2015 SAE AERO DESIGN - WEST COMPETITION MICRO CLASS DESIGN REPORT: L-406 SKYCRANE PUPR Aero Design Polytechnic University of Puerto Rico Team Number: 329 March 9, 2015 1|P age Table of Contents List of Figures and Tables ..........................................................................................................................3 Executive Summary ..................................................................................................................................4 Schedule Summary ...................................................................................................................................5 1. 2. Loads and Environments, Assumptions............................................................................................7 i. Design Loads Derivations .....................................................................................................................7 ii. Environmental Considerations..............................................................................................................8 Design Layout & Trades ..................................................................................................................9 i. Overall Design Layout and Size .............................................................................................................9 ii. Optimization (Sensitivities, System of systems: planform, layout, power plant, etc.).............................. 11 a) Competitive Scoring and Strategy Analysis ...................................................................................... 12 iii. Design Features and Details ............................................................................................................... 13 iv. Interfaces and Attachments ............................................................................................................... 13 3. Analysis ....................................................................................................................................... 14 i. Analysis Techniques........................................................................................................................... 14 a) Analytical Tools.............................................................................................................................. 14 b) Developed Models......................................................................................................................... 14 6.2. Performance Analysis ....................................................................................................................... 15 i. Runway/Launch/Landing Performance............................................................................................ 15 ii. Flight and Maneuver Performance .................................................................................................. 15 iii. Downwash .................................................................................................................................... 16 6.3. iv. Dynamic & Static Stability............................................................................................................... 17 v. Lifting Performance, Payload Prediction, and Margin ....................................................................... 17 Mechanical Analysis ..................................................................................................................... 18 i. Applied Loads and Critical Margins Discussion ................................................................................. 18 ii. Mass Properties & Balance ............................................................................................................. 18 7. Assembly and Subassembly, Test and Integration .......................................................................... 19 8. Manufacturing ............................................................................................................................. 21 9. Conclusion................................................................................................................................... 23 List of Symbols and Acronyms ................................................................................................................. 23 Appendix A – Supporting Documentation and Backup Calculations............................................................ 24 Appendix B – Payload Prediction Graph ................................................................................................... 26 Additional Material ................................................................................................................................. 28 2|P age List of Figures and Tables Figure 1: Aircraft Forces in a Level Turn ..................................................................................................... 7 Figure 2: Selected Airfoil ............................................................................................................................. 9 Figure 3: 3-D Lift Curve Slopes .................................................................................................................. 10 Figure 4: Selected Tail Airfoil .................................................................................................................... 11 Figure 5: Flight Score vs. Payload Fraction ............................................................................................... 12 Figure 6: Downwash vs. Angle of Attack................................................................................................... 