This report is a simulation for a flow over an airfoil "NACA 0009" at Reynolds number equals 1 million for four angles of attack using three different turbulence models and of cause a grid independence solution.
The goal of this study is to apply the knowledge obtained from studying in the university and self-learning in order to solve a specific task of finding the coefficient of drag and lift for the airfoil.
A youtube video made by me explaining how to simulate a flow over an airfoil: https://goo.gl/9VYRFM
Team members:
Ahmed Kamal Shalaby
Ahmed Gaber Ahmed
Esraa Mahmoud Saleh
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.
COMPUTATIONAL FLUID DYNAMIC ANALYSIS OF AIRFOIL NACA0015IAEME Publication
In this chapter we choose standard airfoil NACA 0015. Which is symmetrical airfoil with a 15%thickness to chord ratio was analyzed on ANSYS FLUENT to determine the coefficient of lift,
coefficient of drag and graph of coefficient of lift vs. coefficient of drag. The 2-dimensional crosssectional view was considered. The wind velocity was taken as 17m/s which arecorresponding to232,940 Reynolds number. The airfoil, with an 8 in chord, was analyzed at 0, 5, 10 and 15 degrees.
Parameters viz. Coefficient of lift (Cl), Coefficient of drag (Cd) nd Cl/Cd are calculated and areplotted against different angle of attack.
This report is a simulation for a flow over an airfoil "NACA 0009" at Reynolds number equals 1 million for four angles of attack using three different turbulence models and of cause a grid independence solution.
The goal of this study is to apply the knowledge obtained from studying in the university and self-learning in order to solve a specific task of finding the coefficient of drag and lift for the airfoil.
A youtube video made by me explaining how to simulate a flow over an airfoil: https://goo.gl/9VYRFM
Team members:
Ahmed Kamal Shalaby
Ahmed Gaber Ahmed
Esraa Mahmoud Saleh
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.
COMPUTATIONAL FLUID DYNAMIC ANALYSIS OF AIRFOIL NACA0015IAEME Publication
In this chapter we choose standard airfoil NACA 0015. Which is symmetrical airfoil with a 15%thickness to chord ratio was analyzed on ANSYS FLUENT to determine the coefficient of lift,
coefficient of drag and graph of coefficient of lift vs. coefficient of drag. The 2-dimensional crosssectional view was considered. The wind velocity was taken as 17m/s which arecorresponding to232,940 Reynolds number. The airfoil, with an 8 in chord, was analyzed at 0, 5, 10 and 15 degrees.
Parameters viz. Coefficient of lift (Cl), Coefficient of drag (Cd) nd Cl/Cd are calculated and areplotted against different angle of attack.
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.
Design, Analysis and Testing of Wing Spar for Optimum WeightRSIS International
Aircraft is a complex mechanical structure with flying capability. The structure of an airframe represents one of the finest examples of a minimum weight design in the field of structural engineering. Surprisingly such an efficient design is achieved by the use of simple “strength-of-material” approach. Aircraft has two major components, which are fuselage and wing. For a wing of an aircraft the primary load carrying ability is required in bending. A typical aluminium material 6082-T6 is chosen for the design. A four-Seater aircraft wing spar design is considered in the current study. Wings of the aircraft are normally attached to the fuselage at the root of the wing. This makes the wing spar beam to behave almost like a cantilever beam. Minimum two spars are considered in the wing design. In a conventional beam design approach one will end up in heavy weight for the spar of the wing. In the current project the spar is considered as a beam with discrete loads at different stations. The design is carried out as per the external bending moment at each station. A finite element approach is used to calculate the stresses developed at each station for a given bending moment. Several stress analysis iterations are carried out for design optimization of the spar beam. Linear static analysis is used for the stress analysis. The spar beam is designed to yield at the design limit load. Weight optimization of the spar will be carried out by introducing lightening cut-outs in the web region. The results from the conventional design approach and the optimized design are compared. Weight saving through the design optimization is calculated. Spar will be a built-up structure. A scale-down model of the spar will be fabricated using aluminium alloy 6082-T6 material. Static testing of the spar will be carried out to validate the design and stress analysis results.
INTRODUCTION:
While a helicopter is a far more complex machine than an aeroplane, the fundamental principles of flight are the same.
The rotor blades of a helicopter are identical to the wings of an aeroplane –when air is blown over them, lift is produced.
The crucial difference is that the flow of air is produced by rotating the wings – or rotor blades – rather than by moving the whole aircraft.
When the rotor blades start to spin, the air flowing over them produces lift, and this can cause the helicopter to rise into the air.
So, the engine is used to turn the blades, and the turning blades produce the required lift.
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.
Design, Analysis and Testing of Wing Spar for Optimum WeightRSIS International
Aircraft is a complex mechanical structure with flying capability. The structure of an airframe represents one of the finest examples of a minimum weight design in the field of structural engineering. Surprisingly such an efficient design is achieved by the use of simple “strength-of-material” approach. Aircraft has two major components, which are fuselage and wing. For a wing of an aircraft the primary load carrying ability is required in bending. A typical aluminium material 6082-T6 is chosen for the design. A four-Seater aircraft wing spar design is considered in the current study. Wings of the aircraft are normally attached to the fuselage at the root of the wing. This makes the wing spar beam to behave almost like a cantilever beam. Minimum two spars are considered in the wing design. In a conventional beam design approach one will end up in heavy weight for the spar of the wing. In the current project the spar is considered as a beam with discrete loads at different stations. The design is carried out as per the external bending moment at each station. A finite element approach is used to calculate the stresses developed at each station for a given bending moment. Several stress analysis iterations are carried out for design optimization of the spar beam. Linear static analysis is used for the stress analysis. The spar beam is designed to yield at the design limit load. Weight optimization of the spar will be carried out by introducing lightening cut-outs in the web region. The results from the conventional design approach and the optimized design are compared. Weight saving through the design optimization is calculated. Spar will be a built-up structure. A scale-down model of the spar will be fabricated using aluminium alloy 6082-T6 material. Static testing of the spar will be carried out to validate the design and stress analysis results.
