Design of a High-Altitude Long-Endurance Solar-powered

Design of a high-altitude long-endurance solar-poweredunmanned air vehicle for multi-payload and operationsG Romeo�, G Frulla, and E Cestino

Department of Aerospace Engineering, Turin Polytechnic University, Corso Duca Abruzzi, Turin, Italy

The manuscript was received on 1 June 2006 and was accepted after revision for publication on 22 September 2006.

DOI: 10.1243/09544100JAERO119

Abstract: Several researches are being carried out at the Politecnico di Torino with the aim ofdesigning a high altitude very-long endurance/unmanned air vehicle (HAVE/UAV). Being ableto fly in the stratosphere (15–20 km) and with an endurance of about 4 months offers an advan-tage and possibility that is presently not available with conventional aircraft or satellites.A computer program has been developed to design the platform. The change in solar radiationover a period of a year, the altitude, masses, and efficiencies of the solar and fuel cells, as well asthe aerodynamic, structural, flight mechanics, and aeroelastic performances have all been takeninto account. Extensive use has been made of high modulus graphite/epoxy when designing thestructure in order to minimize the airframe weight, but also to guarantee the required stiffnessand aeroelastic performance.

A blended wing body (BWB) configuration has been selected for solar HAVE aircraft multipayload and operation (SHAMPO) with eight brushless electric motors, as the result of a prelimi-nary design. The BWB solution has been designed according to the conventional procedures andairworthiness regulations. It seems to be the best compromise between performance, availablesurfaces for solar cells and volume for multi-payload purposes, compared to conventionaldesign.

Several profiles and wing plans have been analysed and optimized to achieve the best effi-ciency using the Xfoil and Vsaero computational fluid dynamics (CFD) software. A finite-element method and a classical theoretical analysis was carried out using the Msc/Patran/Nastran code to predict the static and aeroelastic behaviour of the SHAMPO. Aeroelasticanalysis has been performed starting with a classical linear flutter analysis and consideringan undeformed equilibrium condition. Classical linear flutter speed show as the airworthinessrequirements has been achieved in the case of SHAMPO configuration. A preliminary non-linearaeroelastic model is introduced in the design process in order to deal with specific phenomenacorrelated with high static structural deflections occurring during standard flight conditions.Important flutter speed reduction (i.e. up to 42 per cent in special cases) are possible includingsuch kind of phenomena.

Keywords: aircraft design, high altitude very-long endurance, unmanned air vehicle,aerodynamic, stability, structure, aeroelastic, flexible wing

1 INTRODUCTION

Stratospheric platforms, geostationary located ataltitudes of 17–25 km, offer the possibility of

becoming a new generation of infrastructuralelements for future Earth observation and telecom-munication systems. They can act as pseudo-satellites, but with the advantage of being muchcheaper and closer to the ground and they can per-form missions that offer greater flexibility. Theycould be self-launched, easily recovered for mainten-ance, whenever necessary, and moved to coverdifferent regions, if desired. They allow a more

�Correspondence author: Department of Aerospace Engineering,

Turin Polytechnic University, Corso Duca Abruzzi 24, Turin,

10129, Italy. email: [email protected]

SPECIAL ISSUE PAPER 199

JAERO119 # IMechE 2007 Proc. IMechE Vol. 221 Part G: J. Aerospace Engineering

detailed land vision, due to their relative closeness tothe Earth, at a much lower cost than conventionalsatellites; furthermore they are geostationary locatedand do not suffer from typical satellite revisiting timelimits from 12 h up to 20 days. These platforms couldcover a wide range of applications: real-time moni-toring of seismic-risks or volcanic areas, early forestfire detection, border security surveillance, pipeline,and power-line surveys; telecommunication servicessuch as cellular-telephone networks, photogram-metry, hydrographic monitoring system, agriculturemonitoring, etc.

Unmanned an vehicle (UAVs) have been used forover 40 years by the military services for dangerousmissions and surveillance. The technology has beenrelatively expensive and unreliable, as cost andsafety issues were considered of secondary import-ance to performance. However, technology hasadvanced sufficiently for this highly technical area ofthe aeronautical industry to be developed into a newgrowing UAV civilian industry. The United States hasalready begun investing in this field and has initiateda massive drive to develop this new industry. Europealso has the necessary technologies and should nowalso support and encourage industry to take advan-tage of new technological approaches in order tosuccessfully become a leading participant in this field.

In a recent business market study carried out byFrost and Sullivan [1], the global market for UAVsin civil and commercial applications will be close toUS$ 2 billion by 2014. The largest market sharesare expected to pertain to Coastguard and MaritimeSurveillance operations, Border Security, and ForestFire Management.

From a flying altitude of 18 km, an area of about300–400 km in diameter would be covered forcommunication transmission if the onboard antennairradiation diagram is properly chosen. This wouldmean that about seven to eight platforms couldcover the entire south Mediterranean Sea, fromSpain to Turkey (Fig. 1), creating an electronic illegalimmigration barrier as well as a secure bordercontrol system.

Compared to the present cost of airborne systems(5000–7000 E/flight hour), high altitutde longendurance/unmanned air vehicles (HALE/UAVs)offer clear advantages in monitoring missions, ascontinuous observations will be performed through-out the year over the area of interest and all therequired data will usually be available immediately.The main advantage of the very-long endurancesolar-powered autonomous aircraft (VESPAA) isthat this system has less climbing and descendingevents, which is important when considering inter-ference with aviation traffic. Other HALE-UAVconfigurations have a very limited endurance(24–36 h), which would drastically increase any

potential collision risk with civil aviation traffic.Double the number of UAVs would be necessary tocontinuously guarantee the surveillance service,thus the system total life cycle cost would beincreased to a great extent. Other medium altitudelong endurance (MALE)-UAVs have, as a furtherdisadvantage, the fact that a much higher number ofUAVs are necessary to continuously cover the entireMediterranean Sea, since the covered area decreaseswith the square value of the flying altitude (Fig. 2);the total life cycle cost systemwould increase remark-ably with aMALE configuration. Very high endurance,in fact, calls for high mission reliability requirementsof the air vehicle, its systems and payload.

Fig. 1 HeliPlatw Border Surveillance over the

Mediterranean Sea

Fig. 2 HAVE / MALE monitoring area comparison

200 G Romeo, G Frulla, and E Cestino

Proc. IMechE Vol. 221 Part G: J. Aerospace Engineering JAERO119 # IMechE 2007

An endurance of several months would only bepossible using a solar- and hydrogen-poweredplatform. The vehicle would climb to 17–20 kmmainly taking advantage of direct sun radiation.Any electric energy not required for the propulsionand payload operations would be pumped backinto the electrolyzer system that would convert thewater into hydrogen and oxygen fuel; during thenight, the platform would maintain altitude, thanksto the fuel cell system, which would produce electri-city and water as a by-product of the reaction; thegeostationary position would be maintained by alevel turning flight.

