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The rotor blade double-swept-tip geometry can effectively improve the aerodynamic and aeroacoustic behaviors of a rotor. However, the aeroelastic or aeromechanical behaviors of a double-swept rotor...

This paper describes the effects of a composite coupled blade spar on the performance of a slowed-rotation-frequency helicopter rotor in high-speed edgewise flight. Antisymmetric composite coupling...

A comprehensive analysis for evaluating performance and vibratory loads of a coaxial helicopter rotor is developed and validated against existing experimental data. The model is extended from the b...

The present study is in response to increased interest towards assessing the feasibility of a small-scale autonomous helicopter (gross weight less than 1 kg) for Martian exploration. An autonomous rotorcraft may be ideally suited for such an application because of its unique advantages, which include the ability to take off/land vertically on harsh terrain, and greater speed, range, and field of view, when compared to a traditional surface rover. The atmospheric conditions on Mars present a unique set of design challenges. Even though the Martian gravity is only about 38% of Earth’s gravity, the Martian average atmospheric density is about 70 times lower than Earth’s atmospheric density. Therefore, the rotors would be operating at extremely low Reynolds numbers, even lower than 5000 for a small-scale helicopter. However, the Mach number will be significantly higher (M>0.4) because of the higher tip speed required (due to lower density) and because of the fact that the speed of sound on Mars is only about ...

This paper describes a fundamental experimental study, which involved systematic performance and flowfield measurements (PIV) to understand and optimize the hover performance of a MAV-scale helicopter rotor operating at Reynolds numbers lower than 30,000. The rotor parameters that were varied include blade airfoil profile, blade chord, number of blades, blade twist, planform taper and winglets at blade tip. Blade airfoil section had a significant impact on the hover efficiency and among the large number of airfoil sections tested, the ones with the lower thickness to chord ratios and moderate camber (4.5% to 6.5%) produced the highest rotor hover figure of merit. Increasing the solidity of the rotor by increasing the number blades (with constant blade chord) had minimal effect on efficiency; whereas, increasing the solidity by increasing blade chord for a 2-bladed rotor, significantly improved hover efficiency. Moderate blade twist (−10° to −20°) and large planform taper (larger than 0.5) marginally improved rotor efficiency. Rotor blades with small winglets (height ≈ 6% of rotor radius) at the tip also improved hover performance. While using winglets, the flowfield measurements showed a diffused tip vortex, which could reduce the induced aerodynamic losses. Spanwise lift distribution obtained using sectional bound circulation computed from the measured flowfield correlated well with the load cell measurements. The optimal rotor designed based on the understanding gained from the present study produced a figure of merit of 0.67, which is the highest value of FM ever reported in the literature for micro-rotors operating at these low Reynolds numbers.

This paper describes the systematic performance measurements conducted to understand the role of rotor geometry and blade pitching kinematics on the performance of a microscale cycloidal rotor. Key geometric parameters that were investigated include rotor radius, blade span, chord, and blade planform. Because of the flow curvature effects, the cycloidal-rotor performance was a strong function of the chord/radius ratio. The optimum chord/radius ratios were extremely high, around 0.5–0.8, depending on the blade pitching amplitude. Cycloidal rotors with shorter blade spans had higher power loading (thrust/power), especially at lower pitching amplitudes. Increasing the solidity of the rotor by increasing the blade chord, while keeping the number of blades constant, produced large improvements in power loading. Blade planform shape did not have a significant impact, even though trapezoidal blades with a moderate taper ratio were slightly better than rectangular blades. On the blade kinematics side, higher blad...

Effects of trailing-edge flap gaps on rotor performance are investigated using a high-fidelity coupled computational fluid dynamics computational structural dynamics 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 span and chord directions) are examined on an UH-60A rotor at high-speed forward-flight conditions. A novel grid deformation scheme based on the Delaunay graph mapping is developed and implemented to allow the computational fluid dynamics 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 affected with span gaps; the penalty on rotor performance is of the order of 1% compared to the integral flap. On the other hand, the chord...

