Rotor disk for a blade ring

Abstract

A rotor disk for a blade ring of an aircraft engine, in particular for a turbine stage thereof, including an attachment section for attachment to a shaft for torque transfer, a main section that extends radially inwardly from a blade root receptacle in the radial direction, a connecting section that connects the attachment section and the main section, and a plurality of through openings, situated in the circumferential direction, at a radially inner end of the connecting section. A rotor disk whose stresses are reduced in the event of load is provided in that the connecting section has an undercut. An apex of the undercut is radially situated above the through opening in the range of 5%-50% of the axial length of a through opening, and/or is axially spaced apart from an axial end of the through opening at 5%-35% of the axial length of the through opening.

Dependencies are inferred from the claim text.