aerospace

Article Experimental Investigation of the Wake and the of a UAV Model

Pericles Panagiotou, George Ioannidis ID , Ioannis Tzivinikos and Kyros Yakinthos * ID

Laboratory of Fluid Mechanics and Turbomachinery, Department of Mechanical Engineering, Aristotle University of Thessaloniki, 54124 Thessaloniki, Greece; [email protected] (P.P.); [email protected] (G.I.); [email protected] (I.T.) * Correspondence: [email protected]; Tel.: +30-231-099-6411

Received: 5 October 2017; Accepted: 30 October 2017; Published: 1 November 2017

Abstract: An experimental investigation of the wake of an Unmanned Aerial Vehicle (UAV) model using flow visualization techniques and a 3D Laser Doppler Anemometry (LDA) system is presented in this work. Emphasis is given on the flow field at the wingtip and the investigation of the tip vortices. A comparison of the velocity field is made with and without winglet devices installed at the wingtips. The experiments are carried out in a closed-circuit subsonic . The flow visualization techniques include smoke-wire and smoke-probe experiments to identify the flow phenomena, whereas for accurately measuring the velocity field point measurements are conducted using the LDA system. Apart from the measured velocities, and circulation quantities are also calculated and compared for the two cases. The results help to provide a more detailed view of the flow field around the UAV and indicate the winglets’ significant contribution to the deconstruction of -tip structures.

Keywords: UAV; measurements; LDA; vortex; winglets

1. Introduction In recent years, there is an increasing trend in the development and use of fixed-wing Unmanned Aerial Vehicles (UAVs) by both military forces and civilian organizations [1]. They are ideal solutions for a wide range of operations, such as fire detection, search and rescue, coastline and sea-lane monitoring, and security surveillance [2]. Due to the absence of crew on-board, they present several advantages, such as the reduced operational cost, the ability to operate under hazardous conditions, and the increased flight endurance, which is one the most important characteristics when it comes to the aforementioned missions. In the case of a UAV, where the endurance is only limited by the available fuel on-board, it is essentially up to the aerodynamic efficiency of the configuration to maximize the flight time. For a given mission, the required force is defined, therefore the optimization process is essentially keeping the corresponding drag force as low as possible. For a UAV that operates in the low subsonic, incompressible regime, as is the case with most Medium-Altitude-Long-Endurance (MALE) UAVs [3–6], a drag breakdown analysis was conducted in [7], which shows that the part that contributes the most to the total drag force of the aerial vehicle is the main wing. More specifically, when exposed in subsonic, incompressible flow, a finite wing’s drag is a combination of profile drag and induced drag. The profile drag is the drag due to the shape of the body (skin friction + flow effects), whereas the induced drag is a result of the imbalance at the tip of a finite wing between its upper (suction side) and lower (pressure side) surfaces. At the tip though, this imbalance causes the high-pressure air from the lower side to move upwards, where the pressure is lower, leading to the formation of a vortex, i.e., the wingtip vortex (Figure1). This three-dimensional motion alters the flow field above the entire wing as well, thus resulting in the

Aerospace 2017, 4, 53; doi:10.3390/aerospace4040053 www.mdpi.com/journal/aerospace Aerospace 2017, 4, 53 2 of 17 Aerospace 2017, 4, 53 2 of 17 appearancedrag to the of total the induceddrag force drag budget, term [a8 ].lot Due of effort to the has importance been made of in the the induced previous drag decades to the totalto minimize drag forcethis budget, component a lot ofand effort pertain has been the madewingtip in thevortex. previous The most decades notable to minimize achievement this component is probably and the pertainwinglet the concept, wingtip vortex.introduced The mostby Wh notableitcomb achievement [9]. Having isbeen probably studied the by winglet various concept, researchers introduced over the byyears, Whitcomb for issues [9]. Having that have been to studieddo with byaerodynamic various researchers and structural over the issues, years, using for issues both experimental that have to do and withcomputational aerodynamic tools and structuralfor design issues, and optimization using both experimental purposes [10–14], and computational it is essentially tools a sophisticated for design andtype optimization of wingtip purposes fence that [10 alters–14], the it is flow essentially at the tip a sophisticated in a manner that type increases of wingtip aerodynamic fence that alters efficiency the flowand at reduces the tip in the a mannersize of the that tip increases vortex. aerodynamic efficiency and reduces the size of the tip vortex.

Figure 1. Tip vortex formation [7]. Figure 1. Tip vortex formation [7]. A lot of experimental studies exist that deal with the flow over and tip vortices. A lot of experimental studies exist that deal with the flow over wings and tip vortices. Even Even though the configurations vary, the flow development is based on the same fundamental though the configurations vary, the flow development is based on the same fundamental principles, principles, which means that the information from the existing literature can serve as a solid reference which means that the information from the existing literature can serve as a solid reference point. For point. For instance, Shekarriz et al. [15] studied the near field behavior of a vortex, yielding instance, Shekarriz et al. [15] studied the near field behavior of a wing tip vortex, yielding useful useful conclusions regarding the velocity distributions and the roll-up process. Using the data provided conclusions regarding the velocity distributions and the roll-up process. Using the data provided by by Hoffmann and Joubert [16], Nielsen and Schwind [17] divided the vortex into three regions based Hoffmann and Joubert [16], Nielsen and Schwind [17] divided the vortex into three regions based on on an analogy with turbulent boundary layers: the vortex core (inner) region, the logarithmic region, an analogy with turbulent boundary layers: the vortex core (inner) region, the logarithmic region, and the defect law region. Based on the equations describing these regions, Corsiglia et al. [18] and the defect law region. Based on the equations describing these regions, Corsiglia et al. [18] conducted three-dimensional hot-wire anemometer measurements in the far field, providing modified conducted three-dimensional hot-wire anemometer measurements in the far field, providing empirical equations and constants that describe the circulation distributions, as well as information modified empirical equations and constants that describe the circulation distributions, as well as regarding the meandering (lateral movement) of the wingtip vortex core in space as a function of information regarding the meandering (lateral movement) of the wingtip vortex core in space as a time. This phenomenon, known as “vortex wandering”, has also been experimentally investigated by function of time. This phenomenon, known as “vortex wandering”, has also been experimentally various researchers over time [19–23]. Huang and Lin [24] experimentally studied the flow patterns and investigated by various researchers over time [19–23]. Huang and Lin [24] experimentally studied characteristics of vortex shedding on a wing with a NACA 0012 airfoil section using smoke-wire and the flow patterns and characteristics of vortex shedding on a wing with a NACA 0012 airfoil section surface oil-flow techniques for visualization, and hot-wire anemometry to characterize the frequency using smoke-wire and surface oil-flow techniques for visualization, and hot-wire anemometry to domain of the unsteady flow structures. The information on the frequency domain can be useful when characterize the frequency domain of the unsteady flow structures. The information on the frequency it comes to estimating the expected frequencies and setting the acquisition time for reliable mean domain can be useful when it comes to estimating the expected frequencies and setting the values, as is the case in the current work. Pino et al. [25] studied the velocity field behind a wingtip acquisition time for reliable mean values, as is the case in the current work. Pino et al. [25] studied vortex and compared the results with trailing vortex theoretical models, whereas in a recent study the velocity field behind a wingtip vortex and compared the results with trailing vortex theoretical Serrano-Aguilera et al. [26] developed a theoretical model in order to characterize the wingtip vortices models, whereas in a recent study Serrano-Aguilera et al. [26] developed a theoretical model in order in the near field using smoke-wire visualizations in combination with a laser sheet. Their results were to characterize the wingtip vortices in the near field using smoke-wire visualizations in combination used as a first validation in the flow visualizations that were employed in this paper. In a comparative with a laser sheet. Their results were used as a first validation in the flow visualizations that were study, Zheng and Ramaprian [27] conducted a detailed study of the vortex of a rectangular wing, employed in this paper. In a comparative study, Zheng and Ramaprian [27] conducted a detailed comparing the flowfield of a stationary wing to the case where a sinusoidal pitch oscillation is induced. study of the vortex of a rectangular wing, comparing the flowfield of a stationary wing to the case Moreover, Elsayed et al. [28] conducted Particle Image Velocimetry (PIV) studies, at flapped and where a sinusoidal pitch oscillation is induced. Moreover, Elsayed et al. [28] conducted Particle Image unflapped wing configurations, and compared the velocity fields, vorticity contours, and circulation Velocimetry (PIV) studies, at flapped and unflapped wing configurations, and compared the velocity charts. In a very interesting approach, they showed how various vortex characteristics depend directly fields, vorticity contours, and circulation charts. In a very interesting approach, they showed how various vortex characteristics depend directly on the angle of attack, emphasizing on peak tangential velocity, vorticity, and circulation distribution. Finally, Muthusamy [29] carried out experiments in a