The first part of this paper compares three rotor blade structural dynamic formulations: a finite element formulation with modal reduction, a full finite element formulation without modal reduction, and a multibody-based full finite element formulation for arbitrary large deformations. The second part of this paper studies the effect of swashplate dynamics on blade loads and servo-actuator loads. In all cases, measured airloads, damper force, and control pitch angles from the UH-60A flight tests are used to predict and analyze the structural loads. In the first part, the emphasis is on the validation of a multibody formulation, which is first verified with analytical solutions for beams undergoing hypothetical large deformations (elastica), then validated with the Princeton beam large deformation tests, and then finally used to predict the UH-60A structural loads. Two flight conditions are considered: a high-speed, high-vibration flight and a highly loaded dynamic stall flight. Predictions from the multibody analysis are compared with the full finite element and finite element based modal methods. It is observed that the predicted blade loads do not show any significant difference between the three formulations. In the second part, the four-bladed multibody rotor model is coupled to a swashplate-servo model to predict servo loads and to study the effect of swashplate dynamics on blade loads. It is observed that the higher frequencies of servo loads, 8/rev and 12/rev for this rotor, require modeling the swashplate dynamics. The low-frequency component, which is a dominant 4/rev load for this rotor, is less affected by swashplate dynamics and is determined primarily by the accuracy of the 3, 4, and 5/rev pitch-link loads. The 3-5/rev pitch-link loads, and in general the structural loads on the rotor blade, are not affected by swashplate dynamics.

A control method is proposed to reduce vibrations in helicopters using active trailing-edge flaps on the rotor blades. Each blade is controlled independently, taking into account possible blade dissimilarities. The method consists of performing simultaneous system identification and closed-loop control at each time step. For the system identification, different inputs are applied to each blade, and the relationship between the individual blade inputs and the resulting loads in the fixed frame is estimated on-line, assuming a linear-time-periodic model of the helicopter. Closed-loop tests are conducted using a four-bladed Mach-scaled rotor with piezobender trailing-edge flaps. The rotor model is fitted on a bearingless model-scale hub and tested in the Glenn L. Martin wind tunnel. These tests demonstrate the controller's ability to account for blade dissimilarities and generate different optimal inputs for each blade. The 1 and 4/rev components of fixed frame loads are reduced individually by 50 and 60%. Simultaneous reduction of 1 and 4/rev components is also demonstrated (43% reduction). However, vibration increases are noted for some nontarget hub loads.

A computational fluid dynamics (CFD) model is coupled with a computational structural dynamics (CSD) model to improve prediction of helicopter rotor vibratory loads in high-speed flight. The two key problems of articulated rotor aeromechanics in high-speed flight-advancing blade lift phase, and underprediction of pitch link load-are satisfactorily resolved for the UH-60A rotor. The physics of aerodynamics and structural dynamics is first isolated from the coupled aeroelastic problem. The structural and aerodynamic models are validated separately using the UH-60A Airloads Program data. The key improvement provided by CFD over a lifting-line aerodynamic model is explained. The fundamental mechanisms behind rotor vibration at high speed are identified as: 1) large elastic twist deformations and 2) inboard wake interaction. The large twist deformations are driven by transonic pitching moments at the outboard stations. CFD captures 3-dimensional unsteady pitching moments at the outboard stations accurately. CFD/CSD coupling improves elastic twist deformations via accurate pitching moments and captures the vibratory lift harmonics correctly. At the outboard stations (86.5% radius out), the vibratory lift is dominated by elastic twist. At the inboard stations (67.5% and 77.5% radius), a refined wake model is necessary in addition to accurate twist. The peak-to-peak pitch link load and lower harmonic waveform are accurately captured. Discrepancies for higher harmonic torsion loads remain unresolved even with measured airloads. The predicted flap-bending moments show a phase shift of about 10 deg over the entire rotor azimuth. This error stems from 1, 2, and 3/rev lift. The 1/rev lift is unaffected by CFD/CSD coupling. The 2 and 3/rev lift are significantly improved but do not fully resolve the 2 and 3/rev bending moment error.

