Combustion chamber for a turbomachine
View Patent ↗Combustion chamber for a turbomachine, comprising a chamber endwall equipped with fuel injection means and axisymmetric walls connecting the chamber endwall to internal and external flanges for fastening to internal and external casings, these flanges including annular portions pierced by orifices of triangular or approximately triangular shape for the flow of air for cooling downstream components.
1. A combustion chamber for a turbomachine, such as an aircraft turbojet or turboprop, comprising:
a chamber endwall equipped with fuel injection means and axisymmetric walls connecting the chamber endwall to internal and external flanges for fastening to internal and external casings,
said flanges having annular portions pierced by orifices for the flow of air for cooling downstream components,
wherein the orifices of at least one of the flanges are of triangular or approximately triangular shape, and
the triangular orifices of one flange comprise two rows of triangular orifices, in which the orifices each have a base located on a circumference centered on the axis of the chamber, the orifices being reversed from a first row to a second row and the first row is partially imbricated in the second row, wherein the bases of the orifices of the first row are axially offset from the apexes of the orifices of the second row,
wherein the degree of partial imbrication of the two rows is determined by a ratio H 2 /H 1 , where H 1 is a distance between the base of a triangular orifice and its opposite apex and H 2 is the radial distance between the apexes of two successive orifices, opposite their bases, this ratio H 2 /H 1 being equal to or greater than 0.6.
2. The combustion chamber as claimed in claim 1 , wherein the triangular orifices are formed in the internal flange.
3. The combustion chamber as claimed in claim 1 , wherein the triangular orifices are formed in the external flange.
4. The combustion chamber as claimed in claim 2 , wherein triangular orifices are formed in the external flange.
5. The combustion chamber as claimed in claim 1 , wherein the triangular orifices each have a rounded apex.
6. The combustion chamber as claimed in claim 1 , wherein the triangular orifices are isosceles triangles or equilateral triangles.
7. The combustion chamber as claimed in claim 1 , wherein a circumferential distance c between the bases of the triangular orifices of any one row is equal to or greater than 1.5 times a width d of the strip of material that separates two successive orifices.
8. The combustion chamber as claimed in claim 5 , wherein a length L of a straight portion of the strip of material between the rounded apexes of two successive orifices is equal to or less than three times a width d of this strip.
9. The combustion chamber as claimed in claim 1 , wherein an angle α at the apex of the triangular orifices, opposite their bases, is equal to or greater than 20°.
10. A turbomachine, such as an aircraft turbojet or turboprop, which includes a combustion chamber as claimed in claim 1 .
11. The combustion chamber as claimed in claim 1 , wherein the shape of the triangular orifice consists of three flat edges connected by three arcs.
12. The combustion chamber as claimed in claim 1 , wherein the triangular orifices are identical.
13. The combustion chamber as claimed in claim 1 , wherein the ratio H 2 /H 1 is lower than 1.
14. The combustion chamber as claimed in claim 1 , wherein the ratio H 2 /H 1 is equal to about 0.6.
15. The combustion chamber as claimed in claim 1 , wherein the triangular orifices are formed in a frustoconical or radial annular portion of said flange, and
the bases of the orifices are separated from the internal and external edges of said annular portion by strips of material having a circumferential orientation,
wherein a radial width of the strip of material between the base of a triangular orifice and the external edge of the annular flange portion is equal to or greater than a width of the strip of material separating successive orifices, and
a radial width of the strip of material between the base of a triangular orifice and the internal edge of the annular flange portion is equal to or less than a radial distance between the apexes of two successive orifices.
16. A combustion chamber for a turbomachine, such as an aircraft turbojet or turboprop, comprising:
a chamber endwall equipped with fuel injection means and axisymmetric walls connecting the chamber endwall to internal and external flanges for fastening to internal and external casings,
said flanges having annular portions pierced by orifices for the flow of air for cooling downstream components,
wherein the orifices of at least one of the flanges are of triangular or approximately triangular shape, and
the triangular orifices of one flange comprise two rows of triangular orifices, in which the orifices each have a base located on a circumference centered on the axis of the chamber,
the orifices being reversed from a first row to a second row and the first row is partially imbricated in the second row, wherein the bases of the orifices of the first row are axially offset from the apexes of the orifices of the second row, wherein the triangular orifices are formed in a frustoconical or radial annular portion of said flange, and
the bases of the orifices are separated from the internal and external edges of said annular portion by strips of material having a circumferential orientation, wherein a radial width of the strip of material between the base of a triangular orifice and the external edge of the annular flange portion is equal to or greater than a width of the strip of material separating successive orifices, and
a radial width of the strip of material between the base of a triangular orifice and the internal edge of the annular flange portion is equal to or less than a radial distance between the apexes of two successive orifices.