flow past naca 0012 airfoil test

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School of Engineering, Taylors University

Flow past a NACA airfoil test

CFD Project 1 Report

Arthur Saw Sher-Qen

0301339

Bachelor of Engineering (Mechanical)

School of Engineering

Taylors University

MEC4513 Computational Fluid Dynamics

14 October 2012


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Arthur Saw Sher-Qen(0301339) Flow past a NACA airfoil test MEC4513

[School of Engineering, Taylors University] Page 2

Contents Page

1.0 Abstract 3

2.0 Introduction

2.1 Background

2.2Project objectives and problem statement3

3

3

3.0Methodology3.1Geometry3.2Meshing3.3FLUENT setup

44

5

6

4.0Results4.1Velocity4.2Pressure4.3Coefficient of lift and drag

7

7

8

9

5.0 Conclusion 9

6.0 References 10

9.0 Appendix 11


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1.0AbstractThis project main focus is to use computational fluid dynamics as a method of

simulation for a study of an airfoil deign with relation to its angle of attack and also to

explore the capabilities of the FLUENT software as a CFD tool. To reduce the

complexity and computational time of the model, various assumptions are made suchas no surface roughness and inviscid flow. The results are lift and drag coefficientswhich will be compared to published experimental data to validate the results. Also of

interest are the velocity and pressure profiles of each angle of attack which will

demonstrate the changes in lift produced at different angles of attack.

2.0Introduction2.1 Background

An airfoil is generally known as the shape of an airplane wing. What happens is

when air passes across an airfoil the airfoil generates am effect called lift. When theairfoil slices through the fluid medium known as air, the air is split and it flows over

the top and the bottom of the airfoil. On a flat plate or a symmetrical airfoil placed

horizontally or at zero angle of attack, nothing will happen apart from producing dragdue to the resistance between the airfoil and the air particles; the velocity of air across

the top and bottom of the airfoil is the same. As such there is no pressure difference to

cause any movement or force as according to Bernoullisprinciple (Figure 16) [7]. A

force will be generated when there is a pressure difference between the top andbottom section of the airfoil. If the pressure at the bottom is higher than the top due to

the velocity of air passing above the top of the airfoil being higher than the bottom;

then a force called lift will be generated pushing the airfoil upwards; if the oppositehappens, that is the pressure at the top is higher than the bottom then the force called

negative lift or downforce will instead push the airfoil downwards [1], [2].

Airfoil design and sections have come a long way since the earliest seriousdocumented work in the late 1800s. Generally, it is known that a flat plane is able to

produce lift when placed at an angle; the lift generated differs with the angle of

attack; this is because by angling the plane it forces the air to act with more force onone side of the plane thus producing negative or positive lift depending on the angle.

However, it was suspected that curvature shapes, which bear some resemblance to

one of natures ultimate flying machines, birds are also able to perform the job andmore efficiently too. This is because for curvature shapes, even at zero angle of attack,

the velocity of the flow is already different between the top and bottom of the airfoil

thus producing lift at low angles of attack [1].

2.2 Project objectives and problem statement

In the past understanding of the flow on an airfoil requires construction of

expensive prototypes and an expensive wind tunnel. It is cheaper for a small scaledesign, however to go full scale the cost was enormous and time consuming as a full

scale airfoil for example an airplane is huge in design and thus consumes a lot of

resources to build and operate.


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However, with advances in computational technology it has become easier to

model and study the design of airfoils with regards to various parameters, angle of

attack being one of them. This also forms the basis of this project which is to studythe flow, coefficient of lift and subsequently different angles of attack on a fixed

airfoil and also to explore and the about the potential applications of computational

fluid dynamics in modeling various applications.

An angle of attack is defined as the angle where the air at a certain velocity meets

the airfoil. The angle is measured from the chord line of the airfoil. As the angle of

attack increases, the coefficient of lift increases as well up to a certain point, above itis a phenomenon known as stalling in which the airfoil loses all lift; however, this

project will not cover stalling. It is to be noted that with the increase in angle of attack

the drag will also increase [3].

In this project, the angle of attack is fixed at 6 and 10 with all other conditions

as standard using a NACA 0012 airfoil design. The NACA naming sections represent

the camber, max camber and max thickness of the chord line based on the design of aparabola on the chord from the leading edge to the max chord line and back [5].

