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公开(公告)号：US20210262665A1
公开(公告)日：2021-08-26
申请号：US17049442
申请日：2019-04-10
Applicant: SAFRAN AIRCRAFT ENGINES
Abstract: An injection system for a turbine engine annular combustion chamber includes a support configured to support and to center a fuel injector head. The support includes a frustoconical surface connected at its downstream end of smallest diameter to an upstream end of a cylindrical surface. The system further includes a bowl configured to mix air and fuel arranged downstream of the support and at least one axial swirl inducer extending at least in part around the support. Each swirl inducer includes vanes delimiting between them substantially axial channels for the passage of an air flow. The channels open at their upstream ends on said frustoconical surface.
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公开(公告)号：US11125435B2
公开(公告)日：2021-09-21
申请号：US15742447
申请日：2016-07-07
Applicant: Safran Aircraft Engines
Abstract: A turbine engine combustion chamber including: an outer annular housing; a flame tube connected to the outer housing. The flame tube includes an inner annular wall and an outer annular wall and a second axial outlet portion of the flame tube. The flame tube also includes a chamber base located at the inlet of the flame tube; and a fuel injection system configured to inject fuel into the flame tube via the inlet of the flame tube. The injection system includes an injector axis, and an air manifold to move air towards twists in the injection system. The twists are arranged around an implantation axis. The air manifold includes a circular portion around the injector axis. The circular portion, forms an air inlet of the manifold. The opening places the entering air flow in rotation about the implantation axis.
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公开(公告)号：US12025313B2
公开(公告)日：2024-07-02
申请号：US17630284
申请日：2020-07-27
Applicant: SAFRAN AIRCRAFT ENGINES
Inventor： Julien Marc Matthieu Leparoux , Jean-François Cabre , Haris Musaefendic , Romain Nicolas Lunel
CPC classification number: F23R3/343 , F02C7/22 , F23R3/50 , F23R2900/03343
Abstract: A combustion chamber for an aircraft turbomachine includes an annular chamber end wall structure, an annular row of main injection systems mounted in the chamber end wall structure and configured to deliver a sheet of fuel, including a central recirculation region and a corner recirculation region around the central recirculation region, and secondary injection systems each configured to inject an additional flow of air and fuel directly into a corresponding corner recirculation region.
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公开(公告)号：US11933497B2
公开(公告)日：2024-03-19
申请号：US17779789
申请日：2020-11-24
Applicant: SAFRAN AIRCRAFT ENGINES
Inventor： Romain Nicolas Lunel , Haris Musaefendic
Abstract: The invention relates to an air/fuel injection system for a turbomachine, comprising:—an injector comprising a duct and an injection nose, arranged inside said duct, which extends from upstream to downstream along a longitudinal axis;—a mixer device comprising a bowl comprising an annular inlet, forming the inlet of the mixer device, from which there extends a conical portion flared in the downstream direction, said mixer device being arranged downstream of the injection nose; the injection system being characterized in that the injector comprises an air-injection annulus extending from the duct and from which there extends a connection ring comprising a divergent portion, said ring being arranged externally around the annular inlet of the mixer device.
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公开(公告)号：US20230250962A1
公开(公告)日：2023-08-10
申请号：US18014326
申请日：2021-07-02
Applicant: Safran Aircraft Engines
Inventor： Haris Musaefendic , Romain Nicolas Lunel , Jean-François Cabre
Abstract: Annular combustion chamber for an aircraft turbomachine, said chamber having two coaxial annular walls, an inner annular wall and an outer annular wall, respectively, which are connected upstream by an annular bottom wall of the chamber, wherein an injection device passes through an axis and comprises an air injection system and a frustoconical bowl which is flared downstream and has air passage openings, the chamber further having an annular deflector placed downstream of the bottom wall substantially parallel to the latter; and wherein the air injection system, the bottom wall, the deflector and the bowl are integrally formed.
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Patent Agency Ranking
公开(公告)号：US11268699B2
公开(公告)日：2022-03-08
申请号：US17049442
申请日：2019-04-10
Applicant: SAFRAN AIRCRAFT ENGINES
Abstract: An injection system for a turbine engine annular combustion chamber includes a support configured to support and to center a fuel injector head. The support includes a frustoconical surface connected at its downstream end of smallest diameter to an upstream end of a cylindrical surface. The system further includes a bowl configured to mix air and fuel arranged downstream of the support and at least one axial swirl inducer extending at least in part around the support. Each swirl inducer includes vanes delimiting between them substantially axial channels for the passage of an air flow. The channels open at their upstream ends on said frustoconical surface.
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公开(公告)号：US10883718B2
公开(公告)日：2021-01-05
申请号：US16095813
申请日：2017-04-28
Applicant: SAFRAN AIRCRAFT ENGINES
Inventor： Romain Nicolas Lunel , Guillaume Aurelien Godel , Haris Musaefendic , Christophe Pieussergues , Francois Pierre Georges Maurice Ribassin
Abstract: An air intake swirler for a turbomachine injection system includes an upstream wall and a downstream wall, both of revolution about an axis of the air intake swirler, and fins distributed about the axis and connecting the upstream wall to the downstream wall so as to delimit, between the upstream wall and the downstream wall, air inlet channels each having an inlet and an outlet. The swirler includes two aerodynamic deflectors that respectively extend the downstream walls radially outward and that have a concavity oriented upstream. The aerodynamic deflectors extend radially facing the respective inlets of the air inlet channels and thus make it possible to limit the loss of pressure of the air supplied to the air inlet channels.
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公开(公告)号：US11988387B2
公开(公告)日：2024-05-21
申请号：US18014326
申请日：2021-07-02
Applicant: Safran Aircraft Engines
Inventor： Haris Musaefendic , Romain Nicolas Lunel , Jean-François Cabre
Abstract: Annular combustion chamber for an aircraft turbomachine, said chamber having two coaxial annular walls, an inner annular wall and an outer annular wall, respectively, which are connected upstream by an annular bottom wall of the chamber, wherein an injection device passes through an axis and comprises an air injection system and a frustoconical bowl which is flared downstream and has air passage openings, the chamber further having an annular deflector placed downstream of the bottom wall substantially parallel to the latter; and wherein the air injection system, the bottom wall, the deflector and the bowl are integrally formed.
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公开(公告)号：US11840994B2
公开(公告)日：2023-12-12
申请号：US16517863
申请日：2019-07-22
Applicant: SAFRAN AIRCRAFT ENGINES
IPC: F02M55/00
CPC classification number: F02M55/008
Abstract: The fuel injection conduits in a multipoint device surrounding a so-called pilot central injection device include tubes of circumferential orientation. By separating the injection conduits from each other, it is possible to attribute to them different head losses which compensate the differences in length that the fuel has to travel: a uniform flow of fuel may be hoped for, for each of the injection holes. The tubes are individual but joined to form a crown that is unitary or composed of two almost symmetrical unitary portions, which lends itself well to manufacture by addition of material.
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公开(公告)号：US11739936B2
公开(公告)日：2023-08-29
申请号：US17419919
申请日：2019-12-18
Applicant: SAFRAN AIRCRAFT ENGINES
Abstract: An injection system for a turbomachine combustion chamber includes a swirler and a mixing bowl. The mixing bowl includes a converging frustoconical portion and a diverging frustoconical portion. The diverging frustoconical portion is connected to the converging frustoconical portion, forming a continuous aerodynamic profile with the converging frustoconical portion. The diverging frustoconical portion is passed through by vortex holes which each includes a circumferential component around a longitudinal axis of the injection system and an axial component along the longitudinal axis of the injection system.
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