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The paper focuses on understanding the mechanism of force production on a hover-capable flapping-wing system by using a combination of direct force measurements and flowfield studies. The experiments were conducted in air at a Reynolds number of approximately 25,000 (based on mean chord and maximum tip speed), which is the typical operating regime of small flapping-wing micro air vehicles. The forces and moments were measured using a miniature six-component force transducer installed at the wing root. The wing was flapped in air and vacuum at the same frequency and wing kinematics, and the vacuum forces were subtracted from the total forces to obtain the pure aerodynamic forces. Flow visualization and particle image velocimetry were used to characterize the formation, strength, and structure of the leading-edge vortex on the flapping wing. A rapid increase in wing lift coefficient was associated with the growth of the leading-edge vortex and a progressive reduction in lift coefficient with the convection ...

This paper describes the systematic experimental and computational studies performed to investigate the performance of a small-scale vertical-axis wind turbine using dynamic blade pitching. A vertical-axis wind turbine 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. A computational fluid dynamics model was developed, and the model predictions correlated well with test data. Both experimental and computational fluid dynamics studies showed that the turbine efficiency is a strong function of blade pitching amplitude, with the highest efficiency occurring around ±20 to ±25 deg amplitude. The optimum tip-speed ratio depends on the blade pitch kinematics, and it decreases with increasing pitch amplitude for the symmetric blade pitching case. A computational fluid dynamics analysis showed that the blade extracted all the power in the frontal half of the circular trajectory; however...

This paper describes the systematic experimental and computational studies performed to obtain a fundamental understanding of the physics behind the lift and thrust production of a cycloidal rotor (cyclorotor) in forward flight for a unique blade pitching kinematics. The flow curvature effect (virtual camber and incidence due to the curvilinear flow) was identified to be a key factor affecting the lift, thrust, and power of the cyclorotor in forward flight. The experimental study involved systematic testing of a micro air vehicle-scale cyclorotor in an open-jet wind tunnel using a custom built three-component balance. The key parameters varied include rotor chord/radius ratio and blade pitching axis location because these two parameters have a strong impact on flow curvature effects. Because of the virtual camber/incidence effects and the differences in the aerodynamic velocities around the azimuth, the blades produce a small downward lift when they operate in the upper half of the circular trajectory and...

This paper describes the aeroelastic model to predict the blade loads and the average thrust of amicro-air-vehiclescale cycloidal rotor. The analysis was performed using two approaches: one using a second-order nonlinear beam finite element method analysis for moderately flexible blades and a second using a multibody-based largedeformation analysis (especially applicable for extremely flexible blades) incorporating a geometrically exact beam model. An unsteady aerodynamic model is included in the analysis with two different inflow models: single streamtube and double-multiple streamtube inflowmodels. For the cycloidal rotors usingmoderately flexible blades, the aeroelastic analysis was able to predict the average thrust with sufficient accuracy over awide range of rotational speeds, pitching amplitudes, and number of blades. However, for the extremely flexible blades, the thrust was underpredicted at higher rotational speeds, and thismaybebecause of the overprediction of blade deformations. The analysis clearly showed that the reason for the reduction in the thrust-producing capability of the cycloidal rotorwith blade flexibility may be attributed to the large nosedown elastic twisting of the blades in the upper half cylindrical section, which is not compensated by a noseup pitching in the lower half-section. The inclusion of the actual blade pitch kinematics, unsteady aerodynamics, and flow curvature effects was found crucial in the accurate lateral force prediction.

This paper addresses the aerodynamics of insect-based, biomimetic, flapping wings in hover. An experimental apparatus, with a biomimetic flapping mechanism, was used to measure the thrust generated by a number of wing designs at different wing pitch settings. To quantify the large inertial loads acting on the wings, vacuum chamber tests were conducted. Results were obtained for several high-frequency tests conducted on lightweight aluminum and composite 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-deg 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. The lightweight and highly flexible wings used in this study had significant aeroelastic effects which need to be investigated. A finite element based structural analysis of the wing is described, along with an unsteady aerodynamic analysis based on indicial functions. The analysis was 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.

