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This paper describes the systematic experimental studies performed to investigate and understand the aerodynamic performance of a small-scale vertical axis wind turbine (VAWT) with a dynamic blade pitching mechanism. A VAWT prototype was designed, built, and tested in the wind tunnel to understand the role of pitch kinematics on aerodynamic efficiency. The ability of the dynamic blade pitching mechanism to change phase instantaneously in order to adapt to changes in wind direction is the key to maximizing power extraction in space-constrained urban environments. Experimental studies showed that turbine efficiency is heavily dependent upon blade pitch amplitude, with maximum performance occurring around ±20° to ±25° pitch amplitude. Experiments also found that the optimum tip speed ratio (TSR) is also dependent upon pitch amplitude and that it decreases with increasing blade pitch amplitude for symmetric and asymmetric blade pitching. Asymmetric blade pitching was found to have a negative effect upon performance as compared to symmetric blade pitching, as was blade phasing away from the direction of the free-stream. This study clearly indicates that turbine performance is maximized for a symmetric airfoil at a symmetric blade pitch amplitude between ±20° to ±25° with no phasing. A small-scale horizontal axis wind turbine was also designed, built, and tested in the wind tunnel to understand the effect of Reynolds Number on the aerodynamic efficiency of small-scale wind turbines. As the Reynolds Number at the 75% span decreases below a certain value, there is a precipitous drop in turbine aerodynamic performance due to flow separation on the airfoil section. This indicates that perhaps a larger chord or turbine radius holds the potential for improved VAWT performance at the small-scale because of the increase in the blade Reynolds Number.

The quad rotor biplane is a multi-mission vehicle that incorporates four rotors and two high-lift wings for hover and high-speed forward flight. The concept promises improved mission capabilities and low mechanical complexity by relying on the rotors for both propulsion and control, but presents significant flight control challenges due to the unmodelled aerodynamics that arise in the presence of gusts. This paper describes preliminary results from using onboard flow measurements to improve the longitudinal stability of a quad rotor biplane when subjected to vertical gusts. A reduced-degree-of-freedom test stand is constructed with a single wing mounted between a pair of rotors that provide propulsion and control through differential thrust. This modified representation retains the control methodology of the quad biplane and allows the pitch dynamics to be studied without the additional effect of wing-wing or rotor-wing interaction. The pitching dynamics of the test stand is modeled with a focus on the aerodynamic moments induced by external wind, and a pitch control strategy is developed that uses flow measurements to feedback-linearize the system. Ground-based testing of a flow-feedback enhanced autopilot demonstrates the advantage of using onboard flow measurements for improved longitudinal control in the presence of gusts.

Background and Motivation Current requirements for future helicopters as outlined by the Future Vertical Lift, Joint Multirole and DARPA X-plane competitions all call for helicopters to fly at significantly higher speeds than they can currently achieve while maintaining exceptional hovering performance. The Sikorsky X2 has flown faster than 250 kts, demonstrating that the rigid coaxial compound is a viable candidate for future helicopter configurations. In addition to the performance benefits, the rigid coaxial compound provides significant freedom in trim allowing for hover and flight in a variety of pitch attitudes as well as the use of lift offsets in the rotor. This paper presents work that has developed a comprehensive analysis technique for analyzing coaxial rotors and is used to study the impact of coaxial configuration on performance and loads.

Compound helicopter designs utilize a slowed main rotor, which reduces RPM as flight speed increase to maintain sub-critical tip Mach numbers. These concepts are therefore expected to operate at high advance ratios (1.0 or higher). There is a need for experimental data sets in this flight regime to validate and develop predictive tools. Systematic testing in the Glenn L. Martin Wind Tunnel (University of Maryland) was conducted on an instrumented Mach-scale articulated rotor (1.7 m diameter) up to an advance ratio of 1.61. The following measurements were taken: steady and dynamic 6-component hub loads (fixedframe), shaft torque (rotating frame), root flapping angles, pitch link loads, blade torsion and flap bending moments at 5 spanwise stations, and 16 chordwise blade pressures at 30% radius spanwise location. The selected results detailed in this paper revealed insights into high advance ratio, reverse flow phenomenon such as reverse chord dynamic stall and the resulting impact on loads. The combination of rotor advance ratios beyond 1.0, large collective pitch sweeps, and blade surface pressure measurements makes this data set unique.