Several types of high altitude stratospheric plat-forms have already been designed [2–5]. At the endof 1994, NASA started the Environmental ResearchAircraft and Sensor Technology (ERAST) program;one of the four drones is the Pathfinder solar plat-form, which exceeded 24 km of altitude during a15 h flight. In the summer of 2001, the Helios solar-powered platform set a new world record of29.3 km. Although a flight of several days wasplanned for summer 2003, the program suffered adevastating setback when the Helios prototype waslost in June 2003 in a flight mishap during a shake-down mission. The Helios wing bent dramaticallyand oscillated, and then the wing spar broke. Itseems very likely that Helios crashed because ofundamped pitch oscillations that led to a partialstructural failure of the very flexible Kevlar/carbonspar. In the interim status report, the investigationboard stated that the ‘undamped pitch oscillationsmay be related to the complex interactions betweenthe aerodynamic, structural, stability and control,and propulsion systems on a flexible aircraft.’

The main innovative aspects of the VESPAA are thefollowing.

1. The design of the first European aerodynamicsolar-powered platform that would be able toremain continuous in flight for very long periodsof time (4 months).

2. Low noise and zero emission aircraft. The advan-tage of zero emissions is particularly important forenvironmental monitoring.

3. The first European solar-powered platform to vali-date the feasibility and reliability of solar cells, fuelcells and brushless electric motors.

All these features lead to:

(a) reduced cost per flight hour (thanks to a largeincrease in the endurance flight hours);

(b) potentially increased acquisition cost, whilereducing the cost of maintenance and spareparts;

(c) reduced cost – larger area coverage per platform,thus, fewer platforms per area are required;

(d) improved operational safety – due to the factthat the flight would occur above aviation traffic,resulting in limited interference with it, andabove adverse weather;

(e) lower propagation delay compared to satellites.

The Department of Aerospace Engineering at thePolitecnico di Torino (POLITO/DIASP, ScientificResponsible Prof. Romeo) has acquired a great dealof experience, starting from 1994, in the design ofsolar-powered UAV as high altitude very long endur-ance platform positioned in the stratosphere at analtitude of 17–20 km and with an endurance ofabout four months [6–14]. Started under a grantfrom the Italian Space Agency (ASI), a great amountof work was carried out within two Europeanfunded projects (HeliPlatw and Capecon) whosefinal task was to fill the European gap in knowledgeand feasibility of such kinds of innovative systems.

The ESA-STRATOS definition study was alsoperformed for stratospheric platforms, driven bythe possibility of replacing or complementing spacesystems with stratospheric platforms, particularly inthe area of telecommunication applications. Themain objective of this study was to explore the possi-bility of the development and operation of a Euro-pean stratospheric platform based on a soundanalysis of possible service areas, assessment ofavailable and future technologies, and to perform aconceptual design for the most suitable platformconcept that would answer the requirements offuture telecommunication markets (Fig. 3).

Two reference configurations for stratosphericplatforms were selected from a number of studieson stratospheric platforms and some existing designsof stratospheric aircraft: an aerodynamic one and anaerostatic one. Both are electrically driven solar-powered designs. Detailed investigations into theaerodynamic and aerostatic reference design led tothe identification of the critical platform technol-ogies. Most of these technologies are common toboth platform categories. On the basis of the necess-ary technology development, a stepwise approachwas sustained using both types of platforms in theprototype and test phase. This approach indicatesthat an operable stratospheric platform could bedeveloped within 10 years.

There is no clear rule on whether aerostatic oraerodynamic systems are in general the preferablesolution for stratospheric platforms. Because oftheir different characteristics and behaviour, theplatform selection very much depends on theapplications and the corresponding flight envelope.

A comparison of the technical reference designswith the payload requirements from the systemarchitecture showed that the smaller aerodynamicconfiguration (payload of 150 kg and available

High-altitude long-endurance solar-powered unmanned air vehicle 201


power for payload of 1.5 kW) would be suitable as apreliminary demonstrator of all the key technologiesand could also demonstrate some of the servicecapabilities, mainly for Earth observation, while thelarger aerostatic design (payload greater than1000 kg and power of 20 kW) would be required foran operational telecom service scenario).

On the other hand, for a rigorous station-keepingposition throughout the year, a jet stream of up to40 m/s in the wintertime could be difficult to be con-trasted because of the wind sensitivity of aerostaticplatforms. However, because the aerostatic lawimposes UAV dimensions, a higher power energywould be possible as the efficiency of the fuel cellsand solar cells would increase. The high cost ofeach airship, in comparison to the aerodynamicplatform, is another critical point.

Solar-powered stratospheric platforms are alsopreferable to other HALE/MALE platforms because of:

(a) environment-friendly operations (no emission inthe lower stratosphere where the highly sensitiveozone layer is located);

(b) similarity of the technologies to space-based sys-tems (power systems, autonomous operations,and payloads);

(c) similarity in operations to space based systems(launching, very-long operations, etc.).

Regenerative power technologies, such as thin filmphotovoltaic (PV) arrays, fuel cells, electrolyzers,power, and management systems, are the keys toachieve long-endurance regenerative UAV. Sincethe basic power source, the sun, is not availablethroughout the whole day, effective designs formanaging, collecting, storing, and consumingenergy are needed to make the platform a realalternative to satellites for night and day missions.

The design of a vehicle that is able to operate solelyfrom the incoming energy of the sun is a fine balance

between energy collection and energy consumption.This energy balance is influenced by a number of fac-tors such as the operational environment and thecapabilities and efficiencies of the power systemcomponents. Considering the solar radiation that isavailable at 17 km of altitude for a panel surfaceplaced parallel to the Earth’s surface, a maximumspecific solar power of 675 and 475 W/m2, respect-ively, at 368 and 458N would be detectible on theworst day (December 21). During the best optimalday, the maximum specific solar power is about1240 and 1190 W/m2, respectively, at 368 and 458 N.The energy balance for a HeliPlat configurationflying in June at 17 km of altitude is reported inFig. 4. A very large amount of electric power wouldbe available from sun radiation, with a peak at180 kW. However, because of the low solar cellefficiency taken into consideration (21 per cent),just a small amount of power would be availablefor the platform. With an optimised design of the

Fig. 4 Solar power distribution and energy balance

Fig. 3 Aerostatic HAVE-UAV for telecom applications



platform, just a small amount of power would berequired by the motors to fly (about 6.5 kW); theextra energy available from the solar cells would besupplied to the electrolyser for the production ofhydrogen and oxygen for the night flight.

To accurately determine the feasibility of the highaltitude aircraft concept, a detailed model of thewind environment is also necessary. This windmodel would provide mean and 99th percentilewinds along the area of interest for the analysis.

Although the wind does not affect the aircraft’spower generation, it has a significant effect on itsdrag and power consumption. A wind study wastherefore performed at several locations. The upperair data climatic tables recorded by the Italian AirForce from 1963 to 1997 were analysed to obtainthe wind profiles as a function of the altitude (from1000 hPa (111 m) to 10 hPa (31 055 m)) for severallatitudes. Analysing the data (Fig. 5(a)), it was clearthat the jet-streams are at a minimum between thealtitudes of 18 and 22 km, depending on the time ofyear. An average wind speed of 17.1 m/s wasrecorded in January at 18 km in altitude. The record-ing frequency of such high jet streams, from 1963 to1997, was also investigated (Fig. 5(b)). In January,60 per cent of the events recorded were in a lowerspeed range than 15 m/s; a higher speed than35 m/s was recorded just 56 times, in more than2000 events, over 34 years, although in only a fewcases (twice in 34 years), high wind speeds of up to60 m/s were recorded.

