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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
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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
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Wing Geometry
Zero Sweep
Zero Dihedral
Taper Ratio = 0.4
Aspect Ratio = 11.25
Wing Span = 75 inch
Wing Area ≈ 500 sq. inch
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VT VTVT
W W
VT
VT
VT
W
W
L Sc
b S
c
L
S
S
b
HT HTHT
W W
HT
HT
HT
W
W
L Sc
C S
c
L
S
S
C
Volume Coefficient Horizontal Tail
Moment Arm Horizontal Tail
Area Horizontal Tail
Area Wing
Mean Wing Chord
Volume Coefficient Vertical Tail
Moment Arm Vertical Tail
Area Vertical Tail
Area Wing
Wing Span
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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
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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
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XFOIL Compute 2-D Data
Find approximate stall angle
Derive Cla – slope of lift curve
Convert to 3-D
57.31
lL
l
CC
C
eAR
e = span efficiency factor
Cla – 2-D Cl-alpha slope
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Compute Effect of Flap (Brandt pg. 152-153)
..
2
..2 cos
lh
f
Da
lhf
Daa
aLflapsL
S
S
S
S
CC
Ratio of flapped area to total wing area
Sweep angle of flap hinge
2-D change in alpha max
a
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Compute Effect of Tail (Brandt pg. 154-155)
0.25
0.725
21 10 31
7
1a at
oL avg h
h
tL L
C c z
AR l b
SC C
S
St/S - tail area over wing area
De/da - empirical curve fit
CLat – 3-D CL-alpha slope of tail
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Compute Total Lift Coefficient
Compute Total Drag Coefficient
eARk
CCkCC LLDD o
1
1
1
0LflapLtotalLL
awingLflapL
tailLwingLtotalL
CCCC
CC
CCC
64.0045.0178.1 68.0 ARe
CDo = 0.021
k1 = 0.039
e = 0.72
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Wing at Take-Off (Main Wing, Flaps, Tail)
Re: ~100,000
Max alpha: ~9.5-10 [deg]
-5 0 5 10 15-0.5
0
0.5
1
1.5
2
Angle of Attack [deg]
CL
2-D
3-DPlus Tail
Plus Flaps
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Wing at Cruise Conditions (Main Wing, Tail)
Re: ~600,000
-5 0 5 10 15-0.5
0
0.5
1
1.5
2
Angle of Attack [deg]
3-D
Lif
t C
oe
ffic
ien
t
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Drag Polar
Take-Off Speed
Cruise Speed
0.02 0.04 0.06 0.08 0.1 0.12 0.14 0.16 0.180
0.2
0.4
0.6
0.8
1
1.2
1.4
1.6
1.8
2
CD
CL
Stall
0.02 0.04 0.06 0.08 0.1 0.12 0.14 0.16 0.18-0.5
0
0.5
1
1.5
2
CD
CL
Cruise
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More Accurately Define CLmax
Possibly re-design tail so that CD is lower at cruise conditions
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