Safran Aircraft Engines

Haris Musaefendic

Romain Nicolas Lunel

Sébastien Alain Christophe Bourgois

Jean-François Cabre

Julien Marc Matthieu Leparoux

Guillaume Aurelien Godel

Christophe Pieussergues

Alain Rene Cayre

Patrice André Commaret

Francois Pierre Georges Maurice Ribassin

Zakaria El Gnaoui

Guillaume Aurélien Godel

François Pierre Georges Maurice Ribassin

Alain René Cayre

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.

Combustion chamber comprising secondary injection systems, and fuel supply method

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.

Annular combustion chamber for an aircraft turbomachine

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.

Fuel injection system for a turbomachine, combustion chamber comprising such a system, and associated turbomachine

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.

Multipoint fuel injection device

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.

Injection system for turbomachine, comprising a swirler and mixing bowl vortex holes

An annular combustion chamber of a turbomachine is described. The combustion chamber has an axis of revolution and is delimited by coaxial internal and external annular walls joined upstream by a bottom of chamber substantially transverse to the walls. In some embodiments, the chamber includes at least one annular deflector placed in the chamber and substantially parallel to the bottom of chamber. The bottom of chamber may have openings adapted to be traversed by air for cooling the deflector. In some embodiments, the bottom of chamber and the deflector include mounting openings for mounting an annular row of injection devices for injecting a mixture of air and fuel into the chamber. At least a portion of the air for cooling the deflector is conveyed into the chamber through holes in the injection devices.

Combustion chamber of a turbomachine

An injection system for a turbine engine annular combustion chamber includes The system further includes a bowl configured to mix air and fuel arranged downstream of the support and The channels open at their upstream ends on said frustoconical surface.

Injection system for a turbine engine annular combustion chamber

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.

Bent combustion chamber from a turbine engine

To improve air/fuel mixing at low speed, the invention concerns an aerodynamic injection system () for a combustion chamber () of an aircraft turbine engine, the system comprising a central body () comprising a first air circulation space () and a swirler () allowing air to penetrate into the first air circulation space (), the hollow body () also comprising a fuel circuit () arranged at least partially around the first space (), the circuit comprising a second annular space () for a fuel film to circulate and fuel supply structures () communicating with the second space (), the spaces () opening into a third air and fuel mixing space (). According to the invention, the system comprises fuel injection openings () that bring the fuel circuit into communication with the first space ().

Aerodynamic injection system for aircraft turbine engine, having improved air/fuel mixing

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.

Growing community of inventors

Moissy-Cramayel, France

Haris Musaefendic