Design of an Unmanned Aerial Vehicle for Ecological Conservation

American Institute of Aeronautics and Astronautics

1

Design of an Unmanned Aerial Vehicle for Ecological

Conservation

Elsa M. Cárdenas *

Universidad Nacional Experimental Politécnica de Fuerza Armada, Caracas, 1060, Venezuela

Pedro J. Boschetti†, and Andréa Amerio

‡

Universidad Simón Bolívar, Caracas, 89000-1080-A, Venezuela

and

Carlos D. Velásquez §

Universidad Nacional Experimental Politécnica de Fuerza Armada, Maracay, Venezuela

The exploitation of petroleum can cause serious environment problems when oil leakages

occur on the marine or lake surfaces. The constant vigilance over exploitation areas helps to

minimize the adverse impact of such accidents by means of early detection. This article deals

with the activities carried out at present in order to create an unmanned aerial vehicle

designed to patrol the petroleum exploitation zones. Among these activities the preliminary

design of the aircraft, the structural design of a prototype capable of accomplishing the

assigned mission, and the aerodynamic optimization of such a design, are worth mentioning.

A monoplane airplane, twin-boom configuration airplane, with a partially metallic structure

was designed. The aerodynamic optimization process was realized applying theoretical and

experimental methods. In conclusion, the designed vehicle will prove to be satisfactory for

the mission for which it was created, and to be used as a tool for future research.

Nomenclature

CL = Lift Coefficient

CL/CD = Aerodynamic efficiency

MAV Micro Aerial Vehicle

UAV Unmanned Aerial Vehicle

UCAV Unmanned Combat Aerial Vehicle

I. Introduction

T is a well known fact that, in recent years, advanced technology has been applied to the development of aircraft

able to operate without crew, under extreme or limiting conditions for the human being. The use of this

technology has become essential in many countries around the world, where such vehicles are customized and

equipped to carry out specific missions, for instance: aerial surveillance, fumigation of large crop fields,

measurement of atmospheric conditions, among other.

In Venezuela, nowadays, there are certain activities in which an unmanned aerial vehicle system would reduce

operative and maintenance costs, such as environment protection, mainly focused in the detection of petroleum

spillage in offshore facilities. Specifically, the petroleum exploitation in the Lake of Maracaibo basin has been a

source of pollution since the exploitation activities were initiated in the zone. For this reason, Petroleos de

Venezuela, S.A. the state owned petroleum company, maintains an aerial surveillance, which is carried out by means

of manned helicopters, for the early detection of oil leakages from exploitation derricks, oil transporting pipelines

* Instructor Professor, Department of Aeronautical Engineering, 1060, Chuao.

† Research Graduate Student, Direction of Investigation, 1080, Valle de Sartenejas, AIAA Student Member.

‡ Assistant Professor, Department of Industrial Technology, 1080, Valle de Sartenejas.

§ Undergraduate Student, Department of Aeronautical Engineering, Boca de Río.

I


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and other facilities. These can only operate during day time and under good climatic conditions, and their activity is

relatively expensive.

Based on these facts, the need to achieve the complete design of an unmanned aerial vehicle to carry out

ecological conservation missions, day and night, in almost any atmospheric condition, having as first aim the

construction of an early prototype airplane, capable of complying with the expected flight performance, without

advanced electronic devices and, later, the design and development of control, navigation, guidance and positioning

systems needed by the vehicle to carry out the required missions autonomously.

The objective of this paper is to show the activities developed for the design of the Unmanned Airplane for

Ecological Conservation, ANCE, (for its Spanish name). The project was initiated in the year 2002 with the

preliminary design of the aircraft,1 as part of a joint project carried out between Universidad Nacional Experimental

Politécnica de la Fuerza Armada (UNEFA,) and the Universidad Simón Bolívar (USB). Until now the structural

design for a prototype at full scale and an aerodynamic optimization process has been accomplished

II. Preliminary Design

The preliminary design was based on the methodology shown on Ref. 2, with which it was possible to develop a

program consisting of five phases to design an aircraft with a maximum take-off weight of less than 500 kg.

