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Thesis TitleThesis TitleThesis TitleThesis Title
Kwame Nkrumah University of Science andTechnology, Kumasi
Department of Mechanical Engineering
n anc ng ero ynam c er ormanceEstimate in small Aircraft Development
using Object-Oriented Technique
By
YESUENYEAGBE ATSU KWABLA FIAGBE
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Presentation Outline
Background Information
Research Problem Literature Survey
-
Research Goal & Objectives
Research Activities
Aircraft Performance Analysis Results
Conclusion & Recommendations
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Background of Study Future of Small Aircraft Transportation
The SATS Project (1985-Present, NASA and partners)
High-volume operations at airports without control towers or terminal radarfacilities are possible
Available Technologies enabling safe landings at more airports. Integration of Small aircrafts into a higher capacity air traffic control system
Improved single-pilot ability to function competently in evolving, complexnational airspace
Expected Leapfrog in Small Aircraft use Performance Estimate is based on Wing Profile
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A SATS Aircraft Candidate, Hearst Corp. A New Civil Aviation Industry, NASA
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Research Problem Performance Estimate is based on Wing Profile
Lift & Drag coefficients & derived coefficient
Aircraft Design ConceptF=F Fixed, Desi n
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Performance Analysis
Creation of Objects,Numerical wind
Tunnel Development
Subsonic Aircraft Design
Mission Optimized Design ConfigurationObject Oriented Implementationof Design Concept
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USAF: Digital DATCOM developed and used mainly for the controlsystem design
Turevskiy at el, (1999) used combination of free and commercial off-the-shelf(COTS) modeling and simulation software and Digital Datcom for flightvehicle design process.
Raymer, D. (2002) in his doctoral thesis uses Multidisciplinary Optimization(MDO) technique to enhance the conceptual design process of the aircraftdesign. He employed various techniques such as orthogonal steepest descent
-
Literature Survey as Related to Research
Some Review of Related Works
, ,
options with his design code called RDS - selection of various components Neufeld, D et al (2007) developed Multi-Objective Genetic Algorithm
(MOGA) optimizer to assist in the design process for Very Light Jet (VLJ) andUnmanned Aerial Vehicle (UAV). The total lift was projected from the winggeometry.
Trevor S. Ferguson, (2007) Used Microsoft Excel spreadsheet to create anumerical RC design tool (Master thesis)
In all these cases, the aerodynamic analysis is based on the wing geometryand coefficients to determine the projected lift and drag.
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Aircraft Shape & InfluenceConfiguration = Function(Aerodynamic*, Propulsion, Structure, Mission)
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Discipline of Aircraft Design Multidisciplinary Process
Aerodynamics
Structural Mechanics
System Controls
Propulsion
Materials Engineering
o ers
Aircraft Design Stages
Conceptual Design: shape, arrangement of componentsand such features as, size, weight and general performance are
consideredPreliminary Design: specialists in areas e.g.. structures, landing
gear, and control systems will design and analyze aircraft portion
Detail Design: actual pieces to be fabricated are designed30-Mar-10 8
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Object-Oriented Programming
Object-Oriented programming (OOP) is a programming
model that uses "objects" and their interactions to design
applications and computer programs.
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, ,
and sending messages to other objects
Each object is viewed as an independent little machine
with a distinct role or responsibility
Advantages: Makes program discrete units and re-usable,
expandable and maintainable.
