Abstract
Pitch or roll damping of helicopter rotor has been experimentally studied by using model rotors in rocking motion. The rotors have articulated blades with spring constrained hinges and different combinations of Lock number, flapping hinge offset, and hinge constrained stiffness. By considering nonuniform induced velocity distribution a theoretical estimation based on the momentum and blade element theory has shown good coincidence with the experimental results. In contrast with analyses based on the vortex theory the present theory is very simple and does not require complex calculations so that the analytic evaluation and the quick estimation of the dynamic stability derivatives of rotor will be possible. The blade flapping behavior during sinusoidal rocking motion has also been analytically and experimentally analyzed and the mechanism of generation of direct damping and cross coupling moments have been clearly explained.
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