EP2616742B1 - Aerodynamic shroud for the dome of the combustion chamber of a turbomachine - Google Patents

Aerodynamic shroud for the dome of the combustion chamber of a turbomachine Download PDF

Info

Publication number
EP2616742B1
EP2616742B1 EP11773494.7A EP11773494A EP2616742B1 EP 2616742 B1 EP2616742 B1 EP 2616742B1 EP 11773494 A EP11773494 A EP 11773494A EP 2616742 B1 EP2616742 B1 EP 2616742B1
Authority
EP
European Patent Office
Prior art keywords
combustion chamber
shroud
turbomachine
annular
bosses
Prior art date
Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
Active
Application number
EP11773494.7A
Other languages
German (de)
French (fr)
Other versions
EP2616742A1 (en
Inventor
Sébastien Alain Christophe BOURGOIS
Didier Hippolyte Hernandez
Current Assignee (The listed assignees may be inaccurate. Google has not performed a legal analysis and makes no representation or warranty as to the accuracy of the list.)
Safran Aircraft Engines SAS
Original Assignee
Safran Aircraft Engines SAS
Priority date (The priority date is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the date listed.)
Filing date
Publication date
Application filed by Safran Aircraft Engines SAS filed Critical Safran Aircraft Engines SAS
Publication of EP2616742A1 publication Critical patent/EP2616742A1/en
Application granted granted Critical
Publication of EP2616742B1 publication Critical patent/EP2616742B1/en
Active legal-status Critical Current
Anticipated expiration legal-status Critical

Links

Images

Classifications

    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
    • F23R—GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
    • F23R3/00—Continuous combustion chambers using liquid or gaseous fuel
    • F23R3/02—Continuous combustion chambers using liquid or gaseous fuel characterised by the air-flow or gas-flow configuration
    • F23R3/04—Air inlet arrangements
    • F23R3/10—Air inlet arrangements for primary air
    • 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
    • F04D29/42—Casings; Connections of working fluid for radial or helico-centrifugal pumps
    • F04D29/44—Fluid-guiding means, e.g. diffusers
    • F04D29/441—Fluid-guiding means, e.g. diffusers especially adapted for elastic fluid pumps
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
    • F23R—GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
    • F23R3/00—Continuous combustion chambers using liquid or gaseous fuel
    • F23R3/02—Continuous combustion chambers using liquid or gaseous fuel characterised by the air-flow or gas-flow configuration
    • F23R3/26—Controlling the air flow
    • 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
    • F05D2250/00—Geometry
    • F05D2250/50—Inlet or outlet
    • F05D2250/52—Outlet

Definitions

  • the present invention relates to a shroud intended to cover the bottom of an annular combustion chamber in a turbomachine, such as an aircraft turbomachine in particular.
  • the invention also relates to a combustion chamber comprising a fairing of this type, and a turbomachine comprising such a combustion chamber.
  • the invention relates more particularly to a fairing intended to equip the combustion chambers of the turbomachines comprising a centrifugal type compressor arranged upstream of their combustion chamber.
  • An annular turbomachine combustion chamber is usually housed in an annular enclosure downstream of a compressor of the turbomachine and delimited by two coaxial walls of generally cylindrical shape of revolution or frustoconical, these walls being connected to each other substantially at their upstream ends by an annular chamber bottom wall provided with air and fuel injection devices comprising means for supporting fuel injector heads as well as air inlet openings.
  • the coaxial walls of these combustion chambers also have air inlet orifices, sometimes called “primary orifices” when they are arranged around an upstream region of the combustion chamber and “dilution orifices”. when arranged around a downstream region of this chamber, to allow additional injection of air into the chamber.
  • the annular wall of the chamber bottom is generally covered on the upstream side by an annular fairing for guiding a part of the air flow coming from the compressor which is intended to flow downstream into the annular enclosure in which is housed the combustion chamber bypassing the latter, in particular to supply the air inlet holes formed in the coaxial walls of the chamber, another part of this air flow being intended to penetrate inside the the combustion chamber through the air inlet ports of the air and fuel injection devices mounted in the chamber bottom, through openings of the fairing also allowing the passage of the nozzle heads.
  • the fairing covering the bottom of the combustion chambers is intended to reduce the pressure drop experienced by the air flow bypassing the combustion chambers.
  • this fairing generally takes the form of a wall of revolution having a shape substantially in C to concavity downstream when viewed in a half-section on a meridian plane.
  • turbomachines comprising a centrifugal-type compressor upstream of the combustion chamber
  • the air flow coming from this compressor enters the aforementioned chamber via an annular rectifier-diffuser opening into a radially outer region of this chamber. pregnant.
  • the flow of air supplying the air inlet ports of the injection devices and the one bypassing the combustion chamber along the radially inner wall thereof undergo a large radially inward deflection. , of a nature to increase the pressure drop of these air flows.
  • the performance of the air and fuel injection devices can be even higher than the pressure drop within these devices is high, which makes it desirable to reduce the pressure drop upstream of these devices. devices.
  • the Applicant has found that in these centrifugal compressor turbomachines, the flow of air which is intended to bypass the combustion chamber and to flow downstream along the radially inner wall of the combustion chamber, so in particular to feed the air inlet orifices of the coaxial walls of the chamber, presents an increased risk of separation in the vicinity of the shroud and downstream thereof in the radially inner region of the chamber containing the combustion chamber .
  • the invention aims in particular to provide a simple, economical and effective solution to these problems, to avoid at least partly the aforementioned drawbacks.
  • the invention proposes for this purpose an annular fairing, having an inner face intended to cover the bottom wall of an annular combustion chamber of a turbomachine equipped with a centrifugal compressor and an outer face opposite to the inner face. aforementioned, the fairing comprising a plurality of orifices for the passage of fuel injectors supported by the bottom wall of the combustion chamber.
  • the fairing comprises a plurality of bosses projecting from said outer face of the fairing, radially inwards respectively from the respective radially inner edges of said orifices, so that each of said bosses delimits an extension of the corresponding opening open radially outwardly so as to form an air bleed scoop.
  • Such an air sampling bailer makes it possible to improve the supply of air through the corresponding orifice of the shroud by reducing in particular the pressure drop experienced by the air passing through this orifice.
  • the fairings of the fairing make it possible to improve the guiding of the flow of air circulating radially inwards then downstream along the fairing and, in particular, to reduce the risks of detachment of this air flow. .
  • the aforementioned bosses advantageously extend to a radially inner end of the fairing.
  • each of the fairing bosses has a radial plane of symmetry comprising a central axis of said fairing and an injection axis of the corresponding orifice.
  • the injection axis of the orifice corresponds, of course, to the injection axis of an injector when the latter is mounted in said orifice.
  • the fairing according to this first embodiment is particularly advantageous when it is used in a turbomachine in which the flow of air coming from the compressor is devoid of a gyratory component.
  • each of the aforementioned orifices has a protuberance offset circumferentially with respect to an injection axis of the orifice.
  • the axis of injection of the orifice corresponds to the injection axis of an injector mounted in said orifice.
  • the fairing according to this second embodiment is particularly advantageous when it is used in a turbomachine in which the flow of air coming from the compressor has a gyratory component in the direction from the protuberance of the extension of each orifice towards the injection axis of the corresponding injector. This improves the scoop effect produced by these extensions vis-à-vis the air flow from the compressor.
  • each orifice may be parallel to the tangential direction or may be inclined relative to this tangential direction.
  • the inclination of the radially inner edge of the orifices relative to the tangential direction is advantageously such that this edge forms an acute angle with the direction of arrival of the air flow, this angle preferably being a right angle . This maximizes the scoop effect produced by the extensions.
  • the inclination of the radially inner edge of the orifices relative to the tangential direction may be such that this edge forms an obtuse angle with the direction of arrival of the air flow.
  • the invention also relates to an annular combustion chamber intended to be mounted downstream of a centrifugal compressor in a turbomachine, comprising two coaxial walls connected to each other. the other upstream by an annular wall of the chamber bottom, and an annular fairing of the type described above having an inner face covering the chamber bottom wall of the upstream side of the latter.
  • the fairing advantageously comprises two radially inner and outer end edges, which are respectively fixed on the coaxial walls of the combustion chamber and / or on the ends of the bottom wall of this chamber. combustion chamber.
  • the invention also relates to a turbomachine comprising an annular combustion chamber of the type described above and a centrifugal compressor mounted upstream of the combustion chamber.
  • the fairing of the combustion chamber is preferably in accordance with the first embodiment described above.
  • the fairing of the combustion chamber is preferably in accordance with the second embodiment described hereinabove. above.
  • FIGS. 1 to 4 illustrate an annular enclosure 10 in which is housed an annular combustion chamber 12 in a turbomachine 14 according to a first preferred embodiment of the invention.
  • the turbomachine 14 comprises a centrifugal type compressor upstream of the annular enclosure 10, of which only a downstream annular wall 16 is visible on the figures 1 , 3 and 4 .
  • the compressor is connected at the output to a rectifier-diffuser 18 which opens into a radially outer region of the annular enclosure 10.
  • the combustion chamber 12 is delimited by two coaxial walls of generally frustoconical shape, respectively internal 20 and external 22.
  • the inner wall 20 of the combustion chamber is connected to an inner annular wall 24 of the enclosure 10 by an inner annular shroud 26, while the outer wall 22 of the combustion chamber is connected to an outer annular wall 28 of the enclosure 10 by an outer annular ferrule 30.
  • the aforementioned annular ferrules 26 and 30 are provided with orifices 32 for the passage of air ( figure 3 ).
  • the internal and external walls 22 of the combustion chamber are furthermore connected to each other at their upstream end by an annular bottom wall of chamber 33 ( Figures 1 and 2 ) extending substantially in the radial direction and provided with a plurality of air and fuel injection devices 34, each having means 36 for supporting the head 38 of a fuel injector 40 as well as openings air inlet 41 ( figure 3 ), in a manner known per se.
  • annular bottom wall of chamber 33 is covered, on the upstream side, by an annular fairing 42 having generally an axial half-section in the shape of a C with concavity facing downstream ( Figures 1 to 4 ).
  • the fairing 42 thus has an inner face 42i covering the annular bottom wall of chamber 33 and an outer face 42e opposite to the inner face 42i ( figure 4 ).
  • the shroud 42 comprises a median annular portion 44 extending substantially parallel to the annular bottom wall of chamber 33, and two end annular portions, respectively internal 46 and external 48, curved downstream and intended for fixing the fairing 42 on the inner and outer walls 22 of the combustion chamber and on the ends 50 and 52 of the annular bottom wall of the chamber 33 curved upstream ( figure 4 ), for example by bolting ( Figures 1 and 2 ).
  • the median annular portion 44 of the fairing 42 is provided with a plurality of orifices 54 for the passage of the heads 38 of the fuel injectors 40 and the passage of the air 68 ( figure 3 ) for supplying the air inlet openings 41 of the injection devices 34, as will become more clearly apparent in the following.
  • shroud 42 comprises a plurality of bosses 56 formed essentially in its median annular portion 44.
  • each of the bosses 56 extends radially inwardly from a radially inner edge 58 of a corresponding orifice 54 to the annular inner end portion 46 of the fairing 42.
  • each boss 56 defines an upstream extension 60 of the corresponding orifice 54, which extension 60 is open radially outwards ( figures 2 and 3 ).
  • each boss 56 thus forms an air bleed scoop, such as to improve the air supply of the injection devices 34.
  • the bosses 56 each have a radial plane of symmetry comprising a central axis of the fairing 42, not visible in the figures, and an injection pin 64 of the injector 38 of the corresponding injection device 34 ( figure 3 ).
  • the plan of the figure 3 is so plane of symmetry for the boss 56 visible on this figure 3 .
  • each boss 56 is centered with respect to the corresponding injection device 34.
  • the compressor delivers a flow of air 66 ( Figures 3 and 4 ) which splits in the annular enclosure 10 into a central stream 68 supplying the injection devices 34 via the orifices 54 of the shroud 42, and in two bypass flow, respectively internal 70 and external 72, which respectively follow the inner walls 20 and outer 22 of the combustion chamber 12 around the latter, and a part feeds, the case optionally, air inlet orifices formed in these walls 20 and 22 (not visible in the figures), and the remainder of which leaves the annular enclosure 10 through the air passage holes 32 of the inner ferrules 26 and external 30.
  • the air flow 66 from the compressor is substantially devoid of rotating component, so that the conformation of the bosses 56 described above is particularly advantageous.
  • bosses 56 make it possible to reduce the risks of detachment of the air flow 70 bypassing the combustion chamber 12 radially inwards, and thus to reduce the risks of instability of operation of the combustion chamber. 12.
  • This curve represents the pressure drop of the air stream 70 coming from the output of the compressor of the turbomachine 14, between this outlet and the radially internal air passage holes 32 arranged at the downstream end. of the enclosure 10, as a function of a dimensionless ratio between the axial depth of the bosses 56 and a mean radius of the bottom 33 of the combustion chamber 12.
  • the curve is based on a first calculation (point 74) from an annular fairing of known type devoid of bosses equipping a combustion chamber whose bottom has an average radius of 252.75 mm, for which the pressure drop calculated is 1.42%, a second calculation (point 76) from a fairing 42 of the type shown in Figures 1 to 4 and provided with bosses having an axial depth of 7 mm, for which the calculated pressure drop is reduced to 1.36%, and a third calculation (point 78) from a fairing similar to the previous but whose bosses have a depth of 10 mm, and leading to a pressure drop of 1.38%, these three calculations having been performed for identical operating conditions of the turbomachine 14.
  • bosses 56 make it possible, by exerting the scoop function, to reduce the pressure drop experienced by the air flow 68 coming from the compressor outlet of the turbomachine 14 upstream of the air inlet openings. 41 air and fuel injection devices 34, as illustrated by the curve of the figure 6 .
  • This curve represents the pressure drop, obtained by numerical simulation from the three calculations described above, of the air flow 68 coming from the output of the compressor of the turbomachine 14, between this output and the inlet openings.
  • air 41 air and fuel injection devices 34 as a function of a ratio between the axial depth of the bosses 56 and the mean radius of the bottom 33 of the combustion chamber 12.
  • This pressure loss is respectively 0.50%, 0.43% and 0.41% for the three calculations mentioned above.
  • the pressure drop of the air flow 68 supplying the fuel injection devices 34 seems to decrease substantially linearly with the abovementioned dimension ratio ( figure 6 ), while the pressure drop of the airflow 70 bypassing the combustion chamber radially inwards ( figure 5 ) is reduced with moderately deep bosses but seems to be penalized when the aforementioned dimensionless ratio exceeds 2.8%, which can be explained by the fact that the great axial depth of the bosses 56 then induces detachments of this flow. air 70.
  • the figure 7 illustrates a second preferred embodiment of the invention, in which the air stream 66 from the compressor has a gyratory component.
  • the bosses 56 of the fairing 42 are shaped so that the extensions 60 of the orifices 54, formed by these bosses 56, each have a protrusion 80 offset circumferentially with respect to the central injection axis 64 of FIG. the injector 38 of the air injection device and corresponding fuel 34, in a direction such that the air flow 68 feeding these devices meets said protuberance 80 before meeting said injection pin 64.
  • Each boss 56 comprises on either side of its protrusion 80 a relatively small curved portion 84 and a relatively flat portion 86 of relatively large extent, arranged with so that the air stream 68 first encounters the low-extent portion 84 before encountering the wide-area portion 86.
  • each orifice 54 is parallel to the tangential direction ( figure 7 ).
  • this radially inner edge 58 of each orifice 54 may be inclined with respect to the tangential direction, as is shown in FIGS. Figures 8 and 9 .
  • the inclination of the radially inner edge 58 of the orifices 54 with respect to the tangential direction is advantageously such that this edge 58 forms an acute angle 88 with the direction 90 of arrival of the air flow 68.
  • the inclination radially inner edge 58 is preferably such that the edge 58 extends substantially perpendicularly to the direction 90 of arrival of the air stream 68, as shown in FIG. figure 8 . This maximizes the scoop effect produced by the extensions 60.
  • the inclination of the radially inner edge 58 of the orifices 54 with respect to the tangential direction may be such that this edge 58 forms an obtuse angle 92 with the direction 90 of arrival of the air stream 68.

