Active vibration control is a widely implemented method for the helicopter vibration control. Due to the significant progress in microelectronics, this technique outperforms the traditional passive control technique due to weight penalty and lack of adaptability for the changing flight conditions. In this thesis, an optimal controller is designed to attenuate the rotor blade vibration. The mathematical model of the triply coupled vibration of the rotating cantilever beam is used to develop the state-space model of an isolated rotor blade. The required natural frequencies are determined by the modified Galerkin method and only the principal aerodynamic forces acting on the structure are considered to obtain the elements of the input matrix. A linear quadratic regulator is designed to achieve the vibration reduction at the optimum level and the controller is tuned for the hovering and forward flight with different advance ratios.

Helicopters suffer from severe vibration levels compared to fixed-wing aircraft. The main source of vibration in a helicopter is the main rotor which operates in a highly unsteady aerodynamic environment. Active vibration control methods are effective in helicopter vibration suppression since they can adapt to various flight conditions and often involve low weight penalty. One such method is the actively controlled flap (ACF) approach. In the ACF approach, a trailing-edge flap (TEF) located in each rotor blade is deflected at higher harmonics of rotor frequency to reduce vibratory loads at the rotor hub. The ACF approach is attractive because of its simplicity in practical implementation, low actuation power and enhanced airworthiness, since the flap control is independent of the primary control system. Multiple-flaps are better suited to modify the aerodynamic loading over the rotor blade and hence offer more flexibility compared to a single flap. They also provide the advantage of redundancy over single-flap configuration. However, issues like the number, location and size of these individual flaps need to be addressed based on logic and a suitable performance criteria. Preliminary studies on a 4-bladed hingeless rotor using simple aerodynamic and wake models predict that multiple-flaps are capable of 70-75 percent reduction in hub vibration levels. Numerical studies confirm that multiple-flaps require significantly less control effort as compared to single-flap configuration for obtaining similar reductions in hub vibration levels. Detailed studies include more accurate aerodynamic and wake models for the rotor with TEF’s. A simple and efficient flap control algorithm is chosen from literature and modified for use in multiple-flap configuration to actuate every flap near complete authority. The flap algorithm is computationally efficient and performs creditably at both high and low forward speeds. This algorithm works reasonably well in the presence of zero-mean Gaussian noise in hub load data. It is also fairly insensitive to small changes in plant parameters, such as, blade mass and stiffness properties. The optimal locations of multiple TEF’s for maximum reduction in hub vibration are determined using Response Surface methodology. Piezoelectric stack actuators are the most promising candidates for actuation of full-scale TEF’s on helicopter rotors. A major limitation of piezoelectric actuators is their lack of accuracy due to nonlinearity and hysteresis. The hysteresis in the actuators is modeled using the classical Preisach model (CPM). Experimental data from literature is used to estimate the Preisach distribution function. The hub vibration in this case is reduced by about 81-86 percent from baseline conditions. The performance of the ACF mechanism can be further improved by using an accurate hysteresis compensation scheme. However, using a linear model for the piezoelectric actuator or an inaccurate compensation scheme can lead to deterioration in ACF performance. Finally, bench-top experiments are conducted on a commercially available piezostack actuator (APA500L from CEDRAT Technologies) to study its dynamic hysteresis characteristics. A rate-dependent dynamic hysteresis model based on CPM is used to model the actuator. The unknown coefficients in the model are identified using experiments and validated. Numerical simulations show the importance of modeling actuator hysteresis in helicopter vibration control using TEF’s. A final configuration of multiple flaps is then proposed by including the effects of actuator hysteresis and using the response surface approach to determine the optimal flap locations. It is found that dynamic hysteresis not only affects the vibration reduction levels but also the optimal location of the TEF's.

This research investigates on-blade partial span active plain trailing edge flaps (TEFs)with an aim to alleviate the helicopter vibrations. Among all the available smart materials, piezoelectric stack actuator(PEA)has shown its strong candidature for full scale rotor systems. Although, PEAs are quite robust in operation, however, they exhibit rate dependent hysteresis phenomenon and can generate only very small displacements. Dynamic hysteresis is a complex phenomenon which, if not modeled, can lead to drift in the vibration predictions. In this research, a comprehensive experimental analysis is performed on a commercially available piezostack actuator, APA-500L, which is well suited for full scale applications. Rate dependent hysteresis loops are obtained for helicopter operational frequencies. Nonlinear rate-dependent hysteresis loops are modeled using conic section approach and the results are validated with experimental data. Dynamic hysteresis exhibited by the PEA is further cascaded with the helicopter aeroelastic analysis and its effect on helicopter vibration predictions is investigated. PEAs generate high force but are limited by small translational motions. A linear to rotary motion amplification mechanism is required to actuate the TEF for vibration alleviation. A smart flap is designed and developed using computer-aided-design models. A rotor blade test section is fabricated and a lever-fulcrum mechanism (AM-1) is developed for a feasibility study. Smart flap actuation is demonstrated on the rotor blade test section. The conventional motion amplification devices contain several linkages, which are potential sites for structural failure. A novel pinned-pinned post-buckled beam linear-to-rotary motion amplifier (AM-2) is designed and developed to actuate the flaps. A new design of linear-to-linear amplification mechanism (LX-4) is developed and is employed in conjunction with AM-2 to increase the flap angles by an order of magnitude. An analytical model is developed using Mathieu-Hill type differential equations. Static and dynamic tests are conducted on a scaled flap model. Helicopter aeroelastic simulations show substantial reduction in hub loads using AM-2 mechanism. To further enhance the flap angles, an optimization study is performed and optimal beam dimensions are obtained. A new technique is also proposed to actively bias the flaps for both upward and downward motion. Critical flap design parameters, such as flap span, flap chord and flap location influences the flap power requirement and vibration objective function significantly. A comprehensive parametric investigation is performed to obtain the best design of TEFs at various advance ratios. Although, parametric study equips the designer with vital information about various critical system parameters, however, it is a computationally expensive exercise especially when used with large comprehensive helicopter aero elastic codes. A formal optimization procedure is employed to obtain the optimal flap design and location. Surrogate models are developed using design of experiments based on response surface methodology. Two new orthogonal arrays are proposed to construct the second order polynomial response surfaces. Pareto analysis is employed in conjunction with a newly developed computationally efficient evolutionary multi-objective bat algorithm. Optimal flap design and flap locations for dual trailing edge flaps are obtained for mutually conflicting objectives of minimum vibration levels and minimum power requirement to actuate the flaps.