WO2022167373A1 - Ensemble redresseur pour compresseur de turbomachine d'aeronef - Google Patents
Ensemble redresseur pour compresseur de turbomachine d'aeronef Download PDFInfo
- Publication number
- WO2022167373A1 WO2022167373A1 PCT/EP2022/052240 EP2022052240W WO2022167373A1 WO 2022167373 A1 WO2022167373 A1 WO 2022167373A1 EP 2022052240 W EP2022052240 W EP 2022052240W WO 2022167373 A1 WO2022167373 A1 WO 2022167373A1
- Authority
- WO
- WIPO (PCT)
- Prior art keywords
- stator vanes
- shroud
- compressor
- outer shroud
- rectifier assembly
- Prior art date
Links
- 230000003100 immobilizing effect Effects 0.000 claims abstract description 11
- 238000007789 sealing Methods 0.000 claims description 16
- 239000000463 material Substances 0.000 claims description 10
- 238000004519 manufacturing process Methods 0.000 claims description 4
- 238000000034 method Methods 0.000 claims description 4
- 238000011144 upstream manufacturing Methods 0.000 description 4
- 210000003462 vein Anatomy 0.000 description 3
- 230000005540 biological transmission Effects 0.000 description 2
- 230000008878 coupling Effects 0.000 description 2
- 238000010168 coupling process Methods 0.000 description 2
- 238000005859 coupling reaction Methods 0.000 description 2
- 229910052751 metal Inorganic materials 0.000 description 2
- 239000002184 metal Substances 0.000 description 2
- 239000007787 solid Substances 0.000 description 2
- 244000261422 Lysimachia clethroides Species 0.000 description 1
- 229910000831 Steel Inorganic materials 0.000 description 1
- RTAQQCXQSZGOHL-UHFFFAOYSA-N Titanium Chemical compound [Ti] RTAQQCXQSZGOHL-UHFFFAOYSA-N 0.000 description 1
- 230000000712 assembly Effects 0.000 description 1
- 238000000429 assembly Methods 0.000 description 1
- 239000003638 chemical reducing agent Substances 0.000 description 1
- 238000002485 combustion reaction Methods 0.000 description 1
- 230000000295 complement effect Effects 0.000 description 1
- 230000021615 conjugation Effects 0.000 description 1
- 238000013016 damping Methods 0.000 description 1
- 229920001296 polysiloxane Polymers 0.000 description 1
- 239000010959 steel Substances 0.000 description 1
- 239000010936 titanium Substances 0.000 description 1
- 229910052719 titanium Inorganic materials 0.000 description 1
- 238000003466 welding Methods 0.000 description 1
- 210000002268 wool Anatomy 0.000 description 1
Classifications
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
- F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
- F01D9/00—Stators
- F01D9/02—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles
- F01D9/04—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles forming ring or sector
- F01D9/042—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles forming ring or sector fixing blades to stators
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F04—POSITIVE - DISPLACEMENT MACHINES FOR LIQUIDS; PUMPS FOR LIQUIDS OR ELASTIC FLUIDS
- F04D—NON-POSITIVE-DISPLACEMENT PUMPS
- F04D29/00—Details, component parts, or accessories
- F04D29/40—Casings; Connections of working fluid
- F04D29/52—Casings; Connections of working fluid for axial pumps
- F04D29/54—Fluid-guiding means, e.g. diffusers
- F04D29/541—Specially adapted for elastic fluid pumps
- F04D29/542—Bladed diffusers
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
- F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
- F01D11/00—Preventing or minimising internal leakage of working-fluid, e.g. between stages
- F01D11/005—Sealing means between non relatively rotating elements
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
- F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
- F01D9/00—Stators
- F01D9/02—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles
- F01D9/04—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles forming ring or sector
- F01D9/041—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles forming ring or sector using blades
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F04—POSITIVE - DISPLACEMENT MACHINES FOR LIQUIDS; PUMPS FOR LIQUIDS OR ELASTIC FLUIDS
- F04D—NON-POSITIVE-DISPLACEMENT PUMPS
- F04D29/00—Details, component parts, or accessories
- F04D29/08—Sealings
- F04D29/083—Sealings especially adapted for elastic fluid pumps
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
- F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
- F05D2230/00—Manufacture
- F05D2230/20—Manufacture essentially without removing material
- F05D2230/23—Manufacture essentially without removing material by permanently joining parts together
- F05D2230/232—Manufacture essentially without removing material by permanently joining parts together by welding
- F05D2230/238—Soldering
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
- F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
- F05D2230/00—Manufacture
- F05D2230/60—Assembly methods
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
- F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
- F05D2240/00—Components
- F05D2240/10—Stators
- F05D2240/11—Shroud seal segments
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
- F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
- F05D2240/00—Components
- F05D2240/10—Stators
- F05D2240/12—Fluid guiding means, e.g. vanes
- F05D2240/125—Fluid guiding means, e.g. vanes related to the tip of a stator vane
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
- F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
- F05D2240/00—Components
- F05D2240/80—Platforms for stationary or moving blades
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
- F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
- F05D2260/00—Function
- F05D2260/30—Retaining components in desired mutual position
- F05D2260/36—Retaining components in desired mutual position by a form fit connection, e.g. by interlocking
Definitions
- the present invention relates to an aircraft turbomachine compressor.
