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A methodology to develop piezostack-based actuator for trailing edge flap on a full-scale helicopter rotor blade is presented. By a systematic down-selection approach through the experimental validation on piezostack characteristics, one among numerous candidates was chosen for current study, which has higher stiffness as well as good strain/block force capability. The prototype actuator was designed and fabricated using two pieces of selected piezostack incorporated with a double-lever amplification mechanism that extends a level-fulcrum amplification concept. To validate the feasibility of the designed actuator, the prototype actuator was tested in vacuum chamber that simulates the centrifugal force environment. It was shown that the actuator operated with no major degradation in performance up to 600g of centrifugal loading. The present actuator has a capability of oscillating the trailing edge flap to the required deflection/frequency to actively control rotor vibration in the rotating environment. Then, a further design refinement was considered to reduce the drawbacks, which finally results in an improved actuator with peak actuation force of 80 lbs at 38 mils stroke.© (1999) COPYRIGHT SPIE--The International Society for Optical Engineering. Downloading of the abstract is permitted for personal use only.

This paper presents the design and preliminary test results of a piezoelectric stack (piezostack) driven leading-edge flap actuator. The actuator uses six commercially available piezostacks with a curved contact tip for force and stroke profile tailoring. Force requirements were calculated using a quasi-steady aerodynamic model. Stroke amplification is achieved through a single lever. Initial results show that the current prototype is not ready to achieve the desired stroke output, but potential problems are identified and action is being take to rectify these problems.© (1999) COPYRIGHT SPIE--The International Society for Optical Engineering. Downloading of the abstract is permitted for personal use only.

Shape memory alloys (SMAs) have shown promise as high-force, high displacement actuators. Critical issues such as path- dependence, predictability and sensitivity to testing conditions, however, need to be addressed in order to design controllable actuators using SMAs. This paper presents research aimed at addressing some of design issues involving application of SMAs, particularly at actuators. Quasistatic experiments at constant stress, strain and temperature are consolidated on a critical stress-temperature diagram to delineate the regions of stability of the various phases of the material. The critical points from these quasistatic tests are found to be in excellent agreement with each other, and correlate relatively well with the constitutive models for SMA thermomechanical behavior. It is also observed that the state of the material is not unique at points along the transformation, and is dependent on the history of the material before the start of the test, individual test involved, the method of loading, and loading rates. Significant variation of the state of the material with different rates and conditions of loading are shown to further illustrate this point. This behavior is likely to be decisive in determining the dynamic behavior of the material, and underscores the need for approaches incorporating these issues for design of repeatable actuators.© (1999) COPYRIGHT SPIE--The International Society for Optical Engineering. Downloading of the abstract is permitted for personal use only.

This paper investigates the behavior of piezoelectric elements as strain sensors. Strain is measured in terms of the charge generated by the element as a result of the direct piezoelectric effect. Strains from piezoceramic and piezofilm sensors are compared with strains from a conventional foil strain gage and the advantages of each type of sensor are discussed, along with their limitations. The sensors are surface bonded and are calibrated by means of a dynamic beam bending setup over a frequency range of 5 - 500 Hz. Correction factors to account for transverse strain and shear lag effects due to the bond layer are analytically derived and validated experimentally. Additionally, design of signal conditioning electronics to collect the signals from the piezoelectric sensors is addressed. The superior performance of piezoelectric sensors compared to conventional strain gages in terms of sensitivity and signal to noise ratio is demonstrated.

This paper describes the concept, design, fabrication and control of a shape memory alloy (SMA) wire actuator for tracking helicopter rotor blades while in-flight. A NACA 0012 wing section that has a 12 in chord and span was constructed with a trailing-edge tab with a 4 in span and a 2.4 in chord. A shape memory alloy wire actuator was embedded into the wing section. The actuator consists of a wire clamp, a hinge tube and several pre-strained, 0.015 in diameter SMA (Nitinol) wires. It was shown that with SMA wires that have 3.158% initial pre-strain, a tab deflection of 29° could be obtained.

