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Compound rotorcraft can be a viable option to achieve moderately high speeds of 240 knots. However, there is no systematic wind tunnel test data in the public domain to understand the complex interactional aeromechanics involved. This work is part of a test campaign to conduct multiple high-speed wind tunnel tests of various compound configurations. The compound configuration consists of a single main rotor, a single wing on the retreating side for lift augmentation, and a single rear propeller for propulsive augmentation. In this paper, hover test results are presented. Hover tests were conducted for various configurations and at different tip Mach numbers with the maximum value of 0.5. Rotor and wing performances are measured separately using individual balances. Rotor blades are equipped with strain gauges for measuring flap, lag, and torsional moments at [Formula: see text] radial stations. The wing location is [Formula: see text] below the main rotor, and the measured wing download is only 2.8% of the rotor thrust. The wing provides a partial ground effect, making the rotor more efficient and almost compensates for the download with less than 1% reduction in net figure of merit. Blade structural loads correlate satisfactorily between the analysis and measurement except for lag bending moment. This discrepancy is a large 2/rev lag bending moment measured in the test, which needs further investigation. Propeller hover test was conducted up to 2500 revolutions per minute (tip Mach number of 0.25) at a single collective. Prediction correlates well and captures the low-Reynolds-number effect for lower revolutions per minute.

The predictions of an upgraded UMARC comprehensive analysis are compared to experimental lift offset rotor results. The experiments cover a range of collective pitch angles (o) from 2° to 10°, advance ratios (μ) from 0.21 to 0.53, and lift offset from 0% to 20%. The experimental model rotors are from a system of coaxial hingeless rotors, with two blades each, and a first flap frequency of approximately 1.6/rev. The simulation is compared with isolated rotor performance and controls with lift offset, loads, and pitch link forces. Increasing efficiency with increasing lift offset, the impact of lift offset on different loads, and the dependence of pitch link loads on pitch bearing damping are identified in the experiment and correlated with the simulation.

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...

This paper describes the gust rejection study of a twin-cyclocopter micro air vehicle with two cyclorotors and an anti-torque nose rotor. A gust rejection controller relying on velocity feedback and onboard flow sensing was implemented in a closed-loop feedback system. Tethered experiments were conducted with the vehicle mounted on a 6-DOF stand in front of a synthetic gust generation device capable of providing up to 4 m/s step gust input. Planar gusts along the longitudinal and lateral axis of the cyclocopter and crosswinds were systematically studied. Both pitch control (varying nose rotor rpm) and thrust vectoring control strategies were evaluated to counteract perturbation along the longitudinal axis, while the cyclocopter used differential rotational speed of the cyclorotors for lateral gusts. Results showed that thrust vectoring was more effective than pitch control in reducing displacement from gust. Pitch control using the position feedback controller resulted in a maximum gust tolerance of 2.8 m/s with a duration of 3 s. while thrust vector control using the combined flow feedback and position feedback controller was able to withstand 4 m/s step gust input with 1 s duration with only 0.01 m displacement. In addition, flow feedback was more effective than position feedback in minimizing position error. The cyclocopter was also able to mitigate step gusts of 2.8 m/s in magnitude and 3 s duration with crosswind components at 30 deg from the longitudinal axis. The difference in performance can be attributed to the cyclocopter's unique thrust vectoring capability.

A computational study is conducted on thin flat plates to simulate flows of Reynolds numbers at 104 to provide understanding and guidance for micro air vehicles and other low-Reynolds-number airfoi...

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 an investigation of the effects of tip sweep on highly twisted composite rotor blades by acquiring strain data in vacuum. Three parametric test cases were designed for strain me...

In forward flight, slowing down a rotor alleviates compressibility effects on the advancing side blade tip, extending the cruise speed limit, and inducing high-advance-ratio flight regime. Previous...

