Stator component with blade and fin arrangement in turbomachine

Abstract

The invention relates to a stator component (20) comprising: a platform (22); a blade (24, 26) having a maximum radian (526); a fin (28) extending from a root (44) to a head (46), the head (46) extending from a leading edge (31) to a trailing edge (37), the fin (28) and the blades (24, 26) being designed such that:-a maximum radian (526) is located between the leading edge (31) and the trailing edge (37),-a ratio of a distance between the maximum radian (526) and the leading edge (31) to a chord length (54) is between 5% and 95%, and-the fin head defining a recess (45), the recess extends on both sides of the maximum radian (526) with a recess length (47) that is less than or equal to 20% of the chord length (54).

Description

Technical Field

[0001] The invention relates to a stator component of a turbomachine comprising blades situated downstream of a compressor, such as a flow straightener, in particular a straightener with a fixed pitch. Background Art

[0002] In aircraft turbines, particularly in passenger aircraft, the air propelled by the fan and the combustion gases leaving the turbine through the exhaust nozzle apply reverse thrust to the turbine, and reverse thrust is applied to the aircraft by the turbine. The flow of gas through the turbine is affected by rotating blades and fixed vanes. Fixed or stator vanes include outlet guide vanes (also referred to as OGVs), inlet guide vanes (also referred to as IGVs), and vanes with variable pitch (also referred to as variable stator vanes or VSVs). The rectifier vanes of a gas turbine aircraft engine can each have two platforms (inner and outer platforms) attached to the vanes. There is also an open fan architecture that includes rectifier vanes with a single inner platform. In all cases, these rectifier vanes form rows of fixed vanes that guide the gas flow through the engine at an appropriate speed and angle.

[0003] In a flow straightener comprising a plurality of stationary blades, the flow of gas generally occurs between the blades in an upstream to downstream direction.

[0004] However, it is known that secondary aerodynamic flows may occur in the region of the blade root and blade tip.

[0005] For each pair of blades facing each other, the pressure gradient between the pressure side (inner camber side) of the first blade and the suction side (outer camber side) of the second blade creates a cross flow that transports gas to the outer camber side.

[0006] At the blade tip, i.e., at the junction between the blade and the hub or between the blade and the casing, corner separation and corner vortices can occur. This separation creates pressure losses and aerodynamic blockage. This aerodynamic blockage is problematic in terms of operability.

[0007] Fins can be placed on the hub or casing between two blades facing each other to reduce cross flow. The fins can improve the operability of the compressor stator by limiting the cross flow between the vanes.

[0008] The geometry of a cascade of blades (i.e., a group of blades distributed circularly about an axis) can be optimized for high-speed fluid flow (particularly transonic flow near the speed of sound). Adding fins to such an optimized cascade can significantly degrade the operation of these optimized blades.

[0009] Therefore, especially for transonic fluid flows, a new geometry is needed to correct these problems and improve the performance in terms of device efficiency. Summary of the Invention

[0010] An object of the present invention is to provide a novel blade geometry which will improve the performance of the device in terms of efficiency, in particular for transonic flows.

[0011] This object is achieved in the context of the present invention by a stator component for a turbomachine, comprising:

[0012] - a platform defining a wall of the gas flow duct,

[0013] - blades extending radially from the platform relative to the axis of the turbine, the blades having a maximum camber, and

[0014] - a fin extending radially in the duct from a root located on the platform to a tip, the tip extending axially along the axis from the leading edge to the trailing edge and having a chord length connecting the leading edge and the trailing edge in a straight line,

[0015] The fins and blades are designed so that:

[0016] - the maximum arc is located axially between the leading and trailing edges,

[0017] - the ratio of the distance measured axially between the maximum arc and the leading edge to the chord length is less than or equal to 95% and greater than or equal to 5%, and

[0018] The fin height defined between the root and the tip varies along the axis to define a recess in the fin extending axially from one side of the maximum arc to the other side with a recess length less than or equal to 20% of the chord length.