Landscapes

  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Chemical & Material Sciences (AREA)
  • Combustion & Propulsion (AREA)
  • Structures Of Non-Positive Displacement Pumps (AREA)
  • Supercharger (AREA)

Description

DOMAINE TECHNIQUETECHNICAL AREA

La présente invention concerne un carénage destiné à recouvrir le fond d'une chambre annulaire de combustion dans une turbomachine, telle qu'une turbomachine d'aéronef en particulier.The present invention relates to a shroud intended to cover the bottom of an annular combustion chamber in a turbomachine, such as an aircraft turbomachine in particular.

L'invention concerne également une chambre de combustion comprenant un carénage de ce type, ainsi qu'une turbomachine comprenant une telle chambre de combustion.The invention also relates to a combustion chamber comprising a fairing of this type, and a turbomachine comprising such a combustion chamber.

L'invention concerne plus particulièrement un carénage destiné à équiper les chambres de combustion des turbomachines comprenant un compresseur de type centrifuge disposé en amont de leur chambre de combustion.The invention relates more particularly to a fairing intended to equip the combustion chambers of the turbomachines comprising a centrifugal type compressor arranged upstream of their combustion chamber.

ÉTAT DE LA TECHNIQUE ANTÉRIEURESTATE OF THE PRIOR ART

Une chambre annulaire de combustion de turbomachine est habituellement logée dans une enceinte annulaire en aval d'un compresseur de la turbomachine et délimitée par deux parois coaxiales, de forme globalement cylindrique de révolution ou tronconique, ces parois étant raccordées l'une à l'autre sensiblement au niveau de leurs extrémités amont par une paroi annulaire de fond de chambre pourvue de dispositifs d'injection d'air et de carburant comportant des moyens de support de têtes d'injecteurs de carburant ainsi que des orifices d'entrée d'air.An annular turbomachine combustion chamber is usually housed in an annular enclosure downstream of a compressor of the turbomachine and delimited by two coaxial walls of generally cylindrical shape of revolution or frustoconical, these walls being connected to each other substantially at their upstream ends by an annular chamber bottom wall provided with air and fuel injection devices comprising means for supporting fuel injector heads as well as air inlet openings.

En général, les parois coaxiales de ces chambres de combustion comportent également des orifices d'entrée d'air, parfois appelés « orifices primaires » lorsqu'ils sont agencés autour d'une région amont de la chambre de combustion et « orifices de dilution » lorsqu'ils sont agencés autour d'une région aval de cette chambre, pour permettre une injection additionnelle d'air dans la chambre.In general, the coaxial walls of these combustion chambers also have air inlet orifices, sometimes called "primary orifices" when they are arranged around an upstream region of the combustion chamber and "dilution orifices". when arranged around a downstream region of this chamber, to allow additional injection of air into the chamber.

