Showing papers on "Airfoil published in 2007"

Effects of Leading-Edge Protuberances on Airfoil Performance

Abstract: Lift, drag, and pitching moments of airfoils with leading-edge sinusoidal protuberances were measured in a water tunnel and compared with those of a baseline 63 4 -021 airfoil. The amplitude of the leading-edge protuberances ranged from 2.5 to 12% of the mean chord length; the spanwise wavelengths were 25 and 50% of the mean chord length. These ranges correspond to the morphology found on the leading edge of humpback whales' flippers. Flow visualization using tufts was also performed to examine the separation characteristics of the airfoils. For angles of attack less than the baseline stall angle, lift reduction and drag increase were observed for the modified foils. Above this angle, lift of the modified foils was up to 50% greater than the baseline foil with little or no drag penalty. The amplitude of the protuberances had a distinct effect on the performance of the airfoils, whereas the wavelength had little. Flow visualization indicated separated flow originating primarily from the troughs and attached flow on the peaks of the protuberances at angles beyond the stall angle of the baseline foil.

Flexible flapping airfoil propulsion at low Reynolds numbers

Abstract: Water tunnel experiments on a flexible airfoil plunging with constant amplitude have been carried out for Reynolds numbers of 0 to 27000. Peaks in thrust coefficient at intermediate values of airfoil stiffness were observed at both zero and non-zero Reynolds numbers, indicating that a degree of flexibility is beneficial at low Reynolds numbers. Time-averaged velocity fields and momentum flux data revealed a broader, higher-velocity jet in cases of optimum airfoil stiffness. Stronger vortices, separated by a larger lateral distance, characterised the corresponding instantaneous velocity fields. The flexibility causes the airfoil to pitch passively; the phase angle of the pitch was found to lead the plunge. Pitch amplitude and trailing-edge amplitude were found to be single-valued functions of pitch phase angle. The shape characteristics of the airfoil could therefore be described by the pitch phase angle only. Thrust coefficient was found to be a function of only two parameters: Strouhal number and pitch phase angle. For each Strouhal number, a peak in thrust coefficient was observed at a particular value of the pitch phase angle. The optimum pitch phase angle was found to tend to a limit of 105±5 degrees at very large Strouhal numbers. A significant thrust benefit was observed over very stiff airfoils when the optimum flexibility is utilized.

Aeroelastic instability problems for wind turbines

Abstract: This paper deals with the aeroelastic instabilities that have occurred and may still occur for modern commercial wind turbines: stall-induced vibrations for stall-turbines, and classical flutter for pitch-regulated turbines. A review of previous works is combined with derivations of analytical stability limits for typical blade sections that show the fundamental mechanisms of these instabilities. The risk of stall-induced vibrations is mainly related to blade airfoil characteristics, effective direction of blade vibrations and structural damping; whereas the blade tip speed, torsional blade stiffness and chordwise position of the center of gravity along the blades are the main parameters for flutter. These instability characteristics are exemplified by aeroelastic stability analyses of different wind turbines. The review of each aeroelastic instability ends with a list of current research issues that represent unsolved aeroelastic instability problems for wind turbines. Copyright © 2007 John Wiley & Sons, Ltd.

