Conceptual Design Features of the Sun Falcon 2 a Long Duration UAV

Transcription

1 Conceptual Design Features of the Sun Falcon 2 a Long Duration UAV W. Harasani, M. Khalid, and Y Alturki, King Abdul Aziz University, Jeddah, Saudi Arabia K. Hiraoka K. Fukuda and N. Arai, Tokai University, Japan Abstract- This article reports on the study of a small sized UAV Sun Falcon 2, capable of sustained multi-day and night flights powered from rechargeable batteries energized from solar cells installed on the wing surfaces. This project is part of a continuing collaboration between the staff and students of Tokai University of Japan and King Abdul Aziz University which was first reported by Harasani et al in the Royal Society s Aeronautical Journal May 2014 on the successful launch of Sun Falcon 1 capable of long duration day flight. Such vehicles would be extremely useful for continuous surveillance of sensitive areas and other installations of interests to undesired groups or individuals. The present UAV was designed using quick response algorithms using relationship which optimize aerodynamic attributes from framework geometry and weight considerations and then balance the power requirements from battery and solar cell attributes in consideration of the prevailing environment. It was learnt that Saudi climate with many hours of daily sunlight throughout the year was particularly suited for sustained flight of such vehicles. Nomenclature α Angle of attack (deg.) θ Instantaneous inclination of the vehicle (deg.) A Area (m 2 ) AR Aspect Ratio B Full span of the main wing (m) c Airfoil chord length (m) Cl Sectional lift coefficient CL Lift coefficient, Cp Pressure coefficient CD o Profile Drag CD i Induced Drag D Drag Eff Efficiency E Energy (joules) Eb Total required battery energy (joules) Eg Generated energy by solar panel (joules) Es Specification energy of battery module (joules) Eq. Equation g Gravitational constant (ms -2 ) H Height (m) L Lift force per unit length M Mach number Mm Module mass of solar panel (kg) P Power (joule s-1) 1

2 R Gas Constant Rb Constant in temperature model equations Re, Rec Reynolds number, Chord Reynolds number R/C Rate of Climb (dh/dt) (ms -1 ) Rad Radiation (Wm -2 ) Sw Main wing area (m 2 ) Sp Solar panel area (m 2 ) S Semi span (m) V Velocity (ms -1 ) T Thrust (N) T TH Highest temperature ( o C) T TL Lowest Temperature ( o C) V Free stream velocity (ms -1 ) W Weight of airplane (kg) x tr Axial transition location (m) x, y Axis of Coordinate system θ Hour of the day (radians)- where 2π represents a 24 hr cycle ρ Free stream density (kgm -3 ) γ Specific heat μ prop Efficiency of propeller (0.8) μ motor Efficiency of motor (0.85) μ ESC Efficiency of speed controller (0.9) μ mppt Efficiency of MPPT (0.96) μm Efficiency of solar module at 300 o K (0.2) R masssolm Solar module mass ratio (1.3 kg/m 2 ) R massbat Battery mass ratio (0.005 kg/(w*hr) Subscripts ave Average af Airframe bat Battery crit Critical ground Ground Value midair Midair value req Required init Initial solm Solar Module solar Solar I. INTRODUCTION The collaboration between the staff and students of King Abdul Aziz and Tokai Universities, first reported in the Aeronautical Journal of the Royal Aeronautical Society May 2014, has continued towards the design, manufacture 2

