i: ill 4 L s_ L' Control of Helicopter Rotorblade Aerodynamics (original) (raw)
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Piezoelectric servo-flap actuation system for helicopter rotor individual blade control
2000
Thesis (Ph.D.)--Massachusetts Institute of Technology, Dept. of Aeronautics and Astronautics, 2000.Includes bibliographical references (p. 177-186).A novel new actuator for helicopter rotor control, the X-Frame Actuator, was developed, demonstrating superior performance for applications requiring compact, fast acting, large stroke actuation. The detailed experimental characterization of this actuator is described, including bench-top output energy measurements and transverse shake test performance. A Mach scaled rotor blade utilizing the X-Frame actuator to power a trailing edge servo-flap near the tip was also designed, manufactured and tested. A description of the design and composite manufacturing of the rotor blade and servo-flap is presented. Preliminary bench tests of the active blade actuation system are also presented. The hover tests of the active blade provided transfer function identification of the performance of the actuator in producing flap deflections, and the respon...
Development of a Piezoelectric Servo-Flap Actuator for Helicopter Rotor Control
1994
An actuator using a piezoelectric bender to de ect a trailing edge servo-ap for use on a helicopter rotor blade was designed, built, and tested. This actuator is an improvement o ver one developed previously at MIT. The design utilizes a new exure mechanism to connect the piezoelectric bender to the control surface. The e ciency of the bender was improved by tapering its thickness properties with length. Also, implementation of a nonlinear circuit allowing the application of a greater range of actuator voltages increased the resultant strain levels. Experiments were carried out on the bench top to determine the frequency response of the actuator, as well as hinge moment and displacement capabilities. Flap de ections of 11.5 deg were demonstrated while operating under no load conditions
Piezoelectric Servo-Flap Actuator for Helicopter Rotor Control
1994
Master of Science in Aeronautics and Astronautics An actuator using a piezoelectric bender to de ect a trailing edge servo- ap for use on a helicopter rotor blade was designed, built, and tested. This actuator is an improvement over one developed previously at MIT. The design utilizes a new exure mechanism to connect the piezoelectric bender to the control surface. The eciency of the bender was improved by tapering its thickness properties with length. Also, implementation of a nonlinear circuit allowing the application of a greater range of actuator voltages increased the resultant strain levels. Experiments were carried out on the bench top to determine the frequency re-sponse of the actuator, as well as hinge moment and displacement capabilities. Flap de ections of 11.5 deg were demonstrated while operating under no load conditions at 10 Hz. Excessive creep at low frequencies precluded the measurement of achievable hinge moments, but extrapolation from de ection and voltage chara...
In this study, a characterization of hysteresis in a piezoceramic stack actuator similar to those employed in an actively controlled flap (ACF) system is performed to assess the effects of hysteresis on system performance. The effect of unmodeled actuation hysteresis may significantly reduce vibration and noise reduction capabilities. A hysteresis model based on the classical Preisach model has been developed from experimental data. The model displays good agreement with experimental data for input frequencies typical of those used for vibration and noise reduction in full-scale rotors. The hysteresis model has been incorporated into the Active Vibration and Noise Reduction (AVINOR) code, developed at the University of Michigan, so as to evaluate the effect of piezoceramic actuator hysteresis on vibration and noise reduction. The incorporation of hysteresis does not produce a significant performance degradation of the ACF system for vibration reduction at blade-vortex interaction (BVI) and dynamic stall conditions. For BVI noise reduction, hysteresis produces significant differences in flap deflection time histories, demonstrating the importance of hysteresis modeling. However, the overall noise reduction performance of the ACF system is not significantly affected. Nomenclature c Blade chord c c Control surface chord C W Helicopter weight coefficient D Matrix defined to be T T QT + R f (t) Static hysteresis actuator output f − Negative saturation output value F HX4 , F HY 4 , F HZ4 Non-dimensional 4/rev hub shears J(z k , u k) Objective function k Control update index L b Blade length L c Control surface length M HX4 , M HY 4 , M HZ4 Non-dimensional 4/rev hub moments M Time-ordered sequence of maximum input values m Time-ordered sequence of minimum input values M k Maximum input values m k Minimum input values
Control of helicopter rotorblade aerodynamics
1991
The results of a feasibility study of a method for controlling the aerodynamics of helicopter rotorblades using stacks of piezoelectric ceramic plates are presented. A resonant mechanism is proposed for the amplification of the displacements produced by the stack. This motion is then converted into linear displacement for the actuation of the servoflap of the blades. A design which emulates the actuation of the servoflap on the Kaman SH-2F is used to demonstrate the fact that such a system can be designed to produce the necessary forces and velocities needed to control the aerodynamics of the rotorblades of such a helicopter. Estimates of the electrical power requirements are also presented. A Small Business Innovation Research (SBIR) Phase 2 Program is suggested, whereby a bench-top prototype of the device can be built and tested. A collaborative effort between AEDAR Corporation and Kaman Aerospace Corporation is anticipated for future effort on this project.
