klind fogle man aerofoil report

8/10/2019 klind fogle man aerofoil report

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EFFECT OF TRAILING EDGE FLAP ON THE LIFT

AND DRAG OF KLINE FOGLEMAN AIRFOIL

A PROJECT REPORT

Submitted by

PREM ANAND.T.P

RAJAVANNIAN.R

SREEKANTH.A

in partial fulf il lment for the award of the degree

of

BACHELOR OF ENGINEERING

in

AERONAUTICAL ENGINEERING

RAJALAKSHMI ENGINEERING COLLEGE

ANNA UNIVERSITY:CHENNAI -600025

APRIL 2011


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BONAFIDE CERTIFICATE

Certified that this project report EFFECT OF TRAILING EDGE

FLAP ON THE LIFT AND DRAG OF KLINE FOGLEMANAIRFOIL is the bonafide work of

PREM ANAND.T.P (21107101034)

RAJAVANNIAN.R (21107101037)

SREEKANTH.A (21107101049)

During the year 2010-2011 in partial fulfillment for the award of the

BACHELOR OF ENGINEERING degree in AERONAUTICAL

ENGINEERING at RAJALAKSHMI ENGINEERING COLLEGE.

Mr.Yogesh Kumar Sinha Mr.Yogesh Kumar Sinha

Associate Professor, Head of the Department,

Dept. of Aeronautical Engineering, Dept. of Aeronautical Engineering,

Rajalakshmi Engineering college, Rajalakshmi Engineering College,

Rajalakshmi Nagar, Rajalakshmi Nagar,

Chennai 602105. Chennai 602105.


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CERTIFICATE OF EVALUATION

COLLEGE NAME: RAJALAKSHMI ENGINEERING COLLEGE

BRANCH : AERONAUTICAL ENGINEERING

SEMESTER : 8th

SEMESTER

S.NO NAME OF THE

STUDENT

TITLE OF THE

PROJECT

NAME OF THE

GUIDE

1

2

3

PREM ANAND T.P

(21107101034)

RAJAVANNIYAN.R

(21107101037)

SREEKANTH.A

(21107101049)

EFFECT OF

TRAILING EDGE

FLAP ON THE

LIFT AND DRAG

OF KLINE

FOGLEMAN

AIRFOIL

Mr.YOGESH

KUMAR SINHA

The report of the project are submitted by above students in partial

fulfillment for the award of Bachelor of Engineering in Aeronautical

Engineering of Anna University were enabled and confirmed to be

the report work done by the above students and then evaluated.

(INTERNAL EXAMINER) (EXTERNAL EXAMINER)


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ACKNOWLEDGEMENT

We are sincerely grateful to our guide, Mr.Yogesh kumar sinha for guiding us

throughout the course of our project work.

We also thank other faculty members of the Department of Aeronautical

Engineering who have helped us during the review of the project.

PREM ANAND.T.P

RAJAVANNIAN.R

SREEKANTH.A


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TABLE OF CONTENTS

CHAPTER NO TITLE PAGE NO

ABSTRACT i

LIST OF FIGURES ii

LIST OF TABLES iii

1. INTRODUCTION 1

1.1

INTRODUCTION 1

1.2NEED FOR THE PRESENT STUDY 1

1.3

PRESENT STUDY 2

2.

LITERATURE SURVEY 3

2.1 INTRODUCTION 3

2.2 KLINE FOGLENAN AIRFOIL 4

2.3 HIGH LIFT DEVICES 5

2.4

TRAILING EDGE FLAP 6

3. KFm AIRFOIL MODEL FABRICATION 8

3.1

INTRODUCTION 8

3.2 MODEL FABRICATION 8


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4. WIND TUNNEL AND EXPERIMENT 11

4.1

LOW SPEED WIND TUNNEL 11

4.2

EXPERIMENTAL SETUP 12

4.3 EXPERIMENTAL RESULTS 13

4.3.1 AIRFOIL WITHOUT FLAP 13

4.3.2 AIRFOIL WITH FLAP DEFLECTION OF 15 16

4.3.3

AIRFOIL WITH FLAP DEFLECTION OF 25 19


4.3.5


5. ANALYSIS SOFTWARES AND RESULTS 29

5.1

CATIA 29

5.2 GAMBIT 31

5.2.1

PREARING THE MODEL 31

5.2.2

MESHING 32

5.3 FLUENT ANALYSIS 33

5.4

FLUENT ANALYSIS RESULTS 40

5.4.1 AIRFOIL WITHOUT FLAP 41

5.4.2



5.4.4




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5.5

QUALITATIVE RESULTS 56

5.5.1

PRESSURE CONTOURS FOR 15AND 25 56

5.5.2 PRESSURE CONTOURS FOR 30AND 35 60

6.

