LABYRINTH SEALING DEVICE FOR AN AIRCRAFT TURBOMACHINE

Abstract

A labyrinth-type sealing device for an aircraft turbomachine, this device comprising: - a rotor including at least one external annular flap (12a, 12b), and - a stator including an annular coating (16) of abradable material extending around the flap (12a, 12b) and configured to cooperate in operation with the flap (12a, 12b) to form a labyrinth-type seal against a gas flow that flows axially from upstream to downstream through the rotor, the coating (16) comprising: - an internal cylindrical surface (30a, 30b) extending around the flap (12a, 12b) and adapted to come into contact with the flap to form by friction an annular groove, and, upstream of the flap (12a, 12b), a preformed annular groove (32a, 32b). Figure for the summary: Figure 5

Description

Description Title of the invention: SEALING DEVICE FOR THE- BYRINTHE FOR A CHINA TURBOMA AIRCRAFT Technical field of the invention

[0001] — The present invention relates to a labyrinth-type sealing device for an aircraft turbomachine, as well as an aircraft turbomachine comprising such device. Technical background

[0002] The technical background includes, in particular, documents FR-A1-2 825 411, FR- A1-3 071 540, FR-A1-3 072 413.

[0003] — In an aircraft turbomachine, a labyrinth-type sealing device typically includes a rotor comprising at least one external annular blade, and a stator comprising an annular coating of abradable material which extends around the lick and which is configured to cooperate in operation with the lick. The friction of the lick on the coating creates an annular groove in The coating and the lick are designed to be housed in this groove to reduce the play between the rotor and the stator, thus forming a seal between the rotor and the stator by relative to a gas flow that flows axially in the turbomachine and through the rotor.

[0004] — In the present application, upstream and downstream are defined with respect to the direction normal flow of gas streams (from upstream to downstream) in the turbomachine. This The flow occurs along an axis of the turbomachine, which is the axis of rotation of the rotor. The axial direction corresponds to the direction of the turbomachine's axis, and a Radial direction is a direction perpendicular to the axis of the turbomachine and cutting this axis. Similarly, an axial plane is a plane containing the axis of the turbomachine, and a A radial plane is a plane perpendicular to this axis. The adjectives “interior” and “exterior” are used with reference to a radial direction so that the inner part of a element is, along a radial direction, closer to the axis of the turbomachine than the outer part of the same element.

[0005] — An aircraft turbomachine may include one or more devices sealing of the aforementioned type, for example in a compressor or turbine of the turbomachine.