3 of a more challenging solar based UAV model, Sun Falcon 2 capable of longer duration flight lasting many days. The second version would be more suitable for sustained flights to monitor border regions for unwanted intrusions or other illegal activities, or seeking water footprints on vast expanse of Saudi deserts or even monitoring and guiding the activities of large mass of population during the annual Hajj season. Countries with loosely defined borders would be able to scrutinize the movement of people over longer periods of time. Elsewhere such UAV's would be extremely useful for continued observation and assessment of regions suffering from natural disasters such as widespread fire occurrences in large dry region of the world, or monitoring large scale flood activities and other calamities which befall from unexpected radiation leakages or Tsunami driven events. In other applications, one could even envisage the deployment of such appropriately designed vehicles for exploration of Mars atmosphere, where hostile conditions may limit or even restrict the use of other manned vehicles. It would be appropriate at this point to take stock of some of the recent progress in the design of UAV s based on solar power. The credit for the design of first solar power airplane, as documented in Reference [1] goes to the AirVironment s Solar Challenger. The solar powered vehicle weighing 96.4 kg with 50 kg pilot was able to maintain sustained flight for 9 hours at an altitude of 4 km. Other well-known unmanned UAV s include NASA s Pathfinder as discussed in Reference [2} and Helios addressed in Reference [3]. Pathfinder flew to an altitude of some 22 km and proved successful implementation of such enabling technologies as rapid performance and miniaturization in silicon solar cells, power control electronics, electric motors, proton exchange membrane fuel cells, flight control and data acquisition electronics. The Helios with a wing span of m in 2001 set the world record for steady state sustained flight at an altitude of 29.5km. The concept of continuous day-night solar powered flight by a UAV as observed in Reference [4] was first achieved by SoLong in The UAV had a span of 4.75m and stayed airborne continuously for 48 hours. Another solar powered UAV, called Zephyr designed by Qinetiq (See Reference [5]) also set the world record in 2007 for steady state flight lasting over two weeks. References [6,7,8] have provided a very comprehensive outline on design and experimental data on solar powered UAV s. In terms of solar powered piloted vehicles there has been quite a notable progress by Swiss Company Solar Impulse. The latest of its design Solar Impulse 2 is attempting to navigate the globe on solar power. Its voyage starting from Abu Dhabi in March 2015 has already taken it across Asia, Japan, Hawaii and California on April 23 rd, It continued its flight to Phoenix, Arizona on May 2 nd, The paper includes certain aspects of the design including parametric optimization from an algorithm which permutes different aspects of geometric changes (AR, Span, tail design, etc), aerodynamic characteristics, weight considerations, battery and solar performance particularly with reference to day and night operations as well as motor and propeller operations to arrive at a configuration capable of meeting the prescribed long duration flight schedule. The algorithm balances different imperatives of design constraints from different disciplines of the design to conclude if a given configuration would be appropriate. The paper also includes a discussion on the environmental model adopted to design the power requirements of the UAV capable of continuous multi-day operations. 3

4 II. Design of the Sun Falcon 2 Since the aerodynamic aspects of the design have been covered adequately in the first design exercise of Sun Falcon 1detailed in Reference [9], the aerodynamic design details are not discussed in as much detail. The design procedure draws heavily from the design of the first model shown in Figure 1, and reported in Reference[9]. While the wing span wise shape is similar, the inverted T-tail together with double boom fuselage for strength and integrity makes the configurations different. This was a radio controlled solar powered UAV capable of day flight and equipped with an auto-pilot. The essence of the present design relies upon, in determining the precise expenditure of power from the on board battery during the continued day / night operation and its continued replenishment during Figure 1. Test Flight of the Day-Flight Solar Plane, Sun Falcon 1 the sun-light hours. In the overall weight estimates for the complete UAV, the weight of the battery is the single most significant item, as the power requirements at all operations are always optimised against the overall weight considerations. For this type of UAV s, the weight remains virtually constant without any adjustments to fuel consumptions or other payload variations critical for conventional aircraft design. Since the solar panels are equally exposed during day and night operations, it would be important to come up with an appropriate environment model based upon typical day and night temperature distribution throughout the year. The model must accurately simulate the impairment in solar panel efficiency subject to prevailing ambient temperature conditions. The basic aerodynamic design was reached using the same operational procedures as adopted for Sun Falcon 1,(Reference [9]), and will not be repeated at any length here. A dedicated algorithm, Solar Unmanned aircraft Design through Operational Modelling SUNDOME was prepared to conduct the various design cycles for the following design requirements: TABLE I Design requirements of Sun Falcon 2 Description Amount Unit Cruise altitude 500 [m] Payload 200 [g] Cruise speed m/s Rate of Climb R/C 2 [m/s] Day time hours 10 [hr] CL 1 Sea Level Temperature 288 [K] Sea Level Pressure [Pa] 4