Experimental Evaluation of Sinusoidal Leading Edges

Abstract: P REVIOUS studies on increasing airfoil lift and improving stall characteristics have addressed various passive and active approaches to modifying the leading and trailing edge shapes. The passive approaches have covered such methods as rippling the trailing edge, applying serrated-edge Gurney flaps, or modifying the leading-edge (LE) profile [1,2]. Other efforts have effectively eliminated the dynamic stall of an NACA 0012 airfoil by perturbing the LE contour as little as 0.5–0.9%of the chord [3]. Levshin et al. [4] demonstrated that sinusoidal LE planforms on an NACA 63-021 airfoil section decreased maximum lift, but extended the stall angle by almost 9 deg. The larger amplitude sinusoids created “softer” stall characteristics by maintaining attached flow at the peaks despite separated flow in the troughs. These tests were performed to simulate the effects of LE tubercles on humpback whale (Megaptera novaeangliae) flippers. Prior work by the authors also reported wind tunnel measurements for idealized scale models of humpback whale flippers [5]. One model had a smooth leading edge and a secondmodel had sinusoidal bumps (tubercles) along the leading edge for the outer 2 3 of the span. It was found that the addition of tubercles to a 3-D idealized flipper increased the maximum lift coefficient while reducing the drag coefficient over a portion of the operational envelope. It is thought that the tubercles on the flipper leading-edge enhance the whale’s ability to maneuver to catch prey [6]. Though the work to date regarding sinusoidal or serrated leading-edge planforms is largely motivated by marine mammal locomotion, the effects of extending the stall point for lifting surfaces at similar Reynolds numbers (Re) may have application to small-UAV (unmanned aerial vehicle) design and the inevitable laminar stall problems [7]. However other relevant applications might benefit from the effects of simulated tubercles such as stall alleviation/separation control on sailboat centerboards or wind turbines, where an expanded operating envelope could improve the overall effectiveness of the blade [8,9]. In the present work, a better understanding is sought of the mechanism of the improvements measured in previous experiments, with a greater applicability in mind. The authors seek to determine whether the performance improvements resulted from enhancements to the sectional characteristics of wings with tubercles (i.e., essentially 2-D effects), or from Reynolds number effects on a tapered planform, or from other 3-D effects such as spanwise stall progression.

Numerical investigation of the tone noise mechanism over laminar airfoils

Abstract: This paper presents the first numerical investigation via direct numerical simulation of the tone noise phenomenon occurring in the flow past laminar airfoils. This phenomenon corresponds to the radiation of discrete acoustic tones in some specific flow conditions, and has received much attention since the 1970s, and several experimental studies have been carried out to identify and understand the underlying physical mechanisms. However, several points remain to be clarified in order to provide a complete explanation of its origin. The flow around a two-dimensional NACA0012 airfoil is considered in order to have a deeper understanding of the tone noise phenomenon. Consistently with previous experimental studies, it is shown that depending on the Reynolds number and angle of attack, two different types of acoustic spectrum are observed: one which exhibits a broadband contribution with a dominant frequency together with a sequence of regularly spaced discrete frequencies, while the other one is only characterized by a simple broadband contribution. The first configuration is typical of the tone noise phenomenon. The present work shows that in this case, the mean flow on the pressure side of the airfoil exhibits a separation bubble near the trailing edge and the main tone frequency is close to the most amplified frequency of the boundary layer. The mechanism proposed in previous works for the main tone generation – which implies the existence of a separation bubble at the pressure side – is therefore validated by numerical simulation. On the other hand, the analysis of the suction side boundary layer reveals that there is no separation and that the most amplified frequency is different from the main tonal one. However, the suction side boundary layer is highly receptive to the tone frequency. Finally, an original explanation for the existence of the secondary discrete frequencies observed in the radiated pressure spectrum is given. They are associated to a bifurcation of the airfoil wake from a symmetric to a non-symmetric vortex pattern. A possible explanation for the existence of this bifurcation is the interaction between the disturbances which are the most amplified by the suction side boundary layer and those originating in the forcing of the suction side flow by the main tone noise mechanism.

Origin of transonic buffet on airfoils.

Abstract: Buffeting flow on transonic aerofoils serves as a model problem for the more complex three-dimensional flows responsible for aeroplane buffet. The origins of transonic aerofoil buffet are linked to a global instability, which leads to shock oscillations and dramatic lift fluctuations. The problem is analysed using the Reynoldsaveraged Navier–Stokes equations, which for the foreseeable future are a necessary approximation to cover the high Reynolds numbers at which transonic buffet occurs. These equations have been shown to reproduce the key physics of transonic aerofoil flows. Results from global-stability analysis are shown to be in good agreement with experiments and numerical simulations. The stability boundary, as a function of the Mach number and angle of attack, consists of an upper and a lower branch – the lower branch shows features consistent with a supercritical bifurcation. The unstable modes provide insight into the basic character of buffeting flow at nearcritical conditions and are consistent with fully nonlinear simulations. The results provide further evidence linking the transonic buffet onset to a global instability.

