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Instantaneous force and structural deformation experiments were performed on a flexible, structurally characterized, low-aspect-ratio representative flapping wing. A six-component force balance was...

A high-fidelity coupled computational fluid dynamics and comprehensive analysis solver is developed for helicopter flight on Mars. The objectives are accurate prediction of flight loads and fundame...

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 presents a methodology to analyze the coupled structural dynamic response of an elastic airframe and engines of a helicopter in response to main rotor hub loads. Transfer functions of individual components (airframe, engine, mount struts, and torque tube) are coupled together using a substructuring approach to obtain consistent coupled solutions of the entire system. Using this approach, a twin-engine, four-bladed helicopter is analyzed using NASTRAN-based models of the airframe and engines. This efficient substructuring approach is validated against the fully coupled NASTRAN model using forced response studies. Characteristics of the mount properties, i.e., the torque tube stiffness, and aft mount stiffness and damping are systematically varied to study their effect on the engine vibration response. The fore and aft mount element properties for minimizing the 8P engine response are identified without increasing 4P response. A compromise between 4P and 8P response is also identified from parametric studies of rear mount properties, using just three parameters to represent the design space. Using the substructuring approach presented here, future studies can be performed to rapidly match airframe characteristics with available engines at approximately 1000 times the speed of running a detailed finite element model (millions of degrees of freedom), without any reduction in accuracy.

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...

The increasing usage of low-Reynolds-number (10,000–100,000 tip Reynolds number) scale quadrotors for civilian and military applications provides the impetus for the development of reliable design ...

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...

This paper discusses the design, hardware and software methodology, and testing of an ultralight inertial navigation system (Embedded Lightweight Kinematic Autopilot-Revised (ELKA-R)) that can be used as a controller in a wide range of micro air vehicle systems. ELKA-R was designed using the 32-bit low-power ARM Cortex-M4 microprocessor as the microcontroller unit. The microcontroller unit interfaced with state of the art 9 degrees-of-freedom inertial measurement unit using inter-integrated circuit communication (I2C) protocol. A wireless transceiver was also incorporated with serial peripheral interface to wirelessly coordinate pilot inputs and sensor information with a remote basestation. Multiple timer protocols were configured to generate individual driver signals to a wide variety of motor and actuator configurations. The printed circuit board was designed as a four layer layout. ELKA-R weighed 1.7 g with a board area of 4.82 cm2, thus making it one of the smallest and lightest kinematic autopilots in open literature that can be applied to any generic micro air vehicle system. ELKA-R was tested on a variety of micro air vehicle flight demonstrators. Hover stabilization rates of 1000 Hz were achieved which were comparable to the autopilots on larger quad rotor systems such as DJI Phantom and AR-Drone. Oscillations in attitude were reduced by up to 50%–70% when compared with a previous generation lightweight autopilot.

This paper describes the design, controls system development, and hover testing of a 60 -g meso-scale cycloidal-rotor based (cyclocopter) micro air vehicle. The cycloidal rotor (cyclorotor) is a revolutionary vertical take-off and landing concept with a horizontal axis of rotation. The twin-cyclocopter utilizes two optimized cyclorotors and a horizontal tail rotor used to counteract the pitching moment generated by the cyclorotors. An innovative light-weight and high strength-to-weight ratio blade design significantly reduced cyclorotor weight and improved aerodynamic efficiency. In addition, increasing the virtual camber and incidence (by increasing chord-to-radius ratio) and using a symmetric pitching schedule with a maximum ± 45° pitching amplitude also improved rotor efficiency. Due to gyroscopic coupling and inherent instability of the cyclocopter, a closed-loop feedback control system was implemented using a custom autopilot weighing 1.5 g. The 60-g meso-scale twin-cyclocopter successfully demonstrated stable, sustained hover.

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 ...

