FYP Thesis

DESIGN AND CFD ANALYSIS OF FORMULA 1
FRONT WING
Final Year Project Thesis
_____________________________________
Department of Mechanical Engineering KSK Campus
University of Engineering and Technology, Lahore

Group Members
HAZIQ ABDUL JABBAR 2011-ME-301
MUHAMMAD ABDULLAH 2011-ME-316
WAQAS SIDDIQ 2011-ME-322
Final Year Student B.Sc. Mechanical Engineering
University of Engineering and Technology Lahore
A thesis submitted in partial fulfillment of the requirements for the degree of
B.Sc. Mechanical Engineering
Project Supervisor:
Ms. Anam Anwar
Lecturer at
University of Engineering and Technology Lahore
External Advisor Signature: ____________________________________
Thesis Supervisor Signature: ___________________________________

DESIGN AND CFD ANALYSIS OF FORMULA 1
FRONT WING
Internal Examiner
Name: Ms. ANAM ANWAR
Signature:
Dated:
External Examiner
Name:
Signature:
Dated:
Department of Mechanical Engineering - KSK Campus
Universityof Engineering and TechnologyLahore -Pakistan

DEDICATION
To Allah the Almighty
&
To Our Parents and Teachers

i
Table of Contents
Abstract................................................................................................................................................ iii
Acknowledgement................................................................................................................................ iv
List of figures ........................................................................................................................................ v
List of tables......................................................................................................................................... vi
Lists of symbols................................................................................................................................... vii
1 CHAPTER I - LITERATURE REVIEW .................................................................................... 1
1.1 Introduction................................................................................................................................. 1
1.2 Motivation .................................................................................................................................. 1
1.3 Front Wing.................................................................................................................................. 1
1.4 Aerofoil ...................................................................................................................................... 2
Aerofoil Nomenclature........................................................................................................ 2
1.5 NACA-4412................................................................................................................................ 3
NACA Nomenclature .......................................................................................................... 3
1.6 Wing Theory............................................................................................................................... 3
Aerodynamic Force ............................................................................................................. 3
Drag .................................................................................................................................... 4
Lift Induced Drag................................................................................................................ 4
Parasitic Drag...................................................................................................................... 4
Downforce........................................................................................................................... 5
Angle of Attack................................................................................................................... 5
Moment............................................................................................................................... 5
1.7 Historical Review and Evolution of Wing ................................................................................... 5
1.8 Description of the Flow around the Front Wing........................................................................... 8
Different Aspects of Flow over the Wing Span.................................................................... 8
Flow at the Wing Tips ......................................................................................................... 8
1.9 Some Advanced Studies Conducted In Recent Times.................................................................. 9
Boundary Layer Separation Delay Using Flaps and Slats..................................................... 9
Overcome Stall and Separation Losses Using Jets................................................................ 9
Wingtip Vortices ............................................................................................................... 10
Vortex Classification......................................................................................................... 11
Effect of Aspect Ratio on Vortices..................................................................................... 11
2 CHAPTER II – METHODOLOGY AND MODELING........................................................... 13
2.1 Design and Specification of the Model...................................................................................... 13
The Aerofoil...................................................................................................................... 13
Endplates........................................................................................................................... 14
2.2 Meshing.................................................................................................................................... 15
2.3 CFD Simulation ........................................................................................................................ 16
The Solver......................................................................................................................... 16

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2.4 Analysis.................................................................................................................................... 18
Observed Flow Pattern and the Improvements................................................................... 18
3 CHAPTER III – RESULTS AND DISCUSSION...................................................................... 20
3.1 Strain Rate Contours ................................................................................................................. 20
3.2 Static Pressure Contours............................................................................................................ 21
References........................................................................................................................................... 23

iii
ABSTRACT
The aim of our project is to simulate and analyze the front wing of Formula One car using CFD. ANSYS
13.0 is used as the computational tool to perform the task
And Creo Parametric is used to make 3D Models. The area of concern was to the design the wing in such
a way to create down force and to reduce the drag as much as possible. The design also emphases on the
end plates design to fulfill these objectives and to study the flow from them the flow array was premeditated
and the effectiveness of the different devices designed were evaluated. It appears that the behavior of the
flow was greatly subjective by the end plates design and the angle of attack of the wing as expected and
therefore that the design has to take account of these factors to achieve the desired coefficient of lift and
Drag. The turbulence model used in this analysis is k-epsilon model CL and CD values were calculated at
different angles of attack with different End plates designs.

iv
ACKNOWLEDGEMENT
In the name of Allah, the most Gracious and the most Merciful. All praise be to Allah, it is only because
of His permission and blessings that this group has completed this thesis within time.
Our group also would like to express our sincere gratitude to our supervisor Miss Anam Anwar of
Mechanical Engineering Department for her germinal ideas and invaluable guidance, continuous
encouragement and constant support that made this research possible. We are also very thankful to Miss
Syeda Fatima Ali from “Simulation Designs” and Mr. Shahid Akbar from MED UET Lahore (KSK
Campus) for their guidance regarding CFD softwares. We are also thankful to Dr. Shahid Imran from
Mechanical Engineering Department UET Lahore (KSK Campus) as he had always impressed us with his
outstanding professional conduct and his strong conviction for research.
My sincere thanks go to all the faculty members of Mechanical Engineering Department, UET Lahore
(KSK Campus) who helped us in one way or the other for the past four years that enabled us to choose
such a project and complete it on time.
Lastly I would acknowledge my sincere indebtedness and gratitude to my parents for their love, dream and
sacrifice throughout my life.

