Stator component having a fin in a turbine engine

Abstract

The invention relates to a stator component (20) of a turbine engine, the stator component comprising a platform (22), blades (24, 26) extending radially with respect to a central axis (A), and a vane (28) extending radially from a vane root (44) to a vane tip (46), the vane comprising an underside (48) and an upper side (50), each point (100) of the underside or the upper side defining a radial axis (Ar) therethrough, comprising the radial axis defines a lower or upper cross-section (S) in which an angle defined between the root profile and a tangent to the cross-section at an intersection (104) of the cross-section and the root profile is less than or equal to 45 degrees, the cross-section being located between the root profile and the tangent.

Description

Technical Field

[0001] The invention relates to a stator component of a turbine engine, comprising blades, such as a rectifier, in particular a fixed-pitch rectifier, situated downstream of a compressor. Background Art

[0002] In a turbine engine of an aircraft (particularly an aircraft for transporting passengers), the air propelled by the fan and the combustion gases leaving the turbine engine through the exhaust nozzle exert reverse thrust on the turbine engine, and reverse thrust is exerted on the aircraft by the turbine engine. The gas circulation through the turbine engine is affected by rotating blades and fixed blades. In particular, the fixed blades or stator blades include outlet guide vanes (Outlet Guide Vane or OGV), inlet guide vanes (Inlet Guide Vane or IGV) and variable stator vanes (Variable Stator Vane or VSV). The rectifier blades of an aviation gas turbine engine can have two platforms (i.e., an inner platform and an outer platform) applied to the blades, respectively. There is also a shroudless structure including rectifier blades, which has only a single inner platform. In any case, these rectifier blades form multiple rows of fixed blades, which enable the airflow through the engine to be guided at an appropriate speed and angle.

[0003] In a straightener comprising a plurality of stationary blades, the overall gas flow occurs between the blades in an upstream-downstream direction. However, it is known that the region of the blade root may be the location of a secondary aerodynamic flow.

[0004] For each pair of blades facing each other, the pressure gradient between the pressure face (lower side) of the first blade and the suction face (upper side) of the second blade creates a crossflow that transports the gas towards the upper side.

[0005] At the ends of the blades, i.e. at the junction between the cascade and the hub or between the cascade and the casing, corner separation and corner vortex may occur. This separation generates pressure losses as well as aerodynamic blockage. Aerodynamic blockage is problematic in terms of operability. For large angles of attack of the airflow arriving at the rectifier, i.e. when the flow direction of the gas upstream of the rectifier forms a large angle with the direction of the leading edge of the blade, this corner separation is amplified until it causes separation of the boundary layer on the blade, which can no longer provide a deflection of the airflow.

[0006] There is therefore a need for a new geometry that makes it possible to correct these problems and improve the performance in terms of efficiency of the device, in particular at high angles of attack of the airflow entering the straightener. Summary of the invention

[0007] One object of the present invention is to propose a stator component for a turbine engine, the geometry of which improves the flow of the fluid with respect to the prior art.

[0008] This object is achieved within the scope of the invention by a stator component comprising:

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

[0010] - blades extending radially from the platform relative to the central axis of the turbine engine, and

[0011] - an airfoil extending radially into the flow channel from an airfoil root located on the platform to an airfoil tip, the airfoil comprising an underside and an upper side located radially between the airfoil root and the airfoil tip,

[0012] Each point on the lower side or the upper side defines a radial axis passing through the point, and each plane including the radial axis defines a cross-section of the lower side or the upper side, in which the angle defined between the root profile and the tangent of the cross-section at the intersection of the cross-section and the root profile is less than or equal to 45 degrees, and the cross-section is located between the root profile and the tangent.