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Research Goal & Objectives
GoalDevelop Subsonic Aircraft Configurations with Optimal
Performance Capability and Improve estimation ofAerodynamics Performance parameters
Major Objectives:1. Develop Subsonic Aircraft Design Concept
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2. Identify & Exploit Engineering Design Parameters toConstruct Aircraft Configurations with Optimal Capability
3. Develop Object Oriented Program/Code to Implement theDesign Concept
4. Develop Aerodynamic Analysis Tools to Evaluate theintegrated Aircraft Configurations
5. Validate (Individual) Aircraft Subsystems & Tools
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To achieve Objective 1: Develop Subsonic Aircraft Design Concept
parameterDesignparametersFixedx_
_
)(xfF =
),( DragLiftePerformanc
GeometryF
Aircraft Design Concept
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=s
dsfFeperformanc
),( Pff=
Euler Equation
Boundary LayerEquation
kSjLiDF ++=
Drag Lift Slip force
routings
ononstructeometry
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Vertical Tail
LuggageCabin
Tail Boom Aircraft
Wing Fuselage EmpennageLanding
Gear
Mapping Aircraft Reality into OOP Environment
To achieve Objective 3. Develop an Object Oriented Program/Code to Implements ADC
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Nose
Nose-CabinInterface
Wing
Cabin
Horizontal TailPropeller
NoseNoseCabin
InterfaceCabin
LuggageCabin
HorizontalTail
VerticalTail
Tail Boom
OO Implementation of the Design Concept (FORTRAN 95)Functions Development
Module Development
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To achieve Objective 2. Identify & Exploit Engineering Design Parameters to ConstructAircraft Configurations with Optimal Capability
1
2
3
a
65
Design Variables Definition
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NOSE NOSE-CABIN INTERFACE CABIN
8
9
LUGGAGE CABIN
10
Tail Boom
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18 15
To achieve Objective 2. Identify & Exploit Engineering Design Parameters to ConstructAircraft Configurations with Optimal Capability
Design Variables Definition
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17
19
Wing/horizontal tail
Vertical Tail
14
16
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Illustration of Required Design Variables
Component Fixed Parameter ControllingDimension
Design Parameter Symbol Range
Aircraft Length L 1.0
Mach Number M 0.01 0.30
Altitude A Upto 5.0km
Nose Nose tip diameter Nose Length Nose Length to Plane Lengthratio
1 0.05 0.30
Location End Height End Height to Plane Length ratio 2 0.05 0.20
End Width End Width to Plane Length ratio 3 0.05 0.40
1
2
3
a
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Interface
(from Nose)
-4
. .
Width Height to Plane Length ratio 5 0.05 0.40
Offset Height Offset Height to Length ratio 6 0.0 0 0.20
Cabin Start coordinates
(from Nose-Cabin
Interface)
Length Cabin Length to Plane- Length
ratio7 0.10 0.40
Luggage Cabin Start coordinates
(from Cabin)
Length Length to Plane-Length ratio 8 0.1 0.25
End Diameter End Diameter to Plane-Length
ratio9 0.05 0.20
Tail Boom Start coordinates
(from Luggage Cabin)
End Diameter End Diameter to Plane-Length
ratio
10 0.01 0.05
4
6
5
7
8
9
1
0
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Component Fixed Parameter Controlling
Dimension
Design Parameter Symbol Range
Horizontal Tail Profile Span Span to Plane-Length ratio 11 0.15 0.50
Angle of Attack Root Chord Root-Chord to Plane-Length
ratio
12 0.05 0.15
Sweep angle Tip Chord Tip Chord to Root-Chord
ratio
13 0.25 1.00
Dihedral angle
Vertical Tail Profile Span Span to Plane-Length ratio 14 0.07 0.25
Angle of Attack Root Chord Root-Chord to Plane-Length 15 0.05 0.15
Cont
1
7
1
9
1
8
Illustration of Required Design Variables
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ratio
Sweep angle Tip Chord Tip Chord to Root-Chord