v
LIST OF FIGURES
Figure 1 (Chambered aerofoil)................................................................................................................ 2
Figure 2 (Wingspan)............................................................................................................................... 3
Figure 3 (Aerodynamic force and its components) .................................................................................. 4
Figure 4 (Comparison between drag and relative airspeed) ..................................................................... 4
Figure 5 (Form drag for different shapes)................................................................................................ 5
Figure 6 (Lotus 49-B 1968)..................................................................................................................... 6
Figure 7 (Gurney flap)............................................................................................................................ 6
Figure 8 (312 B3 model by Ferrari)......................................................................................................... 7
Figure 9 (Benetton B191) ....................................................................................................................... 7
Figure 10 (Minardi M195) ...................................................................................................................... 8
Figure 11 (Vortices at tip)..................................................................................................................... 10
Figure 12 (Wingtip vortex) ................................................................................................................... 11
Figure 13 (Different vortex locations)................................................................................................... 11
Figure 14 (NACA-4412 inverted profile).............................................................................................. 13
Figure 15 (Preliminary model).............................................................................................................. 14
Figure 16 (Secondary model)................................................................................................................ 15
Figure 17 (Mesh generation)................................................................................................................. 16
Figure 18 (Boundary conditions)........................................................................................................... 18
Figure 19 (Reynolds cell velocity contour)............................................................................................ 18
Figure 20 (Velocity contour)................................................................................................................. 19
Figure 21 (Path line traces) ................................................................................................................... 20
Figure 22 (Strain rate contour).............................................................................................................. 21
Figure 23 (Contours of static pressure).................................................................................................. 21
Figure 24 (Velocity contour)................................................................................................................. 22

vi
LIST OF TABLES
Table 1 (NACA-4412 coordinates) ....................................................................................................... 13
Table 2 (Secondary model)................................................................................................................... 15
Table 3 (Analytical parameters for Fluent)............................................................................................ 16
Table 4 (Air properties) ........................................................................................................................ 17
Table 5 (Results)................................................................................................................................... 18

vii
LISTS OF SYMBOLS
b Span of the wing
c Chord of aerofoil
α Angle of attack
αstall Stall angle
v∞ Relative air velocity
U∞ Free stream velocity
P Pressure
P∞ Free stream pressure
R Vortex core radius
Re Reynolds number
L Lift
D Drag
Cl Coefficient of lift
CD Coefficient of drag
ρ Density of air
S Wing area
Oswald efficiency
Induced drag coefficient
Zero-lift drag coefficient
k-ε K-epsilon
k-ω K-omega

1
1 CHAPTER I - LITERATURE REVIEW
1.1 Introduction
Computerized technological advancement which has armed us with such tools (Software) that can be
utilized not only in achieving optimum design for any typical environment. Software based exploration
distinguishes itself from classical approaches as it allows an engineer to not only understand the effect that
variables have upon the required outcome but also provides a concise method of assuring the design space
encompasses the parameter bounds for optimum performance. Moreover, in addition to detailing the effect
of specific variables, non-active parameters are also highlighted and thus may be fixed in later work to
improve computational time whilst preserving solution validity.
For the modern Formula One car the front wing has an enormous effect upon the overall vehicle
performance and particularly that of the aerodynamic regime. Vehicle stability and the flow fields of all
other components fixed on the front wing making this the most significant aerodynamic device of the entire
vehicle. Further adding to the complexity of this system is the fact that requirements for downforce
production, flow conditioning and drag reduction are mutually exclusive. Of course, of these requirements
the production of downforce is key for increased cornering performance [1].
Our aim to select this project is to study front wing, its design parameters that should be kept in mind while
designing it and parameters that are critical to have a wing with desired properties and experience the use
of softwares in designing and CFD (Computational Fluid Dynamics) analysis of Formula One front wing.
*(To avoid confusion, it should be noted that the lift and lift coefficient in an Aerofoil is counterpart of
downforce and downforce coefficient respectively in a Formula One front wing this is due to the inverted
geometry of Aerofoil in F1)
1.2 Motivation
1.3 Front Wing
The front wing is usually the first part designed by the teams. Failing designing an efficient front wing
may disrupt the whole aerodynamics of a Formula 1 car. One thing that makes its designing extremely
critical is the flexibility to adjust the attack angle of the flap placed on the front wing, indeed the balance
of the car is determined by the load on the front wing.
Nowadays front wing is a multi-elements Aerofoil which is not straight along its span. The main flap is
usually a rectangular plan-form. Additionally a Gurney flap may be fixed at the aft flap trailing-edge to
gain some extra downforce. Since the FIA (Federation International de I’ Automobile) Formula One
technical regulation has restricted the front track width of the cars, wing tip has become very crucial and
complex due to the wheel-wingtip flow interaction [2].

