CFD analysis of Flow across an Aerofoil
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This document details a CFD analysis of airflow around an airfoil at varying angles of attack, specifically focusing on a NACA 0029 aerofoil. Key findings include the computation of lift and drag coefficients, highlighting a linear relationship between lift and angle of attack for low angles, while drag exhibits a non-linear relation. Results indicate that the study can be further validated through wind tunnel testing.
CFD analysis of Flow across an Aerofoil
- 1. CFD ANALYSIS OF FLOW AROUND AEROFOIL FOR DIFFERENT ANGLE OF ATTACKS www.mechieprojects.com
- 2. PRESENTATION OUTLINE • AIM • INTRODUCTION • LITERATURE SURVEY • CFD ANALYSIS OF AEROFOIL • RESULTS • CONCLUSIONS www.mechieprojects.com
- 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). www.mechieprojects.com
- 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 www.mechieprojects.com
- 13. MESH GENERATED IN GAMBIT www.mechieprojects.com
- 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
- 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 www.mechieprojects.com