5 III. DESIGN METHODOLOGY The overall design procedure demands an accurate weight estimation which can be supported comprehensively by the configuration aerodynamics throughout the complete itinerary of the flight mission. An even more demanding challenge is the adequate supply of the power dispensation especially during the sundown hours. The methodology is heavily based on the principle that the on board batteries would be sufficiently charged during the day operation by the on board solar panels to cover the power requirements during the night hours. Obviously there is an iterative process which optimises the weight against the aerodynamic loads and stability as well as the available power demands. In terms of the actual design and performance specifications, the Sun Falcon 2 was estimated to have a 200 g weight with a continuous flight capability lasting at least 5 days (120 hours) with a cruise velocity of 30 km/h. It will have a climb rate of about 2 m/s operated by an electrical motor powered by a battery replenishable by solar panels. It will take-off in a normal fashion from an appropriate ground terrain and remain airborne at an altitude of about 500 m. A typical flight mission would then require the UAV to climb to a maximum height of 500 m, remain airborne continuously for 120 hours loiter at that altitude and ten descend to a prescribed location. The final design will be subject to the safety regulations of the European Aviation Safety Agency and Certification Safety of Very Light Airplanes (CS-VLA). Power Requirements The In order to model the power requirements one must get accurate information on the aerodynamic characteristics of the aerofoil used to design the wing in the aforementioned SUNDOME Excel algorithm. It was ascertained that the aerofoil S8037 as analysed in Reference [10] generates the most desired lift polars (C L vs α) and other pertinent lift to drag ratio (C L /C D ) and pitching moment (C M ) profiles needed for sustained power requirements. A typical set of data for S8037 from Reference 10 is shown in Figure 2. Figure 2. C L and C M vs α. Reference (10) 5

6 An appropriate relationship was established for the lift profile from the zero lift occurrence (α 0 = -2 o ) to a maximum lift condition given by: (1) Beyond this location we may make use of the 3D relationship as contained in Glauret s expression in Reference [11] to arrive at a corresponding relationship for a finite wing as used for the present UAV. Similarly the drag coefficient, which is so critical to the overall power requirements must be obtained from classic friction and induced drag evaluation. Friction drag under laminar and turbulent conditions may be found from the relationships: (2) (3) Total coefficient of friction is the obtained from the Eq: (4) Where b specifies the span location, S W is the wetted wing area and x tr is the transition location estimated from the relationship: (5) The second proponent of the drag, the induced drag of the three dimensional wing may be recovered from the monoplane Eq. with appropriate aspect ratio of the wing, AR and the taper ratio. Drag coefficient of the wet area of the whole plane is thus expressed as: (6) Where the constant 2.6 in Eq. 6 is used to represent the wet wing area and the constant 0.15 portions off the profile drag and interference drag to come up with the total drag. Thus the power required to satisfy the aerodynamic forces may be calculated by first determining the thrust and lift which act upon the aircraft at a given angle of attack shown in Figure 3. Figure 3. Forces acting on the aircraft 6

