Turbo engine

Abstract

The invention relates to a turbo engine, in particular a gas turbine aircraft engine, having a plurality of compressor components, at least one combustion chamber and a plurality of turbine components, wherein at least one support rib (36) is positioned in a flow channel (35) between two turbine components (32, 34) connected one behind the other, wherein the support rib (36) or each support rib (36) has a suction side, a pressure side, a front edge (37) and a rear edge (38), wherein the support rib or each support rib diverts a flow that flows through flow channel (35), and wherein a preferably cylindrical guide element runs in an inside space of the support rib or of each support rib. According to the invention, the turbo engine comprises at least the following features: a) the suction side (39) of the support rib (36) or of each support rib (36) is contoured in such a way that, viewed in the radial direction, a thickness of the respective support rib (36) is enlarged or increases in the direction onto a radially inner boundary wall (42) of flow channel (35) as well as in the direction onto a radially outer boundary wall (43) of flow channel (35); b) the pressure side (40) of support rib (36) or of each support rib (36) is contoured in such a way that, viewed in the radial direction, the thickness of the respective support rib (36) is enlarged or increases at least directly in the region of the radially inner boundary wall (42) of flow channel (35) as well as directly in the region of the radially outer boundary wall (43) of flow channel (35); c) the front edge (37) and the rear edge (38) of support rib (36) or of each support rib (36) are inclined in the meridian direction.

Dependencies are inferred from the claim text.