Design, Development and Testing of Shape Shifting Wing Model

aerospace

Article Design, Development and Testing of Shape Shifting Wing Model

Dean Ninian † and Sam M. Dakka *,† Department of Engineering and Math, Shefﬁeld Hallam University, Howard Street, Shefﬁeld S1 1WB, UK; [email protected] * Correspondence: [email protected]; Tel.: +44-114-225-2085 † These authors contributed equally to this work.

Received: 19 August 2017; Accepted: 30 October 2017; Published: 1 November 2017

Abstract: The design and development of morphing (shape shifting) aircraft wings—an innovative technology that has the potential to increase the aerodynamic efficiency and reduce noise signatures of aircrafts—was carried out. This research was focused on reducing lift-induced drag at the flaps of the aerofoil and to improve the design to achieve the optimum aerodynamic efficiency. Simulation revealed a 10.8% coefficient of lift increase for the initial morphing wing and 15.4% for the optimized morphing wing as compared to conventional wing design. At angles of attack of 0, 5, 10 and 15 degrees, the optimized wing has an increase in lift-to-drag ratio of 18.3%, 10.5%, 10.6% and 4% respectively when compared with the conventional wing. Simulations also showed that there is a significant improvement on pressure distribution over the lower surface of the morphing wing aerofoil. The increase in flow smoothness and reduction in vortex size reduced pressure drag along the trailing edge of the wing as a result an increase in pressure on the lower surface was experienced. A morphing wing reduced the size of the vortices and therefore the noise levels measured were reduced by up to 50%.

Keywords: shape shifting wings; morphing wing; aerodynamics enhancements; experimental aerodynamics; computational ﬂuid dynamics; NACA009

1. Introduction The aircraft industries are constantly looking for newer and more innovative ways to increase aerodynamic efficiency to reduce fuel consumption. In recent years, this has become more challenging as conventional aircraft configurations have been pushed to the very limit of aerodynamic efficiency. This implies that non-conventional technology and design needs to be developed [1] to have a substantial effect of increasing aerodynamic efficiency. In recent years “morphing” has received great interest [2–5]. Morphing implies the change of shape of aircraft structures tailored to optimize aircraft performance. This is because conventional aircraft design is a compromise to meet various flight conditions—for example a design to optimize take-off is not necessarily compatible with the optimal design for cruise mission. The definition of wing morphing is the continuous smooth and flexible change of the wing shape in order to maximize aerodynamic efficiency. A comprehensive review on morphing technologies was conducted by Rodriguez [4]. Two morphing aircraft structure concepts were developed through DARPA funding. The first concept was developed by Lockheed Martin in which the structure has the ability to reduce the wing area while transitioning from loiter to high speed Mach number [6] and the second concept—the NextGen MFX-1 aircraft—is based on flexible skin [7], which provides the capability for the wings to change shapes tailored to the specific mission profile. Wing morphing can be accomplished in multitude of directions, one direction is changing the wing spar and this has been demonstrated by the university of Maryland development of pneumatic telescopic spar for a UAV [8]. The other direction is

Aerospace 2017, 4, 52; doi:10.3390/aerospace4040052 www.mdpi.com/journal/aerospace Aerospace 2017, 4, 52 2 of 24 the change of camber of the wing, this could improve the flying performance of the aircraft including the lift and drag capabilities mainly during take-off and landing manoeuver [9–14]. In the past, a number of approaches have been proposed for sealing the gaps between high lift devices and the wing—a brief overview of these approaches is provided in order to demonstrate the complexities of the proposed designs and the challenges associated with their implementations, which highlights the urgent need for innovative solutions to address this problem. Kunz [15] patented a mechanical device consisting of additional flap connected to the main flap via a sliding interface. Diller and Miller [16] later on, patented mechanical/compliant transition to cover the span-wise and end-chord partial dislocations between the flap and wing. Their design is somehow cumbersome due to compliant Silicone elastomer skins that are reinforced by sliding mechanical rod acting to support the variable length of skin transitions. Caton et al. [17] patented a similar design of two short rods sliding into the elastomer skin as compared to the mechanical one rod mechanism of Diller and Miller. This design change meant the addition of a component to cover the joints which makes the design even more complicated. In 2013, Boeing [18] introduced the concept of multiple ribs that slide in and out of a spanwise rod embedded within the wing—the main feature of the covered elastomer skin sliding ribs is to cover the changing transition length dislocations between the flap and the wing. The patented design introduced the ribs rotating tip concept to extend the rib trailing edge length. Another concept was introduced by NASA Langley in 2012 [19], of wedge elastomer skin to bridge the span-wise gap between the flap and wing. This concept was tested [20] and noise reduction of 3 dB or higher for deflected flap was measured and aerodynamic efficiency enhancement were verified but not quantified in wind tunnel tests. Furthermore, recent advances in designing and proposing solutions based on an active camber morphing—the Fish Bone Active Camber (FishBAC) [21]—concept showed promising results in wind tunnel tests [22] of increased lift-to-drag ratio, making it a good candidate for pairing with other end transition morphing concepts. The FishBAC concept was extended [23] to introduce material structures to seal the gaps created at the end of an active control morphing surface. A Morphing Elastically LofteD (MELD) transition was introduced [23], which is capable of bending and twisting to perform smooth continuous deformation between connection ends of the active morphing camber surface and the fixed wing structure. Initially morphing structures were implemented on military aircrafts, however in recent years the focus has shifted towards small aircrafts such as UAVs [24], this is due to lower aerodynamic loads which makes the development and implementation processes more efficient and diverse. Also, the time line for new product introduction is shorter, due to less rigor certification procedures and fewer testing. A review study [24] discussed actuation systems categorized by the actuation method and implemented on airofoil wing morphing using conventional (lumped) actuators, SMA-like actuators and PZT actuators to induce camber change and thus provide lift augmentation. This paper explores the design and development of a new type of wing technology called shape shifting or morphing aircraft wings. Morphing aircraft wings are based on the dynamics of a bird wing, fundamentally ensuring that flow remains smooth and disruption is minimized. This is accomplished by eliminating the surface dislocations between the wing and the flaps, reducing and delaying the formation of vortices caused by lift-induced drag. These benefits will further increase aircraft performance by reducing take-off distances, landing distances, increasing climb rates, increasing stability and reducing the overall noise generated by the airframe. In addition, the increase in lift can also lead to a reduction in wing size, which implies a reduction in overall weight and a further reduction in fuel consumption.

2. Current Sophisticated Morphing Wing Technology Morphing wing technology had received renewed attention during recent years due to material and actuation systems rapid development. Few papers published primarily focused on mechanical structures and dielectric smart polymers. These are experimental and have a predicted timeline of 50 years until integration—currently the only company in the testing phase is Flexsys and their Aerospace 2017, 4, 52 3 of 24 compliant control surfaces are designed in conjunction with NASA and the US Air Force, however, their research is not published publicly. Currently the only competition for morphing wing technology is a company called Flexsys, they are developing a compliant control surface. Flexsys compliant control surfaces works with a combination of monolithic joint-less mechanisms and flexible material to achieve a flap which changes the camber of the wing rather than a mechanical flap. Flexsys claim their compliant control surface can achieve a 2% reduction in drag when retrofitted and up to 12% with a complete trailing edge redesign, as well as an increase in lift by 14% [25].

2.1. Mechanical Monolithic Joint-Less Mechanism Monolithic joint-less structures for morphing wings are designs [26,27] that uses materials elastic properties to create a function, such as actuation or deflection. A monolithic joint-less mechanism [3] are designed so that the structure doesn’t have any joints and loads are equally distributed throughout the entire structure [26]. This load distribution ensures that no single joint experiences excessive loading, which leads to failure. This structure allows large deflections while minimizing strain. The main advantages to this type of structure are, minimal assembly required, high reliability on repeated action, no joints, therefore no need for lubrication and no friction, can be designed using any resilient material, cyclic fatigue resistant [26]. Issues that can arise with this technology are improper design and external loads. Should the structure not be designed correctly loading may not be distributed equally across the entire structure leading to failure from excessive loading or fatigue. The structure also needs to be designed to ensure it can withstand the external loads experienced throughout flight; these can be considerably high which might lead to sudden unexpected failure [27].

2.2. Rib Based Morphing Trailing Edge The rib based morphing trailing edge design is a complex multi-component structural addition to the ribs already installed in an aircraft wing. It consists of four blocks connected to each other by a rotating joint along the camber line; each block is also attached to a link rod. The blocks are capable of rotating upwards and downward, this allows the blocks to act as a ﬂap control surface [28]. When the blocks are assembled and attached onto the wings ribs an elastomer skin is attached along the entire trailing edge, covering the blocks. This achieves a smooth surface and allows the structure to deform without causing any breaks in the surface of the wing [28]. The advantages of this design are that each rib has a degree of freedom, therefore any block that is stopped from moving causes the trailing edge to not change shape. Should the actuator move then all the blocks will move, causing a change in wing shape. In addition, this design also solves external loading issues as the block assembly is directly attached onto the wings internal main structure, which means loads, will be transferred to wing.

2.3. Dielectric Smart Polymer Morphing Wing Smart polymer, speciﬁcally dielectric smart polymer, is a material that changes shape. These are also known as electroactive smart polymers and they work by passing an electric current though them, this creates an actuation force in the form of material deformation [29]. The advantages of using dielectric smart polymers is that they are very light weight and can have relatively large strains during actuation, they also have very fast response times and large actuation forces [29]. The disadvantage of dielectric smart polymers is that they act as a capacitor when a current is run through them; therefore, they need to be properly insulated to ensure safe operation [29].

