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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 twenty-first century could be called the 'Multifunctional Materials Age'. The inspiration for multifunctional materials comes from nature, and therefore these are often referred to as bio-inspired materials. Bio-inspired materials encompass smart materials and structures, multifunctional materials and nano-structured materials. This is a dawn of revolutionary materials that may provide a 'quantum jump' in performance and multi-capability. This book focuses on smart materials, structures and systems, which are also referred to as intelligent, adaptive, active, sensory and metamorphic. The purpose of these materials from the perspective of smart systems is their ability to minimize life-cycle cost and/or expand the performance envelope. The ultimate goal is to develop biologically inspired multifunctional materials with the capability to adapt their structural characteristics (stiffness, damping, viscosity, etc.) as required, monitor their health condition, perform self-diagnosis and self-repair, morph their shape and undergo significant controlled motion over a wide range of operating conditions.

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.

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.

A viscoelastic solid model, comprising of a combination of linear and non-linear springs and dashpots, is developed to represent an elastomeric damper. A non-linear constitutive differential equation is derived to characterize the damper behaviour in the time domain. A system identification method is presented to determine the spring - dashpot parameters (coefficients of the constitutive equation) from experimental data. The model is able to predict the non-linear amplitude-dependent behaviour of elastomeric dampers under single as well as dual-frequency excitations. A `two-level implicit - implicit' scheme is developed for the integration of the non-linear damper model into a structural dynamic analysis. With increase in amplitude of excitation, softening behaviour of the lead spring results in lesser motion in the Kelvin chain, lower damping levels, and slower decay of initial perturbations. The baseline model is augmented with additional series and parallel non-linear springs to represent the reduction in damping (G'') at very small amplitudes, and the occurrence of limit cycle oscillations.

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 the development of electromechanical actuators based on the concept of mechanical amplification with piezo and electrostrictive stacks as drivers to achieve high force and high displacement actuation. The actuators were designed for two different applications. The first actuator, with an piezo stack, was developed to actuate a `Flaperon' which consisted of a small movable surface to trip the boundary layer, located on the top surface of a wing model with span and chord of 12 in each and of NACA 0012 airfoil. The actuator was designed to produce 8 lbs of force with peak displacement of 10 mils at a maximum frequency of 40 Hz. The second actuator, with an electrostrictive stack as a driver, was designed to move a leading edge droop flap hinged at 25% chord of a wing model with span of 8 in, chord of 4 in and a VR-12 airfoil. This actuator was designed to produce 8.5 lbs of force with peak displacement of 10 mils at a maximum frequency of 45 Hz. Experiments were performed on both stacks to evaluate their important characteristics such as block force, free displacement and stiffness, that were essential in the design of the actuators. The results showed that the block force obtainable from a piezo stack was higher and that of an electrostrictive stack was lower than that specified by the respective manufacturers, while the free displacements are about the same. The dynamic response of the actuators over a frequency range of 33 Hz was evaluated. Results showed that 7 lbs of actuator force was obtainable in both cases, with the flaperon actuator producing 15 mils of dynamic displacement at 15 Hz and the droop flap actuator producing about 6 mils of displacement at 16 Hz. The results were inconclusive beyond 16 Hz due to the setup resonance. The droop flap actuator did not achieve the desired performance because the design calculations were based on the block force listed by the manufacturer which was about 20% higher than the measured value. This led to the conclusion that before the design process begins, the performance of the stack alone should be carefully measured in order to achieve the required performance. Thus, a simple actuator based on a mechanical amplification concept could be effectively designed to produce high force and high displacements.

This paper develops one-dimensional pure bending, coupled bending and extension, and combined bending, extension and torsion models of isotropic beams with induced-strain actuation. A finite thickness adhesive layer between the crystal and beam is included to incorporate shear lag effects. Experimental tests evaluate the accuracy and limitations of the models. The bending and coupled bending and extension models show acceptable correlation with static test results whereas the combined extension, bearing, torsion model predicts the system behavior poorly and needs refinement.

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 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.

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.

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.

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.

