def main():
configs, analyses = full_setup()
simple_sizing(configs, analyses)
configs.finalize()
analyses.finalize()
# mission analysis
mission = analyses.missions.base
results = mission.evaluate()
# load older results
#save_results(results)
old_results = load_results()
# plt the old results
plot_mission(results)
plot_mission(old_results, 'k-')
#plt.show(block=True)
# check the results
check_results(results, old_results)
# print weights breakdown
print_weight_breakdown(configs.cruise)
# ------------------------------------------------------------------
# Vehicle Definition Complete
# ------------------------------------------------------------------
# plot vehicle
plot_vehicle(configs.base, plot_control_points=True)
return
def main():
configs, analyses = full_setup()
simple_sizing(configs)
configs.finalize()
analyses.finalize()
# weight analysis
weights = analyses.configs.base.weights
breakdown = weights.evaluate()
# mission analysis
mission = analyses.missions.base
results = mission.evaluate()
# print weight breakdown
print_weight_breakdown(configs.base,filename = 'weight_breakdown.dat')
# print engine data into file
print_engine_data(configs.base,filename = 'B737_engine_data.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,'B737_parasite_drag.dat')
# print compressibility drag data into file
print_compress_drag(configs.cruise,analyses,filename = 'B737_compress_drag.dat')
# print mission breakdown
print_mission_breakdown(results,filename='B737_mission_breakdown.dat')
# load older results
#save_results(results)
old_results = load_results()
# plt the old results
plot_mission(results)
plot_mission(old_results,'k-')
# check the results
check_results(results,old_results)
## # print some results, for check agaist Aviation paper
## end_segment = results.segments[-1]
## time = end_segment.conditions.frames.inertial.time[-1,0] / Units.min
## distance = end_segment.conditions.frames.inertial.position_vector[-1,0] / Units.nautical_miles
## mass_end = end_segment.conditions.weights.total_mass[-1,0]
## mass_begin = results.segments[0].conditions.weights.total_mass[0,0]
## fuelburn = (mass_begin- mass_end) / Units.lbs
##
## print 'Time: {:5.0f} min , Distance: {:5.0f} nm ; FuelBurn: {:5.0f} lbs'.format(time,distance,fuelburn)
return
def main():
configs, analyses = full_setup()
simple_sizing(configs)
configs.finalize()
analyses.finalize()
# weight analysis
weights = analyses.configs.base.weights
breakdown = weights.evaluate()
# mission analysis
mission = analyses.missions.base
results = mission.evaluate()
# print weight breakdown
print_weight_breakdown(configs.base,filename = 'weight_breakdown.dat')
# print engine data into file
print_engine_data(configs.base,filename = 'B737_engine_data.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,'B737_parasite_drag.dat')
# print compressibility drag data into file
print_compress_drag(configs.cruise,analyses,filename = 'B737_compress_drag.dat')
# print mission breakdown
print_mission_breakdown(results,filename='B737_mission_breakdown.dat')
# load older results
#save_results(results)
old_results = load_results()
# plt the old results
plot_mission(results)
plot_mission(old_results,'k-')
# check the results
check_results(results,old_results)
## # print some results, for check agaist Aviation paper
## end_segment = results.segments[-1]
## time = end_segment.conditions.frames.inertial.time[-1,0] / Units.min
## distance = end_segment.conditions.frames.inertial.position_vector[-1,0] / Units.nautical_miles
## mass_end = end_segment.conditions.weights.total_mass[-1,0]
## mass_begin = results.segments[0].conditions.weights.total_mass[0,0]
## fuelburn = (mass_begin- mass_end) / Units.lbs
##
## print 'Time: {:5.0f} min , Distance: {:5.0f} nm ; FuelBurn: {:5.0f} lbs'.format(time,distance,fuelburn)
return
def main():
configs, analyses = full_setup()
simple_sizing(configs)
configs.finalize()
analyses.finalize()
# weight analysis
weights = analyses.configs.base.weights
breakdown = weights.evaluate()
# mission analysis
mission = analyses.missions.base
results = mission.evaluate()
# print weight breakdown
print_weight_breakdown(configs.base, filename='weight_breakdown.dat')
# print engine data into file
print_engine_data(configs.base, filename='B737_engine_data.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,
'B737_parasite_drag.dat')
# print compressibility drag data into file
print_compress_drag(configs.cruise,
analyses,
filename='B737_compress_drag.dat')
# print mission breakdown
print_mission_breakdown(results, filename='B737_mission_breakdown.dat')
# load older results
#save_results(results)
old_results = load_results()
# plt the old results
plot_mission(results)
plot_mission(old_results, 'k-')
# check the results
check_results(results, old_results)
return
def main():
configs, analyses = full_setup()
simple_sizing(configs)
configs.finalize()
analyses.finalize()
# weight analysis
weights = analyses.configs.base.weights
breakdown = weights.evaluate()
weights.vehicle.mass_properties.center_of_gravity = SUAVE.Methods.Center_of_Gravity.compute_aircraft_center_of_gravity(weights.vehicle, nose_load_fraction=.06)
# mission analysis
mission = analyses.missions.base
results = mission.evaluate()
CM = results.conditions.cruise.stability.static.CM[0][0]
cm0 = results.conditions.cruise.stability.static.cm0[0][0]
cm_alpha = results.conditions.cruise.stability.static.cm_alpha[0][0]
cn_beta = results.conditions.cruise.stability.static.cn_beta[0][0]
static_margin = results.conditions.cruise.stability.static.static_margin[0][0]
# print weight breakdown
print_weight_breakdown(configs.base,filename = 'E170_weight_breakdown.dat')
# print engine data into file
print_engine_data(configs.base,filename = 'E170_engine_data.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,'E170_parasite_drag.dat')
# print compressibility drag data into file
print_compress_drag(configs.cruise,analyses,filename = 'E170_compress_drag.dat')
# print mission breakdown
print_mission_breakdown(results,filename='E170_mission_breakdown.dat')
# plt the old results
plot_mission(results)
return
def main():
configs, analyses = full_setup()
simple_sizing(configs)
configs.finalize()
analyses.finalize()
# weight analysis
weights = analyses.configs.base.weights
breakdown = weights.evaluate()
# mission analysis
mission = analyses.missions.base
results = mission.evaluate()
# print weight breakdown
print_weight_breakdown(configs.base,filename = 'weight_breakdown.dat')
# print engine data into file
print_engine_data(configs.base,filename = 'B737_engine_data.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,'B737_parasite_drag.dat')
# print compressibility drag data into file
print_compress_drag(configs.cruise,analyses,filename = 'B737_compress_drag.dat')
# print mission breakdown
print_mission_breakdown(results,filename='B737_mission_breakdown.dat')
# load older results
#save_results(results)
old_results = load_results()
# plt the old results
plot_mission(results)
plot_mission(old_results,'k-')
# check the results
check_results(results,old_results)
return
def post_process(nexus):
# Unpack data
vehicle = nexus.vehicle_configurations.base
configs = nexus.vehicle_configurations
results = nexus.results
summary = nexus.summary
missions = nexus.missions
# pretty_print(results)
# Static stability calculations
CMA = -10.
