University of California, Davis

Transcription

1 University of California, Davis AMAT 2013 Advanced Modeling Aeronautics Team Advanced Class, Team 215 Team Captains: Ilya Anishchenko Alex Beckerman Logan Halstrom Team Members: Michael Wachenschwanz (DAS lead) Gene Ang Chris Lorenzen Joshua Barram Kelley Lundquist Max Bern Arlene Macias James Dionisopoulos Robyn Murray Louis Edelman Nohtal Partansky Robert Edwards Jason Petersen Hashmatullah Hasseeb Patricia Revolinsky S. Sheida Hosseini Adam Simko Steven Hung Stefan Turkowski Sara Langberg Faculty Advisors: Jean-Jacques Chattot, PhD Stephen K. Robinson, PhD

2 2 Advanced Modeling Aeronautics Team 2013

3 Table of Contents Statement of Compliance... 2 List of Figures and Tables.. 4 Nomenclature Introduction Design Process Design Considerations Aircraft Design Avionics Design Previous Design Experiments and Iterations Prototyping and Material Testing Engine Performance and Propeller Sizing Model Testing Final Selection Aerodynamics and Performance Analysis Aircraft Sizing Selection of Main Wing Selection of Overall Aircraft Configuration Winglet Design Aircraft Stability and Control Longitudinal Stability and Control Lateral Stability and Control Directional Stability and Control Control Surface Sizing Aircraft Performance Payload Analysis Payload Prediction Payload Drop Component Structure and Manufacturing Main Wing Empennage Fuselage Landing Gear Weight Management Conclusion References.. 29 Appendix Technical Plans

4 4 Advanced Modeling Aeronautics Team 2013 List of Figures and Tables Figure 2.3A Mini Plane demonstrating hybrid wing construction. 9 Figure 2.3B Wing bending stress test 9 Figure 2.3C Ideal layup configuration for fuselage and landing gear Figure 2.3D Static thrust test stand with O.S. 46 AXII. 10 Figure 2.3E Static thrust performance of the O.S. 46 AXII Figure 2.3F Prototype Fuselage and tail boom.. 11 Figure 2.3G Experimental polar apparatus Figure 3.1A Non-dimensionalized geometry of the Selig 1223 airfoil.. 13 Figure 3.1B 2D and 3D Polars and Lift Slope Curves of the AMAT 2013 Main Wing Figure 3.1C Effect of Winglet Addition on 2013 AMAT design.. 15 Figure 3.2A Center of gravity and aerodynamic center for entire configuration.. 16 Figure 3.2B Aileron chord required to counteract a range of gust velocities 18 Figure 3.3A Comparison of available power to power required for equilibrium flight. 20 Figure 3.4A Effect of increasing density altitude on maximum payload cabability.. 21 Figure 3.4B Sandbag trajectory prediction for drop from cruise speed. 22 Figure 4.1A Main wing half section.. 23 Figure 4.2A - Empennage without monokote Figure 4.3A Fuselage and landing gear. 25 Figure 4.4A Machining of the airshock. 26 Figure 4.5A - Glass/Epoxy Slurry, Aluminum Fasteners, 5-axis foam mill. 27 Table 2.1A - Summary of 2013 Aero Design Advanced Class Flight Score Criteria.. 6 Table 2.1B Design matrices from AMAT design process. 7 Table 2.2A Lessons from Previous Designs.. 8 Table 2.3A Rear Landing Gear Strut Coupon Mass Comparison.. 10 Table 2.3B Fuselage Wall Coupon Mass Comparison.. 10 Table 2.4A Final Selection of Various Components. 12 Table 3.1A Zephyrus Lifting Surfaces.. 15 Table 3.1B Final Configuration. 15 Table 3.3A Zephyrus Fight Characteristics Table 4.2A Weight (lbf.) Comparison of Zephyrus to 2012 AMAT Design 23 Table 4.5A Zephyrus Component Weight Allocations. 27 Nomenclature Symbols Symbols Cont. Symbols Cont. Main wing chord c Coefficient of _ C Force F Thickness t Equivalent flat plate area f Weight W Main wing span b Density ρ Thrust T Main wing area S Newton s Constant g c Rotation per Minute RPM Main wing aspect ratio AR Angle of attack α Subscripts Center of Gravity CG Deflection angle δ Main wing m Aerodynamic Center AC Velocity V Horizontal stabilizer h Leading Edge LE Rate of climb RC Vertical stabilizer v Trailing Edge TE Radius R Fuselage f Length l Pitch p Aileron a Longitudinal distance x Power P Take off TO Lateral distance y Lift L Initial 0 Moment M Drag D Induced i