INTRODUCTION:
While a helicopter is a far more complex machine than an aeroplane, the fundamental principles of flight are the same.
The rotor blades of a helicopter are identical to the wings of an aeroplane –when air is blown over them, lift is produced.
The crucial difference is that the flow of air is produced by rotating the wings – or rotor blades – rather than by moving the whole aircraft.
When the rotor blades start to spin, the air flowing over them produces lift, and this can cause the helicopter to rise into the air.
So, the engine is used to turn the blades, and the turning blades produce the required lift.
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 presentation is for mechanical engineering/ civil engineering students to help them understand the different type of destructive mechanical testing of materials. The tensile testing, hardness, impact test procedures are explained in detail.
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
Aerodynamic Analysis of Low Speed Turbulent Flow Over A Delta WingIJRES Journal
Delta wing has been a subject of intense research since decades due to decades due to inherent characteristics of generating increased nonlinear lift due to vortex dominated flows. Lot of work has been carried out in order to understand the vortex dominated flows on the delta wing. The delta wing is a wing platform in the form of a triangle. Aerodynamics of wings with moderate sweep angle is recognized by the aerospace community as a challenging problem. In spite of its potential application in military aircraft, the understanding of the aerodynamics of such wings is far from complete. In order to address this situation, the present work is initiated to compute the 3D turbulent flow field over sharp edged finite wings with a diamond shaped plan forms and moderate sweep angle. The detailed flow pattern and surface pressure distribution may further indicate the appropriate kind of flow control during flight operation of such wings. The flow field is computed using an in-house developed CFD code RANS3D.
International Journal of Engineering Research and Applications (IJERA) is a team of researchers not publication services or private publications running the journals for monetary benefits, we are association of scientists and academia who focus only on supporting authors who want to publish their work. The articles published in our journal can be accessed online, all the articles will be archived for real time access.
Our journal system primarily aims to bring out the research talent and the works done by sciaentists, academia, engineers, practitioners, scholars, post graduate students of engineering and science. This journal aims to cover the scientific research in a broader sense and not publishing a niche area of research facilitating researchers from various verticals to publish their papers. It is also aimed to provide a platform for the researchers to publish in a shorter of time, enabling them to continue further All articles published are freely available to scientific researchers in the Government agencies,educators and the general public. We are taking serious efforts to promote our journal across the globe in various ways, we are sure that our journal will act as a scientific platform for all researchers to publish their works online.
CFD and EXPERIMENTAL ANALYSIS of VORTEX SHEDDING BEHIND D-SHAPED CYLINDERAM Publications
The flow around bluff bodies is an area of great research of scientists for several years. Vortex shedding is
one of the most challenging phenomenon in turbulent flows. This phenomenon was first studied by Strouhal. Many
researchers have modeled the various objects as cylinders with different cross-sections among which square and
circular cylinders were the most interested sections to study the vortex shedding phenomenon. The Vortex Shedding
frequency depends on different aspects of the flow field such as the end conditions, blockage ratio of the flow passage,
and width to height ratio. This case studies the wave development behind a D-Shaped cylinder, at different Reynolds
numbers, for which we expect a vortex street in the wake of the D-Shaped cylinder, the well known as von Kármán
Street. This body typically serves some vital operational function in aerodynamic. In circular cylinder flow separation
point changes with Reynolds number but in D-Shaped cylinder there is fix flow separation point. So there is more
wake steadiness in D-Shaped cylinder as compared to Circular cylinder and drag reduction because of wake
steadiness.In the present work CFD simulation is carried out for flow past a D-Shaped cylinder to see the wake
behavior. The Reynolds number regime currently studied corresponds to low Reynolds number, laminar and
nominally two-dimensional wake. The fluid domain is a two-dimensional plane with a D-Shaped cylinder of
dimensions B=90mm, H=80mm and L=200mm. CFD calculations of the 2-D flow past the D-Shaped cylinder are
presented and results are validated by comparing with Experimental results of pressure distribution on cylinder
surface. The experimentation is carried out using small open type wind tunnel. The flow visualization is done by
smoke visualization technique. Results are presented for various B/H ratios and Reynolds numbers. The variation of
Strouhal number with Reynolds number is found from the analysis. The focus of the present research is on reducing
the wake unsteadiness.
Effect of Gap between Airfoil and Embedded Rotating Cylinder on the Airfoil A...CrimsonPublishersRDMS
Effect of Gap between Airfoil and Embedded Rotating Cylinder on the Airfoil Aerodynamic Performance by Najdat Nashat Abdulla* in Crimson Publishers: Peer Reviewed Material Science Journals
Simulation of segregated flow over the 2 d cylinder using star ccm+Burak Turhan
In this thesis numerical simulation for classical case of flow over a cylinder is accomplished for 2D models using commercial CFD code Star CCM+ with k-ϵ model approach. The results are validated by comparing the Drag coefficients to the previously published data. The simulation is carried out to for Reynolds number 3900 to investigate the turbulence modeling on separation from curved surfaces of two different sizes of a circular cylinder, a cylinder with triangular cross section and a rectangular cross section. Investigation of different turbulence models and Mesh convergence is carried out.