This paper investigates the behavior of piezoceramic actuators under different types of excitation and mechanical loading. The research focuses on the application of these actuators to the development of smart rotor systems. The free strain response of the actuators under DC excitation is investigated experimentally along with the associated drift of the strain over time. Effect of tensile stress on the DC response is investigated. Strain response to an AC excitation is also investigated for a free actuator and for a pair of actuators surface bonded to a host structure. The phase lag of the strain response and non-linear hysteretic effects have been observed and the power consumption has been validated by the impedance method. Additionally, depoling of the actuators is discussed, along with the feasibility of recovering performance by repoling in the event of accidental depoling.

This paper presents an analytic model and validation tests of Froude and Mach scaled rotors featuring piezoelectric bender actuated trailing-edge flaps for active vibration suppression. A finite element structural formulation in conjunction with time domain unsteady aerodynamics is used to develop the rotor blade and bender-flap coupled response in hover. The analysis accounts for the aerodynamic, centrifugal, inertial and frictional loads acting on the coupled bender-flap-rotor system in hover. To investigate the feasibility of piezo-bender actuation and validate the analytic model, a 6 foot diameter, two-bladed Froude scaled rotor with piezo-bender actuation is tested on the hover stand. Flap deflections of +/- 4 to +/- 8 degrees, for 1 to 5 /rev bender excitation were achieved at the Froude scaled operating speed of 900 RPM. The trailing-edge flap activation resulted in a 10% variation in the rotor thrust levels at 6 degrees collective pitch. The analytic model shows good correlation with experimental flap deflections and oscillatory hubloads for different rotor speeds and collective settings. Based on the analytic model, two Mach scaled rotor blades with piezo-bender actuation are designed and fabricated. To achieve the desired flap performance at Mach scale, an 8-layered, tapered bender is fabricated. The bender performance is further improved by selectively applying large electric fields in the direction of polarization for individual piezoceramic elements. A radial bearing is incorporated to reduce frictional loads at the blade-flap interface. Preliminary testing of the Mach scaled blades at 900 RPM in a vacuum chamber revealed negligible degradation in flap performance because of centrifugal and frictional loading. The analysis predicts that flap deflections of +/- 5 to +/- 10 degrees for 1 to 4 /rev bender excitation can be achieved at the Mach scaled operating speed of 2100 RPM. Future work will involve hover and wind tunnel testing at Mach scaled operating speeds.

This paper presents the development of an actuator for a trailing edge flap on a full-scale helicopter rotor blade using a high performance piezoelectric stack device. The actuator was designed and constructed using two identical piezostacks that were selected among commercially available actuators through a systematic testing. To calculate the aerodynamic requirement of the flap actuator, an analytical model was formulated using quasi-steady aerodynamics. A new amplification device utilizing double-lever mechanism was designed and fabricated. An amplification factor of 19.4 and the constant response covering up to 8/rev were obtained experimentally under non-rotating condition. It was shown that the present flap actuator is capable of achieving the required flap deflection to actively control rotor vibration.

This paper examines, analytically and experimentally, a shape memory alloy (SMA) wire actuator to be used for in- flight tracking of helicopter rotor blades. The recovery characteristics of SMA's are utilized to detect a trailing edge trim tab. Conceptually, the tab is deflected using SMA wires with initial pre-strains, and any differential recovery displacement due to a thermal field is transformed into an angular motion of the tab. A 1D thermo-mechanical constitutive model is used to predict the constrained recovery characteristics of SMA's and predicted results are validated with measured values. Good correlation between measured and predicted values is obtained.

This paper presents the continued development of an active on-blade vibration-reduction system using smart active blade tips (SABT), that are driven by a piezo-induced bending-torsion coupled actuator. A Vlasov based, specialized one-dimensional finite bar-element is developed to model the (nonrotating) actuator beam and is validated with the free-vibration and static forced response of 8:1 and 2:1 aspect ratio bending-torsion coupled plates. The FEM over-predicts the actuator torsional frequency by 4%, and accurately captures the actuator dynamics up to 60Hz. In hover, at thrust coefficient of 0.0035, and for an activation of 100Vrms the SABT deflection ranges from l.Sdeg at 2/rev to 2.25deg at 4/rev. The results show a distinct coupling of the activation with the first and second flap frequencies of the rotor. The corresponding dynamic thrust, generated by a single active tip, relative to the steady thrust, ranges from 4.5% at 2/rev to 8.3% at 5/rev.