Aerospace 2017, 4, 53 3 of 17 Aerospace 2017, 4, 53 3 of 17 configurationon the angle of with attack, and emphasizing without winglets on peak at a tangential low speed velocity, windtunnel. vorticity, The and study circulation was limited distribution. to force balanceFinally, Muthusamymeasurements [29 ]though, carried and out experimentsno informat inion a was configuration provided with regarding and without the flow winglets development at a low andspeed the windtunnel. effect the winglet The study has wason the limited tip vortex. to force To balance the best measurements of our knowledge, though, no andstudy no exists information where awas direct provided comparison regarding is made the flowin terms development of flow field and development the effect the wingletbetween has a configuration on the tip vortex. with, To and the withoutbest of oura winglet knowledge, device noinstalled study at exists the wingtip, where a and direct few comparison of the existing is made studies in on terms wingtip of flow vortices field dealdevelopment with UAV between platforms. a configuration However, and with, since and in UAV without development a winglet devicethe time installed and budget at the is wingtip,limited whenand few compared of the existing to a commercial studies on , wingtip the vortices wings dealare of with less UAV complex platforms. geometry However, (in terms and of since airfoil in profiles,UAV development twist angle the etc.) time and and rela budgettively simple is limited to manufacture. when compared Thus, to a the commercial tips are highly airliner, loaded the wings and areare ideal of less platforms complex for geometry winglet (instudies terms [11]. of airfoil profiles, twist angle etc.) and relatively simple to manufacture.In the current Thus, paper, the tips an areexperime highlyntal loaded study and of arethe idealflow platformsfield around for a winglet UAV model studies is [presented.11]. The aim,In the and current main contribution, paper, an experimental is to address study the two of the aforementioned flow field around remarks. a UAV That model is, to is investigate presented. theThe flowfield aim, and at main the wing contribution, tip of the is UAV to address and to the quanti twofy aforementioned the effect the winglets remarks. have That on is, the to investigatetip vortex. Thethe flowfieldstudy is part at the of the wing Hellenic tip of the Civil UAV Unmanned and to quantify Air Vehicle the effect(HCUAV) the winglets design study have onthat the was tip carried vortex. outThe at study the isLaboratory part of the Hellenicof Fluid CivilMechanics Unmanned and AirTurbomachinery Vehicle (HCUAV) (LFMT) design at study the thatDepartment was carried of Mechanicalout at the Laboratory Engineering, of Fluid at Mechanicsthe Aristotle and TurbomachineryUniversity of Thessaloniki (LFMT) at the (AUTH) Department [6,7]. of A Mechanical detailed presentationEngineering, of at thethe Aristotleexperimental University studies of Thessalonikiis made, including (AUTH) flow [6,7]. visualization A detailed presentation techniques of and the velocityexperimental measurements. studies is Smoke-wire made, including, laser flow sheet, visualization and smoke-probe techniques techniques and velocity are used measurements. for the flow visualizationSmoke-wire, laserinvestigations, sheet, and smoke-probewhereas point techniques measurements are used by formeans the flowof Laser visualization Doppler investigations,Anemometry (LDA)whereas are point performed measurements in the near by wake. means Two of Laser different Doppler wingtip Anemometry configurations (LDA) are are used, performed one with- in and the onenear without wake. Twothe winglet different device wingtip of the configurations HCUAV, to arecompare used, the one velocity with- and field one for without the two the cases. winglet The wingletdevice ofconfiguration the HCUAV, was to compare a result theof a velocity parametric field CFD for the study two [7] cases. and Theits characteristics winglet configuration are given was in Figurea result 2. of a parametric CFD study [7] and its characteristics are given in Figure2.

FigureFigure 2 2.. HellenicHellenic Civil Civil Unmanned Unmanned Air Air Ve Vehiclehicle (HCUAV) (HCUAV) configuration configuration wi withoutthout and and with with the the winglet winglet installed,installed, including including the the winglet’s winglet’s ch characteristicsaracteristics in the detail [7]. [7].

