SAVITRIBAI PHULE PUNE UNIVERSITY

A Project Stage-II

Project Report
On

“STUDY OF DRAG COEFFECIENT AND

TURBULENCE USING CFD PERSPECTIVE”

Submitted by
Group no. 07

Anshu Shinde (B191090932)

Indraneil Sinha (B191090936)
Shubham Thokare (B191090948)
Kush Tilgule (B191090949)
Siddhant Adak (B191090969)

Under the Guidance of

Dr. Suchit Deshmukh

DR. D. Y. PATIL INSTITUTE OF ENGINEERING MANAGEMENTAND

RESEARCH.

DEPARTMENT OF MECHANICAL ENGINEERING

2023-2024
 DR. D. Y. PATIL INSTITUTE OF ENGYINEERING
MANAGEMENT AND RESEARCH, AKURDI.

DEPARTMENT OF MECHANICAL
ENGINEERING 2023-2024

CERTIFICATE

This is to certify that, Anshu Shinde (B191090932), Indraneil

Sinha(B191090936), Shubham Thokare (B191090948), Kush
Tilgule(B191090949), Siddhant Adak (B191090969) has successful completed
they project entitled “STUDY OF DRAG COEFFECIENT AND
TURBULENCE USING CFD PERSPECTIVE” and it is submitted towards the
partial fulfillment of Bachelor of Engineering (Mechanical Engineering) affiliated
to Savitribai Phule Pune University.

Dr. Suchit Deshmukh External Examiner

Project Guide

Dr. Ganesh Jadhav Dr. A. V. Patil

Head of the Department Principal
 ACKNOWLEDGEMENT

I am fortunate to be a part of Major project initiated by the Department of Mechanical

Engineering, DYPIEMR, Akurdi. I would like to express my sincere gratitude for
shaping this initiative into a successful model. The major project bridges the gap
between an industry and academics. First of all, I would like to thank Dr. Ganesh
Jadhav HOD, Mechanical Department, DYPIEMR, AKURDI and Dr Suchit
Deshmukh (Project Guide), Mechanical Department, DYPIEMR, AKURDI.

I am grateful to acknowledge the contribution of my team members. This project is

one of my big milestones in my career development. I will implement the acquired
skills and knowledge in the best way possible, and will strive to work on its
improvement, to get desired objectives. I gladly hope to cooperation from all of you
in future.

Place: 26-05-2024 Anshu Shinde(B191090932)

Indraneil Sinha (B191090936)
Date: Akurdi Shubham Thokare(B191090948)
Kush Tilgule (B191090949)
Siddhant Adak (B191090969)

MECHANICAL ENGINEERING DEPARTMENT

ABSTRACT
This study presents a numerical investigation into the influence of turbulence on the
drag force characteristics of Aerofoils using computational fluid dynamics (CFD)
simulations. Turbulence is a significant factor affecting aerodynamic performance, yet
its precise impact on drag force remains a subject of intense research and practical
importance.

The research employs advanced numerical methods to simulate turbulent airflow over
a range of Aerofoil geometries and angles of attack. Specifically, Reynolds-Averaged
Navier-Stokes (RANS) and Large Eddy Simulation (LES) models are utilized to
capture turbulent structures and their effects on drag generation with high fidelity.

Through systematic analysis of turbulent flow patterns and associated drag forces, this
study aims to unravel the intricate relationship between turbulence intensity, Aerofoil
geometry, and drag coefficient. The numerical simulations provide valuable insights
into how turbulence influences separation, boundary layer behavior, and overall drag
characteristics.

By focusing solely on numerical methods, this research aims to provide a

comprehensive understanding of turbulence-induced drag effects, offering engineers
and designers crucial information for optimizing Aerofoil performance in turbulent
environments. The findings have implications for diverse applications, including
aircraft design, wind turbine efficiency, and hydrodynamic systems, contributing to
advancements in aerodynamic engineering and sustainable technology development.

The Aerofoil section is the incarnation of a wing or a lifting surface which is very
important in an airplane wing design. The design of Aerofoils is an important aspect of
aerodynamics. Different Aerofoils are used for different flight regimes. While the shape
of the Aerofoil changes, their aerodynamic characteristics also change. This project
deals with a standard symmetrical Aerofoil as reference and the effect of changes in
shape due to minor variations in the coordinates. New Aerofoil shapes have been
produced in this optimization process. At zero angles of attack, asymmetric aero foils
can create lift, but symmetric aero foils have no lift. The design of an Aerofoil has a
significant impact on the lift produced to an aircraft as well as the thrust necessary to
achieve that lift. The aerodynamic characteristic results such as the coefficients of lift
and drag (Cd, Cl), pressure coefficient (Cp), are noted for three different profiles,
produced from the standard NACA 23012, 23014, 63415 Aerofoil with changes in the
chord thickness distance but no change in the maximum thickness in percentage of
chord.

Design Modeler is used to create the geometry for the Aerofoil. NACA 23012, NACA

MECHANICAL ENGINEERING DEPARTMENT

23014, and NACA 63415 are three kinds of Aerofoils that are examined. Ansys Fluent
is used to do the study of the Aerofoil. Constant density is used to compute the
coefficients of lift and drag across the segment. From 0 to 18 degrees AOA, the analysis
is carried out. Using the NACA profiles a 3D model of flight wing is created in CATIA
software. The modus-operandi used in this optimization process is the Computational
Fluid Dynamics (CFD).
Flow changes have been recorded for these Aerofoil shapes and the results are arrived
for finding the best Aerofoil that can be advisable to be used in compressors, turbines,
etc. with reduced flutter and maximum life. A numerical CFD analysis of NACA
profiles is undertaken in this work. ANSYS Fluent software is used to do the analysis.
NACA23012, NACA 23014, and NACA 63415 are considered 2D profiles at first. The
analysis is done up to 18 AOA to find the optimal aero foil. In tabular columns and
graphs, the individual profile drag and lift coefficient are recorded.

MECHANICAL ENGINEERING DEPARTMENT

LIST OF CONTENTS

1. INTRODUCTION…………………………………………...………. 1
1.1.1. Motivation……………………………………………….….1
1.1.2. Background…………………………………………………1
1.1.3. Aerodynamics Forces on Aerofoils………….……………...2
1.1.4. Problem Definition…………..……………………………...4
1.1.5. Aim and Objectives………………………………………....4
2. LITERATURE REVIEW…………………………………………..... 5
2.1.1. Numerical work on Aerofoil wing…………………….……5
2.1.2. Summary……………………………………………………7
3. METHODOLOGY……………..………………………………...….. 7
3.1.1. CAD software………………………………………..…..…8
3.1.2. CAD generation…………………………………………….9
4. CFD…………………………………………………………….....…13
4.1.1. Benefits of CFD……………………………………………13
4.1.2. CFD Process………………………………………...……..13
4.1.3. CFD model Process………………………………………..14
4.1.4. Results of Flight Models………………………………...…21
5. RESULT AND DISCUSSION……………………………………... 22
5.1.1. Drag Coefficient…………………………………………...29
5.1.2. Lift Coefficient………………………………………….…32
5.1.3. Lift: Drag Ratio for Aerofoils……………………….….…35
6. CONCLUSION………………………………………………….…..39
6.1.1. Future Scope……………………………………………….39
7. REFERENCES………………………………………………………40
8. STUDENT AND GUIDE DETAILS…………………………….…. 41
9. PLAGIARISM REPORT………………………………………...….42

