Societe Nationale D'etude Et De Construction De Moteurs D'aviation

Societe Nationale D'etude Et Moteurs D'aviation 's.n.e.c.m.a.'

Philippe M Gastebois

Jacques E Caruel

Gerard E Jourdain

Marc G Loubet

Herve Alain Quillevere

Jean-Francois R Imbault

Alain G Habrard

Jean L Picard

Jean Robert Bedue

Bruno Deroide

Jean-Paul R Gaillac

Simone Coutor

A transition pipe for location between an afterburner pipe and a jet nozzle in a jet pipe assembly for a turbojet engine comprises a funnel shaped inner casing tapering downstream from a circular section at its upstream end to a quadrangular section at its downstream end, a horn shaped outer casing surrounding the inner casing and flared downstream from a circular upstream end section to a downstream end section formed of four joined segments of generally parabolic shape, and a plurality of peripherally spaced, longitudinally extending stiffeners interconnecting the inner and outer casings in the space therebetween.

Transition pipe for a jet pipe assembly of a turbojet engine

A flow mixer and flame stabilizer structure for an afterburner of a turbofan engine is disclosed having a variable mixer to mix a portion of the cold flow with the hot flow gasses and which also forms a flame stabilizer for a secondary afterburner fuel injection manifold which injects fuel into the cold flow air during afterburner operation. The device forms a flame stabilizer around the secondary, annular fuel injection manifold and has a generally 'U' shaped cross-section. One of the legs forming the 'U' shape is fixed, while the opposite leg is movable so as to vary the downstream facing opening between the legs. The movable leg also blocks a portion of the cold flow duct to generate localized increases in pressure upstream of the stabilizer. An air scoop located in this upstream portion communicates with the primary afterburner flame stabilizer to direct pressurized air to cool the primary flame stabilizer.

Flow mixer and flame stabilizer for a turbofan engine

A turbine rotor has composite ceramic blades each having a bulb-shaped root with planar flanks which define an angle of convergence between them not exceeding ten degrees, and by which the root is held in a dovetail section socket in the periphery of the rotor disc by means of wedge-shaped members also disposed in the socket on opposite sides of the root. Each wedge-shaped member is part of an element including a platform which is disposed adjacent the curved portion of the blade radially outwards of the socket and which is joined to the wedge-shaped member by an integral strut. The platforms of all of the elements extend in a circumferential direction and combine to define a continuous ring forming part of the inner boundary of the gas flow path through the turbine.

Turbine rotor

A device for cooling the wall of a combustion chamber, in which air is taken from the space outside the chamber by means of an annular protuberance which extends beyond the outside of the chamber and forms an annular pocket, which is provided with outside air by means of holes and communicates with the interior of the chamber by means of an annular slot between the downstream internal wall of the chamber and a small annular tongue which is an extension of the upstream wall. The air intake holes for the pocket are arranged in the wall of the protuberance which is more or less perpendicular to the wall of the chamber which faces the air flow and have a sufficiently small diameter to make it possible to arrange them in at least three concentric rows, which are offset from one row to the next.

Cooling of combustion chamber walls using a film of air

The invention concerns a combustion chamber for turbojet engines. The comtion chamber is of the annular type and consists of two coaxial flame tubes opening into a common dilution and mixing zone. The inner tube is designed for low operating ratings of the engine, the outer tube for high ratings. Air is injected as far upstream as possible into the dilution zone, to enhance the homogenization of the gaseous flow issuing from the two tubes prior to their passage into the turbine and to assure the optimum radial distribution of temperatures. The combustion chamber according to the invention finds application in a particularly advantageous manner in turbojet engines used in aircraft propulsion because of the reduced emission of pollutants it affords.

Antipollution combustion chamber

A fuel injection nozzle for a turbo engine is mounted at the base of a costion chamber, either tubular or annular, by means of an intermediate inwardly flared member having openings through its sides for admitting cooling air. An annular flange structure around the intermediate member supports the same and defines an annular space with the base of the chamber for a layer of cooling air admitted to that space through holes in the base. Several forms are disclosed in each of which the various holes and air inlets are dimensioned and positioned to provide a stoichiometric mixture of air and fuel under substantially all conditions of operation.

Combustion chamber for turbo engines

Combustion chamber comprising a first combustion zone opening into a wider second combustion zone into which the combustion air is admitted through a practically frontal air entry comprising obstacles which impart a coefficient of blocking of about 60% to 80% to it. The first zone is supplied with air and fuel in substantially stoichiometric proportion in idling. At the other running rates, a supplementary fuel flow is injected into the chamber and burns in the second zone in the wake of the obstacles.

Growing community of inventors

Melun, France

Philippe M Gastebois