def main():
#initialize the problem
nexus = Nexus()
vehicle = vehicle_setup()
nexus.vehicle_configurations = configs_setup(vehicle)
nexus.analyses = Analyses.setup(nexus.vehicle_configurations)
nexus.missions = Missions.setup(nexus.analyses)
#problem = Data()
#nexus.optimization_problem = problem
nexus.procedure = setup()
nexus.sizing_loop = Sizing_Loop()
nexus.total_number_of_iterations = 0
evaluate_problem(nexus)
results = nexus.results
err = nexus.sizing_loop.norm_error
err_true = 0.00975078 #for 1E-2 tol
error = abs((err-err_true)/err)
print 'error = ', error
assert(error<1e-5), 'sizing loop regression failed'
#output=nexus._really_evaluate() #run; use optimization setup without inputs
return
def main():
vehicle = vehicle_setup()
configs = configs_setup(vehicle)
analyses = mission_B737.analyses_setup(configs)
mission = mission_setup(configs,analyses)
configs.finalize()
analyses.finalize()
results = mission.evaluate()
results = results.merged()
plot_results(results)
distance_regression = 4317710.33719722
distance_calc = results.conditions.frames.inertial.position_vector[-1,0]
error_distance = abs((distance_regression - distance_calc )/distance_regression)
assert error_distance < 1e-6
error_weight = abs(mission.target_landing_weight - results.conditions.weights.total_mass[-1,0])
print('landing weight error' , error_weight)
assert error_weight < 1e-6
return
def main():
vehicle = vehicle_setup()
configs = configs_setup(vehicle)
analyses = mission_B737.analyses_setup(configs)
mission = mission_setup(configs, analyses)
configs.finalize()
analyses.finalize()
results = mission.evaluate()
results = results.merged()
plot_results(results)
distance_regression = 4317710.33719722
distance_calc = results.conditions.frames.inertial.position_vector[-1, 0]
error_distance = abs(
(distance_regression - distance_calc) / distance_regression)
assert error_distance < 1e-6
error_weight = abs(mission.target_landing_weight -
results.conditions.weights.total_mass[-1, 0])
print('landing weight error', error_weight)
assert error_weight < 1e-6
return
def main():
# vehicle data
vehicle = vehicle_setup()
configs = configs_setup(vehicle)
# vehicle analyses
configs_analyses = analyses_setup(configs)
aerodynamics = SUAVE.Analyses.Aerodynamics.AVL()
stability = SUAVE.Analyses.Stability.AVL()
aerodynamics.geometry = copy.deepcopy(configs.cruise)
stability.geometry = copy.deepcopy(configs.cruise)
aerodynamics.process.compute.lift.inviscid.training_file = 'base_data_aerodynamics.txt'
stability.training_file = 'base_data_stability.txt'
configs_analyses.cruise.append(aerodynamics)
configs_analyses.cruise.append(stability)
# mission analyses
mission = mission_setup(configs_analyses)
missions_analyses = missions_setup(mission)
analyses = SUAVE.Analyses.Analysis.Container()
analyses.configs = configs_analyses
analyses.missions = missions_analyses
simple_sizing(configs, analyses)
configs.finalize()
analyses.finalize()
# mission analysis
mission = analyses.missions.base
results = mission.evaluate()
# lift coefficient check
lift_coefficient = results.conditions.cruise.aerodynamics.lift_coefficient[
0]
lift_coefficient_true = 0.59571034
print lift_coefficient
diff_CL = np.abs(lift_coefficient - lift_coefficient_true)
print 'CL difference'
print diff_CL
assert np.abs((lift_coefficient - lift_coefficient_true) /
lift_coefficient_true) < 1e-6
# moment coefficient check
moment_coefficient = results.conditions.cruise.stability.static.CM[0][0]
moment_coefficient_true = -0.62167644
print moment_coefficient
diff_CM = np.abs(moment_coefficient - moment_coefficient_true)
print 'CM difference'
print diff_CM
assert np.abs((moment_coefficient - moment_coefficient_true) /
moment_coefficient_true) < 1e-6
return
def main():
# Setup for converging on weight
vehicle = vehicle_setup()
configs = configs_setup(vehicle)
analyses = mission_B737.analyses_setup(configs)
mission = mission_setup(configs, analyses)
configs.finalize()
analyses.finalize()
results = mission.evaluate()
results = results.merged()