for segment in results.base.segments.values():
max_CMA = np.max(segment.conditions.stability.static.cm_alpha[:, 0])
# lamda with a min/max?
if max_CMA > CMA:
CMA = max_CMA
# Dynamics stability calculations
dyn_stab = np.zeros(6)
# j=0
# for element in results.base.segments.conditions.stability.dynamic.values():
#
# max = np.max(element[:, 0])
# min = np.min(element[:, 0])
# avg = np.mean(element[:, 0])
#
# for derivative in
# stability data structure:
# stability
# static
# cm_alpha
# cn_beta
#
#
# dynamic
# cn_r
# cl_p
# 0
# cl_beta
# 0
# cm_q
# cm_alpha_dot
# cz_alpha
#
# summary.static_stability = CMA
#
results = evaluate_field_length(configs, nexus.analyses, missions.base, results)
#
summary.field_length_takeoff = results.field_length.takeoff
# print summary.field_length_takeoff
summary.field_length_landing = results.field_length.landing
# print summary.field_length_landing
#
# #
# pretty_print(nexus)
# throttle in design mission
max_throttle = 0
min_throttle = 0
for segment in results.base.segments.values():
max_segment_throttle = np.max(segment.conditions.propulsion.throttle[:, 0])
min_segment_throttle = np.min(segment.conditions.propulsion.throttle[:, 0])
if max_segment_throttle > max_throttle:
max_throttle = max_segment_throttle
if min_segment_throttle < min_throttle:
min_throttle = min_segment_throttle
summary.max_throttle = max_throttle
summary.min_throttle = min_throttle
# short_w_n, short_zeta, phugoid_w_n, phugoid_zeta = longitudinal(velocity, density, S_gross_w, mac, Cm_q, Cz_alpha, mass, Cm_alpha, Iy, Cm_alpha_dot, Cz_u, Cz_alpha_dot, Cz_q, Cw, Theta, Cx_u, Cx_alpha):
# Fuel margin and base fuel calculations
vehicle.mass_properties.operating_empty += 10e3 # FIXME hardcoded wing mass correction # area scaling?
operating_empty = vehicle.mass_properties.operating_empty
payload = vehicle.mass_properties.payload # TODO fuel margin makes little sense when ejecting aerosol
design_landing_weight = results.base.segments[-1].conditions.weights.total_mass[-1]
design_takeoff_weight = vehicle.mass_properties.takeoff
max_takeoff_weight = nexus.vehicle_configurations.takeoff.mass_properties.max_takeoff
zero_fuel_weight = payload + operating_empty
# pretty_print(nexus.vehicle_configurations)
for i in range(1, len(results.base.segments)): # make fuel burn and sprayer continuous
# print i
results.base.segments[i].conditions.weights.fuel_burn[:, 0] += \
results.base.segments[i - 1].conditions.weights.fuel_burn[-1]
results.base.segments[i].conditions.weights.spray[:, 0] += \
results.base.segments[i - 1].conditions.weights.spray[-1]
summary.op_empty = operating_empty
summary.max_zero_fuel_margin = (design_landing_weight - operating_empty) / operating_empty # used to be (design_landing_weight - zero_fuel_weight) / zero_fuel_weight changed because of aerosol ejection
summary.base_mission_fuelburn = results.base.segments[-1].conditions.weights.fuel_burn[-1] # esults.base.segments[i].conditions.weights.fuel_burn0#results.base.segments.conditions.weights.fuel_burn #design_takeoff_weight - results.base.segments['descent_3'].conditions.weights.total_mass[-1] # - results.base.segments['cruise'].conditions.sprayer_rate
summary.base_mission_sprayed = results.base.segments[-1].conditions.weights.spray[-1]
summary.cruise_range = 0#missions.base.segments.cruise_2.distance # assume we're flying straight
summary.empty_range = results.base.segments['cruise_outgoing'].conditions.frames.inertial.position_vector[:, 0][-1]/1000.
summary.mission_range = results.base.segments['cruise_final'].conditions.frames.inertial.position_vector[:, 0][-1]/1000. # Assuming mission ends at cruise altitude
summary.spray_range = summary.mission_range - summary.empty_range
summary.total_range = results.base.segments[-1].conditions.frames.inertial.position_vector[:, 0][-1]/1000.
summary.main_mission_time = (results.base.segments['descent_final'].conditions.frames.inertial.time[-1] -
results.base.segments[0].conditions.frames.inertial.time[0])
summary.total_mission_time = (results.base.segments[-1].conditions.frames.inertial.time[-1] -
results.base.segments[0].conditions.frames.inertial.time[0])
summary.MTOW_delta = zero_fuel_weight + summary.base_mission_fuelburn - vehicle.mass_properties.takeoff
# summary.engine_weight =
# print results.base.segments.keys()
# print summary.engine_weight
print "zero fuel weight: ", zero_fuel_weight, "kg i.e. (", payload, "+", operating_empty, ")"
print "MTOW selected: ", vehicle.mass_properties.takeoff, "kg, MTOW calculated: ", zero_fuel_weight + summary.base_mission_fuelburn
