def main():
# define the problem
configs, analyses = full_setup()
configs.finalize()
analyses.finalize()
# weight analysis
weights = analyses.configs.base.weights
breakdown = weights.evaluate()
# mission analysis
mission = analyses.missions
results = mission.evaluate()
# run payload diagram
config = configs.base
cruise_segment_tag = "cruise"
reserves = 1750.
payload_range_results = payload_range(config,mission,cruise_segment_tag,reserves)
# plot the results
plot_mission(results)
return
def main():
# define the problem
configs, analyses = full_setup()
configs.finalize()
analyses.finalize()
# weight analysis
weights = analyses.configs.base.weights
breakdown = weights.evaluate()
# mission analysis
mission = analyses.missions
results = mission.evaluate()
# run payload diagram
config = configs.base
cruise_segment_tag = "cruise"
reserves = 1750.
payload_range_results = payload_range(config, mission, cruise_segment_tag,
reserves)
# plot the results
plot_mission(results)
return
def main():
# define the problem
vehicle, mission = full_setup()
# run payload diagram
cruise_segment_tag = "Cruise"
payload_range_results = payload_range(vehicle, mission, cruise_segment_tag)
check_results(payload_range_results)
return
def main():
# define the problem
vehicle, mission = full_setup()
# run payload diagram
cruise_segment_tag = "Cruise"
payload_range_results = payload_range(vehicle,mission,cruise_segment_tag)
check_results(payload_range_results)
return
def main():
# build the vehicle
vehicle = define_vehicle()
# define the mission
mission = define_mission(vehicle)
# evaluate the mission
results = evaluate_mission(vehicle,mission)
# plot results
post_process(vehicle,mission,results)
# compute payload range diagram
cruise_segment_tag = 'CRUISE'
payload_range(vehicle,mission,cruise_segment_tag)
plt.show(True)
return
def main():
# build the vehicle
vehicle = define_vehicle()
# define the mission
mission = define_mission(vehicle)
# evaluate the mission
results = evaluate_mission(vehicle,mission)
# plot results
post_process(vehicle,mission,results)
# compute payload range diagram
cruise_segment_tag = 'CRUISE'
payload_range(vehicle,mission,cruise_segment_tag)
plt.show(True)
return
def main():
# define the problem
configs, analyses = full_setup()
configs.finalize()
analyses.finalize()
vehicle = configs.base
mission = analyses.missions
# run payload diagram
cruise_segment_tag = "Cruise"
reserves = 1750.
payload_range_results = payload_range(vehicle,mission,cruise_segment_tag,reserves)
check_results(payload_range_results)
return
def main():
# define the problem
configs, analyses = full_setup()
configs.finalize()
analyses.finalize()
vehicle = configs.base
mission = analyses.missions
# run payload diagram
cruise_segment_tag = "cruise"
reserves = 1750.
payload_range_results = payload_range(vehicle,mission,cruise_segment_tag,reserves)
check_results(payload_range_results)
return
def main():
# ---------------------------------------------------------------------------------------
# INITIALIZING AIRCRAFT
configs, analyses, vehicle = full_setup()
print 'full setup OK'
simple_sizing(configs)
configs.finalize()
analyses.finalize()
# ---------------------------------------------------------------------------------------
# WEIGHT ANALYSIS
weights = analyses.configs.base.weights
weights.evaluate()
print 'WEIGHTS OK'
# ---------------------------------------------------------------------------------------
# MISSION ANALYSIS
mission = analyses.missions.base
results = mission.evaluate()
print 'MISSION OK'
configs.cruise.conditions = Data()
configs.cruise.conditions = results.segments.cruise.conditions
# print weight breakdown
print_weight_breakdown(configs.base, filename='ATR72_weight_breakdown.dat')
# print parasite drag data into file - define reference condition for parasite drag
ref_condition = Data()
ref_condition.mach_number = 0.3
ref_condition.reynolds_number = 12e6
print_parasite_drag(ref_condition, configs.cruise, analyses,
'ATR72_parasite_drag.dat')
# print compressibility drag data into file
print_compress_drag(configs.cruise,
analyses,
filename='ATR72_compress_drag.dat')
# print mission breakdown
print_mission_breakdown(results, filename='ATR72_mission_breakdown.dat')
state = Data()
state.conditions = SUAVE.Analyses.Mission.Segments.Conditions.Aerodynamics(
)
state.numerics = SUAVE.Analyses.Mission.Segments.Conditions.Numerics()
# ---------------------------------------------------------------------------------------
# PLOT RESULTS
plot_mission(results, 0)
# ---------------------------------------------------------------------------------------
# PAYLOAD RANGE DIAGRAM
config = configs.base
cruise_segment_tag = "cruise"
reserves = [775., 1120., 1040.]
weights.mass_properties.operating_empty = weights.mass_properties.operating_empty - 239.
payloadrange = payload_range(config, mission, cruise_segment_tag, reserves)
# ---------------------------------------------------------------------------------------
# TAKE OFF FIELD LENGTH
# ---- Inputs
analyses.base = analyses.configs.base
airport = mission.airport
clb_grad = 1
altitude = [0., 6000.]
delta_isa = 0. # +23
# 1000, 1100, 1200, 1300, 1400, -, 1500, 1600
# 1200, 1300, 1400, 1500, 1600, 1700
weights_tofl = [[20062, 21044, 21954, 22801, 23000, 23452, 23754, 21200],
[18489, 19342, 20154, 20928, 21671, 22105, 21530]]
# ---- Inputs: FLAP AND SLAT DEFLECTION
flaps = [15., 0.] # Deflections inboard local
slats = [0., 0.]
