Abstract: The invention relates to an assembly for an aircraft turbine engine comprising an integral element (42) consisting of a disk (44) and a plurality of blades (48). Each blade has a connection zone (64) connecting it to the disk, comprising a first part (64a) arranged at the external flowpath delimitation surface (50) provided on the disk (44), and at least one second end part (64b) arranged at a recess (72) formed in the disk along the axial extension of the external surface (50), the average fillet radius defined by the second end part (64b) being larger than the fillet radius defined by said first part (64a). Furthermore, the assembly comprises a flowpath reconstruction part formed in the aerodynamic continuity of said external surface (50), so as to cover said recess (72).