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A set of pure sinusoidal pitch and plunge oscillations for a symmetric airfoil is studied in this paper. A gridless Lagrangian vorticity particle-based solver is used to simulate the unsteady flowfield. Various kinematics chosen here are inspired by the experimental and computational studies presented in earlier works. The Lagrangian tool is used to accurately reproduce the flowfield obtained by earlier experimental results. It follows the experimental trend better compared with some of the grid-based solvers' results from earlier studies. The strength of this tool is its grid-free nature, as the resolution of the grid is a crucial parameter in resolving the unsteady flowfield accurately. This study also investigates the effects of the starting condition in dictating the wake deflection mode at short and long terms and confirms the findings of the earlier works. The effect of mean angle of attack on the wake deflection is also highlighted. One of the main questions considered here is whether there exists a kinematic equivalence between sinusoidal pitch and plunge. Effective angle of attack, quasi-steady criterion, and Theodorsen's criterion of unsteady aerodynamics have been considered here for kinematic equivalence with plunge. The latter two give a reasonably good match both in terms of flowfield and load as was also reported in the literature. This behavior is seen in two different plunge stroke amplitude cases studied here. The success of the quasi-steady and Theodorsen's approach indicates the general wake behavior being an inviscid phenomenon, at least for the chosen range of parameters. DOI: 10.2514/1.J050103

This paper presents an unsteady discrete adjoint algorithm for high-order implicit discontinuous Galerkin discretizations in time-dependent inviscid flow problems. The major function of the adjoint approach is to obtain the sensitivity information in a time-dependent functional output, which in turn is used to drive an unsteady shape- optimization process to deliver a minimum of the objective functional. A gradient-based optimization strategy is investigated, in which the sensitivity derivatives of the objective functional with respect to input variables are formulated in the context of high-order discontinuous Galerkin discretizations, while special emphasis is given to the variations and linearizations of curvilinear boundary elements. Implicit temporal discretizations consisting of a second-order backward Euler scheme and a fourth-order implicit Runge--Kutta scheme are considered exclusively in this work, where the corresponding adjoint problem is required to be solved in a backward time-integration manner due to the associated transpose operation. Two numerical examples for the unsteady shape design techniques are presented to verify the derived sensitivity formulations and to demonstrate the performance of the adjoint approach; the first involves an inverse shape-optimization case by matching a time-dependent target pressure profile for a two-dimensional periodic vortical gust impinging on a RAE-2822 airfoil, and the second considers minimization of the acoustic noise produced by subsonic flow over a NACA0012 airfoil with a 0.03c thick blunt trailing edge. DOI: 10.2514/1.J050444

The efficacy of a single dielectric barrier discharge plasma actuator for controlling turbulent boundary-layer separation from the deflected flap of a high-lift airfoil is investigated between Reynolds numbers of 240,000 (15 m/s) and 750,000 (45 m/s). Momentum coefficients for the dielectric barrier discharge plasma actuator are approximately an order of magnitude lower than those usually employed for such studies, yet control authority is still realized through amplification of natural vortex shedding from the flap shoulder, which promotes momentum transfer between the freestream and separated region. This increases dynamic loading on the flap and further organizes turbulent fluctuations in the wake. The measured lift enhancement is primarily due to upstream effects from increased circulation around the entire model, rather than full reattachment to the deflected flap surface. Lift enhancement via instability amplification is found to be relatively insensitive to changes in angle of attack, provided that the separation location and underlying dynamics do not change. The modulation waveform used to excite low- frequency perturbations with a high-frequency plasma-carrier signal has a considerable effect on the actuator performance. Control authority decreases with increasing Reynolds number and flap deflection, highlighting the necessity for further improvement of plasma actuators for use in realistic takeoff and landing transport aircraft applications. These findings are compared to studies on a similar high-lift platform using piezoelectric-driven zero- net-mass flux actuation. DOI: 10.2514/1.J050452

This paper investigates the impact of small-scale unsteadiness on adjoint-based output sensitivity analysis. In particular, when iterative methods for nonlinear flows fail to converge to a steady state, it is demonstrated that the resulting sensitivity analysis can be highly inaccurate, even when the unsteadiness in the outputs is small. The specific example considered is the viscous subsonic flow around an airfoil over a range of angles of attack. If a strengthened solver is used to solve the adjoint equation (even though the flow equations did not fully converge), it is demonstrated that the sensitivity of the lift with respect to angle of attack can vary significantly, due to linearizing about different solution iterates of the steady flow solver. Further, the unsteady iterates from the time-inaccurate steady-state solver to the time-accurate solution are compared. The unsteadiness of the time-accurate solution is markedly different from the iterate solutions of the steady-state solver. If a strengthened solver is applied to the nonlinear flow equations, steady solutions can be achieved whose lift is significantly different from the time-averaged lift of the time-accurate simulations. Time-accurate unsteady adjoint analysis is then shown to provide accurate sensitivities for the time-averaged lift. DOI: 10.2514/1.J050412

A new far-field method is proposed for the analysis of the aerodynamic force acting on an airfoil in subsonic unsteady flow conditions given a numerical solution of the flowfield. For this purpose, a recent theory proposed by Wu et al. (Wu, J.-Z., Ma, H.-Y., and Zhou, M.-D., Vorticity and Vortex Dynamics, Springer, New York, 2006), relating the aerodynamic force acting on a body to the Lamb vector (the cross product of the vorticity by the velocity) in case of unsteady Navier--Stokes flow, has been extended to the analysis of high-Reynolds-number turbulent flows governed by the Reynolds-averaged Navier--Stokes equations. The applications showed here concern the analysis of numerical solutions around an airfoil in steady flow and in pitching oscillations including a dynamic-stall condition. Both lift and drag coefficients have been computed by means of Lamb vector integrals and have been compared with the reference near-field results (stress integration on the body). The method allows for the decomposition of the aerodynamic force in a contribution associated with the bound vorticity on the body and a contribution related to the free vorticity developing in the airfoil wake. DOI: 10.2514/1.J050326

The transitory response of the flow over a stalled, two-dimensional (NACA 4415) airfoil to pulsed actuation on time scales that are an order of magnitude shorter than the characteristic convective time scale is investigated experimentally (Re = 570, 000). Actuation is effected by momentary [O(1 ms)] pulsed jets that are generated by a spanwise array of combustion-based actuators integrated into the center section of the airfoil. The flowfield in the cross-stream plane above the airfoil and in its near wake is computed from multiple high-resolution particle image velocity images that are obtained phase locked to the actuation waveform and allow for tracking of vorticity concentrations. The brief actuation pulse leads to a remarkably strong transitory change in the circulation about the entire airfoil that is manifested by a severing of the separated vorticity layer and the subsequent shedding of a large-scale clockwise vortex that forms the separated flow domain. The clockwise severed vorticity layer that follows behind this detached vortex has a distinct sharp streamwise edge that grows and rolls up as the layer is advected along the surface. It is shown that the shedding of the severed vortex and the accumulation of surface vorticity are accompanied by a transitory increase in the magnitude of the circulation about the airfoil that lasts 8-10 convective time scales. The attached vorticity layer ultimately lifts off the surface in a manner that is reminiscent of dynamic stall, and the flow separates again. DOI: 10.2514/1.45904