On-blade smart structure actuators 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 with lower weight penalties than the previous actuation methods and without the bandwidth constraints. We cover the development and testing of a new, robust individual blade control methodology for rotor vibration suppression using piezoactuated trailing-edge flaps and active twist tip rotors. The controller uses a neural network to learn to actuate the trailing-edge flap, thus adaptively suppressing the blade or hub vibrations. In this application, no offline training is performed. Instead, a neural network is used in real time to command adaptively the actuator deflections, thus reducing vibrations. Closed-loop experimental tests with piezoactuated-scale rotor systems were conducted on the University of Maryland hover test stand. The results include two different Mach-scale smart rotor systems (trailing-edge flaps and active tip twist) that were controlled by the same adaptive neurocontrol algorithm. These tests demonstrate the controller's robust ability to learn to control successfully the rotor vibrations with no a priori information about the blade/actuator structure or the aerodynamic loading.

Thedevelopmentand open-loop hovertesting ofa smartrotormodelwithtrailing-edgee apsforindividualblade control of helicopter vibration are presented. First, the University of Maryland Advanced Rotorcraft Code was used to sizethetrailing-edgee apandto determinethee ap-dee ectionrequirementsforvibration suppression in the wind tunnel. Next an analytic model for the coupled actuator-e ap-rotor system was used to develop a multilayer piezoelectric bender cone guration that was capable of meeting the e ap-dee ection requirements. Based on this study, a matched set of six smart rotor blades were fabricated in-house. The four-bladed smart rotor model was tested in the open-loop mode in hover using a Bell-412 Mach-scaled hub. Flap dee ections of §4 to §6 deg were recorded in the 1 ‐5/rev frequency range at the model operating speed of 1800 rpm. The e ap dee ection increased to §23 deg at 8/rev because of actuator resonant amplie cation. Rotor collective pitch was found to have negligible impact on actuator performance. The maximum control effectiveness was observed close to the blade e atwise bending and torsion natural frequencies. For 3/rev actuator excitation oscillatory thrust levels of up to §10 lb (45 N) (60% steady rotor blade thrust at 6-deg collective ) were recorded, thereby demonstrating the open-loop control authority of the actuator-e ap system. I. Introduction T HEhelicopterspendsa largeportionofitsoperation in forward e ight, resulting in an aerodynamic asymmetry between the advancingandretreatingsidesoftherotordisk.Asaresultofthis,the rotor e owe eld is extremely complex and can include transonic e ow ontheadvancingbladetip,dynamicstallontheretreatingsideofthe disk, highly yawed and reversed e ows, and blade-wake interactions with tip vortices from preceding blades. This highly complex and nonsteady aerodynamic environment and the dynamic response of the long and e exible rotor blades result in large vibratory forces, which are e ltered through the hub to the fuselage. Typically for an N-bladed rotor, the dominant N i1, N and N C1/rev blade loads are transmitted to the fuselage as an N/rev forcing. Reduction in vibration and noise levels of the helicopter yield benee ts in terms of improved passenger comfort, reduced crew fatigue, improved community acceptance, and increased fatigue life of structuralcomponents. Currently, passive isolators and absorbers are routinely used to reduce vibration. However these devices cause signie cant weight penalties (up to 3% of gross weight ) and rapidly degrade in performance away from the tuned e ight condition. In contrast, active vibration reduction schemes offer the promise of achieving reliable vibration reduction over a wide range of operating conditions with lower weight penalty than conventional passive methods. There are two types of rotating frame active control systems: higher harmonic control (HHC) and individual blade control (IBC). HHC involves excitation of the swashplate at N/rev. The major disadvantage of HHC is the high actuation power required to pitch the rotor blades particularly in extreme e ight conditions and large weight penalties associated with the hydraulic actuators. Another important limitation is that HHC is limited to N/rev excitation of the swashplate. However, for performance improvement, dynamic stall alleviation, acoustic control, and alleviation of gusts and ma

