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Composite Structures 83 (2008) 143–153

Design, manufacturing and testing of a HALE-UAVstructural demonstrator

G. Frulla, E. Cestino *

Politecnico di Torino, Department of Aerospace Engineering, Corso Duca degli Abruzzi 24, 10129 Turin, Italy

Available online 18 April 2007

Abstract

Some main activities have been developed as part of the HELINET research project which was financed in January 2000, by the Euro-pean Commission, (Fifth Framework Program in the IST action), to develop a European project in the field of stratospheric platforms.From the aeronautical point of view the following items were studied: (1) the design of an autonomous high altitude long-enduranceunmanned air vehicle (HALE-UAV) platform capable of remaining aloft for very long periods of time (between 6 and 9 months) usinga solar-powered and fuel cell system and to gain a complete understanding of the feasibility of a near-term aerodynamic HALE concept,in particular as far as stratospheric platforms are concerned (mainly dependent on high efficiency and reliability of solar cells, fuel cellsand electric motors), (2) an evaluation of the production and service costs and an assessment of the safety and regulatory aspects of theplatform, and (3) the manufacturing of a scale-sized technological demonstrator and the execution of static tests on it up to the ultimateload. The first HELIPLAT� (HELIos PLATform) configuration was worked out, on the basis of a preliminary parametric study. Theplatform was a twin-boom tail type monoplane with eight brushless motors, a long horizontal stabilizer and two rudders. A scaled-pro-totype was designed to demonstrate the feasibility of this configuration and to perform some structural static and dynamic tests on it. Themain CFRP structures were manufactured by CASA (Spain): the principal wing and horizontal tail tubular spars, booms, vertical tailspars and some reinforced ribs. These parts were delivered to the Aerospace Engineering Dept. (DIASP) at the Politecnico di Torino(POLITO) and assembled using special joints while some other necessary parts were manufactured by POLITO-DIASP. A parallel activ-ity was performed to define the structural test configurations and structural test frame. The manufacturing activities and the developmentof the structural test system is described in the first part of the paper. Static and dynamic experimental tests were performed in two phases(2003 and 2004) on the prototype and the results of the static tests are presented in this paper and compared with numerical and the-oretical computations.� 2007 Elsevier Ltd. All rights reserved.

Keywords: Carbon-fibre; Assembly; Finite element analysis (FEA); Mechanical testing

1. Introduction

A research was carried out at the Politecnico di Torino,with the aim of designing a HALE-UAV solar-poweredplatform and manufacturing a scale-sized solar-poweredprototype. The first limited financial support was obtainedfrom the Italian Space Agency (ASI) in 1995 [1]. Only asmall part of the research was completed but a great

0263-8223/$ - see front matter � 2007 Elsevier Ltd. All rights reserved.

doi:10.1016/j.compstruct.2007.04.008

* Corresponding author. Tel.: +39 011 5646842; fax: +39 011 5646899.E-mail addresses: [email protected] (G. Frulla), enrico.cesti-

[email protected] (E. Cestino).

amount of experience was gained in this field [1–6]. TheHELINET research project (Network of stratospheric plat-forms for traffic monitoring, environmental surveillanceand broadband services, Coordinator: Politecnico di Tor-ino) has been financed by the European Commission, sinceJanuary 2000, as a part of the 5th Framework Program inthe Information Society Technology Action, to develop aEuropean project in the field of stratospheric platforms.The HELINET project, which involves several Europeanuniversities and partner companies, is based on HALE-UAV HELIPLAT� [6]. The main objectives of the 3-yearproject, from the aeronautical point of view, were:

mailto:[email protected]



144 G. Frulla, E. Cestino / Composite Structures 83 (2008) 143–153

1. To design an automatic HALE-UAV that is capable ofremaining aloft for very long periods of time (about 6–9months) thanks to a solar-powered and fuel cell system.

2. To gain a complete understanding of the feasibility of anear-term aerodynamic HALE concept, especially con-cerning low-Reynolds profiles and propellers.