A reduced-order nonlinear unsteady aerodynamic modeling approach suitable for analyzing pitching/plunging airfoils subject to fixed or time-varying freestream Mach numbers is described. The reduced-order model uses kriging surrogates to account for flow nonlinearities and recurrence solutions to account for time-history effects associated with unsteadiness. The resulting surrogate-based recurrence framework generates time-domain predictions of unsteady lift, moment, and drag that accurately approximate computational fluid dynamics solutions, but at a fraction of the computational cost. Results corresponding to transonic conditions demonstrate that the surrogate-based recurrence framework can mimic computational fluid dynamics predictions of unsteady aerodynamic responses when flow nonlinearities are present. For an unsteady aerodynamic modeling problem considered in this study, an accurate reduced-order model was generated by the surrogate-based recurrence framework approach with significantly fewer computational fluid dynamics evaluations compared to results reported in the literature for a similar problem in which a proper-orthogonal-decomposition-based approach was applied. Furthermore, the results show that the surrogate-based approach can accurately model time-varying freestream Mach number effects and is therefore applicable to rotary-wing applications in addition to fixed-wing applications. DOI: 10.2514/1.J050471

The aeroacoustic sound generation process on a rod-airfoil configuration has been investigated by means of simultaneous particle image velocimetry in the near field and phased-microphone-array measurements in the far field. Up to 20,000 particle image velocimetry snapshots per field of view have been recorded. A coplanar multiplane particle image velocimetry system has been used, providing statistically independent samples of the temporal derivative of the velocity field. Both measurements were conducted in a synchronized manner so as to enable the calculation of the cross-correlation between the acoustic pressure and flow quantities derived from the measured velocity fluctuations. The main idea of the concept used in the study presented here was to use the coefficient matrix obtained from the aforementioned correlation to identify regularities in the near-field fluctuations that are related to the radiated sound field. Additionally, a technique to localize flow structures responsible for the far-field sound by means of the calculation of the near-field intensity has been tested. DOI: 10.2514/1.J050314

Direct numerical simulations have been conducted of the flow over NACA-0006 and NACA-0012 airfoils. For all cases multiple sources of noise have been observed. At low frequencies, the contribution of trailing-edge noise radiation is significant, while for the high frequencies, the radiated noise appears to be due only to flow events in the transition/reattachment region on the suction side. Cross correlations of acoustic and hydrodynamic quantities, combined with ray-acoustic theory, are used to identify the main source locations. While the physical origin of trailing-edge noise can be clearly identified using these methods, the situation is less clear for additional noise sources that radiate at high frequency. Evidence suggests that these additional noise sources are located downstream of transition on the suction side of the airfoil, in the region directly downstream of boundary-layer reattachment. DOI: 10.2514/1.J050345

A novel plasma circulation control technique is proposed to overcome the disadvantage of conventional jet circulation control about the requirement for air source. By employing the plasma actuator on the trailing edge of the airfoil, the plasma-induced wall jet is tangential to the airfoil surface and served as Coanda blowing jet, which makes the trailing-edge detachment point migrate to the lower surface of the airfoil. In this way, the circulation and the lilt of airfoil are directly controlled by the plasma actuator. The effectiveness of plasma circulation control on the aerodynamic characteristics improvement of NCCR 1510-7067N airfoil is demonstrated by solving the Reynolds-averaged Navier--Stokes equations. The plasma actuator is modeled with a phenomenologlcal model by adding the body force source terms to the governing equations. The results indicate that the efficiency of plasma circulation control for lift augmentation is superior to that of conventional jet circulation control method. The lift augmentation efficiency after the location optimization reaches [DELTA][C.sub.L]/[C.sub.[mu]] = 134.86, which is much higher than the maximum lift augmentation efficiency 80 for conventional jet circulation control. The optimal location for the plasma actuator in plasma circulation control is downstream of the boundary-layer separation point with proper distance, where the plasma actuator makes the separated shear layer reattach to the airfoil surface and delays the trailing-edge stagnation by Coanda effect. DOI: 10.2514/1.J050133

Two-and three-dimensional numerical simulations are performed of the flow around a circulation control airfoil (using a Coanda jet blowing over a rounded trailing edge) placed in a rectangular wind-tunnel test section. The airfoil model spans the entire tunnel and the span-to-chord ratio of the model is 3.26. The objective of this numerical study, in which we solve the compressible Reynolds-averaged Navier-Stokes equations in a time-resolved manner (but the solutions eventually converge to steady states), is to investigate the physical mechanisms of wind-tunnel sidewall effects on the flow, especially in the midspan region. The three-dimensional simulations predict that the Coanda jet flow is quasi-two-dimensional until the flow separates from the trailing edge of the airfoil; however, the spanwise ends of this Coanda jet sheet then three-dimensionally roll up on the side walls of the wind tunnel to form two large streamwise vortices downstream. Careful comparisons between the two- and three-dimensional simulations reveal that the wind-tunnel stream goes below the airfoil more in the three-dimensional cases than in the two-dimensional cases due to the presence of these two streamwise vortices downstream. This results in smaller lift and larger drag being produced at the midspan of the airfoil in the three-dimensional cases than in the two-dimensional cases. DOI: 10.2514/1.J050328

A combined experimental and computational study is presented for an airfoil undergoing a combined pitching and plunging motion at Reynolds number 100,000, where transition takes place along laminar separation bubbles. The numerical simulation approach addresses unsteady Reynolds-averaged Navier-Stokes solutions and covers transition prediction for unsteady mean flows. To study the effect of wing flexibility, the aerodynamic computational method is coupled with a structural solver using a Galerkin method. The numerical simulations are validated using high-resolution, phase-locked stereoscopic particle image velocimetry for one flapping case with a reduced frequency of k = 0.2. Hereby both a rigid and a flexible birdlike airfoil are investigated. The flow reveals strong unsteadiness resulting in moving laminar separation bubbles, both well captured by the numerical simulations performed in this study. DOI: 10.2514/1.J050158