2.4. Transition Regions To ensure morphing wings aerodynamic performance, transition regions are implemented between the wing and the flap to ensure smooth flow. Transition regions designs ensure no gaps are created when the flap moves, preventing high pressure air flow towards a low-pressure region. In addition to creating a smoother flow over the wing, transition regions also reduce the size, and delay the formation, of vortices on the flaps, which reduces noise, drag and increases lift. Aerospace 2017, 4, 52 4 of 24

3. Concept Analysis Both design concepts shown in Table1, are excellent choices for a morphing wing, each offer their own advantages and disadvantages but neither of them are equal. The decision on which, is chosen is based on which has the highest score for each of the different subcategories.

Table 1. Concept design analysis.

Concept Feature Dielectric Smart Polymer Elastomer Skin Cost 2 5 Maintainability 5 5 Performance 5 3 Installation Ease 2 5 Ease of assembly 2 3 Weight 5 4 Environment Resistance 5 3 Safety 3 3 TOTAL 29 31

The dielectric smart polymer design offers a lot of benefits but its primary disadvantage is relatively more complex to assemble and install. This implies that it would require a long installation time as well as special training to install and maintain properly, this increases costs related to its operation. Another issue with the dielectric smart polymers is, related to safety concerns which require strict and proper fuel storage insulation. The main advantage of this design is the reduction of aircraft weight by eliminating the requirement for mechanical flaps and actuators. The elastomer skin design is simple but can achieve the same effects as the dielectric smart polymer design. The elastomer skin design would be slightly heavier than the smart polymer design because it would still require the mechanical flap structure. This is counteracted by the low costs, ease of maintenance, ease of installation and overall safety. Installing an elastomer design only requires the wing to be covered in an elastomer skin which is attached using an adhesive. Maintenance would be as simple as visual inspections for cracking and then complete replacement of the skin when the material has reached its fatigue limit. When considering safety, the elastomer skin design is superior because if failure occurs then there is still a working flap to generate lift. If the smart polymer fails due to damage or loss of an electric current then the aircraft won’t have any working high lift devices. Based on the above discussion elastomer skin design had been pursued.

4. Initial Experimental Analysis

4.1. Wind Tunnel Testing Wind tunnel testing was carried out on two of the three models, the conventional NACA 009 wing and the optimized morphing wing. All wings have the same span and chord as well as the same flap dimensions: Wing Span—0.3 m, Aerofoil Chord—0.15 m, Flap Span—0.172 m, Flap Chord—0.055 m, Flap Angle—25◦ (remains the same throughout testing), Wing rotated at a point of 26.7% of chord (thickest point along the chord) to accommodate the fixture bar. This was also applied to CFD simulation. AF100 low speed open loop wind tunnel (TecQuipment Ltd., www.tecquipment.com) equipped with 3 components balance connected to a separate display unit to measure the forces acting on the model was utilized. The wind tunnel test section dimensions, Length = 0.6 m, Width = 0.3 m and Height = 0.3 m. Aerofoil NACA 009 was selected primarily to reduce profile drag as much as possible, as this study is concentrating on vortices being generated by the deployment of flaps. There were also costing and simulation issues to consider, being self-funded cost had to be kept down for 3D printing Aerospace 2017, 4, 52 5 of 24

models.Aerospace The 2017 simulation, 4, 52 was carried out on Ansys academic, therefore there was a limitation5 of 24 of 512,000 cells for meshing. It was noted that cambered aerofoils required a higher cell count, which limitedAerospace512,000 mesh 2017cells refinement., 4 for, 52 meshing. It was noted that cambered aerofoils required a higher cell count, which5 of 24 limitedThe models mesh refinement. were 3D printed using the University printing facility rapid prototyping machine 512,000 cells for meshing. It was noted that cambered aerofoils required a higher cell count, which and thenThe sanded models down were to 3D remove printed the using rough the surface University created printing during facility the 3Drapid printing prototyping process. machine Once the limitedand then mesh sanded refinement. down to remove the rough surface created during the 3D printing process. Once the model is finished, a 12 mm steel bar is installed at the aerodynamic centre, 40 mm from the leading modelThe is modelsfinished, were a 12 3Dmm printed steel bar using is installed the University at the aerodynamic printing facility centre, rapid 40 prototypingmm from the machine leading edge. The finished models are shown in Figure1. Strict fabrication procedures where implemented to andedge. then The sanded finished down models to remove are shown the rough in Figure surface 1. Strict created fabrication during the procedures 3D printing where process. implemented Once the ensuremodelto ensure productions is finished, productions did a 12 meet didmm meet thesteel printthe bar print is specifications installed specifications at the shown aerodynamic shown in in Figure Figure centre, A1 A1.. 40 mm from the leading edge. The finished models are shown in Figure 1. Strict fabrication procedures where implemented to ensure productions did meet the print specifications shown in Figure A1.

(a) (b)

Figure 1. (a) The finished 3D printed conventional wing aerofoil model; (b) The finished 3D printed Figure 1. (a) The finished 3D(a printed) conventional wing aerofoil model;(b) (b) The finished 3D printed optimized morphing aerofoil model. optimizedFigure 1. morphing (a) The finished aerofoil 3D model. printed conventional wing aerofoil model; (b) The finished 3D printed optimizedEach model morphing was installed aerofoil in model. the wind tunnel test section and tested at 0, 5, 10, 15 and 20 degrees angleEach of model attack was with installed a wind speed in the of wind 25 m/s. tunnel The da testta was section collected and tested at set intervals at 0, 5, 10, every 15 and 0.5 s 20 for degrees 20 angles at of allEach attack degrees model with of was aangles wind installed of speed attack, in of the this 25 wind m/s. provides tunnel The dataa test large section was amount collected and of tested data at setatwhich 0, intervals 5, 10,can 15 then everyand be 20 analysed 0.5degrees s for 20 s at allangleto degrees ensure of attack accurate of angles with results. a of wind attack, Liftspeed thisand of providesdrag25 m/s. data The awas largeda tathen was amount used collected to of derive data at set which coefficient intervals can every thenof lift be0.5 and analyseds fordrag. 20 to s at all degrees of angles of attack, this provides a large amount of data which can then be analysed ensureFigure accurate 2 shows results. the difference Lift and dragin coefficient data was of then lift usedbetween to derivethe optimized coefficient morphing of lift and wing drag. and Figure the 2 to ensure accurate results. Lift and drag data was then used to derive coefficient of lift and drag. showsconventional the difference wing. in The coefficient optimized of morphing lift between wing the has optimized an increase morphing in coefficient wing of lift and by the 14.8%, conventional 8.4%, Figure4% and 2 2.6% shows at 0,the 5, difference10 and 15 degreesin coefficient angle of attacklift between respectively. the optimized morphing wing and the wing.conventional The optimized wing. morphing The optimized wing morphing has an increase wing has in an coefficient increase in of coefficient lift by 14.8%, of lift 8.4%,by 14.8%, 4% 8.4%, and 2.6% at 0,4% 5, 10and and 2.6% 15 at degrees 0, 5, 10 and angle 15 ofdegrees attack angle respectively. of attack respectively.

Figure 2. The coefficient of lift on the conventional wing and optimized morphing wing as angle of

attack increases. FigureFigure 2. The 2. The coefﬁcient coefficient of of lift lift on on the the conventional conventional wing and and optimized optimized morphi morphingng wing wing as angle as angle of of attack increases. 4.2.attack Wind increases. Tunnel Results Discussion 4.2. WindWind Tunnel tunnel Results tests showedDiscussion that both the initial morphing wing and the optimized wing have 4.2.significant Wind Tunnel improvements Results Discussion over conventional wing. Although further improvement from the initial morphingWind wingtunnel is testsminor showed and would that causeboth thean increaseinitial morphing in overall wing cost andand complexity the optimized for thewing design. have Wind tunnel tests showed that both the initial morphing wing and the optimized wing have significantThe wind improvements tunnel data overconfirmed conventional that the wingmorphing. Although wing designfurther has improvement the potential from to increase the initial the signiﬁcantmorphinglift generated, improvements wing by is up minor to 14.8% overand wouldat conventional 0 degrees cause anglean wing.increase of attack. Although in overall furthercost and improvement complexity for fromthe design. the initial morphingThe wing wind is tunnel minor data and confirmed would cause that the an increasemorphing in wing overall design cost has and the complexity potential to forincrease the design. the lift generated, by up to 14.8% at 0 degrees angle of attack.

