This research work is concerned with the application of conceptual design of Unmanned Air Vehicle (UAV). UAV is used for surveillance and reconnaissance to serve for the defense as well as national security and intelligence purpose. Here NACA 0012 aerofoil profile is used to design UAV by using CFD (Computational Fluid Dynamics) software. The aim of this research is to investigate the flow patterns and determine the aerodynamic characteristics of NACA 0012 profile by varying the angle of attack and Reynolds Number numerically. The research is carried out with symmetric aerofoil with the chord length of 0.1m. The research work explained different aerodynamic characteristics like lift force and drag force, lift and drag coefficient, pressure distribution over aerofoil etc.

1. DESIGN ANALYSIS OF UAV (UNMANNED AIR VEHICLE) USING NACA 0012 AEROFOIL PROFILE
Alimul Rajib1, Bhuiyan Shameem Mahmud Ebna Hai1 and Md Abdus Salam2
1
Department of Mechanical Engineering, Military Institute of Science and Technology, Dhaka-1216, Bangladesh.
2
Department of Aeronautical Engineering, Military Institute of Science and Technology, Dhaka-1216, Bangladesh.
ABSTRACT
This research work is concerned with the application of conceptual design of Unmanned Air Vehicle (UAV).
UAV is used for surveillance and reconnaissance to serve for the defense as well as national security and
intelligence purpose. Here NACA 0012 aerofoil profile is used to design UAV by using CFD (Computational
Fluid Dynamics) software. The aim of this research is to investigate the flow patterns and determine the
aerodynamic characteristics of NACA 0012 profile by varying the angle of attack and Reynolds Number
numerically. The research is carried out with symmetric aerofoil with the chord length of 0.1m. The research
work explained different aerodynamic characteristics like lift force and drag force, lift and drag coefficient,
pressure distribution over aerofoil etc
Keywords: UAV, CFD, NACA 0012.
The lift force increases almost linearly with the angle of
1. INTRODUCTION
An Unmanned Air Vehicle (UAV) is an unpiloted attack until a maximum value is reached where upon
aircraft. Its aerodynamic characteristics vary with the wing is said to stall. The shape of the drag force vs.
certain parameters like the angle of attack and others. angle of attack is approximately parabolic. It is
Experimental works on UAVs have been conducted in desirable for the wing to have the maximum lift and
many places with various aerofoil profiles but not smallest possible drag.
enough work with the Computational Fluid Dynamics
(CFD) analysis is not available that much till now. The 1.1 BACKGROUND OF THE RESEARCH WORK
present work contains mainly CFD analysis to Designing of UAVs requires designing of aerofoil
determine the flow pattern and the aerodynamic section. Various aerofoil configurations have been
characteristics of an UAV. employed so far and more will be coming. The present
The shape of an aircraft is designed to make the airflow work is carried out numerically with CFD analysis for
through the surface to produce a lifting force in the NACA 0012 symmetric aerofoil profile. Some of the
most efficient manner. In addition to the lift, a force parameters of aerofoil and properties of air have been
directly opposing the motion of the wing through the kept constant and some have been varied.
air is always present, which is called a drag force. The The flow of air over the aerofoil is varied as per
angle between the relative wind and the chord line is requirement. The chord length of the aerofoil is
the angle of attack of the aerofoil. The lift and drag 100mm. The free stream airflow has been kept 12.5 m/s
forces developed by an aircraft will vary with the and the effect of the temperature in the study has been
neglected. The density of air (ρo)= 1.22 kg/m3,
change of angle of attack. The cross sectional shape
obtained by the intersection of the wing with the operating pressure (Po) = 0.101 MPa (1.01 bar) and
absolute viscosity (μ) = 1.789 x 10-5 kg/m-s. The
perpendicular plane is called an aerofoil. Here NACA
0012 symmetric aerofoil profiles have been used for the Reynolds Number has been considered as variable. The
present research work. data have been obtained at different angles of attack
starting from 0o with 1o incremental step.
The various measurement characteristics such as
pressure distribution, pressure contours, Mach number,
etc. around a two dimensional aerofoils of UAV varies
with the angle of attack.
The aerodynamic characteristics of a typical aircraft can
also be experimentally investigated in the wind tunnels.
The surface static pressure is measured from the suction
and the pressure side of the aerofoil through different
pressure tapping points. The aerodynamic
characteristics for different configurations are
determined from the static pressure distribution over
the surfaces of aerofoils at different angles of attack.
Fig 1: Aerodynamic forces on a typical aerofoil