16 Figure 7: Exploded View of Aircraft .......................................................................................................... 20 Figure 8: 3-D Printed Prototype Fuselage ................................................................................................. 21 Figure 9: Assembled Prototype Aircraft.................................................................................................... 22 Figure 10: Lift-to-Drag Ratio vs. Lift & Drag Coefficients (NACA 6409) .................................................... 25 Figure 11: Dynamic Thrust vs. Aircraft Speed ........................................................................................... 25 Figure 12: Dynamic Thrust Equation......................................................................................................... 26 Figure 13: Payload Prediction ................................................................................................................... 27 Figure 14: Cubic Loading vs. Aircraft Empty Weight................................................................................. 28 Table 1: Schedule Summary........................................................................................................................ 5 Table 2: Referenced Documents, References, and Specifications ............................................................. 6 Table 3: General Aircraft Layout ............................................................................................................... 11 Table 4: Performance Margins.................................................................................................................. 15 Table 5: Critical Structural Margins........................................................................................................... 18 Table 6: Level Turn Performance .............................................................................................................. 26 Table 7: Landing Performance .................................................................................................................. 26 3|P age Executive Summary The Micro Class category requires an aircraft weighing less than 10 pounds that fits within a 6” diameter container. The goal is to have the highest payload fraction possible with the lowest empty weight that a design will allow. This type of electric aircraft has to be hand-launched. For this purpose, an aircraft fitting those parameters was designed and manufactured using additive manufacturing. Our team goal for this competition was to reach a high payload fraction: an approximate value of 80%. The innovation that the team developed for this Micro Class Competition was a totally 3-D Printed aircraft. This was done to achieve a better payload fraction by reducing the airplane’s wei ght. In addition, it accelerated the manufacturing process. 4|P age Schedule Summary October November December January February March April Conceptual Design Preliminary Design Airfoil and Wing/Tail Geometry Selection Fuselage Geometry Selection Engineering Analysis Prototype Manufacturing Engineering Analysis Prototype Manufacturing Final Aircraft Design Settled First Prototype Flight Test Design Report Conclusion & Submission Aircraft Assembly Strategy Final Competition Preparations Table 1: Schedule Summary 5|P age Referenced Documents References Estimating R/C Model Aircraft Design: A Aerodynamics and Conceptual Approach; Performance; Nicolai Raymer Specifications Payload dimensions: 1.5” x 1.5” x 5” Mechanics of Flight: SAE Aero Design East and West Second Edition; Warren Desired high payload fraction Rules Phillips Tail Design; Mohammad Aircraft Performance and Aircraft must be assembled in less Sadraey Design; John D. Anderson than 150 seconds Propeller Static & Dynamic The fully packed aircraft system Introduction to Flight; John Thrust Calculator; Gabriel container shall weigh no more than D. Anderson Staples 10 pounds Shigley’s Mechanical Aircraft container must have a Engineering Design maximum diameter of 6” Table 2: Referenced Documents, References, and Specifications 6|P age 1. Loads and Environments, Assumptions i. Design Loads Derivations Given that our aircraft has a non-retractable propeller, it should be landed over grassy areas to reduce the risk of breaking. The aircraft will experience accelerations and decelerations during the flight course, such as when it is clearing the 180° turns, in addition to centripetal forces, shown in the figure below. Figure 1: Aircraft Forces in a Level Turn Here, the aircraft is performing a level turn. It can be seen that the lift is inversely proportional to the bank (roll) angle. In manned flight applications, this is the orthogonal force that the pilot will experience when he is pulling up on the aircraft. For the flight course, operational precautions must be taken into account to reduce this force so as to avoid any structural failures to the aircraft. 7|P age ii. Environmental Considerations Based on our design, several aspects of the location’s weather conditions were taken into consideration. The aircraft was manufactured completely out of PLA using a 3-D printer, and it is suggested that this material should not be exposed to areas of high humidity for long periods of time, since it can absorb the water in the environment, and thus adding more weight to the structure. Due to the mountains that surround the field, lack of air pressure is also being taken into consideration, something our pilot is aware of. The temperature during the time of the event is said to be in an average of 23°C and the modest elevations, there will be no problems with the flight path or the aircraft’s performance. Due to the limited wind information we had available, we decided to test the prototype in the harshest wind conditions in the PR metropolitan area (approximately 20 knots). 