The objective of this study is to demonstrate the concept of active trailing-edge flaps as primary flight control and vibration reduction devices for a typical full-scale helicopter. A comprehensive rotorcraft analysis based on UMARC was developed to analyze 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 a trailing-edge flap system that provides effective control authority within the complete range of advance ratios as well as minimum actuation requirements was achieved. Trailing-edge flaps demonstrated the capability of performing both primary flight 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 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 flight control is 7.1 deg, and an additional 4.7 deg is required for active vibration control.

The effect of actuator dynamics on a helicopter rotor with trailing-edge flaps for vibration control is investigated. Trailing-edge flap, actuator, and elastic rotor blade equations of motion are formulated using Hamilton’s variational principle. The coupled nonlinear, periodic equations are solved using finite elements in space and time. The baseline correlation study is based on wind-tunnel test data for a typical five-bladed bearingless rotor system. Good agreement is seen for the blade flap bending, chord bending, and torsion moments. It is shown that actuator dynamics cannot be neglected for a trailing-edge flap system with torsionally soft actuators. The parametric study performed using both coupled flap/actuator model and prescribed flap motion model indicated that the placement of trailing-edge flaps at 78% radius resulted in minimum flap input for this rotor. The vibration reduction level and trend are close between the predictions of both models at different forward speeds. Control inputs predicted by the coupled model show less sensitivity to the forward speed than that of prescribed model.

A helicopter primary control system with trailing-edge flaps was investigated numerically for its potential to replace a conventional swashplate system. Eliminating the swashplate and associated control system can lead to significant reductions in weight, drag, and cost and an improvement of rotor performance. A comprehensive rotorcraft analysis was developed for analyzing the swashplateless rotor configuration and was implemented to examine the actuation requirements for rotor primary control with trailing-edge flaps. A multicyclic controller was implemented with the swashplateless rotor analysis, and the feasibility of trailing-edge flap performing both primary control and active vibration control was examined. Flap control inputs of a swashplateless rotor are presented at several advance ratios. With optimal selection of blade collective pitch index angle, the flap was shown to be able to trim the rotor with moderate flap inputs. Simulations of flaps performing both primary control and active vibration control were carried out, with the conclusion that trailing-edge flaps are capable of trimming the rotor and minimizing vibratory rotor hub loads simultaneously

A coupled rotor-fuselage vibration analysis is formulated to analyze the effect of rotor system faults on fuselage vibrations and rotor-blade displacements in both hover and forward-flight conditions. Two groups of rotor-system faults, adjustable and nonadjustable component faults, are modeled. Results are presented for an SH-60 helicopter and are compared with available flight-test data. A detailed aeroelastic analysis is carried out where each rotor blade is modeled individually such that rotor dissimilarity can be considered. A fuselage NASTRAN model is incorporated and coupled with the rotor model analysis. The centrifugally tuned, hub-mounted bifilar vibration absorbers are also modeled, and the coupled rotor-fuselage-bifilar nonlinear equations are solved using finite element method in space and time

