Abstract: The gastight sealing device associated with a functional clearance existing between two sectors of a turbomachine includes an annular plate wound up on itself over at least one turn, covering functional clearance.

Abstract: A sealing device is disclosed comprising a least one sealing leaf maintained axially in position by a retaining bracket which comprises a first end bearing on the sealing leaf and a bent-back second end engaged in a groove in a structural element. The sealing leaf is preferably retained radially in position by lugs which are bent so as to engage below lugs provided on the retaining bracket. When the sealing device comprises at least two adjacent sealing leaves placed edge to edge, the device also comprises a cover sheet so disposed between the sealing leaves and the retaining bracket as to cover the joint between the two adjacent leaves.

Abstract: The invention provides a turbomachine including a device for supplying pressurized gas, the gas being bled from the compressor of the turbomachine via a bleed orifice and cooled in a cooling cavity disposed within the compressor rotor upstream of the bleed orifice. This arrangement has the effect of cooling the bled gas by heat transfer through the outer wall of the compressor rotor to the gas stream which is being compressed by the compressor upstream of the bleed orifice, and has the result of reducing the drop in the performance of the compressor as a result of the gas being bled off.

Abstract: A method of reducing the gap between a liner and a turbine distributor of a turbojet engine. The method allows for the improvement of the seal on the inside of the distributor of the high pressure turbine at the lower platform. The method includes the steps of: applying a protective deposit to the lower part of the external surface of the liner, before carrying out a filling by diffusion brazing between the liner and the distributor. The protective deposit, preferably a anti-adherent ceramic, makes it possible for the braze to avoid fixing the liner onto the distributor at the lower platform level. The method also has applications to turbojet engines.

Abstract: A cooling system is disclosed for cooling the platforms of turbine blades fixed to the periphery of a turbine rotor, each turbine blade having an airfoil portion, a platform with opposite longitudinal edges, a root portion affixed to the turbine rotor, and a shank portion connecting the root portion to the platform. Adjacent turbine blades are located such that a longitudinal edge of one platform is adjacent to a longitudinal edge of an adjacent platform. A cavity is formed bounded by a portion of the periphery of the turbine rotor, the shank portions of the adjacent turbine blades, and the platforms of the adjacent turbine blades. A cooling air passage supplies cooling air to the cavity. One or more cooling channels are formed in a side of the platform facing the turbine rotor, each cooling channel having a cooling air collector at an end located adjacent to the air supply passage and communicating with the cavity.

Abstract: This seal (6) running between gaps (6) between two stator sectors of a gas urbine engine is made up of several thin flexible slabs, capable of sliding one over the other, which gives a better seal.

Abstract: Device for cooling the external part of a turbine disk (2) by means of an external fresh air current (circuit II) without mixing with the warmer ventilation from the center of the disk (circuit I). A part (30) through which ducts (31, 32) pass that intersect without interruption, prevents mixing of substantially perpendicular streams in front of a part (22) of the disk (2) and a mixture which reduces the cooling efficiency.