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.
Structural detailing of fuselage of aeroplane /aircraft.PriyankaKg4
This presentation is about the structural detailing of fuselage of aeroplane .The fuselage or body of the airplane, holds all the pieces together. The pilots sit in the cockpit at the front of the fuselage. Passengers and cargo are carried in the rear of the fuselage. Some aircraft carry fuel in the fuselage; others carry the fuel in the wings.
The flow across an airfoil is studied for different angle of attack. The CFD analysis results are documented and studied for different angle of attack using fluent & gambit.
Structural detailing of fuselage of aeroplane /aircraft.PriyankaKg4
This presentation is about the structural detailing of fuselage of aeroplane .The fuselage or body of the airplane, holds all the pieces together. The pilots sit in the cockpit at the front of the fuselage. Passengers and cargo are carried in the rear of the fuselage. Some aircraft carry fuel in the fuselage; others carry the fuel in the wings.
The flow across an airfoil is studied for different angle of attack. The CFD analysis results are documented and studied for different angle of attack using fluent & gambit.
Pressure Distribution on an Airfoil
The team conducted the experiment to determine the effects of pressure distribution on lift and pitching moment and the behavior of stall for laminar and turbulent boundary layers in the USNA Closed-Circuit Wing Tunnel (CCWT) with an NACA 65-012 airfoil at a Reynolds number of 1,000,000. The airfoil was tested in a clean configuration at angles of attack of 0, 5, 8, 10, and 12 degrees. Tape added to the leading edge tripped the boundary layer, and pressure distributions were taken at 8, 10, and 12 degrees angle of attack. Experimental results showed a suction peak at less than 1% of chord, providing a beneficial test article for contrast between smooth and laminar boundary layer behavior at the stall condition. The maximum lift coefficient for the clean airfoil was 0.9 at 10 degrees angle of attack, and tripped airfoil reached a maximum lift coefficient of 1.03 at 12 degrees angle of attack, a 14% increase. These data were 10% lower than the empirical airfoil data found in Theory of Wing Sections from Abbott and von Doenhoff. Pitching moment coefficient about the quarter chord remained near zero below stall as expected for a symmetrical airfoil, but rapidly became negative after stall for experimental and empirical data. The airfoil exhibited a leading edge stall for both laminar and turbulent boundary layers.
What are the elements of aircraft performance?
How much thrust do you need?
How fast and how slow can you fly?
#WikiCourses
http://wikicourses.wikispaces.com/Topic+Performance+of+aerospace+vehicles
Morphing Aircraft Technology – New Shapes for Aircraft Wing DesignMani5436
Morphing aircraft are multi-role aircraft that change their external shape substantially to adapt to a changing mission environment during flight.
Morphing poses several unique challenges when the wing loading is high. Very flexible materials are the designer’s first choice because they are easily reshaped.
he current use of multiple aerodynamic devices (such as flaps and slats) represents a simplification of the general idea behind morphing. Traditional control systems (with fixed geometry and/or location) give high aerodynamic performance over a fixed range and for a limited set of flight conditions.
Nomenclature and classification of controls in an airplane (slide # 3-4).
Which are the aerodynamic forces acting on airplane (slide # 5).
Working principle of an airplane (slide # 6).
How an airplane flies (basic motions of an airplane) (slide # 7).
How controls play their roles in these motions (slide # 8-22).
Simulate a flight in Cessna Skyhawk (slide # 23-28).
References and Questions & answers (slide # 30).
Airfoil properties, shapes & structural dynamical features are described. Nomenclature or the classification types are presented along with the application.
Common methods for analysis of the structural dynamics on a wing or blade are presented along with the possible applications.
Pressure Distribution on an Airfoil
The team conducted the experiment to determine the effects of pressure distribution on lift and pitching moment and the behavior of stall for laminar and turbulent boundary layers in the USNA Closed-Circuit Wing Tunnel (CCWT) with an NACA 65-012 airfoil at a Reynolds number of 1,000,000. The airfoil was tested in a clean configuration at angles of attack of 0, 5, 8, 10, and 12 degrees. Tape added to the leading edge tripped the boundary layer, and pressure distributions were taken at 8, 10, and 12 degrees angle of attack. Experimental results showed a suction peak at less than 1% of chord, providing a beneficial test article for contrast between smooth and laminar boundary layer behavior at the stall condition. The maximum lift coefficient for the clean airfoil was 0.9 at 10 degrees angle of attack, and tripped airfoil reached a maximum lift coefficient of 1.03 at 12 degrees angle of attack, a 14% increase. These data were 10% lower than the empirical airfoil data found in Theory of Wing Sections from Abbott and von Doenhoff. Pitching moment coefficient about the quarter chord remained near zero below stall as expected for a symmetrical airfoil, but rapidly became negative after stall for experimental and empirical data. The airfoil exhibited a leading edge stall for both laminar and turbulent boundary layers.
What are the elements of aircraft performance?
How much thrust do you need?
How fast and how slow can you fly?
#WikiCourses
http://wikicourses.wikispaces.com/Topic+Performance+of+aerospace+vehicles
Morphing Aircraft Technology – New Shapes for Aircraft Wing DesignMani5436
Morphing aircraft are multi-role aircraft that change their external shape substantially to adapt to a changing mission environment during flight.
Morphing poses several unique challenges when the wing loading is high. Very flexible materials are the designer’s first choice because they are easily reshaped.
he current use of multiple aerodynamic devices (such as flaps and slats) represents a simplification of the general idea behind morphing. Traditional control systems (with fixed geometry and/or location) give high aerodynamic performance over a fixed range and for a limited set of flight conditions.