8|P age 2. Design Layout & Trades i. Overall Design Layout and Size The design process is considered a critical activity, because it becomes clear that the manufacturing and cost processes are determined by the decisions made in the initial design stages. By pointing out the stated requirements, the project execution was made possible. To achieve a high payload fraction value, it is desired to decrease the wing loading as much as possible; however, the wing area is constrained by the container’s diameter. To compensate for this, a combination of an airfoil capable of creating the necessary lift with high-lift devices was decided upon. An extensive analysis of different airfoils was conducted at a Reynolds number of approximately 100,000. Figure illustrates the lift curves slopes. The 3-D aerodynamic effects were already taken into consideration in the analysis. The CH10 and E423 airfoils both have a maximum lift coefficient of 2. The NACA 6409 airfoil was selected because, as it can be seen, although it has a moderate maximum lift coefficient, it will not stall immediately at high angles of attack, unlike the other airfoils. This is of great importance since low-speed flight is involved. Figure 2: Selected Airfoil 9|P age 3-D Lift Curve Slopes [All Airfoils] 2.5 LIFT COEFFICIENT 2 1.5 1 0.5 0 -15.00 -5.00 5.00 15.00 25.00 ANGLE OF ATTACK (DEG) NACA 6409 CH10 E423 Figure 3: 3-D Lift Curve Slopes A tapered high wing with an aspect ratio of 8.05 was selected as the final wing configuration due to it being more structurally and aerodynamically efficient than a constant chord wing. A wing of this type would have produced a non-elliptical lift distribution and the bending moments would have been more severe. Also, the addition of wing twist would have increased the volume necessary for the wing to fit in the container. Finally, adding sweep was not considered for many reasons: our design will not operate at very high speeds, and it would not be structurally beneficial. For stability reasons, a symmetrical airfoil with a projected horizontal aspect ratio of 4.68 is selected for the tail. This is desired because symmetrical airfoils have identical upper and lower surfaces, and find applications in V-tail designs, which is the chosen configuration for our aircraft’s tail. To account for stability, a tail sweep of 30° was incorporated to ensure longitudinal control at the high angles of attack that this aircraft will be expected to operate at. The NACA 0012 airfoil was selected for structural and data availability reasons. 10 | P a g e Figure 4: Selected Tail Airfoil The V-Tail configuration was selected for three reasons: ii. Less wetted area, which in turn produce less drag. Less material used due to vertical tail elimination. Less servos and linkages are required for control surface operations. Optimization (Sensitivities, System of systems: planform, layout, power plant, etc.) Wing Airfoil: NACA 6409 Span: 45 inches Reference Area: 259 in2 Aspect Ratio: 8.05 Taper Ratio: 0.4 Tail Airfoil: NACA 0012 Span: 10.4 inches Reference Area: 20.33 in2 (Horiz. Proj.) 6.61 in2 (Vert. Proj.) Aspect Ratio: 4.2 Taper Ratio: 0.4 General Empty Weight: Approx. 3 pounds Taper Ratio: 0.4 Moment Arm: 13.85 in. Aircraft Length: 24.04 in. Fuselage Diameter: 5.3 in. Propeller: 12” diameter x 7” pitch Table 3: General Aircraft Layout 11 | P a g e a) Competitive Scoring and Strategy Analysis According to Section 6.5 in the rule guide, the Final Flight Score is mostly dependent on the payload fraction. Using the Flight Round formula, an interpolation of payload fractions and 4 different container lengths was achieved. The resulting plots were linear in nature and the equations for each container size were obtained. Figure shows the Flight Score versus the Payload Fraction for each of these lengths. Flight Score vs Payload Fraction 95.00 90.00 85.00 80.00 75.00 70.00 10 in Contai ner 15 in Contai ner 18 in Contai ner 20 in Contai ner 65.00 60.00 55.00 50.00 0.50 0.55 0.60 0.65 0.70 0.75 0.80 Figure 5: Flight Score vs. Payload Fraction 12 | P a g e iii. Design Features and Details The design was heavily constrained by the container’s dimensions, so compact, modular design for fast and easy assembly was implemented. For example, ailerons are removable and the airframe was constructed by additive manufacturing; the largest part measuring 14.28 inches. The transmitter was programmed so that two control surfaces have the function as rudders and elevators for the tail, and ailerons and flaps for the wing. These configurations were decided upon in order to maximize the wing and tail area that can be fit inside the container. iv. Interfaces and Attachments Custom-made fittings were designed and 3-D printed for junctures of the V-Tail and wings. Some of the fittings for the 3-D printed parts were constructed with other materials like wood. Tie wraps and nylon screws are considered for use to join the fuselage parts together. 