The design, analysis, and testing of an improved Shape Memory Alloy (SMA)-based tracking tab actuator is described in this paper. The goal of the actuator is to provide in-flight tracking capability for a helicopter rotor in order to minimize 1/rev vibrations due to rotor dissimilarities. Previous SMA-based actuator designs demonstrated the potential for in-flight rotor tracking but admitted drawbacks that led to inconsistent operation under air-loads. The current research builds upon the existing knowledge base and addresses the challenges encountered in previous designs. The objective is to achieve a deflection of 58 with an accuracy of 0:18 under realistic aerodynamic loading conditions. The present actuation concept is based on the bidirectional motion of a pair of antagonistic SMA wires, with a passive friction brake to lock the tab position. A theoretical model of the actuator was developed based on Brinson’s thermomechanical model. The model was used to predict the behavior of the actuator under external loading and applied as a design tool to identify optimal actuator parameters. The actuator was integrated into a NACA 0012 12 in. chord blade section and tested in an open-jet wind tunnel at speeds of up to 120 ft/s (0.107 M) and at angles of attack up to 158. Closed-loop tracking was implemented using a PID controller with gains selected by Ziegler-Nichols tuning. The improved SMA actuator meets the project goals by achieving repeatable tab deflection of up to 58with an average accuracy of 0.058. Position hold under power-off conditions and a duty cycle of 20 cycles/h were also demonstrated.

In this work, a closed-form analysis is performed for the structural response of coupled composite blades with multi-cell sections. The analytical model includes the effects of shell wall thickness, transverse shear, torsion warping and constrained warping. The mixed beam approach based on Reissner's semi-complementary energy functional is used to derive the beam force-displacement relations. The theory is validated against experimental test data and other analytical results for coupled composite beams and blades with single-cell box-sections and two-cell airfoils. Correlation of the present method with experimental results and detailed finite element results is found to be very good.

This paper describes the design and development of a piezoelectric-hydraulic hybrid actuator operating at a high pumping frequency. The actuator is envisaged as a potential actuator for a trailing edge flap on a full scale smart rotor system. While recent research efforts based on the same concept have investigated actuators with large piezoelectric stacks operating at a relatively low pumping frequency, the goal of the present work is to investigate the behavior of the actuator at a low volumetric displacement and high pumping frequency. Preliminary design of the actuator system is carried out, and the dependence of the performance of the actuator on various system parameters is identified. This enables the optimum selection of geometric parameters and piezostack characteristics for a given external load. Challenges to achieving high pumping frequencies were identified and solutions were implemented. The actuator was driven by two piezostacks, of a total length of 36 mm and cross-sectional area 100 mm 2 . The actuator was tested up to a pumping frequency of 1 kHz, developing a maximum no-load velocity of 1.2 in/s and a blocked force of 35 lb in the unidirectional output mode. Bidirectional output performance was also measured, by incorporating a 4-way valve in the hydraulic circuit. At a frequency of 5 Hz, a no-load output displacement with an amplitude 32 mils was measured.

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

The active twist rotor investigated in this research is a derivative of the previously developed smart-activeblade-tip (SABT) rotor. On the SABT rotor, the blade tips are independently pitched, with respect to the main blade. A novel piezo-induced bending ‐torsion coupled actuator beam, located spanwise in the hollow midcell of the main rotor blade, is used to actuate the blade tip. When the blade tip is locked to the main blade, the actuator beam twists the entire blade. A Mach scale rotor with a 1.542 m diameter was hover tested, open loop, to evaluate the control authority for vibration reduction. A nonrotating tip-twist amplitude of 0.78 deg was achieved (below resonance, 150 V rms ). Analysis indicates that no signie cant twist actuation degradation is expected at full rotor speed. In 2000-rpm hover (tip Mach 0.47 ), at 8-deg collective, and for a single blade actuation of 150 V rms at 1, 2, and 3 per revolution, respectively, the measured oscillatory thrust coefe cients are 1 :4 £ 10 3 , 0:55 £ 10 3 , and 0:7 £ 10 3 . The corresponding e nite element model estimated blade twist amplitudes are 0.8, 1.0, and 1.9 deg. Good correlation of the predicted and measured rotor thrust was achieved up to 3 per revolution. The hover test demonstrates the potential of the active twist rotor system using an internal actuation beam and warrants further research for a dedicated next-generation model-scale design and full-scale feasibility study.

