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AAE 451 FALL 2006 TEAM 4 PowerPoint Presentation
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AAE 451 FALL 2006 TEAM 4

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AAE 451 FALL 2006 TEAM 4

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  1. Aerodynamics QDR AAE 451 FALL 2006 TEAM 4 Mark Koch Matt Lossmann Ki Bon Kim Andrew Martin Tung Tran Matt Drodofsky Ravi Patel Nizam Haris

  2. Overview • Aircraft Geometry • Airfoil Selection [wing, vertical tail and horizontal tail] • Wing and Tail Geometry • 3 – view Drawing • Aerodynamic Modeling • Method • Coefficient of Lift and Drag

  3. Airfoil Selection • Wing NACA 4410 • CLmax and CD fits mission requirement • Horizontal Tail NACA 0009 • Minimize Drag • Vertical Tail NACA 0009 • Minimize Drag • Small deflection compensate Propeller torque

  4. Wing Geometry • Wing Geometry • Zero Sweep • Zero Dihedral • Taper Ratio = 0.4 • Aspect Ratio = 11.25 • Wing Span = 75 inch • Wing Area ≈ 500 sq. inch

  5. Tail Geometry Sizing Volume Coefficient Horizontal Tail Volume Coefficient Vertical Tail Moment Arm Horizontal Tail Moment Arm Vertical Tail Area Horizontal Tail Area Vertical Tail Area Wing Area Wing Wing Span Mean Wing Chord

  6. Tail Geometry • Horizontal Tail Geometry • Span = 20 inch • Root Chord = 7 inch • Tip Chord = 4.8 inch • Area ≈ 108 sq. inch • Vertical Tail Geometry • Root Chord = 8 inch • Tip Chord = 4 inch • Area ≈ 53 sq. inch

  7. Aircraft Wetted Area Approximation Fuselage Approximation Lateral area of a cone: (π * R) * [sqrt(R² + H²)] Surface area of a sphere: 4*π*r² Main Wing: 2*SW Horizontal Tail: 2*SHT Vertical Tail: 2*SVT A/C Wetted Area: 1750 sq. in

  8. 3 - View Drawing Aircraft Wetted Area = 1750 sq in

  9. Aerodynamic Modeling Method • XFOIL • Compute 2-D Data • Find approximate stall angle • Derive Cla – slope of lift curve • Convert to 3-D e = span efficiency factor Cla – 2-D Cl-alpha slope

  10. Aerodynamic Modeling Method • Compute Effect of Flap (Brandt pg. 152-153) 2-D change in alpha max Ratio of flapped area to total wing area Sweep angle of flap hinge

  11. Aerodynamic Modeling Method • Compute Effect of Tail (Brandt pg. 154-155) St/S - tail area over wing area De/da - empirical curve fit CLat – 3-D CL-alpha slope of tail

  12. Aerodynamic Modeling Method • Compute Total Lift Coefficient • Compute Total Drag Coefficient CDo = 0.021 k1 = 0.039 e = 0.72

  13. Aerodynamic Modeling Method • Wing at Take-Off (Main Wing, Flaps, Tail) • Re: ~100,000 • Max alpha: ~9.5-10 [deg]

  14. Aerodynamic Modeling Method • Wing at Cruise Conditions (Main Wing, Tail) • Re: ~600,000

  15. Aerodynamic Modeling Method • Drag Polar • Take-Off Speed • Cruise Speed Cruise Stall

  16. Aerodynamic Modeling Improvements • More Accurately Define CLmax • Possibly re-design tail so that CD is lower at cruise conditions

  17. Fast, Sleek and Sexy!