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

501 Societe Nationale D'etude Et De Construction De Meteur

Gerard Y Barbier

Michel A Desaulty

Xavier M Bardey

Rodolphe Martinez

Serge M Meunier

Gerard J Bayle-Laboure

Jerome Perigne

Bruno R Masse

Denis J Sandelis

Guy D'Agostino

Jean-Paul D Alary

Francois J Mirville

Didier Merville

Pierre A Bouillot

Frederic Bruno Beule

Francois Duchene

Jacques G Havard

Ralph M Stelter

Serge R Muccioli

Paul Mackovic

Corine G Delahaye

Gerald J Leboure

Jean M Latour

Michael A Desaulty

Gerard J Laboure

Xavier, M Bardey

Pascal M Truvillot

A separator-diffuser for a gas turbine engine is disclosed wherein the diffuser includes axial walls which function as a guide vane stage for the oxidizer compressor. By incorporating the last, or downstream, guide vane stage into the diffuser structure, the overall axial length and bulk of the gas turbine engine may be significantly reduced. The separator-diffuser has a wall defining a generally annular oxidizer flow passage extending around the central axis of the gas turbine engine, the passage having an oxidizer inlet located downstream of the last rotor stage of the oxidizer compressor so as to enable the oxidizer to pass from the oxidizer compressor into the oxidizer flow passage. A generally annular separator member is located in the oxidizer flow passage downstream of the oxidizer inlet so as to divide the oxidizer flow passage into at least two separate oxidizer flow paths, each flow path having a separate oxidizer outlet. The oxidizer outlets are positioned so as to direct the oxidizer towards one of the annular array of fuel injection heads. A plurality of circumferentially spaced apart, axially extending primary arms connect the separator member to the diffuser walls so as to support the separator within the oxidizer passageway and to serve as a guide vane stage for the oxidizer emanating from the oxidizer compressor.

Gas turbine engine having an integral guide vane and separator diffuser

A generally annular combustion chamber for a gas turbine engine is disclosed in which a plurality of generally cylindrical walls extend forwardly from an upstream end wall of the combustion chamber such that each cylindrical wall defines a cavity which is in communication with the interior of the combustion chamber. A first fuel injection head is located in each of the cylindrical walls so as to inject fuel into the cavity which is mixed with air and passes into the combustion chamber. The first fuel injection heads are located at a first axial position with respect to a longitudinal axis passing through the combustion chamber. A plurality of second fuel injection heads are located adjacent to the upstream end wall of the combustion chamber so as to spray fuel directly into the combustion chamber. The second fuel injection heads are located axially downstream of the axial positions of the first fuel injection heads.

Combustion chamber with axially displaced fuel injectors

A low pollution combustion chamber for a turbojet engine is disclosed in which first and second combustion chambers are oriented in a counter-flow direction and both combustion chambers communicate with a separate and distinct exhaust chamber. The first combustion chamber has a fuel injector, as well as a primary oxidizer intake, and the wall bounding the first combustion chamber defines a plurality of dilution oxidizer intake orifices. The wall also defines a first exhaust orifice located approximately midway between the upstream and downstream ends of the first combustion chamber. A second combustion chamber has a second fuel injector, as well as a second primary oxidizer intake, but the wall defining the second combustion chamber does not define any dilution oxidizer intake orifices. A second exhaust orifice also communicates with the exhaust chamber. The orientations of the first and second combustion chambers are such that the fuel/oxidizer mixtures traveling through the combustion chambers travel in a generally counterflow arrangement. A separate exhaust chamber is included in the combustion chamber assembly and communicates with the first and second combustion chambers such that the burned gases from these chambers enter the exhaust chamber. The exhaust chamber defines a third exhaust orifice through which the burned gases exit the combustion chamber assembly. The exhaust chamber is oriented such that the exhaust gases passing through this exhaust chamber travel in a direction generally parallel to the gases traveling through the second combustion chamber. A wall defining the combustion chamber defines a second plurality of dilution oxidizer intake orifices so that oxidizer may dilute the exhaust gases passing through this chamber.

Low pollution combustion chamber for a turbojet engine

A combustion chamber suitable for use in a gas turbine engine includes a first enclosure with a fuel injector for low-power operation and a primary-oxidizer intake, a second enclosure with a fuel injector for full-power operation and a primary-oxidizer intake, and a third enclosure from which to evacuate burnt gases and which communicates with the first and second enclosures. The wall of the first enclosure includes intake orifices for a dilution oxidizer, but the wall bounding the second enclosure lacks any orifices other than the primary-oxidizer intake orifices.

Lean premixture combustion-chamber comprising a counterflow enclosure to

A fuel injector is disclosed for injecting fuel into a combustion chamber a gas turbine engine which includes an injector nozzle having a central axis, first and second fuel injection paths in which one of the fuel injection paths supplies fuel to the injector nozzle while the other fuel injection path has an outlet displaced from the axis of the injection nozzle, a single fuel supply conduit which supplies fuel to both of the fuel injection paths, and a metering device which controls the flow of fuel from the single fuel supply conduit to the first and second fuel injection paths. The metering device may be a fuel metering valve assembly which has a piston defining a fuel metering orifice biased against a valve seat around an opening which communicates with the single fuel supply conduit. The piston is biased into engagement with the valve seat by a resilient bellows attached between the piston and a fixed structure.