The investigation of the turbulence model of the circular cylinder is carried out by the drag coefficient obtained by four different turbulence models such as K-Epsilon Turbulence, K-Omega Turbulence, Reynolds Stress Turbulence and Spalart-Allmaras Turbulence. Drag coefficient found out by different turbulence model is compared with the experimental value of a previously published data. The Mesh Convergence have been carried out by implementing different base mesh size in a decreasing order and the convergence is obtained when the drag coefficient becomes constant
The International Journal of Engineering and Science (The IJES)theijes
The International Journal of Engineering & Science is aimed at providing a platform for researchers, engineers, scientists, or educators to publish their original research results, to exchange new ideas, to disseminate information in innovative designs, engineering experiences and technological skills. It is also the Journal's objective to promote engineering and technology education. All papers submitted to the Journal will be blind peer-reviewed. Only original articles will be published.
The International Journal of Engineering and Science (The IJES)
CFD analysis of an Airfoil
1. Computational Fluid
Dynamics (CFD) Analysis
of NACA 0012 Airfoil
The City College of New York
Dr. Zhexuan Wang
ME 35600
Mostafa Al Mahmud
05/28/2013
2. M o s t a f a A l M a h m u d | 1
ABSTRACT
In this report, a low-speed airfoil over the NACA 0012 airfoil at 2° and 14° attack angles with
the given inlet velocity of 0.25 m/s, was modeled and computational fluid dynamic (CFD)
analysis were performed using FLUENT in ANSYS. The Reynolds number based on the chord is
roughly 𝑅𝑒 = 2.88 × 106
. The flow was modeled as incompressible and inviscid. All setup and
procedures were done by following the steps provided Cornel University website. Though, mesh
independence was achieved for 2° attack angle, for 14° attack angle; which is more than the stall
angle, mesh independence was not achieved. Lift and drag coefficient increases as the number of
mesh element or the attack angle increases.
3. M o s t a f a A l M a h m u d | 2
TABLE OF CONTENTS
Abstract.........................................................................................................................................................1
Introduction ..................................................................................................................................................4
Incompressible, Inviscid Flow .......................................................................................................................4
Boundary Value Problem..............................................................................................................................4
Boundary Conditions ................................................................................................................................5
Coefficient of Pressure..................................................................................................................................6
2° attack angle with 15000 mesh element ...............................................................................................6
2° attack angle with 40000 mesh element ...............................................................................................6
14o
attack angle with 15000 mesh element..............................................................................................7
14o
attack angle with 40000 mesh element ..............................................................................................7
Comparison of Coefficient of Pressure at 2o
attack angle with 15,000 & 40,000 mesh element .............8
Comparison of Coefficient of Pressure at 14o
attack angle with 15,000 & 40,000 mesh element ...........8
Lift and Drag Coefficient ...............................................................................................................................9
Convergence .................................................................................................................................................9
Conclusion.....................................................................................................................................................9
Appendix .....................................................................................................................................................10
Pressure Coefficient................................................................................................................................10
2° Attack angle and 15,000 mesh element.........................................................................................10
2° Attack angle and 40,000 mesh element.........................................................................................10
14° Attack angle and 15,000 mesh element.......................................................................................11
14o
attack angle with 40000 mesh element ............................................................................................11
Velocity Vector........................................................................................................................................12
2° Attack angle and 15,000 mesh element.........................................................................................12
2° Attack angle and 40,000 mesh element.........................................................................................12
14° Attack angle and 15,000 mesh element.......................................................................................13
14° Attack angle and 40,000 mesh element.......................................................................................13
Velocity Contour .....................................................................................................................................14
2° Attack angle and 15,000 mesh element.........................................................................................14
2° Attack angle and 40,000 mesh element.........................................................................................14
14° Attack angle and 15,000 mesh element.......................................................................................15
Static Pressure.........................................................................................................................................15
2° Attack angle and 15,000 mesh element.........................................................................................15
4. M o s t a f a A l M a h m u d | 3
2° Attack angle and 40,000 mesh element.........................................................................................16
14° Attack angle and 15,000 mesh element.......................................................................................16
14° Attack angle and 40,000 mesh element.......................................................................................17
Stream line..............................................................................................................................................17
2° Attack angle and 14,000 mesh element.........................................................................................17
2° Attack angle and 40,000 mesh element.........................................................................................18
14° Attack angle and 15,000 mesh element.......................................................................................18
14° Attack angle and 40,000 mesh element.......................................................................................19
Convergence ...........................................................................................................................................20
2° Attack angle and 15,000 mesh element.............................................................................................20
2° Attack angle and 40,000 mesh element.........................................................................................21
14° Attack angle and 15,000 mesh element.......................................................................................22
14° Attack angle and 40,000 mesh element.......................................................................................22
Coefficient of drag 𝐶 𝑑 and coefficient of lift 𝐶𝑙 ......................................................................................23
2° Attack angle and 15,000 mesh element.........................................................................................23
2° Attack angle and 40,000 mesh element.........................................................................................23
14° Attack angle and 15,000 mesh element.......................................................................................24
14° Attack angle and 40,000 mesh element.......................................................................................24
Procedure and Setup ..................................................................................................................................25
FLUENT - Flow over an Airfoil .................................................................................................................25
Created by Benjamin J Mullen ............................................................................................................25
5. M o s t a f a A l M a h m u d | 4
INTRODUCTION
The flow over the airfoil is an external flow. It is a kind of flow that flows over the outside the
body of an object; in our case ‘the airfoil.’ These fluid flow moves around the airfoil. Due to
these flow there are forces developed that are normal and parallel to the flow, and these forces
are called drag force and lift force. Drag force is a mechanical force generated by the airfoil
moving through the fluid. And the lift force is the force that helps the airfoil to gain altitude. In
this project, we are considering low speed air flow over the NACA 0012 airfoil at an angle of 2o
and 14o
. For the Reynolds number of 2.88× 106
this flow was modeled as an inviscid and
incompressible flow. Using the Computational Fluid Dynamics (CFD) software “ANSYS”
NACA 0012 airfoil in wind tunnel were simulated for different attack angle and mesh elements.