5 1.0_Introduction The Advanced Modeling Aeronautics Team is proud to present the 2013 AMAT aircraft to the SAE Aero Design West competition. AMAT participates in the Advanced Class, and its mission this year is to design and construct an aircraft with the purpose of aerial delivery of humanitarian aid. For the competition, an aid package is represented by a 3 lbf sandbag, because of its similarity in size, weight, and physical properties to a package of food or supplies. The aircraft is also required to lift a 15 lbf static weight representative of fuel reserve, since on actual humanitarian missions, the aircraft may be required to travel long distances to reach those in need. Aircraft design is also dictated by other factors including an empty weight restriction of 8 lbf and a Data Acquisition System (DAS) capability requirement of real-time altitude measurements and First Person View (FPV) telemetry transmission. To meet this challenge, AMAT has created Zephyrus, a radio-controlled, fixed wing aircraft designed by students at the University of California, Davis. Zephyrus s characteristic features are its composite material construction, its angled lifting tail, and its highly sophisticated stability augmentation system. Each of these traits work in combination with other design features to help the aircraft meet three main design criteria: light structure, high-lift capability, and precise flight control. Maneuverability of the aircraft is a new focus for SAE Aero Design, where the previous objective was pure heavy lifting, and was the justification of AMAT s previous double element wing design. This year s competition provides the interesting challenges of building a stable aircraft, optimizing its control, and strategizing drop trajectories and pilot communication. Accurate delivery of the package is important because the aid must arrive intact and undamaged as well as cause no collateral damage. The primary pilot is not allowed to receive any FPV data from the aircraft they control and must be directed to the target by a secondary pilot. This could be representative a situation where the aircraft is deployed in the field by the primary pilot away from the base at which telemetry data can be received, forcing flight by verbal radio instruction. AMAT has developed Zephyrus for ideal fulfillment of the design requirements and hopes to be a top competitor at SAE Aero Design West. 5

6 2.0 Design Process AMAT utilized past experience, research, and experimentation in its design process to progress from an assessment of the design criteria to an aircraft design idealized for this challenge. 2.1 Design considerations As might be done in the design of a full-size aircraft, AMAT divided the design process into two categories: aircraft systems and electronic systems Aircraft Design AMAT defines a successful airplane design as one that adheres to all stated design objectives, which are in this case determined from the flight scoring below in Table 2.1A: (( ) ) ( ) ( ) Table 2.1A Summary of 2013 Aero Design Advanced Class Flight Score Criteria AMAT incoperated the three components of the flight score FS into three primary design criteria: precise flight control (S 1 ), minimized aircraft weight (S 2 ), and lifting capability of maximum takeoff load W TO = 26 lbf (S 3 ). These criteria drove for weight minimization in all structures and materials and for a focus on aircraft stability, both in design and control. To make selections based on these criteria, AMAT elected to utilize decision matrices, which allowed the team to organize and evaluate the many design options in a systematic manner. This method for meeting the customer s design specifications is a common practice in industry, and its application to the prescribed rules of this competition was appropriate. A design matrix compares design options by ranking them against one another in various categories, such as weight and strength. The importance of each category relative to the design criteria is assigned a percentage, which is used to weigh design option rankings. The points for each design option are then added and the most beneficial design 6

7 determined. One example of importance percentage assignments was the team s favor toward categories that affected flight stability (such as wing surface precision) because of the definite zero FS received for a payload drop not accurate to within 50 feet. Similarly, the team predominantly favored weight reduction because an empty weight of greater than 12 pounds also receives a definite zero FS. Although many such matrices have been completed, only a select number have been shown below. Carbon Fiber and Foam Core Main Wing Design Matrix Double Element Monokote and Balsa Frame Hybrid Carbon Fiber and Foam Core Single Element Monokote and Balsa Frame Hybrid Importance (%) Weight Surface Precision Strength Manufacturing Survivability Previous Experience Score Landing Gear Design Matrix Tricycle Tail Dragger Importance (%) Weight Stability Need to Accelerate at Zero Incidence Feasibility of Shock Absorbers Survivability Potential for Propeller Damage Score Fuselage Design Matrix Flat Plate Enclosed Box Rounded Geometry Importance (%) Weight Strength Compatibility with Enclosed Payload Surface for Attachments Safety for Internal Circuitry Previous Experience Aerodynamics Score Table 2.1B Design matrices from AMAT design process AMAT s rankings in Table 2.1B determined a preliminary selection of aircraft components specialized to the design criteria. This selection was further refined in the team s design process. 7

