GB643474A - Improvements in or relating to damping devices for aircraft rotors - Google Patents

Abstract

643,474. Helicopter rotors. FIRESTONE TIRE & RUBBER CO. June 8, 1947, No. 17987. Convention date, Aug. 16, 1946. [Class 4] [Also in Group XXXII] In an aircraft rotor hub, Fig. 1, each rotor blade 5 is connected to a central axis structure 8 through a sleeve structure 6, fork 7 and the forked end 10 of an arm 9 carried by the structure 8. Each fork 10 carries a vertical pivot 11, which engages the fork 7 through a horizontal pivot to permit flapping motions of the blades. Arms 12, 12a are rigidly secured to each pivot 11 and a hydraulic damper 15 (see Group XXXII) is connected between the end of an arm 12 of one blade and the arm 12a of an adjacent blade to damp vibrations set up by inequalities in the angular spacing of the blades during starting and stopping of the rotor or during flight. The pitch of the blades 5 is controlled by a swash-plate 14 which is connected through arms 6a, 14a and a pushrod 13, to each blade sleeve 6. In another arrangement, Fig. 4, in which a hydraulic damper 45 (see Group XXXII) is pivotally connected between each blade 55 and a hub member 46, each blade 55 is connected through a vertical drag pivot 54 on a terminal fitting 52 of a fork 49, which latter is attached to the hub member 46 by flapping pivots 50, 51. The fitting 52 is attached to the fork 49, for rotational movement about an axis coincident with the axis of the blade 55 to permit changes in blade pitch, under the control of a hydraulic cylinder 56.