Pressure contour, velocity vector, stream line, coefficient of drag and lift of the fluid were
obtained from this simulation. Different boundary conditions were needed to be setup in order to
solve the continuity equation and Navier-Stokes equation for two dimensional flows.
INCOMPRESSIBLE, INVISCID FLOW
An incompressible flow is a kind of flow in which the fluid density remains constant. An
inviscid flow is a flow in which the fluid does not have any viscosity. Drag coefficient is a
dimensionless quantity that is used to quantify the fluid resistance. Since we modeled our flow to
be inviscid or fluid without any resistance, the drag coefficient will always be zero.
BOUNDARY VALUE PROBLEM
We need to set certain boundary condition in the inlet, outlet, velocity magnitude and direction in
order to create the simulation. We define the velocity at the inlet according to our attack angle,
and set the gage pressure at the inlet to be zero. We assume the gage pressure at the outlet is also
zero. Lastly, we treat the air foil as a wall so that no flows can penetrate through the air foil.
Using this boundary condition we have to solve continuity equation, Navier- Stokes equations
for 2-D, steady, incompressible, and inviscid flow.
6. M o s t a f a A l M a h m u d | 5
1. Continuity equation
𝝏𝒖
𝝏𝒙
+
𝝏𝒗
𝝏𝒚
= 𝟎 [
𝝏𝒘
𝝏𝒛
= 𝟎 → 𝟐𝒅 ]
2. x- direction Navier-Stokes equation
𝝆 (𝒖
𝝏𝒖
𝝏𝒙
+ 𝒗
𝝏𝒖
𝝏𝒚
) = −
𝝏𝒑
𝝏𝒙
+ 𝝁 (
𝝏 𝟐
𝒖
𝝏𝒙 𝟐
+
𝝏 𝟐
𝒖
𝝏𝒚 𝟐
)
3. y-direction Navier-Stokes equation:
𝝆 (𝒖
𝝏𝒗
𝝏𝒙
+ 𝒗
𝝏𝒗
𝝏𝒚
) = −
𝝏𝒑
𝝏𝒚
+ 𝝁 (
𝝏 𝟐
𝒗
𝝏𝒙 𝟐
+
𝝏 𝟐
𝒗
𝝏𝒚 𝟐
)
𝝏𝒖
𝝏𝒕
= 𝟎 → 𝐒𝐭𝐞𝐚𝐝𝐲
𝒘
𝝏𝒖
𝝏𝒛
= 𝟎; 𝒘
𝝏𝒗
𝝏𝒛
= 𝟎;
𝝏 𝟐
𝒖
𝝏𝒛 𝟐
= 𝟎 ;
𝝏 𝟐
𝒗
𝝏𝒛 𝟐
= 𝟎 → 𝟐𝑫
𝝆𝒈 𝒙 = 𝟎 ; 𝝆𝒈 𝒚 = 𝟎 → 𝐈𝐧𝐜𝐨𝐦𝐩𝐫𝐞𝐬𝐬𝐢𝐛𝐥𝐞
BOUNDARY CONDITIONS
Inlet velocity 0.25 m/s
For 2o
x-component 0.25 × cos(2 𝑜) = 0.2498477068
y-component 0.25 × sin(2 𝑜) = 0.008724874176
For 14o
x-component 0.25 × cos(14 𝑜) = 0.2425739316
y-component 0.25 × cos(14 𝑜) = 0.0604804739
Gage pressure at the inlet and outlet 0
Airfoil type Wall
Domain C- Mesh.
7. M o s t a f a A l M a h m u d | 6
COEFFICIENT OF PRESSURE
2° ATTACK ANGLE WITH 15000 MESH ELEMENTS
2° ATTACK ANGLE WITH 40000 MESH ELEMENTS
8. M o s t a f a A l M a h m u d | 7
14O
ATTACK ANGLE WITH 15000 MESH ELEMENT
14O
ATTACK ANGLE WITH 40000 MESH ELEMENT
9. M o s t a f a A l M a h m u d | 8
-1
0
1
0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1 1.1
PressureCoefficient
Position (m)
Comparison Between 40000 and 15000 Element
Ele 15000 Ele 40000
COMPARISON OF COEFFICIENT OF PRESSURE AT 2O
ATTACK ANGLE WITH 15,000 &
40,000 MESH ELEMENT
COMPARISON OF COEFFICIENT OF PRESSURE AT 14O
ATTACK ANGLE WITH 15,000 &
40,000 MESH ELEMENT
-8
-7
-6
-5
-4
-3
-2
-1
0
1
2
0.6 0.7 0.8 0.9 1 1.1 1.2 1.3 1.4 1.5 1.6 1.7 1.8
PressureCoefficient
Position (m)
Comparisson Between 40000 and 15000 element
Ele 15000 Ele 40000
10. M o s t a f a A l M a h m u d | 9
For 2o
attack angle there is no difference in pressure coefficient for 15,000 and 40,000 mesh element;
which the model is mesh independent. Although, it was mesh independent 40,000 mesh element took
longer time converge. On the other hand, 14o
attack angle did not converge at all, and there is a slight
difference between the 15,000 and 40,000 mesh element.
