Conceptual and Preliminary Design of a Long Endurance Electric

UAV

Luıs Miguel Almodovar Paradaluisparada2@hotmail.com

Instituto Superior Tecnico, Universidade de Lisboa, Portugal

November 2016

Abstract

The present thesis documents the conceptual and preliminary design of a solar long enduranceUnmanned Aerial Vehicle (UAV). Having in mind surveillance goals, the mission profile requires aninitial climb, at a rate that allows to ascend 1000 m above runway in 10 minutes, followed by cruisewith an endurance of 8 hours and a range 200 km during the equinox. In the conceptual stage,several aircraft configurations were evaluated considering the mission requirements. Resorting to theAnalytic Hierarchy Process methodology, the rear pusher V-tail airplane concept was chosen throughpairwise comparisons between 10 mission criteria (Aerodynamics, Structures an Weight, Solar PanelsIntegration, Propulsion, Manufacturing and Maintenance, Stability and Control, Payload Volume,Remote-Person View Integration, Take-off and Landing and Portable Capabilities) and 8 prospectivecandidate configurations. A UAV market investigation provided an initial estimate of the empty weightfraction and airframe dimensions. Through bibliographic research, the communications system wasdefined beforehand and the propulsion system also received an initial estimate. In the preliminarydesign phase, the airframe was resized by performing a set of fixed point iterations, whose resultingdesign point ensured it is physically possible to fulfill flight requirements. Allying multiple aerodynamicanalyses, performed on a low Reynolds number computational tool, with the ascertained efficiency ofthe propulsion system, power consumption at each mission stage was determined under ideal weatherconditions. The solar energy received by the photo-voltaic arrays installed was determined in a singlelocation applying a theoretical irradiation model in different seasons of the year. It was verified that inthe March equinox endurance reached 7.5 hours, while in September its maximum increased to almost10 hours. An aircraft CAD assembly was modeled without internal airframe detail. The developedparts include a three-piece 5 meter span wing, a fuselage composed by two connectable pieces, totaling1.9 meter length, and two individual V-tail halves. On-board propulsion and avionics components werealso modeled in a simplified way, as solids of uniform density. A flight stability analysis was performed,resulting in a static margin of 5.6% with null Cm in cruise conditions. Looking at a future detaileddesign, the cruise flight envelope was created and the wing pressure distribution was obtained.

Keywords: Solar UAV, Analytic Hierarchy Process, Design Point, Low Reynolds Number, Endurance,Flight Envelope.

1. Introduction

Usage progress of UAVs for civil applications is notsteadily established because, unlike in the defensesector, where air space is properly prepared and thecompletion of the mission is supra to the economicsof the vehicle, for civilian procedures cost sustain-ability and operational complexity are still majorbarriers. Most existing mission profiles for mod-ern UAVs require a platform, or series of platforms,which clearly extend range and endurance beyondthe capability of existing affordable electric vehicles.

At the same time, the use of renewable energies invehicles is becoming more and more demanding by

society due to its growing environmental awareness.Taking this factors into account, it has become clearin the last years that the key to enhance UAVslong endurance capabilities can be found within so-lar aviation. Nowadays, with the relatively easyaccess to numerical and experimental design toolsand with the current efficiency of electric systems(solar panels, motors and batteries) there is a realchance to expand UAVs applications. The aim ofthis thesis is to achieve the preliminary design of asolar powered UAV, meant to be used for civiliansurveillance missions.

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2. Project Background

This work started as a part of a project of aLong Endurance Electric Unmanned Aircraft Ve-hicle (LEEUAV) that is being developed with thecollaboration of IDMEC (Instituto de EngenhariaMecanica), INEGI (Instituto de Ciencia e Inovacaoem Engenharia Mecanica e Engenharia Industrial)and AeroG (Aeronautics and Astronautics ResearchCenter). All units involved are a part of LaboratorioAssociado de Energia, Transportes e Aeronautica(LAETA). The main goal of the project is to de-velop a low cost, small footprint electric UAV, easyto build and maintain, that can be deployed fromshort airfields to perform long endurance surveil-lance missions [5]. To remain airborne for ex-tended periods of time the aircraft must harborgreen power technologies, such as an electric propul-sion system with solar power.

