The present invention concerns the field of turbomachines, in particular the cooling of the ring sectors of a turbine of a turbomachine.
An aeronautical turbomachine conventionally includes a compressor, a combustion chamber and a turbine. The job of the turbine is to rotate the compressor by extracting some of the pressure energy from the hot gases leaving the combustion chamber and converting this energy into mechanical energy.
The turbine, situated downstream of the combustion chamber, is the component of the turbomachine that operates under the harshest conditions. It is, in particular, subjected to high thermal and mechanical stresses generated by the hot gases leaving the chamber.
A turbine conventionally comprises at least one nozzle guide vane assembly, consisting of a cascade of vanes which are fixed with respect to the casing of the turbomachine, and at least one rotor, comprising a set of blades which can be made to rotate.
A rotor is surrounded by a fixed ring secured to the casing. This ring has the function in particular of reconstituting, by way of its inner surface, the outer limit of the aerodynamic flowpath. In most cases, this ring can be made up of a plurality of sectors. Owing to its permanent exposure to the hot gases, a turbine ring has a reduced service life. It is therefore necessary to cool it so that it is able to withstand high temperatures effectively.
A known cooling solution consists in providing the ring with multiperforated plates surrounding said ring. Such plates are described in documents EP0893577, EP1134360, EP0516322, EP1225309 and EP1533478. A multiperforated plate generally has the shape of a bath in the sense that it comprises a bottom and side walls. Orifices are distributed essentially over the bottom of the multiperforated plate in rows. Cooling air passes through the multiperforated plate via the orifices and impacts the wall of the ring so as to cool it. The bottom of the multiperforated plate is spaced from the ring via a gap. All these prior art documents have a constant gap. Documents GB1330892 and US2003/0131980 disclose gaps which are progressive in an axial direction but which nevertheless remain unsatisfactory in terms of cooling.
The temperature is not uniform in all points of the ring. The effect of this is to create deformations therein. When the ring is made up of a plurality of sectors, each sector takes the form of an arc of a circle corresponding to an angular sector of the ring. A multiperforated plate corresponds to each ring sector.
A first type of deformation can be manifested circumferentially by a decambering of the sectors, that is to say a loss in the concentricity of the sectors, and hence of the ring.
A second type of deformation can be manifested axially by an axial canting of the sectors, that is to say an excessive offset in the distance between the upstream part or the downstream part of the sectors and the axis (X) of the turbomachine on which they are fastened.
These two types of deformations can lead to mechanical wear of the ring as a result of frictional engagement by the rotor blades situated opposite. Such wear creates an irreversible clearance between the tips of the blades and the turbine ring. This clearance is detrimental to the efficiency of the turbomachine.
To attenuate these deformations, it is necessary for the temperature of the ring sectors to be made uniform. To achieve this, it is known practice to vary different parameters on a multiperforated plate, such as the diameter of the orifices, the number of orifices per row or the pitch between each row.
USPTO Backgrounds
USPTO Backgrounds