Hover and forward flight capability can be combined in hybrid air vehicle designs such as a quad rotor biplane which is investigated in this paper. The vehicle weighs 240 grams and consists of four propellers with wings arranged in biplane configuration. To measure aerodynamic performance of the vehicle to maintain equilibrium during transition, a wingpropeller system that represents one quarter of the vehicle was used. Wind tunnel tests were performed on a single 6 in, two bladed propeller attached to a wing surface with an aspect ratio of 2.75. Tests were performed first on an isolated propeller at various shaft angles, RPM and forward flight velocity. The combined wing-propeller system was then tested to study key differences in force production with and without the wing surface. Finally trim analysis based on force measurements was performed to extract operating conditions for trimmed flight at flight modes including transition from hover to forward flight. Due to the effect of the wing on propeller slipstream and vice versa, the vertical force was greater, and the horizontal force was lower than that produced by the isolated propeller. A maximum speed of 11 m/s at 0 deg shaft angle was obtained with a cruise speed of about 6 m/s with a lift requirement of 0.6 N. The cruise power was 1.5 W which was one-third of hover power. Free flight testing successfully showed feasibility of vehicle to achieve equilibrium transition flight.

This paper describes the systematic experimental and computational (CFD) studies performed to investigate the performance of a small-scale vertical axis wind turbine (VAWT) utilizing dynamic blade pitching. A VAWT prototype with a simplified blade pitch mechanism was designed, built and tested in the wind tunnel to understand the role of pitch kinematics in turbine aerodynamic efficiency. The ability of the present pitch mechanism to change the blade pitch phasing instantaneously in order to adapt to changes in wind direction is the key to maximizing power extraction in urban environments. A CFD model was developed and the model predictions correlated extremely well with test data. The validated CFD model was used to develop a fundamental understanding of the physics of power extraction of such a turbine. Both experimental and CFD studies showed that the turbine efficiency is a strong function of blade pitching amplitude, with the highest efficiency occurring around ±20◦ to ±25◦ amplitude. The optimum tip speed ratio (TSR) depends on the blade pitch kinematics, and it decreased with increasing pitch amplitude for the symmetric blade pitching case. CFD analysis showed that the blade extracts all the power in the frontal half of the circular trajectory, however, it loses power into the flow in the rear half. One key reason for this being the large virtual camber and incidence induced by the flow curvature effects, which slightly enhances the power extraction in the frontal half, but increases the power loss in the rear half. Fixed-pitch turbine investigated in the present study also showed lower efficiency compared to the variable pitch turbines owing to the massive blade stall in the rear half, caused by the large angle of attack and high reverse camber. Maximum achievable CP of the turbine increases with higher Reynolds numbers, however, the fundamental flow physics remains relatively same irrespective of the operating Reynolds number. This study clearly indicates the potential for major improvements in VAWT performance with novel blade kinematics, lower chord/radius ratio, and using cambered blades.

This paper describes the autonomous hover-capability of a cycloidal-rotor aircraft (Cyclocopter) at Micro Air Vehicle (MAV) scale. Cycloidal rotors have a horizontal axis of rotation in which blades span parallel to the axis and perpendicular to direction of flight. The hybrid Cyclocopter is configured with two cycloidal rotors and a conventional tail rotor. Due to combination of independent RPM control of each motor and thrust vectoring of the cycloidal rotors, the Cyclocopter is highly maneuverable. An eective control strategy was developed to successfully demonstrate pitch, roll, and yaw capabilities. Based on the unique features of cycloidal rotors, the Cyclocopter has a potential to exceed performance of conventional rotary MAVs. Through a feedback controls system implemented by an onboard processor equipped with tri-axial gyros, accelerometer, microprocessor, and wireless communication components, the Cyclocopter achieved autonomous stable hover.

The comprehensive rotorcraft analysis UMARC was modified to calculate the capabilities of leading-edge slats for helicopter performance improvement in moderate to high-speed forward flight. Leading-edge slats were shown to enhance the maximum rotor thrust by 15-30% at advance ratios larger than 0.2 and reduce power requirements by 10-20% at high thrust levels. 20% span slats offered a good compromise between power reductions and adverse effect on vibratory hub loads. The rotor with leading-edge slats could be trimmed at a maximum forward speed that was about 20 knots greater than for the baseline rotor with no slats.

