Stator part with blade and fin arrangement in a turbomachine

Abstract

The invention relates to a stator part (20) comprising a platform (22), a blade (24, 26) that has a maximum camber (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 blade (24, 26) being arranged such that: - the maximum camber (526) lies between the leading edge (31) and the trailing edge (37); - a ratio of a distance between the maximum camber (526) and the leading edge (31) to a fin chord length (54) is between 5% and 95%; and - the fin head defines a recess (45) extending on either side of the maximum camber (526) over a recess length (47) that is less than or equal to 20% of the chord length (54).

Description

[0001] STATOR PART WITH BLADE AND FIN ARRANGEMENT IN A TURBOMACHINE

[0002] FIELD OF THE INVENTION

[0003] The invention relates to the stator parts of a turbomachine comprising a blade such as the flow straighteners located downstream of a compressor and in particular the fixed-pitch straighteners.

[0004] STATE OF THE ART

[0005] In an aircraft turbomachine, and in particular aircraft intended for passenger transport, it is the air propelled by a fan and combustion gases exiting the turbomachine through an exhaust nozzle that exerts a reaction thrust on the turbomachine and, through it, on the aircraft. The circulation of gases through the turbomachine is influenced by rotating blades and fixed blades. The fixed 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"). The stator vanes of a gas turbine aircraft engine may each have two platforms (inner and outer) which are attached to the blades.There are also unducted designs with stator blades that have only one inner platform. In all cases, these stator blades form rows of fixed vanes that guide the gas flow through the engine at an appropriate speed and angle.

[0006] Within a flow straightener comprising a plurality of fixed blades, the flow of gases generally takes place between the blades in an upstream-downstream direction. It is known, however, that the areas of the blade root and tip can be the site of secondary aerodynamic flows.

[0007] 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 passage flow (also known as "crossflow" in English) which transports the gases towards the extrados.

[0008] At the blade tip, i.e. at the junction between the blade and the hub or between the blade and the casing, corner separation (also known as "corner separation" in English) and a vortex (also known as "corner vortex" in English) can occur. This separation generates pressure losses as well as aerodynamic blockage. The latter is problematic in terms of operability. It is possible to place a fin on the hub or on the casing between two blades placed opposite each other so as to reduce the flow. The fins improve the operability of compressor stators by limiting the flow between the blades.

[0009] The geometry of a blade grid, i.e. a set of blades distributed circularly around an axis, can be optimized for high-speed fluid flows, and in particular transonic flows close to the speed of sound. The addition of a fin in such an optimized grid can significantly degrade the operation of these optimized blades.

[0010] There is therefore a need for a new geometry to correct these problems and improve performance in terms of equipment efficiency, particularly for transonic fluid flows.

[0011] STATEMENT OF THE INVENTION

[0012] An aim of the invention is to propose a new geometry of the blades making it possible to improve the performance in terms of equipment efficiency, particularly for transonic flows.

[0013] The aim is achieved within the framework of the present invention thanks to a stator part of a turbomachine comprising:

[0014] - a platform defining a wall of a gas flow vein,

[0015] - a blade extending radially relative to an axis of the turbomachine from the platform, the blade having maximum camber, and

[0016] - a fin extending radially in the vein from a root located on the platform to a head, the head extending axially along the axis from a leading edge to a trailing edge and having a chord length connecting the leading edge and the trailing edge in a rectilinear manner, the fin and the blade being arranged so that:

[0017] - the maximum camber is located axially between the leading edge and the trailing edge,

[0018] - a ratio of a distance measured axially between the maximum camber and the leading edge to the chord length is less than or equal to 95% and greater than or equal to 5%, and

[0019] - a fin height defined between the root and the head varies along the axis so as to define a hollow in the fin, the hollow extending axially on either side of the maximum camber over a hollow length less than or equal to 20% of the chord length.

