Safran Aircraft Engines

Julien Marc Matthieu Leparoux

Haris Musaefendic

Romain Nicolas Lunel

Christophe Pieussergues

Jean-François Cabre

Joseph Jean Marie Burguburu

Zakaria El Gnaoui

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

A combustion chamber for a turbomachine including an annular casing delimiting an inner volume of the combustion chamber and provided with a chimney extending to the outside of the inner volume and delimiting a passage for a penetrating part, and a bushing mounted floating on the chimney, further includes: a main air collection chamber open to the upstream arranged facing the downstream portion of the chimney; and a through-opening of the annular casing, connecting the inner volume to the main air collection chamber. The air collected by the main air collection chamber is injected into the inner volume via the through-opening and can cool an annular casing zone downstream of the chimney.

Combustion chamber comprising means for cooling an annular casing zone downstream of a chimney

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

A combustion chamber for an aircraft turbomachine includes an annular chamber end wall and an annular row of injection systems mounted in the annular chamber end wall. Each injection system includes at least one air inlet swirler and a main fuel injection nozzle to output a fuel stream centered on an injection axis and including a central recirculation zone and a corner recirculation zone extending as an annulus around the central recirculation zone chamber. The combustion chamber also includes a plurality of additional fuel injection devices mounted in the chamber end wall to inject fuel directly into the corresponding corner recirculation zones produced by the corresponding injection systems at an operating speed less than or equal to the idling speed.

Combustion chamber comprising additional injection devices opening up directly into corner recirculation zones, turbomachine comprising such a chamber and fuel supply method for such a chamber

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Paris, France

Julien Marc Matthieu Leparoux