def m0_fn(Ma):
if Ma <= 0.9:
return (2 * math.pi) / numpy.sqrt(1 - Ma**2)
else:
return (2 * math.pi) / numpy.sqrt(1 - 0.9**2)
def Cd_fn(Cl):
return 0.0095 + 0.0040 * (Cl - 0.2)**2
nn = 101
h = 0 # m
v_inf = 70 # m/s
is_SI = True
atm = AtmData(v_inf, h, is_SI)
k = 1.4
R = 287
atm.expand(k, R)
numB = 3
radius = 2 / 2
RPM = 2400
Cl = 0.4
alp0 = numpy.radians(-2)
prop = Propeller(radius, numB, RPM, CP=0, CT=0, CQ=0, Cl=Cl, alp0=alp0)
T_req = 13000 / 8 # N (TOGW/(L/D))
r_vec, c_vec, beta_vec, P_design, T_design, Q_design, eta_P, theta_vec = prop_design(
atm, prop, T_req, m0_fn, Cd_fn)
in_line = [0] * len(r_vec)
plot_propeller_3D(r_vec, c_vec, beta_vec, prop, in_line)
# Propeller
radius = 41 * 0.0254 # m
numB = 3
Cl = 0.4
RPM = 2400
alp0 = numpy.radians(-2)
dens = atm_check.dens
CT = 0.0509
T_req = CT * dens * (RPM / 60)**2 * (radius * 2)**4
v_seq = numpy.arange(v_climb - 30 * 0.3048, v_climb + 126 * 0.3048, 2)
prop_check = Propeller(radius,
numB,
RPM,
eta_P=0,
CP=0,
CT=0,
CQ=0,
Cl=0.4)
r, prop_check.chord, prop_check.bet, P_design, T_design, Q_design, eta_P, prop_check.theta = prop_design(
atm_check, prop_check, T_req, m0_fn, Cd_fn)
ll = numpy.size(v_seq)
J_var = numpy.zeros((ll, ))
P_design_var = numpy.zeros((ll, ))
T_design_var = numpy.zeros((ll, ))
eta_P_var = numpy.zeros((ll, ))
deta_P = numpy.zeros((ll, ))
dT = numpy.zeros((ll, ))
delta_bet = numpy.zeros((ll, ))
T_req = 13000 / 8 / LD #N (TOGW/(L/D))
Cl = 0.4
alp0 = numpy.radians(-2)
for v_inf in range(22, 72, 10):
P_design = [0] * nn
T_design = [0] * nn
Q_design = [0] * nn
eta_P = [0] * nn
for ii in range(0, nn):
atm = AtmData(v_inf, h, is_SI)
atm.expand(1.4, 287)
prop = Propeller(radius,
numB,
RPM[ii],
eta_P,
CP=0,
CT=0,
CQ=0,
Cl=0.4)
[_, _, _, P_design[ii], T_design[ii], Q_design[ii], eta_P[ii],
_] = prop_design(atm, prop, T_req, m0_fn, Cd_fn)
label = "$V_\infty$ = " + str(v_inf) + " (m/s)"
plt.figure(1)
plt.plot(RPM, Q_design, label=label)
plt.figure(2)
plt.plot(RPM, eta_P, label=label)
plt.figure(3)
plt.plot(RPM, P_design, label=label)
# plt.figure(7)
# plt.plot(RPM,T_design,label=label)
# Design condition
T_req = 13000 / (8) * 1.2 * 0.601 # N (TOGW/(L/D))
Cl = 0.4
alp0 = numpy.radians(-2)
v_hover = 2.54
v_cruise = 62
v_design = 30
atm = AtmData(v_design, 0, is_SI)
atm.expand(1.4, 287)
v_seq = numpy.arange(-10, 68, 2)
ll = numpy.size(v_seq)
Design_RPM = 3000
prop = Propeller(radius, numB, Design_RPM, eta_P=0, CP=0, CT=0, CQ=0, Cl=Cl)
r, prop.chord, prop.bet, P_design, T_design, Q_design, eta_P, prop.theta = prop_design(atm, prop, T_req, m0_fn, Cd_fn)
num_len = len(prop.bet)
beta_75_index = int(num_len * 0.75)
x = numpy.linspace(0, 1, num_len)
x_need = x[x >= 0.15]
c_need = prop.chord[x >= 0.15]
AF = 10 ** 5 / 16 * numpy.trapz(c_need / diameter * x_need ** 3, x_need) # activity factor
CL_design = 4 * numpy.trapz(Cl * x_need ** 3, x_need)
print('====================== Propeller Design Parameters ======================')
print("Beta angle at 75% is {:.5f} degree".format(numpy.rad2deg(prop.bet[beta_75_index])))
print("Activity factor is {:.2f}, which corresponds to {:.2f} in openVSP (Source checked)".format(AF, AF * 2))
print("CL_design is {:.2f}".format(CL_design))
vel = 124 * 0.514444 # m/s
c_p = 0.51 / (1980000 * 0.3048) # lb/(hp*hr) to 1/ft to 1/m
eta_P = 0.85
CD_0 = 0.0340
CL = 0.305
CD = 0.0401
# Just to make class work
area = 200 * 0.092903 # ft^2 to m^2
span = 1 / (numpy.pi * 0.0657)
e = 1
atm = AtmData(vel, 'temp', 'pres', 'dens', 'visc', 'k', 'R', 'is_SI')
wing = Wing(area, span, e, 'alpha', 'chord', 'c_bar', CL, 'CL_max', CD, CD_0,
'airfoil')
prop = Propeller('radius', 'RPM', eta_P, 'CP', 'CT', 'CQ', 'Cl = 0.4',
'chord = 1', 'numB = 3', 'alp0 = 0', 'alpha = None',
'beta = None', 'theta = None', 'phi = None')
R, E = propeller_in_cruise(W_initial, W_final, c_p, atm, prop, wing,
const_logic)