2
1.4 Aerofoil
The cross-section of a wing is called an Aerofoil. It is a specially designed shape that alters the flow of air
(fluid) on its two sides by changing respective paths that alter velocities and finally creates a net difference
in pressure on two sides of Aerofoil.
In case of an airplane it will produce lift and a downforce in case of a Formula One car.
Aerofoil Nomenclature
Specified names are assigned to describe the 2-D (two dimensional) as well as 3-D (three dimensional)
aerofoil geometries. Following is the description of these dimensions along with their illustration in fig. 1
and fig. 2 [1].
a) Leading Edge: Section of wing that first come in contact with fluid.
b) Trailing Edge: The last portion of wing opposite to leading edge.
c) Chord: The line joining the leading and trailing edge. It is represented by symbol" ".
d) Mean-Chamber Line: Locus of the points halfway between upper and lower surfaces.
e) Camber: The maximum distance between mean chamber line and chord line.
f) Relative Wind: Direction of incoming air taken as horizontal it velocity is V∞.
g) Angle Of Attack: The angle between relative wind and chord line. It is one of the most important
parameters that have effect on wing performance and the value of downforce and drag. It is represented
by symbol “α”.
h) Wing Span (b): Distance between tips or total breadth of a wing is called wingspan.
i) Wing Area (S): Multiple of chord and wing span is area of wing or Planform area for a non-swept
rectangular wing.
= ∗ (1.1)
j) Aspect Ratio (AR): Ratio of span to chord of a wing is called Aspect Ratio. It vary from plane to plane
and is divided in low, medium and high aspect ratio categories.
. = / (1.2)
Figure 1 (Chambered aerofoil)

3
Figure 2 (Wingspan)
1.5 NACA-4412
There are different types and classes of Aerofoil one of most authorized is NACA (National Advisory
Committee of Aeronautics) air foil that is further divided in different classes such as 4-series, 5-series, 6-
series etc.
The one we are going to use in our design and test belongs to NACA 4-series i.e. NACA-4412
NACA Nomenclature
As discussed in section 1.4 that NACA-4412 would be used in this project. So what does the digits in 4412
represents [3].
1st digit: It tells us about the value of maximum camber as a percentage of chord length.
As in 4412 maximum chamber is (0.04)*(c)
2nd digit: It tells us about the position of maximum camber from leading edge as 10th part of chord.
It will be (4/10)*(c) in this case.
3rd and 4th digit: The last two digits are the value of maximum distance between upper and lower surfaces
in the aerofoil as a percentage of c.
1.6 Wing Theory
Now we would discuss different parameters of wing from theoretical aspect most of the concepts we would
discuss in the wing theory are related to the fluid dynamics domain so the concepts related to fluid
dynamics such as pressure and velocity relation, viscosity and the boundary layer theory and the
application of Bernoulli and Continuity theory are the main ideas that should be kept in mind while
studying the wing theory [4,5].
Aerodynamic Force
Whenever air has a contact with a wing Aerodynamic force is created due to two effects i.e. pressure
difference on the upper and lower side of the wing and the shearing effect along the plane of aerofoil.
This force has two components i.e. Lift and Drag.

4
Figure 3 (Aerodynamic force and its components)
Drag
This component is parallel to flow and arise due to both shear and pressure effects.
= (1.3)
Lift Induced Drag
As the wing redirects the air and creates downforce a drag is induced in result. Its value increases with an
increase in attack angle. It is unavoidable consequence of downforce (or lift). As relative air speed
increases lift induced drag value decreases.
Parasitic Drag
It is a combination of different drag forces that are described in sub-sections (a) and (b). Its effect increases
with the increase in air velocity. It can also be measured as zero-lift drag which has a direct relation with
the drag coefficient which is zero-lift drag coefficient. Zero-lift drag coefficient is difference of drag
coefficient and induced drag coefficient as depicted in equation 1.4 along with relation to total drag in
fig.4.
= − (1.4)
Figure 4 (Comparison between drag and relative airspeed)

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a) Form Drag: It arises due to the broad frontal portion of a body. The value of form drag increases with
speed. It may also be taken as the pull of negative pressure of air as the body leaves some space behind
it just as shown in fig. 5.
b) Skin Friction: Whenever a body’s surface comes in contact with air while in motion there arises
viscous effect between the two and results in a shearing effect.
Figure 5 (Form drag for different shapes)
Downforce
This is perpendicular to flow and is result of pressure difference and is in accordance to Bernoulli’s
principle. Value of lift coefficient varies with varying angle of attack.
= (1.5)
Angle of Attack
It is the incidence angle of the front wing or the angle between the chord line and incoming air.
There is a maximum value for which a wing can be operational is called Stall angle and if we raise the
angle beyond that the wing would not be of any use/operational because of the separation between the
wing surface and air [6].
Moment
Aerodynamic force also produce a moment such as in usual operation in an airplane (during positive attack
angle) there will be a positive moment which will produce a nose-up effect.
1.7 Historical Review and Evolution of Wing
Front wing appeared in F1 on the Lotus 49B at the Monaco Grand prix on 25th of May 1968 [7].

6
Figure 6 (Lotus 49-B 1968)
This trend became popular soon and many F1 teams started to use it. Many wings developed with slight
differences. The wings were installed on the wheel hubs of the car to transfer the load directly to the wheels.
This allows avoiding stiffening the springs of the suspension system that might create instability on certain
tracks. However the wing spans were not enough resistant so they broke sometimes during the race.
Designers decided to enable the drivers to change the attack angle to have greater impact of wings on car
performance so modification was done to enable the driver to change the incidence during drive by the
help of a pedal. But in 1969 CSI (Comité Sportif International) seeing some safety issues modified the
regulation: [8]
 By prohibiting the mobile parts.
 By prohibiting the fixation of the wings struts on the suspension or wheels hubs.
 By limiting the width of wing and chassis height by 1 m.
In 1971, a new improvement appeared [9]. Dan Gurney made a high-lift device called Gurney flap. It is a
flat trailing edge flap perpendicular to the chord and that is not longer than 5% of the chord. It has been
usually used at the front wing trailing edge. It created an understeering effect so was successful especially
in sharp turns and corners.
Figure 7 (Gurney flap)
A major alteration made in 312 B3 model by Ferrari was shifting the wing quite far ahead from the body
that reduced the wing body interaction significantly, but the ground clearance was such that it could not
utilize ground effect benefit as wall as it should be [10].

