STATOR BLADE FOR AN AIRCRAFT TURBOMACHINE

Abstract

Stator vane (40) for an aircraft turbomachine (10), comprising a blade (42) which extends between a platform (44) and fixing lugs (46), the blade (42) comprising a lower surface (42a) and an upper surface (42b) which are connected together at a leading edge (50) and a trailing edge (52) of the blade (42), the leading edge (50) being connected to the platform (44) and being at a distance from the fixing lugs (46), characterized in that the blade (42) comprises a body (60) which is in one piece with the platform (44) and the fixing lugs (46), and a removable bar (70) which is attached and fixed to the body (60), one of the elements among the body (60) and the bar (70) comprising the leading edge (50), and the other of these elements comprising the trailing edge (52). Figure for abstract: Figure 5

Description

Title of the invention: STATOR VANE FOR AN AIRCRAFT TURBOMACHINE Technical field of the invention

[0001] The present invention relates to a stator blade for an aircraft turbomachine, as well as to an aircraft turbomachine comprising such blades. Technical background

[0002] The state of the art includes in particular document FR-A1-3 101 107.

[0003] An aircraft turbomachine comprises a gas generator which conventionally comprises, from upstream to downstream, with reference to the flow of gases in the turbomachine, at least one compressor, an annular combustion chamber and at least one turbine. In the case of a twin-spool turbojet engine, respectively low pressure and high pressure, the gas generator successively comprises a low pressure compressor, a high pressure compressor, the combustion chamber, a high pressure turbine and a low pressure turbine. The gas generator defines a first annular flow vein for a gas flow, called the primary flow, which passes through the compressors, the combustion chamber and the turbines.

[0004] The rotor of the high pressure compressor is connected to the rotor of the high pressure turbine by a high pressure shaft. The rotor of the low pressure compressor is connected to the rotor of the low pressure turbine by a low pressure shaft which passes through the high pressure shaft and which rotates a propulsion propeller generally located upstream of the gas generator.

[0005] When this propeller is shrouded and therefore surrounded by an annular casing, this propeller is called a fan and generates an air flow, called a secondary flow, which flows around the gas generator.

[0006] The secondary flow flows in a vein around the gas generator which includes stator vanes, also called rectifier vanes or OGV vanes which is the acronym for Outlet Guide Vane. These vanes are major parts in the management of engine aerodynamics and in the supply of thrust to the turbomachine. Particular care must be taken with the geometry and condition of these parts in order to properly ensure their aerodynamic role. Due to their position, these parts are subject to numerous external attacks, for example by particles included in the secondary flow, making them extremely sensitive to erosion problems.

[0007] In the current technique, the leading edge of a blade of this type is either included directly in the definition of the part and therefore produced by machining or forging, or can be glued to a composite structure.

[0008] In the event of excessive erosion on the leading edge, a complete replacement of the blade is currently necessary in order to return to the definition and ensure proper aerodynamic operation. This replacement poses a problem because it is expensive, a complete replacement of each part being necessary. In addition, in the case of erosion located on several blades, disassembly and then re-setting of the module containing blades may be necessary if too many blades are removed.

[0009] There is therefore a need to reduce maintenance time and repair costs in the event of erosion of the leading edge of this type of stator blade, and ideally also of the trailing edge of this type of blade.

[0010] The present invention provides a solution to this need, which is simple, effective and economical. Summary of the invention

[0011] The invention relates to a stator blade for an aircraft turbomachine, this blade comprising a blade which extends from a first end connected to a platform to a second end connected to fixing lugs, the blade comprising a lower surface and an upper surface which are connected together at a leading edge and a trailing edge of the blade, the leading edge being connected to the platform at said platform and being at a distance from the fixing lugs at said second end, characterized in that the blade comprises a body which is in one piece with the platform and the fixing lugs, and a removable bar which extends from the first end to the second end and which is attached and fixed to the body, the bar being formed of a plurality of sections arranged end to end along an axis of elongation of the bar, one of the elements among the body and the bar comprising the leading edge, and the other of these elements comprising the trailing edge.

