CFD analysis of air flow across an aeroplane wing is studied using CFD, using commercially available tools of Fluent and Gambit

1. CFD ANALYSIS OF FLOW AROUND
AEROFOIL FOR DIFFERENT ANGLE
OF ATTACKS
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2. PRESENTATION OUTLINE
• AIM
• INTRODUCTION
• LITERATURE SURVEY
• CFD ANALYSIS OF AEROFOIL
• RESULTS
• CONCLUSIONS
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3. AIM
• To understand the aerodynamic flow of air around an airfoil
• To study the change of angle of attack on the lift and drag forces of an airfoil for a
NACA series blade, using CFD (Fluent).
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4. WHAT ARE AIRFOIL?
• Mean Chamber Line: Set of points halfway between upper and lower surfaces
» Measured perpendicular to mean chamber line itself
• Leading Edge: Most forward point of mean chamber line
• Trailing Edge: Most reward point of mean chamber line
• Chord Line: Straight line connecting the leading and trailing edges
• Chord, c: Distance along the chord line from leading to trailing edge
• Chamber: Maximum distance between mean chamber line and chord line
» Measured perpendicular to chord line
• Airfoils are 3-D structures used to create aerodynamic forces

5. • Lift due to imbalance of pressure distribution over top and bottom surfaces of
airfoil (or wing)
– If pressure on top is lower than pressure on bottom surface, lift is generated
2. As V ↑ p↓
– Incompressible: Bernoulli’s Equation
– Compressible: Euler’s Equation
3. With lower pressure over upper surface and higher pressure over bottom surface,
airfoil feels a net force in upward direction → Lift
HOW DOES AN AIRFOIL GENERATE LIFT?
Lift = PA
VdVdp
Vp



 constant
2
1 2
Most of lift is produced in first 20-30% of wing (just downstream of leading edge)

6. RESOLVING THE AERODYNAMIC FORCE
• Relative Wind: Direction of V∞
• Angle of Attack, a: Angle between relative wind (V∞) and chord line
• Total aerodynamic force, R, can be resolved into two force components
– Lift, L: Component of aerodynamic force perpendicular to relative wind
– Drag, D: Component of aerodynamic force parallel to relative wind

7. SYMMETRIC AIRFOIL
Lift(fornow)
Angle of Attack, a
A symmetric airfoil generates zero lift at zero a

8. SAMPLE DATA: CAMBERED AIRFOIL
Lift(fornow)
Angle of Attack, a
A cambered airfoil generates positive lift at zero a

9. NACA FOUR-DIGIT SERIES
• First digit specifies maximum camber in percentage of chord
• Second digit indicates position of maximum camber in tenths of chord
• Last two digits provide maximum thickness of airfoil in percentage of chord
Example: NACA 2415
• Airfoil has maximum thickness of 15%
of chord (0.15c)
• Camber of 2% (0.02c) located 40%
back from airfoil leading edge (0.4c)
NACA 2415

10. NACA 0029
• SYMMETRIC AEROFOIL
• ZERO LIFT AT ZERO ANGLE OF ATTACK
• WIDELY USED BLADE IN WIND TURBINES AND LOW SPEED AIRCRAFTS
• LOT OF PREVIOUS STUDY ON THIS MODEL
• COORDINATES GENERATED IN MS-EXCEL
• COORDINATES EXPORTED TO AUTO-CAD AND PROFILE DRAWN.
NACA 0029- Profile generated in solid works

11. Methodology
• Profile co-ordinates generated in Excel
• Profile drawn in Auto-Cad
• Import geometry in Gambit and Meshing
• Import mesh in Fluent and apply boundary conditions
• Solve
• Post Processing & Results
• Boundary conditions:
– Air
– Velocity: 100m/s (at different angle of attack)
– In-compressible fluid flow analysis (Low Mach)
Velocity =100m/s
Angle(deg.) Angle(rad.) cosθ sinθ X-velocity Y-velocity
0 0 1 0 100 0
5 0.087266 0.996195 0.087156 99.61947 8.715574
10 0.174533 0.984808 0.173648 98.48078 17.36482
15 0.261799 0.965926 0.258819 96.59258 25.8819
20 0.349066 0.939693 0.34202 93.96926 34.20201
45 0.785398 0.707107 0.707107 70.71068 70.71068

12. MESH GENERATED IN GAMBIT
20c
12.5c
25c
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13. MESH GENERATED IN GAMBIT
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19. Flow comparison for 0 & 5 degree of attack

20. 5000
100 150
200 450
Static Pr.
Comparison for
different angle of
attack

21. 5000
100 150
200 450
Velocity
Comparison for
different angle of
attack

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00

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50

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100

25. 5000
100 150
200 450
Pr. Coefficient
Comparison for
different angle of
attack



q
pp
Cp
2
2
1



 



V
pp
q
pp
Cp


26. Angle(deg.) Cl Cd
0 0 0.028
5 0.47 0.049
10 0.83 0.114
15 1.08 0.227
20 1.20 0.358
45 1.26 1.045
Results
Conclusions
1.The flow around an aerofoil is plotted and studied.
2.The Lift & drag coefficients of NACA 0029 aerofoil is computed for different angle of attacks.
3.The lift coefficients is a linear function of angle of attack for lower α.
4.The drag coefficient is a non-linear function of α.
5.The results can be further validated through wind tunnel testing
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