2.2. Materials Materials and and Methods Methods TheThe experiments experiments were were conducted conducted in in a a closed-circuit, closed-circuit, low-speed low-speed windtunnel windtunnel facility facility at at the the LFMT. LFMT. TheThe wind tunnel has an 1810 mm × 600600 mm mm ×× 600600 mm mm test test section section (Figure (Figure 33)) andand isis capablecapable ofof producingproducing a a maximum speed of 50 m/s. The The windtunnel windtunnel blower blower speed speed is is controlled controlled by by a a frequency frequency inverter.inverter. A A smoke smoke generator generator is is placed placed inside inside the the clos closed-looped-loop wind wind tunnel, tunnel, away away from from the the test test section. section. Considering the models used in the study, two scaled variants (1:22) of the HCUAV were used, one without and one with the winglet devices installed at the tip. The models were manufactured by means of 3D printing to ensure accurate representation of the original geometry. Specifically, the main wing has a mean chord of 36.5 mm, a tip chord (c) of 21 mm, an Aspect Ratio of 8, and a NASA NLF(1)-1015 airfoil [30], whereas the wing span is 302 mm. A custom mechanism with a supporting rod is used to place the model at the center of the test section area, away from the windtunnel walls, and ensures that the model is accurately positioned within ± 0.5 deg (Figure 4). The model maximum cross-sectional area at the largest examined angle of attack (15 deg) divided by the area of the wind-tunnel cross-section is equal to 1.05%. Hence, no corrections due to the blockage effect were made [31].

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Figure 3. Schematic of the experimental setup and arrangement of the measurement devices inside the Figure 3. Schematic of the experimental setup and arrangement of the measurement devices inside test section. the test section.

Considering the models used in the study, two scaled variants (1:22) of the HCUAV were used, one without and one with the winglet devices installed at the tip. The models were manufactured by means of 3D printing to ensure accurate representation of the original geometry. Specifically, the main wing has a mean chord of 36.5 mm, a tip chord (c) of 21 mm, an Aspect Ratio of 8, and a NASA NLF(1)-1015 airfoil [30], whereas the wing span is 302 mm. A custom mechanism with a supporting rod is used to place the model at the center of the test section area, away from the windtunnel walls, and ensures that the model is accurately positioned within ±0.5 deg (Figure 4). The model maximumFigure cross-sectional3. Schematic of the area experimental at the largest setup examined and arrangement angle of attackof the measurement (15 deg) divided devices by theinside area of the wind-tunnelthe test section. cross-section is equal to 1.05%. Hence, no corrections due to the blockage effect were made [31].

Figure 4. The HCUAV model inside the windtunnel.

A series of flow visualization experiments were initially conducted on both UAV models for the flow phenomena to be identified and for the measured planes to be defined. More specifically, both smoke-probe and smoke-wire techniques were used to investigate the flow at the wingtip, at a wide range of angles of attack. Regarding the smoke-probe, a 3 mm diameter plastic probe is used, mounted on a set of copper pipes, through which the smoke was channeled. The copper pipes set is attached to a three-axes traversing mechanism featuring high accuracy and repeatability (with 4 μm standard deviation) in order for the smoke probe to be accurately positioned . The mechanism consists of a driving bolt connected Figureto a stepper 4. The HCUAVmotor for model each inside axis. theFinally, windtunnel. a flexible rubber pipe connects the copper pipes with the smokeFigure 4.generator. The HCUAV To modelcarry insideout the the smoke-wire windtunnel. study, a custom frame is used. Specifically,A series of flowa 0.3 visualizationmm diameter experimentswire is mounted were on initially a rigid aluminum conducted frame, on both and UAV a glass models syringe for A series of flow visualization experiments were initially conducted on both UAV models for the isthe adjusted flow phenomena on top, where to the be identifiedglycerin-water and formixtur thee measured was stored. planes The mixture to be defined. consists More of 30% specifically, glycerin flow phenomena to be identified and for the measured planes to be defined. More specifically, both andboth 70% smoke-probe distilled water and and smoke-wire drips on techniquesthe wire thro wereugh usedthe glass to investigate syringe. Proper the flow positioning at the wingtip, of the smoke-probe and smoke-wire techniques were used to investigate the flow at the wingtip, at a wide syringeat a wide ensuresrange the of angles right amount of attack. of Regarding mixture dripping the smoke-probe, over thea wire. 3 mm Α diametern external plastic power probe supply is used, is range of angles of attack. Regarding the smoke-probe, a 3 mm diameter plastic probe is used, usedmounted to charge on a the set wire, of copper thus pipes,creating through a close whichcircuit, the in smokeorder for was the channeled. mixture on The the copper wire to pipes be burnt. set is mounted on a set of copper pipes, through which the smoke was channeled. The copper pipes set is Theattached entire toframe a three-axes is in turn traversing placed in mechanismthe test section, featuring close to high the accuracymodel. The and photographs repeatability were (with taken 4 µm attached to a three-axes traversing mechanism featuring high accuracy and repeatability (with 4 μm usingstandard a Nikon deviation) D3000 inDSLR order camera for the and smoke Nikorr probe 35 tomm, be accuratelyf 1.8 and Nikorr positioned. 18–105 The mm, mechanism f 3.5–5.6 lenses. consists standardRegarding deviation) the LDA in order measurements, for the smoke and probe since to the be flow accurately field is positionedsymmetric,. Theonly mechanism half the wake consists was examined,of a driving so boltthat connected the experiments to a stepper were motor conduct fored each at axis.the right Finally, wing a flexibleof the model.rubber pipeThe connectsrelative the copper pipes with the smoke generator. To carry out the smoke-wire study, a custom frame is

used. Specifically, a 0.3 mm diameter wire is mounted on a rigid aluminum frame, and a glass syringe is adjusted on top, where the glycerin-water mixture was stored. The mixture consists of 30% glycerin and 70% distilled water and drips on the wire through the glass syringe. Proper positioning of the syringe ensures the right amount of mixture dripping over the wire. Αn external power supply is used to charge the wire, thus creating a close circuit, in order for the mixture on the wire to be burnt. The entire frame is in turn placed in the test section, close to the model. The photographs were taken using a Nikon D3000 DSLR camera and Nikorr 35 mm, f 1.8 and Nikorr 18–105 mm, f 3.5–5.6 lenses. Regarding the LDA measurements, and since the flow field is symmetric, only half the wake was examined, so that the experiments were conducted at the right wing of the model. The relative