Suppressing Aeroelastic Instability Using Broadband Passive Targeted Energy Transfers, Part 2: Experiments

Abstract: This paper presents experimental results corroborating the analysis developed in the companion paper, Part 1 (Lee, Y., Vakakis, A., Bergman, L., McFarland, M., and Kerschen G., "Suppression Aeroelastic Instability Using Broadband Passive Targeted Energy Transfers, Part 1: Theory," AIAA Journal, Vol. 45, No. 3,2007, pp. 693-711), and demonstrates that a nonlinear energy sink can improve the stability of an aeroelastic system. The nonlinear energy sink was, in this case, attached to the heave (plunge) degree of freedom of a rigid airfoil which was supported in a low-speed wind tunnel by nonlinear springs separately adjustable in heave and pitch. This airfoil was found to exhibit a limit cycle oscillation at flow speeds above the critical ("flutter") speed of 9.5 m/s, easily triggered by an initial heave displacement. After attachment of a single degree of freedom, essentially nonlinear energy sink to the wing, the combined system exhibited improved dynamic response as measured by the reduction or elimination of limit cycle oscillation at flow speeds significantly greater than the wing's critical speed. The design, application, and performance of the nonlinear energy sink are described herein, and the results obtained are compared to analytical predictions. The physics of the interaction of the sink with the wing is examined in detail.

Proportional Closed-Loop Feedback Control of Flow Separation

Abstract: The aim of this experimental study is the implementation of a practical and efficient closed-loop feedback control of the turbulent flow over a NACA-4412 airfoil equipped with leading-edge zero-net-mass-flux actuators. By using prior computation of correlations between particle image velocimetry data and multiple surface pressure measurements, real-time instantaneous low-dimensional estimates of the velocity field over the wing are then computed from the unsteady surface pressure. From such estimates, a direct knowledge of the state of the flow above the airfoil is obtained (i.e., attached, incipient separation, or fully separated flow). We first show the effectiveness of the low-dimensional modeling approach in extracting and estimating the underlying large-scale structures in a turbulent flow, using the proper orthogonal decomposition and the modified linear/quadratic stochastic measurements. We then show how such an approach is used successfully in a simple, but practical, proportional feedback loop to delay the separation of the flow over the wing at high angles of attack. The benefits of closed-loop vs open-loop control are then discussed. These fundamental results validate the use of low-dimensional modeling techniques for further, more sophisticated, closed-loop feedback control algorithms.

Mechanisms Influencing the Efficiency of Oscillating Airfoil Propulsion

Abstract: A NACA0012 airfoil undergoing pitching and plunging motion at Re = 20,000-40,000 was simulated using a two-dimensional Navier-Stokes flow solver. Results were compared with experimental measurements in the literature and those from an inviscid analytical method and an unsteady panel method code. Although the peak in propulsive efficiency with Strouhal number demonstrated in the experimental results was predicted by the inviscid methods, it was found to be significantly modified by leading-edge vortex shedding and viscous drag at low Strouhal numbers. The occurrence and influence of vortex shedding is controlled by both the motion of the airfoil (amplitudes and phases of plunging and pitching) and the flapping frequency, which limits the time available for vortex formation and convection over the airfoil surface. Thus, Strouhal number alone is insufficient to characterize the efficiency of flapping-foil propulsion.