In order to determine the characteristics of the design, a study of the latest developments in the field of

unmanned aircraft technology was conducted.3 These were classified according to their mission and operation range

in three main categories: MAV, UAV, and UCAV. The UAV were divided into three groups: local, regional, and

endurance. The locals have a maximum range of 50 km, the regional up to 200 km, and the endurance exceeding

200 km.

Based on the designated mission, the physical, technical characteristics, and the performance of the local UAV

were considered, and the approximate data of the aircraft being designed were established accordingly. This will

serve as a reference for successive iterations during the preliminary design.

The first step consists of statistical and analytical calculations for the first estimation of the maximum take-off

weight, wing surface, maximum necessary power at sea level and the selection of the powerplant. In the second step

the final calculations for the maximum take-off weight, wing surface and maximum necessary power at sea level are

made and the use of the above mentioned powerplant was confirmed, and the diameter and the specifications of the

propeller are determined. The third step refers to the wing, determining its location with respect to fuselage, its

planform, dimensions, span and chord, airfoil section and aerodynamic characteristics. The characteristics and

dimensions of the vertical and horizontal stabilizers are determined and calculated on the fourth step. In the fifth step

the study of the position of the horizontal stabilizer with respect to the wing stream, and the definition of the

fuselage shape are made, and using the data obtained in these five steps was possible to draw the general blueprints

of the aircraft.

When the iterative process was finished the design of a monoplane, high wing, twin-boom, and double vertical

stabilizer configuration aircraft, with a length of 4.74 m, wingspan 5.27 m, height 0.93 m, and wing surface of 2.899

m2, was accomplished. The chosen configuration should offer greater stability in all directions. It maximum take-off

weight is 182.06 kg, empty weight 123.73 kg, payload weight 40 kg, and maximum fuel weight 18.33 kg. The

aircraft has an AMW Cuyuna 460 FE-35, of 26.10 kW and 6000 rpm engine located at the fuselage tail, attached to

a Clark Y 5868-9, twin bladed propeller, of fixed pass, with a diameter of 0.915 m. The payload consists of a high

resolution, rotating, dual camera, with capacity for day vision and infrared night vision, with real time transmission,

which is located forward to the wing, attached under the fuselage. Figure 1a) shows a sketch of the design and Fig. 2

the internal vehicle distribution.

Depending on the mission to be accomplished by the aircraft it should be stable or unstable. In view of this,

certain measures were taken in the design steps, such as the wing position with respect to the fuselage, type of wing,

gravity center position, aircraft configuration, and vertical stabilizer size, among other. When the preliminary design

was finished it was necessary to make a theoretical study of the static stability that the aircraft will have, as well as

of the performance that it will develop when its main control surfaces are deflected. For this purpose the

methodology expressed in Ref. 4 was applied. Table 1 shows the results demonstrating that the aircraft is

longitudinally, directionally, and laterally stable.

Other important fact to take into consideration after finishing the preliminary design is to determine the

performance of the aircraft. By means of a detailed theoretical analysis5 the take-off, ascent flight, descent flight,

leveled flight and landing conditions were estimated, based on the available power and necessary power data. The

results are shown in table 2.


3

As graphically shown in Fig. 3, this airplane should be able to take-off from a short asphalt airstrip, or from a

roughly prepared grass or sand airstrip, to climb in a spiral flight to an optimum altitude of 2500 m required by its

mission, which allows it to have a large visual range, to carry out the surveillance in petroleum extraction areas, over

towers, derricks, crude oil transporting pipelines and other similar facilities, for approximately 3 hour time periods.

The camera equipped with an infrared sensor should be able to detect polluting substances on the water from the

flight altitude, which will be possible to be observed by the ground staff in real time, allowing them to take the

corresponding measures immediately. The aircraft shall then land on an airstrip and be made ready for its next

mission.

