Preliminary Design of Solar Powered Unmanned Aerial Vehicle Sumit Jashnani a, Prashant Shaholia b, Ali Khamker c, Muhammad Ishfaq d, and Tarek Nada e Emirates Aviation College, PO Box 53044, Dubai, UAE a sumit.ggs@gmail.com, b pash_shaholia@yahoo.com, c alisajjad110@yahoo.com, d ishfaq.hafiz@gmail.com, e tarek.nada@emirates.com Key Words: Irradiance, Solar Powered Aircraft, Unmanned Aerial Vehicle. Abstract. Applications involving the use of alternate, renewable energy sources are expanding exponentially, and are in high demand. Solar power has long been harnessed for such applications and aviation is no stranger to it with its strong drive towards becoming an environment-friendly industry. This paper describes a straight forward procedure to design and test a solar powered unmanned aerial vehicle that can fly continuously for 24 hours at any day of the year. The paper introduces the modeling and preparation of hardware testing of the propulsion and power sub-system. The main components of this sub-system are solar panels, the electromechanical drive train and the propeller. A design for a thrust stand to measure the performance of the system is also introduced. Nomenclature AR Aspect ratio of the wing E Oswald efficiency factor F Thrust force [N] G Irradiance [W/m 2 ] I Current [A] Component of cell current due to photons [A] I o Minimum saturation current, commonly taken as 1*10-10 [A] Short circuit current [A] K Boltzmann Constant, which is 1.38 I0-23 [J/K] Q Charge of a single photon, which is 1.6*10-19 [coulomb] Wing area [m 2 ] T Cell temperature [K] T a Atmospheric temperature [ C] U Velocity [m/s] V Volt [V] Open circuit volt [V] Airframe weight [kg] Introduction Unmanned Aerial Vehicles (UAV) are being proposed for many applications including surveillance, mapping and atmospheric studies. These applications require a lightweight, low speed, medium to long endurance aircrafts. Due to the weight, speed, and altitude constraints imposed on such an aircraft, solar array generated electric power can be a viable alternative to airbreathing engines for certain missions [1]. Moreover, a great interest has arrived at designing and building solar powered platforms that may replace satellites for civil and military applications [2]. Many solar powered UAV s have been built, which are not only
capable of long endurance flight but also do not require much maintenance [3]. Possible applications of the UAV include military and classified surveillance flights where small aircrafts are difficult to be detected by radars. Scientific applications include ozone monitoring, and collection of data for weather and global warming studies. Commercial applications include aerial surveying, geological and topographical mapping, and communication links. This paper describes a straight forward procedure to design and test a solar powered unmanned aerial vehicle that can fly continuously for 24 hours at any day of the year. The paper introduces the modeling and preparation for hardware testing of the propulsion and power sub-system. A design for a thrust stand to measure the performance of the system is also introduced. The aircraft is aimed to be used in surveillance and road control applications. Aircraft Sizing and Drag Estimation The mission scenario of the aircraft is to takeoff at morning (sunrise time) at any day of the year, then climb to altitude of about 5km, which is the cruise altitude at which the aircraft will fly for few weeks before it lands. This mission requires the design of the aircraft to be based on the coldest day of the year and the power management system to be able to work on a complete cycle over 24 hours. The cruise flight conditions are considered as the design conditions. The following step by step procedure is followed to estimate the aircraft size and generated lift and drag: Assume a value for the solar panel area. This value will be verified later. Find the dimensions of the available solar panels. Select a chord length such that the solar panels width is 90% of the chord length. Practically, 90% of the wing area can be covered by solar panels; hence the wing area can be estimated. For simplicity, use rectangular wing shape. Thus, the wing span and aspect ratio can be estimated. The practical range for aspect ratio in solar powered aircraft is between 12 and 31 [4]. Select an airfoil section and get the lift and drag polar curves for this airfoil. The airfoil used here was WE3.55-9.3, which was specially developed for the Sky-Sailor project [3] As rectangular wing is employed, the Oswald efficiency factor is assigned the minimum value, which is 0.85. At the cruise altitude, find the air density and kinematic viscosity. Select a reasonable cruise speed for the mission and payload, and estimate Reynolds number. Based on Reynolds number, the coefficient of skin friction drag can be estimated using [5]:,. 485000 (1),. 485000 2 From the airfoil data, estimate the 2D lift coefficient, l, at different angle of attack within the workable limit and then convert them to 3D using: 3
