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

Snecma

Snecma Moteurs

Robert Kervistin

Claude C Hallinger

Alain M Lardellier

Georges Mazeaud

Jose Roland Rodrigues

Gilles Louis Rollin

José Roland Rodrigues

Francois E Crozet

Philippe J Lefebvre

Marion De Giacomoni

Gabriel Binetruy

Alain F Teissier

Jean-Marie Desindes

The invention relates to a method and a device for protecting fuel injectors in a turboshaft engine. The protection is carried out by injecting compressed air directly into the injectors. The air thus directly purges the injectors and evacuates the fuel present in them into the combustion chamber.

Method and system for protecting gas turbine fuel injectors

A system for decoupling a fan of a turbojet, the turbojet including a moving fan shaft mounted on bearings each connected to a stationary structure of the turbojet by bearing supports, the bearing supports being fixed to the stationary structure of the turbojet via mechanical links serving, during normal operation of the turbojet, to transmit the forces generated by rotation of the fan, the mechanical links include explosive charges enabling the links to be broken, the explosive charges being controlled by a full authority digital engine control computer on the basis of means for measuring the mechanical stresses in the turbojet and of a computer model simulating the static, dynamic, and thermodynamic behavior of the turbojet.

System for decoupling a fan from a turbojet by means of an explosive charge

A cooling system for an axial flow compressor of a gas turbine engine is disclosed having a device to tap a portion of air from the primary air flow flowing through the axial compressor and a control device associated with the tap to control the amount of air tapped from the primary air flow. The system also includes a cooling device to cool the portion of air tapped from the primary air flow, which cooled air is directed into a cooling air chamber. The cooling air chamber is located between the last rotor stage of the compressor and a cone attached to the last stage and extending downstream from the rotor wheel. Nozzles associated with the cooling air chamber direct cooling air from the chamber onto the rotating rotor wheel such that a portion of the cooling air passes into the peripheral groove defined by the rotor wheel between the wheel and the blade root, while another portion of the cooling air from the nozzles is directed onto the surface of the cone in order to cool the blades and the cone.

Cooling system for a gas turbine engine compressor

The real-time adjustment system according to the invention utilizes an air flow regulating valve in an air conduit circuit activated by an output signal of an electronic computer. The computer determines a desired radial clearance at an operational time T of the gas turbine engine, which may be stored in the computer memory and may be based on a quantified engine model having the mechanical and thermal features of the rotor and stator elements which are to be controlled as function of engine thermodynamic parameters and the geometry of the elements, with the actual radial clearance computed in operation at the time T by the computer from data sensed in real-time and provided to the computer. The system also senses the maximum admissible stator temperature as well as the maximum temperatures and temperature gradients for the rotor. These limits are considered by the computer prior to emitting the output control signal to the valve. The output signal may also be modified by sensing the effect of the radial clearance by the tapping of the air flow from the compressor, by misalignment of the air between the rotor and stator elements and by the effect of the aerodynamic loses caused by the air tapped from the compressor on the specific consumption of the gas turbine engine.

System for adjusting radial clearance between rotor and stator elements

The invention concerns a system for controlling heat expansion and thermal stress in a gas turbine disk. The system includes perforated plates mounted close to and facing at least one surface of the disk, particularly opposite its rim and around its thick inner bored section. Conduits are provided for bleeding cold air and hot air from various stages of the compressor or from the turbine and for injecting cold air toward the rim of the disk and hot air toward its thick inner bored section through the corresponding perforated plates at the beginning of each acceleration, and, preferentially, during deceleration, for injecting hot air toward the said rim and cold air toward said inner bored section. The invention applies particularly to gas turbines used to propel aircraft.

System for controlling heat expansion and thermal stress in a gas

A blade root sockets provided in the discs of a turbine rotor have passages for cooling air which are arranged to be variably blocked by closure means carried by the upstream and downstream ends of the labyrinth carriers disposed between the discs. The labyrinth carriers are of the pin-like type and are able to expand and contract in response to the temperature of the gas flow passing through the turbine, moving the closure means to open or close the passages. The positions of the closure means adjacent the passages are determined by the relative dimensions and the thermal expansion coefficients of the discs and the labyrinth carriers, and may be chosen to provide full cooling air flow during engine acceleration, a reduced flow during stable operation, and a very restricted flow to the first disc only during deceleration.