Nomenclature and classification of controls in an airplane (slide # 3-4).
Which are the aerodynamic forces acting on airplane (slide # 5).
Working principle of an airplane (slide # 6).
How an airplane flies (basic motions of an airplane) (slide # 7).
How controls play their roles in these motions (slide # 8-22).
Simulate a flight in Cessna Skyhawk (slide # 23-28).
References and Questions & answers (slide # 30).
Airfoil properties, shapes & structural dynamical features are described. Nomenclature or the classification types are presented along with the application.
Common methods for analysis of the structural dynamics on a wing or blade are presented along with the possible applications.
This research is concerned with the nature and aerodynamic behavior on cricket balls in flight. It is written to determine the aerodynamic characteristics of a cricket ball with a mass of 156 gm and approximate diameter of 70 mm, where a sample cricket ball was fixed with a shaft in a wind tunnel. The aerodynamic characteristics have been analyzed by varying the rotational rpm of the cricket ball, where the axis of rotation for seam and shaft is same and the pressure difference between the upper and the lower surface of the cricket ball determine by the help of static manometer. Some experimental works have also been carried out and compared with those of the results obtained numerically. The upward pressure, which creates lift, has increased almost linearly with the increase of ball position angle to approximately 15o to 30o and it also decrease with increase of ball rotational speed. Finally some conclusions have been drawn on the basis of the experimental result.
Naca 2415 finding lift coefficient using cfd, theoretical and javafoileSAT Journals
Abstract In this paper we have studied the experimental characteristic graph of NACA 2415.The experimental graphs were taken from the book, “Theory of wing section” by IRA H. ABBOTT. We used these graphs for the validation of our results. Then we use CFD to simulate the experimental flow conditions and check the results and compare them with the experimental results. We meshed the airfoil in ICEM CFD so that the meshing is very precise. We then calculate the NACA 2415 airfoil’s lift at different angle of attack theoretically and using CFD analysis and compare them with the experimental values. We find the errors between experimental and CFD values as well as experimental and theoretical values. We used another simulation software called Javafoil and used it for comparison. Keywords: Experimental, CFD, Theoretical, Javafoil
A comparative flow analysis of naca 6409 and naca 4412 aerofoileSAT Journals
Abstract
In this work, flow analysis of two aerofoils (NACA 6409 and NACA 4412) was investigated. Drag force, lift force as well as the overall pressure distribution over the aerofoils were also analysed. By changing the angle of attack, variation in different properties has been observed. The outcome of this investigation was shown and computed by using ANSYS workbench 14.5. The pressure distributions as well as coefficient of lift to coefficient of drag ratio of these two aerofoils were visualized and compared. From this result, we compared the better aerofoil between these two aerofoils. The whole analysis is solely based on the principle of finite element method and computational fluid dynamics (CFD). Finally, by comparing different properties i.e drag and lift coefficients, pressure distribution over the aerofoils, it was found that NACA 4412 aerofoil is more efficient for practical applications than NACA 6409 aerofoil.
Keywords: NACA, Drag Lift, CFD, ANSYS FLUENT, SolidWorks
Effect of spikes integrated to airfoil at supersonic speedeSAT Journals
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The objective of this is to analyse the flow field over an aerofoil section integrated with spikes at supersonic speed (Mach number
greater than 1). Use of spike integrated with aerofoil changes the flow characteristics over aerofoil and hence aerodynamic lift
and drag. The experiment consists of flow visualization graphs and measurement of coefficient of aerodynamic drag and lift.
Here we are using different shapes of spike like sharp edge and hemi spherical edge. In this we will compare the flow over
aerofoil with spike and without spike. The flow analysis is done by using Computational fluid dynamics (CFD). CFD is the study
of external flow over a body or internal flow through the body. CFD is aiding aero-dynamist to better understand the flow physics
and in turn to design efficient models. In short, CFD is playing a strong role as a design tool as well as a research tool.
Keywords: NACA 651-412 airfoil, spike, Ansys Fluent, Ansys ICEM CFD, Pressure Coefficient
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Abstract The Principle purpose of a nozzle is to accelerate the flow to higher exit velocities. The fluid acceleration is based on the design criteria and characteristics. To achieve good performance characteristics with minimum energy losses a nozzle must satisfy all the design requirements at all operating conditions. This is possible only when the nozzle theory is assumed to be isentropic irrespective of the changes in pressure, temperature and density which is generally caused due to formation of a Shock Wave. The thesis focuses on the design, development and optimization of a Supersonic Convergent-Divergent Nozzle where the analytical results are validated using theory calculations. The simulation work is carried out for CD Nozzles with different angles of divergence keeping the other inputs fixed. The objective of the proposed thesis is to show the best Expansion ratio, Nozzle Pressure ratio (NPR) and Nozzle Area Ratio(NAR) where the thrust obtained by the supersonic nozzle is maximum. The simulation is then repeated for expansion gas the results of which are later compared with standard air to show which possesses better performance characteristics. The Nozzle design chosen is based upon existing literature studies. Key Words: CD Nozzle, Expansion Ratio, Nozzle Pressure Ratio (NPR), Nozzle Area Ratio(NAR),Divergence Angle etc…
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Design Analysis Of Uav (Unmanned Air Vehicle) Using NACA 0012 Aerofoil Profile
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”.