[0006] For example, a sealing device is commonly used on the periphery of a rotor blade of a compressor or turbine (conventional or counter-rotating) for example), as illustrated in [Fig. 1]. This blade 10 has at its outer periphery one or more annular licks 12 which are oriented radially outwards and can be axially inclined. During the rotation of the blade 10, these shims 12 form grooves 14 in the abradable coating 16 fixed to a housing 18 (see Figures 1 and 2). This coating 16 is generally in the form of a honeycomb structure (called "Nida") and has radially oriented cells. This structure provides axial sealing while reducing cutting resistance in the tangential direction. Once the grooves 14 are formed in the coating 16, the clearances between the shims 12 and the coating 16 can be controlled by a housing ventilation system 20 in the case of a turbine, in order to minimize the clearances between the rotor and the stator, and thus reduce leakage and improve the performance of the turbine. When the blades 12 come into contact with the coating 16, during the formation of the grooves 14, a shearing and / or friction force is generated. This tangential force is applied to the blades 22 of the blade assembly 10 and can vary depending on the hardness of the coating 16 and the blades 12, the size of the contact surfaces, and the penetration speed of the blades 12 into the coating 16. The tangential force applied to the blades 12 of the blades 22 causes the blade to bend, resulting in a camber effect and an increase in the radius of the tip of the blades 12. As the radius of the blades 14 increases, the penetration into the coating 16, and therefore the tangential force, also increases. This phenomenon is called self-engagement of the blades 22 and is schematically illustrated by arrow F1 in [Fig. 2]. The geometry of the blades 22, and in particular the angle and curvature of their blades, results in an upstream displacement of the blade tips when a purely tangential force is applied. Thus, in addition to an increase in the radius of the blade tips 14, an advance of these tips relative to the coating 16 occurs when a tangential force is applied (see arrow F2 in [Fig. 3]). The blade tips 12 therefore reach an area of ​​the coating 16 that has not yet been worn, which further increases the applied tangential force. A divergent phenomenon therefore takes place which only stops when the blades 12 are no longer in contact with the re- clothing 16, or when an axial contact with another part stops the advance of the blade 22 (contact for example in C1 at [Fig.4] between a heel 24 of the blade 22 and a distributor 26 located upstream), a break in one or more parts, or a stoppage of the turbomachine.This phenomenon is called "wobble" of blade 10. When wobble occurs, it can lead to damage in the blades 22 and rotor-stator contact between the blades 22 and the distributors 26. This damage results in premature disassembly and replacement of parts, and therefore very costly maintenance operations. Several solutions have been implemented to address this technical problem. but are not entirely satisfactory. The present invention offers a simple, effective and economical solution to this problem. Summary of the invention The invention relates to a labyrinth-type sealing device for an aircraft turbomachine, this device comprising: - a rotor having an axis of rotation and comprising at least one external annular blade extending around said axis, and - a stator extending around the axis, this stator comprising an annular coating of abradable material which extends around the blade and which is configured to cooperate in operation with the blade to form a labyrinth-type seal against a gas flow intended to flow axially from upstream to downstream through the rotor, the coating having an internal cylindrical surface extending around the blade and adapted to come into contact with the blade to form by friction an annular groove extending around the axis, characterized in that said surface comprises, upstream of the slat, a pre-formed annular groove which extends around the axis and which is capable of receiving said slat in case of axial displacement upstream of the rotor with respect to the stator. In this application, a distinction is made between a groove formed by the friction of a squeegee against the coating during the operation of the turbomachine, and a groove pre-formed in the coating during its manufacture. In other words, if the device includes a new coating (never before used in a turbomachine), this coating would include only one or more grooves. If the device is already used in a turbomachine, its coating would include at least one groove and at least one squeegee. The groove formed by a blade, in the rotor's normal operating position relative to the stator, is located directly above the blade or its outer edge. That is to say, the groove and the blade (or its edge) lie in the same plane perpendicular to the axis. Conversely, the groove is located upstream of the groove and the blade and is preferably situated at a predetermined axial distance to avoid or limit the aforementioned wobble. In the aforementioned case where the rotor or part of the rotor were to move axially upstream, the blade would become lodged in the pre-formed groove. This groove would thus interrupt the contact between the blade and the coating and slow down or stop this axial movement, thereby limiting the wobble. In this way, the risk of rotor damage and rotor-stator contact would be avoided. The device according to the invention may include one or more of the following features the following, taken independently of each other or in combination with each other: - the rotor includes at least one upstream scraper and one downstream scraper, the upstream and downstream scrapers being surrounded by the same cylindrical surface of the coating or by two distinct cylindrical surfaces, respectively upstream and downstream, of the coating; - the cylindrical surface which extends around the upstream and downstream sluices includes a pre-formed groove upstream of the upstream sluice, and / or a pre-formed groove upstream of the downstream sluice; - the upstream cylindrical surface includes a pre-formed groove upstream of the upstream slit, and / or the downstream cylindrical surface includes a pre-formed groove upstream of the downstream slit; - the upstream and downstream cylindrical surfaces are stepped and include different diameters; - the groove is axially interposed between the two cylindrical surfaces; - where