# FIXME Fuel burn doesn't change when the operating empty weight changes!
print "Max/Min throttle: ", summary.max_throttle, ", ", summary.min_throttle
print "Take-off field length: ", summary.field_length_takeoff[0], "m"
print "Landing field length: ", summary.field_length_landing[0], "m"
print "Mission Range (must be at least 7000km): ", summary.mission_range, " km"
print "Non-spraying range (must be at least 3500km): ", summary.empty_range, " km"
print "Spraying Range (must be at least 3500km): ", summary.spray_range, " km"
print "Total Range: ", summary.total_range, " km", "(+",summary.total_range-summary.mission_range,")"
print "Mission time: ", summary.main_mission_time[0] * Units['s'] / Units.h, "hours (main) +", \
(summary.total_mission_time - summary.main_mission_time)[0] * Units['s'] / Units.h, "hours (diversion)"
summary.nothing = 0.0
print 'Fuel burn: ', summary.base_mission_fuelburn, " Fuel margin: ", summary.max_zero_fuel_margin
# print 'fuel margin=', problem.all_constraints()
gt_engine = nexus.vehicle_configurations.base.propulsors.turbofan
# print thrust
# FIXME move that to printing/results section
print "Turbofan thrust:", gt_engine.sealevel_static_thrust, " x ", int(
gt_engine.number_of_engines), "engines (tot: ", gt_engine.sealevel_static_thrust * gt_engine.number_of_engines, " N)"
print "Thrust required: ", gt_engine.design_thrust, "N"
print "Estimated engine length: ", gt_engine.engine_length, ", diameter: ", gt_engine.nacelle_diameter, ", wetted area: ", gt_engine.areas.wetted
#FXI
print "Aerosol released: ", summary.base_mission_sprayed[0], " kg\n\n"
# print vehicle.wings.main_wing.chords.root, vehicle.wings.main_wing.spans.projected, vehicle.wings.main_wing.areas.reference
# print "cruise range: ",problem.summary.cruise_range/1000., " km"
hf = vehicle.fuselages.fuselage.heights.at_wing_root_quarter_chord
wf = vehicle.fuselages.fuselage.width
Lf = vehicle.fuselages.fuselage.lengths.total
Sw = vehicle.wings.main_wing.areas.reference
cw = vehicle.wings.main_wing.chords.mean_aerodynamic
b = vehicle.wings.main_wing.spans.projected
Sh = vehicle.wings.horizontal_stabilizer.areas.reference
Sv = vehicle.wings.vertical_stabilizer.areas.reference
lh = vehicle.wings.horizontal_stabilizer.origin[0] + vehicle.wings.horizontal_stabilizer.aerodynamic_center[0] - \
vehicle.mass_properties.center_of_gravity[0]
lv = vehicle.wings.vertical_stabilizer.origin[0] + vehicle.wings.vertical_stabilizer.aerodynamic_center[0] - \
vehicle.mass_properties.center_of_gravity[0]
# when you run want to output results to a file
unscaled_inputs = nexus.optimization_problem.inputs[:, 1] # use optimization problem inputs here
input_scaling = nexus.optimization_problem.inputs[:, 3]
scaled_inputs = unscaled_inputs / input_scaling
problem_inputs = []
# SEGMENTS: need to loop it and add all segments
output_array_segments = np.zeros(4).reshape(4,1)
for i in range(1, len(results.base.segments)):
# print results.base.segments[i].conditions.aerodynamics.lift_coefficient[:, 0]
# print results.base.segments[i].conditions.aerodynamics.angle_of_attack[:, 0] / Units.deg
# print results.base.segments[i].conditions.freestream.dynamic_viscosity[:,0]
# print results.base.segments[i].conditions.freestream.density[:,0]
output_array_i = np.vstack((results.base.segments[i].conditions.aerodynamics.lift_coefficient[:, 0],
results.base.segments[i].conditions.aerodynamics.angle_of_attack[:, 0] / Units.deg,
results.base.segments[i].conditions.freestream.dynamic_viscosity[:,0],
results.base.segments[i].conditions.freestream.density[:,0]))
output_array_segments = np.hstack((output_array_segments,output_array_i))
output_segment_indices = ["CL","alpha","dynamic_visc","air_density"]
# print output_array_segments[segment_output_indexes.index("CL")]
output_array = np.array([
vehicle.wings.main_wing.aspect_ratio,
vehicle.wings.main_wing.areas.reference,
vehicle.wings.main_wing.sweep,
vehicle.wings.main_wing.taper,
vehicle.wings.main_wing.chords.root,
vehicle.wings.main_wing.spans.projected,
vehicle.fuselages.fuselage.effective_diameter,
vehicle.fuselages.fuselage.lengths.total,
vehicle.wings.main_wing.chords.mean_aerodynamic,
2.*vehicle.fuselages.fuselage.origin[1],
configs.takeoff.maximum_lift_coefficient,
configs.landing.maximum_lift_coefficient,
vehicle.maximum_lift_coefficient,
missions.base.segments['descent_final'].air_speed,
missions.base.segments['cruise_1'].air_speed,
gt_engine.number_of_engines,
gt_engine.mass_properties.mass/gt_engine.number_of_engines, #PER ENGINE
gt_engine.mass_properties.mass/1.6/gt_engine.number_of_engines, #PER ENGINE DRY WEIGHT
gt_engine.nacelle_diameter,
summary.base_mission_fuelburn,
zero_fuel_weight + summary.base_mission_fuelburn,
operating_empty,
design_landing_weight,
vehicle.weight_breakdown.wing,
vehicle.weight_breakdown.fuselage,
vehicle.weight_breakdown.landing_gear,
payload,
820., #Fuel density
vehicle.wings.horizontal_stabilizer.sweep,
vehicle.wings.horizontal_stabilizer.spans.projected,
vehicle.wings.horizontal_stabilizer.chords.root,
vehicle.wings.vertical_stabilizer.sweep,
vehicle.wings.vertical_stabilizer.taper,
vehicle.wings.vertical_stabilizer.chords.root
# QUESTIONABLE: CL_alpha, CM_alpha, Lift distribution, CM_delta, CL_delta
])
output_indices = ["A",
"S",
"sweep",
"taper",
"c_r",
"b",
"d_fus",
"l_fus",
"c_mac",
"y_fuselages",
"CL_max_TO",
"CL_max_landing",
"CL_max_clean",
"v_landing",
"v_cr",
"N_engines",
"m_engine_wet",
"m_engine_dry",
"d_engine",
"m_fuel",
"MTOW",
"OEW",
"m_landing",
"m_wing",
"m_fus",
"m_landing_gear",
"m_payload",
"rho_fuel",
"sweep_h",
"b_h",
"c_r_h",
"sweep_v",
"taper_v",
"c_r_v"
]
# print output_array[output_indexes.index("c_r_v")]
# print output_array[-1]
# print vehicle.weight_breakdown
np.save(output_folder+"output_array.npy",output_array)
np.save(output_folder+"output_indices.npy",output_indices)
np.save(output_folder+"output_array_segments.npy",output_array_segments)
np.save(output_folder+"output_segment_indices.npy",output_segment_indices)
for value in unscaled_inputs:
problem_inputs.append(value)
file_out = open(output_folder+'results.txt', 'ab')
file_out.write('fuel weight = ')
file_out.write(str(summary.base_mission_fuelburn))
file_out.write(', inputs = ')
file_out.write(str(problem_inputs))
file_out.write('\n')
file_out.close()
print_weight_breakdown(nexus.vehicle_configurations.base, filename=output_folder + 'weight_breakdown.dat')
#
# # print engine data into file
print_engine_data(nexus.vehicle_configurations.base, filename=output_folder + 'engine_data.dat')
#
# # print parasite drag data into file
# # define reference condition for parasite drag
ref_condition = Data()
ref_condition.mach_number = 0.7 # FIXME
ref_condition.reynolds_number = 7e6 # FIXME
Analyses = Data()
Analyses.configs = nexus.analyses
print_parasite_drag(ref_condition, nexus.vehicle_configurations.cruise, Analyses,
filename=output_folder + 'parasite_drag.dat')
#
# print compressibility drag data into file
# print Analyses
print_compress_drag(nexus.vehicle_configurations.cruise, Analyses, filename=output_folder + 'compress_drag.dat')
# print mission breakdown
print_mission_breakdown(nexus.results.base,
filename=output_folder + 'mission_breakdown.dat') # FIXME fuel weight adds aerosol = wrong!!!!!