13 | P a g e Analysis 3. i. Analysis Techniques a) Analytical Tools Throughout the design process, Creo Parametric was used to simulate our ideas. This helped us to make decisions based on the information we extract from the CAD. This is also an advantage in terms of time and budget regarding the design. For example, we could see tolerance errors with the des ign without building the parts, and procure whether or not the aircraft would fit into the designed container. Microsoft Excel was extensively used for aerodynamic and performance analyses. This permitted the development of data tables and graphs to predict and optimize the behavior of our design. For example, graphs comparing lift coefficients, lift-to-drag ratios and the drag polar for each airfoil were developed to analyze how well one performs compared to the other. b) Developed Models A prototype was built to test the flying qualities of the planform chosen for the wings and V-tail. Also, the programming of the control system (transmitter) was developed using this prototype, thus avoiding the risk of damaging the final aircraft. 14 | P a g e Performance Analysis 4. Aircraft must be hand-launched. Aircraft is required to remain airborne and fly past the designated turn points, perform the two 180° turns in heading, and arrive at the landing zone. The aircraft must take off and land intact to receive points for the flight. All parts must remain attached to the aircraft during flight and during the landing maneuver. Aircraft must land in a designated landing zone measuring 200 feet in length. Table 4: Performance Margins i. Runway/Launch/Landing Performance The aircraft will be hand-launched, according to the stated requirements by SAE. An estimated launch speed of 30 feet per second was assumed. This will give the aircraft the extra push it needs to achieve the pre-analyzed flight performance. Using Anderson’s text, the landing performance was calculated. The ground roll was not taken into consideration since the aircraft does not have a landing gear, and our runway in this case will be grass. Using approximations stated by the book, such as the approach angle, the estimated landing distance from a 50-foot obstacle was determined to be 974 feet. ii. Flight and Maneuver Performance The installed motor will provide approximately 11,160 revolutions per minute (RPM) to the propeller, with dimensions of 12” diameter and 7” pitch. This, in turn, will operate the aircraft at a range of speeds between 40 and 55 miles per hour (MPH). Since one of the competition objectives is to clear two 180° turns, the turn rate needs to be compensated for the load factor so as to avoid wing support failure. The range for unpowered flight was determined assuming that the maximum flying altitude is 50 feet.1 1 See Appendix A for calculations. 15 | P a g e iii. Downwash The downwash angle of a typical wing is a function of its sectional lift coefficient and aspect ratio, and can be approximated by the following equation. 𝜖= 2𝐶𝐿,𝑤 𝜋𝐴𝑅𝑤 Downwash vs. Angle of Attack 12.00 DOWNWASH [DEGREES] 10.00 8.00 6.00 y = 0.3326x + 2.1622 4.00 2.00 0.00 -2.00 -10.00 0.00 10.00 20.00 30.00 ANGLE OF ATTACK [DEGREES] Figure 6: Downwash vs. Angle of Attack It is observed from the above figure that the downwash experienced by the wing is directly proportional to its angle of attack. This is a consequence of the increasing lift in the wing. Too much downwash can create a turbulent airflow over the tail, negatively impacting its performance. 16 | P a g e iv. Dynamic & Static Stability An important measure of the tail effectiveness is the horizontal tail volume coefficient, shown in the following equation. 𝑉𝐻 = 𝑆𝐻 𝑙 𝐻 𝑆𝑐 SH is the horizontal stabilizer planform area, lH is the horizontal stabilizer moment arm, S is the wing planform, and c is the wing chord. For this aircraft, the chosen tail volume coefficients for the horizontal and vertical tails were 0.5 and 0.04, respectively. These values were picked for a homebuilt aircraft.2 Another important parameter required for stability is the location of the aircraft’s center of gravity. The wing was placed on a location that would provide a positive static margin: an approximate value of 10% was obtained. v. Lifting Performance, Payload Prediction, and Margin We researched several heavy-lift airplanes and saw that none of them would exceed a cubic loading of 3.0, and in fact, a payload fraction above 80% was obtained with an airplane with a cubic loading of 2.76. Therefore, we used 80% payload fraction and a maximum cubic loading of 3.0 as our goal using the largest wing area we could fit in the container, that we could add flaperons to during the assembly. Figure 133 illustrates the sensitivity of empty weight to cubic loading and payload fraction. 2 3 Table 6.4, Page 160. Aircraft Design: A Conceptual Approach, Fifth Edition. Graph shown in Additional Material (page 28) 17 | P a g e 5. Mechanical Analysis Aircraft must stay intact during flight and support all dynamic loads. Aircraft must support variable payloads according to SAE requirements. The aircraft must take off and land intact to receive points for the flight. Broken propellers are allowed. Table 5: Critical Structural Margins i. Applied Loads and Critical Margins Discussion During the three rounds of the competition, the aircraft will have to carry an increasing payload every round, to test how well the aircraft is designed. In addition, as mentioned in Section 3.1 of this report, the aircraft needs to support the forces encountered when executing level turns, such as the Gforces. Table 54 shows the calculated level turn parameters for the installed motor’s speed limits. Our airplane was designed to routinely sustain a G-force of 2.0 by assuming a 3.0 ultimate load factor limit. ii. Mass Properties & Balance The weight prediction of the airframe was performed using Creo Parametric. Inputting the values of the material properties of the PLA, a full report was obtained. 