7
Figure 8 (312 B3 model by Ferrari)
Innovations about the front wing were very limited during the end of seventies as teams built wing-cars
which showed some good performance at the time. But as the two elements wings was studied and tested
in early 80’s it proved to be a major breakthrough and revival of front wing. The angle of attack of the
second element was allowed to be modified as it was banned earlier in case of single element wing. Now
the load on the front wing could be altered according to requirement. Size of this second flap increased
over the time along with that the shape also changed by cutting it out near its root thus it had varying chord
over the span with longest at the ends and shortest chord as it joins the nose on the 1989 Ferrari 640 thus
creating more downforce, otherwise the increased attack angle for the second element would produce too
much perturbation in the inside flow to the cooling inlets [11].
In 1990s, raised nose increased the flow rate under the nose cone as in Benetton B191, high nose provided
a solution to avoid wing-body interference and enabled front wing to be faced to the free stream over its
whole span and thus increasing the capacity of wing to produce more downforce [1].
Figure 9 (Benetton B191)
As the air passes over the front wing some of its interacts with the front wing that creates drag so in 1991
air deflectors were mounted behind the front wing tips to bypass air aside the front wheels [1].
It was not until 1994 (death of Ayrton Senna) that FIA Formula One regulation was modified and did not
allow any chassis parts under a minimum ground height[12], so as a result some of the ground effect was
lost but more flow was available for cooling intakes. Designer came with an innovative thought so kept

8
the height different for center and sides and made multi element front wing. Specifically Minardi designed
a wing that was curved at the middle of the span of its M195 car as it was allowed by the regulation [13].
Figure 10 (Minardi M195)
Major change in 1998 regulations was the reduction in the front track width to 1.80 meters as earlier it was
allowed to be 2.0 m. However the width limit of the front wing remained 1.40 m. This resulted in another
problem that the font wing tips have been located in front of the tires. To get the maximum load on the
front wheels the width cannot be reduced so the solution was either to use an endplate that divert the flow
outward or inward of the wheel. Later was preferred as the inside of tires are closer then outboard side but
it provided another advantage that was the increased air for engine cooling ports [13].
1.8 Description of the Flow around the Front Wing
Different Aspects of Flow over the Wing Span
If we study the flow over the wing span it is basically a flow over a rectangular element having a profile
of an inverted wing with having an additional ability to benefit from ground effect. As in section 1.6 we
studied that in 1990s wing span are designed with varying height so specifically at the center portion of
span (as in curved central span) the height is kept such to benefit from ground effect as the regulations
allows to build body devices beneath 0.1 m above the reference plan (so between 0 an d 0.1 m) in an area
that does not exceed 0.25 m from the center line of the car [13].
Span location is also one of the designed parameters that plays part in determining the angle of attack of
first element. It is lower near the tips to keep the vortices lower that may be generated there. It is expected
that the turbulence decreases behind the front wing. Then the first element provides air flow to the second
element lower surface so that the latter’s incidence may be increased at very high angle of attack (it can be
more than 25 degrees) [6].
As we discussed in the historical review that second element chord is usually reduced near the nose or
center of the span, it is done to decrease the deflected flow downstream and to avoid interaction between
the nose and the second element.
As studied in section 1.6 one of the easy methods to increase the downforce in an easily and quickly is the
use of Gurney flap at the flap trailing edge [9]. The Gurney flap secure from a separated flow at an elevated
attack angle by causing a lower pressure area just behind itself which sucks the lower flow closer to the
wing surface. The Gurney caused some extra drag as well, but the wing would produce more downforce
comparatively.
Flow at the Wing Tips
We studied in section 1.6 that as the front wing length was reduced the interaction of tires with air resulted
in massive drag impact so to tackle this problem wing tips were introduced which deflect the air away from

