Stator component of a turbomachine comprising a blade and a vane defining between them a surface decreasing from upstream to downstream according to the direction of gas flow.

Abstract

A stator component of a turbomachine comprising a platform, a blade (24), and a vane (28), the blade (24) and the vane (28) extending from the platform, the platform, an upper surface (25) of the blade (24), and the vane (28) defining between them a gas flow channel, the channel having a cross-section, in a plane normal to an axis of the turbomachine, with an area that decreases continuously from upstream to downstream with respect to a general direction of gas flow through the turbomachine. Figure for the abbreviation: Fig. 4

Description

Title of the invention: Stator part of a turbomachine comprising a blade and a vane defining between them a surface decreasing from upstream to downstream according to the direction of gas flow. 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, in particular, outlet guide vanes (also known as "Outlet Guide Vane" or "OGV"), inlet guide vanes (also known as "Inlet Guide Vane" or "IGV"), and variable stator vanes (also known as "Variable Stator Vane" or "VSV"). Typically, the stator blades of an aircraft gas turbine engine each have two platforms (inner and outer) which are attached to the blade.These rectifier blades form rows of fixed blades 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 pressure losses as well as aerodynamic stall. The latter is problematic in terms of operability. For high incidences of the flow arriving at the straightener, that is to say when the direction of gas flow upstream of the straightener 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] The reduction in compressor performance and operability is all the greater when the s / c ratio between the circumferential distance separating two blades 5 and the chord of a blade c is large. For lightweight motors with a reduced number of blades and made more axially compact by shortened chords, this s / c ratio is even greater, making the effects all the more problematic.

[0007] 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

[0008] An object of the invention is to provide a stator part of a turbomachine whose geometry improves fluid flow compared to the prior art.

[0009] The goal is achieved within the framework of the present invention by means of a stator part of a turbomachine comprising a platform, a blade and a vane, the blade and the vane extending from the platform, the platform an extrados of the blade and the vane defining between them a gas flow channel, the channel having a cross-section in a plane normal to an axis of the turbomachine, having an area which decreases continuously from upstream to downstream with reference to a general direction of gas flow through the turbomachine.

[0010] On the one hand, the proposed winglet limits the flow of passage directed towards the upper surface (extrados). On the other hand, the winglet defines a channel between itself and the upper surface through which the fluid flows. This channel has a cross-section that decreases downstream, so that the cross-section seen by the fluid through this channel narrows. By conserving flow rate in the channel, the fluid flow accelerates downstream in the axial direction. There is therefore an acceleration of the flow on the upper surface, which reduces the thickness of the boundary layer on the upper surface of the blade as well as on the platform. This also reduces the area of ​​low momentum associated with the wedge separation responsible for the aerodynamic stall. This is true over a wide range of angles of attack, and particularly at high angles of attack.

[0011] Such a stator component is advantageously and optionally complemented by the following various features, taken alone or in combination:

[0012] - the upper surface and the winglet are separated in each normal plane by a distance of crescent from upstream to downstream;

[0013] - the fin has in each normal plane an edge contiguous to the channel and presenting an inclination relative to the platform which decreases from upstream to downstream;

[0014] - the fin has a radial dimension which decreases from upstream to downstream;

[0015] - the fin comprises an upstream end, the blade has a point of camber maximum and an axial chord defined as a length of a projection of a chord of the blade along the axis, the upstream end being located axially upstream of the camber point at a distance less than or equal to 30% of the axial chord and downstream of the camber point at a distance less than or equal to 20% of the axial chord; and

[0016] - the blade is a first blade, the stator piece comprising a second blade in looking at the first blade, the fin being located between the first blade and the second blade, each blade comprising a leading edge and a tangent to a line of camber from the blade to the leading edge, the tangents being parallel, for each tangent the upstream end of the fin being located in a plane normal to the tangents at a distance from the tangent greater than or equal to 5% of the axial chord.