Aerospace 2017, 4, 52 6 of 24 Aerospace 2017, 4, 52 6 of 24 Aerospace 2017, 4, 52 6 of 24 TheFigures wind 3–6 tunnel show data standard confirmed deviation that thefor morphingall the coefficient wing design of lift hasand the drag potential data collected to increase during the Figures 3–6 show standard deviation for all the coefficient of lift and drag data collected during liftthe generated,tests at the by wind up to tunnel, 14.8% including at 0 degrees the angle interquartile of attack. range (indicated by green and brown). The the tests at the wind tunnel, including the interquartile range (indicated by green and brown). The greenFigures and brown3–6 show boxes standard indicate deviation the difference for all the be coefficienttween the of75th lift and drag25th datapercentile collected or the during middle the green and brown boxes indicate the difference between the 75th and 25th percentile or the middle tests50% atof thethe winddata’s tunnel, standard including deviation, the interquartilethe maximum range and (indicatedminimum byis also green shown and brown). by the upper The green and 50% of the data’s standard deviation, the maximum and minimum is also shown by the upper and andlower brown whiskers boxes respectively. indicate the differenceThe experimental between data the 75th was andcompared 25th percentile to NACA or 009 the middleprofile 50%wing of with the lower whiskers respectively. The experimental data was compared to NACA 009 profile wing with data’sno flap standard and the deviation,above results the maximum (with 25 degrees and minimum flap deflection) is also shown look byreasonable the upper see and Figure lower A2 whiskers in the no flap and the above results (with 25 degrees flap deflection) look reasonable see Figure A2 in the respectively.Appendix A. The experimental data was compared to NACA 009 profile wing with no flap and the aboveAppendix results A. (with 25 degrees flap deflection) look reasonable see Figure A2 in the AppendixA.

Figure 3. Conventional wing lift coefficient wind tunnel data standard deviation. Figure 3. Conventional wing lift coefficient coefﬁcient wind tunnel data standardstandard deviation.

Figure 4. Conventional wing drag coefﬁcientcoefficient windwind tunneltunnel datadata standardstandard deviation.deviation. Figure 4. Conventional wing drag coefficient wind tunnel data standard deviation.

Aerospace 2017, 4, 52 7 of 24 Aerospace 2017, 4, 52 7 of 24 Aerospace 2017, 4, 52 7 of 24 Aerospace 2017, 4, 52 7 of 24

FigureFigure 5. 5.Optimized OptimizedOptimized wing wingwing lift liftliftcoefﬁcient coefficientcoefficient wiwi windnd tunnel tunnel data data standard standard deviation. deviation. Figure 5. Optimized wing lift coefficient wind tunnel data standard deviation.

Figure 6. Optimized wing drag coefficient wind tunnel data standard deviation. FigureFigure 6. Optimized6. Optimized wing wing drag drag coefﬁcientcoefficient wi windnd tunnel tunnel data data stan standarddard deviation. deviation. Figure 6. Optimized wing drag coefficient wind tunnel data standard deviation. Figures 7–10 illustrates wind tunnel data error bar charts showing the standard deviation of uncertaintyFiguresFigures7– 10for 7–10 illustrates each illustrates angle windof attack.wind tunnel tunnel data data error error barbar chartscharts showing the the standard standard deviation deviation of of uncertaintyuncertainty for for each each angle angle of of attack. attack.

Figure 7. ConventionalConventional wingwing liftlift coecoefficientfficient windwind tunneltunnel errorerror bars.bars. FigureFigure 7. 7.Conventional Conventionalwing wing liftlift coefﬁcientcoefficient wind wind tunnel tunnel error error bars. bars.

Aerospace 2017, 4, 52 8 of 24

Aerospace 2017, 4, 52 8 of 24 Aerospace 2017, 4, 52 8 of 24 Aerospace 2017, 4, 52 8 of 24

Figure 8. Conventional wing drag coefficient wind tunnel error bars. FigureFigureFigure 8. 8.8.Conventional ConventionalConventional wing wingwing dragdrag coe coefﬁcientcoefficientfficient windwind wind tunneltunnel tunnel errorerror error bars.bars. bars.

FigureFigure 9. 9.Morphing Morphing wing wing liftlift coefﬁcientcoefficient wind wind tunnel tunnel error error bars. bars. FigureFigure 9.9. MorphingMorphing wingwing liftlift coefficicoefficientent windwind tunneltunnel errorerror bars.bars.

Figure 10. Morphing wing drag coefficient wind tunnel error bars. FigureFigure 10.10. MorphingMorphing wingwing dragdrag coefficoefficientcient windwind tunneltunnel errorerror bars.bars. Figure 10. Morphing wing drag coefﬁcient wind tunnel error bars. 5. Ansys-Fluent Numerical and Experimental Analysis 5.5. Ansys-FluentAnsys-Fluent NumericalNumerical andand ExperimentalExperimental AnalysisAnalysis 5. Ansys-FluentAnsys-Fluent Numerical simulation and was Experimental carried out at Analysis 0, 5, 10 and 15 degrees angle of attack, at 25 m/s. Ansys-FluentAnsys-Fluent simulationsimulation waswas carriedcarried outout atat 0,0, 5,5, 1010 andand 1515 degreesdegrees angleangle ofof attack,attack, atat 2525 m/s.m/s. Ansys-Fluent simulation was carried out at 0, 5, 10 and 15 degrees angle of attack, at 25 m/s.

Aerospace 2017, 4, 52 9 of 24

Aerospace 2017, 4, 52 9 of 24 5.1. Ansys Solver Set-Up 5.1.Aerospace Ansys 2017 Solver, 4, 52 Set-Up 9 of 24 Solver—Each wing was used in Ansys fluent with pressure based, steady state simulations. Flow5.1. Solver—EachAnsys Domain—The Solver Set-Up wing flow was domainsused in Ansys geometry fluent wi isth shown pressure in based, Figure steady 11, at state the leadingsimulations. edge of the Flow Domain—The flow domains geometry is shown in Figure 11, at the leading edge of the aerofoil isSolver—Each a 1 m radius wing circle was which used in is Ansys revolved, fluent then with extrudedpressure based, 1.5 m. steady This state allows simulations. the formation of aerofoil is a 1 m radius circle which is revolved, then extruded 1.5 m. This allows the formation of vortices whileFlow stayingDomain—The within flow the domains software’s geometry cell limit. is shown The in flow Figure domains 11, at the details leading are edge highlighted of the in vortices while staying within the software’s cell limit. The flow domains details are highlighted in aerofoil is a 1 m radius circle which is revolved, then extruded 1.5 m. This allows the formation of FigureFigure 12. 12. vortices while staying within the software’s cell limit. The flow domains details are highlighted in Figure 12.

Figure 11. The flow domains geometry. Figure 11. Figure 11.The The ﬂowflow domains domains geometry. geometry.

Figure 12. Details of the flow domains sketch. Figure 12. Details of the flow domains sketch. Turbulence Models—TheFigure turbulence 12. Details models of the & ﬂow boundary domains conditions sketch. for all simulations are as follows:Turbulence Models—The turbulence models & boundary conditions for all simulations are as Turbulencefollows: Models—The turbulence models & boundary conditions for all simulations are • Energy—On. • as follows:• Viscous—K-epsilon,Energy—On. Realizable & Non-equilibrium Wall Functions. •• Viscous—K-epsilon, Realizable & Non-equilibrium Wall Functions. • Energy—On. Boundary Conditions—Turbulent Intensity 5% & Turbulent viscosity ratio 10. • Boundary Conditions—Turbulent Intensity 5% & Turbulent viscosity ratio 10. • Viscous—K-epsilon,The turbulence models Realizable for the &wing Non-equilibrium sections were selected Wall Functions. based on information collected from The turbulence models for the wing sections were selected based on information collected from • Boundary(Ansys, 2006, Conditions—Turbulent ANSYS Inc., Canonsburg, Intensity PA, USA). 5% Realizable & Turbulent k-epsilon viscosity was ratioselected 10. because of the (Ansys, 2006, ANSYS Inc., Canonsburg, PA, USA). Realizable k-epsilon was selected because of the possibility of increased accuracy on RNG k-epsilon. It was also selected because of its high accuracy possibility of increased accuracy on RNG k-epsilon. It was also selected because of its high accuracy Thefor high turbulence boundary models layer separation for the wing and large sections vortices. were selected based on information collected from (Ansys,for 2006, high boundary ANSYS Inc.,layer Canonsburg,separation and PA,large USA). vortices. Realizable k-epsilon was selected because of the possibility5.1. Conventional of increased Wing accuracy Ansys Simulation on RNG k-epsilon. It was also selected because of its high accuracy 5.1. Conventional Wing Ansys Simulation for high boundaryFigure 13 shows layer the separation pressure andcoefficients large vortices. of the flow over the conventional aerofoil, the pressure Figure 13 shows the pressure coefficients of the flow over the conventional aerofoil, the pressure of the flow around the flap, vortices and in the vortices wake is slightly lower than the rest of the 5.2. Conventionalof the flow around WingAnsys the flap, Simulation vortices and in the vortices wake is slightly lower than the rest of the

Figure 13 shows the pressure coefficients of the flow over the conventional aerofoil, the pressure of the flow around the flap, vortices and in the vortices wake is slightly lower than the rest of the Aerospace 2017, 4, 52 10 of 24