2.1. Accomplished Tasks

To assist the complete design of the LEEUAV, thereare several tasks from different areas that have beenaddressed beforehand.

Two master students, Hector Vidales and TiagoFerreira, have poured over the propulsion systemdesign for the LEEUAV, as part of their publisheddissertations [9, 3]. Different electric configurationshave been evaluated in terms of performance, over-all weight and cost. The selection of solar panelswas followed by the specification of possible brush-less motors. A hybrid propulsion system ended upbeing established due to the necessity of a secondaryenergy source, in the form of high-density recharge-able batteries, to power climb flight.

A first generation prototype has been built un-der a master’s thesis from UBI written by LuısCandido [2], using advanced model building tech-niques. Construction tests allowed to reach rele-vant conclusions related to applicable materials andmanufacturing processes. Flight tests were also per-formed in radio controlled mode. In spite of notharboring the solar energy collection system and allavionics components, the aircraft allowed to obtainimportant qualitative performance knowledge for aprospective solar UAV configuration.

Regarding aerodynamics, a computational pro-cess enabling high fidelity fluid analysis of differentUAV configurations has been developed by NunoSilva, as part of his published master’s dissertation[8]. As a case study, the aerodynamic behavior ofthe first generation built prototype, a conventionalconfiguration, was evaluated for cruise conditions.

In the field of communications and electronics,the Autopilot and Remote Person View (RPV)hardware and software design have been addressedin two published master thesis, written by DuarteFigueiredo and Pedro Miller, respectively [4, 6].

3. LEEUAV ConceptIn the Conceptual Design different configurationsare evaluated to meet or exceed the mission require-ments in terms of endurance, size and cost.

3.1. Project RequirementsThe particular requirements of the LEEUAVproject were deliberated assuming that atmosphericparameters vary according to the InternationalStandard Atmosphere model. The mission startswith take-off followed by climb until a maximumsteady flight altitude of 2000 m. Climb occurs ata minimum rate of 1.667 m/s during 10 minutes,which allows to ascend 1000 m. Bearing in mindthat flying at 2000 m over most Portuguese loca-tions would not allow the RPV camera to recordground footage with acceptable resolution display,the project service ceiling was fixed at 1300 m.

Figure 1: Scheme of the LEEUAV mission.

Cruise is the longest phase of the mission, repre-sented in Figure 1, in which there must be a timewindow of 8 hours during the equinox to fly overspecified locations and to turn around, reaching arange of 200 km after returning to base. Consider-ing both range and time constraints, the project’scruise speed selected is 7.53 m/s. Occasionally, loi-ter may be needed before landing. The minimumbank angle for the aircraft’s turn is 45o, with anassociated load factor of 1.414.

The LEEUAV must also be able to carry a pay-load of up to 10 N, this includes all components ofthe communications systems and extra equipment,such as measurement sensors. The aircraft’s dimen-sions must allow operations in very small airfields(eventually no airfield will be required) and it hasto possess portable capabilities suitable for smallvehicle ground transportation.

3.2. Conceptual Design SelectionA total of 8 prospective aircraft configurations, il-lustrated in Figure 2, have been selected attendingto the mission profile and project background.

The airframe alternatives were chosen with thepurpose of evaluating conjugated effects of featuresthat did not present clear reasons to be excluded be-forehand. The quadcopter is an exception that was

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chosen to demonstrate how the Analytic HierarchyProcess precludes an evident poor solution.

a: Quadcopter b: Conventional tractor

c: T-tail Tractor d: T-tail Fuselage pusher

e: V-tail Rear pusher f: Double boom pusher

g: High tail double boomtractor

h: Low tail double boomtractor

Figure 2: Selected airframe alternatives.

3.3. Analytic Hierarchy Process Based DecisionThe selection of the aircraft configuration was car-ried out using an Analytic Hierarchy Process (AHP)[7]. With this decision making methodology, a set ofcriteria and alternatives underwent several pairwisecomparisons in order to select the most suitableUAV configuration for the mission requirements.

