Stator component having a fin in a turbine engine

Abstract

The invention relates to a stator component (20) comprising a platform (22), a first blade (24), a second blade (26) and a vane (28) having a leading edge (30) and a trailing edge (32), the vane extending opposite an upper side (124) of the first blade and an underside (126) of the second blade, the leading edge comprising a leading point (34) located on the platform, a tangent of the leading edge at the leading point extending between the first blade and a radial front plane (Pa), the trailing edge includes a trailing point (36) located on the platform, a tangent of the trailing edge at the trailing point extending between the second blade and a radial aft plane (Pf).

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 and aerodynamic blockages. Aerodynamic blockages are 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 blade leading edge, 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] - a first blade,

[0011] - a second blade, and

[0012] - a wing comprising a leading edge and a trailing edge,

[0013] The vane extends radially from the platform into the flow channel relative to the axis of the turbine engine, facing the upper side of the first blade and the lower side of the second blade,

[0014] the leading edge comprises a front point located on the platform, a tangent line to the leading edge at the front point extending from the platform into the flow passage between the first blade and a radial front plane passing through the axis and the front point,

[0015] The trailing edge includes a rear point on the platform, and a tangent line to the trailing edge at the rear point extends from the platform into the flow passage between the second blade and a radial rear plane passing through the axis and the rear point.

[0016] A stator component of this type is advantageously and optionally realized by the following different features, taken alone or in combination:

[0017] - 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 0.4;

[0018] - the first blade and the second blade are angularly spaced apart in the circumferential direction by an angular pitch, the angular spacing of the first blade from the airfoil in the circumferential direction being less than or equal to six tenths of the angular pitch;

[0019] - the first blade comprises a leading edge and a trailing edge separated by a blade chord, the airfoil comprises a plurality of profiles stacked in a radial direction between an airfoil root located on the platform and an airfoil tip, each profile defining a chord between the leading edge of the airfoil and the trailing edge of the airfoil, and each profile defining a maximum thickness of the profile between a lower side of the airfoil and an upper side of the airfoil in a direction perpendicular to the chord line, the absolute value of the ratio of the difference between the tip profile chord and the root profile chord to the blade chord is less than or equal to 0.05; the ratio of the maximum thickness of the root profile to the root profile chord is less than or equal to 0.05; the ratio of the maximum thickness of the tip profile to the root profile chord is less than or equal to 0.05; the ratio of the maximum thickness of the tip profile to the root profile chord is less than or equal to 0.05;

[0020] - the first blade comprises a leading edge and a trailing edge separated by a blade chord, the airfoil comprises a plurality of profiles stacked in a radial direction between the airfoil root and the airfoil tip, each profile defining an airfoil chord between the leading edge and the trailing edge, the ratio of the airfoil root chord to the blade chord being greater than or equal to 0.2 and less than or equal to 1.1; a first tangent from the axis to the mid-camber line of the airfoil root profile at a leading point defines a first orientation angle, a second tangent from the axis to the mid-camber line of the root profile of the first blade at a reference point of the mid-camber line of the root profile of the first blade defines a second orientation angle, the reference point having an axial position of the leading point, a difference between the first orientation angle and the second orientation angle being greater than or equal to -50 degrees and less than or equal to +10 degrees, the airfoil defining an axial coordinate of a mid-chord point of the airfoil profile, the mid-chord point being located at an equal distance from the leading edge of the airfoil and the trailing edge of the airfoil, the axial coordinate being greater than or equal to the axial position of the leading edge 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;

[0021] -The first blade comprises a leading edge and a trailing edge separated by a blade chord, the airfoil comprises a plurality of profiles stacked in a radial direction 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 and defining a median camber line, the ratio of the airfoil root profile chord to the blade chord being greater than or equal to 0.1 and less than or equal to 0.4; the absolute value of the angle between a first tangent to the median camber line of the airfoil root profile at the leading point and a second tangent to the median camber line of the root profile of the first blade at a reference point of the median camber line of the root profile of the first blade being less than or equal to 10 degrees, the reference point having an axial position of the leading point, the airfoil defining an axial coordinate of a median chord point of the airfoil profile, the median chord point being located at an equal distance from the leading edge of the airfoil and the trailing edge of the airfoil, the axial coordinate being greater than or equal to the difference between the axial position of the leading edge of the first blade and one tenth of the blade chord and less than or equal to the sum of the axial position of the leading edge of the first blade and half of the blade chord.