[0019] Advantageously and optionally, this component is supplemented by the following features, alone or in combination:

[0020] - the fin height has a maximum height and a minimum height in the recess, and the ratio of the minimum height to the maximum height is less than or equal to 20%;

[0021] - the ratio of the minimum height to the maximum height is less than or equal to 10%, preferably equal to 5%;

[0022] - the fin heights in the recess are symmetrical relative to the axial position of the maximum arc; and

[0023] - the height of the fins in the recess as a function of the coordinate along the axis, the height of the fins in the recess having a first rate of change upstream of the axial position of maximum arc and a second rate of change downstream of the axial position, the terms "upstream" and "downstream" referring to the main flow direction of the gas in the stator part, the absolute value of the first rate of change being greater than the absolute value of the second rate of change.

[0024] The invention also relates to a turbomachine comprising a stator component as described above and to an aircraft comprising such a turbomachine. BRIEF DESCRIPTION OF THE DRAWINGS

[0025] Other features and advantages of the present invention will become apparent from the following description, which is given by way of illustration only and not limitation, and which should be read with reference to the accompanying drawings, in which:

[0026] Figure 1 is a schematic diagram of a turbine;

[0027] Figure 2 、 Figure 3 and Figure 4 is a schematic diagram of a stator component according to various embodiments. DETAILED DESCRIPTION

[0028] turbines

[0029] Figure 1 A schematic diagram of a turbomachine, more specifically an axial-flow turbofan engine 1, is shown. Other types of turbomachines, in particular open-fan turbomachines, can be improved by the teachings presented herein. The illustrated turbojet engine 1 extends along an axis Δ and comprises, in order, in the main flow direction of the gas in the turbomachine, a fan 2, a compression section that may include a low-pressure compressor 3 and a high-pressure compressor 4, a combustion chamber 5, a turbine section that may include a high-pressure turbine 6 and a low-pressure turbine 7, and an exhaust nozzle.

[0030] 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 .

[0031] In operation, the air flow compressed by the low-pressure compressor 3 and the high-pressure compressor 4 is fed to the combustion chamber 5, and the expansion of the combustion gases in the combustion chamber 5 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 and 7 exert a reaction thrust on the turbojet engine 1 and, through this reaction thrust, on a vehicle or machine, such as an aircraft (not shown).

[0032] stator components

[0033] Downstream of the fan or compression stage, the turbine may include a straightening vane stage. Such a straightening vane stage may include a straightening vane stage as shown in FIG. Figure 2 The stator component 20 is shown.

[0034] The stator component may be one of the stators included in a low-pressure axial compressor or a high-pressure axial compressor. The stator component may also include a cascade of outlet guide vanes, such as in a secondary rectifier on a fan module.

[0035] The stator component 20 or the stator component assembly 20 (if the stator component 20 and the stator component assembly 20 are not integral) has at least one blade 24 , 26 and a platform 22 from which the blade 24 , 26 extends. For example, the stator component may include two adjacent blades 24 , 26 extending from the platform 22 .

[0036] The term "platform" is used herein to refer to any component of the turbine that can be mounted with blades 24, 26. In particular, the platform can be a hub or a housing around the axis of the turbine. The platform can also be sectored, that is, the platform comprises an assembly of components that are angularly distributed around the axis in ideally equal angular sectors. The platform can have a cylindrical surface located at a constant radial distance Δ from the axis of the turbine. The platform 22 has an inner wall or, alternatively, an outer wall against which the air flows, that is, the platform 22 defines the wall of the gas flow duct. The blades 24, 26 extend from the platform 22 into the flow radially outwards away from the turbine axis Δ or radially inwards towards the turbine axis Δ.

[0037] exist Figure 2 In FIG, the axis Δ of the turbine is shown to be oriented in the direction of the main gas flow in the turbine. Figure 2 Also shown is a radial axis r, which is perpendicular to and passes through the axis Δ of the turbine. The radial axis is oriented in a direction away from the axis Δ of the turbine. Figure 2 Also shown is a circumferential axis θ perpendicular to the radial axis r and the axis Δ of the turbine.

[0038] Figure 2 is a schematic diagram of a cross section of the stator component 20 along a plane perpendicular to the radial axis r. More specifically, Figure 2 is a schematic representation of the stator part 20 in a circumferential plane situated at a constant distance Δ from the axis Δ of the turbine. This circumferential plane, parallel to the axis Δ of the turbine, enables the cross-section of the blades 24 , 26 to be defined.