La paroi annulaire de fond de chambre est en général recouverte du côté amont par un carénage annulaire permettant le guidage d'une partie du flux d'air provenant du compresseur qui est destinée à circuler vers l'aval dans l'enceinte annulaire dans laquelle est logée la chambre de combustion en contournant cette dernière, afin notamment d'alimenter les orifices d'entrée d'air formés dans les parois coaxiales de la chambre, une autre partie de ce flux d'air étant destiné à pénétrer à l'intérieur de la chambre de combustion par les orifices d'entrée d'air des dispositifs d'injection d'air et de carburant montés dans le fond de chambre, en passant par des ouvertures du carénage permettant également le passage des têtes d'injecteurs.The annular wall of the chamber bottom is generally covered on the upstream side by an annular fairing for guiding a part of the air flow coming from the compressor which is intended to flow downstream into the annular enclosure in which is housed the combustion chamber bypassing the latter, in particular to supply the air inlet holes formed in the coaxial walls of the chamber, another part of this air flow being intended to penetrate inside the the combustion chamber through the air inlet ports of the air and fuel injection devices mounted in the chamber bottom, through openings of the fairing also allowing the passage of the nozzle heads.

D'une manière générale, le carénage recouvrant le fond des chambres de combustion a pour but de réduire la perte de charge subie par le flux d'air contournant les chambres de combustion. Pour cela, ce carénage prend en général la forme d'une paroi de révolution ayant une forme sensiblement en C à concavité tournée vers l'aval lorsque vue en demi-section selon un plan méridien.In general, the fairing covering the bottom of the combustion chambers is intended to reduce the pressure drop experienced by the air flow bypassing the combustion chambers. For this, this fairing generally takes the form of a wall of revolution having a shape substantially in C to concavity downstream when viewed in a half-section on a meridian plane.

Le document EP562592 décrit un carénage selon le préambule de l'objet de la revendication 1.The document EP562592 discloses a fairing according to the preamble of the subject matter of claim 1.

Cependant, dans les turbomachines comprenant un compresseur de type centrifuge en amont de la chambre de combustion, le flux d'air provenant de ce compresseur pénètre dans l'enceinte précitée en passant par un redresseur-diffuseur annulaire débouchant dans une région radialement externe de cette enceinte. De ce fait, le flux d'air alimentant les orifices d'entrée d'air des dispositifs d'injection et celui contournant la chambre de combustion le long de la paroi radialement interne de celle-ci subissent une importante déviation radialement vers l'intérieur, de nature à accroître la perte de charge de ces flux d'air.However, in turbomachines comprising a centrifugal-type compressor upstream of the combustion chamber, the air flow coming from this compressor enters the aforementioned chamber via an annular rectifier-diffuser opening into a radially outer region of this chamber. pregnant. As a result, the flow of air supplying the air inlet ports of the injection devices and the one bypassing the combustion chamber along the radially inner wall thereof undergo a large radially inward deflection. , of a nature to increase the pressure drop of these air flows.

Or, la performance des dispositifs d'injection d'air et de carburant peut être d'autant plus élevée que la perte de charge au sein de ces dispositifs est élevée, ce qui rend souhaitable une réduction de la perte de charge en amont de ces dispositifs.However, the performance of the air and fuel injection devices can be even higher than the pressure drop within these devices is high, which makes it desirable to reduce the pressure drop upstream of these devices. devices.

De plus, la demanderesse a constaté que dans ces turbomachines à compresseur centrifuge, le flux d'air qui a vocation à contourner la chambre de combustion et à circuler vers l'aval le long de la paroi radialement interne de la chambre de combustion, afin notamment d'alimenter les orifices d'entrée d'air des parois coaxiales de la chambre, présente un risque accru de décollement à proximité du carénage et en aval de celui-ci dans la région radialement interne de l'enceinte contenant la chambre de combustion.In addition, the Applicant has found that in these centrifugal compressor turbomachines, the flow of air which is intended to bypass the combustion chamber and to flow downstream along the radially inner wall of the combustion chamber, so in particular to feed the air inlet orifices of the coaxial walls of the chamber, presents an increased risk of separation in the vicinity of the shroud and downstream thereof in the radially inner region of the chamber containing the combustion chamber .

Or, des décollements de ce flux d'air ne sont pas souhaitables du fait qu'ils sont susceptibles de provoquer des instabilités de fonctionnement de la chambre de combustion.However, detachments of this air flow are undesirable because they are likely to cause operating instabilities of the combustion chamber.

EXPOSÉ DE L'INVENTIONSTATEMENT OF THE INVENTION

L'invention a notamment pour but d'apporter une solution simple, économique et efficace à ces problèmes, permettant d'éviter au moins en partie les inconvénients précités.The invention aims in particular to provide a simple, economical and effective solution to these problems, to avoid at least partly the aforementioned drawbacks.

L'invention propose à cet effet un carénage annulaire, présentant une face interne destinée à recouvrir la paroi de fond d'une chambre annulaire de combustion d'une turbomachine équipée d'un compresseur centrifuge ainsi qu'une face externe opposée à la face interne précitée, le carénage comprenant une pluralité d'orifices destinés au passage d'injecteurs de carburant supportés par la paroi de fond de la chambre de combustion.The invention proposes for this purpose an annular fairing, having an inner face intended to cover the bottom wall of an annular combustion chamber of a turbomachine equipped with a centrifugal compressor and an outer face opposite to the inner face. aforementioned, the fairing comprising a plurality of orifices for the passage of fuel injectors supported by the bottom wall of the combustion chamber.

Selon l'invention, le carénage comprend une pluralité de bossages qui s'étendent en saillie sur ladite face externe du carénage, radialement vers l'intérieur respectivement depuis les bords radialement internes respectifs desdits orifices, de sorte que chacun desdits bossages délimite une extension de l'orifice correspondant ouverte radialement vers l'extérieur de manière à former une écope de prélèvement d'air.According to the invention, the fairing comprises a plurality of bosses projecting from said outer face of the fairing, radially inwards respectively from the respective radially inner edges of said orifices, so that each of said bosses delimits an extension of the corresponding opening open radially outwardly so as to form an air bleed scoop.

Une telle écope de prélèvement d'air permet d'améliorer l'alimentation en air au travers de l'orifice correspondant du carénage en réduisant notamment la perte de charge subie par l'air traversant cet orifice.Such an air sampling bailer makes it possible to improve the supply of air through the corresponding orifice of the shroud by reducing in particular the pressure drop experienced by the air passing through this orifice.

De plus, les bossages du carénage permettent d'améliorer le guidage du flux d'air circulant radialement vers l'intérieur puis vers l'aval le long du carénage et, en particulier, de réduire les risques de décollement de ce flux d'air.In addition, the fairings of the fairing make it possible to improve the guiding of the flow of air circulating radially inwards then downstream along the fairing and, in particular, to reduce the risks of detachment of this air flow. .

A cet effet, les bossages précités s'étendent avantageusement jusqu'à une extrémité radialement interne du carénage.For this purpose, the aforementioned bosses advantageously extend to a radially inner end of the fairing.

Dans un mode de réalisation préféré de l'invention, chacun des bossages du carénage présente un plan de symétrie radial comprenant un axe central dudit carénage et un axe d'injection de l'orifice correspondant.In a preferred embodiment of the invention, each of the fairing bosses has a radial plane of symmetry comprising a central axis of said fairing and an injection axis of the corresponding orifice.

L'axe d'injection de l'orifice correspond bien entendu à l'axe d'injection d'un injecteur lorsque ce dernier est monté dans ledit orifice.The injection axis of the orifice corresponds, of course, to the injection axis of an injector when the latter is mounted in said orifice.

Le carénage selon ce premier mode de réalisation est particulièrement avantageux lorsqu'il est utilisé dans une turbomachine dans laquelle le flux d'air provenant du compresseur est dépourvu de composante giratoire.The fairing according to this first embodiment is particularly advantageous when it is used in a turbomachine in which the flow of air coming from the compressor is devoid of a gyratory component.

Dans un deuxième mode de réalisation de l'invention, l'extension de chacun des orifices précités présente une protubérance décalée circonférentiellement par rapport à un axe d'injection de l'orifice.In a second embodiment of the invention, the extension of each of the aforementioned orifices has a protuberance offset circumferentially with respect to an injection axis of the orifice.

Ici encore, l'axe d'injection de l'orifice correspond à l'axe d'injection d'un injecteur monté dans ledit orifice.Here again, the axis of injection of the orifice corresponds to the injection axis of an injector mounted in said orifice.

Le carénage selon ce deuxième mode de réalisation est particulièrement avantageux lorsqu'il est utilisé dans une turbomachine dans laquelle le flux d'air provenant du compresseur présente une composante giratoire dans le sens allant de la protubérance de l'extension de chaque orifice vers l'axe d'injection de l'injecteur correspondant. Cela permet d'améliorer l'effet d'écope produit par ces extensions vis-à-vis du flux d'air provenant du compresseur.The fairing according to this second embodiment is particularly advantageous when it is used in a turbomachine in which the flow of air coming from the compressor has a gyratory component in the direction from the protuberance of the extension of each orifice towards the injection axis of the corresponding injector. This improves the scoop effect produced by these extensions vis-à-vis the air flow from the compressor.