plot_results(results)
distance_regression = 3804806.720225211
distance_calc = results.conditions.frames.inertial.position_vector[-1, 0]
print('distance_calc = ', distance_calc)
error_distance = abs(
(distance_regression - distance_calc) / distance_regression)
assert error_distance < 1e-6
error_weight = abs(mission.target_landing_weight -
results.conditions.weights.total_mass[-1, 0])
print('landing weight error', error_weight)
assert error_weight < 1e-6
# Setup for converging on SOC, using the stopped rotor vehicle
vehicle_SR, analyses_SR = full_setup_SR()
analyses_SR.finalize()
mission_SR = analyses_SR.mission
results_SR = mission_SR.evaluate()
results_SR = results_SR.merged()
distance_regression_SR = 101649.83535243798
distance_calc_SR = results_SR.conditions.frames.inertial.position_vector[
-1, 0]
print('distance_calc_SR = ', distance_calc_SR)
error_distance_SR = abs(
(distance_regression_SR - distance_calc_SR) / distance_regression_SR)
assert error_distance_SR < 1e-6
error_soc = abs(
mission_SR.target_state_of_charge -
results_SR.conditions.propulsion.battery_state_of_charge[-1, 0])
print('landing state of charge error', error_soc)
assert error_soc < 1e-6
return
def main():
#initialize the problem
nexus = Nexus()
vehicle = vehicle_setup()
nexus.vehicle_configurations = configs_setup(vehicle)
nexus.analyses = Analyses.setup(nexus.vehicle_configurations)
nexus.missions = Missions.setup(nexus.analyses)
problem = Data()
problem_inputs = np.array([
['dummy', 1., (.1, 10.), 1., ' continuous', Units.less],
['dummy2', 2., (.1, 10.), 1., ' continuous', Units.less],
]) #create dummy inputs for optimization to test io
problem.inputs = problem_inputs
nexus.optimization_problem = problem
nexus.procedure = setup()
sizing_loop = Sizing_Loop()
sizing_loop.output_filename = 'sizing_outputs.txt'
nexus.sizing_loop = sizing_loop
#create a fake array of data to test outputs
write_sizing_outputs(sizing_loop, np.array([6.]), [5., 5.])
write_sizing_outputs(sizing_loop, np.array([12.]), [4., 1.])
write_sizing_outputs(sizing_loop, np.array([11.]), [1., 3.])
nexus.total_number_of_iterations = 0
evaluate_problem(nexus)
results = nexus.results
err = nexus.sizing_loop.norm_error
err_true = 0.0008433474527249522 #for 1E-2 tol
error = abs((err - err_true) / err_true)
data_inputs, data_outputs, read_success = read_sizing_residuals(
sizing_loop, problem.inputs)
check_read_res = -0.06803060191281879
error_res = (data_outputs[1][0] - check_read_res) / check_read_res
#remove files for later
os.remove('sizing_outputs.txt')
os.remove('y_err_values.txt')
print('error = ', error)
print('error_res = ', error_res)
assert (error < 1e-4), 'sizing loop regression failed'
assert (error_res < 1e-4), 'sizing loop io failed'
return
def full_setup():
# vehicle data
vehicle = vehicle_setup()
configs = configs_setup(vehicle)
# vehicle analyses
configs_analyses = analyses_setup(configs)
# mission analyses
mission = mission_setup(configs_analyses)
missions_analyses = missions_setup(mission)
analyses = SUAVE.Analyses.Analysis.Container()
analyses.configs = configs_analyses
analyses.missions = missions_analyses
return configs, analyses
def full_setup():
# vehicle data
vehicle = vehicle_setup()
configs = configs_setup(vehicle)
# vehicle analyses
configs_analyses = analyses_setup(configs)
# mission analyses
mission = mission_setup(configs_analyses)
missions_analyses = missions_setup(mission)
analyses = SUAVE.Analyses.Analysis.Container()
analyses.configs = configs_analyses
analyses.missions = missions_analyses
return configs, analyses
def main():
vehicle = vehicle_setup()
configs = configs_setup(vehicle)
analyses = mission_B737.analyses_setup(configs)
mission = mission_setup(configs,analyses)
vehicle.mass_properties.takeoff = 70000 * Units.kg
configs.finalize()
analyses.finalize()
results = mission.evaluate()
results = results.merged()
plot_results(results)
error = abs(mission.target_landing_weight - results.conditions.weights.total_mass[-1,0])
print 'landing weight error' , error
assert error < 1.
