Design of a Solar-powered Unmanned Aerial Vehicle for Surveillance

Design of a Solar-powered Unmanned Aerial Vehicle for Surveillance AHMED HELMI MAAROUFI ZHEN LI Degree Project in Mechanical Engineering, First Cycle(SA108X) Stockholm, Sweden 2014

Design of a Solar-powered Unmanned Aerial Vehicle for Surveillance Ahmed Helmi Maaroufi Zhen Li SA108X Degree Project in Mechanical Engineering, First Cycle KTH Industriell teknik och management Farkost och flyg SE-100 44 STOCKHOLM

SA108X Degree Project in Mechanical Engineering, First Cycle Design of a Solar-powered Unmanned Aerial Vehicle for Surveillance Ahmed Helmi Maaroufi Zhen Li Approved 2014-month-day Abstract Examiner Arne Karlsson Commissioner Arne Karlsson Supervisor Arne Karlsson Contact person Arne Karlsson The degree project we were assigned consists of designing an environment-friendly Unmanned Aerial Vehicle (UAV) that can fly at least one hour. In this case the type of energy used for power system is solar energy as a clean energy. Keywords: Endurance Factor, Airfoil Selection, Airplane Design, Solar Energy, UAV 1

ACKNOWLEDGEMENTS We would like to express our sincere gratitude to our Professor Arne Karlsson for his support and assistance during the degree project, and for his valuable advices. Ahmed Helmi Maaroufi Zhen Li 2

NOMENCLATURE Symbols Re Reynolds number [-] ρ Air density [kg/m 3 ] V max Maximum velocity [m/s] µ Dynamic viscosity [m 2 /s] η Efficiency [-] λ Taper ratio [-] c Chord [m] AR Aspect Ratio [-] K Induced drag parameter [-] e t b Oswald efficiency factor Thickness [mm] Wingspan [m] S ref Reference area [m 2 ] δ Aileron deflection [ ] i w Wing incidence angle [ ] α Angle of attack [ ] P Power [W] P req Required power [W] g Gravitational acceleration constant [m/s 2 ] I Current [A] 3

U E T Voltage [V] Energy [J] Thrust [N] T temp Temperature [K] D L W Drag [N] Lift [N] Weight [N] C Di Induced drag coefficient [-] C D0 Zero-lift drag coefficient [-] CD Drag coefficient [-] CL Lift coefficient [-] CL max Maximum lift coefficient [-] C Fe Skin-friction coefficient [-] A top Top area [m 2 ] A side Side area [m 2 ] S wet Wetted area [m 2 ] R HT Horizontal tail arm [m] V HT Horizontal tail volume [-] V VT Vertical tail volume [-] L VT Vertical tail arm [m] L HT Horizontal tail arm [m] C HT Horizontal chord [m] S HT Horizontal tail area [m 2 ] 4

S VT Vertical tail area [m 2 ] b HT Horizontal tail span [m] b VT Vertical tail span [m] Acronyms NACA UAV National Advisory Committee for Aeronautics Unmanned Aerial Vehicle 5

Table of Contents ACKNOWLEDGEMENTS... 2 NOMENCLATURE... 3 1 INTRODUCTION... 8 1.1 Objectives of the project... 8 2 U.A.V. DESIGN... 9 2.1 Wing planforms... 9 2.1.1 Rectangular Wing... 9 2.1.2 Angle of incidence... 9 2.1.3 Aileron design... 10 2.2 AIRFOIL SELECTION... 11 2.2.1 FX 76-MP-120 and NACA 0015... 11 2.2.2 Estimation of the fuselage length... 15 3 Tail/-Stabilizer design... 16 4 CONCEPTUAL SKEKTCH... 17 5 Aerodynamic calculations... 18 5.1 Lift coefficient C L... 18 5.2 Drag Coefficient C D... 19 5.3 1,5 Endurance factor C L /C D... 20 5.4 Rate of Climb R/C... 21 5.5 Take-off Distance... 22 5.6 Landing Distance... 22 6 UAV PROPULSION AND POWER SYSTEM... 23 6.1 Selection of engine and propeller... 23 6.2 Power source... 23 6.2.1 Solar cell... 23 6.2.2 Battery... 24 6.3 Flight endurance... 25 7 NAVIGATION AND CONTROL SYSTEM... 26 7.1 List of equipment:... 26 8 Camera and recording system... 27 6