From the educational point of view, design of asolar-powered airplane presents an interesting caseto introduce students to the procedures and prac-tices of aircraft design as a means of illustrating theoften conflicting requirements of the many involveddisciplines. In such kind of advanced aircrafts the dif-ficulties with sun-powered flight make an accurateintegration of the structural weight optimisation,aerodynamics, propulsion efficiency, and electronicdevices much more critical. In addition studentsapproach the work with present time problemssuch as non-linear aeroelasticity or design ofadvanced composite structures.

2 CAPECON PROJECT

Seven potential configurations have been evaluatedas part of the EC funded Project ‘CAPECON: CivilUAV Applications and Economic Effectivity of Poten-tial Configuration Solutions’: three HALE (with thePolitecnico di Torino – Department of AerospaceEngineering Scientific Responsible: Prof. ROMEO –as work package leader of solar, modular andblended configuration design, and Task leader forsolar configuration design), two MALE and twoRotary UAV concepts. In addition, a comprehensivecost model has been developed for the first time toassess the economic advantages and disadvantagesof the various configurations. On the basis of theseevaluations, some recommendations were made for

Fig. 5 (a) Average wind speed profile for several altitudes and seasons. (b) Statistic distribution at

18 km in January



further critical technology (CT) and technologyprogram (TP) activities within the sixth framework.

The basic requirements for the CAPECON SolarHALE UAV configuration have been first established.The initial starting point requirements for the SolarHALE UAV configuration have been obtained fromprevious HeliPlat experience know-how. In particu-lar, the platform would accommodate a maximumuser payload mass of 130–150 kg and consume amaximum electrical power of 1.3–1.5 KW. A maxi-mum total mass of 850–900 kg is expected for6-month flight at an altitude of 17 km. The choiceof 17 km is a compromise between less wind andlow air density – that means very low Reynoldsnumbers and solar radiation.

The following tasks have been carried out byPOLITO/DIASP during the HALE UAV solar configur-ation design.

1. Ground rules and assumptions: the solar-poweredHALE UAV has been designed, considering theMediterranean Sea border patrol and forest firemonitoring during summer time as possibleapplications. The platform would fly for 6months, starting on the first of April, at an altitudeof 17 km for a latitude below 44 8N.

2. Preliminary design: including sizing, geometricaldefinition (three views and three-dimensional),external layout, aerodynamic definition, structureconcept, system definition, payload integration,internal layout and weight, and balancecomputation.

3. Aerodynamic analysis (computational fluiddynamics (CFD)).

4. Flight performance analysis: including databaseestablishment (aerodynamic, weight, andpropulsion databases) and performance compu-tations (integral performance, point perform-ance-flight envelopes, and ground performance).

5. Structural analyses finite-element method (FEM)of the chosen configuration and aeroelastic con-siderations of the high aspect ratio/low masswing.

6. Reliability, maintainability, and safety analysis.

3 DESIGN PROCESS

As most applications for stratospheric platformsrequire continuous operation, solar power seems tobe the only adequate option to provide the necessarypower for propulsion and for the payload. A powersystem consisting of solar cells and an energy storagesystem to bridge the night phases would therefore berequired.

Solar power electric energy would be used in theproject to drive the propellers for the requiredflight. It would use electric-motor-driven propellers

to convert electrical power into thrust. Any electricenergy not required for the propulsion and payloadoperations would be pumped back into the fuel cellenergy storage system and, during the night, the plat-form would maintain altitude, thanks to the stored(solar) energy [6–8].

The power collected by the PV panels during theday would be absorbed not only by the electricmotors in order to drive the platform but also bythe control system and by the payload. The powerthat remains would feed the electrolyser, whichwould produce high-pressure hydrogen and oxygenfrom the stored water. The gases would be stored inseparate cylinders and supply the fuel cells duringthe night, so that the electric power needed by theplatform would always be assured. The systemwould be completed by a pump to feed the electroly-ser with water, the gas tanks, the water tank, andsome auxiliaries.

The regenerative fuel cell energy storage systemwould be used to balance the power. Although thepower required to fly the aircraft is roughly constantover time, the power available from sunlight changesduring the day – from zero at sunrise, to peak powerat noon, and back to zero at sunset.

The amount of energy that would need to be storedovernight (energy storage system capacity) dependson the location (Earth latitude) and the time ofyear. In general, winter conditions and higher lati-tudes require more capacity because the nights arelonger and there is less sunlight.

The key figures of merit are the storage system’senergy density (watt-hours of energy stored per kilo-gram of storage system mass) and its efficiency. Thehigh energy density that would be required to keep asolar airplane flying over most locations and duringmost times of the year has forced designers to con-sider more complicated forms of energy storage.Chief among these is the hydrogen/oxygen regenera-tive fuel cell (RFC); an energy storage system, basedon dedicated proton exchange membrane fuel cells,gives an energy density of 450–600 Wh/kg, withrespect to 200–220 Wh/kg of the best lithium bat-teries, and efficiency of about 60 per cent. No otherenergy storage method (outside nuclear power) hasa higher energy density.

Beginning in the early 1990s, significant invest-ments from the automotive industry resulted in aseven-fold increase over a span of 5 years. Rapidlyemerging hydrogen-based technologies, primarilyderived from the automotive industry, could beused to start a new era of propulsion systems forlight aircraft and small commuters and electricalsystem replacement for larger transport aircraft.

The feasibility of this project depends on severalkey-enabling technologies such as fuel cell stack,fuel cell system, hydrogen fuel storage, and a safe



airport hydrogen-fuelling infrastructure. It isestimated that an almost ten-fold increase inenergy density (5 kWh/kg) can be expected withinten years to enable the all-electric flight of smallcommercial aircraft.

Since today’s solar cells only convert a smallpercentage of the sun’s energy into useful work,any stratospheric platform needs a large surfacearea for power generation. The solar cell area doesnot impose a problem on aerostatic platforms dueto their large surface provided by the lift gas hull.For aerodynamic configurations, the sun-poweredlevel flight can only be obtained through a suitabledesign and accurate integration of the highest stan-dards that can be achieved for each involved techno-logical item: structure, aerodynamics, solar cells,energy storage, propulsion efficiency, and avionics

The theoretical efficiency limit of solar energyconversion under idealized conditions is 86 per cent.Present technology, however, limits this percentageto about 50 per cent. One of the most fundamentallimitations on solar cell efficiency is the band gapof the semiconductor from which the cell is made.In order to obtain improvements, researchers andmanufacturers stack different band gap materials inmulti-junction cells. The maximum efficiency thatcan be achieved with a solar cell, made from asingle material, when converting light to electricalpower is about 30 per cent; the best efficiency thathas actually been achieved is about 25 per cent.Many different layers could be stacked to catchphotons at all energies, reaching a better efficiencythan 70 per cent (the best efficiency actually achievedwith multi-junction solar in only a few researchprojects is about 35 per cent).

The two most common types of semiconductormaterials used in PV cell manufacturing are silicon(Si) and gallium arsenide (GaAs).

Most commercial solar cells used in terrestrialapplications are made from wafers of silicon. About85 per cent of the PV market consists of Si solarcells. When used in commercial production,they allow power conversion efficiency close to17 per cent. These thick (200–300 mm) single-crystalsilicon cells are today available at low prices(about 700–1000 E/m2). Higher efficiency (up to25 per cent) very thin (50–70 mm) gallium arsenidecells are also available, although at a higher price(about 25.000 E/m2).