[0039] The blades 24 and 26 each have an inner camber side 624 , 126 and an outer camber side 124 , 626 .

[0040] The blades 24 and 26 each comprise a leading edge 224, 226 on the upstream side and a trailing edge 324, 326 on the downstream side. The terms upstream and downstream are defined relative to the general flow of gas through the turbine, which is from upstream to downstream in the direction and sense of the turbine axis Δ.

[0041] The blades define a blade chord 424 , 426 , which is the length of a segment connecting the leading edge and the trailing edge in a circumferential plane of constant radius or distance from the axis Δ, which may be referred to as a cross-sectional plane.

[0042] Similarly, in the circumferential cross-sectional plane, each blade has a camber line 143, 141, which is a curve equal to the average value between the outer and inner camber curves. More precisely, the camber line is formed by all points located at equal distances from the outer and inner camber sides. The distance from a specific point to the outer camber side (or inner camber side) is defined herein as the minimum distance between the specific point and a point on the outer camber side (or inner camber side).

[0043] Each blade has a maximum arc 526, which corresponds to the maximum distance separating the arc lines 141, 143 from the chords 424, 426. More precisely, the maximum distance corresponds to the maximum value of the distance between a point on the arc line and a corresponding point on the chord when the chord is traversed, measured perpendicular to the chord. The axial coordinate X of the maximum arc C (ie the coordinate along the delta axis of the turbine) corresponds to the axial position of the point on the arc where the distance to the chord is greatest.

[0044] fins

[0045] The stator component 20 further comprises fins 28 extending from the platform 22 in the same direction and in the same direction of extension as the blades 24, 26. The fins 28 extend radially from the platform 22 in the duct relative to the axis Δ of the turbine.

[0046] The fin 28 has an outer arc side 50 that faces the inner arc side 126 of the blade 26 .

[0047] When the component includes two blades 24, 26 facing each other, the fin 28 is located between the blades 24 and 26. More specifically, the fin 28 is located opposite the outer arc side 124 of the first blade 24 and opposite the inner arc side 126 of the second blade 26.

[0048] The fin 28 includes an inner arc side 48 and an outer arc side 50 , the inner arc side 48 facing the outer arc side 124 of the first blade and the outer arc side 50 facing the inner arc side 126 of the second blade 26 .

[0049] The fins extend radially in the duct from a fin root 44 located on the platform 22 to a fin tip 46 .

[0050] The fin 28 can be modeled or represented as a stack of profiles (i.e., cross-sectional surfaces) in the radial direction between a fin root 44, corresponding to a first profile of the fin, and a fin tip 46, corresponding to a final profile of the fin. The fin root 44 is located on the platform 22 and corresponds to the intersection of the fin 28 and the platform 22. The fin tip is located at a distance from the platform 22 in the gas flow duct. Each fin profile extends in a circumferential plane parallel to the axis Δ of the turbine. The circumferential plane, being a section of the fin formed in this circumferential plane at a constant radius or distance from the axis Δ, can be referred to as a cross-sectional plane.

[0051] Each fin profile defines a fin chord located between the leading edge 30 of the fin and the trailing edge 32 of the fin. More specifically, the fin chord is defined between a first point located at the intersection of the leading edge 30 and the cross-sectional plane and a second point located at the intersection of the trailing edge 32 and the cross-sectional plane. The fin chord specifies the length of the segment connecting the first and second points. The chord line specifies the segment connecting the first and second points.

[0052] Figure 2 A cross section of the fin 28 at the fin tip is shown, the fin tip extending axially from the leading edge point 31 of the leading edge 30 to the trailing edge point 37 of the trailing edge 32. The fin tip defines a chord length 54 joining the leading edge point 31 and the trailing edge point 37 in a straight line. When the fin tip is projected onto the axis Δ of the turbine, a projected section is obtained, the two ends of which are the projected portions X of the leading edge 31. BA and the projection X of the trailing edge 37 BF These projections correspond to the axial coordinates X BA and X BF (In other words, the coordinates along the axis Δ of the turbine.) These projections can be used to define a projected tip chord 43 lying on the axis Δ of the turbine, the projected tip chord 43 being |X BA -X BF │ and is less than or equal to the chord length 54 at the tip of the fin.