Aerospace 2017, 4, 52 10 of 24 wingsAerospace wake. 2017 The, 4, lift-induced52 drag reduced the pressure of the wing causing an increase in10 pressure of 24 drag;Aerospacewings this can 2017wake. be, 4, 52 seenThe lift-induced in Figure 13 drag as reduced there is the a significant pressure of pressure the wing differencecausing an betweenincrease in the pressure10 vorticesof 24 generatedwingsdrag; andthiswake. can the The be rest lift-inducedseen of in the Figure wings drag 13 wake.reducedas there The theis a pressu pressuresignificantre of coefficient thepressure wing causingdiffe in therence an centre betweenincrease of the thein pressure vorticesvortices is considerablywingsdrag;generated wake.this lowercan and The be the lift-induced thanseen rest restin of Figure the of drag thewings 13 reduced wake,as wake.there shownthe isThe apressu significant pressu inre Figure ofre the coefficientpressure 14 wing. This causingdiffe in creates therence ancentre increasebetween pressure of the inthe dragvorticespressure vortices on is the wing,drag;generatedconsiderably by pushingthis can and belower thethe seen flaprest than in downwardofFigure rest the of wings 13the as wake, reducing therewake. shown is The a thesignificant pressuin overallFigurere coefficientpressure14. lift This production. creates diffe in therence pressure centre In between addition, ofdrag the the on vortices as vorticesthe the wing, angle is of attackgeneratedconsiderablyby pushing is increased and the lower the flap therest than downward vortexof rest the of sizewings the reducing andwake, wake. intensity shownthe The overall pressu in were Figure liftre furtherproduction. coefficient 14. This increased. creates Inin addition,the pressure centre This as was ofthedrag the corroboratedangle on vortices the of wing,attack is by Figureconsiderablybyis 15 increasedpushing. Figure the lowerthe 15 flap ,vortex illustrates than downward sizerest andof vortices the reducing intensity wake, generated shown thewere overall furt in forFigureher li conventionalft increased. production. 14. This Thiscreates In wing wasaddition, pressure withcorroborated as an dragthe angle angle on by ofthe Figure of attack wing, attack 15. of 0 is increased the vortex size and intensity were further increased. This was corroborated by Figure 15. andby 5Figure degrees pushing 15, respectively, theillustrates flap downward vortices showing reducinggenerated a larger the for vortex overall conventional beinglift production. generated wing with In by addition, an the angle flap as of the theattack angle wing. of of 0attack Thisand had5 isFiguredegrees increased 15, respectively, illustratesthe vortex vortices sizeshowing and generatedintensity a larger were vortexfor conventionalfurt beingher increased. generated wing This with by was the an corroboratedflapangle of of the attack wing. by Figureof This0 and 15.had 5 demonstrated that Ansys is simulating correctly the flow field and therefore ensuring accurate results Figuredegreesdemonstrated 15, respectively, illustrates that Ansys vortices showing is simulating generated a larger correctly forvortex conventional being the flow generated fieldwing and with by therefore thean angleflap ensuring of of the attack wing. accurate of This0 and results had 5 are being produced for further simulation. degreesdemonstratedare being respectively, produced that Ansys forshowing further is simulating a simulation. larger correctlyvortex being the flow generated field and by therefore the flap ensuring of the wing. accurate This results had demonstratedare being produced that Ansys for further is simulating simulation. correctly the flow field and therefore ensuring accurate results are being produced for further simulation. (-) p C (-) p C (-) p C

(a) (b) (b) (a) (a) (b) (-) p C (-) p C (-) p C

(c) (c) Figure 13. The pressure coefficients of the flow over a conventional wing aerofoil, vortices indicated Figure 13. The pressure coefﬁcients of the ﬂow over(c) a conventional wing aerofoil, vortices indicated in Figurein red circle,13. The 0-degree pressure angle coefficients of attack of ( athe) Side flow view; over ( ba )conventional Top view; (c )wing Rear aerofoil,View. vortices indicated red circle, 0-degree angle of attack (a) Side view; (b) Top view; (c) Rear View. Figurein red 13.circle, The 0-degree pressure angle coefficients of attack of (thea) Side flow view; over (ab conventional) Top view; (c wing) Rear aerofoil, View. vortices indicated in red circle, 0-degree angle of attack (a) Side view; (b) Top view; (c) Rear View. (-) p C (-) p C (-) p C

Figure 14. The low-pressure coefficient at the centre of the vortices at 0-degree angle of attack. Figure 14. The low-pressure coefficient at the centre of the vortices at 0-degree angle of attack. Figure 14. The low-pressure coefﬁcient at the centre of the vortices at 0-degree angle of attack. Figure 14. The low-pressure coefficient at the centre of the vortices at 0-degree angle of attack.

(a) (b) (a) (b) Figure 15. Images showing the vortex size for the conventional wing (a) at 0 degrees’ angle of attack (a) (b) Figure(b) at 5 15. degrees’ Images angle showing of attack. the vortex size for the conventional wing (a) at 0 degrees’ angle of attack Figure(b) at 515. degrees’ Images angle showing of attack. the vortex size for the conventional wing (a) at 0 degrees’ angle of attack Figure 15. Images showing the vortex size for the conventional wing (a) at 0 degrees’ angle of attack (b) at 5 degrees’ angle of attack. (b) at 5 degrees’ angle of attack.

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5.3.Aerospace5.2. Initial Initial Morphing2017 Morphing, 4, 52 Wing Wing Ansys Ansys Simulation Simulation 11 of 24 FigureFigure 16 shows16 shows the coefficientthe coefficient of lift of data lift generateddata generated by Ansys, by Ansys, a significant a significant increase increase in coefficient in 5.2. Initial Morphing Wing Ansys Simulation of liftcoefficient when comparing of lift when the comparing initial morphing the initial wingmorphing and conventionalwing and conventional wing. wing. Figure 16 shows the coefficient of lift data generated by Ansys, a significant increase in coefficient of lift when comparing the initial morphing wing and conventional wing.

Figure 16. A graph showing the coefficient of lift results generated from Ansys for the conventional Figure 16. A graph showing the coefficient of lift results generated from Ansys for the conventional wing and initial morphing wing. wing and initial morphing wing. Figure 16. A graph showing the coefficient of lift results generated from Ansys for the conventional wingThe simulation and initial morphing data shows wing. that at 0, 5, 10 and 15 degrees’ angle of attack, there is an increase in liftThe by simulation10.8%, 4.3%, data 3.4% shows and 1% that respectively. at 0, 5, 10 Figure and 15 17 degrees’ is a Cd versus angle Cl of graph. attack, The there data is shows an increase that in lift bythe 10.8%, initialThe simulation morphing 4.3%, 3.4% datawing and shows is 1% more respectively. that aerodynamically at 0, 5, 10 Figure and 15 efficient 17 degrees’ is a Cdthan angle versus the of conventional attack, Cl graph. there Thewing. is an data increaseThe shows initial in that theliftmorphing initial by 10.8%, morphing wing 4.3%, at wing0,3.4% 5, 10 and is and more 1% 15 respectively. aerodynamicallydegrees angle Figure of attack efficient17 is aproduces Cd than versus the10.8%, Cl conventional graph. 4.3%, The 3.4% data wing.and shows 1% The more that initial morphingthelift respectively,initial wing morphing at 0,as 5, well wing 10 andas is a 15moremaximum degrees aerodynamically of angle 2.5% of reduction attack efficient produces in than drag. the 10.8%,Figure conventional 4.3%,18 shows 3.4% wing. the and lift-to-drag The 1% initial more lift morphing wing at 0, 5, 10 and 15 degrees angle of attack produces 10.8%, 4.3%, 3.4% and 1% more respectively,ratio, or glide as well ratio, as this a maximumgraph shows of a 2.5% wings reduction ability to glide in drag. during Figure flight. 18 A shows higher theglide lift-to-drag ratio means ratio, lift respectively, as well as a maximum of 2.5% reduction in drag. Figure 18 shows the lift-to-drag or glidethe aircraft ratio, thiswill be graph able showsto travel a further wings abilitydistance to without glide during losing as flight. much A altitude. higher glide ratio means the ratio,The or glide initial ratio, morphing this graph wing shows has aa wingssignificantly ability tohigher glide glideduring ratio flight. than A higherthe conventional glide ratio meanswing, aircraft will be able to travel further distance without losing as much altitude. especiallythe aircraft at will low be angles able to of travel attack. further This distanceimplies less without engine losing power as muchwill be altitude. required during flight to The initial morphing wing has a significantly higher glide ratio than the conventional wing, maintainThe initialaltitude morphing and therefore wing less has fuel a significantly consumption. higher At 0, glide 5, 10 ratioand 15than degrees the conventional angle of attack wing, the especiallyespeciallyinitial morphing at lowat low angles wing angles has of ofattack. an attack. increase ThisThis in impliesimplies glide ratio less less of engine engine 13.6%, power 5.3%, power 5.8%will will beand berequired 3.1% required respectively. during during flight flight to to maintainmaintain altitude altitude and and therefore therefore less less fuelfuel consumption.consumption. At At 0, 0, 5, 5, 10 10 and and 1515 degrees degrees angle angle of attack of attack the the initialinitial morphing morphing wing wing has has an an increase increase in inglide glide ratioratio of of 13.6%, 13.6%, 5.3%, 5.3%, 5.8% 5.8% and and 3.1% 3.1% respectively. respectively.

Figure 17. A graph showing Coefficient of drag versus coefficient of lift for the initial morphing and conventional wing. FigureFigure 17. 17.A graphA graph showing showing Coefﬁcient Coefficient ofof dragdrag versus co coefﬁcientefficient ofof lift lift for for the the initial initial morphing morphing and and conventionalconventional wing. wing.