On the AHP method pairwise comparisons aremade using a scale that consists of qualitative judg-ments ranging from equal to extreme with corre-sponding numerical values (1 = equal , 3 = moder-ate, 5 = strong, 7 = very strong, 9 = extreme im-portance). A pairwise comparison matrix A was ini-tially filled with the numerical judgments for crite-ria, respecting the reciprocal property (aij = 1/aji).The priorities of the elements were estimated byfinding the principal eigenvector of matrix A, AQ =γmaxQ, where γmax is the maximum eigenvalue ofmatrix A. When vector Q is normalized, it becomesthe priority vector. While building a pairwise com-parison matrix, it is necessary to check consistencyby calculating consistency ratio (CR) as the ratio ofConsistency Index (CI) with Random Index (RI),a tabulated value. Consistency Index is calculatedby CI = (γmax − d)/(d− 1) , where d is the squarematrix dimension. Inconsistencies are tolerable anda reliable result may be expected if CR < 0.1.

The global priority pie chart division (Figure 3)

reveals three major criteria, Aerodynamics, Struc-tures and Weight and Solar Panels Integration,which affect range the most.

Figure 3: Global Priority Vector Results.

The same eigenvalue computing process was usedto determine the comparison matrices of the can-didate configurations with respect to the criteriaand their local priorities. All ten pairwise compar-ison matrices of alternatives, as well as the globalpriority data matrix with respect to criteria, havebeen placed in the main document of this work.The consistency ratio always remained acceptable(CR < 0.1).

Figure 4: AHP Final results (the bar numberingmatches the configurations in figure 2).

The final priority vector was obtained by multi-plying the matrix of local priorities with the globalpriority vector , providing the raking of alternatives,illustrated in Figure 4. After adding up all scores,the overall winner of the AHP was the rear pusherwith V-tail (Figure 2e) that got a preference of 17%,surpassing the second classified, the T-tail tractor,by a margin of 1%. The decisive factor was aero-dynamics, that was the only criteria where the rearpusher excelled over the remaining alternatives.

4. Aircraft Initial SizingWith the concept decided, a first estimation of theLEEUAV’s dimensions and weight was carried out.

4.0.1 Empty Weight Estimation

A market research was conducted to find emptyweight fractions of low wingspan UAVs with dimen-

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sions similar to the first generation prototype. Theresearch focused mainly on model gliders (with elec-tric propulsion or not), due to their high aspect ra-tio and absence of landing gear.

Within the arranged database, three major wingconstruction types were identified: Foam CoredConstruction, Molded Construction and Rib Con-struction.

Figure 5: Empty weight as function of wing areaobtained with model glider scattered data.

The scatter chart in Figure 5 displays the ap-pointed empty weight and wing area of the re-searched aircraft. It is noticeable that models withrib conceived wings tend to be lighter for the samewing area. For that reason, the rib constructiontype should be selected for the wings manufacture,as it had been previously for the first generationLEEUAV.

In spite of the observable gap between values, itwas possible to interpolate an expression relatingempty weight with wing area, while keeping a faircorrelation coefficient.

mairframe = 512.99 e0.01S (1)

Intelligibly, Equation (1) was obtained using solelypoints that correspond to rib constructed wings.

The few solar planes found also have rib con-structed wings, but none was used as an interpo-lation point because their weight values were con-sidered too optimistic. The lack of further exampleswith wing dimensions superior to that of the firstLEEUAV reduces the calculation accuracy, sincethe airframe weight is expected to fall outside therange of interpolation.

4.1. LEEUAV Provisory General Characteristics

The first generation LEEUAV was a guideline toassign general dimensions beforehand. It allowedto determine that there would be wing tip portion

without solar arrays, solely meant for aileron con-trol, while on the remaining wing surface, solar ar-rays would fill as much area as possible. The V-tailwas initially sized as an equivalent conversion fromthe first generation’s conventional tail. Its planformarea and dihedral angle were calculated under theassumption that V-tail’s aspect ratio equals that ofits conventional counterpart’s horizontal tail.

36.9%

46%

6.1%

11%

Hybrid propulsion

Airframe

Extra payload

?Communications and control

Figure 6: LEEUAV’s relative mass distribution.