[0013] A stator component of this type is advantageously and optionally realized by the following different features, taken alone or in combination:

[0014] - the platform is a first platform, the component includes a second platform to define a flow channel between the first platform and the second platform, the flow channel extends radially at a flow channel height, the fins extend radially at a fin height, and the ratio of the fin height to the flow channel height is 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 airfoil comprises a plurality of radially stacked profiles between the airfoil root and the airfoil tip, each profile defining a chord between the leading edge of the airfoil and the trailing edge of the airfoil, 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;

[0016] - each wing profile defines a maximum thickness of the profile between the lower side and the upper side in a direction perpendicular to the chord line, the 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; the ratio of the maximum thickness of the tip profile to the chord of the tip profile being greater than or equal to 0.05 and less than or equal to 0.25;

[0017] - the blade comprises a lower side facing the upper side of the airfoil, the airfoil comprising a trailing edge, the trailing edge comprising a rear point located on the platform, a tangent line to the trailing edge at the rear point extending from said platform, while the tangent line to the trailing edge at the rear point is separated from the axis between the blade and a radial rear plane passing through the axis and the rear point;

[0018] - the trailing edge comprises an asymmetric point such that for any current point of the trailing edge between the asymmetric point and the trailing point, a tangent to the trailing edge at the current point extends from the trailing edge while the tangent to the trailing edge at the current point is separated from the axis between the blade and a radial rear plane passing through the axis and the current point.

[0019] The invention also relates to a turbine engine comprising a stator component as just proposed, and to an aircraft comprising a turbine engine of this type. BRIEF DESCRIPTION OF THE DRAWINGS

[0020] Other characteristics and advantages of the invention will emerge from the following description, which is purely illustrative and non-limiting and must be read with reference to the accompanying drawings, in which:

[0021] - Figure 1 is a schematic diagram of a turbine engine;

[0022] - Figure 2 is a schematic diagram of a stator component according to a first embodiment;

[0023] - Figure 3 is a schematic cross-sectional view of a stator component according to a first embodiment in a plane perpendicular to the radial axis of the turbine engine;

[0024] - Figure 4 is a schematic cross-sectional view of a stator component according to a first embodiment in a plane perpendicular to the axis of the turbine engine. DETAILED DESCRIPTION

[0025] Turbine engine

[0026] refer to Figure 1 , schematically shows a turbine engine, more specifically a twin-flow axial turbojet engine 1. The turbojet engine 1 shown extends along an axis △, and the turbojet engine 1 includes a fan 2, a compression section (the compression section may include a low-pressure compressor 3 and a high-pressure compressor 4), a combustion chamber 5, and a turbine section (the turbine section may include a high-pressure turbine 6, a low-pressure turbine 7 and an exhaust nozzle) in sequence in the gas flow direction in the turbine engine.

[0027] The fan 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 .

[0028] During operation, the air flow compressed by the low-pressure compressor 3 and the high-pressure compressor 4 supports the combustion in the combustion chamber 5, and the expansion of the combustion gases from the combustion chamber drives the high-pressure turbine 6 and the low-pressure turbine 7. The air propelled by the fan 2 and the combustion gases leaving the turbojet engine 1 through the exhaust nozzle downstream of the turbines 6, 7 exert reverse thrust on the turbojet engine 1 and, through the turbojet engine, on a vehicle or machine, such as an aircraft (not shown).

[0029] Stator components

[0030] Downstream of the fan or compression stage, the turbine engine may include a straightening vane stage. This type of straightening vane stage may include a straightening vane stage as shown in FIG. Figure 2 The stator component 20 is shown.

[0031] The stator component 20 or the assembly 20 of stator components, if not of one-piece design, has at least one blade 24, 26 and a platform 22 from which the blade 24, 26 extends. The stator part may comprise two adjacent blades 24, 26 extending from the platform 22, for example.