return
def main():
#initialize the problem
nexus = Nexus()
vehicle = vehicle_setup()
nexus.vehicle_configurations = configs_setup(vehicle)
nexus.analyses = Analyses.setup(nexus.vehicle_configurations)
nexus.missions = Missions.setup(nexus.analyses)
nexus.procedure = setup()
nexus.sizing_loop = Sizing_Loop()
nexus.total_number_of_iterations = 0
evaluate_problem(nexus)
results = nexus.results
err = nexus.sizing_loop.norm_error
err_true = 0.0096907307307155348 #for 1E-2 tol
error = abs((err_true - err) / err_true)
print 'error = ', error
assert (error < 1e-6), 'sizing loop regression failed'
return
def main():
vehicle = vehicle_setup()
configs = configs_setup(vehicle)
analyses = mission_B737.analyses_setup(configs)
mission = mission_setup(configs, analyses)
vehicle.mass_properties.takeoff = 70000 * Units.kg
configs.finalize()
analyses.finalize()
results = mission.evaluate()
results = results.merged()
plot_results(results)
error = abs(mission.target_landing_weight -
results.conditions.weights.total_mass[-1, 0])
print 'landing weight error', error
assert error < 1.
return
def main():
#initialize the problem
nexus = Nexus()
vehicle = vehicle_setup()
nexus.vehicle_configurations = configs_setup(vehicle)
nexus.analyses = Analyses.setup(nexus.vehicle_configurations)
nexus.missions = Missions.setup(nexus.analyses)
nexus.procedure = setup()
nexus.sizing_loop = Sizing_Loop()
nexus.total_number_of_iterations = 0
evaluate_problem(nexus)
results = nexus.results
err = nexus.sizing_loop.norm_error
err_true = 0.0096907307307155348#for 1E-2 tol
error = abs((err_true-err)/err_true)
print 'error = ', error
assert(error<1e-6), 'sizing loop regression failed'
return
def main():
# vehicle data
vehicle = vehicle_setup()
configs = configs_setup(vehicle)
# vehicle analyses
configs_analyses = analyses_setup(configs)
# append AVL aerodynamic analysis
aerodynamics = SUAVE.Analyses.Aerodynamics.AVL()
aerodynamics.process.compute.lift.inviscid.regression_flag = True
aerodynamics.process.compute.lift.inviscid.keep_files = True
aerodynamics.geometry = copy.deepcopy(configs.cruise)
aerodynamics.process.compute.lift.inviscid.training_file = 'cruise_data_aerodynamics.txt'
configs_analyses.cruise.append(aerodynamics)
# append AVL stability analysis
stability = SUAVE.Analyses.Stability.AVL()
stability.regression_flag = True
stability.keep_files = True
stability.geometry = copy.deepcopy(configs.cruise)
stability.training_file = 'cruise_data_stability.txt'
configs_analyses.cruise.append(stability)
# ------------------------------------------------------------------
# Initialize the Mission
# ------------------------------------------------------------------
mission = SUAVE.Analyses.Mission.Sequential_Segments()
mission.tag = 'the_mission'
#airport
airport = SUAVE.Attributes.Airports.Airport()
airport.altitude = 0.0 * Units.ft
airport.delta_isa = 0.0
airport.atmosphere = SUAVE.Attributes.Atmospheres.Earth.US_Standard_1976()
mission.airport = airport
# unpack Segments module
Segments = SUAVE.Analyses.Mission.Segments
# base segment
base_segment = Segments.Segment()
# ------------------------------------------------------------------
# Cruise Segment: constant speed, constant altitude
# ------------------------------------------------------------------
segment = Segments.Cruise.Constant_Speed_Constant_Altitude(base_segment)
segment.tag = "cruise"
segment.analyses.extend( configs_analyses.cruise )
segment.air_speed = 230. * Units['m/s']
segment.distance = 4000. * Units.km
segment.altitude = 10.668 * Units.km
segment.state.numerics.number_control_points = 4
# add to mission
mission.append_segment(segment)
missions_analyses = missions_setup(mission)
analyses = SUAVE.Analyses.Analysis.Container()
analyses.configs = configs_analyses
analyses.missions = missions_analyses
simple_sizing(configs, analyses)
configs.finalize()
analyses.finalize()
# mission analysis
mission = analyses.missions.base
results = mission.evaluate()
# lift coefficient check
lift_coefficient = results.segments.cruise.conditions.aerodynamics.lift_coefficient[0][0]
lift_coefficient_true = 0.6118979131570086
print(lift_coefficient)
diff_CL = np.abs(lift_coefficient - lift_coefficient_true)
print('CL difference')
print(diff_CL)
assert np.abs((lift_coefficient - lift_coefficient_true)/lift_coefficient_true) < 1e-6
# moment coefficient check
moment_coefficient = results.segments.cruise.conditions.stability.static.CM[0][0]