[0012] The invention thus proposes to produce the blade of the vane in two parts. A first part of the blade is integral with the platform and the fixing lugs and therefore represents a main part of the blade and the vane. A second part of the blade is added and fixed on the body to define the leading or trailing edge of the blade. This second part is formed by a removable bar formed of several sections. In the event of wear of the leading or trailing edge, it is therefore sufficient to dismantle all or part of the bar and to replace the worn section(s). The assembly and disassembly of the bar and in particular of its sections are advantageously carried out when the stator vane is mounted in the turbomachine, therefore without dismantling the module comprising this vane and without removing the turbomachine. The assembly and disassembly of the vane can therefore be carried out "under the wing" of the aircraft.

[0013] The stator blade according to the invention may comprise one or more of the following features, taken in isolation from one another or in combination with one another: • the bar is mounted on the body by a male-female engagement, the bar comprising a male, or respectively female, element which extends along an elongation axis of the bar and which cooperates with a female, or respectively male, element of the body which extends between the first and second ends; • the bar is mounted on the body by a sliding or dovetail system; • the bar defines a portion of the intrados and a portion of the extrados of the blade; • the bar comprises a first longitudinal end engaged in a recess in the platform, and a second opposite longitudinal end fixed to the body by screwing or bolting; • the first end of the bar defines a portion of the platform; • the second end of the bar comprises at least one passage hole of a screw or bolt; • the screw or bolt extends parallel or perpendicular to an axis of elongation of the bar; • the sections comprise pins and / or holes configured to cooperate by interlocking with complementary means of the other sections in order to align the sections with each other along the elongation axis; • at least one damping layer is interposed between the bar and the body ; • the damping layer is made of elastically deformable material;

[0014] — the sections all have the same length along said axis; as a variant, at least two of the sections have a length along said axis which is different;

[0015] — at least one piece of equipment such as a sensor is integrated into at least one of the sections;

[0016] — said equipment is chosen from an accelerometer, a temperature sensor, a flow sensor, etc.;

[0017] — said equipment is a heating element and in particular a heating element defrosting;

[0018] — said sections are made of a material chosen from aluminum, titanium, their alloy(s), and a composite material;

[0019] — at least certain sections are produced by additive manufacturing;

[0020] — at least some sections are hollow or cellular, and have for example at least a honeycomb portion;

[0021] — at least some sections comprise serrations (of the tooth type and / or notches) at least on one edge.

[0022] The present invention also relates to an aircraft turbomachine, comprising at least one stator blade as described above.

[0023] The turbomachine may comprise a first flow vein of a primary flow inside a gas generator comprising at least one compressor, a combustion chamber and at least one turbine, and a second flow vein of a secondary flow around the gas generator, a fan being arranged upstream of the first and second veins and stator vanes as described above being arranged in the second vein. Brief description of the figures

[0024] Other characteristics and advantages of the invention will appear during the reading of the detailed description which follows for the understanding of which reference will be made to the appended drawings in which:

[0025] [Fig-1] [Fig.l] is a half schematic view in axial section of a part of a aircraft turbomachine,

[0026] [Fig.2] [Fig.2] is a schematic perspective view of a stator blade;

[0027] [Fig.3] [Fig.3] is a schematic perspective view of a stator blade;

[0028] [Fig.4] [Fig.4] is a schematic perspective view of a part of the body and of a part of the stator blade bar of [Fig.3];

[0029] [Fig.5] [Fig.5] is a schematic perspective view of a stator blade according to one embodiment of the invention;

[0030] [Fig.6] [Fig.6] is a partial schematic perspective view of one end of the stator vane of [Fig.5];

[0031] [Fig.7] [Fig.7] is a schematic perspective view of part of the bar of the stator blade of [Fig.5]. Detailed description of the invention

[0032] [Fig.l] shows a turbomachine 10 for an aircraft, this turbomachine 10 here being a double-flow, double-spool turbojet.

[0033] Axis A designates the longitudinal axis of the turbomachine.

[0034] The turbomachine 10 comprises a gas generator 12 which comprises from upstream to downstream with reference to the flow of gases along the axis A, a low pressure or LP compressor 14, a high pressure or HP compressor 16, an annular combustion chamber, an HP or high pressure turbine and a low pressure or LP turbine.