Active rotorblade tips offer an alternative approach to the challenge of main rotor active vibration control. The tips are pitched with respect to the main blade via a piezo-driven bending-torsion coupled actuator beam that runs down the length of the blade. A Vlasov based, specialized one-dimensional finite beam element is developed to model the rotating actuator beam and is validated with the free-vibration and static forced response of 4:1 and 2:1 aspect ratio, bending-torsion coupled, active and passive plates. A one-eighth scale, reduced tip-speed rotor model (tip Mach 0.26), incorporating the bending-torsion actuator beam, has been previously hover tested (open loop). In these tests, blade tip deflections of the order of 2° (half peak-to-peak) were achieved at 2, 3, 4, 5/rev with corresponding dynamic vertical blade root shear variations of the order of 10-20% of the nominal blade lift at 8° collective (CT/σ = 0.07). The test results are used to validate a coupled actuator and elastic rotorblade model. The correlation of the predicted active blade-tip pitch deflections and the experimental data is within 20%. The predicted values for the active vertical root shears are within the same margin for 4° and 6° collective.

This paper presents the wind tunnel testing of a four-bladed Mach-scaled rotor model with piezoelectric bender actuated trailing-edge flaps. Correctly phased, this flap motion can be used for the active suppression of vibratory hub loads. First, the University of Maryland Advanced Rotorcraft Code was used to conduct a parametric study to determine the optimal flap sizing and location as well as the required deflection amplitudes. Next, a simplified rotor analysis which explicitly models the actuator dynamics was used to design a multi-layer actuator configuration that is capable of achieving the required flap deflections. Based on the above design studies, a matched set of four Mach-scaled rotor blades with piezo-bender actuated trailing-edge flaps were fabricated in-house. This rotor model was operated using a one-seventh scale Bell-412 Mach-scaled rotor hub. Finally, the rotor system was tested in forward flight in the Glenn L Martin wind tunnel. These tests consisted of open-loop single-frequency tests at different rotor speeds and collective settings. Some preliminary closed-loop tests using a neural network control algorithm were also conducted. Presented at the American Helicopter Society 56th Annual Forum, Virginia Beach, VA. Copyright 2000 by the American Helicopter Society, Inc. All rights reserved.

Acomprehensivevibrationanalysisofacoupledrotor/fuselagesystemforatwo-bladedteeteringrotorusinge nite elementmethodsinspaceandtimeisdevelopedthatincorporatesconsistentrotor/fuselagestructural,aerodynamic, and inertial couplings and a modern free-wake model. Coupled nonlinear periodic blade and fuselage equations are transformed to the modal space and solved simultaneously. The elastic line airframe model of the AH-1G helicopter is integrated into the elastic rotor e nite element model. Analytical predictions of rotor control angles, blade loads, hub forces, and vibration are compared with AH-1G operation load survey test data. The blade loads predicted by the present analysis show generally fair agreement with the e ight test data. Calculated 2 and 4 per revolution vertical vibration levels at the pilot seat show fair correlation with the e ight test, but the predicted 2 per revolution lateral vibration level is higher than the measurement, particularly at high advance ratios. Modeling of pylon e exibility is essential in the two-bladed teetering rotor vibration analysis. Ree ned aerodynamics such as free wake and unsteady aerodynamics have an important role in the prediction of vibration. Nomenclature eg = blade center-of-gravity offset from the elastic axis lu = undersling m = blade mass per unit length pfe = temporal fuselage elastic modal displacement vector pfr = temporal fuselage rigid modal displacement vector R = blade radius u;v;w = blade elastic displacements in the axial, lag, and e ap directions x = longitudinal coordinate P x f ; P yf ; P f = fuselage translational velocities

Smart structure activated trailing edge flaps are capable of actively altering the aerodynamic loads on rotor blades. Coupled 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 a potential of lower weight penalties incurred by servo actuation methods. This paper explores new, robust individual blade control (IBC) methodologies for vibration suppression using a piezoactuated trailing edge flap. The controllers employ a single hidden layer neural network, learning in real time, to adaptively cancel the effects of periodic aerodynamic loads on the blades, greatly attenuating the resulting vibrations. Both collocated and noncollocated sensor/actuator pairs are considered. Proofs of the stability and convergence of the proposed neurocontrol strategies are provided, and numerical simulation results for a one-eighth Froude scale blade model are given which demonstrate that the controller can nearly eliminate the blade vibration arising from a wide variety of unknown, periodic disturbance sources.