[0017] 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

[0018] 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:

[0019] [Fig.1] [Fig.1] is a schematic representation of a turbomachine;

[0020] [Fig.2] [Fig.2] is a schematic representation of a stator part according to a first embodiment;

[0021] [Fig.3] [Fig.3] is a schematic cross-sectional view in a plane perpendicular to the axis of the turbomachine of a stator component according to a second embodiment; and

[0022] [Fig.4] [Fig.4] is a schematic representation of a stator part according to the first embodiment in a blade-to-blade plane. DETAILED DESCRIPTION OF THE INVENTION

[0023] With reference to [Fig. 1], a turbomachine is schematically represented, more specifically a twin-flow axial turbojet 1. The illustrated turbojet 1 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.

[0024] 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.

[0025] 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).

[0026] 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].

[0027] The stator piece 20, or the assembly 20 of stator pieces if it is not a single piece, has at least two consecutive blades 24, 26 and a platform 22 from which the blades 24, 26 extend.

[0028] Figure 2 is a schematic cross-sectional representation of the stator component 20 in a plane normal to the axis A of the turbomachine, i.e., a schematic cross-sectional view in a plane perpendicular to the axis of the turbomachine. The axis A is perpendicular to the plane of Figure 2 and directed towards the viewer of Figure 2. The term "platform" here refers to any element of the turbomachine from which blades 24, 26 can be mounted. 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 blades 24, 26 extend from the platform 22 radially outwards or radially inwards. The platform 22 has an inner or outer wall against which the air flows. The stator piece 20 includes a wall 23 located opposite the platform 22.

[0029] The blade 24 has an extrados 25 which faces an intrados of the blade 26. In operation, the air flows through the stator piece in a channel defined by the platform 22, the blades 24 and 26 and the wall 23. The flow is in the direction of the axis A of the turbomachine and from upstream to downstream as in the direction of the axis A directed towards the reader of [Fig.2].

[0030] Fig. 4 is a schematic representation of the stator part 20 in a plane circumferential, that is, at a constant distance from the A axis of the turbomachine. The direction of the A axis is given in [Fig. 4] by the x-axis, whose orientation is the direction of gas flow. The radial axis r is perpendicular to the plane of [Fig. 4] and directed towards the reader of [Fig. 4]. The 0 axis corresponds to the circumferential direction perpendicular- dicular simultaneously to the A axis and the radial axis.

[0031] The blades 24 and 26 each have an intrados and an extrados. The blades 24 and 26 each comprise a leading edge 52, 39 on the upstream side and a trailing edge on the downstream side. The blades define a chord 36, which is the segment connecting the leading edge and the trailing edge. The chord 36 projected onto the direction of the turbomachine axis defines an axial chord 37.

[0032] Each blade has a camber line 41, 43 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 equidistant from the upper and lower surfaces. The distance from a particular point to the upper (or lower) surface is defined here as the minimum distance between the particular point and a point on the upper (or lower) surface.

[0033] On each camber line 41, 43, a maximum camber point is defined (reference 35 on the blade 24). At this point, the length of a segment perpendicular to the chord line and connecting a point on the chord line and a point on the camber line is maximum.

[0034] The coordinate of the point of maximum camber along the x-axis is denoted xO on [Fig.4],

[0035] We also define:

[0036] - a coordinate xl less than the coordinate xO, the length xO-xl being 30% of the axial chord 37;

[0037] - a coordinate x2 greater than the coordinate x0, the length x2-x0 being 20% ​​of the axial chord 37.

[0038] The stator piece 20 also includes a fin 28 extending from the platform in the same direction and with the same sense of extension as the blades 24, 26. The fin is located between the blades 24 and 26. The fin extends over a radial dimension 31 that is less than the height of the blades. In other words, the fin does not extend from the platform 22 to the wall 23 over the entire height of the vein separating the platform 22 from the wall 23. The radial dimension 31 of the fin 28 varies between 1% and 40% of this vein height. The radial dimension 31 depends on the size of an upstream boundary layer.

