Limit cycle torsional motion of helicopter blades due to stall

Abstract

The amplitude of one degree of freedom limit cycle torsional motion of highly loaded helicopter rotor blades in the static thrust condition is shown experimentally to be dependent on the blade mean pitch angle and the reduced frequency of the motion. The origin of limit cycle motion is indicated by consideration of the experimental chordwise pressure variation during a limit cycle. The relationship between limit cycle motion and the damping in pitch in the presence of stall is determined. Finally, the implications of the results in the forward flight condition are discussed and illustrated by reference to flight test records.