Originally posted by probligo
The Cm as I understand it, is the "turning moment" around the C/4 point that the airfoil develops at given angles of attack.
If you have a NACA0008 section at 0* then the moment (tendancy for the wing to rotate downward) will be zero. If you increase the angle to 2* then there will be lift generated but the airfoil will try to twist downward.
Can someone confirm the equation utilization that this Cm value plugs into? Is it analagous to lift equation:
L= 1/2 * rho * V^2 * S * Cl ?
In other words, is it M = 1/2 * rho * V^2 * S * Cm ?
If you use the same units for viscosity, velocity & area in which L results in lbs, does M result in a (foot * lbs) moment value?