Conceptual Design Requirements of a Newly Developed UAV as a ...

Conceptual Design Requirements of a Newly Developed UAV 

as a Research Platform at KFUPM: Structure, Aerodynamics 

and Propulsion 

Amro M. Al-Qutub * 

King Fahd University of Petroleum and Minerals, Dhahran 31261, Saudi Arabia 

The objective of the present paper is to set the structure, 

aerodynamic and propulsion conceptual design requirements of a 

multi-role long endurance UAV. The aircraft is designed to be used 

as a research platform at King Fahd University of Petroleum and 

Minerals, Mechanical engineering department. The maximum 

payload and takeoff weight were set at 10 kg and 65 kg, respectively, 

which are the main design constrains of the system. Further 

specifications include cruise speed of 220 km/hr, 8 hours endurance 

and a ceiling of 18,000 ft. This would allow the UAV to perform 

over 1,700 km of non-stop flight with maximum payload and without 

refueling. The present study was based on available information in 

the literature and published specifications of different UAV systems 

manufacturers. Results included systems weight distribution, and 

limitations on requirements of the aerodynamic performance and 

propulsion system. Design sensitivity analysis resulted in the 

definition of basic performance limitations to achieve the objective 

of the present UAV. This included high aerodynamic performance 

(L/D > 13) which requires a retractable landing gear configuration. 

Also, the propulsion system should be equipped with a pusher 

propeller driven by a 2 stroke IC engine (17 -22 hp) with specific 

power higher than 2.5 hp/kg and low SFC of 0.33 kg/hp.hr. System 

reliability is improved through the selection of higher engine design 

output. Also additional heat-sink system for engine head would 

further improve the service life during hot local season. The overall 

propulsion system performance will be enhanced with a high 

efficiency propeller (> 50%) during cruise. The structure should be 

made of light composite material to comply with limited aircraft dry 

weight, of 46 kg. Wings are separable to allow for testing different 

configurations and for the ease of store and transportation. The 

survivability and reliability of data and system recovery is improved 

by a recovery system composed of a parachute and an airbag in case 

of emergencies. 

* Associate Professor, Mechanical Engineering Department, aqutub@kfupm.edu.sa. 

1 

SSAS UAV Scientific Meeting & Exhibition, Jeddah, Saudi Arabia (June 6, 2006) 

Paper No. SSAS-2006-051

T 

I. Introduction 

HE Unmanned Aerial Vehicle (UAV) is a remotely piloted or self-piloted aircraft that 

can carry cameras, sensors, communications equipment or other payloads. UAV 

technology and usage actually extends beyond Military and defense market. Many 

Universities around the world have their own UAV development program such as 

University of Sydney, Clark University, Cranfield University, Georgia Tech. University 1 , 

Berkeley University, Simon Fraser University 2 , and others. 

For Example University of Sydney UAV fleet 3 is composed of five different flying 

UAV’s: KCEXP-series, Ariel (Figure 1), Brumby, T-Wing, and Bidule miniature Air 

vehicle. This besides all other UAV still under development such as The Solar powered 

High Altitude Endurance (HALE) meteorological UAV. The University of Sydney enjoys 

the funding from different organizations like Booing, and BAE. 

Figure1: Sydney University UAV examples: the Ariel (left) and the Brumby (right). 

King Fahd University of Petroleum and Minerals (KFUPM) is encouraging 

multidisciplinary projects where different department contribute and integrate a system in 

a specific technology field. KFUPM encouraged faculty to work towards a vision that may 

contribute to the well fear of the surrounding society and the enhancement of the 

dissemination of knowledge. The development of UAV in KFUPM has the following 

mission: 

1. Develop know-how of UAV for local applications 

2. Provide experimental platform for multi-disciplinary research related to UAV 

applications 

3. Provide a tool for academic excellence in different areas. 

The specific UAV should satisfy the following: 

• able to accommodate for equipments related to multi-disciplinary research and 

development work. 

• safe to operate 

• reliable for local environment 

• comply with local regulation of communication 

Building a UAV demands a vast knowledge of base vehicle platforms, flight dynamics, 

control theory, and real-time software in a network environment. The flight system enables 

the aircraft platform to serve as the autonomous agent. It requires sophisticated navigation 

2 



sensors and integration methodologies for accurate and reliable autonomous operation. 