Figure 1: An airfoil section with airfoil terminology explained [3]

3.0Methodology3.1 GeometryThe airfoil section basic outline is first outlined in solidworks as it provides batter

control of sketches. It is then loaded into the design modeler in Ansys to be created as asurface. The C mesh is also designed in the design modeler module to create an operating

boundary for the airfoil.

The airfoil is designed in a XY or two dimensional domain for study, this isbecause we are only interested in the flow at the top and bottom section of the airfoil and

also for simplicity.


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Figure 2: The relative size and layout of the airfoil in meters.

As can be seen from the sketch, the NACA 0012 airfoil is symmetrical in design and thus

will only generate lift at an angle.

3.2MeshingTo reduce complexity and size of the project, both 6 and 10 degree angle of attack

airfoil will share the same geometry and mesh model in FLUENT. A 2D mesh model iscreated separated into four sections so that proper bias can be applied. The mesh is a 12.5

meter radius at the C section while the rectangular section also shares the same size.

The bias of the mesh is toward the center of the C mesh domain as we are

interested in the flow around and on the airfoil, as such by biasing the mesh toward the

center it will provide a more accurate result at the area of interest.

Figure 3: MeshAs can be seen from the grid, the element size is heavily biased or concentrated toward

the airfoil in the middle as it is the point of interest and should provide better results. The

mesh size is set at 15000 elements based on Cornells tutorial as aguide [6]. While a

outlet

inlet

airfoil


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more refined mesh (more elements) will produce a more accurate result; however, it will

also increase the complexity of the domain and will also increase the computational time.

As such to keep it simple, all simulation done in this project will retain this mesh.

3.3FLUENT SetupFLUENT is selected as the computational software for this project, the mesh isloaded into FLUENT for study and processing. The following settings is changed

at the FLUENT launcher to ensure a faster processing time; double precision is

used so the results are more accurate and parallel processing is selected with twoprocesses as the machine utilizes a dual core processor so the processing time can

be reduced.

i) SolverAs the operating medium of the airfoil is air, the type of solver chosen is

density based instead of pressure based, as later we are interested to also look

at the pressure distribution around the airfoil. The time is set as steady flowand gravity disabled for simplicity. Other settings are left as default.

ii) ModelsFor simplification, the flow is assumed to be inviscid, that is withoutviscosity. This is assuming an ideal air condition to be able to get the best

result. Nothing else is changed as the others are not considered in this study.

iii) MaterialsThe fluid medium is air and taken at a constant density of 1kg/m

3. This is an

ideal rounded up value of the air density and the solid, is left as default

aluminum.

iv) Boundary ConditionsSince the study is only focused on the airflow, no other boundary conditionshave been set apart from the velocity components at X and Y direction for the

inlet. This is because of the angle of attack; thus the velocity is split into X

and Y components with cosine (angle) for X and sinus (angle) for Y. Whilethe outlet is set as a pressure outlet at 0 gage pressure so that there can be a

flow from the inlet to the outlet due to pressure difference. The airfoil is set as

a wall as that is the solid layer the flow has to pass over and it is assumed

ideal, no slip condition.

v) SolutionThe solution method is set to use second order upwind method. While this

method may take more computational time but it produces a more accurateresult. The convergence of the solution is determined based on the scaled

residuals of the continuity and the momentum (velocity of X and Y)

equations. The solution convergence is set at 1e-6 as the accuracy level.When all the residuals have at least reached this level together the solution is


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considered to be converged. Since the mesh is biased toward the airfoil which

is the wall, better results will be obtained at that area.

Angle X component (m/s) Y component (m/s)

6 0.9945 0.1045

10 0.9848 0.1736Table 1: Angle of attack and the relative X and Y components for inlet boundarycondition

4.0ResultsThe aerodynamic results especially the coefficient of lift are compared to the works of

Chris C. Critzos, Harry H. Heyson, and Robert W. Boswinkle, Jr [4] on their work of theAerodynamic characteristics of NACA 0012 airfoil. The other sections of the results such

as velocity magnitude, pressure coefficients and etc are compared between the different

angles of attack just for study and comparison purposes.4.1 Velocity

Figure 4: 6 velocity vector

Figure 5: 10 velocity vector

As can be seen from the velocity vectors, at 6 angle of attack, the velocity doesnot have much change between the top and bottom section of the airfoil, as such the lift

generated will be lesser than that of the 10 angle of attack. In which it is noticed that that

is a significant different between the velocity at the top and bottom section as such therewill be more pressure difference to generate lift.