CONCLUSION 64

7.

REFERENCES 65


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ABSTRACT

Present study focuses on the effect of a trailing edge Plain flap on

the lift and drag of Kline Fogleman airfoil. Experiments were conducted inthe Low Speed Wind tunnel with the trailing edge plain flap at different

deflection angles degrees with varying angle-of-attack. The test is conducted

at a velocity of 30m/s. Pressure contours had significant variation for the

deflection of flap as against without flap has been observed from the

results. The deflection of flap shows that the area under the pressure

contour diagrams are larger than without deflecting the flap which indicates

the corresponding higher lift coefficient. Result shows that the deflection of

flap increases the maximum lift coefficient by 50% and stalling angle got

reduced for the tested flow velocity. The meshing of the KFm model had

been done in GAMBIT software and it is imported to FLUENT 6.2.16

software. The simulations in FLUENT yielded the best correlations to the

experimental data.

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LIST OF FIGURES

FIGURE NO TITLE PAGE.NO

Figure 2.1 CLvs AOA curve for all types of flaps 5

Figure 2.2 Plain flap 7

Figure 2.3 Effect of trailing edge flap on stalling angle 7

Figure 3.1 3d view of airfoil 9

Figure 3.2 Photograph of model 10

Figure 4.1 Low speed wind tunnel 11

Figure 4.2 Model setup in the wind tunnel 12

Figure 5.1 Kline fogleman airfoil in CATIA workbench 30

Figure 5.2 Kline fogleman airfoil with flap in CATIA 30

Workbench

Figure 5.3 Meshing around airfoil 32

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LIST OF TABLES

TABLE NO TITLE PAGE NO

EXPERIMENTAL

Table 4.1 CLand CDfor airfoil without flap 13

Table 4.2 CLand CDfor airfoil with flap at 15 16




THEORETICAL

Table 5.1 CLand CDfor airfoil without flap 41





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CHAPTER-1

INTRODUCTION

1.1

INTRODUCTION

Ever since the beginning of first flight by mankind there has been a

constant endeavour to enhance the wing performance by various means such

as improvement and refinement of wing design, addition of auxiliary lifting

devices, using light weight materials, Laminar Flow Control (LFC) and other

flow optimization methods. The usage of auxiliary lifting surfaces such as

flaps, slats, leading edge slots has gained prominence in the designing of

wings for aircrafts nowadays along with other developments in wing

designs. The usage of a trailing edge plain flap in a wing enables the wing

to operate at higher angles of attack in situations like landing and take-off

without losing lift.

1.2

NEED FOR PRESENT STUDY

The usage of high lift devices has become a common phenomenon

in the design and developments of aircrafts operating at high angles of

attack. The traling edge devices such as plain, split flap etc are used at

high angle of attack to increase the CLmax are vital during landing and take-

off for aircrafts which essentially operate at shorter runways and it reduces

the stalling speed of aircraft which means that aircraft can fly safely at

lower speeds. The effect of trailing edge device at different deflection

angles on the airfoil performance needs to be investigated and the angle-of-

deflection of flap needs to be determined so as not to increase the drag and

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also to achieve highest possible CLmax without stalling. Besides the stand

alone configuration of Kline Fogleman airfoil with trailing edge flap has not

been reported in the literature.

1.3

PRESENT STUDY

The present study focuses on the effect of a trailing edge plain flap

on the lift and drag of a Kline Fogleman airfoil. The focus is on finding

out the lift coefficient and study the effect of plain flap on the lift and

drag of an airfoil by deflecting the flap to different deflection angles at a

flow velocity of 30m/s and varying angles of attack. The study aims at

finding the maximum lift coefficient for different angles of deflection of

flap and thereby finding the maximum lift coefficient at the stalling angle of

attack of Kline Fogleman airfoil.