7 The forces acting upon an aircraft may be written in terms of the weight, lift, thrust and drag as follows: (7) (8) The necessary power for flight is obtained as follows: (9) (10) (11) Determination of the drag is extremely critical to the solar UAV design. Total drag is the sum of the induced drag and viscose drag. It is given by the Eq.: (12) Where, C Di = C 2 L /(e.ar.π) and C Do = LOG(Re) 2 and e is the Oswald s efficiency factor and Re is the Reynolds number. According to the induced drag and viscous drag calculation, the total drag can be obtained by Eq. 13 (13) Since the required thrust is equal to the total drag, therefore the required power can be calculated from the Eq.: (14) It should be noted that comprehensive CFD simulations and wind tunnel tests were performed to verify the suitability of the airfoil choice for the UAV wings and the tail configuration. It was established that for the present design, an inverted tail joining the double boomed tail provided the most efficient performance. Weight Estimates Along with the aerodynamic performance of the vehicle the weight of the components of vehicle must be obtained with extreme accuracy. The total weight of the vehicle, which must in all cases be less than the aerodynamic lift (L= ρv 2 C L S/2) is given by the Eq.: Total Weight = Weight of batteries+ Payload+ Weight of Solar Cells+ Weight of Airframe (15) The weight of solar cells and the weight of airframe will be calculated from the following two Eqs.: (16) (17) Another important feature of solar based flight is to have an accurate estimate of the energy generated from the solar cells which must exceed the energy requirements of the vehicle. Energy Requirements Energy needed is calculated from adding the energy needed from motor and energy needed to charge the batteries. 7

8 Where energy needed for motors and energy to charge battery can be calculated by the Eq.: (18) (19) (20) The length of day the day time is assumed to be 12 hours in this analysis, and the number of batteries can be obtained from Eq. 22 in the next section. From the batteries specifications in Table 2, the capacity of battery and voltage requirements can be found. The above energy required has to be balanced against the energy available from the solar cells. The energy generated by solar cell is given by Eq. 21: (21) Where Ω rad is the radiation vector obtained from the Jeddah Met Office and S is the wing area and Where minimum capacity and charging current can be found from the battery specifications and the number of batteries can be obtained from Eq. 22 (22) The total energy and single battery energy can be calculated from Eqs. 23 and 24 respectively: (23) (24) The time needed to charge batteries is given by the following Eq: (25) Figure 4 displays an EXCEL based SUNDOME flow chart diagram, in which the individual compartments are updated as the iterative design procedure is advanced to converge towards the final design. It has the ability to iterate between configuration aerodynamics, weight, energy and power requirements as well as the critical time to energy absorption from the daylight operations. While the use of SUNDOME is straightforward, some familiarity and learning curve is required in developing expertise to arrive at a meaningful configuration, the design of which may require appropriate weight balancing, aerodynamics and energy regulation between what is generated from solar cells and what is expended and stored. The mission parameters are introduced into the mission specification module which feeds such information towards the aerodynamics module which uses such basic performance coefficients as the lift, drag and configuration geometry to arrive at the power and energy generations and other motor specifications. This information is in turn 8

9 used in the power plant design to arrive at the solar cell and battery weight, area and power requirements. A final decision module interrogates whether the available energy and weight quantities satisfy the appropriate constraints and meet the critical time needed to replenish the battery charge for night operation in a repetitive manner. If the constraints are not satisfied than the frame geometry in terms of the aspect ratio and airframe weight is updated to repeat the convergence iteration. Figure 4: SUNDOWM Flow Chart Typical configuration geometry, aerodynamics, mission specification, and other energy and power requirements as well as battery and solar power requirements at any instant as they are updated during successive design iterations is shown in Table 2. 9

