def calibrate(self,
obj,
airfoil: Airfoil,
calibrate_x: bool = False,
calibrate_y: bool = True,
calibrate_w: bool = False):
"""
calibrates the splines to match the airfoil as good as possible (lstsq)
"""
mapping, lower_bounds_upper_spline, upper_bounds_upper_spline = self._get_bounds_and_mapping(
calibrate_x, calibrate_y, calibrate_w, upper=True)
mapping, lower_bounds_lower_spline, upper_bounds_lower_spline = self._get_bounds_and_mapping(
calibrate_x, calibrate_y, calibrate_w, upper=False)
bounds_upper_spline = (lower_bounds_upper_spline,
upper_bounds_upper_spline)
bounds_lower_spline = (lower_bounds_lower_spline,
upper_bounds_lower_spline)
upper_array = obj.Proxy.get_upper_array(obj)
lower_array = obj.Proxy.get_lower_array(obj)
new_upper_mat = self._calibrate_one_side(
upper_array, mapping, bounds_upper_spline,
airfoil.get_upper_data()[::-1])
new_lower_mat = self._calibrate_one_side(lower_array, mapping,
bounds_lower_spline,
airfoil.get_lower_data())
obj.upper_array = new_upper_mat.tolist()
obj.lower_array = new_lower_mat.tolist()
def flap_objective(x):
af = Airfoil()
af.read_txt('GA37A315mod_reworked.txt')
flapRatio = 0.3
# lipLen = 0.13 #0.1 - 0.27
# leRad = 0.15 #0.05 - 0.25
# gap = 0.02 #0.005;0.02
# overlap = -0.02 #-0.02 0.02
defl = 35
# vratio1 = 0.73 #0.5 0.95
# vratio2 = 1.3 #1.05 1.5
# vratio3 = 0.55 #0.2 0.9
lipLen = x[0]
leRad = x[1]
gap = x[2]
overlap = x[3]
vratio1 = x[4]
vratio2 = x[5]
vratio3 = x[6]
try:
Flap = flap_geometry(af,flapRatio,lipLen, leRad, gap, overlap, defl,vratio1,vratio2,vratio3)
polar = calc_Jpolar(0.06,3e6,Flap)
clmax = ny.max(polar.cl)
#print clmax
return -clmax,[],[]
except:
return 100,[],[]
def __init__(self, OC, AoA0=0., Sref=1., Lref=1., rel_thick=0.0, sweep=0.0, pcop=0.25, Ka=0.95, grad_active=True):
'''
Constructor for airfoils
@param AoA0:angle of attack of null lift
@param Cd0 : drag
@param Cm0 : pitch moment
@param relative_thickness : relative thickness for addition of a correction of ClAlpha
@param Sref : reference surface
@param Lref : reference length
'''
if rel_thick<0. or rel_thick>0.30:
raise Exception, "Relative thickness must be >0. and <0.30"
Airfoil.__init__(self, OC, Lref=Lref, Sref=Sref, grad_active=grad_active)
self.__AoA0_deg = None
self.__AoA0 = None
self.set_AoA0(AoA0)
self.__Ka = None
self.set_Ka(Ka)
self.__rel_thick = None
self.__rel_thick_grad = None
self.set_rel_thick(rel_thick)
self.__sweep = None
self.__sweep_grad = None
self.set_sweep(sweep)
self.pcop = pcop
def __init__(self, database=None, relative_thickness=0.15, interpolator='2DSpline',Sref=1., Lref=1.):
'''
Constructor for a polar of a simple airfoil.