In this study, a computational fluid dynamics analysis was performed on bio-inspired micro aerial vehicle scale rigid flapping wings. The computational fluid dynamics analysis used a compressible unsteady Reynolds Averaged Navier–Stokes solver with low-Mach number preconditioning to study the complex, highly vortical, three-dimensional flow of low aspect ratio flapping wings at micro aerial vehicle-scale Reynolds numbers. The wing was flapped at a constant 5 Hz flap frequency at a mean chord reference Reynolds number of 25,000. The flapping and pitching kinematics were set to match those of a previous experimental study resulting in a constant flap stroke of 107° at translational pitching angles of 40°, 50°, and 60°. The force and flowfield measurements of the previous flapping-wing experiment were used for the validation of the 3D computational fluid dynamics model. The objectives of this effort were to understand the unsteady aerodynamic mechanisms and their relation to force production and aerodynamic efficiency on a rigid wing undergoing an insect-type flapping motion with passive pitching kinematics. Overall, the computational fluid dynamics results showed good agreement with the measured experimental force data. Additionally, the computational fluid dynamics simulation was able to adequately predict the process of leading edge vortex formation and shedding observed during experimentation. A vorticity summation approach used to calculate the strength of the leading edge vortex from the experimental measurements and from the computational fluid dynamics predicted flowfields showed comparable results. The computational fluid dynamics results were utilized to further analyze the differences in the flowfield and leading edge vortex formation for the three pitch angles tested as well as the instantaneous aerodynamic loads and aerodynamic power.

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

This paper describes 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.

Targeted experiments in parallel with a systematic computational-fluid-dynamics analysis were performed for a micro-air-vehicle-scale rigid flapping wing in forward flight. Two-component time-resolved particle-image-velocimetry measurements were performed in an open-circuit wind tunnel on a wing undergoing pure flap-wing kinematics at a fixed wing-pitch angle. Chordwise velocity fields were obtained at equally spaced spanwise sections along the wing (30 to 90% span) at two instants during the flap cycle (middownstroke and midupstroke) for the reference Reynolds numbers of 15,000. The flowfield measurements were used for the validation of the three-dimensional computational-fluid-dynamics model. The computational-fluid-dynamics analysis used a compressible Reynolds-averaged Navier–Stokes solver to resolve the complex, highly vortical, three-dimensional flow. The objectives of the combined efforts were to understand the unsteady aerodynamic mechanisms and their relation to force production on a rigid wing u...

This work focuses on autonomous operation of rotary-wing Micro Aerial Vehicles (MAVs) in previously unexplored, GPS-denied environments. This type of operation requires the MAV to simultaneously estimate a map of its surroundings and its own position within that map. This work presents a refined algorithm for performing Simultaneous Localization and Mapping (SLAM) which does not rely on revisiting previously mapped areas. The algorithm is first benchmarked against previously published algorithms. Different scenarios, both indoors and outdoors, are then presented, showing the high amount of detail captured by the proposed method. Scenarios were mapped using three different platform types, including a conventional main-tail rotor helicopter, showing that the algorithm is platform-independent and does not require the rotorcraft's dynamic model, which can be complex. Experimental validation of the SLAM method shows an accuracy of approximately 0.1% of the traveled course length. The algorithm is capable of operating in real time on small form factor computers.

This paper discusses the control methodology, flight dynamics identification, and disturbance rejection analysis in hover of a revolutionary horizontal-axis rotary-wing concept: the twin cyclocopter. The vehicle has a gross weight of 500 g (1.25 by 1.67 by 1 ft in dimensions) and comprises two highly optimized cyclorotors along with a tail rotor for pitch control. Stable hover flight required fast control of the rpm and thrust vectoring of the rotors through onboard feedback regulation. A six-degree-of-freedom flight dynamics model of the vehicle was extracted through input excitation and time-domain identification. The longitudinal and heave degrees of freedom were decoupled and independent from the rest of the dynamics. Longitudinal translation damping was higher than in the heave mode, indicating differences in restoring forces at different blade azimuth positions due to the different pitch angles. Strong gyroscopic coupling was observed between lateral and yaw degrees of freedom because the rotors spi...