9 Conclusion and discussion... 28 Division of labor... 29 A. EQUATIONS... 30 B. BIBLIOGRAPHY... 31 7

1 INTRODUCTION The UAV 1 is widely used in a variety of areas including communication, meteorology, agriculture and some civil and military operations with high risk. Flight endurance for UAV varies from hours to days, and it depends on design requirements. Photovoltaic technology is known as a method for generating electric power by using solar cells to convert solar energy into electricity, and this mature technology can be well-used for longtime flying and furthermore, it is Eco-friendly UAV. Using thin-film solar cells can probably make the aircraft weighs less than other kinds of energypowered aircrafts and bring lower expense. Moreover, the efficiency of solar cells is quite high, which provides the aircraft longer flight-time. For safety and simplicity reasons we implemented a continuous design approach in which the battery pack power, aircraft weight and minimum engine power were determined for a given speed and estimated wing loading. In some operations, an aircraft often uses a camera to survey some targets on the ground. Therefor a payload for the UAV is necessary for extra devices. 1.1 Objectives of the project The UAV will be designed to fly at least one hour by using the solar energy as a main source of power, meanwhile monitoring an area and record real-time pictures on the ground continuously during the flight will be held as well. The aircraft is intended to weight 10 kg according to an estimated weight of all the components in the UAV, and its payload is around 2 kg for storing imaging system. I.e. the maximum takeoff weight must be 10 kg, and with consideration of realistic marginal for error the maximum takeoff weight is estimated up to 11 kg. For operating and completing missions we mount a camera of real-time imaging on the UAV s fuselage. The camera lens is exposed towards the ground in order to search and monitor targets on the ground. 1 Unmanned Aerial Vehicle 8

2 U.A.V. DESIGN 2.1 Wing planforms There are three types of straight wing-planforms for low-speed airplane: Rectangular wing, tapered wing and elliptic wing. Trapezoidal or triangular wing provides better aerodynamics but is more difficult to build. Rectangular wing is the easiest to build, simple to install ailerons and flaps. All ribs are identical. It is great on low speed aircraft due to low take off speed (because of the increase of lift). 2.1.1 Rectangular Wing Table 1: Design wing specifications S : Reference area 0.9 m 2 C : Chord 0.3 m = 300 mm AR : Aspect ratio 10 t/c : Max thickness/ chord 12.1% λ : Taper ratio (C tip /C root ) 1 b : Span 3 m Wing side section Figure 1: Airfoil FX 76-MP-120 The maximum thickness 12.1% of airfoil is located at 33.9 % of chord, with a camber 7.6% located at 46.7 % back from the airfoil leading edge. 2.1.2 Angle of incidence The angle of incidence is the angle between the wing (i.e. the chord) and the air flow (i.e. the direction of flight), denoted i w (-or αi). i w = 2 9

2.1.3 Aileron design Table 2: Design aileron specifications Aileron to wing chord ratio 15-25 % C a = 0.15 0.25 C Aileron to wingspan ratio 40-55 % b a = 0.4 0.55 b Aileron deflection angle ± 30 - Ailerons and flaps are typically about 15-25% of the wing chord. (Daniel P.Raymer.P115) From Figure 6.3 Aileron guidelines page 113, we see that the ratio of aileron to wingspan is between 45 to 55%. C a = 0.25 x C = 75 mm b a = 0.4 x b = 1200 mm Figure 2: Aileron (-Wing top view) 10

2.2 AIRFOIL SELECTION 2.2.1 FX 76-MP-120 and NACA 0015 Figure 3: Airfoil selection [1] The selection of an airfoil depends mainly on the type of performance intended from the aircraft. There is a very wide range of profiles, which requires some research work before determining the appropriate one to use. Some other factors participate in the selection of wing profile such as desired CL-value, velocity, stall angle characteristics and size of the UAV. To determine the appropriate wing profile for the UAV, a calculation of the Reynolds number is necessary. The initial calculation has been made at sea level conditions where: The air density is: ρ air = 1.225 Kg/m 3 Pressure: 101 325 Pa Temperature (T): 288.2 K The dynamics viscosity (µ): 1.789 x 10-5 Kg/ (m/s) Speed of sound: 340.3 m/s The designed velocity: 25 m/s 11