Key UAV system technologies were identified to be in: airframes; propulsion units; 

autonomous flight controllers; launch and recovery; navigation and guidance; selfprotection; 

ground control stations; payloads; and data communication, storage, 

processing, and dissemination (Information Technology). Key system technologies cover 

the following items: 

• UAV airframe, structure and aerodynamics 

• Propulsion unit 

• Control system (Flight control, Propulsion control) 

• Telemetry and ground station 

• Instrumentation for performance evaluation and control 

• Support system 

The objective of the present paper is to study the specific conceptual design and 

performance requirements of the UAV structure, aerodynamics and propulsion system. 

II. General Requirements 

1) Structure and Aerodynamics general requirements: 

• Composite structure + Aluminum alloy 

• Payload of 10 kg with gross takeoff weight of 65 kg 

• Reinforced 

• Landing gears retractable 

• Support up to 5g’s 

• Internal fuel tank and equipments 

• Conventional high wing configuration for stability 

2) Flight general requirements: 

• Take off and landing on paved ground 

• 160-220 Km/h cruise speed 

• 8+ hours endurance 

• Ceiling 6 km (18000 ft). 

3) Propulsion general requirements: 

• Internal combustion engine (IC) high performance 

• Propeller driven with minimum diameter and high efficiency during cruise 

• If possible electrical starter + alternator 

• Withstand up to 45 o C, and 55 o C for short duration during takeoff 

• High engine efficiency with low SFC 

• High power/weight ratio 

In the design phase both engine efficiency and power/weight ratio have to be optimized for 

minimum system weight. 

Basically the payload is the most important part in the design phase since it sets the 

limitations on possible missions the UAV can perform. For basic research and surveillance 

purposes 10 kg payload is considered sufficient since continuous developments in 

3 



electronics avails higher performance and lower weights sensors, cameras and processors. 

As shown in Figure 2, compared to conventional UAV's, the payload-to-gross weight ratio 

% Payload/gross Weight 

70 

60 

50 

40 

30 

20 

10 

PAYLOAD vs Endurance 

0 

3 4 5 6 7 8 9 10 

Endurance (h) 

Figure 2: Payload vs Endurance of available Known Conventional 

UAV’s ( 4) 

Ceiling (feet) 

80,000 

70,000 

60,000 

50,000 

40,000 

30,000 

20,000 

10,000 

Present UAV 

Service ceiling of UAV's 

0 

0 10 20 30 40 50 60 

Endurance (h) 

Present UAV 

Figure 3: Ceiling vs. Endurance of available Known Conventional 

UAV’s 4 

is average, however the endurance places the present UAV among the top of the group 4 . 

The ceiling is among the majority of UAV’s of its endurance category, Figure 3. 

4 



A. Aerodynamics and Structure 

Based on the objective of the project the UAV is a long endurance (8 hrs +) with 

reasonable cruise speed of about (220 km/hr). Moreover, takeoff should be possible either 

from a launcher or a runway. This imposes a grate challenge in designing the 

aerodynamics and the structure of the aircraft. Endurance requires: 

− high aerodynamic performance (high L/D, lift to drag ratio) 

− Relatively Large fuel tank ( to improve weight ratio ) 

− Strong structure with high strength to weight ratio materials. 

− Low thrust specific fuel (high efficiency propulsion system) 

Endurance 

To estimate the endurance of an aircraft we may start with the mass reduction of fuel 

during level steady flight: 

dm ⎛ D ⎞ 

= −TSFC 

∗ ⎜ ⎟( 

m • g) 

(1) 

dt 

⎝ L ⎠ 

Where TSFC is the thrust specific fuel consumption of the propulsion system, m is the 

mass of the vehicle, and g is the gravitational acceleration. Assuming constant L/D and 

TSFC through the flight, the endurance of the vehicle can be calculated as: 

1 ⎛ L ⎞ 1 mo 

Δ t = ⎜ ⎟ ln 

(2) 

g ⎝ D ⎠ TSFC mb 

Where mo is the takeoff weight and mb is the fuel empty weight of the UAV. 

