Sealing device for turbomachine

Abstract

Title: Sealing device for a labyrinth seal (22) of an aircraft turbomachine (1), said device comprising a first element and a second element, at least one of the first element and the second element being rotatable about a longitudinal axis (X), the first element comprising an abradable coating (27), the second element comprising a flap (26), the flap (26) and the abradable coating (27) being arranged to achieve a seal during relative rotation between the first element and the second element, characterized in that the first element comprises a frame (29) and a ring (28), the ring (28) being mounted on the frame (29), the abradable coating (27) being arranged on the ring (28). Figure for the abstract: Fig. 3.

Description

Title of the invention: Sealing device for turbomachinery technical field

[0001] The invention relates to the field of turbomachinery and, more particularly, to the field of sealing in turbomachinery. In particular, the invention relates to the field of labyrinth seals in turbomachinery. Prior art

[0002] It is known to form a sealing labyrinth, also called a labyrinth seal, between the rotor and the stator of a turbomachine. Such a labyrinth seal comprises fins, called vanes, generally located on the rotor and directed towards a face of the stator covered by a layer of abradable material. By abradable material, we mean here that the material is designed to wear away by abrasion upon contact with the vanes.

[0003] The flaps provide aerodynamic seals between air chambers under different pressures. For example, flaps are located at the top of turbine blades. These flaps can also provide sealing in various engine cavities, for example, to provide pressure isolation of a cavity under a low-pressure turbine runner, to prevent leaks of air, oil, or an air-oil mixture.

[0004] With wear, it may be necessary to replace the abradable material when it has degraded in order to maintain satisfactory sealing properties of the labyrinth seal. However, such a replacement of the abradable material requires immobilizing the part on which the abradable material to be replaced is mounted. This immobilization time corresponds to the time required for the degraded abradable material to be removed and for a new abradable material meeting the requirements of the corresponding labyrinth seal to be formed on the part.

[0005] Furthermore, depending on the labyrinth seals involved, the process of forming the new abradable material on the workpiece can be lengthy and complex. For example, with thermally sprayed abradable materials, the process of spraying the abradable material onto the workpiece requires complex spraying tooling with specific nozzles. Such a process is also time-consuming to implement. Replacing an abradable material with such a degraded process therefore requires specialized tooling and negatively impacts the downtime of the workpiece being repaired.

[0006] There is therefore a need to overcome at least some of these disadvantages and in particular to facilitate the repair of a labyrinth seal and limit the downtime of a part of such a labyrinth seal. Description of the invention

[0007] One idea underlying the invention is to enable the repair of a labyrinth seal in a turbomachinery in a simple and rapid manner. In particular, one idea underlying the invention is to limit the downtime of a component of said labyrinth seal during its repair. More specifically, one idea underlying the invention is to enable the replacement of the abradable material of a labyrinth seal in a simple and rapid manner. To this end, the invention provides a sealing device for an aircraft turbomachine labyrinth seal, said device comprising a first element, preferably annular, and a second element, preferably annular, at least one of the first element and the second element being rotationally movable about a longitudinal axis, the first element comprising an abradable coating, the second element comprising a slat, the slat and the abradable coating being arranged radially opposite each other so as to achieve a seal during a relative rotation between the first element and the second element, in which the first element comprises a armature and a ring, the ring being mounted on the armature, the abradable coating being arranged on the ring.

[0008] Thanks to these features, the invention provides a labyrinth-type sealing device that can be easily maintained or repaired. Indeed, in the event of degradation of the abradable coating, the invention makes it possible to simply remove the ring on which the abradable coating is mounted from the frame and replace it with another ring having a satisfactory abradable coating.

[0009] Thus, it is not necessary to treat the entire first element for repair when it has a degraded abradable coating. The downtime of the first element for repair can therefore be limited to the time required to remove the ring and insert a new one, without needing the time to replace the abradable coating on the removed ring. Indeed, it is possible to maintain a reserve of rings that can be used for replacement, the repair of a ring with a degraded abradable coating then being independent of the repair of the first element with such a ring.

[0010] Such a replacement of the ring can also be carried out simply and quickly in many workshops, without requiring the tools necessary for the removal and replacement of the abradable material.

[0011] According to different embodiments, such a sealing device may have one or more of the characteristics below, alone or in combination.

