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Design and testing of a deformable wind turbine blade control surface

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2012 Smart Mater. Struct. 21 105019

(http://iopscience.iop.org/0964-1726/21/10/105019)

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Smart Mater. Struct. 21 (2012) 105019 (10pp) doi:10.1088/0964-1726/21/10/105019

Design and testing of a deformable windturbine blade control surfaceS Daynes and P M Weaver

Department of Aerospace Engineering, University of Bristol, Bristol, BS8 1TR, UK

E-mail: stephen.daynes@bristol.ac.uk

Received 7 February 2012, in final form 6 July 2012Published 20 August 2012Online at stacks.iop.org/SMS/21/105019

AbstractWind tunnel tests were conducted on a 1.3 m chord NACA 63–418 blade section fitted with anadaptive trailing edge flap. The 20% chord flap had an aramid honeycomb core covered with asilicone skin and was actuated using servo motors. The honeycomb core had a high stiffness inthe thickness direction but was compliant in chordwise bending. These anisotropic propertiesoffer a potential solution for the conflicting design requirements found in morphing trailingedge structures. Static and dynamic tests were performed up to a Reynolds number of5.4× 106. The tests showed that deflecting the flap from −10◦ to +10◦ changes the bladesection lift coefficient by 1.0 in non-stalled conditions. Dynamic tests showed the flap to becapable of operating up to 9◦ s−1 using a 15 V power supply. A two-dimensional staticaeroelastic model of the morphing flap was developed to analyse strains, predict actuatorrequirements and study fluid–structure interaction effects. The model was used to conductparametric studies to further improve the flap design. Potential applications include windturbine blade load alleviation and increased wind energy capture.

(Some figures may appear in colour only in the online journal)

1. Introduction

The aerodynamic loading on wind turbine blades is highlyvariable and can fluctuate rapidly. This variability in loadingis caused by factors including yawed operation, gusts,wind shear and turbulence [1]. Aerodynamic loads can besignificantly reduced by controlling the pitch of the entireblade [2] but blade inertia restricts the speed at which a bladecan be actuated to pitch in response to rapidly fluctuatingloads. This problem is made worse by the trend with offshorewind turbines to increase the rotor diameter as much aspossible to decrease the costs per kilowatt-hour of windenergy. In addition, blade pitch change cannot control localvariations in aerodynamic loading at different radial positionsalong the length of a blade. There has been significant researchinterest in using trailing edge control surfaces to achievefurther load reduction beyond blade pitch control alone [3–5].Early attempts focused on using ‘conventional’ mechanicalflap devices similar to those used in the aerospace industry [6].However, the use of such devices does raise concerns dueto the added weight, complexity, power requirements and

potential increase in aeroacoustic noise and drag caused bydiscontinuities on the external blade profile [7].

More recent research projects have focused on adaptivetrailing edge devices with novel ‘morphing’ structures toaddress these problems. There are far too many concepts tolist in detail in this work but several good review papers canbe found for both the application of morphing structures toaircraft [8–10] and also more specifically to wind turbineblades [11, 12]. Morphing trailing edges for wind turbineblades have been developed at Risø National Laboratory usingpiezoelectric actuators and wind tunnel tested at 40 m s−1 [13,14]. The use of piezoelectric patches enabled a lightweightflap structure and flap deflections at up to 10 Hz althoughonly a 5◦ flap angle range was achievable. A similar conceptfor wind turbine blades has also been developed at DelftUniversity using a scaled model with piezoelectric patchesembedded in a foam core to achieve a smooth tapered trailingedge [15]. Another adaptive blade solution investigated atthe Risø National Laboratory involved the development ofa rubber trailing edge with internal reinforced voids [16].The deflection of the flap can be pneumatically controlledby varying the pressure in the voids between ±8 bar. A 1 m

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Smart Mater. Struct. 21 (2012) 105019 S Daynes and P M Weaver

chord, 1.9 m span model was wind tunnel tested with a 0.15 mchord flap. A change in lift coefficient of 0.2 was achieved.One of the main benefits of this design is its simplicity andthe absence of moving parts.