10 TABLE 2. UUNIT EXCEL SHEET Mission Calculated Parameter Battery specifications Cruising Velocity [km/hr] Temperature T [K] Nominal Capacity min [mah] 2950 Cruising Altitude H [m] Pressure P [Pa] Nominal Capacity Typical [mah] 3100 Payload [Kg] (Camera and Data Transmitter) 0.12 Density ρ [kg/m^3] 1.17 Approx. Weight [g] 45.5 Rate of Climb R/C [m/s] 2.00 Cruising Velocity [m/s] 9.30 Nominal Voltage [V] 3.6 day time time [h] Wing Area S [m^2] 2.86 Energy for a single battery, [W.h] Chord Length c[m] 0.38 Reynolds Number Re 2.277E+05 Night flight Total Flat Plate Area [m^2] {Wing, Fuselage, Tail, and others 7.44 Night flight hours [h] (=2.6*Wing Area)} Configuration Induced Drag Coefficient Cdi Energy total needed [W.h] 850 Wing Span [m] 7.50 Flat Plate Turbulent Viscous Drag Coefficient Cd Number of Batteries needed 81 AR Total Drag [N] 5.18 Weight of batteries [kg] Solar Module Area [m^2] 2.58 Day flight Night time [h] Solar Energy per [m2, Wh/m2] 989 Energy needed from motor [W.h] 850 Energy from battery needed to charge [W.h] Airframe Energy needed from solar [W.h] Weight of Airframe [kg] (Including Weight of Motor, Controller, Propeller, Electric Circuit, Solar Cell, and Actuator) 5.00 Solar cells area from Energy needed [m2] Given Data Motor and Propeller Cruising CL 1.00 Required Thrust T [N] 5.18 Energy Balance Oswald Efficiency Factor e 0.80 Required needed Power [W] (for High Mounted Wing) Energy generated from solar power [W.h] Viscosity [Pa*s] Weight of Motor [kg] 0.8 Energy needed < Energy generated OK Sea Level Temperature To [K] Charging current [A] 20 Gas Constant R[J/(K*Kg)] Weight Balance Decrease Rate of Temperature a [K/m] Stalling Speed Total Weight W [kg] Sea Level Pressure Po [Pa] Maximum CL 1.00 Lift L [N] Sea Level Density ρo [kg/m^3 ] Stalling Speed at Maximum CL and Sea Level [m/s] 1.69 Lift L [Kg] 15 Efficiency of Motor(0.85) * Propeller(0.8) Airfoil: S8037 Lift must > weight predicted OK Solar Cell Efficiency 0.23 Minimum Average Daily Horizontal Radiation [W*hr/m^2] predicted Weights Time Balance Battery Weight [kg/(w*hr)] weight of solar cells [kg] time to charge battery [hr] Margin Time of Battery Usage [h:m] weight of batteries [kg] 3.69 time to charge < day time OK total weight [kg] The configuration which was used to provide various geometry, weight and other aerodynamic characteristics is shown in Figure 5. The wing airfoil section is based on an S8037 airfoil without a fuselage having an inverted V tail configuration supported by a tail boom which is extended forward for an appropriate c.g location. Detailed geometry and other aerodynamic features of the final design are included in Table 3. 10

11 Figure 5.View of Sun Falcon 2 (all dimensions in mm) TABLE 3. SUNDOME OUTPUT Description Amoun t Unit Wing span length 7.5 [m] Aspect ratio Total amount of global radiation 6006 [W.h/m2] Chord at root 0.4 [m] Cruise speed 33.4 [km/h] Altitude 500 [m] Efficiency of the propeller Efficiency of the motor Efficiency of the solar module Predicted Weigh of the airframe 5 [kg] Predicted Weight of the solar module 2.06 [kg] Predicted Weight of the battery (Lithium ion) 3.69 [kg] Predicted Weight of the plane [kg] Power during cruise 70 [W] The Environmental Model The environmental model must address the region in which the UAV would be deployed. For example, the design parameters for UAV mission in a hot and dry desert climate would be substantially different from cold and wet tropical climate. The increase of temperature or the presence of humidity would degrade the efficiency of the solar cells. Since the primary design exercise was intended for long duration flights in Jeddah Saudi Arabia, it is important to come up with a temperature model appropriate for this locality. 11