[0039] The blade 28 extends along the axis A of the turbomachine from an upstream end 33 to a downstream end, as illustrated in [Fig.4].

[0040] The fin 28 has a side 32 which is opposite the upper surface 25 of the blade 24. The intersection of the side 32 and a plane normal to the axis A of the turbomachine is an edge 29. This edge can be straight or curved.

[0041] The flank 32 of the fin 28 may have a straight edge 29 which allows an inclination 52 to be defined with the platform 22, as shown in [Fig. 3]. This in The inclination is equal to 90° when the edge makes a right angle with the platform. When the platform 22 comprises a cylindrical surface at a constant radial distance to the axis A of the turbomachine, an inclination of 90° of the edge 29 corresponds to an edge that extends along the radial direction.

[0042] The platform 22, the upper surface 25 of the blade 24, and the winglet 28 define a gas flow channel 30. The channel 30 extends from the upper surface 25 to the flank 32 of the winglet 28 in the circumferential direction 0. The edge 29 of the flank 32 of the winglet 28 is contiguous with the channel 30. The channel 30 extends radially from the platform 22 towards the wall 23 over a length equal to the radial dimension 31 of the winglet 28.

[0043] The channel 30 follows the shape of the platform 22, the extrados 25 and the side of the fin 28. The channel 30 does not extend beyond the radial dimension 31 of the fin 28.

[0044] The stator piece is configured so that the channel 30 has a section, in a plane normal to the axis A of the turbomachine, whose area decreases continuously from upstream to downstream.

[0045] In other words, if we choose two planes normal to the axis A of the turbomachine, the two planes comprising a downstream plane and an upstream plane upstream of the downstream plane, the section of the channel 30 in the upstream plane is always greater than or equal to the section of the channel 30 in the downstream plane.

[0046] The continuous decrease of Faire of the section can be obtained in different embodiments which can possibly be combined with each other.

[0047] In a first embodiment, the upper surface 25 and the flank 32 of the fin 28 are separated in each normal plane by a distance that decreases from upstream to downstream. In this case, the radial dimension 31 of the fin can be kept constant and the shape of the edge 39 identical in the different normal planes.

[0048] In a second embodiment, the inclination 52 of the edge 29 with respect to the platform 22 decreases from upstream to downstream. The flank of the fin 28 is then oblique and the angle of the flank with respect to the platform 22 decreases downstream.

[0049] In a third embodiment, the radial dimension 31 of the fin decreases from upstream to downstream. In this case, the distance separating the upper surface 25 and the flank 32 can be kept constant, and the shape of the edge 39 can be identical in the different normal planes. The second and third embodiments can advantageously be combined: the fin decreases in radial dimension downstream, and the inclination of the edge decreases downstream.

[0050] Thanks to the reduction of the cross-section of the channel 30 from upstream to downstream, the area where the gas flow exhibited a low quantity of motion in the prior art is accelerated in the stator piece presented here.

[0051] Furthermore, as the size of the channel decreases, the channel-induced blockage also decreases.

[0052] The boundary layer remains attached longer to the upper surface 25 of the blade 24, which improves its righting efficiency. This effect is significant at high angles of attack, where wedge separation is usually substantial. By limiting stator separation and losses, the flow is better deflected. This helps to minimize the difference between the gas flow and the profile of the stator-exiting stator blades.

[0053] The efficiency of the propulsion assembly formed by the rotor and stator is improved. This effect is visible even at low angles of attack, close to the maximum efficiency point for heavily loaded stators - i.e. for stator rectifiers with a high s / c ratio.

[0054] This allows us to have a more robust stator, which can increase the compressor's operability margin.

[0055] Optionally to the modes previously presented, the upstream end 33 of the fin 28 can be placed in specific areas according to two conditions.

