def solve_nonlinear(self, params, unknowns, resids):
self.r = params['r']
self.chord = params['chord']
self.theta = params['theta']
self.Rhub = params['Rhub']
self.Rtip = params['Rtip']
self.hubHt = params['hubHt']
self.precone = params['precone']
self.tilt = params['tilt']
self.yaw = params['yaw']
self.precurve = params['precurve']
self.precurveTip = params['precurveTip']
self.airfoils = params['airfoils']
self.B = params['B']
self.rho = params['rho']
self.mu = params['mu']
self.shearExp = params['shearExp']
self.nSector = params['nSector']
self.tiploss = params['tiploss']
self.hubloss = params['hubloss']
self.wakerotation = params['wakerotation']
self.usecd = params['usecd']
self.Uhub = params['Uhub']
self.Omega = params['Omega']
self.pitch = params['pitch']
self.ccblade = CCBlade(self.r,
self.chord,
self.theta,
self.airfoils,
self.Rhub,
self.Rtip,
self.B,
self.rho,
self.mu,
self.precone,
self.tilt,
self.yaw,
self.shearExp,
self.hubHt,
self.nSector,
self.precurve,
self.precurveTip,
tiploss=self.tiploss,
hubloss=self.hubloss,
wakerotation=self.wakerotation,
usecd=self.usecd,
derivatives=True)
# power, thrust, torque
self.P, self.T, self.Q, self.M, self.dP, self.dT, self.dQ \
= self.ccblade.evaluate(self.Uhub, self.Omega, self.pitch, coefficients=False)
unknowns['T'] = self.T
unknowns['Q'] = self.Q
unknowns['P'] = self.P
def setUp(self):
# geometry
Rhub = 1.5
Rtip = 63.0
r = np.array([2.8667, 5.6000, 8.3333, 11.7500, 15.8500, 19.9500, 24.0500,
28.1500, 32.2500, 36.3500, 40.4500, 44.5500, 48.6500, 52.7500,
56.1667, 58.9000, 61.6333])
chord = np.array([3.542, 3.854, 4.167, 4.557, 4.652, 4.458, 4.249, 4.007, 3.748,
3.502, 3.256, 3.010, 2.764, 2.518, 2.313, 2.086, 1.419])
theta = np.array([13.308, 13.308, 13.308, 13.308, 11.480, 10.162, 9.011, 7.795,
6.544, 5.361, 4.188, 3.125, 2.319, 1.526, 0.863, 0.370, 0.106])
B = 3 # number of blades
# atmosphere
rho = 1.225
mu = 1.81206e-5
afinit = CCAirfoil.initFromAerodynFile # just for shorthand
basepath = path.join(path.dirname(path.realpath(__file__)), '5MW_AFFiles')
# load all airfoils
airfoil_types = [0]*8
airfoil_types[0] = afinit(path.join(basepath, 'Cylinder1.dat'))
airfoil_types[1] = afinit(path.join(basepath, 'Cylinder2.dat'))
airfoil_types[2] = afinit(path.join(basepath, 'DU40_A17.dat'))
airfoil_types[3] = afinit(path.join(basepath, 'DU35_A17.dat'))
airfoil_types[4] = afinit(path.join(basepath, 'DU30_A17.dat'))
airfoil_types[5] = afinit(path.join(basepath, 'DU25_A17.dat'))
airfoil_types[6] = afinit(path.join(basepath, 'DU21_A17.dat'))
airfoil_types[7] = afinit(path.join(basepath, 'NACA64_A17.dat'))
# place at appropriate radial stations
af_idx = [0, 0, 1, 2, 3, 3, 4, 5, 5, 6, 6, 7, 7, 7, 7, 7, 7]
af = [0]*len(r)
for i in range(len(r)):
af[i] = airfoil_types[af_idx[i]]
tilt = -5.0
precone = 2.5
yaw = 0.0
# create CCBlade object
self.rotor = CCBlade(r, chord, theta, af, Rhub, Rtip, B, rho, mu, precone, tilt, yaw, shearExp=0.2, hubHt=90.0)