De plus, dans ce deuxième mode de réalisation de l'invention, le bord radialement interne de chaque orifice peut être parallèle à la direction tangentielle ou bien être incliné par rapport à cette direction tangentielle.In addition, in this second embodiment of the invention, the radially inner edge of each orifice may be parallel to the tangential direction or may be inclined relative to this tangential direction.

Dans ce dernier cas, l'inclinaison du bord radialement interne des orifices par rapport à la direction tangentielle est avantageusement telle que ce bord forme un angle aigu avec la direction d'arrivée du flux d'air, cet angle étant de préférence un angle droit. Cela permet de maximiser l'effet d'écope produit par les extensions.In the latter case, the inclination of the radially inner edge of the orifices relative to the tangential direction is advantageously such that this edge forms an acute angle with the direction of arrival of the air flow, this angle preferably being a right angle . This maximizes the scoop effect produced by the extensions.

En variante, l'inclinaison du bord radialement interne des orifices par rapport à la direction tangentielle peut être telle que ce bord forme un angle obtus avec la direction d'arrivée du flux d'air.Alternatively, the inclination of the radially inner edge of the orifices relative to the tangential direction may be such that this edge forms an obtuse angle with the direction of arrival of the air flow.

L'invention concerne également une chambre annulaire de combustion destinée à être montée en aval d'un compresseur centrifuge dans une turbomachine, comprenant deux parois coaxiales raccordées l'une à l'autre en amont par une paroi annulaire de fond de chambre, ainsi qu'un carénage annulaire du type décrit ci-dessus ayant une face interne recouvrant la paroi de fond de chambre du côté amont de cette dernière.The invention also relates to an annular combustion chamber intended to be mounted downstream of a centrifugal compressor in a turbomachine, comprising two coaxial walls connected to each other. the other upstream by an annular wall of the chamber bottom, and an annular fairing of the type described above having an inner face covering the chamber bottom wall of the upstream side of the latter.

D'une manière connue en soi, le carénage comprend avantageusement deux bords d'extrémité, respectivement radialement interne et externe, qui sont respectivement fixés sur les parois coaxiales de la chambre de combustion et/ou sur des extrémités de la paroi de fond de cette chambre de combustion.In a manner known per se, the fairing advantageously comprises two radially inner and outer end edges, which are respectively fixed on the coaxial walls of the combustion chamber and / or on the ends of the bottom wall of this chamber. combustion chamber.

L'invention concerne encore une turbomachine comprenant une chambre annulaire de combustion du type décrit ci-dessus ainsi qu'un compresseur centrifuge monté en amont de la chambre de combustion.The invention also relates to a turbomachine comprising an annular combustion chamber of the type described above and a centrifugal compressor mounted upstream of the combustion chamber.

Lorsque le compresseur de la turbomachine est configuré pour délivrer un flux d'air d'alimentation de la chambre de combustion dépourvu de composante giratoire, le carénage de la chambre de combustion est de préférence conforme au premier mode de réalisation décrit ci-dessus.When the compressor of the turbomachine is configured to deliver a supply air flow of the combustion chamber devoid of rotating component, the fairing of the combustion chamber is preferably in accordance with the first embodiment described above.

En revanche, lorsque le compresseur de la turbomachine est configuré pour délivrer un flux d'air d'alimentation de la chambre de combustion présentant une composante giratoire, le carénage de la chambre de combustion est de préférence conforme au deuxième mode de réalisation décrit ci-dessus.On the other hand, when the compressor of the turbomachine is configured to deliver a supply air flow of the combustion chamber having a rotary component, the fairing of the combustion chamber is preferably in accordance with the second embodiment described hereinabove. above.

BRÈVE DESCRIPTION DES DESSINSBRIEF DESCRIPTION OF THE DRAWINGS

L'invention sera mieux comprise, et d'autres détails, avantages et caractéristiques de celle-ci apparaîtront à la lecture de la description suivante faite à titre d'exemple non limitatif et en référence aux dessins annexés dans lesquels :

  • la figure 1 est une vue schématique partielle en perspective et en coupe axiale d'une turbomachine selon un premier mode de réalisation préféré de l'invention ;
  • la figure 2 est une vue schématique partielle en perspective et en coupe axiale d'une chambre de combustion de la turbomachine de la figure 1 ;
  • la figure 3 est une vue schématique partielle de la turbomachine de la figure 1, en coupe axiale selon un plan comprenant l'axe d'un injecteur de carburant ;
  • la figure 4 est une vue schématique partielle de la turbomachine de la figure 1, en coupe axiale selon un plan équidistant de deux injecteurs de carburant consécutifs ;
  • la figure 5 est une courbe représentant la perte de charge d'un flux d'air provenant de la sortie d'un compresseur de la turbomachine de la figure 1, entre cette sortie et la sortie d'une enceinte dans laquelle est logée ladite chambre de combustion, en fonction d'un rapport entre la profondeur axiale de bossages formés dans un carénage du fond de ladite chambre de combustion et un rayon moyen du fond de cette chambre de combustion ;
  • la figure 6 est une courbe représentant la perte de charge du flux d'air provenant de la sortie du compresseur de la turbomachine de la figure 1, entre cette sortie et l'entrée de dispositifs d'injection d'air et de carburant de ladite chambre de combustion, en fonction d'un rapport entre la profondeur axiale des bossages formés dans le carénage du fond de ladite chambre de combustion et le rayon moyen du fond de cette chambre de combustion ;
  • la figure 7 est une vue schématique partielle en perspective d'une turbomachine selon un deuxième mode de réalisation préféré de l'invention, illustrant un carénage du fond de la chambre de combustion de cette turbomachine ;
  • la figure 8 est une vue schématique partielle en perspective d'une turbomachine selon un troisième mode de réalisation préféré de l'invention, illustrant un carénage du fond de la chambre de combustion de cette turbomachine, représenté seul ;
  • la figure 9 est une vue schématique partielle en perspective d'une turbomachine selon un quatrième mode de réalisation préféré de l'invention, illustrant un carénage du fond de la chambre de combustion de cette turbomachine, représenté seul.
The invention will be better understood, and other details, advantages and characteristics thereof will appear on reading the following description given by way of nonlimiting example and with reference to the accompanying drawings in which:
  • the figure 1 is a partial schematic view in perspective and in axial section of a turbomachine according to a first preferred embodiment of the invention;
  • the figure 2 is a partial schematic view in perspective and in axial section of a combustion chamber of the turbomachine of the figure 1 ;
  • the figure 3 is a partial schematic view of the turbomachine of the figure 1 , in axial section along a plane comprising the axis of a fuel injector;
  • the figure 4 is a partial schematic view of the turbomachine of the figure 1 in axial section along an equidistant plane of two consecutive fuel injectors;
  • the figure 5 is a curve representing the pressure drop of a flow of air coming from the output of a compressor of the turbomachine of the figure 1 between this outlet and the outlet of an enclosure in which said combustion chamber is housed, as a function of a ratio between the axial depth of bosses formed in a fairing of the bottom of said combustion chamber and a mean radius of the bottom of the combustion chamber. this combustion chamber;
  • the figure 6 is a curve representing the pressure drop of the air flow coming from the outlet of the compressor of the turbomachine of the figure 1 between this outlet and the inlet of air and fuel injection devices of said combustion chamber, as a function of a ratio between the axial depth of the bosses formed in the bottom fairing of said combustion chamber and the average radius of the bottom of this combustion chamber;
  • the figure 7 is a partial schematic perspective view of a turbomachine according to a second preferred embodiment of the invention, illustrating a shroud of the bottom of the combustion chamber of the turbomachine;
  • the figure 8 is a partial schematic perspective view of a turbomachine according to a third preferred embodiment of the invention, illustrating a shroud of the bottom of the combustion chamber of this turbomachine, shown alone;
  • the figure 9 is a partial schematic perspective view of a turbomachine according to a fourth preferred embodiment of the invention, illustrating a shroud of the bottom of the combustion chamber of the turbomachine, shown alone.

Dans l'ensemble de ces figures, des références identiques peuvent désigner des éléments identiques ou analogues.In all of these figures, identical references may designate identical or similar elements.

EXPOSÉ DÉTAILLÉ DE MODES DE RÉALISATION PREFERESDETAILED PRESENTATION OF PREFERRED EMBODIMENTS

Les figures 1 à 4 illustrent une enceinte annulaire 10 dans laquelle est logée une chambre annulaire de combustion 12 dans une turbomachine 14 conforme à un premier mode de réalisation préféré de l'invention.The Figures 1 to 4 illustrate an annular enclosure 10 in which is housed an annular combustion chamber 12 in a turbomachine 14 according to a first preferred embodiment of the invention.

La turbomachine 14 comprend un compresseur de type centrifuge en amont de l'enceinte annulaire 10, dont seule une paroi annulaire aval 16 est visible sur les figures 1, 3 et 4. Le compresseur est raccordé en sortie à un redresseur-diffuseur 18 qui débouche dans une région radialement externe de l'enceinte annulaire 10.The turbomachine 14 comprises a centrifugal type compressor upstream of the annular enclosure 10, of which only a downstream annular wall 16 is visible on the figures 1 , 3 and 4 . The compressor is connected at the output to a rectifier-diffuser 18 which opens into a radially outer region of the annular enclosure 10.

La chambre de combustion 12 est délimitée par deux parois coaxiales de forme globalement tronconique, respectivement interne 20 et externe 22.The combustion chamber 12 is delimited by two coaxial walls of generally frustoconical shape, respectively internal 20 and external 22.