Rotor hover performance can be enhanced by enclosing it in a shroud. This paper describes the design and testing of a quad rotor vehicle with a shroud structure incorporated, and the resulting improvements in aerodynamic efficiency over a conventional micro quad rotor are demonstrated. Parametric studies were conducted to study the effect of number of blades and root collective on shrouded rotor performance. Comparison of a 2.6” shrouded rotor with an optimized unshrouded rotor showed a 30% improvement in power loading at a power input of 2W. To minimize shroud weight, the structure consisted of a machined foam diffuser with carbon fiber inlets attached. A fully integrated proof of concept micro air vehicle (MAV) was designed with a gross weight of about 100 grams and a shroud structure weight of 12 grams. The vehicle had maximum dimensions of 6” 6”. Successful hover flight testing of the vehicle was achieved using an onboard attitude feedback controller. Comparison with data from another study (9.5” shrouded rotor) revealed that the hover performance of the shrouded rotor scales satisfactorily with size. The shrouded rotor concept appears feasible in improving the performance of a conventional micro quad rotor.

Effects of trailing-edge flap gaps on rotor performance are investigated using a high fidelity, coupled computational fluid dynamics (CFD) – computational structural dynamics (CSD) analysis. Both integral flap (the flap is an integral part of the blade such that there are no physical gaps at the flap ends) and discrete flap (the flap is a separate entity with physical gaps in the spanwise and chordwise directions) are examined on an UH-60A rotor at high speed forward flight condition. A novel grid deformation scheme based on the Delaunay graph mapping is developed and implemented to allow the CFD modeling of the gaps with minimal distortion of mesh around the flap gap regions. This method offers an alternative to the traditional approach of modeling such configurations using overset meshes. The simulation results show that the effectiveness of the flap is minimally lost with spanwise gaps – the penalty on rotor performance is of the order of 1% compared to the integral flap. On the other hand the chordwise gaps significantly degrade the benefits of active flap on rotor performance due to the flow penetration between the upper and lower surfaces of the flap.

A new solution for the SLAM problem is presented which makes use of a scan matching algorithm, and does not rely on bayesian filters. The virtual map is represented in the form of an occupancy grid, which stores laser scans based on the estimated position. The occupancy grid is scanned by means of ray casting to get a scan of the virtual world, called ”virtual scan”. The virtual scan therefore contains data from all previously acquired laser measurements and hence serves as the best representation of the surroundings. New laser scans are matched against the virtual scan to get an estimate of the new position. The scan matching cost function is minimized via an adaptive direct search with boundary updating until convergence. The resulting method is model-free and can be applied to various platforms, including micro aerial vehicles that lack dynamic models. Experimental validation of the SLAM method is presented by mapping a typical office hallway environment with a closed loop, using a manually driven platform and a laser range scanner. The mapping results are highly accurate and the loop closure area appears to be seamless, in spite of no loop closure algorithms and no post-mapping correction processes.

This paper describes a series of systematic experimental studies conducted to optimize the performance of a MAV scale cycloidal rotor. Parametric studies were performed to study the effects of blade airfoil section, blade pitching amplitude, asymmetric pitching, location of blade pitching axis and number of blades (at constant solidity) on the performance of the cycloidal rotor. Higher blade pitch angles were found to improve thrust and increase the power loading (thrust per unit power) of the cycloidal rotor. Asymmetric pitching, with a higher pitch angle at the top than at the bottom of the circular blade trajectory, produced better power loading. The optimum pitching axis location was observed to be around 25-35% of the blade chord. For a constant solidity, the rotor with lesser number of blades produced higher thrust and a 2-bladed rotor had the best power loading. Using the results from these parametric studies, an optimized cycloidal rotor with a significantly improved power loading over a conventional MAV rotor was developed.

