WO2011086305A1 - Distributeur de turbine haute pression d'un turboreacteur - Google Patents
Distributeur de turbine haute pression d'un turboreacteur Download PDFInfo
- Publication number
- WO2011086305A1 WO2011086305A1 PCT/FR2011/050017 FR2011050017W WO2011086305A1 WO 2011086305 A1 WO2011086305 A1 WO 2011086305A1 FR 2011050017 W FR2011050017 W FR 2011050017W WO 2011086305 A1 WO2011086305 A1 WO 2011086305A1
- Authority
- WO
- WIPO (PCT)
- Prior art keywords
- liner
- sleeve
- dispenser according
- vein
- holes
- Prior art date
Links
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
- F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
- F01D5/12—Blades
- F01D5/14—Form or construction
- F01D5/18—Hollow blades, i.e. blades with cooling or heating channels or cavities; Heating, heat-insulating or cooling means on blades
- F01D5/187—Convection cooling
- F01D5/188—Convection cooling with an insert in the blade cavity to guide the cooling fluid, e.g. forming a separation wall
- F01D5/189—Convection cooling with an insert in the blade cavity to guide the cooling fluid, e.g. forming a separation wall the insert having a tubular cross-section, e.g. airfoil shape
-
- 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
- F01D17/00—Regulating or controlling by varying flow
- F01D17/02—Arrangement of sensing elements
- F01D17/08—Arrangement of sensing elements responsive to condition of working-fluid, e.g. pressure
- F01D17/085—Arrangement of sensing elements responsive to condition of working-fluid, e.g. pressure to temperature
-
- 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
-
- 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
- F05D2220/00—Application
- F05D2220/30—Application in turbines
- F05D2220/32—Application in turbines in gas turbines
- F05D2220/321—Application in turbines in gas turbines for a special turbine stage
- F05D2220/3212—Application in turbines in gas turbines for a special turbine stage the first stage of a turbine
-
- 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
- F05D2240/126—Baffles or ribs
-
- 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
- F05D2270/00—Control
- F05D2270/01—Purpose of the control system
- F05D2270/08—Purpose of the control system to produce clean exhaust gases
-
- 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
- F05D2300/00—Materials; Properties thereof
- F05D2300/20—Oxide or non-oxide ceramics
- F05D2300/21—Oxide ceramics
-
- 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
- F05D2300/00—Materials; Properties thereof
- F05D2300/50—Intrinsic material properties or characteristics
- F05D2300/502—Thermal properties
- F05D2300/5021—Expansivity
- F05D2300/50212—Expansivity dissimilar
-
- 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
- F05D2300/00—Materials; Properties thereof
- F05D2300/60—Properties or characteristics given to material by treatment or manufacturing
- F05D2300/603—Composites; e.g. fibre-reinforced
-
- Y—GENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
- Y02—TECHNOLOGIES OR APPLICATIONS FOR MITIGATION OR ADAPTATION AGAINST CLIMATE CHANGE
- Y02T—CLIMATE CHANGE MITIGATION TECHNOLOGIES RELATED TO TRANSPORTATION
- Y02T50/00—Aeronautics or air transport
- Y02T50/60—Efficient propulsion technologies, e.g. for aircraft
Definitions
- the invention relates to a high pressure turbine distributor of a turbojet engine; it relates more particularly to an improvement to limit the emission of harmful and / or polluting gases, particularly during the maneuvering phases on the ground, on airports where such discharges pose the most problems in terms of the environment.
- the high-pressure distributor is a set of fixed blades arranged annularly in the gas flow passage between the outlet of the combustion chamber and the inlet of the high-pressure turbine.
- This subset is therefore the one that is subjected to the highest gas temperatures, exceeding the melting temperature of the materials (metals) used.
- This is typically the case during so-called "hot" operating phases of the engine, typically the take-off phase, the climb phase and the cruise phase. It is therefore necessary to constantly cool this subassembly and in particular the blades.
- the blades extend radially between two annular, segmented platforms defining the flow vein of the gases ejected by the combustion chamber.
- the blades are hollow.
- Each blade comprises at least one and very generally two adjacent cavities, one end of which opens out of the vein, through a platform mentioned above.
- the blades have holes extending between the or each cavity and the vein. For example, there is a leading edge cavity (the side of the combustion chamber) opening radially inwardly and a trailing edge cavity (on the side of the high pressure turbine) opening radially outwardly.
