Analisis del Avión Airbus A380

8/14/2019 Analisis del Avin Airbus A380

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Airbus A380 The SCUFF

Chase Ashton Doug Hillson Dave Simon


2/15

Gate Box Requirement

A380 AR = 7.53

A340 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 had

the same wingspan

max

D

L

AR

AR

106.1

53.7

21.9=


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Airfoil Selection

Supercritical airfoilsconsidered

Whitcomb and others

Transonic Helicopter

blades considered NLR-7223-43

Boeing airfoils considered

Only access to oldairfoils

Company is verysecretive

Choose most recentsupercritical 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 haveairfoils not accessible for

this project Designed for M = .80or greater

Probably work betterthan SC(2)-0714


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NASA SC(2)-0714 AIRFOIL


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NASA SC(2)-0714 AIRFOIL


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Surface Area Calculation

Diagram used to

model the A380 in

CAD

Complex bodies of

revolution modeled inInventor

Planar bodies

measured in AutoCAD


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Calculation using

FRICTION0

DC

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 flow

Wing, Horizontal, Vertical Tail - Modeled as Planer Surfaces

Fuselage, Engines Modeled as Bodies of Revolution

Input Parameters


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0

2DD

CCMIN

=

0D

C

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 =

Cruise

Calculation using

FRICTION


9/15

LAMDES

Used to find

Minimum Drag CG Location

Minimum CLTwist Distribution

Section Cl distribution

Root and Tip Camber

e

Input

Same planform as FRICTION analysis

Mach = 0.85 (Cruise)

CD0 = 0.0163

10 chordwise horseshoe vortices20 spanwise rows


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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)

CLRequired

Best Cruise Altitude at M=0.85

MTOGW

_ Fuel

Zero Fuel


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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 1

y/(b/2)

Twist(degrees)

Linear Theory Twist DistributionTwist Distribution for minimum drag at Cruise CL

Main Wing

Tail


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0.00

0.05

0.10

0.15

0.20

0.25

0.30

0.35

0.40

0.45

0.50

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

Cl

y

Section CL Distribution

Main Wing

Tail


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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)

%camber

Tip Camber

Root Camber

75% Span Camber

25% Span Camber


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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


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References

Airbus Website UIUC Airfoil Database

AIAA-2003-2886

Commercial Aircraft

Software:

LAMDES

FRICTION

VLMpc

Analisis del Avión Airbus A380

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