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Aerospace 2017, 4, 52 12 of 24

Figure 18. A graph showing Lift/drag otherwise known as the glide ratio for the conventional wing and Figure 18. A graph showing Lift/drag otherwise known as the glide ratio for the conventional wing initial morphing wing. andFigure initial 18. morphingA graph showing wing. Lift/drag otherwise known as the glide ratio for the conventional wing and initial morphing wing. Figure 19 shows the pressure coefficients over the lower side of the conventional wing and initial morphingFigureFigure 19 wingshows 19 shows respectively. the pressurethe pressure The coefficients coefficientsconventional over over wings the the lower lowerpressure side side ofatof thethe conventional conventionalflap is slightly wing wing lower and and initialand initial morphingbecamemorphing wing suddenly wing respectively. lowerrespectively. at Thethe flap conventionalThe dislocation conventional wingsfrom wings the pressure wing pressure surface. at the at flap Thisthe is isflap slightly because is slightly lowerof lift-induced lower and became and suddenlydrag,became which lower suddenly results at the lowerin flap sudden dislocationat the pressuflap dislocationre from drop the and from wing therefore the surface. wing increases surface. This is Thisin because pressure is because of drag lift-induced of aslift-induced well as drag, whichreductiondrag, results which in lift sudden results production. pressurein sudden drop pressu andre therefore drop and increases therefore in increases pressure in drag pressure as well drag as reductionas well as in lift production.reductionThe initial in lift morphing production. wing in Figure 19 doesn’t suffer from this, because there is no gap between theThe flap initialThe and initial morphingthe wingmorphing surface, wing wing in the Figurein pressure Figure 19 19 doesn’t distri doesn’tbution suffer suffer is from enhancedfrom this,this, along because the there therelower is is nosurface no gap gap between of between the wing. The pressure on the lower surface of the initial morphing wing is also slightly higher as the flapthe and flap the and wing the wing surface, surface, the pressure the pressure distribution distribution is enhanced is enhanced along along the lowerthe lower surface surface of theof the wing. compared to the conventional wing. The elimination of the gaps between flap and wing surface The pressurewing. The on pressure the lower on surface the lower of the surface initial of morphing the initial wingmorphing is also wing slightly is also higher slightly as compared higher as to caused the pressure along the bottom of the morphing wing to increase and pressure distribution the conventionalcompared to wing. the conventional The elimination wing. of The the gapseliminat betweenion of flapthe gaps and wingbetween surface flap causedand wing the surface pressure enhancementcaused the pressure was experienced along the overbottom the of wings the morphing surface. wingTherefore, to increase aerodynamic and pressure performance distribution as along the bottom of the morphing wing to increase and pressure distribution enhancement was comparedenhancement to conventional was experienced wing was over enhanced the wings due surf to liftace. increase Therefore, and aaerodynamic reduction in performancepressure drag as experienced over the wings surface. Therefore, aerodynamic performance as compared to conventional alongcompared the wing. to conventional wing was enhanced due to lift increase and a reduction in pressure drag wing was enhanced due to lift increase and a reduction in pressure drag along the wing. along the wing. (-) p C (-) p C

(a) (b)

Figure 19. Cont.

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Aerospace 2017, 4, 52 13 of 24

Figure 19. The pressure coefficient (a) Conventional wing upper(d )surface; (b) Initial morphing Wing Figure 19. The pressure coefficient (ac) Conventional wing upper surface; (b) Initial morphing Wing Upper Surface; (c) Conventional wing lower surface; (d) Initial morphing Wing lower Surface. UpperFigure Surface; 19. The (c) pressure Conventional coefficient wing (a lower) Conventional surface; (wingd) Initial upper morphing surface; (b Wing) Initial lower morphing Surface. Wing Upper Surface; (c) Conventional wing lower surface; (d) Initial morphing Wing lower Surface. 5.3. Optimized Morphing Wing Simulation 5.4. Optimized Morphing Wing Simulation 5.3. OptimizedFigure 20 Morphingshows the Wing coefficients Simulation of lift for all the wings, each design has a significant increase in coefficientFigure 20 ofshows lift, the the highest coefficients being the of liftoptimized for all themorphing wings, wing. each designOverall hasat 0, a 5, significant 10 and 15 degrees increase in Figure 20 shows the coefficients of lift for all the wings, each design has a significant increase in coefficientangle of of attack lift, the the highest optimized being morphing the optimized wing have morphing an increase wing. in lift Overall by 15.4%, at 0, 7.5%, 5, 10 5% and and 15 1.8% degrees coefficient of lift, the highest being the optimized morphing wing. Overall at 0, 5, 10 and 15 degrees respectively when compared to the conventional wing. At 0, 5, 10 and 15 degrees the optimized wing angleangle of attack of attack the the optimized optimized morphing morphing wingwing havehave an increase increase in in lift lift by by 15.4%, 15.4%, 7.5%, 7.5%, 5% 5%and and 1.8% 1.8% has an increase of 4.6%, 3.2%, 1.7% and 0.9% respectively, when compared to the initial shape shifting respectivelyrespectively when when compared compared to to thethe conventional conventional wing wing.. At 0, At 5, 0,10 5,and 10 15 and degrees 15 degrees the optimized the optimized wing wing. winghas has an an increase increase of 4.6%, of 4.6%, 3.2%, 3.2%, 1.7% and 1.7% 0.9% and respectively, 0.9% respectively, when compared when compared to the initial to shape the initial shifting shape shiftingwing. wing.

Figure 20. A graph showing coefficient of lift for the conventional wing, initial morphing wing and optimized morphing wing. FigureFigure 20. A20. graph A graph showing showing coefficient coefficient of oflift lift forfor thethe conventionalconventional wing, wing, initial initial morphing morphing wing wing and and optimized morphing wing. optimizedFigure morphing21 shows wing.an efficiency graph for all three wing designs, using coefficient of drag over coefficient of lift. Lift production for each of the morphing wings was enhanced compared with Figure 21 shows an efficiency graph for all three wing designs, using coefficient of drag over conventional wing, while producing less drag. Conventional wing to the optimized morphing wing coefficientFigure 21 ofshows lift. Lift an production efficiency graphfor each for of all the three morphing wing designs,wings was using enhanced coefficient compared of drag with over comparison revealed up to 5% drag reduction was experienced by the optimized morphing wing coefficientconventional of lift. wing, Lift while production producing for less each drag. of theConv morphingentional wing wings to the was optimized enhanced morphing compared wing with dependent on angle of attack. The optimized morphing wing to the initial shape shifting wing conventionalcomparison wing, revealed while up producingto 5% drag reduction less drag. was Conventional experienced by wing the optimized to the optimized morphing morphing wing comparison showed drag reduction by up to 2.7%. Figure 22 illustrates lift-to-drag ratio or glide ratio wingdependent comparison on angle revealed of attack. up to 5%The dragoptimized reduction morp washing experiencedwing to the initial by the shape optimized shifting morphing wing highest magnitude increase is evident for the optimized morphing wing. At angles of attack of 0, 5, comparison showed drag reduction by up to 2.7%. Figure 22 illustrates lift-to-drag ratio or glide ratio wing10 dependent and 15 degrees on the angle optimized of attack. wing The has an optimized increase in morphing lift-to-drag wing ratio of to 18.3%, the initial 10.5%, shape 10.6% shiftingand highest magnitude increase is evident for the optimized morphing wing. At angles of attack of 0, 5, wing4% comparison respectively, showed compared drag with reduction the conventional by up towing. 2.7%. While Figure the optimized22 illustrates wing lift-to-drag compared with ratio or 10 and 15 degrees the optimized wing has an increase in lift-to-drag ratio of 18.3%, 10.5%, 10.6% and glidethe ratio initial highest morphing magnitude wing has increase an increase is evident of 4.7% for, 5.3%, the 4.8% optimized and 0.9% morphing at 0, 5, 10 wing. and 15 At degrees angles of 4% respectively, compared with the conventional wing. While the optimized wing compared with attackrespectively. of 0, 5, 10 and 15 degrees the optimized wing has an increase in lift-to-drag ratio of 18.3%, the initial morphing wing has an increase of 4.7%, 5.3%, 4.8% and 0.9% at 0, 5, 10 and 15 degrees 10.5%, 10.6%Figure and 23 4%shows respectively, the evolution compared of pressure with distri thebution conventional along the wing. lower While surface the of optimized each of the wing respectively. wing configurations. Each wing has a different maximum pressure coefficient on its lower surface, comparedFigure with 23 the shows initial the morphing evolution wing of pressure has an increasedistribution of 4.7%,along 5.3%,the lower 4.8% surface and 0.9% of each at 0, of 5, the 10 and 0.75 for the conventional wing, 0.904 for the initial morphing wing and 0.906 for the optimized 15 degreeswing configurations. respectively. Each wing has a different maximum pressure coefficient on its lower surface, Figure 23 shows the evolution of pressure distribution along the lower surface of each of the wing 0.75 for the conventional wing, 0.904 for the initial morphing wing and 0.906 for the optimized configurations. Each wing has a different maximum pressure coefficient on its lower surface, 0.75 for

Aerospace 2017, 4, 52 14 of 24

Aerospace 2017, 4, 52 14 of 24 the conventionalAerospace 2017, 4, 52 wing, 0.904 for the initial morphing wing and 0.906 for the optimized morphing14 of 24 wing. Thismorphing showsAerospace that 2017 wing., the 4, 52 This optimized shows wingthat the experienced optimized wing a much experienced higher pressure a much alonghigherits pressure lower surfacealong14 itsof than24 the conventionalmorphinglower surface wing. than wing This the and shows conventional the that initial the morphingwing optimized and the wing.wing initial experienced morphing awing. much higher pressure along its lowermorphing surface wing. than Thisthe conventional shows that the wing optimized and the winginitial experienced morphing wing. a much higher pressure along its lower surface than the conventional wing and the initial morphing wing.