9
wheels. As the effect of deflector is studied in accordance to the wheels motion (in straight drive and also
while the car is taking a turn), which is not a part of our project so we would not study the details of it [5].
1.9 Some Advanced Studies Conducted In Recent Times
As in section 1.1 and 1.6 we discussed that purpose of designing a wing in F1 cars is to enhance downforce
and reduce the drag. Engineers have come a long way but still they are struggling to find ways to achieve
maximum lift and minimum drag and the ways to avoid or overcome problems, at least to some extent if
not fully.
In general there are three basic techniques that leads us towards increased lift: [4, 5]
 Change in Camber line
 Span area change
 Boundary layer control
Thin Aerofoil theory shows that both lift and pitching moment increase with increasing aerofoil camber.
Trailing edge flaps and leading edge flaps can also increase aerofoil lift. Deflection of a flap increases the
effective camber of an aerofoil, thus increasing the lift coefficient at a given angle of attack.
Boundary Layer Separation Delay Using Flaps and Slats
In section 1.5.6 we studied about the angle of attack and stall angle. We also discussed in section 1.6 the
“Gurney flap” which was designed to prevent stall and allowed some increased angle of attack. We would
discuss how to delay that stall by some other design innovations.
As the wing theory depicted that the fluid layers that covers the wing produces the lift but what if there is
no layer around the wing so we have to avoid separation but as the attack angle is increased eventually
there will be a separation. Delaying the separation of the boundary layer on the upper surface of an airfoil
increases the lift on the airfoil one method to avoid this is use of leading edge slats and slots. Slats and
slots are different in geometry but they operate in the same way.
Slat is a small airfoil ahead of the leading edge whereas slots are actually cut through an airfoil. In both
cases high pressure air from the lower surface is directed tangential to the upper surface of the airfoil to
increase kinetic energy of that region [14] which will delay separation or stall angle and increase maximum
lift coefficient ( ) without changing zero lift attack angle ( ). A similar effect is used with slotted
trailing edge flaps. The result is delayed boundary layer separation and higher airfoil lift [15,16].
Overcome Stall and Separation Losses Using Jets
Slats and slots usage is one of many methods to overcome stall, many advance studies have been conducted
to find more efficient ways so we would discuss another method which is still under study by engineers
and scientists.
A prominent technique is “active flow control concepts” that insists on improving the efficiency and
stability of aerofoil by controlling flow separation. One way is continuous blowing or suction, which can
produce effective control but is difficult to apply in real applications. The application of synthetic jets to
flow separation control is based on their ability to stabilize the boundary layer by adding/removing
momentum to/from the boundary layer with the formation of vortical structures which promotes boundary
layer mixing consequently there is a momentum exchange between the outer and inner parts of the
boundary layer [17].
Discussing the characteristics of flow over uncontrolled and controlled airfoils we can say:
 The pressure distribution directly indicates the effect of synthetic jets on flow separation, most of the
lift enhancement is achieved in the upstream portion of the airfoil suction surface
 The control effect of synthetic jets on the pressure distribution in the pressure surface is negligible.

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 A typical synthetic-jet actuation with the momentum coefficient of 1.23% produces more than a 70%
increase in the lift coefficient additionally the drag coefficient is found to decrease approximately
15–18% with the synthetic-jet actuation [17].
Wingtip Vortices
Controlling wingtip vortices can be quite relevant in F1 where the high efficiency of wings in particular
and of the overall body in general is required. In sections 1.9.2 and 1.9.3 we discussed vortices takes part
in flow separation. But we can also utilize these vortices to achieve some favorable outcomes. Such as
using vortex generators to avoid or delay flow separation as it would enhance the lift of a wing or would
help in reducing body’s drag [18]. But first we would discuss the production of vortices.
Whenever a wing with some attack angle comes in contact with air and there is some relative velocity
between the two than according to wing theory pressure difference is created between upper and lower
surfaces, in a region near the tip the streamlines on the suction surface are bent towards the root of the
wing whereas the flow on the pressure surface are bent towards the tip. The roll up process is observed to
originate near the leading edge and as the flow is accelerated from the high pressure side to the low pressure
side it tend to circulates around the tip. With this process continue along the tip, the flow coming from the
pressure surface encounters a strong opposing pressure effect which ultimately result in separation of
boundary layer.
The vortex grows in size and strength along the wing. This vortical flow moved downstream to the trailing
edge and it finally forms the trailing vortex that can also be related to the upwash and the downwash which
are resultant of tip vortices. This wake is also entrained into the tip vortex and, at a distance of few chords
from the trailing edge, and seeing from behind the wing we can observe a pair of counter–rotating
axisymmetric trailing vortices in the presence of some smoke/color [19]. Lanchester– Prandtl finite wing
theory [20,21] showed how data from a two–dimensional aerofoil could be used to predict the aerodynamic
characteristics of a wing of finite span and also helped us understand the role of trailing vortices in the lift
generation and formation of vortices itself in which the lifting wing and its wake are replaced by a system
of vortices that divulges to the surrounding air an effect similar to the actual wing flow and generates a
force equivalent to the lift of wing. Utilization of vortices include the enhanced lift for a wing traveling in
wake of some wing that it is proceeding.
The vortex system can be divided into three main parts:
 Starting vortex
 Trailing vortex system
 Bound vortex system
Figure 11 (Vortices at tip)

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We can observe starting vortex and trailing vortex whereas bound vortex system is a hypothetical
arrangement of vortices that replace the real physical wing [5]. Upper and lower surfaces of the wing
carries the boundary layer that are resultant of the upstream segment of the vortex ring which consequently
produces the lift and drag through pressure and shear stress distributions. The distribution of vortex
physically are depicted in the fig. 12 where bound vortex and trailing vortex are illustrated.
Figure 12 (Wingtip vortex)
Vortex Classification
We can categorize the vortex in different zones based on their life each zone depicts certain features [22].
a) Near Field Region: Near field also known as near wake extends from the trailing edge to some chords
behind the aircraft where a complete roll up of the wake is observed and all the circulation is contained
in the vortices. The resulting trailing vortex from both the wing portions forms a sort of couple that
have parallel rotational axis about which the vortex shows a rotational motion and the wake that arises
between these two axes is called downwash whereas the wake out of that region is called upwash
vortex. The portion of wake that lies at a few chords behind the trailing edge is called early-wake. In
this region, the influence of the wing geometry is of great importance and the primary vortex shows a
remarkable asymmetry.
b) Mid Field Region: Similar to the “Near Field Region” “Mild Field Region” is also characterized by
the distance from trailing edge. This region les next to the near field region and extend up to few
hundred chord length. As the vortex has fully developed in prior region here they start decaying but
vortex decays at a relatively small rate in this region based on this characteristic we can also name it
as Diffusion regime.
c) Far Field Region: At a certain distance from the wing, vortices loses their form and shows instability
which remarkably changes their shape as there occur a rapid decay and breakdown of vortex.
Figure 13 (Different vortex locations)
Effect of Aspect Ratio on Vortices
Aspect ratio has a great influence on the vortices. We studied in section 1.9.3 that wingtip vortices are
developed as the flow is directed from lower (high pressure) to upper (low pressure) surface of wing near