III. Structural Design

The structural design of the aircraft was conceived based on the data obtained from the preliminary design.1 In

order to reduce the weight of the structure; the width of the upper part of the fuselage was reduced from 0.47 m to

0.37 m. As a result of this modification, the frontal geometric section of the fuselage is trapezoidal instead of

rectangular, and the geometric wingspan was reduced to 5.18 m. This was the first design variation and it was called

ANCE-X.1, being the first design ANCE-X.0. On Fig. 1a) and 1b) may be seen, comparatively, the frontal and

sideway views of both versions. The main purpose of the structural design is to obtain a quick, efficient, economical

configuration, for easy construction and maintenance, and appropriate to carry out the aircraft specific missions, in

compliance with standard international regulations in the design of aircraft and air space vehicles.

In order to accomplish this, a modular design methodology was developed. The first stage comprises the

collection and analysis of previous data obtained from the preliminary design,1

and the elaboration of the

corresponding modification, such as was expressed above. The second stage comprises the determination of the

loads to be supported by the aircraft through all its operations in air and land, as outlined in Ref. 6-9. The third stage

consists of the selection of materials to be used for the construction of each part and component of the aircraft, and it

was carried out in a totally empiric way, based on the direct observation (on the field) of materials used in similar

aircrafts (in mass, dimensions, types of mission, operation regime, etc.). The 2024-T3 aluminum alloy was used for

parts with main structural importance such as the frames, spars, ribs, stiffeners and skin of the aircraft. Likewise, the

4130 steel alloy was selected for special applications, such as landing gear components and the powerplant support.

For the cargo and service doors, control surfaces, the nose of the aircraft and fairings, the use of composite materials

was chosen, specifically epoxy resin reinforced fiberglass. The fourth stage specifically consists of the structural

calculations of the aircraft. For this purpose different design methods and theories for aircraft and aerial vehicles

were used, as shown in Ref. 6-8, 10.

Specifically, a semi-monocoque structural configuration was selected, because it seemed to be the most

appropriate for this particular air vehicle. The wing was designed reducing the cross section area of its structural

elements alongside the wingspan in order to obtain a lighter and more efficient configuration. The calculation of

spars and stiffeners was made applying the compression member design theories, using the Asymmetrical Flexure

Theory. The frames and bulkheads were designed by applying statistically indeterminate structures analysis and

virtual work methods. The skin was designed considering the predominant effect of shear over the surfaces and,

consequently, the shear stress equations derived from the asymmetrical flexure theory. In addition, the compression

stresses and the combined compression, and shear stresses are considered

With respect to the torsion, the wing design was made considering the effects of the aerodynamic moments

produced, due to the corresponding asymmetrical wing cross section, the torsion loads transmitted by the tail booms

and the contribution of the lateral control surfaces (ailerons), when they are deflected. For this purpose the use of a

central wing torsion box, comprised by the main spar, rear spar, and the skin between them, was taken into

consideration, and the corresponding calculations were made in accordance with the methods outlined in Refs 6-8.

The powerplant support, the tail booms, and the landing gear components, although they are not considered as

part of the structure as such, but rather as independent systems, were also designed considering their effects on the

main structure. In the case of the tail booms and landing gear components, these were designed taking into

consideration the resistance of tubular or prismatic elements, subject to the combined flexion, shear and torsion

loads in the three dimensions, using the design of mechanical elements and compression member theories. For the

design of the powerplant support are also considered the truss member structures design theory and the statically

indeterminate structures theory. The different stress distribution due to different loads, are determined and, by means

of the superimposing method, the most critical effects are obtained. Then, by means of the mechanical elements

design theories, specifically the Maximum Shear Stress Theory and the Deformation Energy Theory, the minimum

allowable diameters and thickness of the components are determined.


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The powerplant support , being a complex reticular type structure, which in addition, is subject to a wide range

of special loads, such as the gyroscopic loads of the propeller shaft, is designed based on the theory of statically

indeterminate structures and virtual work methods, using in addition numerical computation methods to solve the

multiple unknown variables. Tables 3-6 show the results obtained, and Fig. 4 shows the resulting structural

configuration.