Altogether, with the mass of all on-board com-ponents and airframe, an estimate of the completeweight distribution was obtained in advance. Theweight proportions present in Figure 6 are con-verged values that resulted from the completion ofthe preliminary project. The airframe percentageincludes not only the wing and V-tail but also thefuselage. The hybrid proportion system covers allcomponents directly relate with thrust generation(propeller, motor, electronic speed controller), en-ergy storage (batteries) and energy collection (solarpanels and Maximum Power Point Tracker). Asfor the smaller slice, the communications systems,it comprises the video system, the autopilot andall electronic components associated. It should benoted that the sum of the extra payload weight withcommunication components yields a value close tothe maximum payload of 10 N specified on the mis-sion requirements.

5. Preliminary ProjectThe preliminary project aims to define the overalldimensions of the aircraft, as well as the propulsionsystem and general placement of all components.

5.1. AerodynamicsTo enable a swift processing of aerodynamic results,an open source software called XFLR5 was cho-sen. Its capabilities include 2D airfoil direct anal-ysis as well as 3D wing analysis based on severalmethods at choice including the Lifting Line The-ory (LLT), the 3D Panel Method and the Vortex

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Lattice Method (VLM).

5.1.1 Aerodynamic Software Accuracy

Unlike Panel Method and LLT, VLM is capable ofproducing results for wing and tail geometries withlow aspect ratio and high dihedral. Therefore, itwas the chosen 3D analysis method. To verify theaccuracy of VLM, XFLR5 was employed with air-flow characteristics and lifting geometries similar tothe ones used in the high fidelity CFD analysis madefor the first generation LEEUAV in steady flight[8].The results of both software were compared after-wards.

Figure 7: Geometry considered in XFLR5.

In all simulations computed on XFLR5, bothwing and tail are the only aerodynamic geometriespresent (figure 7). Incorporating a fuselage is notrecommended by software developers, as the resultsof its aerodynamic influence are generally not verysatisfactory. Still, XFLR5 results are known for be-ing reasonably accurate for model sailplanes oper-ating at low Reynolds Numbers.

This software allows the user to control the den-sity of the computational grid, always with rectan-gular divisions, within a certain range of panel el-ements. Three grids were generated for this study,having 822, 1600 and 3968 elements.

Figure 8a shows that the lift predicted in XFLR5is rather accurate, when compared with STAR-CCM+ R©, for angles of attack between -5o and 2o.On the other hand, figure 8b confirms that XFLR5underpredicts global drag coefficient real values.This outcome was expected since XFLR5 utilizesinformation within previously XFoil-generated po-lars, at several Reynolds number, to extrapolate the3D viscous drag component. For this same reason,the airplane performance outputs do not include re-sults close to stall angles. In this analysis the com-puted angle of attack has been limited to values be-tween -5o to 5.5o. A closer inspection on the outputfiles revealed that shear and pressure components ofdrag were underpredicted by VLM.

Overall, it was observed that a grid with 1600elements provides the best compromise betweenaccuracy and computation time. Therefore, forthose specific analysis results, in Figure 8, correc-tion functions were calculated, assuming the formZx(α) = Cx(α)STAR−CCM+/Cx(α)XFLR5, where x

(a) CL as function of α.

(b) CD as function of α.

Figure 8: Original LEEUAV cruise performancecharts (U=7.53 m/s) [8].

is either lift (L) or drag (D). To recoup for XFLR5output discrepancies, Zx functions were applied toall XFLR5 3D analysis performed. The practicalpurpose of this correction is achieved by improvingthe XFLR5 output margin of error while rearrang-ing lift and drag predictions, for angles of attackbetween -5o and 5.5o, more conservatively.

5.2. Cruise Stage Computation

The aerodynamic geometries (wing and tail) of theLEEUAV conjectured on the initial sizing wenttrough VLM computational analysis in order to as-sure that in steady flight the maximum take-offweight (MTOW) breaks even with the lift produced.

At the beginning, the geometry update focusedsolely on the wing. As adjustments were beingmade and reinserted in XFLR5, the airframe weightwas recalculated with Equation (1) and the wirelength linking servos updated as well. The wholeprocess, illustrated by the flowchart in figure 9, wasreiterated until MTOW equaled the lift computedaerodynamically.