It is useful to combine hover and forward flight capability of air vehicles into a hybrid design. This paper discusses the development of one such hybrid air vehicle: Quadrotor biplane. The proof-of-concept vehicle weighs 240 grams and consists of four propellers with wings arranged in a biplane configuration. The performance of the propeller-wing was investigated in non-axial flow conditions through a systematic series of wind tunnel experiments. The effect of the wing on propeller slipstream and vice versa significantly changed the magnitude of vertical and horizontal forces when compared with the forces produced by an isolated propeller. Accounting for these effects, trim analysis showed that the maximum speed of 11 m/s at 0 deg shaft angle and a cruise speed of 4 m/s at 18 deg shaft angle was achievable. The cruise power was approximately one-third of that required for hover. Free flight testing of the proof-of-concept vehicle successfully showed feasibility of vehicle to achieve equilibrium transition flight. In order to improve performance, a few key design parameters such as airfoil profile, wing aspect ratio, biplane wing spacing and offset between propeller axis and wing chord were also experimentally investigated. It was found that a careful trade-off between efficiency and compactness affects the final choice of the design.

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

This paper describes the experimental investigation of a quad-shrouded rotor micro air vehicle and focuses on the hover performance improvements over a conventional unshrouded micro quad rotor. The effects of number of rotor blades, blade root pitch angle, and shroud diffuser length on aerodynamic performance were studied. The rotor diameter was 6.6 cm, it had a tip Re of 20,000, and the gross weight of the vehicle was 90 g, with a shroud weight fraction of 13%. With the optimized design, the power loading of the quad-shrouded rotor was about 15% greater than the unshrouded configuration. To completely evaluate the configuration, the performance of the vehicle in edgewise flow was investigated. The drag and pitching moment for the shrouded rotor was about 2.5 times greater than the unshrouded vehicle. However, control-moment measurements suggested that the edgewise gust tolerance was at least 4 m/s. The vehicle prototype was also successfully flight tested in hover using onboard feedback regulation. Base...

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...

Experimental studies were conducted by flapping a rigid rectangular wing with a mechanism that is capable of emulating complex insect wing kinematics, including figure-of-eight motions, in order to explore the fundamental unsteady flow on a flapping wing at micro-air-vehicle-scale Reynolds numbers. Force and moment measurements were obtained from a miniature six-component force transducer installed at the wing root. The wing was flapped in air and vacuum at the same frequency, and wing kinematics, and the resultant forces, were subtracted in order to obtain the pure aerodynamic forces. In the first part of this paper, the forces produced on the wing undergoing single-degree-of-freedom fixed-pitch pure flapping motions (no pitching or out-of-the-plane coning motions) were determined for a variety of pitch angles. The unsteady aerodynamic coefficients measured during these tests were almost six times the steady-state values measured in the wind tunnel. Flow visualization and particle image velocimetry tests...

A cycloidal rotor is an innovative horizontal axis propulsion system, which is utilized in the present study to develop a fully-controllable flying aircraft (known as “Cyclogyro”). The present 535 gram cyclogyro uses a hybrid configuration with two cyclorotors and a horizontal tail rotor. Since a cycloidal rotor relies on cyclic blade pitching for producing thrust and also for control, designing a reliable pitch mechanism that can provide the required blade pitching kinematics (depending on the advance ratio) is extremely important. A novel blade pitch mechanism has been developed, which is passively driven by centrifugal force, and could potentially be adapted to generate the required pitching schedules for efficient operation over a range of advance ratios. A simplified flightworthy version of this mechanism was implemented in the present cyclogyro. The mechanism was able to generate the appropriate blade kinematics and the thrust required for the vehicle to hover. Also, the present mechanism is designed such that it is possible to vary both amplitude and phasing of the cyclic blade pitching. A novel control strategy was developed using blade pitch amplitude (thrust magnitude) control for roll, phasing (thrust direction) for yaw and tail rotor for pitch control. The control strategy was implemented using a three gram onboard processor, which was used to stabilize the vehicle without a pilot, through a closed-loop feedback control system. This is the first flight-capable cyclogyro reported in the literature to utilize cycloidal rotors having both pitch amplitude and phase control. The successful flight for the present vehicle also validates the flightworthiness of this completely passive pitch mechanism design, which has great potential for efficient forward flight.

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...

