Alfred Gessow Rotorcraft Center

Engineering, United States

Alfred Gessow Rotorcraft Center

Engineering, United States
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Hersey S.,University of Maryland University College | Hersey S.,Alfred Gessow Rotorcraft Center | Sridharan A.,University of Maryland University College | Sridharan A.,Alfred Gessow Rotorcraft Center | And 2 more authors.
Journal of Aircraft | Year: 2017

This paper presents the results of a rotorcraft preliminary design problem, solved as a multiobjective design optimization problem. A lift- and thrust-augmented coaxial compound configuration is used to demonstrate the approach.The basic optimization problemis converted into a sequence of approximate optimization problems, inwhich approximate Pareto frontiers are calculated based on response surfaces, obtained from radial basis function interpolation of all the designs analyzed at every step of the sequence.The Pareto frontiers are computed using a genetic algorithm. The designs are analyzed using a high-fidelity rotorcraft analysis that includes nonlinear finite element models of the rotor blade and a free vortex wake model of the rotor inflow. The results presented indicate that 1) the preliminary design problem can be effectively solved using formal numerical design optimization techniques, which therefore can complement classical design methodologies; 2) with appropriate physics-based constraints, the design optimization can be carried out by the computer completely unattended; 3) the optimization methodology is sufficiently robust to deal with multiple local optima and other numerical difficulties; and 4) themethodology is efficient enough to allow the use of high-fidelity analyses throughout the optimization, with the use of graphical processing unit computing (Compute Unified Device Architecture/Fortran) contributing to the computational efficiency. © Copyright 2016 by the authors. Published by the American Institute of Aeronautics and Astronautics, Inc.


Jain R.,HyPerComp, Inc. | Yeo H.,Ames Research Center | Yeo H.,U.S. Army | Chopra I.,University of Maryland University College | Chopra I.,Alfred Gessow Rotorcraft Center
Journal of Aircraft | Year: 2013

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 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. Copyright © 2012 Clearance Center, Inc.


Benedict M.,University of Maryland University College | Benedict M.,Alfred Gessow Rotorcraft Center | Mattaboni M.,Polytechnic of Milan | Mattaboni M.,Alfred Gessow Rotorcraft Center | And 4 more authors.
AIAA Journal | Year: 2011

This paper describes the aeroelastic model to predict the blade loads and the average thrust of a micro-air-vehiclescale cycloidal rotor. The analysis was performed using two approaches: one using a second-order nonlinear beam finite element method analysis for moderately flexible blades and a second using a multibody-based largedeformation analysis (especially applicable for extremely flexible blades) incorporating a geometrically exact beam model. An unsteady aerodynamic model is included in the analysis with two different inflow models: single streamtube and double-multiple streamtube inflow models. 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 nosedown elastic twisting of the blades in the upper half cylindrical section, which is not compensated by a noseup pitching in the lower half-section. The inclusion of the actual blade pitch kinematics, unsteady aerodynamics, and flow curvature effects was found crucial in the accurate lateral force prediction. Copyright © 2011 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved.


Abhishek A.,University of Maryland University College | Datta A.,Eloret Corporation | Datta A.,NASA | Ananthan S.,University of Maryland University College | And 2 more authors.
Journal of Aircraft | Year: 2010

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 fourand 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. Copyright © 2009 by the American Institute of Aeronautics and Astronautics, Inc.


Mayo D.B.,Alfred Gessow Rotorcraft Center | Lankford J.L.,University of Maryland University College | Benedict M.,University of Maryland University College | Chopra I.,Alfred Gessow Rotorcraft Center
Annual Forum Proceedings - AHS International | Year: 2014

Experiments were systematically executed in conjunction with a coupled CFD-CSD-based aeroelastic analysis for a MAV-scale flexible flapping-wing in forward flight. 2-D time-resolved particle image velocimetry (PIV) and force measurements were performed in a wind tunnel at a flow speed of 3 m/s. The flexible wing undergoes pure flapping kinematics held at a fixed wing-pitch angle at the root. Chordwise velocity fields were obtained at equally spaced spanwise sections along the wing (30% to 90% span) at two instants during the flap cycle (mid-downstroke and mid-upstroke) for the reference Reynolds numbers of 15,000. The flowfield measurements and averaged force measurements were used for the validation of the 3-D aeroelastic model. The objectives of the combined efforts were to understand the unsteady aerodynamic mechanisms and their relation to force production on a flexible wing undergoing an avian-type flapping motion. The coupled CFD-CSD analysis combined a compressible Reynolds Averaged Navier Stokes (RANS) solver (OVERTURNS) with a multi-body structural solver (MBDyn) to resolve the complex, highly vortical, three-dimensional flow on a flexible wing. The coupled CFDCSD predicted and experiment flowfields showed comparable results. A vorticity summation approach was used to calculate the circulation of the leading edge vortex (LEV) from the PIV measurements and the numerical simulations to further validate the CFD-CSD code. The time-averaged vertical and horizontal aerodynamic forces measured from the experiment using a miniature force balance were also used to validate the force predictions from the CFD-CSD model. Pertaining to the flow physics, the flapping motion induces large angles of attack along the wing span causing the outboard sections to stall during downstroke. During the upstroke, the outboard sections operate at very low or even negative angles of attack. The present study showed that the dynamic twisting produced by the flexible wing helped in decreasing the effective angle of attack during the downstroke and upstroke. This temporal and spanwise variation of wing pitch angle affected both lift and drag, and primarily helped the wing produce positive thrust during both upstroke and downstroke which is not possible with a rigid wing undergoing pure flap at a constant pitch angle. Both PIV and CFD-CSD studies showed that the LEV stayed attached for a longer duration of the flap cycle during downstroke for the flexible wing compared to the rigid wing, especially towards the outboard sections. Also, during the upstroke, the LEV strength for the flexible wing was significantly higher than that for the rigid wing. ©2014 by the authors except where noted. All rights reserved. Published by the AHS International with permission.