ratio
16 0.25 1.00
Dihedral angle
Wing Profile Span Span to Plane-Length ratio 17 0. 50 2.00
Angle of Attack Root Chord Root-Chord to Plane-Lengthratio
18 0.10 0.25
Sweep angle Tip Chord Tip Chord to Root-Chord
ratio
19 0.25 1.00
Dihedral angle
1
7
1
9
1
8
1
4
1
6
1
5
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Start
Input Data
Wing
NumericalWind Tunnel
Surface
Integration
(Area)
Empenna
ge
Horizontal
Tail
Vertical Tail
Surface
Integration
(Area)
Aircraft Design Flow Chart
Numerical
Wind
Tunnel
Fuselage
Nose
Nose-Cabin
Interface
Surface
Integration
(Area)
Cabin
Luggage-
Cabin
Tail Boom
NumericalWind
Tunnel
L, D
L, D
L, D
L, D
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Create NoseGeometry
Input Data
A
Surface
Create Nose-Cabin
Geometry
B
Surface
Create CabinGeometry
C
Surface
Create Lug-Cabin
Geometry
D
Surface
CreateTailboomGeometry
Surface
E
Aircraft Design Flow Chart
F
Integration
(Area)
EvaluateL & D
G
Integration
(Area)
EvaluateL & D
H
Integration
(Area)
EvaluateL & D
I
Integration
(Area)
EvaluateL & D
Integration
(Area)
EvaluateL & D
L, D
J
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Create Fuselage
Profile
Wind Tunnel
Experiment
A B C D E
Aircraft Design Flow Chart
Surface Properties
(Pressure, Tau)
Nose Surface
Properties
F
Nose-CabinSurface
Properties
G
Cabin SurfaceProperties
H
Lug-CabinSurface
Properties
I
Tailboom
Surface
Properties
J30-Mar-10 19
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Create Wing
Geometry
Aircraft Design Flow ChartInput Data
Create H-Tail
Geometry
Surface
Integration
(Area)
Create V-Tail
Geometry
Surface
Integration
(Area)
Numerical
Wind
Tunnel
n
TunnelExperiment
Surface
Integration(Area)
Surface
Properties
Evaluate
L & D
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Evaluate
L & D
L, D
Evaluate
L & D
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Illustration of Aircraft Configurations Generated
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parametersDesign
parametersFixedx
_
_
)(xfF =
),( DragLiftePerformanc
Geometry
F
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Wetted and Projected Areas
y
z
F Wetted cell of any
orientation
Pro ection onto the
Projection onto xz plane(y=constant): dSy
Projection onto theyz plane (x=constant): dSx
To achieve Objective 4. Develop Aerodynamic Analysis Tools to Evaluate Aircraft Configurations
x
xy plane (z=constant): dSz
))()(( AClBClABllarea =
x
yAn element
A B
C
Herons formula for area of a triangle
2)( CABCABl ++=( ) ( ) ( ){ }
( ) ( ) ( ){ }
( ) ( ) ( ){ }0.5
222
0.5222
0.5222
CBzCByCBx
CAzCAyCAx
BAzBAyBAx
ZZYYXXBC
ZZYYXXAC
ZZYYXXAB
++=
++=
++=
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Surface Integration
Validation
Plane Area
(Integration)
Area (Exact) Error
Surface 2.427975
Truncated Cone
A = ((C1 +C2)s/2)/4
A= (r1+r2)((r1-
r2
)2+h2)0.5/4
= (1.0+0.5)(2.0615)/4
= 2.42870969
0.0007346
9
(0.0302%)
Y-Z 0.5884430
Circle
A = A -A 0.0010056
To achieve Objective 5. Validate (Individual) Aircraft Subsystems & Tools
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=
(r22
r12
)/4= (1.02 0.52)/4
= (0.75)/4
=0.58904862
(0.1707%)
X-Z 1.5
Trapezium
A=(a+b)h/2=(0.5+1.0)(2.0)/2
=1.5
0.0(0.0%)
X-Y 1.5
Trapezium
A=(a+b)h/2
=(0.5+1.0)(2.0)/2=1.5
0.0
(0.0%)
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Surface Integration
Validation PlaneArea
(Integration)
Area (Exact) Error
Surface 3.138356
A=2rh/2
=2
(0.5)(2.0)/(2.0)
=
= 3.141592
0.003236
(0.103%)
To achieve Objective 5. Validate (Individual) Aircraft Subsystems & Tools
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Y-Z 1.000001Rear
1.0
A=LB= (1.0)(1.0)
= 1.0
0.000001(0.0001%)
0.0
X-Z 0.0 A=0.0 0.0(0.0%)
X-Y 2.000001
Rectangular
A=LB
=(2.0)(1.0)
=2.0
0.000001
(0.00005%)
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Forces Acting on Aircraft Lift : The Aerodynamic force component acting perpendicular to free
airstream direction.