Systematic experimental studies were performed to understand the role of two key degrees of freedom, flapping and pitching, in aerodynamic performance of a flapping wing, in both hover and forward flight. Required flapping kinematics is prescribed mechanically, and dynamic pitching/twisting is obtained passively using inertial and aerodynamic forces. Forces produced by the wing are measured at the root using a six-component balance at different flapping frequencies, flapping/pitching amplitudes, and wind speeds. The results clearly show that maximum average thrust over a flap cycle in hover can be achieved using symmetric, high amplitude passive pitching. However, in forward flight, optimum aerodynamic performance (lift and propulsive thrust) is obtained using asymmetric wing pitching with low pitching amplitudes. Furthermore, dynamic twisting (obtained using flexible wings), instead of dynamic pitching, produces better performance in forward flight due to spanwise and temporal modulation of the wing pitch angle. Pure flapping (no pitching) of rigid wings in forward flight at high reduced frequencies and high pitch angles produces a threefold increase in lift coefficient over static values. Maximum average propulsive thrust over a flap cycle in forward flight is obtained using symmetric pitching. To produce high values of both, average lift and thrust, an asymmetry in kinematics along with pitching is required in forward flight. This can be achieved either through asymmetric pitching of rigid wings or dynamic twisting of torsionally flexible wings.

Experimental studieswere conducted to study the response of a shrouded rotormicro air vehicle to edgewise gusts. In edgewise flow, the thrust, drag, and pitching moment produced by three platforms were compared: elliptic inlet, circular inlet shrouded rotor, and an unshrouded rotor. The elliptic inlet shrouded rotor wasmore efficient in hover but had ahigher penalty in drag andpitchingmoment in edgewiseflow.Cyclic pitch variation of a hingeless rotorwas used to counter these adverse pitching moments. The control authority of the shrouded rotors was at least 80–100% higher than the unshrouded rotor, with the elliptic inlet shrouded rotor producing the highest control moments. By optimizing rotor collective settings, it was possible to reduce the deteriorating effect of edgewise flow on control moments. To increase the control authority and gust tolerance of the shrouded rotor, the cyclic pitch travel and blade planform modifications were made. With a careful selection of rotor solidity, planform, operating revolutions per minute, and cyclic pitch travel, it was possible to achieve a gust tolerance of about 3 m=s for the circular inlet shrouded rotor. Free flight tests were then conducted to study the ability of the vehicle to hover in a given position in the presence of gusts generated from pedestal fans with honeycomb flow straighteners. A combination of VICON and an onboard sensor were used for feedback control. The vehicle was satisfactorily able tomaintain hover position in edgewise gusts of up to 3 m=s.

Shrouded rotors are efficient in hover but are quite sensitive to disturbances in external flow. In this paper, the dynamics and control of a shrouded rotor micro air vehicle is studied in hover and when it is subjected to edgewise gust. The importance of incorporating a hingeless rotor in a shrouded rotor configuration was shown and was flight-tested in hover using a proportional-integral attitude feedback controller. In edgewise flow, the shrouded rotor produced up to 300% higher pitching moment than the unshrouded rotor. To counter this pitching moment, the control moments were about 80–100% higher for the shrouded rotor. Time domain attitude dynamics identification of the vehicle, restrained in translation, was conducted with and without the flybar. It was shown to be desirable to incorporate a flybarless rotor for improved maneuverability and hover efficiency. A linear quadratic regulator (LQR) controller was developed based on the extracted attitude dynamics model. Gust disturbance rejection capabilities of the controller were tested with the vehicle in edgewise flow using a spherical gimbal setup. The shrouded vehicle was found to tolerate up to 2 m/s of edgewise gusts, whereas the unshrouded configuration could reject gusts of up to 4.8 m/s.

On-blade active controls with trailing-edge deflection, leading-edge deflection, and active-twist are studied for improvements in rotor aerodynamic efficiency and their influence on structural loads. A full-scale UH-60A Blackhawk rotor at two key flight conditions (high-speed forward flight and high-thrust forward flight) is studied using coupled computational fluid dynamics and computational structural dynamics simulations. A simulationbased trade study is carried out comprising parametric variations of geometric sizing and deployment schedules of the blade morphing. The study shows that active controls improve rotor performance and reduce rotor loads at the sametimewithcarefulselectionofdeploymentscheduleanddesign.Inhigh-speedforward flight,usingtrailing-edge deflection, an improvement of 7.3% in performance and a reduction in the hub vibratory loads of up to 54% is achieved,and using active-twist animprovement of 7.0% in performance and upto 22% reduction in hub vibratory loads is achieved. In high-thrust forward flight, a 15.0% improvement in performance and up to 40% reduction in hub vibratory loads is achieved using leading-edge deflection.