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CHAPTER-2

LITREATURE SURVEY

2.1 INTRODUCTION

The investigation of Kline Fogleman airfoil performance was first

performed by NASA in the year 1960 by Richard KLINE & Floyd

FOGLEMAN wherein he established the CL vs angle of attack (alpha) curve

for various velocities. It was established that the maximum CL is found to

be 0.742 at angle of 9 degrees and stalling angle is 9 degrees.

The possibility of using the auxiliary lifting device is first

demonstrated in late 1919 by NACA where they tested various

configurations of leading and trailing edge devices. Handley page in 1920

explored the possibility of using a trailing edge plain flap with a NACA

0009 airfoil to maximize its lift at lower angles of attack. It was found out

that the lift coefficient value increased by 40% at lower angle of attack and

the stalling angle decreased.

The tandem usage of leading and trailing edge lifting device was

performed on NACA 23012 airfoil with leading edge slot and plain flap in

1920s by Wensinger. He found in the case of cambered airfoil not only

stalling angle is increased but also lift curve slope is increased.

The flaps were found to be ineffective at higher angles of attack due

to increase in drag and the stalling of airfoil was found out from variouswind tunnel investigations of different airfoils forced the deflection of flaps

at lower angle of attack to achieve higher lift coefficient at landing and

take-off conditions.

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The airfoil and the location of flaps are fabricated on the basis of co-

ordinates given for Kline Fogleman airfoil in the website Airfoil

Investigation Database and Theory of Wing Section by Abbot and Von

Doenhoff.

The flow visualization pattern around the airfoil is found out by

placing the airfoil in the mid-section of Hele Shaw apparatus and the

streamline pattern is observed.

2.2 KLINE FOGLEMAN AIRFOIL

The Kline Fogleman airfoils are airfoil shapes for aircraft wings

developed by Richard KLINE and Floyd FOGLEMAN. The Kline Fogleman

airfoil is an airfoil design with single or multiple steps induced along the

length of the wing. Primarily located on the top or bottom side of the wing

to assist with greater lift and stability during flight. The KFm uses the

concept of vortex, which attaches itself to the airfoil behind the step and

becomes part of the airfoil.

CHARACTERISTICS

1.

The KFm airfoils are thicker for the first 50% of the chord which

produces more lift.

2. It is thinner for the rest of the chord portion for travelling faster.

3.

It has a much greater range for center of gravity.

4.

It requires zero degree reflex which reduces the drag.

5.

It is capable of flying without stabilizers and rudders.

6.

It requires no dihedral for stability.

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2.3 HIGH LIFT DEVICES

The study of high lift devices for enhancing the performance of wings

have been on the fore since the beginning of aviation. The type of

operation for which an airplane is intended has a very important bearing on

the selection of the shape and design of the wing for that airplane. Slots,

slats, spoilers, speed brakes and flaps are additions to the wing that perform

a variety of functions related to control of the boundary layer, increase of

the plan form area and reduction of aircraft velocity during landing and

stopping conditions.

Fig 2.1 CLvs AOA for all types of flaps

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2.4 TRAILING EDGE FLAP

Trailing edge flaps are movable aerodynamic devices used on

airplanes. Flaps were first developed by Handley-Page in 1920.

Flaps are hinged surfaces on the trailing edge of the wings of the

aircraft. Extending the flaps increases the camber of the wing airfoil, thus

raising the maximum lift coefficient. This increase in maximum lift

coefficient allows the aircraft to generate a given amount of lift at lower

speeds. Therefore, extending the flaps reduces the stalling speed of aircraft.

Extending flaps also increases the drag. This happens of the higher

induced drag caused by the distorted spanwise lift distribution on the wing

with flap extended. This can be beneficial in the approach and landing phase

because it helps to slow the aircraft

Depending on the type of aircraft, the flaps may be partially extended

during take-off and it may be fully extended during landing to give the

aircraft a lower stalling speed allowing the aircraft to land in a shorter

distance.

Plain flaps when fully deflected increases the wing camber and the

wing area which results in increased lift and drag at a given angle of

attack and increases the maximum CL.

Wings with the flaps deflected usually stall at lower angle of attack

than wings without flaps. This is due to the fact that the pressure gradientsat the CLmax for the cases are roughly equal.