Rotorcraft flight speed is limited by compressibility effects on the advancing blade side and decreasing lift potential on the retreating blade side. It may thus be beneficial to employ a hingeless rotor to generate additional lift with the advancing blade and compensate the resulting rolling moment with a fixed wing on the retreating blade side. This concept is a form of "lift compounding" that appears to show enormous potential. The present paper presents results of a wind tunnel test with a slowed, hingeless rotor and single fixed wing on the retreating blade side. Based on rotor test stand data and flow field measurements, the impact of operational and rotor parameters on system performance and aerodynamics is examined, mutual interaction effects between rotor and fixed wing are analyzed, and dominant flow structures are characterized in the reverse flow region on the retreating blade side. Flow field analysis reveals a reverse flow entrance vortex that freely convects through the reverse flow region and rivals the blade tip vortices in strength. Contrary to previous beliefs, this vortex originates from upstream of the reverse flow region and only its detachment from the rotor blade is related to entering this region. The combination of finite rolling moment trim and aft shaft tilt significantly increases rotor lift coefficient and corresponding peak lift-to-drag ratio of the compound rotorcraft. Results are compared with predictions from a comprehensive rotor analysis that is expanded to cover the main effects of the added fixed wing and is able to reproduce general performance trends of the rotorcraft. The present study highlights that adding a single fixed wing and hingeless rotor to a high-speed rotorcraft could significantly improve its performance.

This paper describes the design and experimental validation of an all-terrain cyclocopter micro air vehicle capable of power-efficient aerial, terrestrial, and aquatic locomotion with seamless transition between the modes. The vehicle has a mass of 1010 g and solely relies on its four cycloidal rotors (cyclorotors) to achieve all modes of locomotion. The cyclorotor rotational speeds and thrust vectors are individually modulated to sustain stable hover in aerial mode. A similar control strategy using aerodynamic forces generated by cyclorotors is also implemented for aquatic locomotion. The wheels are efficiently integrated into the carbon fiber rotor endplates since cyclorotors rotate about the horizontal axis. As a result, the cyclocopter maneuvers in terrestrial mode by directly relying on motor torque. Seamless transition is accomplished using a retractable landing gear system equipped with polystyrene foam pontoons. In aerial mode, the cyclorotors operate at 1550 rpm and consume 232 W to sustain hover. Forward translation at 2 m/s in terrestrial mode requires 28 W, which is a 88% reduction from hover. In aquatic mode, the cyclorotors operate at 348 rpm and consume 19 W, which is a 92% reduction from hover. Overall, a versatile platform capable of multimodal operation is successfully demonstrated with only a modest addition in total mass.

To expand the cruise speed of a compound helicopter, alleviating the compressibility effects on the advancing side with reduced rotor RPM is proved to be an effective design feature, which results in high advance ratio flight regime. To investigate the aerodynamic phenomena at high advance ratios and provide data for the validation of analytical tools, a series of wind tunnel tests were conducted progressively in the Glenn L. Martin Wind Tunnel with a 33.5-inch radius fourbladed articulated rotor. In a recent wind tunnel test, the rotor blades were instrumented with pressure sensors and strain gauges at 30% radius, and pressure data were acquired to calculate the sectional airloads by surface integration up to an advance ratio of 0.8. The experimental results of rotor performance, control angles, blade airloads, and structural loads were compared with the predictions of comprehensive analysis and computational fluid dynamics (CFD) analysis coupled with computational structural dynamics (CSD) structural model. The paper focuses on the data correlation between experimental pressure, airload, and structural load data and the CFD/CSD predicted results at various collective and shaft tilt angles. Overall, the data correlation was found satisfactory, and the study provided some insights into the aerodynamic mechanisms that affect the rotor airload and performance, in particular the mechanisms of backward shaft tilt, the effect of hub/shaft wake, and the formation of dynamic stall in the reverse flow region.

A swashplateless rotor trim using brushless DC motor actuated trailing edge flaps was achieved in the Glenn L. Martin wind tunnel. A 6-ft-diameter, four-bladed articulated rotor with motor–flap system integrated into the NACA 0012 airfoil section was fabricated. A Maxon EC-10 brushless DC motor as an on-blade actuator and a lightweight mechanism were incorporated to actuate the trailing edge flap. The rotor torsion frequency was lowered to 2/rev using soft pitch links, allowing the blade pitch response to a trailing edge flap input. A closed-loop controller was employed to ensure trailing edge flap operation at desired amplitude with correct phase difference and in sync with the rotor azimuth. A trim methodology was implemented, and wind tunnel trim was achieved at 900 and 1200 RPM for a number of advance ratios. Results show that the brushless DC motors can provide sufficient primary control authority and have structural strength to withstand centrifugal loads, while fitting within airfoil profile and incurring minimal weight penalty.