@param Sref : reference surface
@param Lref : reference length
@param database : the dataBase in which the discrete values are stored, or None
'''
Airfoil.__init__(self,Sref,Lref,relative_thickness)
# self.__database=database
# self.__fd_database=Database()
if interpolator == '2DSpline':
print "building new spline 2D Interpolator"
x,y=self.__database.get_all_in_outs(self.DATABASE_ABSCISSA_TAG,self.DATABASE_AERO_COEFS_TAG)
y=array(y).T
x=array(x)
scale=array([1.,1.])
self.__cl_interpolator=Spline2D_interpolator(array(x),array(y[0,:]),scale)
self.__cd_interpolator=Spline2D_interpolator(array(x),array(y[1,:]),scale)
self.__cm_interpolator=Spline2D_interpolator(array(x),array(y[2,:]),scale)
print "Done."
else:
self.__cl_interpolator=aero_coefs_interpolator[0]
self.__cd_interpolator=aero_coefs_interpolator[1]
self.__cm_interpolator=aero_coefs_interpolator[2]
def __init__(self, OC, surrogate_model, relative_thickness=.12, camber=0., Sref=1., Lref=1., sweep=0., surrogate_fcs=None):
Airfoil.__init__(self, OC, Sref, Lref, relative_thickness, sweep, camber)
if not surrogate_fcs:
self.__coefs = SurrogateCoefs(surrogate_model)
else:
self.__coefs = surrogate_fcs
self.__surrogate_model = surrogate_model
def __init__(self, controls=None, proxy=None,
pos = Vector3(0,0,0),
attitude = Quaternion(0.5, -0.5, 0.5, 0.5),
velocity = Vector3(0,0,0),
thrust = 0, ident=None):
global cterrain
Airfoil.__init__(self, pos, attitude, velocity, thrust, cterrain)
ControlledSer.__init__(self, MyAirfoil.TYP, ident, proxy)
self.setControls(controls)
def __init__(self):
super().__init__()
self.af_left = Airfoil()
self.af_right = Airfoil()
self.it_point_list_right = []
self.it_point_list_left = []
self.af_left.data_changed.connect(self.connect_airfoils)
self.af_right.data_changed.connect(self.connect_airfoils)
def __init__(self,
controls=None,
proxy=None,
pos=Vector3(0, 0, 0),
attitude=Quaternion(0.5, -0.5, 0.5, 0.5),
velocity=Vector3(0, 0, 0),
thrust=0,
ident=None):
global cterrain
Airfoil.__init__(self, pos, attitude, velocity, thrust, cterrain)
ControlledSer.__init__(self, MyAirfoil.TYP, ident, proxy)
self.setControls(controls)
def __init__(self,
OC,
surrogate_model,
relative_thickness=.12,
camber=0.,
Sref=1.,
Lref=1.,
sweep=0.,
surrogate_fcs=None):
Airfoil.__init__(self, OC, Sref, Lref, relative_thickness, sweep,
camber)
if not surrogate_fcs:
self.__coefs = SurrogateCoefs(surrogate_model)
else:
self.__coefs = surrogate_fcs
self.__surrogate_model = surrogate_model
def __init__(self, OC, AoA0=0., Cm0=0.0, Sref=1., Lref=1., rel_thick=0.0, sweep=0.0, Ka=0.95):
'''
Constructor for airfoils
@param AoA0:angle of attack of null lift
@param Cd0 : drag
@param Cm0 : pitch moment
@param relative_thickness : relative thickness for addition of a correction of ClAlpha
@param Sref : reference surface
@param Lref : reference length
'''
if rel_thick<0. or rel_thick>0.25:
raise Exception, "Relative thickness must be >0. and <0.25"
Airfoil.__init__(self,OC, Sref,Lref,rel_thick,sweep)
self.__AoA0 = AoA0*np.pi/180.
