Solar UAV. An analysis of literature

Università degli Studi di Salerno Facoltà di Ingegneria Dipartimento di Ingegneria Industriale Tesi di Laurea in Ingegneria Meccanica Corso di Impianti ad Energie Rinnovabili Solar UAV. An Analysis of literature. Relatore: Candidato: Ch.mo Prof. MarcoVillani Gianfranco Rizzo Matr. 0612303360 Anno accademico 2017/2018 Abstract This paper presents the "state of art" of the solar UAV (Unmanned Air Vehicle). Using conventional fuel is a pollutant, have a limited life and costly. So there is a huge demand for using an unlimited non-exhaustible source of energy as a fuel. As solar energy is one of the available renewable energy, it can be used to increase the endurance of UAV. It talks about of the application of Solar UAV, of the key technologies involved in the design and manufacture. The main challenge in a solar air-vehicle design is balancing between the available energy and operational requirements. Since these two represent complex functions of the design variables the design process is based on a Multidisciplinary Optimization Design (MDO) approach. There are various types of electric motors, solar cells, batteries, hybrid operating systems, which will be discussed. An important attention is given to HeliPlat, the solar UAV designed by the Department of Aerospace Engineering of “Politecnico di Torino”. Contents 1. Introduction 5 2. General Results 6 3. Papers 8 3.1. The Development Status and Key Technologies of Solar Powered Unmanned Air Vehicle 8 3.2. Solar Power for Small UAVs 10 3.3. Design of a high altitude long endurance flying­wing solar­powered unmanned air vehicle 11 3.4. Conceptual design of a high-endurance hybrid electric unmanned aerial vehicle 13 3.5. Performance analysis of solar powered Unmanned Aerial Vehicle 18 3.6. Space-Based Solar Power a technical, economic, and operational assessment 24 3.7. Mission Analysis of Solar Powered Aircraft 26 3.8. Conceptual Multidisciplinary Design Optimization (MDO) of Solar Powered UAV 28 3.9. Optimizing Electric Propulsion Systems for Unmanned Aerial Vehicles 34 3.10. Design of a high-altitude long-endurance solar-powered unmanned air vehicle for multi-payload and operations 39 3.11. Design of solar high altitude long endurance aircraft for multi payload & operations 44 3.12. Manufacturing and Design of Lightweight Composite Airplane Structures Part II – Design 48 3.13. Design, manufacturing and testing of a HALE-UAV structural demonstrator 50 3.14. Aeroelastic behavior of a solar-powered high-altitude long-endurance unmanned air-vehicle (HALE-UAV) 53 4. Conclusion 55 5. Bibliography 56 List of figures Figure 1 Planisphere 6 Figure 2 Number of papers in the years 7 Figure 3 Solar UAV 10 Figure 4 Typical Mission 13 Figure 5 Selected T-frame configuration 14 Figure 6 Final UAV design 15 Figure 7 Region of interest for current electric motor/battery powered UAV 16 Figure 8 Region of interest for conceptual hybrid-electric powered UAV 17 Figure 9 Single row of cells - High Aspect ratio wing 20 Figure 10 Two rows of cells - Medium Aspect ratio wing 20 Figure 11 Arrangement and connection of solar cells on the mid-wing 21 Figure 12 Solar High Altitude Powered Platform Power Train 27 Figure 13 The evolution of solar panels technology and efficiency 28 Figure 14 Energy balance for a long endurance solar power vehicle 29 Figure 15 Batteries energy density 30 Figure 16 Solar radiation. Comparison between the model results and meteorological measurements 31 Figure 17 Schematic mission description 33 Figure 18 Motor maximum output power as a function of motor mass 36 Figure 19 Battery capacity as a function of battery mass 37 Figure 20 HeliPlat Border Surveillance over the Mediterranean Sea 40 Figure 21 HeliPlat 40 Figure 22 HAVE / MALE monitoring area comparison 41 Figure 23 Fuel Cell Energy Storage system enables continuos flight through night 43 Figure 24 External Layout SHAMPO 45 Figure 25 Internal layout SHAMPO 46 Figure 26 Propeller thrust vs required thrust 47 Figure 27 FEM ………………………………………………………………………………..49 Figure 28 FEM 49 Figure 29 FEM 49 Figure 30 Testing System 51 Introduction The UAVs are remotely piloted aircraft with remote pilot. As almost everyone knows, drones were born for war, and their first appearance was in the First World War. One of the most common applications of UAV technology is the military as it helps to easily control surveillance issues, but this isn’t the only application; the UAV can also be used for early forest fire detection, telecommunication, immigration control, communication relay, reconnaissance. Places where humans can not enter directly can be easily monitored by an aerial flight unit and photographs of critical locations can also be taken. The first demonstration of solar powered air vehicle was on November 1974. The unmanned Sunrise vehicle flew for 20 minutes in California skies. Five years later the first manned solar aerial vehicle, Solar-Riser, flew a distance of 800m. Solar-Riser was equipped with solar panels that supplied part of the required energy, while the rest was supplied by nickel-cadmium batteries that were charged prior to the flight. The human powered Gossamer Albatross proved the validity of an air-vehicle with low power requirements; a reasonable next step was to use solar power for a similar configuration, which resulted in the Gossamer Penguin. It flew on April 1980 and was the first true solar powered airplane that used solely solar power for its operation. [1] In particular we will talk about the solar UAV, that is a kind of aircraft based on solar radiation as a propulsion energy. During the solar powered UAV flight, the solar energy cell converts the solar energy into DC output, then the Maximum Power Point Tracker (MPPT) controls the maximum output power of the solar cell and converges to the main power supply circuit. The power control device sends the power to the motorpropeller system. Solar radiation is the total energy source of UAV. [2] General Results It is very interesting to know which are the main places of interest of the solar drone and also to know the evolution of the studies of this topic over the years. Interest is shown by most of Europe, USA, Asia and even in Oceania (Fig. 1). Figure SEQ Figure \* ARABIC 1 Planisphere The Fig.2 shows the increase over the years of articles concerning UAV, comparing these with solar UAV (source Google Scholar). Figure SEQ Figure \* ARABIC 2 Number of papers in the years Papers 3.1. The Development Status and Key Technologies of Solar Powered Unmanned Air Vehicle There is a particular type of aircraft based on solar radiation as a propulsion energy; the solar power UAV. There is an important project of Google, “Sky Bender”, that provides to use solar unmanned air vehicle, and maybe in the future with a commercial crew. The technologies used to this aircraft is that glider. Development status of solar power UAV: "Helios" UAV, developed by NASA and the Aeronautical Environment Corporation (UAS), can supply 40 kilowatts of power with its large wings. It is able to reach over 29000 meters in height, with its weight of 720 kg. "Zephyr 7 "UAV. "Zephyr 7" UAV is designed by the United Kingdom Qinetiq Company, with its structure of carbon fiber composite materials. It has a small wingspan. It is able to reach 336 hours of continuous flight at 21.6km, with its weight of 53 kg. "Condor" UAV project proposed by the US Defense Advanced Research Projects Agency, with its rated power of 5kW. The target aircraft flight height is 18 ~ 27 km. The payload can carry is 450kg; is a load-carrying flight. The first phase of the project is the Aurora’s Odysseus, Boeing’s Solar Hawk and Loma's Norma. "Solar Impulse 2" solar aircraft is designed by the sun team of high-altitude long-endurance manned aircraft, whoosh’s flight required energy entirely from solar cells. A large number of new carbon fiber composite materials are used in the aircraft’s manufacture. The wingspan is 72 meters, weighing about 2300kg, can be produce 340 kilowatt hours of electricity for day. "Soarer", UAV is China's first solar aircraft, design and manufacture by Dr. Li Xiaoyang and Professor Zhao Yong. "Soarer" UAV’s body and the wing using carbon fiber and Kevlar, balsa and other materials, the upper wing and horizontal tail are laid monocrystalline silicon solar cell sheet. The energy storage system is the nickel-metal hydride battery pack, propulsion system is a slow speed propeller. "Green Pioneer I", solar aircraft is the green pioneer program’s first generation of sample machine. The whole machine, the lower wing surface set up a high conversion rate of flexible solar arrays as a power and control, task equipment, energy collectors. "Rainbow", is the world's largest solar unmanned aerial vehicle following the United States NASA series. A particular problem is related to efficiency of the system, because it is only about 10%. There are particular types of solar cells that are silicon-based and compound solar cells; in particular the most widely used monocrystalline silicon panels have high photoelectric conversion efficiency (15% -20%), but the structure quality is high, or Amorphous silicon a-Si thin film solar cells as an alternative of monocrystalline silicon, the current conversion efficiency increased to 12% or more. There are also solar cells that have a