Unlike an actuator for a conventional flap, an actuator fora morphing flap structure must not only fight aerodynamicloads but also work against the stiffness of the structure itself.It is therefore important that the morphing flap structure issufficiently compliant so that it does not significantly restrictfree movement of the actuator and limit flap deflections.However, there is also the conflicting need for the flapto be stiff enough so that it does not have detrimentallocal deformation under air loads. Hence a properlydesigned adaptive flap structure must balance the conflictingrequirements of compliance and stiffness [17]. One successfulsolution to this trade-off was demonstrated during the SmartWing Phase 2 program where a compliant trailing edgewas developed with a highly anisotropic honeycomb corecovered with a silicone skin [18, 19]. Honeycomb corestructures are excellent candidates for adaptive trailing edgesbecause they display a high out-of-plane stiffness whichcan resist aerodynamic loads whilst having a low flexuralstiffness so as to not result in a large actuation penalty [20,21]. Gandhi explored the trade-off between these conflictingrequirements with respect to strain and deflection constraintsin the morphing skins of trailing edge structures which aresubject to both actuation and aerodynamic loads [22].

This paper describes the development of a novelmorphing flap design with a honeycomb core specificallydesigned for use on a wind turbine blade as an aerodynamiccontrol device by enabling blade section camber change.A full size flap section model is manufactured, fitted withactuators and wind tunnel tested up to 56 m s−1. The bladegeometry and airspeeds studied are comparable to a 70%blade radial location of a generic 90 m diameter wind turbineoperating at high wind speeds. The aim of the wind tunneltesting is to study the aerodynamic and structural performanceof the flap when it is subject to aerodynamic loading. Previousdevelopment of this morphing flap concept has involvedoperating the device in a ‘bench-top’ environment where theflap was deflected ±15◦ under simulated aerodynamic loadsusing extension springs attached at the trailing edge [23]. It isimportant to include both the actuator and the aerodynamicloads during the design of the morphing flap to selectan appropriate stiffness distribution. A morphing flap willsignificantly deform under aerodynamic loading if it isnot sufficiently stiff. Conversely, a morphing flap whichis too stiff may have prohibitive actuator requirements. Inorder to study this trade-off a static aeroelastic model isused which couples a two-dimensional finite element (FE)model with an aerodynamic solver to enable parametricdesign studies. The following sections provide details of thedesign and manufacture of the morphing flap structure andalso the performance of the actuators and the aerodynamicperformance of the blade section. The paper finishes bypresenting an improved flap design based on the modellingand testing undertaken.

Figure 1. A morphing flap segment with a view of the internalstructure.

2. Trailing edge development

2.1. Initial flap design

The morphing flap consists of three main structural elements:an aramid honeycomb core, an upper carbon fibre reinforcedplastic (CFRP) skin and a lower silicone skin, figure 1.The silicone skin also extends to the side walls ofeach 250 mm long flap segment to create a smooth,seamless, impermeable membrane around the structure. Thehoneycomb core has highly anisotropic properties resultingin a core with a high resistance to out-of-plane loads in thethrough-thickness direction but maintaining a low flexuralstiffness. For simplicity of manufacture the geometry offlap is approximated as a triangular cross-section with thedirection of the honeycomb cells normal to the upper CFRPsurface and at an angle relative to the lower surface. The lowflexural stiffness of the core is important to reduce actuatorforce requirements when the flap is deflected. The hexagonalhoneycomb core is slotted in 20 mm segments in thechordwise direction to suppress the formation of anticlasticcurvature; only 2D camber change is desirable in this study.The upper CFRP skin provides both a smooth aerodynamicsurface and also a hard point into which actuator loads canbe transferred. The CFRP skin has the layup [90/0/90]T with0◦ in the chordwise direction and a total thickness of 0.4 mm.This layup is an additional source of anisotropy enabling thespanwise bending stiffness to be greater than the chordwisebending stiffness.

Actuator loads are transferred to the flap using 2 mmdiameter CFRP rods which pass freely through a rear spar andare bonded to the trailing edge. The function of the rear sparis twofold. Firstly, it separates the structural function of theleading 80% of the blade chord from the morphing functionof the aft 20%. Secondly, actuators are directly attachedto the rear spar resulting in a short load path between theactuators and the flap. Low friction U-shaped PTFE hooksare bonded with epoxy resin into the flap’s honeycomb coreto transfer loads from the CFRP rods into the honeycomb andto prevent the actuator rods from touching the silicone skin.The actuator rods are free to slide over these hooks so onlyforces normal to the rod are transferred. There are four CFRProds per metre of blade span with one for each 250 mm flapsegment, each with its own actuator. The use of CFRP rods

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Smart Mater. Struct. 21 (2012) 105019 S Daynes and P M Weaver

Figure 2. A blade section installed in the 2.1 m× 1.5 m windtunnel with α = 0◦.

to transfer actuator loads into the structure is by no meansthe only solution available; this design was chosen for itssimplicity and low mass. Other potential solutions include theuse of ‘eccentuators’ [18], antagonistically operated sets ofcables [24] or direct actuation of the skin [25]. The total massof the morphing flap in this study, including the rear spar, is5.4 kg m−1 span; 3.2 kg of this is due to the four actuators.

