def det_nose_lg_weight(W_l, L_n, N_nw, N_gear = 3, kneeling_nose_lg = False):
"""
Inputs:
W_l = landing design weight in lb
N_l = ultimate landing load factor = N_gear * 1.5
L_n = length of the nose landing gear in inches
N_mw = number of nose wheels
conditional inputs:
K_np = factor for landing gear (1.15 for kneeling gear, 1 otherwise)
outputs:
nose landing gear weight in lb
"""
W_l = cf.kg_to_pounds(W_l)
N_l = N_gear * 1.5
L_n = cf.meter_to_inch(L_n)
K_np = 1
if kneeling_nose_lg:
K_np = 1.15
nose_lg_weight = 0.032 * K_np * W_l**0.646 * N_l**0.2 * L_n**0.5 * N_nw**0.45
nose_lg_weight = cf.pounds_to_kg(nose_lg_weight)
return nose_lg_weight
def det_main_lg_weight(W_l, L_m, N_mw, N_mss, V_stall, N_gear = 3, kneeling_main_lg = False):
"""
Inputs:
W_l = landing design weight in kg
L_m = length of the main landing gear in m
N_mw = number of main wheels
N_mss = number of main gear shock struts
V_stall = stall speed in m/s
conditional inputs:
K_mp = factor for landing gear (1.126 for kneeling landing gear, 1 otherwise)
outputs:
main landing gear weight in lb
"""
W_l = cf.kg_to_pounds(W_l)
N_l = N_gear * 1.5
L_m = cf.meter_to_inch(L_m)
V_stall = V_stall * 1.94384449
K_mp = 1
if kneeling_main_lg:
K_mp = 1.126
main_lg_weight = 0.0106 * K_mp * W_l**0.888 * N_l**0.25 * L_m**0.4 * N_mw**0.321 * N_mss**-0.5 * V_stall**0.1
main_lg_weight = cf.pounds_to_kg(main_lg_weight)
return main_lg_weight
def det_handling_gear_weight(W_dg):
"""
inputs:
W_dg = design gross weight in kg
outputs:
the total weight of the furnishings in kg
"""
W_dg = cf.kg_to_pounds(W_dg)
handling_gear_weight = 3.0 * 10**-4 * W_dg
handling_gear_weight = cf.pounds_to_kg(handling_gear_weight)
return handling_gear_weight
def det_anti_ice_weight(W_dg):
"""
inputs:
W_dg = design gross weight in kg
outputs:
the total weight of the anti icing system in kg
"""
W_dg = cf.kg_to_pounds(W_dg)
anti_ice_weight = 0.002 * W_dg
anti_ice_weight = cf.pounds_to_kg(anti_ice_weight)
return anti_ice_weight
def det_nacelle_group_weight(N_lt, N_w, W_dg, N_en, S_n, W_engine, pylon_mounted = True, W_ec = 0, propeller = True, thrust_reverser = False):
"""
Inputs:
N_lt = nacelle length in m
N_w = nacelle width in m
W_dg = design gross weight in kg
N_en = number of engines
S_n = nacelle wetted area in m^2
W_engine = weight of each engine in kg
conditional inputs:
K_ng = factor for nacelles (1.017 for pylon mounted nacelle, 1.0 otherwise)
W_ec = weight of the engine and contents in lb
(approximated by 2.331 W_engine^0.901 * K_p * K_tr)
K_p = 1.4 for propeller, 1.0 otherwise
K_tr = 1.18 for jet with thrust reverser, 1.0 otherwise
outputs:
nacelle group weight in kg
"""
N_lt = cf.meter_to_feet(N_lt)
N_w = cf.meter_to_feet(N_w)
N_z = 1.5* lim_load_factor(W_dg)
S_n = cf.metersquared_to_feetsquared(S_n)
W_engine = cf.kg_to_pounds(W_engine)
K_ng = 1
if pylon_mounted:
K_ng = 1.017
if W_ec == 0 and W_engine == 0:
raise NameError("no propulsion??")