The aeroelastic stability of a composite hingeless rotor with elastic couplings is investigated in hover. A two-bladed, Froude-scaled, soft in-plane hingeless rotor model with various isotropic and graphite/ epoxy composite flexures was designed and then tested on a hover stand. The rotor with an isotropic flexure was tested first to examine the experimental procedures. The effects of elastic flapwise bending-torsion and chordwise bending-torsion structural coupling on the aeroelastic stability were then 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 (University of Maryland Advanced Rotorcraft Code). The introduction of negative chordwise bending-torsion coupling has a stabilizing effect on the lag damping for positive collective pitch angles. Predictions agree with measured data.

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 efficient. Flight conditions modeled are helicopter (hover and forward flight), conversion, and airplane modes of operation. Predictions from this analysis for trim controls, response, loads, and aeroelastic stability correlate well with flight test data and predictions from other established analyses.

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.

This paper presents an analytical model and validation tests of a Froude scaled rotor featuring piezoelectric bender (bimorph) actuated trailing-edge flaps for active vibration suppression. The analytical model for the coupled bimorph actuator trailing-edge flap dynamic response in the rotating environment 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 is selected to maximize flap performance in the rotating environment. The bimorph clamping is improved to prevent actuator slippage under high centrifugal loads. The analytical model is validated by carrying out a series of bench tests, vacuum chamber tests and hover tests. In hover, flap deflections of ±6 degrees at 4/rev flap excitation are 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.

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 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 degrees are attainable by the current actuator configurations in forward flight. Although these amplitudes were less than the target value (I to 2 degrees 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.

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.

Applications of smart structures technology to various physical systems are primarily focused on actively controlling vibration, performance, noise, and stability. Applications range from space systems to fixed-wing and rotary-wing aircraft, automotive, civil structures, marine systems, machine tools, and medical devices. Early applications of smart structures technology were focused toward space systems to actively control vibration of large space structures [1] as well as for precision pointing in space (e.g., telescope, and mirrors [2]). The scope and potential of smart structures applications for aeronautical systems have subsequently expanded. Embedded or surface-bonded smart material actuators on an airplane wing or helicopter blade can induce alteration of twist/camber of airfoil (shape change), which in turn can cause variation of lift distribution and may help to control static and dynamic aeroelastic problems. For fixed-wing aircraft, applications cover active control of flutter [3, 4, 5, 6, 7], static divergence [8, 9], panel flutter [10], performance enhancement [11], and interior structure-borne noise [12]. Compared to fixed-wing aircraft, helicopters appear to show the most potential for a major payoff with the application of smart structures technology. Given the broad scope of smart structures applications, developments in the field of rotorcraft are highlighted in a subsequent section. Although most current applications are focused on the minimization of helicopter vibration, there are other potential applications such as interior/exterior noise reduction, aerodynamic performance enhancement that includes stall alleviation, aeromechanical stability augmentation, rotor tracking, handling qualities improvement, rotor head health monitoring, and rotor primary controls implementation (e.g., swashplateless rotors) [13].

The objective of this research was to investigate the potential of rotor morphing to improve the performance of baseline helicopter configurations. The morphing parameters that have been considered in this report are variable rotor speed and variable rotor radius. Both of these parameters are considered in steady state only, so that transients that would be introduced during blade morphing are not considered. Additional morphing concepts including variable twist, chord, camber and tip shape could not be completed due to time constraints. The comprehensive rotorcraft analysis code, UMARC, was modified to accommodate steady variations. The structural model is a full, non-linear, finite element analysis formulation. Each blade is modeled as an articulated, flexible beam with coincident flap and lag hinges. The aerodynamic model is a Weissinger-L type lifting-surface model coupled with 2D airfoil tables, unsteady effects are captured using the Leishman-Beddoes model, and a time accurate free wake model, captures the wake.