self.__Cm0 = Cm0
self.__Ka = Ka
def __init__(self, OC, Sref=1., Lref=1., y_pos=None, grad_active=False):
'''
Constructor for airfoils
@param y_pos : span-wise position for interpolation
@param Sref : reference surface
@param Lref : reference length
'''
Airfoil.__init__(self, OC, Lref=Lref, Sref=Sref, grad_active=grad_active)
self.y_pos = y_pos
self.__y_def_list = None
self.__file_def_list = None
self.__interp_list = None
self.__index_p = None
self.__index_m = None
self.__fact_p = None
self.__fact_m = None
def plot_results():
# xBest =[0.18978,0.24883,0.00824,0.02996,0.50002,1.37683,0.20097]
#xBest =[0.18913399,0.13287908,0.00548619,0.02999233,0.50557791,1.49805884,0.20113115]
xBest=[0.19417077,0.24936138,0.01010853,0.0299996,0.50077753,1.39159708,0.2021984]
af = Airfoil()
af.read_txt('GA37A315mod_reworked.txt')
FlapOpt = flap_geometry(af, 0.3, xBest[0], xBest[1],xBest[2],xBest[3],35,xBest[4],xBest[5],xBest[6])
mainSec, flap, flap0 = FlapOpt
flap1 = Curves.rotate2D(flap,[0.785,-0.135],-35)
mainSec, flap, flap0 = FlapOpt
# write_txt(FlapOpt,'KLA100flap_20121213.txt')
plt.figure(1)
plt.hold(True)
plt.plot(mainSec[:,0],mainSec[:,1])
plt.plot(flap[:,0],flap[:,1])
plt.plot(flap0[:,0],flap0[:,1])
#plt.axis([0,1.1,-.5,.5])
plt.axis('equal')
plt.grid(True)
plt.show()
def generator():
"""
generates airofil data for defined set of airfoils
consider paths
"""
for name in ['mh114', 'mh115', 'mh117']:
airfoil = Airfoil(
'XFOIL',
r'E:/propeller/mh_airofils/{}/{}xfoil.txt'.format(name, name))
generate(airfoil,
r'E:\propeller\qprop\airfoildata\{}.txt'.format(name))
def test_func():
af = Airfoil()
af.read_txt('GA37A315mod_reworked.dat')
print af.up, af.coord
flapRatio = 0.3
lipLen = 0.1643088
leRad = 0.14054948
gap = 0.0157
overlap = -0.02
defl = 35
vratio1 = 0.77033947
vratio2 = 1.39164569
vratio3 = 0.55
Flap = flap_geometry(af, 0.3, lipLen, leRad,gap,overlap,35,vratio1,vratio2,vratio3)
mainSec, flap = Flap
polar = calc_Jpolar(0.06,3e6,Flap)
plt.figure(1)
plt.hold(True)
plt.plot(mainSec[:,0],mainSec[:,1])
plt.plot(flap[:,0],flap[:,1])
plt.axis([0,1.1,-.5,.5])
plt.grid(True)
plt.show()
def get_airfoil(self, obj):
return Airfoil.compute_vandevooren(obj.tau, obj.epsilon,
obj.numpoints * 2 + 1)
class ConnectedAirfoilsModel(QtCore.QObject):
data_changed = QtCore.pyqtSignal()
def __init__(self):
super().__init__()
self.af_left = Airfoil()
self.af_right = Airfoil()
self.it_point_list_right = []
self.it_point_list_left = []
self.af_left.data_changed.connect(self.connect_airfoils)
self.af_right.data_changed.connect(self.connect_airfoils)
def connect_airfoils(self):
if (self.af_left.loaded and self.af_right.loaded):
degrees = list(
set(self.af_left.get_degree_list() +
self.af_right.get_degree_list()))
degrees.sort()
self.it_point_list_right = self.af_right.get_interpolated_point_list(
degrees)
self.it_point_list_left = self.af_left.get_interpolated_point_list(
degrees)
self.data_changed.emit()
def generate_gcode(self, feedrate):
gcode = list()
if (self.af_left.loaded and self.af_right.loaded):
# turn on the spindle (hotwire) at the given temp
gcode.append("M3 S" + str(control_widget.temp_spbox.value()) +
"\n")
gcode.append("M7\n")
gcode.append("F%.2f\n" % feedrate)
# example coordinate for cutting block to size:
# (x, y) -> x = x and u, y = y and z
if (control_widget.cutBlockToSizeCheckbox.isChecked()):