A refined structural model based on a mixed force and displacement method is proposed for the analysis of composite rotor blades with elastic couplings. The present formulation allows the modeling of either open-section or closed-section blades of arbitrary section shape, stacking sequence, and end restraint effects. The theory accounts for the effect of elastic couplings, shell wall thickness, section warping, warping restraint, and transverse shear deformations. A semicomplementary energy functional is used to derive, in a variationally consistent manner, the beam force-displacement relations. Bending and torsion related warpings and shear correction factors are obtained in closed form as part of the analysis. The resulting first-order shear deformation theory (Timoshenko) describes the beam kinematics in terms of the axial, flap and lag bending, flap and lag shear, twist, and torsion-warping deformations. The theory is validated against experimental data and other finite element results for graphite-epoxy composite beams of various cross sections such as I sections, box sections, and two-cell airfoils. Good correlation is achieved for all of the test examples. The influence of wall thickness and transverse shear on the static beam response is also investigated. Wall thickness effects are shown to become significant when the thickness-to-depth ratio of the beam reaches around 20%. The slenderness ratio has a significant effect on the transverse shear behavior of the beam, especially for beams with low slenderness ratios. It is also shown that the layup angle has a nonnegligible effect on the transverse shear behavior of the beam.

In this paper, a refined one-dimensional beam formulation based on a mixed variational approach has been developed for static and free vibration analyses of both rotating and non-rotating composite blades with elastic coupling. The present approach allows the modeling of either open-section or closed-section blades of arbitrary section shape, stacking sequence, and end restraint effects. The theory accounts for the effects of elastic coupling, shell-wall thickness, warping, warping restraint, and transverse shear deformations. A semi-complimentary energy function is used to derive, in a variationally consistent manner, the beam force-displacement relations. The free vibration analysis is validated against experimental test data and other analytical results for composite beams of various cross-sections such as box-section, rectangular solid section, and I-section. Good correlation is achieved for all the test examples. The influence of modeling refinements on the free vibration characteristics of composite blades through the current mixed formulation is also investigated. The wall thickness effects are shown to become significant when the thickness-to-depth ratio of the beam approaches approximately 30%. The effects are seen to be more at higher modes.

A comprehensive physics-based model of the helicopter rotor in forward flight is used to analyze the impact of selected faults on rotor system behavior. The rotor model is based on finite elements in space and time. The helicopter rotor model is used to develop a neural network-based damage detection methodology. Simulated data from the rotor system are contaminated with noise and used to train a feedforward neural network using backpropogation learning. Cases considered for training and testing the neural network include both single and multiple faults on the damaged blade. Results show that the neural network can detect and quantify both single and multiple faults on the blade from noise-contaminated simulated vibration and blade response test data. For accurate estimation of type and extent of damages, it is important to train the neural networks with noise-contaminated response data.

The experimental results are presented of hover test with a one-eighth dynamically scaled (Froude scale) helicopter rotor model with embedded piezoceramic elements as actuators to suppress vibrations. A 6-ft-dlam two-bladed bearingless rotor model was built with banks of piezoelectric torsional actuators capable of manipulating blade twist at harmonics of the rotational speed. The twist performances of several rotor blade configurations were investigated using accelerometers embedded in the blade tip. The change in oscillatory rotor lift due to plezoactuation was measured by a hub balance. Experimental results show that linear twist distributions of up to 0.6 deg, resulting in increases of up to 10% 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 (1-2 deg), it is expected that partial reduction of hub vibration can be achieved with the current smart rotors.

Aeroelastic and sensitivity analyses of the rotor based on a finite element in space and time are linked to an automated optimization algorithm to perform optimization studies for a four-bladed, soft in-plane composite rotor consisting of a two-cell thin-walled beam. The design variables used in this study are the ply angles of the laminated walls of the composite beam. The objective function minimizes the 4/rev hub loads, with constraints on blade frequencies and aeroelastic stability in forward flight. Optimum design solutions show a reduction in the objective function of about 20% due to elastic stiffnesses and an additional 13% due to composite couplings. Starting from an initially infeasible design, the optimum design solution with negative lag bending-torsion coupling results in an increase in lag mode damping of about 140% compared to the baseline layup. c* c* c* ^d-> ^/> ^m CT c