# The parameters tiploss and hubloss toggle Prandtl tip and hub losses repsectively.
# The parameter wakerotation toggles wake swirl (i.e., a′=0).
# The parameter usecd can be used to disable the inclusion of drag in the
# calculation of the induction factors (it is always used in calculations of the
# distributed loads). However, doing so may cause potential failure in the
# solution methodology (see [4]). In practice, it should work fine, but special care
# for that particular case has not yet been examined, and the default implementation
# allows for the possibility of convergence failure.
# All four of these parameters are True by default.
# The parameter iterRe is for advanced usage.
# Referring to [4], this parameter controls the number of internal iterations on the Reynolds number.
# One iteration is almost always sufficient, but for high accuracy in the Reynolds number
# iterRe could be set at 2. Anything larger than that is unnecessary.
m1_rotor = CCBlade(m1_r, m1_chord, m1_theta, m1_af, m1_Rhub, m1_Rtip, B, m1_rho, m1_mu,
precone, tilt, yaw, shearExp, m1_hubHt, nSector)
m4_rotor = CCBlade(m4_r, m4_chord, m4_theta, m4_af, m4_Rhub, m4_Rtip, B, m4_rho, m4_mu,
precone, tilt, yaw, shearExp, m4_hubHt, nSector)
m6_rotor = CCBlade(m6_r, m6_chord, m6_theta, m6_af, m6_Rhub, m6_Rtip, B, m6_rho, m6_mu,
precone, tilt, yaw, shearExp, m6_hubHt, nSector)
m66_rotor = CCBlade(m66_r, m66_chord, m66_theta, m66_af, m66_Rhub, m66_Rtip, B, m66_rho, m66_mu,
precone, tilt, yaw, shearExp, m66_hubHt, nSector)
m66yaw20_rotor = CCBlade(m66_r, m66_chord, m66_theta, m66_af, m66_Rhub, m66_Rtip, B, m66_rho, m66_mu,
precone, tilt, 20.0, shearExp, m66_hubHt, nSector)
m67_rotor = CCBlade(m67_r, m67_chord, m67_theta, m67_af, m67_Rhub, m67_Rtip, B, m67_rho, m67_mu,
precone, tilt, yaw, shearExp, m67_hubHt, nSector)
# convert to RPM
m1_Omega = m1_Uinf*tsr/m1_Rtip * 30.0/pi
m4_Omega = m4_Uinf*tsr/m4_Rtip * 30.0/pi
m6_Omega = m6_Uinf*tsr/m6_Rtip * 30.0/pi
def solve_nonlinear(self, params, unknowns, resids):
self.r = params['r']
self.chord = params['chord']
self.theta = params['theta']
self.Rhub = params['Rhub']
self.Rtip = params['Rtip']
self.hubHt = params['hubHt']
self.precone = params['precone']
self.tilt = params['tilt']
self.yaw = params['yaw']
self.precurve = params['precurve']
self.precurveTip = params['precurveTip']
self.airfoils = params['airfoils']
self.B = params['B']
self.rho = params['rho']
self.mu = params['mu']
self.shearExp = params['shearExp']
self.nSector = params['nSector']
self.tiploss = params['tiploss']
self.hubloss = params['hubloss']
self.wakerotation = params['wakerotation']
self.usecd = params['usecd']
self.V_load = params['V_load']
self.Omega_load = params['Omega_load']
self.pitch_load = params['pitch_load']
self.azimuth_load = params['azimuth_load']
if len(self.precurve) == 0:
self.precurve = np.zeros_like(self.r)
# airfoil files
n = len(self.airfoils)
af = self.airfoils
# af = [0]*n
# for i in range(n):
# af[i] = CCAirfoil.initFromAerodynFile(self.airfoil_files[i])
self.ccblade = CCBlade(self.r, self.chord, self.theta, af, self.Rhub, self.Rtip, self.B,