higher conversion efficiency than silicon cells like those at Gallium arsenide (GaAs) (43% theoretical efficiency, 28% efficiency), but expensive. [2] Solar Power for Small UAVs Nowadays an important problem is to fly to a long time carrying high loads. There is an alternative to extend electric UAV endourance other than adding batteries, making improvements to the electrical systems, or redesigning the UAV to be more aerodynamic. This alternative is proposed by Alta Devices, which has developed an extremely lightweight, flexible, and thin, gallium arsenide (GaAs) solar cell that holds the single and dual junction world record efficiency at 28.8% and 31.6% respectively. This type of cells can produce a Power Output of 260 W/m2 and weigh 130 g/m2; because the solar cells are thin and flexible, they can be adhered directly to a wing or fuselage surface with negligible impact to aerodynamics. It is also possible to place the cells into a carbon fiber or a fiberglass mold, so that the solar cells become an integrated part of the aircraft's structure. With this solar cells it was possible to reach 9 hours of flight time (Fig.3). [3] Figure SEQ Figure \* ARABIC 3 Solar UAV 3.3. Design of a high altitude long endurance flying­wing solar powered unmanned air vehicle The high altitude aircraft that are in the stratosphere (about 20-30 km of altitude) can provide a useful platform for sensors that support a variety of military and civilian surveillance tasks. These may include real-time monitoring many other things as seismic risks or volcanic areas, early forest detection, border security surveillance, conducted surveys and power lines, telecommunication services, agriculture monitoring, etc. This means that the aircraft can see or cover large geographical areas at wide angles and, in addition, the altitude offers some protection in terms of possible interception by hostile vehicles. One objective of the present research work is to investigate the design and optimization of a swept, true flying-wing configuration for application to HALE UAV operations. A swept, flying-wing configuration is to be studied for operation at these flight conditions, designed to meet a particular mission requirement. Specific topics to be considered will be development of a fast and accurate-as-possible multidisciplinary optimization tool able to design and optimize such a flying wing HALE UAV. The methodology which is adopted in order to complete the present project will involve the following steps. 1. Develop a multidisciplinary optimization tool which should have the following characteristics: a) computationally fast as possible with sufficient accuracy of solution; b) employ inviscid and viscous computational fluid dynamics (threedimensional CFD); c) include the influence of structural elasticity and weight prediction; and (d) evaluate stability of the vehicle. 2. Evaluate the initial design according to a mathematical model for a solar powered aircraft to meet particular mission requirements which is 17 kilometer of altitude, 6-month endurance (March 1 to September 1) operating at 31-Nord latitude (south of Iraq) with 50-kilogram payload for early forest fire detection purpose. 3. Apply to the flying-wing configuration for the specified mission profile and find an optimal design solution. This paper presents a design and optimization framework employing aerodynamics and structure influence for a HALE, solar-powered flying wing UAV. By using modified low-order analysis tools that are employed to facilitate efficient computations for the various engineering analyses, a good approximation is achieved compared to the experimental data. Applying the flying-wing con¦guration in this tool leads to an aeroplane that can be trimmed to carry the designed payload for long endurance (6 months) at latitude 31 N and altitude 17 km. As expected, the static longitudinal stability requires the design to be swept and twisted to give a moment arm behind the neutral axis to trim the aeroplane at the reference lift coefficient. [5] 3.4. Conceptual design of a high-endurance hybrid electric unmanned aerial vehicle Unmanned aerial vehicles can be advantageously used for aerial surveillance, with certain missions requiring no detection of the aircraft. To use such an aircraft for clandestine surveillance, it is mandatory to keep the acoustic, visual and even the heat signatures as low as possible. The pacing element in designing a conventional aircraft propulsion system is the sizing of the power plan to the maximum power phase, take-off and climb. During the cruise and surveillance phases, which usually occupy a longer portion of the mission time, the power demand is much lower. It is possible to use an hybrid-electric propulsion system. Specifically, for a surveillance mission, the surveillance can be operated using silent electric propulsion only, with the internal combustion engine contributing to a fast or long-enduring cruise phase. Both systems may be operated together to achieve maximum power, so that both power units may be downsized, thereby saving weight, compared to a conventional system. Current Electric Motor/Battery UAV. The present design is based on the following requirements: vertical take-off and landing capability; conventional short take-off and landing capability; cruise, loitering and hover capability, low cost of manufacture and maintenance, and have a low noise electrical power unit. A summary of a typical mission is given in Fig. 4. Figure SEQ Figure \* ARABIC 4 Typical Mission In addition to the flight phase requirements, the UAV is capable of: 1) Endurance of at least 30 minutes, which includes cruise plus hovering plus take-off and landing. 2) Cruise speed of 20 m/s. 3)Payload of up to 0.5 kg. This will usually consist of several 0.2 kg camera placed on a light-weight gimbal. 4) Operational radius of 15 km. 5)Operational altitude of 10 - 300 m. 6) Collapsing for transportation convenience. For vertical flight mode, there are bi-, tri-, quadro-, hexa-, and octo-copter configuration options. For the current concept it is important that weight is kept to a minimum and hence the number of electric motors should be kept to a minimum. To attain this, and to keep the desired flight maneuverability, a tri-copter arrangement was decided upon. So as to avoid a complex manufacturing process associated with the concept, the choice of a T-frame with a slight modification was made as illustrated in Fig. 5. Figure SEQ Figure \* ARABIC 5 Selected T-frame configuration The airfame and propeller arrangement are summarized in Fig. 6. The UAV has three rotors, two on the main wings to provide thrust for VSTOL and forward flight and one mid-fuselage to provide general stability and yaw. The take-off weight is approximately 6 kg, the wing span is 2 m, the UAV is electrically driven and the propellers have a NACA 4412 cross-section and are 254 mm in diameter. Figure SEQ Figure \* ARABIC 6 Final UAV design Propulsion system. An electric propulsion system consists of the battery (energy storage), electric motor and propulsion device (propellers). Some advantages of such a system are that it is easy to operate and is reliable. Also electric motors are renowned for their precise control characteristics and fast response to throttle input. A disadvantage is that this system is limited by the capacity of the batteries. The maximum thrust required for vertical-take-off (or maneuvering) is estimated to be 𝑅𝑒𝑞𝑢𝑖𝑟𝑒𝑑 𝑇ℎ𝑟𝑢𝑠𝑡 ≈2×𝑊𝑒𝑖𝑔ℎ𝑡 (1) For this given design this works out at 22.8 N thrust per propeller at 100% throttle is required. The electric motor chosen has an out-runner as this was easier to maintain and was quieter and cheaper. The motor is brushless to avoid sparking. When it came to battery selection the three main considerations are, capacity, weight and volume. There are various types of battery on the market as summarized in Table 1. Table SEQ Table \* ARABIC 1 Batteries available An electric motor capable of producing the required thrust at 75% throttle was chosen which required 22.2 V. The power of the motor was 324 W again at 75% throttle, with 75% used as a built in safety margin. This required a LiPo battery pack consisting of six cells in series which had the specification of 8000 mAh. The battery pack, which weighed 1.1 kg, was capable of producing flight for 30 minutes. Difference between electric motor/battery powered UAV and conceptual hybrid-electric powered UAV. This work looks at the feasibility of improving the flight endurance achieved by the current electric UAV shown in Fig. 7 to an improved flight endurance shown in Fig. 8 by the conceptual hybrid-electric UAV. Figure SEQ Figure \* ARABIC 7 Region of interest for current electric motor/battery powered UAV Figure SEQ Figure \* ARABIC 8 Region of interest for conceptual hybrid-electric powered UAV The design of the hybrid-electric UAV demonstrates that an extended cruise endurance hybrid-electric UAV is feasible. The cruise endurance extension, however, comes at the expense of reducing the electric energy onboard. [5] 3.5. Performance analysis of solar powered Unmanned Aerial Vehicle One of the main problems in micro Unmanned Aerial Vehicles (UAV) is endurance or flight time since the general domain aircraft use conventional fuel. In today's world, there are more than 11,000 UAVs in service by