Helicopter Performance and Vibration Enhancement by Twist-Actuated Blades
44th AIAA/ASME/ASCE/AHS/ASC Structures, Structural Dynamics, and Materials Conference, 2003
This paper investigates potential improvements that can be accomplished by integral twist-actuated rotor blades regarding helicopter performance and its vibration. The twist deformation is obtained using anisotropic piezocomposite actuators embedded in the composite blade construction. A four-bladed fullyarticulated Active Twist Rotor (ATR) system was built and tested at Langley Transonic Dynamics Tunnel. From these tests, the integral twist control authority exerted upon the fixed system was determined. Significant control authority in hub vertical shear load component is observed from different blade actuation modes. Similar control authority is found in the other components of the fixed-system loads. Exploiting those authorities, vehicle performance enhancement can be achieved using low-frequency actuation, for example, 0P, 1P and 2P. Payload increase in hover can be obtained with a steady collective actuation of blade twist. Power consumption can be reduced by employing a certain mode of blade actuation at 2P in forward flight. Vehicle pitch and roll moments are generated by longitudinal and lateral blade actuation mode at 1P frequency. On the other hand, actuation at higher frequencies can be used to reduce hub vibratory loads. The closed-loop control algorithm used for this reduction is an improved version of the traditional Higher Harmonic Control. Multiharmonic and multi-mode controller is designed and tested as part of the present study.
In the presented article a comparative analysis of efficiency of the helicopter rotor blades vibrations suppression by active (controlled) and passive (shunted by electric circuit) piezoelectric patches was performed. For obtaining the information about influence of the external load circuit parameters to behavior of kinematically excited actuator PZT patch the harmonic analysis in the frameworks of 1D electro-elasticity for R, L, C and combined RLC type of a load in a longwave approximation is performed. For PZT-5H plate working in d31-mode the dependence of oscillations energy harvesting and change of actuator's elastic properties on the patch's geometrical and material parameters were founded. It is shown that for reasonable values of the load capacitance, inductance and resistance the effective suppression of oscillations energy is marked at frequencies greater than 1 kHz only. A flexible structure with surface-bonded PZT patch actuated by PD controller is modelled using Euler-Bernulli beam theory. Emphasis in this study is given to the effect of actuator size and location on the observability, controllability, and stability of the control system. The hybrid vibration damping system with PZT patches actuators was proposed. In this system the vibration suppression on the first flexural modes is yielded with use of the power PD-controllers, and oscillation damping on the torsional and higher flexural modes -by PZT patches loaded on a tuned RC circuits. Efficiency of the offered solution is illustrated by results of the transient analysis of the beam FE model and by experimental data obtained on a scaled model of the helicopter composite rotor blade with bonded flexural and torsional -operated PZT actuators.
Pseudo-Active Control of Helicopter Blade Vibration Using Piezoelectric Actuators
45th AIAA/ASME/ASCE/AHS/ASC Structures, Structural Dynamics & Materials Conference, 2004
In this paper we develop a closed-loop algorithm for vibration control of helicopter blades incorporating Piezoelectric Fiber Composite actuators for Active Twist Control. A reduced-state sequential feedback controller is designed based on the linear piezoelectric constitutive equations. We use a simple aeroelastic model that includes 4 bending and 4 torsional degrees of freedom for an initial estimation of the effectiveness of the controller in reducing the vibrational components of the loads transmitted to the hub. The influence of the nonlinear material piezoelectric properties of the actuators is studied by taking into account terms quadratic in the electric field in the piezo constitutive equations. The control method is a velocity feedback scheme incorporating a switching mechanism based on a conditioning function inspired by the force balance logic used in sequential semi-active control syntheses for vibration control. A genetic algorithm is employed to select the optimal control gains. Significant reductions of the magnitudes of the periodic components of the root shear force can be achieved in the desired range of frequencies. Accounting for the system nonlinearities implicitly through the optimization process is observed to improve the control system performance.
Bench-top characterization of an active rotor blade flap system incorporating C-block actuators
1998
This paper presents the bench-top testing of a piezoceramic C-block driven active flap system designed to suppress the vibrations of a helicopter rotor blade. The C-block actuators are curved benders designed to generate a larger force output than a straight bender, while providing deflections large enough to eliminate the need for external leveraging systems necessary with stack driven systems. The actuators power a balanced active flap designed to minimize the effect of air speed and rotor speed on flap deflection. Quasi-static experimentation at 1 Hz produced maximum angular flap deflections of 8.4° peak-to-peak. Dynamic tests were conducted over a 40 Hz frequency range demonstrating the ability to generate significant flap deflections both before and after the first natural frequency. Over the 40 Hz range, the flap deflections never dropped below 8° pp, with a first natural frequency of 27 Hz. The flap deflection reached a maximum value of 13.6° pp at 40 Hz. If the applied voltage is increased to the maximum allowable level, it is predicted that flap deflections as large as 20° pp can be achieved.