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Deflection of trailing edge flaps increase the lift at constant geometric

angle of attack, it will also move the CP rearwards and increase both

parasite and induced drag.

Fig 2.2 Plain Flap

Fig 2.3 Effect of trailing edge flaps on stalling angle

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CHAPTER-3

KLINE FOGLEMAN AEROFOIL MODEL FABRICATION

3.1 INTRODUCTION

The NACA airfoil are generally preferred because the symmetric

airfoil works better in small angle of attack and the cambered airfoil works

better in higher angle of attack. But the usage of Kline Fogleman airfoil

compensates for this disadvantage and increases the maximium lift

coefficient when it is used with flaps at the trailing edge. The Kline

Fogleman airfoil has been used only in a paper airplane. The airfoil was

chosen with trailing edge flap for its inherent advantages and ease of

fabrication.

3.2 FABRICATION

The Kline Fogleman airfoil model with trailing edge flap whose

deflection angle can be varied has been fabricated from Balsa wood with

the following specifications.

Chord : 10 cm

Span : 25 cm

Flap location : 20 % of chord (from trailing edge)

Flap : one trailing edge plain flap

Flap deflection : Four different deflection angles (15,25,30 and 35)

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The model is designed in CATIA using the co-ordinates for the Kline

Fogleman airfoil with the trailing edge plain flap being located at 20% of

the chord whose deflection angle can be varied. The angle of attack of

model can be changed by raising and lowering the rod inside the pipe fitted

with screw. A 3d view of model is given in the following figure 3.1.

Fig 3.1 3d view of airfoil

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The fabricated model is shown below the figure

Fig 3.2 Photograph of model

The fabricated model has the location of flap positioned at 20% of

chord from the trailing edge whose deflection angle can be varied and the

AOA of the airfoil model can be changed from -8 degrees to +22 degrees.

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CHAPTER-4

WIND TUNNEL AND EXPERIMENT RESULT

4.1 LOW SPEED WIND TUNNEL

Experiments are performed in the subsonic wind tunnel of test section

Size 30cm length * 30cm width * 30cm height with a maximum speed of

50m/s at a drive speed of 720 rpm as shown in figure 4.1 Wind tunnel is

fitted with a drive panel incorporating various accessories for the speed

control of the fan using the speed control unit, and it also consists of lift,

drag, side force and velocity indicators.

Fig 4.1 Low speed wind tunnel

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4.2 EXPERIMENTAL SETUP

The investigation of the effect of the trailing edge flap on the lift and

drag of the Kline Fogleman airfoil is computed for the configuration of the

flap deflection angles of 15, 25, 30 and 35 degrees. For these four

configurations lift and drag are found out in a flow velocity of 30m/s. The

experiments were repeated for different angles of attack (-8 degrees to +20

degrees).

Fig 4.2 Model setup in the wind tunnel

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4.3 EXPERIMENTAL RESULTS

4.3.1 AIRFOIL WITHOUT FLAP

ANGLE OF ATTACK CL CD L/D

-4 0 0.008341 0

-2 0.0960 0.00776 1.2616

0 0.2917 0.0070 4.226

2 0.3486 0.0073 4.851

4 0.4766 0.0079 6.0909

6 0.5763 0.0088 6.6393

7 0.6332 0.0097 6.5925

7.5 0.6154 0.0101 6.1785

8 0.5940 0.0126 4.7715

10 0.5514 0.0177 3.1632

Table 4.1 CL and CD for airfoil without flap at different angles

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GRAPH

Fig 4.3 CLvs Angle of attack

Fig 4.4 CD vs Angle of Attack

14

-0.2

-0.1

0

0.1

0.2

0.3

0.4

0.5

0.6

0.7

-14 -12 -10 -8 -6 -4 -2 0 2 4 6 8 10 12

Cl

Angle of Attack

CLvs Angle of Attack

Cl

0

0.002

0.004

0.006

0.008

0.01

0.012

0.0140.016

0.018

0.02

-15 -10 -5 0 5 10 15

Cd

Angle of Attack

CDvs Angle of Attack

Cd


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Fig 4.5 L/D vs Angle of Attack

The maximum lift coefficient for airfoil without flap is 0.6332 and the

stalling angle is found to be 13 degrees.