This paper presents an analytical-experimental study of using magnetostrictive actuators in conjunction with an extension-torsion coupled composite tube to actuate a rotor blade trailing-edge flap to actively control helicopter vibration. Thin-walled beam analysis based on Vlasov theory was used to predict the induced twist and extension in a composite tube with magnetostrictive actuation. To validate the analysis, extension-torsion coupled Kevlar®-epoxy tubes or different ply lay-ups were fabricated using an autoclave molding technique. They tubes were first tested under static mechanical loads, and tip twist and axial extension were measured by means or a laser optical system and strain gages, respectively. Good correlation between theory and experiment was achieved. Subsequently, these composite tubes were tested under magnetostrictive actuation. The [11] 2 Kevlar-epoxy tube system generated the maximum twist, 0.19 deg in tension and 0.20 deg in compression. The Kevlar-epoxy systems showed good correlation between measured and predicted twist values. Finally, alternate actuator concepts for these tubes, specifically piezoelectric stacks and electrostrictive actuators, were examined, and a piezoelectric stack actuator was round to induce much larger force and twist (approximately 3 times that created by the magnetostrictive actuator/tube system)

This paper presents an analytical-experimental investigation on structural modeling of coupled composite beams with distributed induced-strain actuators. Analysis based on Vlasov theory is developed to include distributed piezoelectric actuators, either surface mounted or embedded. Salient features of composite open-section beam analysis, like constrained warping and transverse shear deformation, were included. Induced-strain actuation was introduced in the constitutive relations of plate segment of the open-section composite beams. To evaluate the analytical predictions, several bending-torsion and extension-torsion coupled graphite-epoxy solid beams were fabricated using an autoclave molding technique. These were surface mounted with piezoelectric actuators. The actuators were excited to produce local bending moment and axial force on the beam, and the structural response was measured in terms of bending slope, induced twist, and surface strain. Good correlation between analysis and experiment was achieved. Because of the existence of a chordwise actuator moment, the induced twist of bending-torsi on coupled beams was significantly influenced by including the chordwise curvature of the plate segment of beam in the formulation. For [45]i4 solid beams, the chordwise bending of the plate segment of beam was found to increase the tip twist by about

The aeroelastic response, blade and hub loads, and shaft-fixed aeroelastic stability are investigated for a helicopter with elastically tailored composite rotor blades. A finite element analysis including nonclassical effects such as transverse shear, torsion related warping, and in-plane elasticity is integrated with the University of Margland Advanced Rotorcraft Code. The 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%

This paper presents an analytical-cum-experimental study of the structural response of composite rotor blades with elastic couplings. Vlasov theory is expanded to analyze two-cell composite rotor blades made out of general composite laminates including the transverse shear deformation of the crosssection. Variation of shear stiffness along the contour of the section is included in the warping function . In order to validate this analysis, two-cell graphite-epoxy composite blades with extension-torsion coupling were fabricated using matched-die molding technique. These blades were tested under tip bending and torsional loads, and their structural response in terms of bending slope and twist was measured with a laser optical system. Good correlation between theory and experiment is achieved. Axial force induced twist rate of the order of 0.2 degree per inch length can be realized in extensiontorsion coupled blades with a hygrothermally stable [20/-70]2s layup for potential applications in the design of tilt rotors.

This paper presents a theoretical-cum-experimental study on the static structural response of composite I-beams with elastic couplings. A Vlasov type linear theory is developed to analyze composite open section beams made out of general composite laminates, where the transverse shear deformation of the beam cross-section is included. In order to validate this analysis, graphite-epoxy and kevlar-epoxy symmetric I-beams were fabricated using an autoclave molding technique. The beams were tested under tip bending and torsional loads, and their structural response in terms of bending slope and twist measured with a laser optical system. Good correlation between theoretical and experimental results is achieved. A 630 percent increase in the torsional stiffness due to constrained warping is noticed for graphite-epoxy beams with slenderness ratio of 30. Also extension-twist coupling 'B16' of flanges of these I-beams increases the bending-torsion coupling stiffness of