12 The daily global horizontal radiation changes throughout the year as produced by El-Sebaii et al in Reference [12] is represented as a bar chart depicted in Figure 6. Figure 6. Temperature and Global Horizontal Radiation of Jeddah The maximum and minimum radiations are 7,451 (W/m 2 day) in June to 4,140 (W/m 2 day) in December respectively. Ordinarily, the minimum radiation value should be used to determine if the solar based UAV would remain airborne during all days throughout the year. However, for Saudi climate it would not be too far from design to use an average value of radiation as the design parameter. Towards this end an average radiation value of 6,003 (W/m 2 daytime) based on aforementioned data was selected as the daily global horizontal radiation Rad sun for the current design purposes. The other annual distributions of maximum and minimum modelling temperatures are graphed as dotted lines in the same Figure 6. The extreme temperatures vary in between a maximum of 39.4 o C to a minimum of 27.6 o C. It was mentioned earlier that the efficiency of a solar cell deteriorates with the rise of temperature, so, for the modelling purpose these two extreme for our modelling parameters of T TH and T TL. The Temperature Model Using the two extreme temperatures in the atmosphere it is possible to write down the ground temperature as a trigonometric function of these two values: (26) 12

13 These equations actually model the variation of the temperature during the day. The day is divided up into three separate segments and each segment is dealt with separately. The first equation deals with the portion starting from midnight until the time when the lowest temperature of 300 o K is reached. The second portion deals with the midday temperatures starting from about 6 a.m to around 2 p.m. The last segment addresses the variation of temperatures from 2 a.m. back to midnight. The temperature on the ground as obtained through this expression is given by the curve shown in Figure 7. Figure 7. Temperature Distribution Curve (27) (28) (29) We can obtain the temperature at some height using the atmospheric Eqs. 27 to 29. Here ΔT, with a value of about is the incremental lapse rate for the temperature grounds up, and h in above expressions corresponds to the height above ground. Then the actual efficiency of the solar module, at a given altitude and time can be calculated by first evaluating a solar cell lapse rate using a known efficiency value at given temperature μ m and known altitude (say at 300 o K at ground ) and then obtaining a time distribution value (θ) using a time function relationship as shown in Eq.30: (30) Now using a similar time function relationship for time distribution value we have: (31) 13

14 The solar power supplied by the solar cell can be represented by the trigonometric function shown in Figure 7. The maximum power can now be obtained from the above mentioned daily global horizontal radiation through the use of the similarity rule. It is given by the Eq.: (32) cos Figure 7. Trignometric distribution of supplied power In above expression the value of the parameter p max is given by the formula: (33) Where is the area density of global horizontal radiation as shown in Figure 7. Critical Energy Requirements for Multi- day Flight The solar module must absorb sufficient energy to continuously charge the battery for night time operations when the source of energy replenishment is no longer available. The reservoir of energy is designated is S 2 where the energy is deposited during the day and where the leftover amount is more than sufficient to feed the night time requirements. The energy schedule for night time operations is shown in Figure 8. 14

15 Figure 8. Energy requirements for night time flight. Three types of energy requirements for the night time can be inferred from the graph in Figure 8. S eb represents the minimum energy requirements for the night, S 1 is that portion of energy that is used up from S 2 the total saved from the day long flight. The maximum power needed is calculated from, p smax = μ solm μ mppt p max in the case when μ solm remains constant. In a more accurate analysis which follows, the quantity μ solm is allowed to vary as function of time. The maximum power needed for flight with constant velocity and altitude in a day has limitations which must satisfy the relationship: (34) Where the critical power is recovered from the relationship: = (35) And the critical value is indeed found from the relation: (36) Therefore, with constant cruise flight conditions, the solar UAV would satisfy a power condition: (37) in order to meet the multi-day requirement of flight. In the present conceptual flight design, this preceding limitation was used as a minimum power restriction for flight while μ solm is assumed to be a time dependent function. The Minimum Battery Capacity For a given multi day and night flight operation a minimum battery capacity must be determined. Using values of p critical, θ c1,θ c2 in the following Eq. for p, θ 1 and θ 2 respectively in the Eq. for S eb : 15