Drag: The Aerodynamic force component in free airstream direction
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WeightLift
DragThrust
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Aircraft Aerodynamics Analysis
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Possible Solution/Analysis Methods
Full NS Solution using CFD/COTS ToolsCoupled 3D Euler & Boundary Analysis Tools (COTS)Coupled 2D Euler & Boundary Analysis Tools
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Full System of Navier Stoke Equation
Solving the following Coupled Equations
Mass Equation:
Momentum Equation:0)( =+
V
t
viscousxx Ffx
pu
t
u)()(
)(++
=+
V
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Energy Equation:
Boundary Conditions & Grids
viscouszz
viscousyy
Ffz
pw
t
w
Ff
y
v
t
)()()(
)()(
++
=+
++
=+
V
V
+++=
++
+
ViscousViscous WQpq
Ve
Ve
t)()(
22
22
VfVV
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Solving the following Coupled Equations:
Mass Equation:
Momentum Equation:
0)( =+
V
t
pu
u =+
)(V
Euler Equation & Boundary Analysis
Tools
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Energy Equation:
Boundary Conditions, BLA & Grids
z
pw
t
w
y
pv
t
v
xt
=+
=+
)()(
)()(
V
V
)(22
22
VV pqV
eV
et
=
++
+
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2D Euler Equation For Incompressible flow and
Irrotational steady flow
Mass E uation
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0= V
Momentum Equation
Defining the Stream function, , the two equations reduced to
02
= 0
2
2
2
2
=
+
yx
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Method of Implementation
02
=
30-Mar-10 302D Sliced-Aero-Model with Grids
02
2
2
2
=
+
yx
To achieve Objective 4 Develop Aerodynamic Analysis Tools to Evaluate Aircraft Configurations
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Aerodynamic Analysis (NWT)
u = K1
b = K5
Left = K3
To achieve Objective 4. Develop Aerodynamic Analysis Tools to Evaluate Aircraft Configurations
x
y
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2 = 0
Boundary conditions: = Constant
Wind Tunnel
x
y
L = K2 Right = K4
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Implementation
Boundary Fitted Grid method Laplace equation with the variable (, ) on the transformed or
computational grid and (x, y) on the physical grid
0
0
2
2
2
2
2
2
2
2
=
+
=
+
yx
Physical plane (x, y) Computational plane (, ) = (x, y), = (x, y).
and inverse (, ) (x, y)x = x(, ) and y = y(, ).