the Military for various purposes. Unmanned Aerial Vehicles (UAVs) are the ones which can fly either remotely or autonomously. In spite of their usage in various ap-plications, they lack in performance due to power restrictions and also nowadays the ability to fly without using conventional fossil fuels is primarily focused both in application point of view and as well as in scientific fields, since the major concerns are an increase in global warming and a decrease in natural resources. Since then, the use of electric aircraft have been widespread but here the crucial issue is their high power consumption when compared to their limited energy storage capability. So if either by increasing the size of the battery or by incorporating more batteries will only lead to increasing the weight of the plane which directly affects the flight time of UAV. One of the possibilities to increase the flight time is by using unlimited solar energy through solar cells. So, the possible solution to enhance the endurance is by using solar-powered aircraft driven by electric-based propulsion systems in which the power is supplied continuously throughout the day by solar energy which can eliminate fuel and also solve the limited energy storage capa-bility problem. When it comes to the performance of solar cells, the perfor-mance of a Photovoltaic system depends not only on its basic characteristics but also on the environmental issues. One such environmental issue is the ambient temperature which plays an important role in the photovoltaic conversion process. The solar cell efficiency is usually measured under standard test conditions (STC), with PV cell temperature of 25 °C, irradiance of 1000 W/m2 and air mass ratio AM= 1.5, but these conditions are rarely met at outdoor installations, as the ambient temperature and wind speed affect the performance of the module for that particular locality. The open-circuit voltage decreases significantly with increasing PV module temperature (values are up to -0.45%/K for crystalline silicon) whereas the short circuit current increases only slightly (values range between 0.04 and 0.09%/K). There is a method to estimate the number of solar cells and how they must be arranged on the wing The main principle is to make use of available solar energy by converting it into electricity through solar cells. Here the cells are arranged in series on top of the wing to get the required voltage in order to safely charge a 3S battery (3 Lipo batteries are connected in series and be used as a single unit) and from there the battery power is supplied to the motor for throttling during constant level flight. The main objective is to improve the duration of flight time. In Table 2 there are the Mission specifications, about the plane and the weather, while in Table 3 there are the features of C60 solar cell. Table SEQ Table \* ARABIC 2 Mission specifications Table SEQ Table \* ARABIC 3 Specifications of sun power C60 solar cell For charging the 3S battery, it requires a constant safe charging voltage of about 12.4 V. The solar cells selected for this design were Sun power C-60 monocrystalline photovoltaic cells having single cell efficiency of 22%, these cells were more efficient than most silicon-based solar cells of 15% efficiency. According to Table 2, each solar cell gives out 0.57 V which means 22 solar cells are required to meet the targeted voltage of 12.4 V, here two more cells are added for safe side and finally designed for 24 cells. These 24 cells should be connected in series to achieve the required voltage. From the number of cells, I calculate the minimum wing area. There are two possible layout to arrange the solar cells on the wing, Type-1 and Type-2 as shown in Fig. 9 and Fig.10 respectively. Figure SEQ Figure \* ARABIC 9 Single row of cells - High Aspect ratio wing Figure SEQ Figure \* ARABIC 10 Two rows of cells - Medium Aspect ratio wing In the design shown in the Fig.11 there are some problems: There is a problem bound to weight; the solar cells are arranged in a single row, the wing is too long which requires a tough structure leading to increasing in weight of the plane. There is a problem of stability; controlling the wings to stabilize the plane during flight will be difficult without of ailerons. There is a problem of power efficiency; while connecting the electronic circuit of solar cells with tapping wire the two ends of the wire are too far to close the circuit which results in loss of power efficiency. While in the design shown in the Fig.10, the problems presented on the Fig.9 are solved as the solar cells are arranged only on the middle portion of the wing in two rows (each row has 12 solar cells) which helps in reducing the total length of the wing to accommodate the solar cells. The side wings (ailerons) are used for stability purpose having a polyhedral angle of about 7° on both sides for easy control of the plane. As each solar cell is of 0.125 m in length, for 12 solar cells in a row a wing span (middle portion) of 1.5 m is required. But, apart from this length, extra space is required for soldering purpose and also at the ends at least 2-3 cm should be left because there will be abnormal forces at the junction. The aspect ratio (which is the ratio of wing span to the mean chord) of the entire wing is 9,36. Figure SEQ Figure \* ARABIC 11 Arrangement and connection of solar cells on the mid-wing And also from Fig. 11 it is observed that in Type-1 more tapping wire is used in order to close the circuit which means more loss while transferring the current but in Type-2 this problem is nullified as we can close the circuit with the usage of very less wire. After finalizing the wing design the airfoil is selected based on the following requirements: high coefficient of lift (CL), low coef-ficient of drag (CD) and less camber for cell placement. Power required for flight. The lift force (L) generated by the wing compensates the weight of the plane (W) and the thrust (T) pro-duced by the propeller compensates the drag force (D). The lift force generated by the wing was calculated as : The lift force balance the weight force then The and thrust T balances the drag force D: then Where: m is the mass of the plane (kg); g is gravity (m/s^2); r is the air density (kg/m^3); S is the surface area of the wing (m^2); V is the cruise velocity of the plane (m/s); CL is the coefficient of lift, CTD is total drag coefficient. The total drag on the wing is the sum of induced drag (CID) and profile drag of airfoil (CD). The induced drag is a consequence of the produced lift that is not aligned with the force of the gravity. The induced drag and the total drag coefficient can be calculated by: Oswald's efficiency (e): Where AR is the Aspect Ratio. Table SEQ Table \* ARABIC 4 Comparing the four selected airfoils Airfoil-4 WE3.55 was chosen for its best performance, as it has high CL/CD ratio which is the same airfoil designed for Sky-Sailor. As the mass of the entire plane was initially fixed as 2 kg, the total weight becomes 19.6 N. From Eq. (1) the cruise velocity of the plane can be found out as 7.96 m/s. And the total drag for the entire plane is1.96 N. From the total drag and the cruise velocity, the power required (Preq) for a level flight can be calculated as But considering the different efficiencies (motor, propeller, etc.) the Power required for cruise is . So, the minimum solar irradiance required to maintain level flight can be calculated as: So, 631.4 W/m2 of solar irradiance is required in order to maintain level flight. By analyzing different airfoils the suitable airfoil which has less camber for solar cell placement was chosen by XFLR5 analysis, in which airfoil WE3.55 was chosen for its best performance as it has high CL/CD ratio of 78.5. With this airfoil the wing span and chord was decided according to the battery requirement, in which here for 3S battery a safe charge voltage of 12.4 V was needed. Then energy and power analysis was done theoretically to check whether the design would be feasible or not, which gives a basic foundation in order to design a solar-powered UAV. Then the integration of solar cells on to the wing was per-formed by considering various factors such as the deflection, sta-bility and circuit connectivity. The deflection of the central wing was tested experimentally in which for a uniform load of 4 Kg a deflection of 1.7 cm was observed for a wingspan of 1.63 m. Then the solar wing was tested on ground level on 15th of April and May from morning 9 a.m. to evening 5 p.m. (IST), in which until 3.45 p.m. more power (on average 55 W) was obtained from the solar cells than the required power (40 W) for the level flight, which shows that with this power the plane can fly with only solar energy continuously for more than 6h. Then the entire fabricated plane was tested for the whole day in which we observed the inverse relationship between the cell temperature and the output voltage of the cell. [6] 3.6. Space-Based Solar Power a technical, economic, and operational assessment There is three type of assessment: Technical Assessment: This section introduces the basic concept of space-based solar power (SBSP). Finally, the section examines the critical technologies required for the successful development of the space, ground, and support elements of the system. Economic