- Document EP 2 093 383 A1 describes a compressor in which the stator vanes are fixed to the inner shroud.
- the outer shroud is subjected to great mechanical stresses, in particular in turbomachine architectures where it is located on the path of the main force of the thrust. Part of these mechanical stresses originate from the stator vanes which are fixed to this outer shroud.
- An object of the present invention is to reduce the mechanical stresses in an aircraft turbomachine.
- the invention provides a rectifier assembly for an aircraft turbomachine compressor, comprising: • an internal shroud,
- stator vanes in which the stator vanes are attached only to the inner shroud and are in non-immobilizing mechanical contact with the outer shroud; wherein the outer shroud includes a groove receiving radially outer ends of the stator vanes; characterized in that the groove extends, axially, to a downstream end of the outer shroud.
- stator vanes are fixed only on the inner shroud, which makes it possible to avoid the places of concentration of mechanical stresses on the outer shroud.
- Contact with the outer shroud is non-immobilizing, i.e. it does not involve immobilization of the stator vanes relative to the outer shroud.
- Such contact avoids transmission of force between the blade and the outer shroud via the radially outer end of the blade, while also avoiding air leaks between the outer shroud and the radially outer end of the dawn.
- the groove of the outer shroud which extends axially to a downstream end of the outer shroud, allows the mounting of the blades to be particularly easy.
- the inner shroud is attached to the other elements of the turbomachine via the stator vanes and the outer shroud, a person skilled in the art would therefore not think of removing the attachment to the outer shroud.
- the inner shroud is provided to be attached to the other elements of the turbomachine by other means. These means are preferably more rigid than in the prior art (generally in constrained supports). The force transmission chain (turbomachine/inner shroud/blade) is thus more rigid than in the prior art.
- stator vanes are welded to the inner shroud.
- the welding allows a particularly solid fixing.
- Another fixing for example by bolting and/or riveting, is possible, while remaining within the scope of the invention.
- the outer shroud comprises a sealing element in a flexible material in contact with radially outer ends of the stator vanes.
- the sealing element prevents leaks between the radially outer ends of the stator vanes and the outer shroud.
- the flexible material preferably has a Young's modulus of less than 10 GPa.
- the flexible material may for example be silicone.
- the sealing element is preferably at least partially in the groove.
- the sealing element may comprise several separate parts while remaining within the scope of the invention.
- the sealing element is located, at least in part, at a radially outer position with respect to the radially outer ends of the stator vanes and extends, at least in part, axially along radially outer ends of the stator vanes.
- the radially outer ends of the stator vanes can slide on the sealing element while remaining in contact with it.
- the sealing element comprises a gasket.
- the seal is preferably located at an upstream end, or at a downstream end of the groove. The radially outer ends abut against it.
- stator vanes comprise, at their radially outer end, a platform extending downstream.
- a sealing element in the form of a gasket is particularly advantageous in this case.
- the inner shroud is in one piece. According to another embodiment, the inner shroud consists of a plurality of sectors forming a ring.
- the invention further provides an aircraft turbine engine comprising a first compressor having a rectifier assembly according to one embodiment of the invention.