[0032] Here, the term "platform" means any element of the turbine engine on which the blades 24, 26 can be mounted. In particular, the platform can be a hub or a housing around the axis of the turbine engine. The platform can have a cylindrical surface that is at a constant radial distance from the axis Δ of the turbine engine. The platform 22 has an inner wall or an outer wall against which the air circulates, that is, the platform 22 defines the wall of the gas flow channel. The blades 24, 26 extend radially outward from the platform 22 into the flow channel while separating the blades 24, 26 from the axis △ of the turbine engine; or the blades 24, 26 extend radially inward from the platform 22 into the flow channel while being close to the axis △ of the turbine engine.

[0033] Figure 2 is a schematic perspective view of a stator component 20. It is shown that the axis Δ of the turbine engine is oriented positively in the direction of gas flow in the turbine engine. Figure 2 Also shown are the radial axis r and the circumferential axis θ that pass through the point 34 of the platform 22. At each point in space, for example at the point 34 of the platform 22, a radial axis r can be defined that is perpendicular to the axis △ of the turbine engine and passes through the point and the axis △ of the turbine engine. The radial axis is oriented in a direction that separates it from the axis △ of the turbine engine. A circumferential axis θ can also be defined that passes through the point and is perpendicular to the radial axis r and the axis △ of the turbine engine. The circumferential axis is oriented in a direction that separates it from the axis △ of the turbine engine.

[0034] exist Figure 2In the example of FIG. 2 , the blades 24 , 26 extend radially from the platform 22 , while the blades 24 , 26 are separated from the axis of the turbine engine, but the invention is not limited to this case.

[0035] Figure 3 is a schematic representation of the stator part 20 in a circumferential plane passing through the platform 22 and at a constant distance from the axis Δ of the turbine engine. Such a circumferential plane, parallel to the axis Δ of the turbine engine, enables the cross section of the blades 24, 26 to be defined.

[0036] The direction of the axis △ is Figure 3 The radial axis r is perpendicular to Figure 3 plane and points to Figure 3 The axis θ corresponds to the circumferential direction and is perpendicular to both the axis △ and the radial axis.

[0037] Each of the blades 24 and 26 has an underside 624 , 126 and an upper side 124 , 626 .

[0038] Each of the blades 24 and 26 comprises a leading edge 224, 226 located on the upstream side, and a trailing edge 324, 326 located on the downstream side. The terms "upstream" and "downstream" are defined relative to the general gas flow through the turbine engine, which occurs from upstream to downstream in the direction and orientation of the axis Δ of the turbine engine.

[0039] The blade defines a blade chord 424 which is the length of a line segment connecting the leading edge and the trailing edge in a circumferential plane having a constant radius or a constant distance from the axis Δ, which circumferential plane may be considered a cross-sectional plane.

[0040] Likewise, in the circumferential cross-sectional plane, each blade has a median camber line 43, 41, which is a curve equal to the average of the upper side curve and the lower side curve. More precisely, the median camber line is formed by all points that are at an equal distance from the upper side and the lower side. Here, the distance from a specific point to the upper side (or lower side) is defined as the minimum distance between the specific point and the point on the upper side (or lower side).

[0041] Wings

[0042] The stator part 20 also comprises a vane 28 extending from the platform 22 in the same direction and in the same extension orientation as the blades 24, 26. The vane 28 extends radially from the platform 22 into the flow channel relative to the axis Δ of the turbine engine.

[0043] The airfoil 28 includes an upper side 50 that faces the underside 126 of the blade 26 .

[0044] When the component comprises two blades 24, 26, the fin 28 is located between the blade 24 and the blade 26. More precisely, the fin 28 is located facing the upper side 124 of the first blade 24 and the lower side 126 of the second blade 26.

[0045] The airfoil 28 comprises an underside 48 facing the upper side 124 of the first blade and an upper side 50 facing the underside 126 of the second blade 26 .

[0046] The airfoil 28 includes a leading edge 30 and a trailing edge 32 , the leading edge 30 being located upstream of the trailing edge 32 .

[0047] The leading edge 30 comprises a front point 34 situated on the platform 22. The front point 34 corresponds to the intersection of the leading edge 30 with the platform 22. A radial front plane Pa is defined, passing through the axis Δ of the turbine engine and the front point 34.