IV. Drag Cleanup and Aerodynamic Optimization

The objective is to evaluate the design aerodynamic characteristics and to optimize the aerodynamic efficiency in

cruise flight. This process consisted of the evaluation of the drag coefficient and the lift coefficient using theoretical,

analytical, numerical and experimental tools, and then, based on literature,11-13

to make changes in the aircraft shape

to reduce the drag and increase efficiency.

The theoretical procedure started with the estimation of the aircraft lift coefficient by means of a numerical code,

specially created for this particular case, based on the Prandtl’s Lift Line Theory,14

and the computing methods

developed by Glauert and Lotz.8 After this, the profile drag and the induced drag were estimated. The estimation of

the profile drag was carried out by means of a calculation procedure expressed on Ref. 15, which is based on

analytical and statistical data.

The experimental procedure was carried out by means of wind tunnel tests. A subsonic wind tunnel of close

circuit, close throat Rollab SWT-009, located at the facilities of the Department of Aeronautical Engineering of the

UNEFA was used. This unpressurized atmospheric wind tunnel has a velocity range between 14 and 47 m/s, and has

a square test section of 0.32m x 0.32m. The tests were carried out with models exactly alike to the sketch, at scale

1:20.2, made of fiberglass and polyester resin, one of them can be seen in Fig. 5. The tests were made at different

speeds and angles of attack. Buoyancy, blockage, tare-and-interferences, and scale effect corrections were applied.16

From these studies resulted the location of fairings over the landing gear, and an aerodynamic twit between wing

stations 2.32 and 2.40, measured in meters, which helped to reduce the induced drag and improve wing efficiency.13

Figures 1b) to 1d) are shown all the successive changes suffered by the design during this process (ANCE X-2, and

X-3), compared with ANCE X-1, and Fig. 6 shows ANCE X-3 in a three-dimensional drawing. Figure 7 shows the

lift curve as a function of drag coefficient for versions X-1, X-2, and X-3, obtained by means of wind tunnel tests.

See Ref. 17 for a full description of theoretical calculations, experimental procedures, and discussion of the results

related with the aerodynamic optimization process

V. Conclusion

This paper presents an overview of the critical features of the ANCE development program to date. By means of

an analytical and statistical study a versatile design was obtained, which is expected to comply appropriately with its

mission. The obtained structural design gives a preliminary configuration that satisfies the initial requirements,

which however could be subject to later modifications to improve performance. The drag cleanup process, carried

out using tools such as theoretical analysis and wind tunnel testing, derived in successive improvements to reach a

sufficiently efficient version, capable of carrying out the mission.

Finally, the vehicle is expected to be a very valuable tool, both to be used as test bench for advanced research

studies as well as for the specific tasks for which it was conceived.

Acknowledgments

The authors wish to acknowledge the financial support of Direction of Investigation, Universidad Simon Bolivar,

Caracas, and FUNDACITE Aragua, Maracay, both in Venezuela.

We also thank to Prof. Eng. Ganimeh Diaz, from Department of Aeronautical Engineering of Universidad

Nacional Experimental Politécnica de la Fuerza Armada, Maracay, for allowing the use of the aerodynamic

laboratory facilities.

References 1 Boschetti, P. and Cárdenas, E., “Diseño de un Avión no Tripulado de Conservación Ecológica,” Theses, Department of

Aeronautical Engineering, Universidad Nacional Experimental Politécnica de la Fuerza Armada, Maracay, Venezuela, 2003. 2 Salmerón, R., and Salvatore, M., “Diseño de un Blanco Aéreo para Evaluar los Sistemas Automáticos de Tiro de las

Fragatas Tipo ‘Mariscal Sucre’,” Theses, Department of Aeronautical Engineering, Instituto Universitario Politécnico de la

Fuerza Armada Nacional, Maracay,Venezuela, 1987. 3 Wallops Flight Facility - Unmanned Aerial Vehicles web site, “Unmanned Aerial Vehicles Web Site,” [online database],

URL: http://uav.wff.nasa.gov/Main.cfm [cited 21 October 2002].


