def alpha_inv_analysis(airfoil, alpha_i, alpha_f, alpha_step, max_iter, id):
"""Inviscid analysis over range of angle of attacks."""
# Initializes airfoil and assigns NACA
xf = XFoil()
xf.naca(airfoil)
xf.max_iter = max_iter
# Collects values
a, cl, cd, cm, cp = xf.aseq(alpha_i, alpha_f, alpha_step)
x, cp_0 = xf.get_cp_distribution()
# Plots all the data
plot(a, cl, cd, cm, cp, x, cp_0, id)
def cl_visc_analysis(airfoil, cl_i, cl_f, cl_step, re, mach, max_iter, id):
"""Viscous analysis over range of lift coefficients."""
# Initializes airfoil and assigns NACA
xf = XFoil()
xf.naca(airfoil)
xf.max_iter = max_iter
xf.Re = re
xf.M = mach
# Collects values
a, cl, cd, cm, cp = xf.cseq(cl_i, cl_f, cl_step)
x, cp_0 = xf.get_cp_distribution()
# Plots all the data
plot(a, cl, cd, cm, cp, x, cp_0, id)
def getCoefficients(naca,
reynolds=1e6,
iterations=20,
angle=10,
angle_step=.5):
xf = XFoil()
xf.naca(f"{naca:04d}")
xf.Re = reynolds
xf.max_iter = 20
a, cl, cd, cm, cp = xf.aseq(0, angle, .5)
ratio = cl / cd
max_idx = np.nanargmax(ratio)
return (ratio[max_idx], a[max_idx]), a, cl, cd, max_idx