[0022] 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

[0023] 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:

[0024] - Figure 1 is a schematic diagram of a turbine engine;

[0025] - Figure 2 is a schematic diagram of a stator component according to a first embodiment;

[0026] - Figure 3 and Figure 4 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. DETAILED DESCRIPTION

[0027] Turbine engine

[0028] 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 Δ, and the turbojet engine 1 comprises, in order in the direction of gas flow in the turbine engine, 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).

[0029] The fan 2 and the low-pressure compressor 3 are driven to rotate by the low-pressure turbine 7 via a first transmission shaft 9 , while the high-pressure compressor 4 is driven to rotate by the high-pressure turbine 6 via a second transmission shaft 10 .

[0030] 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).

[0031] Stator components

[0032] 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.

[0033] The stator component 20 or the assembly 20 of stator components, if not of one-piece design, has at least two adjacent blades 24 , 26 and a platform 22 from which the blades 24 , 26 extend.

[0034] Here, the term "platform" denotes any element of the turbine engine on which 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, i.e. 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 Δ 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 Δ of the turbine engine.

[0035] 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 A radial axis r and a circumferential axis θ passing through a point 34 of the platform 22 are also shown. At each point in space, for example at a point 34 of the platform 22, a radial axis r can be defined, perpendicular to the axis Δ of the turbine engine and passing through the point and the axis Δ of the turbine engine. The radial axis is oriented positively in a direction separating it from the axis Δ of the turbine engine. A circumferential axis θ can also be defined, passing through the point and perpendicular to the radial axis r and the axis Δ of the turbine engine. The circumferential axis is oriented positively in a direction separating it from the axis Δ of the turbine engine.

[0036] exist Figure 2 In the example of FIG. 2 , the blades 24 , 26 extend radially from the platform 22 while separating the blades 24 , 26 from the axis of the turbine engine, but the invention is not limited to this case.

[0037] 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 sectors of the blades 24, 26 to be defined.

[0038] The direction of the axis Δ 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 Δ and the radial axis.

[0039] Each of the blades 24 and 26 has an underside 624 , 126 and an upper side 124 , 626 .

[0040] 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.

[0041] 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.

[0042] 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).

[0043] Wings

[0044] The stator component 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.

[0045] The fin 28 is located between the blades 24 and 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.

[0046] 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 .

[0047] 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 .

[0048] 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.

[0049] Any radial plane includes the axis Δ of the turbine engine.

[0050] 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 rear point 36 and includes the axis Δ of the turbine engine.

[0051] The most general embodiment of the present invention corresponds to the following two features.

[0052] The airfoil 28 has a tangent to the leading edge 30 at a front point 34, which extends from the platform 22 into the flow path between the first blade 24 and the radial front plane Pa. In other words, in a radial plane normal to the axis Δ of the turbine engine and passing through the front point, the tangent to the leading edge approaches the first blade 24 while separating from the platform 22 on the flow path side of the platform. Expressed in another different way, the leading edge 30 is inclined towards the first blade 24 at the front point 34. Furthermore, it should be noted that this tangent to the leading edge may have a non-zero projection on the axis Δ of the turbine engine.

[0053] 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 may have a non-zero projection on the axis Δ of the turbine engine.

[0054] The inclination towards the first blade on the leading edge side enables a kink effect to be produced at the leading edge, which better guides the cross flow.

[0055] The inclination towards the second blade on the trailing edge side 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 towards the upper side of the second blade 26. This is a positive rarefaction effect. Then, the cross flow is strongly driven back to the lower side of the second blade 26.

[0056] The specific form of the fins proposed here makes it possible to improve the operability of the compressor rectifier by blocking the cross flow, which limits the development of angular separation. This has the effect of limiting the losses and blockages associated with the angular separation phenomenon, which improves the stability of the rectifier, especially in cases where the rectifier is subjected to high angles of attack. The specificity of the fins obtained relies on the guidance of the cross flow and on the positive rarefaction effect that improves the way the cross flow is driven back.

[0057] It should be noted that the airfoil may have a straight leading edge 30 and thus merge with a tangent line to the leading edge 30 at the front point 34 .

[0058] 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 .

[0059] Wing height

[0060] As described above, the platform 22 defines an inner radial wall or an outer radial wall of the gas flow passage.

[0061] When the stator component 20 corresponds to a shrouded structure, the stator component includes a second platform positioned radially facing the first platform 22, the second platform defining an outer radial wall or an inner radial wall of the gas flow channel. Then, the gas flow channel radially passes between the first platform 22 and the second platform, and the channel extends radially over a certain channel height.

[0062] 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.

