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This study evaluates the potential benefits that fixed but deployable trailing-edge flaps may have on a low-twist hovering rotor using a blade-element method combined with two-dimensional and three-dimensional computational fluid dynamics. Low blade twist is beneficial in terms of forward-flight performance, which, combined with a deployable flap, could also offer good hover performance, resulting in an overall better design in comparison with a plain twisted blade. To evaluate this concept, a parametric study of flap configurations was conducted using a simple blade-element method based on computational-fluid-dynamics-generated two-dimensional aerodynamics. This simple model indicated that up to 6 deg of blade twist could be recovered at high rotor thrust by using an optimized flap configuration. Performance improvements were also obtained for outboard slotted-flap configurations. The optimum slotted-flap designs were also evaluated using three-dimensional computational fluid dynamics, which confirmed that an inboard flap combined with a low-twist (-7deg) rotor blade matched the performance of a plain blade with -13deg of twist. A blended-flap configuration was also applied and evaluated. This simpler design demonstrated the equivalent performance of a clean rotor with 10 deg of blade twist, providing further evidence of the potential of the fixed inboard flap concept.

An experimental study of a rotatable tail mechanism applied to a small unmanned aerial vehicle was performed using a six-component wind-tunnel balance in the U.S. Air Force Institute of Technology low-speed wind tunnel. Attributes of the control and stability characteristics of the original vehicle, which were documented in an earlier study, are compared with those of a unique control methodology, a tail consisting of a single surface, with controllable elevation and rotation. An advantage of this change is a reduction in the storage length of the vehicle. Because there are similarities in the rotatable tail mechanism and the tail of many birds, the rotatable tail reflects a biomimetric feature. Measured force and moment coefficient measurements for the actual vehicle at a typical flight speed indicated that a rotatable tail provides a sufficient yaw moment for turning. For example, yaw moment coefficients [C.sub.n], ranging from -0.02 to +0.02, which is typical for a rudder, were achievable as long as the absolute value of the tail elevation angle was large. The dependence of the yaw moment coefficient on the elevator angle and angle of attack, in addition to the tail rotation angle, indicates that there would be significant challenges in applying a robust flight control scheme with the current actuator configuration. An additional feature of the tail design is that by deflecting the tail upward, it could also function effectively as an air brake. A more than twofold increase in drag coefficient for constant angle of attack was measured when the tail elevation angle was increased to nearly 70 deg. DOI: 10.2514/1.24192

Aerodynamic and acoustic experiments were carried out to investigate the induced inflow distribution and noise characteristics of a scissors rotor in hover. The aerodynamic experiment was conducted on a model rotor rig and the induced inflows for two different scissors rotor configurations (the L configuration and U configuration) were measured. The experimental results indicate that the induced velocity near the lower blades is consistently larger than that near the upper blades. A free-wake analysis was used to predict the aerodynamic characteristics. The variation of the induced velocity with different scissors angles was simulated analytically and the results were compared with test data. The acoustic experiment was conducted in an anechoic aeroacoustic laboratory and the noise levels for both the scissors rotor configuration and a conventional configuration were measured. The effects of different scissors angles on rotor noise were investigated. It was shown that the noise level of a scissors rotor depends on both the azimuthal and vertical spacing and is not necessarily less than that of a conventional configuration.

A simple semi-empirical model for the aerodynamic behavior of a low-aspect ratio pararotor in autorotation at low Reynolds numbers is presented. The paper is split into three sections: Sec. II deals with the theoretical model derivation, Sec. III deals with the wind-tunnel measurements needed for tuning the theoretical model, and Sec. IV deals with the tuning between the theoretical model and the experimental data. The study is focused on the effect of both the blade pitch angle and the blade roughness and also on the stream velocity, on the rotation velocity, and on the drag of a model. Flow pattern visualizations have also been performed. The value of the free aerodynamic parameters of the semi-empirical model that produces the best fit with the experimental results agrees with the expected ones for the blades at the test conditions. Finally, the model is able to describe the behavior of a pararotor in autorotation that rotates fixed to a shaft, validated for a range of blade pitch angles. The movement of the device is found to be governed by a reduced set of dimensionless parameters. DOI: 10.2514/1.21435

Results from research into active structural load alleviation are presented for a tilt-rotor aircraft. The work formed part of the European Commission's Fifth Framework 'critical technology' Rotorcraft Handling, Interactions, and Loads Prediction project. Results are presented from an analysis of the structural loads of a tilt-rotor aircraft maneuvering in airplane mode. The potential for suppression of structural loads through active control is assessed. The study has addressed modeling aspects, particularly the nature of the buildup of dynamic loads during maneuvers. The Eurocopter EUROTILT configuration has been used as the test aircraft and a model is developed within the FLIGHTLAB simulation environment. Mathematical analysis and simulation of the rotor in-plane moments and gimbal flapping during longitudinal and lateral maneuvers are presented. The problem also addressed the torque split on the interconnect drive shaft in airplane mode for maneuvers involving large roll and yaw rates. Attempts were made to design an 'ideal' controller that minimizes the gimbal flap excursions and the loads using a linear quadratic Gaussian formulation. In this preliminary conceptual study, robustness, actuator requirements, or the availability of measurements were not considered. The main objective of the controller design study was to investigate the effectiveness of cyclic controls to achieve simultaneous suppression of the loads and flapping. Evaluations performed using the controller show that use of cyclic controls is very effective in suppressing the buildup of in-plane loads, gimbal flapping, and interconnect drive shaft torque split.