MECHANICAL ENGINEERING DEPARTMENT

LIST OF FIGURES

Figure 1. Lift and Drag forces Acting upon an aerofoil………………………...………2

Figure 2 The development of Wingtip Vortices……………………………….……….3
Figure 3 NACA Aerofoil 23012…………………………………………………….…9
Figure 4 NACA Aerofoil 23-415…………………………………………….…….…..9
Figure 5 Aerofoil Coordinates…………………………………………………….….10
Figure 6 Imported Coordinate file……………………………………………………10
Figure 7 NACA Aerofoil 23012 curve………………………………………………..11
Figure 8 NACA Aerofoil 23012 3D model…………………………….…………….11
Figure 9 Ansys Workbench ………………………………………………………….14
Figure 10 Ansys Fluid Fluent………………………………………..……………….14
Figure 11 NACA Aerofoil 23012 3D model axes……………………………………14
Figure 12 NACA Aerofoil 23012 3D model geometry…………………………..…..15
Figure 13 NACA Aerofoil 23012 3D model with geometry…………………………15
Figure 14 NACA Aerofoil 23012 boundary box…………………………………..…16
Figure 15 NACA Aerofoil 23012 boundary box (2)………………………………….16
Figure 16 NACA Aerofoil 23012 boundary conditions……………………………....16
Figure 17 NACA Aerofoil 23012 boundary conditions…………………………...…17
Figure 18 NACA Aerofoil 23012 meshing……………………………………….…..17
Figure 19 NACA Aerofoil 23012 results…………………………………………..…17
Figure 20 Geometry model and air domain……………………………………….….18
Figure 21 Structured Mesh on Aerofoil domain………………………………….…..18
Figure 22 Aerofoil with detail view…………………………………………….…….19
Figure 23 Boundary Conditions on air………………………………………….…....19
Figure 24 Turbulent Model selected for analysis…………………………………….20
Figure 25 Velocity distribution around Aerofoil NACA 23012 at 0° AOA………….21
Figure 26 Pressure distribution around Aerofoil NACA 23012 at 0° AOA………….22
Figure 27 Velocity distribution around Aerofoil NACA 23012 at 18° AOA……......22
Figure 28 Pressure distribution around Aerofoil NACA 23012 at18° AOA……..….23
Figure 29 Velocity distribution around Aerofoil NACA 23014 at 0° AOA………....23
Figure 30 Pressure distribution around Aerofoil NACA 23014 at 0° AOA………....24
Figure 31 Velocity distribution around Aerofoil NACA 23014 at 18° AOA…….….24
Figure 32 Pressure distribution around Aerofoil NACA 23014 at18° AOA………...25
Figure 33 Velocity distribution around Aerofoil NACA 63415 at 0° AOA……....…25
Figure 34 Pressure distribution around Aerofoil NACA 63415 at 0° AOA………....26
Figure 35 Velocity distribution around Aerofoil NACA 63415 at 18° AOA…….….26
Figure 36 Pressure distribution around Aerofoil NACA 63415 at18° AOA………...27
Figure 37 Drag coefficient plot for NACA 23012…………………………………...28
Figure 38 Lift coefficient plot for NACA 23012………………………………….…29
Figure 39 Drag coefficient plot for NACA 23014………………………….………..30
Figure 40 Lift coefficient plot for NACA 23014………………………………….....31

MECHANICAL ENGINEERING DEPARTMENT

Figure 41 Drag coefficient plot for NACA 63415………………………………..…32
Figure 42 Lift coefficient plot for NACA 63415………………………………..….33

MECHANICAL ENGINEERING DEPARTMENT

 LIST OF TABLES
Table 1: Drag Coefficient plot for NACA Aerofoil 23012……………………………28
Table 2: Lift Coefficient plot for NACA Aerofoil 23012…………………………....29
Table 3: Drag Coefficient plot for NACA Aerofoil 23014……………………….….30
Table 4: Lift Coefficient plot for NACA Aerofoil 23014……………………….…...31
Table 5: Drag Coefficient plot for NACA Aerofoil 63415…………………….…….32
Table 6: Lift Coefficient plot for NACA Aerofoil 63415…………………………....33
Table 7: Aerofoil Drag coefficient comparison……………………………………...34
Table 8: Lift coefficient of NACA 23012, 23014, 63415…………………………....35
Table 9: Lift/Drag ratio of NACA 23012, 23014, 63415……………………………36

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 LIST OF GRAPHS

Graph 1: Drag coefficient of NACA 23012, 23014, 63415………………………….34

Graph 2: Lift coefficient NACA 23012, 23014, 63415…………………………...…35
Graph 3: Lift/Drag ratio of NACA 23012,23014,63415…………………...………...36

MECHANICAL ENGINEERING DEPARTMENT

 INTRODUCTION
1.1 Motivation
Now days, rise in new kinds of technologies made humans to prefer luxurious life. From
moving one place to another place, we have to depend on transport. Generally, the
modes of transport are road, air and water. To travel long distance with in short time,
air transport is one of the sources. So that domestic and international flights are the best
option. The flight design depends upon the aerodynamic characters such as drag and
lift. The Engine performance is increased when the drag force is reduced. The flights
can travel from ground level to certain altitude. When the flight flying, the drag and lift
forces will change. The analysis of these aerodynamic forces is depending on ground
effect. This type of study is advanced and trending topic for researchers to design flight.
For many years, researches were going on for the ground effect. Finite element analysis
is a good option than the manual calculation and for experimental point of view. The
flight is a complex shape to find out the aerodynamic characters of flight and numerical
analysis is best option to reduce the time for solving and to obtain precise results. This
type of study is helpful for flight manufacturing industries. Before constructing a flight,
the drag acting on the flight is known before using CAD software and will give
approximate results. Wing is one of the main parts of the flight. On these wings,
aerodynamic forces with ground effect by using simulation software by considering
different level of wing length to height ratio is planned in this work.

1.2 Background
There are two different types of flights are available these days, they are fixed wing
flight and rotating wing flights. For long distance travelling and high speed, fixed wing
flight has huge demand. But the flight controlling is mainly depending on the wing and
wing pitch angle. For this type of flights, when the wing size and structure geometry
increases, the drag force effect will be more. The resisting force is high that the engine
performance is needed more than usual. Because of high usage of engine performance,
the fuel cost increases and passenger ticket cost also increases. For this type of flights,
the reduction of drag force is a big task for aircraft designers. There are various kinds
of wings shapes that are available in the market with different series with NACA
profile. When the flight wing’s pitch angle changes it is necessary for the industries to
check on the change in drag force. If the drag force is reducing, the engine performance
is drastically improving and there would not be much use of engine power. So, to
identify the particular drag force, sensors are needed and real time set up is need and it
is a highly complicated task for researcher and manufacturers. So here the alternative
method called finite volume techniques is applied and analysis can be preceded.
Because of this, working time is reduced, equipment’s and other components cost can
be avoided completely as the work is completely depending on the software. In the
software, aerodynamic resistance force, dynamic force results are gained can be
performed using software. Based on this, drag can be reduced models can be formed.

MECHANICAL ENGINEERING DEPARTMENT 1

1.3 Aerodynamic Forces on Aero foils
1.3.1 Flight
Any object under moving condition and stationary condition if the lifting forces are
produced, they are called lifting body. These lifting bodies are available and depend on
size and shape. Aerofoil is one of the lifting bodies. This Aerofoil body can be
characterized by depending on lift and drag ratios. This Aerofoil should produce high
amount of lift and lower amount drag, only then the body is known as efficient body.
If Aerofoil moved in air, due to the air movement around the Aerofoil depends on the
Aerofoil shape, a low pressure will create. At the top surface of the Aerofoil and high
amount of pressure will create at lower surface of Aerofoil. Due to this pressure
difference, the body can travel from high pressure to low pressure and body gains the
lifting forces and Aerofoil is pushed toward upward direction to tilt back slightly.