2.2. Wind tunnel testing of the initial flap design

The full scale wind turbine blade section model is shownin figure 2. The model has a chord length of 1.3 m, a spanof 1 m and a NACA 63-418 profile [26]. The model wasmanufactured to demonstrate the ability of the 20% chordflap to deflect in the range β = ±10◦ when subject to typicalin-service airspeeds and angles of attack. The morphing flapspans the entire 1 m long model and each flap segment isattached to the blade at 80% chord using detachable rearspars. The CFRP rear spars have compliant flanges so theycan be individually inserted or removed from the blade in thechordwise direction without removing the entire flap.

The main section of the wind tunnel model up to 80%chord is constructed by forming a beech plywood skin overa plywood stringer–rib internal structure. End plates are thenattached to each end of the blade section to reduce induceddrag. An empirical correction factor is introduced later in theresults section to compensate for induced drag to calculatethe 2D section characteristics of the morphing blade. Theblade section is installed in the wind tunnel vertically tominimize blockage effects and to make it easier to measureaerodynamic forces and moments. Forces and moments aremeasured using an OR6-7 2000 force platform from AdvanceMechanical Technology, Inc., mounted at the base of themodel. The 25% chord location of the wind tunnel model’sblade section is positioned to be aligned with the verticalaxis of the force platform about which pitching moments aremeasured. The base of the force platform rests on a metal platewith bolt holes positioned in it so that the force platform canbe rotated through ±18◦ in 3◦ increments. The force platformoutputs six analogue voltage signals corresponding to forcesand moments in all three orthogonal axes. These analoguevoltage signals from the force platform are interfaced with a

PC using a dSPACE ControlDesk unit. Calibration of the sixsignals is performed using Simulink by MathWorks enablinglift, drag and pitching moment about the blade’s 25% chord tobe calculated.

As well as aerodynamic forces, aerodynamic pressuredistributions are also measured. The section static pressuredistribution around the blade’s surface is measured using32 pressure tappings placed around the blade at 30% spanfrom the base. The pressure tappings consist of brass tubeswith an outer diameter of 1.59 mm and a wall thickness of0.36 mm. The pressure tappings are linked to piezoresistivepressure transducers to acquire the pressure distributionselectronically. The measurement and data acquisition system(Scanivalve Corporation) uses an electronic pressure scanner(ZOC22B32PxX2) connected via plastic tubes to the pressuretappings with an analogue to digital converter (RAD3200)connected to a PC. Data are acquired via ‘Radlink’ softwareat a rate of 10 Hz for 5 s.

The morphing flap is able to be deflected at up to 9◦ s−1.This is equivalent to 0.45 Hz at±5◦. The primary focus of thisresearch is the aerodynamic and structural performance of theflap and not the method of actuation which is an ‘off-the-shelf’solution. The actuators selected are i00600 heavy duty rotaryservos from Invenscience LC which have an operating voltageof 15 V and a maximum torque of 11 N m when in continuousoperation. As well as sufficient torque, one of the other criticaldesign requirements is finding an actuator which is smallenough to be inserted into the blade at 80% chord. This placesa through-thickness dimensional constraint of 70 mm onactuator size. The servos are controlled from a PC via a servomotor controller which interfaces the USB communicationfrom the PC with the servo communication which operates onpulse width modulation. In total four servo motors are usedto actuate the morphing flap model with the flap segmentedinto four 250 mm lengths which can operate independently.Segmentation of the flaps is desirable in the current designbecause it distributes actuator loads over the morphing flap.Also, it is possible that a similar in-service wind turbine flapwould need to be segmented into such lengths to account forthe bending and twisting of the wind turbine blade structurewhen subject to operational loads.

Wind tunnel tests are performed in Bristol University’slarge low speed, closed-circuit tunnel which has an octagonal2.1 m × 1.5 m working section. The maximum dynamicpressure which can be tested with this size of model isq = 1.9 kPa. This is equivalent to 56 m s−1 airspeed atsea level and a Reynolds number of 5.4 × 106. Pitot tubesare used to measure static and total pressures in the windtunnel’s test section and temperature is also recorded foreach test to correct for air density. Three of the possibleflap configurations which are achievable with this morphingtrailing edge are shown in figure 3 prior to wind tunnelinstallation. Of greatest interest in this current research isfor all of the flaps to deflect an equal amount to achievetwo-dimensional flow over the blade section, see figure 3(a).The modelling approach used in this work is only valid forthis two-dimensional flow regime. However, a segmented flapenables additional functionality by including ‘twist’ modes

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Smart Mater. Struct. 21 (2012) 105019 S Daynes and P M Weaver

Figure 3. Trailing edge geometries: (a) uniform 5◦ deflection, (b) ‘twist’ mode, (c) ‘braking’ mode.

where the effective camber of the blade is varied as afunction of span, see figure 3(b). ‘Braking’ modes are alsopossible whereby the flaps are deflected in opposing directionsresulting in little change in total lift but a large increase indrag, figure 3(c). Such a configuration could potentially beused in wind turbine blades for controlling the turbine’s speedof rotation.

