LASER AIAA PAPER (1)
- 1. 1 American Institute of Aeronautics and Astronautics L.A.S.E.R. – LightAircraft Solar Extended Range Rohan S. Sharma1* , Austin Gerber┬ , Christopher Giuffre╩ , Hao Guo* , Matthew Gustafson┬ , Adam Harper2* , Yen- Chen Liu╩ , Jan Lopez╩ , Matthew Poetting╩ and Kyle Suther╩ Iowa State University, Aerospace Engineering Department, Ames, IA, 50014 Parker Oltrogge┬ and Paul Uhing3* Iowa State University, Electrical Engineering Department, Ames, IA, 50014 With the aviation sector becoming a notable contributor to atmospheric pollution, the industry is experiencing a quantum-shift in the technology of air travel. A radical re-think is occurring in not only how aircrafts are designed but also what materials are used to construct them. L.A.S.E.R. (Light Aircraft Solar Extended Range) is a team of multi- disciplinary undergraduate students focused on designing, building, and testing a UAV (Unmanned aerial vehicle) which employs non-fossil fuel sources of energy to accomplish various missions. The goals of this research are to claim the record for the farthest distance travelled in a straight line by a solar-powered UAV (Unmanned Aerial Vehicle) (F5-SOL Category, FAI), to design and fly aircraft which use both hydrogen fuel cells and solar technology as a means to power onboard propulsion and electronic systems, and to continue to design and test eco-friendly UAVs that can accomplish a variety of missions. The team is currently constructing L.A.S.E.R. 5, a fully composite sailplane design, and testing circuit designs which safely charge the onboard battery using two power source systems, hydrogen fuel cells and a solar array. This report will heavily focus on the design process of this aircraft and the design of the various circuit boards onboard. This project advances the state-of-the-science and demonstrates the practicality and feasibility of renewable sources of energy for aircraft to address varied missions. Nomenclature 𝑉 –Velocity of the aircraft with respect to the ground frame of reference 𝑔 –Acceleration due to gravity 𝑆𝐹𝐶 –Specific Fuel Consumption 𝐶𝐿 –Coefficient of lift 𝐶 𝐷 –Coefficient of drag 𝑊𝑖𝑛𝑖𝑡𝑖𝑎𝑙 –Initial weight of the aircraft 𝑊𝑓𝑖𝑛𝑎𝑙 –Final weight of the aircraft 𝑉𝐻 –Horizontal tail volume coefficient 𝑆 𝐻 – Horizontal tail area 𝐿 𝐻 – Distance from the horizontal tail’s aerodynamic center to the aircraft’s center of gravity 𝑆 𝑊 – Main Wing area 𝑐̅ – Mean aerodynamic chord 𝑉𝑉 – Vertical tail volume coefficient 𝑆 𝑉 – Vertical tail Area 1 srohan@iastate.edu, Team Lead, Aerodynamics and Configuration Lead 2 Structures Design and Analysis Lead 3 Electronics and Electrical Systems Lead ┬ Freshman ╩ Junior * Senior
- 2. 2 American Institute of Aeronautics and Astronautics 𝐿 𝑉 – Distance from the vertical tail’s aerodynamic center to the aircraft’s center of gravity 𝑏 – Main wing span 𝑃 – Power 𝐼 – Current 𝑉 – Voltage I. Introduction The Unmanned Aerial Vehicle discussed in this paper is referred to as L.A.S.E.R. The team consists of multi- disciplinary undergraduate students from Iowa State University. The primary goal of the project is to design, build and fly solar-powered remote-controlled UAVs which have the capability of long range flight. The current short-term goal for L.A.S.E.R. is to break the world record for the farthest distance travelled by a solar-powered UAV in the F5- SOL Category under the FAI. In order to be operate the F5-SOL Category, certain regulations and requirements must be met. These regulations and requirements will be discussed in the section titled Conceptual Design Optimization. Another equally important goal of L.A.S.E.R. is to design, construct, and test an aircraft made from composite materials while also housing a solar array charging system as well as a hydrogen fuel-cell charging system. Future variants of L.A.S.E.R. will address various mission requirements that such as tradeoffs between speed, payload, range, VSTOL design. L.A.S.E.R missions could even include operational support such as reconnaissance or search-and- rescue. This is very important because it shows the research’s scope for mission versatility. The project currently has two different areas of research and development which are occurring simultaneously. The area is focused on the aerospace engineering side of the project, which primarily focuses on improving the aerodynamics of the plane and reducing the weight of the aircraft without compensating structural integrity. The other research area involves optimizing the electronic components. It involves designing, building and testing circuit boards which will manage the power supply from the solar-panels to the batteries and ensure the most effective use of power onboard. The two research areas work independently but regularly meet to ensure the aircraft abides by all the requirements. This project started in 2011, and since the group has designed, built, and flown many aircraft with different aerodynamic characteristics to