Date of Graduation

1996

Document Type

Dissertation/Thesis

Abstract

Vibration and Air-Resonance stability characteristics of helicopters in forward flight are studied using rotor-body dynamic interaction. The rotor blade and fuselage are considered as flexible structures. The blade undergoes flap, lag, and axial motions and twist about the elastic axis. The air loads are based on quasi-steady aerodynamics and linear inflow model. The governing equations are generated by expanding the generalized forces in the Hamiltonian as a first-order Taylor series. The zeroeth and n/rev. vibratory loads are predicted directly in the fixed frame. This is expected to be more accurate than the force or modal summation methods. The blade controls and fuselage orientation angles are determined such that the average equilibrium of the helicopter is satisfied. An iterative numerical scheme for vibration analysis is devised by shifting the rotor inertia forces (from {dollar}\\delta{dollar}T) to fuselage external forces (to {dollar}\\delta{dollar}W). For the stability analysis, selected modal bases of the rotor blade and fuselage are combined using multiblade coordinate transformation to generate the helicopter model. The stability characteristics of this system are studied using Floquet theory. The stability results obtained from this study are validated with existing experimental results for a ground-resonance problem. The blade response and the hub loads are compared with results from a linearized theory. The comparison indicates that the linearized theory cannot properly capture the higher harmonic content in the response. Extensive numerical results are obtained at various forward speeds for the vibration analysis by considering the fuselage rigid as well as elastic. The results indicate that the rigid fuselage model is insufficient to properly account for coupling effects. The fuselage coupling changes the 3rd, 4th, and 5th harmonic blade responses by as much as 80%. The influence on vibratory hub loads are as high as 60% and on the hub moments about 5%. The damping levels in the blade lag cyclic modes in the fixed frame changes considerably (40%-80%) with the inclusion of fuselage elastic modes.

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