This paper predicts, analyzes, and validates main rotor airloads, structural loads, and swashplate servo loads in a prescribed high-g pull-up maneuver. A multibody finite element structural model is coupled to a transient lifting-line aerodynamic model. The structural model includes a swashplate model to calculate servo loads. The lifting-line model combines airfoil tables, Weissinger-L near wake, time marching free wake, and a semi-empirical dynamic stall model. The maneuver is the Army/NASA UH60A Airloads Program Flight Counter 11029. The primary objective of this paper is to isolate the eects of structural dynamics, free wake, dynamic stall, and pitch control angles, to determine the key loads mechanisms in this flight. The structural loads are first calculated using airloads measured in flight. The measured airloads are then replaced with a lifting-line coupled analysis ‐ that is ideally suited to isolate the eects of free wake and dynamic stall. It is concluded that the maneuver is almost entirely dominated by stall with little or no wake induced eect on blade loads ‐ even though the wake cuts through the disk twice during the maneuver. At the peak of the maneuver, almost 75% of the operating envelope of a typical airfoil lies beyond stall. The mechanism of dynamic stall, in the analysis, consists of multiple cycles within a wide disk area. The peak-to-peak structural loads prediction from the lifting-line analysis show an under-prediction of 10%‐20% in flap and chord bending moments and 50% in torsion loads. The errors stem from the prediction of 4 and 5/rev stall loads. Swashplate dynamics appears to have a significant impact on the servo loads - unlike in level flight ‐ with more than 50% variation in peak loads.

In this study, a mixed-based approach that combines the classical stiffness and flexibility methods is developed to analyze the rotating dynamic behavior of thin-walled, tapered composite blades with built-in twist angles. The analysis model includes the effects of elastic couplings, shell wall thickness, transverse shear couplings, torsion warping, and constrained warping. The resulting theory describes the beam kinematics in terms of the axial, flap and lag bending, flap and lag shear, torsion, and torsion-warping deformations. The equations of motion for the beam are obtained by using Hamilton's principle. The analysis is validated against experimental test data and detailed finite element analysis results for pretwisted and tapered beams with different material distributions and cross-section shapes. Good correlation is achieved for various configuration cases considered in this study. The influence of pretwist, taper and fiber orientation angles on the structural dynamic response of rotating composite beams are investigated.

An aeroelastic analysis for hover-capable, bio-inspired flapping wings was developed. Bio-Inspired flapping-pitching mechanisms reported in literature, usually operate in oil or water at very low flapping frequencies. In contrast, the mechanism used in this study operated in air, at relatively high frequencies. Because of this, the light-weight and highly flexible wings used in this study had significant aeroelastic eects which need to be investigated. In this paper, a finite element based structural analysis of the wing is described, along with an unsteady aerodynamic analysis based on indicial functions. Analytical results for aluminum-mylar and composite-mylar wings were compared with the thrust measured on the bioinspired flapping-pitching mechanism. A parametric study was also carried out to investigate the eect of wing stiness on thrust generation. This parametric study showed that a large increase in thrust could be obtained by suitably adjusting the wing stiness.

Helicopter rotor blades and aircraft wings can be modeled as thin-walled beams that are non-uniform and pretwisted. The non-uniformities are due to the taper of the beam arising from the fact that the thickness-to-chord ratios of the airfoils are higher at the root and lower at the tip. In addition, the wall thicknesses can be higher at the root than at the tip. Taper introduces interactions between axial strain and transverse shear strains. Pre-twist introduces a coupling between axial strain and torsion even for beams made of isotropic materials. Additional couplings can be introduced by the use of fiber reinforced composites. In order to fully model a thin-walled composite beam that has taper and pre-twist, it is necessary to model the interactions between the deformations due to axial, bending, transverse shear, and (St.Venant and Vlasov) torsion. Apart from some finite element codes, published analyses that capture these interactions are limited. In the present paper, a variationally consistent method is proposed that can analyze beams with the following features: thin- and thick-walled beams, open- or closed cross-sections, single- or multiple cells, pre-twist, taper in the cross-section profile and in wall thickness, and incorporate elastically coupled composite materials. The analysis considers the wall of the beam as a composite Reissner - Mindlin shell and includes the influence of elastic couplings and bending and transverse shear stiffnesses of the shell wall. The global beam model includes axial deformation, bending in two planes, St. Venant and constrained torsion, and transverse shear in two planes. No assumptions are made on cross-section deformability and the influences of hoop stress and hoop moment are included in the analysis. A mixed method based on a modified Hellinger-Reissner variational principle is used to model the influence of transverse shear stresses and cross-section deformation. The advantage of the present method is that it does not make any ad hoc assumptions on the cross-section warpings due to torsion and transverse shear effects. These warping functions are generated as part of the analysis, leading to a more accurate model. The analysis is applicable to thick- and thin-walled beams with open- or closed cross-sections. The cross-section stiffness matrix is obtained in closed-form. As an example, the cross-section matrix for a single-cell composite thin-walled beam is presented. It is shown that, for elastically-coupled composite beams, the transverse shear stiffness terms can influence the other responses and that ignoring these terms could lead to erroneous results. The analysis is validated by comparisons with results from elasticity theory, finite element analyses, and experiments for a variety of cross sections. These examples examine the influence of pre-twist, taper, and crosssection deformations on the axial, bending, transverse shear and torsion response of: uniform and pre-twisted solid beams with rectangular and elliptic cross-sections, uniform and tapered I-beams with pre-twist and single- and twocell box beams with pre-twist and elastic couplings.

