axis A of the turbomachine is shown oriented positively in the direction of the gas flow 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 on platform 22, we can define a radial axis r that is perpendicular to the axis A of the turbomachine and passes through the point and the axis A of the turbomachine. The radial axis is oriented positively in the direction away from the axis A of the turbomachine. We can also define a circumferential axis 0 that passes through the point and is perpendicular to the radial axis r and to the axis A of the turbomachine. The circumferential axis is oriented positively in the direction away from the axis A of the turbomachine.

[0029] In the example of [Fig.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.

[0030] 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 allows a cross-section of the blades 24, 26 to be defined.

[0031] The direction of axis A is given in [Fig. 3] by the x-axis, the orientation of which is the direction of gas flow. The radial axis r is perpendicular to the plane of [Fig. 3] and directed towards the reader of [Fig. 3]. Axis 0 corresponds to the circumferential direction perpendicular simultaneously to axis A and the radial axis.

[0032] The blades 24 and 26 each have an intrados 624, 126 and an extrados 124, 626.

[0033] The blades 24 and 26 each comprise a leading edge 224, 226 on the upstream side and a trailing edge 324, 326 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.

[0034] 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 A, a circumferential plane which can be described as a cutting plane.

[0035] Similarly, in a circumferential section plane, each blade exhibits a camber line 43, 41 which is the curve equal to the average of the curve on the upper surface (extrados) and the curve on the lower surface (intrados). More precisely, the camber line is formed by all points located equidistant from the upper surface (extrados) and the lower surface (intrados). The distance from a particular point to the upper surface (or lower surface) is defined here as the minimum distance between the particular point and a point on the upper surface (or lower surface). Fin

[0036] 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 A of the turbomachine from the platform 22.

[0037] The winglet 28 includes an extrados 50 which is opposite the intrados 126 of the blade 26.

[0038] 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.

[0039] The fin 28 comprises 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.

[0040] 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.

[0041] 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 A of the turbomachine and the attack point 34.

[0042] Any radial plane includes the axis A of the turbomachine.

[0043] 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 A of the turbomachine and the trailing point 36.

[0044] Figure 4 is a schematic representation in a Pr plane of certain parameters of the fin profile.

[0045] The most general embodiment of the invention corresponds to the following two characteristics in relation to figures 2 and 4.

[0046] 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.

[0047] Each plane Pr comprising the radial axis Ar defines a section S of the intrados 48 or respectively of the extrados 50.

[0048] 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 may or may not be a radial plane.

[0049] By defining this plane Pr, a cutting plane of the intrados 48 is defined 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.

[0050] With reference to [Fig.4], an angle 106 is defined in the Pr plane between the foot profile 44 and a tangent T to the section S, the tangent being constructed at the point at the intersection 104.

[0051] This angle 106 is less than or equal to 45 degrees.

[0052] 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.

[0053] Any section S as defined above is located between the 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:

[0054] - a first part of the passage flow located upstream of the fin in guiding this first part without separation along the intrados of the winglet, and on

[0055] - a second part of the flow passage 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.

[0056] Corner separation is then greatly reduced.

[0057] Furthermore, the upstream portion of the fin has a relatively gentle slope relative to the platform 22. This reduces the risk of aerodynamic stalling even when this upstream portion is located in the region of smallest cross-section of the airflow most prone to stalling. This is particularly advantageous at high Mach numbers where excessive stalling can induce shock waves.

[0058] The slow evolution of the height prevents the initial part of the flow from separating: the gases follow the lower surface of the blade and are thus guided to the trailing edge of the blade. Guiding this initial part of the flow can locally cancel the transverse pressure gradient, thereby limiting stator blade stall.

[0059] 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.

[0060] Fig. 4 represents the situation where section S includes a common point with the head profile 46, but this is not necessarily the case. Fin height

[0061] The platform 22 as described so far defines an inner radial wall or respectively outer radial wall of the gas flow vein.