16 Where S c is the area of the solar cells.θ 1 and θ 2 could first be evaluated for preliminary start up values using: (38) (39) Then a minimum energy capacity is obtained from: (40) Now if the energy always remains at about 20% of the capacity, then the actual required capacity E b is given by: (41) The Mass Prediction The mass of battery M bat and solar module M solm are obtained from the relationships: (42) (43) where and are mass rate of battery and solar module as itemized in Table 4. The total mass hence is computed from the individual sum of all the components as: (44) where M af is the mass of airframe including mass of motor, gear etc. The mass of airframe itself was calculated using Noth s procedure of Ref. [5] derived from manned gliders and radio controlled planes. Accordingly then, (45) Simulation Method for the Final Design Convergence The iterative procedure adopted to refine the configuration towards the most efficient design the flow chart diagram depicted in Figure 4 was meticulously executed. Various variables and their respective iterative steps are identified in Table 3. The decision whether a particular set of results produced a flyable configuration was left to the satisfactory outcome from the final flow chart condition. The Excel based algorithm contains all of the above mentioned aerodynamic, weight assessment, aerodynamic performance and design equations. Results 16

17 Figure 9 outlines the distribution of power per unit area for various aspect ratios and cruise velocities for a prescribed span length of 4 m. The increase in power requirements with the increase in cruise velocity is clearly demonstrated. The curving nature of the graphs is brought about by the variation of the extent of the laminar and transition region. The model uses a fixed approach for arriving at a linear transition location, whereas in reality the expanse of these regions fluctuates according to a given 3D geometry and the immediate flow conditions. This being the case, the graphs do point to a minimum power distribution requirement at each cruise velocity and aspect ratio. Figure 9. Power Requirements vs Aspect ratio and Cruise velocities 17

18 Figure 10- Weight and Minimum lift distribution vs Aspect Ratio Figure 10 shows the weight and minimum lift distribution plotted against the aspect ratio for a span length of 4 m while the airframe mass is kept M af = 0.66 M af(noth). It is further noted that as the aspect ratio increases the requirement for minimum lift decreases as indeed does the requirement for minimum weight. 18

19 Figure 11. Regions of flyable configurations for multiple days Through repeated iterations of the algorithm at a variety of aspect ratio, span and velocity conditions, it was observed that the number of flyable conditions (represented by green squares) amount to about 161 at C L2d = 0.6 and 1268 for C L2d = 0.8. The outcome is duly shown in Table 4. TABLE 4. FLY TYABLE AND MANUFACURABLE CONFIGGURATIONS C L(2D) Flyable Manufacurable Cases 0 Cases 0.8 1,218 Cases 84 Cases It should be mentioned that many of the cases included here represent configurations where the weight of the airframe is too light for practical purposes. It is difficult to declare configurations flyable where the stiffness does not measure up to the loading conditions. In order to determine whether a given weight is sufficient, one may make use of the Noth s weight criteria as itemized in Reference [5]: 19

20 (54) It is argued that if this condition is satisfied then the plane would have sufficient force to support the weight of the frame with the prescribed stiffness. Table 5 show that there was no flyable configuration at C L2d = 0.6, however there were as many as 84 configurations which were deemed as flyable at C L2d = 0.8. The various specifications, which in theory could meet the conditions supported by these 84 cases are identified in Table 5. The final design arrived at after careful consideration of all the design constraints is shown in Figure 12. The configuration finally selected is a double tail boom design supporting a simple fuselage housing the battery and other instrumentation hardware and joined at the rear end by an inverted V tail. The configuration is shown in Figure 12. TABLE 5: FLYABLE AND MANUFACTURE ABLE CONFIGURATIONS Span [m] AR [.] TR [.] Velocity [km/hr] Mass [Kg] Figure 12: Sun Falcon 2 test flight The double tail boom is used to provide a better structural integrity to the large span wing which must accommodate the battery as well facilitate the installation of solar panels spread across the entire surface of the wing. The inverted V- tail is adopted for superior aerodynamics at velocities of around 38 m/s. The Sun Falcon 2 made its maiden flight in June 2015 and remained airborne for about 20 minutes where its systems for utilizing solar panels for transferring energy to batteries were successfully checked and the porotype has now flown again to validate it for a more sustained flight on 16 February Under civil air space provision regulations, the students 20