Assembly for controlling the flow of cooling air in an engine turbine

An assembly to control automatically the clearance of a labyrinth seal of the turbo machine, including a stator member is provided which includes an annular chamber supplied with a hot air flow through orifices and a colder air flow supplied through apertures formed in the radially inner part of the stator member constituting the annular carrier of a wear seal member which forms with tips carried by the one part of the rotor the labyrinth seal, these apertures being disposed upstream of the seal. The apertures lie in the zone of corresponding parts of the carrier and of the rotor part which are so shaped as to create a annular convergent-divergent annular nozzle forming a throat. The clearance of the seal is maintained constant due to relative variations effected in the air flows and, in the sense of heating up in the case of a reduction of clearance and in the sense of cooling down in the case of an increase in clearance.

Automatic control device of a labyrinth seal clearance in a turbo-jet

Apparatus associated with a labyrinth seal of a turbo machine for automatically maintaining a desired clearance between lip members and a wear seal of the labyrinth seal over a range of operational states of the turbo machine. The wear seal comprises a first part made of honeycomb material and a second smooth part spaced axially from the first part. An annular carrier supports the wear seal, and a stator supports the carrier, thereby defining between the carrier and the stator an annular chamber having radially outer orifices for admission of an air supply to the chamber. An annular sheet member disposed within the annular chamber has a plurality of holes and is spaced at a small distance from the carrier. Whereas acceleration and deceleration of the turbo machine normally would tend to alter the spacing of the wear seal and the lip members, the structure recited above corrects for this effect, thereby maintaining a desired clearance between the wear seal and the lip members.

Automatic control device of a labyrinth seal clearance in a turbo jet

A clearance control device in an inter-shaft turbojet engine bearing journal placed between a high pressure trunnion and a low pressure trunnion and including inserted between an inner ring and an outer ring, the high pressure trunnion including a series of longitudinal grooves and a series of longitudinal holes which are supplied with oil from an intake and an annular collector in the low pressure trunnion, from which the oil moves by centrifugation through radial holes into an annular recess in the downstream tightening nut of the outer ring, which in turn communicates with the grooves through holes formed in the nut. Clearance in the bearing journal is controlled by cooling the high pressure trunnion at the maximum operating speeds of the turbine via circulating oil. By virtue of an appropriate circuit, the oil supplied by the collector lubricates the bearing and cools the roller bearing journal.

Means for controlling clearance in an intershaft bearing journal of a

A device for adjusting the clearance between moving blades and the ring of a turbine wherein the ring includes a cylindrical sleeve, a perforated partition, and a wall, which together form an enclosure. Boreholes receive tubular elements which fit together with coaxial boreholes provided in opposite walls of a distribution chamber. The chamber communicates with pipes which admit heat regulating air. Alternating tubular elements traverse the perforated wall and the chamber to evacuate exhaust gas through pipes.

Device for adjusting the clearance between moving turbine blades and the

Apparatus is disclosed which is intended to measure the temperature of a flow traversing a blade grid, for example the valve blades of a turbojet turbine. This apparatus comprises a temperature sounder lodged within a gas escape circuit provided in a blade. In-flight monitoring of the gas temperature in an aircraft turbojet is possible with the present invention.

Apparatus for measuring the temperature of a gas flow traversing a grid

A device for automatically regulating the clearance between the tips of the otor blades of a turbine, particularly of a gas turbine, and the adjacent wall of a turbine stator, comprising means for directing a gaseous flow against the said wall to regulate its temperature. This gaseous flow is obtained from a proportioner including at least two inlet passages connected to sources of gas at different temperatures and controlled by the thermal expansion of an obturator responsive to the temperature of fluid passing through the turbine, so as to reduce the temperature of the said gaseous flow as the obturator expands, in order to cool the stator in stable conditions and to heat it in transitory conditions.

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Le Mee sur Seine, France

Robert Kervistin