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

Pierre A Glowacki

Gerard M Mandet

Michel R Ayache

Josette Benoist

Jean-Claude Asselin

Daniel J Martin

Alain P Daguet

Claude G Corbin

Joel Foucard

Herve Lassignardie

A casing structural arm of a turbo-engine is fixed between an external ferrule and a hub of the casing and is formed as a composite material. The arm has one leading edge, one trailing edge, one radially external edge and one radially internal edge. These four edges define the four corners of the arm and securing of the arm to the ferrule and said hub is ensured by two pairs of linking braces respectively secured to the four corners of the arm.

Structural arm of the casing of a turbo-engine

An aircraft having a turbojet engine disposed within its fuselage wherein towards the rear end thereof it is provided with a sealing assembly between the rear end of the fuselage and the outer flaps of the nozzle of the turbojet engine, the sealing assembly including a central carrier ring, preferably made of a carbon-epoxy material, secured to the turbojet engine by supporting legs, a hollow front seal at the front end of the carrier ring engaging the inner surface of the fuselage, and a hollow rear seal at the rear end of the carrier ring engaging each of the outer flaps, the carrier ring also engaging the inner surface of the fuselage at a position spaced rearwardly from the front seal.

Sealing assembly between an aircraft fuselage and the outer nozzle flaps

An axial flow turbomachine rotor comprises a disc having a plurality of annular flanges extending outwardly from its periphery, a plurality of blades secured to the disc and extending radially outwardly from the periphery thereof, each of the blades having its root notched to define heels which are received in the grooves formed between the annular flanges of the disc, and a plurality of separate platforms which are secured to the disc and are disposed between the blades to define the inner boundary of the fluid flow path through the rotor, the blades each being fixed to the disc by two parallel pins which are equidistance from the axis of the disc and are received in corresponding bores passing through the annular flanges and the heels of the blade, the pins also serving to secure the platforms to the disc.

Axial flow turbomachine rotor

A blow-off device for a gas turbine bypass engine having a primary air flow path separated from an outer secondary air flow path by a partition includes a plurality of discharge passages passing through the partition from the primary air flow path to the secondary air flow path, each of the discharge passages having a first hinged flap in the outer wall of the primary air flow path, a second hinged flap in the inner wall of the secondary air flow path, and a connecting rod mechanically interlinking the first and second flaps, the connecting rod being hinged at one of its end to the first flap and at its other end to the second flap.

Blow-off device for a bypass gas turbine engine

An aircraft turbo-engine having an intake grille formed by radial struts disposed between inner and outer rings, each of said radial struts being in two parts, an upstream, fixed structural first part, and a downstream second part pivoted on said first part and forming an adjustable flap which constitutes an intake guide vane, all of said struts having the same cross-section as each other and at least some of the struts incorporating a pipe for the passage of oil, those oil pipe carrying struts each having its structural first part divided by a radial portion into an upstream chamber for the passage of hot air and a downstream chamber which is open along the length of its trailing edge and which receives the oil pipe.

Turbo-engine air intake grill

A casing structure for an annular combustion chamber of a turbojet engine is disclosed in which air is tapped from an upstream end portion of the casing, while at the same time preventing contamination of such tapped air from the injected fuel. The casing has annular inner and outer walls disposed on either side of the combustion chamber. The casing has an upstream end which is attached to the downstream end of the exhaust gas turbine. The outer annular casing wall defines circumferentially spaced apart fuel injector orifices and circumferentially spaced apart air tapping orifices which are circumferentially alternating with the fuel injector orifices. A partition wall extending between an air diffuser and the inner surface of the outer annular wall divides the radial space separating the diffuser and the outer annular wall into first cavities which communicate with the air tapping orifices and second cavities which communicate with the interior space between the inner and outer annular walls. The partition serves to completely separate the cavities preventing contamination of the tapped air by the injected fuel.

Casing for a turbojet engine combustion chamber

A nose cowl or spinner is provided for a fixed, non-rotatable central hub of a turbojet engine having assembly means that facilitate the attachment and detachment of the nose cowl to the central hub, and provides a readily visible indication of whether or not the nose cowl has been properly assembled to the hub. The nose cowl has a smooth outer surface to provide undisturbed air flow into the turbojet engine compressor and defines means to circulate heated air to prevent icing of the external surface. An inner cowl member is attached to a bearing cover of the central hub in a non-rotatable manner, an outer cowl member is axially locked to the central hub so as to be concentrally arranged about the inner cowl member, and a locking device is used to lock the inner and outer cowl members together to prevent inadvertent rotation of the outer cowl member.

Nose cowl for a turbojet engine shaft

The present invention concerns the cooling of turbine rotors, especially those of aircraft turboreactors, and discloses a turbine disc equipped with two sets of channels bored respectively close to each of the sides of the disc and in conformity with its profile, in which flows the cooling air of the turbine blades in order to superficially cool the disc.

Turbine rotor cooling

A portion of the stator of a gas turbine known as turbine shroud, which is ituated radially opposite the mobile vanes of a turbine stage includes an 'abradable' member supported by a first shell braced by a second shell and connected with turbine casing 2 by a thin, i.e. flexible, annular wall member embedded in a casing. Cooling air inflates the second shell so as to brace the exterior shell and circulates between the two shells so as to cool them and also cools the 'abradable' member by impingement. The present invention is applicable to aviation turbojets.

Temperature control device for gas turbines

A labyrinth seal mounted between a shaft and the casing of a gas turbine combustion chamber, wherein the support of an abradable sealing surface in the seal which operates in conjunction with rapidly rotating fins is cooled by an upstream airflow parallel to the envelope surface of the extremities of such fins and opening into an upstream space. The cooling air is taken in from the casing by means of a conduit which is connected with an overflow valve tied to the gas turbine load. The invention is applicable to aviation turbojets.

Labyrinth seals

The rim of a turbine disc is provided with a circular collar upon which is ngaged the heel of an annular flange having a U-shaped cross section with unequal sectors, i.e., a principal piece forming a circlip and a smaller piece forming a key which completes the flange following the installation of the circlip; the key is retained in its position by cuts of the ends of said key.

Rotor flanges of turbine engines

A turbine housing for a jet engine is provided with a ring of an abradable orous material. A supporting and mounting device consisting of a cylindricoconical monoblock having a conical internal jacket provides for the convection cooling of the external surface and a cylindrical jacket for cooling air is held over the abradable ring by means of a perforated ferrule integral with the rest of the device. The abradable ring carries transverse partitions to prevent axial leakage.

Porous abradable shroud with transverse partitions

A device to secure and insure the tightness of vanes of a turbine is an alar plate which carries on the side toward its external edge a plurality of studs and toward its internal edge a plurality of hangers. The disk of the rotor has on its lower edge a series of teeth or indentations, the top of the teeth having a configuration to cooperate with the inner surface of the hangers. The vanes have on their rims a plurality of grooves. Two adjacent vanes form a groove into which the studs enter when the hangers cooperate with the teeth. The studs are borne by tongues cut radially into the annular plate.

Growing community of inventors

Fontaine le Port, France

Pierre A Glowacki