# (0, 0)
gcode.append("G01 X%.3f Y%.3f U%.3f Z%.3f\n" % (0, 0, 0, 0))
# (0, block height)
gcode.append("G01 X%.3f Y%.3f U%.3f Z%.3f\n" %
(0, control_widget.block_height_spbox.value(), 0,
control_widget.block_height_spbox.value()))
# (block length, block height)
gcode.append("G01 X%.3f Y%.3f U%.3f Z%.3f\n" %
(control_widget.block_length_spbox.value(),
control_widget.block_height_spbox.value(),
control_widget.block_length_spbox.value(),
control_widget.block_height_spbox.value()))
# (block length, 0)
gcode.append("G01 X%.3f Y%.3f U%.3f Z%.3f\n" %
(control_widget.block_length_spbox.value(), 0,
control_widget.block_length_spbox.value(), 0))
# (block length, block height)
gcode.append("G01 X%.3f Y%.3f U%.3f Z%.3f\n" %
(control_widget.block_length_spbox.value(),
control_widget.block_height_spbox.value(),
control_widget.block_length_spbox.value(),
control_widget.block_height_spbox.value()))
# (0, block height)
gcode.append("G01 X%.3f Y%.3f U%.3f Z%.3f\n" %
(0, control_widget.block_height_spbox.value(), 0,
control_widget.block_height_spbox.value()))
# (0, 0)
gcode.append("G01 X%.3f Y%.3f U%.3f Z%.3f\n" % (0, 0, 0, 0))
# Cut the wing profiles
gcode.append("G01 X%.3f Y%.3f U%.3f Z%.3f\n" %
(self.af_right.start[0], self.af_right.start[1],
self.af_left.start[0], self.af_left.start[1]))
for i in range(len(self.it_point_list_right[0])):
gcode.append("G01 X%.3f Y%.3f U%.3f Z%.3f\n" %
(self.it_point_list_right[0][i],
self.it_point_list_right[1][i],
self.it_point_list_left[0][i],
self.it_point_list_left[1][i]))
gcode.append("G01 X%.3f Y%.3f U%.3f Z%.3f\n" %
(self.af_right.end[0], self.af_right.end[1],
self.af_left.end[0], self.af_left.end[1]))
return gcode
def foilProps(self):
# -------------------------------------------------------------------------------
# section to define airfoils and local properties
# -------------------------------------------------------------------------------
def foilname(t):
""" return airfoil name for given thickness, in % chord """
# thicknesses = np.array([9.8, 11, 13, 14.7, 16.2])
# names = ['mh117','mh115', 'mh114','mh113','mh112' ]
thicknesses = np.array([.05])
names = ['mh117']
return 'mh117'
a = np.insert(thicknesses, 0, 0)
thresholds = thicknesses - 0.5 * (thicknesses - a[:-1])
for i in range(len(thresholds) - 1):
if t < thresholds[i + 1] and t > thresholds[i]:
return names[i]
return names[i + 1]
self.airfoils = []
self.airfoilsdata = [] # use this to load preevaluated airofil data
# list of airfoil objects for each radial postion
for i in range(self.N_stations):
# print(foilname(self.t[i] / self.chord[i]*100))
self.airfoils.append(
Airfoil('XFOIL',
r'E:/propeller/mh_airofils/{}/{}xfoil.txt'.format(
foilname(self.ts[i] * 100),
foilname(self.ts[i] * 100)),
origin=0))
for name in ['mh112', 'mh113', 'mh114', 'mh115', 'mh117']:
self.airfoilsdata.append(
np.loadtxt(
r'E:\propeller\qprop\airfoildata\{}.txt'.format(name)))
def returnX(name):
i = 0
for nm in ['mh112', 'mh113', 'mh114', 'mh115', 'mh117']:
if nm == name:
return self.airfoilsdata[i]
i += 1
# arrays of airofil properties
self.Cl0 = np.zeros(self.N_stations)
self.Cla = np.zeros(self.N_stations)
self.Cd0 = np.zeros(self.N_stations)
self.ClCd0 = np.zeros(self.N_stations)
self.Cd2u = np.zeros(self.N_stations)
self.Cd2l = np.zeros(self.N_stations)
self.m = self.rs * 2 * self.rpm * np.pi / 60. / 340.