Smart structure activated trailing edge flaps are capable of actively altering the aerodynamic loads on rotor blades. With a suitable feedback control law, such actuators could potentially be used to counter the vibrations induced by periodic aerodynamic loading on the blades, without the bandwidth constraints and with lower weight penalties incurred by previous actuation methods. This paper explores a new, robust individual blade control (IBC) control methodology for vibration suppression using a piezo actuated trailing edge flap. The controller uses a neural network, learning in real time, to adaptively cancel the effects of periodic aerodynamic loads on the blades, greatly attenuating the resulting vibrations. A complete proof of the stability and convergence of the proposed neurocontrol strategy is provided, and numerical simulation results for a one-eighth Froude scale blade model are given which demonstrate that the controller can virtually eliminate the blade vibration arising from a wide variety of unknown, periodic disturbance sources.

A smart structure involves distributed actuators and sensors, and one or more microprocessors that analyze the responses from the sensors and use distributed-parameter control theory to command the actuators to apply localized strains to minimize system response. A smart structure has the capability to respond to a changing external environment (such as loads or shape change) as well as to a changing internal environment ( such as damage or failure). It incorporates smart actuators that allow the alteration of system characteristics ( such as stiffness or damping) as well as of system response (such as strain or shape) in a controlled manner. Many types of actuators and sensors are being considered, such as piezoelectric materials, shape memory alloys, electrostrictive materials, magnetostrictive materials, electro-rheological fluids and fiber optics. These can be integrated with main load-carrying structures by surface bonding or embedding without causing any significant changes in the mass or structural stiffness of the system. Numerous applications of smart structures technology to various physical systems are evolving to actively control vibration, noise, aeroelastic stability, damping, shape and stress distribution. Applications range from space systems, fixed-wing and rotary-wing aircraft, automotive, civil structures and machine tools. Much of the early development of smart structures methodology was driven by space applications such as vibration and shape control of large flexible space structures, but now wider applications are envisaged for aeronautical and other systems. Embedded or surface-bonded smart actuators cn an airplane wing or helicopter blade will induce alteration of twist/camber of airfoil (shape change), that in turn will cause variation of lift distribution and may help to control static and dynamic aeroelastic problems. Applications of smart structures technology to aerospace and other systems are expanding rapidly. Major barriers are: actuator stroke, reliable data base of smart material characterteristics, non-availability of robust distributed parameter control stratgies, and non-existent mathematical modeling of smart systems. The objective of this paper is to review the state-of-the-art of smart actuators and sensors and integrated systems and point out the needs for future research.

This paper presents the testing of a Froude scaled rotor featuring piezoceramic bimorph actuated trailing-edge flaps for active vibration suppression. An analytical model for the coupled bimorph actuator trailing-edge flap dynamic response in the rotating environment was developed. The analysis takes into account the aerodynamic, centrifugal, inertial and frictional loads acting on the actuator-flap system. The linkage arm length associated with the mechanical amplification mechanism was selected in order to maximize flap performance in the rotating environment. The bimorph clamping was improved to prevent actuator slippage under high centrifugal loads. The analytical model was validated by carrying out a series of bench tests, vacuum chamber tests and hover tests. In hover, flap deflections of plus or minus 6 degrees at 4/rev flap excitation were achieved at 900 rpm, thus demonstrating the potential of the piezoceramic bender as a lightweight and compact actuation system for individual blade control purposes. This paper also includes a feasibility study for piezo-bimorph actuation of a trailing-edge flap for a Mach scaled rotor model.