LIFT AND DRAG COEFFICIENT
Angle of Attack Mesh Element Drag Coefficient, 𝑪 𝒅 Lift Coefficient, 𝑪𝒍
2o 15,000 0.0018339245 0.2289141
40,000 0.0043382516 0.23306806
14o
15,000 0.03843407 0.87291313
40,000 0.063931345 1.2510649
CONVERGENCE
Mesh element Iteration Continuity x-velocity y-velocity
15,000 2,540 4.5319 × 10−7
9.3167 × 10−7
4.4451 × 10−7
40,000 3,584 5.3753 × 10−7
9.9396 × 10−7
3.0979 × 10−7
CONCLUSION
It is evident form the data obtained from these simulation is that, both drag and lift coefficient will
increase as the angle of attack increases. However, the drag coefficient does not increase as significantly
as the lift coefficient. We modeled the flow to be inviscid; which means the drag coefficient should have
been zero. Although, it is not zero, comparing to the lift coefficient the drag coefficient is much smaller.
Lift and drag coefficient increases as the number of mesh element or the attack angle increases.
11. M o s t a f a A l M a h m u d | 10
APPENDIX
PRESSURE COEFFICIENT
2° ATTACK ANGLE AND 15,000 MESH ELEMENT
2° ATTACK ANGLE AND 40,000 MESH ELEMENT
12. M o s t a f a A l M a h m u d | 11
14° ATTACK ANGLE AND 15,000 MESH ELEMENT
14O
ATTACK ANGLE WITH 40000 MESH ELEMENT
13. M o s t a f a A l M a h m u d | 12
VELOCITY VECTOR
2° ATTACK ANGLE AND 15,000 MESH ELEMENT
2° ATTACK ANGLE AND 40,000 MESH ELEMENT
14. M o s t a f a A l M a h m u d | 13
14° ATTACK ANGLE AND 15,000 MESH ELEMENT
14° ATTACK ANGLE AND 40,000 MESH ELEMENT
15. M o s t a f a A l M a h m u d | 14
VELOCITY CONTOUR
2° ATTACK ANGLE AND 15,000 MESH ELEMENT
2° ATTACK ANGLE AND 40,000 MESH ELEMENT
16. M o s t a f a A l M a h m u d | 15
14° ATTACK ANGLE AND 15,000 MESH ELEMENT
STATIC PRESSURE
2° ATTACK ANGLE AND 15,000 MESH ELEMENT
17. M o s t a f a A l M a h m u d | 16
2° ATTACK ANGLE AND 40,000 MESH ELEMENT
14° ATTACK ANGLE AND 15,000 MESH ELEMENT
18. M o s t a f a A l M a h m u d | 17
14° ATTACK ANGLE AND 40,000 MESH ELEMENT
STREAM LINE
2° ATTACK ANGLE AND 14,000 MESH ELEMENT
19. M o s t a f a A l M a h m u d | 18
2° ATTACK ANGLE AND 40,000 MESH ELEMENT
14° ATTACK ANGLE AND 15,000 MESH ELEMENT
20. M o s t a f a A l M a h m u d | 19
14° ATTACK ANGLE AND 40,000 MESH ELEMENT
21. M o s t a f a A l M a h m u d | 20
CONVERGENCE
2° ATTACK ANGLE AND 15,000 MESH ELEMENT
22. M o s t a f a A l M a h m u d | 21
2° ATTACK ANGLE AND 40,000 MESH ELEMENT
23. M o s t a f a A l M a h m u d | 22
14° ATTACK ANGLE AND 15,000 MESH ELEMENT
14° ATTACK ANGLE AND 40,000 MESH ELEMENT
24. M o s t a f a A l M a h m u d | 23
COEFFICIENT OF DRAG 𝐶 𝑑 AND COEFFICIENT OF LIFT 𝐶𝑙
2° ATTACK ANGLE AND 15,000 MESH ELEMENT
2° ATTACK ANGLE AND 40,000 MESH ELEMENT
25. M o s t a f a A l M a h m u d | 24
14° ATTACK ANGLE AND 15,000 MESH ELEMENT
14° ATTACK ANGLE AND 40,000 MESH ELEMENT
26. M o s t a f a A l M a h m u d | 25
PROCEDURE AND SETUP
FLUENT - FLOW OVER AN AIRFOIL
CREATED BY BENJAMIN J MULLEN
https://confluence.cornell.edu/display/SIMULATION/FLUENT+-+Flow+over+an+Airfoil
27. Problem Specification
In this tutorial, we will show you how to simulate a NACA 0012 Airfoil at a 6 degree angle of attack placed in a
wind tunnel. Using FLUENT, we will create a simulation of this experiment. Afterwards, we will compare
values from the simulation and data collected from experiment.