Reynolds number: Main wing: Re = ρ.v.c μ 513 555 Tail wing: Re tail = ρ.v.c HT μ 239 660, see section (3) for horizontal chord The Reynolds number range interval for the main wing is between] 500 000 3 000 000[(figure 3). The airfoil type in this case should be one of Wortmann airfoil. Another alternative is to choose among NACA-airfoils, available on Illinois university website, directed for low Reynolds number. Figure 4: Cl vs. α Figure 5: Endurance factor Figure 6: Airfoil Profiles In the present project the selection of airfoils is based on two main factors CL-value with good stall angle characteristic, and a good endurance factor. The endurance factor would affect the value of the minimum power required in steady and level flight. The selected airfoil for the main wings was Wortmann FX 76-MP-120. 12

For the tail wings, Reynolds number is less than 500 000, which could match any of Selig Figure 7: Cl vs. α Figure 8: Endurance factor Figure 9: Airfoil Profiles For the tail wings, Reynolds number is less than 500 000, which could match any of Selig Airfoils as well as NACA profiles for low speed aircraft. The airfoil chosen is NACA 0015. 2.2.1.1 Verified Stall angle by Ansys-Fluent The above plot of CL against angle of attack (figure 4) does not show accurate value of stall angle. The stall angle according to Xflr5 is between 10 and 14 degree. Therefore, we need to investigate the real stall angle of the airfoil using Ansys-Fluent. 13

Figure 10: Velocity contour at 10º angle of attack Figure 11: Velocity contour at 12º angle of attack Figure 12: Velocity contour at 14º angle of attack 14

Ansys-fluent has given the following values of Lift force 192.074 N, 189.390 N and 147.509 N for respective angle of attack 10º, 12º and 14º. From XFLR5 we did obtain the following lift coefficient 1.74, 1.73 and 1.71 for respective angle of attack 10º, 12º and 14º. Both Ansys and Xflr5 has confirmed that the stall angle is at 10º with high value of lift force and lift coefficient. 2.2.2 Estimation of the fuselage length From the table 3 below, we see that an acceptable fuselage length is about 60% to 80 % of airplane wingspan. The values shown in the table were made on some military UAVs. We can also estimate the fuselage length by using the following equation (according to Raymer): L fuse = 1.35 W 0.23 The estimated length of the fuselage: 2.3 m (76.67 % of wingspan) Table 3: Fuselage to wing ratio UAV Wingspan [m] Length [m] Ratio [%] Watchkeeper Tactical UAV(WK180), 6 4.43 73.8 United Kingdom Bayraktar Mini Unmanned Aerial 2 1.2 60 Vehicle, Turkey Hunter RQ-5A / MQ-5B/C UAV, United 8.84 7.01 79.18 States of America KZO Reconnaissance and Target 3.42 2.28 66.67 Acquisition UAV, Germany IAI Malat, Israel 8.55 5.85 68.42 Shadow 200 RQ-7 Tactical Unmanned 4.27 3.4 79.63 Aircraft System, United States of America SKYLARK I-LE, ISRAEL 2.9 2.2 64.7 15

3 Tail/-Stabilizer design A conventional T-tail is selected for the U.A.V Table 4: Candidate horizontal and vertical tail V HT Sail plane 0.50 0.02 Homebuilt 0.50 0.04 GA-Single engine 0.7 0.04 GA-Twin engine 0.8 0.07 Horizontal tail L HT = 1.725 m; V HT =0.5; S ref = 0.9 m; C ref = 0.3 m V HT = S HTL HT S ref C ref S HT = 0.0783 m 2 Typically Aspect ratio for HT should be 3 < AR HT < 5 Tail span: b HT = AR HT S HT = 4 x 0.0783 0.56 m Chord for horizontal tail: V VT S HT = b HT C HT C HT = 0.1397 m 14 cm Vertical tail L vt = 1.725 m; V vt =0.04; S ref = 0.9 m; C ref = 0.3 m V VT = S VTL VT S ref b ref S VT = 0.0626 m 2 Typically Aspect ratio for VT should be 0.9 < AR VT < 2 Tail span: b VT = AR VT S VT = 2 x 0.0626 0.354m Chord for vertical tail: S VT = b VT C VT C VT = 0.1769 m 17.7 cm 16

4 CONCEPTUAL SKEKTCH 17

5 Aerodynamic calculations 5.1 Lift coefficient C L The lift coefficient is defined by the following equation: C L = L q S, q = 1 ρ v2 2 Figure 13: CL Vs. A.O.A The value of the lift coefficient will be used later to determine the minimum speed value, known as stall velocity. If the angle of attack exceeds the stall angle, the airplane wing will no longer generate sufficient Lift force to sustain the flight. W V stall = = 10.75 m/s 0.5 x Cl max x ρ x S 18