This paper presents research into developing an adaptive nonlinear neural network control algorithm that can be used with smart structure actuators and sensors to control the vibrations of rotor blades. The dynamic equations of motion for a blade have the same form as a multilink manipulator (robot arm) and adaptive nonlinear control algorithms have proven successful in active control of these manipulators. The recent development of neural network control algorithms has provided the ability to adaptively learn in real time a set of parameters that will approximate external forces operating on the blades. The controller combines these two control techniques enabling the controller to adapt its parameters in response to changes in blade properties such as its mass or stiffness and to also learn the parameters necessary to account for the unknown but bounded, periodic disturbance forces such as those caused by the unsteady, periodic aerodynamic forces in the rotor system. Current efforts have been directed at testing the control algorithm on real beams with piezoceramic actuators and sensors. The initial test results have shown that vibration reduction and desired beam motion tracking can be achieved even under the influences of periodic disturbances.

The objective of this research is to develop a smart rotor with active control of blade twist using embedded piezoceramic elements as sensors and actuators to minimize rotor vibrations. A 1/8-th Froude-scale (dynamically-scaled) bearingless helicopter rotor model was built with banks of torsional actuators capable of manipulating blade twist at frequencies from 5 to 100 Hz. To assess the effectiveness of the torsional actuators and vibration suppression capabilities, systematic wind tunnel testing was conducted in the Glenn L. Martin Wind Tunnel. Using accelerometers embedded in the blade tip, the oscillatory blade twist response was measured. The changes in rotor vibratory loads due to piezo- induced twist were determined using a rotating hub balance located at the rotor hub. Experimental test results show that tip twist amplitudes on the order of 0.5 deg are attainable by the current actuator configurations in forward flight. Although these amplitudes were less than the target value (1 - 2 deg for complete vibration suppression control), test results show that partial vibration reduction is possible. Using open-loop phase shift control of blade twist at the first four rotor harmonics, changes in rotor thrust of up to 9% of the steady-state values were measured, resulting in up to 3 and 8% reductions in rotor pitching and rolling moments, respectively. It is expected that the hub load control authority of the smart rotor can be improved in future models with refined actuator configurations and implementation of closed-loop feedback controls.© (1997) COPYRIGHT SPIE--The International Society for Optical Engineering. Downloading of the abstract is permitted for personal use only.

Coupled rotor/fuselage vibration model is formulated incorporating consistent structural, aerodynamic and inertial coupling effects. The elastic line model of the AH-1G helicopter in conjunction with a typical 4-bladed hingeless rotor is used in the analysis. The coupled nonlinear, periodic equations of elastic rotor, elastic body and rigid body trim are solved using finite element methods in space and time. For calculation of inflow and blade loads, a pseudo-implicit free wake model and the time-domain unsteady aerodynamics are incorporated. Rotor/fuselage coupling effects usually increase oscillatory hub loads and these are significant at low and high advance ratios. Unsteady aerodynamics plays an important role at high advance ratio and free wake is essential in the prediction fuselage vibration at all forward speeds. CT rotor thrust coefficient

A consistent plate finite element model is formulated for coupled composite plates with induced strain actuation and validated with test data obtained from cantilevered isotropic and anisotropic plates. Actuators are modeled as additional plies fully integrated into substrate laminae, and the formulation is based on modified thin classical laminated-plate theory. The analysis is formulated for a generic anisotropic plate with a number of piezoactuators of arbitrary size, surface-bonded or embedded at arbitrary locations. Composite plates with extension-twist and bending-twist couplings were built and tested. Two rows of piezoceramic elements are surface mounted on both top and bottom surfaces near the root. Static tests are carried out using induced strain actuation, and mechanical loading and measured data are correlated with predictions for bending and twist distributions. For an extension-twist coupled plate, the agreement between predicted and measured induced twist due to extensional strain with piezoactuation is excellent. For the strongly bending-twist coupled composite plate, the predicted induced twist due to bending strain with piezoactuation agreed well in trends, but magnitudes were underpredicted by a maximum of 20% from measured values. For the weakly bending-twist coupled composite plate, the predicted induced-twist angle agreed extremely well with measured data. The modeling and validation results show the usability of the piezoactuation in the field of plate shape control.