[0012] The first element and the second element can be any element of a turbomachine jointly forming at least one labyrinth-type joint. According to In one embodiment, the first element is a turbomachine stator, or part of a turbomachine stator. In another embodiment, the second element is a turbomachine rotor, or part of a turbomachine rotor, for example, a rotor shaft with one or more blades arranged at its end. In another embodiment, the first element is a turbomachine rotor, or part of a turbomachine rotor. In another embodiment, the second element is a turbomachine stator, or part of a turbomachine stator.

[0013] The armature of the first element can also take many forms. According to a preferred embodiment, the armature is a turbomachine ejection tube support. Such a turbomachine ejection tube support comprises, for example, two cylindrical portions of different diameters and a frustoconical portion connecting the two cylindrical portions. Thus, in the context of such a turbomachine ejection tube support with an abradable coating, if said abradable coating deteriorates, it is possible to repair the ejection tube support simply by removing the ring on which the deteriorated abradable coating is mounted and installing a new ring with a satisfactory abradable coating on the ejection tube support.

[0014] The abradable coating can take many forms. Thus, according to one embodiment, the abradable coating is a sprayed-type abradable coating, for example of the AlSi-polyester and / or ANi type. According to another embodiment, the abradable material is of the honeycomb type.

[0015] The interaction between the nozzle and the abradable coating is of the labyrinth seal type. That is to say, the nozzle penetrates the abradable coating in such a way as to form a groove in said coating and force a flow through the turbomachine to take a longer path. For example, in the context of a sealing device between engine cavities of a turbomachine, the nozzle, together with the abradable coating, ensures the seal between these engine cavities during the flight phase.

[0016] The ring is preferably fixed to the armature. In other words, the ring and the armature are rotationally fixed around the longitudinal axis, while the former is itself rotationally free. This fixing of the ring can be achieved by any means ensuring rotational rigidity between the ring and the armature.

[0017] According to a preferred embodiment, the ring is mounted on the armature by shrink fitting. In other words, the ring is a shrink-fitted ring. Such shrink fitting ensures the absence of relative movement between the ring and the armature. Furthermore, such shrink fitting allows for the quick and easy removal of the ring to be replaced by simply pulling on said ring.

[0018] According to one embodiment, the ring is mounted radially inside the frame and preferably is mounted by shrink fitting in the frame.

[0019] In one embodiment, the armature has a support surface, the ring being mounted on said support surface. Preferably, this support surface, i.e., the surface of cooperation with the ring, is circular. In one embodiment, this support surface is coaxial with the longitudinal axis of the turbomachine. In another embodiment, the support surface extends over a portion of, and preferably the entire circumference of, the turbomachine. Such a support surface may be continuous or discontinuous. Preferably, the support surface extends over at least one-third of the circumference of the armature.

[0020] According to one embodiment, the support surface is an internal surface of the reinforcement, that is, one defining an internal diameter of the reinforcement. In this case, the ring has an external surface defining an external diameter of the ring, the ring being mounted on the reinforcement by cooperation between the support surface and said external surface of the ring. For example, the ring is mounted on the reinforcement by interference between the external surface of the ring and the support surface. In this case, the abradable coating is mounted on an internal surface of the ring, that is, one radially opposite to the external surface of the ring.

[0021] According to one embodiment, the ring has a groove, the abradable coating being housed in said groove.

[0022] Such a groove allows for a satisfactory thickness of abradable coating to interact with the lick without generating excess thickness on the ring.

[0023] According to one embodiment, the abradable coating has a radial thickness equal to the radial depth of the groove. In the case of an abradable coating arranged on the inner surface of the ring, the abradable coating is flush with the inner surface of the ring adjacent to the groove. Thus, the presence of the abradable coating does not generate disturbances and, in cooperation with the nozzle, ensures a good seal in the turbomachine.

[0024] According to one embodiment, one end, preferably axial, of the ring has a collar.

[0025] Such a collar provides a satisfactory gripping surface for installing or removing the ring from the armature. Indeed, simply using a suitable gripping tool to grasp the collar is sufficient for manipulating the ring.

[0026] According to one embodiment, the collar is offset along the longitudinal axis relative to the reinforcement, and more particularly relative to the support surface. Preferably, the collar extends radially from the axial end of the ring and protrudes radially beyond a portion of the reinforcement forming the support surface. In other words, the collar is housed outside the reinforcement so as to allow easy access with a ring manipulation tool. By For example, in the context of a support surface formed by an internal surface, the collar develops radially outwards and protrudes radially outwards beyond the portion of the reinforcement forming the support surface.

[0027] According to one embodiment, one end, preferably radial, of the collar has a rib projecting axially towards the reinforcement.