Assessment: This section examines SBSP system cost estimates from a variety of sources. It then compares these costs to competing alternative energy solutions such as terrestrial-based photovoltaic power plants. The section also addresses regulatory factors that may affect the development and operation of SBSP systems as well as current international Operational Assessment: This section explores the strategic considerations for SBSP systems within the general context of national space operations. It then examines potential garrison-level applications and compares these with the current plans of the Army’s Energy Initiatives Task Forces to integrate terrestrial based photovoltaic power into the energy systems of several major installations. Finally, the section briefly explores possible SBSP applications to support remote operating locations. Energy security is an issue that continues to become more acute as global populations grow and limited fossil fuel reserves shrink. The promises of space-based solar power for clean and unlimited energy for all humankind are certainly appealing. But the reality is that such systems always seem to be seen as just 10 years away by their advocates. While significant progress continues in the enabling technologies that will make SBSP systems economically viable and competitive power generators, there is no compelling evidence that such systems will provide the best energy solution. Considering the austerity of current federal budgets, the Army’s evolutionary approach to incorporating a diverse portfolio of renewable energy sources in distributed locations seems more prudent than placing significant amounts of resources in highrisk ventures such as SBSP systems. Perhaps in a decade or so, there will be technological breakthroughs that will fully support practical SBSP systems. But it is also possible that within that decade there may be breakthroughs such as fusion energy exploitation, which will make SBSP systems obsolete before they are even fielded. [7] Mission Analysis of Solar Powered Aircraft This is the third of a series of studies on solar powered aircraft. The first examined the feasibility of solar powered aircraft and focused on the identification of critical technologies that must be further developed if solar-powered aircraft must become a reality. The complete methodology developing the first plan established for the feasibility of solar powered aircraft for certain missions. The second study examined various structural schemes and identified critical structural elements technologies that need to be developed. The present study emphasizes the general concept of mission operations of a solar HAPP. The main product is a chronology of preliminary missions that identifies the events that it must take place to realize the whole mission. Events in the time line and interface evaluation, data and control links and peripheral equipment provide information on the operation of the generic class of long-haul high-altitude aircraft. This relationship is organized along the lines followed in conducting the study. A brief review of the previous one related work following this introduction, describes the HAPP Solar which includes a functional arrangement of subsystems and general system performance. Solar HAPP System description This type of aircraft was nominally designed to fly for one year at 20 km over the Great Centeal Valley of California with a 110 kg of payload. Analysis of board equipment (Fig. 12) 1. Energy collection and storage subsystem comprising solar cells, electrolyzer, reactant storage, fuel cell, and power conditioner and associated controls. 2. Rotating components including motor, gearbox and propeller. 3. Autopilot and system monitor including electronics and actuators necessary for flight control and system monitoring functions. 4. Navigation subsystem. 5. Payload, sensors and equipment necessary to process and transmit mission data to ground stations. Figure SEQ Figure \* ARABIC 12 Solar High Altitude Powered Platform Power Train A power conditioner (also known as a line conditioner or power line conditioner) is a device intended to improve the quality of the power that is delivered to electrical load equipment. Conditioners specifically work to smooth the sinusoidal A.C. wave form and maintain a constant voltage over varying loads. Ground equipment is divided into two groups. The first is the system monitor and control made up of equiment required for launch, mission control system monitoring and recovery. The second is the fata subsystem which comprises equipment and software required to receive and process data from the airplane to prepare it for display and/or interrogation.[8] 3.8. Conceptual Multidisciplinary Design Optimization (MDO) of Solar Powered UAV The main challenges of solar powered air-vehicle design are high efficiency propulsion system along with very efficient air-vehicle, i.e. low power requirements. These two challenges make the design of a solar powered air-vehicle a candidate for multidisciplinary design optimization (MDO) effort. Design Challenges. To make a solar powered vehicle successful, namely having sufficient design margins, two technological hurdles should be considered: the first one is the efficiency of the solar panels and the second one is the energy storage system (i.e. the battery pack). The Fig.13 presents solar panel efficiencies of several leading technologies that have been developed during recent years. It is noticeable that there is a consistent progress, but still efficiency is not exceeding 40%. Moreover, because of practical considerations (cost, technology maturity, weight, etc.) the actual efficiency is in the range of 20%- 25%. Figure SEQ Figure \* ARABIC 13 The evolution of solar panels technology and efficiency The Figure 14, shows energy balance for a 24-hrs solar powered vehicle. At night the solar power is zero and the vehicle uses its energy reservoir. Shows the importance of energy storage for such vehicles. The most common and reliable type of energy storage means are electro-chemical batteries. Figure SEQ Figure \* ARABIC 14 Energy balance for a long endurance solar power vehicle The Figure 15, shows energy densities of several battery types. Nowadays the practical value is about 200÷250 Wh/kg and one should remember that 100% usage of this energy capacity is not possible. In most cases only 80% ÷ 90% of the stored energy is available for usage. Beyond that, after using most of the stored energy, the battery voltage drops sharply and may risk further operation of the air-vehicle. The two technological challenges (solar panels efficiency and energy storage capability) result in a large vehicle that allows for a sufficient solar panels area, but at the same time this vehicle must have a relatively low power demand (thus reducing the weight of the batteries). Figure SEQ Figure \* ARABIC 15 Batteries energy density Analisis Model. The design process uses several analysis models: solar radiation model, air-vehicle model, and energy balance. Based on the air-vehicle geometry, its drag polar and weight are estimated. The location, date, and time of day determine the available amount of solar radiation. These data (drag polar, weight, and solar radiation) are then used by the energy balance module taking into account the flight conditions of the vehicle. The energy balance model makes sure that the mission can be accomplished, thus determines if the considered configuration is feasible for the specific mission. The Solar Radiation Model is used to to find the available solar power to enable the air-vehicle flight. The model takes into consideration the date and time, longitude and latitudes, and the radiation atmospheric absorption. The model was implemented using Matlab and was validated versus available meteorological data. The Figure 16 shows the radiation in Israel during a representative August day. The figure contains the radiation level outside of Earth's atmosphere, total radiation at sea level, and the radiation at a surface horizontal to the earth. The difference between the radiations on the ground and outside the atmosphere is due to the atmospheric absorption. During darkness the air-vehicle is supported by its internal power source (batteries), thus accurate sunrise and sunset times are essential for a proper simulation of the air-vehicle mission. Figure SEQ Figure \* ARABIC 16 Solar radiation. Comparison between the model results and meteorological measurements The Air Vehicle Model model includes two sub-models: aerodynamic model and mass estimation model. These models define how the design parameters affect the ability of the vehicle to accomplish its mission. i.e. Based on the air-vehicle geometric characteristics, the air-vehicle models estimate its aerodynamic and mass characteristics. The aerodynamic model is based on drag bookkeeping. A conventional main wing and cross tail are considered. Using the air-vehicle dimensions, geometric parameters such as aspect ratio and wetted areas are calculated. Using semi-empirical models, the vehicle drag polar is estimated. The mass of the vehicle is estimated using a bookkeeping of the following vehicle components: Structure, Propulsion system, Batteries, Solar panels and Payload. Energy Balance Model As mentioned above one of the main challenges in solar