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AERODYNAMICS PDR 2

AERODYNAMICS PDR 2. TEAM 4 Jared Hutter, Andrew Faust, Matt Bagg, Tony Bradford, Arun Padmanabhan, Gerald Lo, Kelvin Seah November 18, 2003. OVERVIEW. Concept Review Aircraft CL and CM Updated Wing Size Aircraft Plots Follow-Up Actions. CONCEPT REVIEW. Empennage

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AERODYNAMICS PDR 2

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  1. AERODYNAMICSPDR 2 TEAM 4 Jared Hutter, Andrew Faust, Matt Bagg, Tony Bradford, Arun Padmanabhan, Gerald Lo, Kelvin Seah November 18, 2003

  2. OVERVIEW • Concept Review • Aircraft CL and CM • Updated Wing Size • Aircraft Plots • Follow-Up Actions

  3. CONCEPT REVIEW Empennage Horizontal and Vertical Tails sized using modified Class 1 Approach (per D & C QDR 1) High Wing S = 47.8 ft2 b = 15.5 ft, c = 3.1 ft AR = 5 Twin Booms 3 ft apart; 7.3 ft from Wing MAC to HT MAC Twin Engine 1.8 HP each Avionics Pod 20 lb; can be positioned front or aft depending on requirements

  4. AIRCRAFT LIFT COEFFICIENT • Lift Coefficient CL = CL* + CLe*elevator + CL0 Matlab script based on Roskam Vol VI Ch8: CL = 5.41(rad-1)* + 0.4675(rad-1)*elevator + 0.3086 Predator Codes from AAE 565 CL = 5.473(rad-1)* + 0.454(rad-1)*elevator + 0.3113

  5. AIRCRAFT PITCHING MOMENT • Moment Coefficient CM = CM* + CMe*elevator + CM0 Matlab Script based on Roskam CM = -2.0496 (rad-1)* + (-0.1771)(rad-1)* elevator + 0.0425 Predator Codes CM = -2.2682 (rad-1)* + (-1.058)(rad-1)* elevator - 0.2785

  6. AIRCRAFT CL AND CM • AAE 565 Predator code similar to Roskam • Roskam uses graphs in his book • Predator has the graphs hard coded into the program • Predator will be more accurate Update Constraint Diagram • Need Maximum CL for Constraint Diagram • Roskam Code solves for Maximum CL • .06 difference between Roskam Code and Predator for CL

  7. MAXIMUM LIFT COEFFICIENT • Need section cl along wing span • Increase Angle of Attack and find new section cls • Repeat until the wing begins to Stall • That is the stall angle • Integrate section cl’s to find Maximum CL

  8. AIRCRAFT CL AND CM • Three Major Codes: Predator, CL max, and Constraint Diagram • Predator: • Input: Aircraft Geometry • Output: CL and CM equations • Maximum Lift Coefficient • Input: Main Wing and Horizontal Tail Geometry • Output: CL Max and Alpha at CL Max • Constraint Diagram • Input: Flight Conditions, CL at 0 Alpha, CL max, Engine Info • Output: Wing Area and Required Power

  9. AIRCRAFT CL AND CM Predator Code • Iterative loop can be used • Used old constraint diagram values for initial guess • Used Wing Area as the Control variable Constraint Code Max CL Code

  10. AIRCRAFT CL AND CM • Lift Coefficient CL = CL* + CLe*elevator + CL0 CL = 5.931(rad-1)* + 0.59(rad-1)*elevator + 0.2809 • Moment Coefficient CM = CM* + CMe*elevator + CM0 CM = -3.6947(rad-1)* + (-1.058)(rad-1)* elevator -.3956 Reduce CM0 for clean flight

  11. AIRCRAFT CL AND CM • CM0 main contribution is from the Incidence angle of Horizontal Tail (-2.51 degrees) • Using the Iterative Loop, ran over a range of Horizontal Tail Incident angles • Found Incident Angle that reduced CM0 the most

  12. AIRCRAFT CL AND CM • Lift Coefficient CL = CL* + CLe*elevator + CL0 CL = 6.0339(rad-1)* + 0.6201(rad-1)*elevator + 0.4237 • Moment Coefficient CM = CM* + CMe*elevator + CM0 CM = -4.0421(rad-1)* + (-1.058)(rad-1)* elevator + 0.00 Incident Angle=.23 degrees Does not seem right, may be caused by Downwash from the Main Wing

  13. AIRCRAFT PARAMETERS • Wing Area= 34.5 ft^2 • Wing Span= 13.1 ft • Max CL= 1.8034 @ 12.88 Degree Angle of Attack • CD= .0339 @ 0 Angle of Attack

  14. TRIM DIAGRAM AT CRUISE CL=.4327

  15. Drag Polar Based on Roskam Vol VI Ch 4

  16. FOLLOW-UP ACTIONS • Verify CD calculations • Triple Check CM0 and Incident Angle of the Horizontal Tail • React to changes from D+C, Propulsion, and Structures

  17. Questions?

  18. Appendix Lift Curve Slope CL = f(CLW, CLHT, HT, w)  HT = Ratio of dynamic pressure. Mostly caused by propeller wash and velocity Downwash, w = Caused by main wing’s vortex flow on tail. Changes effective angle of attack for the tail. Negative Positive

  19. AIRCRAFT PARAMETERS • Lift Curve Slope for Elevator Deflection CLe = f(elevator size, horizontal tail planform) • Zero Angle of Attack Lift Coefficient CL0 = f(CL0W, CL0HT, HT, incident angles)  HT = Ratio of dynamic pressure. Mostly caused by propeller wash and velocity Incident angles are for both main wing and horizontal tail

  20. AIRCRAFT PARAMETERS • Moment Coefficient CM = CM* + CMe*elevator + CM0 CM = -0. 0225(deg-1)* + (-0.0027)(deg-1)* elevator + 0.0280 • Moment Curve Slope CM = f(dCM/dCL, CL)  dCM/dCL = f(CG, Aerodynamic Center of Aircraft)

  21. AIRCRAFT PARAMETERS • Zero Angle of Attack Moment Coefficient CM0 = f(CM0_W, CM0_HT [both about the CG]) LIFT Aerodynamic Center WEIGHT

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