[0028] Such a rib allows for better gripping of the collar with a suitable tool, and therefore simplified and reliable installation or removal of the ring.

[0029] The invention also provides a method for manufacturing a sealing device for an aircraft turbomachine labyrinth seal, said sealing device comprising a first element and a second element movable in relative rotation about a longitudinal axis, the first element comprising a reinforcement, the second element comprising a flap, the method comprising the steps of - Provide a ring with an abradable coating. - Attach the ring to the frame so that the abradable coating either radially opposite the slit and so that the slit creates a seal with the abradable coating.

[0030] Such a method may further comprise one or more of the features below, taken alone or in combination. Moreover, the sealing device manufactured via this method may comprise one or more of the features of the sealing device as described above, alone or in combination.

[0031] According to one embodiment, the step of fixing the ring on the frame includes a step of shrink-fitting the ring into the frame.

[0032] According to one embodiment, the shrink-fitting of the ring in the frame comprises successive steps of: - to thermally contract, for example by immersing in liquid nitrogen, the ring - insert the ring in a thermally contracted state into the armature, and - bring the ring back to a temperature, for example to room temperature, at which the ring has a state of thermal contraction in which the cooperation between the ring and the armature takes place with an interference of material capable of fixing, and therefore securing in rotation, the ring on the armature.

[0033] According to one embodiment, the method further comprises a step of projecting abradable material onto the ring, for example in a groove of the ring.

[0034] According to one embodiment, the process further comprises a step of applying an undercoat to the ring, for example by projection, the abradable coating being applied, for example by projection, to said undercoat.

[0035] Such an underlayer offers a satisfactory roughness to allow the fixation of the abradable coating, in particular in the context of a sprayed abradable coating.

[0036] The invention also provides an aircraft turbomachine comprising a sealing device such as above. Brief description of the drawings

[0037] The invention and its advantages will be better understood upon reading the following detailed description of various embodiments of the invention, given by way of non-limiting examples. This description refers to the accompanying figure pages, on which:

[0038] [Fig-1] Fig. 1 represents a schematic longitudinal cross-sectional view of a turbomachine,

[0039] [Fig.2] Fig.2 represents a schematic longitudinal and partial cross-sectional view of a low-pressure turbine of the turbomachine of Fig.1,

[0040] [Fig.3] Fig.3 represents a schematic longitudinal cross-sectional view of a first labyrinth seal element according to the invention, said first element comprising an ejection tube support on which a shrink-fitted ring is mounted,

[0041] [Fig.4] [Fig.4] represents a schematic longitudinal and partial cross-sectional view of a first labyrinth joint element as illustrated in [Fig.3] in cooperation with a second element of said labyrinth joint,

[0042] [Fig.5] The [Fig.5] represents a schematic perspective view of a labyrinth seal ring according to an embodiment of the invention.

[0043] [Fig.6] The [Fig.6] a schematic perspective view of the labyrinth seal ring of the [Fig.3]. Detailed description of the invention

[0044] In this description, the terms "inside," "internal," "outside," and "external" are defined with respect to a longitudinal axis, typically the longitudinal axis of the turbomachine around which elements of said turbomachine are rotated. Thus, an element described as "internal" or "internal" is close to or oriented toward said longitudinal axis relative to, or in contrast to, an element described as "external" or "external." Similarly, the terms "axial," "radial," and their derivatives are defined with respect to the principal axis of the turbomachine. Furthermore, the terms "upstream" and "downstream" are subsequently defined with respect to the direction of gas flow through a turbomachine, indicated by arrow G in Figures 1 and 2.

[0045] Fig. 1 illustrates a double-flow turbomachine 1 comprising, in a known manner from upstream to downstream successively, at least one blower 2, an engine part comprising successively at least one stage of low-pressure compressor 3, high-pressure compressor 4, a combustion chamber 5, at least one stage of high-pressure turbine 6 and low-pressure turbine 7.

[0046] Rotors, rotating around the main axis X of the turbomachine 1 and able to be coupled together by different transmission and gear systems, correspond to these different elements.

[0047] In a known manner, a fraction of air is taken from the high-pressure compressor 4 and is conveyed through a cooling duct 8 in order to cool hotter areas of the turbomachine 1, in particular the high-pressure turbine 6 and the low-pressure turbine 7.

[0048] Fig. 2 is an enlargement of an area of ​​the turbomachine 1, illustrating in a simplified way the upstream part of the low pressure turbine 7, the high pressure turbine 6 not being shown.