8 2.1.2 Avionics Design Advanced Modeling Aeronautics Team 2013 The 2013 Advanced Class competition calls for number of Data Acquisition System (DAS) functions. To meet these requirements, AMAT set a goal that the DAS would be able to display the altitude of the plane in real time and log the altitude when the payload is dropped. Additional DAS goals were to record GPS and air speed data as well as provide a real time stream from a camera to assist the dynamic payload dropping. Flight score component S 1 drove the team to also incorporate an aircraft control system that would be able to assist flight with negative feedback from gyroscopes as well assist payload dropping using telemetry data. Previously, an Arudino Uno microcontroller was used to run a simple DAS. This year, to achieve the more demanding DAS capabilities, the team upgraded to an Ardupilot 2.5, which has built in gyroscopes, accelerometers, a digital compass, a GPS, and a high resolution altimeter. Considering the precision nature of this competition, an off board GPS module, the ublox LEA-6h, was chosen over the onboard Mediatek GPS for higher accuracy. AMAT also incorporated a pitot tube to be able to use airspeed data in drop calculations. 2.2 Previous Design Though the current Advanced Class mission is considerably different than that of last year, AMAT was able to use the knowledge it derived from its 2012 design to intelligently improve the 2013 Zephyrus. Before other research, the team first considered which features of the previous design should be further developed and which would not be effective in the new competition, as summarized in Table 2.2A: Table 2.2A Lessons from Previous Design Effective Previous Features Features to Improve or Replace Lifting tail Complex and heavy composite double element Tricycle landing gear Flat plate, open fuselage Winglets Heavy carbon fiber tail boom Composite materials Heavy composite empennage 8

9 2.3 Experiments and Prototyping Advanced Modeling Aeronautics Team 2013 AMAT devised experiments to test the feasibility of the designs determined in section 2.1. These tests included the construction of a prototype mini plane, a wing stress test, carbon fiber coupon sampling, and engine testing Prototyping and Material Testing The team decided to test the hybrid wing design innovation by constructing a miniature aircraft, called the mini plane, so that qualitative analysis could be performed on its practicality. The result was a flying aircraft that proved Figure 2.3A Mini Plane demonstrating manufacturing such a wing was possible and that the connections between hybrid portions were acceptably strong. AMAT also chose to stress test a full scale section of what would be the hybrid wing for strength. The team reasoned that the hybrid monokote/composite skin would provide adequate torsional resistance, so the wing would be most likely to fail first in bending. AMAT represented wing loading as point forces applied to the main spar, simulating the bending moment caused by lift force. The result was total failure at 135 lbf of applied point force, corresponding to a bending moment of ft*lbf, which is more than the spar will ever experience under the design conditions. The team Figure 2.3B Wing bending stress test considered this result a validation of the hybrid wing design. To be able to meet the weight requirements, AMAT also experimented with various composite material arrangements of foam, wood, and carbon fiber to determine the lightest and strongest custom materials to use on the aircraft. The fuselage was to be made out of a wood/composite sandwich for 9

10 rigidity, and the rear landing gear was to be made out of a foam/composite sandwich to achieve the strength of last year s gear at a fraction of the weight. The team conducted layups as detailed in Tables 2.3A and 2.3B to determine the area density of each material, and qualitatively tested each for strength. Table 2.3A Rear Landing Gear Strut Coupon Mass Comparison Layup Order m/a (g/in^2) [0 /90, 45 /-45, 0 /90, foam]s [0 /90, 45 /-45, 0 /90, 45 /-45, foam]s [0 /90, 45 /-45, 0 /90, 45 /-45, 0 /90, foam]s Table 2.3B Fuselage Wall Coupon Mass Comparison Layup Order Wood (finish) m/a (g/in^2) [0 /90, 45 /-45, wood, 0 /90 ]s 1/8 Base (Mylar) [0 /90, 45 /-45, wood, 0 /90 ]s 1/8 Balsa (Mylar) [0 /90, 45 /-45, wood, 0 /90 ]s 1/8 Balsa (peel ply) [0 /90, 45 /-45, 0 /90, wood]s 1/8 Balsa (peel ply) [0 /90, 45 /-45, 0 /90, wood]s 1/16 Balsa (Mylar) Along with composite layer structure variation, the team also designed these coupons to determine the effect of wood types and finishes of either smooth from impermeable Mylar or rough from permeable peel ply. The tests showed the rigidity differences between wood types is negligible (unlike the weight differences), with balsa being superior, and peel ply finishes allow the maximum amount of epoxy removal from the layup, and thus result in lighter materials. In both cases, AMAT selected the material with the least area density (Figure 2.3C) since it was decided that each would be sufficiently strong for their application. Figure 2.3C Ideal layup configuration for fuselage and landing gear Figure 2.3D Static thrust test stand with O.S. 46 AXII 10