A method is presented to obtain reduced-order models of multiparametric, steady, transonic aerodynamic flows around two-dimensional airfoils. The method is based on proper orthogonal decomposition and appropriate projection on the proper orthogonal decomposition manifold of the governing equations and boundary conditions, and it includes various ingredients that are new in this field. Proper orthogonal decomposition modes are obtained from a set of snapshots calculated by computational fluid dynamics based on the compressible Navier-Stokes equations and a turbulence model, but projection on the proper orthogonal decomposition manifold is made using the inviscid Euler equations, which makes the method independent of the computational fluid dynamics scheme; projection is made using only a limited number of mesh points, obtaining quite good results. Shock waves are treated specifically in the proper orthogonal decomposition description, to avoid the need of using a too-large number of snapshots. The method is checked and discussed on a specific aerodynamic problem. DOI: 10.2514/1.J050153

Earlier research for the NACA 643-618 airfoil has shown that lift-curve and stall behavior change noticeably when the chord Reynolds number is reduced significantly below its design value. For the same airfoil and a Reynolds number based on chord of Re = 64, 200, separation control by pulsed vortex generator jets and a two-dimensional volume forcing that was motivated by a plasma actuator were investigated. Pulsed vortex generator jets with moderate jet velocities and the two-dimensional volume forcing were found to introduce coherent spanwise structures that are highly effective for separation control. A linear stability analysis of the time-averaged flow showed that spanwise disturbances are amplified by an instability mechanism, which explains why both control strategies are highly efficient. Angled pulsed vortex generator jets with large jet velocities were found to introduce streamwise vortices and to result in earlier transition. Although the control is highly effective, it is not very efficient, as streamwise vortices are not supported by the flow. DOI: 10.2514/1.J050051

Vortex shedding and the associated noise radiation from a trailing edge were experimentally investigated for a leading-edge slat of a multi-element airfoil at stowed chord Reynolds number Re < 5.9 x [10.sup.5]. A particular focus was on the competition between the instability of the slat boundary layer excited by acoustic feedback and the absolute instability of the wake. Periodic vortex shedding was observed to occur from the slat trailing edge at the Reynolds numbers examined. For Re < 1.9 x [10.sup.5], the vortex shedding is governed by the absolute instability of the laminar wake of the slat without any distinct tonal noise radiation. For Re > 2.1 x [10.sup.5], however, acoustic feedback becomes pronounced between the trailing-edge noise and the boundary-layer instability waves on the suction surface, so that multiple spectral peaks appear both in the velocity fluctuations and sound pressure. At and around the Reynolds number for the first appearance of tonal noise, Re = 1.9 x [10.sup.5], both of the instability modes coexist. Beyond Re = 2.1 x [10.sup.5], the boundary-layer instability waves excited by the acoustic feedback evolve into high-intensity vortices before reaching the trailing edge and suppress the absolute instability of the wake through diminishing the reversed-flow region in the wake. DOI: 10.2514/1.45777

A theoretical and experimental aeroelastic study of a typical airfoil section with control surface free play for nonzero angle of attack in low subsonic flow is presented. The study includes the flutter and limit cycle oscillation behavior and also the linear and nonlinear aeroelastic responses excited by periodic gust loads. The theoretical approach uses Peters's finite state airloads model. The experimental investigation has been carried out in the Duke University wind tunnel using a rotating slotted cylinder gust generator. The theoretical and experimental results show that the serf-excited aeroelastic limit cycle oscillation is sensitive to the effect of initial pitch angle. For the gust response, the effect of initial pitch angle is smaller for the plunge and pitch responses and is larger for the flap response. The fair to good quantitative agreement between theory and experiment verifies that the present analytical approach has reasonable accuracy and good computational efficiency for nonlinear aeroelastic response analysis of such systems. DOI: 10.2514/1.44538

Large eddy simulation of flow around a NACA-0012 airfoil at a Reynolds number of 50,000 has been used to study the behavior of a laminar separation bubble near stall The effects of the subgrid-scale model and explicit filtering were studied for a test case in which direct numerical simulation results were available. It was found that a method incorporating a mixed-time-scale model in addition to explicit filtering gave improved results compared with a method with filtering alone. Statistical results as well as snapshots of the flow below stall exhibit good agreement with the direct numerical simulations. For a configuration near stall, the effect of the spanwise domain width was investigated by increasing the spanwise length from 20 to 50% chord. Two-point velocity correlations showed a significant improvement for the wider computational domain, in which the simulation was able to capture a low-frequency flow oscillation, in which intermittent bursting of the bubble was observed. The bubble bursting observed here is more irregular than in experiments at higher Reynolds number. The amplitude and frequency are compared with experimental results and with an unsteady viscons-inviscid interaction method which is shown to be capable of capturing unsteady behavior during stall. DOI: 10.2514/1.44298

This paper presents a reduced-order-modeling approach for nonlinear, multi-degree-of-freedom aerodynamic systems using multi-input Volterra theory. The method is applied to a two-dimensional, 2 degree-of-freedom transonic airfoil undergoing simultaneous forced pitch and heave harmonic oscillations. The so-called Volterra cross kernels are identified and shown to successfully model the aerodynamic nonlinearities associated with the simultaneous pitch and heave motions. The improvements in accuracy over previous approaches that effectively ignored the cross kernels by using superposition are demonstrated. DOI: 10.2514/1.38964

Experiments aimed at delaying flow separation through discrete jets pointing in the direction of streaming and sweeping side to side along the span were conducted on a V-22 airfoil with and without deflected trailing-edge flaps. The results indicated substantial drag reduction and lift increase at moderately low inputs of mass and momentum. Additional experiments were carried out on a semispan V-22 wing/nacelle combination, and they too provided an increase in lift-to-drag ratio LID of approximately 60 % (although active flow control was applied to the wing only). The effectiveness of the sweeping jets on reducing the download force acting on a V-22 full-span powered model in hover was also examined. A 29% reduction in download was realized using the embedded sweeping jets, corresponding approximately to a 2000 lb increase in hover lift. DOI: 10.2514/1.43663

Knowledge of local heat transfer coefficients is essential to understanding and successful modeling of the ice accretion process on airfoils. Ice can change airfoil shape, alter the aerodynamic behavior, and impact aircraft performance and control. This paper reports measurements of local heat transfer coefficients (Frossling number) around a 28% thickness ratio airfoil covering 90% of both the top and bottom surfaces. Measurements were conducted at a constant wall temperature for angles of attack from -8.5 to 19.5 deg, freestream Math numbers from 0.055 to 0.246, and chord Reynolds numbers from 2.5 x [10.sup.5] to 10.85 x [10.sup.5]. Static pressures were measured simultaneously at the same operating conditions. The results show that the heat transfer coefficient starts high at the stagnation point, decreases with the developing laminar boundary layer, and then increases through a region of transition to a peak at the onset of the turbulent boundary layer. Thereafter, the heat transfer coefficient slowly decreases as the thickness of the turbulent boundary layer increases. The location of the onset of turbulence moves toward the leading edge on the top surface and toward the trailing edge on the bottom surface with increasing angle of attack at a constant Reynolds number. This transition is also affected by the Reynolds number. The measured surface static pressure coefficients provide additional evidence to the changing trend of the transition locations; that is, that the location where the abrupt jump in the Frossling number occurs closely coincides with the state change from a negative pressure gradient to a positive gradient. DOI: 10.2514/1.44093