self.rho, self.mu, self.precone, self.tilt, self.yaw, self.shearExp, self.hubHt,
self.nSector, self.precurve, self.precurveTip, tiploss=self.tiploss, hubloss=self.hubloss,
wakerotation=self.wakerotation, usecd=self.usecd, derivatives=True)
# distributed loads
Np, Tp, self.dNp, self.dTp \
= self.ccblade.distributedAeroLoads(self.V_load, self.Omega_load, self.pitch_load, self.azimuth_load)
# concatenate loads at root/tip
unknowns['loads_r'] = self.r
# conform to blade-aligned coordinate system
unknowns['loads_Px'] = Np
unknowns['loads_Py'] = -Tp
unknowns['loads_Pz'] = 0*Np
# return other outputs needed
unknowns['loads_V'] = self.V_load
unknowns['loads_Omega'] = self.Omega_load
unknowns['loads_pitch'] = self.pitch_load
unknowns['loads_azimuth'] = self.azimuth_load
# 1 ----------
precone = 0.0
precurve = np.linspace(0, 4.9, len(r))**2
precurveTip = 25.0
# create CCBlade object
rotor = CCBlade(r,
chord,
theta,
af,
Rhub,
Rtip,
B,
rho,
mu,
precone,
tilt,
yaw,
shearExp,
hubHt,
nSector,
precurve=precurve,
precurveTip=precurveTip)
# 1 ----------
plt.plot(precurve, r, 'k')
plt.plot(precurve, -r, 'k')
plt.axis('equal')
plt.grid()
def get_loads(Uinf,
yaw,
azimuth,
yawcorrection=False,
coupled=False,
return_bounds=False,
return_az_variation=False,
station_idx=0):
fname = 'H%02d%04d0' % (Uinf, yaw)
fid = open(fname + '.HD1', 'r')
fid.readline()
parts = fid.readline().split(',')
nchan = int(parts[0])
nsamp = int(parts[1])
fid.close()
th_idx = np.zeros(5)
tq_idx = np.zeros(5)
q_idx = np.zeros(5)
with open(fname + '.HD1', 'r') as inF:
for num, line in enumerate(inF):
if 'Thrust coefficient at 30% span' in line:
th_idx[0] = int(num) - 2
elif 'Thrust coefficient at 47% span' in line:
th_idx[1] = int(num) - 2
elif 'Thrust coefficient at 63% span' in line:
th_idx[2] = int(num) - 2
elif 'Thrust coefficient at 80% span' in line:
th_idx[3] = int(num) - 2
elif 'Thrust coefficient at 95% span' in line:
th_idx[4] = int(num) - 2
if 'Torque coefficient at 30% span' in line:
tq_idx[0] = int(num) - 2
elif 'Torque coefficient at 47% span' in line:
tq_idx[1] = int(num) - 2
elif 'Torque coefficient at 63% span' in line:
tq_idx[2] = int(num) - 2
elif 'Torque coefficient at 80% span' in line:
tq_idx[3] = int(num) - 2
elif 'Torque coefficient at 95% span' in line:
tq_idx[4] = int(num) - 2
elif 'QNORM30' in line:
q_idx[0] = int(num) - 2
elif 'QNORM47' in line:
q_idx[1] = int(num) - 2
elif 'QNORM63' in line:
q_idx[2] = int(num) - 2
elif 'QNORM80' in line:
q_idx[3] = int(num) - 2
elif 'QNORM95' in line:
q_idx[4] = int(num) - 2
elif 'Blade azimuth angle' in line:
az_idx = int(num) - 2
elif 'CONE' in line:
cone_idx = int(num) - 2
elif 'RHOTUN' in line:
rho_idx = int(num) - 2
elif 'RPM' in line:
rpm_idx = int(num) - 2
data = get_eng_data(fname + '.ENG', nsamp, nchan)
az_vec = data[:, az_idx]
idx = np.nonzero(np.abs(az_vec - azimuth) < 1)
cone_vec = data[:, cone_idx]
cone = np.mean(cone_vec[idx])
rho_vec = data[:, rho_idx]
rho = np.mean(rho_vec[idx])
rpm_vec = data[:, rpm_idx]
rpm = np.mean(rpm_vec[idx])
r_exp = [0.3, 0.47, 0.63, 0.8, 0.95]
c_exp = [0.711, 0.62492758, 0.54337911, 0.457, 0.381]
q_exp = np.zeros(5)
