airfoil selection , roy

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?