An experimental-theoretical study of Shape Memory Alloys (SMA) wires was undertaken in order to use the SMA wires in the active tuning of rotating composite shafts. Four constitutive models are compared and validated with experimental results. The four models compared were developed by Tanaka, Liang and Rogers, Boyd and Lagoudas, and Brinson. Careful testing of SMA wires was carried out to determine the necessary coefficients for the four constitutive models. Coefficients such as the Young's modulus for the martensite and austenite phase and the four phase transformation temperatures were determined. All models showed good qualitative trends; however quantitatively, all except one fared unsatisfactorily. The model developed by Brinson showed good correlation with experimental data.

This paper presents an active moving blade tip for a smart rotor with torsional actuation via a piezo-induced bending-torsion coupled composite beam. A novel spanwise variation in the beam layup and piezoceramic element phasing is used to maximize the twist response, while minimizing the bending response. The composite beam is located lengthwise within the blade with surface bonded piezoceramic elements excited to induce spanwise and chordwise bending of the beam. The bending results in an induced twist via the structural coupling and this twist deflects the moving tip. A proof-of-concept actuator beam was developed and tested. Tip twist with essentially zero tip bending deformation was demonstrated. Two operational 1/8th scale model rotor blades were manufactured for hover testing. In non-rotating calibration tests, at a 60 Hz actuator excitation, a peak to peak moving tip deflection of 4 deg was achieved.© (1996) COPYRIGHT SPIE--The International Society for Optical Engineering. Downloading of the abstract is permitted for personal use only.

This paper presents the results of research in using piezoelectric stacks to power a trailing edge flap for use as an individual blade control device to suppress vibrations in helicopter rotor blades. A flap actuator was designed and constructed using two piezoelectric stacks arranged in series. Mechanical amplification was provided through a hinged L-arm. An analytical model was formulated to calculate the predicted flap deflection using quasi-steady aerodynamics. The model showed that the high force piezoelectric stacks were capable of maintaining flap deflection over a wide range of freestream velocities. An experimental model consisting of an 8 inch chord NACA 0012 airfoil was constructed incorporating a single 4-inch span, 1.6 inch chord trailing edge flap. The flap performance of the model was tested in no load conditions and in the wind tunnel under freestream velocities of up to 118 ft/sec. The flap deflections were lower for higher velocities and below analytical predictions.© (1996) COPYRIGHT SPIE--The International Society for Optical Engineering. Downloading of the abstract is permitted for personal use only.

A smart structure involves distributed actuators and sensors, and one or more microprocessors that analyze the responses from the sensors and use distributed-parameter control theory to command the actuators to apply localized strains to minimize system response. A smart structure has the capability to respond to a changing external environment (such as loads or shape change) as well as to a changing internal environment (such as damage or failure). It incorporates smart actuators that allow the alteration of system characteristics (such as stiffness or damping) as well as of system response (such as strain or shape) in a controlled manner. Many types of actuators and sensors are being considered, such as piezoelectric materials, shape memory alloys, electrostrictive materials, magnetostrictive materials, electro- rheological fluids and fiber optics. These can be integrated with main load-carrying structures by surface bonding or embedding without causing any significant changes in the mass or structural stiffness of the system. Numerous applications of smart structures technology to various physical systems are evolving to actively control vibration, noise, aeroelastic stability, damping, shape and stress distribution. Applications range from space systems, fixed-wing and rotary-wing aircraft, automotive, civil structures and machine tools. Much of the early development of smart structures methodology was driven by space applications such as vibration and shape control of large flexible space structures, but now wider applications are envisaged for aeronautical and other systems. Embedded or surface-bonded smart actuators on an airplane wing or helicopter blade will induce alteration of twist/camber of airfoil (shape change), that in turn will cause variation of lift distribution and may help to control static and dynamic aeroelastic problems. Applications of smart structures technology to aerospace and other systems are expanding rapidly. Major barriers are: actuator stroke, reliable data base of smart material characteristics, non-availability of robust distributed parameter control strategies, and non-existent mathematical modeling of smart systems. The objective of this paper is to review the state-of-the-art of smart actuators and sensors and integrated systems and point out the needs for future research.© (1996) COPYRIGHT SPIE--The International Society for Optical Engineering. Downloading of the abstract is permitted for personal use only.