[0053] The fin tip has a curved line between points A and A′, which is a curve equal to the average value between the curves of the outer curved side 50 and the curves of the inner curved side 48 in the cross-sectional plane.

[0054] Figure 3 and Figure 4 corresponds to a cross section of the fin along a meridian plane, the meridian plane corresponding to the axial direction r which passes through when the arc is traversed Figure 2 The point of the curve AA' in the cross-sectional plane. Figure 3 and Figure 4 The abscissa axis in corresponds to the abscissa along the curve AA', and the axial coordinate can also correspond to the curve AA' by projecting it onto the axis Δ of the turbine. Figure 3and Figure 4 In FIG, the ordinate axis corresponds to the radial direction r and to a distance Δ from the axis of the turbine.

[0055] Leading edge 30 includes a leading edge point 34 located on platform 22. Leading edge point 34 corresponds to the intersection of leading edge 30 and platform 22.

[0056] Trailing edge 32 includes a trailing edge point 36 located on platform 22. Trailing edge point 36 corresponds to the intersection of trailing edge 32 and platform 22.

[0057] The shape of the leading edge 30 and the trailing edge 32 is not critical. These edges can be steep (i.e., oriented in the radial direction) or inclined (i.e., oriented in a direction that forms a non-zero angle with the radial direction). The leading edge 30 and the trailing edge 32 can have any shape, depending on the stacking pattern that forms the profile of the fin 28.

[0058] The fins 28 and blades 24 and / or 26 are arranged so that the maximum curvature 526 is located axially between the leading edge 31 and the trailing edge 37. In other words, when the maximum curvature and the fin tip profile are projected onto the axis Δ of the turbine, the projection of the maximum curvature is located within the projection section between the two ends of the projection of the fin tip profile, the two ends corresponding to the projection of the leading edge 31 and the trailing edge 37. In other words, the axial coordinate X C Located at the axial coordinate X BA and X BF In other words, with respect to the main gas flow in the turbine, the maximum camber is located downstream of the leading edge 31 of the fin tip 46 and upstream of the trailing edge 37 of the fin tip 46 .

[0059] Furthermore, the ratio of the axially measured distance between the maximum camber 526 and the leading edge 31 of the fin tip 46 to the projected chord length 43 of the fin tip 46 is less than or equal to 95% and greater than or equal to 5%. This means that by calculating the ratio between the distance separating the projection of the maximum camber 526 on the axis Δ and the projection of the leading edge 31 on the axis and the projection of the chord 54 of the fin tip 46 on the axis, a ratio between 5% and 95% is obtained. This condition can also be calculated using the axial coordinate X C 、X BA and X BF Written in the following form:

[0060] 0.95≥│X BA -X C │ / │X BA -X BF │≥0.05

[0061] This condition is equivalent to the following condition: the ratio of the axially measured distance between the maximum camber 526 and the trailing edge 37 of the fin tip 46 to the projected chord length 43 of the fin tip 46 is less than or equal to 95% and greater than or equal to 5%. This means that by calculating the ratio between the distance separating the projection of the maximum camber 526 on the axis Δ and the projection of the trailing edge 37 on the axis Δ and the projection of the chord 54 of the fin tip 46 on the axis, a ratio between 5% and 95% is obtained. This condition can also be calculated using the axial coordinate X C 、X BA and X BF Written in the following form:

[0062] 0.95≥│X BF -X C │ / │X BA -X BF │≥0.05.

[0063] Fin height

[0064] The fins 28 extend radially from a root profile 44 to a tip profile 46 , with a fin height h(x).

[0065] It should be noted that this height is defined relative to the platform 22 , which is considered to be located at a constant distance from the axis of the turbine.

[0066] The fin height h(x) varies as a function of the axial coordinate x (in other words, along the axis Δ). The height h(x) varies as a function of the axial coordinate x (in other words, along the axis Δ). Figure 3 More precisely, h(x) is the fin height along arc line AA'.