[0020] Such a part is advantageously and optionally completed by the following different characteristics taken alone or in combination:

[0021] - the fin height has a maximum height and a minimum height in the hollow, a ratio of the minimum height to the maximum height being less than or equal to 20%;

[0022] - the ratio of the minimum height to the maximum height is less than or equal to 10% and preferably equal to 5%;

[0023] - the fin height in the hollow is symmetrical with respect to an axial position of the maximum camber; and

[0024] - the fin height in the hollow as a function of a coordinate along the axis has a first rate of variation upstream of an axial position of the maximum camber and a second rate of variation downstream of the axial position, the terms upstream and downstream referring to a main flow direction of the gases in the stator part, the first rate of variation being greater in absolute value than the second rate of variation.

[0025] The invention also relates to a turbomachine comprising a stator part such as has just been presented and to an aircraft comprising such a turbomachine.

[0026] DESCRIPTION OF FIGURES

[0027] Other characteristics and advantages of the invention will emerge from the following description, which is purely illustrative and non-limiting, and must be read in conjunction with the appended drawings in which:

[0028] Figure 1 is a schematic representation of a turbomachine;

[0029] Figure 2, Figure 3 and Figure 4 are schematic views of a stator part according to different embodiments.

[0030] DETAILED DESCRIPTION OF THE INVENTION

[0031] Turbomachine

[0032] With reference to Figure 1, a turbomachine is shown schematically, more specifically a double-flow axial turbojet 1. Other types of turbomachines can be improved by the teaching presented here and in particular unducted turbomachines. The turbojet 1 illustrated extends along an axis A and successively comprises, in the main flow direction of the gases in the turbomachine, a fan 2, a compression section which may comprise a low-pressure compressor 3 and a high-pressure compressor 4, a combustion chamber 5, and a turbine section which may comprise a high-pressure turbine 6, a low-pressure turbine 7 and an exhaust nozzle.

[0033] The blower 2 and the low pressure compressor 3 are driven in rotation by the low pressure turbine 7 via a first transmission shaft 9, while the high pressure compressor 4 is driven in rotation by the high pressure turbine 6 via a second transmission shaft 10.

[0034] In operation, a flow of air compressed by the low and high pressure compressors 3 and 4 feeds 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 engine 1 through an exhaust nozzle downstream of the turbines 6, 7 exert a reaction thrust on the turbojet engine 1 and, through it, on a vehicle or machine such as an aircraft (not shown).

[0035] Stator part

[0036] Downstream of the fan or a compression stage, the turbomachine may comprise a stage of straightening vanes. Such a stage of straightening vanes may comprise a stator part 20 as shown with reference to FIG. 2.

[0037] The stator part may be one of the stators included in one of the low pressure or high pressure axial compressors. The stator part may also include a grid of outlet guide vanes as in the secondary vein rectifier on the blower module.

[0038] The stator part 20, or the set 20 of stator parts if it is not a single piece, has at least one blade 24, 26 and a platform 22 from which the blade 24, 26 extends. The stator part may for example comprise two adjacent blades 24, 26 which extend from the platform 22.

[0039] The term "platform" here designates any element of the turbomachine from which blades 24, 26 are capable of being mounted. The platform may in particular be a hub or a casing which surrounds the axis of the turbomachine. The platform may also be sectorized, that is to say that it comprises a set of parts distributed angularly all around the axis according to ideally equal angular sectors. The platform may have a cylindrical surface at a constant radial distance from the axis A of the turbomachine. The platform 22 has an internal wall or an external wall against which the air circulates, that is to say that the platform 22 defines a wall of a gas flow vein. The blades 24, 26 extend from the platform 22 in the vein either radially outwards away from the axis of the turbomachine A or radially inwards approaching the axis of the turbomachine A.

[0040] In Figure 2, the axis A of the turbomachine is shown oriented positively in the direction of the main gas flow in the turbomachine. Figure 2 also shows a radial axis r perpendicular to the axis A of the turbomachine and which passes through the axis A of the turbomachine. The radial axis is oriented positively in the direction away from the axis A of the turbomachine. Figure 2 also shows a circumferential axis 0 which is perpendicular to the radial axis r and to the axis A of the turbomachine.