3. To design the entire advanced composite wing (about75 m long), payload housing, booms and tail structures.

4. To verify the production costs of each platform.5. To manufacture a 1:3 scale-sized technological demon-

strator and perform static tests on it.6. To assess the safety and regulatory aspects of the

platform.

Details of the project can be found in the cited references[5–7], and only a summary is given here. There is currentlya large request for automatic high altitude (17–25 km) fly-ing platforms capable of remaining aloft for very long peri-ods of time [HALE/UAV]. They could play the role ofartificial satellites, with the advantage of being muchcheaper, closer to the ground and more flexible. They couldin fact be self-launched and be easily recovered for mainte-nance, whenever necessary, and could be moved to coverdifferent regions if necessary. The missions of such plat-forms would cover a very large field of applications: atmo-spheric pollution control and meteorological monitoring,real-time monitoring of seismic-risk areas, coastal surveil-lance, telecommunication services such as cellular–tele-phone networks, video-surveillance, photogrammetry,hydrographical monitoring systems, agriculture monitor-ing and so on. Several types of high altitude solar-poweredplatforms (HASP) were designed in the past [1–6,10,11]. Afuel cell system and a brushless motor and solar cell systemare being designed by the Department of Energy and theDepartment of Electrical Engineering at our University,respectively. The aim of this paper concerns point 5 inthe list of the main objectives.

Fig. 1. Heliplat� scaled-size structural

2. Scaled-size prototype design and manufacturing

A scaled-size technological demonstrator was prelimin-ary designed, with the aim of obtaining a structural andtechnological proof-of-concept aircraft [6,8,9,12–14]. Thefinal structural project of the full-scale platform wasdesigned and built with the same technologies and on thebasis of the scaled-prototype test results. Taking the preli-minary HeliPlat configuration [6,7], with its oversized hor-izontal tail span as a reference point in the design of thescaled-prototype, the wing and tail spans were reduced to1:3 of their final HeliPlat dimensions and the chord wisedimensions to about 1:2 , while the thickness and the layoutof the different elements were kept the same as the real con-figuration. The overall architecture of the scaled-size proto-type and particular sections of the wing, horizontal tail,and booms are plotted in Fig. 1. The total aircraft weightis about 312 kg with a wing surface of about 27 m2 and aspan of 24 m. The HeliPlat applied loads, including bothmanoeuvre and gust (resultant bending, torsion momentsand resultant shear), take into account the effective aerody-namic and mass distributions according to the JAR-VLAairworthiness requirements. A maximum design strain levelof 1000 le would guarantee the integrity of the structure upto the ultimate loads. The cross-section of the wing wasdefined in detail according to the preliminary structuraldesign program and FE comparison. The wing and hori-zontal tail spars are made of different materials and havedifferent layouts (±20�/0�n/�45�/core)s in order to intro-duce a level of optimization inside the iterative calculations(Fig. 2). CFRP (M55J/epoxy) data were introduced withreference to a lamina thickness of 0.135 mm and a densityof 1580 kg/m3. The inertial loads (structural and non struc-tural) were estimated and taken into account according tothe preliminary configuration design. An estimate of theexpected wing weight of the single components arereported: covering 10.8 kg, wing main spar 36.5 kg (super-

configuration; construction details.

Fig. 2. Sandwich tube layout definition.

Fig. 3. Main spar special fittings.

G. Frulla, E. Cestino / Composite Structures 83 (2008) 143–153 145

position and reinforcements included), ribs 4.3 kg, rearspar 6.5 kg, and special fittings 14.5 kg.