An improved aeroelastic formulation for the advanced geometry blades, involving variable sweep, droop, pretwist, and planform, is presented. The blade is modeled as a series of arbitrarily-oriented elastic segments with each segment divided into finite elements. Inter-element compatibility relations governing non-Eulerian moderate rotations of the finite elements are also presented. Fuselage dynamic interaction with the advanced geometry blades is included in the formulation. The nonlinear partial differential equations of motion are discretized in space and time using Hamilton's principle. Selective results are presented in hover and forward flight. Results indicate that sweep, and droop in particular, can have a strong influence on both the rotor aeroelastic stability and the rotorcraft aeromechanical stability. Droop can be considerably stabilizing. Sweep increases the blade torsional loads, but is not detrimental to flap and lag vibratory loads.

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 article presents a free-vibration analysis of coupled composite I-beams with couplings under rotation. A linear analysis based upon Vlasov theory was developed to obtain coupled flap-lag-torsion equations of motion for I-beams made out of general composite laminates. Constrained warping and transverse shear effects were included. Free-vibration characteristics were obtained by solving these equations using Galerkin's method. In order to validate the theory, graphite-epoxy and kevlar-epoxy I-beams with bending-torsion coupling were fabricated using an autoclave molding technique and tested in an in vacuo rotor test facility for their vibration characteristics. Induced-strain actuation by piezoceramic elements was used to excite the I-beams in the rotating frame. Strain gauges were used to measure the response of these beams over a range of rotational speeds up to 1000 rpm. Natural frequencies and strain mode shapes were determined by carrying out signal analysis using a spectrum analyzer. Good correlation between theory and experiment was achieved. About 600% increase in torsional frequency due to constrained warping occurs for graphite-epoxy beams with a slenderness ratio of 18. Nomenclature A = cross-sectional area of blade b, h = semiwidth and semiheight of I-beam Eh Et — Young's moduli in principal directions of plies of beam Glt - shear modulus of plies in principal plane Ixx, Iyy = blade cross-sectional moment of inertia about x and y axes, respectively K] - stiffness matrix for beam Kf] = flap bending stiffness matrix Kft] — flap bending-torsion coupling stiffness matrix K] = lag bending stiffness matrix K!t] - lag bending-torsion coupling stiffness matrix [Kt] = torsion stiffness matrix KA = effective polar radius of gyration of blade cross-section V(/vv + IXX)IA

The aeromechanical stability, including air resonance in hover, air resonance in forward flight, and ground resonance, of a helicopter with elastically tailored composite rotor blades is investigated. Five soft-inplane hingeless rotor configurations, featuring elastic pitch-lag, pitch-flap and extension-torsion couplings, are analyzed. Elastic couplings introduced through tailored composite blade spars can have a powerful effect on both air and ground resonance behavior. Elastic pitch-flap couplings (positive and negative) strongly affect body, rotor and dynamic inflow modes. Air resonance stability is diminished by elastic pitch-flap couplings in hover and forwrad flight. Negative pitch-lag elastic coupling has a stabilizing effect on the regressive lag mode in hover and forward flight. The negative pitch-lag coupling has a detrimental effect on ground resonance stability. Extension-torsion elastic coupling (blade pitch decreases due to tension) decreases regressive lag mode stability in both airborne and ground contact conditions. Increasing thrust levels has a beneficial influence on ground resonance stability for rotors with pitch-flap and extension-torsion coupling and is only marginally effective in improving stability of rotors with pitch-lag coupling.

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%

Piezoceramic actuators and sensors have found wide application in a variety of fields such as aerospace, automotive and biomedical. The coupled electro-mechanical behavior exhibited by these materials enables them to be used either as actuators or as sensors. The material has a high bandwidth, can be manufactured in a number of different shapes, and can be easily integrated into structures. This chapter focuses on the behavior of piezoceramic actuators and sensors. The electro-mechanical constitutive equations are discussed, and the static as well as dynamic response of the material to applied electric fields is described. Nonlinear phenomena are also described, and the power consumption of piezoceramic actuators is explained. The use of the material as a strain sensor is also described in detail, along with the required signal conditioning electronics and calibration factors. This information is intended to serve as a reference for the design of a structure incorporating piezoceramic materials. Keywords: piezoceramics; actuators; sensors; power consumption; free strain; nonlinear effects

This paper describes the aeroelastic model to predict the blade loads and the average thrust of a micro air vehicle (MAV) scale cycloidal rotor. The analysis was performed using two approaches, one using a non-linear FEM analysis for moderately flexible blades and second using a multibody based largedeformation analysis (especially applicable for extremely flexible blades). An unsteady aerodynamic model is included in the analysis with two different inflow models, uniform inflow and a double-multiple streamtube inflow model. For the cycloidal rotors using moderately flexible blades, the aeroelastic analysis was able to predict the average thrust with sufficient accuracy over a wide 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 this may be because of the overprediction of blade deformations. The analysis clearly showed that the reason for the reduction in the thrust producing capability of the cycloidal rotor with blade flexibility may be attributed to the large nose-down elastic twisting of the blades in the upper half cylindrical section which is not compensated by a nose-up pitching in the lower half section. The inclusion of the actual blade pitch kinematics and unsteady aerodynamics was found crucial in the accurate lateral force prediction.