This paper presents a methodology for preliminary sizing of unconventional rotorcraft using a physics-based approach to estimate the weight of primary load-carrying members and rotor efficiencies. The methodology is demonstrated for a quadrotor biplane tailsitter, a tilt-body configuration that can operate in both helicopter and airplane mode. A beam lattice framework for the airframe structure is iteratively adjusted in the sizing loop to accommodate the limit loads. A similar semianalytical approach is followed to size and estimate weight of the rotor blades. Using this analysis, a consistent combination of vehicle macrodimensions (rotor radius, wing span) and tip speed as well as detailed design parameters (spar height, skin thickness, and cross-section weight) are obtained simultaneously. To compare the effectiveness of various power plants within a weight class, the sizing methodology was modified to identify the payload for three different vehicle takeoff weights: 20, 50, and 1000 lb. To enable operation within constrained urban canyons, the effect of restricting maximum vehicle dimensions to 10 ftfor the 1000-lb designs is also examined. An electric transmission model is used in these designs owing to its relative insensitivity of transmission efficiency to the operating RPM. A variable-pitch and variable-RPM rotor design allows for control redundancy within each rotor.

Owing to its ability to alleviate the compressibility effect on the advancing side, the slowed rotor operating at high advance ratios is a key feature in high-speed compound rotorcraft. A series of wind tunnel tests were conducted in the Glenn L. Martin Wind Tunnel with a four-bladed Mach-scaled articulated rotor. The objective of the tests was to gain a basic understanding of unique features of high-advance-ratio aerodynamic phenomena, such as thrust reversal and dynamic stall in the reverse flow region. In this study, high-advance-ratio tests were carried out with highly similar, noninstrumented blades and on-hub control angle measurements, to minimize possible error due to blade structural dissimilarity and pitch angle discrepancy. The tests were conducted at 900 and 1200 RPM, advance ratios of 0.3–0.9, and a shaft tilt study was conducted at±4°. Pitch and flap motion at the blade roots, rotor performance, and vibratory hub loads were investigated during the test. The test data were then compared with those of previous tests and with predictions from comprehensive analysis. The airload results were investigated using comprehensive analysis to gain insights into the influences of advance ratio and shaft tilt angle on rotor performance and hub vibratory loads. Results indicate that the thrust benefit from backward shaft tilt is dependent on the change in the inflow condition and the induced angle of attack increment, and the reverse flow region at high advance ratios is the major contributor to changes in shaft torque and horizontal force.

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.

During the past two decades, there has been major growth of small-unmanned aerial vehicles (sUAV) for hobbyists and rapidly expanding commercial and military applications. The impetus for this dramatic expansion has been due to the explosion of mobile technology in terms of microelectronics, data processing and transmission capability, high-energy-density batteries, miniaturized integrated programmable chips, and innovations in computer vision and videography/photography. However, there are many challenges to overcome before these small unmanned aerial systems (sUAS) can be used for routine commercial and military applications, which include reliability, sizable payload and range, stringent navigation/guidance requirements, and precision takeoff/landing and robust autonomous flight in constrained and low-altitude gusty environment. The objective of this paper is to cover state-of-the-art of sUAS and delivery drones, identify technology gaps and key scientific barriers, and present future research needs for high payoff applications.