return nexus
def main():
# ---------------------------------------------------------------------------------------
# INITIALIZING AIRCRAFT
configs, analyses, vehicle = full_setup()
# analyses.configs.cruise.aerodynamics.settings.drag_coefficient_increment = 0.001
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')
# # print engine data into file
# print_turboprop_data(configs.cruise,filename = 'ATR72_engine_data.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., 1100.]
weights.mass_properties.operating_empty = weights.mass_properties.operating_empty - 239.
# payloadrange = payload_range(config, mission, cruise_segment_tag, reserves)
return
def post_process(nexus):
# Unpack data
vehicle = nexus.vehicle_configurations.base
configs = nexus.vehicle_configurations
results = nexus.results
summary = nexus.summary
missions = nexus.missions
# pretty_print(results)
# start comment
# # Static stability calculations
# CMA = -10.
# for segment in results.base.segments.values():
# max_CMA = np.max(segment.conditions.stability.static.cm_alpha[:, 0])
# # lamda with a min/max?
# if max_CMA > CMA:
# CMA = max_CMA
#
# # Dynamics stability calculations
# # j=0
#
# cn_r = np.array([])
# cl_p = np.array([])
# cl_beta = np.array([])
# cm_q = np.array([])
# cm_alpha_dot = np.array([])
# cz_alpha = np.array([])
# cy_phi = np.array([])
# cl_r = np.array([])
#
# cl_psi = np.array([])
# cz_u = np.array([])
# cz_alpha_dot = np.array([])
# cz_q = np.array([])
# cx_u = np.array([])
# cx_alpha = np.array([])
#
# eigen_modes = np.array([])
#
# for segment in results.base.segments.values():
# # max_CMA = np.max(segment.conditions.stability.static.cm_alpha[:, 0])
#
# # print segment.conditions.stability.dynamic.keys()
# # pretty_print(segment.conditions.stability)
# stability_model = segment.conditions.stability.stability_model
#
# cn_r = np.append(cn_r,segment.conditions.stability.dynamic.cn_r[:,0])
# # cl_p = np.append(cl_p,segment.conditions.stability.dynamic.cl_p)
# # cl_beta = np.append(cl_beta,segment.conditions.stability.dynamic.cl_beta)
# cm_q = np.append(cm_q,segment.conditions.stability.dynamic.cm_q[:, 0])
# cm_alpha_dot = np.append(cm_alpha_dot,segment.conditions.stability.dynamic.cm_alpha_dot[:, 0])
# cz_alpha = np.append(cz_alpha,segment.conditions.stability.dynamic.cz_alpha[:, 0])
#
# cy_phi = np.append(cy_phi,segment.conditions.stability.dynamic.cy_phi[:,0])
# cl_r = np.append(cl_r,segment.conditions.stability.dynamic.cl_r[:,0] )
#
# cl_psi = np.append(cl_psi,segment.conditions.stability.dynamic.cl_psi[:,0] )
# cz_u = np.append(cz_u,segment.conditions.stability.dynamic.cz_u[:,0] )
# cz_alpha_dot = np.append(cz_alpha_dot,segment.conditions.stability.dynamic.cz_alpha_dot[:,0] )
# cz_q = np.append(cz_q,segment.conditions.stability.dynamic.cz_q[:,0] )
# cx_u = np.append(cx_u,segment.conditions.stability.dynamic.cx_u[:,0] )
# cx_alpha = np.append(cx_alpha,segment.conditions.stability.dynamic.cx_alpha[:,0] )
#
#
# eigen_modes = np.append(eigen_modes,
# np.array([stability_model.dutch_roll.natural_frequency,
# stability_model.dutch_roll.damping_ratio,
# stability_model.spiral_tau,
# stability_model.roll_tau,
# stability_model.short_period.natural_frequency,
# stability_model.short_period.damping_ratio,
# stability_model.phugoid.natural_frequency,
# stability_model.phugoid.damping_ratio])
# )
#
# # stability_model = results.stability_model
# #
# # eigen_modes = np.array([stability_model.dutch_roll.natural_frequency,
# # stability_model.dutch_roll.damping_ratio,
# # stability_model.spiral_tau,
# # stability_model.roll_tau,
# # stability_model.short_period.natural_frequency,
# # stability_model.short_period.damping_ratio,
# # stability_model.phugoid.natural_frequency,
# # stability_model.phugoid.damping_ratio])
#
# dyn_stab = np.array([cn_r,cl_p,cl_beta,cm_q,cm_alpha_dot,cz_alpha,
# cy_phi,
# cl_r,
# cl_psi,
# cz_u,
# cz_alpha_dot,
# cz_q,
# cx_u,
# cx_alpha,
# eigen_modes])
# # lol = vehicle.stability_model.dutch_roll
# stop comments
# stability_model = self.stability_model
# cn_r
# cl_p
# cl_beta
# cm_q
# cm_alpha_dot
# cz_alpha
# lamda with a min/max?
# if max_CMA > CMA:
# CMA = max_CMA
# lst=[]
#
# for element in results.base.segments.conditions.stability.dynamic.values():
#
# lst.append(element)
#
# for segment in results.base.segments.values():
#
#
#
# max_CMA = np.max(segment.conditions.stability.static.cm_alpha[:, 0])
# # lamda with a min/max?
# if max_CMA > CMA:
# CMA = max_CMA
#
#
#
# for derivative in
# stability data structure:
# stability
# static
# cm_alpha
# cn_beta
#
#
# dynamic
# cn_r
# cl_p
# 0
# cl_beta
# 0
# cm_q
# cm_alpha_dot
# cz_alpha
#
# summary.static_stability = CMA
#
results = evaluate_field_length(configs, nexus.analyses, missions.base, results)