Performance and flowfield measurements were conducted on a small-scale cyclorotor for application to a micro air vehicle. Detailed parametric studies were conducted to determine the effects of the number of blades, rotational speed, and blade pitching amplitude. The results showed that power loading and rotor efficiency increased when using more blades; this observation was found over a wide range of blade pitching amplitudes. The results also showed that operating the cyclorotor at higher pitching amplitudes resulted in improved performance, independently of the number of blades. A momentum balance performed using the flowfield measurements helped to quantify the vertical and sideward forces produced by the cyclorotor; these results correlated well with the force measurements made using load balance. Increasing the number of blades increased the inclination of the resultant thrust vector with respect to the vertical because of the increasing contribution of the sideward force. The profile drag coefficient of the blade sections computed using a momentum deficit approach correlated well with typical values at these low chord Reynolds numbers. Particle image velocimetry measurements made inside the cage of the cyclorotor showed that there are rotational flows that, when combined with the influence of the upper wake on the lower half of the rotor, explain the relatively low efficiency of the cyclorotor.

This paper predicts and analyzes 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 with a transient lifting-line aerodynamic model. The structural model includes a swashplate model to calculate servo loads. The lifting-line model combines airfoil tables, a Weissinger-L near-wake time-marching free wake, and a semiempirical dynamic stall model. The maneuver data were taken from the Army/NASA UH-60A Airloads Program Flight Counter 11029. The primary objective of this paper is to isolate the effects of structural dynamics, free wake, dynamic stall, and pitch control angles in order 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, which is ideally suited to isolate the effects of free wake and dynamic stall. It is concluded that the maneuver is almost entirely dominated by stall, with little or no wake-induced effect 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 shows an underprediction of 10-20% in flap and chord bending moments and 50% in torsion loads. The errors stem from the prediction of four-and five-revolution stall loads. Swashplate dynamics appear to have a significant impact on the servo loads (unlike in level flight), with a more than 50% variation in peak loads.

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.

This paper addresses the aerodynamics of insect-based, biomimetic, flapping wings in hover. An experimental apparatus, with a biomimetic flapping mechanism, was used to measure the thrust generated by a number of wing designs at different wing pitch settings. To quantify the large inertial loads acting on the wings, vacuum chamber tests were conducted. Results were obtained for several high-frequency tests conducted on lightweight aluminum and composite wings. The wing mass was found to have a significant influence on the maximum frequency of the mechanism because of a high inertial power requirement. All the wings tested showed a decrease in thrust at high frequencies. In contrast, for a wing held at 90-deg pitch angle, flapping in a horizontal stroke plane with passive pitching caused by aerodynamic and inertial forces, the thrust was found to be larger. To study the effect of passive pitching, the biomimetic flapping mechanism was modified with a passive torsion spring on the flapping shaft. Results of some tests conducted with different wings and different torsion spring stiffnesses are shown. A soft torsion spring led to a greater range of pitch variation and produced more thrust at slightly lower power than with the stiff torsion spring. The lightweight and highly flexible wings used in this study had significant aeroelastic effects which need to be investigated. A finite element based structural analysis of the wing is described, along with an unsteady aerodynamic analysis based on indicial functions. The analysis was validated with experimental data available in literature, and also with experimental tests conducted on the biomimetic flapping-pitching mechanism. Results for both elastic and rigid wing analyses are compared with the thrust measured on the biomimetic flapping-pitching mechanism.

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.

