Airbus A380 The SCUFF

Chase Ashton Doug Hillson Dave Simon

Gate Box Requirement

A380 AR = 7.53A340 AR = 9.21

is proportional to andR is proportional to

Range and L/D wouldbe %10.6 greater

max

⎟⎠⎞

D

L

If the A380 and A340 hadthe same wingspan

max

⎟⎠⎞

D

LAR

AR

106.153.7

21.9 =

Airfoil Selection

• Supercritical airfoilsconsidered– Whitcomb and others

• Transonic Helicopterblades considered– NLR-7223-43

• Boeing airfoils considered– Only access to old

airfoils– Company is very

secretive• Choose most recent

supercritical airfoil– SC(2)-0714

• The SC(2)-0714 airfoilappears relatively easy tomanufacture– Depending on t/c

• Ran on Tsfoil2– Normal Mach number

greater than 1.3• Airbus and Boeing have

airfoils not accessible forthis project– Designed for M = .80

or greater– Probably work better

than SC(2)-0714

NASA SC(2)-0714 AIRFOIL

NASA SC(2)-0714 AIRFOIL

Surface Area Calculation• Diagram used to

model the A380 inCAD

• Complex bodies ofrevolution modeled inInventor

• Planar bodiesmeasured in AutoCAD

Calculation usingFRICTION0D

C

21.6

29.1

22.3

41.5

231

Reference Length (ft)

0.6303138 (each)Engines (4)

0.0932990Vertical Tail

0.0814211Horizontal Tail

0.08218284Wing

0.11916050Fuselage

Thickness/ChordWetted Area (ft2)Component

Reference Area: 9380 ft2

All surfaces assumed 100% turbulent flowWing, Horizontal, Vertical Tail - Modeled as Planer SurfacesFuselage, Engines – Modeled as Bodies of Revolution

Input Parameters

02 DD CC

MIN⋅=

0DC

Results

0.601

0.601

0.605

0.602

0.600

0.611

0.689

CL

0.0322

0.0324

0.0326

0.0324

0.0322

0.0334

0.0424

CDmin

0.0161

0.0162

0.0163

0.0162

0.0161

0.0167

0.0212

CD0

18.650.9030000

18.620.8930000

18.510.8530000

18.600.8025000

18.680.7520000

18.320.5010000

16.260.105000

L/D MaxMach NumberAltitude

0DL CeARC ⋅⋅⋅= π

Cruise

Calculation usingFRICTION

LAMDES

Used to findMinimum Drag CG LocationMinimum CLTwist DistributionSection Cl distributionRoot and Tip Camber‘e’

InputSame planform as FRICTION analysisMach = 0.85 (Cruise)CD0 = 0.016310 chordwise horseshoe vortices20 spanwise rows

0.00

0.05

0.10

0.15

0.20

0.25

0.30

0.35

0.40

0.45

0.50

0 5000 10000 15000 20000 25000 30000 35000 40000

Altitude (feet)

CL

Req

uir

ed

Best Cruise Altitude at M=0.85

MTOGW

_ Fuel

Zero Fuel

0.0

0.5

1.0

1.5

2.0

2.5

3.0

0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1y/(b/2)

Tw

ist

(deg

rees

)

Linear Theory Twist DistributionTwist Distribution for minimum drag at Cruise CL

Main Wing

Tail

0.00

0.05

0.10

0.15

0.20

0.25

0.30

0.35

0.40

0.45

0.50

-125-115-105-95-85-75-65-55-45-35-25-15Cl

y

Section CL Distribution

Main Wing

Tail

Root and Tip Mean Camber Lines

0

0.5

1

1.5

2

2.5

0 0.2 0.4 0.6 0.8 1 1.2

y/(b/2)

% c

amb

er

Tip Camber

Root Camber

75% Span Camber

25% Span Camber

Stability ‘e’NP – Neutral Point, aft CG limit for stability

(107.6 ft aft of LE of Fuselage)10% Stable – CG 10% forward of NP

(fraction of mean chord)23% Stable – LAMDES Minimum Drag Solution

NP10% Stable

4 Feet

9.2 Feet

23% Stable

‘e’ = 0.74

References• Airbus Website

• UIUC Airfoil Database

• AIAA-2003-2886Commercial Aircraft

• Software:» LAMDES

» FRICTION

» VLMpc