Since the solar cell efficiency closely depends onthe surface temperature of the PV array, a low oper-ational temperature of the solar cell is beneficial. Forthis application, the ambient temperature is animportant parameter, because the estimated temp-erature at the flying level (17 km) is about 558C, butthe solar cells specification are given for the celltemperature of þ288C (standard test conditions).

Good efficiency could be obtained through theproper design of the electric brushless motor(95 per cent) and of the propellers (85 per cent),which have a high diameter and a speed of about a1000 min; these elements should be designed tofulfil the requirements of low mass and highreliability [15]. Particular attention should be paidto the automation of the full propulsive system; fur-thermore, the power that is available from the sunshould be properly split between the motors andenergy storage system.

A multi-disciplinary optimisation computer pro-gram has been developed to design the platform,taking into account wind speed up to an altitude of27 km, the solar radiation changes over a period of1 year, altitude, latitude, solar cell efficiency andweight, fuel cell efficiency and energy density, aero-dynamic profile drag, etc. The preliminary design ofa Solar HALE Aircraft Multi Payload and Operation(SHAMPO) (Fig. 6) flying at high altitude and lowReynolds numbers, has led to the main featuresreported in Table 1.

4 INTERNAL AND EXTERNAL LAYOUT

The power system of a stratospheric platformconsists of the three main subsystems:

(a) power source;(b) energy storage;(c) electrical propulsion system.

which should be managed by an electronic controlsystem to make optimum use of the availableenergy. The architectural analysis and design of thepower conditioning (DC/DC converter) system forthe fuel cells, the solar power supply and the propul-sion system (motor and inverter) interface of theplatform has been performed with the assistance ofnumerical simulation [16].

The PV energy source module converts the solarenergy into electrical energy and provides it to theDC link in a suitable way. The module consists ofthe solar cells, the DC/DC converter, and the maxi-mum power point tracker (MPPT) controller. TheMPPT controller and the DC/DC are used to obtainthe maximum power from the solar cells.

The FC energy storage module acts as an energystorage system. It consists of the fuel cell array, thehydrolyser and the hydrogen, oxygen, and watertanks. The gases are stored at high pressure (120–250 bar) and used during the night to feed the fuelcells that supply the motors; the water feeds theelectrolyser to close the cycle.

The electric motor module, consisting of theelectric motor and the inverter, converts electricalenergy into mechanical energy.



The mixed supply System, with solar cells and fuelcells, requires a common DC link in order to adjustthe voltage level between the different sources.Solar cells have the additional requirement of loadimpedance adjustment to allow the maximumpower point to vary according to the ambient con-ditions, like temperature or solar irradiance.

The core of the full scale HeliPlat Power SupplyDrive (PoSuDri) should be based on a DC link of360 V. This voltage has been chosen as a balancedvalue in order to limit the size and weight of the elec-tric cables, which in the present design of the solaraircraft must supply power over a 70 m length ofwingspan. Therefore, all electrically powered subsys-tems become part of the electric power distributionsystem, which unites all the electrical sources andloads of the aircraft by means of the high voltagepower distribution bus. A three-dimensional CAD

model of the preliminary internal layout has beenset up using CATIA V5 (Fig. 6a).

A blended wing body (BWB) configuration witheight brushless electric motors has been chosen asa result of the preliminary study. The weight ofeach motor is about 6.5 Kg, including the carbon-fibre propeller blade.

The BWB solution, compared to the conventionaldesign [6–12], seems to be the best compromisebetween performance, available surface for solar-cellsand volume for multi-payload purposes; the fuselageblended zone in fact presents a volume of about 38 m3.

TheHeliplat propeller blade (blade radius ¼ 1.15 m),[6–8], has been used for the preliminary design ofSHAMPO. Propeller characteristics are obtainedusing two in-house FORTRAN developed tools. Thefirst tool is able to compute the maximum efficiencypropeller geometrical characteristics giving as inputthe design condition, blade number, and airfoilsdatabases. The output of the first tool are the chordand pitch distributions. The second tool gives per-formances in terms of efficiency, thrust coefficient,and torque coefficient and are used to evaluate thecharacteristic far from the design point.

The thrust supplied by this propeller in cruise con-ditions is sufficient for the SHAMPO design point(Fig. 7), and it is possible to increase the airspeed byusing the 5000 W of surplus power available for 2 h.

A flight power of about 6700 W supplied to theelectric motors have been computed in cruise con-ditions (Z ¼ 17000 m, 1st April 38 8N) flying at25 m/s; supplying a maximum available power ofabout 9800 W, to the motors the airplane couldreach 31 m/s and in the case of emergency or to con-trast wind jet-streams, it would be able to reach36 m/s, using 5000 W reserve power.

Rolling control on SHAMPO is accomplished byusing ailerons located on the outboard section of thewing. The restrictions imposed by roll control forgeneral aviation aircraft are not severe [17–19]. The

Table 1 The main characteristics of SHAMPO

Description Symbol Value

Flight altitude Z 17 000 mMaximum power available for

payloadPPL 1300 W

Avionic mass WAV 32 kgMaximum payload mass WPL 130 kgStructural mass WSTR 430 kgSolar cell mass WSC 127 kgTake-off weight WTO 924 kgPower available for the avionic PAV 325 WCruise flight Power supplied to the

electric motorsPFLY 6700 W

Sun power (38 8N April) PSUN 11 560 WEfficiency energy storage system

(Fuel cellþ A1)hFC 0.6

Density energy storage system(Fuel cellþ A1)

WFC 550 Wh/kg

Efficiency solar cells hSC 0.21Density solar cells WSC 0.6 kg/m2

Efficiency electric motor hM 0.95Efficiency propeller hPROP 0.85Number of motors N 8Cruise airspeed TAS 25 m/s

Fig. 6 Aileron segments and internal layout



ailerons are positioned along the wing span at adistance of 18.8 to 30.8 m, and the total surface is sub-divided into three parts (Fig. 6b), each one controlledby an electric motor in order to have a triple redun-dancy system. The control surfaces have been intro-duced in the Vsaero CFD aerodynamic model. Theinput file has therefore been modified to take intoaccount several rotations of the control surfaces.

The CATIA V5 external layout (Fig. 8) has also beendeveloped and used for aerodynamic and structuralanalysis models and to estimate the aerodynamicderivatives.

The main geometrical characteristics of the finalconfiguration are:

bw ¼ 73 m; Sw ¼ 192 m2; ARw ¼ 28; cmean

¼ 3.89 m; Lw ¼ 58; bht ¼ 12 m; Sht ¼ 25 m2;ARht ¼ 5.76; and Lht ¼ 8.88.

5 AERODYNAMIC ANALYSIS (CFD)

The design procedure that has been followed in thepreliminary analysis is based on the energy balance

equilibrium between the available solar power andthe required power, the former being dependenton the solar cell area installed on the wing and stabil-izer, the latter depending on the velocity and totaldrag of the platform.

The power required, per unit wing area, (Preq/Sw)to maintain horizontal flight is given by

Preq/Sw ¼ (2/r)0:5(Wtot/Sw)3=2(CD/C

3=2L )

hprop

(1)

where the propulsion efficiency hprop includes pro-peller, motor, and inverter efficiency.