A novel method of computing the flutter boundary for an isolated airfoil based on a high-fidelity computational fluid dynamics model reveals unusual behavior in a critical transonic range. Inviscid and viscous predictions of the flutter boundary for the two airfoils examined differ substantially in this critical region and become sensitive to Mach number and grid topology due to complicated shock/boundary-layer interactions. Computational fluid dynamics predictions of the flutter boundary for a NACA 0012 section airfoil are also compared with previous experimental results. DOI: 10.2514/1.30703

The paper presents a method for the design and optimization of a multi-element airfoil-based inertial particle separator. To facilitate design in an interactive manner, a MATLAB graphical user interface was developed with the following capabilities: 1) design of individual airfoil elements; 2) orientation and arrangement of the airfoils to form a multi-element airfoil-based inertial particle separator system by employing translational, sealing, and rotational functions; 3) analysis of the inertial particle separator system; and 4) optimum design of the inertial particle separator system using an evolutionary algorithm-based direct-search optimization technique. The design of individual airfoils was achieved through the use of the PROFOIL code, a state-of-the-art multipoint inverse airfoil design program. Multiple airfoils were arranged to form a multi-element airfoil-based inertial particle separator system subject to geometric constraints. A two-dimensional analysis tool for the multi-element airfoil was linked to the synthesis component of the graphical user interface to carry out flow and particle trajectory analysis component of the program. For the optimization component of the program, a computational objective function was implemented and coupled with a pattern search optimizer of the Direct Search Toolbox in MATLAB. Design and optimization was achieved using multivariable optimization. The paper presents two design examples to demonstrate the design method. DOI: 10.2514/1.40921

In supersonic flight, airfoils generate strong sonic boom and wave drag accompanied by shock waves and expansion waves. The Busemann biplane is representative of an airfoil that has the possibility of realizing low drag. With the aim of realizing a new concept of supersonic transport, aerodynamic design and analysis are discussed based on computational fluid dynamics. Traditional biplane airfoils were extended to three-dimensional wings with a design Mach number of 1.7. Euler simulations of three-dimensional biplane wings of several configurations were conducted. Because of the existence of wing tips, three-dimensional biplane wings do not perform as well as two-dimensional biplane airfoils. This is because the areas affected by Mach cones originated from the wing tips, which precludes the occurrence of an appropriate pressure wave interaction. Thus, the wing tip area has a large drag coefficient. To overcome these problems, a tapered wing is herein considered. After parametric studies, a wing planform was determined, the taper ratio and aspect ratio of which were 0.25 and 5.12, respectively. Aerodynamic design of wing section shapes of the three-dimensional tapered biplane wing was conducted using an inverse-problem method. The designed biplane wing shows a better lift-to-drag ratio performance than the two-dimensional flat-plate airfoil in the range where the lift coefficient is more than 0.17. Tapered wings were also found to have several span sections that achieve better performances than two-dimensional biplane airfoils. DOI: 10.2514/1.40313

This investigation addresses the simulation of the unsteady separated flows encountered by a plunging airfoil under low-Reynolds-number conditions ([Re.sub.c] [less than or equal to] 6 x [10.sup.4]). The flowfields are computed employing an extensively validated high-fidelity implicit large-eddy simulation approach. Calculations are performed first for a SD7003 airfoil section at an angle of attack [[alpha].sub.0] = 4 deg plunging with reduced frequency k = 3.93 and amplitude [h.sub.0]/c = 0.05. Under these conditions, it is demonstrated that, for [Re.sub.c] = [10.sup.4], transitional effects are not significant. For [Re.sub.c] = 4 x [10.sup.4], the dynamic-stall vortex system is laminar at its inception, however, shortly afterward, it experiences an abrupt breakdown due to the onset of spanwise instabilities. A description of this transition process near the leading edge is provided. As a second example, the suppression of stall at high angle of attack ([[alpha].sub.0] = 14 deg) is investigated using high-frequency, small-amplitude vibrations (k = 10, [h.sub.0]/c = 0.005). At [Re.sub.c] = 6 x [10.sup.4], separation is completely eliminated in a time-averaged sense, and the mean drag is reduced by approximately 40%. For larger forcing amplitude ([h.sub.0]/c = 0.04, [Re.sub.c] = [10.sup.4]), a very intriguing regime emerges. The dynamic-stall vortex moves around and in front of the leading edge and experiences a dramatic breakdown as it impinges against the airfoil. The corresponding phased-averaged flow displays no coherent vortices propagating along the airfoil upper surface. This new flow structure is also characterized in the mean by the existence of a strong jet in the near wake which produces net thrust. This study demonstrates the importance of transition for low-Reynolds-number maneuvering airfoils and the suitability of the implicit large-eddy simulation approach for exploring such flow regime. DOI: 10.2514/1.43038

Stall flutter is a nonlinear aeroelastic phenomenon that can affect several types of aeroelastic systems such as helicopter rotor blades, wind turbine blades, and highly flexible wings. Although the related aerodynamic phenomenon of dynamic stall has been the subject of many experimental studies, stall flutter itself has rarely been investigated. This paper presents a set of experiments conducted on a NACA0012 airfoil undergoing stall flutter oscillations in a low-speed wind tunnel. The aeroelastic responses are analyzed with the objective of characterizing the local bifurcation behavior of the system. It is shown that symmetric stall flutter oscillations are encountered as a result of a subcritical Hopf bifurcation, followed by a fold bifurcation. The cause of these bifurcations is the occurrence of dynamic stall, which allows the transfer of energy from the freestream to the wing. A second bifurcation occurs at the system's static divergence airspeed. As a consequence, the wing starts to undergo asymmetric stall flutter bifurcations at only positive (or only negative) pitch angles. The dynamic stall mechanism itself does not change but the flow only separates on one side of the wing. DOI: 10.2514/1.39571

A highly efficient local-piston theory is presented for the prediction of inviscid unsteady pressure loads at supersonic and hypersonic speeds. A steady mean flow solution is first obtained by an Euler method. The classical piston theory is modified to apply locally at each point on the airfoil surface on top of the local mean flow to obtain the unsteady pressure perturbations caused by the deviation of the airfoil surface from its mean location without the need of performing unsteady Euler computations. Results of two- and three-dimensional unsteady air loads and flutter predictions are compared with those obtained by the classical piston theory and an unsteady Euler method to assess the accuracy and validity range in airfoil thickness, flight Mach number, and angle of attack and with the presence of blunt leading edges. The local-piston theory is found to offer superior accuracy and much wider validity range compared with the classical piston theory, with the cost of only a fraction of the computational time needed by an unsteady Euler method. DOI: 10.2514/1.37750