Aerospace 2017, 4, 53 5 of 17 of a driving bolt connected to a stepper motor for each axis. Finally, a flexible rubber pipe connects the copper pipes with the smoke generator. To carry out the smoke-wire study, a custom frame is used. Specifically, a 0.3 mm diameter wire is mounted on a rigid aluminum frame, and a glass syringe is adjusted on top, where the glycerin-water mixture was stored. The mixture consists of 30% glycerin and 70% distilled water and drips on the wire through the glass syringe. Proper positioning of the syringe ensures the right amount of mixture dripping over the wire. An external power supply is used to charge the wire, thus creating a close circuit, in order for the mixture on the wire to be burnt. The entire frame is in turn placed in the test section, close to the model. The photographs were taken using a Nikon D3000 DSLR camera and Nikorr 35 mm, f 1.8 and Nikorr 18–105 mm, f 3.5–5.6 lenses. Regarding the LDA measurements, and since the flow field is symmetric, only half the wake was examined, so that the experiments were conducted at the right wing of the model. The relative position for each measurement plane is presented in Figure5. The first plane lies half a tip chord (X/c = 0.5), the second plane one tip chord (X/c = 1), and the third planes lies two tip chords (X/c = 2) downstream of the wingtip, respectively. The measurement grid at each of the planes is the same, and has 625 nodes in a 25 × 25 arrangement, covering a 50 mm × 50 mm area. In order to increase the accuracy, the grid is denser at the location of the vortex, as indicatively presented for the regular wingtip case in Figure6. The starting point of the coordinate system at each measurement was located at tip of the wing/winglet for all measurements. To conduct the measurements a complete 3-component Aerospace 2017, 4, 53 5 of 17 LDA system from Dantec Dynamics (Skovlunde, Denmark) is used, which consists of an Ar-ion positionBeam Generator, for each measurement an optical system plane with is presented 3D component in Figure Probes, 5. The a first 3-axes plane traverser, lies half and a tip a chord dedicated (X/c =commercial 0.5), the second software plane (BSA one Software tip chord© by(X/ Dantecc = 1), and Dynamics the third A/S). planes The lies Traverser two tip onchords which (X the/c =3D 2) downstreamcomponent Probes of the arewingtip, mounted, respec istively. placed The next measurement to the test-section, grid at with each the of Probes the planes facing is thethe model.same, andUsing has the 625 smoke nodes generator, in a 25 × 25 a smoke arrangement, cloud is covering generated a throughout50 mm × 50 themm entire area. pathIn order of the to increase windtunnel, the accuracy,thus inducing the grid the particlesis denser in at the the flow location that are of neededthe vortex, for the as LDAindicatively Measurements. presented Since for nothe apparent regular wingtipfrequencies case were in Figure expected 6. The in starting the lower point frequency of the coordinate range [18 ,system21,22], theat each acquisition measurement time was was set located at 5 s atfor tip each of node, the wing/winglet adequately long for for all providing measuremen reliablets. To mean conduct values. the measurements a complete 3- componentFinally, LDA regarding system the from inlet Dantec conditions, Dynamics for the (Skovl LDAunde, measurements Denmark) theis used, freestream which velocityconsists atof thean 3 Ar-iontest section Beam is 13.2Generator, m/s, held an optical constant system within with 1%, which3D component gives a Reynolds Probes, Numbera 3-axes equaltraverser, to 33 and× 10 a dedicatedbased on thecommercial mean aerodynamic software (BSA chord, Software© resembling by Dantec existing Dynamics wingtip A/S). vortex-related The Traverser studies on [which26,28]. theThe 3D freestream component Probes are intensity mounted, was is placed measured nextat to 0.7%.the test-section, For theflow with visualizationthe Probes facing studies, the model.the freestream Using velocitythe smoke is 2 generator, m/s to make a smoke sure that cloud the smokeis generated is clearly throughout visible [32 ].the Due entire to the path difference of the windtunnel,in freestream thus velocity inducing and, consequently,the particles in in the the flo correspondingw that are needed Reynolds for the number, LDA Measurements. no direct comparison Since nocan apparent be made frequencies between the were flow expected visualization in the results lower andfrequency the LDA range measurements. [18,21,22], the The acquisition visualization time wasstudies set areat 5 critical s for each though, node, to adequately get a first view long of for the providing flowfield reliable and accurately mean values. position and size the LDA measurements grids. The lateral velocity components are insignificant in all studies.

Figure 5. Measurement planes for the Laser Doppler Anemometry (LDA) measurements around the Figure 5. Measurement planes for the Laser Doppler Anemometry (LDA) measurements around the Hellenic Civil Unmanned Air Vehicle (HCUAV) model. Hellenic Civil Unmanned Air Vehicle (HCUAV) model.

Figure 6. High resolution grid for the 3rd measurement plane (regular wingtip). The trailing edge is marked with dark red.

Finally, regarding the inlet conditions, for the LDA measurements the freestream velocity at the test section is 13.2 m/s, held constant within 1%, which gives a Reynolds Number equal to 33 × 103 based on the mean aerodynamic chord, resembling existing wingtip vortex-related studies [26,28].

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position for each measurement plane is presented in Figure 5. The first plane lies half a tip chord (X/c = 0.5), the second plane one tip chord (X/c = 1), and the third planes lies two tip chords (X/c = 2) downstream of the wingtip, respectively. The measurement grid at each of the planes is the same, and has 625 nodes in a 25 × 25 arrangement, covering a 50 mm × 50 mm area. In order to increase the accuracy, the grid is denser at the location of the vortex, as indicatively presented for the regular wingtip case in Figure 6. The starting point of the coordinate system at each measurement was located at tip of the wing/winglet for all measurements. To conduct the measurements a complete 3- component LDA system from Dantec Dynamics (Skovlunde, Denmark) is used, which consists of an Ar-ion Beam Generator, an optical system with 3D component Probes, a 3-axes traverser, and a dedicated commercial software (BSA Software© by Dantec Dynamics A/S). The Traverser on which the 3D component Probes are mounted, is placed next to the test-section, with the Probes facing the model. Using the smoke generator, a smoke cloud is generated throughout the entire path of the windtunnel, thus inducing the particles in the flow that are needed for the LDA Measurements. Since no apparent frequencies were expected in the lower frequency range [18,21,22], the acquisition time was set at 5 s for each node, adequately long for providing reliable mean values.

Figure 5. Measurement planes for the Laser Doppler Anemometry (LDA) measurements around the

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Figure 6. High resolution grid for the 3rd measurement plane (regular wingtip). The trailing edge is markedFigure 6. with High dark resolution red. grid for the 3rd measurement plane (regular wingtip). The trailing edge is marked with dark red. 3. Results Finally, regarding the inlet conditions, for the LDA measurements the freestream velocity at the test Atsection first, is the 13.2 results m/s, atheld the constant tip of the within model 1%, with which the regulargives a Reynolds wingtip (winglet Number device equal removed)to 33 × 103 arebased shown, on the followed mean aerodynamic by the results chord, at the resembling tip of the model existing with wingtip the winglet vortex-related installed. studies Concluding, [26,28]. a comparison of the flow field between the two configurations is made. The contour plots show the flow field in the wake of the HCUAV right wing, assuming that the view point of the reader is downstream of the aerial vehicle, whereas line plots and other charts are also shown to give a better understanding of the phenomena inside the core. The normal and spanwise axes have been normalized using the tip chord of the model (c), and the time-averaged velocities have been normalized using the freestream velocity (Uinf). Moreover, the vorticity and circulation quantities are normalized by appropriately using the tip chord and freestream velocity.