[0041] Figure 2 is a schematic representation of the stator part 20 according to a section in a plane perpendicular to the radial axis r. More precisely, Figure 2 is a schematic representation of the stator part 20 in a circumferential plane which is at a constant distance from the axis A of the turbomachine. Such a circumferential plane parallel to the axis A of the turbomachine makes it possible to define a section of the blades 24, 26.

[0042] The blades 24 and 26 each have a lower surface 624, 126 and an upper surface 124, 626.

[0043] 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 in relation to the general flow of gases through the turbomachine which takes place from upstream to downstream in the direction and sense of the axis A of the turbomachine.

[0044] The blades define a blade chord 424, 426 which is the length of the segment connecting the leading edge and the trailing edge in a circumferential plane of constant radius or constant distance from the axis A, a circumferential plane which can be described as a cutting plane.

[0045] Similarly, in a circumferential section plane, each blade has a camber line 143, 141 which is the curve equal to the average between the curve of the extrados and the curve of the intrados. More precisely, the camber line is formed by all the points located at equal distance from the extrados and the intrados. The distance from a particular point to the extrados (or intrados) is defined here as the minimum distance between the particular point and a point on the extrados (or intrados).

[0046] Each blade has a maximum camber 526 which corresponds to the maximum distance separating the camber line 141, 143 and the chord 424, 426. The maximum distance corresponds more precisely to the maximum of the distances measured perpendicular to the chord between a point on the camber line and the corresponding point on the chord when traveling along the chord. The axial coordinate xc - that is to say the coordinate along the axis A of the turbomachine - of the maximum camber corresponds to the axial position of the point on the camber line for which the distance to the chord is maximum. Winglet

[0047] The stator part 20 also comprises a fin 28 which extends from the platform 22 in the same direction and the same sense of extension as the blade(s) 24, 26. The fin 28 extends in the vein radially relative to the axis A of the turbomachine from the platform 22.

[0048] The fin 28 comprises an extrados 50 which is opposite the intrados 126 of the blade 26.

[0049] When the part comprises two blades 24, 26 facing each other, the fin 28 is located between the blades 24 and 26. More precisely, the fin 28 is located facing the extrados 124 of the first blade 24 and the intrados 126 of the second blade 26.

[0050] The fin 28 comprises a lower surface 48 which is opposite the upper surface 124 of the first blade and an upper surface 50 which is opposite the lower surface 126 of the second blade 26.

[0051] The fin extends in the vein radially from a fin root 44 located on the platform 22 to a fin head 46.

[0052] The fin 28 can be modeled or represented as a stack of profiles, i.e. cutting surfaces, in a radial direction between a fin root 44 corresponding to a first profile of the fin and a fin head 46 corresponding to a last 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 head is located at a distance from the platform 22 in the gas flow path. Each fin profile extends in a circumferential plane parallel to the axis A of the turbomachine, like a section of the fin made in this circumferential plane with constant radius or constant distance from the axis A, a circumferential plane which can be described as a section plane.

[0053] Each winglet profile defines a winglet chord between the leading edge

[0054] 30 of the fin and the trailing edge 32 of the fin. More precisely, the fin chord is defined between, on the one hand, a first point at the intersection of the leading edge 30 and the cutting plane and, on the other hand, a second point at the intersection of the trailing edge 32 and the cutting plane. The fin chord designates the length of the segment connecting the first point and the second point. The chord line designates the segment connecting the first point and the second.

[0055] In Figure 2, the fin 28 is shown in its section at the fin head which extends axially from a point of attack 31 of the leading edge 30 to a trailing point 37 of the trailing edge 32. The fin head defines a length of chord 54 connecting in a rectilinear manner the point of attack

[0056] 31 and the vanishing point 37. When the fin head is projected onto the axis A of the turbomachine, a projection segment is obtained whose two ends are the projection XBA of the leading edge 31 and the projection XBF of the trailing edge 37. These projections correspond to axial coordinates XBA and XBF - that is to say coordinates along the axis A of the turbomachine. From these projections, it is possible to define the projected tip chord 43 on the axis A of the turbomachine which is equal to | XBA - XBF | and which is less than or equal to the chord length 54 of the fin head.