Pre-Analysis & Start-Up
Boundary Conditions
One of the simple things we can think about before we set up the simulation is begin planning the boundary
conditions of the set up. One of the popular meshes for simulating a airfoil in a stream is a C-Mesh, and that is
what we will be using. At the inlet of the system, we will define the velocity as entering at a 6 degree angle of
attack (as per the problem statement), and at a total magnitude of 1. We will also define the gauge pressure at
the inlet to be 0. As for the outlet, the only thing we can assume is that the gauge pressure is 0. As for the airfoil
itself, we will treat it like a wall. Together, these boundary conditions form the picture below:
Open ANSYS Workbench
Now that we have the pre-calculations, we are ready to do a simulation in ANSYS Workbench! Open ANSYS
Workbench by going to Start > ANSYS > Workbench. This will open the start up screen seen as seen below
28. To begin, we need to tell ANSYS what kind of simulation we are doing. If you look to the left of the start up
window, you will see the Toolbox Window. Take a look through the different selections. We will be using
FLUENT to complete the simulation. Load the Fluid Flow (FLUENT) box by dragging and dropping it into the
Project Schematic.
Geometry
Download the Airfoil Coordinates
In this step, we will import the coordinates of the airfoil and create the geometry we will use for the simulation.
Begin by downloading this file here and saving it somewhere convenient. This file contains the points of a
NACA 0012 airfoil.
Launch Design Modeler
Before we launch the design modeler, we need to specify the problem as a 2D problem. Right click
and select Properties. In the Properties of Schematic A2: Geometry Window,
29. select Analysis Type > 2D. Now, double click to launch the Design Modeler.
When prompted, select Meters as the unit of measurement.
Airfoil
First, we will create the geometry of the airfoil. In the menu bar, go to Concept > 3D Curve. In the Details
View window, click Coordinates File and select the ellipsis to browse to a file. Browse to and select the
geometry file you downloaded earlier. Once you have selected the desired geometry file, click to
create the curve. Click to get a better look at the curve.
Next, we need to create a surface from the curve we just generated. Go to Concepts > Surfaces from Edges.
Click anywhere on the curve you just created, and select Edges > Apply in the Details View Window. Click
to create the surface.
30. Create C-Mesh Domain
Now that the airfoil has been generated, we need to create the meshable surface we will use once we begin to
specify boundary conditions. We will begin by creating a coordinate system at the tail of the airfoil - this will
help us create the geometry for the C-mesh domain. Click to create a new coordinate system. In the Details
View window, select Type > From Coordinates. For FD11, Point X, enter 1.
Click to generate the new coordinate system. In the Tree Outline Window, select the new coordinate
system you created (defaulted to Plane 4), then click to create a new sketch. This will create a sketching
plane on the XY plane with the tail of the airfoil as the origin. At the bottom of the Tree Outline Window, click
the Sketching tab to bring up the sketching window.
The first action we will take is create the arc of the C-Mesh domain. Click . The first click
selects the center of the arc, and the next two clicks determine the end points of the arc. We want the center of
the arc to be at the tail of the airfoil. Click on the origin of the sketch, making sure the P symbol is showing
31. For the end points of the arc, first select a point on the vertical axis above the origin (a C symbol will show),
then select a point on the vertical axis below the origin. You should end up with the following:
To create the right side of the C-Mesh donain, click . Click the following points to create
the rectangle in this order - where the arc meets the positive vertical axis, where the arc meets the negative
vertical axis, then anywhere in the right half plane. The final result should look like this:
Now, we need to get rid of necessary lines created by the rectangle. Select Modify in the Sketching Toolboxes
window, then select . Click the lines of the rectangle the are collinear with the positive and negative
vertical axises. Now, select the Dimensions toolbox to dimension the C-Mesh domain. Click ,
followed by the arc to dimension the arc. Assign the arc a value of 12.5. Next, select . Click the
vertical axis and the vertical portion of the rectangle in the right half plane. Also assign the horizontal
dimension a value of 12.5.
32. Click here to enlarge the image
Next, we need to create a surface from this sketch. To accomplish this, go to Concept > Surface From
Sketches. Click anywehere on the sketch, and select Base Objects > Apply in the Details View Window. Also,
select Operation > Add Frozen. Once you have the correct settings, click . The final step of creating
the C-Mesh is creating a surface between the boundary and the airfoil. To do this, go to Create > Boolean. In
the Details View window, select Operation > Subtract. Next, select Target Bodies > Not selected, select the
large C-Mesh domain surface, then click Apply. Repeat the same process to select the airfoil as the Tool Body.
When you have selected the bodies, click
Selecting the Airfoil Body
Because the C-Mesh domain and the airfoil overlap, once you click in the vicinity of the airfoil ANSYS will select the C-
Mesh domain but give you the option of selecting multiple layers
Select the layer that corresponds to the airfoil and the airfoil will be highlighted.
Create Quadrants
In the final step of creating the geometry, we will break up the new surface into 4 quadrants; this will be useful
for when we want to mesh the geometry. To begin, select Plane 4 in the Tree Outline Window, and click .
Open the sketching menu, and select . Draw a line on the vertical axis that intersects the entire C
mesh. Trim away the lines that are beyond the C-Mesh, and you should be left with this:
33. Next, go to Concepts > Lines from Sketchs. Select the line you just drew and click Base Objects > Apply,
followed by . Now that you have created a vertical line, create a new sketch and repeat the process
for a horizontal line that is collinear to horizontal axis and bisects the geometry.
Now, we need to project the lines we just created onto the surface. Go to Tools > Projection. Select Edges press
Ctrl and select on the vertical line we drew (you'll have to select both parts of it), then press Apply. Next, select
Target and select the C-Mesh surface, then click Apply.