air-vehicles design is to maintain a positive energy balance, thus the required energy to operate the vehicle is equal or less than the available energy absorbed from the solar radiation (Fig. 10). Three flight conditions are considered: straight and level, climb, and descent. During a straight and level flight at a constant speed the drag force, D, is equal to the thrust force, T, while the lift force, L, is equal to the vehicle weight, W. The available power from the solar radiation is estimated according to the available radiation, Q, multiplied by the solar panels area, A, and factorized by the solar panels efficiency, ηSolar: During the air-vehicle mission, in phases that have higher available power than required power, the surplus power is used to charge the batteries. Note that while the available solar power is below the required power (e.g. at night time), the deficit is compensated for by the charged batteries. Thus a total too low battery mass, namely a too low energy storage capacity, may prevent accomplishment of the mission. Design definition and implementation. A mission analysis can be conducted using the above models. The mission analysis uses the models to simulate the air-vehicle behavior along a specific flight pattern. For a given air- vehicle geometry (e.g. wing span, b, aspect ratio, AR) and mission specifications (e.g. day loiter altitude, night loiter altitude), the vehicle characteristics are estimated and a mission cycle is simulated. If the energy balance is negative, the specific mission is not feasible with the specific vehicle geometry. The design problem is then defined as a math programming problem solved with Matlab A representative design case. Representative design cases of a solar UAV conceptual sizing are considered. In these cases a payload is carried to a mission at altitude of 70,000 ft. To enable a "never ending" aloft mission, during night time the airvehicle is allowed to descent to lower altitudes and stay there using battery power which was charged during the previous daytime. The mission is defined as takeoff at sea-level with fully charged batteries, at sunrise time. Climb to 70,000 ft at a rate-of-climb of 200 ft/min and staying at that altitude until night time. Then descent to a lower altitude, stay there for the night time, and at sunrise climb again to 70,000 ft. Figure 17 shows a schematic description of the mission definition. Figure SEQ Figure \* ARABIC 17 Schematic mission description The radiation model is based on a well-established formulation that includes the dependency on the time-of year, geographic location, and time-of-day. This model is validated versus actual measurements. The aerodynamic model is based on a simple drag bookkeeping which takes into account the various components of the vehicle. The weight estimation is based also on bookkeeping of the various air-vehicle components. The most challenging mass estimation model is the structural mass model. It is based on statistical regression of existing vehicles, depending on the wing geometry (e.g. span and aspect ratio). Nowadays most designers are aware that MDO is an important aeronautical design tool, especially during the conceptual sizing phase. For a solar vehicle design the importance is even higher. The challenge of enabling a positive energy balance under additional complex cross-influences (e.g. solar panels area and wing area), makes solar air-vehicle design a truly complex multidisciplinary design problem that will be difficult to carry out without using optimization tools. [1] Optimizing Electric Propulsion Systems for Unmanned Aerial Vehicles Most of these UAVs are equipped with electric motors that contribute to the simplicity of operation and significantly reduce their noise signature. The propulsion systems of these small UAVs (batteries, motor, propeller, etc.) account for as much as 60% of the vehicle weight. The electric propulsion system of a typical UAV includes the following components: propeller; electric motor, energy source, gear box (optional), driver, wiring, plugs and connectors, and cooling system (optional). It will concentrate only on propeller, electricmotor and the batteries. The Betz theory is used for propeller design; this approach is based on optimizing the propeller’s geometry at a certain specific operating condition (a certain combination of airspeed, altitude, and propeller rotational speed), such that the power, which is required to obtain a certain propulsive force at these operating conditions, is minimized (or, alternatively, the thrust produced by a certain power is maximized). During MDO, all of the different design goals and constraints are addressed simultaneously. Thus, the best compromise between contradictory design goals will be reached. This kind of approach has been already applied to rotary wing designs. Most of the previous investigations were limited to two disciplines: either aerodynamic and structural analyses, mostly for helicopter rotors, or aerodynamic and acoustic design of propellers. Most of the previous investigations used a very limited number of design variables, rather than the full range of design parameters that are under the authority of the designer. Previous investigations also included a limited number of constraints and did not consider the entire propulsion system (namely, the coupled system: propeller, gearbox, engine, and energy source). For example, does not take into account the engine characteristics during the propeller design. Only after the isolated propeller optimization are these characteristics used to calculate the performance of the entire propulsion system, which is the measure of the design quality. Recently, a new comprehensive MDO design tool for propeller-based propulsion systems was presented by the authors that offer savery high flexibility in choosing the cost function, design variables, and constraints. It is clear from previous studies on electric UAVs and other propulsion-based propulsion systems that an optimization of the electric propulsion system of a UAV should include a simultaneous consideration of: propeller, electric motor and battery. The performance and characteristics of the vehicle depend on the strong interaction between these three. The aim of the present study is to represent the ideal of a particular electric propulsion system for a given UAV, but rather to investigate trends and obtain information on the interactions between the various components of the system. To do this, theoretical models of these components are needed. Although the helix model of the present study is based on the well known theory of the blade / moment element, a particular effort is directed to the modeling of the electric motor and the battery. These models are based on a comprehensive survey of existing engines and batteries, followed by derivations of representations of some intermediate parameters. All the various models are combined with various optimization schemes to form a complete MDO object that allows you to manage a number and a variety of variables and design constraints. There is also great flexibility in choosing the cost function: that is to say, the goal of design. Thus, the MDO object was used to optimize propulsion systems for different design objectives (cost functions) and constraints. Common tools for the structural analysis of blades are finite element models. To reduce the computations, a more efficient rod model, together with a transfer-matrix formulation, is used. The rod structural model describes the propeller blades as a series of straight segments located along the blade’s elastic axis. The structural cross-sectional properties are uniform along each segment andequaltothestructural properties ofarepresentativecrosssection of that segment. The transfer-matrix formulation is applied using the boundary conditions of a cantilevered rod (clamped root and free tip). The solution procedure is very efficient and the results are very accurate. Electric Motor Model. There is a relation between a motor maximum output power and its size/weight (Fig.18) This figure shows 3 types of electric motors: Contains manufacturers of electric motors for heavy / high-voltage applications. contains manufacturers of high-performance industrial electric motors. contains the manufacturers of aeromodel / hobby electric motors. Heavy duty engines have a low weight / power ratio, while model aircraft engines have a high weight / power ratio. It is also possible that the aeronautical engine offers a high power-to-weight ratio, so model aircraft engines are natural candidates for UAV applications. However, UAV engines are expected to exhibit much better reliability and endurance than aeromodel engines. This aspect, as well as the experience with existing UAV engines, leads to the conclusion that Group II is a better representative of UAV engines. Figure SEQ Figure \* ARABIC 18 Motor maximum output power as a function of motor mass Battery Model. An important component of an electric propulsion system is the battery pack. The battery often represents one of the heaviest components of the entire vehicle. One of the most common types of batteries is the lithium polymer (LiPo), which offers a relatively high energy capacity along with low weight. The Figure 19 shows the battery energy capacity EB as a function of the