The endurance parameter has in particular to besatisfied to minimize the power required for ahorizontal flight. When minimizing the parameterCD/CL

3/2, it has also been preferred to reduce thedrag coefficient CD instead of only increasing thelift coefficient CL, because of the lower structuralload. The total drag coefficient is obtained from thesum of the wing and tail profile drag, the wing andtail induced drag and the parasite drag (fuselage,booms, hard interaction with wing and tails). Rey-nolds numbers of less than 700 000 should be con-sidered in most cases.

The aerodynamic calculations were made using theVSAERO code [20], a computer programused to calcu-late thenon-linear aerodynamic characteristicsof arbi-trary configurations in a subsonic flow. Non-lineareffects of wake shape are treated in an iterative wakerelaxation procedure, and the effects of viscosity aretreated in an iterative loop, while coupling the poten-tial flow and integral boundary layer calculations.

The aircraft configuration has been changedduring the design process. The first unswept blendedconfiguration, which was worked out during the firstiteration, presented some problems from the longi-tudinal stability point of view. In order to move thecenter of gravity forward, with the constrains ofinstallation volumes, a classical solution with a fuse-lage has been tested. This solution was later dis-carded because it was too penalizing in terms ofaerodynamic efficiency.

To increase the stability margin without any sig-nificant aerodynamic efficiency reduction, a wingwithout fuselage and with a small angle of sweep(about 58) has been chosen in order to have a staticmargin of about 9–10 per cent.

A second investigation has been then performed,with a wing sweep angle of 58, to optimize the wingairfoils and the shape of the tip, with the aim of redu-cing the contribution of the induced drag.

The main factor responsible for total platformcoefficient drag (0.0266 at a 68 angle of attack) isthe wing and in particular the outer part. Induceddrag corresponds to about 63 per cent of the totalFig. 8 Shampo configuration, external view

Fig. 7 Propeller thrust versus required thrust



drag, whereas viscous drag accounts for 24 per centand pressure drag for 13 per cent.

The use of a round wing tip appears to have apositive role in minimizing tip vortex formationand thus minimizing induced drag.

When using the same airfoil DAE21 for the wingand blended fuselage part, a large zone ofdepression, which was responsible for performanceslosses, is formed behind the wing. For this reason,new SH118 profiles, which are capable of improvingthe pressure recovery in the fuselage zone, have beendeveloped; with this airfoil, the wake leaves the bodyin a more uniform way (Fig. 9), reducing downwashin the blended zone.

This configuration presents a maximum efficiencyof about 41.6 at CL ¼ 0.85 (Fig. 10) and a maximumendurance parameter of about 43 at CL ¼ 1.2. The

lift and pressure distribution along the wing spanhas also been computed (Fig. 11).

The configuration also presents a good stabilitymargin of about 10 per cent of the mean aerody-namic chord. This configuration has been chosenas a starting point for the structural design.

6 FLIGHT PERFORMANCE

A multidisciplinary design to determine the staticand dynamic behaviour of an aircraft in the earlystage can lead to significant weight and cost savingand improvement in configuration and performance[6–8, 12]. The computational fluid dynamics solver(VSAERO CFD) has been used to directly calculatethe stability and control derivatives of the newSHAMPOconfiguration. TheCFD forces andmomentshave been differentiated with respect to the angle ofattack, angle of sideslip, and aircraft shape parametersin order to determine these derivatives. A completedatabase of the main static and dynamic derivativeswere worked out using this method.

The developed VSAERO model has a wing andhorizontal tail, but lacks a vertical tail, therefore clas-sical formulas have been used [17–19]. The effect ofa roll angular velocity on the pressure distributionand wake shape is shown in Fig. 12.

Aircraft flying qualities and stability requirementsare defined from many sources. The requirementsfor longitudinal static stability simply require thatthe pitching moment at zero-lift angle of attack,Cm0, should be positive and that the derivative ofthe pitching moment with respect to the angle ofattack, Cma, must be negative. These two require-ments are mainly determined by the wing configur-ation. The new SHAMPO configuration satisfiedboth conditions.

Fig. 9 Downwash blended zone

Fig. 10 Aerodynamic efficiency

Fig. 11 Upper surface pressure distribution at 78 angleof attack



Similar to the longitudinal stability requirements,lateral-directional stability is based on both staticand dynamic requirements. For static stability, theaircraft’s yaw stiffness, Cnb, should be positive andits roll stiffness, Clb, should be negative. Theserequirements ensure that the aircraft generates apositive yawing moment and a negative rollingmoment for a positive sideslip, and it can thus stayin controllable flight conditions. From a first esti-mation of the SHAMPO derivatives, considering avalue of natural dihedral of 2.78, the static conditionhas been verified with the calculated valuesCnb ¼ 0.012 Clb ¼ 20.083.

A flight dynamic analysis has been performed, inthe preliminary design phase, considering thatrigid-body degrees of freedom and elastic degreesof freedom are decoupled; however, this is notalways possible. The frequency separation betweenthe rigid-body motions and elastic degrees of free-dom for advanced aircraft (such as solar UAV, verylight weight UAV or advanced aircraft with activecontrol systems) may not be sufficient to permitthe typical treatment of the vehicle dynamics. Inparticular for flexible HALE-UAV, it is not alwayssafe to treat rigid body effects and flexibility effectsseparately as demonstrated by the Helios prototypemishap.

The intention of this preliminary stage has been tohave a first estimation of the airplane handlingqualities and a rigid mathematical model has beentherefore used to describe the flight dynamics ofthe aircraft.

MIL-F-8785C contains a number of criteria withwhich the modes of a piloted aircraft can be com-pared to predict the expected flying quality levels.The closest comparison category, for the SHAMPO-UAV, could be that of a small light aircraft, which isa Class I category aircraft. The approach used to ana-lyse the dynamic behaviour of the airplane has been

of using the uncoupled, linearized equations ofmotion solutions. A Matlab–Simulink [21] flightdynamic software has been used to obtain theequations.

All the longitudinal modes are stable for theSHAMPO-UAV cruise condition. With reference tothe short period criteria within MIL-F-8785C, forthe Category B Flight phase, a minimum dampingratio of 0.2 is obtained for Level 2 and a limit of0.15 for Level 3. The SHAMPO UAV longitudinaldamping is 0.203.

There are two aperiodic modes (one stable and oneunstable) and one stable periodic mode for theSHAMPO-UAV cruise condition in the lateralmotion.

The Dutch roll is a coupled oscillation betweenyaw/sideslip and roll, a complex mode usually deter-mined by the characteristics of the rolling andyawing moments with sideslip. The Dutch rollmode is stable, in SHAMPO-UAV, with reference toMIL-F-8785C. The roll-mode flying quality criteriaare also well within the specified requirementsgiven in MIL-F-8785C.

The remaining mode, the spiral mode, has alreadybeen established as being unstable. A certain degreeof spiral instability can be accepted for differentflying quality levels. For Level 1 Cat B, a time-to-double amplitude greater than 20 s is acceptableaccording to MIL-F-8785C. SHAMPO-UAV has avalue of time-to-double amplitude of about 45 s.

It can be concluded that the aircraft is staticallystable and that it is able to stabilize longitudinaloscillations on its own; it has a spiral instabilitywith a time-to-double amplitude that is howeveracceptable.