A comprehensive rotor aeroelastic analysis based on finite element theory in space and time is coupled with a three-dimensional transonic small disturbance finite difference analysis to investigate three-dimensional aerodynamic effects on blade response and loads in forward flight. Each blade is assumed to be an elastic beam undergoing flap bending, lag bending, elastic twist, and axial deflections. The blade steady response is calculated from nonlinear periodic normal mode equations using a finite element in time scheme. For induced inflow distributions on the rotor disk, a free wake model is used. Dynamic stall and reverse flow effects are also included. Vehicle trim and rotor elastic response are calculated as one coupled solution using a modified Newton method. The blade loads and structural bending are calculated for two blade configurations: a straight-tip blade and a 30-deg swept-back tip blade. Calculated results are correlated with flight-test data obtained from the Gazelle helicopter (with a straight-tip blade) for two level flight speeds. Results then are calculated for this rotor with a swept-tip configuration and the effects of three-dimensi onal aerodynamics are assessed. Considerable three-dimensional aerodynamic effects are observed in the swept-tip blade.

A sensitivity study of blade stability in forward flight for a hingeless rotor with respect to design variables is carried out using a direct analytical method. Design variables include nonstructural mass distribution (spanwise and chord wise), chordwise offset of center of gravity, and blade bending stiffnesses (flap, lag, and torsion). The formulation for blade steady response is based on a finite-element method in space and time. The vehicle trim and blade steady response are calculated iteratively as one coupled solution using a modified Newton method (coupled trim analysis). Eigenvalues corresponding to different blade modes are calculated using Floquet transition matrix theory. The formulation for the derivatives of the eigenvalues with respect to the design variables is implemented using a direct analytical approach and constitutes an integral part of the regular stability analysis. For the calculation of the stability derivatives with respect to a total of 30 design variables, there is an 85% reduction in CPU time using the direct analytical approach compared to the frequently adopted finite-difference approach. A parametric study showed that nonstructural mass and chordwise blade e.g. offset of outboard elements, and lag bending stiffness of inboard elements, have powerful influence on blade stability.

A sensitivity study of blade response and oscillatory hub loads in forward flight for a hingeless rotor with respect to structural design variables is examined. Structural design variables include nonstructural mass distribution (spanwise and chord wise), chordwise offset of center of gravity, and blade bending stiffnesses (flap, lag, and torsion). The blade is discretized into a number of beam elements, and response equations are transformed into the normal mode space. The nonlinear, periodic, steady response of the blade is calculated using a finite element in time approach. Then, the vehicle trim and blade steady response are calculated iteratively as one coupled solution using a modified Newton method. The formulation for the derivatives of blade response and hub loads is implemented, using a direct analytical approach (chain rule differentiation), and forms an integral part of the basic response analysis. For calculation of the sensitivity derivatives, a 96% reduction of CPU time is achieved by using the direct analytical approach, compared with the finite-difference approach.

The flutter behavior of a thin panel resting on a linear elastic foundation is studied, using an analogy with the flutter of a panel of the same geometry without spring support. With quasi-steady aerodynamic loading, it is observed that there is a change not only in the flutter frequency, but also in the total damping of the system, due to the elastic foundation. Furthermore, it is found that for small values of aerodynamic damping and spring stiffness, the quasi-static approach is reasonable, whereas for high values of aerodynamic damping or of spring stiffness, quasi-static theory is inadequate. For high values of spring stiffness, the flutter mode of the panel exhibits traveling wave motion.

T nonlinear equations of motion for an isolated 3-deg flapping-lagging-feathering rotor blade are derived using Lagrange's equations for arbitrary large angular deflections. The aerodynamic forces and the tower interaction are not considered here. A consistent set of nonlinear equations is obtained by using nonlinear terms up to third order. The limit cycle analysis for forced oscillations and the principal parametric resonance of the flap-lag blade under a periodic gravity field are studied using the harmonic balance method. For a relatively small initial coning angle (about 9 deg), the nonlinearity becomes of the softening spring type, and large coupled responses are possible for rotational frequencies significantly lower than the lagging frequency.