numB = 3
radius = 1.78 / 2
RPM = numpy.linspace(500, 3000, nn)
LD = 15
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)
if __name__ == '__main__':
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)
# Test A, constant Cl and airspeed, numbers from lecture 01.27.2021
const_logic = [True, True, False]
W_initial = 2500 * 4.4482216153 # lbf to N
W_final = 2215 * 4.4482216153 # lbf to N
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!
return (2 * numpy.pi) / numpy.sqrt(1 - Ma**2)
else:
return (2 * numpy.pi) / numpy.sqrt(1 - 0.9**2)
def Cd_fn(Cl):
return 0.0095 + 0.0040 * (Cl - 0.2)**2
# Atm
# P_eng = 177 * 745.7 # W
v_des = 154 * 0.514444 # m/s
v_climb = 87 * 0.514444 # m/s
h = 8000 * 0.3048 # m
is_SI = True
is_HP = True
atm_check = AtmData(v_des, h, is_SI)
atm_check.expand(1.4, 287)
# 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,
radius = diameter / 2 LD = 15 # Assumed is_SI = True # 3 Blades numB = 3 # 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
from STOL_initial_sizing import STOL_initial_sizing
from std_atm import std_atm
from Class130 import AtmData
v_stall_TO = 61 * 0.514444 # m/s, stall airspeed
sigma = 1
vel = v_stall_TO
h = 0 # m, sea level
is_SI = True
k = 1.4
R = 287
S_TO = 1000 # m
S_L = 1000 # m
sigma = 1
CL_max = 2.0
print(v_stall_TO)
temp, pres, dens, c_sound, visc, g = std_atm(h, is_SI)
atm_TO = AtmData(vel, temp, pres, dens, visc, k, R, is_SI)
atm_L = AtmData(vel, temp, pres, dens, visc, k, R, is_SI)
wing_load_TO, wing_load_L, weight_power_TO = STOL_initial_sizing(
v_stall_TO, sigma, CL_max, S_TO, S_L, atm_TO, atm_L)
print(wing_load_TO)
print(wing_load_L)
print(weight_power_TO)
def m0_fn(Ma):
if numpy.any(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
v_inf = 156 * 0.514444
h = 2438.4
k = 1.4
R = 287
is_SI = True
atm_check = AtmData(v_inf, h, is_SI)
atm_check.expand(k, R)
dens = atm_check.dens
radius = 41 * 0.0254
RPM = 2400
CT = 0.0509
T_req = CT*dens*(RPM/60)**2*(radius*2)**4
# T_req = 425.6896 * 4.44822
Cl = 0.4
numB = 3
alp0 = numpy.radians(-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)
# Fixed pitch check
# 02.14.2021: Created by XT # 02.15.2021: Reviewed by TVG. # Changed line 32 to reflect changes in Wing class (added alpha and CD as dummy values). # Not verified relative to literature values # atm data v_inf = 87 * 0.514444 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
"""
Calculates saturation pressure of water vapor
:param T: float, (deg C), temperature of interest
:return: (Pa), saturation pressure
"""
T_k = T + 273.15 # convert from C to K
# from https://www.engineeringtoolbox.com/water-vapor-saturation-pressure-air-d_689.html
exp_power = (77.3450 + 0.0057 * T_k - 7235 / T_k) / (T_k**8.2) # empirical
return np.exp(exp_power)
if __name__ == '__main__':
# needed inputs
vel = 121 * 0.514444 # airspeed, 121 knots to m/s
h = 6000 * 0.3048 # altitude, 6000 ft to m
atm = AtmData(vel, h, is_SI=True)
x_loc = 0.0005 # m, location of interest (just 0.5 mm past leading edge)
C_liquid = 4184 # J/(kg K), specific heat capacity of liquid water
C_ice = 2093 # J/(kg K), specific heat capacity of ice
L_f = 3.34e5 # J/kg, latent heat of water fusion
T_target = 6 # deg C, target heating temperature
T_inf = atm.temp - 273.15 # deg C, environment temperature (assume under standard atmosphere)
T_skin = T_target # deg C, temperature at surface
k_air = 0.0227 # W/(mK), thermal conductivity of air at 255.3 K
R_h = 1 # Non-dimensional, relative humidity, actual / saturated vapor pressure
C_p_air = 1000 # J/(kg K), specific heat capacity under constant pressure for air (at 300K)
L_e = 2.257e6 # J/kg, latent heat of water evaporation
rho_LWC = 4.5e-4 # kg/m^3, liquid water content, mass of supercooled water per volume (assume stratocumulus)
thickness = 0.14166 * 1.26491 # m, maximum airfoil thickness, using mean chord length
span = 12.65 # m, airfoil span-wise extension for total wing
v_wall = 0 # m/s, (Seems to be) surface layer velocity (assumed 0 due to laminar flow)
if __name__ == '__main__':
from Class130 import AtmData, Propeller, Wing
#****** The range and endurance is too low for our mission
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
numB = 3
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)
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)
# P_aero vs. v_inf
plt.figure(figsize=[20, 15])
for i in range(Cd_bar_num):
Cd_bar = Cd_bar_vec[i]
P_aero_vec = [0] * num
for j in range(num):
atm_trial = AtmData(vel_vec[j], h, is_SI)
P_aero_vec[j], _ = rotor_hover_power(Weight, gamma, num_rotor, f_body, atm_trial, prop_trial, Cd_bar=Cd_bar)
plt.plot(vel_vec, P_aero_vec, label='Cd_bar = {:.2f}'.format(Cd_bar))
title = 'Total Climb Power versus. Climb Speed'
plt.title(title)
plt.xlabel(r'Climb Speed $V_\infty$ (m/s)')
plt.ylabel(r'Total Power $P_{aero}$ (W)')
plt.grid()
plt.legend(loc='upper right')
plt.show()
from Class130 import Airfoil, AtmData # Use XFOIL to test the stall angle of NLF 0414 F airfoil foilpath = '../Sizing/OpenVSP/v1.5_03_13_2021/Airfoil/NLF 0414F.dat' foilname = 'NLF0414' # Atm h = 6000 / 3.28084 # ft vel = 62 # m/s is_SI = True atm = AtmData(vel, h, is_SI) c = 1 Re = (atm.dens * atm.vel * c) / atm.visc alf_start = 0 alf_end = 80 num_alfs = 50 iter_num = 100 foil = Airfoil(foilname) foil.set_iter_num(iter_num) foil.set_num_alfs(num_alfs) foil.add_geom_file(foilpath) foil.get_polar(Re, alf_start, alf_end) foil.lift_curve() foil.drag_polar()