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.

Because of nonlinear and complex interactions involved with the dynamic and aeroelastic stability of a rotor system, detailed and accurate knowledge of the stability characteristics are essential to prove design safety and to validate theoretical analysis. The task of determining these characteristics for a rotor from tests becomes complicated because of the presence of substantial amplitudes of the undamped harmonics, the possibility of close modes, and the complexity associated with the excitation of the modes in the rotating environment. Time domain modal parameter estimation methods are very useful for damping estimation when there are close modes, but they are very sensitive to noise. Subspace methods substantially improve the time domain estimates for noisy data, but they require higher computation time. A simple method is developed that retains the low variance estimation property of the subspace methods, but which is comparable in computation cost to methods that do not use the subspace approach. Its performance is evaluated for multi-output and single-output implementations and compared to the standard sparse time domain method. It is found that the modified method is more accurate in terms of the bias and standard deviation of the damping estimates, and it is faster when the number of modes is much less than the order of the data matrix.

This paper presents a theoretical-cum-experimental study of the free vibration characteristics of thin-walled composite box beams with bending-twist and extension-twist coupling under rotating conditions. The governing equations in generalized displacements were derived using a Newtonian approach. The composite structural model in the derivation used a solid-section approach and contained transverse shear-related couplings and appropriate cross-section warping. The free vibration characteristics of composite box beams were determined by the Galerkin method. In order to validate the theory, glass-epoxy, kevlar-epoxy and graphite-epoxy symmetric and antisymmetric box beams were fabricated using an autoclave molding technique, and tested in an in-vacuo rotor test facility for their vibration characteristics. Beam excitation in the rotating condition was effected by means of induced-strai n actuation with the help of piezoceramic bending elements. Strain gages were used to measure the response of the first three modes over a range of rotational speeds up to 1000 rpm. It was determined that the experimental frequencies and mode shapes correlated satisfactorily with the theoretical results. It is shown also that bending-shear coupling influences the flexural vibration frequencies of antisymmetric box beams significantly. Extension-shear coupling, on the other hand, does not influence the flexural-torsion vibration frequencies of symmetric box beams significantly.

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.

The structural behavior of coupled, thin-walled, composite beams of open as well as closed section was analyzed using Vlasov theory and then the results were validated by experiment. The analysis modeled the walls of beams as general composite laminates and accounted for the transverse shear deformation of the cross-section. The out-of-plane warping deformation of the cross-section was included implicitly in this formulation. In order to validate the analysis, graphite-epoxy beams of various cross-sections such as solid rectangular, I-section, single-cell rectangular and two-cell airfoil were fabricated and tested for their structural response under tip bending, torsional and extensional loads. Specialized bending-torsion and extension-torsion couplings were introduced in these beams using proper ply lay-ups. Good correlation between theoretical and experimental results was achieved. Transverse-shear-related couplings were found to influence the structural response of open- as well as closed-section beams. For blades with hygrothermally stable lay-ups, bending-transverse shear coupling increased the bending flexibility by about 50%. The in-plane-bending coupling stiffness [ B ] of the walls of the beam generally influenced the structural response of the beams quite significantly; this effect was expecially large for I-beams. The influence of constraining the warping deformation was found to be substantial on the structural response of open-section beams as compared to closed-section beams. A 630% increase in the torsional stiffness due to constrained warping was noticed for graphite-epoxy I-beams of slenderness ratio 30. The feasibility of achieving the desired levels of bending-torsion and extension-torsion couplings in two-cell rotor blades was demonstrated.

We envision situational awareness developed through warfighters deployment of a system of diverse mobile, communicating platforms that cooperate to provide full coverage of interior and exterior spaces. The goal of the ARL-MAST Center on Microsystem Mechanics is to perform the fundamental research that will enable flying and ambulating platforms to achieve the required mobility for the proposed missions and environments. In this paper the fundamental issues and challenges associated with achieving this goal will be discussed.

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.