11 2.3.2 Engine Performance and Propeller Sizing Advanced Modeling Aeronautics Team 2013 This year, AMAT selected the O.S. 46 AXII engine as its power plant. To efficiently size and configure its aircraft, AMAT needed an estimation of engine performance. The team derived the results in Figure 2.3E by running propellers of various diameter and pitch at the maximum safe RPM attainable by the engine and recording static thrust with a fish scale attached to a sliding platform on which the engine was mounted, seen in Figure 2.3D on the previous page. The resulting maximum thrust measurement allowed AMAT to predict takeoff equilibrium conditions and determine that a gearbox would not be necessary for sufficient thrust. Figure 2.3E Static thrust performance of the O.S. 46 AXII, yielding maximum of T 0,max =8.5 lbf Model Testing Through the use of a model fuselage, the team adequately sized the location of various components. This produced a more efficient layout that minimizes unused space while reducing fuselage material. Figure 2.3F Prototype Fuselage and tail boom. Figure 2.3G Experimental polar apparatus 11

12 The team also had great ambitions for obtaining an accurate experimental lift slope curve and drag polar, but failed to obtain the proper sensors. The planned test involved attaching a model wing segment to a car with the apparatus seen in Figure 2.3G so as to measure lift and drag at various speeds and various angles of attack. The wing mounts would put pressure on sensors in both the vertical and horizontal direction. 2.4 Final Selection With the results from the design matrices validated through testing, the team finalized the selection of the various aircraft components. The results are summarized in Table 2.4A: Table 2.4A Final Selection of Various Components Single element, hybrid wing Box fuselage Angled truss tail boom Lifting tail Tricycle landing gear Airshock front gear Composite rear gear Single engine, Un-geared AMAT selected these features to combine into a design that is competitively optimized for the current Advanced Class mission. The hybrid wing strikes an ideal balance between weight, strength, and airfoil precision. The addition of winglets improves the efficiency of the untwisted, rectangular main wing, bringing it very close to the ideal elliptic wing efficiency. The lifting tail configuration allows the standard tailplane design to have the aerodynamic efficiency of a canard on takeoff, where the tail lifts with the wing. The tail boom angle provides the added benefit of increased horizontal tail lifting capability by elevating it above the downwash of the main wing ahead of it. The tricycle landing gear and angled tail boom allow for a large degree of rotation on takeoff, which means the plane can accelerate at low incidence to minimize induced drag. This efficiency in takeoff reduces the need for additional power, so the gearbox design was rejected due to weight restrictions. The culmination of these finalized selections results in an aircraft that is best suited for the 2013 design criteria. 12

13 3.0 Aerodynamic Analysis This section describes the analytical processes the team used to size the aircraft, ensure its stability, and predict its performance in payload transport and delivery. 3.1 Aircraft Sizing To size the aircraft, AMAT set a primary design criterion of achieving the mission-required lifting capability with the lowest weight configuration possible Selection of Main Wing AMAT opted for a single element wing equipped with the Selig 1223 airfoil. The team calculated the viscous polar of this airfoil with XFOIL at Reynolds number 200,000. The maximum lift coefficient was found to be C L,max =2.1 as seen in Figure 3.1A. The team began the sizing process by using a rapid prototyping code to dimension the main wing, since it generates the Figure 3.1A Non-dimensionalized geometry of the Selig 1223 airfoil majority of the overall configuration s lift. Optimization was achieved by iterating takeoff performance for successively increasing chord lengths c with a given wingspan b until the desired takeoff condition was achieved. This process ensured minimum wing volume and thus weight. Inputs included C L,max of the airfoil, static thrust, and thrust slope of the engine. In the code, lift and induced drag were calculated, given zero-lift drag of the wing and other aircraft components C D,0, and the best chord was selected when a 3 climb angle (to ensure the aircraft can clear the runway) was achieved on takeoff. The team then created a 3D viscous polar for the finite design wing using Prandtl lifting-line theory, which applies the 2D viscous polar along the span. Figure 3.1B shows the effect of induced drag on the 3D polar, which is shifted further along the drag axis than its 2D counterpart. Airfoil C L,max = 2.1 and aircraft C L,TO = 1.7 and C L,cruise = 0.62 can be seen in Figure 3.1.B, and Figure 3.1A illustrates the more efficient L/D value in cruise compared to takeoff. 13

14 Figure 3.1B 2D and 3D Polars and Lift Slope Curves of the AMAT 2013 Main Wing Selection of Overall Aircraft Configuration With the main wing geometry constrained, AMAT proceeded to configure the empennage using equilibrium analysis of takeoff, the most demanding stage in the mission. The lifting tail design requires that the center of gravity location x cg is behind the main wing center of pressure at takeoff, so that the aircraft weight is balanced between the two lifting surfaces, as seen in Figure 3.2A. This is accomplished by moving backward the aircraft aerodynamic center AC, on which CG location is dependent for static stability. The team moved the AC aft by increasing the horizontal tails s aerodynamic moment contribution with a larger tail lifting area and a longer tail boom moment arm. Empennage sizing was governed by minimizing weight increases so that the tail would be able to at least lift its own weight in the final takeoff configuration. For stall stability, AMAT constrained the tail to have an aspect ratio AR less than main wing, so that first stall would occur at the forward lifting surface and create a nose-down pitching moment, ending the stall. As with rapid prototyping, designs were iterated and trimmed for takeoff with desired main wing C L,TO and climb angle. AMAT validated each iteration with an acceleration analysis of takeoff roll to determine that takeoff velocity could indeed be reached. The results of sizing are given in Tables 3.1A and 3.1B on the following page, and it can be seen that the tail indeed lifts its own weight. 14