La paroi interne 20 de la chambre de combustion est reliée à une paroi annulaire interne 24 de l'enceinte 10 par une virole annulaire interne 26, tandis que la paroi externe 22 de la chambre de combustion est reliée à une paroi annulaire externe 28 de l'enceinte 10 par une virole annulaire externe 30. Les viroles annulaires 26 et 30 précitées sont pourvues d'orifices 32 de passage d'air (figure 3).The inner wall 20 of the combustion chamber is connected to an inner annular wall 24 of the enclosure 10 by an inner annular shroud 26, while the outer wall 22 of the combustion chamber is connected to an outer annular wall 28 of the enclosure 10 by an outer annular ferrule 30. The aforementioned annular ferrules 26 and 30 are provided with orifices 32 for the passage of air ( figure 3 ).

Les parois interne 20 et externe 22 de la chambre de combustion sont en outre raccordées l'une à l'autre au niveau de leur extrémité amont par une paroi annulaire de fond de chambre 33 (figures 1 et 2) s'étendant sensiblement selon la direction radiale et pourvue d'une pluralité de dispositifs d'injection d'air et de carburant 34, comportant chacun des moyens 36 de support de la tête 38 d'un injecteur de carburant 40 ainsi que des ouvertures d'entrée d'air 41 (figure 3), d'une manière connue en soi.The internal and external walls 22 of the combustion chamber are furthermore connected to each other at their upstream end by an annular bottom wall of chamber 33 ( Figures 1 and 2 ) extending substantially in the radial direction and provided with a plurality of air and fuel injection devices 34, each having means 36 for supporting the head 38 of a fuel injector 40 as well as openings air inlet 41 ( figure 3 ), in a manner known per se.

La paroi annulaire de fond de chambre 33 est recouverte, du côté de l'amont, par un carénage annulaire 42 ayant globalement une demi-section axiale en forme de C à concavité tournée vers l'aval (figures 1 à 4).The annular bottom wall of chamber 33 is covered, on the upstream side, by an annular fairing 42 having generally an axial half-section in the shape of a C with concavity facing downstream ( Figures 1 to 4 ).

Le carénage 42 présente ainsi une face interne 42i recouvrant la paroi annulaire de fond de chambre 33 et une face externe 42e opposée à la face interne 42i (figure 4).The fairing 42 thus has an inner face 42i covering the annular bottom wall of chamber 33 and an outer face 42e opposite to the inner face 42i ( figure 4 ).

De plus, le carénage 42 comprend une partie annulaire médiane 44 s'étendant sensiblement parallèlement à la paroi annulaire de fond de chambre 33, et deux parties annulaires d'extrémité, respectivement interne 46 et externe 48, recourbées vers l'aval et destinées à la fixation du carénage 42 sur les parois interne 20 et externe 22 de la chambre de combustion et sur des extrémités 50 et 52 de la paroi annulaire de fond de chambre 33 recourbées vers l'amont (figure 4), par exemple par boulonnage (figures 1 et 2).In addition, the shroud 42 comprises a median annular portion 44 extending substantially parallel to the annular bottom wall of chamber 33, and two end annular portions, respectively internal 46 and external 48, curved downstream and intended for fixing the fairing 42 on the inner and outer walls 22 of the combustion chamber and on the ends 50 and 52 of the annular bottom wall of the chamber 33 curved upstream ( figure 4 ), for example by bolting ( Figures 1 and 2 ).

La partie annulaire médiane 44 du carénage 42 est pourvue d'une pluralité d'orifices 54 destinés au passage des têtes 38 des injecteurs de carburant 40 et au passage de l'air 68 (figure 3) destiné à alimenter les ouvertures d'entrée d'air 41 des dispositifs d'injection 34, comme cela apparaîtra plus clairement dans ce qui suit.The median annular portion 44 of the fairing 42 is provided with a plurality of orifices 54 for the passage of the heads 38 of the fuel injectors 40 and the passage of the air 68 ( figure 3 ) for supplying the air inlet openings 41 of the injection devices 34, as will become more clearly apparent in the following.

Par ailleurs, le carénage 42 comprend une pluralité de bossages 56 formés essentiellement dans sa partie annulaire médiane 44.Furthermore, the shroud 42 comprises a plurality of bosses 56 formed essentially in its median annular portion 44.

Plus précisément, chacun des bossages 56 s'étend radialement vers l'intérieur depuis un bord radialement interne 58 d'un orifice 54 correspondant jusque la partie annulaire d'extrémité interne 46 du carénage 42.More specifically, each of the bosses 56 extends radially inwardly from a radially inner edge 58 of a corresponding orifice 54 to the annular inner end portion 46 of the fairing 42.

De cette manière, chaque bossage 56 délimite une extension vers l'amont 60 de l'orifice 54 correspondant, laquelle extension 60 est ouverte radialement vers l'extérieur (figures 2 et 3). De plus, chaque bossage 56 forme ainsi une écope de prélèvement d'air, de nature à améliorer l'alimentation en air des dispositifs d'injection 34.In this way, each boss 56 defines an upstream extension 60 of the corresponding orifice 54, which extension 60 is open radially outwards ( figures 2 and 3 ). In addition, each boss 56 thus forms an air bleed scoop, such as to improve the air supply of the injection devices 34.

Dans le premier mode de réalisation décrit sur les figures 1 à 4, les bossages 56 présentent chacun un plan de symétrie radial comprenant un axe central du carénage 42, non visible sur les figures, ainsi qu'un axe d'injection 64 de l'injecteur 38 du dispositif d'injection 34 correspondant (figure 3). Le plan de la figure 3 est ainsi plan de symétrie pour le bossage 56 visible sur cette figure 3. De ce fait, chaque bossage 56 est centré par rapport au dispositif d'injection 34 correspondant.In the first embodiment described on the Figures 1 to 4 , the bosses 56 each have a radial plane of symmetry comprising a central axis of the fairing 42, not visible in the figures, and an injection pin 64 of the injector 38 of the corresponding injection device 34 ( figure 3 ). The plan of the figure 3 is so plane of symmetry for the boss 56 visible on this figure 3 . As a result, each boss 56 is centered with respect to the corresponding injection device 34.

En fonctionnement, le compresseur délivre un flux d'air 66 (figures 3 et 4) qui se divise dans l'enceinte annulaire 10 en un flux central 68 alimentant les dispositifs d'injection 34 via les orifices 54 du carénage 42, et en deux flux de contournement, respectivement interne 70 et externe 72, qui longent respectivement les parois interne 20 et externe 22 de la chambre de combustion 12 autour de cette dernière, et dont une partie alimente, le cas échéant, des orifices d'entrée d'air formés dans ces parois 20 et 22 (non visibles sur les figures), et dont le reste sort de l'enceinte annulaire 10 au travers des orifices de passage d'air 32 des viroles interne 26 et externe 30.In operation, the compressor delivers a flow of air 66 ( Figures 3 and 4 ) which splits in the annular enclosure 10 into a central stream 68 supplying the injection devices 34 via the orifices 54 of the shroud 42, and in two bypass flow, respectively internal 70 and external 72, which respectively follow the inner walls 20 and outer 22 of the combustion chamber 12 around the latter, and a part feeds, the case optionally, air inlet orifices formed in these walls 20 and 22 (not visible in the figures), and the remainder of which leaves the annular enclosure 10 through the air passage holes 32 of the inner ferrules 26 and external 30.

Dans le premier mode de réalisation décrit sur les figures 1 à 4, le flux d'air 66 provenant du compresseur est sensiblement dépourvu de composante giratoire, de sorte que la conformation des bossages 56 décrits ci-avant est particulièrement avantageuse.In the first embodiment described on the Figures 1 to 4 , the air flow 66 from the compressor is substantially devoid of rotating component, so that the conformation of the bosses 56 described above is particularly advantageous.

D'une manière générale, les bossages 56 permettent de réduire les risques de décollement du flux d'air 70 contournant la chambre de combustion 12 radialement vers l'intérieur, et donc de réduire les risques d'instabilités de fonctionnement de la chambre de combustion 12.In general, the bosses 56 make it possible to reduce the risks of detachment of the air flow 70 bypassing the combustion chamber 12 radially inwards, and thus to reduce the risks of instability of operation of the combustion chamber. 12.

La réduction des risques de décollement du flux d'air 70 se traduit par une réduction de la perte de charge subie par ce flux d'air entre la sortie du redresseur-diffuseur 18 et les orifices de passage d'air 32 prévus à l'extrémité aval de l'enceinte annulaire 10, comme l'illustre la courbe de la figure 5.The reduction of the risks of separation of the air flow 70 results in a reduction in the pressure drop experienced by this airflow between the outlet of the diffuser-diffuser 18 and the air passage holes 32 provided for in FIG. downstream end of the annular enclosure 10, as illustrated by the curve of the figure 5 .

Cette courbe, obtenue par simulation numérique, représente la perte de charge du flux d'air 70 provenant de la sortie du compresseur de la turbomachine 14, entre cette sortie et les orifices de passage d'air 32 radialement internes agencés à l'extrémité aval de l'enceinte 10, en fonction d'un rapport sans dimension entre la profondeur axiale des bossages 56 et un rayon moyen du fond 33 de la chambre de combustion 12.This curve, obtained by numerical simulation, represents the pressure drop of the air stream 70 coming from the output of the compressor of the turbomachine 14, between this outlet and the radially internal air passage holes 32 arranged at the downstream end. of the enclosure 10, as a function of a dimensionless ratio between the axial depth of the bosses 56 and a mean radius of the bottom 33 of the combustion chamber 12.