[0063] The vanes 28 extend radially over the vane height.

[0064] 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 0.4.

[0065] Circumferential wing-first blade distance

[0066] 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 26 is given by this angle and the radius to the axis Δ at which it is desired to estimate the distance.

[0067] 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 .

[0068] According to a second optional variant of the most general embodiment and its first variant, the angular spacing of the first blade 24 from the wing 28 in the chosen circumferential direction is chosen to be less than or equal to six tenths of the angular pitch 42 .

[0069] Wing profile

[0070] As previously seen, the first blade 24 defines a blade chord 424 between its leading edge 224 and its trailing edge 324 .

[0071] 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.

[0072] Figure 4 It is a schematic diagram of some parameters of the airfoil profile in the circumferential plane.

[0073] refer to Figure 3 and Figure 4 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.

[0074] 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.

[0075] Each airfoil profile makes it possible to define a camber line 64, 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 64 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.

[0076] 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.

[0077] The third optional variation of the most general embodiment and its first optional variation and second optional variation include the following three features:

[0078] - the ratio of the blade root profile chord to the blade chord 424 is greater than or equal to 0.2 and less than or equal to 1.1;

[0079] - the ratio of the maximum thickness of the root profile to the chord of the root profile is less than or equal to 0.05; and

[0080] - The ratio of the maximum thickness of the tip profile to the chord of the root profile is less than or equal to 0.05.

[0081] The "metal" corners of the fins and first blade

[0082] For the first blade 24 and the second blade 26, it can be defined that:

[0083] the metal blade rake angle, which is the angle between the tangent to the 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

[0084] - The metal blade clearance 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.

[0085] For each wing profile, refer to Figure 4 , we can define:

[0086] the metal wing rake angle, which is the angle between the tangent 61 of the 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

[0087] The metal foil trailing angle is the angle between the tangent 63 of the camber line at the foil trailing edge and the axis Δ, the angle being oriented from the axis Δ towards the tangent 63 .

[0088] 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 .

[0089] For this purpose, refer to Figure 4 An axial position and a radial position can be associated with each point of each camber line 64 of each airfoil profile. An airfoil point is associated with a reference point of the camber line 66 of the profile of the first blade 24, which reference point has an axial position and a radial position of the airfoil point.

[0090] The fin point may be located, for example, at the leading edge 30 at an axial position 68: the associated reference point is at Figure 4 It is marked as 72.

[0091] The flap point may be located, for example, at the trailing edge 32 at an axial position 70: the associated reference point is at Figure 4 It is marked as 74.

[0092] The local camber angle 56 may be associated with the metal airfoil rake angle, the local camber angle 56 being the angle oriented from the axis Δ toward a tangent line 57 of a camber line 66 of the profile of the first blade 24 at an associated reference point 72. The metal airfoil rake angle may be described with reference to the local camber angle as a difference 60 between the metal angle and the local camber angle 56.

[0093] The local camber angle 58 may be associated with the metal airfoil back angle, the local camber angle 58 being the angle oriented from the axis Δ toward a tangent line 59 of a camber line 66 of the profile of the first blade 24 at an associated reference point 74. The metal airfoil back angle may be described with reference to the local camber angle as a difference 62 between the metal angle and the local camber angle 58.

[0094] Typically, the airfoil described in the present invention has metal angles close to the local camber angles at the leading and trailing edges of the tip profile and at the trailing edge of the root profile. More precisely:

[0095] the absolute value of the angle 60 formed by a first tangent line 61 of the camber line 64 of the airfoil tip profile at the leading edge 30 and a second tangent line 57 of the camber line 66 of the profile of the first blade at a reference point 72 associated with the leading edge of the airfoil tip profile is less than or equal to 10 degrees;

[0096] the absolute value of the angle 62 formed by the third tangent line 63 of the camber line 64 of the airfoil tip profile at the trailing edge 32 and the fourth tangent line 59 of the camber line 66 of the profile of the first blade at the reference point 74 associated with the trailing edge of the airfoil tip profile is less than or equal to 10 degrees; and

[0097] The absolute value of the angle formed by a fifth tangent to the camber line of the airfoil root profile at the rear point 36 and a sixth tangent to the camber line of the first blade profile at a reference point associated with the rear point 36 is less than or equal to 10 degrees.

[0098] 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.

[0099] Axial position of the wing relative to the first blade

[0100] To characterize the position of the wing in the direction of the axis Δ of the turbine engine, refer to Figure 4 The axial coordinate 78 of the mid-chord point 76 of the airfoil profile is used.

