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it is about wing design process
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Arnab Roy
Chengzhi Qi
Wing Design Process
High Wing Geometry.
Increases the dihedral effect.
It makes the aircraft laterally more
stable. (fuselage will also make contribution)
Eases and facilitates to maintenance.
Rolling landing/Rolling take-off: Rotor blade and ground interactions
Cl vs alpha Clmax: Required stall speed mainly governs the Clmax ( Clmax gives
Cd, but Clmax gives better flight envelope)
αs: stall angle ( : 12° - 16° flight safety )
α0: Zero lift angle of attack ( {more negative value}: Leaves the capacity for more lift at 0 AOA)
Cl0: Cl at zero angle of attack ( : Implies we can create more lift at 0 AOA)
Clα: Affects the transition ( : Less power used in the rotor)
Cli : Ideal lift coefficient (Clcruise should be close to this to have minimum drag)
Stall Behavior: An airfoil with a gentle drop in lift after the stall is more desired
Cl vs alpha (Continued) Req & Assumption:
Cl max around 1.3
zero lift angle of attack (negative, with flag should around - 5-10 degree)
stall angle > 12 degree better around 15
zero angle of attack, no requirement, but as good as high it goes.
Cm1/4 Vs. α & Cmac Vs. α The slope of Cm Vs. alpha at ¼ chord relates to the
stability of the airplane (a reasonable negative slope is required)
Size of the tail, elevators are governed by Cm value.
More negative Cm results in larger tail = Higher drag,
heavier aircraft, higher costs.
Req & Assumption:
Cm vs alpha slope is negative
Cm at AC is around -.02 to -.05
Cd vs Cl Cd minimum as low as possible, reduce fuel required
At minimum slope: (Cd/Cl)min = (Cl/Cd)max
During 240 knots (Cruise): Cl should be Cl (ideal)
During 180 knots (loiter): Cl should be Cl (design)
Req & Assumption:
Cdmin about .003 to .006
Thickness Lift curve slope :Cla=1.8*pi*(1+0.8tmax/c)
Strength to support torque by rotors
Storing fuels
Enough space for rotation motion of the rotor
Reduce flutter
Req & Assumption:
t/cmax is about 15% to 20%
Airfoil Selection Criteria
Comparison of airfoils
Airfoils Choices: NACA 43018: ATR 42
Sm 701: High Lift Airfoil
NACA 64(4) 421: Fokker F-27
NACA 65(3) 218: Airtech Cn-235
Airfoil Design Objectives
Airfoil Stall
Angle
(12-16)
α0 (More
negative
exp.-2)
Clmax Clideal
≈Clcruise
Clα Stall
Behavior
Cm Vs. Cl Cm Vs.
α
(Cl/Cd
)
Thickness
NACA
43018
15° -3.2° 2.0 .85 .108 Smooth -.017 + 155 18.02%
Sm 701 15° -5.0° 1.8 .8 .12 Not
Smooth
-.137 + 150 15.99%
NACA
64(4)421
18.5° -2.95° 1.22 .55 .06 Smooth -.078 - 130 20.96%
NACA
65(3)218
13.5° -1.8° 1.0 .2 .075 Smooth -.041 - 80 18%
Evaluation of the performance
Design Objectives WEIGHT NACA 43018 Sm 701 NACA 64(4)421 NACA 65(3)218
Stall Angle (12-16) 10% 8.5 8.5 10 6
α0 (More negative ex.-2) 4% 8.5 10 7.5 6
Clmax(High, assumed 1.3) 15% 10 9.5 8 7
Clideal ≈Clcruise 7% 10 9 6 4
Clα (High) 10% 9 10 7 8
Stall Behavior 10% 9 2 10 10
Cm Vs. Cl (Low const Cm) 12% 10 4 8 9
Cm Vs. α 7% 0 0 10 10
High (Cl/Cd) 10% 10 9.5 8 6
Thickness 15% 9 8 10 9
Total Score 100% 8.74 7.135 8.58 7.7
NACA 43018
SM 701
NACA 64(4) 421
NACA 65(3) 218
Final Airfoil: NACA 43018
Aspect Ratio Justification:
Upcoming Analysis Using XFLR 5:
Questions:
Elevator Defection Airfoil Section
Area of Elevator
Deflection vs. V
CL of design elevator
NACA 0009 vs. NACA 0012
Airfoil Thickness Cm Clmax Cl/Cd Stall Angle
NACA 0009 9% 0.04 1.2 77 13
NACA 0012 12% 0.027 0.7 34 9
NACA 0009
Area of Elevator
Elevator Chord Se/Sh= .254
Assumed be/bh = 1
Ce/Ch=.254
Ce=.86 ft
Elevator Deflection vs. V
NACA 0009 with design elevator
Questions?