The induced drag coefficient is obtained using: 4 For the respective, l, find the C d, and then add C d to C Di to get the total drag coefficient C D. Multiply both C D and C L into (0.5*ρ*V 2 *S) to get the generated lift and drag at each angle of attack. The drag due to fuselage can be ignored when compared to wing drag. Some designs consider flying wing geometry. Using reasonable mounting angle for the wing, the angle of attach for horizontal flight, the generated lift and drag at this angle of attack are identified. For steady horizontal flight, the lift should be equal to the weight and the thrust should be equal to the drag. The required thrust and cruise speed are now known; hence the propulsive power can be estimated. The following section will validate the assumed solar panel area by estimating the available solar power and compare it with the required propulsive power. Solar Panels Dubai is positioned 25 13 North Latitude, 55 16 East Longitude and is 16 meters above sea level. At certain irradiance during the day, the current produced by a single cell is function of the cell voltage [6]: 1 5 The cell temperature, T, is the temperature of the solar cell that changes with the amount of irradiance across the duration of the entire day. Each cell has its limiting short circuit current,, and open circuit voltage,. This means that operating outside this envelope may damage the cell. The short circuit current is equal to the current due to light intensity,. To a very good approximation, the cell current is directly proportional to the cell irradiance. Thus, if the cell current is known under standard test conditions, G = 1 kw/m, at air mass, AM = 1.5, then the cell current at any other irradiance, G, is given by: 6 Similarly by setting the current to zero in Eq. 5, open circuit voltage can be estimated using the formula: 7 For variations in ambient temperature and irradiance, the cell temperature, in C, can be estimated quite accurately with the linear approximation:. 8 The Nominal Operating Cell Temperature, NOCT, is the temperature that the cells reaches when operated at open circuit in an ambient temperature of 20 C at air mass = 1.5, irradiance conditions, G=0.8kW/m 2, and a wind speed less than 1 m/s [6]. Information such as the I sc, V oc, and NOCT is provided by the cell manufacturer. The resulting power at different irradiance and voltage is plotted in Fig. 1 for the selected cell (Q6lmxp3, manufactured by Q-
cells). This figure shows the power produced by a panel of 30 cells (common number of cells in a single panel) for different levels of irradiance. The minimum and maximum irradiance coincides with the actual available irradiance available in Dubai over the year. Also, it is clear from Fig. 1 that the maximum power at any irradiance occurs at voltage in the range from 17-18V. Figure 2 depicts the hourly energy received from the sun on a horizontal surface in Dubai for the hottest and coldest days of the year [7]. The hottest day is June 8 th, while the coldest day is December 31 st. The total available solar power on the coldest day can be estimated by evaluating the area under the curve in Fig. 2. The drive train behind the propeller consists of the gearbox, motor, inverter, battery, maximum power point tracker, and charge controller. An estimate for the efficiency of each component would be employed. An overall efficiency of about 12% including the solar panel is a typical value [3]. Thus, the required solar panel area can be estimated and compared to the previous assumed value. Iterations may be required until convergence is achieved. Power [W] 160 140 120 100 80 60 40 20 0 200G= 285G= 400G= 600G= 840G= 0 2 4 6 8 10 12 14 16 18 20 22 Voltage [V] Figure 1 Variation of power with voltage at different irradiance levels Solar energy (MJ/m^2) 5.0 4.0 3.0 2.0 1.0 0.0 Hotest day Coldest day 0 2 4 6 8 10 12 14 16 18 20 22 24 Hours of the day Figure 2 Available solar energy at the hottest and coldest days of the year Electromechanical Drive Train For a given amount of irradiance, the available technology allows only a limited amount of solar energy to be extracted from the solar cells. It is however desirable to operate these cells at their upper limit in order to extract the maximum amount of energy from this renewable
source. The Maximum Power Point Tracker (MPPT) ensures that the solar panels are working at their maximum power points that can be seen in Fig. 1 [8]. The battery voltage is selected to be 18 V, and the MPPT will change the current accordingly to ensure maximum generated power. This obtained power is used firstly to supply the propulsion group and the onboard electronics, and secondly to charge the battery with the surplus energy. The solar panel should be able to provide the battery with enough charge during the day for it to run the propeller through the night. Fig. 3 demonstrates the power management over a full day. The number of hours available to charge the battery will be maximum if the UAV takes-off at sunrise. This number decreases as the takeoff time is delayed and thus more energy has to be initially stored in the battery at the time of takeoff. Thus, sunrise is the recommended takeoff time. The limit line marked in Fig.3 indicates the maximum power needed for the battery to run the propeller for the entire night and contain the exact same initial charge value at the sunrise time of the next day. The stored power in the battery starts with about 50W at takeoff time, reaches maximum of 