[0056] A first condition is that the upstream end 33 can be located axially, that is to say in the direction of the axis A of the turbomachine, upstream of the camber point 35 at a distance less than or equal to 30% of the axial chord 37 and downstream of the camber point 35 at a distance less than or equal to 20% of the axial chord 37.

[0057] In other words, the upstream end 33 is located between the lines with equations x=xl and x=x2, with the coordinates xl and x2 introduced previously. The lines x=xl and x=x2 are represented by dashed lines in [Fig.4].

[0058] In addition to this first condition, the upstream end 33 can be located, according to a second condition, at particular distances from tangents of the camber lines 41, 43 of the blades 24, 26. More precisely, the tangent T1 to the camber line 41 of the blade 26 at its leading edge 52 is defined, and the tangent T2 to the camber line 43 of the blade 24 at its leading edge 39.

[0059] These two tangents T1 and T2 are parallel and a plane can be defined that is simultaneously normal to the two tangents T1, T2. According to the second condition, the upstream end 33 is located at a distance from each of the tangents greater than or equal to 5% of the axial chord 37.

[0060] Figure 4 illustrates a distance d equal to 5% of the axial chord 37. The lines K1, K2 are parallel to the tangents T1, T2. The line K1 is at a distance d from the tangent T1, the line K1 being closer to the blade 24. The line K2 is at a distance d from the tangent T2, the line K2 being closer to the blade 26.

[0061] The lines Kl and K2 define a zone between them and if the upstream end 33 of the fin 28 is in this zone, the second condition is verified.

[0062] Furthermore, the axial position of the downstream end of the fin can be the axial position from the trailing edge of blades 24, 26.

[0063] The two conditions make it possible to optimize the position of the fin according to the maximum curvature zone of the blades and to optimize the separation control effect on the downstream part of blade 24, while reducing the disadvantages of adding a fin.

Claims

Demands

1. Stator part (20) of a turbomachine comprising: - a platform (22), - a blade (24), and - a vane (28), the blade (24) and the vane (28) extending from the platform (22), the platform (22), an extrados (25) of the blade (24) and the vane (28) defining between them a gas flow channel (30), the channel (30) having a cross-section, in a plane normal to an axis (A) of the turbomachine, having an area which decreases continuously from upstream to downstream with reference to a general direction of gas flow through the turbomachine.

2. Stator piece according to claim 1, wherein the extrados (25) and the fin (28) are separated in each normal plane by a decreasing distance from upstream to downstream.

3. Stator piece according to claim 1 or 2, wherein the fin (28) has in each normal plane an edge (29) contiguous to the channel (30) and having an inclination (33) with respect to the platform (22) which decreases from upstream to downstream.

4. Stator part according to any one of the preceding claims, wherein the fin (28) has a radial dimension (31) which decreases from upstream to downstream.

5. Stator part according to any one of the preceding claims, wherein: - the fin (28) comprises an upstream end (33), - the blade (24) has a maximum camber point (35) and an axial chord (37) defined as a length of a projection of a chord (36) of the blade (24) along the axis, - the upstream end (33) is located axially upstream of the camber point (35) at a distance less than or equal to 30% of the axial chord (37) or downstream of the camber point (35) at a distance less than or equal to 20% of the axial chord (37).

6. A stator piece according to claim 5, wherein the blade (24) is a first blade, the stator piece comprising a second blade (26) opposite the first blade (24), the fin (28) being located between the first blade (24) and the second blade (26), each blade (24, 26) comprising a leading edge (52, 39) and a tangent (T1, T2) to a camber line (41, 43) of the blade (24, 26) to the leading edge (52, 39), the tangents (Tl, T2) being parallel, for each tangent (Tl, T2) the upstream end (33) of the fin (28) being located in a plane normal to the tangents (Tl, T2) at a distance from the tangent greater than or equal to 5% of the axial chord (37).

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

8. Aircraft comprising a turbomachine according to the preceding claim.

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