[0049] The low-pressure turbine 7 illustrated here comprises a plurality of turbine stages 9, 10. A first stage 9, as well as the stages 10 located downstream of it, each comprise a set of fixed distributors 11 and 12, respectively. Each stage 9, 10 further comprises a movable disk 13 on which is mounted a set of blades 14 driven in rotation by the movable disk 13. In the example illustrated in [Fig. 2], the distributor 11 forms a single piece with a housing 15 constituting the turbine and is hollow to allow the passage of cooling air, exiting via an injection device 16 associated with the distributor 11, comprising a plurality of injectors. The movable disk 13 is rotationally fixed to a low-pressure shaft 17 extending along the axis XX, while each distributor 12 is connected to the housing 15.

[0050] In the embodiment illustrated in [Fig. 2], a fraction of air taken from the high-pressure compressor 4 flows into the cooling duct 8, then into the hollow distributor 11. The direction of flow of the air fraction through the hollow distributor 11 is illustrated by the arrows 18. The air fraction is then injected via the injection device 16 into a sub-flow cavity 19. The distributed air serves, in particular, to cool the turbine discs 13, as illustrated by the arrows 20.

[0051] The cooling air injected by the injection device 16 also allows for the purging of hot air present in the low-pressure turbine 7, thus ensuring its cooling. More precisely, the cooling air taken from the high-pressure compressor 4 and routed to the cavity under the flow 19 constitutes a pressure barrier, or purge, preventing hot air from the combustion chamber and flowing in the main air circulation flow of the turbines, i.e., in the primary air circulation flow of the turbomachine 1, from entering the cavity under the flow 19. The purging of hot air from the low-pressure turbine 7 is symbolized here by arrow 21. The risks of overheating of the turbine rotors are thus limited. In particular, by preventing air from the primary flow to enter the cavity under vein 19, this cavity is less hot than the vein, and the turbine rotor can therefore withstand higher centrifugal forces and be dimensioned on lower limit stresses.

[0052] A malfunction in the cooling of the turbine 7 can have several causes. One cause of the cooling malfunction may result from excessive wear or rupture of one or more seals, or dynamic seals, of the low-pressure turbine 7. For example, a malfunction in the cooling of the turbine 7 results from the failure of a labyrinth seal 22 that provides pressure isolation of the cavity under the flow 19 of the low-pressure turbine 7. Labyrinth seals 22 are installed at various locations in the turbomachine 1 to provide sealing between different engine cavities and to prevent leaks of air, oil, or an air / oil mixture. Thus, a labyrinth seal 22 is illustrated in [Fig. 2] at an ejection tube support receiving a rotor shaft.Such an ejection tube support allows for the support of an ejection tube which passes through the inside of an arm of an exhaust housing and blows air from the low-pressure compressor 3 or the high-pressure compressor 4 to create a suction in order to evacuate the air into an upstream cavity at the level of the low-pressure compressor 3 to prevent the oil in this area from accumulating and coking.

[0053] Figure 3 illustrates a first labyrinth seal element 22 according to an embodiment of the invention. This first element comprises a frame 29 and a ring 28. In the embodiment illustrated in Figure 3, the frame 29 is formed by an ejection tube support as illustrated in Figure 2. The ring 28 is mounted on the frame 29. More particularly, the ejection tube support has an internal tubular portion 25 for receiving a rotor shaft (see Figure 3), and the ring 28 is press-fitted into this internal tubular portion 25. The first element further comprises an abradable coating.

[0054] Figure 4 illustrates a labyrinth seal 22 formed by a first element of the turbomachine 1, for example a stator part 23 comprising an ejection tube support as illustrated in Figure 3, and a second element of the turbomachine 1, for example a rotor shaft 24, said first and second elements being mounted in relative rotation about the main axis X of the turbomachine 1. In the embodiment illustrated in Figures 3 and 4, the first element comprises an ejection tube support of the turbomachine 1, but the invention applies to any labyrinth seal 22 present in the turbomachine 1 and comprising a stator and / or rotor part.

[0055] As illustrated in [Fig. 3], the labyrinth seal 22 has a wiper 26 / abradable coating 27 interface. More specifically, the first element of the labyrinth seal 22 is formed by the stator portion 23 on which the coating is mounted abradable 27 and the second element of the labyrinth seal 22 is formed by the rotor part 24 on which the blades 26 are mounted.