Figure 1: Lift and Drag Forces Acting upon an Aero foil

Due to this motion two different forces will act on Aerofoil which are lift and drag
force. The force which is perpendicular to wind direction is called lift force. The parallel
direction is called drag force. Bernoulli’s theory will be helpful to calculate how the
flight will fly.

1.3.2 Lift
Lift is a force which is acting on Aerofoil and it can be calculated with non- dimensional
parameter named lift coefficient.

Where Cl = coefficient of lift

L= length
ρ = density
V=volume
S= swept area

MECHANICAL ENGINEERING DEPARTMENT 2

1.3.3 Drag
It is a main considerable parameter on Aerofoil and it can be characterized by two types
which are skin friction drag and pressure drag.
Skin friction drag:
Skin friction drag is created by resulting from shear stresses acting on the Aerofoil
body. It depends on the geometry and will remain constant with ground effect

1.3.4 Pressure Drag:

Pressure drag is equal to the rate of change of air particles' linear momentum normal to
the local surface in a local surface's co-moving inertial frame minus pressure forces.
The pressure drag is developed only when the pressure acting on Aerofoil profile while
the air flows around the Aerofoil. This type of pressure drag is further classified into
wave drag, induced drag and boundary layer pressure drag.

Figure 2: The Development of Wingtip Vortices

In pressure drag, induced drag is main factor to consider. The maximum forces
available on Aerofoil in the above three types of drags. Higher pressure will act at the
front and bottom of the Aerofoil. The lower pressure will act at tip of Aerofoil. Due to
this, circular flow can show at tip of the wings. This circular flow of air at the tip of the
wing and it further moves down. It is known as downwash. While the air is circulating,
it will create small amount of lift force in to drag force. The drag force can be calculated
with non-dimension parameter coefficient of drag.

Where Cl = coefficient of lift

L= length
ρ = density
V=volume
S= swept area

MECHANICAL ENGINEERING DEPARTMENT 3

1.3.5 Aerofoil
It is a structure with curved surfaces designed to give the most favorable ratio of lift to
drag in flight, used as the basic form of the wings, fins, and horizontal stabilizer of most
aircraft. To the super critical airflow, a separate Aerofoil is needed. In between 1960s
to 1970s, National Aeronautics had developed a new model of Aerofoil. It is described
as increase nose radius, flat upper surface, blunt tailing gate, modern supercritical
Aerofoil and cambered rear. CFD package is more useful to develop these kinds of
models. The purpose of designing of Aerofoil shape is to reduce the shock waves and
drag force. After the model is developed, it gains attractive demand for commercial
aircraft companies. Due to this, different types of models are developed by National
Aeronautics

1.4 Problem Definition

Day by day, the advancement in technologies is improving. Based on these, people’s
mindset and comfort levels are also changing. To travel from one place to another in
regular basis, the automobile vehicles, trains and flights plays a key role in
transportation. To run such type of vehicles, it needs fuels or electricity. Using this input
source, engine or motor are run and provides mechanical energy to provide forward and
backward movement. When the flight moves, wind will resist that movement which is
called drag force. Some lifting forces will induce due to the presence of air. When the
flight wing structure is improper, the drag force will act high on the wing. Due to this,
to overcome such effects, the engine performance should be more than usual. So, the
design of aircraft wings is an important factor for flight manufacturers. It depends on
aerodynamic characteristics on Wind in Ground effect planes. In this work,
aerodynamic analysis on NACA profile is planned through CFD. Lift force, drag force
and pressure readings are planned to measure with the help of ground clearance.

1.5 Aim and Objective

The main aim of this work is to analyses the aerodynamic characteristics like drag, lift
and pressure reading on aero foil. Once the wings angle of attack is changes, how the
lift force and drag force is changed and how to avoid the drag force on flight wing is a
main aim to consider and further work is extended based on this. The study on
importance of drag, lift coefficient and dynamic pressures are to be studied on real time
model. According to previous articles, new geometry models have to be planned to
reduce the resisting force on flight. The total project is depending upon ANSYS
FLUENT software. So, the main concentration is shown to simulation software named
FLUENT. In this, different types of analysis are available but among them, steady state
analysis is required to run such type of analysis. After performing analysis for all types
of models, output result plotting is required to easily understand and helpful for
comparison of results.

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 LITERATURE REVIEW
In this chapter, the details about previous research articles with regards to this work are
discussed. In past few years, work related to Aerofoil for both numerical and
experimental works were performed.

2.1 Numerical Work on Aerofoil Wing

The numerical research has been the primary topic of ground effect analysis over the
beginning of observing this occurrence. Saunders demonstrated contortions in
numerical and experimental analysis under the ground effect. Over different models, he
revealed that when the boundary condition is steady and unchanged, the numerical
value and the experimental value will differ. He repeated exact experimental methods
that maybe applied which are intended to validate the numerical values. The methods
that are determined to be ahead of fixed ground plane based on him are lineal
visualization with the towing technique and Aerofoil. In the end of 1970s, some
research studies were done base on fluid mechanics which is related to ground effect.
In this study, the more focus is on the high Reynolds number with low Mach numbers.
In this study, they worked in blunt body which is subjected to ground effect condition
by considering various high to chord ratios. It is mainly concentrated on extreme and
normal ground effect. In the extreme ground effect, lower lifting is acting on the wing
surface due to the absence of air.

➢ Coulliette.C study states that steady state aero foil ground effect is studied both
numerically and analytically. Discrete vortex and linear vortex panel methods are
applied to a parabolic arc and symmetric Joukowski aero foil, respectively. A single
vortex model for the flow over the parabolic arc aero foil is developed. The single
vortex model and other analytical solutions, valid either near or far from the ground,
are compared with the numerical results. The numerical results are used to delineate the
influences of angle of attack, camber and thickness. For small values of camber and
angle of attack, normalized lift is enhanced near the ground and reduced far from it. For
a fixed distance above the ground, normalized lift decreases with increasing angle of
attack and camber. Thickness reduces lift at all heights above the ground. The effect of
camber, attacking angle, thickness of Aerofoil with ground effect is important study
and many researches were done by Potkin. He concluded that lift forces increase as
ground clearance decreases and also the lift forces increase with the increase in Aerofoil
thickness. The vortices shed from Aerofoil in ground effect are a good study which was
done in 2003 by Fonseca. The author discussed about the aerodynamic effect and
pressure distribution with respect to ground.

➢ Nuhait performed the numerical analysis on steady and unsteady flow by

considering the ground effect which was done in 1989. The aerodynamic coefficient
was higher in unsteady flow when compared to steady flow. Based on this, some
numerical approaches were performed by using unsteady flow by considering the

MECHANICAL ENGINEERING DEPARTMENT 5

ground effect. This research was performed while flight landing and take-off positions
and aerodynamic coefficients are studied.

➢ Hsiun had done work on NACA 4412 profile Aerofoil with the help of
computational research. The CFD analysis is performed to analyse the effect of
Reynolds number on Aerofoil which is operating in ground effect. For CFD analysis,
PHEONICS code is used. In this work, turbulent model is considered with K-e turbulent
model. Initially, the work is taken from experimental work which is done by If the
Reynolds number increases, lift force increases. When the distance between the ground
and Aerofoil decreases, the drag force was decreased.

➢ Chang used experimental work and CFD. The CFD analysis was performed to
validate the experimental work. In this work, steady and moving ground plane effects
on the Aerofoil is considered to obtain accurate aerodynamic forces as same as the
experimental work. The lift force was same for two operating conditions and some
difference was observed due to the drag force. The drag force was higher when the
ground is in motion condition.

➢ Barber said that to study on the ground effect, ground plane should be in moving
condition. In his study, four types of analysis were done with same wind speed by
considering the ground plane as stationary, moving, slip and no slip. Among all the four
conditions, ground motion with same wind speed gives more accurate as same as the
experimental work.