Plus précisément, la courbe se fonde sur un premier calcul (point 74) à partir d'un carénage annulaire de type connu dépourvu de bossages équipant une chambre de combustion dont le fond présente un rayon moyen de 252.75 mm, pour lequel la perte de charge calculée est de 1,42%, un deuxième calcul (point 76) à partir d'un carénage 42 du type représenté sur les figures 1 à 4 et pourvu de bossages ayant une profondeur axiale de 7 mm, pour lequel la perte de charge calculée est réduite à 1,36%, et un troisième calcul (point 78) à partir d'un carénage semblable au précédent mais dont les bossages ont une profondeur de 10 mm, et conduisant à une perte de charge de 1,38%, ces trois calculs ayant été réalisés pour des conditions identiques de fonctionnement de la turbomachine 14.More specifically, the curve is based on a first calculation (point 74) from an annular fairing of known type devoid of bosses equipping a combustion chamber whose bottom has an average radius of 252.75 mm, for which the pressure drop calculated is 1.42%, a second calculation (point 76) from a fairing 42 of the type shown in Figures 1 to 4 and provided with bosses having an axial depth of 7 mm, for which the calculated pressure drop is reduced to 1.36%, and a third calculation (point 78) from a fairing similar to the previous but whose bosses have a depth of 10 mm, and leading to a pressure drop of 1.38%, these three calculations having been performed for identical operating conditions of the turbomachine 14.

Par ailleurs, les bossages 56 permettent, en exerçant la fonction d'écope, de réduire la perte de charge subie par le flux d'air 68 provenant de la sortie du compresseur de la turbomachine 14 en amont des ouvertures d'entrée d'air 41 des dispositifs d'injection d'air et de carburant 34, comme l'illustre la courbe de la figure 6.Furthermore, the bosses 56 make it possible, by exerting the scoop function, to reduce the pressure drop experienced by the air flow 68 coming from the compressor outlet of the turbomachine 14 upstream of the air inlet openings. 41 air and fuel injection devices 34, as illustrated by the curve of the figure 6 .

Cette courbe représente la perte de charge, obtenue par simulation numérique à partir des trois calculs décrits ci-dessus, du flux d'air 68 provenant de la sortie du compresseur de la turbomachine 14, entre cette sortie et les ouvertures d'entrée d'air 41 des dispositifs d'injection d'air et de carburant 34, en fonction d'un rapport entre la profondeur axiale des bossages 56 et le rayon moyen du fond 33 de la chambre de combustion 12.This curve represents the pressure drop, obtained by numerical simulation from the three calculations described above, of the air flow 68 coming from the output of the compressor of the turbomachine 14, between this output and the inlet openings. air 41 air and fuel injection devices 34, as a function of a ratio between the axial depth of the bosses 56 and the mean radius of the bottom 33 of the combustion chamber 12.

Cette perte de charge est respectivement de 0,50%, de 0,43% et de 0,41% pour les trois calculs précités.This pressure loss is respectively 0.50%, 0.43% and 0.41% for the three calculations mentioned above.

Ainsi, la perte de charge du flux d'air 68 alimentant les dispositifs d'injection de carburant 34 semble décroître sensiblement linéairement avec le rapport sans dimension précité (figure 6), tandis que la perte de charge du flux d'air 70 contournant la chambre de combustion radialement vers l'intérieur (figure 5) est réduite avec des bossages de profondeur modérée mais semble être pénalisée lorsque le rapport sans dimension précité dépasse 2,8%, ce qui peut s'expliquer par le fait que la grande profondeur axiale des bossages 56 induit alors des décollements de ce flux d'air 70.Thus, the pressure drop of the air flow 68 supplying the fuel injection devices 34 seems to decrease substantially linearly with the abovementioned dimension ratio ( figure 6 ), while the pressure drop of the airflow 70 bypassing the combustion chamber radially inwards ( figure 5 ) is reduced with moderately deep bosses but seems to be penalized when the aforementioned dimensionless ratio exceeds 2.8%, which can be explained by the fact that the great axial depth of the bosses 56 then induces detachments of this flow. air 70.

La figure 7 illustre un deuxième mode de réalisation préféré de l'invention, dans lequel le flux d'air 66 provenant du compresseur présente une composante giratoire.The figure 7 illustrates a second preferred embodiment of the invention, in which the air stream 66 from the compressor has a gyratory component.

Dans ce deuxième mode de réalisation, les bossages 56 du carénage 42 sont conformés de sorte que les extensions 60 des orifices 54, formées par ces bossages 56, présentent chacune une protubérance 80 décalée circonférentiellement par rapport à l'axe central d'injection 64 de l'injecteur 38 du dispositif d'injection d'air et de carburant 34 correspondant, dans un sens tel que le flux d'air 68 alimentant ces dispositifs rencontre ladite protubérance 80 avant de rencontrer ledit axe d'injection 64. Chaque bossage 56 comprend de part et d'autre de sa protubérance 80 une partie incurvée 84 de relativement faible étendue et une partie sensiblement plane 86 de relativement grande étendue, disposées de sorte que le flux d'air 68 rencontre d'abord la partie de faible étendue 84 avant de rencontrer la partie de grande étendue 86.In this second embodiment, the bosses 56 of the fairing 42 are shaped so that the extensions 60 of the orifices 54, formed by these bosses 56, each have a protrusion 80 offset circumferentially with respect to the central injection axis 64 of FIG. the injector 38 of the air injection device and corresponding fuel 34, in a direction such that the air flow 68 feeding these devices meets said protuberance 80 before meeting said injection pin 64. Each boss 56 comprises on either side of its protrusion 80 a relatively small curved portion 84 and a relatively flat portion 86 of relatively large extent, arranged with so that the air stream 68 first encounters the low-extent portion 84 before encountering the wide-area portion 86.

Par ailleurs, le bord radialement interne 58 de chaque orifice 54 est parallèle à la direction tangentielle (figure 7).Moreover, the radially inner edge 58 of each orifice 54 is parallel to the tangential direction ( figure 7 ).

En variante, ce bord radialement interne 58 de chaque orifice 54 peut être incliné par rapport à la direction tangentielle, comme cela est représenté sur les figures 8 et 9.As a variant, this radially inner edge 58 of each orifice 54 may be inclined with respect to the tangential direction, as is shown in FIGS. Figures 8 and 9 .

Dans ce cas, l'inclinaison du bord radialement interne 58 des orifices 54 par rapport à la direction tangentielle est avantageusement telle que ce bord 58 forme un angle aigu 88 avec la direction 90 d'arrivée du flux d'air 68. L'inclinaison du bord radialement interne 58 est de préférence telle que le bord 58 s'étende sensiblement perpendiculairement à la direction 90 d'arrivée du flux d'air 68, comme illustré sur la figure 8. Cela permet de maximiser l'effet d'écope produit par les extensions 60.In this case, the inclination of the radially inner edge 58 of the orifices 54 with respect to the tangential direction is advantageously such that this edge 58 forms an acute angle 88 with the direction 90 of arrival of the air flow 68. The inclination radially inner edge 58 is preferably such that the edge 58 extends substantially perpendicularly to the direction 90 of arrival of the air stream 68, as shown in FIG. figure 8 . This maximizes the scoop effect produced by the extensions 60.

En variante, l'inclinaison du bord radialement interne 58 des orifices 54 par rapport à la direction tangentielle peut être telle que ce bord 58 forme un angle obtus 92 avec la direction 90 d'arrivée du flux d'air 68.As a variant, the inclination of the radially inner edge 58 of the orifices 54 with respect to the tangential direction may be such that this edge 58 forms an obtuse angle 92 with the direction 90 of arrival of the air stream 68.

Claims (8)

  1. An annular shroud (42), having an internal face (42i) intended to cover the back end wall (33) of an annular combustion chamber (12) of a turbomachine (14) fitted with a centrifugal compressor, together with an external face (42e) opposite said internal face (42i), where said shroud includes multiple apertures (54) intended to allow fuel injectors (38, 40) supported by said back end wall (33) of the combustion chamber (12) to pass through, where said shroud is characterised in that it includes multiple bosses (56) which extend as projections on said external face (42e) of the shroud, radially towards the interior respectively from the respective radially internal edges (58) of said apertures (54), such that each of said bosses (56) delimits an extension (60) of the corresponding aperture (54) open radially towards the exterior so as to form an air intake scoop.
  2. An annular shroud according to claim 1, characterised in that said bosses (56) extend as far as a radially internal end of said shroud (42).
  3. An annular shroud according to claim 1 or 2, characterised in that each of said bosses (56) has a radial plane of symmetry including a central axis of said shroud (42) and an axis of injection (64) of the corresponding aperture (54).
  4. An annual shroud according to claim 1 or 2, characterised in that said extension (60) of each of said apertures (54) has a protrusion offset circumferentially relative to an axis of injection (64) of the aperture (54).
  5. An annular combustion chamber (12) intended to be installed downstream from a centrifugal compressor in a turbomachine (14), including two coaxial walls (20, 22) connected to one another upstream by a back-of-chamber annular end wall (33), characterised in that it includes an annular shroud (42) according to any one of the previous claims, where said shroud (42) has an internal face (42i) covering said back-of-chamber end wall (33) on the upstream side of the latter.
  6. A turbomachine (14), characterised in that it includes an annular combustion chamber (12) according to claim 5, together with a centrifugal compressor installed upstream from said combustion chamber (12).
  7. A turbomachine according to claim 6, characterised in that said compressor is configured to deliver an airstream (66) feeding said combustion chamber (12) which has no gyratory component, and in that the shroud (42) of the combustion chamber (12) is in accordance with claim 3.
  8. A turbomachine according to claim 7, characterised in that said compressor is configured to deliver an airstream (66) feeding said combustion chamber (12) which has a gyratory component, and in that the shroud (42) of the combustion chamber (12) is in accordance with claim 4.
EP11773494.7A 2010-09-14 2011-09-13 Aerodynamic shroud for the dome of the combustion chamber of a turbomachine Active EP2616742B1 (en)