7 STRUCTURAL CONCEPTS AND ANALYSES

A preliminary structural concept was defined accord-ing to the estimated inertial and aerodynamic loads.High structural efficiency is required to minimize theairframe weight and increase the payload mass. Thisis obtained through:

(a) extensive use of high modulus graphite/epoxymaterial to obtain a very light-high stiffenedstructure with good aeroelasticity behaviour;

(b) numerical structural analysis using the FEM.

In spite of the weight reduction, in high aspectratio wings like VESPAA, aeroelastic phenomenaplay an important role that has to be satisfied andverified.

The total platform weight, per wing area, is deter-mined by adding the weight of the propulsion, thesolar cells, avionics and payload to the airframe;the weight of the propulsion system includes

Fig. 12 Effect of roll angular velocity



brushless motors, inverter, conditioning system, andpropellers.

For the adopted configuration, a good structuralwing solution to enhance the bending stiffnesscould be represented by an opportunely reinforcedhigh modulus carbon fibre leading edge wing-box(Fig. 13). The main spar could be subdivided intofive parts, which could be connected to each otherby metal fittings.

The aero-structural loads, including bothmanoeuvre and gust, have been defined accordingto the JAR-VLA 333 airworthiness regulations, witha maximum limit load n ¼ 3 for the diving speed.

A computer program has been developed to evalu-ate and define the structural characteristics of themain wing/tail spars of the platform consideringstrains under the value of 1000 m1

˙for all flight limit

loads. Different materials and layouts have beenconsidered in order to introduce a level of optimiz-ation to the iterative calculations. CFRP data havebeen introduced referring to a Young’s modulus inthe fibre direction of about 275 GPa, ultimatecompressive strain limit of about 1500 m1, a laminathickness of 0.135 mm and a density of 1580 kg/m3.The inertial loads (structural and non structural)have been estimated and taken into account accord-ing to the preliminary configuration design. The ulti-mate strain behaviour along the wing span isreported in Fig. 14.

An FE model of the SHAMPO main spar configur-ation has been developed at the preliminary stageusing the Patran/Nastran code, based on preliminarydesign structural concept. The model consists of onehalf-wing, which is considered clamped in the sym-metry plane. The leading-edge wing-box was simu-lated using shell elements (Quad4). The mesh onthese surfaces has been made respecting the exactradial and longitudinal position of the ply lay-up,which has been determined with the preliminary

design structure program. The load distribution hasbeen determined from an estimated lift and massdistribution according to the preliminary data andaerodynamic calculations in order to obtain anequivalent load condition. The load conditionreported in this work only concerns the D point ofthe flight envelope (n ¼ 3, VEAS ¼ 47 km/h, accord-ing to airworthiness regulation). A comparisonbetween the theoretical and Nastran results isreported in terms of ultimate strains (Fig. 14). Anew and improved FE model of the entire SHAMPOairframe has been developed, in order to evaluatethe structural behaviour of the whole optimizedSHAMPO configuration (Fig. 15). The structure hasbeen designed to have a very stiff low weight air-frame. Wing and horizontal tail spars have differentlay-up (+208/0n8/+458/core)s along the span inorder to minimize the weight of the structure whileribs, for modelling simplicity, are considered allmade by the same material with the same lay-up.The conformal fuselage also had a variable lay-up.The results of the FE model showed a wing weight(wing-box and ribs) of about 314 kg; the fuselageweight of about 58 kg, and there was a total airframeweight without skin of 412 kg. These results are in agood agreement with weight estimation formulasfor man-powered airplanes, high-altitude airplanes,or ultra-light cantilever twin-boom tails. Some FEresults are displayed in Fig. 15; a maximum deflec-tion of about 6 m at the tip of the wing has beendetermined at point D in the n–V diagram, consider-ing a safety factor of 1.5 (N ¼ 4.5). No particularstress concentration has been revealed.

8 AEROELASTIC CONSIDERATIONS

A preliminary aeroelastic analysis has been carriedout on SHAMPO UAV. The first preliminary modelconsisted of a beam-wise structure of one half-wingconsidered clamped in the symmetry plane,Fig. 13 Structural configuration

Fig. 14 Strain behaviour along wing span



positioned along wing-box elastic axis as determinedin section 8. The stiffness and inertia properties havebeen introduced, according to the preliminary struc-tural design concepts.

The motors, fuel cell system, payload, and vesselshave been considered as concentrated masses withestimated properties, according to the preliminarylayout. The other non-structural masses have beendistributed along a beam wise rear spar with onlyan inertial effect. No specific structural details havebeen introduced into the model, such as junctionsor rotating propellers. The aerodynamic wing loadshave been defined according to the strip theory andthe Theodorsen function [22].

In this preliminary analysis, the flutter calculationhas been performed considering an undeformedequilibrium condition; typical V-g plots have beenobtained and critical velocities have been deter-mined as the damping parameters tend to be posi-tive (Fig. 16(d)). No structural damping or other

external damping has been introduced. Some ofthe linear wing modes considered in itsun-deformed equilibrium state are also plotted(Figs 16(a) to (c)).

No critical speeds have been detected up to 100 m/susing the linear calculation. Compared to the designdiving speed, reported in previous sections, it ispossible to conclude that the airworthiness target issatisfied by the proposed configuration. Moreover,when the airplane shows a highly flexible wing,an advanced aeroelastic behaviour should be con-sidered as stated in references [13], [23], and [24].Moderate-to-large bending deflections, due to aerody-namic loads, cause non-linear coupling betweenedge-wise bending and torsion. Therefore, the trimcondition and the natural frequencies of flightdynamics could be significantly affected by theinherent flexibility and non-linear deformations ofthe wing structure, leading to an overall change inthe wing’s aeroelastic characteristics [23, 25].

Fig. 15 FE model and wing and tails deflection



A non-linear aeroelastic model has beendeveloped and proposed as a design/diagnosticanalysis tool [24, 25]. A geometrically consistentnon-linear structural model for large deformations,developed by Hodges and Dowell [26] and modifiedaccording to Da Silva’s second-order geometricalnon-linear terms [27], has been implemented. Thestructural model has been coupled to an unsteadyaerodynamic model for an incompressible flowfield, based on the Wagner aerodynamic indicialfunction, in order to obtain a non-linear aero-elastic model [28]. Linear and non-linear staticand dynamic aeroelastic responses have beenconsidered.

The introduction of a small dynamic perturbationabout a non-linear static equilibrium has beenapplied in order to solve the system of governingequations and to study the sub-critical and super-critical aeroelastic response and the flutter bound-aries. The in-plane, out-of-plane, and torsionaldisplacements have been considered to be a sum-mation of the static and dynamic components inthe un-deformed reference system.

The partial differential equations governing thedynamics of the flexible beam, were reduced to asystem of ordinary differential equations [24, 25,29] using a series discretization technique [30],along with Galerkin’s method, to obtain the aeroelas-tic governing equations. Galerkin’s shape functionshave been chosen from a set of linearly independentand complete functions derived from a clamped-freebeam case.