# ---- Inputs: FACTOR CLMAX
configs.takeoff.max_lift_coefficient_factor = 1.189
configs.takeoff.V2_VS_ratio = 1.143
# ---- Open output file
fid = open('TOFL.txt', 'w') # Open output file
# ---- Run
for j, h in enumerate(altitude):
airport.altitude = h * Units.ft
airport.delta_isa = delta_isa
fid.write('Altitude: %4.0f ft \n' % (h))
fid.write(
'TOFL CLIMB GRADIENT THRUST L/D L/Dv2 CDasym CDwindm CL CD CG_CORRECT\n'
)
tofl = np.zeros(len(weights_tofl[j]))
secsegclbgrad = np.zeros(len(weights_tofl[j]))
thrust = np.zeros(len(weights_tofl[j]))
l_over_d = np.zeros(len(weights_tofl[j]))
l_over_d_v2 = np.zeros(len(weights_tofl[j]))
asymmetry_drag_coefficient = np.zeros(len(weights_tofl[j]))
windmilling_drag_coefficient = np.zeros(len(weights_tofl[j]))
clv2 = np.zeros(len(weights_tofl[j]))
cdv2 = np.zeros(len(weights_tofl[j]))
secsegclbgrad_corrected = np.zeros(len(weights_tofl[j]))
CLmax_ind = 0
# configs.takeoff.maximum_lift_coefficient = maximum_lift_coefficient[CLmax_ind]
configs.takeoff.wings[
'main_wing'].flaps.angle = flaps[CLmax_ind] * Units.deg
configs.takeoff.wings[
'main_wing'].slats.angle = slats[CLmax_ind] * Units.deg
for i, TOW in enumerate(weights_tofl[j]):
configs.takeoff.mass_properties.takeoff = TOW * Units.kg
tofl[i], secsegclbgrad[i], thrust[i], l_over_d[i], l_over_d_v2[i], asymmetry_drag_coefficient[i], \
windmilling_drag_coefficient[i], clv2[i], cdv2[i], secsegclbgrad_corrected[
i] = estimate_take_off_field_length(configs.takeoff,
analyses, airport,
clb_grad)
if secsegclbgrad_corrected[i] < 0.024:
# CLmax_ind = CLmax_ind + 1
# if CLmax_ind > 1:
# CLmax_ind = 1
print CLmax_ind, CLmax_ind, CLmax_ind, CLmax_ind, CLmax_ind, CLmax_ind
configs.takeoff.wings[
'main_wing'].flaps.angle = flaps[CLmax_ind] * Units.deg
configs.takeoff.wings[
'main_wing'].slats.angle = slats[CLmax_ind] * Units.deg
tofl[i], secsegclbgrad[i], thrust[i], l_over_d[i], l_over_d_v2[i], asymmetry_drag_coefficient[i], \
windmilling_drag_coefficient[i], clv2[i], cdv2[i], secsegclbgrad_corrected[
i] = estimate_take_off_field_length(configs.takeoff,
analyses, airport,
clb_grad)
fid.write(
'%4.2f %4.4f %4.4f %4.4f %4.4f %4.4f %4.4f %4.4f %4.4f %4.4f \n'
% (tofl[i], secsegclbgrad[i], thrust[i], l_over_d[i],
l_over_d_v2[i], asymmetry_drag_coefficient[i],
windmilling_drag_coefficient[i], clv2[i], cdv2[i],
secsegclbgrad_corrected[i]))
fid.write('\n')
fid.close()
# ---------------------------------------------------------------------------------------
# LANDING FIELD LENGTH
# ---- Inputs
airport.delta_isa = 0
altitude = [0, 6000]
weights_lfl = [[17000, 18000, 19000, 20000, 21000, 22000, 22350],
[17000, 18000, 19000, 20000, 21000, 22000, 22350]]
flaps = [30., 15.]
slats = [0., 0.]
configs.landing.landing_constants = Data()
configs.landing.landing_constants[0] = 250. * 0.25
configs.landing.landing_constants[1] = 0.
configs.landing.landing_constants[2] = 2.485 / 9.81
configs.landing.wings['main_wing'].flaps.angle = flaps[0] * Units.deg
configs.landing.wings['main_wing'].slats.angle = slats[0] * Units.deg
configs.landing.max_lift_coefficient_factor = 1.31
fid = open('LFL.txt', 'w') # Open output file
for j, h in enumerate(altitude):
airport.altitude = h * Units.ft
lfl = np.zeros(len(weights_lfl[j]))
fid.write('Altitude: %4.0f ft \n' % (h))
for i, TOW in enumerate(weights_lfl[j]):
configs.landing.mass_properties.landing = TOW * Units.kg
lfl[i] = estimate_landing_field_length(configs.landing, analyses,
airport)
fid.write('%4.2f \n' % (lfl[i]))
fid.write('\n')
fid.close()
return