moment_coefficient_true = -0.6267104237340391
print(moment_coefficient)
diff_CM = np.abs(moment_coefficient - moment_coefficient_true)
print('CM difference')
print(diff_CM)
assert np.abs((moment_coefficient - moment_coefficient_true)/moment_coefficient_true) < 1e-6
return
def setup():
nexus = Nexus()
problem = Data()
nexus.optimization_problem = problem
# -------------------------------------------------------------------
# Inputs
# -------------------------------------------------------------------
# [ tag , initial, [lb,ub], scaling, units ]
problem.inputs = np.array([
['wing_area', 124.8, (70., 200.), 124.8, Units.meter**2],
['wing_aspect_ratio', 10.18, (5., 20.), 10.18, Units.less],
['wing_sweep', 25., (0., 35.), 25., Units.degrees],
['wing_thickness', 0.105, (0.07, 0.20), 0.105, Units.less],
['design_thrust', 52700., (10000., 70000.), 52700., Units.N],
['MTOW', 79090., (20000., 100000.), 79090., Units.kg],
['MZFW_ratio', 0.77, (0.6, 0.99), 0.77, Units.less],
['flap_takeoff_angle', 10., (0., 20.), 10., Units.degrees],
['flap_landing_angle', 40., (0., 50.), 40., Units.degrees],
['short_field_TOW', 64030., (20000., 100000.), 64030., Units.kg],
['design_TOW', 68520., (20000., 100000.), 68520., Units.kg],
['noise_takeoff_speed_increase', 10.0, (10., 20.), 10.0, Units.knots],
['noise_cutback_altitude', 304.8, (240., 400.), 304.8, Units.meter],
])
# -------------------------------------------------------------------
# Objective
# -------------------------------------------------------------------
problem.objective = np.array([
['noise_cumulative_margin', 17, Units.less],
])
# -------------------------------------------------------------------
# Constraints
# -------------------------------------------------------------------
# [ tag, sense, edge, scaling, units ]
problem.constraints = np.array([
['MZFW consistency', '>', 0., 10, Units.less],
['design_range_fuel_margin', '>', 0., 10, Units.less],
['short_field_fuel_margin', '>', 0., 10, Units.less],
['max_range_fuel_margin', '>', 0., 10, Units.less],
['wing_span', '<', 35.9664, 35.9664, Units.less],
['noise_flyover_margin', '>', 0., 10., Units.less],
['noise_sideline_margin', '>', 0., 10., Units.less],
['noise_approach_margin', '>', 0., 10., Units.less],
['takeoff_field_length', '<', 1985., 1985., Units.meters],
['landing_field_length', '<', 1385., 1385., Units.meters],
['2nd_segment_climb_max_range', '>', 0.024, 0.024, Units.less],
['2nd_segment_climb_short_field', '>', 0.024, 0.024, Units.less],
['max_throttle', '<', 1., 1., Units.less],
['short_takeoff_field_length', '<', 1330., 1330., Units.meters],
['noise_cumulative_margin', '>', 10., 10., Units.less],
])
# -------------------------------------------------------------------
# Aliases
# -------------------------------------------------------------------
problem.aliases = [
[
'wing_area',
[
'vehicle_configurations.*.wings.main_wing.areas.reference',
'vehicle_configurations.*.reference_area'
]
],
[
'wing_aspect_ratio',
'vehicle_configurations.*.wings.main_wing.aspect_ratio'
],
[
'wing_incidence',
'vehicle_configurations.*.wings.main_wing.twists.root'
],
[
'wing_tip_twist',
'vehicle_configurations.*.wings.main_wing.twists.tip'
],
[
'wing_sweep',
'vehicle_configurations.*.wings.main_wing.sweeps.quarter_chord'
],
[
'wing_thickness',
'vehicle_configurations.*.wings.main_wing.thickness_to_chord'
],
['wing_taper', 'vehicle_configurations.*.wings.main_wing.taper'],
[
'wing_location',
'vehicle_configurations.*.wings.main_wing.origin[0]'
],
[
'horizontal_tail_area',
'vehicle_configurations.*.wings.horizontal_stabilizer.areas.reference'
],
[
'horizontal_tail_aspect_ratio',
'vehicle_configurations.*.wings.horizontal_stabilizer.aspect_ratio'
],
[
'vertical_tail_area',
'vehicle_configurations.*.wings.vertical_stabilizer.areas.reference'
],
[
'vertical_tail_aspect_ratio',
'vehicle_configurations.*.wings.vertical_stabilizer.aspect_ratio'
],
[
'design_thrust',
'vehicle_configurations.*.propulsors.turbofan.thrust.total_design'
],
[
'MTOW',
[
'vehicle_configurations.*.mass_properties.takeoff',
'vehicle_configurations.*.mass_properties.max_takeoff'
]
],
['design_TOW', 'vehicle_configurations.base.mass_properties.takeoff'],
[
'short_field_TOW',
'vehicle_configurations.short_field_takeoff.mass_properties.takeoff'
],
[
'flap_takeoff_angle',
[
'vehicle_configurations.takeoff.wings.main_wing.control_surfaces.flap.deflection',