Aerodynamics of a NACA4412 airfoil in ground effect

Abstract: The flow characteristics over a NACA4412 airfoil are studied in a low turbulence wind tunnel with moving ground simulation at a Reynolds number of 3.0 x 105 by varying the angle of attack from 0 to 10 deg and ground clearance of the trailing edge from 5% of chord to 100%. The pressure distribution on the airfoil surface was obtained, velocity survey over the surface was performed, wake region was explored, and lift and drag forces were measured. To ensure that the flow is 2-D, particle image velocimetry measurements were performed. A strong suction effect on the lower surface at an angle of attack of 0 deg at the smallest ground clearance caused laminar separation well ahead of the trailing edge. Interestingly, for this airfoil, a loss of upper surface suction was recorded as the airfoil approached the ground for all angles of attack. For angles up to 4 deg, the lift decreased with reducing ground clearance, whereas for higher angles, it increased due to a higher pressure on the lower surface. The drag was higher close to the ground for all angles investigated mainly due to the modification of the lower surface pressure distribution.

Competing Broadband Noise Mechanisms in Low-Speed Axial Fans

Abstract: This paper compares two broadband noise mechanisms, the trailing-edge noise or self-noise, and the leading-edge noise or turbulence-ingestion noise, in several blade technologies. Two previously developed analytical models for these broadband contributions are first validated with well-defined measurements on several airfoils embedded in an homogeneous flow at low-Mach number. Each instrumented airfoil is placed at the exit of an open jet anechoic wind tunnel with or without a grid generating turbulence upstream of it. Sound is measured in the far field at the same time as the wall-pressure fluctuations statistics close to the airfoil trailing edge and the inlet velocity fluctuation statistics impacting the airfoil leading edge. The models are then compared in some practical cases representative of airframes, wind turbines, and automotive engine cooling modules. The airfoil models of the two mechanisms are then extended to a full rotating machine in open space. The model predictions of both mechanisms are compared with in-flight helicopter measurements and automotive engine cooling modules measurements. In both instances, the turbulence-ingestion noise is found to be a dominant source over most of the frequency range. The self-noise only becomes a significant contributor at high angles of attack close to flow separation.

Shock-Wave/Boundary-Layer Interaction Control Using Three-Dimensional Bumps for Transonic Wings

Abstract: Three-dimensional bumps have been developed and investigated on transonic wings, aiming to fulfill two major objectives of shock-wave/boundary-layer interaction control, that is, drag reduction and buffet delay. An experimental investigation has been conducted for a rounded bump in channel flow at the University of Cambridge and a computational study has been performed for a spanwise series of rounded bumps mounted on a transonic aerofoil at the University of Stuttgart. In both cases wave drag reduction and mild control effects on the boundary layer have been observed. Control effectiveness has been assessed for various bump configurations. A double configuration of narrow rounded bumps has been found to perform best, considerably reducing wave drag by means of a well-established X-shock structure with little viscous penalty and thus achieving a maximum overall drag reduction of about 30%, especially when significant wave drag is present. Counter-rotating streamwise vortex pairs have been produced by some configurations as a result of local flow separation. On the whole a large potential of three-dimensional control with discrete rounded bumps has been demonstrated both experimentally and numerically.

Numerical and Experimental Flow Analysis of Moving Airfoils with Laminar Separation Bubbles

Abstract: Experimental measurements and unsteady Reynolds-averaged Navier-Stokes simulations of the low-Reynolds-number flow past an SD7003 airfoil with and without plunge motion at Re = 60 k are presented, where transition takes place across laminar separation bubbles. The experimental data consist of high-resolution, phase-locked particle image velocimetry measurements in a wind tunnel and a water tunnel. The numerical simulation approach includes transition prediction which is based on linear stability analysis applied to unsteady mean-flow data. The numerical results obtained for steady onflow are validated against particle image velocimetry data and published force measurements. Good agreement is obtained for specific turbulence models. Flows with plunge motion reveal strong effects of flow unsteadiness on transition and the resulting laminar separation bubbles which are well captured in the simulations.