self.re = self.m * 340 * self.chords / 1.42e-5
print('\nr [m]\tchord\tfoil\tmach\tRe')
for i in range(self.N_stations):
print('{:.2f}\t{:.2f}\t{}\t{:.2f}\t{:.1f}'.format(
self.rs[i], self.chords[i], foilname(self.ts[i] * 100),
self.m[i], self.re[i]))
for i in range(self.N_stations):
# first option when dont have airfoil data
# second uses preevaluated airofil polar data
# self.Cl0[i], self.Cla[i], self.Cd0[i], self.ClCd0[i], self.Cd2u[i], self.Cd2l[i] = self.airfoils[i].qpropData(self.m[i], self.re[i])
self.Cl0[i], self.Cla[i], self.Cd0[i], self.ClCd0[i], self.Cd2u[
i], self.Cd2l[i] = readQPolar(
self.re[i],
X=returnX(foilname(self.ts[i] / self.chords[i] * 100)))
self.Clmin = np.ones(self.N_stations) * .2
self.Clmax = np.ones(self.N_stations) * 1
self.Reref = self.chords * self.rs
import numpy as np
import matplotlib.pyplot as plt
wind_speed = knots(10)
blade_speed = rpm(140)
ax = plt.subplot(1, 1, 1)
ax.set_xlabel('Wind Speed (knots)')
ax.set_ylabel('Blade Speed (rpm)')
import time
start_time = time.time()
i = 0
path = glob(r'C:\Users\mlip814\airfoil_data\airfoil_data\*')[56]
airfoil = Airfoil.from_folder(path, 100000, 9)
for n in (2, 3, 6, 12):
blade = Blade(airfoil, n)
torque = resistance(blade_speed) / n # (per blade)
blade.design(knots(10), torque, rpm(140))
wind_speeds = np.linspace(8, 12, 20)
results = [
blade.solve(knots(wind_speed), system) for wind_speed in wind_speeds
]
blade_speeds = [rads(result[0]) for result in results]
ax.plot(wind_speeds, blade_speeds, label=str(n))
ax.legend()
def genAirfoil(self, t=0):
""" generate airfoil"""
return Airfoil('XFOIL', self.xfoil, t=t)
def get_airfoil(self, obj, numpoints=50, curvature_factor=0.5):
coordinates = self.discretize(obj, numpoints, curvature_factor)
return Airfoil(coordinates)
def get_airfoil(self, obj):
if hasattr(self, "_airfoil") and self._airfoil:
return self._airfoil
else:
self._airfoil = Airfoil.import_from_dat(obj.filename)
return self._airfoil
def testFcn():
path = paths.CFD_paths('GA_opt')
cruise = Flight_conditions(1500, 48.92)
af = Airfoil()
af.CST_airfoil(ny.array([0.178551,0.273254,0.268906,0.226346]),ny.array([-0.178551,-0.101338,-0.255260,-0.043527]))
af.write_airfoil_txt('optAirfoil.dat')
af.create_af_CAT(save = path.file_igs)
Airfoil_mesh(path,cruise)
Airfoil_mesh.yplus_wall = 1.0
fluent = Solver(path,cruise)
fluent.turb_model = 'ke-realizable'
for alpha in range(10, 21, 2):
fluent.run_fluent(alpha)
af.polar.Re = cruise.Re
af.polar.M = cruise.Mach
af.polar.alpha = fluent.alpha
af.polar.CL = fluent.cl
af.polar.CD = fluent.cd
af.polar.CM = fluent.cm
#af.calc_Jpolar(0.1463,2999951,ny.array([-2,20,2.]))