5.2 Drag Coefficient C D Figure 14: C D vs. α Figure 15: Cl vs. C D The drag coefficient has multiple origins and first two main coefficients that appear in the formula characterizing the Drag force (D = (C D0+ C Di ). ½ ρv 2 S 2 ), called: zero-lift drag coefficient represented by the first term C D0, and induced drag C Di, which is represented by the second term of the relation below: C D = C D0 + C Di, C D0 = C Fe S wet S C Di = C L 2 π.e.ar =K.Cl2 It is clear to note from the relationship above, that the induced drag can be reduced by increasing the aspect ratio, and the parasitic drag depends on frictions coefficient caused by the friction of the air over the entire surface of the plane. Friction coefficient decreases when the speed the plane increases. From the Cl vs. Cd plot above, we note that the both the lift coefficient and drag coefficient continue to increase until both reach the stall angle, afterward the Cl-value begin to decrease (Lift force decreases), while the drag coefficient increases rapidly. 19

The drag coefficient value at steady state flight: V max 25 m/s A.O.A 2 Cl 1.163 S wet 3.4207 m 2 K 0.0421 C Fe 0.0055 C D = 0.0209 + 0.0569 = 0,0778 The C D value at critical angle of attack corresponds to C D = 0.148. 5.3 Endurance factor C L 1,5 /C D This ratio is further development of lift-to-drag ratio and known as endurance factor. The power required for the UAV depends on the value of C L 1,5 means reduced power consumption based on the comparison. C D.Thus, a bigger value of C 1,5 L C D Figure 16: Endurance factor 20

5.4 Rate of Climb R/C The rate of climb at maximum velocity: R/C = η prp eng W C D0ρV 3 W S 1 K 2(W/S) ρv (14) Where P eng = 500 W; η pr = 0.91; V max = 25 m/s; W = 10 9.81 = 98.1 N Density is 1.1786 kg/m 3 at 400 m according to Swedish laws for UAV 2. R/C = 3.49 m/s Figure 17: Rate of Climb V.s. Velocity 2 http://www.transportstyrelsen.se/tsfs/tsfs_2009_88.pdf 21

5.5 Take-off Distance During the takeoff roll, the aircraft accelerates on the runway before takeoff speed reaches. The take-off ends in altitude of 10 meter above the ground. Lift-off velocity: V L.O = 1.2V S = 1.2 W ½ρV 2 SC L max Takeoff distance includes ground roll distance and distance before the aircraft clear an obstacle of height 10 m: W S L.O = S g + S a = ln T µw (C D +µc L )gρs 5.6 Landing Distance T 1.21 C D µw(1 1.21 C L C L C max L max + ) HW T 1.325 C D C L max W = 116 m 3 The Regulatory landing distance is the distance traveled by the airplane from the passage of 15 meter ( 50 feet) to the full stop on the runway. 2 W 1 W /S 2 Sg = jn + j = 127m ρ S CLmax ρ gclmax + TR /W + [ D / W + µ r(1 L/W) ] 0.7VTD j = 1.15, N = 3 4 3 http://www.kokpit.com/takeoff_distance.html 4 http://people.clarkson.edu/~pmarzocc/ae429/ae-429-13.pdf 22

6 UAV PROPULSION AND POWER SYSTEM 6.1 Selection of engine and propeller First of all we calculate the engine power and then choose an electric engine that has appropriate output power. The minimum power required: P r,min 1/2 1/3 3/2 2 2 KC D0 W = 4 = 81,1 W ρ 3 S The minimum power required for minimum thrust required: P r,tr,min 1,140P == 92,5 W r,min An electric engine (Type: AXI 5330/F3A GOLD LINE) is chosen from online shop Model Motors 5 and fits the criteria of UAV. The maximum efficiency of the motor is 91%. Moreover, an appropriate propeller (two bladed, fan-shaped 20 x13 propeller) that fits best for the motor is recommended and the whole set will be more efficient. Table 5: Engine specifications Motor AXI 5330/F3A GOLD LINE Weight 652 g Dimensions Diameter: 63 mm Length: 64 mm Axis diameter: 8 mm Efficiency 0.91 6.2 Power source 6.2.1 Solar cell The most efficient and compact solar cell found on the market is A 300 solar cell 6 by Sunpower corporation and the efficiency can reach up to 21.5%. 5 http://www.modelmotors.cz/index.php?page=61&product=5360&serie=20&line=gold 6 https://www.cs.wmich.edu/~sunseeker/files/a-300%20data%20sheet.pdf 23