Chordwise there were not any changes on thewing, because increasing the mean chord would ad-mittedly enlarge the existing gap between MTOWand produced lift at cruise speed. Conversely, rais-ing the wingspan beyond the mark of five meterscould narrow down that difference, however, thatpossibility was not embraced since it would add fur-ther structural and transportation issues.

Instead, the approach followed was to modify thewing’s airfoil. Formerly it had a curvature of 5.3%,which was renewed to 7.5% so that its lift coeffi-cient raised. Envisioning difficulties related to the

5

Initial stateLEEUAV

-

- ∆mwiring

-Equation(1)

?MTOW

@@R

Lift -VLM Analysis

XFLR5 3D -

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6

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- Air data- Ucruise

6

GeometryUpdate6

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Figure 9: Initial Aerodynamic iterative scheme forcruise stage.

Figure 10: Updated wing airfoil of the LEEUAV,t/c = 13%.

prototype’s construction, the airfoil relative thick-ness was increased to 13%, as seen in figure 10.

Initially, the wing had an constant incident angleof attack of 5o, which was modified with washout.It has been proved, in [8], that, at higher angles ofattack, this feature may reduce flow separation inthe wing region where the ailerons are located. Thetorsion applied diminished the angle of attack in 4o,starting on the aileron sections, which resulted in awing angle of attack of 1o on the tip and 5o on thesections with constant chord.

The definitive lift and drag data, computed atcruise velocity, is shown in figure 11 for all simu-lated angles of attack. A total of twenty two anglesof attack were simulated from -5o to 5.5o, with in-crements of 0.5o on XFLR5, using VLM method.

In steady flight (αairplane = 0o, αwing = 5o), thelift coefficient is rounded to 1.153 . Due to computa-tional limitations mentioned in Section 5.1.1 it wasnot possible to determine both maximum lift andminimum drag coefficients, so those values were es-timated, on a conservative way, whenever necessary.To maximize the airplane’s range, the aircraft mustfly at at angles of attack close to 0o.

5.3. Stability and Control

To support airworthiness towards long endurancemissions at low speed, natural longitudinal stabilitywas targeted by fulfilling two design prerequisites.Assuming a longitudinal axis with origin upstreamand pointing towards the tail, the center of gravity(CG) was placed behind of the neutral point (NP).Which means the static marginKn was assured pos-

(a) CL as function of AoA.

(b) CL/CD as function of CL.

Figure 11: LEEUAV aerodynamic performancecharts (U=7.53 m/s).

itive, that is

Kn =xNP − xCGMAC

> 0 (2)

, where xNP and xCG are the longitudinal positionsof the neutral point and center of gravity, respec-tively, and MAC is the mean aerodynamic chord.

At the same time, the V-tail was arranged tocounteract the moment that comes from the wing,so that the overall moment coefficient is kept closeto null in steady flight conditions.

Figure 12: LEEUAV assembly drawn in CAD soft-ware SolidWorks R©

Resorting to CAD software SolidWorks, the sec-ond generation LEEUAV and the majority of on-board components were modeled, as shown in figure12, and the respective mass centers positioning pin-pointed from there. Then, the airplane’s updatedinertia was transferred to XFLR5, where compo-nents and parts, excluding wing and tail-plane, werereplaced by punctual masses.

The VLM analysis loop was resumed with incre-mental modifications in the weight distribution andV-tail characteristics until the results indicated thatthe longitudinal stability prerequisites were accom-plished. Figure 13 illustrates the whole procedure,which also includes the previous aerodynamic loopfor cruise (figure 9) to attest that lift keeps match-ing with flying weight.

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(Solidworks R©)hypothesis

GravityCenter of - XFLR5 VLM analysis

?��

��Kn > 0 ?

CM (α = 0) ' 0 ?

��@@R����No

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cycle (fig 9)Aerodynamic

6

6

→ mwiring, mactuators

→ Tail geometry and/or trim→ Weight distributionUpdate:

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Figure 13: Static stability iterative scheme.