- the first compressor may for example be the low pressure compressor or the high pressure compressor of the turbomachine.
- the relative positioning of the outer shroud with respect to the inner shroud does not use the blade but by one or more element(s) of the turbomachine external(s) to the rectifier assembly.
- the invention is particularly suitable for a turbomachine comprising a reduction gear between the shaft and the fan, because the presence of the latter generates particularly high mechanical forces on the outer shroud.
- the turbomachine comprises a second compressor, downstream of the first compressor.
- it is the most upstream compressor of the two which comprises the rectifier assembly according to the invention.
- stator vanes fixed only to the inner shroud and in non-immobilizing mechanical contact with the outer shroud are the stator vanes furthest downstream of the first compressor. This makes it possible to fix the inner shroud downstream of the first compressor in a simpler way than if the inner shroud on which the stator vanes are fixed was axially in the middle of the first compressor.
- the turbomachine comprises an intermediate support casing located, preferably directly, downstream of the first compressor, the inner shroud being fixed to the intermediate support casing or being integral with the casing intermediate support. This makes fixing the inner shroud particularly easy and solid.
- the invention also relates to an assembly comprising the intermediate support casing and the rectifier assembly.
- the outer shroud is fixed to the intermediate support casing.
- the invention further proposes an aircraft comprising a turbomachine according to the invention.
- the invention further provides a method of manufacturing a rectifier assembly, comprising the steps of:
- - Figure 1 is an axial section of a turbine engine according to one embodiment of the invention
- - Figure 2 illustrates a stator vane according to one embodiment of the invention
- FIG. 4 is a flowchart of a method for manufacturing a rectifier assembly according to one embodiment of the invention
- Figure 6 is the equivalent of Figure 5b in the case of a half-shell outer shroud.
- FIG. 1 illustrates an aircraft turbine engine 100 which may include a rectifier assembly 1 according to the invention. It may also be called a “stator assembly”.
- the aircraft turbomachine 100 is, for example, a dual-flow axial turbomachine successively comprising along the engine axis X, a fan 110, a first compressor 120 (or low-pressure compressor), a second compressor 130 (or high-pressure compressor ), a combustion chamber 160, a high pressure turbine 140 and a low pressure turbine 150.
- the mechanical power of the low 150 and high 140 pressure turbines is transmitted respectively via shafts 101 and 102 to the compressors low 120 and high 130 pressure, as well as to the fan 110 via a reducer 111 interposed at the level of the shaft 101.
- the fan 110 makes it possible to generate a primary flow 106 passing through the aircraft turbomachine 100 in a primary aerodynamic vein and a secondary flow 107 externally around the compressors 120, 130 and the turbines 140, 150.
- the first compressor 120 is provided with at least one row of rotor blades 122 followed directly downstream by a row of stator vanes 10, each row of stator vanes 10 forming a rectifier assembly 1.
- the invention can be applied to any or all of the rectifier assemblies of the first compressor 120, and in particular to the most downstream rectifier assembly of the first compressor 120.
- the aircraft turbine engine 100 comprises an inlet support casing 181 which extends around the inlet of the primary stream (in which the primary stream 106 passes), downstream of the fan 110.
- the turbine engine aircraft 100 also includes an intermediate support casing 40 which extends circumferentially between the first 120 and second 130 compressors.
- This intermediate support casing 40 comprises an annular sleeve preferably having a gooseneck profile and delimiting the primary aerodynamic vein between the first 120 and second 130 compressors. It is preferably provided with structural arms 184 extending radially through the primary vein.
- FIG. 2 illustrates a stator vane 10 of a rectifier assembly 1 according to one embodiment of the invention.
- the stator vane 10 is fixed, preferably by a weld 11, by its radially inner end 12, to an inner shroud 20.
- the fixing between the stator vane 10 and the inner shroud 20 prevents any relative movement.
- the stator vane 10 is in non-immobilizing mechanical contact, for example via a sealing element 31, via its radially outer end 13, with an outer shroud 30.
- the element seal 31 is located, at least partially, in a groove 35, preferably circumferential, in the outer shroud 30.
- the groove 35 preferably receives the radially outer ends 13 of all the stator vanes 10 of the rectifier assembly 1 .
