Enhancing Aerodynamic Performance Estimate in small Aircraft Development using Object-Oriented Technique

8/9/2019 Enhancing Aerodynamic Performance Estimate in small Aircraft Development using Object-Oriented Technique

1/52

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


2/52

Presentation Outline

Background Information

Research Problem Literature Survey

-

Research Goal & Objectives

Research Activities

Aircraft Performance Analysis Results

Conclusion & Recommendations

30-Mar-10 2


3/52

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

30-Mar-10 3

A SATS Aircraft Candidate, Hearst Corp. A New Civil Aviation Industry, NASA


4/52

Research Problem Performance Estimate is based on Wing Profile

Lift & Drag coefficients & derived coefficient

Aircraft Design ConceptF=F Fixed, Desi n

30-Mar-10 4

Performance Analysis

Creation of Objects,Numerical wind

Tunnel Development

Subsonic Aircraft Design

Mission Optimized Design ConfigurationObject Oriented Implementationof Design Concept


5/52

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.

30-Mar-10 5


6/52

Aircraft Shape & InfluenceConfiguration = Function(Aerodynamic*, Propulsion, Structure, Mission)

30-Mar-10 6


7/52


8/52

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


9/52

Object-Oriented Programming

Object-Oriented programming (OOP) is a programming

model that uses "objects" and their interactions to design

applications and computer programs.

30-Mar-10 9

, ,

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.


10/52

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

30-Mar-10 10

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


11/52

To achieve Objective 1: Develop Subsonic Aircraft Design Concept

parameterDesignparametersFixedx_

_

)(xfF =

),( DragLiftePerformanc

GeometryF

Aircraft Design Concept

30-Mar-10 11

=s

dsfFeperformanc

),( Pff=

Euler Equation

Boundary LayerEquation

kSjLiDF ++=

Drag Lift Slip force

routings

ononstructeometry


12/52

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

30-Mar-10 12

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


13/52

To achieve Objective 2. Identify & Exploit Engineering Design Parameters to ConstructAircraft Configurations with Optimal Capability

1

2

3

a

65

Design Variables Definition

30-Mar-10 13

NOSE NOSE-CABIN INTERFACE CABIN

8

9

LUGGAGE CABIN

10

Tail Boom


14/52

18 15

To achieve Objective 2. Identify & Exploit Engineering Design Parameters to ConstructAircraft Configurations with Optimal Capability

Design Variables Definition

30-Mar-10 14

17

19

Wing/horizontal tail

Vertical Tail

14

16


15/52

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

30-Mar-10 15

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


16/52

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

30-Mar-10 16

ratio


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


ratio

19 0.25 1.00

Dihedral angle

1

7

1

9

1

8

1

4

1

6

1

5


17/52

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

End30-Mar-10 17


18/52

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


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

30-Mar-10 18


19/52

Create Fuselage

Profile

Wind Tunnel

Experiment

A B C D E


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


20/52

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

30-Mar-10 20

Evaluate

L & D

L, D

Evaluate

L & D


21/52

Illustration of Aircraft Configurations Generated

30-Mar-10 21

parametersDesign

parametersFixedx

_

_

)(xfF =

),( DragLiftePerformanc

Geometry

F


22/52

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

++=

++=

++=

30-Mar-10 22


23/52

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

30-Mar-10 23

=

(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%)


24/52

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%)


30-Mar-10 24

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%)


25/52

Forces Acting on Aircraft Lift : The Aerodynamic force component acting perpendicular to free

airstream direction.

Drag: The Aerodynamic force component in free airstream direction

30-Mar-10 25

WeightLift

DragThrust


26/52

Aircraft Aerodynamics Analysis

30-Mar-10 26

Possible Solution/Analysis Methods

Full NS Solution using CFD/COTS ToolsCoupled 3D Euler & Boundary Analysis Tools (COTS)Coupled 2D Euler & Boundary Analysis Tools


27/52

Full System of Navier Stoke Equation

Solving the following Coupled Equations

Mass Equation:

Momentum Equation:0)( =+

V

t

viscousxx Ffx

pu

t

u)()(

)(++

=+

V

30-Mar-10 27

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


28/52

Solving the following Coupled Equations:

Mass Equation:

Momentum Equation:

0)( =+

V

t

pu

u =+

)(V

Euler Equation & Boundary Analysis

Tools

30-Mar-10 28

Energy Equation:

Boundary Conditions, BLA & Grids

z

pw

t

w

y

pv

t

v

xt

=+

=+

)()(

)()(

V

V

)(22

22

VV pqV

eV

et

=

++

+


29/52

2D Euler Equation For Incompressible flow and

Irrotational steady flow

Mass E uation

30-Mar-10 29

0= V

Momentum Equation

Defining the Stream function, , the two equations reduced to

02

= 0

2

2

2

2

=

+

yx


30/52

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


31/52

Aerodynamic Analysis (NWT)

u = K1

b = K5

Left = K3

To achieve Objective 4. Develop Aerodynamic Analysis Tools to Evaluate Aircraft Configurations

x

y

30-Mar-10 31

2 = 0

Boundary conditions: = Constant

Wind Tunnel

x

y

L = K2 Right = K4


32/52

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)

30-Mar-10 32

02

02

2

22

2

2

2

22

2

2

=

+

=

+

yyy

xxx

22

22

+

=

+

=

+

=

yx

yyxx

yx


33/52

Grid Transformation

Boundary 3

Boundary 2

Object

Region

Boundary 1

Boundary 4

AB

y y

30-Mar-10 33

Boundary 2

Boundary 1

Boundary 4Boundary 3

A

B

Bl

A1

Region

x

x


34/52

The Governing Equation of interest is transformed

02

2

2

2

=

+

yx

Implementation

Boundary Fitted Grid method

022

22

2

2

=

+

30-Mar-10 34

at the object boundary

at far field boundary

0),( =

SinxUCosyU ),(),(),(

=


35/52

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


36/52

Velocity Estimation

+

+

+=

)()(

))(1(1

212

xygfxygfgJn

f

30-Mar-1036

+

+

+=

= )()(

))(1(1

212

xygxyggJn

V


37/52

Pressure Solution to Euler equation

),( Pff=

+

= PV

VVP

22

12

30-Mar-10 37

)(int

1

NspofoNumber

P

P

N

i

i

avg

=

=


38/52


39/52

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

30-Mar-10 39



40/52

Validation

Velocity Distribution

Velocity Distribution, Ref @

30-Mar-10 40


41/52

Sample Velocity Distributions

30-Mar-10 41

To achieve Objective 2. Identify & Exploit Engineering Design Parameters to Construct AircraftC fi i i h O i l C bili


42/52

ResultsResultsResultsResults

Configurations with Optimal Capability

30-Mar-1042

ComparativeComparativeComparativeComparative Analysis of 5 AircraftAnalysis of 5 AircraftAnalysis of 5 AircraftAnalysis of 5 Aircraft

ConfigurationsConfigurationsConfigurationsConfigurations

Aircraft Sample Design Parameters


43/52

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

30-Mar-10 43

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


44/52


45/52

#6

#7

30-Mar-10 45

#9

#10


46/52

Object Performance

30-Mar-10 46


47/52

Object Performance

30-Mar-10 47


48/52

Cabin Height Analysis

30-Mar-10 48

Luggage Cabin Analysis


49/52

Luggage Cabin Analysis

30-Mar-10 49


50/52

30-Mar-10 50


51/52

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)

30-Mar-10 51


52/52

ThankThankThankThank YouYouYouYou

30-Mar-1052

Documents

Enhancing Aerodynamic Performance Estimate in small Aircraft Development using Object-Oriented Technique