15 Table 3.1A Final Lifting Surfaces Main Wing Horizontal Tail Vertical Tail Span [in] Chord [in] Winglet Design Advanced Modeling Aeronautics Team 2013 Table 3.1B Final Configuration CG Percentage of c m 51.55% Aircraft AC Percentage of c m 88.65% Main and Horizontal LE Separation in Lift Contribution of Tail at Takeoff 1.36 lbf 1.36F To reduce the amount of induced drag at takeoff, AMAT incorporated winglets into the 2013 design. With inputs of C L,TO, V TO, and the 2D viscous polar, the team used an optimization code to design winglets that are 12% of half of the wingspan in height and mounted so as to have a toe-in angle of 6.5. The toe-in angle causes the symmetric profile winglet to develop a circulation distribution that forces a constant downwash on the main wing that is less than that of the elliptic wing for the same global lift. Figure 3.1C shows the loading redistribution effect of winglets on the 2013 AMAT wing. Figure 3.1C Effect of Winglet Addition on 2013 AMAT design The ideal distribution of circulation is an elliptical profile, and is represented as a dashed line. To simulate this distribution, winglets shift the AMAT wing from the red circulation curve to the blue, maintaining the area under the curve by giving an effective elongation to the wingspan. The result is a 10% increase in efficiency at takeoff, which allows the plane to reach takeoff speed sooner. 15

16 3.2 Aircraft Stability and Control Advanced Modeling Aeronautics Team 2013 The controllability of the aircraft and the accuracy of the payload drop are extremely dependent its stability, which is defined into two categories. Static stability refers to a natural tendency of the aircraft to produce a restoring moment once displaced from equilibrium. Dynamic stability refers to the ability of these restoring moments and forces to return the aircraft to equilibrium over some time Longitudinal Stability and Control Longitudinal stability refers to an aircraft s stability about the pitch axis. Pitching is rotation that causes variance in the angle of the attack. The center of gravity CG is highly influential on the natural stability of the aircraft, and proper placement of its location was required for Zephyrus to be stable. The team used the equilibrium code to determine the optimum locations of center of gravity x cg and aerodynamic center x ac (shown in Figure 3.2A) for static stability. Figure 3.2A Center of gravity and aerodynamic center for the entire configuration. Lift locations at takeoff. To be longitudinally stable, the aircraft must respond to any disturbing pitching moment with an induced pitching moment in the opposite direction. The slope of the pitching moment coefficient is the product of the lift curve slope of the aircraft, which is positive for positive chamber airfoils, and the static margin, as in Eqn 1. 16

17 ( ) (1) For longitudinal stability, SM should be positive, meaning the CG must be ahead of the AC. The SM of Zephyrus was optimized to a value of 8% of the fuselage length as a reference area to give a desirable amount of stability without severe impediment of maneuverability. Because payload delivery could potentially alter x cg of the aircraft, AMAT chose to place the dynamic payload as close to the CG location as possible to minimize its disturbance of stability Lateral Stability and Control Lateral stability refers to the aircraft s stability about the roll axis. Many parameters influence roll stability including the height of the CG, dihedral angle, vertical tail size, rudder size, and ailerons. It is unstable to locate the CG above the AC, as this will create a moment response in the direction of a roll moment perturbation, increasing the perturbation. For this reason, AMAT chose to place the wings, and therefore the AC, above the fuselage and all components. This will ensure the CG is far below the AC, contributing a great deal of roll stability. Dihedral angle, or the angle at which the wings meet the fuselage, can also increase the roll stability of an aircraft, but due to the complications in manufacturing and the added weight of increased mounting mechanisms, the team decided to not utilize this form of stabilization. The team compensated for any shortcomings in lateral static stability by incorporating ailerons, which are roll-producing control surfaces that can be used to actively stabilize the aircraft with commands from the pilot and control system Directional Stability Directional stability refers to the aircraft s stability in the yaw direction. The team decided to use a single vertical tail and rudder for yaw stability. For any deflection from trim on the yaw axis, the vertical tail will become angled relative to the wind, which will induce a lifting force that creates an opposing moment to the yaw deflection. The winglets also provide some stability in this manner. 17

