Gregg Abate

Flight Dynamics of a Flapping-Wing Air Vehicle

The research and development efforts presented in this paper address the flight dynamics of a flapping-wing air vehicle (ornithopter). The 74-cm-wing-span ornithopter was equipped with an automatic flight control system that provides stability augmentation and navigation of the vehicle and flight data acquisition. Wind tunnel tests were conducted with the control surfaces fixed in the trimmed position and flapping motion of the wings activated by a motor at a constant throttle setting. Coefficients of vertical and horizontal force, and pitching moment were determined at a free stream velocity of 7.25 m/sec, and the angle of the stroke plane varied from 0 to 40 degrees. A series of flight tests were conducted with fixed controls, demonstrating ornithopter stability in all axes. Proportional control laws were programmed into the autopilot for the closed-loop controls. A number of flights of the autonomous ornithopter were conducted with the telemetry acquisition. During the autonomous flights, the ornithopter performed waypoint and altitude navigation, demonstrating stable performance.

Dynamic structure of confined shocks undergoing sudden expansion

Abstract The gas dynamic phenomenon associated with a normal shock wave within a tube undergoing a sudden area expansion consists of highly transient flow and diffraction that give rise to turbulent, compressible, vortical flows. These interactions can occur at time scales typically ranging from micro- to milliseconds. In this article, we review recent experimental and numerical results to highlight the flow phenomena and main physical mechanisms associated with this geometry. The topics addressed include time-accurate shock and vortex locations, flowfield evolution and structure, wall-shock Mach number, two- vs. three-dimensional sudden expansions, and the effect of viscous dissipation on planar shock-front expansions. Between axisymmetric and planar geometries, the flow structure evolves very similarly early on in the sudden expansion process (i.e., within the first two shock tube diameters). Both numerical and experimental studies confirm that the trajectory of the vortex formed at the expansion corner is convected into the flowfield faster in the axisymmetric case than the planar case. The lateral propagation of the vortices correlates very well between axisymmetric and planar geometries. In regard to the rate of dissipation of turbulent kinetic energy (TKE) for a two-dimensional planar shock undergoing a sudden expansion within a confined chamber, calculations show that the solenoidal dissipation is confined to the region of high strain rates arising from the expansion corner. Furthermore, the dilatational dissipation is concentrated mainly at the curvature of the incident, reflected, and barrel shock fronts. The multiple physical mechanisms identified, including shock–strain rate interaction, baroclinic effect, vorticity generation, and different aspects of viscous dissipation, have produced individual and collective flow structures observed experimentally.

Validation of a Low Reynolds Number Aerodynamic Characterization Facility

The characterization of a new wind tunnel at the University of Florida’s Research and Engineering Education Facility is presented. The tunnel is specifically designed to operate in the low Reynolds number regime in which Micro Air Vehicles operate. The wind tunnel is driven by a 60 inch, 50 hp axial blower controlled by a variable frequency drive. The tunnel entrance consists of a flow conditioning section and a 8:1 area contraction ratio that results in a 42” square entrance to the open jet test section. The enclosure surrounding the test section has a volume of nearly 2000 ft with an axial length of 10 ft. The free stream velocities in the tunnel range from nominally 0 to 22 m/s by altering the frequency on the variable frequency drive. Flow uniformity studies for free stream velocities of 2 and 15 m/s were conducted demonstrating a potential core throughout the test section of at least 60% of the 1.14 m contraction exit. Experiments using hot wire anemometry were also performed and the turbulence intensities were found to be less 0.22% for free stream velocities greater than 1m/s.

An Experimental Study of Flexible Membrane Wings in Flapping Flight

An experimental study was conducted to assess the aerodynamic benefits of using flexible membrane wings for the development of flapping-wing Micro-Air-Vehicles (MAVs). The overall aerodynamic performances (i.e. time-averaged lift and thrust/drag generation) of two flexible membrane wings with different skin flexibility (i.e., a flexible nylon wing and a more flexible latex wing) were compared with those of a conventional rigid wing in order to assess the effects of skin flexibility (rigidity) of the tested wings on their aerodynamic performances for flapping flight applications. The measurement results revealed clearly that, for all the tested wings, flapping motion would bring significant aerodynamic benefits when the flapping flight is in unsteady state regime with advance ratio of the flapping flight being smaller than 1.0. The aerodynamic benefits of flapping flight were found to decay exponentially with the increasing advance ratio. Flapping motion was found to become detrimental for high speed flight applications. The skin flexibility (rigidity) of the tested wings was found to have considerable effects on their aerodynamic performances for both soaring flight and flapping flight applications: The flexible membrane wings were found to have better overall aerodynamic performance (i.e., lift-to-drag ratio) over the rigid wing for soaring flight, especially for high speed soaring flight applications. The rigid wing was found to have better lift production performance for flapping flight in general. The latex wing, which is the most flexible among the three tested wings, was found to have the best thrust generation performance for flapping flight. The less flexible nylon wing, which has the best aerodynamic performance for soaring flight applications, was found to be the worst for flapping flight applications.