Applications Claiming Priority (2)

Application Number Priority Date Filing Date Title
FR1057319A FR2964725B1 (en) 2010-09-14 2010-09-14 AERODYNAMIC FAIRING FOR BOTTOM OF COMBUSTION CHAMBER
PCT/FR2011/052084 WO2012035248A1 (en) 2010-09-14 2011-09-13 Aerodynamic shroud for the bottom of a combustion chamber of a turbomachine

Publications (2)

Publication Number Publication Date
EP2616742A1 EP2616742A1 (en) 2013-07-24
EP2616742B1 true EP2616742B1 (en) 2018-10-31

Family

ID=44063986

Family Applications (1)

Application Number Title Priority Date Filing Date
EP11773494.7A Active EP2616742B1 (en) 2010-09-14 2011-09-13 Aerodynamic shroud for the dome of the combustion chamber of a turbomachine

Country Status (8)

Country Link
US (1) US8661829B2 (en)
EP (1) EP2616742B1 (en)
CN (1) CN103080652B (en)
BR (1) BR112013006037B1 (en)
CA (1) CA2811163C (en)
FR (1) FR2964725B1 (en)
RU (1) RU2572736C2 (en)
WO (1) WO2012035248A1 (en)

Families Citing this family (9)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
FR2943403B1 (en) 2009-03-17 2014-11-14 Snecma TURBOMACHINE COMBUSTION CHAMBER COMPRISING IMPROVED AIR SUPPLY MEANS
FR2945854B1 (en) 2009-05-19 2015-08-07 Snecma MIXTURE SPINDLE FOR A FUEL INJECTOR IN A COMBUSTION CHAMBER OF A GAS TURBINE AND CORRESPONDING COMBUSTION DEVICE
FR3003632B1 (en) 2013-03-19 2016-10-14 Snecma INJECTION SYSTEM FOR TURBOMACHINE COMBUSTION CHAMBER HAVING AN ANNULAR WALL WITH CONVERGENT INTERNAL PROFILE
US9650916B2 (en) 2014-04-09 2017-05-16 Honeywell International Inc. Turbomachine cooling systems
FR3035481B1 (en) 2015-04-23 2017-05-05 Snecma TURBOMACHINE COMBUSTION CHAMBER COMPRISING A SPECIFICALLY SHAPED AIR FLOW GUIDING DEVICE
US10619856B2 (en) 2017-03-13 2020-04-14 Rolls-Royce Corporation Notched gas turbine combustor cowl
US10816213B2 (en) 2018-03-01 2020-10-27 General Electric Company Combustor assembly with structural cowl and decoupled chamber
US10907831B2 (en) * 2018-05-07 2021-02-02 Rolls-Royce Corporation Ram pressure recovery fuel nozzle with a scoop
US10982852B2 (en) 2018-11-05 2021-04-20 Rolls-Royce Corporation Cowl integration to combustor wall

Family Cites Families (25)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
GB239127A (en) * 1925-03-25 1925-09-03 Stephen Edward Beeson Improvements in or relating to sawing, cutting and similar machines
BE795867A (en) * 1972-03-01 1973-06-18 Gen Electric DEVICE FOR UNIFORMISING THE FLOW OF AIR IN A GAS TURBINE
FR2559856B1 (en) * 1984-02-17 1987-06-19 Caillau Ets TIGHTENING COLLAR AND MANUFACTURING METHOD THEREOF
FR2686683B1 (en) * 1992-01-28 1994-04-01 Snecma TURBOMACHINE WITH REMOVABLE COMBUSTION CHAMBER.
CA2089272C (en) 1992-03-23 2002-09-03 James Norman Reinhold, Jr. Impact resistant combustor
US5279126A (en) * 1992-12-18 1994-01-18 United Technologies Corporation Diffuser-combustor
DE10159668A1 (en) * 2001-12-05 2003-06-18 Rolls Royce Deutschland Combustion chamber head has at least one turbulence-creating element on flow surface of cover
GB2391297A (en) * 2002-07-24 2004-02-04 Rolls Royce Plc Gas supply assembly
US7222488B2 (en) * 2002-09-10 2007-05-29 General Electric Company Fabricated cowl for double annular combustor of a gas turbine engine
US6952927B2 (en) * 2003-05-29 2005-10-11 General Electric Company Multiport dome baffle
FR2856467B1 (en) * 2003-06-18 2005-09-02 Snecma Moteurs TURBOMACHINE ANNULAR COMBUSTION CHAMBER
RU2250415C1 (en) * 2003-08-05 2005-04-20 Открытое акционерное общество "Научно-производственное объединение "Сатурн" (ОАО "НПО "Сатурн") Annular combustion chamber of gas-turbine engine
FR2885201B1 (en) * 2005-04-28 2010-09-17 Snecma Moteurs EASILY DISMANTLING COMBUSTION CHAMBER WITH IMPROVED AERODYNAMIC PERFORMANCE
FR2888631B1 (en) 2005-07-18 2010-12-10 Snecma TURBOMACHINE WITH ANGULAR AIR DISTRIBUTION
FR2897145B1 (en) * 2006-02-08 2013-01-18 Snecma ANNULAR COMBUSTION CHAMBER FOR TURBOMACHINE WITH ALTERNATE FIXINGS.
FR2897144B1 (en) * 2006-02-08 2008-05-02 Snecma Sa COMBUSTION CHAMBER FOR TURBOMACHINE WITH TANGENTIAL SLOTS
FR2897417A1 (en) * 2006-02-10 2007-08-17 Snecma Sa ANNULAR COMBUSTION CHAMBER OF A TURBOMACHINE
FR2910597B1 (en) * 2006-12-22 2009-03-20 Snecma Sa FURNITURE FOR BOTTOM OF COMBUSTION CHAMBER
FR2911668B1 (en) * 2007-01-18 2009-03-20 Snecma Sa COMBUSTION CHAMBER OF A TURBOMACHINE
FR2914399B1 (en) * 2007-03-27 2009-10-02 Snecma Sa FURNITURE FOR BOTTOM OF COMBUSTION CHAMBER.
FR2921464B1 (en) 2007-09-24 2014-03-28 Snecma ARRANGEMENT OF INJECTION SYSTEMS IN A COMBUSTION CHAMBER BOTTOM OF AN AIRCRAFT ENGINE
FR2929690B1 (en) * 2008-04-03 2012-08-17 Snecma Propulsion Solide COMBUSTION CHAMBER SECTORIZED IN CMC FOR GAS TURBINE
FR2943403B1 (en) * 2009-03-17 2014-11-14 Snecma TURBOMACHINE COMBUSTION CHAMBER COMPRISING IMPROVED AIR SUPPLY MEANS
FR2945854B1 (en) 2009-05-19 2015-08-07 Snecma MIXTURE SPINDLE FOR A FUEL INJECTOR IN A COMBUSTION CHAMBER OF A GAS TURBINE AND CORRESPONDING COMBUSTION DEVICE
FR2975465B1 (en) 2011-05-19 2018-03-09 Safran Aircraft Engines WALL FOR TURBOMACHINE COMBUSTION CHAMBER COMPRISING AN OPTIMIZED AIR INLET ORIFICE ARRANGEMENT

Non-Patent Citations (1)

* Cited by examiner, † Cited by third party
Title
None *

Also Published As

Publication number Publication date
CN103080652A (en) 2013-05-01
US8661829B2 (en) 2014-03-04
RU2013117008A (en) 2014-10-20
FR2964725B1 (en) 2012-10-12
BR112013006037B1 (en) 2020-11-17
BR112013006037A2 (en) 2016-06-07
CN103080652B (en) 2015-05-06
EP2616742A1 (en) 2013-07-24
WO2012035248A1 (en) 2012-03-22
RU2572736C2 (en) 2016-01-20
FR2964725A1 (en) 2012-03-16
US20130160452A1 (en) 2013-06-27
CA2811163C (en) 2018-10-23
CA2811163A1 (en) 2012-03-22