This study explores the effect of chordwise flexible deformation on unsteady aerodynamic characteristics for biplane counter-flapping dual NACA0014 airfoils with various combinations of Reynolds number and reduced frequency. Unsteady laminar viscous flows over duel rigid and flexible airfoils executing counter-plunge motion are computed with time-dependent two-dimensional laminar Reynolds-averaged Navier-Stokes equations coupled with conformal hybrid meshes. The tested Reynolds number with an airfoil characteristic chord length is [l0.sup.2], [10.sup.3], and [10.sup.4], and the reduced frequency ranges from 0.5 to 3.5. The dynamic mesh technique is applied to illustrate the flapping deformation modes of the flexible airfoils. To investigate the influence of the chordwise flexure extent on the aerodynamic performance of the flapping airfoils, the present study considers various different curvature deformations of flapping foils with flexure extent ranging from 0 to 0.3 times the chord length at 0.05 intervals. The visualized particle-tracing paths clearly revealed the formation and evolution of leading-edge vortices along the body of the flexible airfoil as it undergoes biplane counter-plunge motion. The generation of thrust-indicative wake structure or the drag-indicated wake structure behind the flexible airfoils depended on the degree of flexure extent of the airfoil at a fixed range of reduced frequency. The thrust force for each airfoil with biplane counter-flapping mode will be enhanced 6.32% more than that for a rigid single flapping airfoil. Present results show that flexible airfoils with flexure extent 0.25 times the chord length in counter-plunge flapping motion could get maximum propulsive efficiency and produce about 64.65 % more than that of biplane rigid airfoils. The outcome indicates that appropriate flexible biplane flapping flight could not only increase the thrust force, but also boost the propulsive performance.

The design and analysis of a new airfoil to be employed on a ultralight aircraft with short takeoff and landing is presented. An inverse design philosophy has been applied and is described; the numerical analysis performed used XFOIL, MSES, and TBVOR computational codes and the effects of airfoil shape on complete aircraft performances were taken into account. A high-lift configuration, including slat and single-slotted flap geometries, has been developed and is illustrated in this paper. Exhaustive wind.tunnel tests were performed at the Department of Aerospace Engineering and the experimental results are described here. To validate numerical results and to analyze the effect of laminar bubbles on airfoil performance, the pressures on airfoil surface and in the wake were measured and flow visualizations were done using fluorescent oil. The landing configuration was also tested and an experimental optimization of flap and slat positions was carried out to obtain a high maximum lift coefficient.

This work investigates the potential of short-wave instability developing near blade tip vortex as the main source of blade/wake-interaction noise. In this aim, a numerical simulation of these instabilities is performed considering a simplified wake geometry. The aerodynamic fields are extracted to determine the blade pressure response using the unsteady compressible airfoil theory. Finally a loading-noise computation is performed. All the results obtained from these computations are compared with the experimental data and a good qualitative agreement is found.

This paper presents an experimental and computational study of a passive version of the air-jet vortex-generator flow control system for the delay or suppression of trailing-edge boundary-layer separation and subsequent stall. The results show that a passive vortex-generating jet flow control system, which uses the high pressure front the leading-edge attachment line to feed the upper-surface air jets, can effectively delay trailing-edge separation and subsequent stall to higher angles of attack, without the need for any active energy input and without significant drag penalty. DOI: 10.2514/1.41986

Reynolds-averaged Navier--Stokes simulations are presented for a circulation control airfoil. Comparisons with measured load characteristics are presented. The effects of turbulence models and the spatial accuracy of the simulations on the predictions are examined. It is observed that increasing the spatial accuracy from third-order to seventh-order had only a minor effect on the solution accuracy on the fine grid simulations reported here. Two-equation turbulence models, and particular the [kappa]-[omega]/[kappa]- [epsilon] blended baseline model, tended to perform better than the one-equation model. The predictions were found to be sensitive to the turbulent kinetic energy level of the jet. Inclusion of the plenum chamber and the jet nozzle allowed the physics of the flow to be captured in greater detail than modeling the jet only at the jet slot, but had only a minor influence on the overall loads. DOI: 10.2514/1.41638

Unsteady solutions for the aerodynamic coefficients of a thin airfoil in compressible sub- or supersonic flows were studied. The lift, the pitch moment, and the pressure coefficients were obtained numerically for the following motions: the indicial response (unit step function) of the profile, that is, a sudden change in the angle of attack; a thin profile penetrating into a sharp edge gust (for several gust velocity ratios); a thin profile penetrating into a one- minus-cosine gust, that is, a typical gust used in commercial aircraft design; the oscillating airfoil; and the interaction of the profile with a convected (from convection phenomenon) vortex passing under the profile, a phenomenon known in the literature as blade vortex interaction or, for a profile case, airfoil--vortex interaction. The present work uses a numerical approach based on vortex singularity. The numerical model was created through the profile discretization in uniform segments, and the compressible flow vortex singularity was used. The results available in the literature are based on approximated exponential equations or computed via computational fluid dynamics. Thus, the purpose of this method is to obtain results as accurate as those from approximated equations and significantly faster than those done via computational fluid dynamics. DOI: 10.2514/1.40610

Through experiments using two-dimensional particle-image velocimetry (PIV), this paper examines the nature of transition in a separation bubble and manipulations of the resultant breakdown to turbulence through passive means of control. An airfoil was used that provides minimal variation in the separation location over a wide operating range, with various two-dimensional modifications made to the surface for the purpose of manipulating the transition process. The study was conducted under low-freestream-turbulence conditions over a flow Reynolds number range of 28,000-101,000 based on airfoil chord. The spatial nature of the measurements has allowed identification of the dominant flow structures associated with transition in the separated shear layer and the manipulations introduced by the surface modifications. The Kelvin--Helmholtz (K-H) instability is identified as the dominant transition mechanism in the separated shear layer, leading to the roll-up of spanwise vorticity and subsequent breakdown into small-scale turbulence. Similarities with planar free-shear layers are noted, including the frequency of maximum amplification rate for the K-H instability and the vortex-pairing phenomenon initiated by a subharmonic instability. In some cases, secondary pairing events are observed and result in a laminar intervortex region consisting of freestream fluid entrained toward the surface due to the strong circulation of the large-scale vortices. Results of the surface-modification study show that different physical mechanisms can be manipulated to affect the separation, transition, and reattachment processes over the airfoil. These manipulations are also shown to affect the boundary-layer losses observed downstream of reattachment, with all surface-indentation configurations providing decreased losses at the three lowest Reynolds numbers and three of the five configurations providing decreased losses at the highest Reynolds number. The primary mechanisms that provide these manipulations include: suppression of the vortex-pairing phenomenon, which reduces both the shear-layer thickness and the levels of small-scale turbulence; the promotion of smaller-scale turbulence, resulting from the disturbances generated upstream of separation, which provides quicker transition and shorter separation bubbles; the elimination of the separation bubble with transition occurring in an attached boundary layer; and physical disturbance, downstream of separation, of the growing instability waves to manipulate the vortical structures and cause quicker reattachment. [DOI: 10.1115/1.2978997]