[0048] Any radial plane includes the axis Δ of the turbine engine.

[0049] The trailing edge 32 comprises a rear point 36 situated on the platform 22. The rear point 36 corresponds to the intersection of the trailing edge 32 with the platform 22. A radial rear plane Pf is defined, which passes through the axis Δ of the turbine engine and the rear point 36.

[0050] Figure 4 It is a schematic diagram of some parameters of the airfoil profile in plane Pr.

[0051] The most general embodiment of the present invention corresponds to the reference Figure 2 and Figure 4 The following two characteristics.

[0052] Each point 100 of the underside 48 of the airfoil 28 or of the upper side 50 of the airfoil 28 defines a radial axis Ar passing through the point 100. The radial axis is orthogonal to the central axis Δ of the turbine engine and passes through the central axis Δ of the turbine engine.

[0053] Each plane Pr including the radial axis Ar defines a section S of the lower side 48 or of the upper side 50 .

[0054] 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 Δ of the turbine engine, or other directions. The plane Pr may or may not be a radial plane.

[0055] When defining this plane Pr, if the point 100 belongs to the lower side 48, a section plane of the lower side 48 is defined, or if the point 100 belongs to the upper side 50, a section plane of the upper side 50 is defined. The section plane then defines a section of the lower side 48 or of the upper side 50. The section S passes through the point 100 and through a point at the intersection 104 of the root profile 44 with the lower side 48 or with the upper side 50. This point is also located at the intersection 104 of the section S with the root profile 44.

[0056] refer to Figure 4 , in the plane Pr, an angle 106 is defined between the root profile 44 and a tangent T to the section S, which is formed at the intersection 104 .

[0057] The angle 106 is less than or equal to 45 degrees.

[0058] In plane Pr, section S is situated between root profile 44 and tangent line T, ie each point of section S defines a radial axis on which it is situated between points of root profile 44 and tangent line T.

[0059] Any section S as defined above is located between the platform 22 and the tangent line T as described above. Since the tangent line is relatively close to the platform, the airfoil has a squat pyramidal shape relative to the value of the angle 106. This shape enables:

[0060] - acting on a first part of the cross flow upstream of the foil by directing the first part of the cross flow along the underside of the foil without causing separation, and

[0061] - acting on the second part of the cross flow, which is closest to the blade (the underside of which faces the wing) by confining it between the underside of the blade and the upper side of the wing.

[0062] Thus, the angular separation is greatly reduced.

[0063] Furthermore, the upstream portion of the wing has a relatively gentle slope relative to the platform 22. This reduces the risk of aerodynamic blockage, even when this upstream portion is located in the region of smallest cross section of the flow channel, which is most susceptible to blockage. This is particularly advantageous at high Mach numbers, where a blockage of a too large body would cause shock waves.

[0064] The slow variation of the height makes it possible not to cause the first part of the flow to separate: the gas is directed along the underside of the airfoil and is therefore directed to the trailing edge of the airfoil. The fact of directing this first part of the cross flow makes it possible to locally counteract the transverse pressure gradient, which limits the stall of the stator blades.

[0065] It should be noted that if the point 100 is located on the leading edge 30 or the trailing edge 32 of the airfoil 28, the point 100 forms part of both the lower side 48 and the upper side 50. The features described above are then verified for both the lower side and the upper side.

[0066] Figure 4 The case is shown where the section S includes points in common with the tip profile 46 , but this is not necessarily the case.

[0067] Wing height

[0068] As described above, the platform 22 defines an inner radial wall or an outer radial wall of the gas flow passage.

[0069] When the stator component 20 corresponds to a structure with a shroud, the stator component includes a second platform positioned radially facing the first platform 22, and the second platform defines an outer radial wall or an inner radial wall of the gas flow channel. Therefore, the gas flow channel passes radially between the first platform 22 and the second platform, and the flow channel is formed by Figure 4 The mark Hv in the figure extends radially at a certain flow channel height.