try and find the most effective characteristics for long range–capable aircraft. . The original aircraft was built based on a very simple design. It is currently used as a test bed aircraft for eventual solar flights. It was engineered to be very stable and easy to modify, creating the ideal test aircraft to measure the effectiveness of the solar cells and the hydrogen fuel-cell battery protection system. The second iteration of L.A.S.E.R was designed based on a very similar structure as the first iteration but with added wingtip dihedrals to test the effectiveness of reducing the aerodynamic induced drag. The results from this aircraft played a great part into designing the current iteration of the aircraft. The fourth iteration of L.A.S.E.R. included major design change which were made to increase the range of the aircraft. For this iteration, the most striking change was the addition of a canard. The decision to add a canard was an attempt to extend not only the range but also the endurance and further add extra surface area to implement solar panels onto the aircraft. This iteration taught the team that a rear horizontal stabilizer was not necessary for the aircraft. There were also several changes to body of the aircraft in an attempt to reduce the skin drag. The most recent iteration of the aircraft makes improvements upon the previous iterations of L.A.S.E.R. Gone is the canard from past, and in its place, the wing has been increased to almost 14ft. in span and the tail configuration has been changed dramatically. The other improvement to the aircraft is that it features an innovative composite airframe. The outcomes of this project have produced innovation in fields not limited to aviation and electronics. This project advances the state-of-the-science along with demonstrating the practicality and feasibility of alternative energy for world contexts. II. Conceptual Design Optimization The design of an engineering system is a complicated and elaborate process. This section outlines the methodology behind our design process. The design process employed is very similar to a global systems optimization process with constraints. The constraints which must be applied are the regulations set forth by the FAI F5-SOL Category. These regulations are as follows: 1. Electrical motor propulsion. 2. Radio controlled flight without the help of any telemetry. 3. Maximum upper surface area of 1.5 m2 .
- 3. 3 American Institute of Aeronautics and Astronautics 4. Only Solar Cells are permitted as the propulsion system power source. Dimensions – 12” x 96” These regulations have had a crucial part in the design of L.A.S.E.R. 5 due to the nonconventional needs associated with implementing renewable energy into the aircraft. To elaborate with the regulations, the second prevents the use of an autopilot system to control the path of the aircraft throughout its flight. The third regulation restricts the size of the shadowed area to 1.5 m2 or less. This area collectively includes the main wing, horizontal stabilizer and the fuselage. The flowchart depicted below shows the design process of the aircraft, illustrating how we progress and regress through the steps as the design is continuously refined moving towards an optimized solution. Various software suites are used throughout the different processes in the design procedure depicted above in Figure 4.1. In the Basic Configuration Process, a general SolidWorks4 model is generated. This model is then carried forward to be analyzed and edited through the next stages of the design procedure. Next is the Aerodynamics Analysis stage where a variety of programs are used to analyze our designs. XFLR55 and StarCCM+6 are used to determine the airfoil(s) for all lifting surfaces on the aircraft, verify data, and justify any design changes as they occur. The software employed also allows us to accurately assess the performance of the aircraft and ensure it meets our design specifications. . The majority of the Structural Design stage is done using SolidWorks. This is done to most accurately represent the structure of the wing, i.e. ribs and spars, fuselage and tail. This refined model is then analyzed in the Structural Analysis stage using ANSYS 14.57 . The overall configuration finally develops as this process is repeated numerous times until the desired state of the model is reached. III. Low-Reynold’s Number High Aspect Ratio Wing A. Airfoil Selection In order for LASER to break the world record for this type of aircraft and to achieve many future goals, an efficient airfoil design was needed. For our purposes, we needed an airfoil that could generate a high lift to drag ratio due to its direct correlation with gliding performance, via the Breuget Range Equation shown below. 