This paper presents the experimental results of hover tests with a 1/8th dynamically-scaled (Froude scale) helicopter rotor model with embedded piezoceramic elements as actuators to suppress vibrations. A 6-ft diameter four-bladed bearingless rotor model was built with banks of specially-shaped piezoelectric actuators embedded at +/- 45 deg on the top and bottom surfaces to manipulate blade twist at harmonics of the rotational speed. The twist performance of rotors with single-layer and dual-layer actuator configurations was measured using accelerometers located at the blade tip. The change in oscillatory rotor lift due to piezoactuation was measured by a rotating balance. Experimental results show that linear twist distributions on the order of 0.3 deg, resulting in increases of up to 9 percent of the nominal rotor lift, are possible with existing piezoceramic technology. Although the twist amplitudes attained in this experiment were less than the target value (2 deg for complete vibration suppression), partial reduction of rotor vibrations can be achieved with the current smart rotors. (Author)

A comprehensive aeroelastic analysis is developed to predict the performance, vibratory loads, and aeroelastic stability of composite-coupled advanced-geometry tilt-rotors. Elastic motion of the rotor blades, wing, and fuselage and gimbal motion are modeled. Full wingspan and twin rotors are modeled. Finite element modeling is used along with normal mode reduction to make the analysis robust and efe cient. Flight conditions modeled are helicopter (hover and forward e ight ), conversion, and airplane modes of operation. Predictions from this analysis for trim controls, response, loads, and aeroelastic stability correlate well with e ight test data and predictions from other established analyses.

The aeroelastic stability of a composite hingeless rotor with elastic couplings is investigated in hover. A four-bladed, Froude-scaled, soft in-plane hingeless rotor model with various isotropic and graphite/epoxy composite flexures was designed and tested on a hover stand. By changing the flexure setting angle, the rotor with an isotropic flexure was tested to examine the effects of flap-lag coupling on stability. The effect of elastic flapwise bending-torsion and chordwise bending-torsion structural coupling on the aeroelastic stability were investigated using elastically coupled composite flexures. Lag mode stability test results are compared with the theoretical predictions obtained using the comprehensive aeroelastic rotor analysis, UMARC. The introduction of either negative chord-wise bending torsion coupling or negative flap-lag coupling has a stabilizing effect on the lag damping for positive collective pitch angles. Predictions agree with measured data. (Author)

The objective of this paper is to present the feasibility of applying Vlasov's theory of thin walled bars to model the behavior of open section beams actuated in torsion and bending by skewed piezoactuators. It is shown that piezoactuators placed at an angle with the longitudinal axis of a beam induce bending and twisting by subjecting the beam to a system of bending moments applied at a distance from the pole axis. A simple analytical model based on Vlasov's theory is presented and validated against experimental results.

Main rotor blades with plain trailing edge flaps are investigated as a potential means of vibration reduction using a comprehensive rotorcraft analysis. The analysis is modified to incorporate an unsteady aerodynamic model including flap effects. Predicted blade and pitch link loads show adequate correlation with experimental data. An initial open loop study shows that, for a typical articulated four bladed rotor, significant reductions in fixed system 4/rev vertical shears and hub moments are possible with open loop flap deflections at 3/rev and 4lrev. The reductions are most closely associated with changes in the response of the blade third flatwise bending mode and torsional deflections. A reduction in blade torsional stiffness reduced the effectiveness of the flap in reducing vibration.

A tiltrotor model is formulated to predict the dynamics and aeroelastic stability of composite-coupled advanced geometry rotors. Advanced geometry rotor blades are modeled with arbitrary sweep, anhedral, pretwist and planform taper variation with span. The effect of advanced geometry blades on the tiltrotor response and whirl flutter stability in airplane mode is studied. Blade tip sweep and tip droop increase the flutter speed while forward tip sweep, tip anhedral and planform taper decrease it. Predicted wing beam mode damping correlates satisfactorily with test data.