15

-2

-1

0

1

2

3

4

5

6

7

8

-15 -10 -5 0 5 10 15

L/D

Angle of Attack

L/D vs Angle of Attack

L/D


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4.3.2 AIRFOIL WITH FLAP DEFLECTION OF 15


-4 0.149 0.0953 1.567

-2 0.263 0.0853 3.083

0 0.430 0.0889 4.84

2 0.508 0.0924 5.5

4 0.584 0.1003 5.822

6 0.747 0.1152 6.686

8 0.839 0.1351 6.210

9 0.784 0.1387 5.656

10 0.777 0.1479 5.256

Table 4.2 CL and CDfor airfoil with flap at 15

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GRAPH

Fig 4.6 CL vs Angle of Attack

Fig 4.7 CDvs Angle of Attack

17

00.1

0.2

0.3

0.4

0.5

0.6

0.7

0.8

0.9

-10 -5 0 5 10 15

Cl

Angle of Attack

CL

vs Angle of Attack

Cl

0

0.02

0.04

0.06

0.08

0.1

0.12

0.14

0.16

0.18

-10 -5 0 5 10 15

Cd

Angle of Attack


Cd


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The maximum lift coefficient for airfoil with flap at 15 is 0.839 and the


18

0

1

2

3

4

5

6

7

8

-10 -5 0 5 10 15

L/D

Angle of Attack


L/D


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-6 0.3813 0.0544 7.006

-4 0.3984 0.0562 7.088

-2 0.4695 0.0569 8.25

0 0.5442 0,0586 9.272

2 0.7385 0.0718 10.277

4 0.8544 0.0758 11.660

6 0.9597 0.0853 11.241

8 0.9782 0.0928 10.536

10 0.9986 0.0946 10.563

11 1.0138 0.0964 10.516

12 0.9676 0.1006 9.611

Table 4.3 CL and CDfor airfoil with flap at 25

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GRAPH

Fig 4.9 CLvs Angle of Attack


20

0

0.2

0.4

0.6

0.8

1

1.2

-10 -5 0 5 10 15 20

Cl

Angle of Attack


Cl

0

0.02

0.04

0.06

0.08

0.1

0.12

0.14

-10 -5 0 5 10 15 20

Cd

Angle of Attack


Cd


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21

0

2

4

6

8

10

12

14

-10 -5 0 5 10 15 20

L/D

Angle of Attack


L/D


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-6 0.5165 0.0114 4.594

-4 0.5485 0.0116 4.818

-2 0.6054 0.0119 5.157

0 0.7449 0.0146 5.183

2 0.9156 0.0149 6.247

4 0.9939 0.0160 6.321

6 1.011 0.0182 5.661

8 1.059 0.0196 5.494

10 1.075 0.0209 5.231

12 1.055 0.0224 4.753

14 0.988 0.0240 4.196

Table 4.4 CLand CDfor airfoil with flap deflection at 30

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GRAPH



23

0

0.2

0.4

0.6

0.8

1

1.2

-10 -5 0 5 10 15 20

Cl

Angle of Attack


Cl

0

0.005

0.01

0.015

0.02

0.025

0.03

-10 -5 0 5 10 15 20

Cd

Angle of Attack


Cd


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24

0

1

2

3

4

5

6

7

-10 -5 0 5 10 15 20

L/D

Angle of Attack


L/D


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-6 0.7506 0.2276 3.296

-4 0.8039 0.2241 3.587

-2 0.8445 0.2312 3.652

0 0.9050 0.2419 3.741

2 0.9391 0.2454 3.826

4 1.004 0.2575 3.900

6 1.0014 0.2646 3.838

8 1.0053 0.2717 3.874

10 0.9569 0.2774 3.448

12 0.7648 0.2860 2.674

Table 4.5 CL and CDfor airfoil at flap at 35

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GRAPH

Fig 4.15 CL vs Angle ofAttack


26

0

0.2

0.4

0.6

0.8

1

1.2

-8 -6 -4 -2 0 2 4 6 8 10 12 14

Cl

Angle of Attack


Cl

0

0.05

0.1

0.15

0.2

0.25

0.3

0.35

-8 -6 -4 -2 0 2 4 6 8 10 12 14

Cd

Angle of Attack


Cd


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27

0

0.5

1

1.5

2

2.5

3

3.5

4

4.5

-10 -5 0 5 10 15

L/D

Angle of Attack


L/D


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COMPARISON OF CL FOR ALL DEFLECTION ANGLES

Fig 4.18 Comparison of CL

28

-0.2

0

0.2

0.4

0.6

0.8

1

1.2

-15 -10 -5 0 5 10 15 20

Cl

Angle of Attack

Comparison of CL

WITHOUT FLAP

WF15

WF 25

WF 30

WF 35


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CHAPTER-5

ANALYSIS SOFTWARES AND RESULTS