[0035] Although not visible in [Fig.l], the rotor of the HP compressor 16 is connected to the rotor of the HP turbine by a high pressure shaft, and the rotor of the compressor BP 14 and connected to the rotor of the BP turbine by a low pressure shaft which passes through the high pressure shaft and which drives in rotation a propulsion propeller, called fan 23, located upstream of the gas generator 12 and which is surrounded by an annular casing called fan casing 24.

[0036] The fan casing 24 is connected to the gas generator 12 by an intermediate casing 26 which comprises a central hub 28 and a series of radial arms 30 connecting the hub 28 to the fan casing 24.

[0037] The gas generator 12 defines a main annular vein V1 for the flow of a first air flow, called primary flow FL. The gas generator 12 is surrounded by a secondary annular vein V2 for the flow of a second air flow, called secondary flow F2.

[0038] The air flow F entering the fan 23 is divided into a part forming the primary flow FL. The air of this primary flow FL is compressed in the LP 14 and HP 16 compressors, then mixed with fuel and burned in the combustion chamber. The combustion gases of the primary flow are then expanded in the HP and LP turbines and finally flow into an exhaust nozzle.

[0039] The other part of the air flow entering the fan 23 forms the secondary flow F2 and is intended to be mixed with the primary flow F1 downstream of the nozzle. This secondary flow F2 successively passes through stator blades 40 and the radial arms 30 of the intermediate casing 26.

[0040] [Fig.2] shows a stator blade 40 of the current technique.

[0041] This blade 40 comprises a blade 42 which extends from a first end connected to a platform 44 to a second end connected to fixing lugs 46.

[0042] The blade 42 thus has a generally elongated shape with an elongation axis denoted B which is substantially radial to the axis A (see [Fig. 1]). As can be seen in [Fig. 1], the platform 44 is located at the radially internal end of the blade 42 and makes it possible to fix the blade 40 to the gas generator 12 and for example to the intermediate casing 26. The fixing lugs 46 are located at the radially external end of the blade 42 and make it possible to fix the blade 40 to the fan casing 24.

[0043] The platform 44 and the legs 46 can be fixed by screws or bolts for example, which pass through corresponding orifices 48 of the platform 44 and the legs 46.

[0044] The blade 42 has an aerodynamic profile and comprises a lower surface 42a and an upper surface 42b which are connected together at a leading edge 50 and a trailing edge 52 of the blade.

[0045] The platform 44 may extend over the entire length of the chord of the blade 42. This means that the leading 50 and trailing 52 edges of the blade 42 are connected to the platform 44.

[0046] On the contrary, the fixing lugs 46 preferably have precise and punctual positions along the chord of the blade 42, as illustrated in the drawing.

[0047] The number of legs 46 can be two or four for example.

[0048] In the example shown, a first double tab 46a extends in a transverse direction in the vicinity of the trailing edge 52 of the blade and comprises two orifices 48 for the passage of screws or bolts. A second double tab 46b extends in a transverse direction, between the first double tab 46a and the leading edge 50, and also comprises two orifices 48 for the passage of screws or bolts. The orifices 48 of each of the tabs 46a, 46b are located respectively on the side of the intrados 42a and the extrados 42b of the blade 42.

[0049] It is therefore understood that the leading edge 50 of the blade 42 is connected to the platform 44 at the radially inner end, and is at a distance from the fixing lugs 46 at the radially outer end. The trailing edge 52 of the blade 42 is connected to the platform 44 at the radially inner end, and may be at a distance or not from the double lug 46a at the radially outer end.

[0050] Figures 3 and following illustrate a solution which makes it possible in particular to facilitate the maintenance of a stator blade 40.

[0051] According to the invention, the blade 42 comprises on the one hand a body 60 which is in one piece with the platform 44 and the fixing lugs 46, and on the other hand a bar 70 which is removable.

[0052] The particularity of the bar 70 is that it extends between the ends of the blade 42 and is attached and fixed to the body 60. One of the elements among the body 60 and the bar 70 comprises the leading edge 50, and the other of these elements comprises the trailing edge 52.