each groove is located at an axial distance upstream of the corresponding lick, which is less than or equal to half the axial distance between the two licks; -- Alternatively, the groove or each groove is located at an axial distance upstream of the corresponding lick, which is greater than half the axial distance between the two licks; -- the axial distance is between 1 and 20mm, and preferably between 1 and 5mm; or even more on some large motors; - where each groove has a depth or radial dimension less than or equal to a thickness or radial dimension of an area of ​​the coating in which this groove is formed; - where each groove has an axial dimension greater than or equal to half of an axial thickness of the corresponding slit, this thickness being measured at the level of an external periphery of this slit; - the number of pre-formed throats is less than or equal to the number of licks; -- the number of grooves and channels is greater than the number of licks; - where each groove is located at an axial distance from an upstream axial end of the cylindrical surface in which this groove is formed; - the rotor is a compressor or turbine blade, and the stator is a sealing ring carried by a casing of this compressor or turbine. The invention also relates to an aircraft turbomachine, comprising at least one device as described above. Brief description of the figures Other features and advantages of the invention will become apparent upon reading the detailed description that follows, for an understanding of which reference should be made to the attached drawings in which: [Fig.1] Fig.1 is a partial schematic axial cross-sectional view of a sealing device for an aircraft turbomachine, according to the prior art of the present invention: [Fig.2] The [Fig.2] is a view similar to that of the [Fig.1] and illustrates the formation of grooves by licks of the device in an abradable coating of the device; [Fig.3] The [Fig.3] is a view similar to that of the [Fig.1] and illustrates the upstream axial displacement of the rotor relative to the stator of the device; [Fig.4] The [Fig.4] is a view similar to that of the [Fig.1] and illustrates the axial contact between the rotor of the device and a stator of the turbomachine; [Fig.5] The [Fig.5] is a partial schematic axial cross-sectional view of a sealing device of an aircraft turbomachine, according to a first embodiment of the invention; [Fig.6a] [Fig.6a] is a view similar to that of [Fig.5] and illustrates a resting position of the device; [Fig.6b] The [Fig.6b] is a view similar to that of the [Fig.5] and illustrates a first operating position of the device; [Fig.6c] The [Fig.6c] is a view similar to that of the [Fig.5] and illustrates a second operating position of the device; [Fig.6d] The [Fig.6d] is a view similar to that of the [Fig.5] and illustrates a third operating position of the device; [Fig.6e] The [Fig.6e] is a view similar to that of the [Fig.5] and illustrates a fourth operating position of the device; [Fig.7] The [Fig.7] is a view similar to that of the [Fig.5] and illustrates a second embodiment of the invention; [Fig.8] The [Fig.8] is a view similar to that of the [Fig.5] and illustrates a third embodiment of the invention; [Fig.9] The [Fig.9] is a view similar to that of the [Fig.5] and illustrates a fourth embodiment of the invention; [Fig.10] The [Fig.10] is a view similar to that of the [Fig.5] and illustrates a fifth embodiment of the invention; [Fig.11] The [Fig.11] is a view similar to that of the [Fig.5] and illustrates a sixth embodiment of the invention. Detailed description of the invention Figures 1 to 4 have been described above. Figure 5 illustrates a first embodiment of a sealing device according to the invention and Figures 6a to 6e illustrate the operation of this device. Fig. 5 is a partial view of a turbomachine and shows a compression stage of a compressor or an expansion stage of a turbine of that turbomachine. This stage includes a stator blade called distributor 26 and a rotor blade 10 which is located downstream of the distributor 26. The distributor 26 is fixed to a housing 18 which has an annular shape and extends around the stage. The rotor blade 10, also called the "rotor," is free to rotate about an axis that is not visible in the drawing. The blade 10 comprises a plurality of blades 22 and has at its outer periphery at least one annular scraper 12 oriented radially outwards. In the example shown, the rotor has two scrapers 12 located at an axial distance from each other and designated respectively upstream scraper 12a and downstream scraper 12b. The housing 18 extends around the blade 10 and carries a sealing ring 28 which can be divided into sections. This ring 28 includes an internal annular coating 16 which surrounds the outer periphery of the blade 10 and thus extends around the blades 12. The coating 16 is made of abradable material and includes, for example, a honeycomb structure, that is, a structure with cells that are preferably oriented radially with respect to the aforementioned axis. This type of structure is well known to those skilled in the art. Alternatively, the coating 16 could be solid. The coating 16 includes at least one internal cylindrical surface 30a, 30b which extends around the slats 12. The number of surfaces 30a, 30b can be equal to the number of slats 12, and can be two as in the example shown, so that each of the surfaces 30a, 30b extends around one of the slats 12. The coating 16 thus includes an upstream cylindrical surface 30a around the upstream slat 12a and a downstream cylindrical surface 30b around the downstream slat 12b. Each of the slats 12 is oriented radially outwards and can be inclined, for example from downstream to upstream radially outwards, as in the example shown. During operation, the swabs 12 can come into contact with the coating 16 and their surfaces 30a, 30b and create annular grooves 14 by friction and wear of the abradable material 16. The grooves 14 are therefore formed during operation and in particular during the first operation of the turbomachine. Surfaces 30a and 30b can be stepped and have different diameters. In the example shown, the upstream surface 30a has a smaller diameter than the downstream surface 30b. The coating 16 includes a groove (upstream) 14a which is formed in the surface upstream 30a, and a groove (downstream) 14b which is formed in the downstream surface 30b. In the normal operating position as illustrated in [Fig.5], the grooves 14a, 14b are located respectively at the level of the slats 12a, 12b