18 3.2.4 Control Surface Sizing Advanced Modeling Aeronautics Team 2013 For an aircraft to make a turn or climb and stray from it natural equilibrium, it must use control surfaces to induce repositioning moments. Zephyrus uses a combination of ailerons, rudder, and elevator to create these moments. To be effective, ailerons must be large enough to create a roll moment equal in magnitude to the roll moment created from a gust of wind on the vertical tail. The force of the wind is found from the time rate of change of the wind moment (Eqn 1) and is non-dimentionalized and set equal to the coefficient of roll of the aileron (Eqn 2):. ( ) (2) ( ) (3) To solve for the ratio of aileron chord to main wing chord τ, all other values were determined using equilibrium analysis or through assumption. The results are shown in Figure 3.2B. 7 Chord of Aileron Required for Various Perpendicular Gust Velocities 6 5 Aileron Chord, in Velocity of Perpendicular Gust, ft/s Figure 3.2B Aileron chord required to counteract a range of gust velocities. The team assumed a max perpendicular gust velocity equal to the forward velocity of the aircraft, approximately 32 ft/s. This resulted in τ = and the chord of the aileron equal to 5.0 in, given a factor of safety of

19 3.3 Aircraft Performance Advanced Modeling Aeronautics Team 2013 AMAT used equilibrium analysis as in part to calculate aircraft performance at important flight conditions like takeoff and cruise, as seen in Table 3.3A. Table 3.3A Zephyrus Fight Characteristics Velocity [ft/s] Thrust [lbf] Lift [lbf] Viscous Drag [lbf] Induced Drag [lbf] Takeoff Cruise The high lift state of Zephyrus s takeoff causes it to experience 150% more induced drag force than in cruise. This is due the high C L required to create the necessary lift force at low speed. In cruise, the thrust is exactly equal to the drag, and aircraft is traveling almost twice as fast as in takeoff. The required value of C L decreases to a point of larger L/D on the 3D viscous polar in Figure 3.1B. AMAT determined the limits of Zephyrus s aerodynamic performance by comparing predicted engine power performance with calculated overall drag behavior of the aircraft. First, the team calculated the drag contributions of the lifting surfaces as a sum of each surface s zero-lift drag C D0 and the induced drag C D,i. C D0 is a combination of drag from friction, profile, and interference: ( ( ) ) (4) The team calculated induced drag C D,i from its dependency on the total lift coefficient C L, which was weighed proportionally for each lifting surface. The vertical stabilizer had no lift and therefore no induced drag in the trim flight: ; ; (5), (6), & (7) The effect of the winglets on induced drag occurs in Eqn (5), where the increase in efficiency e causes a decrease in C D,i. The total drag of each wing was calculated as in Eqn (6). The fuselage and tail boom were considered to have negligible lift, so only C D0 was calculated for each. The greater length of these bodies allowed the formation of longer boundary layers that became turbulent, necessitating the use of the following relationship: 19

20 ( Advanced Modeling Aeronautics Team 2013 ( ) ) (8) The team treated the remaining external components (landing gear and engine) as bluff bodies, so that drag depended on frontal area and tabulated drag coefficients. Equivalent flat plate areas f were calculated to represent each component s drag in a form that could be directly added, and then a range of total drag forces were found by multiplying this sum by various dynamic pressures. In flight, thrust opposes drag, and performance is a direct function of their difference. The team assumed propeller power output P to be constant at low speeds and dependent on an experimentally measured static thrust T 0 =8.5 lbf found for a propeller of radius R=0.5 ft and pitch p=4 in. The team also determined the required power P req to sustain equilibrium by drag balance: ( ) (9) The team calculated thrust T to decrease linearly from its static maximum according to the following: ( ) (10) Figure 3.3A Comparison of available power to power required for equilibrium flight. ; ; 20

21 The power performance demonstration in Figure 3.3A shows the feasible range of flight conditions for Zephyrus. The top of the range is limited to the intersection of the two curves at V max, where the entire power of the propulsion system is needed to maintain cruise. Figure 3.3A also allows the calculation of rate of climb RC, with the maximum value corresponding to the greatest excess power available ΔP max : ( ) (11) 3.4 Payload Analysis Zephyrus is tasked with two main payload-related requirements: general lifting and delivery accuracy Payload Prediction The maximum lifting performance requirement for Zephyrus is the combination of weight of the aircraft, payload for delivery, and payload representative of fuel. Since fuel capacity directly affects the operational range of the aircraft, AMAT performed a payload prediction analysis to demonstrate mission capability in a variety of conditions. UC Davis Advanced Modeling Aeronautics Team Figure 3.4A Effect of increasing density altitude on maximum payload cabability. Team