#
summary.field_length_takeoff = results.field_length.takeoff
# print summary.field_length_takeoff
summary.field_length_landing = results.field_length.landing
# print summary.field_length_landing
#
# #
# pretty_print(nexus)
# throttle in design mission
max_throttle = 0
min_throttle = 0
for segment in results.base.segments.values():
max_segment_throttle = np.max(segment.conditions.propulsion.throttle[:, 0])
min_segment_throttle = np.min(segment.conditions.propulsion.throttle[:, 0])
if max_segment_throttle > max_throttle:
max_throttle = max_segment_throttle
if min_segment_throttle < min_throttle:
min_throttle = min_segment_throttle
summary.max_throttle = max_throttle
summary.min_throttle = min_throttle
# short_w_n, short_zeta, phugoid_w_n, phugoid_zeta = longitudinal(velocity, density, S_gross_w, mac, Cm_q, Cz_alpha, mass, Cm_alpha, Iy, Cm_alpha_dot, Cz_u, Cz_alpha_dot, Cz_q, Cw, Theta, Cx_u, Cx_alpha):
# Fuel margin and base fuel calculations
vehicle.mass_properties.operating_empty += 0e3 # FIXME hardcoded wing mass correction # area scaling?
operating_empty = vehicle.mass_properties.operating_empty
payload = vehicle.mass_properties.payload # TODO fuel margin makes little sense when ejecting aerosol
design_landing_weight = results.base.segments[-1].conditions.weights.total_mass[-1]
design_takeoff_weight = vehicle.mass_properties.takeoff
max_takeoff_weight = nexus.vehicle_configurations.takeoff.mass_properties.max_takeoff
zero_fuel_weight = payload + operating_empty
# pretty_print(nexus.vehicle_configurations)
clmax = 0
for segment in results.base.segments.values():
cl = np.max(segment.conditions.aerodynamics.lift_coefficient[:, 0])
if cl > clmax:
clmax = cl
summary.clmax = clmax
for i in range(1, len(results.base.segments)): # make fuel burn and sprayer continuous
# print i
results.base.segments[i].conditions.weights.fuel_burn[:, 0] += \
results.base.segments[i - 1].conditions.weights.fuel_burn[-1]
results.base.segments[i].conditions.weights.spray[:, 0] += \
results.base.segments[i - 1].conditions.weights.spray[-1]
summary.op_empty = operating_empty
summary.max_zero_fuel_margin = (
design_landing_weight - operating_empty) / operating_empty # used to be (design_landing_weight - zero_fuel_weight) / zero_fuel_weight changed because of aerosol ejection
summary.base_mission_fuelburn = results.base.segments[-1].conditions.weights.fuel_burn[
-1] # esults.base.segments[i].conditions.weights.fuel_burn0#results.base.segments.conditions.weights.fuel_burn #design_takeoff_weight - results.base.segments['descent_3'].conditions.weights.total_mass[-1] # - results.base.segments['cruise'].conditions.sprayer_rate
summary.base_mission_sprayed = results.base.segments[-1].conditions.weights.spray[-1]
summary.cruise_range = 0 # missions.base.segments.cruise_2.distance # assume we're flying straight
summary.empty_range = results.base.segments['cruise_outgoing'].conditions.frames.inertial.position_vector[:, 0][
-1] / 1000.
summary.mission_range = results.base.segments['cruise_final'].conditions.frames.inertial.position_vector[:, 0][
-1] / 1000. # Assuming mission ends at cruise altitude
summary.spray_range = summary.mission_range - summary.empty_range
summary.total_range = results.base.segments[-1].conditions.frames.inertial.position_vector[:, 0][-1] / 1000.
summary.main_mission_time = (results.base.segments['descent_final'].conditions.frames.inertial.time[-1] -
results.base.segments[0].conditions.frames.inertial.time[0])
summary.total_mission_time = (results.base.segments[-1].conditions.frames.inertial.time[-1] -
results.base.segments[0].conditions.frames.inertial.time[0])
summary.MTOW_delta = zero_fuel_weight + summary.base_mission_fuelburn - vehicle.mass_properties.takeoff
# summary.power_draw = results.base.segments.
# summary.engine_weight =
# print results.base.segments.keys()
# print summary.engine_weight
print "zero fuel weight: ", zero_fuel_weight, "kg i.e. (", payload, "+", operating_empty, ")"
print "MTOW selected: ", vehicle.mass_properties.takeoff, "kg, MTOW calculated: ", zero_fuel_weight + summary.base_mission_fuelburn
