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Path: ACPro/ACProAerodynamics
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Computes the ideal lift coefficient from 0 through supersonic.
The zero mach lift coefficient is
cL = cL0 + cLAlpha*alpha;
Also applies the Prandtl Glauert compressibility correction. Does not
account for stall.
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Form:
cL = LiftCoeff( alpha, cLAlpha, cL0, m, delta )
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Inputs
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alpha (1,:) Angle of attack
cLAlpha (1,1) Coefficient dcL/dAlpha
cL0 (1,1) Lift at zero angle of attack
m (1,:) Mach number
delta (1,1) If entered use the transonic damping
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Outputs
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cL (1,:) Lift coefficient
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See also PrandtlGlauertCorrection
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ACPro: ACProAerodynamics/PrandtlGlauertCorrection Common: Graphics/Plot2D
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