def get_max_cl(Re, r):
"""
Analyze airfoil at a fixed Re,
changing aoa from 10 to 15 by 0.1
and returns cl, cd, aoa that makes maximum cl
"""
xf = XFoil()
if r <= 0.175:
xf.airfoil = naca6409
else:
xf.airfoil = naca2412
xf.Re = Re
xf.Re = Re
xf.max_iter = 200
xf.n_crit = 9.00
xf.xtr = [1.00, 1.00]
xf.M = 0
a_seq, cl_seq, cd_seq, cm_seq, cp_seq = xf.aseq(10, 15, 0.1)
# ignore nan by making it 0
cl_seq = np.nan_to_num(cl_seq)
# find the maximum cl
cl_maxi = np.max(cl_seq)
# index of the maximum cl
idx = np.argmax(cl_seq)
return round(cl_maxi, 2), round(a_seq[idx], 2), round(cd_seq[idx], 2)
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 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
# XFOIL #######
from xfoil import XFoil
xf = XFoil()
# Import an airfoil
from xfoil.test import XXXXX
xf.airfoil = XXXXX
# Setting up the analysis parameters
xf.Re = re
xf.max_iter = 100
xf.M = 0.7
# Obtaining the angle of attack, lift coefficient, drag coefficient and momentum coefficient of the airfoil
a, cl, cd, cm = xf.aseq(0, 30, 0.5)
#!/usr/bin/python3
from xfoil import XFoil
import matplotlib.pyplot as plt
xf = XFoil()
xf.load("revclarky.dat")
xf.Re = 1e5
xf.M = 0
xf.max_iter = 100
a, cl, cd, cm, cp = xf.aseq(-2, 2, 0.5)
plt.plot(a, cl)
plt.title("alfa vs cl")
plt.show()
#plt.plot(xf.airfoil.x,xf.airfoil.y)
#plt.axis('equal')
#plt.show()
klist = [1]
for k in range(1):
print(k)
x = aifoils[k]['x'].values
y = aifoils[k]['y'].values
ang_low = -32
ang_high = 32
ang_spacing = 2
xfc = XFoil()
xfc.airfoil = Airfoil(x, y)
xfc.Re = re[k]
xfc.max_iter = 40
ac, clc, cdc, cmc, cpminc = xfc.aseq(ang_low, ang_high, ang_spacing)
df = aifoils_cfd[k].where(aifoils_cfd[k]['alpha [deg]'] >= aoa_l)
df = df.where(aifoils_cfd[k] <= aoa_h)
df = df.dropna() #.values
cl_panel = []
cdp_panel = []
for aoa in df['alpha [deg]']:
data_xy = aifoils[k].values
CL, CDP, Cp, pp = panel(data_xy[:, :], alfader=aoa)
cl_panel.append(CL)
cdp_panel.append(CDP)
cl_panel = np.array(cl_panel)
for line in range(len(aa)):
aux = [aa[line], b[line]]
string = str(aux).replace("[",
" ").replace(",",
" ").replace("]",
" ").replace("'", " ")
archive3.writelines(string + '\n')
aux.clear
# Calculates the aerodynamic coeff using Xfoil
archive3.close()
xf.load('lt.txt')
xf.Re = 1e3
xf.max_iter = 100
ann, cl, cd, cm, co = xf.aseq(1, 11, 1)
an = np.linspace(1, 10, 10)
desired = np.zeros(31)
desired[0:10] = cl
desired[10:20] = cd
desired[20:30] = cm
Surf, Iben, Itor, Xcg, Ycg = Struct.ArfoilProperties(1.0, Xwing_struct1)
desired[30] = Xcg
#2.Initialize Aero and Structural Models
#2.a. Aero Model
Elements, Xc, Xn, Xt, DL, A = Aero.InitializeAeroModel(Xwing_aero)
AoA, Vnorm = 0.0, 6.0
Vinf = np.array([np.cos(AoA / 57.3), np.sin(AoA / 57.3)]) * Vnorm #wind speed
#Vinf=Vinf*Vnorm
Vc = Xc * 0.0
from xfoil import XFoil from xfoil.test import naca0012 xf = XFoil() xf.airfoil = naca0012 xf.Re = 1e6 xf.max_iter = 40 a, cl, cd, cm, co = xf.aseq(-20, 20, 0.5)