[0057] The fin head has a camber line between points A and A' which is the curve equal to the average between the curve of the extrados 50 and the curve of the intrados 48 in the section plane.

[0058] Figures 3 and 4 correspond to a section of the fin along a meridian plane which corresponds to the axial directions r which pass through the points of the curve AA' in the section plane of Figure 2 as one travels along this camber line. The abscissa axis in Figures 3 and 4 corresponds to a curvilinear abscissa following the curve AA' to which one can also make correspond an axial coordinate in projection on the axis A of the turbomachine. In Figures 3 and 4, the ordinate axis corresponds to the radial direction r and corresponds to a distance from the axis A of the turbomachine.

[0059] The leading edge 30 includes a point of attack 34 located on the platform 22. The point of attack 34 corresponds to the intersection of the leading edge 30 and the platform 22.

[0060] The trailing edge 32 includes a vanishing point 36 located on the platform 22. The vanishing point 36 corresponds to the intersection of the trailing edge 32 and the platform 22.

[0061] The shape of the leading edge 30 and the trailing edge 32 is not important here. These edges can be steep, that is to say oriented in a radial direction or oblique, that is to say oriented in a direction which makes 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 law of the profiles which constitute the fin 28.

[0062] The fin 28 and the blade 24 and / or 26 are arranged so that the maximum camber 526 is located axially between the leading edge 31 and the trailing edge 37. In other words, when the maximum camber and the fin head profile are projected onto the axis A of the turbomachine, the projection of the maximum camber is located inside the projection segment between the two ends of the projection of the fin head profile which correspond to the projections of the leading edge 31 and the trailing edge 37. In other words, the axial coordinate xc is between the axial coordinates XBA and XBF. In other words, relative to the main flow of gases in the turbomachine, the maximum camber is located downstream of the leading edge 31 of the fin head 46 and upstream of the trailing edge 37 of the fin head 46.Furthermore, a ratio of an axially measured distance between the camber maximum 526 and the leading edge 31 of the winglet head 46 to the projected chord length 43 of the winglet head 46 is less than or equal to 95% and greater than or equal to 5%. This means that by calculating the ratio of the distance separating the projection on the axis A of the camber maximum 526 and the projection on the axis A of the leading edge 31 to the projection on the chord axis 54 of the winglet head 46, a ratio of between 5% and 95% is obtained. This condition can also be written using the axial coordinates xc, XBA and XBF in the form: 0.95. 0.05

[0063] This condition is equivalent to the following condition: an axially measured distance between the camber maximum 526 and the trailing edge 37 of the fin head 46 over the projected chord length 43 of the fin head 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 on the axis A of the camber maximum 526 and the projection on the axis A of the trailing edge 37 over the projection on the chord axis 54 of the fin head 46, a ratio of between 5% and 95% is obtained. This condition can also be written using the axial coordinates xc, XBA and XBF in the form:

[0064] 0.95 0.05.

[0065] Fin height

[0066] The fin 28 extends radially from the root profile 44 to the head profile 46 over a fin height h(x).

[0067] It should be noted here that this height is defined in relation to a platform 22 considered at a constant distance from the axis of the turbomachine.

[0068] The fin height h(x) varies as a function of the axial coordinate x, i.e. along the axis A. The height h(x) is illustrated schematically in Figure 3. More precisely, h(x) is the fin height along the camber line AA'.

[0069] In particular, the height h(x) as a function of the axial coordinate x has a local minimum, the local minimum being located axially close to the axial coordinate xc of the maximum camber 526.

[0070] In other words, the fin has a hollow 45 which extends axially on either side of the maximum camber 526.