In order to nullify the moment coefficient atcruise AoA of 0o, the V-tail, featuring a modifiedNACA0008 with 1mm trailing edge thickness, wastrimmed with an incidence negative angle of 3o withrespect to the fuselage.

The static margin was calculated with the follow-ing equation

CMα= −KnCLα <=> Kn =

CMα

CLα. (3)

Where CLα and CMα correspond to the derivativesof the lift (CL) and moment (CM ) coefficients withrespect to the angle of attack.

In steady flight it is observed that Kn = 5.6%,corresponding to, according to a referential with ori-gin in the leading edge, a CG position at 38.86% ofthe MAC, which was considered acceptable.

Incidentally, a flaw that is highlighted from thisstability study approach has to do with the absenceof fuselage weight modifications following dimen-sion updates.

5.4. PropulsionShort-term climbing flight was the first to be exam-ined. Attending to its short duration with respectto the whole mission, it is always assumed that thesolar energy received during that stage is negligible.

At this point, while climb drag is determined witha computational VLM simulation at the respectivevelocity Uclimb, thrust is computed as as Tclimb =Wsin(ϕ) +D, where ϕ is the climb angle and W isthe aircraft weight.

The raw power required in every part of themission was determined knowing the efficiency ofthe propulsive system components. The propeller’swas studied with more detail since it is the onlycomponent whose performance is directly affectedby aerodynamics. Computed performance data,supplied by a manufacturer, provided estimates ofthrust, torque and propeller efficiency over a broadrange of model speeds and motor revolutions per

minute(RPM). All values of any parameter con-tained within a certain propeller data file can berepresented with a single three dimensional graphicas function of two other listed parameters.

Figure 14: Efficiency graphical calculation of the15” x 8” propeller with climb regime data.

The propeller efficiency computation is illus-trated in Figure 14, where the manufacturer datais the curved surface and the climb regime input isdefined by a plane of constant thrust together witha perpendicular plane that assumes constant air-craft velocity. The resulting intersection from thosethree surfaces yields a point in space that corre-sponds to the propeller efficiency for the respectiveflight regime.

ηESC ηMotor ηPropeller ηprop

Climb 95 % 89 % 44.8 % 37.9 %

Cruise 95 % 89 % 66.1 % 55.9 %

Table 1: Propulsion sub-system overall efficiency.

Although this study was not extensive, it allowedto select the 15” x 8” composite propeller model,whose efficiency surpasses, in both climb and cruise,all other propellers considered so far. Based on pre-vious propulsion studies [9], a moderate efficiencyvalue of 89% is attributed to the motor. Data re-garding the propulsion subsystem efficiency is pre-sented in Table 1, where the ESC, battery and ca-bles receive typical ratings and ηprop is the totalnetwork efficiency.

The flowchart in Figure 16 sums up how propul-sion and energy studies referring to climb are inte-grated in the design iterative process. In the eventof any flight predicament the motor must be ableto provide considerably more than enough power toclimb, that is Pmotor >> Pprop. It was imposed thatthe battery maximum storage (Ebattery) exceeds theenergy needed for the stipulated climb (Eclimb) inat least 20%. Having this constraint in mind, it wasverified that two 3S LiPo batteries, with 4200 mAheach, connected in parallel contain enough energyfor a 10 minute climb.

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Figure 15: Three view drawing of the LEEUAV, dimensions in mm.

�

(Advanced Iterations)→Stability: fig.13

(Early Iterations)→Aerodynamics: fig.9Previous Cycles

?

→ Thrust→ Drag -

ηprop@@R

InterpolationPropellerMATLAB

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Hypothesis

System→ Propulsive

→ Battery

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Pprop = T.Ustage/ηprop HHHHHj

?Epropclimb = Pprop.tclimb

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�� � Pmotor >> Pprop ?�� � 1.2 · Eclimb ≤ Ebattery ? -

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Figure 16: Propulsion and climb energy iterativeflowchart.

5.5. Flight EnvelopeThe operational framework in relation to air speedand load factor, consisting on the maneuvering andgust envelopes combined in Figure 17, defines thelimits for a balanced detailed design of internalstructures. The current choice for the load fac-tor upper and lower bounds has been supported byFederal Aviation Regulations (FAR) Part 23 andan available flight envelope of small UAVs with aresembling size, weight and maneuverability [1].