5.1 ANSYS

CATIA (Computer Aided Three-dimensional Interactive Application) is a

multi-platform CAD/CAM/CAE commercial software suite developed by the

French company Dassault Systems and marketed worldwide by IBM. Written in

the C++ programming language, CATIA is the cornerstone of the Dassault

Systems product lifecycle management software suite.

The software was created in the late 1970s and early 1980s to develop

Dassault's Mirage fighter jet, and then was adopted in the aerospace, automotive,

shipbuilding, and other industries.

CATIA competes in the CAD/CAM/CAE market with Siemens NX,

Pro/ENGINEER, Autodesk Inventor and Solid Edge.

Commonly referred to as a 3D Product Lifecycle Management software

suite, CATIA supports multiple stages of product development (CAx), from

conceptualization, design (CAD), manufacturing (CAM), and engineering (CAE).

CATIA can be customized via application programming interfaces (API).

V4 can be adapted in the Fortran and C programming languages under an API

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called CAA. V5 can be adapted via the Visual Basic and C++ programming

languages, an API called CAA2 or CAA V5 that is a component object model

(COM)-like interface.

CATIA is widely used throughout the engineering industry, especially in the

automotive and aerospace sectors. CATIA V4, CATIA V5, Pro/ENGINEER, NX

(formerly Unigraphics), and SolidWorks are the dominant systems

Fig 5.1 Kline Fogleman airfoil in CATIA workbench

Fig 5.2 Kline Fogleman airfoil with plain flap in CATIA workbench

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5.2 GAMBIT

5.2.1 PREPARING THE MODEL

The model is prepared in the GAMBIT software by importing the co-

ordinates as a dat file and the geometry is created around the model to

make it a valid CFD model.

An important thing in this is creating the mesh surrounding the object.

This needs to be extended in all the directions to get the physical properties

of the surrounding fluid. The mesh and the edges must also be grouped in

order to set the necessary boundary conditions.

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5.2.2 MESHING

An environment consisting of 2 squares and 1 semicircle surrounds the

KFm airfoil. The mesh is constructed to be very fine at regions close to the

airfoil. For this airfoil a structured quadratic mesh was used. The grid size

of the mesh is given as 0.20.

Fig 5.3 Meshing around the airfoil

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5.3 FLUENT ANALYSIS

Start Fluent2D and load the mesh file as follows:

FileReadCaseGrid fin.

GridCheck.

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DefineModelsSolverSegregated, 3D, Absolute, Cell-Based, Implicit,

Steady, Superficial Velocity.

ModelsEnergy equationOFF.

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ModelsViscous modelInvisid

Models-Materials-Create/Change-Density-Constant(1.2256kg/m3)-Close

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Operating Conditions-Operating Pressure-Constant(101325 pa)

Boundary Conditions-Velocity Specification Method-Components-Set x and

y values

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Solve-Control-Solution-PRESTO under Pressure-Second Order Upwind

under Momentum

Solve-Initialize-Initialize-Inlet under Compute from-ok

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Solve-Monitors-Residual-Set Convergence value-Ok

Solve-Monitors-Force-Give values for lift and drag-Apply

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Report-Reference Values-Compute from-Inlet-Ok

Solve-Iterate.