22 Using equilibrium analysis with the trim setting for highest takeoff lift available, the team found maximum payload values at various densities, which are plotted in Figure 3.4A. A linear fit was applied to the plot to determine a maximum payload prediction equation for the 2013 AMAT aircraft. This relationship predicts that a gain in 4000 ft of density altitude from sea level at standard conditions will reduce Zephyrus s payload by less than 7.5%, showing that the aircraft will be mission-capable in a variety of environments and conditions that could require humanitarian aid Payload Drop The succeess of AMAT s mission also depends on the aircraft s ability to deliver its humanitarian aid package accurately, so that its contents arrive intact and no damage is caused by its arrival, which is the basis for the accuracy score S 1. AMAT used a viscous free-fall code to simulate the package s trajectory and predict the optimal conditions for its release, as shown in Figure 3.4B. These calculations were incorperated into the DAS so that it will be able to supervise the payload drop in real time. When the payload release is armed, the Figure 3.4B Sandbag trajectory prediction for drop from cruise speed. Ardupilot will calculate the expected impact locations and advise the secondary pilot on when to release the payload. The secondary pilot will use this data along with video stream to guide the primary pilot and determine when to drop the payload. 22

23 4.0 Component Structure and Manufacturing To meet the design intent, AMAT employed a number of structural and manufacturing innovations in the construction of the aircraft. 4.1 Main Wing Zephyrus s main wing is a balsa rib and composite hybrid wing that combines the lightness of traditional balsa construction with the strength and precision of composite materials. The team laser cut all balsa airfoil components to ensure exact shape. These were attached to the primary wing spar, which was made of sandwiched base wood and bi-ply carbon fiber with layers oriented at 0 /90 and 45 /-45 to provide axial and shear strength under bending. To reduce weight, the team cut holes from the spar center, where the material has less bending resistance. AMAT produced a sharp trailing edge TE from hotwire-cut foam attached to the wing with an underneath composite panel to provide torsional resistance. Only a single side is composite because the team determined that a fully enclosed crosssection of twice the weight would be a strength overdesign. The ailerons are fully enclosed in composite to guarantee rigidity so that shape deformation does not occur when they are actuated. Tape secures the ailerons to the wing and allows rotation. Figure 4.1A Main wing half section 23

24 The team milled winglets from plates of composite/foam sandwich material for simplicity and lightness. The flat plate satisfied the symmetric airfoil design condition. The TE s were sharpened to produce a Kutta condition, where the streamlines leave tangent to the TE, reducing form drag. The leading edges LE were made of rounded foam to reduce the front curvature at the airfoil, improving boundary layer attachment. Bumpers were extended downward from the winglets for wingtip strike protection, and these areas were strengthened with an additional layer of carbon fiber. Drag from these small extensions is negligible. 4.2 Empennage AMAT constructed the tail lifting surfaces with the same streamlined flat plate shape as the winglets and a similar hybrid structure design as the main wing. The TE design of the main wing was used on the horizontal and vertical stabilizers to give rigid shape to the empennage control surfaces. This year s hybrid design and reduced thickness of the stabilizers compared to AMAT s previous empennage has resulted in a significantly lighter tail, as summarized in Table 4.2A. Figure 4.2A - Empennage without Monokote The tail boom structure has also changed significantly from the heavy carbon/epoxy tube in AMAT s previous design. The team s current light-weight solution is a structure of three small carbon/epoxy rods joined at intervals by wood spacers, which, in combination, provide a cross-section that is highly 24

25 resistive in bending. Torsion is controlled by a Monokote covering. This design also simplified tail mounting, with low-profile composite brackets in the tail and direct insertion into the fuselage. This allowed for easy angling of the tail boom and reduced weight as seen in Table 4.2A. Table 4.2A Weight (lbf.) Comparison of Zephyrus to 2012 AMAT Design Main Wing Horizontal Tail Vertical Tail Tail Boom AMAT Zephyrus Difference Table 4.2A shows weight reduction in all listed components of Zephyrus despite each having larger dimensions than its predecessor. 4.3 Fuselage AMAT s fuselage is a box shape made of balsa/composite sandwich material detailed in Figure 2.3C that encloses the aircraft s operational components to reduce drag. The team chose a box cross-section so that minimal material would be required to provide adequate bending resistance. Further weight reduction was accomplished by removing excess center material. Figure 4.3A Fuselage and landing gear The fuselage provides mounting for the landing gear, engine, and tail boom and contains the avionics, fuel tank, and payloads. The static payload is mounted in slots in the fuselage walls to allow for CG 25