Jet Effects on Coflow Jet Airfoil Performance

Abstract: A control volume analysis is presented in this paper to analyze the jet effect on the coflow jet airfoil with injection and suction and on the airfoil with injection only. The formulations to calculate the duct's reactionary forces that must be included for the lift and drag calculation are given. The computational fluid dynamics solutions based on the Reynolds-averaged Navier-Stokes model are used to provide the breakdowns of lift and drag contributions from the airfoil surface force integral and jet duct's reactionary forces. The results are compared with experiment for validation. The duct reactionary forces are also validated with the result of a 3-D computational fluid dynamics calculation of the complete airfoil with jet ducts and wind tunnel walls. The study indicates that the suction occurring on the airfoil suction surface of the coflow jet airfoil is more beneficial than the suction occurring through the engine inlet such as the airfoil with injection only. For the airfoil with injection only, the drag actually acted on the aircraft, or the equivalent drag, is significantly larger than the drag measured by the wind tunnel balance due to the ram drag and captured area drag when the jet is drawn from the freestream. For a coflow jet airfoil, the drag measured by the wind tunnel balance is the actual 2-D drag that the aircraft will experience. A coflow jet airfoil does not have the ram drag and captured area drag. For a coflow jet airfoil, the suction penalty is offset by the significant circulation enhancement The coflow jet airfoil with both injection and suction yields stronger mixing, larger circulation, more filled wake, higher stall angle of attack, less drag, and lower energy expenditure.

High-Performance Airfoil Using Coflow Jet Flow Control

Abstract: T O ACHIEVE revolutionary aircraft performance, advanced technologies should be pursued to drastically reduce the weight of aircraft and fuel consumption and significantly increase aircraft mission payload and maneuverability. Both the military and commercial aircraft will benefit from the same technology. Flow control is a promising technology to break through the limits of the conventional aerodynamic concepts. Recently, a novel active airfoil flow control concept with zero-net mass flux, the coflow jet (CFJ) airfoil, has been developed by Zha et al. [1–5]. The CFJ airfoil achieves three effects simultaneously in a dramatic fashion: lift augmentation, stall margin increase, and drag reduction. The energy expenditure of the CFJ airfoil is low [1], and the CFJ airfoil concept is straightforward to implement. The CFJ airfoil may create a new concept of an “engineless” airplane, which uses the CFJ to generate both lift and thrust without conventional propulsion systems of propellers or jet engines [6]. A CFJ airfoil [1–5] uses an injection slot near the leading edge (LE) and a suction slot near the trailing edge (TE) on the airfoil suction surface. Similar to tangential blowing, the LE jet is in the same direction of the main flow, but the same amount of mass flow that is injected is removed via suction near the TE, resulting in zeronet mass-flux flow control. A proposed fundamental mechanism [2] is that the severe adverse pressure gradient on the suction surface strongly augments turbulent mixing between the main flow and the jet [7]. The mixing then creates the lateral transport of energy from the jet to the main flow and enables the main flow to overcome the large adverse pressure gradient and remain attached at high angle of attack (AOA). The stall margin is hence significantly increased. At the same time, the high-momentum jet drastically increases the circulation, which significantly augments lift, reduces drag, or even generates thrust (negative drag). The objective of this paper is to demonstrate the high performance of the CFJ airfoil with the windtunnel test results.

Flight testing of Mission Adaptive Compliant Wing

Abstract: This paper describes flight test results of a “Mission Adaptive Compliant Wing” (MACWing) variable geometry trailing edge flap in conjunction with a natural laminar flow airfoil. The MAC-Wing technology provides lightweight, low power, variable geometry reshaping of the upper and lower flap surface with no seams or discontinuities. In this particular program, the airfoil-flap system is optimized to maximize the laminar boundary layer extent over a broad lift coefficient range for endurance aircraft applications. The expanded “laminar bucket” capability allows the endurance aircraft to significantly extend their range (15% or more) by continuously optimizing the wing L/D throughout the mission. The wing was tested at full-scale dynamic pressure, full scale Mach, and reduced-scale Reynolds Numbers on Scaled Composites’ White Knight aircraft. Test results confirmed laminar flow regime up to approximately 60% chord for much of the lift range. Analysis and test results suggest significant fuel savings, weight savings and a higher control authority. Preliminary drag results, future aerodynamic applications and vehicle performance projections are discussed.