Interchanging the independent and dependent variables, we havetransformed elliptic equation given by Thompson et al, (1974)
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02
02
2
22
2
2
2
22
2
2
=
+
=
+
yyy
xxx
22
22
+
=
+
=
+
=
yx
yyxx
yx
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Grid Transformation
Boundary 3
Boundary 2
Object
Region
Boundary 1
Boundary 4
AB
y y
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Boundary 2
Boundary 1
Boundary 4Boundary 3
A
B
Bl
A1
Region
x
x
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The Governing Equation of interest is transformed
02
2
2
2
=
+
yx
Implementation
Boundary Fitted Grid method
022
22
2
2
=
+
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at the object boundary
at far field boundary
0),( =
SinxUCosyU ),(),(),(
=
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NACA 2414 at AoA=0 NACA 2414 at AoA=3
Win
gApplica
tion
30-Mar-10 35Fuselage at AoA=0 Fuselage at AoA=3FuselageA
pplicatio
n
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Velocity Estimation
+
+
+=
)()(
))(1(1
212
xygfxygfgJn
f
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+
+
+=
= )()(
))(1(1
212
xygxyggJn
V
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Pressure Solution to Euler equation
),( Pff=
+
= PV
VVP
22
12
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)(int
1
NspofoNumber
P
P
N
i
i
avg
=
=
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Force due to Pressure, P
Force due to Shear stress, FFF prrr
+=
}
Aerodynamic Force Evaluation
rrrrrr
kGjLiDFr
rrr
++=
( ) ++===
===
Szyxavg
Savg
S
S
zyxavg
S
avg
S
p
kdSjdSidSSdSdFr
rrrrr
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To achieve Objective 5. Validate (Individual) Aircraft Subsystems & Tools
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Validation
Velocity Distribution
Velocity Distribution, Ref @
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Sample Velocity Distributions
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To achieve Objective 2. Identify & Exploit Engineering Design Parameters to Construct AircraftC fi i i h O i l C bili
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ResultsResultsResultsResults
Configurations with Optimal Capability
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ComparativeComparativeComparativeComparative Analysis of 5 AircraftAnalysis of 5 AircraftAnalysis of 5 AircraftAnalysis of 5 Aircraft
ConfigurationsConfigurationsConfigurationsConfigurations
Aircraft Sample Design Parameters
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No. Sample #1 #2 #3 #4 #5
1 Nose Length to Plane Length Ratio 0.100 0.100 0.100 0.100 0.100
2 Nose End Height Plane Length Ratio 0.100 0.100 0.100 0.150 0.120
3 Nose End Width Plane Length Ratio 0.120 0.120 0.100 0.150 0.130
4 Nose Cabin Length Plane Length Ratio 0.100 0.100 0.100 0.100 0.150
5 Nose Cabin Offset Height Length Ratio 0.030 0.010 0.000 0.000 0.060
6 Nose Cabin End Width Plane Length Ratio 0.140 0.140 0.100 0.150 0.150
7 Cabin Length Plane Length Ratio 0.250 0.250 0.250 0.400 0.300
8 Luggage Cabin Length Plane Length Ratio 0.300 0.200 0.200 0.200 0.100
9 Luggage Cabin End Diameter Plane Length Ratio 0.060 0.020 0.060 0.150 0.050
Aircraft Sample Design Parameters
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10 Tail End Diameter Plane Length Ratio 0.010 0.010 0.010 0.010 0.010
11 H_Root Chord Plane Length Ratio 0.100 0.100 0.100 0.100 0.100
12 H_Tip To Root Chord Ratio 0.500 0.500 0.500 0.500 0.500
13 H_Span To Plane Length Ratio 0.200 0.200 0.200 0.200 0.300
14 V_Root Chord Plane Length Ratio 0.100 0.100 0.100 0.100 0.100
15 V_Tip To Root Chord Ratio 0.500 0.500 0.500 0.500 0.50016 V_Span To Plane Length Ratio 0.100 0.100 0.100 0.100 0.100
17 Wing Root Chord to Plane Length Ratio 0.200 0.300 0.300 0.300 0.300
18 Wing Tip To Root Chord Ratio 0.500 0.700 1.000 0.300 0.300
19 Wing Span To Plane Length Ratio 1.500 1.500 1.500 1.500 1.500
20 Wing Profile NACA1412
NACA
2412
NACA
2424
NACA
2410
NACA
2410
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#6
#7
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#9
#10
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Object Performance
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Object Performance
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Cabin Height Analysis
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Luggage Cabin Analysis
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Luggage Cabin Analysis
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Conclusion & Recommendation
ConclusionDesign Tool is Developed capable for more accurate aerodynamic
performance estimateAngle of Attack between 2o and 4o for Small aircraft
Limitations: Incompressible flow regime, Small aircraft
RecommendationPhysical Experimental validation of Luggage cabin length
parameter
Complementary areas need to be developed (structural design,Propulsion)
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ThankThankThankThank YouYouYouYou
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