Takeoff requirements 

The UAV has to be able to takeoff from a runway, or from a launcher. Taking off from 

a runway requires landing gear system that reduces L/D if extended during flight. So, 

retractable landing gear should be used to maintain high aerodynamic performance. The 

major penalty in this case is increased structural weight (by only few percent). A 

reasonable takeoff velocity of the UAV is about 25 m/s. Based on the objective of the 

project, the maximum payload is 10kg. Referring to figure 3, the aircraft takeoff gross 

weight can be chosen as 65 kg, which corresponds to payload to gross weight ratio of 

15.4%. To calculate the wing platform area, we may use the lift equation: 

ρ 2 = V C S 

(3) 

2 

L L 

With plain flap system and large aspect ratio (>9) of the wing, maximum lift 

coefficient can reach a value over 1.4. Using equation (3), and considering the takeoff 

speed to be grater than stall speed by 20%, the required wing area for the UAV for takeoff 

5 



will be 1.66 m 2 at standard atmosphere. Higher atmospheric temperature and/or altitude 

would require slightly higher takeoff speed due to reduced density. For example at 50 C o , 

sea level, the takeoff speed will be 26.5 m/s, only a 6% increase compared to standard 

atmosphere conditions. 

Structure 

The Present UAV has a conventional aerodynamic configuration with pusher-type 

propeller See figure 4. The main fuselage has a maximum diameter of about 350 mm to 

contain cameras similar to the BAI 5 as well as large fuel tank and Payload: 

Propeller 

Fuselage 

Camer 

a 

Figure 4: Basic configuration of the UAV. 

The UAV is to be made from composites (glass fiber and/or Carbon fiber) and 

Aluminum alloy, to provide high strength to weight ratio. Skin parts and some of the 

skeleton supposed to be made from composite material while hard points and joints to be 

mainly from aluminum alloy for improved reliability. Part of the wing is supposed to be 

separable during storage and transportation for convenience of use and handling. For the 

systems contained in the fuselage they have to abide by the dimensional restrictions and 

maintain a reasonable static margin for stability purposes. Systems layout in the fuselage is 

illustrated in figure 5. Engine, landing gear and camera are supposed to have fixed 

locations. Fuel will be stored in multiple tanks, close to center of gravity for stability 

during flight and flexibility of payload as well as static margin consideration. The payload 

bay should contain flexible mount system for different parts and configuration of 

communication, control and payloads. The camera is (which is a part of the payload) 

mounted at the bottom of the aircraft as shown in figure 5. Recovery Parachute and airbag 

has to be on the center of gravity. 

6 



Fuel 

Tanks 

Propulsio 

n 

Cruise performance 

Figure 5 : Layout for Main systems in the Fuselage 

The UAV is expected to operate at a ceiling altitude of 18,000 ft. At the cruise speed of 

220 km/hr the lift coefficient is CL = 0.31 (at gross weight). This is a reasonable lift 

coefficient for high aerodynamic performance. Lower speed will produce higher endurance 

and this can be optimized according to the type of the UAV mission. Moreover, the cruise 

speed of the present UAV is considered among the highest in its category as can be seen in 

table 1 6,7 . 

Table 1: Different UAV performance and power requirements 6,7 . 

UAV Gross 

Weight 

(kg) 

Parachute Payload bay 

Ceiling 

(ft) 

Airbag 

Endurance 

(hrs) 

Camera 

Main Landing Gear Nose Gear 

Cruise 

speed 

(km/hr) 

Power 

(hp) 

From Table 2 the weight to power ratio of UAV's of similar category as the proposed is 

in the range of 2.4 to 5.2 kg/hp. The variation of designed hp on propulsion system 

depends on mainly on: 

7 


Paper No. SSAS-2006-051 

W/P ratio 

(kg/hp) 

Huntair 127 17,000 7.5 184 38 3.3 

Raptor 854 65,000 8 - 50 17.1 

Prowler 91 21,000 6 100 38 2.4 

STM-5B 114 16,000 6 144 25.5 4.47 

Pectre II 100 23,000 6 220 26 3.85 

Skyeye 355 18,000 10 126 46 7.72 

Hunter 727 15,000 12 162 120 6.06 

Model 410 818 30,000 12 184 160 5.11 

Shadow 272 17,000 14 129 52 5.23

− Weight of the aircraft 

− Flight altitude 

− Design clime rate 

− Flight speed 

In the present work it is proposed to use an engine with 17 to 22 hp, which corresponds 

to a weight to power ratio of 3.82 to 3. This will put proposed UAV among the top of the 

UAV's list in terms of engine power and performance. Actually the main reasons of that 

are the high cruise speed and the high operating temperature in the local environment, 

which is part of the objective for this project. To calculate the endurance of the UAV the 

weight ratio has to be known. Table 2 is a list of most recent UAV engines specification 

available in the world market that can be used for the present project 8-11 . Based on present 

technology, the propulsion system is expected to weigh about 9 kg, including engine, 

propeller and generator. 