A single main rotor helicopter's maximum forward speed is limited due to the compressibility effects on the advancing side and reverse flow and dynamic stall on the retreating side. Compound helicopters can address these issues with a slowed rotor and lift compounding. There is a scarcity of test data on compound helicopters, and the present research focuses on a systematic wind tunnel test on lift compounding. Slowing down the rotor increases the advance ratio and, hence, the reverse flow region, which does not produce much lift. The lift is augmented with a wing on the retreating side. A hingeless rotor hub helps to balance the rolling moment with lift offset. Wind tunnel tests were carried out on this configuration up to advance ratios of 0.7 at two different wing incidence angles. Rotor performance, controls, blade structural loads, and hub vibratory loads were measured and compared with in-house comprehensive analysis, UMARC. A comparison between different wing incidences at constant total lift provided many insights into the lift compounding. It increased the vehicle efficiency and reduced peak-to-peak lag bending moment and in-plane 4/rev hub vibratory loads. The only trade-off was steady rotor hub loads and rolling moment at the wing root carried by the fuselage.

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

A computational study has been conducted on various airfoils to simulate flows at Reynolds numbers (Re) primarily between 104 and 105 to provide understanding and guidance for MAV and other low-Rey...

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.

In this work, the aerodynamic performance and optimization of a vertical-axis wind turbine with a high tip-speed ratio are theoretically studied on the basis of the two-dimensional airfoil theory. By dividing the rotating plane of the airfoil into the upwind and downwind areas, the relationship among the angle of attack, azimuth, pitch angle, and tip-speed ratio is derived using the quasi-steady aerodynamic model, and aerodynamic loads on the airfoil are then obtained. By applying the polynomial approximation to functions of lift and drag coefficients with the angle of attack for symmetric and asymmetric airfoils, respectively, explicit expressions of aerodynamic loads as functions of the angle of attack are obtained. The performance of a fixed-pitch blade is studied by employing a NACA0012 model, and influences of the tip speed ratio, pitch angle, chord length, rotor radius, incoming wind speed and rotational speed on the performance of the blade are discussed. Furthermore, the optimization problem based on the dynamic-pitch method is investigated by considering the maximum value problem of the instantaneous torque as a function of the pitch angle. Dynamic-pitch laws for symmetric and asymmetric airfoils are derived.

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.

The paper focuses on understanding the mechanism of force production on a hover-capable flapping-wing system by using a combination of direct force measurements and flowfield studies. The experiments were conducted in air at a Reynolds number of approximately 25,000 (based on mean chord and maximum tip speed), which is the typical operating regime of small flapping-wing micro air vehicles. The forces and moments were measured using 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 vacuum forces were subtracted from the total forces to obtain the pure aerodynamic forces. Flow visualization and particle image velocimetry were used to characterize the formation, strength, and structure of the leading-edge vortex on the flapping wing. A rapid increase in wing lift coefficient was associated with the growth of the leading-edge vortex and a progressive reduction in lift coefficient with the convection ...

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...

One challenge that plagues the high-speed operation of rotorcraft is the inherent increase in size of the reverse flow region on the retreating side of the rotor disk. The present work experimentally investigated a portion of the reverse flow region on a 1.7- m-diameter sub-scale slowed rotor at advance ratios up to $$\mu =0.9$$ and three shaft tilt angles $${\alpha _{\text {s}}=-\ 4^\circ ,\,0^\circ ,\,4^\circ }$$ . Two-component time-resolved particle image velocimetry was used to characterize the flow field around a blade element in the reverse flow region, nominally positioned at $$\psi =270^\circ $$ and $${y/R=0.4}$$ . Four dominant flow structures were observed: the reverse flow entrance vortex, the blunt trailing edge wake sheet, the reverse flow dynamic stall vortex, and the tip vortex. Analyses are focused on the reverse flow dynamic stall vortex due to its influence on unsteady airloads, and similarities to that of a leading edge vortex over a rotating or surging wing at high angle of attack. The number of semi-chords traveled by the blade element through the reverse flow region directly affects the strength, trajectory, and estimated vortex-induced pitching moment of the reverse flow dynamic stall vortex. Shaft tilt angle also has a strong effect on the evolution of the vortex, with forward shaft tilt resulting in dramatically increased strength and size. The results of this characterization and sensitivity study serve as a reference for evaluation of numerical simulations and aid in the development of low-order aerodynamic models of the reverse flow region and resulting force histories. Collectively, these tools will be used to better predict airloads, wake interactions, and vibrations of rotors operating at high advance ratios.

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...