# FIXME Fuel burn doesn't change when the operating empty weight changes!
print "Max/Min throttle: ", summary.max_throttle, ", ", summary.min_throttle
print "Take-off field length: ", summary.field_length_takeoff[0], "m"
print "Landing field length: ", summary.field_length_landing[0], "m"
print "Mission Range (must be at least 7000km): ", summary.mission_range, " km"
print "Non-spraying range (must be at least 3500km): ", summary.empty_range, " km"
print "Spraying Range (must be at least 3500km): ", summary.spray_range, " km"
print "Total Range: ", summary.total_range, " km", "(+", summary.total_range - summary.mission_range, ")"
print "Mission time: ", summary.main_mission_time[0] * Units['s'] / Units.h, "hours (main) +", \
(summary.total_mission_time - summary.main_mission_time)[0] * Units['s'] / Units.h, "hours (diversion)"
summary.nothing = 0.0
print 'Fuel burn: ', summary.base_mission_fuelburn, " Fuel margin: ", summary.max_zero_fuel_margin
print "CL_max: ", summary.clmax
# print 'fuel margin=', problem.all_constraints()
gt_engine = nexus.vehicle_configurations.base.propulsors.turbofan
# print thrust
# FIXME move that to printing/results section
print "Turbofan thrust:", gt_engine.sealevel_static_thrust, " x ", int(
gt_engine.number_of_engines), "engines (tot: ", gt_engine.sealevel_static_thrust * gt_engine.number_of_engines, " N)"
print "Thrust required: ", gt_engine.design_thrust, "N"
print "Estimated engine length: ", gt_engine.engine_length, ", diameter: ", gt_engine.nacelle_diameter, ", wetted area: ", gt_engine.areas.wetted
# FXI
print "Aerosol released: ", summary.base_mission_sprayed[0], " kg\n\n"
# print vehicle.wings.main_wing.chords.root, vehicle.wings.main_wing.spans.projected, vehicle.wings.main_wing.areas.reference
# print "cruise range: ",problem.summary.cruise_range/1000., " km"
hf = vehicle.fuselages.fuselage.heights.at_wing_root_quarter_chord
wf = vehicle.fuselages.fuselage.width
Lf = vehicle.fuselages.fuselage.lengths.total
Sw = vehicle.wings.main_wing.areas.reference
cw = vehicle.wings.main_wing.chords.mean_aerodynamic
b = vehicle.wings.main_wing.spans.projected
Sh = vehicle.wings.horizontal_stabilizer.areas.reference
Sv = vehicle.wings.vertical_stabilizer.areas.reference
lh = vehicle.wings.horizontal_stabilizer.origin[0] + vehicle.wings.horizontal_stabilizer.aerodynamic_center[0] - \
vehicle.mass_properties.center_of_gravity[0]
lv = vehicle.wings.vertical_stabilizer.origin[0] + vehicle.wings.vertical_stabilizer.aerodynamic_center[0] - \
vehicle.mass_properties.center_of_gravity[0]
# when you run want to output results to a file
unscaled_inputs = nexus.optimization_problem.inputs[:, 1] # use optimization problem inputs here
input_scaling = nexus.optimization_problem.inputs[:, 3]
scaled_inputs = unscaled_inputs / input_scaling
problem_inputs = []
# SEGMENTS: need to loop it and add all segments
output_array_segments = np.zeros(6).reshape(6, 1)
for i in range(0, len(results.base.segments)):
# print results.base.segments[i].conditions.aerodynamics.lift_coefficient[:, 0]
# print results.base.segments[i].conditions.aerodynamics.angle_of_attack[:, 0] / Units.deg
# print results.base.segments[i].conditions.freestream.dynamic_viscosity[:,0]
# print results.base.segments[i].conditions.freestream.density[:,0].shape
# print results.base.segments[i].conditions.freestream.mach_number[:, 0].shape
# print results.base.segments[i].conditions.freestream.reynolds_number[:, 0]
output_array_i = np.vstack((results.base.segments[i].conditions.aerodynamics.lift_coefficient[:, 0],
results.base.segments[i].conditions.aerodynamics.angle_of_attack[:, 0] / Units.deg,
results.base.segments[i].conditions.freestream.dynamic_viscosity[:, 0],
results.base.segments[i].conditions.freestream.density[:, 0],
results.base.segments[i].conditions.freestream.mach_number[:, 0],
results.base.segments[i].conditions.freestream.reynolds_number[:, 0]))
output_array_segments = np.hstack((output_array_segments, output_array_i))
output_segment_indices = ["CL", "alpha", "dynamic_visc", "air_density", "mach", "reynolds_number"]
C_L_des = np.average(output_array_segments[0, 112:176]) # Average CL over the spraying cruise
# print C_L_des
# print output_array_segments[segment_output_indexes.index("CL")]
# print vehicle.wings.main_wing.areas.wetted + vehicle.wings.horizontal_stabilizer.areas.wetted +\
# vehicle.wings.vertical_stabilizer.areas.wetted + vehicle.fuselages.fuselage.areas.wetted
SUAVE.Input_Output.D3JS.save_tree(vehicle, output_folder + 'tree.json')
output_array = np.array([
vehicle.wings.main_wing.aspect_ratio,
vehicle.wings.main_wing.areas.reference,
vehicle.wings.main_wing.sweep,
vehicle.wings.main_wing.taper,
vehicle.wings.main_wing.chords.root,
vehicle.wings.main_wing.spans.projected,
vehicle.fuselages.fuselage.effective_diameter,
vehicle.fuselages.fuselage.lengths.total,
vehicle.wings.main_wing.chords.mean_aerodynamic,
2. * vehicle.fuselages.fuselage.origin[1],
configs.takeoff.maximum_lift_coefficient,
configs.landing.maximum_lift_coefficient,
vehicle.maximum_lift_coefficient,
missions.base.segments['descent_final'].air_speed,
missions.base.segments['cruise_1'].air_speed,
gt_engine.number_of_engines,
gt_engine.mass_properties.mass / gt_engine.number_of_engines, # PER ENGINE
9000, # PER ENGINE DRY WEIGHT
gt_engine.nacelle_diameter,
summary.base_mission_fuelburn,
vehicle.mass_properties.max_takeoff,
operating_empty,
design_landing_weight,
vehicle.weight_breakdown.wing,
vehicle.weight_breakdown.fuselage,