Desirable Airfoil Features for Smaller-Capacity Straight-Bladed VAWT:

Abstract: In the small scale wind turbine market, the simple straight-bladed Darrieus type vertical axis wind turbine (SB-VAWT) is very attractive for its simple blade design. A detailed aerodynamic performance analysis was conducted on a smaller capacity fixed-pitch SB-VAWT. Brief analyses of the main aerodynamic challenges of this type of wind turbine were first discussed and subsequently the authors conducted further literature survey and computational analysis to shortlist aerodynamic characteristics of a desirable airfoil for a self-starting and better performing SB-VAWT. The required geometric features of the desirable airfoil to achieve the short listed characteristics were also discussed. It has been found out that conventionally used NACA symmetric airfoils are not suitable for smaller capacity SB-VAWT. Rather, it is advantageous to utilize a high-lift and low-drag asymmetric thick airfoil suitable for low speed operation typically encountered by SB-VAWT.

Numerical investigations of injection-slot-size effect on the performance of coflow jet airfoils

Abstract: Three co-flow jet (CFJ) airfoils with injection slot size differed by two times consecutively are calculated by using a RANS CFD solver with 1-equation Spalart-Allmaras model. At the same angle of attack(AoA), the twice larger injection slot size airfoil passes about twice larger jet mass flow rate with the momentum coefficients also nearly doubled. The CFJ airfoil with the largest slot size has the least stall angle of attack(AoA). When the injection slot size is reduced by half, the stall AoA and the maximum lift coefficient is increased. However, when the injection slot size is further reduced by half, the stall AoA is still increased, but the maximum lift coefficient is lower due to the smaller momentum coefficient. The trend of the stall AoA and maximum lift agree with the experiment. At low AoA, both the computed lift and drag agree fairly well with the experiment. At high AoA, both the lift and drag are underpredicted. The reason may be that the RANS model can not handle the turbulence mixing well at high AoA.

Autonomous Sensing and Control of Wing Stall Using a Smart Plasma Slat

Abstract: DOI: 10.2514/1.24057 The concept of a self-governing smart plasma slat for active sense and control of flow separation and incipient wing stall is presented. The smart plasma slat design involves the use of an aerodynamic plasma actuator on the leading edge of a two-dimensional NACA 0015 airfoil in a manner that mimics the effect of a movable leading-edge slat of a conventional high-lift system. The self-governing system uses a single high-bandwidth pressure sensor and a feedback controller to operate the actuator in an autonomous mode with a primary function to sense and control incipient flow separation at the wing leading edge and to delay incipient stall. Two feedback control techniques are investigated. Wind tunnel experiments demonstrate that the aerodynamic effects of a smart actuator are consistent with the previously tested open-loop actuator, in that stall hysteresis is eliminated, stall angle is delayed by 7 deg, and a significant improvement in the lift-to-drag ratio is achieved over a wide range of angles of attack. These feedback control approaches provide a means to further reduce power requirements for an unsteady plasma actuator for practical air vehicle applications. The smart plasma slat concept is well suited for the design of low-drag, quiet, highlift systems for fixed-wing aircraft and rotorcraft.