This paper presents an aeroelastic analysis for hover-capable, biomimetic flapping wings. Biomimetic flapping-pitching mechanisms reported in literature, usually operate in oil or water at very low flapping frequencies. In contrast, the mechanism used in this study operates in air, at relatively high frequencies. Because of this, the light-weight and highly flexible wings used in this study may have significant aeroelastic eects which need to be investigated. In this paper, a finite element based structural analysis of the wing is described, alongwith an unsteady aerodynamic analysis based on indicial functions. The analysis is validated with experimental data available in literature, and also with experimental tests conducted on the biomimetic flapping-pitching mechanism. Results for both elastic and rigid wing analyses are compared with the thrust measured on the biomimetic flapping-pitching mechanism.

A helicopter trailing edge flap with a piezoceramic actuated servotab is investigated experimentally. The system is designed to amplify the small stroke of the piezoceramic using a combination of mechanical, electrical and aerodynamic factors. Flap deflections occur in response to the aer odynamic hinge moments created by the oscillations of the servotab. A model of the dynamic behavior of the flap -actuator -servotab is developed for a fixed airfoil section. The design of the experimental two dimensional wing model is described, al ong with modifications to reduce the inertia of the system. Results of tests conducted in an open -jet wind tunnel are presented, including flap response to servotab deflections, and the amplitude of servotab response to voltage excitation. Correlation of the data with the model prediction s are discussed.

Results were obtained for several high frequency tests conducted on biomimetic, flapping-pitching wings. The wing mass was found to have a significant influence on the maximum frequency of the mechanism because of a high inertial power requirement. All the wings tested showed a decrease in thrust at high frequencies. In contrast, for a wing held at 90◦ pitch angle, flapping in a horizontal stroke plane with passive pitching caused by aerodynamic and inertial forces, the thrust was found to be larger. To study the effect of passive pitching, the biomimetic flapping mechanism was modified with a passive torsion spring on the flapping shaft. Results of some tests conducted with different wings and different torsion spring stiffnesses are shown. A soft torsion spring led to a greater range of pitch variation and produced more thrust at slightly lower power than with the stiff torsion spring. Some flow visualization images have also been obtained using the passive pitching wings.

The objective of this study is to demonstrate the concept of active trailing-edge flaps as primary rotor control and vibration reduction devices for a typical full-scale helicopter. A comprehensive rotorcraft analysis based on UMARC is developed to obtain the results for the swashplateless rotor. A parametric study of various key design variables involved in the trailing-edge flap design was carried out. An optimal design of trailing-edge flap system that provides effective control authority within the complete flight envelope as well as minimum actuation requirements was achieved. Trailing-edge flaps demonstrated the capability of performing both primary control and active vibration control functions. At a high forward speed (advance ratio of 0.32), the 4/rev vertical force, and roll and pitch moments at hub are successfully eliminated (by about 90%), and the 4/rev in-plane hub forces are reduced by more than 40%. The half peak-to-peak value of the trailing-edge flap deflection for primary control is about 7.1 ‐ and additional 4.7 ‐ is required for active vibration control.