3. The fluid–structure interaction (FSI) model

The morphing flap demonstrator presented in section 2 wasdesigned based on the results from an FSI model. In order totailor the stiffness of the morphing flap there is a requirementto model the influence of both aerodynamic and actuator loadson the flap [27–32]. This section describes details of the staticaeroelastic models developed for this task and includes anon-linear detailed analysis of individual flap designs and aless accurate linear analysis for the faster parametric analysisof multiple design variables.

3.1. The non-linear FSI model and design variables

The non-linear FSI model couples a two-dimensionalABAQUS FE structural model with a two-dimensionalviscous XFOIL panel code aerodynamic model. The FE andXFOIL simulations are run separately with the solution of onepassed to the other using a fully automated MATLAB script.Static aeroelastic equilibrium is reached when the differencebetween the aerofoil section lift coefficient at iteration n(cln ) and the previous iteration n − 1 (cln−1 ) falls belowa given tolerance. In this work the required tolerance is|cln − cln−1 | ≤ 0.01 which typically took between 3 and 6iterations to achieve. The analysis in this paper is restrictedto two dimensions where only the chordwise distribution ofaerodynamic loads and structural deformations is considered.This assumption is valid only for the flap operating with

uniform deflection (figure 3(a)) and is not valid during ‘twist’or ‘braking’ deflections (figures 3(b) and (c)) where the flowregime is three-dimensional.

The structural and aerodynamic models used havedifferent meshes so loads and geometries are passed betweenthe models using interpolation functions in the MATLABscript. The fact that the models have different meshes enableseach of the meshes to be independently discretized and refinedin the areas of most physical interest. In the case of the XFOILmodel it is important to model the flow field in detail at theleading edge since this is the region with the highest pressuregradients. The XFOIL model uses the default 140 nodes in thesoftware with the ‘PANE’ command used to create a higherresolution panel node distribution in areas of high curvature.

The mesh distribution of the FE model is substantiallydifferent from the XFOIL model. Firstly, only the flap regionis modelled instead of the entire aerofoil section. Secondly,the mesh is refined in areas of high stress gradients. Thehoneycomb core is modelled using a 5 × 49 mesh consistingof plane strain quadrilateral elements with the elementsdecreasing in size towards the trailing edge, see figure 4. Thelength of the honeycomb core is 260 mm with its thicknessincreasing from zero at the trailing edge to 70 mm at theflap root end. Close to the tip triangular elements are usedto model the core since it proved difficult to mesh the tipwith quadrilateral elements. The upper CFRP skin and lowersilicone skin are both modelled with beam elements with theirrespective material properties assigned. All of these beam andshell elements are assigned a width of 250 mm to model oneof the flap segments. The CFRP actuator rod is modelled asa separate part using 110 equally spaced beam elements witha 2 mm diameter circular cross-section and a total length of330 mm. Interaction between the honeycomb core and theactuator rod is modelled using coupling constraints whichconstrain the distance between the rod and the lower surface to4 mm. The end of the actuator rod is constrained to the upperCFRP skin close to the flap tip and actuator movement, u, is

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Smart Mater. Struct. 21 (2012) 105019 S Daynes and P M Weaver

Figure 4. FE model showing the coupling constraints and boundaryconditions (elements shrunk by 30% for clarity).

Figure 5. The linear FSI parametric analysis procedure.

simulated with a displacement constraint applied at the otherend of the rod. A clamped boundary condition is placed at the80% chord location to model the rear spar. Aerodynamic loadsare applied using 40 equally spaced pressure loads distributedover the two aerodynamic surfaces.