and in particular their external peripheries, which means that the groove 14a and the slat 12a are located substantially in the same plane P1 perpendicular to the axis, and that the groove 14b and the slat 12b are located substantially in the same other plane P2 perpendicular to the axis. To avoid or limit the aforementioned weaving phenomenon, the invention proposes to preform at least one annular groove 32a, 32b in the coating 16, upstream of the or each slat 12. It is thus understood that this type of groove 32a, 32b is formed in the coating 16 during its manufacture and is therefore not generated during the operation of the turbomachine unlike the grooves 30a, 30b. In the embodiment of [Fig. 5], the coating 16 comprises two grooves 32a, 32b. That is to say, the number of grooves 32a, 32b is equal to the number of slits 12a, 12b. The coating 16 comprises an (upstream) groove 32a upstream of the upstream slit 12a and which is formed in the upstream surface 30a, and a (downstream) groove 32b upstream of the downstream slit 12b and which is formed in the downstream surface 30b. Figure 6a is similar to Figure 5 and illustrates the axial position of the rotor relative to the stator of the turbomachine at rest, before its first use. The blade 10 is radially spaced from the liner 16. During the initial operation of the turbomachine, the scrapers 12a, 12b rub against the coating 16 and wear it down (Figures 6b and 6c), thus creating the aforementioned annular grooves 14a, 14b. Figures 6a and 6b show that the grooves 32a, 32b are already present in the coating 16 and that the grooves 14a, 14b are distinct from these grooves 32a, 32b. During a weaving phenomenon, the blades 22 of the blade 10 tend to move upstream ([Fig. 6d]). The scrapers 12a, 12b then widen the grooves 14a, 14b, further wearing down the coating 16 in the axial direction. The grooves 32a, 32b are located upstream and at a predetermined distance from the grooves 14a, 14b and the scrapers 12a, 12b so as to stop the axial movement of the rotor as early as possible. When the slats 12a, 12b reach the grooves 32a, 32b, there is no longer any contact between the slats 12a, 12b and the coating 16, which slows down and stops the axial displacement of the rotor, thus preventing axial contact between the rotor and the stator ( [Fig.6e]). The grooves 32a, 32b therefore have the function of receiving the scrapers 12a, 12b in case of axial displacement upstream of the rotor with respect to the stator, in order to avoid or limit the phenomenon of swaying. Preferably, the groove or each groove 32a, 32b is located at an axial distance L1 upstream of the corresponding slit 12a, 12b, which is less than or equal to half the axial distance L2 between the two slits 12a, 12b (see [Fig. 6a]). L1 is preferably between 1 and 5 mm. Each groove 32a, 32b has an axial dimension El which is preferably greater than or equal to half of an axial thickness E2 of the corresponding lick 12a, 12b, this thickness E2 being measured at the level of an external periphery of this lick (cf. [Fig.6a]). Each groove 32a, 32b can have any shape and for example have a square or rectangular cross-section, as in the illustrated example, or be round, oval, etc. The bottom of the or each groove 32a, 32b, that is to say the wall located at the bottom of the or each groove, can have any shape, flat, curved, or otherwise. Figures 7 and 8 illustrate alternative embodiments of the device in which the number of pre-formed grooves in the coating 16 is less than the number of slats 12a, 12b, and is in this case one. In [Fig. 7], the coating 16 includes an upstream groove 32a formed in the upstream surface 30a upstream of the upstream slat 12a. In [Fig. 8], the coating 16 includes a downstream groove 32b formed in the downstream surface 30b upstream of the downstream slat 12b. These variants show that a single groove can suffice to receive one of the slats and prevent the wobbling phenomenon. Indeed, even if only one of the two licks 12a, 12b is received in a pre-formed groove, removing the contact between this lick and the coating 16 may be enough to stop the weaving phenomenon even if the other lick remains in contact with the coating 16. Figures 9 and 10 illustrate alternative embodiments of the device in which the dimensions of the grooves 32a, 32b are adjusted. In the case of [Fig. 9], the axial dimensions E3 of the grooves 32a, 32b are larger than those of the grooves in [Fig. 6a]. E3 is, for example, greater than or equal to twice E2. In the case of [Fig. 10], the depths H1, H2, or radial dimensions of the grooves 32a, 32b, are greater than those of the grooves in [Fig. 6a]. Furthermore, the grooves here extend across the entire thickness of the coating 16. Since the surfaces 30a, 30b do not have the same diameter, and the coating 16 includes an external cylindrical surface of constant diameter surrounding these surfaces 30a, 30b, the coating 16 does not have the same radial thickness along its entire axial extent. The groove 32a has a depth H1 greater than the depth H2 of the groove 32b. In the case of [Fig. 10], it can be seen that the coating 16 can be formed by the arrangement of three independent and successive annular blocks B1, B2, and B3, placed axially one behind the other and at an axial distance from each other corresponding to the axial dimension of the grooves 32a, 32b. Block B1 comprises Block B1, B2, and B3 comprise an upstream portion of surface 30a, block B2 includes a downstream portion of surface 30a and an upstream portion of surface 30b, and block B3 includes a downstream portion of surface 30b. Blocks B1, B2, and B3 can be divided into sectors. Figure 11 differs from Figure 10 in that grooves 32a and 32b are positioned even further upstream compared to the grooves in Figure 10. Groove 32b is thus located at the upstream axial end of surface 30b and is axially interposed between surfaces 30a and 30b. Groove 32a is located at the upstream axial end of surface 30a. It can be seen that the coating 16 can be formed by the arrangement of two independent and successive annular blocks B1 and B2 placed axially one behind the other and at an axial distance from each other corresponding to the axial dimension of the groove 32b. Block B1 comprises the surface 30a and block B2 comprises the surface 30b. Blocks B1 and B2 can be divided into sectors. The groove(s) 32a, 32b can be formed in a coating by machining, for example. In the case of Figures 10 and 11, this machining step can be eliminated by using and positioning the aforementioned blocks.