4 See Appendix B. 18 | P a g e 6. Assembly and Subassembly, Test and Integration The aircraft will be divided into 3 assembly points each: fuselage, wings and tail. Fuselage The fuselage was printed in 5 different sections; two of them will be permanently joined. Each section will be joined with nylon screws. The payload will be carried inside the center section. The frontal section will contain the motor and propeller. The rear section will hold the tail assembly piece and will hold them together with 2 nylon screws. Also the wiring for all the electrical components will mostly be inside the fuselage. Wings The wings will be divided into a total of 12 pieces: 3 for one wing and 3 for the “flaperons”, and two sections will be permanently joined together. These will be attached to the fuselage using rectangular spars. Each wing has two channels in which the spars will be passed from one wing to the other, passing through the top part of the middle section of the fuselage. The “flaperons” for each wing will be attached with hinges to improve the stability. Tail The tail will be divided into 6 pieces. The control surfaces on the tail wings will be attached the same way as for the wings. When these are attached, the tail wings will be placed in between 2 pieces that will hold the fuselage together with 2 nylon screws. 19 | P a g e Electronic components 4 servos will be installed for each control surface: two for the “flaperons” and two for the V-tail. A receiver, antenna, 11.1 volt lithium polymer battery, controller with BEC system, outrunner brushless motor will be the primary electronic components used. Figure 7: Exploded View of Aircraft 20 | P a g e 7. Manufacturing The aircraft was fabricated using additive manufacturing. This was decided because when using wood, the manufacturing of each piece would have required 2 to 3 weeks. When using 3-D printing, the manufacturing of the aircraft took nearly 23 hours. The material the team chose was PLA because it is cost efficient, easier to manufacture, and lighter than wood. This also gives us the advantage of lighter structures throughout the airframe. Figure 8: 3-D Printed Prototype Fuselage The wings were manufactured in 6 sections per wing. Each section of the wing was printed at a length of 7.15”. The wings contain 2 spars, one measuring 45” long and the other measuring 27” long, each crossing from one wing to the other. Each section of the wings will have an interlocking attachment to help in the assembly process and also to resist axial loads that might be applied to the wings. These attachments were primarily made to be able to connect each section of the wing to each other and the fuselage. 21 | P a g e The base of the tail, which has a 0.50” diameter and 1” length tube on one of its faces, was inserted into the fuselage. The base’s dimensions are 2”diameter and a 2.40” length. The V-tail with each tail wing have dimensions of 5.50” of width, and 1.95” in depth. Figure 9: Assembled Prototype Aircraft 22 | P a g e 8. Conclusion The PUPR Aero Design team has conducted a complete conceptual design, performed a thorough engineering analysis, and completed the construction of a final design that will meet the requirements laid out by the Society of Automotive Engineers for the Aero Design West competition. With a low empty weight and a smooth, streamlined body, the “L-406 Skycrane” is more than prepared to take to the skies in the April competition. The aircraft is extremely lightweight, aerodynamically efficient, and stable. List of Symbols and Acronyms 23 | P a g e AR Aspect ratio W Aircraft weight α Angle of attack CL Lift coefficient MAC Mean aerodynamic chord λ Taper ratio D Total drag L Total lift V Velocity S Wing area c Wing chord b Wingspan α0 Zero-lift angle of attack CD 3D Polar Drag Appendix A – Supporting Documentation and Backup Calculations 24 | P a g e 18.00 0.160 16.00 0.140 14.00 0.120 0.100 10.00 0.080 CD L/D RATIO 12.00 8.00 0.060 6.00 0.040 4.00 0.020 2.00 0.00 0.000 0.500 0.000 1.500 1.000 CL Lift-to-Drag Ratio Drag Polar Figure 10: Lift-to-Drag Ratio vs. Lift & Drag Coefficients (NACA 6409) 9.00 8.00 7.00 Thrust, F (lbf) 6.00 F = -0.3656*V0 + 8.2425 5.00 4.00 3.00 2.00 1.00 0.00 0.00 5.00 10.00 15.00 20.00 25.00 Aircraft Airspeed, V0 (mph) Figure 11: Dynamic Thrust vs. Aircraft Speed V [mph] Load Factor nmax Roll Angle φ (degrees) Turn Radius (feet) Turn Rate (degrees/s) 25 | P a g e 40 45 50 55 1.08 1.37 1.69 2.05 22 43 54 61 259 145 123 113 13.0 26.1 34.3 40.8 Table 6: Level Turn Performance Required Thrust (Drag) [pounds] Lift-toDrag Ratio Thrust-toWeight Ratio Glide Angle (degrees) Sink/Climb Rate @ 50 mph [feet/s] Range (50 foot obstacle) [feet] 0.70 0.64 0.58 0.54 0.56 0.57 0.84 1.06 1.27 12.78 14.04 15.40 16.79 16.01 15.80 10.66 8.52 7.08 0.078 0.071 0.065 0.060 0.062 0.063 0.094 0.117 0.141 4.48 4.08 3.72 3.41 3.57 3.62 5.36 6.70 8.04 5.72 5.21 4.75 4.36 4.57 4.63 6.85 8.55 10.25 639 702 770 840 801 790 533 426 354 Table 7: Landing Performance Figure 12: Dynamic Thrust Equation Appendix B – Payload Prediction Graph 26 | P a g e Payload Prediction Graph at Maximum Velocity 6.80 6.60 PW = -0.0004hdensity + 6.7744 Payload Weight [lbf] 6.40 6.20 6.00 5.80 5.60 5.40 300 700 1100 1500 1900 2300 2700 3100 Density Altitude [slug/ft3] Figure 13: Payload Prediction 27 | P a g e Additional Material Figure 14: Cubic Loading vs. Aircraft Empty Weight 17.00 16.00 Lift-to-Drag Ratio 15.00 14.00 13.00 12.00 11.00 10.00 30.00 40.00 50.00 60.00 70.00 80.00 90.00 Stall Speed [mph] 28 | P a g e