The chordwise Young’s modulus of the honeycomb coreis experimentally measured as 0.4 MPa. The stiffness of thecore material in the thickness direction is 100 MPa. Thematerial properties of the honeycomb core are orientated inthe FE model so the honeycomb cells are normal to theupper aerodynamic surface. All Poisson’s ratios are set to zeroin the core to model the segmented honeycomb core whichis experimentally shown to have no observable anticlasticcurvature in bending. The silicone skin has a thickness of0.5 mm and a Young’s modulus of 1 MPa and the CFRP rodhad a longitudinal Young’s modulus of 150 GPa and a radialYoung’s modulus of 10 GPa. The [0/90/0]T upper CFRP skinhas a flapwise bending stiffness of 2.5 × 104 N mm2. Themain design variables during the design of the manufactureddemonstrator were the bending stiffness of the upper CFRPskin, the chordwise bending stiffness of the honeycomb core,the diameter of the CFRP actuator rod, the chordwise flaplength and the flap segment span since these were foundto have the greatest influence on the flap’s fluid–structureresponse and actuator requirements.

A convergence procedure based on lift coefficient waschosen for this model since it is convenient to work

Figure 6. Experimental and FE force–displacement characteristics(the direction of the hysteresis loops is clockwise).

with a single variable. Both the manufactured demonstratorand the FE model have the same geometry and materialproperties. It will be shown later in the paper that thesesimilarities result in both models having very similar stiffnesscharacteristics (figure 6) and deflected shapes (figure 8).Comparison between the experimental and numerical pressuredistributions in figure 8 shows that there is close agreementwhich supports this analysis approach.

3.2. The linear FSI model for parametric analyses anddesign refinements

The non-linear FSI model was used for the design ofthe manufactured wind tunnel model. However, sincethis time, a second FSI model has been developed toconduct parametric analyses. The objective of the parametricanalysis model is to develop an automated tool for thestiffness tailoring of morphing flaps. Development of themanufactured concept was primarily based on ‘trial anderror’ whereby the non-linear FSI model was used to find aflap design with an acceptable trade-off of small deflectionsunder aerodynamic loading without prohibitive actuatorrequirements or prohibitive bending strains. A procedurebased on two iterative loops used in the parametric analysisis shown in figure 5 and described in this section.

The procedure involves deflecting the flap until a targetlift coefficient, cltarget , is achieved. The first step (n = 1) in theprocedure is to calculate the aerofoil section lift coefficient(cl1 ) in XFOIL prior to the flap being deflected (u1 = 0).Here u is used to define the actuator stroke, see figure 4.In a second step the flap is deflected to an initial guessposition (u2 = 10 mm.) At this stage aerodynamic loads arenot yet applied to the FE model (q = 0). After this stage aniterative procedure between XFOIL and ABAQUS begins forn iterations with each iteration successively calculating clnand un. Convergence is achieved by modifying un using asimple Newton–Raphson procedure until cln reaches a targetlift coefficient, cltarget . In this work the angle of attack of theblade is set at 5◦ and cltarget = 1.5. This represents a maximumtypical ‘flap-down’ operating condition. Once convergence isreached a second iterative loop is initiated with aerodynamic

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Smart Mater. Struct. 21 (2012) 105019 S Daynes and P M Weaver

loads now applied to the ABAQUS model (q 6= 0) withthe actuator position held. Successive iterations stop oncea converged fluid–structure solution is found. This secondloop in the analysis procedure enables the flap geometry,actuator loads and actuator rod strains to be compared withand without aerodynamic loads applied. Typically the wholeprocess requires approximately 6–9 iterations and takes 15 sto compute on a typical modern PC (Intel Core i5). Usingthe viscous flow solver in XFOIL would not take appreciablylonger and would result in more accurate aerodynamic loaddistributions. However, this procedure frequently resulted innon-converged solutions or stalling whereby cltarget could notbe achieved. For this reason it was less convenient to use aviscous solver in conjunction with the FE model.

The results generated by the FE model are furtheranalysed by the MATLAB script upon finding a convergedsolution. The MATLAB script reads the ABAQUS resultfile (*.dat) and records the maximum bending strain in theactuator rod and the required actuator force at the end of theactuator rod. The change in lift coefficient (cln − cltarget) dueto the application of aerodynamic loads to the FE model isalso stored. Flap deflection under aerodynamic loading is notdirectly considered as important in this analysis; instead thechange in lift due to FSI is a key design consideration sincevariability in lift characteristics is of primary importance andnot for the flap to maintain a predefined shape.

4. Comparison between the non-linear FSI modeland wind tunnel results of the initial flap design

In this section four different sets of results are presented:(1) stiffness characterization of the flap using a bench-topmodel; (2) steady state aerodynamic characteristics of theblade section for various flap deflections and various anglesof attack; (3) the steady state pressure distributions aroundthe blade for different angles of attack and flap deflections;(4) actuation requirements and also the strains in the actuatorrods.