➢ Saunders.G.H said the vehicle aerodynamics concerns the effects arising due to
motion of the vehicle through, or relative to, the air. Its importance to road vehicles
became apparent when they started to achieve higher speeds. The automobile as we
know it came onto the scene in the last decade of the nineteenth century. Its beginnings
roughly coincided with the advent of powered flight, and perhaps for this reason, it
became of interest to aerodynamicists right from the start. One of the first attempts to
apply aerodynamic principles to road vehicles was the streamlining.

➢ Tuck, E.O study says that flow induced by a body moving near a plane wall is
analysed on the assumption that the normal distance from the wall of every point of the
body is small compared to the body length. The flow is irrotational except for the vortex
sheet representing the wake. The gap-flow problem in the case of unsteady motion is
reduced to a nonlinear first-order ordinary differential equation in the time variable. In
the special case of steady flow, some known results are recovered and generalized. As
an illustration of the unsteady theory, the problem is solved of a flat plate falling toward
the ground under its own weight, while moving forward at uniform speed.

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2.2 Summary:
Hence by considering and studying the above journals we got an idea of how to use the
different aero foil profiles. In this era, we have got a clear idea on the manufacturing
the required aero foil profile to maintain minimum drag and maximum lift. In this
chapter, our detailed research on various research papers has been described. The inputs
we took from these papers help us in designing and fabricating the model according to
our requirements. These will be discussed in detail in subsequent chapters.

METHODOLOGY

Model Generation

CFD Simulation

Graph Generation

Results

MECHANICAL ENGINEERING DEPARTMENT 7

 3D MODELLING OF AEROFOIL USING SOLIDWORKS

Throughout the history of our industrial society, many inventions have been patented
and whole new technologies have evolved. Perhaps the single development that has
impacted manufacturing more quickly and significantly than any previous technology
is the digital computer. Computers are being used increasingly for both design and
detailing of engineering components in the drawing office.
Computer-aided design (CAD) is defined as the application of computers and graphics
software to aid or enhance the product design from conceptualization to documentation.
CAD is most commonly associated with the use of an interactive computer graphics
system, referred to as a CAD system. Computer-aided design systems are powerful
tools and in the mechanical design and geometric modelling of products and
components. There are several good reasons for using a CAD system to support the
engineering design

Function:
➢ To increase the productivity
➢ To improve the quality of the design
➢ To uniform design standards
➢ To create a manufacturing data base
➢ To eliminate inaccuracies caused by hand-copying of drawings andinconsistency
between
➢ Drawings

3.1 CAD Software

Software allows the human user to turn a hardware configuration into a powerful design
and manufacturing system. CAD/CAM software falls into two broad categories, 2-D
and 3-D, based on the number of dimensions are called 2-D representations of 3-D
objects is inherently confusing. Equally problem has been the inability of
manufacturing personnel to properly read and interpret complicated 2-D
representations of objects. 3-D software permits the parts to be viewed with the 3-D
planes-height, width, and depth-visible. The trend in CAD/CAM is toward 3-D
representation of graphic images. Such representation approximates the actual shape
and appearance of the object to be produced; therefore, they are easier to read and
understand.

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 CAD GENERATION

Creating an Aerofoil in SolidWorks using XYZ points from an Excel file involves
several steps. Below is a detailed guide to accomplish this:
Step 1: Generate or Obtain NACA Aerofoil Coordinates
1. Generate or download NACA Aerofoil coordinates:
We can generate the coordinates using various online tools or software. Ensure we have
a set of X, Y, Z coordinates (Z will be zero for a 2D Aerofoil).

Figure 3: NACA Aerofoil 23012

Figure 4: NACA Aerofoil 63-415

Step 2: Prepare Excel File

1. Create an Excel file:
- Arrange the Aerofoil coordinates in three columns: X, Y, and Z.

MECHANICAL ENGINEERING DEPARTMENT 9

 - Save the Excel file in a location accessible from SolidWorks.

Figure 5: Aerofoil Coordinates

Step 3: Import Coordinates into SolidWorks

1. Open SolidWorks.
2. Start a new part:
- Go to `File` > `New` > `Part`.
3. Insert a 3D sketch:
- Go to `Insert` > `3D Sketch`.
4. Import points from Excel:
- Go to `Tools` > `Curve Through XYZ Points`.
- In the dialogue box, click `Browse` to locate and select our Excel file.
- Ensure the data is mapped correctly (check the preview to ensure points are correct).
- Click `OK` to import the points as a curve.

Figure 6: Imported coordinate file

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Step 4: Create the Aerofoil Shape
1. Convert the imported curve to a sketch:
- Select the imported curve.
- Go to `Convert Entities` to create a sketch based on the curve.
- Alternatively, you can manually sketch over the imported points using a spline.

Figure 7: NACA Aerofoil 23012 curve

2. Refine the Aerofoil shape:

- Use the `Spline` tool to create a smooth curve through the imported points if
necessary.
- Make adjustments to the spline to ensure the Aerofoil shape is accurate.

Step 5: Create the 3D Model (Optional)

1. Extrude the Aerofoil shape:
- If you want a 3D Aerofoil, you can extrude the 2D sketch.
- Go to `Features` > `Extruded Boss/Base`.
- Set the desired extrusion depth and direction

Figure 8: NACA Aerofoil 23012 3-D model

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2. Complete the model:
- Add any additional features or details required for your Aerofoil design.

Step 6: Save our Work

1. Save the part:
- Go to `File` > `Save As`.
- Choose a location and file name for your SolidWorks part file.

Example Summary
1. Generate or obtain NACA Aerofoil coordinates.
2. Prepare an Excel file with X, Y, Z coordinates.
3. Open SolidWorks and start a new part.
4. Insert a 3D sketch and import the coordinates.
5. Convert the imported curve to a sketch and refine the shape.
6. Extrude the sketch if a 3D model is needed.
7. Save the work.

MECHANICAL ENGINEERING DEPARTMENT 12

 COMPUTATIONAL FLUID DYNAMICS

A computation fluid dynamic analysis on NACA 4412 Aerofoil model is studied in

this work; the geometrical parameters of Aerofoil wing are taken from National
Aeronautics Once the geometry model is done, the aerodynamic simulation is done to
find out the drag and lift coefficient. Here, 2D wing are taken and simulation is done
by changing the angle of attack of the flow. The simulation software is trending now
days in various industries which are mainly construction, shipping, automobile etc.
This simulation software is developed by different companies depending on CAD.
The companies are choosing and purchasing depending on cost, services and
maintenance. But this software can be classified into three categories such as drafting,
3D modelling design and analysis software. In this work, ANSYS software is used.

4.1 Benefits of CFD

To solve experimentally for fluid flow related problem is a complicated task. To solve
such problems, various kinds of sensors are required and at the same time air cannot be
seen. But in CFD it is possible. To test experimentally on Aerofoil, forming geometry
is complex work and also to find out the aerodynamic characters such as to detect lift
and drag is required many sensing sensors. This will cause huge burden to the
companies. So CFD is one of the sources to provide best results.

4.2 CFD Process

From the beginning of geometry till gaining results; there are different types of steps
included such as
1. Geometry formation related to the current work.
2. Applying meshing to the selected geometry.
3. Defining the materials
4. Choosing type of study either steady state or transient
5. Assigning material properties to cell zones
6. Applying boundary condition at inlet and outlet.
7. Requesting the output reports
8. Run the analysis
9. Gaining output results such as pressure, velocity, lift, drag and moments.