- a jacket is embedded in the or each cavity of such a blade.
- This jacket has a plurality of holes on its entire surface extending between the annular platforms.
- the jacket has a bottom wall, closed at one end and opens out of the vein, on the same side as the cavity that encloses it.
- Relatively fresh air is drawn upstream of the combustion chamber at the outlet of a stage of the compressor. This air is injected into the shirts through the platforms. Fresh air enters the interior of the jacket to cool the blade from the inside by impact. Then, escaping into the vein via the holes of the blade, the air of Cooling creates a relatively cooler air barrier film flowing along the outer surface of the blade.
- the flow section defined by the holes is calculated to define a ventilation capable of suitably cooling the blade when the engine is in a "hot" operating phase as defined above.
- the disadvantage of this system is that the fresh air taken from the compressor, bypassing the combustion chamber, does not participate in the combustion, which contributes to increase pollution (including compounds CO, CH and NOx).
- the invention follows from the following analysis:
- the turbojet engine operates with lower flue gas temperatures in the combustion chamber. This is particularly the case of maneuvers on the ground, idling, that is to say, the majority of the maneuvers that take place on the airport, specifically where it is most interesting to reduce pollution.
- the invention proposes a system for automatically regulating the flow of cooling air passing through the distributor, which is well adapted to the temperatures encountered under different operating conditions.
- the basic idea is to reduce the flow of cooling air during so-called "cold" phases of the engine, in particular to reduce pollutant emissions, while reducing the amount of air taken from the compressor.
- the invention relates to a high-pressure turbine distributor of a turbojet, comprising a ring of fixed and hollow blades arranged between two coaxial annular platforms delimiting a gas flow channel and in which each blade comprises at least one cavity housing a first sleeve, one end of which opens out of said vein through a platform and has a bottom wall at the other end, said first sleeve being pierced with a plurality of holes opening into said cavity, characterized in that a second liner of material having a coefficient of expansion different from that of said first liner is engaged within each liner and shaped so that its outer wall is substantially in contact with the inner liner of said first liner shirt and in that this second The liner is pierced with a plurality of holes substantially in correspondence with those of said first liner, one end of said second liner being attached to said bottom of said first liner.
- the holes made in each first liner and those made in the corresponding second liner are in maximum correspondence for a nominal operating temperature corresponding to a hot operating phase of the turbojet engine.
- each blade has a first leading edge cavity liner opening out of the vein through a platform and a first trailing edge cavity liner opening out of the vein through the other platform.
- Each first leading edge and trailing edge cavity liner houses a second aforesaid liner of corresponding shape and size.
- said first jackets are made of metal while said second jackets are made of composite material with a low coefficient of expansion. This coefficient of expansion is significantly lower than that of the metal used to form said first folders.
- said second shirts are ceramic.
- each first liner and the second liner therein are apairées.
- the bores can be made simultaneously hot. More particularly, the apairées shirts can be pierced together in a chamber brought to a aforementioned nominal temperature corresponding to a hot operating phase of the turbojet engine.
- FIG. 1 is a partial external perspective view of the dispenser according to the invention.
- FIG. 1 is a partial perspective view inward of the same distributor
- FIG. 3 illustrates a leading edge cavity metal jacket
- FIG. 4 illustrates a liner of corresponding composite material
- Figure 5 is a perspective view of a metal jacket of a trailing edge cavity
- FIG. 6 is a perspective view of the corresponding composite material jacket.
- a high-pressure turbine distributor of a turbojet engine is constituted by a fixed blade crown 13 made by side-by-side assembly of several segments 14 according to FIGS. 1 and 2.
- the blades are hollow and arranged between two annular platforms 16, 18 coaxial formed by the side-by-side assembly of the segments 14.
- the annular platforms thus delimit a gas flow stream 20 in which are located the blades 13 regularly distributed angularly between the platforms 16,18.
- Each hollow blade here comprises a leading edge cavity 24 opening out of the vein through the inner platform 18 and a trailing edge cavity 26 opening out of the vein through the outer platform. 16.
- each hollow blade is divided into two adjacent cavities, one 24 located on the leading edge side and the other 26 on the trailing edge side. These cavities communicate with the vein by rows of radially extending holes. It will be understood, for example, that the rows of holes 31 extend between the leading edge cavity and the vein while the rows of holes extend between the trailing edge cavity and the vein.