The deflected equilibrium condition has also beenaffected by a change in the overall flight dynamiccharacteristics of the aircraft due to wing flexibility.It has been possible to implement the procedure ina preliminary aeroelastic evaluation as the nonlinearequilibrium system has been included in the analy-sis: the actual static deflection has to be connectedto a specific flight condition. If, for example, we con-sider that at each speed a level flight condition has tobe satisfied, the flutter speed and frequency areobtained with the subsequent main steps: (a) choos-ing a flight speed and the correct angle of attack tosatisfy the vertical equilibrium condition, (b) calcu-lating the static equilibrium deformed shape at the

Fig. 16 Linear natural modes and flutter identification



flight speed, (c) linearizing around the deformedshape, (d) calculating the eigenvalues of the linear-ized system, and (e) checking for stability. If the flut-ter speed and frequency are stable, increase the flightspeed and repeat all the preceding steps untilinstability speed is reached.

The flutter speed and frequency are defined as thelowest airspeed and corresponding frequency atwhich a given structure flying in a specific atmos-phere will exhibit sustained simple harmonic oscil-lations [22]. The flutter condition is a borderlinesituation, or neutral stability, due to the fact thatsmall motions must be stable at a speed below flutterspeed, whereas divergent oscillations occur in arange of speed above flutter speed.

Using the in-house developed tool linear flutterspeed can be computed assuming the equilibriumstatic configuration as zero with no coupling effect.Including equilibrium terms, non-linear flutterspeed calculations can be performed. Results on atypical HALE configuration are reported in refer-ences [24], [25], and [29] showing the flutter speedbehaviour closely depends on tip deflection. With astatic tip displacement of 1 m, the nonlinear flutterspeed reduces to 70 per cent of that obtained in theun-deformed equilibrium condition, while for a wtip

of 2.5 m a flutter speed reduction of about 42 percent respect to linear flutter has been obtained.

Another design indication that should be takeninto account [13, 23, 25] is that the stiffness ratiobetween in-plane stiffness and out-of-plane stiffnesscould play an important role in increasing the flutterboundaries; the positive effect of increased flutterspeed when a higher value of the stiffness ratio isintroduced into the airplane structure design hasbeen pointed out [23, 25].

Stiffness characteristics are determined accordingto an equivalent Euler-Bernoulli beam. Otherextended models [31] are under development inorder to have a better structural approximation.

The linear flutter results imply that smalldisturbances will grow exponentially for velocitieshigher than the flutter speed. However, as the ampli-tude of oscillations grows and additional non-linearstiffness is introduced, the vibrations do not growto infinity but instead converge to a limit cycle oscil-lation (LCO). The amplitude of the LCO gives an ideaof the amount of stress/strain on the structure and itsstudy could be useful for analysis and design.Whether such non-linear effects are favourable ornot will greatly depend on the particular circum-stances and physical parameters that are involved.Nevertheless, it is clear that non-linear effects oftenlead to LCOs. In their absence, the alternativewould be a catastrophic flutter and consequentlystructural failure.

The amplitude of the LCO can only be determinedby time-marching the non-linear governingequations of the aeroelastic system, while theLCO stability can be done via a Lyapunov-basedapproach [28].

An example of supercritical aeroelastic behaviour(out-of plane tip-deflection wtip), detected abovethe non-linear flutter speed has been obtained con-sidering a small initial disturbance and by analysingthe response via a time marching integrationscheme, as reported in Fig. 17; the non-linearMathematicaw solver has been used.

Traditionally, flying aircraft have been treatedwithin the confines of flight dynamics which, inmost cases, deals with rigid aircraft. On the otherhand, flexible aircraft fall into the domain of aero-elasticity. Although some attempts have been madeto include aircraft rigid body degrees of freedom,aeroelasticity is concerned mostly with the vibrationand flutter instability of wings fixed at the root. Inreality, all aircraft possess somemeasure of flexibilityand carry out rigid body manoeuvres, so that thequestion arises as to whether flexibility can be con-sidered in the real aircraft flight dynamics behaviour

Fig. 17 Limit cycle oscillations (LCO) in the supercritical regime



and whether it should already be considered in thepreliminary design phase [25].

The flight dynamics of a flexible-model similar toSHAMPO has been considered, including the geome-try of the aircraft, aerodynamic data and mass, flex-ural rigidity, and torsional rigidity distributions[31]. The wing of the aircraft consists of a right-halfwing and left-half wing modelled as beams under-going bending and torsion. The remaining membersare assumed to be rigid. The equations of motion areobtained by means of the Lagrangian equations inquasi-coordinates. The resulting equations are non-linear ordinary differential equations for the rigidbody translations and rotations, and the boundary-value problems for the elastic deformations. The dis-tributed variables are discretized in space using theGalerkin’s method. Using a perturbation approach,the resulting set of non-linear ordinary differentialequations are separated into a set of zero-orderequations for the rigid body variables and a set offirst-order equations for the perturbations in therigid body variables and elastic deformations. Thezero-order equations are used to design the desiredmanoeuvres. The first-order equations are used toassess the stability of the aircraft on the chosenflight paths and to design feedback controls to main-tain stable flights.

In the numerical example, steady level flights atthree different forward velocities have been con-sidered [31]. The required control inputs thatpermit the flight and the resulting static deflectionshave been computed. The stability of the aircraftabout these trims has been dealt with by inspectingthe eigenvalues of the first-order (perturbation)equations. The aircraft has been found to be unstablein one of the trims and marginally stable in the othertwo. These results have been compared with theresults of the quasi-rigid aircraft and the restrainedaircraft models. Moreover, it has been shown that acontrol design based on the quasi-rigid aircraftmodel can destabilize the aircraft and that the fullmodel must be used for a successful feedbackcontrol.

9 CONCLUSION

The results of this preliminary study show that itcould be possible to obtain a very long endurancehigh altitude platform for Earth observation and tele-communication applications, at least for low latitudesites in Europe and for several months of continuousoperation.

A BWB configuration of Solar HALE Aircraft hasbeen obtained as a result of the preliminary design.The BWB solution seems to be the best compromisebetween performance available surface for solar-

cells and volume for multi-payload purposes. Thenumerical aerodynamic results obtained show theavailability of suitable airfoils for operation at lowReynolds numbers.

The structural concept layout has been definedand designed according to regulation requirements.The preliminary analytical and FE results showed agood correlation. The assumed structural concept,sandwich wing-box spar, demonstrates low struc-tural mass and satisfactory classical flutter beha-viour. According to airworthiness regulations apreliminary flight dynamic and flutter analysis hasbeen carried out. Moreover, when the airplaneshows highly flexible wings an advanced aeroelasticbehaviour should be considered. Considering thecase of a typical HALE UAV, a flutter speed reductionof up to 42 per cent with respect to classical analysishas been obtained considering a 2.5 m wing tipdeflection.

ACKNOWLEDGEMENTS

The authors acknowledge the important contri-bution made by the following funding programmes.

1. HELIPLAT w. ‘Design of High Altitude Solar-pow-ered Platform for Telecommunication and EarthObservation Applications’. ASI (Italian SpaceAgency). Contracts 1995-99. Coordinator: Prof.G. Romeo

2. HELINET: (Heliplat Network). ‘Network of theStratospheric Platforms for Traffic Monitoring,Environmental Surveillance and Broadband Ser-vices’. EC 58 FP – Contract IST-1999-11214,2000-03; Coordinator: Polito. WP Leader of Heli-platw Design and Manufacturing: Prof. G. Romeo

3. CAPECON. ‘Civil UAV Applications and EconomicEffectivity of Potential Configuration Solutions’.EC 58 FP – Contract GRD1-2001-40162, 2002-05.Coordinator: IAI. - WP Leader of 3 HALEsDesign: Prof. G. Romeo.