12
the wing tip. That will cause an early separation so this phenomenon can be reduced if we are using a wing
with high aspect ratio
The concept of high aspect ratio producing lesser wingtip vortices can be understood easily by imagining
the wing as a cylindrical beam of air that is moving down, this downward motion would result in creation
of vortices at the tips so if the wing has a higher aspect ratio it would have a lower tip area comparing to
the tip area of a high aspect ratio wing with same span area, so it would create lesser tip vortices [23].
In section 1.9 we have discussed different problems with their possible solutions but till today a large
number of research papers has been published and many experiments have been conducted on the wingtip
vortex despite all that time spent to understand them by scientists there are still many unanswered questions
and ambiguities. Nowadays, CFD methods have reached a state of reliability so utilization of these
techniques can give us detailed descriptions of complex flows; in particular it is now common to find
simulations on wings, and the main objective is the prediction of the development and decay of the trailing
vortices and their interaction with other bodies that can be achieved by an accurate representation of the
vortex formation. Seeing their importance we will analyze the surface pressure measurements which is one
of the important parameters that effects the vortices.

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2 CHAPTER II – METHODOLOGY AND MODELING
We have discussed the basics of wing profile, its nomenclature, history and improvements done in the front
wing of formula one and also had an idea what we are going to do in our project. We also discussed a little
about the softwares that are used. In this chapter we will discuss the methodology adopted for modeling
and analysis of front wing.
The model was designed from scratch, we have studied the effect of downforce on the front wing and now
we would observe its effect by the help of analysis.
2.1 Design and Specification of the Model
This portion is about the design of wing we created along with its profile and different features it carry.
We will discuss different aspects including how it was modeled.
The Aerofoil
Basic element of a front wing is an Aerofoil. We are going to use two elements i.e. wing and endplate.
We are designing from the scratch and having a simplified approach, as we did not get any Formula One
aerofoil data so we are choosing a simple inverted NACA 4412 aerofoil [24]. To draw its profile we used
“PROFSCAN” as a tool in which we allocated 101 points as shown in fig. 14 along with some of these
coordinates in table 1.
Figure 14 (NACA-4412 inverted profile)
Table 1 (NACA-4412 coordinates)
Point X Y Point Y X Point X Y
1 983.0 1.238 40 84.960 -58.547 67 172.290 27.602
5 939.964 -10.714 45 26.285 -31.475 72 283.251 22.441
10 830.418 -36.833 50 0.756 -4.934 82 552.229 11.107
15 695.325 -62.941 51 0.00 0.00 87 695.325 6.104
20 552.229 -83.171 52 0.756 4.629 92 830.418 2.929
30 283.251 -94.623 57 26.285 21.311 97 939.964 1.513
35 172.290 -81.303 62 84.960 28.261 101 983.0 1.238

14
The models of the front wing were generated using the NACA-4412, the wing was produced just by
extrusion of the Aerofoil imported in Creo-parametric so it has a constant chord throughout. Later the
endplates were attached by setting the constraints in Creo-parametric. Using the basic NACA profile we
generated two different wings. Which are different comparing the type of end plates used and the effect of
nose.
Endplates
Endplates have to be located properly between 10 cm and 30 cm above the reference plane. The space
can also be utilizes to have an additional downforce impact by deflecting the flow inboard the wheels
with the help of some additional features or parts.
We studied in sections 1.6 and 1.7 the function and a little design feature. To deflect the flow we are
using the same NACA-4412 Aerofoil here as endplate that will be installed vertically at an incidence of
12 degrees.
The profile of 4412 Aerofoil was generated by the help of “PROFSCAN”.
Now we would discuss the two models we designed for CFD analysis.
a) Preliminary Model
The first model was made without the nose cone. The end plates are not curve and a single airfoil
wing inclined at 20 degrees. The airfoil have been premeditated to direct the flow inboard. The
NACA 4412 airfoil is used as the wing. The aim here is to refract the flow around the tires and to
reduce the turbulence and the effect of vortices produced around the front of the wing with
minimum area. Illustration of CAD model can be seen in fig. 15.
Figure 15 (Preliminary model)
b) Secondary Model (with Nose)
The second model was improved by making the endplate curved and the nose cone was added to
the front not only to create the additional down force but also to make the body streamline the angle
of attack of the wing was 12 and the Basic NACA airfoil was used the model was the curved
generated in the plate not only decrease the turbulence and the vortices formation but also the CL
is improved. The features of model are depicted in the fig. 16 and table 2.

15
Table 2 (Secondary model)
Profile
Chord
(c)
(mm)
Span
(b)
(mm)
Planform
(S)
(mm2)
Aspect
Ratio
Incidence
angle
(degrees)
Max.
Chamber
(mm)
Max.
Chamber
from
leading
edge
(mm)
Max.
Thickness
of
Aerofoil
(mm)
NACA-
4412 217.221 1185.09 257,426.43489 5.455 12 8.688 86.88 26.06
Figure 16 (Secondary model)
2.2 Meshing
Meshing is breaking of physical problem that might be 2-D or 3-D into simpler element i.e. triangles,
quadrilaterals, tetrahedral or hexahedral to make the solution easier and more accurate. The denser the
meshing the more accurate the results will be but at the same time it becomes more complex and difficult
to solve. For the best results the mush should be refined at the edges.