Hrishikeshavan V.,University of Maryland University College | Chopra I.,University of Maryland University College | Chopra I.,Alfred Gessow Rotorcraft Center
Journal of Aircraft | Year: 2012

Experimental studies were conducted to study the response of a shrouded rotor micro 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 was more efficient in hover but had a higher penalty in drag and pitching moment in edgewise flow. Cyclic pitch variation of a hingeless rotor was 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 to maintain hover position in edgewise gusts of up to 3 m/s. Copyright © 2011 by Luis Delgado.


Mayo D.B.,University of Maryland University College | Lankford J.L.,University of Maryland University College | Benedict M.,University of Maryland University College | Chopra I.,University of Maryland University College | Chopra I.,Alfred Gessow Rotorcraft Center
Journal of Aircraft | Year: 2015

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-imagevelocimetry 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 threedimensional 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 undergoing an avian-type flapping motion. Overall, the computational-fluiddynamics results showed good agreement with the experimental data for resolution of the overall highly unsteady and vortical flowfield. A control-volume approach used to calculate the strength of the leading-edge vortex from the particle-image-velocimetry measurements and from the computational-fluid-dynamics-generated flowfields showed comparable results. A hybrid momentum-based method was used to estimate the sectional vertical force coefficient from the particle-image-velocimetry-measured flowfield, which agreed well with the computational-fluid-dynamics force prediction over a range of flapping frequencies and wing-pitch angles. In general, it was observed that the flow over the wing was highly susceptible to changes in induced angle of attack resulting from the flapping motion and variations in reduced frequency, which manifested in the predicted airloads. Based on the computational analysis, the spanwise flow component was not significant, except near the wing tip, and therefore most of the vertical force and propulsive thrust produced could be explained using the magnitude and direction of the sectional lift and drag forces acting on the wing. For the present wing kinematics, most of the upward vertical force was produced during the downstroke and positive propulsive thrust during the upstroke, which shows the need for appropriate temporal and spanwise pitch modulation of the wing along with flapping to produce positive vertical force and propulsive thrust during the entire flap cycle. © 2014 by David Mayo. Published by the American Institute of Aeronautics and Astronautics, Inc.


Benedict M.,University of Maryland University College | Jarugumilli T.,University of Maryland University College | Chopra I.,University of Maryland University College | Chopra I.,Alfred Gessow Rotorcraft Center
Journal of Aircraft | Year: 2013

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, dependingonthe blade pitching amplitude. Cycloidal rotors with shorterblade spans had higher power loading (thrust/power), especiallyatlower 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 blade pitching amplitudes were foundtoimprove the power loadingofthe cycloidal rotor. Asymmetric pitching with a 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. The power loading of the optimized cycloidal rotor was higher than that of a conventional microrotor. © 2012 by the American Institute of Aeronautics and Astronautics, Inc.


Seshadri P.,University of Cambridge | Benedict M.,University of Maryland University College | Benedict M.,Alfred Gessow Rotorcraft Center | Chopra I.,University of Maryland University College
Journal of Aircraft | Year: 2013

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 singledegree- 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 were also conducted, which showed that the key reason for the force increase on the flapping wing is due to a strong leading-edge vortex for which the strength varied throughout the flapping cycle. In the second part of this paper, complete three-degree-of-freedom (flapping, pitching, and coning) insect wing kinematics were investigated for different pitching and coning variations. The aerodynamic forces obtained in these tests were compared with coefficients obtained from the single-degree-of-freedom flapping tests and wind-tunnel tests. © 2012 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved.


Benedict M.,University of Maryland University College | Ramasamy M.,University of Maryland University College | Ramasamy M.,NASA | Chopra I.,University of Maryland University College | Chopra I.,Alfred Gessow Rotorcraft Center
Journal of Aircraft | Year: 2010

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.Amomentum 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 deflcit 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. Copyright © 2010 by the American Institute of Aeronautics and Astronautics.

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