2. 1.2 OBJECTIVES 2.2 AEROFOIL DESIGN
a. Designing of NACA 0012 aerofoil section and The vertices obtained from the C program were used to
investigation of the flow pattern with the help of draw the profile line which was as follows:
CFD software.
b. Determination of the surface static pressure
distribution, pressure contours, Mach number on
the aerofoils in the biplane configuration.
c. Determination of the aerodynamic characteristics
from the static pressure distributions.
d. Discussion on the computational results of the
Fig 3: NACA 0012 aerofoil section.
CFD analysis.
The boundary was then given.
2. WORKING PRINCIPAL
The computation and graphical plotting involves the
Table 1: Values of boundary vertices for NACA 0012
following sequence:
aerofoil profile.
a. Programming to get vertices for aerofoil section
using governing equation.
Label X Y Z
b. Working with vertices using GAMBIT software.
A 0.1 1.25 0
c. Working with FLUENT software.
B 2.1 1.25 0
C 2.1 0 0
2.1 DESIGN METHOD
D 2.1 -0.25 0
The early NACA aerofoil series, the 4-digit was
E 0.1 -0.25 0
generated using analytical equations that describe its
F -1.15 0 0
geometrical feature.
G 0.1 0 0
The boundaries were chosen such to get uniform
meshing. Line and face both the meshing were
employed here. To mesh, interval counts and successive
ratios were used here.
After employing boundary and meshing the following
meshed geometry was found.
Fig 2: NACA aerofoil geometrical construction.
The first digit specifies the maximum camber (m) in
percentage of the chord, the second indicates the
position of the maximum camber (p) in tenths of chord,
and the last two numbers provide the maximum
thickness (t) of the aerofoil in percentage of chord. So,
our concerned NACA 0012 aerofoil means 0% camber
at 0 (zero) position (as there is no camber) and
thickness of .012m. The thickness distribution above
(+) and below (-) the mean line was calculated by
plugging the value of t into the following equation for
each of the x coordinates.
 
t
Fig 4: Meshing of aerofoil section (2D).
 yt  x  0 . 1260 x  0 . 3516 x 2  0 . 2843 x 3  0 . 1015 x 4
0 . 2969
0 .2
It was recommended that the boundaries around the
The equation was solved here by a C program to find
aerofoil were far enough.
vertices for the aerofoil line. Approximately, 10,000
vertices were used.
2.3 Analysis of Data in Fluent
The mesh file was imported to the Fluent and it
required certain features. First of all, 2-D mode was
selected. The parameters which needed to be constant
were: pressure (atmospheric pressure = 101325 Pa), air
velocity (v = 12.5 m/s), density of air (ρ = 1.225
kg/m3), absolute viscosity (μ = 1.789 x 10-5 kg/m-s).
The Reynolds number and Mach number were kept
constant and sometimes varied as per requirement.

