Damping of Vibration and Noise

4 Damping of Vibration and Noise Fixed-Wing Aircraft 

Integration of SM A wires into the composite skin of a fin of 0.5 m in height has been realised within the EU-funded project ADAPT [70]. SMA wires of 150 pm thickness were integrated into the glass fibre reinforced composite skin which could then be actuated. An initial simple test showed that tip deflection amplitudes could be reduced by around a half in a vibration test once the SMA wires had been heated up to an austenitic condition (Fig. 8.18). 

Adaptronics for rotorcraft was very much driven by concepts in the early 1990s looking at rotor blade twist to improve aerodynamic performance as well as reduction of the negative drawbacks of lead lag damping. Individual blade control has become key in rotorcraft technology and it became quickly 

Validation of such flap actuation solutions have been performed in wind tunnel tests on a one-seventh downscaled Bell-412 Mach-scaled rotor hub [87]. It has been shown that trailing edge deflections of 4° to 5° can be achieved at up to 1800 rpm which allowed suppression of vibratory bending moments imder an open loop control condition. Even some preliminary closed-loop tests using a neural network controller were performed which however required simultaneous actuation of all four blades. In [88] an induced-shear piezoelectric actuator has been described to actuate trailing edge flaps 

Piezoelectric actuators have also been used in a similar collaboration between EADS Corp. Research and Eurocopter both in Germany and France to develop smart struts which serve to decouple the helicopter cabin from neighbouring gear boxes [91] of which the principle is shown in Fig. 8.20. Control mechanisms and ways on how to simulate waves transmitted through these struts have been described in [92]. 

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