[0056] The stator part 23 of the second element includes the ring 28 mounted on the armature 29, hereinafter referred to as the ejection tube support 29 in the context of the example illustrated in Figures 3 to 6. The abradable coating 27 is carried by the stator 23 via the ring 28, as explained in more detail below.

[0057] In such a labyrinth seal 22, the slit 26 penetrates the abradable coating 27 to create a groove in said abradable coating 27 in order to force the air passing through the labyrinth seal 22 to take a longer path. The slit 26 is also close to and opposite, typically directly opposite, the abradable coating 27 during the aircraft's cruise flight phase to limit air leaks and / or ensure sealing between the engine cavities of the turbomachine 1, thus improving the performance of the turbomachine 1 and / or preventing oil leaks by sealing cavities immersed in oil.

[0058] The ejection tube support 29, and more particularly the internal tubular portion 25, has a cylindrical support surface 30 of revolution defining an internal diameter 31. The ring 28 is press-fitted onto the ejection tube support 29 by being mounted on said support surface 30. For this purpose, the ring 28 has an external surface 32 having, at ambient temperature, an external diameter 33 greater than the internal diameter 31 of the support surface 30. In other words, the ring 28 is mounted on the support surface 30 with interference between the support surface 30 and the external surface 32. This mounting with interference between the ring 28 and the ejection tube support 29 ensures that the ring 28 is fixedly mounted on the ejection tube support 29 and can only be removed with suitable tools.

[0059] In this embodiment, the abradable coating 27 is mounted on a circular cylindrical internal surface 34 of the ring 28.

[0060] In order to mount the ring 28 in the ejection tube support 29, the ring 28, which carries the abradable coating 27, is immersed in liquid nitrogen so that, in a contracted state, it has an external diameter 33 smaller than the internal diameter 31 of the support surface 30. This simplifies the insertion of the ring 28 into the ejection tube support 29. When the ring 28 returns to its normal external diameter 33, typically at room temperature, the external surface 32 of the ring 28 interferes with the support surface 30.

[0061] To facilitate the installation and removal of the ring 28 on the ejection tube support 29, the ring 28 has a flange 35. This flange 35 extends radially outwards from an axial end 36 of the ring 28 arranged outside the ejection tube support 29 in the mounted state of the ring 28. This flange 35 further has a rib 37 projecting along the longitudinal direction X towards the surface support 30. The collar 35 and the rib 37 allow simple and quick cooperation with a gripping tool (not shown) allowing manipulation of the ring 28 when inserting it into or removing it from the ejection tube support 29.

[0062] The abradable coating 27 may exhibit over time different forms of degradation, for example flaking or an excessive groove depth, i.e. exceeding the required specifications.

[0063] The press-fit ring 28 allows, in the event of degradation of the abradable coating 27, to simply remove the press-fit ring 28 and send this single ring 28 to the workshop for replacement of the abradable coating 27. In addition, the ejection tube support 29 on which the press-fit ring 28 was installed can accommodate as a replacement another press-fit ring having a satisfactory abradable coating 27, thus limiting the immobilization time of said ejection tube support 29 to only the time necessary for the removal of the degraded press-fit ring 28 and the insertion of the replacement press-fit ring.

[0064] The arrangement of the abradable coating 27 on the ring 28 is described in more detail below with reference to Figures 4 and 5. As illustrated in [Fig. 4], the inner surface 34 of the ring 28 has a groove 38. This groove 38 accommodates the abradable coating 27. The abradable coating 27 is flush with a portion 39 of the inner surface 34 located outside the groove 38. In other words, a radial depth of the groove 38 is equal to a radial thickness of the abradable coating layer 27. Alternatively, the groove may have a radial depth greater than the radial thickness of the abradable coating.

[0065] The abradable coating 27 can take many forms, in particular depending on the location of the labyrinth seal 22 in the turbomachine 1.

[0066] Thus, according to one embodiment, the abradable coating 27 is of the honeycomb type. This honeycomb-type abradable coating 27 is mounted in the groove 38 by a brazing process which consists of placing metal powder between the abradable coating 27 and the internal surface 34 and melting this metal powder at a low temperature in a furnace to weld the abradable coating 27 to the internal surface 34.

[0067] According to another embodiment, this abradable coating 27 is of the sprayed type, for example of the AlSi-polyester and / or ANi type. Such an abradable coating 27 is directly sprayed onto the internal surface 34 via spray nozzles (not shown). A primer may be pre-applied to the internal surface 34 to ensure a certain roughness in order to improve the adhesion of the abradable coating 27 to said internal surface 34.