battery mass mB. Figure SEQ Figure \* ARABIC 19 Battery capacity as a function of battery mass An extensive effort was made to optimize electric propulsion systems for UAVs. The method is based on an MDO approach that includes tools for aerodynamic, structural, electrical and performance analysis. These analysis tools are combined with three different optimization schemes to achieve optimal design based on various design goals, using various design variables and based on various constraints. Although the aerodynamic and structural models of the propeller have been presented elsewhere, the derivation of the electric motor model has been presented in detail. This is a simplified model suitable for optimization: a large number of analyzes are performed and therefore a numerically efficient model is required. This model includes four parameters that are defined on the basis of the examination of a large number of engine data from various manufacturers. A model of the energy capacity of the battery pack according to its mass is similarly derived. In this case, a new identification system was applied for a mini electric UAV. The design includes propeller, motor and battery. There are two important performance indicators for this vehicle: the loitering time, which is directly related to its main reconnaissance task and should be maximized, and the climb rate, which is directly related to the vulnerability and safety of the vehicle. The study begins with optimal designs for a single helix objective for maximum dwell time (which produces a zero climb rate due to propeller constraints) and a propeller for maximum climb speed (which leads to zero dwell time due to zero battery mass). The optimization is canceled with the maximum support for the masses of memory and battery. Although these two projects are impractical, they are important because they give an idea of ​​the two extreme trends in design. Thus, for example, because the propeller radius is limited, the maximum climb speed results in a relatively low propeller efficiency. To achieve the high thrust required for a high climb rate, the design is able to easily design the propeller design for maximum dwell time. It can be concluded, based on the sensitivity study, that the battery density and maximum power-to-mass ratio have the largest influences on the design and performance of the system. The other parameters have a relatively minor influence on thevehicle performance, but they may influence the optimal design. The present study shows once more that when optimizing a propeller-based propulsion system, it is essential to simultaneously consider all the components of this system: propeller, motor, and energy source, as well as the vehicle’s characteristics. [9] Design of a high-altitude long-endurance solar-powered unmanned air vehicle for multi-payload and operations Unmanned air vehicles (UAVs) have been used for over 40 years by the military services for dangerous missions and surveillance. The technology has been relatively expensive and unreliable, as cost and safety issues were considered of secondary importance to performance. However, technology has advanced sufficiently for this highly technical area of the aeronautical industry to be developed into a new growing UAV civilian industry. The United States has already begun investing in this field and has initiated a massive drive to develop this new industry. Europe also has the necessary technologies and should now also support and encourage industry to take advantage of new technological approaches in order to successfully become a leading participant in this field. he largest market shares are expected to pertain to Coastguard and Maritime Surveillance operations, Border Security, and Forest Fire Management. In 1995, with the funding of the Italian Space Agency (ASI), the Department of Aeronautical and Space Engineering of the Polytechnic of Turin (Scientific Coordinator Prof. G. Romeo) proposed the HeliPlat solar energy UAV aircraft as a high altitude platform very long autonomy. The project continued in the context of projects funded by the European Union and is currently still underway with the aim of designing an HAVE / UAV (High Altitude Very Long Endurance / Unmanned Air Vehicle) aircraft. From a flying altitude of 18 km, an area of about 300–400 km in diameter would be covered for communication transmission if the onboard antenna irradiation diagram is properly chosen. This would mean that about seven to eight platforms could cover the entire south Mediterranean Sea, from Spain to Turkey (Fig. 20-21), creating an electronic illegal immigration barrier as well as a secure border control system. Compared to the present cost of airborne systems (5000–7000 E/flight hour), high altitutde long endurance/unmanned air vehicles (HALE/UAVs) offer clear advantages in monitoring missions, as continuous observations will be performed throughout the year over the area of interest and all the required data will usually be available immediately. The project has a total cost of about 4 million euros, of which three quarters are covered by European funding. Figure SEQ Figure \* ARABIC 20 HeliPlat Border Surveillance over the Mediterranean Sea Figure SEQ Figure \* ARABIC 21 HeliPlat The main advantage of the very-long endurance solar-powered autonomous aircraft (VESPAA) is that this system has less climbing and descending events, which is important when considering interference with aviation traffic. Other HALE-UAV configurations have a very limited endurance (24–36 h), which would drastically increase any potential collision risk with civil aviation traffic. Double the number of UAVs would be necessary to continuously guarantee the surveillance service, thus the system total life cycle cost would be increased to a great extent. Other medium altitude long endurance (MALE)-UAVs have, as a further disadvantage, the fact that a much higher number of UAVs are necessary to continuously cover the entire Mediterranean Sea, since the covered area decreases with the square value of the flight altitude, if there are low-altitude UAVs, more units are needed; Figure 22 shows an example of an MALE-UAV and a HALE-UAV with 2 different altitude. With a “MALE” configuration, the total cost would increase because more unit are needed. Figure SEQ Figure \* ARABIC 22 HAVE / MALE monitoring area comparison An endurance of several months would only be possible using a solar- and hydrogen-powered platform. The vehicle would climb to 17–20 km mainly taking advantage of direct sun radiation. Any electric energy not required for the propulsion and payload operations would be pumped back into the electrolyzer system that would convert the water into hydrogen and oxygen fuel; during the night, the platform would maintain altitude, thanks to the fuel cell system, which would produce electricity and water as a by-product of the reaction; the geostationary position would be maintained by a level turning flight. The main innovative aspects of the VESPAA are the following. 1. The design of the first European aerodynamic solar-powered platform that would be able to remain continuous in flight for very long periods of time (4-6 months). 2. Low noise and zero emission aircraft. The advantage of zero emissions is particularly important for environmental monitoring. 3. The first European solar-powered platform to validate the feasibility and reliability of solar cells, fuel cells and brushless electric motors. All these features lead to: a) reduced cost per flight hour (thanks to a large increase in the endurance flight hours); b) potentially increased acquisition cost, while reducing the cost of maintenance and spare parts; c) reduced cost – larger area coverage per platform, thus, fewer platforms per area are required; d) improved operational safety – due to the fact that the flight would occur above aviation traffic, resulting in limited interference with it, and above adverse weather; e) lower propagation delay compared to satellites. Fuel cell energy storage system (Fig. 23) Day cicle. Sun energy converted to electricity by solar cells. Half of electricity goes to motor to propel plane. Other half of electricity goes to electrolyzer to converter water into hydrogen and Oxygen fuel. Night cycle. Oxygen and Hydrogen combine in fuel cell to produce electricity to propel plane.Water from oxygen and hydrogen stored until next day. Figure SEQ figure \* ARABIC 23 Fuel Cell Energy Storage system enables continuos flight through night The results of this preliminary study show that it could be possible to obtain a very long endurance high altitude platform for Earth observation and telecommunication applications, at least for low latitude sites in Europe and for several months of continuous operation. A BWB (Blended Wing Body) configuration of Solar HALE Aircraft has been obtained as a result of the preliminary design. The BWB solution seems to be the best compromise between performance available surface for solar cells and volume for multi-payload purposes. The numerical aerodynamic results obtained show the availability of suitable airfoils for operation at low Reynolds numbers. The structural concept layout has been defined and designed according to regulation requirements. The preliminary analytical and FE results showed a good correlation. The assumed structural