Initially, a 2-dimensional simulation is planned on Aerofoil so that NACA 23012

profile geometry is selected. To prepare the geometry, it needs x, y and z coordinate
data. These data are gathered from National Aeronautics. The data is initially in the
form of x, y. The preparing of geometry is planned in ANSYS software. To import the
data in terms of points, it should be set of formats such as group, point number, x, y and
z direction coordinates. Once entered into geometry, there is 3D point option in the
geometry. Through this option data can be imported with set of point. Using this data
points, a 3D curve is prepared. The profile chord length is 1m and maximum thickness
is 12%.

MECHANICAL ENGINEERING DEPARTMENT 13

ANSYS FLUENT PROCEDURE
Step 1: Open Ansys Workbench

Figure 9: Ansys Workbench

Step 2: Click ANSYS fluent

Figure 10: Ansys Fluid Fluent

Figure 11: NACA Aerofoil 23012 3-D model axes

MECHANICAL ENGINEERING DEPARTMENT 14

Step 3: Add Geometry

Figure 12: NACA Aerofoil 23012 3-D model

Figure 13: NACA Aerofoil 23012 3-D model

MECHANICAL ENGINEERING DEPARTMENT 15

Step 4: Make boundary box

Figure 14: NACA Aerofoil 23012 boundary box

Figure 15: NACA Aerofoil 23012 boundary box(2)

Step 5: Make mesh

Figure 16: NACA Aerofoil 23012 boundary conditions

MECHANICAL ENGINEERING DEPARTMENT 16

 Step 6: Add boundary condition

Figure 17: NACA Aerofoil 23012 boundary box (2)

Figure 18: NACA Aerofoil 23012 meshing

Figure 19: NACA Aerofoil 23012 results

MECHANICAL ENGINEERING DEPARTMENT 17

4.3.1 Geometric Model and Air Domain
Once the Aerofoil is formed in closed boundary, it is converted to surface. Air domain
geometry is critical to gain structure type of mesh. The front portion of Aerofoil is
created as semicircular arc and after that lines are prepared at rectangular shape. The
air domain length is taken as 20 times of chord length. The height of the Aerofoil is
12.5 times of Aerofoil chord length. The profile is closed and made in to surface. Using
Boolean operation, subtract option and Aerofoil surface is subtracted from air domain
boundary which is shown in the below image.

Figure 20: Geometry model and air domain.

4.3.2 Structured Mesh on Aerofoil Domain

Once the geometry is done, the next step is meshing. Meshing is an important
consideration in this work. Different shapes of element types are available in the
software. Depending on processing time, accurate result, the element shape must be
selected. The main concentration is on the around Aerofoil boundary shape. Using these
considerations, the meshing is planned. In this structured mesh, meshing is done to form
higher order quadratic element. Initially edge meshing is done to prepare mesh planning
and bios control is used with the factor of 20 because to place maximum concentration
on Aerofoil bounded area.

Figure 21: Structured mesh on Aerofoil domain.

MECHANICAL ENGINEERING DEPARTMENT 18

4.3.3 Aero foil With Detail View

Figure 22: Aero foil with detail view.

Once the planning of meshing is done, assigning the boundary is an important task. In
these boundaries, inlet name is given to send the air flow and the outlet name is given
to send the air exist from the boundary. The Aerofoil wall name is given to Aerofoil
region because the main work is to gain the aerodynamic results around the Aerofoil.
The lower part of the boundary is named as ground and the upper part of the boundary
is named as free surface.

4.3.4 Boundary Conditions on Air Domain

Figure 23: Boundary conditions on air domain.

Once meshing and naming for boundary condition is done, fluid simulation is the next
procedure. They are various kind of software available to performed fluid simulation.
Simulation software such as ANSYS FLUENT, ANSYS CFX, OPENFORM are
available in the market. Depending on the availability ANSYS FLUENT software is
selected. In this software, 2D simulation is performed. This software is divided into
general, model, cell zones, material, boundary condition, report, report request,
initialization and run analysis.

MECHANICAL ENGINEERING DEPARTMENT 19

4.4.4 Turbulent Model
Once entered in FLUENT software, in general; pressure based 2D planner is selected.
There is no gravity consideration is taken. The pressure-based solver is considered due
to the drag and lift produced due pressure variation around the Aerofoil. For other tasks,
gravity is important. But for this particular problem there is no requirement of gravity.
The 2D problems are characterized into two types. They are axis-symmetrical model
and planner mode. In axis-symmetric, round shaped air domain section is taken for the
analysis. But here, full model is considered from front portion of the Aerofoil to outside
of the Aerofoil as air domain and not considering the axis. The domain is taken as
planner.
In the model portion, turbulent model option is selected and energy is not taken as
consideration as this work is not related to thermal problem. Different types of turbulent
model available in the software. Depending upon the accurate results are obtained at
vertices points and processing time. In this analysis, K-E turbulence model provides
good agreement with experimental results.

Figure 24: Turbulent model selected for analysis.

The air is taken as a material of the fluid and velocity of 30m/s is given to the inlet
condition which is taken as constant throughout the entire work. Using this velocity
coefficient of lift and drag is also calculated. The outlet boundary condition is taken as
pressure outlet. Here the pressure value is 0 Pa applied at the pressure outlet condition.
The software will work on the pressure difference; in inlet velocity is given that is why
there will be more pressure at the inlet region. The negligible pressure is place at the
outlet. Using these differences, air flows from inlet to outlet. The Aerofoil wall is places
between inlet and outlet, the air will hit the foil, through this, and pressure variations
along the boundaries are induced. Using report option in FLUENT we can gain the
pressure results. In reference condition, reference area, density, velocity, length and
pressure values are entered. These parameters will affect aerodynamic forces. At the
inlet, initialization is started and the problem is solved in steady state condition

MECHANICAL ENGINEERING DEPARTMENT 20

pressure. Nearly 500 iterations were run. As there is structured meshed, at 100
iterations, the problem is solved. Once these steps are completed, problem must
converge. To converge, for pressure, moment and velocity, convergence criteria of 10-
5 is given to obtain accurate results. The convergence criteria mean error residuals
between two iterations.

4.5 Results of Flight Model (Distributions)

The simulation was done by considered the ground motion. The bottom wall of the air
domain is named as ground. The problem is steady state problem, the wind velocity and
ground moving wall velocity is taken same as 30m/s. The pressure outlet of the domain
is taken as 0. As same as 2D analysis, drag force and lift force for various ground
clearance distance are analysed. The problem is solved until iterated reached the
convergence 10-5.

4.5.1 Distributions of NACA 23012 at 0 AOA

4.5.1.1 Velocity Distribution

Figure 25: Velocity distribution around Aerofoil NACA 23012 at 0° AOA.

The image above is velocity distribution around Aerofoil at 0 AOA is shown above.
The red colour in the image indicates maximum velocity distribution and the blue
colour indicates minimum velocity distribution. The maximum velocity distribution is
observed at the top surface of the Aerofoil. The minimum velocity distribution is
observed at the tip end of the air foil. When the air flows through the air foil, the velocity
is high at the top and after the tip end of the Aerofoil, the turbulence flows through it
continuously. The maximum velocity value is 3.82e+01 and the minimum velocity
value is 0.00e+00.

4.5.1.2 Pressure Distribution

The pressure distribution around Aerofoil at 0o angle of attack is shown in the image
above. The maximum pressure distribution is displayed in red colour, whereas the
minimum pressure distribution is displayed in blue colour. From the image, the
maximum velocity is noticed at the starting tip of the Aerofoil where the pressure
distribution value is 3e+02.

MECHANICAL ENGINEERING DEPARTMENT 21

 Figure 26: Pressure distribution around aero foil NACA 23012 at 0° AOA.

