FLow over Airfoil
I am doing case study on Flow over airfoil at Low Reynolds number. i took NACA 4412 airfoil. I found all the equations but how can i found co-efficient of lift and Drag?
Can any one please tell me?
You can calculate Cl and Cd.
Lift = 0,5*rho*V^2*S*Cl
Drag = 0,5*rho*V^2*S*Cd
rho is air density (kg/m³)
V = Velocity (m/s)
S = Area (m²)
And the Lift and Drag you can get from the post processor.
Thank you Sir,
But How can i get Cl and Cd that is main difficulty i am facing in this case study.
Are the Cl and Cd depending on Airfoil or any other shape?
you should calculate Cl and Cd from Cf and Cp. you can find it in "Fundamentals Of Aerodynamics" by Anderson section 1.5. If you can not find it, send me an e-mail to sent this book to you.
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