cn_exp = np.zeros(5)
ct_exp = np.zeros(5)
th_mean = np.zeros(5)
th_max = np.zeros(5)
th_min = np.zeros(5)
tq_mean = np.zeros(5)
tq_max = np.zeros(5)
tq_min = np.zeros(5)
th_vec_all = np.zeros((5, nsamp))
tq_vec_all = np.zeros((5, nsamp))
for i in range(5):
cth_vec = data[:, th_idx[i]]
ctq_vec = data[:, tq_idx[i]]
q_vec = data[:, q_idx[i]]
th_vec = cth_vec * q_vec * c_exp[i] / np.cos(np.radians(cone_vec))
tq_vec = ctq_vec * q_vec * c_exp[i] / np.cos(np.radians(cone_vec))
th_mean[i] = np.mean(th_vec[idx])
th_max[i] = np.max(th_vec[idx])
th_min[i] = np.min(th_vec[idx])
tq_mean[i] = np.mean(tq_vec[idx])
tq_max[i] = np.max(tq_vec[idx])
tq_min[i] = np.min(tq_vec[idx])
th_vec_all[i, :] = th_vec
tq_vec_all[i, :] = tq_vec
Np_exp = th_mean
Tp_exp = tq_mean
Np_min = th_min
Tp_min = tq_min
Np_max = th_max
Tp_max = tq_max
# geometry
Rhub = 0.508
Rtip = 5.029
# 0.508,
# r = np.array([0.660, 0.883, 1.008, 1.067, 1.133, 1.257, 1.343, 1.510, 1.648,
# 1.952, 2.257, 2.343, 2.562, 2.867, 3.172, 3.185, 3.476, 3.781, 4.023, 4.086, 4.391,
# 4.696, 4.780, 5.000]) # , 5.305, 5.532]
# # 0.218,
# chord = np.array([0.218, 0.183, 0.349, 0.441, 0.544, 0.737, 0.728, 0.711, 0.697,
# 0.666, 0.636, 0.627, 0.605, 0.574, 0.543, 0.542, 0.512, 0.482, 0.457, 0.451, 0.420,
# 0.389, 0.381, 0.358]) # , 0.328, 0.305])
# # 0.0,
# theta = np.array([0.0, 0.0, 6.7, 9.9, 13.4, 20.04, 18.074, 14.292, 11.909, 7.979,
# 5.308, 4.715, 3.425, 2.083, 1.15, 1.115, 0.494, -0.015, -0.381, -0.475, -0.920,
# -1.352, -1.469, -1.775]) # , -2.191, -2.5])
# af_idx = [0, 0, 0, 0, 0, 3, 3, 4, 5, 6, 7, 7, 7, 7, 7, 7, 7, 7, 7, 7, 7, 7, 7, 7]
r = np.array([
0.6060, 0.8801, 1.2321, 1.5087, 1.7099, 1.9278, 2.1457, 2.3469, 2.5480,
2.7660, 2.9840, 3.1850, 3.3862, 3.6041, 3.8220, 4.0232, 4.2244, 4.4004,
4.5764, 4.7776, 4.9536
])
chord = np.array([
0.219, 0.181, 0.714, 0.711, 0.691, 0.668, 0.647, 0.627, 0.606, 0.584,
0.561, 0.542, 0.522, 0.499, 0.478, 0.457, 0.437, 0.419, 0.401, 0.381,
0.363
])
theta = np.array([
0.000, -0.098, 19.423, 14.318, 10.971, 8.244, 6.164, 4.689, 3.499,
2.478, 1.686, 1.115, 0.666, 0.267, -0.079, -0.381, -0.679, -0.933,
-1.184, -1.466, -1.711
])
af_idx = [0, 2, 3, 4, 5, 6, 7, 7, 7, 7, 7, 7, 7, 7, 7, 7, 7, 7, 7, 7, 7]
B = 2 # number of blades
# # atmosphere
# rho = 1.225
# rho = 1.224385
mu = 1.81206e-5
# import os
afinit = CCAirfoil.initFromAerodynFile # just for shorthand
basepath = 'af_files2' + os.path.sep
airfoil_types = [0] * 8
airfoil_types[0] = afinit(basepath + 'cylinder.dat')
airfoil_types[1] = afinit(basepath + 'S809_CLN_129.dat') # .648741
airfoil_types[2] = afinit(basepath + 'S809_CLN_185.dat') # .930365
airfoil_types[3] = afinit(basepath + 'S809_CLN_242.dat') # 1.217018
airfoil_types[4] = afinit(basepath + 'S809_CLN_298.dat') # 1.498642
airfoil_types[5] = afinit(basepath + 'S809_CLN_354.dat') # 1.780266
airfoil_types[6] = afinit(basepath + 'S809_CLN_410.dat') # 2.06189
airfoil_types[7] = afinit(basepath + 'S809_CLN_Outboard.dat')
# ere is a cylindrical section at the root that extends from 0.508 to 0.883 m. The airfoil transition begins at approximately the 0.883- m radial station.