Table 6: Solar cells specifications Sol cells A-300 solar cell Weight 11g/unit Amount 68 units Total weight 11 x 68 = 748 g Dimension 125 x 125 mm Efficiency 21 % Power (P) 3.1 W/unit 6.2.2 Battery Panasonic 18650 Li-Ion rechargeable battery is powerful with capacity 3.4Ah. 7 Figure 17: Panasonic 18650 Li-Ion Table 6: Battery specifications Battery Capacity Max Voltage Min Voltage Average Voltage Weight Energy Dimension Total Total Capacity Panasonic 3400 mah Li-Ion 4.2 V 2.5 V 3.7 V 46 g/unit 12.2 Wh 65 mm 18 unit 219,6 Wh 7 http://www.fasttech.com/product/1141100-panasonic-ncr18650b-rechargeable-3400mah-3-7v 24

6.3 Flight endurance We assume that the UAV flies in a clear weather with sunshine and solar cells are exposed in the sunshine. Total power consumption for all electric components and equipment during stable and level flight: 92,5 W + 12 W + 12 W = 116,5 W The batteries endurance for the UAV s flying: 219,6 Wh 116,5 W 1,9 h 114 min Total power of solar cells: 68 3,1 W = 210,8 W In comparison with data above we can simply see that total power of solar cells is much more than total power consumption, i.e. the UAV will have infinite flight time as long as it is exposed for sunshine. Besides, the battery can supply the UAV 114 minutes extra flight time when there is no good sunlight. 25

7 NAVIGATION AND CONTROL SYSTEM 7.1 List of equipment: FPV TX/RX Set Black 8 : Transmitter and receiver are both 433MHZ and the receiver is just 8 grams and very compact. Range of use is up to 20 km and power adjustment is 1.2W/2W. Figure 18: Transmitter (black) and receiver (green) IG-500N 9 : The world smallest GPS enhanced Attitude and Heading Reference System (AHRS 10 ). Due to SmartFusion 11 the unit can also adjust automatically to difference flight models such as Dynamic model and specific application constraints. IG-500N is a perfect solution for stabilization and navigation of large to miniature UAV. Figure 19: GPS unit Table 7: GPS specifications Physical Dimensions OEM 27 30 14 mm Weight 10 g Specified temperature -40 to 85 C Electrical Operating voltage 3.3V to 30V Power consumption 800 mv @ 5.0V 8 http://www.goodluckbuy.com/433mhz-20km-remote-control-power-adjustable-transmitter--receiver-fpv-tx-rxset-black-1.html 9 http://www.sbg-systems.com/products/ig500n-miniature-ins-gps 10 AHRS : See Wikipedia 11 http://www.sbg-systems.com/news/smartfusion-new-data-fusion-technology 26

8 Camera and recording system Most UAVs are equipped with a gimbaled system 12 and a camera for transmitting and recording real-time video to the ground station. We have two alternative designs for imaging and recording system i.e. an alternative with two cameras mounted respectively in the front of the UAV as bird s eye and under the fuselage, the other alternative with only one camera mounted under the fuselage of the UAV. We choose the latter due to less payload and space requirements and no need of the bird s eye for front-view. In this case it s easy for maintenance and upkeep as well. On the other hand the UAV need to complete operations such as day and night observation and the camera should be power-saving for long flight. The gimbal CM100 camera 13 is best choice and it s easily mounted on the UAV as it is very compact, light-weight, and energy-saving. It can detect and follow 5 moving objectives at once by automatically tagging up. SYSTEM SPECIFICATION Table 8: Gimbal capabilities: Position Accuracy 0,05586 Elevation +/-115 Slew Rate 105 /sec(1,83 rad/s) Power 12W Voltage 9-36V Table 9: Physical dimension Weight 695 grams Dimensions D x H : 100 x 129 mm Temperature -20 C to 50 C 12 Gimbal : See Wikipedia 13 http://www.uavvision.com/product/cm100-html/ 27

9 Conclusion and discussion Our UAV has a great and simple design of wings and the relatively high aspect ratio rectangular wing (10:1) and the airfoil with favorable lift-to drag coefficient are basis of stability and long endurance for a UAV with low speed. Moreover the motor will provide different required speeds for the UAV in different phase and the solar cells will guarantee sufficient energy for long endurance and required power for motor etc. Aileron design will improve maneuverability in rolling in terms of stable and level flight. 28

Division of labor Both of us as project participants have been active through whole project phases and contributed to the accomplishment of this degree project. In every section of the rapport we search information and put forward solutions together and there is no clear division of labor on the whole. 29