[0067] In particular, the height h(x) as a function of the axial coordinate x has a local minimum located axially near the axial coordinate x of the maximum arc 526. C .

[0068] In other words, the fin has recesses 45 extending axially on either side of the maximum curvature 526 .

[0069] The recess has a recess length 47 defined along the axis Δ.

[0070] The recess length 47 may be defined, for example, between an upstream point 41 of the fin tip 46 and a downstream point 49 of the fin tip 46:

[0071] the upstream point 41 of the fin tip 46 is located upstream of the recess, ie between the recess 45 and the leading edge 31 of the tip 46, and

[0072] The downstream point 49 of the fin tip 46 is situated downstream of the recess, ie between the recess 45 and the trailing edge 37 of the tip 46 .

[0073] The upstream point 41 may, for example, correspond to the first point on the arc of the fin tip 46 where the height of the fin decreases as one traverses the fin from upstream to downstream.

[0074] Downstream point 49 may, for example, correspond to a first point on the arc of fin tip 46 downstream of maximum arc 526 for which the height of the fin remains constant or reaches the value of the fin height at upstream point 41 as the fin traverses the fin from upstream to downstream.

[0075] The length 47 of the recess is selected to be less than or equal to 20% of the length of the chord 54 .

[0076] Preferably, the length 47 of the recess may also be selected to be less than or equal to 15% or 10% of the length of the chord 54 .

[0077] Preferably, the length 47 of the recess may also be selected to be greater than or equal to 2% of the length of the chord 54 .

[0078] The presence of the recesses in the fins limits the aerodynamic blocking effects encountered on the stator at high speeds (especially in the transonic regime), while maintaining the blockage of the cross flow.

[0079] In the transonic regime, a portion of the flow has a flow velocity greater than the speed of sound and a portion of the flow has a flow velocity less than the speed of sound.

[0080] The regions where the velocity drops from above the speed of sound to below it correspond to shocks. This is particularly true as the flow cross section increases, which is accompanied by a decrease in flow velocity. In these regions, there are sharp drops in velocity and sharp changes in pressure, which can be highly spatially localized (corresponding to shocks). These shocks can lead to separation of the hub boundary layer and also to separation of the outer camber sides of the blades involved in the inter-blade passages. This shock is contrary to the main effect of the fins we would expect, which is to limit angular separation by acting on the crossflow.

[0081] In stator blade cascades for transonic operation, these regions are located inside the blade cascades, their position depending in particular on the inclination and thickness of the blades. However, the position is correctly approximated by the axial coordinate of the maximum curvature of the blades.

[0082] By locating the recess axially at the maximum arc of the blade, the recess's axial position corresponds to the axial location of the shock associated with aerodynamic blockage. Thus, the recess, thus positioned, removes material from the area of the flow duct where the shock associated with aerodynamic blockage occurs. If this material were retained, it could cause or exacerbate the effects of the shock. This material removal limits the effects of aerodynamic blockage.

[0083] Compared with non-recessed shapes, the advantages of this fin shape are as follows:

[0084] - at high incidence (e.g. at partial speeds), the fins retain their primary effect of blocking the cross flow and thus reducing the secondary flow at the stator root: the operability of the compressor is improved, and

[0085] In regulating operation (for example at the cruise point), the recess limits the influence of the fins on the blockage associated with the surges and therefore limits the losses: the maximum efficiency of the compressor is increased.

[0086] Crossflow over the rectifier is effective at low speeds (subsonic), but also at high speeds (transonic).

[0087] As an alternative to the minimum fin height, the fin height has a maximum height h in the recess 45. 最大 and minimum height h 最小 , the ratio of minimum height to maximum height is less than or equal to 20%.

[0088] In particular, the maximum height may be achieved upstream of the recess 45 and downstream of the recess 45. For example, the fin height is equal to the maximum height on the camber at the fin tip 46 between the leading edge point 31 and the upstream point 41 and between the downstream point 49 and the trailing edge point 37.

[0089] Minimum height h in recess 45 最小 The local height minimum in the region corresponding to the concave portion. In particular, the axial coordinate X of the maximum arc 526 can be C Get the minimum height, that is, h(X C )=h 最小 .