af = [0] * len(r)
for i in range(len(r)):
af[i] = airfoil_types[af_idx[i]]
tilt = 0.0
precone_ccblade = -cone
yaw_ccblade = yaw
shearExp = 0.05
hubHt = 12.192
nSector = 8
# create CCBlade object
aeroanalysis = CCBlade(r,
chord,
theta,
af,
Rhub,
Rtip,
B,
rho,
mu,
precone_ccblade,
tilt,
yaw_ccblade,
shearExp,
hubHt,
nSector,
yawcorrection=yawcorrection,
coupled=coupled)
# # set conditions
# Uinf = 10.0
# tsr = 7.55
pitch = 3.0
Omega = rpm # Uinf*tsr/Rtip * 30.0/pi # convert to RPM
# azimuth_ccblade = (360.0 - azimuth) % 360 # opposite definition
azimuth_ccblade = azimuth
# evaluate distributed loadsp
Np, Tp = aeroanalysis.distributedAeroLoads(Uinf, Omega, pitch,
azimuth_ccblade)
q = 0.5 * rho * (Uinf + Omega * r * 3.14 / 30)**2
q_exp = np.interp(r_exp, r / Rtip, q)
# cn = Np / (q * chord)
# ct = Tp / (q * chord)
# cn_exp = th_mean / (q_exp * c_exp)
# ct_exp = tq_mean / (q_exp * c_exp)
rstar = r / Rtip
if return_bounds:
return rstar, Np, Tp, r_exp, Np_exp, Tp_exp, Np_min, Tp_min, Np_max, Tp_max
elif return_az_variation:
Np_vec_exp = th_vec_all[station_idx, :]
Tp_vec_exp = tq_vec_all[station_idx, :]
n = 200
az_vec_sim = np.linspace(0, 360, n)
Np_vec = np.zeros(n)
Tp_vec = np.zeros(n)
for iii in range(n):
Np, Tp = aeroanalysis.distributedAeroLoads(Uinf, Omega, pitch,
az_vec_sim[iii])
Np_vec[iii] = np.interp(r_exp[station_idx], r / Rtip, Np)
Tp_vec[iii] = np.interp(r_exp[station_idx], r / Rtip, Tp)
return az_vec, Np_vec_exp, Tp_vec_exp, az_vec_sim, Np_vec, Tp_vec
else:
return rstar, Np, Tp, r_exp, Np_exp, Tp_exp
for i in range(len(r)):
af[i] = airfoil_types[af_idx[i]]
# 2 ----------
# 3 ----------
# create CCBlade object
rotor = CCBlade(r,
chord,
theta,
af,
Rhub,
Rtip,
B,
rho,
mu,
precone,
tilt,
yaw,
shearExp,
hubHt,
nSector,
derivatives=True)
# 3 ----------
# 4 ----------
# set conditions
Uinf = 10.0
tsr = 7.55
pitch = 0.0
def CCrotor(
Rtip=base_R,
Rhub=1.5,
hubHt=90.0,
shearExp=0.2,
rho=1.225,
mu=1.81206e-5,
path_to_af="5MW_AFFiles",
):
r = (Rtip / base_R) * np.array([
2.8667,
5.6000,
8.3333,
11.7500,
15.8500,
19.9500,
24.0500,
28.1500,
32.2500,
36.3500,
40.4500,
44.5500,
48.6500,
52.7500,
56.1667,
58.9000,
61.6333,
])
chord = (Rtip / base_R) * np.array([
3.542,
3.854,
4.167,
4.557,
4.652,
4.458,
4.249,
4.007,
3.748,
3.502,
3.256,
3.010,
2.764,
2.518,
2.313,
2.086,
1.419,
])
theta = np.array([
13.308,
13.308,
13.308,
13.308,
11.480,
10.162,
9.011,
7.795,
6.544,
5.361,
4.188,
3.125,
2.319,
1.526,
0.863,
0.370,
0.106,
])
B = 3 # number of blades
# In this initial version, hard-code to be NREL 5MW
afinit = CCAirfoil.initFromAerodynFile # just for shorthand
basepath = path_to_af
# load all airfoils
airfoil_types = [0] * 8
airfoil_types[0] = afinit(path.join(basepath, "Cylinder1.dat"))
airfoil_types[1] = afinit(path.join(basepath, "Cylinder2.dat"))
airfoil_types[2] = afinit(path.join(basepath, "DU40_A17.dat"))