Figure 21. A graph showing the wing efficiency using coefficient of drag over coefficient of lift for Figure 21. A graph showing the wing efficiency using coefficient of drag over coefficient of lift for each Figureeach of 21.the A wings. graph showing the wing efficiency using coefficient of drag over coefficient of lift for of theeachFigure wings. of the 21. wings. A graph showing the wing efficiency using coefficient of drag over coefficient of lift for each of the wings.

Figure 22. A graph showing lift-to-drag ratios or glide ratios for all three wings. Figure 22. A graph showing lift-to-drag ratios or glide ratios for all three wings. Figure 22. A graph showing lift-to-drag ratios or glide ratios for all three wings. Figure 22. A graph showing lift-to-drag ratios or glide ratios for all three wings. (-) p (-) C p C (-) p C

(a) (b) (b) (a) (a) (b)

Figure 23. Cont.

Aerospace 2017, 4, 52 15 of 24

(e) (f)

Figure 23. The pressure coefficient distribution (a) along the upper surface of the conventional wing; Figure 23. The pressure coefficient distribution (a) along the upper surface of the conventional wing; (b) along the upper surface of the initial morphing wing; (c) along the upper surface of optimized (b) along the upper surface of the initial morphing wing; (c) along the upper surface of optimized morphing wing; (d) along the lower surface of the conventional wing; (e) along the lower surface of morphing wing; (d) along the lower surface of the conventional wing; (e) along the lower surface of the initial morphing wing; (f) along the lower surface of optimized morphing wing. the initial morphing wing; (f) along the lower surface of optimized morphing wing. Comparison pressure evaluation of the lower surface of the optimized wing with the initial and conventionalComparison wing pressure indicated evaluation significant of differences the lower in surface the overall of pressure the optimized distribution. wing At with the trailing the initial andedge conventional of the wing wing (left indicated side), low significant pressure areas differences are significantly in the overall reduced, pressure as well distribution. as much higher At the trailingpressure edge at of the the leading wing (leftedge side), (right lowside). pressure The increase areas in are pressure significantly created reduced, higher lift as which well aswas much higherexperienced pressure by at thethe leadingoptimized edge wing, (right as well side). as Thethe increase increase in in flow pressure smoothness created due higher to the lift curvature which was of the transition regions. Furthermore, pressure drag is significantly reduced on the optimized wing experienced by the optimized wing, as well as the increase in flow smoothness due to the curvature because of regions with extremely low pressure on the conventional wing are almost eliminated. of the transition regions. Furthermore, pressure drag is significantly reduced on the optimized wing Upper surfaces comparison of the conventional, initial morphing and optimized morphing becausewings of are regions illustrated with in extremely Figure 24. lowThe pressure along on the the conventional trailing edge of wing the wing are almost (left side) eliminated. increased significantly,Upper surfaces due comparisonto vortex size of thereduction conventional, as well as initial a delay morphing in the andvortex optimized formation. morphing This effect wings are illustratedreduced the inpressure Figure drag24. Theand caused pressure significant along the changes trailing to the edge pressure of the in wing the wake (left side)of the increasedwing. significantly,The conventional due to wing vortex (top size left, reduction Figure 24a) as is wellcompared as a delayto a morphing in the vortex wing (top formation. right and Thisbottom effect reducedmiddle, the Figure pressure 24b,c), drag the and morphing caused wings significant experien changesced less to chaotic the pressure behaviour in theand wakethe pressure of the is wing. Themore conventional equalized wing resulting (top in left, less Figure drag. 24Onea) isof comparedthe major benefits to a morphing of the optimized wing (top wing right is andthat bottomthe middle,increased Figure curvatures 24b,c), the on morphing the flap transition wings experienced regions allowed less chaotic smoother behaviour flow to occur, and the which pressure reduced is more equalizedthe pressure resulting drag, in vortex less drag. size and One caused of the a major delayed benefits formation of the of optimizedthe vortices wing on the is wing. that theFigure increased 25 curvaturesshows the on pressure the flap transitioncoefficients regions above and allowed below smootherthe surface flow of the to wings, occur, the which red circles reduced indicate the pressure the pressure of the vortex which is being generated by the flaps. The conventional wings vortex (top left, drag, vortex size and caused a delayed formation of the vortices on the wing. Figure 25 shows the Figure 25a) is significantly lower in pressure and much larger in size than the morphing wings (top pressure coefficients above and below the surface of the wings, the red circles indicate the pressure of the vortex which is being generated by the flaps. The conventional wings vortex (top left, Figure 25a) is significantly lower in pressure and much larger in size than the morphing wings (top right and

Aerospace 2017, 4, 52 16 of 24

Aerospace 2017, 4, 52 16 of 24 bottom centre, Figure 25b,c), the optimized shape shifting wing has the highest-pressure vortex and Aerospace 2017, 4, 52 16 of 24 thereforeright an and overall bottom higher centre, wing Figure wake 25b,c), pressure the optimize and reducedd shape pressureshifting wing drag. has the highest-pressure Figurevortexright and and26 showsbottom therefore thecentre, an pressure overall Figure distributionhigher 25b,c), wing the optimizewake along pressure thed shape upper and shifting and reduced lower wing pressure surface has the drag. of highest-pressure each wing at the locationvortex whereFigure and vorticestherefore26 shows are anthe overall generatedpressure higher distribution on wing the flap.wake along The pressure the conventional upper and and reduced lower wing pressure surface experienced ofdrag. each wing a significant at the droplocation in pressureFigure where 26 on shows vortices both the the are pressure topgenerated and distribution bottom on the surfaceflap along. The the of conventional upper the wing and lower as wing flow surface experienced approached of each a wingsignificant the at trailing the edgedrop (chordlocation in pressure length where =vortices on 0.04 both to are the 0.11 generatedtop m). and This bottom on is the due surfac flap to. the eThe of largethe conventional wing vortex as flow that wing approached is generatedexperienced the at trailinga thissignificant location, edge it resulted(chorddrop in the length pressure pressure = 0.04 on to both to drop 0.11 the which m).top andThis consequently bottom is due surfacto the increased elarge of the vortex wing pressure asthat flow is drag. generatedapproached Comparison at the this trailing location, between edge it the conventionalresulted(chord length the wing pressure = to 0.04 the toto initial 0.11drop m). andwhich This optimized consequently is due to morphing the increased large wing,vortex pressure revealedthat drag.is generated a Comparison major differenceat this between location, between the it conventional wing to the initial and optimized morphing wing, revealed a major difference between the upperresulted and the lower pressure surface to drop pressures. which consequently With the morphing increased wings, pressure as the drag. flow Comparison approached between the trailing the the upper and lower surface pressures. With the morphing wings, as the flow approached the trailing it increasedconventional instead, wing this to is the because initial and of the optimized reduction mo inrphing vortex wing, size revealed and delay a major in formation. difference between itthe increased upper and instead, lower thissurface is be pressures.cause of the With reduction the morp inhing vortex wings, size andas the delay flow in approached formation. the trailing it increased instead, this is because of the reduction in vortex size and delay in formation. (-) p C (-) p C

(a) (b)

(a) (b) (-) p C (-) p C

(c)

Figure 24. Images showing pressure coefficient for(c the) flow over (a) the conventional wing; (b) Initial Figure 24. Images showing pressure coefﬁcient for the ﬂow over (a) the conventional wing; (b) Initial morphingFigure 24. wing;Images (c showing) optimized pressure morphing coefficient wing allfor at the 5 degrees’flow over angle (a) the of conventionalattack. wing; (b) Initial morphing wing; (c) optimized morphing wing all at 5 degrees’ angle of attack. morphing wing; (c) optimized morphing wing all at 5 degrees’ angle of attack. (-) p C (-) p C

(b) (a) (a) (b)

Figure 25. Cont.

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Aerospace 2017, 4, 52 17 of 24 (-) p C (-) p C

(c)

Figure 25. Images showing pressure cut plots at the location where vortices are generated on each Figure 25. Images showing pressure cut plots at the location(c) where vortices are generated on each wings wings flap at 5 degrees’ angle of attack (a) the conventional wing; (b) Initial morphing wing; (c) ﬂap at 5 degrees’ angle of attack (a) the conventional wing; (b) Initial morphing wing; (c) optimized Figureoptimized 25. Images morphing showing wing. pressure cut plots at the location where vortices are generated on each morphingwings wing.flap at 5 degrees’ angle of attack (a) the conventional wing; (b) Initial morphing wing; (c) optimizedFrom Figure morphing 26 there wing. is a clear evidence of differences in pressure behaviour, the conventional Fromwing is Figure much 26more there chaotic is a than clear the evidence morphing of wi differencesngs which infurther pressure support behaviour, the main theexplanation conventional for From Figure 26 there is a clear evidence of differences in pressure behaviour, the conventional wingthe is much efficiency more increase. chaotic than the morphing wings which further support the main explanation for wing is much more chaotic than the morphing wings which further support the main explanation for the efﬁciency increase. the efficiency increase.

