Snecma

Safran Aircraft Engines

Herakles

Snecma Propulsion Solide

Romain Nicolas Lunel

Didier Hippolyte Hernandez

Haris Musaefendic

Thomas Olivier Marie Noel

Patrice Andre Commaret

Denis Jean Maurice Sandelis

Mathieu Dakowski

Christophe Pieussergues

Patrice André Commaret

Stephane Pierre Guillaume Blanchard

Fabrice Marcel Noel Garin

Thomas Daris

Matthieu Francois Rullaud

Sylvain Duval

Sébastien Alain Christophe Bourgois

Guillaume Aurelien Godel

Jacques Marcel Arthur Bunel

Alain Dominique Gendraud

Guillaume Sevi

David Locatelli

Laurent Bernard Cameriano

Michel Pierre Cazalens

Sebastien Alain Christophe Bourgois

Michel Andre Albert Desaulty

Jean-François Cabre

François Xavier Chapelle

Yvan Yoann Guezel

Daniel Georges Plona

Antoine Robert Alain Brunet

Dominique Raulin

Clement Roussille

Clément Roussille

Thomas Olivier Noel

Alain Rene Cayre

Guy Lapergue

Benjamin Frantz Karl Villenave

Jeremy Giordan

Jean-Christophe Leininger

Dan-Ranjiv Joory

Clement Bernard

Julien Marc Matthieu Leparoux

Laurent Pierre Elysee Gaston Marnas

Pierre André Gabriel Malbois

Laurent Pierre Marnas

Galadriel Dancie

Sebastien Loval

Olivier Kreder

Dan Ranjiv Joory

Daniel Olivier Bainachi

Clément Jean Henri Bechet

Emmanuel Jean André Greslin

Jérémy Giordan

Denis Jean Sandelis

Bruno Andre Jean Varin

Thierry Lequitte

Damien Bonnefoi

Fabrice Marcel Noël Garin

Laurent Marnas

Pascale Rollet

Dominique Guezengar

Daniel Francis Paul Bernier

Luc Namer

Francois Pierre Georges Maurice Ribassin

Matthieu François Rullaud

Frederic Dos Santos

Aurelien Lionel Tedesco

David Pinchon

Guillaume Aurélien Godel

François Pierre Ribassin

Fabrice Marcel Garin

Jacques Marcel Bunel

Daniel Francis Bernier

Antoine Robert Brunet

Michel Andre Desaulty

Didier Hyppolyte Hernandez

Didier-Hippolyte Hernandez

François Pierre Georges Maurice Ribassin

Sébastien Loval

Thomas Olivier Marie NOËL

Clément Bernard

Frédéric Dos Santos

Denis Sandelis

Aurélien Lionel Tedesco

Michel Desaulty

Alain René Cayre

Thomas Noël

Daniel Bernier

Pascale Valérie Cothilde Rollet

Bruno André Jean Varin

A diffusion cone for the rear part of a jet engine has a base and, on an opposite side, a truncated tip defining an annular trailing edge, and comprises a flame-holder ring with a cross section comprising an external branch and an internal branch which are connected to one another on the base side and such that the external branch extends around the internal branch, whereby the external and internal branches delimit between them an internal space opening on the opposite side to the base. One of the branches is made up of the annular trailing edge. The flame-holder ring comprises at least one fuel inlet designed to let fuel into the internal space. Such a flame-holder ring makes it possible to ensure and to control the propagation of the flame within the afterburner channel.

Diffusion cone for the rear part of a jet engine incorporating a flame-holder ring at the trailing edge

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

The invention concerns a turbo engine comprising a combustion chamber () arranged inside the outer housing () and comprising an internal revolution wall () and an external revolution wall (). First stop parts () and second stop parts () fixed respectively to the outer housing and to the combustion chamber are provided, the first and second stop parts being adapted to come to a substantially axial stop two by two by forming stop pairs ().

Combustion module for a gas turbo engine with chamber bottom stop

A combustion chamber for a turbine engine, in particular for an aircraft turbojet engine or turboprop engine. The combustion chamber includes a radially outer annular shroud, a radially inner annular shroud coaxial with the radially outer shroud, and an end wall connecting the radially outer shroud and the radially inner shroud. The combustion chamber further includes a first annular sealing member coaxial with said radially inner and outer shrouds. The first annular sealing member is radially interposed between the end wall and the radially outer shroud.

Turbine engine combustion chamber

The provision of air passage holes through a wall of a gas turbomachine combustion chamber. Multi-perforations are virtually positioned and distributed, even in a first safety zone without air passage openings. Multi-perforations with a virtual inlet or outlet in this first security zone are virtually removed. According to certain criteria, at least some of said removed multi-perforations are then virtually reintegrated, and, from then on a perimeter passing through the virtual inlets and outlets of all the multi-perforations present is defined, in the direction of a primary or dilution hole to be installed, a modified safety zone is defined, then, respecting around said hole and with the freedom to reposition it within this limit, the shape of this hole is redefined.

Configuring and positioning air passage holes in a combustion chamber wall

The invention relates to a bottom wall () of a gas turbomachine combustion chamber. This bottom wall comprises openings () for mounting combustion air supply systems and holes () for the passage of cooling air between at least one inlet () and at least one outlet of said holes. The holes () extend inwardly along the bottom wall with the outlet port located closer to the opening () adjacent to it than the inlet hole.

Turbine engine combustion chamber bottom

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 combustion chamber for a turbomachine having a bottom wall, at least one mixing bowl for promoting the mixing of air and fuel, mounted in an opening in the bottom wall, at least one annular baffle mounted axially downstream of the bottom wall, with respect to the direction of the gas flow within the combustion chamber, around the opening. The baffle is produced in one piece with the mixing bowl so as to form a one-piece assembly that has at least one channel for the flow of cooling air. The channel has an air inlet located upstream of the bottom wall and an air outlet located downstream of the bottom wall. The air outlet is located radially opposite the baffle.

Combustion chamber for a turbomachine

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

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.

Air intake swirler for a turbomachine injection system comprising an aerodynamic deflector at its inlet

Growing community of inventors

Montereau sur le Jard, France

Romain Nicolas Lunel