[0062] When the stator piece 20 corresponds to a streamlined 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 channel. The gas flow channel therefore passes radially between the first platform 22 and the second platform, the channel extending radially over a certain channel height designated by the reference Hv on [Fig.4].

[0063] When the stator part 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 blades of the stator part 20, blades which project radially outwards from the platform 20.

[0064] The fin 28 extends radially over a fin height Ha indicated in [Fig.4].

[0065] 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.

[0066] It may be advantageous for the fin described above to remain of relatively small height so as not to block the flow in the vein.

[0067] Circumferential distance fin - first blade

[0068] The first blade 24 and the second blade 26 are separated in a circumferential direction 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 as and having the same axial position as blades 24 and 26, and which are located all around the axis A of the turbine. The distance separating the first blade 24 and the second blade is given by this angle and the radius to the axis A at which this distance is to be measured.

[0069] We can also 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.

[0070] 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

[0071] As seen previously, the first blade 24 defines a blade chord 424 between its leading edge 224 and its trailing edge 324.

[0072] The fin 28 can be modeled or represented as a stack of profiles along a radial direction between a fin root 44 and a fin tip 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 tip is located at a distance from the platform 22 in the vein gas flow. Each fin profile extends in a circumferential plane parallel to the axis A of the turbomachine, like a cross-section of the fin made in this circumferential plane at a constant radius or constant distance from the axis A, a circumferential plane which can be described as a cutting plane.

[0073] With reference to [Fig. 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 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 point and the second point. The chord line designates the segment connecting the first point and the second point.

[0074] 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.

[0075] Each fin profile allows for the definition of a camber line, which is the curve equal to the average of the curve on the upper surface 50 of the fin and the curve on the lower surface 48 of the fin. More precisely, in a given fin profile, the camber line is formed by all points located equidistant 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 fin camber line can be chosen to be close to the camber line of the first blade 24.

[0076] 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 head profile.

[0077] A second optional variant of the most general embodiment and / or its variant comprises the following three features:

[0078] - the chord of the head profile is less than the chord of the foot profile;

[0079] - 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;

[0080] - 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.

[0081] 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.

[0082] A third optional variant of the most general embodiment and / or its variants comprises the following two features:

[0083] - 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

[0084] - 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.

[0085] The fin is then relatively thick, which facilitates its manufacture and mechanical strength, particularly in the event of ingestion of particles or foreign bodies by the engine.

[0086] "Metal" anterior parts of the fin and the first blade

[0087] For the first blade 24 and the second blade 26, we can define:

[0088] - a blade attack metal angle as the angle between the tangent to the line of camber at the leading edge of the blade and the A axis of the turbomachine, the angle being oriented from the A axis towards the tangent, and

[0089] - a blade trailing metal angle as the angle between the tangent to the line of camber at the trailing edge of the blade and axis A, the angle being oriented from axis A towards the tangent.

[0090] For each fin profile, we can define:

[0091] - a fin attack metal angle such as the angle between the tangent 61 to the line of camber at the leading edge of the fin and the A axis of the turbomachine, the angle being oriented from the A axis towards the tangent 61, and

[0092] - a fin trailing metal angle such as the angle between the tangent 63 to the line of camber at the trailing edge of the fin and axis A, the angle being oriented from axis A towards tangent 63.

[0093] 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.

[0094] To this end, an axial position and a radial position can be associated with each point on each camber line 64 of each fin profile. The fin point is associated with a reference point on a camber line 66 of a profile of the first blade 24, the reference point having the axial position and the radial position of the fin point.

[0095] The metal leading edge angle of the fin can be associated with the local camber angle, which is the angle oriented from axis A to the tangent to the camber line of the profile of the first blade 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.

[0096] The local camber angle, which is the angle oriented from axis A to the tangent to the camber line of the first blade's profile 24 at the associated reference point, can be associated with the metal angle of the fin. The metal 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.