[0070] When the stator part 20 corresponds to a structure without a shroud, the stator part comprises only a single platform 22 defining the inner radial wall of the gas flow channel. The channel extends radially over a specific channel height defined by the height of the blades of the stator part 20, which project radially outwards from the platform 22.

[0071] Wing 28 in Figure 4 The fins shown extend radially over a height Ha.

[0072] According to a first optional variant of the most general embodiment, the ratio of the vane height to the flow channel height is greater than or equal to 0.01 and less than or equal to 0.25.

[0073] It may be advantageous for the aforementioned fins to remain relatively low in height so as not to block the flow path.

[0074] Circumferential wing-first blade distance

[0075] The first blade 24 and the second blade 26 are separated in the circumferential direction by a pitch 42. The pitch 42 separating the blades is the angle separating the radial direction of the first blade 24 from the radial direction of the second blade 26. The pitch is determined by the total number of blades having the same function and the same axial position as the blades 24 and 26, which are all located around the axis Δ of the turbine engine. The distance separating the first blade 24 from the second blade is given by this angle and the radius to the axis Δ, at which it is desired to estimate the distance.

[0076] It is also possible to define an angle that separates the first blade 24 from the airfoil 28 in the circumferential direction, or to define an angle that separates the radial direction of the first blade 24 from the radial direction of the airfoil 28 .

[0077] The angular spacing of the first blade 24 and the winglet 28 in the circumferential direction can be freely selected to be less than or equal to the angular pitch 42. In other words, the winglet can be located between the first blade and the second blade at any distance from the first blade.

[0078] Wing profile

[0079] As previously seen, the first blade 24 defines a blade chord 424 between its leading edge 224 and its trailing edge 324 .

[0080] The airfoil 28 can be modeled or represented as a stack of profiles in the radial direction between an airfoil root 44 and an airfoil tip 46. The airfoil root 44 is located on the platform 22 and corresponds to the intersection of the airfoil 28 with the platform 22. The airfoil tip is located at a distance from the platform 22 in the gas flow path. Each airfoil profile extends in a circumferential plane parallel to the axis Δ of the turbine engine, like an airfoil section formed in this circumferential plane, which has a constant radius or a constant distance from the axis Δ, and the circumferential plane can be regarded as a section plane.

[0081] refer to Figure 3 Each airfoil profile defines an airfoil chord 54 between the leading edge 30 of the airfoil and the trailing edge 32 of the airfoil. More precisely, the airfoil chord 54 is defined between a first point located at the intersection of the leading edge 30 with the cross-sectional plane on the one hand and a second point located at the intersection of the trailing edge 32 with the cross-sectional plane on the other hand. The airfoil chord represents the length of the line segment connecting the first point and the second point. The chord represents the line segment connecting the first point and the second point.

[0082] Each airfoil profile also defines a maximum thickness 52 between an underside 48 of the airfoil and an upper side 50 of the airfoil in a direction perpendicular to the chord line.

[0083] Each airfoil profile makes it possible to define a camber line, which is a curve equal to the average of the curves of the upper side 50 of the airfoil and the curves of the lower side 48 of the airfoil. More precisely, in a given airfoil profile, the camber line is formed by all the points equidistant between the intersection of the upper side 50 with the cross-sectional plane on the one hand and the intersection of the lower side 48 with the cross-sectional plane on the other hand. The camber line of the airfoil can be chosen to be close to the camber line of the first blade 24.

[0084] The airfoil has a continuous variation of the profile in the radial direction, ie all parameters of camber, chord and thickness vary continuously from the root profile to the tip profile.

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

[0086] -The tip profile chord is smaller than the root profile chord;

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

[0088] - The 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 enables blocking of cross flows over a large portion of the blade chord length.