'vehicle_configurations.short_field_takeoff.wings.main_wing.control_surfaces.flap.deflection'
]
],
[
'flap_landing_angle',
'vehicle_configurations.landing.wings.main_wing.control_surfaces.flap.deflection'
],
[
'slat_takeoff_angle',
[
'vehicle_configurations.takeoff.wings.main_wing.control_surfaces.slat.deflection',
'vehicle_configurations.short_field_takeoff.wings.main_wing.control_surfaces.slat.deflection'
]
],
[
'slat_landing_angle',
'vehicle_configurations.landing.wings.main_wing.control_surfaces.slat.deflection'
],
[
'wing_span',
'vehicle_configurations.base.wings.main_wing.spans.projected'
],
['noise_approach_margin', 'summary.noise_approach_margin'],
['noise_sideline_margin', 'summary.noise_sideline_margin'],
['noise_flyover_margin', 'summary.noise_flyover_margin'],
['static_stability', 'summary.static_stability'],
[
'vertical_tail_volume_coefficient',
'summary.vertical_tail_volume_coefficient'
],
[
'horizontal_tail_volume_coefficient',
'summary.horizontal_tail_volume_coefficient'
],
['wing_max_cl_norm', 'summary.maximum_cl_norm'],
['design_range_fuel_margin', 'summary.design_range_fuel_margin'],
['takeoff_field_length', 'summary.takeoff_field_length'],
['landing_field_length', 'summary.landing_field_length'],
['short_takeoff_field_length', 'summary.short_takeoff_field_length'],
[
'2nd_segment_climb_max_range',
'summary.second_segment_climb_gradient_takeoff'
],
[
'2nd_segment_climb_short_field',
'summary.second_segment_climb_gradient_short_field'
],
['max_throttle', 'summary.max_throttle'],
['short_field_fuel_margin', 'summary.short_field_fuel_margin'],
['max_range_fuel_margin', 'summary.max_range_fuel_margin'],
['max_range', 'missions.max_range_distance'],
['MZFW consistency', 'summary.MZFW_consistency'],
['MZFW_ratio', 'MZFW_ratio'],
['noise_takeoff_speed_increase', 'noise_V2_increase'],
[
'noise_cutback_altitude',
'missions.takeoff.segments.climb.altitude_end'
],
['noise_cumulative_margin', 'summary.noise_margin'],
['weighted_sum_objective', 'summary.weighted_sum_objective'],
]
# -------------------------------------------------------------------
# Vehicles
# -------------------------------------------------------------------
vehicle = vehicle_setup()
nexus.vehicle_configurations = configs_setup(vehicle)
# -------------------------------------------------------------------
# Analyses
# -------------------------------------------------------------------
nexus.analyses = Analyses.setup(nexus.vehicle_configurations)
# -------------------------------------------------------------------
# Missions
# -------------------------------------------------------------------
nexus.missions = Missions.setup(nexus.analyses)
# -------------------------------------------------------------------
# Procedure
# -------------------------------------------------------------------
nexus.procedure = Procedure.setup()
# -------------------------------------------------------------------
# Summary
# -------------------------------------------------------------------
nexus.summary = Data()
return nexus
def main():
# initialize the vehicle
vehicle = vehicle_setup()
for wing in vehicle.wings:
wing.areas.wetted = 2.0 * wing.areas.reference
wing.areas.exposed = 0.8 * wing.areas.wetted
wing.areas.affected = 0.6 * wing.areas.wetted
# initalize the aero model
aerodynamics = SUAVE.Analyses.Aerodynamics.Fidelity_Zero()
aerodynamics.geometry = vehicle
aerodynamics.initialize()
#no of test points
test_num = 11
#specify the angle of attack
angle_of_attacks = np.linspace(-.174, .174, test_num)[:,
None] #* Units.deg
# Cruise conditions (except Mach number)
state = SUAVE.Analyses.Mission.Segments.Conditions.State()
state.conditions = SUAVE.Analyses.Mission.Segments.Conditions.Aerodynamics(
)
state.expand_rows(test_num)
# --------------------------------------------------------------------
# Initialize variables needed for CL and CD calculations
# Use a pre-run random order for values
# --------------------------------------------------------------------
Mc = np.array([[0.9], [0.475], [0.05], [0.39], [0.815], [0.645], [0.305],
[0.22], [0.56], [0.73], [0.135]])
rho = np.array([[0.8], [1.], [0.5], [1.1], [0.4], [1.3], [0.6], [0.3],
[0.9], [0.7], [1.2]])
mu = np.array([[1.85e-05], [1.55e-05], [1.40e-05], [1.10e-05], [2.00e-05],
[8.00e-06], [6.50e-06], [9.50e-06], [1.70e-05], [1.25e-05],
[5.00e-06]])
T = np.array([[270.], [250.], [280.], [260.], [240.], [200.], [290.],
[230.], [210.], [300.], [220.]])