The minimum pressure distribution is noticed at the top surface of the Aerofoil at 0o
angle of attack and the minimum pressure value is -3.85 e+02. When the air passes
through the Aerofoil at 0 AOA, the pressure moved up throughout the top surface of
the Aerofoil. Which means, the air foil moves upward at 0o angle of attack.

4.3.2 Distributions of NACA 23012 at 18 AOA

4.3.2.1 Velocity Distribution

Figure 27: Velocity distribution around Aerofoil NACA 23012 at 18° AOA.

The velocity distribution on Aerofoil at 18 AOA is observed at above image. The

maximum velocity displayed in red colour which is at half of the top surface. The
minimum velocity displayed in blue colour and it is noticed at the bottom tip of the
Aerofoil and tip end of the Aerofoil. When the air flows through the Aerofoil, the
velocity is moves up and the slowly turbulence reaches the end of the Aerofoil tip and
moves continuously. The maximum velocity value is 4.26 e+01. The minimum velocity
value is 0.00e+00.

MECHANICAL ENGINEERING DEPARTMENT 22

4.3.2.2 Pressure Distribution

Figure 28: Pressure distribution around aero foil NACA 23012 for 18° AOA.

The image above represents the pressure distribution around Aerofoil at 18o angle of
attack. The red colour in the images indicates the maximum pressure distribution;
whereas the blue colour in the image indicates minimum pressure distribution. The
maximum velocity is noticed at the bottom front tip of the Aerofoil where the pressure
distribution value is 4.16 e+02. The minimum pressure distribution is noticed at the top
front tip surface of the Aerofoil at 18o angle of attack and the minimum pressure value
is -9.29e+03. When the air flows through the Aerofoil at 18 AOA, the pressure slightly
moved up at the starting top surface of the Aerofoil. That means, the air foil moves
slightly upward and holds at 18o angle of attack.

4.3.3 Distributions of NACA 23014 at 0 AOA

4.3.3.1 Velocity Distribution
The image above is velocity distribution around Aerofoil at 0 AOA is shown above.
The red colour in the image indicates maximum velocity distribution and the blue
colour indicates minimum velocity distribution.

Figure 29: Velocity distribution around Aerofoil NACA 23014 at 0° AOA.

MECHANICAL ENGINEERING DEPARTMENT 23

The maximum velocity distribution is observed at the top surface of the Aerofoil. The
minimum velocity distribution is observed at the tip end of the air foil. When the air
flows through the air foil, the velocity is high at the top and after the tip end of the
Aerofoil, the turbulence flows through it continuously. The maximum velocity value is
3.37e+01 and the minimum velocity value is 0.00e+00.

4.3.3.2 Pressure Distribution

Figure 30: Pressure distribution around aero foil NACA 23014 at 0° AOA.

The pressure distribution around Aerofoil at 0o angle of attack is shown in the image
above. The maximum pressure distribution is displayed in red colour, whereas the
minimum pressure distribution is displayed in blue colour. From the image, the
maximum velocity is noticed at the starting tip of the Aerofoil where the pressure
distribution value is 3.96e+02. The minimum pressure distribution is noticed at the top
surface of the Aerofoil at 0o angle of attack and the minimum pressure value is -4.25
e+02. When the air passes through the Aerofoil at 0 AOA, the pressure moved up
throughout the top surface of the Aerofoil. Which means, the air foil moves upward at
0o angle of attack.

4.3.4 Distributions of NACA 23014 at 18 AOA

4.3.4.1 Velocity Distribution

Figure 31: Velocity distribution around Aerofoil NACA 23014 at 18° AOA.

MECHANICAL ENGINEERING DEPARTMENT 24

The velocity distribution on Aerofoil at 18 AOA is observed at above image. The
maximum velocity displayed in red colour which is at half of the top surface. The
minimum velocity displayed in blue colour and it is noticed at the bottom tip of the
Aerofoil and tip end of the Aerofoil. When the air flows through the Aerofoil, the
velocity is moves up and the slowly turbulence reaches the end of the Aerofoil tip and
moves continuously. The maximum velocity value is 3.7 e+01. The minimum velocity
value is 0.00e+00

4.3.4.2 Pressure Distribution

The figure 4.3.4.2 represents the pressure distribution around Aerofoil at 18o angle of
attack. The red colour in the images indicates the maximum pressure distribution;
whereas the blue colour in the image indicates minimum pressure distribution. The
maximum velocity is noticed at the bottom front tip of the Aerofoil where the pressure
distribution value is 4.4 e+02. The minimum pressure distribution is noticed at the top
front tip surface of the Aerofoil at 18o angle of attack and the minimum pressure value
is 0. When the air flows through the Aerofoil at 18 AOA, the pressure slightly moved
up at the starting top surface of the Aerofoil. That means, the air foil moves slightly
upward and holds at 18o angle of attack.

Figure 32: Pressure distribution around aero foil NACA 23014 for 18° AOA.

4.3.5 Distributions of NACA 63415 at 0 AOA

4.3.5.1 Velocity Distribution

Figure 33: Velocity distribution around Aerofoil NACA 63415 at 0° AOA

MECHANICAL ENGINEERING DEPARTMENT 25

The velocity distribution on Aerofoil at 0 AOA is observed at above image. The
maximum velocity displayed in red colour which is at half of the top surface. The
minimum velocity displayed in blue colour and it is noticed at the bottom tip of the
Aerofoil and tip end of the Aerofoil. When the air flows through the Aerofoil, the
velocity is moves up and the slowly turbulence reaches the end of the Aerofoil tip and
moves continuously. The maximum velocity value is 3.76 e+01. The minimum velocity
value is 0.00e+00.

4.3.5.2 Pressure Distribution

Figure 34: Pressure distribution around aero foil NACA 63415 at 0° AOA.

The pressure distribution around Aerofoil at 0o angle of attack is shown in the image
above. The maximum pressure distribution is displayed in red colour, whereas the
minimum pressure distribution is displayed in blue colour. From the image, the
maximum velocity is noticed at the starting tip of the Aerofoil where the pressure
distribution value is 3.63e+02. The minimum pressure distribution is noticed at the top
surface of the Aerofoil at 0o angle of attack and the minimum pressure value is –
3.72e+02. When the air passes through the Aerofoil at 0 AOA, the pressure moved up
throughout the top surface of the Aerofoil. Which means, the air foil moves upward at
0o angle of attack.

4.3.6 Distributions of NACA 63415 at 18 AOA

4.3.6.1 Velocity Distribution

Figure 35: Velocity distribution around Aerofoil NACA 63415 at 18° AOA.

MECHANICAL ENGINEERING DEPARTMENT 26

The 00-velocity distribution on Aerofoil at 18 AOA is observed at above image. The
maximum velocity displayed in red colour which is at half of the top surface. The
minimum velocity displayed in blue colour and it is noticed at the bottom tip of the
Aerofoil and tip end of the Aerofoil. When the air flows through the Aerofoil, the
velocity is moves up and the slowly turbulence reaches the end of the Aerofoil tip and
moves continuously. The maximum velocity value is 4.44 e+01. The minimum velocity
value is 0.00e+00.

4.3.6.2 Pressure Distribution

Figure 36: Pressure distribution around aero foil NACA 63415 at 18° AOA.

The pressure distribution around Aerofoil at 18o angle of attack is shown in the image
above. The maximum pressure distribution is displayed in red colour, whereas the
minimum pressure distribution is displayed in blue colour. From the image, the
maximum velocity is noticed at the starting tip of the Aerofoil where the pressure
distribution value is 4.86e+02. The minimum pressure distribution is noticed at the top
surface of the Aerofoil at 18o angle of attack and the minimum pressure value is -
1.07 e+03. When the air passes through the Aerofoil at 18 AOA, the pressure moved
up throughout the top surface of the Aerofoil. Which means, the air foil moves upward
at 0o angle of attack.