We report on an experimental study of a Clark-Y airfoil subjected to both flow injection and suction in the manner of the coflow-jet flow control method. The study was conducted at Reynolds numbers of 7.4 x [10.sup.4] and 1.3 x [10.sup.5] for three different jet momentum coefficients: [C.sub.[mu]] = 0.02, 0.04, and 0.16. Some numerical computations are used to assess the extent to which they are adequate for engineering analysis and design of the coflow-jet method. The results clearly demonstrate the lift augmentation capability, particularly at high incidence angles. The increase in lift and stall margin depends on the momentum coefficient, and for the range tested, the increase in the integral of the airfoil static pressure distribution varies from 40 to 100%. The results suggest that the coflow-jet airfoil is more effective when both injection and suction are used simultaneously, compared with the case in which they are individually used. A first-order analysis of the combined merits of injection and suction indicate that these processes are not strongly coupled to each other and thus may possibly be analyzed separately.

An experimental study was conducted to investigate the flow behavior around a bioinspired corrugated airfoil compared with a traditional streamlined airfoil and a flat plate at the chord Reynolds number of Re = 34, 000 to explore the potential application of such bioinspired corrugated airfoils for micro air vehicle applications. The experiments were conducted in a low-speed wind tunnel. A high-resolution particle image velocimetry system was used to conduct detailed flowfield measurements to quantify the transient behavior of vortex and turbulent flow structures around the studied airfoils. The particle image velocimetry measurement results demonstrated clearly that the corrugated airfoil has better performance over the streamlined airfoil and the flat plate in preventing large-scale flow separation and airfoil stall at low Reynolds numbers. It was found that the protruding corners of the corrugated airfoil would act as turbulators to generate unsteady vortex structures to promote the transition of the separated boundary-layer flow from laminar to turbulent. The unsteady vortex structures trapped in the valleys of the corrugated cross section would pump high-speed fluid from outside to near-wall regions to provide sufficient kinetic energy for the boundary layer to overcome adverse pressure gradients, thus discouraging large-scale flow separations and airfoil stall. Aerodynamic force measurements further confirmed the possibility of using such bioinspired corrugated airfoils in micro air vehicle designs to improve their flight agility and maneuverability.

The aerodynamics of a Gurney-flap-equipped airfoil has been explored by means of low-speed wind-tunnel experiments performed at a chord Reynolds number of 1.0 x [10.sup.6]. Various chordwise locations and sizes of Gurney flaps were tested. Surface-pressure distributions and the wake momentum deficit were measured and used to determine lift, pitching moment, and drag. Compared with the clean airfoil, the measured maximum lift coefficient can be increased by nearly 30% with these simple devices. The amount of lift increase has a nearly linear dependency on the chordwise location and size of the Gurney flap. Minimum drag is primarily affected by the flap size and, to a lesser extent, by the chordwise location. The Gurney flap increases in maximum lift are obtained by increasing the lower-surface pressures over the aft part of the airfoil. At the same time, the magnitude of pressure peak on the upper surface near the leading edge is reduced such that the upper-surface pressures over the middle parts of the airfoil are reduced and the separation point is moved aft by the reduced pressure-recovery gradients. As expected, this increases the aft loading and results in an increased nose-down pitching moment. As the angle of attack is decreased, the influence of a Gurney flap extending from the lower surface likewise decreases as the flap is increasingly immersed in the thickening boundary layer. A Gurney flap mounted to the upper surface behaves in the opposite way: increasing the negative lift at low angles of attack and having less and less influence as the angle of attack is increased. Although Gurney flaps result in significantly higher drags for airfoils with extensive runs of laminar flow, this disadvantage disappears as the amount of turbulent boundary-layer flow increases, as is the case with fixed transition near the leading edge of the airfoil.

Three coflow jet airfoils with twice-doubled injection-slot sizes are calculated using a Reynolds-averaged Navier-Stokes computational fluid dynamics solver with the one-equation Spalart--Allmaras model. At the same angle of attack, the twice-larger injection-slot-size airfoil passes the (about twice-greater) jet mass flow rate, with the momentum coefficients also nearly doubled. The coflow jet airfoil with the largest slot size has the least stall angle of attack. When the injection-slot size is reduced from the maximum by half, the stall angle of attack and the maximum lift coefficient are increased. When the injection-slot size is further reduced by half, the stall angle of attack is still increased, but the maximum lift coefficient is lower due to the smaller momentum coefficient. The trends of the stall angle of attack and maximum lift coefficient agree with the experiment. At low angles of attack, both the computed lift and drag coefficients agree fairly well with the experiment. At high angles of attack, the lift and drag are underpredicted. The reason may be that the Reynolds-averaged Navier--Stokes model cannot handle the turbulence mixing at high angles of attack.

The transient process of flow separation and reattachment to a concave surface was investigated experimentally on the rear ramp of a modified Glauert Laminar Airfoil Section II. The measurements were carried out at low incompressible speeds using miniature time-resolved piezoresistive pressure transducers (Endevco model). These measurements were supplemented by flow visualization and by exploratory use of a particle image velocimeter. It was observed that the concavity of the surface impeded flow reattachment, and therefore a much higher momentum coefficient was required to partially attach the flow to a concave surface than to a straight one when both surfaces were supposed to turn the flow by the same amount. Streamwise vortices were observed to exist on the concave ramp of the modified Glauert Laminar Airfoil Section II, and thus the interaction of streamwise and spanwise vortices and their impact on flow control are discussed. Lift augmentation by periodic excitation is mainly achieved by the entrainment of the upstream flow into the pulsating jet. For the on-off application of control, the transient time normalized by the time of flight between the actuator and the trailing edge varies between 15 and 20 over the upper and lower surfaces and between 5 to 15 over the ramp. The possibilities of gradually controlling the airfoil performance are also considered, and the phase delays of the surface-pressure perturbation are discussed.