[0068] Such a shrink-fitted ring can be produced by any known manufacturing method, for example by additive manufacturing.

[0069] Although the present invention has been described with reference to specific embodiments, it is evident that modifications and changes can be made to these examples without departing from the general scope of the invention as defined by the claims.

[0070] Thus, Figures 3 to 5 illustrate the invention in the context of a turbomachine ejection tube support in which a shrink-fitted ring carrying an abradable coating is inserted. However, the invention also applies to other parts, such as the rotor or stator, of a turbomachine.

[0071] Similarly, the ring is mounted on an internal surface of a stator, i.e. turned towards the longitudinal axis of said stator, but said ring could also be mounted on an external surface of a stator or rotor.

[0072] Furthermore, the labyrinth joint as illustrated in [Fig.3] comprises a plurality of slits cooperating with the abradable coating, however the labyrinth joint could comprise one or a different number of slits whose shapes could also be variable.

[0073] The shrink-fit ring is further illustrated within an abradable coating that extends over the entire circumference of the turbomachine; however, the labyrinth seal may extend only over a portion of the turbomachine's circumference. Preferably, the ring cooperates with the support surface over at least one-third of the turbomachine's circumference. This circumference of cooperation between the ring and the workpiece ensures satisfactory positioning and retention of the labyrinth seal.

[0074] Furthermore, individual characteristics of the different embodiments illustrated / mentioned can be combined in additional embodiments.

[0075] Therefore, the description and drawings should be considered in an illustrative rather than restrictive sense.

Claims

Demands

1. A sealing device for a labyrinth seal (22) of an aircraft turbomachine (1), said device comprising a first element (23) and a second element (24), at least one of the first element (23) and the second element (24) being rotatable about a longitudinal axis (X), the first element (23) comprising an abradable coating (27), the second element (24) comprising a slat (26), the slat (26) and the abradable coating (27) being arranged radially opposite each other so as to achieve a seal during a relative rotation between the first element (23) and the second element (24), characterized in that the first element (23) comprises a frame (29) and a ring (28), the ring (28) being mounted on the frame (29), the abradable coating (27) being arranged on the ring (28).

2. Sealing device according to claim 1, wherein the ring (28) is mounted radially inside the frame (29) and preferably is mounted by shrink fitting in the frame (29).

3. Sealing device according to claim 1 or 2, wherein the frame (29) is a turbomachine ejection tube support (1), the turbomachine ejection tube support (1) comprising two cylindrical portions of different diameter and a frustoconical portion which connects the two cylindrical portions together.

4. Sealing device according to any one of claims 1 to 3, wherein the ring (28) has a groove (38), the abradable coating (27) being housed in said groove (38).

5. Sealing device according to claim 4, wherein the abradable coating (27) has a radial thickness equal to a radial depth of the groove (38).

6. Sealing device according to any one of claims 1 to 5, wherein one end (36) of the ring (28) has a collar (35).

7. Sealing device according to claim 6, in which one end of the collar (36) has a rib (37) projecting axially in the direction of the reinforcement (29).

8. Aircraft turbomachine (1) comprising a sealing device according to any one of claims 1 to 7.

9. Method of manufacturing a sealing device for a labyrinth seal (22) of an aircraft turbomachine (1), said sealing device comprising a first element (23) and a second element (24) movable in relative rotation about a longitudinal axis (X), the first element (23) comprising a frame (29), the second element (24) comprising a slat (26), the method comprising the steps of - Providing a ring (28) comprising an abradable coating (27), - Fixing the ring (28) on the frame (29) so that the abradable coating (27) is radially opposite the slat (26) and the slat (26) makes a seal with the abradable coating (27).

10. Method of manufacturing a sealing device according to claim 9, wherein the step of fixing the ring (28) on the frame (29) includes a step of shrink-fitting the ring (28) in the frame (29).

11. A method for manufacturing a sealing device according to claim 10, wherein the shrink-fitting of the ring (28) in the frame (29) successively comprises the steps of: - thermally contracting the ring (28); - inserting the ring (28) in a thermally contracted state into the frame (29), and - bringing the ring (28) back to a temperature at which the ring (28) is in a state of thermal contraction in which the cooperation between the ring (28) and the frame (29) is achieved with an interference of material suitable for fixing the ring (28) on the frame (29).

12. Method of manufacturing a sealing device according to any one of claims 9 to 11, further comprising a step of spraying an abradable material (27) onto the ring (28).

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