pressure = np.array([[100000.], [190000.], [280000.], [370000.], [460000.],
[550000.], [640000.], [730000.], [820000.], [910000.],
[1000000.]])
re = np.array([[12819987.97468646], [9713525.47464844], [599012.59815633],
[12606549.94372309],
[5062187.10214493], [29714816.00808047], [9611290.40694227],
[2112171.68320523], [8612638.72342302], [14194381.78364854],
[9633881.90543247]])
air = Air()
a = air.compute_speed_of_sound(T, pressure)
re = rho * a * Mc / mu
state.conditions.freestream.mach_number = Mc
state.conditions.freestream.density = rho
state.conditions.freestream.dynamic_viscosity = mu
state.conditions.freestream.temperature = T
state.conditions.freestream.pressure = pressure
state.conditions.freestream.reynolds_number = re
state.conditions.aerodynamics.angle_of_attack = angle_of_attacks
# --------------------------------------------------------------------
# Surrogate
# --------------------------------------------------------------------
#call the aero model
results = aerodynamics.evaluate(state)
#build a polar for the markup aero
polar = Data()
CL = results.lift.total
CD = results.drag.total
polar.lift = CL
polar.drag = CD
# --------------------------------------------------------------------
# Test compute Lift
# --------------------------------------------------------------------
#compute_aircraft_lift(conditions, configuration, geometry)
lift = state.conditions.aerodynamics.lift_coefficient
lift_r = np.array([
-1.21776694, -0.48854139, -0.43529644, -0.34913149, -0.20409426,
0.11639443, 0.35889527, 0.58542263, 0.88028747, 1.24354956, 1.26293499
])[:, None]
print('lift = ', lift)
lift_test = np.abs((lift - lift_r) / lift)
print('\nCompute Lift Test Results\n')
#print lift_test
assert (np.max(lift_test) <
1e-6), 'Aero regression failed at compute lift test'
# --------------------------------------------------------------------
# Test compute drag
# --------------------------------------------------------------------
#compute_aircraft_drag(conditions, configuration, geometry)
# Pull calculated values
drag_breakdown = state.conditions.aerodynamics.drag_breakdown
# Only one wing is evaluated since they rely on the same function
cd_c = drag_breakdown.compressible['main_wing'].compressibility_drag
cd_i = drag_breakdown.induced.total
cd_m = drag_breakdown.miscellaneous.total
# cd_m_fuse_base = drag_breakdown.miscellaneous.fuselage_base
# cd_m_fuse_up = drag_breakdown.miscellaneous.fuselage_upsweep
# cd_m_nac_base = drag_breakdown.miscellaneous.nacelle_base['turbofan']
# cd_m_ctrl = drag_breakdown.miscellaneous.control_gaps
cd_p_fuse = drag_breakdown.parasite['fuselage'].parasite_drag_coefficient
cd_p_wing = drag_breakdown.parasite['main_wing'].parasite_drag_coefficient
cd_tot = drag_breakdown.total
print('cd_m =', cd_m)
(cd_c_r, cd_i_r, cd_m_r, cd_m_fuse_base_r, cd_m_fuse_up_r, cd_m_nac_base_r,
cd_m_ctrl_r, cd_p_fuse_r, cd_p_wing_r, cd_tot_r) = reg_values()
drag_tests = Data()
drag_tests.cd_c = np.abs((cd_c - cd_c_r) / cd_c)
for ii, cd in enumerate(drag_tests.cd_c):
if np.isnan(cd):
drag_tests.cd_c[ii] = np.abs(
(cd_c[ii] - cd_c_r[ii]) / np.min(cd_c[cd_c != 0]))
drag_tests.cd_i = np.abs((cd_i - cd_i_r) / cd_i)
drag_tests.cd_m = np.abs((cd_m - cd_m_r) / cd_m)
## Commented lines represent values not set by current drag functions, but to be recreated in the future
# Line below is not normalized since regression values are 0, insert commented line if this changes
# drag_tests.cd_m_fuse_base = np.abs((cd_m_fuse_base-cd_m_fuse_base_r)) # np.abs((cd_m_fuse_base-cd_m_fuse_base_r)/cd_m_fuse_base)
# drag_tests.cd_m_fuse_up = np.abs((cd_m_fuse_up - cd_m_fuse_up_r)/cd_m_fuse_up)
# drag_tests.cd_m_ctrl = np.abs((cd_m_ctrl - cd_m_ctrl_r)/cd_m_ctrl)
drag_tests.cd_p_fuse = np.abs((cd_p_fuse - cd_p_fuse_r) / cd_p_fuse)