concept, sandwich wing-box spar, demonstrates low structural mass and satisfactory classical flutter behaviour. According to airworthiness regulations a preliminary flight dynamic and flutter analysis has been carried out. [10] Design of solar high altitude long endurance aircraft for multi payload & operations Research is being carried out at the Turin Polytechnic University with the aim of designing a HAVE/UAV (High Altitude Very-long Endurance/Unmanned Air Vehicle). The vehicle should be able to climb to an altitude of 17–20 km by taking advantage of direct sun radiation and maintaining a level flight; during the night, afuel cells energy storagesystem would be used. A computer program hasbeen developedtocarryout a parametric study for the platform design. The solar radiation changes over one year, altitude, masses and efficiencies of the solar and fuel cells, as well as the aerodynamic performances have all been taken into account. The parametric studies have shown how the efficiency of the fuel and solar cells and mass have the most influence on the platform dimensions. High modulus CFRP has been used in designing the structure in order to minimize the airframe weight. A Blended Wing Body (BWB) configuration of Solar HALE Aircraft Multi Payload & Operation (SHAMPO) with 8 brushless electric motors has been developed, as a result of the parametric study. The BWB solution, compared with conventional designs, seems to provide the best compromise between performance, availability of surfaces for solar-cells, and volume for multi-payload purposes. In particular the platform will accommodate a maximum user payload mass of 100– 150 kg and consume a maximum electrical power of 1–1.5KW. Several profiles and wing plans have been analyzed using the CFD software Xfoil and Vsaero. The airfoil coordinates at the root and along the wing span as well as the wing planform were optimised to achieve the best efficiency. A FEM analysis was carried out using the Msc/Patran/Nastran code to predict the static and dynamic behaviour of the UAV structure. Within the European project "CAPECON: Civil UAV Applications & Economic Effectivity of Potential Configuration Solutions", various configurations of non-piloted aircraft with high altitude and high autonomy have been studied, taking into consideration conventional propulsion systems (Modular configuration & Blended) configuration), is an innovative configuration with solar energy called SHAMPO (Fig. 24), which is considered to be the best compromise between performance, available area for solar cells and volume available for loading. Figure SEQ Figure \* ARABIC 24 External Layout SHAMPO Internal and external layout The electrical energy system is composed of three modules: 1) The photovoltaic energy source module: it converts solar energy into electrical energy. During the day, the solar power is used to supply the electric motors of the propellers and to feed an electrolyser which produces hydrogen and oxygen. 2) The electric motor module: it converts electrical energy into mechanical energy. It consists of an electric motor and an inverter. 3) The fuel cells energy storage module: it acts as energy storage. It consists of the fuel cell array, the hydrolyser as well as hydrogen, oxygen and water tanks. The gases are stored at high pressure (120 bar) and used during the night to feed the fuel cells which supply the motors; the water feeds the electrolyser to finish the cycle. Fig. 25 show a possible installation of this module into the blended part of the wing. Figure SEQ Figure \* ARABIC 25 Internal layout SHAMPO The Photovoltaic Energy Source Module is composed of 21% efficient solar cells installed on 95% of lifting surfaces. A Blended Wing Body configuration with 8 brushless electric motors was developed, as a result of the parametric study. Each motor weight is about 6.5 kg including the carbon-fibre propeller blade. The BWB solution, compared with conventional design seems to be the best compromise between performance, availability of surfaces for solar-cells, and volume for multi-payload purposes. In fact the fuselage blended zone presents a volume of about 38 m3. The Heliplat propeller blade (blade radius = 1.15 m), which was developed during the HELINET Project, has been used for the preliminary design of SHAMPO. The thrust supplied by this propeller in cruise condition is sufficient for the SHAMPO design point as reported in Fig. 26. Figure SEQ Figure \* ARABIC 26 Propeller thrust vs required thrust Using the 5000 W of surplus power available for 2 hours it is possible to increase the airspeed. Considering the maximum power available to the motors is about 9800 W (max propeller thrust NO RES) the airplane can reach 31 m/s. In the case of an emergency or to contrast wind jet-stream, using its reserves, it will be able to reach 36 m/s (max propeller thrust RES). Work is still in progress to define a proper solution for the landing gear system. A retractable tricycle configuration, with one wheel for each leg has only been chosen as a first hypothesis. The nose landing gear has a wheel that is capable of being steered. The main gear acts as a solid spring absorbing the shock due to landing by means of its own deformation. [11] Manufacturing and Design of Lightweight Composite Airplane Structures Part II – Design Structural analyzes FEM (finite element method). FEM analyzes are important when carrying out the analysis of complex structures or engineering the behavior of mechanical systems and machines. Through a geometric grid defined mesh, the FEM method divides the geometric model into many small elements that are easy to calculate. The final solution is found by the system "adding" all the partial solutions calculated for each element. A front of the application conditions, an FEM analysis allows to obtain displacements, deformations and tensions present in a structural system. (Fig. 27-28-29) -A very expensive technology is actually used to manufacture military airplane. - New technlogies has to be improved to reduce manufacturing cost. -The production and operational costs of composite structures must be competitive with metal structures. Weight savings are a bonus, cost is the driver. -Mechanical properties, however, are strongly directional. -A good analytical knowledge is necessary in order to properly exploit their excellent mechanical properties, especially in the detail design of joints, cut-outs and discontinuities, buckling and posto-buckling, etc. - FE analyses have been applied successfully to most composite design. In region of high stress gradients (cut-outs, stiffener drop-offs) a fine mesh must be used. -Concurrent engineering approach: design and production teams to improve comunications. By resolving problems early, costs are reduced. -Design and cerification of composite more conservative- expensive than for metal. Figure SEQ Figure \* ARABIC 27 FEM Figure SEQ Figure \* ARABIC 28 FEM Figure SEQ Figure \* ARABIC 29 FEM Design, manufacturing and testing of a HALE-UAV structural demonstrator The first HELIPLAT (HELIos PLATform) configuration was worked out, on the basis of a preliminary parametric study. The platform was a twin-boom tail type monoplane with eight brushless motors, a long horizontal stabilizer and two rudders. A scaled-pro-totype was designed to demonstrate the feasibility of this configuration and to perform some structural static and dynamic tests on it. The main CFRP structures were manufactured by CASA (Spain): the principal wing and horizontal tail tubular spars, booms, vertical tail spars and some reinforced ribs. These parts were delivered to the Aerospace Engineering Dept. (DIASP) at the Politecnico di Torino (POLITO) and assembled using special joints while some other necessary parts were manufactured by POLITO-DIASP. A parallel activ-ity was performed to define the structural test configurations and structural test frame. The manufacturing activities and the development of the structural test system is described in the first part of the paper. Static and dynamic experimental tests were performed in two phases (2003 and 2004) on the prototype and the results of the static tests are presented in this paper and compared with numerical and the-oretical computations. The HELINET research project (Network of stratospheric plat-forms for traffic monitoring, environmental surveillance and broadband services, Coordinator: Politecnico di Tor-ino) has been financed by the European Commission, since January 2000, as a part of the 5th Framework Program in the Information Society Technology Action, to develop a European project in the field of stratospheric platforms. The HELINET project, which involves several European universities and partner companies, is based on HALE-UAV HELIPLAT. A research was conducted at the Polytechnic University of Turin, with the aim of designing a HALE-UAV solar energy platform and producing a solar-scale prototype. The first limited financial support was obtained by the Italian Space Agency (ASI) in 1995. Only a small part of the research has been completed, but a great deal of experience has been gained in this field. The research project HELINET (Network of stratospheric platforms for traffic monitoring, environmental surveillance and broadband services, Coordinator: Polytechnic University of Tor-ino) was funded by the European Commission, since January 2000, as part of the Fifth Framework Program in the technological action of the information society, to develop a European project in the field of stratospheric platforms. The HELINET project, which involves several European universities and partner companies, is based on HELIPLAT HALE-UAV. The main objectives of the three-year project, from the aeronautical point of view, were: 1) Design an automatic HALE-UAV able to stay up for very long periods of time (about 6-9 months) thanks to a solar cell and fuel cell system. 