𝑅𝑎𝑛𝑔𝑒 = 𝑉 𝑔 1 𝑆𝐹𝐶 𝐶 𝐿 𝐶 𝐷 ln ( 𝑊 𝑖𝑛𝑖𝑡𝑖𝑎𝑙 𝑊 𝑓𝑖𝑛𝑎𝑙 ) (Eq 3.1) Throughout the duration of the flight the aircraft will be operating with at a low velocity and angle of attack, which was used as our criteria for optimization. After searching for such designs, we settled on two options: the Selig 7075 or the Clark Y airfoils. The Seilig 7075 has a maximum thickness of 9% and camber of 2.8% at approximately 29% and 46.8% of the chord length respectively; whereas the Clark Y has a maximum thickness of 11.7% and camber of 3.4% at about 28% and 42% of the chord length respectively. The Selig 7075 was eventually chosen due to its exceptional performance under the constraints of our objective, and it was thinner than the Clark Y, thus smaller potential weight and drag. This airfoil has a substantially low Reynold’s number performance indicating higher laminar flows over the wing. This is important for our goals as it improves the efficient use of our fuel source by eliminating some of the viscosity or turbulent flows the aircraft may experience, thus lowering the amount of energy needed to sustain flight. It also exhibited the necessary high lift to drag ratio, in which a ratio of 88.812 was calculated for an angle of attack of two degrees and a velocity of 35.5 meters per second. With a maximum lift to drag ratio occurring at a low angle of attack, the glide ratio is much higher and the aircraft no longer requires higher angles of attack to sustain an effective duty cycle flight. Furthermore, our analysis of the Selig 7075 airfoil can be seen below: 4 Computer Aided Design (CAD) Software. 5 An analysis tool for airfoils, wings and aircraft operating at low Reynold’s Numbers. 6 (CCM standing for computational continuum mechanics) A high level computational fluid dynamic software. 7 Finite element analysis software for structural physics that simulates static, dynamic, and thermal problems.
- 4. 4 American Institute of Aeronautics and Astronautics B. Planform Modulation and Selection Many requirements and preferences needed to be considered while designing the planform and dimensions of the wing. The major constraints that limited the design space are highlighted in the previous section titled Conceptual Design Optimization. Successful full scale sailplanes feature high aspect ratio tapered wings, so such a design was the starting point in the design process. As the root chord length must be at least as long as the width of the solar array, a root chord length of 13 inches was decided so as to enable the aspect ratio to be as high as possible while allowing some room for error To make the wing as efficient as possible, a planform with an elliptical lift distribution is desired. However, such a planform was not a good option for us, as they can be exceedingly difficult to construct and some of the benefits are lost at the low Reynolds numbers. Therefore, a planform that approximates an elliptical lift distribution was the best option. There are a numerous amount of designs that do this, but after analyzing the options, it was determined that a planform with a rectangular center panel and tapered tips would be the best option. This planform is easy to build and the departure from an elliptical lift distribution is small. It also helps to mitigate the chances of tip stalling because the center portion of the wing is slightly overloaded and will tend to stall first. The most efficient taper ratio for this wing is approximately 0.5, but having that much taper would lead to a dangerously small tip chord. A very small tip chord is undesirable because it would result in lower Reynolds numbers that decrease performance and encourage stalling. At the same time, a small tip has the benefit of reducing wingtip vortices, which decreases induced drag. In the end, the optimum taper ratio was determined to be 0.7, making the tip chord 9 inches. By tapering both the leading and trailing edges the correct amount, the quarter chord line goes straight across the wing, which has both structural and manufacturing advantages. To determine the span, we had to consider the area of the fuselage and horizontal stabilizer to ensure the area restriction was not exceeded. A span of 14 feet allowed the wing to have a high aspect ratio of approximately 14 while keeping the total shadowed area slightly below 1.5𝑚2 . We considered adding either winglets or Hoerner tips to the wing to decrease induced drag and increase climbing performance, among other things. After doing more analysis, we decided that the added performance would not justify the increase in weight. Wingtip devices have been proven to be beneficial, but depending on the design, only in a small region of the flight envelope. We did not have the time or resources to Figure 3a.1: Coefficient of Lift vs angle of attack and Lift over Drag ratio vs angle of attack plots for the Selig 7075 (9% thickness. Generated using XFLR5