MECHANICAL ENGINEERING DEPARTMENT 27

4.4 Results of Flight Model (Coefficients)
4.4.1 Drag Coefficient Plot for NACA 23012

Figure 37: Drag coefficient plot for NACA 23012

Table 1: Drag coefficient plot for NACA 23012

AOA NACA 23012

0 0.02
2 0.01
4 0.01
6 0.00
8 -0.01
10 -0.03
12 -0.04
14 -0.05
16 -0.05
18 -0.06

The above table shows the drag coefficient plot for NACA 23012. In this plot the x axis
shows iteration number and y axis shows drag coefficients. At the starting of the
processing time, air is hitting the flight and will create turbulence and a curve
fluctuation is observed at the beginning of few iterations. After some iteration, it
becomes idle. It continues till the problem is converged. The coefficient of drag reading
0.015 is noted.

MECHANICAL ENGINEERING DEPARTMENT 28

4.4.2 Lift Coefficient Plot for NACA 23012

Figure 38: Lift coefficient plot for NACA 23012

Table 2: Lift coefficient plot for NACA 23012

AOA ACA 23012

0 0.12
2 0.23
4 0.33
6 0.42
8 0.51
10 0.59
12 0.65
14 0.70
16 0.74
18 0.75

The lift coefficient plot for NACA 23012 is shown in the above table. In this plot the x
axis represents iteration number and y axis represents lift coefficients. At the starting
of the processing time, air is hitting the flight and will create turbulence and a curve
fluctuation is observed at the beginning of few iterations. After some iteration, it

MECHANICAL ENGINEERING DEPARTMENT 29

becomes idle and proceeds till the problem is converged. The coefficient of lift reading
0.12 is noted.

4.4.3Drag Coefficient Plot for NACA 23014

Figure 39: Drag coefficient plot for NACA 23014

Table 3: Drag coefficient plot for NACA 23014

AOA NACA 23014

0 0.02
2 0.01
4 0.01
6 0.00
8 -0.01
10 -0.03
12 -0.04
14 -0.05
16 -0.06
18 -0.06

The above table shows the drag coefficient plot for NACA 23014. In this plot the x axis
shows iteration number and y axis shows drag coefficients. At the starting of the
processing time, air is hitting the flight and will create turbulence and a curve
fluctuation is observed at the beginning of few iterations. After some iteration, it
becomes idle. The coefficient of drag reading 0.015 is noted.

MECHANICAL ENGINEERING DEPARTMENT 30

4.4.4 Lift Coefficient Plot for NACA 23014

Figure 40: Lift coefficient plot for NACA 23014

Table 4: Lift coefficient plot for NACA 23014

AOA NACA 23014

0 0.15
2 0.25
4 0.34
6 0.44
8 0.52
10 0.60
12 0.66
14 0.71
16 0.74
18 0.76

The lift coefficient plot for NACA 23014 is shown in the above table. In this plot the x
axis represents iteration number and y axis represents lift coefficients. At the starting
of the processing time, air is hitting the flight and will create turbulence and a curve
fluctuation is observed at the beginning of few iterations. After some iteration, it
becomes idle and proceeds till the problem is converged. The coefficient of lift reading
0.12 is noted.

MECHANICAL ENGINEERING DEPARTMENT 31

4.4.5 Drag Coefficient Plot for NACA 63415

Figure 41: Drag coefficient plot for NACA 63415

Table 5: Lift coefficient plot for NACA 63415

AOA NACA 63415

0 0.02
2 0.01
4 0.00
6 -0.02
8 -0.03
10 -0.05
12 -0.07
14 -0.08
16 -0.09

The above table shows the drag coefficient plot for NACA 63415. In this plot the x axis
shows iteration number and y axis shows drag coefficients. At the starting of the
processing time, air is hitting the flight and will create turbulence and a curve
fluctuation is observed at the beginning of few iterations. After some iteration, it
becomes idle. It continues till the problem is converged. The coefficient of drag reading
-0.09 is noted.

MECHANICAL ENGINEERING DEPARTMENT 32

4.4.6 Lift Coefficient Plot for NACA 63415

Figure 42: Lift coefficient plot for NACA 63415

Table 6: Lift coefficient plot for NACA 63415

AOA NACA 63415

0 0.29
2 0.41
4 0.52
6 0.62
8 0.72
10 0.79
12 0.85
14 0.91
16 0.92
18 0.90

The lift coefficient plot for NACA 63415 is shown in the above table. In this plot the x
axis represents iteration number and y axis represents lift coefficients. At the starting
of the processing time, air is hitting the flight and will create turbulence and a curve
fluctuation is observed at the beginning of few iterations. After some iteration, it
becomes idle and proceeds till the problem is converged. The coefficient of lift reading
0.9 is noted.

MECHANICAL ENGINEERING DEPARTMENT 33

 RESULTS AND DISCUSSION

A numerical CFD analysis of NACA profiles is undertaken in this work. ANSYS Fluent
software is used to do the analysis. NACA23012, NACA 23014, and NACA 63415 are
considered 2D profiles at first. The analysis is done up to 18° AOA to find the optimal
aero foil. In tabular columns and graphs, the individual profile drag and lift coefficient
are recorded.

5.1 Drag Coefficient

Table 7: Aerofoil drag coefficient comparison

AOA NACA 23012 NACA 23014 NACA 63415

0 0.02 0.02 0.02
2 0.01 0.01 0.01
4 0.01 0.01 0.00
6 0.00 0.00 -0.02
8 -0.01 -0.01 -0.03
10 -0.03 -0.03 -0.05
12 -0.04 -0.04 -0.07
14 -0.05 -0.05 -0.08
16 -0.05 -0.06 -0.09
18 -0.06 -0.06 -0.09

The table shown in the above is the drag coefficient of AOA for three different profiles.
10 Angle of attack for three different NACA profiles which are NACA 23012; NACA
23014 and NACA 63415 are shown. The angle of attack ranges from 0 to 18.

0.04
0.02
0.00 NACA 23012
10 15 20
0
-0.02 NACA 23014
-0.04 NACA 63415
-0.06
-0.08
AOA

Graph.1: Drag coefficient of NACA 23012, 23014, 63415.

The graphical representation shown above is the drag coefficient for different range of
angle of attack for three different NACA profiles. The graph shows that NACA 23012

MECHANICAL ENGINEERING DEPARTMENT 34

is indicated in blue colour, the NACA 23014 is indicated in red colour and finally, the
NACA 63415 is indicated in green colour. The x direction indicates angle of attack and
y axis indicates drag coefficient. From the graph, it is seen that the drag coefficient will
higher at low angle of attack and low at higher angle of attack. The maximum drag
coefficient is same for all profiles which is 0.02. At AOA for 18, the minimum drag
coefficient value is -0.06 for NACA 23012 and NACA 23014 and -0.09 for NACA
63415.

5.2 Lift Coefficient

Table 8: Lift coefficient of NACA 23012, 23014, 63415

AOA NACA 23012 NACA 23014 NACA 63415

0 0.12 0.15 0.29
2 0.23 0.25 0.41
4 0.33 0.34 0.52
6 0.42 0.44 0.62
8 0.51 0.52 0.72
10 0.59 0.60 0.79
12 0.65 0.66 0.85
14 0.70 0.71 0.91
16 0.74 0.74 0.92
18 0.75 0.76 0.90

The table shown in the above is the lift coefficient of AOA for three different profiles.
10 Angle of attack for three different NACA profiles which are NACA 23012; NACA
23014 and NACA 63415 are shown. The angle of attack ranges from 0 to 18.

