1. Design of a Composite Wing with Leading Edge Discontinuity Daniel Hult AerE 423 Project December 12, 2009

2. Overview Background Project Goals Design Computational Analysis Fabrication Testing Results & Conclusions Future Work

3. Background Purpose – Discontinuity causes vortex to form, keeping flow attached to outer wing and ailerons – Improved stability and performance at high α – Spin prevention Cirrus Aircraft Company

4. Project Goals Determine the structural feasibility of a composite, single-piece wing with a discontinuous leading edge. Design, build and structurally test a single- piece composite wing.

5. Design Phases: – Aerodynamic Analysis – Structural Design Purpose of project is structural Aerodynamics only to get accurate loads

6. Aerodynamics XFLR5 Analysis – Open source aerodynamics for R/C gliders – Uses Vortex Lattice Method – Allows low Reynolds Number analysis of any wing

7. Structural Design Laminate Study – Analysis of laminate geometry with comp_core – Varied combinations of 0/90  plies and ±45  plies – Loading Tension and Bending Compression and Bending – Laminates with more ±45  plies performed better in bending

8. Structural Design Final Laminate – 6 plies of 0.002 in. thick bi-weave fiberglass – 4 plies at 0 and 90 degrees – 2 plies at +45 and -45 degrees

9. Computational Analysis ANSYS used for Finite Element Analysis Three cases tested – Isotropic material (aluminum) – Graphite-Epoxy composite – Fiberglass-Epoxy composite 300 N distributed load at tip – Loading from XFLR5 – Depicted test to be performed

10. Computational Analysis Fiberglass – Max Stress= 587 Mpa – Max disp = 1.25 cm

11. Fabrication Mold – Airfoil sections cut out of particle board – Used as stencils to hotwire blue foam – 2 sections joined and handle added to root

12. Fabrication

13. Lay-up – Hand lay-up around mold – Wrapped and cured with vacuum assistance.

14. Testing A Successful test would clearly accomplish project goals Wing anchored at root with load applied at tip Load added to tip until failure

15. Testing Screw MethodClamp Method

16. Testing Wood mount failed along screws (20 lb) Fiberglass failed along clamped shims (40 lb)

17. Results & Conclusions Wing failed at clamped root at small load ANSYS predicted stress concentrations and therefore failure at discontinuity The results were inconclusive, necessitating further testing

18. Future Work Better fabrication techniques and materials – Two-piece wing – Carbon Fiber – VARTM or Pre-Preg Better testing and mounting methods – Metal or composite mounting plate and insert – Metal or composite tip insert for loading

19. References Abbott, Ira H. and Albert E. von Doenhoff. Theory of Wing Sections: Including a Summary of Airfoil Data. New York: Dover Publications, Inc. c1959. “CAPS™ and Stall/Spin.” Cirrus Aircraft Company. Accessed 12 October 2009.. Deperrois, André. “About XFLR5 calculations and experimental measurements” August 2008.. Deperrois, André. “Guidelines for XFLR5 V4.16.” April 2009.. Goyer, Robert. “Airplane on a Mission: Created for use in the humanitarian field, the Quest Kodiak delivers raw utility at a great price.” Flying Magazine. February 2009. “Kodiak Features.” Quest Aircraft Company. Accessed 29 September 2009. http://www.questaircraft.com/index.php?filename=features.php Meschia, Francesco. “Model analysis with XFLR5.” RC Soaring Digest. February 2008: p27-51. NASA Langley Research Center. “Spin Resistance” Updated 17 October 2003..

20. Acknowledgments Dr. Vinay Dayal, Professor Chunbai Wang & Peter Hodgell, TA’s AerE 462 group, especially Robert Grandin for ideas and support Iowa State University, Department of Aerospace Engineering

21. Questions?

22. Background Uses – Messerschmitt Bf-109 – Large commercial jets – NASA Spin Prevention Tests – Cirrus SR20 – Quest Kodiak Airliners.net Cirrus Airliners.net

23. NASA Spin Prevention Figures from NASA Langley report

24. Aerodynamics Airfoil Design – NACA 2412 chosen for basis – Common, well-known low-speed airfoil – Discontinuity created by extending NACA 2412

25. Aerodynamics Wing Design – Basic wing designed to be fabricated and tested structurally – NACA 2412 inner section (0.3 m) – Modified airfoil outer section (0.2 m) – b/2=0.5 m – cr=0.25 m

26. Laminate Study TestMaterialLaminate T1T300/5208 Graphite-Epoxy{[0,90];6}s T2T300/5208 Graphite-Epoxy{[0.90];2,[45,-45];1,[0.90];3}s T3T300/5208 Graphite-Epoxy{[0,45,90,0,-45,90];2}s T4T300/5208 Graphite-Epoxy{[0,90,45,-45];3}s A1AS/3501 Graphite-Epoxy{[0,90];6}s A2AS/3501 Graphite-Epoxy{[0.90];2,[45,-45];1,[0.90];3}s A3AS/3501 Graphite-Epoxy{[0,45,90,0,-45,90];2}s A4AS/3501 Graphite-Epoxy{[0,90,45,-45];3}s Table A1: Test Laminates PlanarBending TestEx (Gpa)Ey (Gpa)Gxy (Gpa)PoissonEx (Gpa)Ey (Gpa)Gxy (Gpa)Poisson T195.99 7.170.0302106.785.287.170.0339 T288.51 13.740.105896.7479.6813.90.1187 T379.84 20.310.193488.6970.3119.970.2103 T469.68 26.880.29683.2571.3321.850.2333 A173.83 6.90.036781.9565.736.90.0413 A268.34 11.680.108574.5861.6711.80.1214 A362.02 16.470.190971.9551.0515.890.2116 A454.69 21.250.286564.6955.6617.590.2286

27. Laminate Study TensileCompressive TestX (Mpa)Y (Mpa)Z (MPa)X (Mpa)Y (Mpa)Z (Mpa) T1-7.5939.91.22E-08-8.2339.91.22E-08 T2-57.939.4-1.89-58.639.4-1.89 T3-95.138.7-6.33-95.738.7-6.33 T4-13238.5-2.3-13238.5-2.3 A1-9.7548.11.25E-08-10.448.11.25E-08 A2-59.747-2.22-60.447-2.21 A3-95.245.6-7.68-95.845.6-7.67 A4-13045.3-2.82-13145.2-2.82

28. Computational Analysis Isotropic – Max Stress= 654 Mpa – Max disp = 1.34 cm

29. Computational Analysis Carbon Fiber – Max Stress= 600 Mpa – Max disp = 1.25 cm

30. Mold

31. Finished Laminate

32. Testing