Claims

Demands

1. Labyrinth-type sealing device for a turbomachine of aircraft, this device comprising: - a rotor having a rotation axis and comprising at least one blade outer annular (12a, 12b) extending around said axis, and - a stator extending around the axis, this stator comprising a re- annular garment (16) of abradable material that extends around the lick (12a, 12b) and which is configured to cooperate in operation with the lick (12a, 12b) to form a seal of the type to the- byrinth in relation to a gas flow that is intended to flow axially from upstream to downstream through the rotor, the coating (16) comprising an internal cylindrical surface (30a, 30b) extending around of the lick (12a, 12b) and able to come into contact with the lick for forming, through friction, an annular groove extending around the axis, characterized in that said surface (30a, 30b) comprises, upstream of the lick (12a, 12b), a preformed annular groove (32a, 32b) which extends around the axis and is suitable for receiving said lick (12a, 12b) in the event of axial displacement upstream of the rotor with respect to the stator

2. Device according to claim 1, wherein the rotor comprises at minus one upstream lick (12a) and one downstream lick (2b), the licks upstream and downstream (12a, 12b) being surrounded by the same cy- surface lindric (30) of the coating (16) or by two cylindrical surfaces distinct, respectively upstream (30a) and downstream (30b), of the coating (16).

3. Device according to claim 2, wherein: - in the case where the upstream and downstream licks (12a, 12b) are surrounded by a single cylindrical surface (30) of the coating (16), this cy- surface Lindrique (30) includes a pre-formed groove (32a) upstream of the upstream lick (12a), and / or a pre-formed gorge (32b) upstream of the downstream lick (12b), or - in the case where the upstream and downstream licks (12a, 12b) are surrounded by two distinct cylindrical surfaces, respectively upstream (30a) and downstream (30b) of the coating (16), the upstream cylindrical surface (30a) includes a pre-formed groove (32a) upstream of the upstream lick (12a), and / or the downstream cylindrical surface (30b) includes a groove (32b) preformed upstream of the downstream lick (12b).

4. Device according to claim 2 or 3, wherein the surfaces cy- The upstream and downstream indrigs (30a, 30b) are stepped and include different diameters.

5. Device according to claim 4, wherein the groove (32b) is in- axially intercalated between the two cylindrical surfaces (30a, 30b).

6. Device according to any one of claims 2 to 5, wherein the or each gorge (32a, 32b) is located at an axial distance (LI) upstream of the corresponding lick (12a, 12b), which is less than or equal to the half of the axial distance (L2) between the two licks (12a, 12b).

7. Device according to any one of the preceding claims, wherein the or each groove (32a, 32b) has a depth or radial dimension (H1) in- less than or equal to a thickness or radial dimension of a zone of the re- garment (16) in which this groove (32a, 32b) is formed.

8. Device according to any one of the preceding claims, in which the or each groove (32a, 32b) has an axial dimension (El) greater than or equal to half of an axial thickness (E2) of the lick (12a, 12b) corresponding, this thickness being measured at the level of a pe- external rim of this lick.

9. Device according to any one of the preceding claims, in which the The number of pre-formed grooves (32a, 32b) is less than or equal to the number of licks (12a, 12b).

10. A device according to any one of claims 1 to 3, and 5 to 7, wherein the or each groove (32a, 32b) is located at an axial distance from one end upstream axial of the cylindrical surface (30a, 30b) in which is formed this gorge (32a, 32b).

11. Device according to any one of the preceding claims, wherein the The rotor is a compressor or turbine blade (10), and the stator is a sealing ring (28) carried by a housing (18) of this compressor or of this turbine.

12. Aircraft turbomachine, comprising at least one device according to one any of the preceding claims.

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