- 5. 5 American Institute of Aeronautics and Astronautics research an optimal design, so blindly making one would have likely resulted in minimal performance gains with an increase in weight and cost. Given the constraints of the wing, the planform described above is the optimal design for a range optimized small scale motor glider. Induced drag has been reduced by tapering the wing, which is vital because at the flight condition of maximum lift to drag ratio (and maximum glide ratio), induced drag is the largest component of drag. Importantly, the whole solar array can fit on the wing. Structurally, the design does not present any problems and meshes well with the rib and spar design while being a relatively simple wing to manufacture. C. Structural Design and Material Selection For an aircraft, the wing structure can account for half of the total structural weight. To meet this challenge, we designed a composite multi-rib wing to minimize the weight. Implementing this rigid wing box system, added structural integrity and minimized weight, but there are design tradeoffs. Due to the majority of the wing box being made of carbon fiber, the cost of this design is considerably more than the foam equivalent of previous design iterations. The foam planes allowed for easy repair when failure occurred but despite these tradeoffs, the implementation of the rigid wing system allowed for better performance of L.A.S.E.R. 05 and so this design was pursued further. The wing system can be broken up into three components: spar, ribs, and skin of the wing. The first component discussed are the ribs. The ribs are constructed from a 1.7 mm carbon fiber plate. The carbon fiber ribs allows us to stiffen our wing to resist global buckling and provide adequate mounting for the skin. The ribs are designed to resist a 10g inertial load only deforming a few millimeters. The next component discussed is the spar system. The spar system consists of a forward and an aft spar. The forward spar is a 0.5 in cylindrical carbon fiber rod and the aft spar is a 0.25 in square extruded carbon fiber rod. The spar system functions as the main load bearing structure for our wing, absorbing and distributing the span- wise bending moments. During finite element analysis through ANSYS, the spars were found to only deform a few millimeters during normal flight conditions, concluding that these rods produce more than enough structural rigidity for the purpose of the flight envelope. The final section of the wing system is the skin system. The skin is a layer of MonoKote with balsa wood stiffeners installed along each of the ribs as well as at the leading and trailing edge. This skin systems allows us to minimize the weight penalty added to the wing by the skin itself, thus optimizing the flight performance and range. The wing box is constructed out of thin carbon fiber parts which help to decrease the weight penalty. The wing box provides the needed buckling and torsional stability to maintain structural integrity of the aircraft. Figure 3d.1 shown above details the geometry of the wing box structure. IV. Fuselage Design A. Design and Dimensional Optimization The overall design of the fuselage is very similar to fuselages on full scale sailplanes. The fuselage is very long and streamlined, with a slight bulbous section in the front. This larger section is required for the cockpit on full scale planes, but for us the primary purpose of the fuselage is to house necessary electronic systems. To perform solar powered flights the fuselage also needs to house the circuit boards that contain the power supply management systems. The boards were designed with the fuselage size in mind, but we still had to take into consideration where they will fit when designing a new fuselage. The design of the fuselage is such so that the majority of the components will be centered around the quarter chord of the wing, helping us to maintain a constant center of gravity. By doing this we improve upon the stability of the aircraft, which in turn leads to an airplane that is easier to fly long distances and one that causes less pilot fatigue. The bulbous and streamlined design has aerodynamic properties that are very beneficial for a range-optimized sailplane. The smooth curves mitigate flow separation and help keep drag from the fuselage to a minimum. The high wing sits flush with the top of the fuselage, helping to keep the flow as smooth as possible over both the fuselage and wing. The general shape has a drag coefficient of approximately 0.2, which is lower than other