The objective of this research is to develop a dynamically scaled (Froude scale) helicopter rotor blade with embedded piezoceramic elements as sensors and actuators to control blade vibrations. A 6-ft-diameter two-bladed bearingless rotor model was built, where each blade is embedded with banks of piezoelectric actuators at +/- 45-degree angles with respect to the beam axis on the top and bottom surfaces. A twist distribution along the blade span is achieved through in-phase excitation of the top and bottom actuators at equal potentials, while a bending distribution is achieved through out-of-phase excitation. In order to fix design variables and to optimize blade performance, a uniform strain beam theory is formulated to analytically predict the static bending and torsional response of composite rectangular beams with embedded piezoelectric actuators. Parameters such as bond thicknesses, actuator skew angle, and actuator spacing are investigated by experiments and then validated by theory. The static bending and torsional response of the rotor blades is experimentally measured and correlated with theory. Dynamic torsional and bending responses are experimentally determined for frequencies from 2-120 HZ to assess the viability of a vibration reduction system based on piezoactuation of blade twist. Although the magnitudes of blade twist attained in this experiment were small, it is expected that future models can be built with improved performance.© (1994) COPYRIGHT SPIE--The International Society for Optical Engineering. Downloading of the abstract is permitted for personal use only.

We develop electromechanical actuators based on the concept of mechanical amplification with piezo and electrostrictive stacks as drivers to achieve high force and high displacement actuation. The actuators were designed for two different applications. Experiments were performed on both stacks to evaluate their important characteristics such as block force, free displacement, and stiffness that were essential in the design of the actuators. The results showed that the block force obtainable from piezo stack was higher and that of electrostrictive stack was lower than that specified by the respective manufacturers while the free displacements are about the same. The dynamic response of the actuators over a frequency range of 33 Hz was evaluated. Results showed that the actuation force of 7 lbs was obtainable in both cases with flaperon actuator producing 15 mils of dynamic displacement at 15 Hz and droop flap actuator producing about 6 mils of displacement at 16 Hz. The results were inconclusive beyond 16 Hz due to the setup resonance. Droop flap actuator did not achieve the desired performance because the design calculations were based on block force listed by the manufacturer which was about 20% higher than the measured value. This led to the conclusion that before design process begins, the stack alone performance should be carefully measured to achieve required performance. Thus, a simple actuator based on mechnaical amplification concept could be effectively designed to produce high force and high displacements.© (1994) COPYRIGHT SPIE--The International Society for Optical Engineering. Downloading of the abstract is permitted for personal use only.

A nonlinear viscoelastic solid model, comprising a combination of linear and nonlinear springs and dashpots, is developed to represent an elastomeric damper. The nonlinear constitutive differential equation obtained from the model completely characterizes the damper behavior. A method is presented to determine the spring-dashpot parameters (coefficients of the constitutive equation) from experimental data. A quartic softening spring, in series with linear Kelvin chain, is used to match experimental data. Nonlinear hysteresis cycles at different equilibrium positions are examined. The model is able to predict behavior of elastomeric dampers under dual-frequency excitations. A `two-level implicit-implicit' scheme is developed for the integration of the nonlinear damper model into a structural dynamic analysis. With the increase in amplitude of oscillatory force, the energy dissipation by the nonlinear viscoelastic damper is found to decrease, as compared to a linearized perturbation model. With increase in initial perturbation, transient decay is slower.© (1994) COPYRIGHT SPIE--The International Society for Optical Engineering. Downloading of the abstract is permitted for personal use only.