K_p = 1
K_tr = 1
if propeller:
K_p = 1.4
if thrust_reverser:
K_tr = 1.18
if W_ec == 0:
W_ec = 2.331 * W_engine**0.901 * K_p * K_tr
nacelle_group_weight = 0.6724 * K_ng * N_lt**0.10 * N_w**0.294 * N_z**0.119 * W_ec**0.611 * N_en**0.984 * S_n**0.224
nacelle_group_weight = cf.pounds_to_kg(nacelle_group_weight)
return nacelle_group_weight
def det_furnishings_weight(N_c, W_c, S_f):
"""
inputs:
N_c = number of pilots
W_c = maximum cargo weight in kg
S_f = fuselage wetted area in m^2
outputs:
the total weight of the furnishings in kg
"""
W_c = cf.kg_to_pounds(W_c)
S_f = cf.metersquared_to_feetsquared(S_f)
furnishings_weight = 0.0577 * N_c**0.1 * W_c**0.393 * S_f**0.75
furnishings_weight = cf.pounds_to_kg(furnishings_weight)
return furnishings_weight
def lim_load_factor(MTOM):
"""
Inputs:
MTOM = Maximum Take-Off Mass in kg
Output:
Limit Load Factor (Dimensionless)
"""
MTOM = cf.kg_to_pounds(MTOM)
if MTOM <= 4100:
n_lim = 3.8
if 4100 < MTOM <= 50000:
n_lim = 2.1 + (24000/(MTOM + 10000))
if MTOM > 50000:
n_lim = 2.5
return n_lim
def det_hor_tail_weight(F_w, B_h, W_dg, S_ht, L_t, quarter_chord_sweep_ht, AR_ht, S_e, all_moving_unit = False, K_y = 0.3):
"""
Inputs:
F_w = fuselage width at horizontal tail intersection in m
B_h = horizontal tail span in m
W_dg = design gross weigth in kg
S_ht = horizontal tail area in m^2
L_t = tail length (wing quarter MAC to tail quarter MAC) in m
quarter_chord_sweep_ht = sweep angle of the horizontal tail at the quarter chord line in degrees
AR_ht = aspect ratio of the horizontal tail
S_e = elevator area in m^2
conditional inputs:
all_moving_unit = is the horizontal tail moving as a unit (True) or only part of it (False)
all_moving_unit = True -> 1.143
all_moving_unit = False -> 1
K_y = Aircraft pitching radius of gyration in ft (usually 0.3 * L_t)
output:
Horizontal tail weight in kg
"""
F_w = cf.meter_to_feet(F_w)
B_h = cf.meter_to_feet(B_h)
N_z = 1.5* lim_load_factor(W_dg)
W_dg = cf.kg_to_pounds(W_dg)
S_ht = cf.metersquared_to_feetsquared(S_ht)
L_t = cf.meter_to_feet(L_t)
quarter_chord_sweep_ht = quarter_chord_sweep_ht * (np.pi/180)
S_e = cf.metersquared_to_feetsquared(S_e)
K_uht = 1
if all_moving_unit:
K_uht = 1.143
if K_y == 0.3:
K_y = 0.3 * L_t
hor_tail_weight = 0.0379 * K_uht * (1 + F_w/B_h)**-0.25 * W_dg**0.639 * N_z**0.10 * S_ht**0.75 * L_t**-1 * K_y**0.704 * np.cos(quarter_chord_sweep_ht)**-1 * AR_ht**0.166 * (1 + S_e/S_ht)**0.1
hor_tail_weight = cf.pounds_to_kg(hor_tail_weight)
return hor_tail_weight
def det_ver_tail_weight(W_dg, L_t, S_vt, quarter_chord_sweep_vt, AR_vt, t_c_root_vt, K_z = 1, TTail = True):
"""
Inputs:
W_dg = design gross weigth in kg
L_t = tail length (wing quarter MAC to tail quarter MAC) in m
S_vt = vertical tail area in m^2
quarter_chord_sweep_vt = sweep angle of the vertical tail at the quarter chord line in degrees
AR_vt = aspect ratio of the vertical tail
t_c_root_vt = thickness to chord ratio at the root of the vertical tail
conditional inputs:
K_z = Aircraft yawing radius of gyration in ft (usually L_t)
outputs:
vertical tail weight in kg
"""
if TTail:
H_ht = 1
H_vt = 1
else:
H_ht = 0
H_vt = 1
N_z = 1.5* lim_load_factor(W_dg)
W_dg = cf.kg_to_pounds(W_dg)
L_t = cf.meter_to_feet(L_t)
S_vt = cf.metersquared_to_feetsquared(S_vt)
quarter_chord_sweep_vt = quarter_chord_sweep_vt * (np.pi/180)
if K_z == 1:
K_z = L_t
ver_tail_weight = 0.0026 * (1 + H_ht/H_vt)**0.225 * W_dg**0.556 * N_z**0.536 * L_t**-0.5 * S_vt**0.5 * K_z**0.875 * np.cos(quarter_chord_sweep_vt)**-1 * AR_vt**0.35 * t_c_root_vt**-0.5
ver_tail_weight = cf.pounds_to_kg(ver_tail_weight)
return ver_tail_weight
def det_fuselage_weight(W_dg, L, S_f, taper, B_w, quarter_chord_sweep, L_D_ratio, cargo_doors = 1, fuselage_mounted_lg = True):
"""
Inputs:
W_dg = design gross weigth in kg
L = fuselage structural length in m (no radome/nosecone and tail cap)
S_f = fuselage wetted area in m^2
taper = taper ratio
B_w = wing span in m
quarter_chord_sweep = sweep angle at the quarter chord line in degrees
L_D_ratio = the lift over drag ratio of the aircraft
conditional inputs:
K_door = factor for the cargo doors (1 for no cargo door, 1.06 for one side cargo door, 1.12 for 2 side cargo doors)
K_lg = factor for landing gear (1.12 for fuselage mounted landing gear, 1 otherwise)
outputs:
fuselage weight in kg
"""
N_z = 1.5* lim_load_factor(W_dg)
W_dg = cf.kg_to_pounds(W_dg)
L = cf.meter_to_feet(L)
S_f = cf.metersquared_to_feetsquared(S_f)
B_w = cf.meter_to_feet(B_w)
quarter_chord_sweep = quarter_chord_sweep * (np.pi/180)
K_door = 1 + cargo_doors * 0.06
K_lg = 1
if fuselage_mounted_lg:
K_lg = 1.12
K_ws = 0.75 * (1 + 2 * taper)/(1 + taper) * (B_w * np.tan(quarter_chord_sweep)/L)
fuselage_weight = 0.3280 * K_door * K_lg * (W_dg * N_z)**0.5 * L**0.25 * S_f**0.302 * (1 + K_ws)**0.04 * L_D_ratio**0.10
fuselage_weight = cf.pounds_to_kg(fuselage_weight)
return fuselage_weight