[0090] Minimum height h 最小 With the maximum height h 最大 The ratio is selected to be less than or equal to 20%, that is, (h 最大 / h minimum)≤0.2.

[0091] More advantageously, the minimum height h 最小 With the maximum height h 最大 The ratio is selected to be less than or equal to 5%, that is, (h 最大 / h 最小 )≤0.05.

[0092] These limited minimum heights are therefore sufficient to maintain a blocking effect in the lower part of the boundary layer at partial velocities in situations where the stator is subjected to high incident fluid.

[0093] According to a first embodiment, the fin height in the recess is symmetrical with respect to the axial position of the maximum arc. For example, the curve h(x) of the fin height along the arc line AA' is symmetrical in the recess. This means that, on the one hand, the fin height decreases from the upstream point 41 downstream and reaches a minimum height, and, on the other hand, the fin height increases from the minimum height downstream and reaches a downstream point 49, which are mirror images of each other. The minimum height is thus arranged halfway between the upstream point 41 and the downstream point 49. This symmetry can be evaluated in particular along the arc line or by an axial projection on this arc line. Preferably, the symmetry is evaluated along the arc line. This first method is in Figure 3 Shown in.

[0094] According to a second embodiment, the fin height in the recess is symmetrical with respect to the axial position of the maximum arc. In particular, the fin height in the recess (as a function of the coordinate along the axis Δ of the turbine) has a first rate of change upstream of the axial position of the maximum arc and a second rate of change downstream of the axial position, the absolute value of the first rate of change being greater than the absolute value of the second rate of change. This means that, firstly, the rate of change at which the fin height decreases from the upstream point 41 downstream and reaches the minimum height is greater than the rate of change at which the fin height increases from the minimum height downstream and reaches the downstream point 49. Therefore, the minimum height is arranged closer to the upstream point 41 than to the downstream point 49. This second method is in Figure 4 Shown in.

[0095] Thus, as one traverses the fin from upstream to downstream along the chord, the concavity decreases rapidly in height initially and then increases slowly in height. This difference allows the flow to be directed longer at the fin peak and potentially improves the wall effect against crossflow.

Claims

1. A stator component (20) for a turbine, comprising: - a platform (22) defining a wall of the gas flow duct, - blades (24, 26) extending radially from the platform (22) relative to the axis (Δ) of the turbine, the blades having a maximum camber (526), and a fin (28) extending radially in the duct from a root (44) located on the platform (22) to a tip (46), the tip (46) extending axially along the axis (Δ) from a leading edge (31) to a trailing edge (37) and having a chord length (54) connecting the leading edge (31) and the trailing edge (37) in a straight line, The fins (28) and the blades (24, 26) are designed such that: - said maximum arc (526) is located axially between said leading edge (31) and said trailing edge (37), - the ratio of the distance measured axially between the maximum arc (526) and the leading edge (31) to the chord length (54) is less than or equal to 95% and greater than or equal to 5%, and - a fin height defined between the root and the tip that varies along the axis (Δ) to define a recess (45) in the fin, the recess (45) extending axially from one side of the maximum arc (526) to the other side over a recess length (47) that is less than or equal to 20% of the chord length (54).

2. The stator component (20) according to claim 1, wherein The fin height has a maximum height (h 最大 ) and minimum height (h 最小 ), the ratio of the minimum height to the maximum height is less than or equal to 20%.

3. The stator component (20) according to claim 2, wherein: The ratio of the minimum height to the maximum height is less than or equal to 10%, preferably equal to 5%.

4. The stator component (20) according to any one of claims 1 to 3, wherein: The fin heights in the recess are symmetrical relative to the axial position of the maximum arc.

5. The stator component (20) according to any one of claims 1 to 3, wherein: The fin height in the recess has a first rate of change as a function of the coordinate along the axis (Δ), upstream of the axial position of the maximum arc, and a second rate of change downstream of the axial position, the terms upstream and downstream referring to the main flow direction of the gas in the stator part, the absolute value of the first rate of change being greater than the absolute value of the second rate of change.

6. A turbomachine comprising a stator component (20) according to any one of claims 1 to 5.

7. An aircraft comprising a turbine according to claim 6.

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