Table.2: Some UAV Engines available in the international market. 

Engine 

Output 

Weight Power /Weight ratio 

(Brand) 

(hp) 

(kg) 

(hp/kg) 

A200B 

17 5.1 / 6 kg (est.) 

2.83 

Quadra 

300W Gen. 

3W-200i B2TSQS 

20 4.8/6 kg (est.) 

3.67 

3W Motoren 

W.Gen 

313 

22 8 

2.75 

Zanzotterra 

W. 300w Gen. 

L275 E 

25 7.7 / 9 kg (est.) 

2.78 

Limbach 

W. Gen 

AR741 

37 

10.7 /12kg (est.) 

3.1 

UAV Engine Ltd. (rotary) 

W. Gen. 

Based on the calculated size of wing and fuselage, the airframe of the UAV is expected 

to have a weight between 10 and 16 kg, depending on design and material used for the 

construction of the system. Table 3 provides the design weight of different systems in the 

UAV. 

Fuel consumption can be estimated using figure 6, which is an example for typical 

turboprop engine 12 . From this figure we can use a conservative thrust specific fuel 

consumption of 0.06 kg/N.hr. Using equation 2 with the assumption of having an 

aerodynamic performance of L/D=15, the UAV endurance will be 11.14 hrs. This satisfies 

the objective of the project. 

Airframe Propulsion Control & 

communication 

Table 3: Weights of different system in the UAV. 

Landing 

Gear 

Recovery 

System 

Payload 

(Max.) 

Fuel 

with 

max 

payload 

8 



Gross 

Weight 

12 kg 9 kg 3 kg 4 kg 4 kg 10 kg 23 kg 65 kg

Figure 6: Typical Engines TSFC (12) 

The calculated endurance is based on several assumptions regarding structural 

technology, aerodynamic technology and propulsion system performance. Figures 7 to 9 

Are the results of sensitivity analysis to illustrate the effect of each of the mentioned 

parameter on the endurance of the UAV. The total structure mass includes airframe, 

propulsion, communication and control, landing gear, and recovery system. Reducing the 

structure weight to 26 kg will result an increased endurance of over 15 hrs. If all typical 

values of aerodynamic performance and propulsion performance are met, the total 

structural mass has to be limited to 36 kg in order to meet the 8 hrs flight endurance Figure 

7. Aerodynamic performance also proved to be critical for endurance. Increasing 

aerodynamic performance will increase the endurance. The limit to meet the objective here 

is L/D ≥13, if all other variables are typical. Reducing the (TSFC) improves the flight 

endurance linearly, Figure 9. Improving the TSFC down to 0.045 will increase the 

endurance to 14.85 hrs. Based on available technology, the TSFC has to be maintained 

below 0.06 during cruise conditions. Figure 10 illustrates the effect of different payload 

mass on flight endurance. Reducing payload to 5 kg (half of max.) will increase flight 

endurance to 12.3 hrs, a 1.2 hrs increase over full payload (10kg). Detailed discussion on 

the propulsion system is at the next section. 

In conclusion, there are three main parameters that affect the flight endurance: 

− Structure design and material 

− Aerodynamic design 

− Propulsion system performance 

Structure and aerodynamics are to be optimized in the design process. Propulsion 

performance is considered as a major design constrain. 

9 



In conclusion the following are the aerodynamic and structure specifications: 

Aerodynamic: 

− High aerodynamic performance L/D>13 

− Pay load = 10 kg max 

− Gross weight 65 kg 

− Conventional configuration 

− Takeoff speed = 25 m/s (standard atmosphere) 

− Wing Area 1.6 m 2 approx. 

− Max cruise 220 km/hr 

− Endurance 8 - 12 hrs. 

− Max ceiling 18,000 ft 

− Flight range > 1,800 km 

Structure: 

− Made of composites (fiberglass-carbon) with Aluminum Alloy. 

− Separable wings. 

− Retractable landing gear 

− Recovery system (parachute + airbag) 

− Total Structure weight < 36 kg 

− Multiple Fuel tanks 

− Payload rack for flexibility of mounting with min dim. of (200X200X400) mm. 