3. Manufacturing and assembling activity

3.1. Wing: manufacturing and technological feasibility

The cross-section of the wing is plotted in Fig. 1. Themain graphite/epoxy tubular spar should be able to carryall of the shear/bending/torsion loads applied to the wing,while the leading edge and trailing ribs have (in most cases)a profile shape function and were manufactured in verylight rigid foam (Rohacell 51, with a density of 0.051 kg/dm3) bonded to the spar. The leading edge between thetwo booms consists of a hand lay-up glass fibre sandwichpanel. The scaled wing structure can be divided into a11.2 m long constant chord inner-wing and two 6.57 mlong tapered chord outer wings, one for each side. The con-stant chord wing part consists of two inner-wing tubularspar elements, each 4.18 m long, connect to a 4.2 m cen-tre-wing-box. The inner-wing-boxes enter the centre onefor a distance of 0.44 m and are joined by two pairs of bolts(one on the front and one on the rear web, as in Fig. 6). Aseries of specific metal reinforcements were designed inorder to reinforce the joint position between the differenttubular spars and to allow easy assembling and deploymentof the scaled-prototype. Different joint shapes were definedin correspondence to the specific joint position and tubedimension taking into consideration the difference of eachlocal curvature and thickness. The design of these specialjunctions was dedicated to dealing with the problem ofbearing in CFRP. A good result was obtained, as shownin Refs. [4,10,11] . A typical joint system is reported inFig. 3. The joints shown in Fig. 3 are in the connectionzone between the centre-wing-box and inner-wing-box.Two 6.57 m long parts make up the tapered sections, whichare connected in a similar way. The main spar is composedof sandwich reinforced tubes made of M55J graphite/epoxypre-preg tape and Korex or Nomex honeycomb materials.The tools were made of aluminium and were slightly coni-cal to allow for demoulding. Tools for an autoclave cannot

be massive and the tool was positioned on a flat beam toavoid deformation of the parts due to deflection. Differentsteps of the process, such as the core lay-up positioning and0� layering were performed by CASA and are explained inRef. [15]. Each tubular spar was autoclave cured at 120 �Cand 0.3 MPa pressure in a single-cure cycle. The effectiveand local dimensions and lay-up of these elements are plot-ted in Fig. 4.

Some of the ribs were bonded to the tubular spars in theindicated positions, particularly in correspondence to thefittings and engine mounts. Carbon-fibre-reinforced foamribs were mounted onto some sections, as plotted by ‘‘rr’’in Fig. 1, while special carbon-fibre-reinforced plastic(CFRP) honeycomb ribs were mounted in correspondenceto the fittings between the several wing spars and, aboveall, in correspondence to the booms. All of the ribs werebonded to the spars with a micro-sphere-reinforced glassepoxy. The CFRP wing-boxes were manufactured usingM55J graphite/epoxy pre-preg tape. This high-modulusgraphite fibre greatly increases the wing flexural stiffnessand reduces the wing tip deflection at the maximum appliedlimit load. The lamina used to design the wing-boxes havethe following main mechanical properties:

E1 ¼ 279:3 GPa; E2 ¼ 5:84 GPa; G12 ¼ 4:05 GPa, m12 ¼0:36; rt1R ¼ 1036 MPa; rc1R ¼ �381 MPa; et1R ¼ 0:37%;and q ¼ 1:58 kg=dm3. The lay-up of each portion of the sin-gle tube was chosen by introducing the previously mentionedmechanical properties into the structural design computerprogram and it can be seen in the corresponding drawings.

Fig. 4. Heliplat� wing main spar dimensions.


3.2. Horizontal and vertical tail: manufacturing and

technological feasibility

The scaled horizontal tail structure consists of a 11.2 mlong constant chord element (Fig. 1). It was manufacturedin three pieces: the 4.0 m long inner part and two 3.9 mlong outer parts. The cross-section consists of a mainCFRP sandwich tubular spar that can carry all the shear/bending/torsion loads applied to the horizontal tail. A sec-ondary CFRP rear spar was positioned aft of the main oneso as to have a suitable structural support for the elevatorsystems. The leading and trailing ribs were manufacturedin very light rigid foam and were bonded to the box. Thescaled double vertical tail structure consists of a 1.685 mlong tubular spar for each tail, with an 88 mm externaldiameter circular cross-section. The two spars were madeof a CFRP sandwich construction, using M55J graphite/epoxy pre-preg tape and Nomex honeycomb.