[0071] The hollow has a hollow length 47 defined along axis A. This hollow length 47 can for example be defined between:

[0072] - an upstream point 41 of the fin head 46 located upstream of the hollow, that is to say between the hollow 45 and the leading edge 31 of the head 46, and

[0073] - a downstream point 49 of the fin head 46 located downstream of the hollow, that is to say between the hollow 45 and the trailing edge 37 of the head 46.

[0074] The upstream point 41 may correspond, for example, to the first point on the camber line of the fin head 46 for which the height of the fin decreases when moving along the fin from upstream to downstream.

[0075] The downstream point 49 may correspond, for example, to the first point of the camber line of the fin head 46 located downstream of the maximum camber 526, for which, when the fin is traveled from upstream to downstream, the height of the fin remains constant or reaches the value of the height of the fin at the upstream point 41.

[0076] The hollow length 47 is chosen to be less than or equal to 20% of the chord length 54.

[0077] Preferably, it is also possible to choose that the hollow length 47 is less than or equal to 15% or 10% of the chord length 54.

[0078] Preferably, it is also possible to choose that the hollow length 47 is chosen to be greater than or equal to 2% of the chord length 54.

[0079] The presence of a hollow in the fin makes it possible to limit the aerodynamic obstruction effects encountered on stators at high speed and in particular in transonic regime while maintaining a blockage of the passage flow.

[0080] In transonic flow, part of the flow has a flow velocity greater than the speed of sound and part of the flow has a flow velocity less than the speed of sound.

[0081] The areas of reduced speed going from above the speed of sound to below it correspond to a shock. This is particularly the case when the flow section increases, which is accompanied by a reduction in the flow speed. In these areas there is a very strong decrease in speed and a very strong variation in pressure which can be spatially very well localized, corresponding to the shock. These shocks can cause a detachment of the hub boundary layer but also on the extrados of the blade involved in an inter-blade channel. This shock corresponds to the opposite of the main effect expected from a blade, namely to limit corner separations by acting on the flow of passage. In a stator grid intended for use in transonic regimes, these areas are located inside the grid and their positions depend in particular on the inclination and thickness of the blades.However, it turns out that this position is correctly approached by the axial coordinate of the maximum camber of the blades.

[0082] The hollow, being axially located at the level of this maximum camber of the blades, has an axial position which corresponds to that of the shock linked to the aerodynamic obstruction. The hollow thus placed therefore corresponds to a removal of material from a place in the flow vein where the shock linked to the aerodynamic obstruction occurs. This material, if retained, could produce a shock or worsen the effects linked to a shock. Its removal makes it possible to limit the effects of aerodynamic obstruction.

[0083] The advantages of this fin shape, compared to the non-hollowed shape, are:

[0084] - under high incidence (for example at partial speed), the fin retains its main effect of blocking the passage flow and therefore reduces secondary flows at the stator foot: the operability of the compressor is improved, and

[0085] - in adapted operation (for example at cruising point), the hollow limits the effect of the fin on the blocking linked to the shock and therefore limits the losses: the maximum efficiency of the compressor is increased.

[0086] The flow passing over the rectifiers is then effective at low speed (subsonic) but also at high speed (transonic).

[0087] As an option regarding the minimum fin height, the fin height has a maximum height hmax, and a minimum height hmin in the hollow 45, a ratio of the minimum height to the maximum height being less than or equal to 20%.

[0088] The maximum height can in particular be reached upstream of the hollow 45 and downstream of the hollow 45. For example, the fin height is equal to the maximum height on the camber line at the fin head 46 between the attack point 31 and the upstream point 41, then between the downstream point 49 and the trailing point 37.

[0089] The minimum height hmin in the hollow 45 corresponds to the local minimum height in the hollow area. This minimum height can be obtained in particular for the axial coordinate xc of the maximum camber 526, i.e. h(xc)=h m in.