Figure 26. A graph showing the pressure distribution along each wing at the location where vortices are generated on the flap at 5 degrees’ angle of attack. Figure 26. A graph showing the pressure distribution along each wing at the location where vortices Figure 26. A graph showing the pressure distribution along each wing at the location where vortices 5.4. areWind generated Tunnel onTesting the flap and at Si5 mulationdegrees’ angle Comparison of attack. (Validation) are generated on the flap at 5 degrees’ angle of attack. 5.4. WindFigure Tunnel 27, illustratesTesting and the Si mulationcoefficient Comparison of lift versus (Validation) angle of attack for wind tunnel and simulation 5.5. Winddata. TunnelIt is observed Testing andthat Simulationthe results Comparisonare very similar. (Validation) For the conventional wing at 0, 5, 10 and 15 degreesFigure Angle 27, illustrates of Attack the(AOA), coefficient the percentage of lift versus differences angle of between attack for the wind simulation tunnel andand windsimulation tunnel data.Figuretests Itare is27 1.87%,observed, illustrates 0.04%, that the 1.8% the coefficient resultsand 0.86% are of respectivelyvery lift versussimilar. for angleFor coefficient the of conventional attack of forlift. windTherefore, wing tunnel at 0,results 5, and 10 simulationobtainedand 15 data.degreesfrom It is observed Ansys Angle simulation of that Attack the results(AOA),are accurate arethe verypercentage within similar. 1.87% di Forfferences of thethe conventionalwind between tunnel the tests. simulation wing For at 0,the 5,and optimized 10 wind and 15tunnel degreeswing Angleteststhe of percentageare Attack 1.87%, (AOA), 0.04%,difference the 1.8% percentage between and 0.86% the differences respectivelysimulation betweenand for windcoefficient thetunnel simulation of tests lift. isTherefore, 1.3%, and 0.8%, wind results 0.8% tunnel obtained and tests 0.8% are 1.87%,fromrespectively. 0.04%, Ansys 1.8% simulation This and implies 0.86% are therespectivelyaccurate optimized within for wing 1.87% coefficient is accurate of the of wind lift.within Therefore,tunnel 1.3%. tests. results For the obtained optimized from wing Ansys simulationthe percentage are accurate difference within between 1.87% the of thesimulation wind tunnel and wind tests. tunnel For thetests optimized is 1.3%, 0.8%, wing 0.8% the and percentage 0.8% differencerespectively. between This the implies simulation the optimized and wind wing tunnel is accurate tests within is 1.3%, 1.3%. 0.8%, 0.8% and 0.8% respectively. This implies the optimized wing is accurate within 1.3%.

Aerospace 2017, 4, 52 18 of 24 Aerospace 2017, 4, 52 18 of 24

Aerospace 2017, 4, 52 18 of 24

FigureFigure 27. 27.A graphA graph showing showing coefficients coefficients ofof lift for Ansys Ansys simulation simulation and and wind wind tunnel tunnel testing. testing. FigureFigure 28 27.shows A graph the coefficientshowing coefficients of drag dataof lift foforr Ansysboth simulation simulation andand wind experimental tunnel testing. data for the conventionalFigure 28 shows wing theand coefficientoptimized ofmorphing drag data wing. for At both 0, 5, simulation 10 and 15 anddegrees experimental angle of attack data the for the Figure 28 shows the coefficient of drag data for both simulation and experimental data for the conventionalpercentage wingdifference and optimizedbetween simulation morphing and wing. measur Ated 0, data 5, 10 for and the 15conventional degrees angle wing ofis attack1.04%, the conventional wing and optimized morphing wing. At 0, 5, 10 and 15 degrees angle of attack the percentage6.75%, 1.43% difference and 4.19% between respectively. simulation For andthe optimised measured wing data the for percentage the conventional difference wing is 0.49%, is 1.04%, percentage difference between simulation and measured data for the conventional wing is 1.04%, 6.75%,3.91%, 1.43% 1.24% and and 4.19% 2.08% respectively. respectively. For Based the optimisedon this, the wingdrag data the percentage is accurate within difference 6.75%. is 0.49%,Based on 3.91%, 6.75%, 1.43% and 4.19% respectively. For the optimised wing the percentage difference is 0.49%, the data shown in Figures 18 and 19 it can be determined that the data produced from Ansys is 1.24%3.91%, and 1.24% 2.08% and respectively. 2.08% respectively. Based on Based this, theon this, drag the data drag is accuratedata is accurate within within 6.75%. 6.75%. Based Based on the on data accurate to real world conditions by 1.87% for coefficient of lift and 6.75% for coefficient of drag. shownthe indata Figures shown 18 in and Figures 19 it can18 and be determined19 it can be thatdetermined the data that produced the data from produced Ansys from is accurate Ansys tois real worldaccurate conditions to real by world 1.87% conditions for coefficient by 1.87% of for lift coef andficient 6.75% of for lift coefficient and 6.75% for of drag.coefficient of drag.

Figure 28. A graph showing coefficient of drag for wind tunnel testing and Ansys Simulation.

5.5. AnsysFigureFigure and 28. Wind28.A graphA graphTunnel showing showing Results coefﬁcient coefficient Discussion ofof dragdrag for wind wind tunnel tunnel testing testing and and Ansys Ansys Simulation. Simulation.

5.5. AnsysThe simulation and Wind Tunnel results Results showed Discussion the morphing wings have a significant increase in aerodynamic 5.6.efficiency Ansys and compared Wind Tunnel to the Results conventional Discussion wing. The most efficient wing design was the optimized The simulation results showed the morphing wings have a significant increase in aerodynamic morphing wing, which because of increased flow smoothness allowed for higher pressure efficiencyThe simulation compared results to the showed conventional the morphing wing. Th wingse most haveefficient a significant wing design increase was the in optimized aerodynamic distribution over a larger area of the lower wing surface. This resulted in higher lift production and efficiencymorphing compared wing, which to the conventionalbecause of increased wing. Theflow most smoothness efficient wingallowed design for higher was the pressure optimized reduced pressure drag. morphingdistribution wing, over which a larger because area ofof increasedthe lower flowwing smoothnesssurface. This allowedresulted forin higher higher lift pressure production distribution and The initial and optimized morphing wing also had the effect of reducing the size and delaying reduced pressure drag. overthe a largerformation area of of vortices the lower in the wing wake surface. of the Thiswing, resulted which reduced in higher pressure lift production drag. This andoccurred reduced The initial and optimized morphing wing also had the effect of reducing the size and delaying pressurebecause drag. the vortices are no longer generated on the flap and the pressure drop at the points where the formation of vortices in the wake of the wing, which reduced pressure drag. This occurred theThe flaps initial would and usually optimized dislocate morphing from the wing wing also surfac hade theis no effect longer of reducingavailable; thethis sizecan be and seen delaying in because the vortices are no longer generated on the flap and the pressure drop at the points where the formationFigure 24. of vortices in the wake of the wing, which reduced pressure drag. This occurred because the flaps would usually dislocate from the wing surface is no longer available; this can be seen in the vorticesThe arepressure no longer data provided generated in onFigures the flap 25 and and 26 the reinforced pressure the drop success at the of pointsthe morphing where wing the flaps Figure 24. woulddesign. usually They dislocate showed no from significan the wingt pressure surface drop is no along longer the available;trailing edge this of canthe morphing be seen in wings Figure and 24 . The pressure data provided in Figures 25 and 26 reinforced the success of the morphing wing therefore,The pressure leading data to a provided reduction inin Figurespressure 25drag and as 26well reinforced as a reduction the successin vortex of size. the morphing wing design. They showed no significant pressure drop along the trailing edge of the morphing wings and design.therefore, They leading showed to no a reduction significant inpressure pressure dropdrag as along well theas atrailing reduction edge in vortex of the size. morphing wings and therefore, leading to a reduction in pressure drag as well as a reduction in vortex size.

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6. Noise One source of aircraft noise is generated by turbulent flow over the airframe. A considerable noise signatureAerospace is 2017 generated, 4, 52 when an aircraft flaps are deployed, this is due to violent disruption of19 of airflow 24 and the creation of turbulence. Therefore, if lift-induced drag vortices are reduced by the morphing 6. Noise wing then noise is reduced. To test this, noise level readings were taken during wind tunnel tests to determineOne any source noise of level aircraft reductions. noise is generated by turbulent flow over the airframe. A considerable noiseEach signature wing has is a generated noise level when reading an aircraft taken flaps at 0, 5,are 10, deployed, 15 and 20this degrees is due to angle violent of attack.disruption Based of on the measuredairflow and data the showncreation inof Tableturbulence.2 and EquationTherefore, (1)if lift-induced after performing drag vortices baseline are wind reduced tunnel by the noise morphing wing then noise is reduced. To test this, noise level readings were taken during wind level correction, the sound pressure level produced by the wing only was obtained. The actual noise tunnel tests to determine any noise level reductions. levels of the wing as a function of AOA are shown in Figure 29. Each wing has a noise level reading taken at 0, 5, 10, 15 and 20 degrees angle of attack. Based on the measured data shown in Table 2 and Equation (1) after performing baseline wind tunnel noise P1 level correction, the sound pressureNoise level level produced(dB) = by20 the log wing10 only, was obtained. The actual noise (1) P0 levels of the wing as a function of AOA are shown in Figure 29. where: P1 = Sound pressure Pa (What the microphone detects), P0= Reference Pressure (Threshold of −5 Noise level dB =20log , (1) hearing 2 × 10 Pa).

where: P1 = Sound pressure Pa (What the microphone detects), P0 = Reference Pressure (Threshold of Table 2. Noise levels produced by the conventional wing and morphing wing at different angles hearing 2 × 10−5 Pa). of attack. Table 2. Noise levels produced by the conventional wing and morphing wing at different angles of attack. Angle of Attack Conventional Wing Shape Shifting Wind Tunnel Noise Angle of (AOA)Attack ConventionalNoise Wing Level ShapeWing Shifting Noise Wing Level WindLevel Tunnel (All AOA) Noise (AOA) Noise Level Noise Level Level (All AOA) 0 68 dB 63 dB 0 5 68 dB 68 dB 63 65dB dB 5 10 68 dB 69 dB 65 67dB dB 62 dB 10 15 69 dB 70 dB 67 67dB dB 62 dB 15 20 70 dB 71 dB 67 69dB dB 20 71 dB 69 dB Figure 29 shows there is a signiﬁcant drop in noise level for the optimized wing, up to 40% Figure 29 shows there is a significant drop in noise level for the optimized wing, up to 40% dependentdependent on angleon angle of attack.of attack. Although Although this this datadata willwill not be be 100% 100% accurate accurate because because the the wind wind tunnel tunnel nevernever maintains maintains at constantat constant speed speed so so the the noise noise levellevel varies. The The actual actual noise noise reduction reduction value value will will be be betweenbetween 30% 30% and and 40% 40% less less than than the the conventional conventional wingwing value.