- Figures 3a to 3c illustrate three embodiments of the invention, which differ, on the one hand, by the fixing of the inner shroud 20 to the intermediate support casing 40, and, on the other hand, by the mechanical coupling between the stator vane 10 and the outer ring 30.
- Those skilled in the art will understand that all the ways of fixing the inner ring 20 to the intermediate support casing 40 are compatible with all the mechanical couplings between the blade stator 10 and outer shroud 30.
- the groove 35 extends to the downstream end 32 of the outer shroud 30. It is filled with a flexible material in contact with radially outer ends 13 of the stator vanes 10 , and which forms the sealing element 31 . This is located at a radially outer position relative to the radially outer ends 13 of the stator vanes 10 and extends axially along the radially outer ends 13 of the stator vanes 10.
- the downstream end 22 of the inner shroud 20 is fixed to the intermediate support casing 40 by fixing means 52, for example screws.
- the groove 35 extends to the downstream end 32 of the outer shroud 30. It is filled with a flexible material in contact with radially outer ends 13 of the stator vanes 10 , and which forms the sealing element 31. This is located at a radially outer position with respect to the radially outer ends 13 of the stator vanes 10 and extends axially along the radially outer ends 13 of the stator vanes 10.
- the downstream end 22 of the inner shroud 20 is integral with the intermediate support casing 40.
- the groove 35 extends to the downstream end 32 of the outer shroud 30.
- a seal 60 for example an o-ring, is located at an upstream end 37 of the groove 35.
- stator vane 10 The upstream end of the stator vane 10 abuts on it. It forms the sealing element 31.
- stator vane 10 comprises, at its radially outer end 13, a platform 15 extending downstream and abutting against the outer shroud 30.
- downstream end 22 of the inner shroud 20 is integral with the intermediate support casing 40.
- downstream end 32 of the outer shroud 30 is fixed to the intermediate support casing 40 by fixing means 51, for example screws. Furthermore, the downstream end 22 of the inner shroud 20 is fixed to the intermediate support casing 40 or is integral with it. Consequently, in these three embodiments, the positioning of the inner shroud 20 relative to the outer shroud 30 does not stress the junction between the stator vanes 10 and the outer shroud 30 because this junction allows relative movement.
- the positioning of the inner shroud 20 with respect to the outer shroud 30, which absorbs the structural and operating forces of the turbomachine, is ensured by the junction of the inner shroud 20 with respect to the intermediate support casing 40, and of the intermediate support 40 relative to the outer shroud 30.
- Figures 4, 5a to 5d and 6 illustrate certain steps of a method 200 of manufacturing a rectifier assembly 1 according to the invention, and of its assembly with the intermediate support casing 40.
- a block of metal, for example titanium, 201 is machined 202 so as to form the inner shroud 20, preferably with holes 301 for fixing means 52.
- the inner shroud 20 is then fixed 203 to the stator vanes 10 ( Figure 5a).
- Inserts 302 are preferably inserted into the holes 301.
- the stator vanes 10 and the outer shroud 30 are positioned 204 so as to leave a space between them which will be filled with a suitable material for a non-immobilizing mechanical contact (FIGS. 5b and 6).
- FIG. 5b schematically shows a lifting tool 305 making it possible to lift the annular outer shroud 30 .
- the arrow 306 in FIG. 6 indicates that the radial flange of the half-shell outer ring 30 is located higher.
- the suitable material for a non-immobilizing mechanical contact is then deposited 205 at the junction between the stator vanes 10 and the outer shroud 30, for example using a mold 307, which is preferably such that said material does not adhere to it.
- the mold 307 can be fixed to the support tool 304.
- a rectifier assembly 1 is then obtained, which is turned over and assembled 206 to the intermediate support casing 40.
- the fixing means 51 can comprise screws 51 a and nuts 51 b.