This paper describes in-flight measurements in low- and high-speed experiments to investigate laminar-turbulent transition on airfoils. The low-speed experiments were carried out using a laminar wing glove for a sailplane. High-speed experiments on an modified A320 hybrid laminar fin were conducted in close collaboration with Airbus. Different surface sensor measuring techniques (piezofoil, surface hot wires, wall microphones) were used in the low-speed experiments. The piezofoil technique was applied in the high-speed in-flight experiments to prove the operation robustness, the functional capability, durability, and reliability under atmospheric conditions in cruise flight on modern airplanes. All measuring techniques were successfully implemented for the particular task and the laminar-turbulent transition was detected with a high degree of certainty. For the high-speed investigations a complete piezofoil sensor array was glued directly to the fin without causing any interference between the boundary layer flow and the sensor foil. This technique therefore offered the good potential to detect transition in the boundary layer under flight conditions especially in high-speed tests.

This research presents some common features of oscillatory airfoils, and the method for indicating dynamic stall onset for the unsteady process. Under deep stall conditions, the stall-onset angle in oscillation is independent of the mean angle of the oscillatory motion, and by combining the reduced frequency and the amplitude of the oscillatory motion, the equivalent reduced pitch rate is an analog of this motion to the constant reduced pitch rate of the ramp-up motion. By correlating with the measured data, and with the ramp-up results, the equivalent reduced pitch rate can be defined as a representation for the oscillatory motion. Accordingly, the triple-parameter problem of an oscillation (mean angle, reduced frequency, and amplitude) degrades into the single-parameter problem (equivalent reduced pitch rate). Based on these foundations, an extension of the stall-onset criterion is then made for oscillatory airfoils: a method of extracting the stall-onset parameters directly from oscillatory test data, and an indication of stall onset for the oscillatory airfoils. The results from the new proposed method have shown the consistency with the data of Glasgow University and the public data.

The investigation focuses on the analysis of the airfoil segment performances along rotor blades in the parked configuration. In this research, wind tunnel experiments on two twisted blade geometries with different airfoils played a dominant role. These measurements were carried out by the Swedish Aeronautical Research Institute, former FFA, and by the American National Renewable Energy Laboratories (NREL) during the Unsteady Aerodynamic Experiment. The spans of the blades were 2.375 m and 5 m, the STORK 5 WPX and the NREL Phase VI blade, respectively. Five span locations (inboard, midspan, outboard, and tip regions) were considered and compared with the 2D airfoil characteristics. Wing model experiments with similar blade aspect ratio were included in the research. Furthermore, the commercial computational fluid dynamics code FLUENT was used for the validation and analysis of the spanwise lift and drag coefficients at four different pitch settings, 20 deg, 30 deg, 45 deg, and 60 deg. The computed pressure distributions compared reasonably well, but the derived lift and drag showed quite some differences with the blade measurements. The lift coefficients for the sections beyond the leading-edge stall angle of the STORK blade were larger than for the NREL blade and were close to that of a wing model with similar airfoil and aspect ratio. Lift and drag coefficients for the sections of the two blades were always much smaller than the 2D results. The drag values for both blades showed quite some agreement, and airfoil and blade dependency seemed to be small. [DOI: 10.1115/1.2931507]

Naphthalene sublimation experiments were conducted to study heat transfer for flow through blockages with holes in an internal cooling passage near the trailing edge of a gas turbine airfoil. The cooling passage was modeled as two rectangular channels whose heights decreased along the main flow direction. The air made a right-angled turn before passing through two blockages with staggered holes in each channel and left the channel through an exit that was partially blocked by periodic lands with rounded leading edges. There were ten holes along each blockage and all of the holes had rounded edges. Local heat (mass) transfer was measured and overall heat (mass) transfer results were obtained, on the exposed surfaces of one of the walls downstream of the two blockages, for Reynolds numbers (based on the hydraulic diameter of the channel at the upstream surface of the first blockage) between 5000 and 36,000. The results showed that the blockages with the larger hole-to-channel cross-sectional area ratio in one of the two test sections enhanced the heat (mass) transfer downstream of the blockages more than the blockages with the smaller open area ratio in the second test section. For the geometric configurations and flow conditions studied, the average heat (mass) transfer was higher downstream of the second blockage than downstream of the first blockage. The configurations of the inlet channel and the exit slots considered in this study did not significantly affect the local heat (mass) transfer distributions or the average heat (mass) transfer downstream of the blockages. Keywords: turbine blade trailing edge cooling, naphthalene sublimation, mass transfer, blockages with holes

Experiments were performed on a flat-plate airfoil and an Eppler E338 airfoil at very low flight Reynolds numbers (3000 [less than or equal to] Re [less than or equal to] 50,000), in which dielectric barrier discharge plasma actuators were employed at the airfoil leading edges to effect flow control. The actuators were driven in a high-frequency (kilohertz) steady mode and a pulsed mode in which pulse frequency and duty cycle were varied in a systematic fashion. Optimum reduced frequencies for generating poststall lift were approximately between 0.4 and 1, and this was broadly consistent with zero-mass-flux slot-blowing data acquired at Reynolds numbers that were approximately 200 times higher. Nevertheless, profound differences in the response to reduced frequency and duty cycle were observed between the flat-plate and E338 airfoils. In general, actuation produced considerable performance improvements, including an increase in maximum lift coefficient of 0.4 to 0.8 and maintained elevated endurance at significantly higher lift coefficients. Actuation in the steady mode resulted in circulation control at Re = 3000. Pulsed actuation also exerted a significant effect on the wake at prestall angles of attack, in which control of the upper-surface flat-plate bubble shedding produced significant differences in wake spreading and vortex shedding. The flat plate was also tested in a semispan-wing configuration (AR = 6), and the effect of control was comparable with that observed on the airfoil.

An experimental study was conducted to characterize the transient behavior of laminar flow separation on a NASA low-speed GA (W)-1 airfoil at the chord Reynolds number of 70,000. In addition to measuring the surface pressure distribution around the airfoil, a high-resolution particle image velocimetry (PIV) system was used to make detailed flow field measurements to quantify the evolution of unsteady flow structures around the airfoil at various angles of attack (AOAs). The surface pressure and PIV measurements clearly revealed that the laminar boundary layer would separate from the airfoil surface, as the adverse pressure gradient over the airfoil upper surface became severe at AOA [greater than or equal to] 8.0 deg. The separated laminar boundary layer was found to rapidly transit to turbulence by generating unsteady Kelvin-Helmholtz vortex structures. After turbulence transition, the separated boundary layer was found to reattach to the airfoil surface as a turbulent boundary layer when the adverse pressure gradient was adequate at AOA < 12.0 deg, resulting in the formation of a laminar separation bubble on the airfoil. The turbulence transition process of the separated laminar boundary layer was found to be accompanied by a significant increase of Reynolds stress in the flow field. The reattached turbulent boundary layer was much more energetic, thus more capable of advancing against an adverse pressure gradient without flow separation, compared to the laminar boundary layer upstream of the laminar separation bubble. The laminar separation bubble formed on the airfoil upper surface was found to move upstream, approaching the airfoil leading edge as the AOA increased. While the total length of the laminar separation bubble was found to be almost unchanged (~20% of the airfoil chord length), the laminar portion of the separation bubble was found to be slightly stretched, and the turbulent portion became slightly shorter with the increasing AOA. After the formation of the separation bubble on the airfoil, the increase rate of the airfoil lift coefficient was found to considerably degrade, and the airfoil drag coefficient increased much faster with increasing AOA. The separation bubble was found to burst suddenly, causing airfoil stall, when the adverse pressure gradient became too significant at AOA > 12.0 deg.