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5.4 FLUENT ANALYSIS RESULTS

5.4.1 AIRFOIL WITHOUT FLAP

40


-8 0.095 0.0314 3.025478

-6 0.156 0.0354 4.40678

-4 0.21 0.037 5.675676

-2 0.287 0.0388 7.396907

0 0.35129 0.039927 8.798307

2 0.4482 0.0477 9.291405

4 0.55 0.0588 9.353741

6 0.6088 0.0735 8.282993

8 0.7051 0.0932 7.565451

10 0.891 0.11533 7.023324

12 0.81 0.1442 6.178918

14 0.9769 0.1751 5.579098

16 1.056 0.20869 5.060137

18 1.1296 0.2446 4.618152

20 1.2037 0.2829 4.25486


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22 1.281 0.33 3.881818

23 1.346 0.366 3.677596

24 1.329 0.42 3.164286

25 1.296 0.452 2.867257

26 1.25 0.47 2.659574

27 1.222 0.48 2.545833

Table 5.1 CL and CD for airfoil without flap

GRAPH


41

0

0.2

0.4

0.6

0.8

1

1.2

1.4

1.6

-10 -8 -6 -4 -2 0 2 4 6 8 10 12 14 16 18 20 22 24 26 28 30

Cl

Angle of Attack


Cl


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The maximum lift coefficient for airfoil without flap is 1.346 and the