R *= 0.000539957 # m to nmi
E /= 3600 # s to hr
#print(R) # should be ~500 nmi
#print(E) # should be ~4.03 hr
# Checked!
# Test B, constant Cl and altitude, numbers from lecture 01.27.2021
const_logic = [True, False, True]
W_initial = 2500 * 4.4482216153 # lbf to N
temp = 288 pres = 101250 dens = 1.225 visc = 3E-6 k = 1.4 R = 287 is_SI = True atm_check = AtmData(v_inf, temp, pres, dens, visc, k, R, is_SI) # prop. data radius = 41 * 0.0254 RPM = 1854.4805 Cl = 0.4 numB = 3 alp0 = numpy.radians(-2) prop_check = Propeller(radius, RPM, Cl, "chord", numB, alp0) # wing data area = 16 span = 5 e = 0.9 CL_max = 2.1 wing_check = Wing(area, span, e, "alpha", "chord", 1, 1, CL_max, "CD", "CD_0", "airfoil") # function call m = 6000 mot_distr = [1,1] print(numpy.size(mot_distr)) dist_to = 2000 power_vec = STOL_power_req(m, dist_to, mot_distr, atm_check, wing_check, prop_check)
# 3 Blades
numB = 3
# Design condition
T_req = 13000 / 8 * 1.2 * 0.601 # N (TOGW/(L/D))
Cl = 0.4
alp0 = np.radians(-2)
v_design = 30
atm = AtmData(v_design, 0, is_SI)
atm.expand(1.4, 287)
Design_RPM = 3000
prop = Propeller(radius, numB, Design_RPM, eta_P=0, CP=0, CT=0, CQ=0, Cl=Cl)
output = prop_design(atm, prop, T_req, m0_fn, Cd_fn, num=401)
r, prop.chord, prop.bet, P_design, T_design, Q_design, eta_P, prop.theta = output
# Point reduction and interpolation
# Adjustment
num_len = len(prop.bet)
beta_75_index = int(num_len * 0.75)
x = np.linspace(0, 1, num_len)
x_need = x[x >= 0.15]
c_need = prop.chord[x >= 0.15]
AF = 10**5 / 16 * np.trapz(c_need / diameter * x_need**3,
x_need) # activity factor
CL_design = 4 * np.trapz(Cl * x_need**3, x_need)
print(
g = 9.81 # m/s^2
W_initial = 13000 # N
m_loss = 5 # this affects the output tremendously
W_final = 13000 - (m_loss * g) # N
W_diff = W_initial - W_final
# atm data
h = 6000 * 0.3048 # m
vel = 62 # m/s
is_SI = True
atm = AtmData(vel, h, is_SI)
atm.expand(1.4, 287)
# prop data
eta_P = 0.82
prop = Propeller('radius', 'numB', 'RPM', eta_P=eta_P)
const_logic = [False, True, True] # [CL, v_inf, h]
area = 16 # m^2
span = (4.7046 + 1.62) * 2
CL = (W_initial + W_final) / (atm.dens * atm.vel**2 * area)
# CL = 0.4128
CD = 0.018 # from drag polar with no flap, 62 m/s, 6000 ft
CD_0 = 0.012
wing = Wing(area, span, e=0.7, CL=CL, CD=CD, CD_0=CD_0)
# c_p
P_fuelcell_cont = 100 * 1000 # W
radius = 1.78 / 2
RPM = numpy.linspace(500, 2950, nn)
LD = 15
T_req = 13000 / 8 # N (TOGW/(L/D))
Cl = 0.4
alp0 = numpy.radians(-2)
for v_inf in range(20, 70, 10):
P_design = [0] * nn
T_design = [0] * nn
Q_design = [0] * nn
eta_P = [0] * nn
for ii in range(0, nn):
atm = AtmData(v_inf, h, is_SI)
atm.expand(1.4, 287)
prop = Propeller(radius, numB, RPM[ii], eta_P, CP=0, CT=0, CQ=0, Cl=0.4)
[_, _, _, P_design[ii], T_design[ii], Q_design[ii], eta_P[ii], _] = prop_design(atm, prop, T_req, m0_fn,
Cd_fn)
label = "$V_\infty$ = " + str(v_inf) + " (m/s)"
plt.figure(4)
plt.plot(RPM, Q_design, label=label)
plt.figure(5)
plt.plot(RPM, eta_P, label=label)
plt.figure(6)
plt.plot(RPM, P_design, label=label)
# plt.figure(7)
# plt.plot(RPM,T_design,label=label)
plt.figure(4)
plt.legend()
plt.xlabel("RPM")
import matplotlib.pyplot as plt
## ******The resulting graph does not make sense to XT
Weight = 13000 # N
f_body = 0.001 # m^2, almost neglected
gamma = 0 # deg
num_rotor = 8
# prop
radius = 1.78 / 2 # m
numB = 3
c_bar = 0.1 # m, assumed
RPM = 2500 # assumed
prop_trial = Propeller(radius, numB, RPM, c_bar=c_bar)
# atm
h = 500 # m
is_SI = True
num = 101
vel_start = -10
vel_end = 20
vel_vec = numpy.linspace(vel_start, vel_end, num)
Cd_bar_start = 0.01
Cd_bar_end = 0.05
Cd_bar_num = 5
Cd_bar_vec = numpy.linspace(Cd_bar_start, Cd_bar_end, Cd_bar_num)