[0090] A third optional variant of the most general embodiment and / or its variants comprises the following two features:

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

[0092] - The ratio of the maximum thickness of the tip profile to the chord of the tip profile is greater than or equal to 0.05 and less than or equal to 0.25.

[0093] The fins are then relatively thick, which facilitates the manufacture of the fins and the mechanical strength of the fins, particularly in the event of an ingestion of particles or foreign matter by the engine.

[0094] The "metal" corners of the fins and first blade

[0095] For the first blade 24 and the second blade 26, it can be defined that:

[0096] the metal blade rake angle, which is the angle between the tangent to the median camber line at the leading edge of the blade and the axis Δ of the turbine engine, the angle being oriented from the axis Δ towards the tangent, and

[0097] - The metal blade rear angle, which is the angle between the tangent to the camber line at the blade trailing edge and the axis Δ, the angle being oriented from the axis Δ towards the tangent.

[0098] For each wing profile, it is possible to define:

[0099] the metal wing rake angle, which is the angle between the tangent 61 of the median camber line at the leading edge of the wing and the axis Δ of the turbine engine, the angle being oriented from the axis Δ towards the tangent 61 , and

[0100] The metal wing trailing angle is the angle between the tangent 63 of the camber line at the trailing edge of the wing and the axis △, the angle being oriented from the axis △ towards the tangent 63 .

[0101] In the specific case of airfoils described in the present invention, it may be chosen to estimate the metal angle based on the local camber angle of the blade 24 or 26 .

[0102] To this end, an axial position and a radial position may be associated with each point of each camber line of each vane profile. A vane point is associated with a reference point of the camber line 43 of the profile of the first blade 24 having its axial position and radial position.

[0103] A local camber angle may be associated with the metal airfoil rake angle, the local camber angle being the angle oriented from the axis Δ toward a tangent line at an associated reference point to a mid-camber line of the profile of the first blade 24. The metal airfoil rake angle may be described with reference to the local camber angle as the difference between the metal angle and the local camber angle.

[0104] A local camber angle may be associated with the metal airfoil relief angle, the local camber angle being the angle oriented from the axis Δ toward a tangent line at an associated reference point to a mid-camber line of the profile of the first blade 24. The metal airfoil relief angle may be described with reference to the local camber angle as the difference between the metal angle and the local camber angle.

[0105] Typically, the airfoil described in the present invention has metal angles at the leading and trailing edges of the tip and root profiles that are close to the local camber angles. More precisely:

[0106] the absolute value of the angle formed by a first tangent line to the camber line of the profile of the airfoil tip at the leading edge 30 and a second tangent line to the camber line of the profile of the first blade at a reference point associated with the leading edge of the profile of the airfoil tip is less than or equal to 10 degrees;

[0107] the absolute value of the angle formed by a third tangent line to the camber line of the profile of the airfoil tip at the trailing edge 32 and by a fourth tangent line to the camber line of the profile of the first blade at a reference point associated with the trailing edge of the profile of the airfoil tip is less than or equal to 10 degrees;

[0108] the absolute value of the angle formed by a fifth tangent line to the camber line of the profile of the airfoil root at the front point 34 and by a sixth tangent line to the camber line of the profile of the first blade at a reference point associated with the front point 34 is less than or equal to 10 degrees; and

[0109] The absolute value of the angle formed by a seventh tangent to the camber line of the airfoil root profile at the rear point 36 and an eighth tangent to the camber line of the profile of the first blade at a reference point associated with the rear point 36 is less than or equal to 10 degrees.

[0110] The tip profile and the root profile then have a metal angle at the trailing edge, the value of which is close to the local camber angle of the blade. In this way, the airfoil directs the flow at the trailing edge, just like a blade.

[0111] Axial position of the wing relative to the first blade

[0112] In order to characterize the position of the airfoil in the direction of the axis Δ of the turbine engine, the axial coordinate of the mid-chord point of the airfoil profile is used.

[0113] The mid-chord point is located at an equal distance from the leading edge 30 of the airfoil and the trailing edge 32 of the airfoil.