vehicle.weight_breakdown.landing_gear,
payload,
vehicle.weight_breakdown.horizontal_tail,
vehicle.weight_breakdown.vertical_tail,
vehicle.weight_breakdown.systems,
vehicle.weight_breakdown.rudder,
820., # Fuel density
vehicle.wings.horizontal_stabilizer.sweep,
vehicle.wings.horizontal_stabilizer.spans.projected,
vehicle.wings.horizontal_stabilizer.chords.root,
vehicle.wings.vertical_stabilizer.sweep,
vehicle.wings.vertical_stabilizer.taper,
vehicle.wings.vertical_stabilizer.chords.root,
vehicle.wings.main_wing.areas.wetted + vehicle.wings.horizontal_stabilizer.areas.wetted + \
vehicle.wings.vertical_stabilizer.areas.wetted + vehicle.fuselages.fuselage.areas.wetted
# Oswald
# Wetted area
# QUESTIONABLE: CL_alpha, CM_alpha, Lift distribution, CM_delta, CL_delta
])
output_indices = ["A",
"S",
"sweep",
"taper",
"c_r",
"b",
"d_fus",
"l_fus",
"c_mac",
"y_fuselages",
"CL_max_TO",
"CL_max_landing",
"CL_max_clean",
"v_landing",
"v_cr",
"N_engines",
"m_engine_wet",
"m_engine_dry",
"d_engine",
"m_fuel",
"MTOW",
"OEW",
"m_landing",
"m_wing",
"m_fus",
"m_landing_gear",
"m_payload",
"m_htail",
"m_vtail",
"m_systems",
"m_rudder",
"rho_fuel",
"sweep_h",
"b_h",
"c_r_h",
"sweep_v",
"taper_v",
"c_r_v",
"S_wet"
]
print output_array[output_indices.index('l_fus')]
fuel_burn_sec = np.zeros(1)
for i in range(1, len(results.base.segments)):
fuel_burn_sec = np.hstack((fuel_burn_sec, results.base.segments[i - 1].conditions.weights.fuel_burn[-1]))
wing_loading = (
operating_empty + summary.base_mission_fuelburn - fuel_burn_sec) * 9.81 / vehicle.wings.main_wing.areas.reference
# print wing_loading
# print output_array[output_indexes.index("c_r_v")]
# print output_array[-1]
# print vehicle.weight_breakdown
if numpy_export:
np.save(output_folder + "wing_loading.npy", wing_loading)
np.save(output_folder + "output_array.npy", output_array)
np.save(output_folder + "output_indices.npy", output_indices)
np.save(output_folder + "output_array_segments.npy", output_array_segments)
np.save(output_folder + "output_segment_indices.npy", output_segment_indices)
# np.save(output_folder + "dynamic_stability.npy",dyn_stab)
for value in unscaled_inputs:
problem_inputs.append(value)
file_out = open(output_folder + 'results.txt', 'ab')
file_out.write('fuel weight = ')
file_out.write(str(summary.base_mission_fuelburn))
file_out.write(', inputs = ')
file_out.write(str(problem_inputs))
file_out.write('\n')
file_out.close()
# print vehicle.weight_breakdown.wing
print_weight_breakdown(nexus.vehicle_configurations.base, filename=output_folder + 'weight_breakdown.dat')
#
# # print engine data into file
print_engine_data(nexus.vehicle_configurations.base, filename=output_folder + 'engine_data.dat')
#
# # print parasite drag data into file
# # define reference condition for parasite drag
ref_condition = Data()
ref_condition.mach_number = 0.71 # FIXME
ref_condition.reynolds_number = 9e6 # FIXME
Analyses = Data()
Analyses.configs = nexus.analyses
print_parasite_drag(ref_condition, nexus.vehicle_configurations.cruise, Analyses,
filename=output_folder + 'parasite_drag.dat')
#
# print compressibility drag data into file
# print Analyses
print_compress_drag(nexus.vehicle_configurations.cruise, Analyses, filename=output_folder + 'compress_drag.dat')
# print mission breakdown
print_mission_breakdown(nexus.results.base,
filename=output_folder + 'mission_breakdown.dat') # FIXME fuel weight adds aerosol = wrong!!!!!
return nexus
def plot_mission(nexus, id=0, line_style='bo-'):
results = nexus.results.base
axis_font = {'fontname': 'Arial', 'size': '14'}
configs = nexus.vehicle_configurations
analyses = Data()
analyses.configs = Data()
analyses.configs = nexus.analyses
# print mission breakdown
print_mission_breakdown(results,
filename='mission_breakdown' + str(id) + '.dat')
# print engine data into file
# print_engine_data(configs.base, filename='engine_data'+str(id)+'.dat')
# print weight breakdown
print_weight_breakdown(configs.base,
filename='weight_breakdown' + str(id) + '.dat')
# print compressibility drag data into file
print_compress_drag(configs.cruise,
analyses,
filename='compress_drag' + str(id) + '.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,
'parasite_drag' + str(id) + '.dat')
# ------------------------------------------------------------------
# Aerodynamics
# ------------------------------------------------------------------
fig = plt.figure('Mission Parameters' + str(id), figsize=(8, 12))
for segment in results.segments.values():
time = segment.conditions.frames.inertial.time[:, 0] / Units.min
Thrust = segment.conditions.frames.body.thrust_force_vector[:,
0] / Units.lbf
eta = segment.conditions.propulsion.throttle[:, 0]
mdot = segment.conditions.weights.vehicle_mass_rate[:, 0] * 60 * 60
mdot2 = segment.conditions.weights.vehicle_mass_rate[:,
0] / Units['lb/h']
try:
sfc = segment.conditions.propulsion.psfc
pitch = segment.conditions.propulsion.pitch_command[:, 0]
P = segment.conditions.propulsion.propeller_power[:,
0] * 1.34102 / 1000
gas_turbine_power = segment.conditions.propulsion.gas_turbine_power[:,
0] * 1.34102 / 1000
etap = segment.conditions.propulsion.etap[:, 0]
mdot3 = segment.conditions.propulsion.vehicle_mass_rate
sfc2 = mdot2 / P / 2.
except:
sfc = mdot2 / Thrust / 2.
pitch = segment.conditions.propulsion.throttle[:, 0] * 0.
P = segment.conditions.propulsion.throttle[:, 0] * 0.
gas_turbine_power = segment.conditions.propulsion.throttle[:,
0] * 0.
etap = segment.conditions.propulsion.throttle[:, 0] * 0.