Figure 29. A graph showing the difference in noise level between the conventional wing and Figure 29. A graph showing the difference in noise level between the conventional wing and optimised optimised morphing wing. morphing wing. The noise was measured off the wing section using a sound level meter; it was placed inside the windThe noisetunnel was 10 cm measured behind offthe thewings wing trailing section edge using and a10 sound cm abov levele. As meter; the wind it was tunnel placed does inside not the windmaintain tunnel a 10 set cm speed, behind multiple the wings decibel trailing readings edge were and taken 10 and cm above.an average As of the the wind data tunnelwas used does for not maintaineach angle a set of speed, attack. multiple decibel readings were taken and an average of the data was used for each angleThe of Figures attack. 30–33 show the standard deviation and error bar charts for the data collected from wing noise. The wind tunnel does not maintain a set speed therefore noise is not constant, five decibel readings were taken per angle of attack shown to gain an average noise level.

Aerospace 2017, 4, 52 20 of 24 Aerospace 2017, 4, 52 20 of 24 Aerospace 2017, 4, 52 20 of 24 Aerospace 2017, 4, 52 20 of 24 The figures below show standard deviation for noise data collected during tests conducted at the windThe figurestunnel, belowincluding show the standard interquartile deviation rang efor (indicated noise data by collected green and during brown). tests The conducted green and at TheThe Figures figures 30 below–33 show show the standard standard deviation deviation for no andise errordata collected bar charts during for the tests data conducted collected at from brownthe wind boxes tunnel, indicate including the difference the interquartile between rang the 75the (indicated and 25th by percentile green and or brown). the middle The 50%green of andthe wingthe noise. wind Thetunnel, wind including tunnel doesthe interquartile not maintain rang a sete (indicated speed therefore by green noise and isbrown). not constant, The green ﬁve and decibel data’sbrown standardboxes indicate deviation, the difference the maximum between and the mi 75thnimum and is25th also percentile shown byor thethe middleupper and50% lowerof the readingsbrown were boxes taken indicate per the angle difference of attack between shown the to 75th gain and an average25th percentile noise level.or the middle 50% of the whiskersdata’s standard respectively. deviation, the maximum and minimum is also shown by the upper and lower data’swhiskers standard respectively. deviation, the maximum and minimum is also shown by the upper and lower whiskers respectively.

Figure 30. A graph showing the standard deviation of noise level for the conventional wing. FigureFigure 30. 30.A graphA graph showing showing the the standard standard deviationdeviation of of noise noise level level for for the the conventional conventional wing. wing. Figure 30. A graph showing the standard deviation of noise level for the conventional wing.

Figure 31. A graph showing the standard deviation of noise level for the shape shifting wing. Figure 31. A graph showing the standard deviation of noise level for the shape shifting wing. FigureFigure 31. 31.A A graph graph showing showing thethe standardstandard deviation deviation of of noise noise level level for for the the shape shape shifting shifting wing. wing.

Figure 32. Conventional wing noise error bars. Figure 32. Conventional wing noise error bars.

FigureFigure 32.32. Conventional wing noise error bars. Conventional wing noise error bars.

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FigureFigure 33. 33. ShapeShape shifting shifting wing wing noise noise error error bars. bars.

7. Conclusions The figures below show standard deviation for noise data collected during tests conducted at theThe wind main tunnel, objective including of this paper the interquartile was to design range and (indicateddevelop a morphing by green andwing brown). technology The based green onand implementation brown boxes indicate of transition the difference regions between between the the wing 75th and and the 25th flap, percentile which was or the simpler middle than 50% the of currentthe data’s sophisticated standard deviation, technology, the ideally maximum reducing and th minimume timescale is alsofor operational shown by theuse. upper To achieve and lower this, wingwhiskers models respectively. were created and tested through CFD simulation to determine the aerodynamic performance of the wing. Also, wind tunnel tests were carried out to measure lift and drag coefficients7. Conclusions and to determine noise level reduction compared to conventional wing. SimulationThe main objective was carried of this out paper on Ansys was Fluent to design for andeach develop of the models. a morphing Simulation wing technologyrevealed a 10.8% based coefficienton implementation of lift increase of transition for the initial regions morphing between wing the and wing 15.4% and thefor the flap, optimized which was morphing simpler wing than asthe compared current sophisticated to conventional technology, wing design. ideally At angles reducing of attack the timescale of 0, 5, 10 for and operational 15 degrees use. the Tooptimized achieve wingthis, winghas an models increase were in lift-to-drag created and ratio tested of 18.3%, through 10.5%, CFD 10.6% simulation and 4% to respectively determine thewhen aerodynamic compared withperformance the conventional of the wing. wing. Also, Simulations wind tunnel also tests showed were carried that there out to is measure a significant lift and improvement drag coefficients on pressureand to determine distribution noise over level the reduction lower surface compared of th toe conventionalmorphing wing wing. aerofoil. The increase in flow smoothnessSimulation and wasreduction carried in out vortex on Ansys size reduced Fluent for pres eachsure of drag themodels. along the Simulation trailing edge revealed of the a 10.8%wing whichcoefficient caused of lift an increaseincrease forin pressure the initial on morphing the lower wing surface. and 15.4%A morphing for the wing optimized reduced morphing the size wing of the as vorticescompared and to therefore conventional the noise wing levels design. of Atthe anglesmorphing of attack wings of based 0, 5, 10 on and data 15 collected, degrees the was optimized reduced bywing up hasto 50%. an increase in lift-to-drag ratio of 18.3%, 10.5%, 10.6% and 4% respectively when compared with the conventional wing. Simulations also showed that there is a significant improvement on Acknowledgments: The research was self-funded. The authors would also like to thank ANSYS, Inc. for the pressure distribution over the lower surface of the morphing wing aerofoil. The increase in flow possibility to use their software. smoothness and reduction in vortex size reduced pressure drag along the trailing edge of the wing Authorwhich causedContributions: an increase Deanin Ninian pressure designed on the the lower experiments; surface. pe Arformed morphing the experiments; wing reduced and the analysed size of the data;vortices Dean and Ninian therefore and Sam the noiseDakka levels wrote ofthe the paper. morphing The research wings was based performed on data collected,in the frame was work reduced of Dean by Ninian’sup to 50%. Engineering Degree under the supervision and direction of Dr. Sam Dakka. Conflicts of Interest: The authors declare no conflict of interests. Acknowledgments: The research was self-funded. The authors would also like to thank ANSYS, Inc. for the possibility to use their software. Abbreviations Author Contributions: Dean Ninian designed the experiments; performed the experiments; and analysed the data; TheDean following Ninian andabbreviations Sam Dakka are wrote used the in paper.this manuscript: The research was performed in the frame work of Dean Ninian’s Engineering Degree under the supervision and direction of Sam Dakka. P1 sound pressure, Pa Conflicts of Interest: The authors declare no conflict of interests. P0 reference pressure (threshold of hearing 2 × 10−5 Pa), Pa

Cp Pressure Coefficient,- dB Noise Level, decibels AOA Angle of Attack NACA National Advisory Committee on Aeronautics DARPA The Defense Advanced Research Projects Agency UAV Unmanned Air Vehicle SMA Shape Memory Alloy PZT Piezoelectric Transducer

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Abbreviations The following abbreviations are used in this manuscript:

P1 sound pressure, Pa −5 P0 reference pressure (threshold of hearing 2 × 10 Pa), Pa Cp Pressure Coefﬁcient dB Noise Level, decibels AOA Angle of Attack NACA National Advisory Committee on Aeronautics DARPA The Defense Advanced Research Projects Agency UAV Unmanned Air Vehicle SMA Shape Memory Alloy PZT Piezoelectric Transducer Aerospace 2017, 4, 52 22 of 24 CFD Computational Fluid Dynamics RNGCFD Re-Normalization Computational Group Fluid Dynamics RNG Re-Normalization Group Appendix A Appendix A

(a)

(b)

Figure A1. (a) Conventional aerofoil with flap, units in mm; (b) Initial shape shifting wing design Figure A1.with(a )flap, Conventional units in mm. aerofoil with ﬂap, units in mm; (b) Initial shape shifting wing design with ﬂap, units in mm.

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FigureFigure A2. A2.Conventional Conventional aerofoil aerofoil with with flap, flap, Initial Initial shapeshape shifting shifting wing wing design design with with flap flap and and NACA009 NACA009 with no flap [30]. with no flap [30].

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