3.3. Boom: manufacturing and technological feasibility

The scaled boom-structure consists of a 5.98 m longtapered tubular member, with a circular 235 mm diametercross-section near the wing-box and a 150 mm diameter

Fig. 5. Wing–boom

cross-section in correspondence to the vertical tail junction.The 235 mm diameter cross-section is constant up to a dis-tance of 2.55 m from the wing-box. The two booms weremade of a CFRP sandwich construction, using M55J graph-ite/epoxy pre-preg tape and Nomex honeycomb. A specialconnection, based on the introduction of a structural rib,was designed to join the boom to the wing-box. The junctionis reported in Fig. 5 and is made up of three elements: the first(in correspondence to the wing-box spar) is a specificallyshaped plate connected to the spar with specials joints (asreported); the second (positioned at 0.7 m aft of the wing-box spar) consists of a pin-bolted connection; and the thirdwas positioned in correspondence to the rear spar.

3.4. Verification of the delivered CFRP parts and metal

fitting

A preliminary verification activity was performed on theCFRP parts and delivered metal fittings. The differenttubular CFRP parts received from CASA were subjectedto a specific identification phase, in order to check theexternal dimensions (overall, diameters, thickness etc.)and to verify the correspondence of the joints and metal fit-ting positions. A series of measurements was performed.

special fitting.

Fig. 6. Assembling activity.


The diameters and lengths were determined, compared tothe expected ones, and used in order to define the connec-tion joints and their bonding. Some details that did not cor-respond to the delivered drawings were detected and aresummarized as follow: (1) the reinforced ribs did not havea perfect correspondence to the drawings; (2) the aft part ofthe booms was missed and a new joint position, that wasdifferent from that of the design, was determined beforeassembling. The vertical tail–boom-join was moved backby about 230 mm and the main wing–boom-join by about300 mm in order to maintain the focal distance between thewing and the horizontal tail as unchanged as possible; (3)the tubular spars showed some wrinkles and (4) the holesin the horizontal tail external tubes were rotated by 90�.In this way, the spar caps become positioned in the neutralaxis and thus change the structural behaviour of the taildrastically. All the weights of the different parts were deter-mined to compare them with the expected ones and todefine a well behaved analytical model for the static testcomparison. The measured thickness showed a slightly dif-ferent mean ply thickness from the theoretical value of0.135 mm. The theoretical model was modified introducingthe measured lamina thickness and an M55j density of1680 kg/m3.

3.5. Scaled-size prototype assembly

CASA Espacio Spain – EADS made each of the CFRPelements of the complete structure (wing, horizontal andvertical tail tubular spars and reinforced ribs, booms), plusthe metal fittings by machine manufacturing. The Depart-ment of Aerospace Engineering (Politecnico di Torino)assembled the various structural components (wing, hori-zontal and vertical tails, booms), as well as the entire air-craft. The assembling was performed by the Italiancompany ARCHEMIDE Advanced Composites, under

the supervision, and with the collaboration of, the platformdesign and testing group at the Politecnico di Torino. Afterall the different parts had been positioned in their properplaces so as to define the bonding tree for the junction, theywere bonded to the main wing spar (Fig. 6). A rear tubularspar was manufactured by ARCHEMIDE to complete thewing and tail structures. A reinforced leading edge wasdesigned and manufactured for the centre part of the wingto increase torsional stiffness (Fig. 6). The lifting surfacewas then completed through the positioning of a coveringskin.