4. STRATOS. ‘Stratospheric platform – a definitionstudy for an ESA system’. ESA Contract, 2002–2004. Coordinator: Airobotics -WP Leader Aerody-namic Platform Study: Prof. G. Romeo.

5. UAVNET. ‘Civilian UAV Thematic Network: Tech-nologies, Applications, Certification’. EC 58FP,G4RM-CT-2001-05. Coordinator: IAI

REFERENCES

1 Cohen, D. 2005 Global markets for civil and commer-cial UAVs. Frost & Sullivan Interactive Briefing,UAVNET, available from http://www.uavnet.com.

2 Anon. Eternal airplane. The lift-off of unmanned solar-electric wings. Pop. Sci., 1994, 100, 70–75.



3 Hall, D. W., Watson, D. A., Tuttle, R. P., and Hall, S. A.Mission analysis of solar powered aircraft. NASA CR-172583, 1985.

4 Hall, D. W. and Hall, S. A. Structural sizing of a solarpowered aircraft. NASA CR-172313, 1984.

5 Youngblood, J. W., Talay, T. A., and Pegg, R. J. Designof long endurance-endurance unmanned airplanesincorporating solar and fuel cell propulsion. AIAApaper 84-1430, 1984.

6 Romeo, G. Design proposal of high altitude very-longendurance solar powered platform for earth obser-vation and telecommunication applications. 48th Inter-national Astronautical Congress, Turin, October 1997,paper no. IAF-97-M.2.05.

7 Romeo, G., Frulla, G., Cestino, E., and Corsino, G.HELIPLAT: design, aerodynamic, structural analysis oflong-endurance solar-powered stratospheric platform.J. Aircr., 2004, 41(6), 1505–1520.

8 Romeo, G. and Frulla, G.HELIPLAT: high altitude very-long endurance solar powered UAV for telecommuni-cation and earth observation applications. Aeronaut.J., 2004, 108(1084), 277–293.

9 Romeo, G. Manufacturing and testing of graphite-epoxy wing box and fuselage structures for a solar-pow-ered UAV-HAVE. 21st ICAS Congress, Melbourne, Sep-tember 1998, paper A98-31591.

10 Romeo, G. Design of high altitude very-long endurancesolar powered platform for earth observation and tele-communication applications. Aerotecnica Missili ESpazio, 1998, 77(3–4), 88–99.

11 Romeo, G. and Frulla, G. HELIPLAT: aerodynamic andstructural analysis of HAVE solar powered platform. 1stAIAA Technical Conference and Workshop onUnmanned aerospace vehicles, systems, technologiesand operations, Portsmouth, USA, 20–23 May 2002.

12 Romeo, G., Frulla, G., and Fattore, L. HELIPLAT: highaltitude very long endurance solar powered UAV fortelecommunication application. FEM analysis,manufacturing and testing of 21 m long CFRP wingbox. In Proceedings of the Applied Vehicle TechnologyPanel Symposium on Unmanned vehicle for aerial,ground and naval military operations, NATO, RTO,Ankara, Turkey, October 2000, paper no. 12.

13 Frulla, G. Aeroelastic behaviour of a solar-poweredhigh-altitude long endurance unmanned air vehicle(HALE-UAV) slender wing. Proc. Instn Mech. Engrs,Part G: J. Aerospace Engineering, 2004, 218(G3), 179–188.

14 Romeo, G., Frulla, G., Cestino, E., and Corsino, G.HALE-UAV solar configuration: aerodynamic and struc-ture analysis, 2nd Iteration. CAPECON consortiumreport ID4.2.2/2e, January 2004.

15 Frulla, G. Preliminary reliability design of a solar pow-ered high altitude very long endurance unmanned airvehicle. Proc. Instn Mech. Engrs, Part G: J. AerospaceEngineering, 2002, 216, 189–196.

16 Bojoi, R., Lazzari, M., Profumo, F., and Tenconi, A.Digital field-oriented control for dual three-phaseinduction motor driver. IEEE Trans. Ind. Appl., 2003,39(3), 752–760.

17 Etkin, B. and Reid, L. Dynamics of flight: stability andcontrol, 1996 (J. Wiley & Sons, New York).

18 Perkins, C. andHage, R. Airplane performance stabilityand control, 1949 (John Wiley & Sons, New York).

19 Raymer, D. P. Aircraft design: a conceptual approach,1999 (AIAA Education Series, New York).

20 Natham, J. K. A code for calculating the nonlinear aero-dynamic characteristic of arbitrary configuration,VSAERO user’s manual version 6.2, 2001 (AnalyticalMethods, Inc., Washington, USA).

21 Matlab-Simulink User’s Manual, copyright 1984–1999(The MathWorks, Inc., Natick, MA, USA).

22 Bisplinghoff, R. L., Ashley, H., and Halfman, R. L.Aeroelasticity, 1996, p. 880 (Dover Science, New York).

23 Dowell, E., Edwards, J., and Strganac, T. Nonlinearaeroelasticity. J. Aircr., 2003, 40(5), 857–874.

24 Romeo, G., Frulla, G., Cestino, E., Marzocca P., andTuzcu I. Nonlinear aeroelastic modelling and exper-iments of flexible wings. 47th AIAA/ASME/ASCE/AHS/ASC SDM Conference, Newport, Rhode Island,USA, 2006, p. 15.

25 Cestino, E. Design of very long-endurance solar pow-ered UAV. PhD Thesis, Department of AerospaceEngineering, Politecnico di Torino, 2006.

26 Hodges, D. H. and Dowell, E. H. Non-linear equationsof motion for the elastic bending and torsion of twistednon uniform rotor blades. NASA TN D-7818, 1974.

27 Crespo Da Silva, M. R. M. Non-linear flexural-flexural-torsional-extensional dynamics of beams – I. Formu-lation. Int. J. Solids Struct., 1988, 24(12), 1225–1234.

28 Sedaghat, A., Cooper, J. E., andWright, J. R. Predictionof non linear aeroelastic instabilities. ICAS Congress,Harrogate, UK, 2000.

29 Romeo, G., Frulla, G., Cestino, E., and Marzocca, P.Non-linear aeroelastic behavior of highly flexibleHALE wings. In Proceedings of 25th ICAS Congress,Hamburg, Germany, 3–8 September 2006, p. 11.

30 Meirovitch, L. Fundamentals of vibrations, 2001(McGraw Hill, New York).

31 Tuzcu, I., Marzocca, P., Cestino, E., Romeo, G., andFrulla, G. Stability, control, and simulation of high-altitude-long-endurance UAVs. 47th AIAA/ASME/ASCE/AHS/ASC SDM Conference, Newport, RhodeIsland, USA, 2006, p. 16.

APPENDIX

Notation

ARht, ARw horizontal tail and wing aspect ratiobht, bw horizontal tail and wing spancmean mean aerodynamic chordCD, CL drag and lift coefficientsN limit load factorPreq required power for the horizontal

flightRe Reynolds number based on the mean

aerodynamic chordSht, Sw horizontal tail and wing areaV platform airspeedWaf, Wtot airframe and total platform weightW wing deflection



Z altitude

a angle of attackhprop propulsion efficiency

Lht, Lw horizontal tail and wing sweepangle

l taper ratior air density



Documents

Design of a High-Altitude Long-Endurance Solar-powered