Fuel injector for a gas turbine engine

An oxidizer intake flow control system is disclosed for a gas turbine engine combustion chamber. The chamber has an oxidizer intake assembly which extends through the wall of the chamber such that an inlet aperture of the oxidizer intake and an exhaust aperture of the oxidizer intake are located on opposite sides of the wall defining the combustion chamber. An oxidizer intake sleeve defines a central passage through which oxidizer may pass into the combustion chamber. The intake passage extends along a central axis and the intake assembly is attached to the wall of the combustion chamber such that it is rotatable about this central axis. The inlet aperture and/or the exhaust aperture is located in a plane extending non-perpendicularly to the central axis. The combustion chamber may have several oxidizer intake assemblies extending through the combustion chamber walls. The intake assemblies may be interconnected such that they may be simultaneously rotated with respect to the combustion chamber wall.

Gas turbine engine combustion chamber with oxidizer intake flow control

The present invention relates to a combustion chamber having a counter-flow arrangement in which a first chamber portion mixes the fuel with the primary oxidizer, which mixture is ignited and burned in the first chamber portion. The exhaust gases from the first chamber portion pass into a second chamber portion in which dilution oxidizer air is added to the burned gases. The orientations of the first chamber portion and the second chamber portion are such that the gases pass through the chamber portions in substantially opposite directions, in a counter-flow arrangement. The first chamber portion in which the combustion takes place, provides only primary oxidizer to the fuel mixture, there being no further communication between the interior of the first chamber portion and the oxidizer air supply. This prevents the leaning of the fuel/air mixture in order to reduce the emissions of nitrogen oxide. The mixing of secondary oxidizer air with the exhaust gases takes place solely in the second chamber portion.

Counter flow combustion chamber for a gas turbine engine

A combustion chamber wall structure is formed of a plurality of segments, each segment having an upstream portion and a downstream portion. The upstream portion is attached to an outer surface of an adjacent wall segment at its downstream portion in an overlapping fashion such that the overlapping portions form a cooling air chamber. The main cooling air holes are defined by the upstream portion so as to communicate with the cooling chamber, which has an exit to direct the cooling air along an inner surface of the downstream portion of the wall segment. A primary air hole is defined by the downstream portion of the segment and is located adjacent to the main cooling air holes. This location maximizes the distance between the primary air hole and an upstream portion of the same segment, which, in turn, maximizes the distance along the inner surface of the segment on which the cooling air may flow without being disrupted by air passing through the primary air hole. The guidance of the primary/dilution air streams can also be carried out by a portion of the downstream air cooling film.

Gas turbine combustion chamber wall structure for minimizing cooling

An improved swirler system for a turbojet engine combustion chamber is disclosed in which stationary first vanes and moveable second vanes are oriented at different angles with respect to a radius of the fuel injector system. The different angles allow the angle of the incoming air to be adjusted between full power and idle conditions so as to vary the conial shape of the atomized air-fuel mixture according to the operating conditions of the engine. Radial and axial inlet duct portions serve to attenuate the wakes of the air flow generated by the vanes prior to the air passing into the combustion chamber.

Air-fuel injection system for a turbojet engine

A combustion chamber for a turbojet engine is disclosed which utilizes double walled construction. Hot walls are attached to adjacent cold walls solely by an orifice member which defines an orifice to supply primary of dilution air to the combustion chamber. The attachment allows relative radial movement of the hot walls with respect to the cold walls so as to accommodate the thermal expansion and contraction of the hot walls. The sleeves forming the walls also cooperate so as to provide an internal cooling film on the interior surface of the hot walls. Convection cooling may also be utilized between the walls to prevent structural damage due to high temperatures.

Combustion chamber structure for a turbojet engine

A swirler for a gas turbine fuel injection apparatus is disclosed in which the intake air is directed both axially and radially to maximize the efficiency of the gas turbine during all modes of operation. Circumferentially alternating axial and radial passages are defined by the swirler member surrounding the fuel injection nozzle. A diaphragm control is provided to selectively control the amount of air passing through the axial and radial passages, depending upon the operational mode of the gas turbine.

Axial-centripetal swirler injection apparatus

A system is disclosed for controlling the flow of air through the air-fuel mixing device for a turbojet engine. The system simultaneously controls the amount of air passing through turbulence generating baffles and that passing into the combustion chamber depending upon the operating conditions of the turbojet engine.

Variable flow air-fuel mixing device for a turbojet engine

A fuel and air injection system for a turbojet engine is disclosed wherein a bowl-shaped member surrounding the fuel injection nozzle defines an impact cooling chamber divided into four sectors. A pair of diametrically opposite sectors have openings to permit air to pass through the cooling chamber into the combustion chamber, while opposite diametrically opposed sectors have openings of larger diameter which also allow communication between the cooling chamber and the combustion chamber. A diaphragm control system allows the air passing into the sectors having the larger diameter openings to be modulated depending upon the throttle setting of the turbojet engine.

Growing community of inventors

Morangis, France

Gerard Y Barbier