AAE 451 AERODYNAMICS PDR 2 TEAM 4 Jared Hutter, Andrew Faust, Matt Bagg, Tony Bradford, Arun Padmanabhan, Gerald Lo, Kelvin Seah November 18, 2003

TEAM4 OVERVIEW Concept Review Aircraft CL and CM Updated Wing Size Aircraft Plots Follow-Up Actions

TEAM4 CONCEPT REVIEW High Wing S = 47.8 ft 2 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 Empennage Horizontal and Vertical Tails sized using modified Class 1 Approach (per D & C QDR 1)

TEAM4 AIRCRAFT LIFT COEFFICIENT Lift Coefficient C L = C L  *  + C L  e *  elevator + C L0 Matlab script based on Roskam Vol VI Ch8: C L = 5.41(rad -1 )*  (rad -1 )*  elevator Predator Codes from AAE 565 C L = 5.473(rad -1 )*  (rad -1 )*  elevator

TEAM4 AIRCRAFT PITCHING MOMENT Moment Coefficient C M = C M  *  + C M  e *  elevator + C M0 Matlab Script based on Roskam C M = (rad -1 )*  + ( )(rad -1 )*  elevator Predator Codes C M = (rad -1 )*  + (-1.058)(rad -1 )*  elevator

TEAM4 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

TEAM4 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

TEAM4 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

TEAM4 AIRCRAFT CL AND CM Iterative loop can be used Used old constraint diagram values for initial guess Used Wing Area as the Control variable Constraint Code Predator Code Max CL Code

TEAM4 AIRCRAFT CL AND CM Lift Coefficient C L = C L  *  + C L  e *  elevator + C L0 C L = 5.931(rad -1 )*  (rad -1 )*  elevator Moment Coefficient C M = C M  *  + C M  e *  elevator + C M0 C M = (rad -1 )*  + (-1.058)(rad -1 )*  elevator Reduce C M0 for clean flight

TEAM4 AIRCRAFT CL AND CM C M0 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 C M0 the most

TEAM4 AIRCRAFT CL AND CM Lift Coefficient C L = C L  *  + C L  e *  elevator + C L0 C L = (rad -1 )*  (rad -1 )*  elevator Moment Coefficient C M = C M  *  + C M  e *  elevator + C M0 C M = (rad -1 )*  + (-1.058)(rad -1 )*  elevator Incident Angle=.23 degrees Does not seem right, may be caused by Downwash from the Main Wing

TEAM4 AIRCRAFT PARAMETERS Wing Area= 34.5 ft^2 Wing Span= 13.1 ft Max CL= Degree Angle of Attack 0 Angle of Attack

TEAM4 TRIM DIAGRAM AT CRUISE CL=.4327

TEAM4 Drag Polar Based on Roskam Vol VI Ch 4

TEAM4 FOLLOW-UP ACTIONS Verify CD calculations Triple Check C M0 and Incident Angle of the Horizontal Tail React to changes from D+C, Propulsion, and Structures

AAE 451 Questions?

TEAM4 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. Positive Negative

TEAM4 AIRCRAFT PARAMETERS Lift Curve Slope for Elevator Deflection C L  e = f(elevator size, horizontal tail planform) Zero Angle of Attack Lift Coefficient C L0 = f(C L0W, C L0HT,  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

TEAM4 AIRCRAFT PARAMETERS Moment Coefficient C M = C M  *  + C M  e *  elevator + C M0 C M = (deg -1 )*  + ( )(deg -1 )*  elevator Moment Curve Slope C M  = f(dC M /dC L, C L )  dC M /dC L = f(CG, Aerodynamic Center of Aircraft)

TEAM4 AIRCRAFT PARAMETERS Zero Angle of Attack Moment Coefficient C M0 = f(C M0_W, C M0_HT [both about the CG]) LIFT WEIGHT Aerodynamic Center