4. Static test activities

Static and dynamic tests are required by airworthinessregulations during an airplane certification process. In thepast, static tests were performed by applying wing loadsusing sand bags or metal bricks. If a high quality level oftest is required, or load conditions are very complex toreproduce, this method presents many problems and it ispreferred to use alternative solutions. During the testingphase, the airplane is usually supported on fixtures, andhigh-performance servo hydraulic actuators are used topush and pull on certain areas. Strain gauges are attachedto these areas and the information is passed via cable tocomputers. Static and fatigue tests can be performed apply-ing loads by means of rigid whiffletrees or enforced dis-placements at the hinges to simulate, for example,horizontal stabilizer deflections. The main problem inreproducing a load condition on an isolated part of an air-craft concerns the realization of the constraint between theairplane and the fixed test structure. The connection mustreproduce the real connection as much as possible. Statictests were carried out on the whole Heliplat� scaled-sizeprototype. A static test on the complete model was consid-ered to be more meaningful than several tests on each ele-

Table 1Load factors

Regulation Max value Min value

JAR-VLA nmax ¼ 3:8 nmin ¼ �1:5RAI-UAV nmax ¼ 3 nmin ¼ �1


ment of the structure because several interaction effects canarise when the different elements are tested whileassembled:

(1) The booms were simultaneously subjected to loadsapplied by the wing as well as by the horizontal andvertical tails. It would be very difficult to isolate theformer without introducing some simplifications,which do not represent the proper boundaryconditions.

(2) The highest applied loads on the wing derive from thehorizontal tail during a pull-up manoeuvre. (D-D1)

(3) The behaviour of the horizontal tail is greatly influ-enced by the constraints on the vertical tail.

Mechanical equipment was designed and manufacturedto perform a shear–bending–torsion test on the completescaled-size prototype and verify the theoretical behaviour.A steel supporting structure was defined and manufacturedby POLITO in order to sustain the scaled-prototype. Thetree-beam systems and the hydraulic jack are shown inFig. 7. The purpose of the tree-beam systems is to reachthe defined loading conditions during tests. Beams wereconnected by steel chains to avoid spurious loads duringthe test. Two trolleys were used to support the tree-beamsystem and the position was adapted according to thescaled behaviour. A dummy fuselage was designed to apply

Fig. 7. Testing system.

Fig. 8. Theoretical and appli

the expected loads to the centre of the model. A generalview of the testing system is reported in Fig. 7. The distri-bution of the load, according to the described behaviour,was obtained using a hydraulic jack and a system of leversconnected to eight loading stations on the wing. Straingauges and transducers were mounted along the wingand the horizontal tail main spars, to estimate deforma-tions and wing deflection. JAR-VLA 333 concerns on-flightenvelope definition. Compliance to the strength require-ments of this subpart must be shown at any combinationof airspeed and load factor on and within the boundariesof a flight envelope that represents the envelope of theflight loading conditions specified by the manoeuvringand gust criteria of sub-paragraphs (b) and (c) of this par-agraph, respectively. The selected flight airspeeds aredefined by JAR-VLA 335. Limiting manoeuvring load fac-tors are defined by JAR-VLA-337 (a) and (b). The scaled-prototype design was drawn up according to the ItalianRAI-UAV 337: nmin and nmax are shown in Table 1. ForPoint A1 and A tests, the horizontal tail balancing loadnecessary to maintain equilibrium in any specified flightcondition with no pitching acceleration, was computedaccording to JAR-VLA-421 and simulated through weightspositioned on the boom. During the shear/bending tests onthe wing, in stationary flight conditions at the A1, A and Dpoints, the loads on the booms, on the jack and the reac-tion on the beams, vary linearly. The ultimate load was cal-culated as 1.5 times the D load point of the flight envelopediagram (representative the limit load condition). The tree-beam system was designed to obtain the desired bendingmoment behaviour. The A1, A and D condition loads weredifferent but still proportional, so the same tree-beam sys-tem can be used to simulate them. The theoretical and sim-ulated shear and bending moment acting on the scaledwing in the D condition are reported in Fig. 8. JAR-VLA423 indicates that each horizontal tail surface must be

ed loads at D condition.


designed to manoeuvre loads imposed under several condi-tions. The first step is to compute the horizontal tail loadincrement (Eq. (1)), due to a sudden deflection of the eleva-tor which would cause the normal acceleration to changefrom an initial value to a final value. The aeroplane is ini-tially in level flight, and its altitude and airspeed do notchange; the loads are balanced by inertia forces and theaerodynamic tail load increment is given by :