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4.2 Pressure

Figure 6: 6 pressure contour

Figure 7: 10 pressure contour

From Figure 6 and 7 it is more clearly illustrated that the 10 angle of attack iscapable of producing a higher angle of attack. As it is noted that the pressure at the

bottom of the 2nd

airfoil is clearly higher that the top as compared to the 1stairfoil. So

Bernoullis principle would dictate that at 10 attack angle more lift would be produceddue to higher pressure [7]. Although, it is noted that the tip or leading edge of the airfoil

the pressure is the highest, this is to be expected as that point interacts with the air firstbefore the other parts of the airfoil.

Figure 8: Graph of pressure coefficient at 6 angle of attack


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Figure 9: Graph of pressure coefficient at10 angle of attack

4.3 Coefficient of lift and drag

Angle Project value Experimental value

6 -0.005 0.48

10 0.629 0.91

Table 1: Lift coefficient

Angle Project value Experimental value

6 0.041 0.02

10 0.054 0.05

Table 2: Drag coefficient

From the comparison between the project value and the experimental value, it is clear

that lift coefficients are quite far off the mark while the drag coefficients are closer to the

experimental values. These errors can be attributed to the test conditions that wereapplied. The experimental values were obtained from a wind tunnel testing with surface

roughness and low turbulence wind tunnel system. While this project assumed ideal

conditions; inviscid flow and no surface roughness or gravity applied. Also a finer meshsystem could produce a more accurate result. While perhaps mistakes in setting up the

boundary conditions or mesh could also lead to errors in the simulation. However, as can

be seen from the data, as the angle of attack increases, the accuracy increases as well.

5.0ConclusionIn conclusion, based on the original objectives of this project to explore the use of

computational fluid dynamics to model various applications and to study about the

aerodynamic characteristics of an airfoil; the objective has been met.

By the use of the software FLUENT, it is possible to model the airfoil in variousangles of attack and determine its lift, drag, flow/ velocity around the airfoil. This

proves the usefulness of CFD modeling in the aerospace applications allowing

engineers to test any airfoil in any condition without resorting to expensive prototypesand wind tunnel testing at least in the initial stage.


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This project also proves the assumption that an increase in the angle of attack of

an airfoil its coefficient of lift will also increase as well as the drag partly due to more

surface area being exposed to the attacking air.

Also if accuracy in relative to real world applications is of a concern, then it is

recommended not to include so many ideal assumptions as the test result may varyquite a lot depending on the condition. As such, for example it is better to use laminarflow instead of inviscid and set the viscosity according to the Reynolds number of the

testing condition.

6.0 References

1. Adg.stanford.edu (n.d.)Airfoil History. [online] Available at:http://adg.stanford.edu/aa241/airfoils/airfoilhistory.html [Accessed: 10 Oct 2012].

2. Boeing.com (n.d.) What is an airfoil?. [online] Available at:http://www.boeing.com/companyoffices/aboutus/wonder_of_flight/airfoil.html[Accessed: 11 Oct 2012].

3. Centennialofflight.gov (n.d.)Angle of Attack. [online] Available at:http://www.centennialofflight.gov/essay/Dictionary/angle_of_attack/DI5.htm[Accessed: 12 Oct 2012].

4. Critzos, C. et al. (1955)AERODYNAMIC CHARACTERISTICS OF NACA 0012AIRFOIL SECTION AT ANGLES OF ATTACK FROM 0 TO 180. [report]

Washington: NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS,

p.23.

5. Desktop.aero (n.d.)Airfoil Geometry. [online] Available at:http://www.desktop.aero/appliedaero/airfoils1/airfoilgeometry.html [Accessed: 12

Oct 2012].

6. Confluence.cornell.edu (2011)ANSYS WB - Airfoil - Problem Specification -Simulation - Confluence. [online] Available at:https://confluence.cornell.edu/display/SIMULATION/ANSYS+WB+-+Airfoil+-

+Problem+Specification [Accessed: 09 Oct 2012].

7. Hyperphysics.phy-astr.gsu.edu (1998)Pressure. [online] Available at:http://hyperphysics.phy-astr.gsu.edu/hbase/pber.html [Accessed: 14 Oct 2012].


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Figure 14: 6 angle of attack full pressure coefficient

Figure 15: 10 angle of attack full pressure coefficient

Figure 16: Bernoullis principle


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Graph 1: Graph of angle of attack vs coefficient of lift [4]


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Graph 2: Graph of angle of attack vs coefficient of drag [4]

flow past naca 0012 airfoil test

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