26 adjustment but does not contribute to the airframe s structure, which will be structurally sound upon static payload removal. The dynamic payload bay is a suspension member that supports the payload in tension and is located at the CG for reasons discussed in section The dynamic payload is held by a pin that is removed by a servo at the time of drop. 4.4 Landing Gear Zephyrus is equipped with a tricycle landing gear, which allows it to accelerate to takeoff speed with minimum induced drag and then rotate into the required takeoff incidence. The rear landing gear are tapered foam/composite struts that descend at 30 o from vertical to provide the best compromise between ground clearance, gear load distribution, and stable stance (layup in Figure 2.3C). The foam was cut with a 5-axis CNC mill allowing for complex geometry and rapid repeatability. AMAT placed the rear gear near and aft of the CG to balance the aircraft weight on the landing gear and to minimize the moment required from the tail for rotation. For increased control of landing rollout, the team designed minimalist brakes for the rear gear consisting of composite pads that are pushed against the rear wheels with servos. For the front landing gear, the team manufactured a custom air shock with spring and damper characteristics so that landing impact is softened and the resulting oscillations are impeded. The front gear is made primarily of aluminum, making it lighter than any metal shock available for purchase, and it can be rotated with a servo for steering on the tarmac. AMAT purchased and manufactured foam wheels for all gear to minimize weight. Figure 4.4A Machining of the airshock 26

27 4.5 Weight Management Advanced Modeling Aeronautics Team 2013 Fight score component S 2 had a major influence on the structural design of Zephyrus. To manage final weight, AMAT set allocations (shown in Table 4.5A) for structural systems, so that each could be minimized individually, rather than trying to meet the empty weight requirement all at once. Table 4.5A Zephyrus Component Weight Allocations Weight Allocation (lbf) % of Total Main Wing % Vertical Stabilizer % Horizontal Stabilizer % Fuselage % Tail Boom % Engine % Landing Gear % Avionics & Misc % To meet these strict weight restrictions, AMAT devised a number of weight-reducing innovations in addition to the structural designs already mentioned. For instance, AMAT manufactured fasteners out of aluminum, which is a third of the weight of steel. The team further reduced fastener weight by drilling out center material. Composite/foam layups were lightened by applying a slurry of epoxy and glass microballoons to foam surfaces. The addition of glass to epoxy effectively reduces its density without severely reducing its strength. Foam pores absorb a significant amount of epoxy, so filling them with slurry reduces the total amount of epoxy and therefore weight in the layup. Figure 4.5A - From left to right: Glass/Epoxy Slurry, Aluminum Fasteners, 5-axis foam mill used to machine lightweight landing gear and wheels. 27

28 5.0 Conclusion AMAT s aircraft design is optimized for the 2013 SAE Aero Design Advanced Class mission. The team followed a thorough design procedure to select aircraft features that, in combination, provide ideal aircraft functionality under the design constraints. AMAT then configured the aircraft design using codes for rapid prototyping, equilibrium, and acceleration analysis. This configuration was analyzed for stability and control as well as performance to confirm that it was suitable for the challenge. With this determined, manufacturing began with an emphasis on minimum weight construction. For competitive flight capability, the team designed a control system to simplify the pilot s approach to the target and supervise the conditions at which the payload is dropped. AMAT designed Zephyrus with the intention of creating an ideal combination of aerodynamics, structure, and electronics, and the team is excited to have the opportunity to see how its product performs at competition. 28

29 6.0 References 1. Bauchau, O.A. and Craig, J.I., Structural Analysis with Applications to Aerospace Structures, New York: Springer Science+Business, Chattot, J. J., Glider and Airplane Design for Students, Int. J. Aerodynamics, Vol. 1, No. 2, Etkin, Bernard, and Lloyd Duff Reid. Dynamics of Flight, Stability and Control. New York: John Wiley & Sons, Inc., Mattingly, Jack D, Elements of Propulsion: Gas Turbines And Rockets. AIAA Moran, Jack, An Introduction to Theoretical and Computational Aerodynamics, New York: Dover Nelson, Robert C. Flight Stability and Automatic Control. New York: McGraw-Hill, Inc., Roskam, Jan. Airplane Design Part1: Sizing of Airplane Kansas: Design, Analysis and Research Corporation, Roskam, Jan. Airplane Aerodynamics and Performance Kansas: Design, Analysis and Research Corporation, Shigley, Joseph E. et al. Mechanical Engineering Design: Seventh Edition, McGraw Hill, Inc

30 Empty CG Loaded CG D 2.01 Appendix A - Technical Plans D Wingspan In Approximate Empty Weight 7.9lbf Engine OS 46 AXII C Cargo Bay Volume 66.0 cu. in C Ballast for Empty CG Balance 15.0 lbf. placed in behind Main Wing Leading Edge B B A SolidWorks Student Edition. For Academic Use Only UNLESS OTHERWISE SPECIFIED: DIMENSIONS ARE IN INCHES TOLERANCES: FRACTIONAL 1/32 TWO PLACE DECIMAL.02 THREE PLACE DECIMAL.005 MATERIAL FINISH DO NOT SCALE DRAWING DRAWN CHECKED ENG APPR. MFG APPR. Q.A. COMMENTS: NAME DATE Edelman 3/3/2013 TITLE: University of California, Davis AMAT #215 AMAT Zephyrus SIZE B DWG. NO. 1 SCALE: 1:22 REV SHEET 1 OF 1 A