- 6. 6 American Institute of Aeronautics and Astronautics plausible fuselage designs by a considerable margin. The geometry also provides a small amount of lift while in flight, but this is negligible when compared to the lift generated from the wing. The long length of the fuselage allows the area of tail stabilizers to be reduced. This increases performance because there is a net decrease in drag when extending the fuselage and decreasing tail area. B. Structural Design and Material Selection The fuselage of L.A.S.E.R. 05 consists of a foam core sandwich panel design. In this design, we have a 1/4 in BlueCor™ Foam core hand-wrapped with a bi-axial carbon fiber sock with fibers oriented at a 0 and 90 degree orientation. This design gives us the much needed structural rigidity to resist harsh landings expected from landing on the aircraft’s fuselage, a feature previous iterations could not claim. During ANSYS analysis, we tested a worst case scenario by applying a 10g load to the end of the fuselage and noted that it only had a displacement of 3.48 mm. V. Tail Design A. Tail Design Selection and Dimensional Optimization In the case of a glider, the design of the tail is crucial to achieving the mission specifications and plays a major role in factors such as glide ratio, total drag, total lift, etc. all of which affect the aircraft’s range capability. There is a plethora of different tail configurations, each designed for optimal operation in different flight situations. Using the data gathered in Table 5.1, the team was able to assess which tail configuration would be most suitable for the aircraft. Although the twin tail design (H-Tail) features favorable yaw authority, it was immediately ruled out as its larger structural weight and drag would limit the overall range of the aircraft. The V-tail, with fewer surfaces than a conventional three-aerofoil Tail, is lighter, and ideally has a smaller wetted surface area producing less overall drag. However, NACA (National Advisory Committee for Aeronautics) studies have indicated that V-tail surfaces must be larger than design projections would suggest, such that the total wetted area is roughly
- 7. 7 American Institute of Aeronautics and Astronautics constant. Essentially, overall drag reduction is due to the reduced interference drag not parasitic drag. While the V- tail configuration does produce an adverse-roll moment this can be minimized by increasing the length of the fuselage (increases the moment arm) and reducing the span of the V-tail making the adverse roll negligible. Although the Inverted V-tail produces pro-verse yaw moments, both the V-tail and Inverted V-tail were ruled out due to their complex control system and stress placed on the aft portion of the fuselage. While the T-tail yields a better glide ratio than most other tail configurations and is featured on some popular gliders in the market, from a manufacturing standpoint, the joint between the vertical stabilizer and horizontal stabilizer would have to support any moments generated by the horizontal stabilizer, requiring a heavier structure . Additionally the T-tail configuration can succumb to flutter forces when subjected to a turbulent flow. For these reasons, the T-tail design was ruled out. The conventional tail design was chosen as it has been tested and verified multitude of times and can be optimized and designed with relative ease. Through careful selection of materials, the problem of a high structural weight of a conventional tail can be avoided. In design, sizing the tail is a very subjective process, one which involves conflicting requirements of center of gravity, range, stability, control, and desired aircraft handling characteristics. The dimensionalization of the horizontal stabilizer and vertical stabilizer in a conventional tail configuration is imperative to the lateral stability, yaw (normal axis) stability, and damping of the aircraft. The size of these lifting surfaces determines how effectively the aircraft can dampen itself into an equilibrium position when the system is perturbed. When dimensioning the two surfaces it is imperative to consider two very important design checks, the Horizontal Stabilizer Sizing Criteria (Volume) – Pitch Stability (VH), and the Vertical Stabilizer Sizing Criteria (Volume) – Yaw Damping and Rudder Power (VV). Both these tail volume coefficients relate the surface area, the displacement of this area from the center of gravity, main wing net area, mean aerodynamic main wing chord, and main wing span. Aircraft with the same volume coefficients tend to have similar static stability characteristics. For aileron thermal duration gliders, values of Horizontal Stabilizer Sizing Criteria range from 0.3 – 0.6 and values of Vertical Stabilizer Sizing Criteria range from 0.015 – 0.025. By abiding by these ranges, this design process can be refined by finding the optimal dimensions of the tail surfaces. The goal of this optimization procedure is to find the minimum dimensions required to operate slightly above the specified stability criteria. The formulas involved in this optimization process are given below. 𝑉𝐻 = 𝑆 𝐻 𝐿 𝐻 𝑆 𝑊 𝑐̅ (Eq 5.1) 𝑉𝑉 = 𝑆 𝑉 𝐿 𝑉 𝑆 𝑊 𝑏 (Eq 5.2) The team developed an optimization toolset using a Microsoft Excel Workbook. By continuously varying the length of the fuselage and dimensions of the tails, different volume coefficients are calculated until the optimum values are reached. These optimum values are determined keeping structural weight, control authority, and drag penalties in consideration. Screenshots of this optimization tool with the finalized tail dimensions are shown in figures.
