AAE 451 FALL 2006 TEAM 4

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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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AAE 451 FALL 2006 TEAM 4

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