Transonic Aerodynamic Analysis of Lattice Grid Tail Fin Missiles

2 ) at Mississippi State University. The impetus for the study was the increased use of lattice grid tail fins for control devices in tactical and strategic missiles, and aerodynamic data from free flight tests of a sub-scale model performed at the Aeroballistic Research Facility (ARF) at Eglin AFB, Florida. The CFD results elucidated the flowfield development in the lattice grid fin cells as Mach numbers increased from 0.744 to 1.19. At a critical transonic Mach number a normal shock forms at the back of some of the lattice grid cells, and causes the flow to choke with a significant increase in overall drag. The comparisons to ARF test data and analytic predictions for choked flow were favorable. Nomenclature

Aerodynamics of Cambered Membrane Flapping Wings

The aerodynamics of cambered membrane flapping wings is the focus of this paper. A cambered airfoil was introduced into the wing by shaping metal ribs attached to the membrane skin of the 25-cm-wing-span model. Tests in still air of the flapping wings oriented horizontally and vertically with respect to the gravitational field show no effects on generated aerodynamic forces. The thrust force generated by a 9% camber wing is found to be 30% higher than that of the same size flat wing. The aerodynamic forces and pitching moment generated by flapping wings were measured in a wind tunnel with the flapping wing angle of attack varying from horizontal to vertical. Cambered wings show significantly higher lift and thrust in comparison with flat wings. Adding a dihedral angle to the wings and keeping the flapping amplitude constant improved the cambered wing’s performance even further. The aerodynamic coefficients are defined using a reference velocity as a sum of two components: a free stream velocity and a stroke-averaged wing tip flapping velocity. The lift, drag, and pitching moment coefficients obtained using this procedure collapse well for studied advance ratios, especially at lower angles of attack.

Flight dynamics of flapping-wing air vehicle

The research and development efforts presented in this paper address the flight dynamics of a flapping-wing air vehicle (ornithopter). The 74-cm wing span ornithopter was equipped with the automatic control system that provides the stability augmentation and navigation of the vehicle, and flight data acquisition. Wind tunnel tests were conducted with the control surfaces fixed in neutral position and flapping motion of the wings activated by a motor at a constant throttle setting. Coefficients of a lift, drag, and pitching moment were determined at a free stream velocity of 7.2 m/sec and the angle of attack varied from 0 to 41 degrees. In addition, variations of derivatives of aerodynamic coefficients with the freestream velocity were investigated. A series of flight tests were conducted with fixed controls demonstrating ornithopter stability in all axes. Proportional control laws were programmed into the autopilot for the closed-loop controls. A number of test flights of the autonomous ornithopter were conducted with the telemetry acquisition. During the autonomous flights, the autopilot performed waypoint and altitude navigation demonstrating stable performance.

Longitudinal aerodynamics of rapidly pitching fixed-wing micro air vehicles

Fixed-wing vertical-takeoff-and-landing (VTOL) micro air vehicles (MAVs) can transition between two flight modes: forward-flight and near-hover. This study was conducted to improve the transition-performance of such vehicles. Our test subject consists of a rigid Zimmerman wing with an S-shaped airfoil and a propulsion system with contrarotating propellers. The wing had an aspect ratio of 1.9 and was subjected to freestream and slipstream flows while being rapidly pitched about its aerodynamic center (AC). Data was acquired for non-dimensional pitching rates ranging from -0.031 to 0.031 at an average freestream Reynolds number of 86K. Two elevator deflections were used, 0 and -17 deg, as well as three different propulsive-settings corresponding to propulsion-off and average advance ratios of 0.47 and 0.60. Under steady conditions, with throttle-setting constant, propeller rotation-rate decreases linearly with angle of attack (AOA) from 20 to 70 degrees, while both advance-ratio and thrust-coefficient increase linearly. Normal force coefficient was found to increase with AOA. Under steady conditions, higher throttle-setting results in greater stall delay, which causes maximum lift coefficient to increase significantly. When throttle-setting increases, lift and drag coefficients increase throughout the tested AOA-domain. Interestingly, wing aerodynamic efficiency is virtually independent of throttle-setting between 30 and 70 degrees. Rapid-pitching tests showed that nose-up pitching delays stall and nose-down pitching hastens it. Lift and drag coefficients increase with positive pitching-rate and decreased with negative pitching-rate. Unsteady efficiency curves converge on steady curves near 30 degrees AOA and remained converged thereafter. The pitch-damping observed in flight tests was confirmed by acquired data. The propulsions system was not sensitive to rapid-pitching; thrust, normal force and propulsive moment were relatively unchanged.

