Post on 30-May-2018
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Airbus A380 The SCUFF
Chase Ashton Doug Hillson Dave Simon
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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
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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