3.1. Regular Wingtip Figure7 presents the results of the smoke-probe investigations (top view), whereas Figure8 presents the results of the combined smoke-wire/laser sheet method, at one tip chord downstream of the wing (X/c = 1). The corresponding planes and direction of freestream velocity are also given for each case. As is evident, at 0 deg of angle of attack, the tip vortex is practically non-existent, since the lift production is negligible and the pressure imbalance between the two sides is very small. At −5 and 5 deg, the vortex is more clearly defined, especially in the laser sheet visualization. At −5 deg of angle of attack the vortical structure is upside down, since the lift production is inverse. At 10 deg of angle of attack, the tip vortex appears to be even larger in size and robust, whereas at 15 deg, the vortical structure retains its size but fades quickly as the flow moves downstream of the wing. In general, the flow visualization studies may not directly provide numbers about the examined phenomenon, but they helped a long way in identifying the key-characteristics of the vortex at each angle of attack, i.e., size and core location. Thus, the LDA measurement grids could be properly sized with greater accuracy, at a reduced time. Aerospace 2017, 4, 53 7 of 17 Aerospace 2017, 4, 53 7 of 17

FigureFigure 7. 7.Smoke-probe Smoke-probe visualization visualization studies studies at at the the wingtip. wingtip.

FigureFigure 8. 8.Laser Laser sheet sheet visualization visualization studies studies at at the the wingtip. wingtip.

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The LDA results are presented in Figures9 and 10 in contour plots and in Figures 11–13 in line plots. Note that the vortex core radius is defined as the radius at which the swirl velocity component maximizes, whereas the axial vorticity was calculated using the experimental data from the swirl velocity components, and more specifically

dW dV Ω = − (1) dΥ d where V is the normal and W the spanwise, time-averaged velocity components. In Figure9, the time-averaged axial, normal and spanwise velocity components and axial vorticity are presented in contour plots at X/c = 1 and three different angles of attack, whereas in Figure 10, the same variables are shown for the three measurement levels at 10 deg of angle of attack in contour plots at the three measurement locations. The freestream velocity direction is identical the one presented in Figure8. When observing the axial velocity component (U), it is evident that the wake at the wingtip can be divided at two separate regions. One region is behind the main wing, which corresponds to the typical wake profile observed behind a blunt body, as a result of shear layer separation, and the other corresponds to the wingtip vortex. In the wing-wake region, U is of smaller order of magnitude when compared to the vortex core one, with the exact values varying greatly between the different angles of attack and measurement levels. Regarding the time-averaged normal and spanwise velocity components (V and W), two maxima of opposite sign are observed around a point where V and W equal to zero. This point coincides with the one where the local minimum of the U component is located (U0) at every measurement location. Hence, that point was defined as the center of the vortex. At every measurement plane, the V magnitude increases considerably as the angle of attack grows (Figure9), but for the same angle of attack, as the flow moves downstream of the tip, the difference in magnitude is smaller (Figure 10). The same observations can be made for the spanwise velocity component. Overall, due to the fact that the W component itself does not add any new information on the phenomenon, and since the axial vorticity includes both of the V and W components, it was decided to not present the spanwise component in the rest of the figures of this work. The time-averaged axial velocity component inside the vortex core can be more clearly seen in Figure 11, where line plots show the normalized values of U. It should be noted that for each measurement the line plot sections pass through the corresponding vortex center. The time-averaged axial velocity component appears to form an axisymmetric wake-like profile. This axisymmetric behavior is also observed in Figures 12 and 13, which indicate that the roll-up process is complete for X/c < 1, as also reported in [13,25]. Namely, Figure 12 presents the V component in line plot representation. At 15 deg of angle of attack, the range of the swirl velocity magnitude is approximately three times higher than the one measured at 0 deg, indicating that the tip vortex grows in strength. Regarding the change in magnitude between different longitudinal positions, an average decrease of 10% was calculated between X/c = 1 and X/c = 0.5 for every examined angle of attack, whereas between X/c = 2 and X/c = 1, the percentage average is approximately 13%. Moreover, it should be noted that behind the main wing geometry, the normal velocity component has negative values outside the vortex core. This is probably due to the downwash effect and the corresponding velocity component that appears in the normal axis. The third parameter in the contour plots, i.e., the axial vorticity, is also presented in line plots in Figure 13. At every measurement level the vorticity maximum is approximately three times larger at 15 deg than the one at 0 deg. Between 10 and 15 deg of angle of attack, though, the difference in peak vorticity is very small, indicating that the lift increases at a much lower rate and, possibly, that stalling occurs. As a general observation, the peak vorticity values follow the trend of the normal velocity, having a 10–15% difference from level to level, at each angle of attack, an observation that is in agreement with the data from [28]. However, the line plots provide the information in a more detailed way, so that a direct comparison with another experimental or computational data, can be made. Aerospace 2017, 4, 53 99 of of 17 17

Figure 9. Time-averaged axial, normal and spanwise velocity, and vorticity contours at the wingtip. Figure 9. Time-averaged axial, normal and spanwise velocity, and vorticity contours at the wingtip. Comparison between various angles of attack. Comparison between various angles of attack.

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Figure 10. Time-averaged axial, normal and spanwise velocity and vorticity contours at the wingtip. FigureComparison 10. Time-averaged between the axial,various normal measurement and spanwise levels. velo velocity city and vorticity contours at the wingtip. Comparison between the variousvarious measurement levels.levels.

FigureFigure 11. 11.Line Line plots plots of of time-averaged time-averaged axial axial velocity velocity at at the the vortex vortex core. core. Figure 11. Line plots of time-averaged axial velocity at the vortex core.

Aerospace 2017, 4, 53 11 of 17 Aerospace 2017,, 4,, 5353 11 of 17

FigureFigure 12. 12.Line LineLine plots plotsplots of ofoftime-averaged time-averagedtime-averagednormal normnormal velocity velocity at atthe thevortex vortex core. core.

FigureFigure 13. 13.Line LineLine plots plotsplots vorticity vorticityvorticityfor forforthe thethethree threethreemeasurement measurementmeasurement planes. planes.planes.