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4 Nelson, R. C., “Static Stability and Control,” Flight Stability and Automatic Control, 2nd ed., McGraw-Hill, Boston, IL,

1998, Chapter 2. 5 Ordoñez, C., Aerodinámica, tomo IV, edited by UTEHA, México D.F., 1963, pp. 278-369. 6 Bruhn, E.F, Analysis Design of Flight Vehicle Structures, 2nd ed.,. S.R. Jacobs & Associates, Inc., Indianapolis, IN, 1973,

Chapters A, B, D. 7 Niu, M. C-Y., Airframe Structural Design, 2nd ed., Hong Kong Conmilit Press LTD., Victoria, Hong Kong, 1999, Chapters

3-11,14,16. 8 Peery, D., Aircraft Structures, edited by Mc Graw-Hill Book Company, New York, 1950, Chapters 1-17. 9 The Federal Aviation Administration, “Part 23 - Airworthiness Standards: Normal, Utility, Acrobatic, and Commuter

Category Airplanes,” Federal Aviation Regulations [online database], URL:

http://www.airweb.faa.gov/Regulatory_and_Guidance_Library/rgFAR.nsf/MainFrame?OpenFrameSet [cited November 2004]. 10 Shigley, J. and Mischke, Ch., Diseño en Ingeniería Mecánica, 5th ed., Mc Graw-Hill / Interamericana de México, S.A. de

C.V. México D.F., 2001, Chapters 1-3,5,6,8,9,18. 11 Bushnell, B. M., “Aircraft Drag Reduction – a review,” Journal of Aerospace Engineer, Vol. 217, No 1, 2003, pp. 1-18. 12 Coe, P., “Review of Drag Cleanup Test in Langley Full – Scale Tunnel (from 1935 to 1945) Applicable to Current General

Aviation Airplanes,” NASA TN D-8206, Jun. 1976. 13 Phillips, W. F., Fugal, S. R. and Spall, R., “Minimizing Induced Drag with Geometric and Aerodynamic Twist, CFD

Validation,” AIAA Paper 2005-1034, Jan. 2005. 14 Prandtl, L., “Applications of Modern Hydrodynamics to the Aeronautics,” NACA TR-116, 1921. 15 Hoerner, S. F. Résistance á L’avancement dans les Fluides, edited by Gauthier Villars Editeurs, Paris, France, 1965,

Chapter XIV. 16 Rae, W. H., and Pope, A., Low-Speed Wind Tunnel Testing,2nd ed, edited by Wiley Interscince Publication, New York,

1984, pp. 198-205, 419-424, 457-464. 17 Boschetti, P. J., Cárdenas, E. M., and Amerio, A., “Aerodynamic Optimization of an UAV Design,” AIAA Paper 2005-

7399, Sep. 2005.


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Table 1 Static Stability.

Longitudinal Stability

Pitching moment of wing -0.00486

Pitching moment of tail -0.145715

Pitching moment of fuselage 0.0813565

Pitching moment of airplane -0.10007

Directional Stability

Yawing moment of interaction wing – fuselage -0.01881

Yawing moment of tail 0.01969118

Roll Stability

Rolling moment -0.0004

Roll control moment -0.00153302

Table 2 Performance characteristics.

Cruise speed 46.77 m/s

Stall speed at sea level 28.66 m/s

Stall speed at 2500 m 25.84 m/s

Maximum speed at sea level 62.91 m/s

Maximum speed at 2500 m 68.93 m/s

Range 344.35 km

Maximum Endurance 11h 12’ 44’’

Service ceiling 3625.46 m

Maximum rate of climb at sea level 4.21 m/s

Maximum rate of climb at 2500 m 1.16 m/s

Take off field length 377.3 m

Landing field length 428.5 m

Table 3 Wing design results.

