Hirofumi Igarashi

An Experimental Investigation on Aerodynamic Hysteresis of a Low-Reynolds Number Airfoil

( ) An experimental study was conducted to investigate the aerodynamic characteristics of a NASA low speed GA(W)-1 airfoil at the chord Reynolds number of ReC=160,000. Aerodynamic hysteresis was observed for the angles of attack close to the static stall angle of the airfoil. In addition to mapping surface pressure distribution around the airfoil, a high-resolution PIV system was used to make detailed flow field measurements to quantify the occurrence and behavior of laminar boundary layer separation and transition on the airfoil when aerodynamic hysteresis occurs. The flow field measurements were correlated with the airfoil surface pressure measurements to elucidate underlying fundamental physics. For the same angle of attack in hysteresis loop, the flow obtained along the increasing angle branch was found to result in an almost attached flow with small unsteadiness, higher lift and lower drag, whereas the one with decreasing angle of attack branch was associated with large unsteadiness, lower lift, and higher drag. The hysteresis was found to be closely related to the behavior of the laminar boundary layer separation and transition on the airfoil. The ability of the flow to remember its past history is believed to be responsible for the hysteretic behavior.

Aerodynamic hysteresis of a low-Reynolds-number airfoil

L OW-REYNOLDS-NUMBER airfoil aerodynamics is important for bothmilitary and civilian applications. The applications include propellers, sailplanes, ultralight man-carrying/man-powered aircraft, high-altitude vehicles, wind turbines, unmanned aerial vehicles (UAVs), and micro air vehicles (MAVs). For the applications just listed, the combination of small length scale and low flight velocities results in airfoils operating at low chord Reynolds numbers of Re < 500; 000. It is well known that many significant aerodynamic problems occur below chord Reynolds numbers of about 500,000. Hysteresis phenomena have been found to be relatively common for roundnosed airfoils at low Reynolds numbers. Aerodynamic hysteresis of an airfoil refers to airfoil aerodynamic characteristics as it becomes history dependent, i.e., dependent on the sense of change of the angle of attack, near the airfoil stall angle. The coefficients of lift, drag, and moment of the airfoil are found to be multiple-valued rather than single-valued functions of the angle of attack. Aerodynamic hysteresis is of practical importance because it produces widely different values of lift coefficient and lift-to-drag ratio for a given angle of attack. It could also affect the recovery from stall and/or spin flight conditions. Whereas aerodynamic hysteresis associated with the pitchingmotion of airfoils (also known as dynamic stall) has been investigated extensively as summarized in the review article of McCorskey [1], hysteresis phenomena observed for static stall of an airfoil have received much less attention. Mueller [2] investigated the aerodynamic characteristics of Lissaman 7769 and Miley M06-13-128 airfoils at low Reynolds numbers, and found both airfoils produced hysteresis loops in the profiles of measured lift and drag forces when they operated below chord Reynolds numbers of 300,000. Based on qualitative flow visualization with smoke, he suggested that airfoil hysteresis is closely related to laminar boundary-layer transition and separation on the airfoils. Hoffmann [3] studied the aerodynamic characteristics of a NACA 0015 airfoil at a chord Reynolds number of 250,000, and hysteresis loopwas observed in themeasured coefficients of drag and lift. He also found that hysteresis was observed for low-freestream turbulence cases but disappeared for high-freestream turbulence cases. More recently, Mittal and Saxena [4] conducted a numerical study to predict the aerodynamic hysteresis near the static stall angle of a NACA 0012 airfoil in comparison with the experimental data of Thibert et al. [5]. In the present study, we report the measurement results of an experimental study to investigate aerodynamic hysteresis near the static stall angle of a low-Reynolds-number airfoil. In addition to mapping surface pressure distribution around the airfoil with pressure sensors, a high-resolution particle image velocimetry (PIV) system was used to make flowfield measurements to quantify the occurrence and behavior of boundary-layer transition and/or separation on the airfoil when aerodynamic hysteresis occurs. To the best knowledge of the authors, this is the first effort of its nature. The primary objective of the present study is to gain further insight into fundamental physics of aerodynamic hysteresis. In addition, the quantitative flowfield measurements will be used as the database for the validation of computational fluid dynamics (CFD) simulations of such complex phenomena for the optimum design of low-Reynoldsnumber airfoils.