airfoil_types[3] = afinit(path.join(basepath, "DU35_A17.dat"))
airfoil_types[4] = afinit(path.join(basepath, "DU30_A17.dat"))
airfoil_types[5] = afinit(path.join(basepath, "DU25_A17.dat"))
airfoil_types[6] = afinit(path.join(basepath, "DU21_A17.dat"))
airfoil_types[7] = afinit(path.join(basepath, "NACA64_A17.dat"))
# place at appropriate radial stations
af_idx = [0, 0, 1, 2, 3, 3, 4, 5, 5, 6, 6, 7, 7, 7, 7, 7, 7]
af = [0] * len(r)
for i in range(len(r)):
af[i] = airfoil_types[af_idx[i]]
tilt = -5.0
precone = 2.5
yaw = 0.0
nSector = 8 # azimuthal discretization
rotor = CCBlade(
r,
chord,
theta,
af,
Rhub,
Rtip,
B,
rho,
mu,
precone,
tilt,
yaw,
shearExp,
hubHt,
nSector,
)
return rotor
def solve_nonlinear(self, params, unknowns, resids):
self.ccblade = CCBlade(params['r'], params['chord'], params['theta'], params['airfoils'], params['Rhub'], params['Rtip'], params['B'], params['rho'], params['mu'], params['precone'], params['tilt'], params['yaw'], params['shearExp'], params['hubHt'], params['nSector'])
Uhub = np.linspace(params['control_Vin'],params['control_Vout'], self.n_pc)
P_aero = np.zeros_like(Uhub)
Cp_aero = np.zeros_like(Uhub)
P = np.zeros_like(Uhub)
Cp = np.zeros_like(Uhub)
T = np.zeros_like(Uhub)
Q = np.zeros_like(Uhub)
M = np.zeros_like(Uhub)
Omega = np.zeros_like(Uhub)
pitch = np.zeros_like(Uhub) + params['control_pitch']
Omega_max = min([params['control_maxTS'] / params['Rtip'], params['control_maxOmega']*np.pi/30.])
# Region II
for i in range(len(Uhub)):
Omega[i] = Uhub[i] * params['control_tsr'] / params['Rtip']
P_aero, T, Q, M, Cp_aero, _, _, _ = self.ccblade.evaluate(Uhub, Omega * 30. / np.pi, pitch, coefficients=True)
P, eff = CSMDrivetrain(P_aero, params['control_ratedPower'], params['drivetrainType'])
Cp = Cp_aero*eff
# search for Region 2.5 bounds
for i in range(len(Uhub)):
if Omega[i] > Omega_max and P[i] < params['control_ratedPower']:
Omega[i] = Omega_max
Uhub[i] = Omega[i] * params['Rtip'] / params['control_tsr']
P_aero[i], T[i], Q[i], M[i], Cp_aero[i], _, _, _ = self.ccblade.evaluate([Uhub[i]], [Omega[i] * 30. / np.pi], [pitch[i]], coefficients=True)
P[i], eff = CSMDrivetrain(P_aero[i], params['control_ratedPower'], params['drivetrainType'])
Cp[i] = Cp_aero[i]*eff
regionIIhalf = True
i_IIhalf_start = i
unknowns['V_R25'] = Uhub[i]
break
if P[i] > params['control_ratedPower']:
regionIIhalf = False
break
def maxPregionIIhalf(pitch, Uhub, Omega):
Uhub_i = Uhub
Omega_i = Omega
pitch = pitch
P, _, _, _ = self.ccblade.evaluate([Uhub_i], [Omega_i * 30. / np.pi], [pitch], coefficients=False)
return -P
# Solve for regoin 2.5 pitch
options = {}
options['disp'] = False
options['xatol'] = 1.e-2
if regionIIhalf == True:
for i in range(i_IIhalf_start + 1, len(Uhub)):
Omega[i] = Omega_max
pitch0 = pitch[i-1]
bnds = [pitch0 - 10., pitch0 + 10.]