[0114] The axial coordinate of the mid-chord point may be selected to be greater than or equal to the axial position of the leading edge 224 of the first blade and less than or equal to the sum of the axial position of the leading edge of the first blade and the blade chord 424 .

[0115] In particular, the mid-chord point of the airfoil tip profile may be used as the mid-chord point.

[0116] For its part, the blade root profile is axially positioned relative to the tip profile due to the stacking law of the blades, which gives the relative positioning of the blade tip profile relative to the blade root profile.

[0117] For this relative positioning, the mid-chord point is chosen as the positioning reference for each profile of the stack. Two axes are defined:

[0118] - a first axis "t" oriented in the chord direction of the root profile and oriented from upstream to downstream in the same direction as the axis Δ of the turbine engine;

[0119] A second axis “n” perpendicular to the first axis “t” and oriented towards the second blade 26 in the same direction as the circumferential axis θ of the first blade 24 .

[0120] Define two angles:

[0121] - the sweep angle, which is the angle between the axis "t" and the direction defined by the mid-chord point of the root profile and the mid-chord point of the tip profile;

[0122] - Angle of inclination, which is the angle between the axis "n" and the direction defined by the mid-chord point of the root profile and the mid-chord point of the tip profile.

[0123] In the particular case of the airfoil described here, as just defined, the sweep angle and the inclination angle take values ​​greater than or equal to -10 degrees and less than or equal to +10 degrees. These angles are small enough that the axial position of the airfoil is considered to be well described by the axial coordinate 78 of the mid-chord point 76 of the airfoil profile.

[0124] Active rarefaction effect at the trailing edge

[0125] A second embodiment according to the most general embodiment or its first variation includes the following features.

[0126] The airfoil 28 has a tangent to the trailing edge 32 at a rear point 36, which extends from the platform 22 into the flow path between the second blade 26 and the radial rear plane Pf. In other words, in a radial plane normal to the axis Δ of the turbine engine and passing through the rear point, the tangent to the trailing edge approaches the second blade 26 while separating from the platform 22 on the flow path side of the platform. Expressed in another different way, the trailing edge 32 is inclined towards the second blade 26 at the rear point 36. Furthermore, it should be noted that this tangent to the trailing edge can have a non-zero projection along the axis Δ of the turbine engine.

[0127] The inclination on the trailing edge side toward the second blade enables a rarefaction effect to be produced at the trailing edge, which more effectively blocks the cross flow. This makes it possible to avoid the cross flow originating from the second blade 26 from rising to the upper side of the fin 28 and crossing the fin to connect to the lower side of the first blade 24. Therefore, the fin "lies down" on the trailing edge side to have a wall inclined toward the lower side of the second blade 26. This is a positive rarefaction effect. Therefore, the cross flow is strongly driven back to the lower side of the second blade 26.

[0128] It should be noted that the airfoil may have a straight trailing edge 32 and thus merge with a tangent line to the trailing edge 32 at the rear point 36 .

[0129] Positive rarefaction effect downstream of the first asymmetric point on the trailing edge

[0130] More generally, the airfoil 28 may include an asymmetric point and the following features: at any current point of the trailing edge between the asymmetric point and the rear point 36, a tangent line of the trailing edge extending from the platform 22 into the flow channel is located 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 approaches the second blade 26 while separating from the current point toward the asymmetric point. Stated another way, the trailing edge 32 is inclined toward the second blade 26 between the rear point 36 and the asymmetric point.

[0131] The trailing edge upstream of the asymmetry point may be symmetrical as such, ie each tangent line to the trailing edge is contained in a radial plane including the axis of the turbine engine.

[0132] The leading edge may also be symmetrical, ie each tangent line to the leading edge is contained in a radial plane including the axis of the turbine engine.

[0133] In this case, the foil can be considered to be symmetrical upstream of the asymmetry point and asymmetrical downstream.