Similar Documents

Publication Publication Date Title
EP2616742B1 (en) Aerodynamic shroud for the dome of the combustion chamber of a turbomachine
EP2488739B1 (en) Air intake for a as turbine engine within a nacelle
EP1881179B1 (en) System for ventilating the wall of a combustion chamber in a turbomachine
EP2286065B1 (en) Compressor rotor for a turbomachine including centripetal air bleeding means
EP1862644B1 (en) Air-flow guiding device for a turbomachine, and corresponding turbomachine and diffuser
CA2754419C (en) Turbine engine combustion chamber comprising improved air supply means
EP2678610B1 (en) Annular combustion chamber for a turbine engine including improved dilution openings
WO2011054880A2 (en) Fuel mixing device for a turbine engine combustion chamber including an improved air supply means
CA2740660A1 (en) Oil separator rotor for turbomachine
EP3039341B1 (en) Turbomachine combustion chamber comprising air deflection means for reducing wake created by igniter
FR2977909A1 (en) Blade for rotor of low-pressure turbine of e.g. turbojet of aircraft, has knife edge comprising internal part and external part that is tilted relative to axis, where knife edge extends from outer end to outside in direction of leading edge
WO2014122403A1 (en) Turbine engine impeller
WO2013060985A1 (en) Aircraft turbomachine combustion chamber module and method for designing same
EP3039342B1 (en) Combustion chamber for gas turbine with homogeneous air inlet through the fuel injection systems
FR2960603A1 (en) Radial diffuser for turboshaft engine e.g. turbojet of airplane, has slots whose shape and positioning are complementary to that of projecting portions such that projecting portions are embedded into slots
FR2951502A1 (en) Turbomachine e.g. jet engine, for aircraft, has connection unit with air flow annular channel centered on longitudinal axle, where channel is traversed by oblique torque arms connecting internal units with streamline annual cases
FR3029271B1 (en) ANNULAR DEFLECTION WALL FOR TURBOMACHINE COMBUSTION CHAMBER INJECTION SYSTEM PROVIDING EXTENSIVE FUEL ATOMIZATION AREA
EP3449185B1 (en) Turbomachine injection system comprising an aerodynamic deflector at its inlet and an air intake swirler
FR2973479A1 (en) Revolution wall e.g. external revolution wall, for combustion chamber of turbomachine of commercial plane, has circumferential row of primary air holes whose regions are located away from plane along row of dilution holes
FR2951504A1 (en) Gas turbine engine and nacelle assembly for e.g. helicopter, has secondary deflecting channel shaped such that flow velocity of air increases from upstream to downstream, where channel has outlet with opening leading into wall of nacelle
EP3969813B1 (en) Combustion chamber comprising means for cooling an annular casing zone downstream of a chimney
FR2998330A1 (en) Single piece part i.e. casting part, for intermediate casing hub of e.g. turbojet engine, of aircraft, has deflecting surface whose radial internal end partially defines separation nozzle, where surface is extended to external end
FR3022597A1 (en) TRIPLE FLUX DIFFUSER FOR TURBOMACHINE MODULE COMPRISING AIR PIPING DEVICES BETWEEN THE TWO WALLS OF THE DIFFUSER SEPARATION
FR2957659A1 (en) Air and fuel injecting system for base of combustion annular chamber of turbine engine of aircraft, has fuel ejection opening with ejection axis passed in downstream of downstream end of partition wall with reference to flow of air stream
FR3039202A1 (en) AIRCRAFT COMPRISING A TURBOMACHINE INTEGRATED WITH REAR FUSELAGE WITH VARIABLE POWER SUPPLY

Legal Events

Date Code Title Description
PUAI Public reference made under article 153(3) epc to a published international application that has entered the european phase

Free format text: ORIGINAL CODE: 0009012

17P Request for examination filed

Effective date: 20130321

AK Designated contracting states

Kind code of ref document: A1

Designated state(s): AL AT BE BG CH CY CZ DE DK EE ES FI FR GB GR HR HU IE IS IT LI LT LU LV MC MK MT NL NO PL PT RO RS SE SI SK SM TR

DAX Request for extension of the european patent (deleted)
GRAP Despatch of communication of intention to grant a patent

Free format text: ORIGINAL CODE: EPIDOSNIGR1

STAA Information on the status of an ep patent application or granted ep patent

Free format text: STATUS: GRANT OF PATENT IS INTENDED

INTG Intention to grant announced

Effective date: 20180405

RAP1 Party data changed (applicant data changed or rights of an application transferred)

Owner name: SAFRAN AIRCRAFT ENGINES

GRAS Grant fee paid

Free format text: ORIGINAL CODE: EPIDOSNIGR3

GRAA (expected) grant

Free format text: ORIGINAL CODE: 0009210

STAA Information on the status of an ep patent application or granted ep patent

Free format text: STATUS: THE PATENT HAS BEEN GRANTED

AK Designated contracting states

Kind code of ref document: B1

Designated state(s): AL AT BE BG CH CY CZ DE DK EE ES FI FR GB GR HR HU IE IS IT LI LT LU LV MC MK MT NL NO PL PT RO RS SE SI SK SM TR

REG Reference to a national code

Ref country code: CH

Ref legal event code: EP

Ref country code: GB

Ref legal event code: FG4D

Free format text: NOT ENGLISH

REG Reference to a national code

Ref country code: AT

Ref legal event code: REF

Ref document number: 1059852

Country of ref document: AT

Kind code of ref document: T

Effective date: 20181115

REG Reference to a national code

Ref country code: DE

Ref legal event code: R096

Ref document number: 602011053458

Country of ref document: DE

REG Reference to a national code

Ref country code: IE

Ref legal event code: FG4D

Free format text: LANGUAGE OF EP DOCUMENT: FRENCH

REG Reference to a national code

Ref country code: SE

Ref legal event code: TRGR

REG Reference to a national code

Ref country code: NL

Ref legal event code: MP

Effective date: 20181031

REG Reference to a national code

Ref country code: LT

Ref legal event code: MG4D

REG Reference to a national code

Ref country code: AT

Ref legal event code: MK05

Ref document number: 1059852

Country of ref document: AT

Kind code of ref document: T

Effective date: 20181031

PG25 Lapsed in a contracting state [announced via postgrant information from national office to epo]

Ref country code: LV

Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT

Effective date: 20181031

Ref country code: IS

Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT

Effective date: 20190228

Ref country code: AT

Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT

Effective date: 20181031

Ref country code: FI

Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT

Effective date: 20181031

Ref country code: BG

Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT

Effective date: 20190131

Ref country code: NO

Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT

Effective date: 20190131

Ref country code: HR

Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT

Effective date: 20181031

Ref country code: PL

Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT

Effective date: 20181031

Ref country code: ES

Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT

Effective date: 20181031

Ref country code: LT

Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT

Effective date: 20181031

PG25 Lapsed in a contracting state [announced via postgrant information from national office to epo]

Ref country code: GR

Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT

Effective date: 20190201

Ref country code: NL

Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT

Effective date: 20181031

Ref country code: RS

Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT

Effective date: 20181031

Ref country code: PT

Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT

Effective date: 20190301

Ref country code: AL

Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT

Effective date: 20181031

PG25 Lapsed in a contracting state [announced via postgrant information from national office to epo]

Ref country code: CZ

Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT

Effective date: 20181031

Ref country code: DK

Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT

Effective date: 20181031

REG Reference to a national code

Ref country code: DE

Ref legal event code: R097

Ref document number: 602011053458

Country of ref document: DE

PG25 Lapsed in a contracting state [announced via postgrant information from national office to epo]

Ref country code: SK

Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT

Effective date: 20181031

Ref country code: RO

Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT

Effective date: 20181031

Ref country code: EE

Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT

Effective date: 20181031

Ref country code: SM

Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT

Effective date: 20181031

PLBE No opposition filed within time limit

Free format text: ORIGINAL CODE: 0009261

STAA Information on the status of an ep patent application or granted ep patent

Free format text: STATUS: NO OPPOSITION FILED WITHIN TIME LIMIT

26N No opposition filed

Effective date: 20190801

PG25 Lapsed in a contracting state [announced via postgrant information from national office to epo]

Ref country code: SI

Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT

Effective date: 20181031

PG25 Lapsed in a contracting state [announced via postgrant information from national office to epo]

Ref country code: TR

Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT

Effective date: 20181031

PG25 Lapsed in a contracting state [announced via postgrant information from national office to epo]

Ref country code: MC

Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT

Effective date: 20181031

REG Reference to a national code

Ref country code: CH

Ref legal event code: PL

PG25 Lapsed in a contracting state [announced via postgrant information from national office to epo]

Ref country code: LU

Free format text: LAPSE BECAUSE OF NON-PAYMENT OF DUE FEES

Effective date: 20190913

Ref country code: IE

Free format text: LAPSE BECAUSE OF NON-PAYMENT OF DUE FEES

Effective date: 20190913

Ref country code: CH

Free format text: LAPSE BECAUSE OF NON-PAYMENT OF DUE FEES

Effective date: 20190930

Ref country code: LI

Free format text: LAPSE BECAUSE OF NON-PAYMENT OF DUE FEES

Effective date: 20190930

REG Reference to a national code

Ref country code: BE

Ref legal event code: MM

Effective date: 20190930

PG25 Lapsed in a contracting state [announced via postgrant information from national office to epo]

Ref country code: BE

Free format text: LAPSE BECAUSE OF NON-PAYMENT OF DUE FEES

Effective date: 20190930

PG25 Lapsed in a contracting state [announced via postgrant information from national office to epo]

Ref country code: CY

Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT

Effective date: 20181031

PG25 Lapsed in a contracting state [announced via postgrant information from national office to epo]

Ref country code: HU

Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT; INVALID AB INITIO

Effective date: 20110913

Ref country code: MT

Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT

Effective date: 20181031

PG25 Lapsed in a contracting state [announced via postgrant information from national office to epo]

Ref country code: MK

Free format text: LAPSE BECAUSE OF FAILURE TO SUBMIT A TRANSLATION OF THE DESCRIPTION OR TO PAY THE FEE WITHIN THE PRESCRIBED TIME-LIMIT

Effective date: 20181031

PGFP Annual fee paid to national office [announced via postgrant information from national office to epo]

Ref country code: IT

Payment date: 20230822

Year of fee payment: 13

Ref country code: GB

Payment date: 20230823

Year of fee payment: 13

PGFP Annual fee paid to national office [announced via postgrant information from national office to epo]

Ref country code: SE

Payment date: 20230822

Year of fee payment: 13

Ref country code: FR

Payment date: 20230822

Year of fee payment: 13

Ref country code: DE

Payment date: 20230822

Year of fee payment: 13