Landscapes
- Engineering & Computer Science (AREA)
- Mechanical Engineering (AREA)
- General Engineering & Computer Science (AREA)
- Structures Of Non-Positive Displacement Pumps (AREA)
Abstract
Description
Claims
Priority Applications (3)
Application Number | Priority Date | Filing Date | Title |
---|---|---|---|
CN202280012639.4A CN116802405A (zh) | 2021-02-02 | 2022-01-31 | 飞行器涡轮发动机压缩机的定子桨叶组件 |
US18/263,943 US20240117747A1 (en) | 2021-02-02 | 2022-01-31 | Stator vane assembly for an aircraft turbine engine compressor |
EP22704334.6A EP4288668A1 (fr) | 2021-02-02 | 2022-01-31 | Ensemble redresseur pour compresseur de turbomachine d'aeronef |
Applications Claiming Priority (2)
Application Number | Priority Date | Filing Date | Title |
---|---|---|---|
BEBE2021/5074 | 2021-02-02 | ||
BE20215074A BE1029074B1 (fr) | 2021-02-02 | 2021-02-02 | Ensemble redresseur pour compresseur de turbomachine d'aeronef |
Publications (1)
Publication Number | Publication Date |
---|---|
WO2022167373A1 true WO2022167373A1 (fr) | 2022-08-11 |
Family
ID=74572577
Family Applications (1)
Application Number | Title | Priority Date | Filing Date |
---|---|---|---|
PCT/EP2022/052240 WO2022167373A1 (fr) | 2021-02-02 | 2022-01-31 | Ensemble redresseur pour compresseur de turbomachine d'aeronef |
Country Status (5)
Country | Link |
---|---|
US (1) | US20240117747A1 (fr) |
EP (1) | EP4288668A1 (fr) |
CN (1) | CN116802405A (fr) |
BE (1) | BE1029074B1 (fr) |
WO (1) | WO2022167373A1 (fr) |
Citations (4)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
US3778184A (en) | 1972-06-22 | 1973-12-11 | United Aircraft Corp | Vane damping |
EP2093383A1 (fr) | 2008-02-19 | 2009-08-26 | United Technologies Corporation | Aubes statoriques et ensemble des aubes statoriques |
FR2950116A1 (fr) * | 2009-09-15 | 2011-03-18 | Snecma | Redresseur de compresseur pour turbomachine, comprenant des tetes d'aubes montees a l'aide d'un materiau amortisseur de vibrations sur la virole exterieure |
EP2799721B1 (fr) | 2013-05-03 | 2016-09-07 | Safran Aero Booster S.A. | Redresseur de turbomachine axiale avec aubes auxiliaires en pieds d'aubes |
Family Cites Families (2)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
US4861229A (en) * | 1987-11-16 | 1989-08-29 | Williams International Corporation | Ceramic-matrix composite nozzle assembly for a turbine engine |
US10577951B2 (en) * | 2016-11-30 | 2020-03-03 | Rolls-Royce North American Technologies Inc. | Gas turbine engine with dovetail connection having contoured root |
-
2021
- 2021-02-02 BE BE20215074A patent/BE1029074B1/fr active IP Right Grant
-
2022
- 2022-01-31 EP EP22704334.6A patent/EP4288668A1/fr active Pending
- 2022-01-31 CN CN202280012639.4A patent/CN116802405A/zh active Pending
- 2022-01-31 US US18/263,943 patent/US20240117747A1/en active Pending
- 2022-01-31 WO PCT/EP2022/052240 patent/WO2022167373A1/fr active Application Filing
Patent Citations (4)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
US3778184A (en) | 1972-06-22 | 1973-12-11 | United Aircraft Corp | Vane damping |
EP2093383A1 (fr) | 2008-02-19 | 2009-08-26 | United Technologies Corporation | Aubes statoriques et ensemble des aubes statoriques |
FR2950116A1 (fr) * | 2009-09-15 | 2011-03-18 | Snecma | Redresseur de compresseur pour turbomachine, comprenant des tetes d'aubes montees a l'aide d'un materiau amortisseur de vibrations sur la virole exterieure |
EP2799721B1 (fr) | 2013-05-03 | 2016-09-07 | Safran Aero Booster S.A. | Redresseur de turbomachine axiale avec aubes auxiliaires en pieds d'aubes |
Also Published As
Publication number | Publication date |
---|---|
CN116802405A (zh) | 2023-09-22 |
BE1029074A1 (fr) | 2022-08-25 |
EP4288668A1 (fr) | 2023-12-13 |
US20240117747A1 (en) | 2024-04-11 |
BE1029074B1 (fr) | 2022-08-29 |
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