AA directdirect comparison between between the the three three measuremen measurementtt levelslevels levels isis mademade is made inin FiguresFigures in Figures 1414 andand 14 15.15.and MoreMore 15. Morespecifically, specifically, Figure Figure 14 presents 14 presents the thetime-averaged time-averaged U andandU and V components,components,V components, vorticityvorticity vorticity andand and circulationcirculation circulation atat atthethe the vortexvortex vortex corecore core atat at anan an angleangle angle ofof of attackattack attack ofof of 1010 10 deg.deg. deg. ReRe Regardinggarding the the velocities, itit isis evidentevident thatthatthat the thethe normal normalnormal componentcomponent alongalong withwith thethe axialaxial vorticity vorticity havehave aa much much smaller smaller change change in in magnitude magnitude thanthan the the axial axial velocity,velocity, between between the the three three streamwise streamwise locations. locations. Namely, Namely, a a 10% 10% and and 13% 13% change change in inV Vand andandvorticity vorticityvorticity waswas measured, measured, respectively, respectively, from from one one level level to to the the next next one, one, which which is is in in agreement agreement with with the the comments comments mademade inin thethe previousprevious paragraph.paragraph. The vortex vortex core core is is lo locatedcated at at aa radius radius of of 0.07·c, 0.07·c, around around the the center center of ofthethe the vortex,vortex, vortex, thethe the samesame same forfor for allall all threethree three measurementmeasurement measurement planesplanes planes at this at this angle angle of attack. of attack. The The circulation circulation was was also alsocalculated calculated for forcomparison comparison purposes, purposes, as asthe the flux flux of of vorticity, vorticity, through integration.integration. TheThe followingfollowing equationequation was was used: used: Γ = Ω·dS (2) =x S (2)(2) where Ω is the axial vorticity from Equation (1) and S is the surface. Since it was more convenient, where Ω is the axial vorticity from Equation (1) and S is the surface. Since it was more convenient, polarwhere coordinates Ω is the axial were vorticity employed from for Equation this type of(1) chart, and S with is the the surface. starting Since point it of was the more coordinate convenient, point polar coordinates were employed for this type of chart, with the starting point of the coordinate point beingpolar thecoordinates center of were the vortex. employed The typicalfor this three-regiontype of chart, distinction with the star canting be madepoint [of14 the,16 ],coordinate with the slope point being the center of the vortex. The typical three-region distinction can be made [14,16], with the slope inbeing the logarithmicthe center of region the vortex. decreasing The typical as the three-regi vortex getson distinction carried downstream. can be made The [14,16], changes with are the in slope the in the logarithmic region decreasing as the vortex gets carried downstream. The changes are in the samein the order logarithmic of magnitude region with decreasing the ones as at the normal vortex velocity gets carried and vorticity, downstream. that is betweenThe changes 10% are and in 15% the same order of magnitude with the ones at normal velocity and vorticity, that is between 10% and 15% fromsame level order to of level. magnitude with the ones at normal velocity and vorticity, that is between 10% and 15% fromfrom levellevel toto level.level. InIn FigureFigure 15,15, thethe smallersmaller valuevalue ofof axialaxial velocityvelocity insideinside thethe vortexvortex corecore ((U00)) isis shownshown forfor thethe threethree measurement planes at every angle of attack. At each level, a linear interpolation can be made, indicatingindicating thatthat U00 decreasesdecreases asas thethe angleangle ofof attackattack growsgrows at every measurement plane. The magnitude of U00,, though,though, increasesincreases asas thethe planesplanes movemove downstream,downstream, whereaswhereas thethe linearlinear slopeslope dU00//dα decreases.decreases.

Aerospace 2017, 4, 53 12 of 17 Aerospace 2017, 4, 53 12 of 17 Aerospace 2017, 4, 53 12 of 17

Figure 14. Time-averaged axial (a) and normal velocity (b), vorticity (c), and circulation (d), without Figurewinglet. 14.FigureTime-averaged 14. Time-averaged axial (a )axial and ( normala) and normal velocity velocity (b), vorticity (b), vorticity (c), and ( circulationc), and circulation (d), without (d), without winglet. winglet.

Figure 15. Time-averaged axial velocity in the center of the vortex (U0) at the three measurement Figure 15. Time-averaged axial velocity in the center of the vortex (U0) at the three measurement levels, levels,withoutFigure without winglet. 15. Time-averaged winglet. axial velocity in the center of the vortex (U0) at the three measurement levels, without winglet. 3.2. Winglet Installed In Figure 15, the smaller value of axial velocity inside the vortex core (U ) is shown for the 3.2. Winglet Installed 0 threeFigure measurement 16 shows planes the time-averaged at every angle axial of attack. and normal At each velocity level, aand linear axial interpolation vorticity in contour can be made, plots atindicating X/c = Figure1 for that 5, 16U 10,0 showsdecreases and 15 the deg astime-averaged of the angle angle of of attack. attackaxial and The grows normal freestream at every velocity measurementvelocity and axialdirection plane.vorticity is identical The in magnitudecontour to the plots oneof Uat presented0 ,X though,/c = 1 for increasesin 5, Figure 10, and as8. 15 theBased deg planes onof anglethe move results of downstream, attack. of this The study, freestream whereas two theseparate velocity linear swirling slopedirectiondU motions0 /isd identicalα decreases. can tobe the observed,one presented one around in Figure the vortex 8. Based core, on at thethe resultstip of the of winglet,this study, and two another separate around swirling the entire motions winglet can be configuration.3.2.observed, Winglet Installed one The around normal the velocity vortex magnitudecore, at the tipis about of the two winglet, times and larger another inside around the “vortex” the entire region. winglet However,configuration.Figure a 16second shows The region the normal time-averaged can bevelocity clearly magnitude axialobserved and normalaroundis about velocitythe two entire times and winglet larger axial vorticityconfiguration,inside the in “vortex” contour from plots nowregion. onat XHowever,referred/c = 1 forto a as 5,second 10,the and“winglet” region 15 deg can region. ofbe angle clearly In ofterms observed attack. of vo The aroundrticity, freestream the entiredifference velocity winglet isdirection even configuration, greater, is identical since from the tonow vorticitytheon one referred presented maximum to as in insidethe Figure “winglet” the8. Basedvortex region. on core the Inis results termsat least ofof three thisvorticity, study, times the twoas differencehigh, separate compared swirlingis even to greater, motionsthe vorticity since can the valuesbevorticity observed, of the maximum “winglet” one around insideregion. the the vortex Hence, vortex core, the core “winglet at theis at tip le” ast ofregion thethree winglet,indicates times as and that high, anotheranother compared aroundvortical to thethestructure entirevorticity wingletexists,values which configuration. of the may “winglet” be far The weaker region. normal inHence, velocity terms the of magnitude “wingletmagnitude” isregion aboutbut covers indicates twotimes a larger that larger another area, inside since vortical the it “vortex”spreads structure exists, which may be far weaker in terms of magnitude but covers a larger area, since it spreads