4.1. Stiffness characterization

In total five flap segments were manufactured based onthe initial flap design; four were used on the windtunnel model and one was used as a bench-top model tocharacterize the stiffness and deflected shape of the flap.The bench-top model allowed initial investigations into theactuator force–displacement performance required to deflectthe flap model prior to wind tunnel testing and actuatordevelopment [23]. The flap segment is orientated verticallyon a test rig to enable a direct connection with an Instron1 kN load–displacement test machine, see figure 6. The testrig consists of a rectangular aluminum frame with the rearspar of the model clamped to the upper horizontal member.The actuator rod passes through this horizontal aluminummember and is connected to the moving cross-head on the testmachine. The lower horizontal member is bolted to the baseof the test machine.

The experimental force–displacement characteristics ofthe flap are shown in figure 6. It was found that the flapangle, β, in degrees can be approximated as linearly relatedto the actuator stroke in millimetres, u, by the expressionu = 1.0 × (β − 4.6). The test machine was programmed tooperate cyclically between ±10◦ at two different frequencies:0.025 Hz representing a quasi-static test and the maximumspeed of the test machine at 0.4 Hz. A hysteretic response inthe force–displacement results can be observed for both setsof experimental results. Hysteresis in the results is caused byfriction between the carbon actuator rod and the U-shapedPTFE restraints which are bonded into the core. PTFE wasused for these restraints due to its extremely low coefficientof friction to minimize any hysteresis. The non-linear FEforce–displacement results for the manufactured flap designare also shown in figure 6 in blue. These results are in closeagreement with the experimental results and also display alinear relationship between the stroke u and the angle β.

4.2. The steady state aerodynamic characteristics of theblade section

The experimental wind tunnel results are shown in figure 7for section lift, cl, drag, cd, and pitching moment about 1/4chord, cm1/4 . These results were obtained using the forceplatform at an airspeed of 40 m s−1 (q = 1 kPa). Theresults displayed correspond to time averaged data measuredover an 8 s period at a sampling frequency of 100 Hz. Nosignificant flap oscillations due to dynamic fluid–structureinteraction effects were apparent from the data obtained. Bycomparing published data for the lift curve of the NACA63-418 section [26] it was clear that the raw experimental dataobtained showed that the wind tunnel model was experiencinginduced drag with the gradient of the experimental lift curvesreduced. Therefore, an empirical correction of −1.0c2

l wasmade to the drag data to compensate for induced drag and acorrection of −sin−1(1.0cl) was made to compensate for thecorresponding induced downwash [33]. The correction factorfor the angle of incidence was determined by making theexperimental and numerical lift curves (∂cl/∂α) equal whenα = β = 0.

The corrected lift curves in figure 7(a) show that thelift changes with flap deflection by approximately ∂cl/∂β =

0.05/deg in the linear regions of the lift curves. The changein cl with β tends to diminish for larger flap angles mostlikely due to the onset of flow separation. In the linear partof figure 7(a) the gradient of the lift curves is approximately∂cl/∂α = 0.1/deg. The lift curve results generated by thenon-linear FSI model are also displayed in figure 7(a). The FSIlift results are in close agreement although the experimentallyobserved change in lift with flap deflection is slightly less thanpredicted. This discrepancy is not surprising since XFOILdoes not account for the possibility of earlier flow separationduring wind tunnel testing as a result of the surface roughnessof the model and similar experimental model imperfections.Efforts were made to make the model as smooth as possible.However, visible imperfections were clear in the painted woodfinish and there was also a tape covered seam where the flapinterfaced with the blade which was not completely flush.

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Smart Mater. Struct. 21 (2012) 105019 S Daynes and P M Weaver

Figure 7. Blade section characteristics: (a) cl versus α, (b) cd versus cl, (c) cmc/4 versus α.

Figure 8. Pressure distributions around the blade section for (a) β = −10◦ and (b) β = 10◦.

It should also be noted that XFOIL predictions for anglesof attack beyond the maximum and minimum lift angles arenot accurate so the performance of the aerofoil at anglesof attack outside of the range ±10◦ should be thoughtof as less reliable and non-conservative. Equally, it proveddifficult to acquire experimental lift data outside the angleof attack range of ±10◦ owing to the large size of the windtunnel model and the space restrictions we had in the windtunnel. As a first approximation the section lift coefficientof the aerofoil for angles of attack outside this range can beapproximated using cl = 2π sin(α). This solution is derivedusing conformal mapping and the Joukowski transformationfor flat aerofoils [34].