Experimental Data for Micro Air Vehicles with Pliant Wings in Unsteady Conditions

This paper describes a new system for dynamic wind-tunnel testing for the purpose of estimating the dynamic stability derivatives of micro air vehicles. The research is presented for two geometrically identical Zimmerman wing configurations: rigid and flexible. A two degrees-of-freedom rig permits the measurement of the two individual components of the rotary damping moment by manipulating the model in a pure plunging motion or a combined pitching/plunging motion. A modern design of experiments methodology was used to elucidate the correlation between the wings’ structural deformations and the aerodynamic damping characteristics of the rigid and flexible wings. In the case of a flexible wing, the latex membrane elastic deformations are measured using visual image correlation whereby the pre-tension strain state is characterized prior to the aerodynamic tests. Comparisons of lift and drag between the rigid and flexible wings in static and dynamic conditions showed that the flexible wing with the elastic membrane skin with medium tension provided the best lift-to-drag ratio. Specific patterns in the aerodynamic coefficients were observed in the presence of dynamic changes in angle of attack or pitch angle. The increase of α& and θ & is shown to be a significant factor in the response of the MAV wing, particularly on the pitching moment coefficient.

An experimental and numerical investigation of fast transient gas dynamics

An experimental and numerical investigation has been performed to study the evolution of shock waves undergoing a sudden expansion in one direction while restricted in the second. Experimental data are gathered and studied for shock waves undergoing the sudden 4:l area expansion in air for Mach numbers of 1.5 and 2.0. Detailed, time-accurate measurements of the shock wave and vortex core location as well as wall pressure data are presented. In addition, the evolving flow structure through the time-accurate flowfield imagery is also presented. The results of these experiments are compared to twodimensional numerical simulations specific to the Mach 1.5 and 2.0 initial conditions and geometry. The direct comparisons of the experimental work and numerical simulations provide insight into flowfield phenomena such as viscous dissipation and sh&&/vortex interaction. The data presented in this effort further elucidates key modeling questions by providing time-accurate flow visualization and pressure data of a two-dimensional shock wave undergoing a sudden expansion in a confined chamber. * Aerospace Engineer, Senior Member AIAA ’ Senior Principal Consultant, Associate Fellow AIAA ’ Research Engineer, Member AIAA ’ Professor & Department Chairman, Associate Fellow AIAA ‘I Associate Professor ’ Associate Professor, Senior Member AIAA ‘* Assistant Professor This paper is declared a work of the U.S. Government and is not subject to copyright protection in the United States. Background Gas dynamics of fast, transient characteristics includes issues such as shock wave reflections, vorticity production, shock-vortex interaction, energy exchange between shock and turbulence, and shock focusing (Chang and Kim, 1995). Figure 1 illustrates the sudden area expansion of a shock wave within a confined domain indicating these interesting fluid dynamic phenomena. These phenomena occur at time scales ranging from microseconds to milliseconds. This paper presents a focused look at sudden shock wave expansion within a confined chamber and the interaction of that shock wave with other fluid dynamic features. Both experimental and numerical tools have been adopted. Recent investigations of this type of gas dynamic flow, given by Jiang et al. (1997), present experimental and numerical simulations of a circular shock tube generated flow undergoing a sudden expansion. Figure 1 highlights the salient flow features of such a sudden expansion. Here it is seen that as the initial shock wave travels down the shock tube (Figure 1) it will undergo a sudden area expansion into a confined area (Figure lb-d), the flowfield gives rise to two important features. First, a “starting jet”, characterized by the formation of a vortex ring, Mach disk, and shear layer’ appear. Second, this precursor shock reflects off the confining walls of the expansion chamber and interacts with the starting jet flowfield. It is clearly seen in Figure 1 that this reflected shock splits once it interacts with the primary vortex ring, which gives rise to secondary shock waves. This shock/vortex/shear layer interaction is also shown in Figure 1. Here the reflected shock splits when it interacts with the primary vortex ring. It was also American Institute of Aeronautics and Astronautics (c)2000 American Institute of Aeronautics & Astronautics or published with permission of author(s) and/or author(s)’ sponsoring organization. observed by Jiang et al. (1997) that once the reflected shock wave passes through the developing shear layer that the shear layer splits, it is moved into the jet flow, and it is convected downstream. This allows another shear layer to form from the jet exit forming another weaker vortex ring. This is referred to as shear layer splitting and is largely an inviscid phenomenon (Jiang et al., 1997) Another phenomenon noted by Jiang et al. (1997) is the shock/shear layer interactions. It is noted that a vortex ring in the shear layer induced a shock wave and the shock wave separated the vortex ring from the shear layer. This is most often noted in strongly expanded free jets. Due to the strong influence of vortex dynamics and compressibility, the turbulence structure can be fruitfully assessed via an evaluation of viscous dissipation.