Aerospace 2017, 4, 53 13 of 17

Aerospace 2017 4 approximately, , 53 all over the winglet wake region. Furthermore, both V and vorticity magnitude13 of 17 present very small differences between 10 and 15 deg, indicating that the lift generation has reached region.its peak, However, resembling a secondthe corresponding region can betrends clearly without observed the winglet around installed the entire at wingletthe tip. Regarding configuration, the fromaxial nowvelocity, on referred and in a to similar as the “winglet”fashion to region.the regu Inlar terms wingtip of vorticity, case, the the wake difference can be divided is even greater, at two sinceseparate the regions, vorticity i.e., maximum the “blunt inside body the wake” vortex region core and is at the least “vortex” three timesones. asThe high, axial compared velocity in to both the vorticityregions is values reduced of thein magnitude “winglet” as region. the angle Hence, of attack the “winglet” increases. region indicates that another vortical structureMoreover, exists, a whichcomparison may be of farthe weakeraforementioned in terms vari of magnitudeables inside but the covers vortex a is larger made area, in Figure since 17 it spreadsfor all of approximately the examined allangles over of the attack. winglet The wake streamwise region. Furthermore, and normal time-averaged both V and vorticity velocity magnitude and the presentaxial vorticity very small line differencesplots provide between more 10 detail and 15about deg, the indicating development that the of lift these generation quantities has reachedinside the its peak,vortex resembling at the tip of the the corresponding winglet and can trends be used without in order the winglet for a comparison installed at with the tip. computational Regarding the data axial to velocity,be made. and It should in a similar be noted fashion that to the the peak regular values wingtip of the case, normal the wake velocity canbe component divided at and two the separate axial regions,vorticity i.e.,are the considerably “blunt body smaller wake” regionthan the and ones the “vortex” presented ones. in The Figures axial velocity11–13. A in more bothregions detailed is reducedcomparison in magnitude between the as two the anglecases ofis made attack in increases. Section 3.3.

Figure 16. Time-averagedTime-averaged axial axial velocity, velocity, normal normal velocity, velocity, and and vorticity contours at at the wingtip, with winglet.

Moreover, a comparison of the aforementioned variables inside the vortex is made in Figure 17 for all of the examined angles of attack. The streamwise and normal time-averaged velocity and the axial vorticity line plots provide more detail about the development of these quantities inside the vortex at the tip of the winglet and can be used in order for a comparison with computational data to be made. It should be noted that the peak values of the normal velocity component and the axial vorticity are considerably smaller than the ones presented in Figures 11–13. A more detailed comparison between the two cases is made in Section 3.3.

Figure 17. Time-averaged axial velocity (a), time-averaged normal velocity (b), and vorticity (c), with winglet.

Aerospace 2017, 4, 53 13 of 17

approximately all over the winglet wake region. Furthermore, both V and vorticity magnitude present very small differences between 10 and 15 deg, indicating that the lift generation has reached its peak, resembling the corresponding trends without the winglet installed at the tip. Regarding the axial velocity, and in a similar fashion to the regular wingtip case, the wake can be divided at two separate regions, i.e., the “blunt body wake” region and the “vortex” ones. The axial velocity in both regions is reduced in magnitude as the angle of attack increases. Moreover, a comparison of the aforementioned variables inside the vortex is made in Figure 17 for all of the examined angles of attack. The streamwise and normal time-averaged velocity and the axial vorticity line plots provide more detail about the development of these quantities inside the vortex at the tip of the winglet and can be used in order for a comparison with computational data to be made. It should be noted that the peak values of the normal velocity component and the axial vorticity are considerably smaller than the ones presented in Figures 11–13. A more detailed comparison between the two cases is made in Section 3.3.

AerospaceFigure2017, 416., 53 Time-averaged axial velocity, normal velocity, and vorticity contours at the wingtip,14 of 17 with winglet.

FigureFigure 17.17. Time-averagedTime-averaged axialaxial velocityvelocity ((aa),), time-averagedtime-averaged normalnormal velocityvelocity ((bb),), andand vorticityvorticity ((cc),), withwith winglet. winglet. Aerospace 2017, 4, 53 14 of 17

3.3. Comparison Having measuredmeasured and and analyzed analyzed the the flowfield flowfield around around both withoutboth without and with and the with winglet the installed,winglet ainstalled, direct comparison a direct comparison between the between two cases the can two be case made.s can More be made. specifically, More asspecifically, is evident fromas is theevident two previousfrom the sections,two previous there sections, is a considerable there is a decrease considerable in the decrease size and in magnitude the size and of the magnitude wingtip vortex,of the whenwingtip the vortex, winglet when device the is installedwinglet atdevice the tip. is installe This decreased at the can tip. be This qualitatively decrease observed can be qualitatively in Figure 18, whereobserved a top in Figure view of 18, the where smoke-probe a top view visualization of the smoke-probe experiments visualization is presented experiments at the tip is of presented the wing, withat the and tip withoutof the wing, the wingletwith and installed, without attheX /wingletc = 1 and installed, 10 deg ofat X angle/c = 1 of and attack. 10 deg of angle of attack.

Figure 18. Smoke-probe visualization studiesstudies at the wingtip, without (leftleft)) andand withwith wingletwinglet ((rightright).).

This comparison is also made in Figure 19 using LDA contour plots for time-averaged axial This comparison is also made in Figure 19 using LDA contour plots for time-averaged axial velocity and axial vorticity and in Figure 20 using line plots. The angle of attack of 10 deg at velocity and axial vorticity and in Figure 20 using line plots. The angle of attack of 10 deg at X/c = 1 X/c = 1 was chosen as a reference point, since at that angle, the vortex appears to be strong and well- was chosen as a reference point, since at that angle, the vortex appears to be strong and well-defined defined for both cases. In Figure 20a, the time-averaged axial velocity is shown, followed by the time- for both cases. In Figure 20a, the time-averaged axial velocity is shown, followed by the time-averaged averaged normal velocity component in Figure 20b. It is clear that, even though the axial component normal velocity component in Figure 20b. It is clear that, even though the axial component is of the is of the same order of magnitude between the two cases, with a percentage difference smaller than same order of magnitude between the two cases, with a percentage difference smaller than 5%, the swirl 5%, the swirl component peak values are reduced with the addition of the winglet by a factor of 3. component peak values are reduced with the addition of the winglet by a factor of 3. That can also be That can also be seen in Figure 20c, where the peak axial vorticity with winglet equals to 35.5% of the seen in Figure 20c, where the peak axial vorticity with winglet equals to 35.5% of the corresponding corresponding peak vorticity without winglet. In other words, the addition of the winglet seems to peakreshape vorticity the wake without and tip winglet. vortex, In causing other words, the swirl the velocity addition components of the winglet and seems peak tovorticity reshape magnitude the wake andinside tip the vortex, vortex causing core to the reduce. swirlvelocity Finally, componentsin Figure 20d and the peak slope vorticity of the curve magnitude inside inside the logarithmic the vortex coreregion to is reduce. reduced Finally, to 28% in with Figure the 20 additiond the slope of th ofe thewinglet, curve whereas inside the the logarithmic maximum regionvalue is is almost reduced five to 28%times with smaller. the addition of the winglet, whereas the maximum value is almost five times smaller.