1.00

0.80

0.60
NACA 23012
0.40
NACA 23014
0.20 NACA 63415

0.00
10 15 20

Graph 2: Lift coefficient of NACA 23012, 23014, 63415

MECHANICAL ENGINEERING DEPARTMENT 35

The graphical representation shown above is the drag coefficient for different range of
angle of attack for three different NACA profiles. The graph shows that NACA 23012
is indicated in blue colour, the NACA 23014 is indicated in red colour and finally, the
NACA 63415 is indicated in green colour. The x direction indicates angle of attack and
y axis indicates drag coefficient. From the graph, the lift coefficient increases with
increase in AOA for all profiles. The maximum lift coefficient is observed in NACA
63415. Minimum lift coefficient is observed in NACA 23012.

5.3 L/d RATIO FOR AERO FOILS

The lift by drag coefficient for different angle of attack for three different profiles is
shown in the table 5.3. The three different NACA profiles are NACA 23012, NACA
23014 and NACA 63415 are shown table 5.3.

Table 9: L/d ratio of NACA 23012, 23014, 63415

NACA 23012 NACA 23014 NACA 63415

0 8.04717104 8.046310488 19.02597
2 19.3250654 17.08244074 48.87556
4 60.5619835 43.99621285 -203.997
6 -123.512971 -292.035932 -36.6264
8 -35.6405261 -40.1755196 -20.9754
10 -22.1409285 -22.9681397 -14.9986
12 -17.1322925 -16.8783572 -12.5147
14 -15.0100945 -14.6914398 -11.0869
16 -13.6322563 -13.0152946 -10.3758
18 -13.1640365 -12.2056188 -9.93901

100
50

10 15 20 NACA 23012
-50
NACA 23014
0
NACA 63415
-100
-150
-200
-250
-300
Graph 3: L/d ratio of NACA 23012, 23014, 63415

MECHANICAL ENGINEERING DEPARTMENT 36

The graphical represent shown above is the l/d ratio with respect to angle of attack for
three different NACA profile is shown above. The graph shows that NACA 23012 is
indicated in blue colour, the NACA 23014 is indicated in red colour and finally, the
NACA 63415 is indicated in green colour. The x direction indicates angle of attack and
y axis indicates l/d ratio. From the image above, the maximum l/d ratio is seen in NACA
23012 when compared to other and minimum l/d ratio is seen in NACA 23014 when
compared to others.

MECHANICAL ENGINEERING DEPARTMENT 37

 CONCLUSIONS
The selected profiles 23012, 23014, 63415 are designed in CATIA software and
simulated in ANSYS software. The final conclusions made after the above simulation
and analysis of the selected profiles are -:
•The form of the Aerofoil is regarded one of the most essential aspects affecting an
aircraft's basic flight, it's critical that the Aerofoil utilized in its design delivers more
lift than drag.
•The lift and drag supplied by the aero foil are significantly reliant on the volume of air
that passes through its form, hence the lift and drag coefficients of the aero foil are
studied under various NACA configurations.
•The NACA 63415 Aerofoil offers higher lift than the NACA 23012 and NACA 23014
Aerofoils, according to the calculations and measurements above. For the airplane
employing an Aerofoil design for its wing to be more efficient, the drag should be
reduced.
•The drag produced by the NACA 63415 Aerofoil is less than that of the NACA 23012
and NACA 23014 Aerofoils. According to this argument, the NACA 63415 Aerofoil is
superior to the NACA 23012 series Aerofoil.
•The lift and drag forces over a wing of the NACA type's Aerofoil shape were estimated
via fluent testing.
• Secondly, the lift and drag coefficients were calculated across a range of angles of
attack, providing a comprehensive understanding of the aerofoil's performance
characteristics. By analyzing the lift-to-drag ratio, we identified the optimal angle of
attack for maximizing aerodynamic efficiency, crucial for applications such as aircraft
design and performance optimization.
• In conclusion, this CFD analysis has provided valuable insights into the aerodynamic
performance of the aerofoil, enhancing our understanding of its behavior under
different flow conditions. The findings presented here can serve as a foundation for
further research and development in aerodynamics, with potential applications in
aerospace engineering, wind energy, and beyond.

6.1 Future Scope

•Model analysis can be performed on wings with various profiles to determine the
natural frequencies of the wings.
•CFD flow analysis on Aerofoil profiles with variable tail sections can be study to
meet the practical take-off and landing conditions of the planes.

MECHANICAL ENGINEERING DEPARTMENT 38

 REFERENCES
1. Coulliette, C. and A. Plotkin. “Aerofoil ground effect revisited.” Aeronautical
Journal, February 1996, pp. 65-74.

2. Nuhait, A.O., and M.F. Zedan. “Numerical Simulation of Unsteady Flow

Induced by a Flat Plate Moving Near Ground.” Journal of Aircraft, Vol. 30, No.
5, 1993, pp. 611-617.

3. C. K. Chen and Hsiun, C.H., “Aerodynamic Characteristics of a Two-

dimensional Aerofoil with Ground Effect.” Journal of Aircraft, Vol. 33, No. 2,
1996, pp. 386-392.

4. Steinbach, D. “Comment on Aerodynamic Characteristics of a Two-

Dimensional Aerofoil with Ground Effect.” Journal of Aircraft, Vol. 34, No. 3,
1997, pp. 455-456.

5. R.H. Chang. And Chun, H.H. “Turbulence Flow Simulation for Wings in
Ground Effect with Two Ground Conditions: Fixed and Moving Ground.”
International Journal of Maritime Engineering, 2003, pp. 211-228.

6. J. Leonardi, Barber, T. E., and R.D. Archer. “A Technical Note on the

Appropriate CFD Boundary Conditions for the Prediction of Ground Effect
aerodynamics.” The Aeronautical Journal, 1999, pp. 545-547.

7. Saunders, G.H. “Aerodynamic Characteristics of Wings in Ground Proximity.”

Canadian Aeronautics and Space Journal, 1965, pp. 185-192.

8. Tuck, E.O. “A Nonlinear Unsteady One-Dimensional Theory for Wings in

Extreme Ground Effect.” Journal of Fluid Mechanics, Vol. 98, Part 1, 1980, pp.
33- 47.

9. Handbook,F.,2007.AvStop[Online]Available
http://avstop.com/ac/flighttrainghandbook/images4r.jpg

10. .Scott,2005.AerospaceWeb.[Online]Available
http://www.aerospaceweb.org/question/nature/q0237.shtml[

MECHANICAL ENGINEERING DEPARTMENT 39

 STUDENTS AND GUIDE DETAILS

Student Information:

Name: Anshu Kumar Shinde

Seat No: B191090932
Class: BE (B)
PRN: 72159433G
Contact No: 7273047677
Email ID: anshushinde9868@gmail.com

Name: Indraneil Arup Sinha

Seat No: B191090936
Class: BE (B)
PRN: 72159460D
Contact No: 8308811859
Email ID: neil7041@gmail.com

Name: Shubham Kishor Thokare

Seat No: B191090948
Class: BE (B)
PRN: 72159500G
Contact No: 8830382874
Email ID: shubhamthokare18@gmail.com

Name: Kush Chandrakant Tilgule

Seat No: B191090949
Class: BE (B)
PRN: 72159503M
Contact No: 8999354215
Email ID: ktilgule@gmail.com

Name: Siddhant Sonbhau Adak

Seat No: B191090969
Class: BE (A)
PRN: 72158975G
Contact No: 9082777251
Email ID: siddhantadak@gmail.com

Project Guide:
Name: Suchit Deshmukh
Contact No: 9096393605
Email ID: suchit.deshmukh@dypiemr.ac.in

MECHANICAL ENGINEERING DEPARTMENT 40

 PLAGIARISM REPORT

MECHANICAL ENGINEERING DEPARTMENT 41