af.write_polar_txt(r'D:\3. Projects\afOpt\experimental data\GAopt.pol')
#def batch_calculation():
designs = open(r'D:\3. Projects\afOpt\results\CFD_20120719\Designs.txt','rt')
Au = ny.zeros([4,1])
Al = ny.zeros([4,1])
lines = designs.readlines()
designs.close()
landing = Flight_conditions(0, 32.5)
path = paths.CFD_paths()
results = r'D:\3. Projects\afOpt\results\CFD_20120719'
for N,line in enumerate(lines):
path.set_name(('DoE_%d'%N))
segLine = line.split()
Au[0] = float(segLine[0])
Au[1] = float(segLine[1])
Au[2] = float(segLine[2])
Au[3] = float(segLine[3])
Al[0] = -float(segLine[0])
Al[1] = float(segLine[4])
Al[2] = float(segLine[5])
Al[3] = float(segLine[6])
af = Airfoil()
af.name = ('DoE %d'%N)
af.CST_airfoil(Au,Al)
af.create_af_CAT(save = path.file_igs)
Airfoil_mesh(path,landing)
Airfoil_mesh.yplus_wall = 1.0
fluent = Solver(path,landing)
fluent.turb_model = 'ke-realizable'
alphaSeq = ny.array([10, 12, 14, 16, 18, 20])
for alpha in alphaSeq:
fluent.run_fluent(alpha)
af.polar.Re = landing.Re
af.polar.M = landing.Mach
af.polar.alpha = fluent.alpha
af.polar.cl = fluent.cl
af.polar.cd = fluent.cd
af.polar.cm = fluent.cm
af.write_airfoil_txt(('%s\\airfoil_%d_coord.dat'%(results,N)))
af.write_polar_txt(('%s\\airfoil_%d.pol'%(results,N)))
from airfoil import Airfoil
if __name__ == "__main__":
af = Airfoil()
af.angle = 30
af.angle = 30
af.nodes =200
af.refinement = 3
af.samples = 10
af.viscosity = 0.0001
af.speed = 10.0
af.time = 1
af.execute()
def get_airfoil(self, obj):
return Airfoil.compute_joukowsky(obj.real_part + 1j * obj.imag_part,
obj.numpoints * 2 + 1)
def create_propeller(x,afOld):
diameter = 1.75
hubDiameter = 0.3
hubX = hubDiameter / diameter
baseAfPath = 'clark-Y.txt'
chordX = array([hubX,0.35,0.5,0.75,0.95])
pitchX = array([hubX,0.35,0.5,0.75,0.95])
thickX = array([hubX,0.35,0.5,0.75,0.95])
xnew = linspace(hubX,0.95,10)
betaSet = pitchX[3]
n1,n2,n3 = len(chordX), len(pitchX), len(thickX)
chordY = x[0:n1]
pitchY = x[n1:n1+n2]
thickY = x[n1+n2:]
method = 'spline'
chord = _interpolate(chordX,chordY,xnew,method)
pitch = _interpolate(pitchX,pitchY,xnew,method)
thick = _interpolate(thickX,thickY,xnew,method)
prop = propeller()
if afOld==None:
af = list()
for tc in thick:
afnew = Airfoil()
afnew.read_txt(baseAfPath)
afnew.analyze_geometry()
afnew.scale_thickness(tc,False)
afnew.build_aero_table(alphaSeq=[-40.0,40.0,1.0])
af.append(afnew)
prop.airfoil = af
else:
prop.airfoil = afOld
prop.name = 'MDO_project'
prop.diameter = diameter
prop.hubDiameter = hubDiameter
prop.chord = chord
prop.beta = pitch
prop.x = xnew
prop.r = xnew*diameter/2.0
prop.radius = diameter/2.0
prop.numBlades = 3.0
prop.thickness = thick
prop.numStations = len(xnew)
prop.analyze_geometry()
prop.set_beta(betaSet)
# if show:
# fig = plt.figure(1)
# ax1 = fig.add_subplot(311)
# ax1.grid(True)
# ax1.hold(True)
# ax1.plot(xnew,chord)
# ax2 = fig.add_subplot(312)
# ax2.grid(True)
# ax2.plot(xnew,pitch)
# ax3 = fig.add_subplot(313)
# ax3.grid(True)
# ax3.plot(xnew,thick)
# plt.show()
return prop