Figure 19. Time averaged axial-velocity (top) and axial vorticity (bottom) comparison without and with winglet (α = 10°, X/c = 1).

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3.3. Comparison Having measured and analyzed the flowfield around both without and with the winglet installed, a direct comparison between the two cases can be made. More specifically, as is evident from the two previous sections, there is a considerable decrease in the size and magnitude of the wingtip vortex, when the winglet device is installed at the tip. This decrease can be qualitatively observed in Figure 18, where a top view of the smoke-probe visualization experiments is presented at the tip of the wing, with and without the winglet installed, at X/c = 1 and 10 deg of angle of attack.

Figure 18. Smoke-probe visualization studies at the wingtip, without (left) and with winglet (right).

This comparison is also made in Figure 19 using LDA contour plots for time-averaged axial velocity and axial vorticity and in Figure 20 using line plots. The angle of attack of 10 deg at X/c = 1 was chosen as a reference point, since at that angle, the vortex appears to be strong and well- defined for both cases. In Figure 20a, the time-averaged axial velocity is shown, followed by the time- averaged normal velocity component in Figure 20b. It is clear that, even though the axial component is of the same order of magnitude between the two cases, with a percentage difference smaller than 5%, the swirl component peak values are reduced with the addition of the winglet by a factor of 3. That can also be seen in Figure 20c, where the peak axial vorticity with winglet equals to 35.5% of the corresponding peak vorticity without winglet. In other words, the addition of the winglet seems to reshape the wake and tip vortex, causing the swirl velocity components and peak vorticity magnitude inside the vortex core to reduce. Finally, in Figure 20d the slope of the curve inside the logarithmic region is reduced to 28% with the addition of the winglet, whereas the maximum value is almost five Aerospacetimes smaller.2017, 4, 53 15 of 17

AerospaceFigureFigure 2017, 19. 419., 53Time Time averaged averaged axial-velocity axial-velocity( top(top)) and and axial axial vorticity vorticity ( bottom(bottom)) comparison comparison without without and and15 of 17 withwith winglet winglet ( α(α= = 10 10°,◦, X X//c = 1).

FigureFigure 20. Time-averagedTime-averaged axial axial ( (aa)) and and normal normal velocity velocity ( (bb),), vorticity vorticity ( (cc)) and and circulation circulation ( (dd),), with with and and without winglet.

4.4. Conclusions Conclusions TheThe near near flow flow field field around around a a UAV model was ex experimentallyperimentally investigated by by means means of of flow flow visualization and 3D LDA point measurements. The The experiments experiments were were carried out at a a wide wide range range of of anglesangles of of attack to ensure that all the operational conditions are examined and emphasis was given on the flow at the wingtip area. Two configurations, one without and one with winglets installed, were used. The flow visualization studies were conducted using the smoke-probe, smoke-wire, and laser sheet methods. The LDA measurements yielded the three time-averaged velocities, whereas the axial vorticity and circulation quantities were in turn calculated using the experimental data. Thus, a complete experimental investigation of the wingtip vortex of a UAV could be made, whereas a direct comparison of the flowfield without and with a winglet installed is made, serving as the main contribution of this study. Considering the flow at the wingtip, a comparison at various angles of attack for three locations downstream of the model was made. The tip vortex does not appear to have a defined structure until 5 deg, whereas as the angle of attack increases the vortex grows in strength, the swirling components increase in magnitude, and the vorticity at the core increases as well. The wake seems to form two distinctive regions: one affected by the presence of the vortex, and another affected by the shear layer separation at the wake of the model. At every angle of attack, an average difference of 10–15% was measured, between the three levels for the time-averaged normal velocity and axial vorticity. In the “winglet installed” case, the trends resemble those of the “winglet removed” case, regarding the changes between the various angles of attack. A second swirling motion, weaker in strength, was also identified, which in contrary to the wingtip vortex exists all over the winglet wake region. Overall, the addition of the winglet reshapes the tip vortex, causing the swirl velocity

Aerospace 2017, 4, 53 16 of 17 on the flow at the wingtip area. Two configurations, one without and one with winglets installed, were used. The flow visualization studies were conducted using the smoke-probe, smoke-wire, and laser sheet methods. The LDA measurements yielded the three time-averaged velocities, whereas the axial vorticity and circulation quantities were in turn calculated using the experimental data. Thus, a complete experimental investigation of the wingtip vortex of a UAV could be made, whereas a direct comparison of the flowfield without and with a winglet installed is made, serving as the main contribution of this study. Considering the flow at the wingtip, a comparison at various angles of attack for three locations downstream of the model was made. The tip vortex does not appear to have a defined structure until 5 deg, whereas as the angle of attack increases the vortex grows in strength, the swirling components increase in magnitude, and the vorticity at the core increases as well. The wake seems to form two distinctive regions: one affected by the presence of the vortex, and another affected by the shear layer separation at the wake of the model. At every angle of attack, an average difference of 10–15% was measured, between the three levels for the time-averaged normal velocity and axial vorticity. In the “winglet installed” case, the trends resemble those of the “winglet removed” case, regarding the changes between the various angles of attack. A second swirling motion, weaker in strength, was also identified, which in contrary to the wingtip vortex exists all over the winglet wake region. Overall, the addition of the winglet reshapes the tip vortex, causing the swirl velocity components and peak vorticity magnitude inside the vortex core to reduce. More specifically, the maximum vorticity magnitude in this study is reduced by a factor of three with the winglet installed. The above observations help to acquire a better idea about the tip vortex of a UAV and about the effect the winglet device has on its dispersion. Concerning future studies, a turbulence budget investigation could also be conducted to provide even more details about the vortex core, whereas a flow visualization using a high-speed camera is suggested in order to acquire more information about the wandering of the vortex and the effect the winglet has on this phenomenon.

Acknowledgments: The work presented in this paper is a part of the 11SYNERGASIA_6_629 “Hellenic Civil Unmanned Air Vehicle—HCUAV” research project, implemented within the framework of the National Strategic Reference Framework (NSRF) and through the Operation Program “Competitiveness & Entrepreneurship—SYNERGASIA 2011”. The research project is co-financed by National and Community Funds, 25% from the Greek Ministry of Education and Religious Affairs-General Secretariat of Research and Technology and 75% from E.U.—European Social Fund. Author Contributions: Pericles Panagiotou conceived, designed and performed the experiments, analyzed the data, and wrote the paper; George Ioannidis and Ioannis Tzivinikos performed the experiments and analyzed the data; Kyros Yakinthos supervised the experimental procedure, analyzed the results, and wrote the paper. Conflicts of Interest: The authors declare no conflict of interest.

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