Helicopter vibration reduction using fixed-system auxiliary moments

Reductions in helicopter rotor hub vibratory loads that can be obtained through the introduction of steady auxiliary pitching and rolling moments in the fixed system are examined. The auxiliary moments allow changes in vehicle attitude and rotor pitch control inputs, such that vibrations can be minimized, while vehicle equilibrium is still satisfied. Because auxiliary moments can be generated by variation of the pitch of the right and left halves of the horizontal tail, the weight penalty, complexity, and performance losses associated with auxiliary lift and thrust (conventional compounding) would be alleviated. A comprehensive rotorcraft aeroelastic analysis was modified to include the fixed system auxiliary moments in the trim equations, and vibration reductions were examined for two different cases: 1) a light BO-105 type helicopter with a four-bladed hingeless rotor at high flight speeds (advance ratio of 0.35) and 2) a medium-weight UH-60 type helicopter with a four-bladed articulated rotor at a more moderate flight speed (advance ratio of 0.25). For the BO-105 type aircraft, a nosedown auxiliary pitching moment and a roll-left auxiliary rolling moment could reduce the in-plane vibratory hub forces by 20-30% when used individually and by 90% when used in combination. For the UH-60 type aircraft, a nosedown auxiliary pitching moment and a roll-right auxiliary rolling moment produced reductions in the range of 15-65% in all components of vibratory hub loads. For both aircraft, the nosedown auxiliary pitching moment increased the nosedown pitch attitude of the aircraft and reduced the rotor longitudinal cyclic flapping and the longitudinal cyclic pitch input. For the BO-105, the roll-left auxiliary moment increased the leftward shaft tilt and no longer required the tip path plane to tilt left relative to the hub plane. This decreased the lateral cyclic pitch input. However, for the UH-60, the roll-right auxiliary moment actually reversed the roll attitude of the aircraft (shaft now tilting right). This required a large leftward tilt of the tip path plane relative to the hub plane and increased the lateral cyclic pitch requirement. When implementation of the auxiliary moments through collective and differential pitch inputs to the horizontal tail was examined, it was observed that the expected reductions in vibration were indeed achieved, but that the auxiliary roll moment generated was smaller than expected due to one-half of the horizontal tail being stalled.