drag_tests.cd_p_wing = np.abs((cd_p_wing - cd_p_wing_r) / cd_p_wing)
drag_tests.cd_tot = np.abs((cd_tot - cd_tot_r) / cd_tot)
print('\nCompute Drag Test Results\n')
print('cd_tot=', cd_tot)
for i, tests in list(drag_tests.items()):
assert (np.max(tests) < 1e-4), 'Aero regression test failed at ' + i
def main():
vehicle = vehicle_setup() # Create the vehicle for testing
for wing in vehicle.wings:
wing.areas.wetted = 2.0 * wing.areas.reference
wing.areas.exposed = 0.8 * wing.areas.wetted
wing.areas.affected = 0.6 * wing.areas.wetted
aerodynamics = SUAVE.Analyses.Aerodynamics.Supersonic_Zero()
aerodynamics.geometry = vehicle
aerodynamics.settings.drag_coefficient_increment = 0.0000
vehicle.aerodynamics_model = aerodynamics
vehicle.aerodynamics_model.initialize()
test_num = 11 # Length of arrays used in this test
# --------------------------------------------------------------------
# Test Lift Surrogate
# --------------------------------------------------------------------
AoA = np.linspace(-.174, .174, test_num)[:, None] # +- 10 degrees
# Cruise conditions (except Mach number)
state = SUAVE.Analyses.Mission.Segments.Conditions.State()
state.conditions = SUAVE.Analyses.Mission.Segments.Conditions.Aerodynamics(
)
state.expand_rows(test_num)
# --------------------------------------------------------------------
# Initialize variables needed for CL and CD calculations
# Use a seeded random order for values
# --------------------------------------------------------------------
random.seed(1)
Mc = np.linspace(0.05, 0.9, test_num)
random.shuffle(Mc)
AoA = AoA.reshape(test_num, 1)
Mc = Mc.reshape(test_num, 1)
rho = np.linspace(0.3, 1.3, test_num)
random.shuffle(rho)
rho = rho.reshape(test_num, 1)
mu = np.linspace(5 * 10**-6, 20 * 10**-6, test_num)
random.shuffle(mu)
mu = mu.reshape(test_num, 1)
T = np.linspace(200, 300, test_num)
random.shuffle(T)
T = T.reshape(test_num, 1)
pressure = np.linspace(10**5, 10**6, test_num)
pressure = pressure.reshape(test_num, 1)
state.conditions.freestream.mach_number = Mc
state.conditions.freestream.density = rho
state.conditions.freestream.dynamic_viscosity = mu
state.conditions.freestream.temperature = T
state.conditions.freestream.pressure = pressure
state.conditions.aerodynamics.angle_of_attack = AoA
# --------------------------------------------------------------------
# Surrogate
# --------------------------------------------------------------------
#call the aero model
results = aerodynamics.evaluate(state)
#build a polar for the markup aero
polar = Data()
CL = results.lift.total
CD = results.drag.total
polar.lift = CL
polar.drag = CD
# --------------------------------------------------------------------
# Test compute Lift
# --------------------------------------------------------------------
#compute_aircraft_lift(conditions, configuration, geometry)
lift = state.conditions.aerodynamics.lift_coefficient
# Truth value
lift_r = np.array([
-2.42489437, -0.90696416, -0.53991953, -0.3044834, -0.03710598,
0.31061936, 0.52106899, 0.77407765, 1.22389024, 1.86240501, 1.54587835
])[:, None]
lift_r = lift_r.reshape(test_num, 1)
lift_test = np.abs((lift - lift_r) / lift)
print '\nCompute Lift Test Results\n'
print lift_test
assert (np.max(lift_test) <