DLHT ¼ DnMgX CG

lt

� SHT

SaHT

a1� de

da

� �� q0

2

SHTaHTlt

M

� ��

ð1Þwhere Dn is the load factor increment, M is the mass of theairplane in kg, g is the acceleration due to gravity, XCG isthe longitudinal distance of the aeroplane centre of gravityaft of the aerodynamic centre of the aeroplane less the hor-izontal tail (m), S and SHT are the horizontal tail area andwing surface, respectively (m2), a and aHT are the slope ofthe wing and the horizontal tail lift curve per radiant, de

da isthe rate of change of the downwash angle with the angle ofattack, lHT is the tail arm and q0 is the density of air at sea-level (kg/m3). In the case of the scaled-prototype, the high-est applied loads on the wing derive from the (D-D1)manoeuvre when the airplane is flying at VD and experienc-ing a load factor increment of Dn ¼ �2 passing from a Dcondition where the load factor is 3, to D1 where n = 1.The theoretical and applied loads on the leverage systemfor the D-D1 manoeuvre are reported Fig. 9. In (D-D1)manoeuvring conditions, the wing loads are not propor-tional to the stationary flight conditions, but they can be

Fig. 9. Theoretical and applied

Fig. 10. Wing strain gauges distributio

still simulated via the same tree-beam system used in theprevious cases.

5. Static test results

A first experimental test activity was performed in 2003.Various tests were considered in different conditions: wingbending/shear in the D condition, horizontal tail bending/shear in the D-D1 condition and wing bending/shear/tor-sion in the D-D1 condition. The maximum load reachedduring this first test was reduced in order to perform thedifferent conditions and to avoid unexpected failure duringthe tests. In so doing, various conditions had to be com-pared with the calculations to validate the preliminarydesign program and numerical models. This activityproved useful for the full-scale design. The bending teston the wing was conducted loading the wing with thedescribed system, until the A load condition of themanoeuvre diagram (cruise speed = 58 km/h andn = 2.5 g), and 50% of the manoeuvring load (D-D1: fromD point where V = 82 km/h and n = 3 to D1 whereV = 82 km/h, n = 1 and Dn = �2) were reached. Straingauges and transducers were mounted along the main sparof the wing in order to estimate the deformations and thewing deflection (Fig. 10). The strain gauges were mountedlongitudinally in correspondence to the main spar caps: theodd numbers indicate the SG located along the upper cap,while the even ones are on the lower cap. The strain anddisplacement results for the A condition are reported inFig. 12a and b. A good correlation between the tests, the

loads at D-D1 condition.

n and wing transducers position.


theoretical and the FE analysis was obtained. Static testswere also performed in manoeuvring conditions withincreasing loads until 50% of the manoeuvring D-D1 con-dition was reached. The experimental deflection is shown inFig. 11 compared with the FEA deflection in the D-D1condition, and a maximum tip deflection of about764 mm resulted. Manoeuvring is a combined bending/tor-sion test. During this test, the wing torsion moment due tothe horizontal tail load was simulated by applying a weight

Fig. 11. FE and experimental

Fig. 12. 2003 t

Fig. 13. Horizontal tail strain gage

of about 12.5 kg on every boom at a distance from the wingspar of about 2780 mm. The results, in terms of strains anddeflection, are also reported in Fig. 12c and d. A bendingtest on the horizontal tail was also performed, with a max-imum load of 100 kg (50% D-D1 condition) along the HTspan. In order to perform the test, the HT was suspendedfrom the extremities. The deformations were determinedfrom strain gauges mounted longitudinally along the HTmain spar lower cap (see Fig. 13). The deflections were

results at D-D1 condition.

ests results.

s position and loads position.