3. 3. RESULTS AND GRAPHS 3.4 Angle of attack Vs Drag coefficient:
3.1 Lift coefficient (CL)
Sample results of lift coefficient CL with variable angle
0.025
of attack α and Reynolds number Re are as follows:
Table3. Values of CL: 0.02
Lift coefficient
Re CL 0.015
α 3.6 x 105 7 x 105 1 x 106 2 x 106 5 x 106 0.01
0 0.0000 0.0000 0.0000 0.0000 0.0000 0.005
2 0.2200 0.2200 0.2200 0.2200 0.2200 0
4 0.4400 0.4400 0.4400 0.4400 0.4400 0 2 4 6 8 10 12 14
6 0.6600 0.6600 0.6600 0.6600 0.6600 Angle of attack
8 0.8542 0.8800 0.8800 0.8800 0.8800 Re = 360000 Re = 700000 Re = 1000000
10 0.9811 1.0343 1.0512 1.0727 1.1000 Re = 2000000 Re = 5000000
12 0.9132 1.0390 1.1212 1.2072 1.2673
14 - 0.6284 0.8846 1.1614 1.3423 Fig 6: Angle of attack Vs Drag coefficient for several
Reynolds number (Re).
3.2 Drag coefficient (CD)
Table2. Values of CD. 3.5 Pressure Distribution over NACA 0012 Aerofoil
Re CD
α 3.6 x 105 7 x 105 1 x 106 2 x 106 5 x 106
0 0.0079 0.0067 0.0065 0.0064 0.0064
2 0.0084 0.0070 0.0068 0.0066 0.0066
4 0.0098 0.0083 0.0078 0.0073 0.0072
6 0.0125 0.0108 0.0101 0.0090 0.0081
8 0.0153 0.0128 0.0119 0.0105 0.0092
10 0.0184 0.0159 0.0147 0.0128 0.0106
12 0.0217 0.0195 0.0180 0.0155 0.0130
14 - 0.0236 0.0222 0.0191 0.0159
These values are used to find the graphs: (a) lift
coefficient vs. angle of attack and (b) drag coefficient
vs. angle of attack graphs for various Reynolds number.
3.3 Lift coefficient Vs Angle of attack:
Fig 7: The pressure distribution for the NACA 0012
aerofoil under free stream condition for Mach number
1.6
0.7 and angle of attack 4o.
1.4
1.2
Lift coefficient
1
0.8
0.6
0.4
0.2
0
0 5 10 15 20
Angle of attack
Re = 360000 Re = 700000 Re = 1000000
Re = 2000000 Re = 5000000
Fig 5: Lift coefficient Vs Angle of attack for several
Reynolds number (Re).
This graph shows that maximum lift coefficient is not
constant for NACA 0012 aerofoil, it increases with the
increasing Reynolds number with angle of attack.
Fig 8: The pressure distribution for the NACA 0012
aerofoil for inviscid flow for Mach number 0.8 and
angle of attack 2o.

4. 7. NOMENCLATURE
Symbol Meaning Unit
α Angle of attack Degree
Re Reynolds number None
CL Lift coefficient None
CD Drag coefficient None
kg/m3
ρo Density of air
Po Operating pressure MPa
μ Absolute viscosity kg/m-s
v air velocity m/s
Fig 9: The pressure distribution for the NACA 0012
aerofoil for viscid flow for Mach number 0.7 and angle
of attack 4o.
4. DISCUSSION
From Fig 5 it is seen that at zero degree angle of attack
the lift coefficient is zero and it increases linearly with
the increase of angle of attack. After reaching at a peak
point, the lift coefficient decreases sharply with the
increase of angle of attack and the values also vary with
different Reynolds number.
One major feature of drag coefficient is that for zero
degree angle of attack it is not zero and so thus the drag
force. From Fig 6, parabolic curves are found as were
expected.
5. CONCLUSION
This research work has been carried out to observe the
characteristics of UAV NACA 0012. This mainly
involved the conceptual design for better design and
economical construction. The design concept is a better
approach to choose among various types of UAV
(Unmanned Air Vehicle). Mainly, this work has brought
some important aerodynamic characteristics of
aerofoils. These results found in two dimensional
designs may vary with the three dimensional.
6. REFERENCES
1. Sq Ldr GM Jahangir Alam, 2007, “Investigation Of
The Aerodynamic Characteristics Of The Biplane
Configurations Using NACA 0024 Profile”, M. Sc.
Thesis, BUET, Dhaka, Bangladesh.
2. J. N. Reddy, 2005, Finite Element Method, Texas A
& M University, Texas.
3. Byron S. Gottfried, 1996, Schaum’s Outline Of
Theory And Problems Of Programming With C
4. Rajesh Bhaskaran, Fluent Tutorials.
5. James C. Date and Stephen R. Turnock, 2002,
“Computational Evaluation Of The Periodic
Performance Of A NACA 0012 Fitted With A
Gurney Flap”.
6. Bertin J.J, Smith M.L., “Aerodynamics for
Engineers”, 3rd – Ed, Prentice Hall.
7. Clancy, L. J., “Aerodynamics”.