[0090] The ratio of the minimum height hmin to the maximum height hmax is chosen to be less than or equal to 20%, i.e. (hmin / h max) < 0.2. More advantageously, the ratio of the minimum height hmin to the maximum height hmax can be chosen to be less than or equal to 5%, i.e. (hmin / h max) - 0.05.

[0091] These minimum heights thus limited are sufficient to maintain the blocking effect, in the lower part of the boundary layer, at partial speed in the case where the stator is subjected to a high incidence fluid.

[0092] According to a first embodiment, the fin height in the hollow is symmetrical with respect to an axial position of the maximum camber. For example, it is the curve h(x) of the fin height along the camber line AA' which has a symmetry at the hollow. This means that on the one hand the decrease in fin height from the upstream point 41 towards the downstream and up to the minimum height and on the other hand the increase in fin height from the minimum height towards the downstream and up to the downstream point 49 are mirror images of each other. The minimum height is then placed in the middle between the upstream point 41 and the downstream point 49. This symmetry can be evaluated in particular along the camber line or by axially projecting this camber line. Preferably the symmetry is evaluated along the camber line. This first mode is represented in figure 3.

[0093] According to a second embodiment, the fin height in the hollow is not symmetrical with respect to an axial position of the maximum camber. In particular, the fin height in the hollow as a function of a coordinate along the axis A of the turbomachine has a first rate of variation upstream of an axial position of the maximum camber and a second rate of variation downstream of the axial position, the first rate of variation being greater in absolute value than the second rate of variation. This means that on the one hand the decrease in the fin height from the upstream point 41 towards the downstream and up to the minimum height occurs according to a greater rate of variation than the increase in the fin height from the minimum height towards the downstream and up to the downstream point 49. The minimum height is then placed closer to the upstream point 41 than to the downstream point 49. This second mode is shown in FIG. 4.

[0094] Thus, when traversing the fin chordwise from upstream to downstream, the trough begins with a rapid decrease in height followed by a slower increase in height. This difference allows for longer guidance of the fin crest flow, and potentially improves the wall effect against the passing flow.

Claims

CLAIMS 1. Stator part (20) of a turbomachine comprising: - a platform (22) defining a wall of a gas flow vein, - a blade (24, 26) extending radially relative to an axis (A) of the turbomachine from the platform (22), the blade having a maximum camber (526), ​​and - a fin (28) extending in the vein radially from a foot (44) located on the platform (22) to a head (46), the head (46) extending axially along the axis (A) from a leading edge (31) to a trailing edge (37) and having a chord length (54) connecting in a rectilinear manner the leading edge (31) and the trailing edge (37), the fin (28) and the blade (24, 26) being arranged so that: - the maximum camber (526) is located axially between the leading edge (31) and the trailing edge (37), - a ratio of a distance measured axially between the maximum camber (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 head varies along the axis (A) so as to define a hollow (45) in the fin, the hollow (45) extending axially on either side of the maximum camber (526) over a hollow length (47) less than or equal to 20% of the chord length (54).

2. Stator part (20) according to claim 1 in which the fin height has a maximum height (hmax), and a minimum height (hmin) in the hollow (45), a ratio of the minimum height to the maximum height being less than or equal to 20%.

3. Stator part (20) according to claim 2 in which the ratio of the minimum height to the maximum height is less than or equal to 10% and preferably equal to 5%.

4. Stator part (20) according to any one of claims 1 to 3 in which the fin height in the hollow is symmetrical with respect to an axial position of the maximum camber.

5. Stator part (20) according to any one of claims 1 to 3 in which the fin height in the hollow as a function of a coordinate along the axis (A) has a first rate of variation upstream of an axial position of the maximum camber and a second rate of variation downstream of the axial position, the terms upstream and downstream referring to a main flow direction of the gases in the stator part, the first rate of variation being greater in absolute value than the second rate of variation.

6. Turbomachine comprising a stator part (20) according to one of claims 1 to 5.

7. Aircraft comprising a turbomachine according to claim 6.

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