[0053] Since the bar 70 is removable, it is understood that it is sufficient to dismantle and replace the bar 70 to replace the leading edge 50 or trailing edge 52 of the blade 42 in the event of wear, for example.

[0054] The following description focuses on the replacement of a bar 70 at the leading edge 50 of the blade 42 but it is understood that by analogy the invention applies in the case of a bar 70 at the trailing edge 52 of the blade 42.

[0055] The bar 70 is preferably mounted on the body 60 by a male-female engagement as illustrated in the drawings. The bar 70 then comprises a male element 72, or respectively female, which extends along an axis C of elongation of the bar 70 and which cooperates with a female element 74, or respectively male, of the body 60 which extends between the ends of the blade 42. This assembly and this cooperation can be similar to a slide system 76, the bar 70 and the body 60 being mounted on each other by sliding. Alternatively, this assembly and this cooperation could be similar to a dovetail system.

[0056] As can also be seen in the drawings, the bar 70 preferably defines a portion of the intrados 42a and a portion of the extrados 42b of the blade 42. Advantageously, these portions extend in the extension of the intrados and extrados portions of the body 60 so as to have aerodynamic continuity of these faces of the blade 42.

[0057] Figures 3 to 7 show characteristics of the stator blade 40 according to the invention in which the bar 70 comprises a first longitudinal end 70a engaged in a recess 80 of the platform 44 (Figures 4 and 5), and a second opposite longitudinal end 70b fixed to the body by screwing or bolting ([Fig.5]).

[0058] The first end 70a of the bar 70 can also be fixed to the body 60 by screwing or bolting ([Fig.6]).

[0059] Figures 4 and 5 show that the first end 70a of the bar 70 can define a portion of the platform 44, and in particular a portion of a connecting fillet 82 between the blade 42 and the platform 44.

[0060] [Fig. 6] shows that the second end 70b of the bar 70 comprises at least one orifice 84 for the passage of a screw 86 or a bolt. The screw 86 or the bolt extends parallel to the axis C of elongation of the bar 70.

[0061] In the example shown, a single screw 86 makes it possible to fix the bar 70 to this end. This screw 86 is preferably axially offset relative to the leading edge 50, along the axis C, so as to be outside the flow vein V2 of the secondary flow F2.

[0062] This is also the case for the screw 88 for fixing the bar 70 to the platform 44, the head of which is located radially inside the platform 44 with respect to the axis A.

[0063] In the example shown, the orifice 84 is formed in a finger 90 of the bar 70 which extends perpendicular to the axis C of elongation of the bar 70, here in the direction of the trailing edge 52 of the blade 42.

[0064] Figures 5 to 7 illustrate an embodiment of the stator blade 40 according to the invention which has the particularity that its bar 70 is formed of a plurality of sections 100 arranged end to end along the axis C of elongation of the bar 70.

[0065] The sections 100 comprise pins 102 and / or holes 104 configured to cooperate by interlocking with complementary means of the other sections 100 in order to align the sections with each other along the elongation axis C.

[0066] Furthermore, [Fig. 6] shows that the second end of the bar 70 comprises two orifices 84 for the passage of screws 86 or bolts, which extend perpendicular to the axis C of elongation of the bar 70.

[0067] The production of the bar 70 by sections is advantageous because it allows come and change only certain sections and therefore specific areas of the leading edge. It is then possible to reduce the areas changed in order to reduce the costs associated with changing parts. It is also possible to prioritize areas that undergo more regular erosion (upper part of blade 40 for example) and thus direct greater production for corresponding sections.

[0068] At least one damping layer 110 may be interposed between the bar 70 and the body 60, preferably over the entire length or extent of the bar 70 along its axis C. This damping layer 110 is preferably made of elastically deformable material.

[0069] In the event of erosion on the leading edge, it is possible to remove the blade, change all or part of the bar and reinstall it. This procedure offers the advantage of not changing the setting of the gap between the blades.

[0070] The sections 100 may all have the same length along the axis C. Alternatively, at least two of the sections 100 could have a length along the axis C which is different.