- the first leading-edge cavity liner 34 opens out of the vein through a platform, here the inner platform 18 while the first liner 36 of the trailing-edge cavity opens to the outside. outside the vein through the other platform, here the outer platform 16.
- relatively fresh air is taken from the compressor and led on both sides of the distributor, that is to say both outside the outer platform 16 and inside the inner platform 18.
- the cooling air can therefore enter the jackets to cool the inner walls of the blades by impact effect and then flow into the vein through the holes of said blades 13 to create a cooling film around each of them.
- a second jacket 44, 46 of material having a coefficient of expansion different from that of said first jacket 34, 36 is engaged inside each of them and shaped so that its outer wall or substantially in contact with the inner wall of said corresponding first sleeve.
- This second sleeve 44, 46 is pierced with a plurality of holes 47 substantially in correspondence with those of said first sleeve.
- one end 48a, 48b of the second liner is attached to a bottom 38a, 38b, closed, of said first liner, respectively, that is to say opposite the opening by which the air fresh is introduced.
- each first jacket and those in the corresponding second jacket are in maximum correspondence (i.e. ensuring a maximum flow section) for a nominal operating temperature corresponding to hot operating phases of the turbojet engine.
- the second shirts are made of composite material having a coefficient of weak expansion, in particular significantly lower than that of the metal which constitutes said first folders.
- said second shirts are ceramic.
- the second jacket 44, 46 of composite material is slightly “longer” radially than the metal jacket that contains it.
- the holes are only partially matched, two by two, and the passage section is reduced.
- the amount of air drawn will be relatively low, which is acceptable since in such a phase of operation the distributor hardly requires to be cooled.
- the rise in the temperature of the combustion gases causes an expansion of the metal jackets 34, 36, that is to say their elongation in the direction radial. This results in a larger passage section for cooling air. The larger cooling air flow therefore effectively cools the dispenser.
- each first sleeve 34, 36 and the second sleeve 44, 46 contained therein are apairées.
- the two apairées shirts can be nested one inside the other and drilled together in a chamber brought to a high nominal temperature as defined above.
- the seal between the two shirts 38a, 48a - 38b, 48b is as good as possible at all phases of operation of the engine.
- the second jacket, made of composite material, is set cold in the first metal jacket and the ribs of these two parts have been defined so that the insertion of the one into the other is done with zero or extremely low. Therefore, the seal is guaranteed during the cold running phases of the engine.
- the relative expansion of the metal jacket relative to the composite material jacket, in the width direction is of the order of 4 hundredths of a millimeter, which is very low and can be considered as having no influence on the tightness between the shirts.
Landscapes
- Engineering & Computer Science (AREA)
- Mechanical Engineering (AREA)
- General Engineering & Computer Science (AREA)
- Turbine Rotor Nozzle Sealing (AREA)
Abstract
Description
Claims
Priority Applications (2)
Application Number | Priority Date | Filing Date | Title |
---|---|---|---|
US13/522,163 US9328618B2 (en) | 2010-01-14 | 2011-01-06 | High-pressure turbine nozzle for a turbojet |
GB1211902.0A GB2488958B (en) | 2010-01-14 | 2011-01-06 | A high-pressure turbine nozzle for a turbojet |
Applications Claiming Priority (2)
Application Number | Priority Date | Filing Date | Title |
---|---|---|---|
FR1050229A FR2955145B1 (fr) | 2010-01-14 | 2010-01-14 | Distributeur de turbine haute pression d'un turboreacteur |
FR1050229 | 2010-01-14 |
Publications (1)
Publication Number | Publication Date |
---|---|
