import mesh
import mesh_c
import mesh_o
import mesh_su2
from potential import potential_flow_o, velocity, pressure,\
lift_n_drag, streamlines, potential_flow_o_n
import util
########################
# PERFIL AERODINAMICO
########################
# importar desde archivo
airfoil_points = 549
c = 1
filename = 'nombre_archivo_nube_de_puntos.csv'
perfil = airfoil.airfoil(c)
perfil.create(filename)
# perfil NACA 4
m = 0
p = 0
t = 12
c = 1
perfil = airfoil.NACA4(m, p, t, c)
perfil.create_sin(airfoil_points)
#########
# FLAP
#########
# importar desde archivo
filename_flap = 'nombre_archivo_nube_de_puntos_flap.csv'
def from_txt_mesh(filename='./garbage/mesh_own.txt_mesh'):
'''
importa malla de archivo txt_mesh. formato propio
'''
if filename[-8:] != 'txt_mesh':
print('WARNING!')
print('La extensión del archivo a importar no coincide con la \
extensión utilizada por este programa')
print('FINALIZANDO EJECUCIÓN')
exit()
return
with open(filename, 'r') as f:
mesh = f.readlines()
tipo = str(mesh[0].split()[-1])
d_eta = float(mesh[1].split()[-1])
d_xi = float(mesh[2].split()[-1])
R = float(mesh[3].split()[-1])
M = int(mesh[4].split()[-1])
N = int(mesh[5].split()[-1])
airfoil_alone = str(mesh[6].split()[-1])
airfoil_join = int(mesh[7].split()[-1])
if airfoil_alone == 'True':
airfoil_alone = True
elif airfoil_alone == 'False':
airfoil_alone = False
airfoil_boundary = np.fromstring(mesh[9], sep=',')
X = np.zeros((M, N))
Y = np.zeros((M, N))
i = 0
line = 11
for line_ in range(line, M + line):
X[i, :] = np.fromstring(mesh[line_], sep=',')
i += 1
line += M + 1
i = 0
for line_ in range(line, M + line):
Y[i, :] = np.fromstring(mesh[line_], sep=',')
i += 1
perfil = airfoil.airfoil(c=1)
perfil.x = X[:, 0]
perfil.y = Y[:, 0]
if tipo == 'O':
# mesh = mesh_o.mesh_O(R, M, N, perfil)
mesh = mesh_o.mesh_O(R, N, perfil)
elif tipo == 'C':
# mesh = mesh_c.mesh_C(R, M, N, perfil, from_file=True)
mesh = mesh_c.mesh_C(R, N, perfil, from_file=True)
# se asignan atributos a malla
mesh.d_eta = d_eta
mesh.d_xi = d_xi
mesh.R = R
mesh.M = M
mesh.N = N
mesh.airfoil_alone = airfoil_alone
mesh.airfoil_join = airfoil_join
mesh.airfoil_boundary = airfoil_boundary
mesh.X = X
mesh.Y = Y
return mesh
#!/usr/bin/python
import airfoil
reload(airfoil)
myairfoil = airfoil.airfoil()
myairfoil.load('clarky.dat')
myairfoil.createSketch()
from airfoil import airfoil from meshsolv import meshgen, solver camber = 0 position = 0 thickness = 12 AoA = 0 points = 10 #airfoil(camber,position,thickness,AoA(degree),half_the_points_on_airfoil) airfoil(camber, position, thickness, AoA, points) # Input: airfoil data points #meshgen() # Output: cgns_unstr # Input: cgns_unstr #solver() # Output: solution file along with multiple files.
#########################################################################################################
########################################### Input data ##################################################
#########################################################################################################
N = 101 # Airfoil number of points
M = 50 # 20
c = 1. # Chord [m]
m = 2. # Is the maximum camber (100 m is the first of the four digits),
p = 5. # Is the location of maximum camber (10 p is the second digit in the NACA xxxx description).
t = 12. # Is the maximum thickness as a fraction of the chord
# (so t gives the last two digits in the NACA 4-digit denomination divided by 100)
#########################################################################################################
#########################################################################################################
#########################################################################################################
af.airfoil(c, N, m, p, t)
cordxy = np.loadtxt('perfilCoordenadas.txt', float)
[X, Y] = mh.mesh(1., M, cordxy)
########################### Plot mesh & airfoil ###############################
plt.plot(X, Y, "b", X.T, Y.T, "g")
plt.title('Airfoil NACA %d%d%d' % (m, p, t))
plt.xlabel('x')
plt.ylabel('y')
plt.grid()
plt.axis("equal")
##################### Coordinates, Mesh & Boundary ############################
X = X[:-1, :] # remove the last element from the column
niter = f['step 1']['Niter']
#loop through the flap deflection range
for i in range(N_df):
#set the flap deflection for this loop iteration
df = df_min + float(i) * delta_df
#determine the filename
if df < 0.:
sgn = 'n'
else:
sgn = 'p'
filename = '{}{:0>2d}'.format(sgn,int(abs(df))) # filename cannot be greater than 6 characters long
#create the airfoil and run xfoil
myaf = af.airfoil(naca=NACA,xf=xf,flapType=0,delta=df)
m.add_airfoil(myaf)
m.write_xfoil(filename = filename+'.txt')
deflection_set = xfoil.generate_machup_from_xfoil(airfoils = [filename+'.txt'],
alpha_min = aoa_min,
alpha_max = aoa_max,
delta_alpha = delta_aoa,
reynolds_number = Re,
grid_size = grid,
niter = niter,
ncrit = ncrit,
xtrt = xtr,
xtrb = xtr,
flap = df,
xf = xf,
yf = 0.0)
def f_one_delta(self, x, naca, xf, flapType, alpha, targetCL):
self.myaf = af.airfoil(naca=naca, xf=xf, flapType=flapType, delta=x)
CL, CmLE, dummy = self.single(alpha)
return (CL - targetCL)