A Stereoscopic PIV Study of a Near-field Wingtip Vortex

An experimental study was conducted to investigate the wing-tip vortex generated from a rectangular NACA0012 airfoil model. A high-resolution Stereoscopic Particle Image Velocimetry (SPIV) system was used to conduct detailed flow field measurements to quantify the transient behavior of the wing-tip vortex in the near field. The characteristics of the vortex wandering phenomena, i.e., the slow side-to-side movement of the wing-tip vortex core, were revealed in great detail based on the SPIV measurement results.

An Experimental Study of Unsteady Vortex Structures in the Wake of a Piezoelectric Flapping Wing

An experimental study was conducted to explore the potential applications of compact, gearless, piezoelectric flapping wings with the wing size, stroke amplitude and flapping frequency within the range of actual insect characteristics for the development of novel insect-sized, flapping-wing-based Nano-Air-Vehicles (NAVs). Unlike most of previous studies with 2-D flapping airfoil models, a fix-rooted 3-D piezoelectric flapping wing was used in the present study with the consideration of more practical configurations usually used in NAV designs. The experimental study was conducted in a low-speed wing tunnel with the test parameters of chord length of C = 12.7mm, chord Reynolds number of Re = 1,200, flapping frequency of f = 60 Hz, reduced frequency of k = 3.5, and non-dimensional flapping amplitude at wingtip h = A/C = 1.3. The corresponding Strouhul number of the root-fixed 3-D piezoelectric flapping wing is Str = 0.30, which is within the optimal range of 0.2 < Str < 0.4 usually used by flying insects and swimming fishes. A digital particle image velocimetry (PIV) system was used to achieve phased-locked and time-averaged flow field measurements to quantify the formation and separation processes of the Leading Edge Vortex (LEV) structures on the upper and lower surfaces of the flapping wing in relation to the phase angle (i.e., the positions of the flapping wing) during upstroke and down stroke flapping cycles. The evolutions of the wake vortex structures in the chordwise cross planes at different wingspan locations of the rootfixed flapping wing were compared quantitatively to elucidate underlying physics for better understanding of the unsteady aerodynamics of the flapping flight and to explore/optimize design paradigms for the development of novel insect-sized, flapping-wing-based NAVs.

Wind tunnel effects on wingtip vortices

The influence of wind tunnel boundaries on a wing tip vortex is studied experimentally and by computation. Computations show that vortex is similar to that which would exist without tunnel walls. The experimental data are corrected for meandering. Then computations are in close agreement to data. A computation with the wall moved closer to the tip shows the primary effect to be displacement of the vortex core by the image effect. Wind tunnel experiments have studied the dynamics of the initial rollup from a stationary wing, and the subsequent development of axial and tangential velocities and turbulence with downstream distance. Devenport et al. (1996) reported that the flow outside the vortex core was dominated by the remainder of the wing wake, which wound into an ever increasing spiral, and that the turbulence stress levels varied along the wake spiral in response to the varying rate of strain imposed by the vortex. Chow et al. (1997) investigated the wing-tip vortex of a NACA 0012 airfoil model. They indicated a high level of axial velocity, in excess of 1.7U ∞ at all measurement locations. They also reported that the turbulence intensity in the vortex can be as high as 24%, but it decayed quickly with streamwise distance because of the stabilizing effect of the nearly solid-body rotation within the vortex-core. Many other detailed properties of the tip vortex have been measured in wind tunnel experiments. These studies have uncovered a good deal of useful information. However, the majority of them used point-wise flow measurement techniques. A common shortcoming of point-wise measurements is the inability to provide spatial structure of the unsteady vortices. Full field measurements are needed to effectively reveal the transient behavior of the wing-tip vortex structures. Temporally synchronized and spatially resolved flow field measurements are highly desirable in order to elucidate underlying physics. In a companion paper we describe the instrumentation for our current experiments. It consists of high resolution, stereoscopic PIV. Here we report preliminary data and CFD comparisons. It is well known that wind tunnel walls interfere with the flow around a test model. That is a major source of uncertainty in simulating free flight conditions. Various correction methods have been proposed for flows with concentrated vortices. For instance, Wang and Coton (2000) suggested combining a low-order panel method and a prescribed wake vortex into a coupled model to assess the basic affect of wind tunnel walls on wind turbine flow and performance. They reported that the numerical results from the coupled model compared well with the wind tunnel test data, although some discrepancies were noted. However, very little can be found in the literature to evaluate wind tunnel confinement effects on wing-tip vortices. Tunnel effects include test model blockage, inlet flow non-uniformities and inlet turbulence levels. Questions about how similar measured characteristics are to those occurring in free flight remain. Confinement effects provide one focus of the present experimental and numerical study.

