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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
4.3.4 AIRFOIL WITH FLAP DEFLECTION OF 30 22
4.3.5
AIRFOIL WITH FLAP DEFLECTION OF 35 25
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
AIRFOIL WITH FLAP DEFLECTION OF 15 43
5.4.3 AIRFOIL WITH FLAP DEFLECTION OF 25 46
5.4.4
AIRFOIL WITH FLAP DEFLECTION OF 30 49
5.4.5 AIRFOIL WITH FLAP DEFLECTION OF 35 52
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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
Table 4.3 CLand CDfor airfoil with flap at 25 19
Table 4.4 CLand CDfor airfoil with flap at 30 22
Table 4.5 CLand CDfor airfoil with flap at 35 25
THEORETICAL
Table 5.1 CLand CDfor airfoil without flap 41
Table 5.2 CLand CDfor airfoil with flap at 15 44
Table 5.3 CLand CDfor airfoil with flap at 25 46
Table 5.4 CLand CDfor airfoil with flap at 30 49
Table 5.5 CLand CDfor airfoil with flap at 35 52
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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
ANGLE OF ATTACK CL CD L/D
-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
CDvs Angle of Attack
Cd
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Fig 4.8 L/D vs Angle of Attack
The maximum lift coefficient for airfoil with flap at 15 is 0.839 and the
stalling angle is found to be 11 degrees.
18
0
1
2
3
4
5
6
7
8
-10 -5 0 5 10 15
L/D
Angle of Attack
L/D vs Angle of Attack
L/D
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4.3.3 AIRFOIL WITH FLAP DEFLECTION OF 25
ANGLE OF ATTACK CL CD L/D
-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
Fig 4.10 CDvs 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
CLvs 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
CDvs Angle of Attack
Cd
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Fig 4.11 L/D vs Angle of Attack
The maximum lift coefficient for airfoil with flap at 25 is 1.0138 and the
stalling angle is found to be 10 degrees.
21
0
2
4
6
8
10
12
14
-10 -5 0 5 10 15 20
L/D
Angle of Attack
L/D vs Angle of Attack
L/D
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4.3.4 AIRFOIL WITH FLAP DEFLECTION OF 30
ANGLE OF ATTACK CL CD L/D
-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
Fig 4.12 CLvs Angle of Attack
Fig 4.13 CDvs Angle of Attack
23
0
0.2
0.4
0.6
0.8
1
1.2
-10 -5 0 5 10 15 20
Cl
Angle of Attack
CLvs 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
CLvs Angle of Attack
Cd
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Fig 4.14 L/D vs Angle of Attack
The maximum lift coefficient for airfoil with flap at 30 is 1.075 and the
stalling angle is found to be 9 degrees.
24
0
1
2
3
4
5
6
7
-10 -5 0 5 10 15 20
L/D
Angle of Attack
L/D vs Angle of Attack
L/D
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4.3.5 AIRFOIL WITH FLAP DEFLECTION OF 35
ANGLE OF ATTACK CL CD L/D
-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
25
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GRAPH
Fig 4.15 CL vs Angle ofAttack
Fig 4.16 CD vs Angle of Attack
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
CLvs 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
CDvs Angle of Attack
Cd
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Fig 4.17 L/D vs Angle of Attack
The maximum lift coefficient for airfoil with flap at 35 is 1.0053 and the
stalling angle is found to be 7 degrees.
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 vs 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
ANGLE OF ATTACK CL CD L/D
-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
Fig 5.4 CLvs Angle of Attack
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
CLvs Angle of Attack
Cl
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Fig 5.5 CDvs Angle of Attack
Fig 5.6 L/D vs Angle of Attack
The maximum lift coefficient for airfoil without flap is 1.346 and the
stalling angle is found to be 23 degrees.
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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
CDvs 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 vs Angle of Attack
L/D
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5.4.2 AIRFOIL WITH FLAP DEFLECTION OF 15
43
ANGLE OF ATTACK CL CD L/D
-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
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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
Fig 5.7 CLvs Angle of Attack
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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
CLvs Angle of Attack
Cl
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Fig 5.8 CDvs Angle of Attack
Fig 5.9 L/D vs Angle of Attack
The maximum lift coefficient for airfoil with flap at 15 is 1.4709 and the
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
CDvs 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 vs Angle of Attack
L/D
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5.4.3 AIRFOIL WITH FLAP DEFLECTION OF 25
ANGLE OF ATTACK CL CD L/D
-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
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GRAPH
Fig 5.10 CLvs Angle of Attack
Fig 5.11 CDvs Angle of Attack
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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
CLvs 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
CDvs Angle of Attack
Cd
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Fig 5.12 L/D vs Angle of Attack
The maximum lift coefficient for airfoil with flap at 30 is 1.7418 and the
stalling angle is found to be 14 degrees
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0
1
2
3
4
5
6
7
8
9
-10 -5 0 5 10 15 20
L/D
Angle of Attack
L/D vs Angle of Attack
L/D
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5.4.4 AIRFOIL WITH FLAP DEFLECTION OF 30
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
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GRAPH
Fig 5.13 CL vs Angle of Attack
Fig 5.14 CDvs 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
CLvs 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
CDvs Angle of Attack
Cd
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Fig 5.15 L/D vs Angle of Attack
The maximum lift coefficient for airfoil with flap at 30 is 1.7472 and the
stalling angle is found to be 10 degrees.
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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 vs Angle of Attack
L/D
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5.4.5 AIRFOIL WITH FLAP DEFLECTION OF 35
ANGLE OF ATTACK CL CD L/D
-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
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GRAPH
Fig 5.16 CLvs Angle of Attack
Fig 5.17 CD vs Angle of Attack
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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
CLvs 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
CDvs Angle of Attack
Cd
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Fig 5.18 L/D vs Angle of Attack
The maximum lift coefficient for airfoil with flap at 35 is 1.8475 and the
stalling angle is found to be 8 degrees
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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 vs Angle of Attack
L/D
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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
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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
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2
4
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6
8
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10
12
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5.5.2 PRESSURE CONTOURS FOR 30 AND 35
ANGLE
OF
ATTACK
30 35
0
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2
4
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6
8
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10
12
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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.
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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.