Patent ID: 7568350
Filing Date: 2009-08-04
Classification: F23R

Abstract:
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