[0101] The mid-chord point 76 is located at an equal distance from the leading edge 30 of the airfoil and the trailing edge 32 of the airfoil.

[0102] In particular, the mid-chord point of the airfoil tip profile may be used.

[0103] 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.

[0104] For this relative positioning, the mid-chord point is chosen as the positioning reference for each profile of the stack. Two axes are defined:

[0105] 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;

[0106] 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 .

[0107] Define two angles:

[0108] - 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;

[0109] - 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.

[0110] 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.

[0111] "Long Fence" Example

[0112] The second embodiment of the present invention is a specific case of the most general embodiment and its variations, and the second embodiment includes the following features:

[0113] - the ratio of the blade root profile chord to the blade chord 424 is greater than or equal to 0.2 and less than or equal to 1.1;

[0114] - the difference between the metal airfoil rake angle at the airfoil root profile and the local camber angle of the root profile of the first blade is greater than or equal to -50 degrees and less than or equal to +10 degrees;

[0115] an axial coordinate 78 of a mid-chord point 76 of the airfoil profile, the mid-chord point being located at an equal distance from the leading edge 30 of the airfoil and from the trailing edge 32 of the airfoil, the axial coordinate 78 being greater than or equal to the axial position 80 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.

[0116] With regard to the metal airfoil angle of attack at the airfoil root profile, as previously described, this is a first orientation angle defined by a first tangent line from the axis Δ to the camber line of the airfoil root profile at the front point 34 of the leading edge 30. The local camber angle of the root profile of the first blade is a second orientation angle defined by a second tangent line from the axis Δ to the camber line of the root profile of the first blade at a reference point having an axial position of the front point 34.

[0117] In this second embodiment, the fin is relatively long, which enables the physical effect of limiting the cross flow between the upper side 50 of the fin 28 and the underside 126 of the second blade 26 to occur.

[0118] The near-wall flow is strongly kinked, i.e. the flow has a gradient that increases as it separates from the platform. In fact, the gas does not flow in the same way near the platform and far from it. Near the platform, a cross flow is dominant, so that the fluid flows mainly in a direction from the underside of the second blade 26 towards the upper side of the first blade 24; this cross flow is also slow, i.e. it corresponds to a small amount of movement. Far from the platform, the flow mainly proceeds from upstream to downstream, while being guided by the blades, and this flow is faster.

[0119] In this second mode, the leading edge of the wing has a metal kink. Thus, the metal angle of the wing root profile can be selected to be smaller than the metal angle of the wing tip to adapt to the slow airflow near the platform, and the metal angle of the wing tip profile can be adapted to the tip flow in the acoustic flow region.

[0120] "Upstream sharp" embodiment

[0121] The third embodiment of the present invention is a specific case of the most general embodiment and its variations, and the third embodiment includes the following features:

[0122] - the ratio of the blade root profile chord to the blade chord 424 is greater than or equal to 0.1 and less than or equal to 0.4;

[0123] - the absolute value of the difference between the metal airfoil rake angle on the airfoil root profile and the local camber angle of the root profile of the first blade is less than or equal to +10 degrees;

[0124] an axial coordinate 78 of a mid-chord point 76 of the airfoil profile, the mid-chord point 76 being located at an equal distance from the leading edge 30 of the airfoil and from the trailing edge 32 of the airfoil, the axial coordinate 78 being greater than or equal to the difference between the axial position 80 of the leading edge 224 of the first blade 24 and one-tenth of the blade chord 424 and less than or equal to the sum of the axial position 80 of the leading edge of the first blade 24 and half of the blade chord 424.

[0125] With regard to the metal airfoil rake angle on the airfoil root profile, as previously described, this is a first orientation angle defined by a first tangent line from the axis Δ to the camber line of the airfoil root profile at the front point 34 of the leading edge 30. The local camber angle of the root profile of the first blade is a second orientation angle defined by a second tangent line from the axis Δ to the camber line of the root profile of the first blade at a reference point having an axial position of the front point 34.

[0126] In this third embodiment, the vanes are relatively short and are placed on the upstream portion of the inter-vane passage defined between the vanes. This will enable the generation of corner vortices at the vane tips 28, which can excite the flow near the platform 22. This makes it possible to straighten the cross flow and ultimately limit the corner separation.

[0127] It should be noted that a larger metal front angle can also make the corner vortex more intense, and is conducive to mixing and accelerating the flow on the upper side of the wing. Thus, the corner vortex will mix the acoustic flow and the accelerated flow in the boundary layer of the hub.