1550W at noon, and finally ends with the same 50W at the sunrise time of the next day. This guarantees that the system can repeat this cycle many times and its capability will increase if working in a hotter day. For 18V battery, the corresponding current ranges from 2.7 to 86A. This range of current can be easily achieved. In the following section, the total aircraft weight will be estimated and compared to the generated lift. 1.80 1.60 1.40 output power from the battery Power stored in the battery Power [kw] 1.20 1.00 0.80 0.60 0.40 Battery charging limit Takeoff 0.20 0.00 Hours of the day Figure 3 Power management over 24 hours Weight Estimation The following step by step procedure can be followed to estimate the weight of aircraft components and subsystems, and then the total weight can be summed. The payload of the mission is defined and its weight is determined Due to the sheer number of solar powered aircrafts operational today and the peculiarities involved in the structure of such an aircraft; a unified weight estimation model is hard to find. However, a statistical model has been formulated using data obtained from large number of solar powered UAV [3]:
0.44.. :, 9 The density of Silicon is 0.3651Kg/m 2, and hence, the weight of solar panels can be estimated from the known area. A 20% weight is added to account for coating and accessories [6]. The maximum expected stored energy can be concluded from Fig. 3. The energy density of batteries employed in solar aircrafts can reach 350Wh/kg such as that used in reference [4]. Using this value along with the maximum expected stored energy, the weight of the battery can be estimated. The propulsion group is composed of control electronics, motor, gearbox and propeller. A model was introduced in reference [3] to estimate the weight of the propulsion group based on curve fitting of data from existing commercial components: 78.48 10 The total weight is the sum of all above weights, and this weight should be equal to or less than the generated lift that was estimated before. If the weight is more than the generated lift iteration will be required. Following the above procedures, the weight and size of the aircraft were estimated. Also, the required thrust to be generated by the propeller was obtained. Table 1 summarizes the obtained values. Table 1 Summary of estimated values Parameter Value Parameter Value Solar panel area 3.2 m 2 Airframe weight 75.4 N Solar panel width 0.5 m Weight of solar panels 13.7 N Wing chord 0.55 m Battery weight 43.5 N Wing area 3.55 m 2 Weight of propulsion group 5.1 N Wing span 6.4 m Payload weight 5 N Wing aspect ratio 11.5 Total weight 142.7 N Cruise speed 12.0 m/s Cruise thrust 5.4 N Propeller Design The propeller is the final component of the mechanical drive train and is responsible for generating thrust required for flight. The concepts of propeller design have been explained by various theories developed over the years [9, 10]. However, the blade element theory is the most sophisticated and detailed of all. Using this theory for propeller design facilitates the selection of an optimum diameter-rpm setting at maximum efficiency and a detailed geometry which includes propeller geometric pitch, blade span-wise chord, and twist angle. This procedure starts with the selection of appropriate section airfoils which are thin, possess a high lift-to-drag ratio and have a large range for the stall angle of attack. The desired ondesign flight condition is the cruise condition. The details of the design procedure can be found in [9]. The propeller efficiency at different RPM for various diameters is plotted in Fig. 4. It is clear from Fig. 4 that with increase in diameter, the propeller efficiency increases due to the reduction in required power for a specific thrust value. Selecting 20in 1900RPM design set is impractical as it produces the highest efficiency but at very high RPM. However, selecting
30in 1100RPM design set over 40in 800RPM is more practical as the difference in efficiency is minimal and is easily cancelled out by the weight and size penalty imposed by the latter. Also the power required is well within the available limit and so is the RPM as maximumm allowable from the motor is 1330RPM. 80 70 Efficiency 60 50 40 30 0 500 1000 1500 D = 20in D = 30in D = 40in 2000 2500 30000 RPM Figure 4 RPM versus efficiency for different propeller diameters at cruise conditions The tip Mach number and the structural/ /mechanical loading limits on the propeller must also be addressed. In the case of this level of thrust, they fall out of the investigated range. Thus, the design RPM is 1100, which corresponds to thrust of 5.4N at 12.0m/s cruise speed. This requires a power of 100W output from the battery at 5.56A and 18V. Thrust Stand It is required to test the output thrust for continuous operation over 24 hours. There are many techniques to simulate and test the propeller operation [11, 12]. After reviewing many techniques, it was decided to employee a simple however reliable method such that described in [13]. The thrust stand is designed and manufactured with the necessary components; namely the motor, S-type load cell and the propeller in the assembly as shown in Fig. 5. Figure 5 Schematic drawing of the thrust stand ( 3D CAD).