As mentioned in section 1, one of the key conflictingrequirements for structural morphing is the trade-off betweenstructural compliance and stiffness. For a given flap deflectionand angle of attack there was no measurable change in cl forthe range of airspeeds achievable by the wind tunnel (q ≤1.9 kPa). However, figure 7(a) shows that when the non-linear

FSI model is subject to a freestream dynamic pressure ofq = 3 kPa then there is some variability in cl. Variability in thepitching moment coefficient, cm1/4, with dynamic pressure iseven more discernable in figure 7(c).

The wind tunnel model’s drag characteristics are plottedin figure 7(b) after corrections are made for induced drag.Comparison of the β = 0 results with published data forthe NACA 63-418 section [26] shows that there is generalagreement with the shape of the drag profile although theexperimentally measured section drag coefficients are higherby an offset of approximately 0.015. The most likely causeof this discrepancy between the two sets of results is due tothe drag forces exerted on the end plates and on the forceplatform as well as skin roughness. Figure 7(b) also showsthat deflecting the flaps to 10◦ causes the largest increase indrag when the four flap segments are deflected by an equalamount. If the flaps are not deflected by an equal amount,but by ±10◦ in the case of the ‘braking’ mode in figure 3(c),then there is a further increase in drag. In this configuration

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Smart Mater. Struct. 21 (2012) 105019 S Daynes and P M Weaver

Figure 9. Quasi-static actuation characteristics: (a) force requirements and (b) maximum rod strains.

figures 7(a) and (c) show that the equal and opposite flapdeflections have only a small effect on either cl or cm1/4 . Theseaerodynamic coefficients should be interpreted as averagedvalues taken across the 1 m span of the model. The abilityto aerodynamically ‘brake’ a wind turbine blade could be auseful additional function that a multiple flapped blade has toslow down the rotation of a wind turbine.

The pitching moment coefficients, cm1/4, about 25%chord in figure 7(c) are similar for both the non-linearFSI model and the wind tunnel model, although the FSIresults tend to calculate more variability in pitching momentwith flap deflection compared to the results from the windtunnel model. Both sets of results show that the blade’spitching moment becomes increasingly positive for negativeflap angles. This could be of concern for blades with flapsover large percentages of their spans since positive pitchingmoments could lead to aeroelastic divergence.

4.3. Steady state pressure distributions

A selection of distributions of the section pressure coefficient,cp, are shown in figure 8. Of the 32 pressure tappings placedaround the blade section five of these were installed in themorphing flap using flexible plastic tubes embedded into thehoneycomb core: two of these tappings were connected to theupper surface, two to the lower surface and one was connectedto the trailing edge. All the pressure tappings are located at33% span from the base of the wind tunnel model. Figure 7shows that the pressure distributions around the entire sectionare affected by flap deflection. The experimentally measuredpressure distributions are compared with the viscous FSImodel with q = 4 kPa and also with the case when there isno FSI when q approaches zero. The FSI results show thatthere is very little variability in the pressure distributions inthe range 0 kPa < q ≤ 4 kPa.

4.4. Actuation

As well as the aerodynamic performance, another objective ofthe wind tunnel model is to quantify actuation performancerequirements under aerodynamic loading. The variation ofactuator load against angle of attack found experimentallyat an airspeed of 40 m s−1 and for several flap angles isshown in figure 9(a). Actuator force is calculated using a

calibrated strain gauge placed on the moment arm of theservo motor situated in the blade at 62.5% span. The datafrom the strain gauge are logged using a StrainSmart unitfrom Vishay Precision Group. The results show that thereis no discernable increase in actuator loads as the angle ofattack is increased. This apparent independence of load withangle of attack may be the result of the morphing flap havingincreasing camber towards the trailing edge when deflectedunlike a conventional hinged flap which deflects at a constantangle. Being able to bias out-of-plane deflection towards thetrailing edge is desirable since the aerodynamic pressureson both surfaces tend towards the freestream value at thetrailing edge. Similar ‘amplification’ effects have previouslybeen investigated as a means of reducing actuator workrequirements [35]. The FSI results in figure 9(a) for q = 1 kPaare in close agreement with these experimentally observedtrends. An unusual result in figure 9(a), however, is the smallincrease in actuator force between β = 5◦ and 10◦ in theexperimental results. Previous bench-top studies have shownthe relationship between quasi-static actuation force and flapdeflection to be linear, figure 6. This discrepancy may bethe result of the instrumented morphing flap segment beingdamaged towards the end of manufacture which resulted inpart of the honeycomb core de-bonding from the carbon fibreskin near the trailing edge.

The peak total strains found in the actuator rod from theFSI analyses are shown in figure 9(b). The trends in thesestrain results are very similar to those of the actuator forceresults with strains not varying significantly with angle ofattack. The highest strain in the actuator rod during windtunnel testing at q = 1 kPa is predicted to be 0.8% and upto 1.1% when q = 3 kPa. It will be shown in section 5 thatthese strains can be reduced significantly using the linear FSImodel to redesign the flap.