def get_airfoil(self, obj):
return Airfoil.compute_trefftz_kutta(
obj.real_part + 1j * obj.imag_part, obj.tau, obj.numpoints * 2 + 1)
def get_airfoil(self, obj):
return Airfoil.compute_naca(obj.naca_digits, obj.numpoints)
def _hitGround(self):
Airfoil._hitGround(self)
if self._velocity:
self.__play_tire = True
self.markChanged(full_ser=True)
def _hitGround(self): Airfoil._hitGround(self) if self._velocity: self.__play_tire=True self.markChanged(full_ser=True)
def flap_geometry2(x=None,args=1):
if x==None:
x = ny.array([0.15, 0.3, 0.05, 0.5, 0.1, 0.5, 0.02, 0.02])
flapRatio = 0.7
defl = 35.0
airfoil = Airfoil()
airfoil.read_txt('aiaa_cfd_final_optimum.txt')
offset = 0.005 # main section flap offset
lipTe = 0.0033 # trailing edge thickness
rTe = 0.01 # trailing edge radius
exitLipLen = x[0]
leRadPos = x[1]
entryLipLen = x[2]
v1 = x[3]
v2 = x[4]
v3 = x[5]
gap = x[6]
overlap = x[7]
upNode = ny.zeros([4,2])
loNode = ny.zeros([4,2])
upCurve = Curves.xyCurve(airfoil.upPts)
loCurve = Curves.xyCurve(airfoil.loPts)
thickness = upCurve(flapRatio)-loCurve(flapRatio) #local thickness
leRad = v2 * thickness # V2 length
leRadPos = leRadPos * thickness + loCurve(flapRatio)
upNode[0,0] = flapRatio + exitLipLen
upNode[0,1] = upCurve(upNode[0,0])
tan0 = upCurve.tan(upNode[0,0])
upNode[1,0] = upNode[0,0] - exitLipLen*v1
upNode[1,1] = -(upNode[0,0]-upNode[1,0])*tan0 + upNode[0,1]
loNode[0,0] = flapRatio + entryLipLen
loNode[0,1] = loCurve(loNode[0,0])
tan1 = loCurve.tan(loNode[0,0])
loNode[1,0] = loNode[0,0] - entryLipLen*v3
loNode[1,1] = loNode[0,1] - (loNode[0,0]-loNode[1,0])*tan1
upNode[2,0] = flapRatio
loNode[2,0] = flapRatio
upNode[3,0] = flapRatio
loNode[3,0] = flapRatio
upNode[3,1] = leRadPos
loNode[3,1] = leRadPos
upNode[2,1] = upNode[3,1] + leRad
loNode[2,1] = loNode[3,1] - leRad
curve3 = Curves.BezierCurve(upNode,10)
curve4 = Curves.BezierCurve(loNode,10)
curve4e = ny.flipud(curve4)
curve4e[:,2] = curve4[:,2]
curve4e = curve4e[1:]
curve4e[:,2] = curve4e[:,2]+1.0
curve11 = ny.vstack([curve3,curve4e])
curve11offset = Curves.xytCurveOffset(curve11,-offset)
curve6 = Curves.xyCurveSplit(upCurve.pts,upNode[0,0])
curve7 = Curves.xyCurveSplit(loCurve.pts,0,loNode[0,0])
curve8 = Curves.xyCurveSplit(loCurve.pts,loNode[0,0])
curve9, curve10 = Curves.trimTe(airfoil.upPts,curve11offset,lipTe)
curve7, curve10 = Curves.splitCurveSimple(curve7,curve10,[0,0])
curve12, circle1, curve13 = Curves.roundCorner(curve7,curve10,rTe)
mainSec = ny.vstack([ny.flipud(curve9),curve12,circle1,ny.flipud(curve13[:,0:2])])
flapSec1 = ny.vstack([ny.flipud(curve6)[:-1,:],curve11[:-1,0:2],curve8])
#flapSec0 = flapSec1
flapSec1 = Curves.rotate2D(flapSec1,[0.7,-.1],defl)
flapAdd = Curves.rotate2D(curve11,[0.7,-.1],defl)
refPt = mainSec[-1]
moveX1 = ny.min(flapSec1[:,0]) - refPt[0] + overlap
flapSec1[:,0] = flapSec1[:,0] - moveX1
flapAdd[:,0] = flapAdd[:,0] - moveX1
#gap
dy = Curves.adjustGap(flapAdd,refPt,gap)
flapSec1[:,1] = flapSec1[:,1] + dy
return mainSec, flapSec1
def genAirfoil(self, t=0):
return Airfoil('XY', x=self.x, y=self.y, t=t)