- 8. 8 American Institute of Aeronautics and Astronautics Figure 5.2: Design Checks B. Structural Design and Material Selection Similar to the fuselage design methodology, the tail of L.A.S.E.R. 05 consisted of a foam core sandwich panel design. The tail also comprised of a BlueCor™ Foam core hand-wrapped with a single layer of bi-axial carbon fiber with the fibers oriented such that the fibers lay at a 0 and 90 degree orientation. This design generates enough structural rigidity to resist the gust loads expected within the designed flight envelope. Due to the strength of the carbon fiber covering, we expect the tail to sustain little to no damage when undergoing a harsh belly landing. VI. CFD Analysis A flow simulation was done for LASER 5 to validate the design using Solidworks Flow Simulation. The following plots give a visual representation of the pressure distribution on and around the aircraft. The included figures show the pressure distribution for flow conditions of 25 mph free stream velocity at sea level with the plane flying at a 7 degree angle of attack. Figure 6.1: Pressure Distribution about the y-x plane of the aircraft Figure 6.2: Pressure Distribution on the wing surface, fuselage, and tail section.
- 9. 9 American Institute of Aeronautics and Astronautics The plots show the expected distribution of high pressure below the wing and on the leading edge. The figures indicate that the most lift is generated from the center rectangular portion of the wing, but the reduced lift from the tapered portions is not dramatic. The fuselage does seem to interfere with the flow over the part of the wing that is attached to it, but this is unavoidable and the overall effect is minimal. The pressure distribution on the fuselage is relatively constant (except at the wing), which suggests a proper design for a streamlined sailplane fuselage. VII. Onboard Electronics A. Electronics Necessary for Basic Flight Operation Choosing the correct RC (Remote-Control) components is a vital step in the design and manufacturing process. A high efficiency 800 kV brushless motor with a maximum RPM of 7250 was chosen for the electric motor and a 60 A ESC was determined to be necessary for that particular motor. The battery used is a three cell, 11.1 V, lithium polymer with a power output of 3300 mAh, and continuous discharge rate of 25 C. A total of four servos are used to actuate the control surfaces. The optimum propeller for this type of plane is a 14 X 6 folding propeller. Using a folding propeller is a huge advantage, as the two blades will fold on to the fuselage during the large portions of the flight in which the plane is gliding. As the figures indicate, the drag penalty from a stationary propeller in gliding flight is dramatic. While the specific increase in drag caused can vary depending on fuselage geometry and the propeller used, a folding propeller will most certainly provide a significant increase in performance during gliding flight. B. Solar Power Converter The main goal of the non-RC electronics is to control the flow of power from the solar array to the battery and the rest of the system. There are two requirements for this system. The first requirement is that the power converter keeps the solar array at its maximum power point. This maximum point occurs at about 15.4V for the PowerFilm solar array. While this point will yield the maximum power transfer from the solar array to the system it is not acceptable the charge our 11.1V 3 cell Lithium Polymer battery. Unlike other types of battery chemistry that can take a large amount of abuse and only shorten the life of the battery, Lithium Polymer batteries when forced outside of their designed operating parameters they tend to get very hot and eventually start on fire. Therefore the solar power converter circuit must convert the 15.4V input from the solar array to the battery voltage. It would have been most efficient in terms of both power and time to use a commercial maximum power point tracker (MPPT). Unfortunately there were no MPPTs commercially available that meet our power, voltage, size, and cost requirements. Therefore we had to switch to a simpler system to accomplish our goals within our requirements. We decided to use boost-buck controller to meet the design requirements. By choosing a LM5118 boost-buck DC-DC converter from National Semiconductor, we can accomplish most of our goals. This boost-buck can create our desired output voltage of 11.1V from a range of inputs voltages, 1.23V to 75V in part because it can switch operating modes and duty cycle on the fly. The circuit was optimized by using National Semiconductor’s WEBENCH® application for our voltage requirements. While this circuit will not be as efficient as an MPPT it should still be about 90% efficient. C. Battery Protection System Although we have designed a solar power converter, we still need to ensure that the battery is not being overly stressed by overcharging, discharging the battery too quickly and excessive heat generation. Specifications of the battery are shown in Table 6.1. Figure 7a.1: Coefficient of drag values for various fuselage implementations.