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

This article examines the distribution of strain energy in the various component materials of actuated plates and investigates the manner in which the strain energy distribution is influenced by the actuated span-to-thickness ratio and the thickness of the adhesive bond layer. Furthermore, the article investigates the effect of modeling choices (e.g., kinematic assumptions and mesh density) on the predicted magnitude and mode of the dominant strain energy form in each component material. These computed parameters can be used to quantify the overall efficiency of an actuated plate in addition to aiding the understanding of the local mechanics that govern the process. The focus problem consists of a square aluminum plate with a single symmetric pair of surface-mounted piezoceramic actuators that are used to produce in-plane extension or bending in the aluminum plate. The behavior of the actuated plate is examined over a range of plate thicknesses and adhesive bond layer thicknesses by using a series of finite element models that feature different levels of kinematic complexity and different levels of two-dimensional (2-D) mesh density. The results of the study emphasize the need for discrete layer kinematics in determining the magnitude and mode of the dominant strain energy form in each constituent material; however, these computed parameters are shown to be rather insensitive to changes in 2-D mesh density. Most importantly, the study confirms the existence and quantifies the magnitude of the local kinematic effect, whereby a portion of the available actuation energy is diverted to the production of localized transverse shear deformation and transverse normal deformation, thus reducing the amount of actuation energy available to produce in-plane deformation in the structural substrate.

This article discusses finite element modeling of progressive damage in the adhesive bond layers of actuated plates and investigates the reduction in actuation capacity caused by the damaged bond layers. The primary challenge posed by this class of problems stems from the vast range of geometric scales that are represented, with the thickness of the adhesive layer representing the smallest scale, the overall thickness of the actuated plate representing the intermediate scale, and the in-plane dimensions of the plate representing the largest scale. In multiscale problems, the overall efficiency of the numerical methodology is of paramount importance, thus model development is guided by the need to obtain a sufficiently accurate solution at an acceptable computational expense. In this study, this goal is achieved through the use of a hierarchical, displacement-based, 2-D finite element model that includes the first-order shear deformation (FSD) model, Type I layerwise models (LW1) and Type II layerwise models (LW2) as special cases. Both the LW1 layerwise model and the more familiar FSD model use a reduced constitutive matrix that is based on the assumption of zero transverse normal stress; however, the LW1 model includes discrete layer transverse shear effects via in-plane displacement components that are C0 continuous with respect to the thickness coordinate. The LW2 layerwise model utilizes a full 3-D constitutive matrix and includes both discrete layer transverse shear effects and discrete layer transverse normal effects by expanding all three displacement components as C0 continuous functions of the thickness coordinate. The hierarchical finite element model incorporates a 3-D continuum damage mechanics model that predicts local orthotropic damage evolution and local stiffness reduction at the geometric scale represented by the individual material ply or, in the case of layerwise models, by the individual numerical layer. The results clearly demonstrate that the resulting model can efficiently simulate progressive damage in the adhesive layers. For rectangular actuator patches, the adhesive damage is highest near the corners of the actuator and is driven primarily by local concentrations in the transverse normal and transverse shear stresses. In contrast to previous studies that have shown that the inclusion of discrete layer transverse normal stress does not significantly influence the predicted global deformations, the present study shows that the transverse normal stress has a very significant effect in the initiation and progression of localized damage in the adhesive layers.

A quasi-static model for NiMnGa magnetic shape memory alloy (MSMA) is formulated in parallel to the Brinson and Tanaka thermal SMA constitutive models. Since the shape memory effect (SME) and pseudoelasticity exist in both NiTi and NiMnGa, constitutive models for SMAs can serve as a basis for MSMA behavioral modeling. The simplified, linear, quasi-static model for NiMnGa was characterized by nine material parameters identified by conducting a series of uniaxial compression tests in a constant field environment. These model parameters include free strain, Young's modulus, fundamental critical stresses, fundamental threshold fields, and stress-influence coefficients. The Young's moduli of the material in both its field and stress preferred configurations were determined to be 450 MPa and 820 MPa respectively, while the free strain was measured to be 5.8%. These test data were used to assemble a critical stress profile that is useful for determining model parameters and for understanding the dependence of critical stresses on magnetic fields. Once implemented, the analytical model shows good correlation with test data for all modes of NiMnGa quasi-static behavior, capturing both the magnetic shape memory effect and pseudoelasticity. Furthermore, the model is also capable of predicting partial pseudoelasticity, minor hysteretic loops and stress–strain behaviors. To correct for the effects of magnetic saturation, a series of stress influence functions were developed from the critical stress profile. Although requiring further refinement, the model's results are encouraging, indicating that the model is a useful analytical tool for predicting NiMnGa actuator behavior.