Laminar boundary-layer separation and separated shear layer development on a NACA 0025 airfoil at low Reynolds numbers were studied experimentally. Flow visualization, hot-wire velocity measurements, and time-resolved surface pressure measurements were employed in this investigation. The results for two distinct flow regimes, namely, flow separation without subsequent shear layer reattachment and separation bubble formation, are discussed in detail. For both flow regimes, the transition occurs due to the amplification of natural flow disturbances in the separated shear layer. Initially, disturbances within a band of frequencies centered at some fundamental frequency are amplified. Further downstream, nonlinear interactions set in, leading to a breakdown to turbulence. Based on the spectral and correlation analysis of velocity and surface pressure fluctuations, it is demonstrated that the amplification of disturbances and the attendant fluctuations in the flow velocity give rise to distinct surface pressure fluctuations at the fundamental frequency. Thus, time-resolved surface pressure measurements can be employed to estimate important unsteady characteristics of the separated flow region on an airfoil operating in low Reynolds flows.

In this paper, we compare the global accuracy of different strategies to build response surfaces by varying sampling methods and modeling techniques. The aerodynamic test functions are obtained by deforming the shape of a transonic airfoil. For comparisons, a robust strategy for model fit using a new efficient initialization technique followed by a gradient optimization was applied. First, a study of different sampling methods proves that including a posteriori information on the function to sample distribution can improve accuracy over classical space-filling methods such as Latin hypercube sampling. Second, comparing kriging and gradient-enhanced kriging on two- to six-dimensional test cases shows that interpolating gradient vectors drastically improves response-surface accuracy. Although direct and indirect cokriging have equivalent formulations, the indirect cokriging outperforms the direct approach. The slow linear phase of error convergence when increasing sample size is not avoided by cokriging. Thus, the number of samples needed to have a globally accurate surface stays generally out of reach for problems considering more than four design variables.

A static extended trailing edge attached to a NACA0012 airfoil section is studied for achieving lift enhancement at a small drag penalty. It is indicated that the thin extended trailing edge can enhance the lift, whereas the zero-lift drag is not significantly increased. Experiments and calculations are conducted to compare the aerodynamic characteristics of the extended trailing edge with those of the Gurney flap and the conventional flap. The extended trailing edge, as a simple mechanical device added on a wing without altering the basic configuration, has a good potential to improve the cruise flight efficiency. DOI: 10.2514/1.31995

This paper considers the unsteady reestablishment of fully attached flow of an airfoil from the stalled condition while undergoing ramp-down motions. The main purpose of the work was to extend the applicability of the Leishman-Beddoes dynamic-stall model during the return from stalled state to the low Mach and Reynolds numbers of the Glasgow unsteady database. This was done by the adoption of the Beddoes modeling philosophy and the technique of the lagging angle of attack that Sheng et al. ['A New Stall-Onset Criterion for Low Speed Dynamic-Stall,' Journal of Solar Energy Engineering, Vol. 128, Nov. 2006, pp. 461-471; and 'Improved Dynamic-Stall-Onset Criterion at Low Mach Numbers,' Journal of Aircraft (to he published)] proposed for developing a new dynamic-stall onset criterion. It is concluded that a more secure assessment was obtained of the nondimensional time taken to convect the stalled flow over and away from the airfoil, compared with those of Niven et al. ('Analysis of Reattachment During Ramp Down Tests,' Vertica, Vol. 13, No. 2, 1989, pp. 187-196) and Green and Galbraith ('Dynamic Recovery to Fully Attached Aerofoil Flow from Deep Stall,' AIAA Journal, Vol. 33, No. 8,1995, pp. 1433-1440). This, combined with the appropriate lift-curve slope, yielded an improved data reconstruction over that of the Leishman-Beddoes model. The present depiction and modeling of the reestablishment of fully attached flow from stalled state is a valid enhancement to the Leishman-Beddoes model. DOI: 10.2514/1.30554

Motivated by the discovery that living creatures may be able to enhance their locomotion capability through passive or active deformations of their wings or fins, we carry out a fully coupled fluid-structure interaction study to investigate the flapping motion of a foil with either chordwise or spanwise flexibility. We employ a fluid-structure interaction model, which accounts for fluid dynamics using the boundary-element method and structural dynamics using a two-dimensional nonlinear thin-plate model. With this approach, we numerically investigate the effect of structural deformation on the performance of the foil when it is immersed in two different fluids, a low-density fluid (e.g., air), in which the deformation is determined primarily by the inertia of the foil, and a high-density fluid (e.g., water), in which the fluid loading has a significant impact. In the first scenario, we find that the chordwise flexibility reduces both the thrust and the propulsion efficiency and the spanwise flexibility increases the thrust without efficiency reduction within a small range of structural parameters. In the second scenario, we find that the chordwise flexibility increases the efficiency. The spanwise flexibility, on the other hand, compromises the performance of the foil by diminishing both the thrust and the efficiency. Possible applications of these findings and their relation to designs in aquatic animals are discussed. DOI: 10.2514/1.28565

In this work, the effect of wall interference on steady and oscillating airfoils in a subsonic wind tunnel is studied. A variety of approaches including linear theory, compressible inviscid and viscous computations, and experimental data are considered. Integral transform solutions of the linearized potential equations show an augmentation of the lift magnitude for steady flows when the wall is close to the airfoil surface. For oscillating airfoils, lift augmentation is accompanied by a significant change in the phase of the lift response. Idealized compressible Euler calculations are seen to corroborate the linear theory under conditions that are sufficiently away from acoustic resonance. Further, the theory compares well with compressible Reynolds-averaged Navier-Stokes calculations and experimental measurements over a wide range of attached flows at subsonic Mach numbers. The present methodology can thus be used to predict wall interference effects and also to help extrapolate linear and nonlinear (dynamic stall) wind tunnel data to free-air conditions. DOI: 10.2514/1.28143