pitch_regionIIhalf = minimize_scalar(lambda x: maxPregionIIhalf(x, Uhub[i], Omega[i]), bounds=bnds, method='bounded', options=options)['x']
pitch[i] = pitch_regionIIhalf
P_aero[i], T[i], Q[i], M[i], Cp_aero[i], _, _, _ = self.ccblade.evaluate([Uhub[i]], [Omega[i] * 30. / np.pi], [pitch[i]], coefficients=True)
P[i], eff = CSMDrivetrain(P_aero[i], params['control_ratedPower'], params['drivetrainType'])
Cp[i] = Cp_aero[i]*eff
if P[i] > params['control_ratedPower']:
break
def constantPregionIII(pitch, Uhub, Omega):
Uhub_i = Uhub
Omega_i = Omega
pitch = pitch
P_aero, _, _, _ = self.ccblade.evaluate([Uhub_i], [Omega_i * 30. / np.pi], [pitch], coefficients=False)
P, eff = CSMDrivetrain(P_aero, params['control_ratedPower'], params['drivetrainType'])
return abs(P - params['control_ratedPower'])
if regionIIhalf == True:
# Rated conditions
def min_Uhub_rated_II12(min_params):
return min_params[1]
def get_Uhub_rated_II12(min_params):
Uhub_i = min_params[1]
Omega_i = Omega_max
pitch = min_params[0]
P_aero_i, _, _, _ = self.ccblade.evaluate([Uhub_i], [Omega_i * 30. / np.pi], [pitch], coefficients=False)
P_i,eff = CSMDrivetrain(P_aero_i, params['control_ratedPower'], params['drivetrainType'])
return abs(P_i - params['control_ratedPower'])
x0 = [pitch[i] + 2. , Uhub[i]]
bnds = [(pitch0, pitch0 + 10.),(Uhub[i-1],Uhub[i+1])]
const = {}
const['type'] = 'eq'
const['fun'] = get_Uhub_rated_II12
params_rated = minimize(min_Uhub_rated_II12, x0, method='SLSQP', tol = 1.e-2, bounds=bnds, constraints=const)
U_rated = params_rated.x[1]
if not np.isnan(U_rated):
Uhub[i] = U_rated
pitch[i] = params_rated.x[0]
else:
print('Regulation trajectory is struggling to find a solution for rated wind speed. Check rotor_aeropower.py')
Omega[i] = Omega_max
P_aero[i], T[i], Q[i], M[i], Cp_aero[i], _, _, _ = self.ccblade.evaluate([Uhub[i]], [Omega[i] * 30. / np.pi], [pitch0], coefficients=True)
P_i, eff = CSMDrivetrain(P_aero[i], params['control_ratedPower'], params['drivetrainType'])
Cp[i] = Cp_aero[i]*eff
P[i] = params['control_ratedPower']
else:
# Rated conditions
def get_Uhub_rated_noII12(pitch, Uhub):
Uhub_i = Uhub
Omega_i = min([Uhub_i * params['control_tsr'] / params['Rtip'], Omega_max])
pitch_i = pitch
P_aero_i, _, _, _ = self.ccblade.evaluate([Uhub_i], [Omega_i * 30. / np.pi], [pitch_i], coefficients=False)
P_i, eff = CSMDrivetrain(P_aero_i, params['control_ratedPower'], params['drivetrainType'])
return abs(P_i - params['control_ratedPower'])
bnds = [Uhub[i-1], Uhub[i+1]]
U_rated = minimize_scalar(lambda x: get_Uhub_rated_noII12(pitch[i], x), bounds=bnds, method='bounded', options=options)['x']