16
Figure 17 (Mesh generation)
2.3 CFD Simulation
The CFD simulation were carried out with ANSYS FLUENT 13.0 which is a Navier-Stocker equation
[25].
. ⃗ +
⃗
= ⃗ − ⃗ + ∆ ⃗ (2.1)
The Solver
The run first simulations segregated solver was used and they were steady flow calculation with absolute
reference frame.
a) Viscous Solver
The aim was to carry out viscous simulation to get evidence of the strong impact of a moving ground. The
Spalart-Allmaras viscous solver seemed to be the most suitable solver for this kind of study. However this
kind of solver did not converge for all types of models. So we decided first to run different models with the in-
viscid or k-ε (with standard wall functions and then sometimes with a little enhanced) solver. After it was done
we moved to Spalart-Allmaras. It was not such a difficulty as it matches the grid generation and adaptation
refinement strategy. We tried multiple times with the k-ε (enhanced wall functions with the pressure gradients
options as well as the standard one), k-ω (SST with the translational flow) and Spalart-Allmaras [26,27].
b) Calculation Solver:
First order simple model was used to perform all the calculations. Second order were not attempted
because simulation with the second order was taking too long because of the mesh size selected.
Table 3 (Analytical parameters for Fluent)
Enabled values in Fluent setup Assumed parameters
Space Three dimensional
Motion Type Stationary
Time dependence Steady
Material Aluminum
Material density (kg/m3) 2719
Flow Type Absolute Flow
Equation of state Constant density
Viscous regime Turbulent
Turbulence Model Realizable k-epsilon Model
Optimal physical Model None

17
c) K-epsilon Turbulent Model
Two-equation turbulence model is such that it solves two separate transport equations so it allow
the determination of both, a turbulent length and time scale. We used ANSYS Fluent, standard
model in ANSYS Fluent belongs to this class of models and so it has become very popular in
engineering flow calculations. It is efficient, economical and reasonably accurate for a wide range
of turbulent flows explain its popularity in industrial flow and heat transfer simulations [28].
Our problem has a semi-empirical model, and the derivation of the model equations relies on
phenomenological considerations and experimentation. The standard model is a model based on
model transport equations for the turbulence kinetic energy and its degeneracy rate. The model
transport equation for wing is derived from the exact equation, while the model transport equation
for wing was obtained using physical reasoning and bears little resemblance to its mathematically
exact counterpart. In the derivation of the model fully turbulent flow was assumed and the effects
of molecular viscosity was considered to be negligible. Seeing that we can say that validity of
standard model is only for fully turbulent flows. After we analyzed all the strength and weakness
of the standard model, we introduced some modifications to improve its performance.
Table 4 (Air properties)
Temperature (K) 293
Air Density (kg/m3) 1.22499
Specific Heat (KJ/(Kg.K)) 1.005
Thermal Conductivity (W/(m. K)) 0.0257
Kinematic Viscosity (m2 /s) 1.78*10-05
Turbulence kinetic energy (J/Kg) 0.8
Prandtl Number 1
d) The Boundary Conditions
Basically there are three basic parameters to consider while setting the boundary conditions i.e.
 Inlet
 Interior air
 Outlet
 The wing
 Wall air
Currently the average on-track speed of a Formula One car over a lap is between 150km/h to 240
km/h so we have selected 216km/h that is 60m/s for our analysis.
The boundary conditions has been set as follows:
 Air at entrance is 60m/s
 Turbulence intensity is set to standard 2%
 Front wing is set as a “wall”
 The ground is set as a moving wall having the same speed as that of the incoming air i.e. 60m/s.
 Outlet of the domain is set as the outflow

18
Figure 18 (Boundary conditions)
2.4 Analysis
The analysis was done on the flow visualization obtained using ANSYS FLUENT 13.0 that provided us
the values of the Cl and Cd in addition to the flow pattern. The values obtained by analysis are depicted in
table 5.
Table 5 (Results)
Model Solver type Coefficient of Lift Coefficient of Drag
No1. (Preliminary
Model)
K-epsilon -0.49 0.19
No. 2 (Secondary
Model)
Spalart-Allmaras -0.41 0.17
Observed Flow Pattern and the Improvements
The velocity profile of both the model was the same as it was expected that the speed of fluid on the lower
portion of the wing was very high and it was slow on the upper side as predicted by the velocity contours
of the wing and hence according to the Bernoulli’s equation the force is produced that tries to push it
downward that helps to create the down force and hence stabilize the car.
Figure 19 (Reynolds cell velocity contour)

19
The pattern shows that the angle of attack should be increased up to 25-26 degrees so that the flow can be
moved up from the body of the car. To allow more air to be moved up to the tires and outward to the body
of the car the end plates design should be improved by adding an additional curved vane at the end.
Figure 20 (Velocity contour)

20
3 CHAPTER III – RESULTS AND DISCUSSION
The thesis includes the designing geometries and analyzing it. The designing section of the wing was the
complicated task as it need to be done with the great care with the minute details so that the required test
can be performed on the model and no extra edges are created during the mesh generation process. The
simulation should be run by the complete knowledge of the solver specially the turbulence models and the
relaxation factors, Reynolds number etc.
The thesis focuses on the two model one was the nose cone and the without it also the effect of endplates
on the model was studied. To capture the effect of speed and turbulence the meshing process is the key to
gain the perfect results the so the body of influence should be made at the sites where the effects are critical
and mesh sizing should be adjusted.
Another important parameter in the designing is the setting of angle of attack that the wing experiences
and the height of the wing from the base is key factor in it.
Figure 21 (Path line traces)
3.1 Strain Rate Contours
The maximum strain occur at the endplates sides where the small vortex are formed the strain can be
minimize by adding another vane or a flat plate that further targets the flow inboard.