A. EQUATIONS Aspect Ratio: AR = b2 S = b c (1) Mach number: M = V, a= 340 m/s (2) a Lift coefficient: Cl= L q S, q = 1 2 ρ v2 (3) Drag Coefficient: CD = D q S, q = 1 2 ρ v2 (4) Straight-Wing Aircraft: e 0 = 1.78 (1 0.045 AR 0.68 ) 0.64 (5) K = 1 π AR e If t/c < 0.05 S wet,c = 2.003 S exposed (6) If t/c > 0.05 S wet,c = [1.977 + 0.52( t c )] S exposed Fuselage wetted area and similar components according to Raymer: S wet = a 1 2 (A top + A side ), (7) a = π, for circular cross section, there a = 4, for rectangular cross section a = 3.4, for common cross section somewhere between circular and rectangular C D0 = C Fe S wet S (8) C Di = K.Cl 2 (11) CD = C D0 + C Di Propeller efficiency: J = V nd (13) Rate of climb: R/C = V sin γ = (T D)V W = P pr DV W = η prp eng C W D0 1 2 ρv3 W S 1 K 2( W C ) cos2 γ ρv (14) sin γ = η prp eng C W V D0 1 2 ρv2 W S 1 K 2(W C ) cos2 γ (15) ρv 2 30

B. BIBLIOGRAPHY [1]. http://itlims.meil.pw.edu.pl/zsis/pomoce/bipol/bipol_1_handout_8a.pdf [2]. http://wwwmdp.eng.cam.ac.uk/web/library/enginfo/aerothermal_dvd_only/aero/atmos/atmos.html [3]. http://aerospace.illinois.edu/m-selig/uiuc_lsat/low-speed-airfoil-data-v5.pdf [4]. http://airfoiltools.com/airfoil/details?airfoil=s9000-il [5]. https://www.grc.nasa.gov/www/k-12/airplane/winglets.html [6]. http://theflyingengineer.com/flightdeck/winglets-and-sharklets/ [7]. http://www.hobbyic.com/products.php?product=remote-control-433mhz-for-fpv-rc- Airplane [8]. http://www.hobbyic.com/products.php?product=zero-uav-ys%252dx4%252dp- Autopilot%2850-waypoints%29 [9]. http://adg.stanford.edu/aa241/performance [10]. http://www.kokpit.com/takeoff_distance.html [11]. http://people.clarkson.edu/~pmarzocc/ae429/ae-429-13.pdf [12]. Karlsson, A.: Cruise Performance, 2004. [13]. Bill Gunston. The Cambridge Aerospace Dictionary. Cambridge University Press, 2 nd edition, 2009. [14]. Karl Nickel and Michael Wohlfahrt. Tailless Aircraft in Theory and Practice. AIAA Educational Series, 1994. Translated by Capt. E. M. Brown RN. [15]. Daniel P. Raymer. Aircraft Design: A Conceptual Approach. AIAA Education Series. Air Force Institute of Technology [16]. Wright-Patterson Air Force Base, Ohio,5th edition, 2012. [17]. Snorri Gudmundsson. General Aviation Aircraft Design Applied Methods and Procedures 2014 Elsevier Inc. [18]. Mohammad H. Sadraey. Aircraft Performance Analysis. VDM Verlag, 2011. [19]. Mohammad H. Sadraey. Aircraft Design: A Systems Engineering Approach. Wiley,2013. [20]. Bill Gunston. The Cambridge Aerospace Dictionary. Cambridge University Press, 2 nd edition, 2009. [21]. Karl Nickel and Michael Wohlfahrt. Tailless Aircraft in Theory and Practice. AIAA Educational Series, 1994. Translated by Capt. E. M. Brown RN. [22]. Daniel P. Raymer. Aircraft Design: A Conceptual Approach. AIAA Education Series, 5th edition, 2012. [23]. W. Austyn Mair and David L. Birdsall. Aircraft Performance. Cambridge UP, 1992. 31

[24]. Arne Karlsson. How to estimate CD0 and K in the simple parabolic drag polar C D = C D0 + K C 2 L. 19 th January 2013 [25]. Arne Karlsson. The aeroplane-some basics. 29 th December 2012 [26]. Arne Karlsson. Steady an level flight of an aeroplane with propeller propulsion. 18 th February 2013 [27]. Arne Karlsson. Thelift-todrag ratios. 10 th March 2004 [28]. Arne Karlsson. Steady climb performance with propeller propulsion. 10 th February 2013 32