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

Significant advances in the development of a Froude scaled helicopter rotor model featuring a trailing-edge flap driven by piezoceramic bimorph actuators for active vibration suppression are discussed. A quasisteady aerodynamic analysis used to determine flap size and actuator requirements is presented. The block force and stroke of the current actuators are evaluated using two theories and compared with experimental results. The dynamic performance of the actuator as well as the actuator - flap assembly is examined. Earlier hover tests showed severe degradation in flap deflections with increasing rotor speed, and flap deflections were too small to be effectively utilized for significant vibration control. To investigate the causes of the performance degradation, new blades are constructed and tested in vacuo to isolate the effects of centrifugal loading on the actuator - flap system. A beam model of the piezo bimorph including propeller moment effects is formulated to better illustrate the physical mechanisms affecting the system in a rotating environment. The cause of the reduced deflections is traced to frictional forces created at the junction where the flap is supported during rotation of the blades. The use of a thrust bearing was found to alleviate this problem and subsequent hover tests showed a dramatic increase in flap deflection at high excitation frequencies.

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 2-bladed bearingless rotor model was built where each blade is embedded with banks of piezoelectric actuators at 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 piezo-actuation of blade twist. To assess the performance of the piezo-actuators in rotation, hover tests were conducted where accelerometers embedded in the blades were used to resolve the tip twist amplitudes. Although the magnitudes of blade twist attained in this experiment were small, it is expected that future models can be built with improved performance.

Sensitivity derivatives of blade loads and aeroelastic stability of a helicopter rotor in forward flight are calculated as an integral part of a basic aeroelastic analysis using a direct analytical approach. Design variables include nonstructural mass and its placement, chordwise offset of blade center of gravity and aerodynamic center from the elastic axis, blade bending stiffnesses (flap, lag, torsion), and tip geometry (sweep, anhedral, pretwist and planform taper). By means of a sensitivity study, the importance of different design variables on oscillatory hub loads and damping of blade modes is examined. Aeroelastic and sensitivity analyses of the rotor based on a finite element method in space and time are linked with automated optimization algorithms to perform optimization studies of rotor blades. Optimum design solutions, calculated for a four-bladed, soft-inplane hingeless rotor achieved a reduction of 25-60 percent of all 4/rev loads.

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 discusses the systematic experimental and vehicle design/development studies conducted at the University of Maryland which culminated in the development of the first flying Cyclocopter in the history. Cyclocopter is a novel Vertical Take-Off and Landing (VTOL) aircraft, which utilizes cycloidalrotors (cyclorotors), a revolutionary horizontal axis propulsion concept, that has many advantages such as higher aerodynamic efficiency, maneuverability and high-speed forward flight capability when compared to a conventional helicopter rotor. The experimental studies included a detailed parametric study to understand the effect of rotor geometry and blade kinematics on cyclorotor hover performance. Based on the experimental results, higher blade pitch angles were found to improve thrust and increase the power loading (thrust per unit power) of the cyclorotor. Asymmetric pitching with higher pitch angle at the top than at the bottom produced better power loading. The chordwise optimum pitching axis location was observed to be around 25-35% of the blade chord. Because of the flow curvature effects, the cycloidal rotor performance was a strong function of the chord/radius ratio. The optimum chord/radius ratios were extremely high, around 0.5-0.8, depending on the blade pitching amplitude. A flow field investigation was also conducted using Particle Image Velocimetry (PIV) to unravel the physics behind thrust production of a cyclorotor. PIV studies indicated evidence of a stall delay as well as possible increases in lift on the blades from the presence of a leading edge vortex. The goal of all these studies was to understand and optimize the performance of a micro-scale cyclorotor so that it could be used in a flying vehicle. An optimized cyclorotor was used to develop a 200 gram cyclocopter capable of autonomous stable hover using an onboard feedback controller.