Morphing wing micro-air-vehicles via macro-fiber-composite actuators

Abstract: *† ‡ § ** The design and implementation of a morphing unmanned aircraft using smart materials is presented. Articulated lifting surfaces and articulated wing sections actuated by servos are difficult to instrument and fabricate in a repeatable fashion on thin, composite wing microair-vehicles. Assembly is complex and time consuming. A type of piezoceramic composite actuator commonly known as Macro Fiber Composite is used for wing morphing. The actuation capability of this actuator on fiberglass unimorph was quantified by experimentation. Wind tunnel tests were performed to compare conventional trailing edge control surface effectiveness to an MFC actuated wing section. The continuous surface of the MFC actuated composite wing produced lower drag and wider actuation bandwidth. The MFC actuators were implemented on a 0.76 m wingspan aircraft. The remotely piloted experimental vehicle was flown using two MFC patches in an elevator/aileron (elevon) configuration. Preliminary testing has proven the stability and control of the design. I. Introduction This study investigates the use of macro fiber composites (MFCs) to control the roll and pitch maneuvers of micro air vehicles (MAVs). Design, manufacturing, and the control of MAVs in unsteady aerodynamic loading remain an active area of interest to researchers. The authors aim to understand the behavior of MFC actuated micro air vehicles under low speed, quasi steady air flow. Wind tunnel tests were conducted to quantify the effectiveness of MFC actuators. Results show that MFC actuation does have improved efficiency over a conventional control surface. The main goal for these experiments was to show the improved performance of a variable camber airfoil compared to a conventional control surface at low Reynolds Numbers. An experimental MAV designed and built by the authors was used as a test platform for the morphing wing concept. This paper first covers the background of the research. The next section presents the wind tunnel experimentation setup and test results. Next, the experimental aircraft design and initial flight results are presented. The paper concludes with a summary of results and discussion of future work.

Two-Dimensional Wind Tunnel and Computational Investigation of a Microtab Modified Airfoil

Abstract: *† ‡ A computational and wind tunnel investigation into the effectiveness of a microtab-based aerodynamic load control system is presented. The microtab-based load control concept consists of a small tab, with a deployment height on the order of 1% of chord, which emerges approximately perpendicular to a lifting surface in the vicinity of the trailing edge. Lift mitigation is achieved by deploying the tabs on the upper (suction) surface of a lifting surface. Similarly, lift enhancement can be attained by tab deployment on the lower (pressure) surface of a lifting surface. A sensitivity analysis using Reynolds-averaged NavierStokes methods was conducted to determine optimal sizing and positioning of the tabs for active load control at a chord Reynolds number of 1.0 million for the S809 baseline airfoil. These numerical simulations provide insight into the flow phenomena that govern this promising load control system and guided tab placement during the wind tunnel study of the S809 airfoil. The numerical and experimental results are largely in agreement and demonstrate that load control through microtabs is viable. Future efforts will include a study of the unsteady load variations that occur during tab deployment and retraction, and three-dimensional issues involving spanwise tab placement and tab gaps.

Wind tunnel test on wind turbine airfoil with adaptive trailing edge geometry

Abstract: ,A wind tunnel test of the wind turbine airfoil Ris oe-B1-18 airfoil equipped with an Adaptive Trailing Edge Geometry (ATEG) was carried out. The ATEG was made by piezo electric actuators attached to the trailing edge of a non-deformable airfoil and controlled by an amplifier. The airfoil was tested at Re = 1.66x10 6 . Steady state and dynamic tests were carried out with prescribed deflections of the ATEG. The steady state tests showed that deflecting the ATEG towards the pressure side (posi tive β) translated the lift curve to higher lift values and deflecting the ATEG towards the suc tion side (negative β β β β) translated the lift curve to lower lift values. Furthermore, cd was almost unaffected by the ATEG actuation. Testing the airfoil for a step change of the ATEG f rom β=-3.0 to +1.8 showed that the obtainable Δcl was 0.10 to 0.13 in the linear part of the lift cu rve. Modeling the step response with an indicial function formulation showed that t he time constant in the step change and in sinusoidal deflections in dimensionless terms was T0* =0.6. Testing the ability of the ATEG to cancel out the load variations for an airfoil in si nusoidal pitch motion showed that it was possible to reduce the amplitude with around 80% from Δ Δ Δ Δcl=0.148 to Δcl=0.032.