- 10. 10 American Institute of Aeronautics and Astronautics Based off the battery parameters we determined how strong of a battery protection protocol required. At the maximum power point of our solar array, 15.4V and 1400mA. Then assuming our best converter efficiency of 90% and our worst case battery voltage of 9V the worst case charging current can be calculated using equations 6.1-6.3. 𝑃𝑖𝑛 = 1400𝑚𝐴 ∗ 15.4𝑉 = 21.56𝑊 (Eq 6.1) 𝑃𝑜𝑢𝑡 = 0.9 ∗ 𝑃𝑖𝑛 = 19.40𝑊 (Eq 6.2) 𝐼𝑜𝑢𝑡(max) = 𝑃 𝑜𝑢𝑡 𝑉 𝑚𝑖𝑛 = 2156𝑚𝐴 (Eq 6.3) Based off the maximum current output of the solar power converter, the input current should not reach above the maximum charging rate of the battery. Due to this we need not worry about limiting the input current and because we are using a battery commonly used for RC aircraft the system shouldn’t draw too much current from the battery. Therefore we mainly want monitor the battery conditions. The battery conditions will be monitored by a gas gauge circuit. We use a BQ2084-v143, which allows us to monitor battery conditions including battery charge state, remaining capacity, discharge rates, battery health, voltage, current, and temperature. All of this information will be sent to the telemetry system and then relayed to the base station. This will allow the pilot and ground team to measure the current battery conditions. Many of these battery parameters are measured by BQ29312A, which is the analog front end sister component of the gas gauge. This circuitry converts the analog values measured from the battery to digital values. It can also isolate the battery in an event of any battery faults. D. Flight Data and Telemetry The flight data and telemetry system was originally designed as a custom component that was designed to operate at 3.3V. This significantly reduces the amount of power consumed over the traditional 5V logic used by many RC aircraft electronics. It also included sensors to measure pitch, roll, and the altitude of the aircraft. After building the flight data and telemetry board and debugging for about 3 months with no progress on implementing the design the decision was made to look for alternative systems to accomplish a similar task with a reduced implementation time. This led us to consider an Ardupilot. While the class of RC aircraft we are building doesn’t allow for automated control of the aircraft the Ardupilot has many advantages over our previous Flight data implementation. Firstly the Ardupilot is a proven piece of hardware so any issues with implementation can be traced back to the system software. Secondly, the Ardupilot already has all of the sensors that were implemented on our previous design and some of these are improved compared to our previous design. The Ardupilot will also allow for implementation of an I²C data bus to retrieve the battery data from the gas gauge circuit and has enough analog inputs for measuring the solar array voltage and current as is already implemented on the solar array board. Using an Ardupilot will not require an extensive redesign of most of our systems. It will only require a small change of a resistor connected to the low voltage supply of the flight data and telemetry systems as the Ardupilot requires a 5V input. This means that the system will draw more power than the original design but it will be easier to implement. If testing determines that this difference is too great we will need to redesign a new flight data board. This change most likely will not make a significant difference to power draw. The Ardupilot also easily uses our Xbee Pro 900MHz radio to send data back to the base station. The final advantage of the Ardupilot is the ease of which a GPS unit can be integrated with the system. This GPS will increase the power draw of the system and is required to record our flight path. Overall the Ardupilot is a good option Figure 7c.1: Battery Parameters