21
Figure 22 (Strain rate contour)
3.2 Static Pressure Contours
The contours of static pressure show that the static pressure is same at the whole wing except a few areas
where the little red spots appear at the front of the nose cone and at the leading edge of the wing.
Figure 23 (Contours of static pressure)

22
Figure 24 (Velocity contour)

23
REFERENCES
[1] Pierre Ménard, The Great Encyclopedia of Formula 1, 1950-1999: 50 years of Formula 1., Constable
and Robinson, London, 2000
[2] F. Mortel, CRANFIELD TEAM F1: THE FRONT WING, Introduction, 2003, p. 2-3.
[3] Marzocca, Pier. "The NACA airfoil series" (PDF). Clarkson University. Retrieved 2009
[4] Abbott Ira., Theory of Wing Sections: Including a Summary of Airfoil Data. New York: Dover
Publications, 1959, p. 115-128
[5] John D. Anderson Jr., "Fundamentals of aerodynamics", McGraw-Hill; 5th edition, 2010.
[6] A.G Chervonenko, “Effect of attack Angle on the Non stationary Aerodynamic Characteristics and
Flutter Resistance of a Grid of Bent Vibrating Compressor Blades”, Ukrainian Academy of Sciences,
Plenum Publishing Corporation, Ukraine, Volume 39, No. 10,1991, pp. 78-81.
[7] John Krewson, "Fast, Present, Future: 1967 Lotus 49 vs. 2013 Corvette ZR1" , Road and Track,
retrieved 2013
[8] Pritchard, Anthony, “Directory of Formula One Cars: 1966-1986”, Aston Publications Limited.
United Kingdom, 1986, p. 223.
[9] Howard K., “Gurney Flap.”, Edition of Motorsport magazine, England. Cited in: Dan Gurney’s All
American Drivers, 2000
[10] Acerbi, Leonardo, “Il Nuovo Tutto Ferrari”, Nada 2008.
[11] Henry, Alan, “AUTOCOURSE 1989-90”, Hazleton Publishing Ltd., 1989, pp. (76, 81).
[12] “A History of Safety in Formula One". formulaone.com. Formula One Management.
Retrieved 2011.
[13] "F1 rules and stats 1990-1999". f1technical.net, January 2009. Retrieved 11 July 2011.
[14] David F. Rogers, “EFFECT OF SLATS AND FLAPS O N A FINITE WING,” Experiment V,
2010.
[15] D. You, P. Moin, “Active control of flow separation over an airfoil using synthetic jets,” Journal of
Fluids and Structures 24, 2008, pp. 1349 – 1357.
[16] Gostelow, J. P., Blunden, A. R., Walker, G. J., ‘‘Effects of Free Stream Turbulence and Adverse
Pressure Gradients on Boundary Layer Transition,’’ ASME J. Turbomach, 116, 1994, pp. 392–404.
[17] Gad-el Hak, “Flow Control-Passive, Active, and Reactive Flow Management,” 1st edition,
Cambridge University Press: Cambridge, UK, 2000; pp. 150–203.
[18] Clancy, L.J., Aerodynamics, Pitman Publishing Limited, London, 1975.
[19] G. Bramesfeld, M. D. Maughmer, "Effects of Wake Rollup on Formation-Flight
Aerodynamics", Journal of Aircraft, Vol. 45, No. 4 (2008), pp. 1167-1173.
[20] Lanchester, Frederick W., Constable, ed. Aerodynamics, 1907.
[21] Prandtl, Ludwig, Königliche Gesellschaft der Wissenschaften zu Göttingen, ed. Tragflügeltheorie.
1918.
[22] Batchelor G.K., Axial flow in trailing line vortices. J. Fluid Mech. 20, 1964, 645–658.
[23 ] Clancy, L.J., Aerodynamics, John Wiley & Sons, 2010, section 5.15.

24
[24] Ahsan, Galas, Thaivalappil, Abdul Hameed, ANALYSIS OF NACA 4412 AIRFOIL: ANALYSIS
USING FLUENT & PANEL METHOD AND VERIFICATION WITH EXPERIMENTALDATA, LAP
LAMBERT Academic Publishing, 2010.
[25] Rogers, S. E. and Kwak, D., “An Upwind Differencing Scheme for the Time Accurate
Incompressible Navier-Stokes Equations,” AIAA Journal, vol. 28, 1990, pp. 253–262.
[26] Menter, F. R., "Zonal Two Equation k-ω Turbulence Models for Aerodynamic Flows", AIAA
Paper, 1993, 93-2906.
[27] Menter, F. R., "Two-Equation Eddy-Viscosity Turbulence Models for Engineering Applications",
AIAA Journal, vol. 32, no 8, 1994, pp. 1598-1605.
[28] Yu, N. J., Allmaras, S. R., Moschetti, K. G., “Navier-Stokes Calculations for Attached and
Separated Flows Using Different Turbulence Models,” AIAA Paper, 1991, 91-1791.

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