[0134] Positive rarefaction effect downstream of the second asymmetric point of the leading edge

[0135] When the asymmetric point is located on the tip profile, it may be a first asymmetric point and the airfoil may comprise a second asymmetric point, this time located on the leading edge.

[0136] At any current point of the leading edge between the asymmetric point and the tip profile, a tangent line of the leading edge extending from the leading edge into the flow channel is between the second blade 26 and a radial plane passing through the current point. In other words, in the radial plane passing through the current point, the trailing edge approaches the second blade 26 while separating from the current point toward the tip profile. Stated differently, the leading edge 30 is inclined toward the second blade 26 between the second asymmetric point and the tip profile.

[0137] The airfoil can then be considered symmetrical upstream of the second asymmetry point and asymmetrical downstream.Possibly, the second asymmetry point can be the front point 34, for which the entire airfoil can be considered asymmetric.

Claims

1. A stator component (20) of a turbine engine, the stator component comprising: - a platform (22) defining a wall of a gas flow channel, - blades (24, 26) extending radially from the platform (22) relative to a central axis (Δ) of the turbine engine, and an airfoil (28) extending radially into the flow channel from an airfoil root (44) located on the platform (22) to an airfoil tip (46), the airfoil comprising an underside (48) and an upper side (50) located radially between the airfoil root (44) and the airfoil tip (46), Each point (100) of the lower side (48) or the upper side (46) defines a radial axis (Ar) passing through the point (100), each plane (Pr) including the radial axis (Ar) defines a section (S) of the lower side (48) or the upper side (46), the 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 than or equal to 45 degrees, the section (S) being located between the root profile (44) and the tangent (T).

2. The stator component according to claim 1, wherein: The platform (22) is a first platform, and the component includes a second platform to define the flow channel between the first platform (22) and the second platform, the flow channel extends radially at a flow channel height (Hv), and the fins (28) extend radially at a fin height (Ha), and the ratio of the fin height (Ha) to the flow channel height (Hv) is 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) includes a leading edge (224, 226) and a trailing edge (324, 326) separated by a blade chord (424, 426), the airfoil (28) including a plurality of radially stacked profiles between the airfoil root (44) and the airfoil tip (46), each profile defining a chord (54) between a leading edge (30) of the airfoil (28) and a trailing edge (32) of the airfoil (28), The tip contour chord is smaller than the root contour chord; The ratio of the tip profile chord to the blade chord is 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 is greater than or equal to 0.3 and less than or equal to 1.

1.

4. The stator component according to claim 3, wherein: Each wing profile defines a maximum thickness (52) of the profile between the lower side (48) and the upper side (50) in a direction perpendicular to the chord line, The ratio of the maximum thickness of the root profile to the chord of the root profile is greater than or equal to 0.05 and less than or equal to 0.25; A ratio of a maximum thickness of the tip profile to a chord of the tip profile is greater than or equal to 0.05 and less than or equal to 0.

25.

5. A stator component according to any one of the preceding claims, wherein: The blade comprises a lower side (126) facing an upper side (50) of the airfoil (28), the airfoil comprising a trailing edge (32), the trailing edge comprising a rear point (36) located on the platform (22), a tangent line of the trailing edge (32) at the rear point (36) extending from the platform (22), while the tangent line of the trailing edge at the rear point is separated from the axis (△) between the blade (26) and a radial rear plane (Pf) passing through the axis (△) and the rear point (36).

6. The stator component according to claim 5, wherein: The trailing edge (32) includes an asymmetric point such that for any current point of the trailing edge (32) between the asymmetric point and the rear point (36), a tangent to the trailing edge (32) at the current point extends from the trailing edge (32), while the tangent to the trailing edge at the current point is separated from the axis (△) between the blade (26) and a radial rear plane passing through the axis (△) and the current point.

7. Turbine engine comprising a stator component (20) according to any one of claims 1 to 6.

8. Aircraft comprising a turbine engine according to claim 7.

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