Main spar

(30% of the chord)

Rear spar

(80% of the chord) Station,

m Thickness, m Thickness, m

Skin

maximum

thickness, m

0 0.0022352 0.015875 0.0014224 0.0667 0.0022352 0.015875 0.0014224

0.2056 0.0022352 0.0115824 0.0014224

0.3445 0.0004064 0.009144 0.001016

0.37225 0.0004064 0.009144 0.001016

0.4 0.0004064 0.009144 0.0012192

0.5667 0.0004064 0.0067056 0.001016

0.7334 0.0002032 0.0065532 0.001016

0.9 0.0002032 0.0041656 0.001016 1.0667 0.0002032 0.0028448 0.001016

1.2334 0.0002032 0.0020066 0.0008128

1.4 0.0002032 0.001016 0.0008128

1.5667 0.0002032 0.0004064 0.0008128 1.7334 0.0002032 0.0002032 0.0008128

1.9 0.0002032 0.0002032 0.0006096

2.0667 0.0002032 0.0002032 0.0006096 2.2334 0.0002032 0.0002032 0.0006096

2.4 0.0002032 0.0002032 0.0004064


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Table 4 Fuselage design results.

Station, m Frame

thickness, m

Main spar

thickness, m

Stiffeners

thickness, m

Skin maximum

thickness, m

0 0.00127 0.0004064 0.00127 0.0008128 0.34 0.0002032 0.0004064 0.00127 0.0018034

0.71 0.0002032 0.0004064 0.00127 0.002032

1.12 0.002286 0.0004064 0.00127 0.0039624

1.505 0.015875 0.0004064 0.00127 0.0028448

1.735 0.0127 0.0004064 0.00127 0.0004064

1.825 0.00635 0.0004064 0.00127 0.0006096

Table 5 Horizontal tail design results.

Station, m Main spar

thickness, m

Rear spar

thickness, m

Skin maximum

thickness, m

- 0.7775 0.0006096 0.0006096 0.00635 - 0.7221 0.0006096 0.0006096 0.00635 - 0.5556 0.0006096 0.0006096 0.00635

- 0.5 0.0006096 0.0006096 0.00635 - 0.3332 0.0006096 0.0006096 0.00635

- 0.1667 0.0006096 0.0006096 0.00635

0 0.0006096 0.0006096 0.00635 0.1667 0.0006096 0.0006096 0.00635

0.3332 0.0006096 0.0006096 0.00635 0.5 0.0006096 0.0006096 0.00635

0.5556 0.0006096 0.0006096 0.00635 0.7221 0.0006096 0.0006096 0.00635

0.7775 0.0006096 0.0006096 0.00635

Table 6 Vertical tail design results.

Station,

m

Main spar

thickness, m

Rear spar

thickness, m

Skin maximum

thickness, m

- 0.3 0.0002032 0.0002032 0.0004064

- 0.16406 0.0002032 0.0002032 0.0004064

- 0.025 0.0002032 0.0002032 0.0004064

0 0.0002032 0.0002032 0.0004064

0.025 0.0002032 0.0002032 0.0004064

0.16406 0.0002032 0.0002032 0.0004064

0.3 0.0002032 0.0002032 0.0004064


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Figure 1: Port side and front side views of the ANCE versions a) X-0, b) X-1, c) X-2, and d) X-3.

Figure 2: Internal view of the ANCE, and general equipment.


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Figure 3: General outline of the mission of ANCE.

Figure 4: General view of the internal structure components of ANCE X-1 and subsequent versions,

generated by CAD tools.


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Figure 5: Wind tunnel model in the test section.

Figure 6: General view of ANCE X-3, generated by

CAD tools.

Figure 7: Lift coefficient as a function of angle of

attack of different version of ANCE.

Figure 8: Lift coefficient as a function of drag

coefficient of different version of ANCE.

Design of an Unmanned Aerial Vehicle for Ecological Conservation

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