Once you click , you'll notice that the geometry is now composed of two surfaces split by the line we
selected. Repeat this process to create 2 more projections: one projection the line left of the origin onto the left
34. surface, and one projecting the right line on the right surface. When you're finished, the geometry should be
split into 4 parts.
The geometry is finished. Save the project and close the design modeler, as we are now we are ready to create
the mesh for the simulation.
Mesh
Mapped Face Meshing
First, we will apply a mapped face meshing control to the geometry. In the Outline window, click on Mesh to
bring up the Meshing Toolbar. In the Meshing Toolbar, select Mesh Control > Mapped Face Meshing.
Making sure the face selection filter is selected , select all four faces by holding down the right mouse
button and dragging the mouse of all of the quadrants of the geometry. When all of the faces are highlighted
green, in the Details view Window select Geometry > Apply. Next, select
Edge Sizing
Next, we will apply edge sizing controls to all of the edges of the mesh. To begin, go to Mesh Control >
Sizing. Next, click the edge selection filter . Select the following 4 edges buy holding Ctrl and using the left
mouse button:
35. Click here to enlarge
Once the edges are selected, in the Details View Window select Geometry > Apply. Next, select Type >
Number of Divisions. Change the Number of Divisions to 50. Select Behavior > Hard. We also want the mesh
to have a bias, so select the first bias type: Bais > ----- — - -, and give the edge sizing a Bias Factor of 150. The
Edge sizing should now look like this:
Notice that the element sizes get smaller towards the airfoil. This will give us a better resolution around the
airfoil where the flow gets more complicated. Create a new edge sizing with the same parameters, but choose
the 4 remaining straight edges (see figure below). The number of divisions will still be 50, but now will be
selecting a different biasing type by selecting the second Bias option: Bias > - - — -----. Again, set the Bias
Factor to 150
36. Edge Bias
It is important to make sure that the edge divisions to this point are biased towards the center of the mesh:
otherwise you may run into some problems later. If your mesh does not match the pictures in the tutorial, make sure
to change the parameters of the mesh to make sure that they do: this might mean choosing different edges for the
different biasing types than those outlined in this tutorial.
Finally, create a third edge sizing, and select the rounded edges as the geometry. Again, select Type > Number
of Divisions, and change Number of Divisions to 100. Select Behavior > Hard. This time, we will not bias the
edges.
37. Now, select Mesh > Generate to generate the mesh. It should look like this.
Click here to enlarge
Named Selections
Now will assign names to some of the edges to make creating boundary conditions for the mesh easier. Let's
recall the boundary conditions we planned in the Pre-Analysis Step:
The edges highlighted blue are the inlet, the edges highlighted red are the outlet, and the airfoil is highlighted
white in the middle. Now we are ready to name the sections. In the Outline window, select geometry - this will
make seeing the edges a little easier. Again make sure the edge selection tool is selected. Now, select the
two vertical edges on the far right side of the mesh. Right click, and select Create Named Selections. Name the
edges outlet. Next, select the edges that correspond to the inlet of the flow as defined by the picture above.
38. Again, right click and select Create Named Selections and this time name the selection inlet. Finally, select
the two edges making up the airfoil, and name the selection airfoil
Setup(Physics)
Launch the Solver
In this step, we will open fluent and define the boundary conditions of the problem. If you haven't already, close
the meshing window to return to the Project Outline window. Now, click . This will load the
mesh into FLUENT. Now, double click Setup. The Fluent Launcher Window should open. Check the box
marked Double Precision. To make the solver run a little quicker, under Processing Options we will select
Parallel and change the Number of Processes to 2. This will allow users with a double core processor to utilize
both.
Select the Solver
Click OK to launch Fluent. The first thing we will do once Fluent launches is define the solver we are going to
use. Select Problem Setup > General. Under Solver, select Density-Based.
Models and Materials
Next, we will define the model we are going to use. We do this by going Problem Setup > Models > Viscous-
Laminar. Then press Edit... This will open the Viscous Model Menu Window. Select Inviscid and press OK.
Now, we will specify characteristics of the fluid. Because we specified the fluid as inviscid, we will only have
to define the density of the fluid. To make matters even simpler, we are only looking for non-dimensionalized
values like pressure coefficient, so we will define the density of our fluid to be 1 kg/m^3. To define the density,
click Problem Setup > Materials > (double click) Air. This will launch the Create/Edit Materials window.
39. Under Properties, ensure that density is set to Constant and enter 1 kg/m^3 as the density. Click
Change/Create to set the density.
Boundary Conditions
Inlet
Now that the fluid has been described, we are ready to set the boundary conditions of the simulation. Bring up
the boundary conditions menu by selecting Problem Setup > Boundary Conditions. In the Boundary
Conditions window, look under Zones. First, let's set the boundary conditions for the inlet. Select Inlet to see
the details of the boundary condition. The boundary condition type should have defaulted to velocity-inlet: if it
didn't, select it. Now, click Edit to bring up the Velocity-Inlet Window. We need to specify the magnitude and
direction of the velocity. Select Velocity Specification Method > Components. Remember, we want the flow to
enter the inlet at an angle of 6 degrees since the angle of attack of the airfoil is 6 degrees; thus, the x velocity
will be , and the y velocity will be . Specify X-Velocity as 0.9945 m/s and Y-Velocity as 0.1045
m/s. When you have finished specifying the velocity as entering the inlet at 6 degrees (the same thing as having
an angle of attack of 6 degrees), press OK
40. Outlet
In the Boundary Conditions window, look under Zones. Select Outlet to see the details of the boundary
condition. The boundary condition type should have defaulted to pressure-outlet: if it didn't, select it. Click
Edit, and ensure that the Gauge Pressure is defaulted to 0. If it is, you may close this window.
