Abstract: A retractable damping device for the blades of a turbojet, located between two consecutive blades against the lower face of the platforms of said blades. The damping element, mounted to rotate by one of its ends in the middle part of the space formed between the blades and a rotor disk positioned along an axis perpendicular to a radial axis of the disk, exhibits an upper part which constitutes a projecting dihedral that takes the form of the protruding dihedral consisting of the platforms of the two consecutive blades under the action of centrifugal force.

Abstract: The device has two ring segments (13) held in a circumferential score (11) provided in one edge of the upper surface (8) of the rotor disk. These ring segments consist of a portion of a cylindrical band (17) on whose edges are provided two collars (18, 19) of unequal thickness. The thicker collar (18) is placed between the upper wall of the transverse score (7) made in the projection (6) and the upper wall (8) of the rotor (1). The cylindrical band (17) works in conjunction with the bottom of the circumferential score (11). The thinnest collar (19) leans against the upper wall of the groove. The two segments placed end to end leave between them a space corresponding to the width of one axial notch. By displacing the ring segments in the score, the space is brought opposite the blade to be changed. The locking ring has two pins (21) which engage between the ends of the ring segments and block them in position. Other types of applications are described.