21 were not permitted to fly it through the night hours but all system checks during its flight confirmed its viability as a long duration solar powered UAV. Conclusion Sun Falcon 2 has been designed from the lessons learnt from the successful designs of Sun Falcon 1. Suitable temperature models have been used to assess the functions of the solar cells and their inevitable impact on the power /unit area distribution and the weight estimates. Meticulous design procedures with fast turnaround times, were devised to arrive at the most optimum design for the multi day operation of the Sun Falcon 2. Sun Falcon 2 made its first test flight in June 2015 and a more sustained demonstrative flight on February 16, For the test flights conducted to date, all systems pertaining to sustained flight such as power requirements, solar cell energy recovery, battery replenishment and rates of power consumption and storage have been thoroughly checked to validate the mission success. ACKNOWLEDGMENT This project was funded by the Deanship of Scientific Research (DSR) King Abdulaziz University, Jeddah, under the grant No. (431 / 009), the authors, therefore, acknowledge with thanks DSR technical and financial support, furthermore the authors would like to express their gratitude and appreciation to Tokai University for their technical help and support. REFERENCES [1] P. MacCready, P., Lissman, W. Morgan, and J. Burke, Sun powered aircraft designs, Journal of Aircraft, vol. 22, no. 10, pp , [2] K. Flittie, and B. Curtin, Pathfinder solar powered aircraft flight performance, vol AIAA, [3] T. Noll, J., Brown, M. Perez-Davis, S. Ishmael, G. Tiffany and M. Gaier, Invertigation of the Helios prototype aircraft mishap report, NASA [4] Solar Plane Breaks Two-Night Flight Barrier Renewable Energy World, July 5, 2005, [5] J. J. Amos, Eternal plane returns to Earth. BBC News, 2010, [6] A. Noth, W. Engel, and R. Siegwart, Design of an ultra-lightweight autonomous solar airplane for continuous flight, Field and Service Robotics, vol. 25, pp , [7] A. Noth, M. Engel, and R. Siegwart, Flying solo and solar to Mars, IEEE International Conference on Robotics and Automation, vol. 13, no. 3, pp , [8] A. Noth, R. Siegwart, and W. Engel, Autonomous solar UAV for sustainable flights, in Advances in Unmanned Aerial Vehicles, ser. Intelligent Systems, Control and Automation: Science and Engineering. Springer Netherlands, 2007, vol. 33, pp [9] W. Harasani, M. Khalid, M., N. Arai, K. Fakuda, and K. Hiroaka, Initial Design and Wing Aerodynamic Analysis of a Solar Powered UAV, The Aeronautical Journal, May [10] C. A. Lyon, A. P. Broeren, P. Giguère, A. Gopalarathnam, and M. S.., Summary of Low-Speed Airfoil Data, Vol. 3, SoarTech Publications, Virginia Beach, VA, 1998, 418 pages. [11] H. Glauert, The Elements of Aerofoil and Air Screw Theory, second edition, Cambridge University Press, 1947 [12] A. A. El-Sebaii, F. S. Al-Hazmi, A. A. Al-Ghamdi, and S. J. Yaghmour, Global direct and diffused solar radiation on horizontal and tilted surfaces in Jeddah, Saudi Arabia, Applied Energy, vol.. 87, Issue 2, February 2010, Pages , Elsevier. [13] A. Noth,, Design of Solar Powered Airplanes for Contjinuous Flight, P.47, ETH, Zurich,