Design and Wind-Tunnel Verification of Low-Noise Airfoils for Wind Turbines

Abstract: A method for the prediction of the airfoil trailing-edge far-field noise is presented. The model employs the airfoil analysis code XFOIL to determine the initial and boundary conditions for a subsequent boundary-layer analysis using the finite-difference code EDDYBL featuring a Reynolds stress turbulence model that finally provides the input data for the noise prediction by a modified TNO Institute of Applied Physics model. The prediction scheme was applied in the European silent rotors by acoustic optimization project to design new, quieter airfoils for the outer blade region of three different wind turbines in the megawatt class. The objective was to reduce the airfoil self-noise without loss in aerodynamic performance

Separation Control by Self-Activated Movable Flaps

Abstract: Separation control is an important issue in the physiology of birdflight. Here, the adaption of the separation control mechanism by bird feathers to the requirements of engineering applications is described in detail. Self-activated movable flaps similar to artificial bird feathers represent a high-lift system for increasing the maximum lift of airfoils. Their effect on the unsteady flow around a two-dimensional airfoil configuration is investigated by a joint numerical and experimental study. First, attention is paid to the automatic opening and closing mechanism of the flap. Following this, its beneficial effect on lift is investigated for varying incidences and flap configurations. In-depth analysis of experimental and numerical results provides a detailed description of the important phenomena and the effect of self-adjusting flaps on the flow around the airfoil. In the second part of this paper, a contribution is made to verification of the applicability of unsteady Reynolds-averaged approaches using statistical turbulence models for unsteady flows with particular attention to turbulent time scales with comparison to the results of a hybrid simulation based on unsteady Reynolds-averaged Navier-Stokes equations and large-eddy simulation. Finally, flight experiments are described using an aircraft with movable flaps fitted on its laminar wing.

Design and Testing of a Pneumatic Telescopic Wing for Unmanned Aerial Vehicles

Abstract: This paper discusses the design, development, and testing of an unmanned aerial vehicle pneumatic telescopic wing that permits a change in the wingspan, while simultaneously supporting structural wing loads. The key element of the wing is a pressurized telescopic spar able to undergo large-scale spanwise changes while supporting wing loadings in excess of 15 lb/ft 2 . The wing cross section is maintained by NACA 0013 rib sections fixed at the end of each element of the telescopic spar. Hollow fiberglass shells are used to preserve the spanwise airfoil geometry and ensure compact storage and deployment of the telescopic wing. A full-scale telescopic wing assembly was built and tested in the Glenn L. Martin Wind Tunnel at the University of Maryland. These tests included aerodynamic measurements at a variety of Reynolds numbers. The telescopic wing was tested in three different configurations and experimental results are compared with finite wing theory and results obtained on a rigid fixed-wing counterpart Preliminary aerodynamic results were promising for the variable wingspan telescopic wing. As expected, the telescopic wing at maximum deployment incurred a slightly larger drag penalty and a reduced lift-to-drag ratio when compared to its solid fixed-wing counterpart. However, the penalty was minimal and thus the development of an unmanned aerial vehicle with a pneumatic variable span wing is feasible.

Vortex lock-in phenomenon in the wake of a plunging airfoil

Abstract: The flow over a NACA0012 airfoil, oscillated sinusoidally in plunge, is simulated numerically using a two-dimensional Navier-Stokes solver at a Reynolds number of 20,000. The wake of the airfoil is visualized using a numerical particle tracing method for high reduced frequencies (1.0

Jet switching phenomenon for a periodically plunging airfoil

Abstract: An experimental investigation has been carried out on rigid and flexible airfoils oscillating in still fluid. It was found that the vortex pairs generated by the oscillating airfoil move at an angle to the chordwise direction. The deflection angle of the induced jet was observed to change periodically in time. The switching period was found to increase with increasing airfoil stiffness and to decrease with increasing heave frequency and increasing amplitude. Over the range of frequency, amplitude, and stiffness tested, the switching period was found to be two orders of magnitude greater than the heave period. The development of the vorticity field for upward and downward deflected jets, as well as the transition between the two modes, was captured with the particle image velocimetry measurements. The deflection of the jet, and thus the jet switching effect, was found to diminish with increasing free stream velocity (decreasing Strouhal number).