An Experimental Study of Stall Hysteresis of a Low-Reynolds-Number Airfoil

An experimental study was conducted to investigate static stall hysteresis of a NASA GA(W)-1 airfoil at the chord Reynolds number of Re = 162,000. In addition to mapping the surface pressure distribution around the airfoil, a digital PIV system was used to make detailed flowfield measurements to quantify the occurrence and behavior of laminar boundary layer separation and transition on the airfoil when static stall hysteresis occurs. The measurement results revealed clearly that, for a same angle of attack in the hysteresis loop, incoming flow streams were found to be able to attach to the airfoil upper surface in general with a thin separation bubble formed near the airfoil leading edge when the angle is at the increasing branch of the hysteresis loop. The attached flow pattern resulted in higher lift and lower drag acting on the airfoil as well as a lower Reynolds stress level and smaller unsteadiness in the wake of the airfoil. When the angle of attack is at the decreasing branch of the hysteresis loop, the laminar boundary layer was found to separate from the airfoil upper surface for good at a location very near to the airfoil leading edge. The turbulence transition of the separated laminar boundary layer was found to take place rapidly accompanied by periodical shedding of strong Kelvin-Helmholtz vortex structures in the wake of the airfoil. Large-scale flow separation was found to take place on almost the entire upper surface of the airfoil. As a result, the lower lift and higher drag acting on the airfoil were found with Reynolds stress and turbulence kinetic energy levels in the wake of the airfoil increased significantly. The present study elucidates quantitatively that static stall hysteresis of a low-Reynolds number airfoil is closely related to the behavior of laminar boundary layer separation and transition on the airfoil. The ability of the flow to remember its past history is responsible for the stall hysteretic behavior.

The Effects of Wind Tunnel Walls on the Near-field Behavior of a Wingtip Vortex

An experimental study was conducted to investigate the behavior of the wing-tip vortex structures generated by a square-tipped, rectangular NACA0012 wing. A Stereoscopic Particle Image Velocimetry (SPIV) system was used to conducted detailed flow field measurements to elucidate the key features of the wing-tip vortex structures in the near field. One of the great advantages of the present SPIV measurements over the classical measurement technique is that the vortex wandering can be removed directly by tracking the center of the wingtip vortex in the instantaneous measurement frames. By tracking the center of the wingtip vortex, the wandering and turbulence in the vortex can be decoupled completely. This method was applied to investigate the effects of the angle of attack of the test wing and wind tunnel wall on the evolution of the wingtip vortex in the near field. In order to decouple the effects of vortex wandering, Devenport et al. (1996) suggested an analytical method to predict the wandering free velocity profile. The velocity profile predicted by the Devenport et al. (1996) method was compared with the SPIV re-centered velocity profile quantitatively.

AN EXPERIMENT STUDY OF WALL SLOT JETS PERTINENT TO TRAILING EDGE COOLING OF TURBINE BLADES

An experimental study was conducted to quantify the flow characteristics of wall jets pertinent to trailing edge cooling of turbine blades. A high-resolution stereoscopic PIV system was used to conduct detailed flow field measurements to quantitatively visualize the evolution of the unsteady vortex and turbulent flow structures in cooling wall jet streams and to quantify the dynamic mixing process between the cooling wall jet streams and the main stream flows. The detailed flow field measurements are correlated with the adiabatic cooling effectiveness maps measured by using pressure sensitive paint (PSP) technique to elucidate underlying physics in order to improve cooling effectiveness to protect the critical portions of turbine blades from the harsh ambient conditions.Copyright