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Trying to process raw aerodynamic data measured directly in the wind tunnel on airfoil and wings. The width of the

tunnel test section is 12 inches. That would be the span of the infinite wing. So I have S [inch^2 / m^3] of 48.

The static air data is P= 29.32 and temperature of 74.4 degrees F. I believe Pa would then be 31397.331 and T (degrees C / K 27.6C / 296.71 K .

I think air density p [ kg/m^3] is 1.1658

I have a series of lift and drag results as follows


-4 -0.240 0.1100

-2 0.1000 0.0700

0 0.2600 0.1040 etc

In the end we have

CL= L [kgf] x 9.81 (for newtons)


q [Pa] x S [m2)

for coefficient of lift - I also will calculate coefficient of drag

I don't know what q would be {I think Pa 31397.331) would S be S above? (it is the surface area) and m^2?

Eventually I want to graph CL v AOA and CD vs AOA

Thank you

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