This paper develops a one-dimensional model to predict the coupled extension, bending and torsion response of a beam subject to piezoceramic strain actuation. The effects of cross-sectional warping are shown to be negligible for thin rectangular isotropic beams. The impact of adhesive shear lag, on the other hand, is measurable, especially in the torsional response. Experimental test results show that the models are accurate up to a 45O actuator orientation with respect to the beam axis even though detailed strain data indicate that the local strain state is highly two-dimensional . Nomenclature

The effect of a new 'differential equation based' elastomeric lag damper model is examined on shaft-fixed flaplag stability in forward flight. The nonlinear elastomeric damper is modeled using a combination of linear and nonlinear springs and dashpots, and is characterized completely by a nonlinear constitutive differential equation. Since damper description is not based on the frequencyand amplitudedependent complex modulus components, it can conveniently be used in a variety of operating conditions, including multi-frequency excitations, varying rotor speeds, and different equilibrium conditions. For propulsive trim calculations, the nonlinear damper equation is solved simultaneously with the rotor and fuselage equilibrium relations, with the equilibrium damper response being accurately determined. Flaplag stability is calculated in the rotating frame about the propulsive trim condition using Floquet theory. For this, the perturbation damper equation is solved along with the linearized blade flap and lag equations, with damper-force appearing as a separate state. The elastomeric damper has a stabilizing influence on blade lag damping. Neglecting the equilibrium damper response overpredicts lag damping by 100%. Neglecting the cyclic components of equilibrium damper response (with collective included), overpredicts hingeless blade lag damping, but can accurately predict articulated blade lag damping up to advance ratios of 0.3. Lag damping decreases as disk load increases, more so for hingeless blades.

This paper presents an experimental study on the development of a Froude scale helicopter rotor blade with trailing edge flap as a vibration reduction device actuated by piezoelectric crystals. A fixed wing model with NACA 0012 airfoil, 3.0 inch blade chord and 20% trailing edge flap is fabricated and tested in the open-jet tunnel to determine the dynamic flap response at various blade angles of attack and excitation frequencies. The results show that for a given velocity, flap response does not change appreciably with the excitation frequency and the blade angle of attack.© (1993) COPYRIGHT SPIE--The International Society for Optical Engineering. Downloading of the abstract is permitted for personal use only.

The objective of this research is to develop a dynamically-scaled helicopter rotor blade with embedded piezoceramic elements as sensors and actuators to control blade vibrations. A 6 ft diameter 2-bladed Froude-scale bearingless rotor model is built where each blade is embedded with banks of specially-shaped piezoelectric crystals at +/- 45 degree angles on the top and bottom surfaces. A twist distribution along the blade span is achieved through in-phase excitation of the top and bottom crystals at equal potentials. The non-rotating static torsional response of the piezoceramic blade is experimentally determined and then correlated with the prediction by theory.

This paper develops one-dimensional pure bending, coupled bending and extension, and combined bending, extension and torsion models of isotropic beams with induced strain actuation. A finite thickness adhesive layer between the crystal and beam is included to incorporate shear lag effects. Experimental tests evaluate the accuracy and limitations of the models. The bending and coupled bending and extension models show acceptable correlation with static test results whereas the combined extension, bending, torsion model poorly predicts the system behavior and needs refinement.© (1993) COPYRIGHT SPIE--The International Society for Optical Engineering. Downloading of the abstract is permitted for personal use only.

The aeroelastic response, blade and hub loads, and shaft-fixed aeroelastic stability is investigated for a helicopter with elastically tailored composite rotor blades. A new finite element based structural analysis including nonclassical effects such as transverse shear, torsion related warping and inplane elasticity is integrated with the University of Maryland Advanced Rotorcraft Code. The structural dynamics analysis is correlated against both experimental data and detailed finite element results. Correlation of rotating natural frequencies of coupled composite box-beams is generally within 5-10 percent. The analysis is applied to a soft-inplane hingeless rotor helicopter in free flight propulsive trim. For example, lag mode damping can be increased 300 percent over a range of thrust conditions and forward speeds. The influence of unsteady aerodynamics on the blade response and vibratory hub loads is also investigated. The magnitude and phase of the flap response is substantially altered by the unsteady aerodynamic effects. Vibratory hub loads increase up to 30 percent due to unsteady aerodynamic effects.