The gear ratio used is 1:1, which is only done to relocate the position of the main shaft connected to the propeller. This also allows hassle-free placement of the strain gauge. The bearings hold the shaft in place and rest on the pedestal, which is connected to the load cell. As the propeller produces thrust, the shaft tends to move forward and thus makes the pedestal move along. This movement consequently is transferred to the load cell in the form of tension and the readings can be taken. This tension felt by the strain gauge is equal to the effective thrust force being produced by the propeller minus friction force. The amount of friction on the pedestal roller has been reduced greatly using rollers and addition lubrication. Moreover, a calibration is done using dead weights. Conclusions A straight forward procedure for preliminary design of solar powered unmanned aerial vehicle was introduced. The estimation of available daily solar energy over the year in Dubai was done and the required solar panel area and number of cells were evaluated for the desired design conditions. The electromechanical drive train was designed to efficiently manage the available power and provide the propeller with constant power over 24 hours. Finally, the thrust stand that is needed to simulate and test the operation of the propulsion and power subsystem was designed. The system will be integrated and tested in the future and the results of the testing phase will be published later. References [1] Colozza A. J., Scheiman D. A., and Brinker D. J., GaAs/Ge Solar Powered Aircraft, NASA/TM 1998-208652, October 1998. [2] Symolon W. E., High-Altitude, Long-Endurance UAV s vs. Satellites: Potential Benefites for US Army Application, MSc thesis, MIT 2009. [3] Noth A., Design of Solar Powered Airplanes for Continuous Flight, PhD thesis, ETH ZÜRICH, 2008. [4] Mehdi Hajinmaleki, Conceptual Design Method for Solar Powered Aircrafts, AIAA 2011-165, 49 th AIAA Aerospace Sciences Meeting, January 2011, Orlando, Florida, USA. [5] Rizzo E., and Frediani A., A model for solar powered aircraft preliminary design, The Aeronautical Journal, Vol. 112, No. 2, February 2008, pp. 57-78. [6] Messenger. R and Ventre. J, Photovoltaic System Engineering, second ed., CRS Press, Taylor and Francis E-library, 2005. [7] Duffie.J and Bechkman.W, Solar Engineering of Thermal Processes, third ed., John Wiley and Sons, 2006. [8] Cohen J. M., Peak Power Tracking for a Solar Buck Charger, MSc thesis, MIT 2010. [9] Carpenter H., Aerodynamics, Butterworth Heinemann Inc, 2003. [10] GurO., and Rosen A., Comparison between blade-element models of propellers, The Aeronautical Journal, Vol. 112, No. 12, December 2008, pp. 689-704. [11] Merchant M. P., Propeller Performance Measurement for Low Reynolds Number Unmanned Aerial Vehicle Applications, MSc thesis, Wichita State University, 2004 [12] Asson K. M., and Dunn P. F., Compact Dynamometer System That Can Accurately Determine Propeller Performance, Journal of Aircraft, Vol. 29, No 1, 1991. [13] Ash R. L., Miley S. J., Landman D., and Hyde K. W., Evolution of Wright Flyer Propellers between 1903 and 1912, AIAA 2001-0309, the 39 th Aerospace Sciences Meeting and Exhibit, 2001.