WO2011086305A1 true WO2011086305A1 (fr) | 2011-07-21 |
Family
ID=42371219
Family Applications (1)
Application Number | Title | Priority Date | Filing Date |
---|---|---|---|
PCT/FR2011/050017 WO2011086305A1 (fr) | 2010-01-14 | 2011-01-06 | Distributeur de turbine haute pression d'un turboreacteur |
Country Status (4)
Country | Link |
---|---|
US (1) | US9328618B2 (fr) |
FR (1) | FR2955145B1 (fr) |
GB (1) | GB2488958B (fr) |
WO (1) | WO2011086305A1 (fr) |
Families Citing this family (11)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
US9151173B2 (en) * | 2011-12-15 | 2015-10-06 | General Electric Company | Use of multi-faceted impingement openings for increasing heat transfer characteristics on gas turbine components |
US10012092B2 (en) * | 2015-08-12 | 2018-07-03 | United Technologies Corporation | Low turn loss baffle flow diverter |
FR3086329B1 (fr) | 2018-09-26 | 2020-12-11 | Safran Aircraft Engines | Distributeur ameliore de turbomachine |
US11702941B2 (en) * | 2018-11-09 | 2023-07-18 | Raytheon Technologies Corporation | Airfoil with baffle having flange ring affixed to platform |
US10774657B2 (en) | 2018-11-23 | 2020-09-15 | Raytheon Technologies Corporation | Baffle assembly for gas turbine engine components |
FR3094034B1 (fr) * | 2019-03-20 | 2021-03-19 | Safran Aircraft Engines | Chemise tubulaire de ventilation pour un distributeur de turbomachine |
FR3094743B1 (fr) | 2019-04-03 | 2021-05-14 | Safran Aircraft Engines | Aube améliorée pour turbomachine |
FR3099793B1 (fr) | 2019-08-06 | 2022-07-29 | Safran Aircraft Engines | tronçon de distributeur de turbine comportant un chemisage interne |
PL431184A1 (pl) * | 2019-09-17 | 2021-03-22 | General Electric Company Polska Spółka Z Ograniczoną Odpowiedzialnością | Zespół silnika turbinowego |
FR3107733B1 (fr) * | 2020-02-28 | 2022-07-08 | Safran Aircraft Engines | Pale de distributeur haute ou basse pression pour turbomachine, distributeur et turbomachine comportant de telles pales |
FR3109961B1 (fr) | 2020-05-06 | 2022-05-13 | Safran Aircraft Engines | Distributeur en CMC amélioré pour turbine de turbomachine |
Citations (5)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
EP0266235A1 (fr) * | 1986-10-01 | 1988-05-04 | Societe Nationale D'etude Et De Construction De Moteurs D'aviation, "S.N.E.C.M.A." | Turbomachine munie d'un dispositif de commande automatique des débits de ventilation de turbine |
EP1284338A2 (fr) * | 2001-08-13 | 2003-02-19 | General Electric Company | Insert avec sortie tangentielle de fluide de refroidissement |
EP1840331A1 (fr) * | 2006-03-29 | 2007-10-03 | Snecma | Ensemble d'une aube et d'une chemise de refroidissement, distributeur de turbomachine comportant l'ensemble, turbomachine, procédé de montage et de réparation de l'ensemble |
EP1936468A1 (fr) * | 2006-12-22 | 2008-06-25 | Siemens Aktiengesellschaft | Bilames d'ajustement d'un canal de refroidissement |
GB2457073A (en) * | 2008-02-04 | 2009-08-05 | Rolls-Royce Plc | Modulating cooling airflows |
Family Cites Families (1)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
EP1806203A1 (fr) * | 2006-01-10 | 2007-07-11 | Siemens Aktiengesellschaft | Méthode de fabrication d'un trou |
-
2010
- 2010-01-14 FR FR1050229A patent/FR2955145B1/fr active Active
-
2011
- 2011-01-06 US US13/522,163 patent/US9328618B2/en active Active
- 2011-01-06 GB GB1211902.0A patent/GB2488958B/en active Active
- 2011-01-06 WO PCT/FR2011/050017 patent/WO2011086305A1/fr active Application Filing
Patent Citations (5)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
EP0266235A1 (fr) * | 1986-10-01 | 1988-05-04 | Societe Nationale D'etude Et De Construction De Moteurs D'aviation, "S.N.E.C.M.A." | Turbomachine munie d'un dispositif de commande automatique des débits de ventilation de turbine |
EP1284338A2 (fr) * | 2001-08-13 | 2003-02-19 | General Electric Company | Insert avec sortie tangentielle de fluide de refroidissement |
EP1840331A1 (fr) * | 2006-03-29 | 2007-10-03 | Snecma | Ensemble d'une aube et d'une chemise de refroidissement, distributeur de turbomachine comportant l'ensemble, turbomachine, procédé de montage et de réparation de l'ensemble |