Two main additional scenarios are presented, atthe dive speed the aircraft structure cannot with-stand gusts that surpass 2.58 m/s, nevertheless, if

Figure 17: Combined V-n diagram of the LEEUAV,h=1300 m.

velocity does not surpass the maneuvering speed theadmissible gust raises to 5.02 m/s.

5.6. Design Point

In this work, the design point is intended to indi-cate the working condition at which the motor canbe adapted to its use in the UAV to match the en-durance requirement. Therefore, cruise was chosenas the representative situation of the design point.

Observing Table 2, it is clear that the represen-tative design point does not allow the LEEUAV tofulfill climb requirements, thereupon the climb per-formance specifications have been defined as an off-design switch point. The final overall dimensionsof the converged aircraft are presented in Figure15. Upstream the wing, the fuselage was resized toharbor all components, including the RPV cameraupfront. The ailerons and ruddervators were not

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drawn since their conceiving involves assembly so-lutions not addressed at this stage of the project.

W/S 35.26 N/m2

(T/W)climb 0.34

(P/W)climb 6.96 W/N

(T/W)cruise 0.07

(P/W)cruise 1.04 W/N

Table 2: Performance specifications.

6. Solar Energy ManagementThe design point is highly influenced by the solarenergy collection sub-system response to the en-durance requirement, that was over-leaped in theiterative process. The efficiency of the photo-voltaiccells (ηpanels) and Maximum Power Point Tracker(ηMPPT ), shown in table 3, was estimated based onmanufacturer data and experimental tests carried in[3].

ηpanels ηcamber ηMPPT ηenergy

20 % 90 % 92% 16.7 %

Table 3: Energy collection sub-system efficiency.

Additionally, an efficiency factor of 90%, ηcamber,related to the panels camber, was added to copewith the effect the cells with smaller elevation anglehave on the whole group of arrays.

A daily irradiation model, explained in more de-tail in the main thesis document, was arranged foreach day of the year in any specified location. Theconstant set of geographical coordinates chosen cor-respond to Pampilhosa da Serra, that also served asreference to decide the cruise altitude of 1300 m.

By applying the overall efficiency of the solar en-ergy collection system, ηenergy, to the irradianceprofile, the solar power, Psolar, harnessed throughsolar panels area, Asolar, was obtained. There-after, knowing the battery energy at the begin-ning of cruise, Ebattery, and the power consump-tion, Pcruise, flight endurance, ∆tcruise, could becomputed. The flowchart in Figure 18 illustrateshow solar energy calculations have been integratedon the iterative design process.

Both spring and autumnal equinox dates were in-serted in the loop and since it is not specified forwhich one the endurance requirement must be ful-filled, the design is considered converged if cruisesurpasses the 8 hour mark in at least one of them.In the advent of not being able to accomplish theendurance requirement, the airplane design wouldhave to be completely reiterated in order to increasethe wing area available for solar panels. Fortu-itously, the solar energy management loop satisfiedthe endurance imposition at the first attempt.

LocationFlight -

ImplementationMATLAB r.sun�

DateEquinox

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Ee[W/m2/hour]

Daily Irradiation Model

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→ MPPT→ Asolar?

Psolar = Ee.Asolar.ηenergy

Network daily solar power

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Cruise endurance-

6

(Fig. 16)energy consumption

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→Stability: fig.13 (Advanced Iterations)→Aerodynamics: fig.9 (Early Iterations)

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Figure 18: Cruise energy management iterativeflowchart.

Figure 19 presents the daily solar net power avail-able in both equinoxes plotted in MATLAB againstthe power required to cruise. The number of 3SLiPo batteries considered in data treatment rangesfrom two, that is the standard airborne amount, tothree, assuming that the weight of the aircraft doesnot change.