2) To obtain a complete understanding of the feasibility of a short-term aerodynamic HALA concept, in particular for Reynolds low-profile profiles and propellers. 3) To design the entire advanced composite wing (about 75 m long), the payload housing, the arms and the tail structures. 4) To verify the production costs of each platform. 5) Produce a 1: 3 scale technological demonstrator and perform static tests on it. 6) Evaluate the security and regulatory aspects of the platform. Mechanical equipment was designed and manufactured to perform a shear–bending–torsion test on the complete scaled-size prototype and verify the theoretical behaviour. A steel supporting structure was defined and manufactured by POLITO in order to sustain the scaled-prototype. The tree-beam systems and the hydraulic jack are shown in Fig. 30. Figure SEQ Figure \* ARABIC 30 Testing System The structural design and assembling of the scaled-pro-totype of the HELIPLAT UAV is presented. The manufac-turing of the main tubular spars and metal fitting was performed by "CASA" while the advanced spar joints, specif-ically designed for the project, were developed and manu-factured by POLITO. Another important contribution was made by the Politecnico di Torino (DIASP) and its subcontractor with the assembling of several parts of the prototype and the manufacturing of all the different parts that complete the aircraft, as shown in the figures. The manufacturing of the testing facility and the execution of test activities were also made by the authors, including the positioning of the prototype and all the sensors and equipment necessary for the test. A subsequent torsion test on the main wing spar was also performed in order to completely validate the spar itself and the special advanced joint configuration. The final failure test was carried out to evaluate the margin from the ultimate design load condi-tion, and satisfactory behaviour was obtained. [12] Aeroelastic behavior of a solar-powered high-altitude long-endurance unmanned air-vehicle (HALE-UAV) A preliminary dynamic and aeroelastic investigation is performed in a simplified equivalent configuration derived from the HELIPLAT HAVE-UAV (high-altitude very-long endurance unmanned air vehicle) wing structure. The presence of slender wing structures introduces in the structural design a certain degree of non-linearity, forcing the designer to deal with specific phenomena, one of which is described. A simple dynamic system is used to point out the important effect due to the inclusion of the deformed equilibrium configuration into the dynamic behaviour of the system. A modified wing structure configuration is used to point out this effect in the flutter behaviour of slender structures. The preliminary design of the HELIPLAT HAVE-UAV wing structure, included in the HELINET research project, has indicated very low weight and high aspect ratio requirements. The presence of such types of slender structures introduces in the structural design a certain degree of non-linearity, forcing the designer to deal with specific phenomena. A typical aspect ratio of about 30 is obtained after the preliminary design, with a gross weight of about 700–1000kg. As this new kind of aircraft copes with two contrasting requirements, low weight and high aspect ratio, an improved study is needed. Considering that a thickness–chord ratio of about 11 per cent is necessary for aerodynamic reasons, it is possible to imagine that the ratio between the main structure transversal dimension to wing span could be in the order of 10-2. A linear classical beam theory (Euler–Bernoulli) is representative enough for structures that demonstrate deflection lower than transversal dimensions, while a non-linear beam model is necessary for structures that exhibit deflections of the same order of transversal dimensions: the elastic beam model has to be considered with deflection of the order of the length. The dynamic effect of a non-linear structural configuration is pointed out in a very simple dynamical system where the relative importance of the considered stiffness changes the system behaviour. The aeroelastic behaviour of high aspect ratio wings has been performed including this non-linear effect. A simplified wing has been considered and derived from the scaled prototype design of the HELIPLAT project. The relative importance of the different stiffnesses of the wing has been pointed out when a very simple case is analysed. Different behaviours are determined for different relative stiffness levels. Noticeable from a design point of view is the fact that the chord-wise stiffness seems to play a central role in the dynamic and flutter behaviour of a specific high aspect ratio wing when the equilibrium configuration is quite different from the undeformed condition. Some preliminary design indications are assumed to be useful for the full-scale HELIPLAT design. [13] Conclusion With this thesis work, various aspects concerning UAV have been dealt. Over the years there has been an increase in articles witnessed by Google Scholar concerning solar UAV. With the Solar UAV there is the possibility of use for border Security Surveillance, immigration control, early forest fire detection. The Solar UAV can be seen as an “pseudo satellite”, with advantage of being much cheaper, close to the ground (more detailed land vision) and more flexible than a real satellite. For an increase the performance they are used composite materials (for the weight) and the propulsion system is optimized thanks to solar cells at Gallium-Arsenide with its conversion efficiency of 28%, Li-Po batteries for their high ratio of 200 ÷ 250 Wh/kg and an electric motor with a power/weight ratio of about 200 W/kg. It is hoped that in the near future, with the conversion efficiency of solar cells, the power density of the energy storage system and the efficiency of the propulsion system, advanced aircraft design technology, development and application of new technological materials, the true meaning of the permanent flight will be able to achieve. Bibliography [1] “Conceptual Multidisciplinary Design Optimization (MDO) of Solar Powered UAV”, Avi Ayele, Ohad Gur, ResearchGate 290566271 [2] “The Development Status and Key Technologies of Solar Powered Unmanned Air Vehicle”, Li Sai, Zhou Wei, Wang Xueren , IOP Conference Series: Materials Science and Engineering 187 012011 [3] “Solar Power for Small UAVs, AltaDevices”, PB-UAV-1217-001-EN [4]“Design of a high altitude long endurance flying­-wing solar­powered unmanned air vehicle”, Alsahlani, A, Johnston, LJ and Atcliffe, PA , Progress in Flight Physics 9 (2017) 3-24 [5] “Conceptual design of a high-endurance hybrid electric unmanned aerial vehicle”, Martin Jaeger, Desmond Adair, Materials Today: Proceedings 4 (2017) 4458–4468 [6] “Performance analysis of solar powered Unmanned Aerial Vehicle”, Karthik Reddy B.S. , Aneesh Poondla, Renewable Energy 104 (2017) 20-29 [7] “Space-Based Solar Power a technical, economic, and operational assessment”, Jeffrey L. Caton, Title 17, United States Code, Sections 101 and 105 [8] “Mission Analysis of Solar Powered Aircraft”, David W. Hall, David A. Watson, Robert P. Tuttle, Stanley A. Hall , NASA Contractor Report 172583 [9] “Optimizing Electric Propulsion Systems for Unmanned Aerial Vehicle”, Ohad Gurand, Aviv Rosen, JOURNAL OF AIRCRAFT Vol. 46, No. 4, July–August 2009 [10] “Design of a high-altitude long-endurance solar-powered unmanned air vehicle for multi-payload and operations”, G. Romeo, G. Frulla, E. Cestino , Proc. IMechE Vol. 221 Part G: J. Aerospace Engineering [11] “Design of solar high altitude long endurance aircraft for multi payload & operations”, Enrico Cestino, Aerospace Science and Technology 10 (2006) 541–550 [12] “Manufacturing and Design of Lightweight Composite Airplane Structures Part II – Design”, Prof. Eng. Giulio Romeo - Dr. Eng. Giacomo Frulla, UAV-NET Meeting N.2, Capua(Italy), 11-12 February 2002 [13] “Aeroelastic behaviour of a solar-powered high-altitude long endurance unmanned air vehicle (HALE-UAV) slender wing”, G. Frulla, Proc. Instn MechE Vol. 218 Part G: J. Aerospace Engineering Acknowledgements I finally reached the end of my bachelor degree. First of all, I would like to thank my thesis supervisor, prof. Gianfranco Rizzo, in particular for his availability and his way of involving me in the thesis work. I also would like to thank my family, without them everything would not have begun. During my course of study I’ve met a lot of new friends and collegues, which helped and motivated me also during difficult moments. Some of them are Leucio, Pietro, Giovanni, Antonio, Giuseppe, Mario and the friends of the sailing team like Alberto, Andrea, Mario, Alexander, Francesco and Flavio which allowed me to grow a lot. 48