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Design Analysis of UAV (Unmanned Air Vehicle) using NACA 0012 Aerofoil Profile

Alimul Rajib
B.Sc in Mechanical Engineering.
Military Institute of Science and Technology, Dhaka-1216, Bangladesh

Bhuiyan Shameem Mahmud Ebna Hai
Lecturer
Department of Mechanical Engineering.
Military Institute of Science and Technology, Dhaka-1216, Bangladesh.
e-mail: [email protected]

Wing Commander Md Abdus Salam
Ag. Head of Department of Aeronautical Engineering.
Military Institute of Science and Technology, Dhaka-1216, Bangladesh
e-mail: [email protected]

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.

1. INTRODUCTION
An Unmanned Air Vehicle (UAV) is an unpiloted aircraft. Its aerodynamic characteristics vary with certain parameters like the angle of attack and others. Experimental works on UAVs have been conducted in many places with various aerofoil profiles but not enough work with the Computational Fluid Dynamics (CFD) analysis is not available that much till now. The present work contains mainly CFD analysis to determine the flow pattern and the aerodynamic characteristics of an UAV.
The shape of an aircraft is designed to make the airflow through the surface to produce a lifting force in the most efficient manner. In addition to the lift, a force directly opposing the motion of the wing through the air is always present, which is called a drag force. The angle between the relative wind and the chord line is the angle of attack of the aerofoil. The lift and drag forces developed by an aircraft will vary with the change of angle of attack. The cross sectional shape obtained by the intersection of the wing with the perpendicular plane is called an aerofoil. Here NACA 0012 symmetric aerofoil profiles have been used for the present research work.

Fig 1: Aerodynamic forces on a typical aerofoil

The lift force increases almost linearly with the angle of attack until a maximum value is reached where upon the wing is said to stall. The shape of the drag force vs. angle of attack is approximately parabolic. It is desirable for the wing to have the maximum lift and smallest possible drag.

1.1 BACKGROUND OF THE RESEARCH WORK
Designing of UAVs requires designing of aerofoil section. Various aerofoil configurations have been employed so far and more will be coming. The present work is carried out numerically with CFD analysis for NACA 0012 symmetric aerofoil profile. Some of the parameters of aerofoil and properties of air have been kept constant and some have been varied.
The flow of air over the aerofoil is varied as per requirement. The chord length of the aerofoil is 100mm. The free stream airflow has been kept 12.5 m/s and the effect of the temperature in the study has been neglected. The density of air (?o)= 1.22 kg/m3, operating pressure (Po) = 0.101 MPa (1.01 bar) and absolute viscosity (?) = 1.789 x 10-5 kg/m-s. The Reynolds Number has been considered as variable. The 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.

1.2 OBJECTIVES
a. Designing of NACA 0012 aerofoil section and investigation of the flow pattern with the help of 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 CFD analysis.

2. WORKING PRINCIPAL
The computation and graphical plotting involves the following sequence:
a. Programming to get vertices for aerofoil section using governing equation.
b. Working with vertices using GAMBIT software.
c. Working with FLUENT software.

2.1 DESIGN METHOD
The early NACA aerofoil series, the 4-digit was generated using analytical equations that describe its geometrical feature.

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.

The equation was solved here by a C program to find vertices for the aerofoil line. Approximately, 10,000 vertices were used.

2.2 AEROFOIL DESIGN
The vertices obtained from the C program were used to draw the profile line which was as follows:

Fig 3: NACA 0012 aerofoil section.

The boundary was then given.

Table 1: Values of boundary vertices for NACA 0012 aerofoil profile.

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 4: Meshing of aerofoil section (2D).

It was recommended that the boundaries around the aerofoil were far enough.

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 (? = 1.225 kg/m3), absolute viscosity (? = 1.789 x 10-5 kg/m-s). The Reynolds number and Mach number were kept constant and sometimes varied as per requirement.

3. RESULTS AND GRAPHS
3.1 Lift coefficient (CL)
Sample results of lift coefficient CL with variable angle of attack ? and Reynolds number Re are as follows:

Table3. Values of CL:

3.2 Drag coefficient (CD)

Table2. Values of CD.

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 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.

3.4 Angle of attack Vs Drag coefficient:

Fig 6: Angle of attack Vs Drag coefficient for several Reynolds number (Re).

3.5 Pressure Distribution over NACA 0012 Aerofoil

Fig 7: The pressure distribution for the NACA 0012 aerofoil under free stream condition for Mach number 0.7 and angle of attack 4°.

Fig 8: The pressure distribution for the NACA 0012 aerofoil for inviscid flow for Mach number 0.8 and angle of attack 2°.

Fig 9: The pressure distribution for the NACA 0012 aerofoil for viscid flow for Mach number 0.7 and angle of attack 4°.

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”.



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