Retractable damping device for blades of a turbojet
AssigneeSnecma
InventorsJean G. Bouiller, Jean-Claude L. Delonge, Marcel L. A. Rigo, Didier G. Zietek
Life of this patent
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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.
Claims4
A turbojet rotor assembly comprising:
- a rotor disk having a rotational axis;
- a plurality of blades mounted in a rim of said disk, each of said blades having a platform, platforms of adjacent ones of said blades defining a dihedral, and a space defined between each said dihedral and said desk rim;
- a damping element in each said space, each said damping element having an upper part forming a projecting dihedral whose shape corresponds to said dihedral defined by said platforms, each said damping element further comprising means for pivoting said upper part about an axis extending transverse to said rim axis and in a plane of said disk;
- and
- an elastic element for elastically biasing said damping element to pivot said upper part such that said projecting dihedral contacts and presses against said dihedral defined by said platforms.
Dependencies are inferred from the claim text.