[0097] Generally speaking, the fins described in this invention have metal angles at the leading and trailing edges of the tip and foot profiles close to the local angle of camber. More precisely:

[0098] - the angle formed by a first tangent to the line of camber of the head profile of fin 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;

[0099] - the angle formed by a third tangent to the line of camber of the head profile of fin at 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;

[0100] - the angle formed by a fifth tangent to the line of camber of the foot profile of fin at the attack point 34 and a sixth tangent to the camber line of the profile of the first blade at the reference point associated with the attack point 34 is less than or equal in absolute value to 10 degrees; and

[0101] - the angle formed by a seventh tangent to the line of camber of the foot profile of fin at vanishing point 36 and an eighth tangent to the camber line of the profile of the first blade at the reference point associated with vanishing point 36 is less than or equal in absolute value to 10 degrees.

[0102] The tip and foot 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 like the blades.

[0103] Axial position of the fin relative to the first blade

[0104] To characterize the position of the fin in the direction of the axis A of the tur-bomachine, an axial coordinate of a point at mid-chord of a fin profile is used.

[0105] The mid-chord point is located at an equal distance from the leading edge 30 of the fin and the trailing edge 32 of the fin.

[0106] 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.

[0107] In particular, the mid-chord point of the fin head profile can be used as the mid-chord point.

[0108] The foot profile of the fin is positioned axially relative to the head profile thanks to the stacking law of the fin which gives the relative positioning of the fin head profile relative to the foot profile of the fin.

[0109] For this relative positioning, the mid-chord point is chosen as the positioning reference for each stacking profile. Two axes are defined:

[0110] - a first axis "t" directed along the direction of the chord of the foot profile and oriented in the same direction as axis A of the turbomachine from upstream to downstream;

[0111] - a second axis "n" perpendicular to the first axis "t" and oriented in the same direction that the circumferential axis 0 of the first blade 24 towards the second blade 26.

[0112] Two angles are defined:

[0113] - 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;

[0114] - 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.

[0115] In the particular 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

[0116] A second embodiment depending on the most general embodiment or its first variant includes the following feature.

[0117] The blade 28 has a tangent to the trailing edge 32 at the trailing point 36, extending in the flow 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 A 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 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 A of the turbomachine

[0118] 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 the trailing edge side so as to present a surface inclined 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.

[0119] 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.

[0120] Positive dihedral effect downstream of a first asymmetry point of the trailing edge

[0121] More generally, the fin 28 may include an asymmetry point and the ca 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 vein from the trailing edge 32 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.

[0122] 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.

[0123] The leading edge can also be symmetrical, that is to say that each tangent to the leading edge is contained in a radial plane which includes the axis of the turbomachine.

[0124] In this situation, the fin can be described as symmetrical upstream of the point of asymmetry and asymmetrical downstream.

[0125] Positive dihedral effect downstream of a second asymmetry point of the leading edge

[0126] When the asymmetry point is located on the head profile, this asymmetry point can be a first asymmetry point and the fin can include a second asymmetry point included this time on the leading edge.

[0127] 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 in the groove from the leading edge 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.

[0128] 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 edge 34, in which case the entire fin can be described as asymmetrical.

Claims

Demands

1. A stator component (20) of a turbomachine comprising: - a platform (22) defining a wall of a gas flow duct, - a blade (24, 26) extending radially from the platform (22) to a central axis (A) of the turbomachine, and - a blade (28) extending radially in the duct from a blade root (44) located on the platform (22) to a blade tip (46), the blade comprising an intrados (48) and an extrados (50) situated radially between the blade root (44) and the blade 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 the extrados (46),an angle (106) defined in the plane (Pr) between the foot profile (44) and a tangent (T) to the section (S) at an intersection (104) of the section (S) and the foot profile (44) being less than 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. A 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 piece according to claim 3, wherein each fin profile defines a maximum thickness (52) of the profile between the lower surface (48) and the upper surface (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 top profile to the chord of the top 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 piece (20) according to any one of claims 1 to 6.

8. Aircraft comprising a turbomachine according to claim 7.

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