Fig. 14. HT displacements and strains comparison.


measured using a metric stick (Fig. 13). The load stationsare displayed in Fig. 13b. Each station was loaded with10 kg , using a local suspended weight . The displacementand strain comparison is shown in Fig. 14. The strainresults show an increase in strain for the outer HT spar.Some important differences from the design were in factdetermined in that position consisting of the wrong posi-tioning of the spar itself due to the wrong location of thefitting holes as previously indicated. The load was notincreased any further in order to prevent reaching failurein the tail.

A second test activity was performed in 2004. Thetested flight conditions were: 100% D-D1 manoeuvre, ulti-mate load at the D point (computed as 1.5 the D Pointlimit load), torsion test and static failure test. The D-D1test showed a good repeatability compared to the 2003test. Slight differences occurred which probably due to adifferent application of the 154 N boom weight which sim-ulated the HT load in this condition. During the 2003tests, the boom weight was applied to the boom at a dis-

Fig. 15. 2004

tance of 2780 mm from the wing–boom joint location.During the 2004 tests, a different solution was adoptedin order to have a safer boom load carriage zone. In2004, the boom weight was applied directly to thewing–boom joint location using a back-pulley system.At 100%, the D-D1 load resultant was about 4159 N.The results are reported in Fig. 15a and b. The D pointtest was performed until ultimate load where the totalhydraulic-jack load was about 3658 N and the boomweight about 1130 N. In this condition, the structurewas able to support the ultimate load without failure.The maximum deflection and strains at the ultimate loadare reported in Fig. 15c and d. The strains on the maintubular spar computed by FEA and the deflection, arealso plotted in Fig. 16. A torsional test was performedin order to test the wing–boom joint. The fuselage wasclamped to the load structure and a load was applied,step-by-step, using several metal bricks in the boom-verti-cal junction zone (Fig. 17). The maximum applied tor-sional moment was 650 kg m. The torsion angle and a

results.

Fig. 16. Main spar strain and deflection (D point – limit load) (FEA).

Fig. 18. 2004 failure test.


comparison with the FEM analysis and test are reportedin Fig. 17. The junction showed good behaviour, butthere was a certain difference between the FEM and theexperimental results at a high torsion angle due to theconstraint. The last test was carried out to test the proto-type up to failure. The test was performed without apply-ing boom weights, starting from the D-D1 condition andincreasing loads in two phases: a first phase up to a loadof about 630 kg, where the first signals of failure had beendetected, and then in a second phase, the load wasreduced to 500 kg and the final increase was up to850 kg where an explosive failure occurred in the centre/inner-wing-box connection zone. The model at its maxi-mum experimental wing deflection and details of the fail-ure zone are shown in Fig. 18.

Fig. 17. 2004 t

6. Conclusions

The structural design and assembling of the scaled-pro-totype of the HELIPLAT UAV is presented. The manufac-turing of the main tubular spars and metal fitting wasperformed by CASA while the advanced spar joints, specif-ically designed for the project, were developed and manu-factured by POLITO. Another important contributionwas made by the Politecnico di Torino (DIASP) and itssubcontractor with the assembling of several parts of theprototype and the manufacturing of all the different partsthat complete the aircraft, as shown in the figures. Themanufacturing of the testing facility and the execution oftest activities were also made by the authors, includingthe positioning of the prototype and all the sensors andequipment necessary for the test. The static tests performedat points A and D of the flight envelope showed a fair cor-respondence with the numerical and preliminary analysisconfirming the good evaluation of the structural behaviourof the system. Some differences that were encountered

orsion test.


could be due to some defects and imprecision in the man-ufacturing of the main tubular spars. A subsequent torsiontest on the main wing spar was also performed in order tocompletely validate the spar itself and the special advancedjoint configuration. The final failure test was carried out toevaluate the margin from the ultimate design load condi-tion, and satisfactory behaviour was obtained.

Acknowledgements

The authors would firstly like to thank Prof. Giulio Ro-meo for his essential contribution during all the projectphases. Special thanks also due to Eng. Guido Corsinoand Eng. Fabio Borello for their important contributionduring the assembling and test activity.

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