(a) March equinox. (b) September equinox

Figure 19: Power required for steady cruise flightand daily solar net power available during equinoxesin Pampilhosa da Serra (h=1300 m)

On the March equinox, it was established thatcruise would start at the hour of the day in whichsolar power equals the minimum required for steadyflight, such instant corresponds to the first inter-section between line plots in Figure 19a. In the

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long run, all exceeding solar energy can be storedwith 2 3S LiPo batteries and used afterwards dur-ing the period of solar energy shortage, which leadsto an endurance of 7 hours and 34 minutes. TheLEEUAV only reaches the 8 hour mark with a thirdbattery in line that is not used for solar energy stor-age but allows to extend endurance throughout theperiod of solar energy deficit.

During the September equinox, there is an higheramount of solar power when compared with theMarch equinox, explained by an increase on the an-gle of incidence of solar beam irradiation. Althoughthe registered surplus can be contained within 2empty 3S LiPo batteries, the energy remaining inthe storage system after climb would reduce thestorage volume available. To prevent extra en-ergy losses, the cruise starting hour was delayedso that when solar power alone allows steady flightthe whole storage system is virtually empty. Hence,cruise is initially powered by batteries along with in-creasing solar power, which maximizes energy stor-age afterwards. With this, steady flight is prolongedto 9 hours and 58 minutes.

7. ConclusionsIn this work the project of a long endurance solarairplane has been carried out, from the conceptualdesign to the preliminary sizing. Although struc-tural design did not fall under the scope of thisthesis, the pressure coefficient wing distribution ob-tained at this point in XFLR5 can be transferredto a future Finite Element Method (FEM) analysisperformed in prospective wing structure proposals.The purpose of the structural project on the overalliterative design procedure is to replace the airframeestimation model, that was given in the initial siz-ing by Equation (1). Supposing that the detaileddesign imposes an airframe weight that surpassesthe estimate, there is still an error margin assuredby the extra payload, visible in Figure 6. Thus,the design point of the LEEUAV is not expected tosuffer reiterations due to the detailed design.

As soon as the structural project complete, ad-vanced construction methodologies will have to befound and applied. Experimental techniques canalso be arranged to verify the accuracy of FEMcomputational analysis and to study the residualresistance of the prototype in the presence of de-fects. At the same time, the take-off system mustbe engineered. With the prototype built, a set ofplanned flight tests can be progressively carried on.

Moreover, wind tunnel testing may be used to de-termine experimentally propulsion efficiency ratesthat so far have only been estimated, namely themotor’s. Regarding aerodynamics, it is possible tovirtually eliminate the uncertainty associated to theresults obtained in XFLR5 if a CFD computationalanalysis is performed on the updated airframe ge-

ometry. As long as the drag correction applied donot underestimate real values roughly, the currentdesign will still be able to fulfill the mission require-ments. Climb performance is expected to end upmore changed than cruise’s because it depends notonly on drag polar computations but also on theestimate made for stall at higher angles of attack.

References[1] Z. Bessoni and M. Junior. Structural design

of mini-uavs with conventional and alternativematerials. In 14th Brazilian Congress of Ther-mal Sciences and Engineering. University ofBrasilia, November 2012.

[2] L. Candido. Projeto de um UAV Solar de grandeautonomia. Master’s thesis, Universidade daBeira Interior, October 2014.

[3] T. Ferreira. Hybrid Propulsion System of a LongEndurance Electric UAV. Master’s thesis, Insti-tuto Superior Tecnico, November 2014.

[4] D. Figueiredo. Autopilot and Ground ControlStation for UAV. Master’s thesis, Instituto Su-perior Tecnico, November 2014.

[5] A. C. Marta. Long Endurance Electric UAV.Technical report, Instituto Superior Tecnico,2013.

[6] P. Miller. Remote Person View (RPV) for aLong Endurance Electric UAV. Master’s thesis,Instituto Superior Tecnico, May 2015.

[7] T. L. Saaty. The Analytic Hierarchy Process:Planning, Priority Setting, Resource Allocation.McGraw-Hil, 1980.

[8] N. Silva. Parametric Design, AerodynamicAnalysis and Parametric Optimization of a So-lar UAV. Master’s thesis, Instituto SuperiorTecnico, June 2014.

[9] H. Vidales. Design, Construction and Test ofthe Propulsion System of a Solar UAV. Master’sthesis, Instituto Superior Tecnico, March 2013.

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