Several different Mach scale model rotors with composite tailored blades were designed for reduction of vibratory hub loads. Comprehensive aeroelastic analysis shows that a non-uniform spanwise segmented composite flap-bending/torsion coupling distribution can lead to reduced vibratory hub loads in forward flight. Two Mach scale rotor sets with pre-twisted composite tailored blades (baseline and optimum coupling) were fabricated. Static bench-top tests and dynamic tests were carried out to validate the structural analysis of these composite blades. The rotors were tested up to the nominal rotor speed of 2300rpm (with tip Mach number 0.65) on the hover stand. Predicted rotor thrust matched the test data in hover.

This paper presents the material modeling and experimental characterization of SMA rod and tube actuators undergoing torsional deformations. The investigation of the torsional characteristics of SMAs was carried out in order to gain fundamental understanding of the behavior of torsional actuators for actively altering the twist distribution of a tiltrotor blade between hover and forward flight. A torsional model involving the extension of the one-dimensional formulation of Brinson’s model for prediction of SMA behavior is presented. The model is shown to have good correlation with experiment over a wide range of thermomechanical conditions including constant temperature torque-angle and actuation tests against torsional springs. Experimental phenomena associated with the torsional behavior of the SMA actuator such as the effects of heat treatment, twist rate and loading pattern are examined, and limitations of the current model pointed out. It is shown that the theoretical model demonstrates good agreement with the experimental data over a wide thermo-mechanical range of conditions.

The lack of polar symmetry in two-bladed rotors leads to equations of motion with periodic coefficients in axial flight, which is contrary to three or more bladed rotors that result in constant coefficient equations. With periodic coefficients, the analysis becomes involved, as a result very few studies have been directed towards the analysis of two-bladed rotors. In this paper, the aeroelastic stability of two-bladed proprotor/pylon/wing combinations is examined in high speed axial flight. Several parametric studies are carried out to illustrate the special nature of two-bladed proprotors and to better understand the mechanism of whirl-flutter in such rotors. The wing beam bending mode for two-bladed rotors is found to be stable over the range of parameters examined, a behaviour very different from three-bladed rotors. Also, the wing torsion mode exhibits a new type of instability similar to a wing torsional divergence scouring at I/rev frequency. This type of behaviour is not seen in three and more bladed rotors. The interaction between wing chordwise bending and torsion modes is found to be much greater in the case of two-bladed rotors and, over the range of parameters considered, these two modes govern the stability of the system.

This paper formulates a new control methodology for vibration reduction at the rotor hub by controlling trailing edge flaps. The novelty of the proposed methodology lies in its ability to control each rotor blade separately and optimally taking into account blade-to-blade dissimilarities, while using exclusively fixed frame measurements. The controller is formulated in the time domain, and adaptively generates in real time individual control inputs to the trailing edge flaps to achieve vibration reduction. Numerical simulations using a hingeless rotor model show that the controller generates control inputs to each blade, taking blade dissimilarities into account, and successfully minimizes vertical hub vibrations.

This paper presents the design, fabrication and open-loop testing of an active rotor system with trailing-edge flaps for individual blade control of helicopter vibration. First, the University of Maryland Advanced Rotorcraft Code (UMARC) was used to size the trailing-edge flap and to determine the flap deflection requirements for vibration suppression in the wind tunnel. Next an analytic model for the coupled actuator-flap-rotor system was used to design a multilayer piezoelectric bender configuration that was capable of meeting the flap deflection requirements. Based on the design study, a matched set of six Machscaled active rotor blades were fabricated in-house. The 4-bladed rotor model was tested in the open-loop mode in hover and then in the wind tunnel using a Bell-412 Mach-scaled hub. Flap deflections of ±4 to ±6 deg were recorded in the 1-5/rev frequency range at the Mach-scaled operating speed of 1800 RPM. The flap deflection increased to ±23 deg at 8/rev due to actuator resonant amplification. Rotor collective pitch and advance ratio were found to have negligible impact of actuator performance. The maximum control effectiveness was observed close to the blade flapbending and torsion natural frequencies. For 3/rev actuator excitation, oscillatory thrust levels of up to ±36 Ibs (60% steady rotor thrust at 6 deg collective) were recorded, thereby demonstrating the open-loop control authority of the actuator-flap at Mach-scale.

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 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.