sfc2 = sfc
mdot3 = mdot2
axes = fig.add_subplot(5, 1, 1)
axes.plot(time, Thrust, line_style)
axes.set_ylabel('Thrust (lbf)', axis_font)
axes.grid(True)
axes = fig.add_subplot(5, 1, 2)
axes.plot(time, P, 'ko-', label='Propeller')
axes.plot(time, gas_turbine_power, 'ro-', label='Gas Turbine')
axes.set_ylabel('Power (HP)', axis_font)
axes.grid(True)
axes = fig.add_subplot(5, 1, 3)
axes.plot(time, eta, line_style)
axes.set_xlabel('Time (min)', axis_font)
axes.set_ylabel('Throttle', axis_font)
axes.grid(True)
axes = fig.add_subplot(5, 1, 4)
axes.plot(time, pitch * 180 / 3.14159, line_style)
axes.set_xlabel('Time (min)', axis_font)
axes.set_ylabel('Pitch Command', axis_font)
axes.grid(True)
axes = fig.add_subplot(5, 1, 5)
axes.plot(time, etap, line_style)
axes.set_xlabel('Time (min)', axis_font)
axes.set_ylabel('Propeller Efficiency', axis_font)
axes.grid(True)
plt.savefig("engine" + str(id) + ".png")
# ------------------------------------------------------------------
# Aerodynamics
# ------------------------------------------------------------------
fig = plt.figure("Aerodynamic Coefficients", figsize=(8, 10))
for segment in results.segments.values():
time = segment.conditions.frames.inertial.time[:, 0] / Units.min
CLift = segment.conditions.aerodynamics.lift_coefficient[:, 0]
CDrag = segment.conditions.aerodynamics.drag_coefficient[:, 0]
aoa = segment.conditions.aerodynamics.angle_of_attack[:, 0] / Units.deg
l_d = CLift / CDrag
axes = fig.add_subplot(3, 1, 1)
axes.plot(time, CLift, line_style)
axes.set_ylabel('Lift Coefficient', axis_font)
axes.grid(True)
axes = fig.add_subplot(3, 1, 2)
axes.plot(time, l_d, line_style)
axes.set_ylabel('L/D', axis_font)
axes.grid(True)
axes = fig.add_subplot(3, 1, 3)
axes.plot(time, aoa, 'ro-')
axes.set_xlabel('Time (min)', axis_font)
axes.set_ylabel('AOA (deg)', axis_font)
axes.grid(True)
plt.savefig("aero" + str(id) + ".png")
# ------------------------------------------------------------------
# Aerodynamics 2
# ------------------------------------------------------------------
fig = plt.figure("Drag Components", figsize=(8, 10))
axes = plt.gca()
for i, segment in enumerate(results.segments.values()):
time = segment.conditions.frames.inertial.time[:, 0] / Units.min
drag_breakdown = segment.conditions.aerodynamics.drag_breakdown
cdp = drag_breakdown.parasite.total[:, 0]
cdi = drag_breakdown.induced.total[:, 0]
cdc = drag_breakdown.compressible.total[:, 0]
cdm = drag_breakdown.miscellaneous.total[:, 0]
cd = drag_breakdown.total[:, 0]
if line_style == 'bo-':
axes.plot(time, cdp, 'ko-', label='CD parasite')
axes.plot(time, cdi, 'bo-', label='CD induced')
axes.plot(time, cdc, 'go-', label='CD compressibility')
axes.plot(time, cdm, 'yo-', label='CD miscellaneous')
axes.plot(time, cd, 'ro-', label='CD total')
if i == 0:
axes.legend(loc='upper center')
else:
axes.plot(time, cdp, line_style)
axes.plot(time, cdi, line_style)
axes.plot(time, cdc, line_style)
axes.plot(time, cdm, line_style)
axes.plot(time, cd, line_style)
axes.set_xlabel('Time (min)')
axes.set_ylabel('CD')
axes.grid(True)
plt.savefig("drag" + str(id) + ".png")
# ------------------------------------------------------------------
# Altitude, sfc, vehicle weight
# ------------------------------------------------------------------
fig = plt.figure("Altitude_sfc_weight", figsize=(8, 10))
for segment in results.segments.values():
time = segment.conditions.frames.inertial.time[:, 0] / Units.min
aoa = segment.conditions.aerodynamics.angle_of_attack[:, 0] / Units.deg
mass = segment.conditions.weights.total_mass[:, 0] / Units.lb
altitude = segment.conditions.freestream.altitude[:, 0] / Units.ft
mdot = segment.conditions.weights.vehicle_mass_rate[:, 0]
thrust = segment.conditions.frames.body.thrust_force_vector[:, 0]
sfc = (mdot / Units.lb) / (thrust / Units.lbf) * Units.hr
axes = fig.add_subplot(3, 1, 1)
axes.plot(time, altitude, line_style)
axes.set_ylabel('Altitude (ft)', axis_font)
axes.grid(True)
axes = fig.add_subplot(3, 1, 3)
axes.plot(time, sfc, line_style)
axes.set_xlabel('Time (min)', axis_font)
axes.set_ylabel('sfc (lb/lbf-hr)', axis_font)
axes.grid(True)
axes = fig.add_subplot(3, 1, 2)
axes.plot(time, mass, 'ro-')
axes.set_ylabel('Weight (lb)', axis_font)
axes.grid(True)
plt.savefig("mission" + str(id) + ".png")
# plt.show(block=True)
return
def main():
configs, analyses, vehicle = full_setup()
simple_sizing(configs)
configs.finalize()
analyses.finalize()
# weight analysis
weights = analyses.configs.base.weights
breakdown = weights.evaluate()
# analyses.configs.base.weights.mass_properties.operating_empty = 20736. * Units.kg
# mission analysis
mission = analyses.missions.base
results = mission.evaluate()
# print weight breakdown
print_weight_breakdown(configs.base,filename = 'E170_weight_breakdown.dat')
# print engine data into file
print_engine_data(configs.base,filename = 'E170_engine_data.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,'E170_parasite_drag.dat')
# print compressibility drag data into file
print_compress_drag(configs.cruise,analyses,filename = 'E170_compress_drag.dat')
# print mission breakdown
print_mission_breakdown(results,filename='E170_mission_breakdown.dat')
state = Data()
state.conditions = SUAVE.Analyses.Mission.Segments.Conditions.Aerodynamics()
state.numerics = SUAVE.Analyses.Mission.Segments.Conditions.Numerics()
# ---------------------------------------------------------------------------------------
# PAYLOAD RANGE DIAGRAM
config = configs.base
cruise_segment_tag = "cruise"
reserves = 1600.
# payload_range_results = payload_range(config, mission, cruise_segment_tag, reserves)
# ---------------------------------------------------------------------------------------
# PLOT RESULTS
# plot_mission(results)
# ---------------------------------------------------------------------------------------
# TAKE OFF FIELD LENGTH
# ---- Inputs
analyses.base = analyses.configs.base
airport = mission.airport
airport.delta_isa = 0.0
airport.altitude = 8000.0 * Units.ft
clb_grad = 1
weights_tofl = [20000,
21009.319,
22008.12425,
23017.44325,
24005.73477,
25009.79689,
25998.08841,
27427.95699,
28999.76105,
29988.05257,
30997.37157,
32994.98208,
35008.3632,
37005.97372,
38588.29152]
np.linspace(vehicle.mass_properties.operating_empty, vehicle.mass_properties.max_takeoff, 10)
tofl = np.zeros(len(weights_tofl))
secsegclbgrad = np.zeros(len(weights_tofl))
# ---- Run
for i, TOW in enumerate(weights_tofl):
configs.takeoff.mass_properties.takeoff = TOW * Units.kg
tofl[i], secsegclbgrad[i] = estimate_take_off_field_length(configs.takeoff, analyses, airport, clb_grad)
print tofl[i]
import pylab as plt
title = "TOFL vs W"
plt.figure(1)
plt.plot(weights_tofl, tofl, 'k-')
plt.xlabel('TOW (kg)')
plt.ylabel('Takeoff Field Length (m)')
plt.title(title)
plt.grid(True)
plt.show(True)
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