if not np.isnan(U_rated):
Uhub[i] = U_rated
else:
print('Regulation trajectory is struggling to find a solution for rated wind speed. Check rotor_aeropower.py')
Omega[i] = min([Uhub[i] * params['control_tsr'] / params['Rtip'], Omega_max])
pitch0 = pitch[i]
P_aero[i], T[i], Q[i], M[i], Cp_aero[i], _, _, _ = self.ccblade.evaluate([Uhub[i]], [Omega[i] * 30. / np.pi], [pitch0], coefficients=True)
P[i], eff = CSMDrivetrain(P_aero[i], params['control_ratedPower'], params['drivetrainType'])
Cp[i] = Cp_aero[i]*eff
for j in range(i + 1,len(Uhub)):
Omega[j] = Omega[i]
if self.regulation_reg_III == True:
pitch0 = pitch[j-1]
bnds = [pitch0, pitch0 + 15.]
pitch_regionIII = minimize_scalar(lambda x: constantPregionIII(x, Uhub[j], Omega[j]), bounds=bnds, method='bounded', options=options)['x']
pitch[j] = pitch_regionIII
P_aero[j], T[j], Q[j], M[j], Cp_aero[j], _, _, _ = self.ccblade.evaluate([Uhub[j]], [Omega[j] * 30. / np.pi], [pitch[j]], coefficients=True)
P[j], eff = CSMDrivetrain(P_aero[j], params['control_ratedPower'], params['drivetrainType'])
Cp[j] = Cp_aero[j]*eff
if abs(P[j] - params['control_ratedPower']) > 1e+4:
print('The pitch in region III is not being determined correctly at wind speed ' + str(Uhub[j]) + ' m/s')
P[j] = params['control_ratedPower']
T[j] = T[j-1]
Q[j] = P[j] / Omega[j]
M[j] = M[j-1]
pitch[j] = pitch[j-1]
Cp[j] = P[j] / (0.5 * params['rho'] * np.pi * params['Rtip']**2 * Uhub[i]**3)
P[j] = params['control_ratedPower']
else:
P[j] = params['control_ratedPower']
T[j] = 0
Q[j] = Q[i]
M[j] = 0
pitch[j] = 0
Cp[j] = P[j] / (0.5 * params['rho'] * np.pi * params['Rtip']**2 * Uhub[i]**3)
unknowns['T'] = T
unknowns['Q'] = Q
unknowns['Omega'] = Omega * 30. / np.pi
unknowns['P'] = P
unknowns['Cp'] = Cp
unknowns['V'] = Uhub
unknowns['M'] = M
unknowns['pitch'] = pitch
self.ccblade.induction_inflow = True
a_regII, ap_regII, alpha_regII,_, _ = self.ccblade.distributedAeroLoads(Uhub[0], Omega[0] * 30. / np.pi, pitch[0], 0.0)
# Fit spline to powercurve for higher grid density
spline = PchipInterpolator(Uhub, P)
V_spline = np.linspace(params['control_Vin'],params['control_Vout'], num=self.n_pc_spline)
P_spline = spline(V_spline)
# outputs
idx_rated = list(Uhub).index(U_rated)
unknowns['rated_V'] = U_rated
unknowns['rated_Omega'] = Omega[idx_rated] * 30. / np.pi
unknowns['rated_pitch'] = pitch[idx_rated]
unknowns['rated_T'] = T[idx_rated]
unknowns['rated_Q'] = Q[idx_rated]
unknowns['V_spline'] = V_spline
unknowns['P_spline'] = P_spline
unknowns['ax_induct_cutin'] = a_regII
unknowns['tang_induct_cutin'] = ap_regII
unknowns['aoa_cutin'] = alpha_regII