[0071] In a variant not shown, at least one piece of equipment such as a sensor is integrated into at least one of the sections 100. The equipment may be chosen from an accelerometer, a temperature sensor, a flow sensor, etc. The equipment may be a heating element and in particular a defrosting element.

[0072] The sections 100 may be made of a material chosen from aluminum, titanium, their alloy(s), and a composite material.

[0073] At least some sections 100 can be produced by additive manufacturing.

[0074] In a variant not shown, at least certain sections 100 are hollow or al veolar, and have for example at least one honeycomb portion

[0075] In a variant not shown, at least certain sections 100 comprise serrations, i.e. a series of teeth and / or notches, on part or all of at least one edge of the or each section.

[0076] The aims of this invention are multiple, they allow in particular:

[0077] - to limit the costs of replacing a stator blade in the event of wear of the type erosion for example,

[0078] - to limit the scrapping of parts in good condition through the use of removable portions (bar, sections),

[0079] - not to interfere with the assembly of adjacent parts.

Claims

Claims

1. Stator vane (40) for an aircraft turbomachine (10), this vane (40) comprising a blade (42) which extends from a first end connected to a platform (44) to a second end connected to fixing lugs (46), the blade (42) comprising a lower surface (42a) and an upper surface (42b) which are connected together at a leading edge (50) and a trailing edge (52) of the blade (42), the leading edge (50) being connected to the platform (44) at said first end and being at a distance from the fixing lugs (46) at said second end, characterized in that the blade (42) comprises a body (60) which is in one piece with the platform (44) and the fixing lugs (46), and a removable bar (70) which extends from the first end to the second end and which is attached and fixed on the body (60),the bar (70) being formed from a plurality of sections (100) arranged end to end along an axis (C) of elongation of the bar (70), one of the elements among the body (60) and the bar (70) comprising the leading edge (50), and the other of these elements comprising the trailing edge (52).,

2. Stator vane (40) according to claim 1, in which the bar (70) is mounted on the body (60) by a male-female engagement, the bar (70) comprising a male (72), or respectively female, element which extends along the axis (C) of elongation of the bar (70) and which cooperates with a female (74), or respectively male, element of the body (60) which extends between the first and second ends.

3. A stator vane (40) according to claim 2, wherein the bar (70) is mounted on the body (60) by a slide or dovetail system.

4. Stator blade (40) according to one of the preceding claims, in which the bar (60) defines a lower surface portion and an upper surface portion of the blade (42).

5. Stator blade (40) according to one of the preceding claims, in which the bar (70) comprises a first longitudinal end (70a) engaged in a recess (80) of the platform (44), and a second longitudinal end (70b) opposite fixed on the body (60) by screwing or bolting.

6. A stator blade (40) according to any preceding claim, wherein the first end (70a) of the bar (70) defines a portion of the platform (44).

7. Stator blade (40) according to claim 5 or 6, wherein the second end (70b) of the bar (70) comprises at least one orifice (84) for the passage of a screw (86) or a bolt.

8. Stator blade (40) according to claim 7, wherein the screw (86) or bolt extends parallel or perpendicular to the axis (C) of elongation of the bar (70).

9. Stator blade (40) according to one of the preceding claims, in which the sections (100) comprise pins (102) and / or holes (104) configured to cooperate by interlocking with complementary means of the other sections (100) in order to align the sections (100) with respect to each other along the elongation axis (C).

10. Stator blade (40) according to one of the preceding claims, in which at least one damping layer (110) is interposed between the bar (70) and the body (60).

11. A stator blade (40) according to claim 10, wherein the damping layer (110) is made of an elastically deformable material.

12. Aircraft turbomachine (10), comprising at least one stator blade (40) according to one of the preceding claims.

13. Turbomachine (10) according to claim 12, comprising a first flow vein (VI) of a primary flow (F1) inside a gas generator (12) comprising at least one compressor (14, 16), a combustion chamber and at least one turbine, and a second flow vein (V2) of a secondary flow (F2) around the gas generator (12), a fan (23) being arranged upstream of the first and second veins (VI, V2) and stator vanes (40) according to one of claims 1 to 11 being arranged in the second vein (V2).

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