Flight Endurance ( hrs ) 

16 

14 

12 

10 

8 

6 

4 

2 

0 

26 29 32 35 38 41 

Structure mass ( kg ) 

Figure 7: Effect of Structure Mass on Flight Endurance 

10 





18 

16 

14 

12 

10 

8 

6 

11 13 15 17 19 21 

Aerodynamic Performance (L/D) 

Figure 8: Effect of Aerodynamic Performance on Flight Endurance 

18 

16 

14 

12 

10 

8 

6 

0.045 0.05 0.055 0.06 0.065 0.07 0.075 

TSFC ( kg/ N.hr ) 

Figure 9: Effect of thrust Specific Fuel Consumption on Flight Endurance 

11 




15 

14.5 

14 

13.5 

13 

12.5 

12 

11.5 

11 

10.5 

10 

Propulsion system 

The objective of the project requires a reliable and efficient propulsion system 

specifically during cruise condition (180 - 220 km/hr) with a ceiling of 18,000 ft. In 

addition, the UAV has to be practical and easy to maintain. 

Type of System 

10 8 6 4 2 1 

Payload Mass ( kg ) 

Figure 10: Effect of Payload Mass on Flight Endurance, Fixed Fuel mass of 23 kg. 

For the UAV application similar to present UAV category, propulsion systems are 

usually limited to either turbojet or propeller driven by an IC engine. Propeller UAV are 

more popular due to lower cost and higher efficiency at relatively low speeds (Ma

Engine Requirements 

The engine has to be able to operate in the local environment. So the basic 

requirements are: 

Cruise 

- Power to meet takeoff at 9,000 ft altitude (Abha-Saudi Arabia) 

- Low fuel consumption to meet endurance requirements 

- Tolerate high temperature and dusty environment 

- Tolerate long periods of flights >20 hrs. 

- Tolerate high g forces during catapult launch. 

At level flight with a speed of V, the required propulsion power can be calculated as: 

Power P 

⎛ D ⎞ 

= m g ⎜ ⎟ V 

(4) 

⎝ L ⎠ 

For cruise at 18,000 ft and 220 km/hr, with gross weight of 65 kg and minimum L/D 

of 11, the required propulsion power will be Powerp = 3.54 kW ( 4.74 hp ). Typically, IC 

engines lose about 7% of output power per 3000 ft altitude. Taking this into consideration 

and assuming that the engine will operate at 67% of rated power during cruise, also 

assuming propeller efficiency of 60%, the required engine shaft power at standard 

atmospheric conditions should be about 20 hp, to meet the cruise condition. The extra 33% 

of engine power is for maneuver and climb requirement. 

High altitude takeoff 

For takeoff at 9,000 ft the power available will be 79% of the rated engine power. To 

maintain a reasonable weight to power ratio of less than 5.2 kg/hp, the engine rated shaft 

power should be > 15.82 hp. This is even less than the engine power requirement for cruise 

condition. 

Propeller Efficiency 

Based on endurance calculations, the required TSFC < 0.06 kg/N.hr. Most UAV engine 

suppliers claim specific fuel consumption (SFC) for their product as less than 0.33 

kg/hp.hr. Diesel engines recently developed for aviation application can reach SFC as low 

as 0.19 13 . The TSFC is affected by both SFC and propeller efficiency (ηp) : 

V 

TSFC = 0. 

00134 SFC 

(5) 

η 

13 



P

Typical propeller efficiency can be as high as 85%, Figure 11. Usually each propeller 

design is based on a given flight speed. Designers should select propeller according to 

specific design requirement: 

- Power matching with engine 

- efficiency at cruise 

- static and low speed thrust 

- noise 

- diameter 

Propeller Efficiency (%) 

TSFC (kg/N.hr) 

80 

60 

40 

20 

100 200 300 400 500 

Flight Speed MPH 

Figure 11: Typical Propeller Efficiency (12) 

0.08 

0.07 

0.06 

0.05 

0.04 

0.03 

0.02 

0.01 

0 

0.4 0.5 0.6 0.7 0.8 

Propulsion Efficiency 

Figure 12: Effect of Propulsion Efficiency on Thrust Specific Fuel Consumption 

14 



For the present project the propeller selected should have maximum efficiency close to 

a speed of 200 km/hr. Figure 12 illustrates the required TSFC variation with propulsion 

efficiency for airspeed of 220 km/hr and SFC of 0.33 kg/hp.hr. 