42

0

0.1

0.2

0.3

0.4

0.5

0.6

-10 -5 0 5 10 15 20 25 30

Cd

Angle of Attack


Cd

0

1

2

3

4

5

6

7

8

9

10

-10 -5 0 5 10 15 20 25 30

L/D

Angle of Attack


L/D


55/77


43


-8 0.5143 0.06 8.571667

-6 0.6521 0.0641 10.17317

-4 0.8682 0.0789 11.0038

-2 0.9654 0.0845 11.42485

0 1.1009 0.0895 12.30056

2 1.1841 0.09355 12.6574

4 1.2585 0.12095 10.40513

6 1.3228 0.15087 8.767813

8 1.3753 0.18248 7.536716

10 1.4147 0.21504 6.578776

12 1.441 0.247 5.834008

14 1.4574 0.2951 4.938665

16 1.4609 0.3111 4.695918

18 1.4709 0.3446 4.268427

19 1.458 0.3468 4.204152


56/77

20 1.4125 0.3782 3.734796

21 1.3782 0.3923 3.513128

22 1.3502 0.4235 3.188194

Table 5.2 CL and CD for airfoil with flap at 15

GRAPH


44

0

0.2

0.4

0.6

0.8

1

1.2

1.4

1.6

-10 -5 0 5 10 15 20 25

Cl

Angle of Attack


Cl


57/77




stalling angle is found to be 18 degrees

45

0

0.05

0.1

0.15

0.2

0.25

0.3

0.35

0.4

0.45

-10 -5 0 5 10 15 20 25

Cd

Angle of Attack


Cd

0

2

4

6

8

10

12

14

-10 -5 0 5 10 15 20 25

L/D

Angle of Attack


L/D


58/77



-6 0.956 0.1345 7.107807

-4 1.056 1.456 7.252747

-2 1.258 0.1678 7.49702

0 1.4512 0.1799 8.066

2 1.5165 0.1822 8.3232

4 1.5735 0.2143 7.342

6 1.6234 0.2524 6.431

8 1.6745 0.2915 5.743

10 1.712 0.3297 5.192

12 1.7338 0.3675 4.717

14 1.7418 0.4039 4.312

16 1.7418 0.4417 3.944

17 1.7423 0.4601 3.782

Table 5.3 CLand CD for airfoil with flap at 25

46


59/77

GRAPH



47

0

0.2

0.4

0.6

0.8

1

1.2

1.4

1.6

1.8

2

-10 -5 0 5 10 15 20

Cl

Angle of Attack


Cl

0

0.05

0.1

0.15

0.2

0.25

0.3

0.35

0.4

0.45

0.5

-10 -5 0 5 10 15 20

Cd

Angle of Attack


Cd


60/77




48

0

1

2

3

4

5

6

7

8

9

-10 -5 0 5 10 15 20

L/D

Angle of Attack


L/D


61/77


ANGLE OF

ATTACK

CL CD L/D

-6 0.9651 0.1043 9.253116

-4 1.0954 0.1186 9.236088

-2 1.268 0.1296 9.783951

0 1.6195 0.1479 10.94997

2 1.6605 0.183 9.07377

4 1.6955 0.2195 7.724374

6 1.7267 0.2578 6.697828

8 1.739 0.2966 5.863115

10 1.7472 0.332 5.262651

12 1.7157 0.361 4.752632

Table 5.4 CLand CD for airfoil with at 30

49


62/77

GRAPH

Fig 5.13 CL vs Angle of Attack


50

0

0.2

0.4

0.6

0.8

1

1.2

1.4

1.6

1.8

2

-8 -6 -4 -2 0 2 4 6 8 10 12 14

Cl

Angle of Attack


Cl

0

0.05

0.1

0.15

0.2

0.25

0.3

0.35

0.4

-8 -6 -4 -2 0 2 4 6 8 10 12 14

Cd

Angle of Attack


Cd


63/77




51

0

2

4

6

8

10

12

-8 -6 -4 -2 0 2 4 6 8 10 12 14

L/D

Angle of Attack


L/D


64/77



-6 1.054 0.1329 7.930

-4 1.2143 0.1471 8.254

-2 1.456 0.1598 9.111

0 1.7234 0.1691 10.1916

2 1.7726 0.2066 8.579

4 1.8128 0.2433 7.450

6 1.8367 0.2853 6.437

8 1.8475 0.3248 5.688

9 1.8461 0.329 5.445

10 1.8432 0.3435 5.365

Table 5.5 CLand CDfor airfoil with flap at 35

52


65/77

GRAPH



53

0

0.2

0.4

0.6

0.8

1

1.2

1.4

1.6

1.8

2

-8 -6 -4 -2 0 2 4 6 8 10 12

Cl

Angle of Attack


Cl

0

0.05

0.1

0.15

0.2

0.25

0.3

0.35

0.4

-8 -6 -4 -2 0 2 4 6 8 10 12

Cd

Angle of Attack


Cd


66/77




54

0

2

4

6

8

10

12

-8 -6 -4 -2 0 2 4 6 8 10 12

L/D

Angle of Attack


L/D


67/77

COMPARISON OF CL FOR ALL DEFLECTION ANGLES

55

0

0.2

0.4

0.6

0.8

1

1.2

1.4

1.6

1.8

2

-10 -5 0 5 10 15 20 25

Cl

Angle of Attack

Comparison of CL

WF15

without flap

WF25

WF 30

WF35


68/77

5.5 QUALITATIVE RESULTS

This table shows the different contour outputs from the simulations.

Inviscid models of airfoil with flap at different deflection angles are

compared in separate column

5.5.1 PRESSURE CONTOURS OF 15 AND 25

ANGLE

OF

ATTACK

15 25

0

56


69/77

2

4

57


70/77

6

8

58


71/77

10

12

59


72/77

5.5.2 PRESSURE CONTOURS FOR 30 AND 35

ANGLE

OF

ATTACK

30 35

0

60


73/77

2

4

61


74/77

6

8

62


75/77

10

12

63


76/77

CHAPTER 6

CONCLUSION

The usage of flaps at different deflection angles reveal that the

flap deflection does alter the lift and drag on the airfoil and compared to

the 15, 25 and 30 deflection, the 35 deflection does not contribute favorably

on the stalling angle due to the flow separation and resulting in excessive

drag. The optimal deflection of flap is found to be within the 15, 25 flap

deflection which yields maximum lift coefficient at lower angles of attack

and higher stalling angle compared to the other deflection angles in the

experiment.

The result reveals that the deflection of flap increases the

maximum lift coefficient by nearly 50% and reduced the stalling angle by 4

degrees respectively.

64


77/77

CHAPTER 7

REFERENCES

1.

Aerodynamic performance of an airfoil with step-induced vortex for

lift augmentation by Fathi Finaish, journal of Aerospace engineering,

vol no.11, 1998.

2.

Kline-Fogleman airfoil comparison study for scratch-bulit foam

airslanes by Rich Thompson, Feb 15, 2008.

3. Design of the low-speed NLF(1)-0414F and the high-speed

HSNLF(1)-0213 airfoils with high-lift systems by J.K.Viken, 1983.

4.

Introduction to flight by J.D.Anderson, Edition 3. McGraw-Hillpublishers, 1989.

5. Aerodynamics, Flight mechanics and Stability by Mac cormick,

McGraw-Hill publishers, 1998.

6.

Fundamentals of Aerodynamics by John.D.Anderson, Edition 4.

McGraw-Hill, 2007.

klind fogle man aerofoil report

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