To overcome many of the problems associated with conventional articulated rotor systems, new rotor systems are being contemplated. In this paper, the state-of-art technology of advanced rotor systems is assessed. Advanced rotors include hingeless, bearingless, composite, circulation control, tilt and advanced geometry rotors. The paper reviews mathematical modelling, analysis methods, past and recent developments, potential limitations and future research needs in each system. Also, the potential of incorporation of structural optimization methodology and smart structures technology in rotors to improve the efficiency and capabilities of rotorcraft is discussed.

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.

An analytical study, based on an advanced Higher Harmonic Control (HHC) analysis for helicopter rotor systems, is carried out to investigate HHC application for rotor performance enhancement. The effects of HHC on stall characteristics of rotor and blade pitch-link loads when the system is configured to suppress vibration are also examined. For vibration control, simulated results indicate that HHC may promote early blade stall. Effects of blade torsion frequencies on HHC performance are moderate, and torsionally stiff blades require less actuator power than torsionally soft blades. For rotor performance improvement, a 3 to 5 percent reduction in rotor shaft power can be achieved with 2 deg of two-per-rev blade pitch control.

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 structural optimization analysis of a hingeless helicopter rotor is developed and applied with the objective of reducing oscillatory hub loads in forward flight. The aeroelastic analysis of the rotor is based on a finite element method in space and time and is linked with automated optimization algorithms. Two types of structural blade representations are used: a generic stiffness-distribution beam and a single-cell, thin-walled beam. For the generic beam representation the design variables are nonstructural mass and its placement, chordwise center of gravity offset from the elastic axis, and structural stiffness (flap, lag, and torsion). For the second type of structural representation, spar width, height, and thickness are used as design variables instead of blade stiffness. Constraints on frequency placement, autorotational inertia, and aeroelastic stability of the blade are included. Sensitivity derivatives are efficiently calculated using a direct analytical approach, with a resulting 80% reduction in total CPU time required to obtain an optimum solution compared with a commonly used finite-difference approach. Optimum solutions resulted in reductions of 25-77% for the generic blade, and 30-50% for the box-beam blade relative to baseline values of the objective function that was comprised of all six components of hub load.

Symmetric and antisymmetric layup graphite-epoxy composite beams with thin-walled rectangular cross sections are fabricated using an autoclave molding technique and tested under bending, torsional, and extensional loads. The bending slope and elastic twist at a station are measured using an optical system, and the results correlated with predicted values from a simple beam analysis as well as a refined finite element analysis. A symmetric ply layup results in bending-twist coupling whereas an antisymmetric layup causes extension-twist coupling. Simple analytical results with plane-stress assumption agree better with measured data as well as finite element predictions than with plane-strain assumption. For symmetric layup beams, the bending-induced twist and torsion-induced bending slope are predicted satisfactorily using simple analytical solution. Correlations with measured data, however, are generally improved using a finite element solution. For antisymmetric beams, axial force-induced twist is predicted satisfactorily by both methods.

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.

A servo-tab mechanism was designed for a full-scale rotor blade section, and tested in an open-jet wind tunnel. The intent was to use aerodynamic forces and moments to amplify the stroke and force provided by a compact piezoelectric bender actuator. The design was constrained by the stroke-force characteristics of the piezoceramic benders previously developed for model scale applications. The flap chord was chosen as 25% of the blade chord, and the tab chord was 5% of the blade chord. The flap response to the aerodynamic moment generated by the servotab was predicted using blade element momentum and thin airfoil theory. Testing in an open-jet wind tunnel explored the response of the flap in both steady and oscillating tab conditions.