EP1936468A1 (fr) * | 2006-12-22 | 2008-06-25 | Siemens Aktiengesellschaft | Bilames d'ajustement d'un canal de refroidissement |
GB2457073A (en) * | 2008-02-04 | 2009-08-05 | Rolls-Royce Plc | Modulating cooling airflows |
Also Published As
Publication number | Publication date |
---|---|
FR2955145A1 (fr) | 2011-07-15 |
FR2955145B1 (fr) | 2012-02-03 |
GB201211902D0 (en) | 2012-08-15 |
US20130051980A1 (en) | 2013-02-28 |
US9328618B2 (en) | 2016-05-03 |
GB2488958B (en) | 2016-04-13 |
GB2488958A (en) | 2012-09-12 |
Similar Documents
Publication | Publication Date | Title |
---|---|---|
WO2011086305A1 (fr) | Distributeur de turbine haute pression d'un turboreacteur | |
CA2567878C (fr) | Fond de chambre de combustion avec ventilation | |
EP2042806B1 (fr) | Chambre de combustion d'une turbomachine | |
EP1612374B1 (fr) | Aube fixe de turbine à refroidissement amélioré | |
EP2472067B1 (fr) | Intégration d'un échangeur de chaleur surfacique avec débit d'air régulé dans un moteur d'avion | |
CA2782661C (fr) | Chambre de combustion pour turbomachine | |
FR3113732A1 (fr) | Capteur de température d’air | |
EP1445421B1 (fr) | Dispositif de ventilation d'un rotor de turbine à haute pression d'une turbomachine | |
FR3068128A1 (fr) | Capteur de temperature d'air | |
FR3051016A1 (fr) | Dispositif de degivrage d'un bec de separation de turbomachine aeronautique | |
CA2925565C (fr) | Chambre de combustion de turbomachine pourvue de moyens de deflection d'air pour reduire le sillage cree par une bougie d'allumage | |
WO2016156741A1 (fr) | Module de soufflante de turbomachine comprenant un système de dégivrage d'un cône d'entrée de turbomachine et procédé de dégivrage | |
EP2643069B1 (fr) | Dispositif d'evacuation d'huile et turbomachine comprenant un tel dispositif | |
FR2999249A1 (fr) | Compresseur pour turbomachine dote de moyens de refroidissement d'un joint tournant assurant l'etancheite entre un redresseur et un rotor | |
FR3009747A1 (fr) | Chambre de combustion de turbomachine pourvue d'un passage d'entree d'air ameliore en aval d'un orifice de passage de bougie | |
FR2973479A1 (fr) | Paroi pour chambre de combustion de turbomachine comprenant un agencement optimise d'orifices d'entree d'air et de passage de bougie d'allumage | |
FR3080406A1 (fr) | Distributeur de turbine ameliore pour turbomachine | |
EP4146913B1 (fr) | Distributeur en cmc amélioré pour turbine de turbomachine | |
EP3969813B1 (fr) | Chambre de combustion comprenant des moyens de refroidissement d'une zone d'enveloppe annulaire en aval d'une cheminée | |
FR3087840A1 (fr) | Capot de nacelle pour ensemble propulsif d'aeronef | |
EP3262348B1 (fr) | Chambre de combustion de turbomachine comportant une pièce pénétrante avec ouverture | |
WO2022223916A1 (fr) | Dispositif d'injection de carburant pour postcombustion de turboreacteur | |
WO2023180668A1 (fr) | Module pour turbomachine d'aéronef | |
WO2021148740A1 (fr) | Dispositif amélioré de détection d'anomalie de refroidissement pour turbomachine d'aéronef | |
FR3136016A1 (fr) | Anneau accroche-flammes pour postcombustion de turboreacteur comprenant un conduit pour chauffer un segment angulaire de l'anneau |
Legal Events
Date | Code | Title | Description |
---|---|---|---|
121 | Ep: the epo has been informed by wipo that ep was designated in this application |
Ref document number: 11704256 Country of ref document: EP Kind code of ref document: A1 |
|
ENP | Entry into the national phase |
Ref document number: 1211902 Country of ref document: GB Kind code of ref document: A Free format text: PCT FILING DATE = 20110106 |
|
WWE | Wipo information: entry into national phase |
Ref document number: 1211902.0 Country of ref document: GB |
|
NENP | Non-entry into the national phase |
Ref country code: DE |
|
WWE | Wipo information: entry into national phase |
Ref document number: 13522163 Country of ref document: US |
|
122 | Ep: pct application non-entry in european phase |
Ref document number: 11704256 Country of ref document: EP Kind code of ref document: A1 |