From Figure 12 it is clear that to meet the endurance requirements the propulsion 

efficiency during cruise condition should be ηp > 45 %. The recommended efficiency of 

propeller in the present study is 50% or better at cruise condition. This is consider a 

reasonable estimate compared to low speed propellers. 13,14 Altitude and flight speed will 

change the propeller size and configuration to meet the efficiency requirement. So, more 

than one propeller type should be used to meet different altitudes and flight speed for 

optimum performance. The propulsion requirements are as the following: 

a) Engine : 

− 20 -25 hp Two stroke engine with: 

− SFC < 0.33 kg/hp.hr (at cruise) 

− Power to weight ratio > 2.5 kg/hp (including generator) 

− Ability to operate to temperatures up to 45 o C 

− Ability to operate in dusty environment (equipped with intake filter) 

− Able to with stand axial acceleration > 10 g. 

− Operate efficiently at max altitude 18,000 ft 

− Self priming ignition system with min. 100 W generator. 

b) Propeller: : 

− Pusher type 

− Meets engine power 

− 70-80 cm diameter, minimum possible size 

− Efficiency ≥ 50% at cruise condition 

III. Conclusion 

The major objective of the multi-role UAV is to be capable of long endurance (8-12 

hrs) flights with a maximum payload of 10 kg. The maximum expected cruise speed is 220 

km/hr, and the ceiling is 18,000 ft. This would allow the UAV to perform over 1,700 km 

of non-stop flight with maximum payload and without refueling. Compliance to 

specifications limits of different systems is expected to assure achievement to the 

objectives of this project. The present study was based on available information in the 

literature and published specifications of manufacturers. Design constrains are identified 

as: 

− Available engine performance and specifications 

− Available control system specifications (size – weight-power) 

− Communication System 

Other major specifications can be met through careful design and execution. Based on 

design sensitivity analysis the UAV is expected to have a conventional aerodynamic 

configuration with pusher propeller driven by a 2 stroke IC engine with > 20 hp output for 

improved reliability. Special heat sinks for engine heads are recommended for improved 

reliability of the system during hot local season. Also, the structure should be made of light 

15 



composite material to comply with limited aircraft dry weight of 36 kg. This would 

provide a tolerance of 4 kg for the structural components of the aircraft in the design and 

construction phase compared to the typical design in table 3. Wings should be separable 

for ease of store and transportation. Moreover, aerodynamic performance (L/D) have to be 

grater than 13, which requires a retractable landing gear configuration. In addition, the 

engine has to have high specific power > 2.5 hp/kg, with low SFC< 0.33 kg/hp.hr. Also, 

the propeller should have high efficiency during cruise (> 50%). Survivability of the UAV 

is improved by a recovery system composed of a parachute and an airbag in case of 

emergencies. 

Acknowledgment 

The author acknowledges the support of King Fahd University of Petroleum and 

Minerals during this research. 

References 

1 Johnson, E. N., Hart, M. G., and Christophersen, H.B., “Development of an Autonomous Aerial 

Reconnaissance System at Georgia Tech,” AIAA's 1 st Technical Conference and Workshop on 

Unmanned Aerospace Vehicles, Systems, Technologies, and Operations, 20-23 May 2002 

2 Hennessey, C., “Autonomous Control of a Scale Airplane,” School of Engineering Science, Simon 

Fraser University, 14 April 2000 

3 Wang K.C." UAV Design Activities in a University Environment" presented at the 9th Australian 

International Aerospace Congress, Canberra, Australia, 6-8 March 2001 (PDF 4.0 file) 

4 Amro M. Al-Qutub and Sami El Ferik "System Definition for Unmanned Aircraft Development" 

KFUPM Internal Report Jun, 2003. 

5 BAI Aero systems. Payloads and Sensors. www.baiaerosystems.com/payloads.html 

6 UAV forum: http://www.uavforum.com/index.shtml 

7 Association for Unmanned Vehicle Systems International: http://www.auvsi.org/ 

8 UAV Engines Ltd.: http://www.uavenginesltd.co.uk 

9 Zanzottera Engines: http://www.zazoterteraengines.com 

10 Limbach Engines: http://www.limflug.de 

11 3W Modellmotoren GmBH Engines: http://www.3w-modellmotoren.com 

12 Hill and Peterson "Mechanics and Thermodynamics of Propulsion", Addition-Wesley, 2 nd edition, 

1992 

13 Shaila L. Jaszlics and Rianer Stemme " Possibility of the Next Generation of High-Altitude Long- 

Endurance Unmanned Aerial Vehicles" Technical Publication/ HALE-UAV. 

14 F 2 Craig D. Paxton, Peter J. Gryn, Eerisa Hines, Ulises Perez, and Ge-Cheng Zha " High 

Efficiency Foreword Swept Propellers at Low Speed" AIAA 2003-1069. 

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