- 11. 11 American Institute of Aeronautics and Astronautics for our current needs. If it is determined that the aircraft requires a custom flight data and telemetry onboard, the Ardupilot will give us adequate data to compare with any custom system designed to replace it. VIII. Discussion Current plans for the L.A.S.E.R project in its immediate future consist of completing the final touches in the latest iteration of the design, L.A.S.E.R 5. Once the aircraft is cleared to fly, flight tests will begin and any necessary troubleshooting will be conducted. Final preparations are being done on solar flight testing as well. The team is looking to soon attempt setting the world record for the farthest distance travelled by a solar-powered UAV in the F5-SOL Category under the FAI. The team will continue to work on improvements to the design, further expanding the aircraft’s capabilities. The team hopes to expand LASER’s functionality as an “eco-friendly” UAV through the implementation of hydrogen fuel cell technology. This additional fuel source could lead to more efficient flight, and providing more flexibility for longer ranged missions. Furthermore, unconducive weather conditions such as overcast can limit the use of a solar powered vehicle, but with an additional power source of hydrogen fuel cells this limitation could be bypassed while still remaining “eco-friendly”. Other than working on advancements in propulsion technology, the team is very interested in improving the material composition of our aircraft. L.A.S.E.R 5 is the first aircraft, in the team’s fleet, whose primary structures are composite materials, such as carbon fiber. Eventually the goal is to construct an airplane out of lighter and stronger materials, which combined with our advanced power sources, can help achieve a vast variety of mission objectives. Finally, we aim to create a semi-autonomous UAV with an implemented advanced control system that is user friendly, and the inclusion of an onboard camera. Overall, a vast array of avenues can be explored with this project in both the near future and further down the road. IX. Conclusion Over the past three and a half years, L.A.S.E.R. has been working hard towards the goal of setting the FAI range record in its category. During this time, we have focused on many different areas of research that have allowed us to soon attempt the record. The L.A.S.E.R. aerodynamics team has refined the fuselage and wings to create a high aspect ratio, long range and highly efficient aircraft. The structures group has succeed in creating a minimal weight, high strength fuselage out of composite materials and a foam core. This fuselage will not only be the fuselage to attempt the world record flight, but will also serve as a platform for years to come. The electronics team is in the final phase of testing and is looking to the future to create the next generation of electronics for L.A.S.E.R. With these three groups working together with the end goal in mind, L.A.S.E.R. is rapidly advancing towards the range record X. Acknowledgments The authors of this paper thank Iowa State University, Aerospace Engineering Department, for providing invaluable resources to enable the design and production of L.A.S.E.R. Specifically, the Make to Innovate (M:2:I) program initiated by the department. The authors would like to thank Matthew Nelson and Professor Ambar Mitra for their help. L.A.S.E.R would like to acknowledge the several companies who have supported this project. These companies include The Boeing Company, PowerFilm, Alliance Pipeline, Advanced Circuitry, and Rockwell Collins. XI. References Simons, Martin, Model Aircraft Aerodynamics, Nexus Special Interests, Swanley, England, UK, 1999. Print Lennon, Andy, R/C Model Aircraft Design, Air Age Inc., Ridgefield, Connecticut, 1996. Print Anderson, John David. Aircraft Performance and Design. Boston, MA: WCB/McGraw-Hill, 1999. Print. Anderson, John David. Fundamentals of Aerodynamics. 5th ed. New York: McGraw-Hill, 1984. Print. Nelson, Robert C. Flight Stability and Automatic Control. New York: McGraw-Hill, 1989. Print. Hurt, Hugh H. Aerodynamics for Naval Aviators. Washington, D.C.: Office of the Chief of Naval Operations, Aviation Training Division;, 1965. Print. Megson, T. H. G. Aircraft Structures for Engineering Students. London: Edward Arnold, 1972. Print.