5. Linear FSI parametric analysis for designimprovements

The linear FSI analysis enables a large number of flap designsto be investigated without the need to manually create andanalyse multiple models. The design variables in this sectionare the Young’s modulus of the core and the bending stiffnessof the upper skin. The skin bending stiffness is varied fromclose to zero to 1 N m2 and the core Young’s modulus is

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Smart Mater. Struct. 21 (2012) 105019 S Daynes and P M Weaver

Figure 10. Parametric flap design study (q = 3 kPa, α = 5◦ andcltarget = 1.5).

varied from 0.2 to 0.6 MPa. Core chordwise stiffnesses below0.2 MPa are not modelled since this is taken as a lower boundfor feasible manufactured aramid core materials. The designcase is for the flap to be deflected until a target coefficient oflift is obtained, cltarget = 1.5, when α = 5◦ before the dynamicpressure is increased from 0 to 3 kPa. A contour plot is thengenerated which contains results for the maximum actuatorrod strain, the force in the actuator rod and the change insection lift coefficient, 1cl, when the dynamic pressure isincreased from 0 to 3 kPa, figure 10. Strains in the siliconeskin and honeycomb core are not considered as importantdesign considerations since the strains in these materials inthe flap are typically below 20% even for large deflections.Experimental testing has shown that both materials are able towithstand such strains without permanent deformation. It canbe seen in figure 10 that the interdependence of the selectedvariables is quite complex although several general trendscan be observed. It can be seen that actuator loads tend todecrease with the decrease in both core stiffness and skinbending stiffness. These observations were made during theoriginal design of the manufactured flap using the non-linearFSI model prior to this parametric analysis.

The manufactured flap design used for wind tunneltesting is highlighted in figure 10 at point A. It can be seenthan an improved flap design occurs at point B where the corechordwise stiffness is reduced to its minimum and the bendingstiffness is increased to 5 × 105 N mm2. At this locationthere are predicted to be both minimal actuator forces withF ≈ 240 N and also |1cl| ≈ 0 with maximum rod strainsin the region of 0.5%. Under aerodynamic loading the flaptip can both deflect vertically and rotate. At location B theinteraction of the aerodynamic loads and chordwise stiffnessdistribution results in a combination of vertical deflection androtation, and a corresponding redistribution of aerodynamicloads, which results in no overall change in total lift.

6. Conclusion

This paper summarizes the findings related to the structuraland aerodynamic performance of a novel morphing flapdesign. A NACA 63-418 blade section with a 20% chordmorphing trailing edge was wind tunnel tested at up to56 m s−1, airspeeds comparable to those experienced bywind turbine blade sections of a similar chord length on ageneric 90 m diameter wind turbine. The trailing edge wasconstructed with an aramid honeycomb core covered with asilicone skin. The blade section tested had a 1.3 m chord anda 1 m span. Downwash was significantly reduced using endplates attached to the ends of the blade section. 2D aerofoilsection characteristics were then obtained using an empiricalcorrection for the remaining induced drag. Four servo motoractuators were attached to the flap to study both the steadystate and the dynamic performance of the design with the flapdeflecting ±10◦. In non-stalled flow regimes deflection of theflap resulted in a change in lift coefficient of 0.05/deg. Theflap was successfully operated at up to 9◦ s−1 with a maximumcurrent draw of 10 A m−1 span and a 15 V DC supply.

A fluid–structure interaction model was developed whichcould accurately describe the bench-top and wind tunnelmodel results. Parametric fluid–structure interaction modelswere then used to further improve the morphing flap designto reduce the variability in cl with dynamic pressure, reduceactuator forces and also to reduce strains in the structure. Anew flap design was then proposed based on this modellingwhich is an improvement on the manufactured design withrespect to all three of these design features.

The wind tunnel tests have shown that it is feasibleto control the loads on a model comparable in size andaerodynamic loading to a typical wind turbine blade section.However, one of the limitations of these wind tunnel testsand the FSI models developed is that they cannot fullysimulate the dynamic loading environment experienced bywind turbine blades which are subject to forces includingcyclic gravitational loading and centrifugal forces. Also,more understanding is needed concerning the maximum flapoperating speed, operating costs and fatigue and maintenancerequirements if the flap is to be used to its full potential.

Acknowledgments

The authors would like to thank Vestas Wind Systems forfunding of this work as part of the ‘morphing compositesfor wind turbine flap applications’ project, which is gratefullyacknowledged.

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