Slideshare uses cookies to improve functionality and performance, and to provide you with relevant advertising. If you continue browsing the site, you agree to the use of cookies on this website. See our User Agreement and Privacy Policy.

Slideshare uses cookies to improve functionality and performance, and to provide you with relevant advertising. If you continue browsing the site, you agree to the use of cookies on this website. See our Privacy Policy and User Agreement for details.

Successfully reported this slideshow.

Like this document? Why not share!

- NACA 8414 PROFİLİNE SAHİP UÇAK KANA... by Kenan Çal 4918 views
- Uçak Kanadında Titreşim Analizi-Mod... by tst34 5743 views
- Aer 101 chapter 5 by anashalim 4967 views
- CFD analysis of Flow across an Aero... by JJ Technical Solu... 4852 views
- CFD analysis of aerofoil by Neel Thakkar 7914 views
- Fluid motion in the posterior chamb... by Dr. Bhuiyan S. M.... 3498 views

No Downloads

Total views

14,100

On SlideShare

0

From Embeds

0

Number of Embeds

39

Shares

0

Downloads

98

Comments

4

Likes

7

No notes for slide

- 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”.

No public clipboards found for this slide

Login to see the comments