Numerical calculation of Cl, Cd, Cm from Cf and Cp
I have a boundary layer code coupled with an inviscid solver that gives me the pressure coefficient and the skin friction coefficient in "s" (a coordinate along the airfoil that starts from the stagnation point).
My question is how can I correctly integrate Cp and Cf in oder to get Cd, Cl and Cm. I've found some formulas in "Fundamentals of Aerodynamics" by Anderson. However, these formulas are in terms of Cartesian coordinates and suggest the integration from the leading edge to the trailing edge.
1. Is it correct to assume that these formulas must be applied on the two parts of the airfoil obtained after we find the stagnation point and split the airfoil in two parts ?
2. Or should I transfer Cp and Cf back to Cartesian coordinates and integrate on the geometrical lower and upper side ?
Thanks for your understanding and answers for what could be a basic question.
You could use the formulas from Anderson to integrate Cp and Cf on the airfoil in a simpler way:
1. Walk the airfoil from the TE, lower side, LE, upper side, TE. This is the way classical panel method store the airfoil geometry.
2. Keep only the terms with "u" indices from the Anderson formulas. This simplification will avoid the need to split the data in upper/lower side.
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