The results of a study to determine the optimum two-per-revolution blade control inputs required to minimize rotor power, maximize thrust, or minimize rotor speed, using formal numerical optimization techniques, are presented. The impact of improved rotor modeling on the prediction of the effects of two-per-revolution input on rotor power is also addressed, in particular, the effect of using a free wake inflow model, and a flexible rotor blade model. The main conclusions of the present study are as follows: The optimization for optimum two-per-revolution input is best carried out by solving the trim problem for every value of the inputs proposed by the optimizer, as opposed to integrating the trim calculations as equality constraints in the optimization. An appropriate two-per-revolution input can increase the rotor maximum thrust at high speed by about 11%. The additional thrust is generated at the front and the rear of the rotor disk. The two-per-revolution input can also further reduce the rotor speed compared with a case without such input. The physical mechanism is the same as in the optimization for maximum thrust, and the maximum value of [C.sub.T]/[sigma] is also the same. Although the power predictions obtained using a free wake model are different from those obtained with a simpler linear inflow model, the accuracy of the optimum two-per-revolution input predicted using the simpler model is adequate. The best practical strategy to optimize the two-per-revolution input is to conduct an initial study with a simple linear inflow model and then refine it with the more sophisticated one. The same conclusions hold for more sophisticated flexible blade models, compared with simpler rigid blade ones.

The maximum blade forced response amplitudes for mistuned turbomachinery rotors are generally much greater than those of their tuned counterparts. However, it is known that the ratio of mistuned to tuned maximum vibration amplitudes, the amplitude magnification, is often largest at a relatively small level of mistuning. Increasing the level of mistuning beyond this critical value actually leads to a decrease in the amplitude magnification. This suggests that it might be beneficial to introduce some level of mistuning into the nominal design of the system intentionally. In this study, the effectiveness of this intentional mistuning strategy is investigated. Intentional mistuning is introduced into the rotor design by varying the nominal blade stiffnesses in harmonic and square-wave patterns. In addition, the unavoidable, random mistuning of the blades is included in the model as usual. The statistics of the forced response are examined for lumped parameter models, as well as for a finite element based reduced-order model of an industrial rotor. For the cases considered, it is found that intentional mistuning can greatly reduce a rotor's sensitivity to random mistuning.

The active twist rotor investigated in this research is a derivative of the previously developed smart-active-blade-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 significant 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 coefficients are 1.4 x [10.sup.-3], 0.55 x [10.sup.-3], and 0.7 x [10.sup.-3]. The corresponding finite 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.

Computational and experimental methods in assessing a leading-edge slat, a deformable leading-edge and upper-surface blowing revealed the utility of these concepts in enhancing the lift, drag and pitching moment features, thereby improving rotor maneuver capability. Despite the utility of these concepts in delaying dynamic stall, further tests on the impacts of Mach numbers and high Reynolds numbers are needed to pave the way for efficient usage of these concepts in rotorcraft.

Roforcraft design engineers make use of rotorcraft simulations to check new designs of airplanes and analyze high risk operational and aerodynamic areas. This helps in avoiding the dangers accompanying a real flight test. Sikorsky Aircraft's Full Mission Simulation (FMS) analyzes the technology used in the design of an aircraft through mathematical models. The general mission equipment package (GENMEP) helps pilots in visualizing combat scene and provides him with solutions to tackle the problems. Simulator networks make rotorcraft testing and checking of new weapons system easier.

The rotorcraft industry is now utilizing advancements in the computational structures technology (CST) to enhance the design of rotorcraft structures. Through CST, larger and more detailed models are now being used in the structural analysis of rotorcraft parts. Highly advanced computers and computer programs are also enabling engineers to predict the behavior of airframes under various conditions.

The probabilistic structural analysis and design of a silicon micro-turbo-generator rotor are presented. This rotor was designed to have a tip speed of 500 m/s. Three-dimensional finite element analysis, fracture strength characterization of single crystal silicon, and structural failure probability calculations were performed. The results show that the design of micro-turbine-generator rotor, although highly stressed, is feasible. However, it is important to note that this feasibility is very dependent on the etching process achieving a high surface quality. The overall approach and tests employed are equally applicable to other highly stressed microelectromechanical systems.

The vibratory load reduction at a rotor hub using smart materials and closed-loop control is investigated. The principal load-carrying member in the blade is represented by a composite box beam, of arbitrary thickness, with surface bonded self-sensing actuators. A comprehensive theory is used to model the smart box beam. The theory, which is based on a refined displacement field, is a three-dimensional model that approximates the elasticity solution so that the beam cross-sectional properties are not reduced to one-dimensional beam parameters. Both in-plane and out-of-plane warpings are included automatically in the formulation. Next, an integrated rotor vibratory load analysis procedure is developed by coupling an unsteady aerodynamic model with the rotor blade dynamic model based on the smart composite box beam theory. The dynamic deformations of the blade in all three directions, flap, lead lag, and torsion, are included in the analysis. The finite-state, induced-flow model is used for predicting the dynamic loads. The procedure results in significant reductions in the amplitudes of rotor dynamic loads with closed-loop control. Detailed parametric studies are presented to assess the influence of number of actuators and their locations in vibratory hub load reduction.