Airfoil
In the Boundary Conditions window, look under Zones and select airfoil. Select Type > Wall if it hasn't been
defaulted.
Reference Values
The final thing to do before we move on to solution is to acknowledge the reference values. Go to Problem
Setup > Reference Values. In the Reference Values Window, select Compute From > Inlet. Check the
reference values that appear to make sure they are as we have already set them.
Solution
Methods
First, go to Solution > Solution Methods. Everything in this section should have defaulted to what we want, but
let's make sure that under Flow the selection is Second Order Upwind. If this is the selection, we may move on.
Monitors
Now we are ready to begin solving the simulation. Before we hit solve though, we need to set up some
parameters for how Fluent will solve the simulation.
Let's begin by going to Solution > Monitors. In the Monitors Window, look under Residuals, Statistic, and
Force Monitors. Select Residuals - Print,Plot and press Edit. In the Residual Monitors Window, we want to
change all of the Absolute Criteria to 1e-6. This will give us some further trust in our solution.
Initial Guess
Now, we need to initialize the solution. Go to Solution > Solution Initialization. In the Solution Initialization
Window, select Compute From > Inlet. Ensure the values that appear are the same values we inputted in Step
5. If the are, initialize the solution by clicking Initialize.
41. Solve
Once the solution has been initialized, we are ready to solve the simulation. Go to Solution > Run Calculation.
Change Number of Iterations to 3000, then double click Calculate. Sit back and twiddle your thumbs until
Fluent spits out a converged solution.
Results
Velocity
First, we will look at the velocity vectors of the solution to see if the make intuitive sense. To plot the velocity
vectors, go to Results > Graphics and Animations. In the Graphics and Animations Window, select Vectors
and click Set Up.... This will bring up the Vectors Menu.
Make sure the settings of the menu match the figure above: namely Vectors of > Velocity, Color by > Velocity,
and set the second box as Velocity Magnitude. To see the velocity vectors, press Display.
42. Pressure Contours
To view the pressure contours over the entire mesh, go to Results > Graphics and Animations again, and in the
Graphics and Animations Window, select Contours.
Click Set Up... to bring up the Contours Menu. Check the box next to Filled. Under Contours Of, ensure that
the two boxes that are selected are Pressure... and Static Pressure.
Once these parameters are set, press Display to see the pressure contours.
Streamlines
To view the streamlines, keep the Contours window open, and change the Contours Of box to Velocity, and the
box below to Stream Function. Change Levels to 100. Also, uncheck the box marked Auto Range, and set
Min(kg/s) to 13.11, and Max(kg/s) to 14.16
43. To view the streamlines, press Display
Pressure Coefficient
Next, we will plot the pressure coefficient along the surface of the airfoil. Click on Results > Plots to open up
the Plots Window. Under Plots, select XY Plot, and click Set Up.... In the window that pops up, change the
settings Y-Axis Function > Pressure, and change the second box to Pressure Coefficient. Ensure X-Axis
Function > Direction Vector. Under Surfaces, select airfoil. See the figure below for help.
44. When all the settings are correct, press Plot to plot the data to the command window. To save the data to a text
file, check the box next to Write to File. You'll notice that the Plot button has been replaced by a button marked
Write..., click it. Change the file type to All Files and save the file name as Pressure_Coefficient.txt
Coefficients of Lift and Drag
To find the Coefficients of Lift and Drag, click Results > Reports to bring up the Reports Window. In the
Reports Window, select Forces and click Set Up.... This will bring up the Force Reports menu
We need to set the parameters so drag across the airfoil (keep in mind, which is at an angle) will be displayed.
In the Force Reports window change the Direction Vector such that X > .9945 and Y > .1045. Click Print to
print the drag coefficient to the command window. To print the lift coefficient, in the Force Reports window
change the Direction Vector such that X > -.1045 and Y > .9945. Again, press Print.
Verification and Validation
45. Verification
One of the ways we can verify our data is by refining the mesh. Open up the mesh, and increase the Number of
Divisions for Edge Sizing and Edge Sizing 2 to 100. Click Mesh in the Outline window, and in the Details
window, expand statistics. The number of elements should now be 40000.
Click here to enlarge
Exit out of the mesher. First, right click Setup and select Reset. Then click in the project
schematic. Open up the solver, and solve the simulation using the same solver and boundary conditions (you'll
have to input them again), but this time change the number of iterations to 5000. Again, calculate the force
coefficients and graph the pressure coefficient.
Validation
To validate our data, we will take a compare the data from actual experiment.
Unrefined Mesh Refined Mesh Experimental Data
Lift Coeffient 0.6315 0.6670 0.6630
Drag Coefficient 0.0122 0.0063 0.0090
Below is a graph displaying the comparing Coefficient of Pressure along the airfoil for the experimental data
and the CFD simulation. The data is from Gregory & O'Reilly, NASA R&M 3726, Jan 1970.
46. Click here to see an enlarged image
As we can see from the table and the graph, the CFD matches the data fairly well. There are inaccuracies from
factors like our assumption that the flow was inviscid, but we still managed to extract some meaningful
information from the simulation.