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, - a first blade (24), - a second blade (26), and - an airfoil (28) comprising a leading edge (30) and a trailing edge (32), The vane (28) extends radially from the platform (22) into the flow channel relative to the axis (Δ) of the turbine engine, facing the upper side (124) of the first blade and the lower side (126) of the second blade (26), The leading edge (30) includes a front point (34) located on the platform (22), a tangent line to the leading edge (30) at the front point (34) extending from the platform (22) into the flow channel between the first blade (24) and a radial front plane (Pa), the radial front plane passing through the axis (Δ) and the front point (34), The trailing edge (32) includes a rear point (36) located on the platform (22), and a tangent line to the trailing edge (32) at the rear point (36) extends from the platform (22) into the flow channel between the second blade (26) and a radial rear plane (Pf), the radial rear plane passing through the axis (Δ) and the rear point (36).

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 the flow channel height, the fins (28) extend radially at the 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.

4.

3. The stator component according to claim 1 or 2, wherein: The first blade (24) and the second blade (26) are angularly spaced apart by an angular pitch (42) in a circumferential direction, and the angular spacing between the first blade (24) and the airfoil (28) in the circumferential direction is less than or equal to six tenths of the angular pitch (42).

4. A stator component according to any one of the preceding claims, wherein: The first blade (24) includes a leading edge (224) and a trailing edge (324) separated by a blade chord (424), the airfoil (28) including a plurality of profiles stacked in a radial direction between an airfoil root (44) and an airfoil tip (46) located on the platform (22), each profile defining a chord (54) between a leading edge (30) of the airfoil (28) and a trailing edge (32) of the airfoil (28), and each profile defining a maximum thickness (52) of the profile between a lower side (48) of the airfoil and an upper side (50) of the airfoil in a direction perpendicular to the chord line, The absolute value of the ratio of the difference between the tip profile chord and the root profile chord to the blade chord is less than or equal to 0.05; The ratio of the maximum thickness of the root profile to the chord of the root profile is less than or equal to 0.05; and The ratio of the maximum thickness of the tip profile to the chord of the root profile is less than or equal to 0.

05.

5. A stator component according to any one of the preceding claims, wherein: The first blade (24) includes a leading edge (224) and a trailing edge (324) separated by a blade chord (424), The airfoil (28) includes a plurality of profiles stacked in a radial direction between an airfoil root (44) and an airfoil tip (46), each profile defining an airfoil chord (54) between the leading edge (30) and the trailing edge (32), and a ratio of the airfoil root profile chord to the blade chord (424) is greater than or equal to 0.2 and less than or equal to 1.1; A first tangent line from said axis (Δ) to the camber line of the blade root profile at said front point (34) defines a first orientation angle, a second tangent line from the axis (Δ) to the camber line of the root profile of the first blade at a reference point of the camber line of the root profile of the first blade, the reference point having the axial position of the front point (34), defines a second orientation angle, The difference between the first orientation angle and the second orientation angle is greater than or equal to -50 degrees and less than or equal to +10 degrees, The airfoil (28) defines an axial coordinate (78) of a mid-chord point (76) of the airfoil profile, the mid-chord point (76) being located equidistant from a leading edge (30) of the airfoil and a trailing edge (32) of the airfoil, the axial coordinate (78) being greater than or equal to an axial position (80) of a 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).

6. A stator component according to any one of the preceding claims, wherein: The first blade (24) includes a leading edge (224) and a trailing edge (324) separated by a blade chord (424), The airfoil (28) comprises a plurality of profiles stacked in a radial direction between an airfoil root (44) and an 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) and defining a camber line (64), a ratio of the airfoil root profile chord to the blade chord (424) being greater than or equal to 0.1 and less than or equal to 0.4, the absolute value of the angle between a first tangent line to the camber line of the root profile of the airfoil at the front point (34) and a second tangent line to the camber line of the root profile of the first blade at a reference point to the camber line of the root profile of the first blade is less than or equal to 10 degrees, the reference point having the axial position of the front point (34), The airfoil (28) defines an axial coordinate (78) of a mid-chord point (76) of the airfoil profile, the mid-chord point (76) being located at an equal distance from the leading edge (30) of the airfoil and the trailing edge (32) of the airfoil, the axial coordinate (78) being greater than or equal to a difference between an axial position (80) of the leading edge (224) of the first blade and one-tenth of the blade chord (424) and less than or equal to a sum of the axial position (80) of the leading edge (224) of the first blade and half of the blade chord (424).

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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