1e-4), 'Supersonic Aero regression failed at compute lift test'
# --------------------------------------------------------------------
# Test compute drag
# --------------------------------------------------------------------
#compute_aircraft_drag(conditions, configuration, geometry)
# Pull calculated values
drag_breakdown = state.conditions.aerodynamics.drag_breakdown
# Only one wing is evaluated since they rely on the same function
cd_c = drag_breakdown.compressible['main_wing'].compressibility_drag
cd_i = drag_breakdown.induced.total
cd_m = drag_breakdown.miscellaneous.total
cd_m_fuse_base = drag_breakdown.miscellaneous.fuselage_base
cd_m_fuse_up = drag_breakdown.miscellaneous.fuselage_upsweep
cd_m_nac_base = drag_breakdown.miscellaneous.nacelle_base['turbo_fan']
cd_m_ctrl = drag_breakdown.miscellaneous.control_gaps
cd_p_fuse = drag_breakdown.parasite['fuselage'].parasite_drag_coefficient
cd_p_wing = drag_breakdown.parasite['main_wing'].parasite_drag_coefficient
cd_tot = drag_breakdown.total
print cd_c
print cd_i
print cd_m
print cd_m_fuse_base
print cd_m_fuse_up
print cd_m_nac_base
print cd_m_ctrl
print cd_p_fuse
print cd_p_wing
print cd_tot
# Truth values
(cd_c_r, cd_i_r, cd_m_r, cd_m_fuse_base_r, cd_m_fuse_up_r, cd_m_nac_base_r,
cd_m_ctrl_r, cd_p_fuse_r, cd_p_wing_r, cd_tot_r) = reg_values()
cd_c_r = cd_c_r.reshape(test_num, 1)
cd_i_r = cd_i_r.reshape(test_num, 1)
cd_m_r = cd_m_r.reshape(test_num, 1)
cd_m_fuse_base_r = cd_m_fuse_base_r.reshape(test_num, 1)
cd_m_fuse_up_r = cd_m_fuse_up_r.reshape(test_num, 1)
cd_m_nac_base_r = cd_m_nac_base_r.reshape(test_num, 1)
cd_m_ctrl_r = cd_m_ctrl_r.reshape(test_num, 1)
cd_p_fuse_r = cd_p_fuse_r.reshape(test_num, 1)
cd_p_wing_r = cd_p_wing_r.reshape(test_num, 1)
cd_tot_r = cd_tot_r.reshape(test_num, 1)
drag_tests = Data()
drag_tests.cd_c = np.abs((cd_c - cd_c_r) / cd_c)
drag_tests.cd_i = np.abs((cd_i - cd_i_r) / cd_i)
drag_tests.cd_m = np.abs((cd_m - cd_m_r) / cd_m)
# Line below is not normalized since regression values are 0, insert commented line if this changes
drag_tests.cd_m_fuse_base = np.abs(
(cd_m_fuse_base - cd_m_fuse_base_r
)) # np.abs((cd_m_fuse_base-cd_m_fuse_base_r)/cd_m_fuse_base)
drag_tests.cd_m_fuse_up = np.abs(
(cd_m_fuse_up - cd_m_fuse_up_r) / cd_m_fuse_up)
drag_tests.cd_m_ctrl = np.abs((cd_m_ctrl - cd_m_ctrl_r) / cd_m_ctrl)
drag_tests.cd_p_fuse = np.abs((cd_p_fuse - cd_p_fuse_r) / cd_p_fuse)
drag_tests.cd_p_wing = np.abs((cd_p_wing - cd_p_wing_r) / cd_p_wing)
drag_tests.cd_tot = np.abs((cd_tot - cd_tot_r) / cd_tot)
print '\nCompute Drag Test Results\n'
#print drag_tests
for i, tests in drag_tests.items():
assert (np.max(tests) <
1e-4), 'Supersonic Aero regression test failed at ' + i
def main():
# initialize the vehicle
vehicle = vehicle_setup()
for wing in vehicle.wings:
wing.areas.wetted = 2.0 * wing.areas.reference
wing.areas.exposed = 0.8 * wing.areas.wetted
wing.areas.affected = 0.6 * wing.areas.wetted
# initalize the aero model
aerodynamics = SUAVE.Analyses.Aerodynamics.Fidelity_Zero()
aerodynamics.process.compute.lift.inviscid_wings = SUAVE.Analyses.Aerodynamics.Lifting_Line()
aerodynamics.geometry = vehicle
aerodynamics.initialize()
#no of test points
test_num = 11
#specify the angle of attack
angle_of_attacks = np.linspace(-.174,.174,test_num)[:,None] #* Units.deg
# Cruise conditions (except Mach number)
state = SUAVE.Analyses.Mission.Segments.Conditions.State()
state.conditions = SUAVE.Analyses.Mission.Segments.Conditions.Aerodynamics()
state.expand_rows(test_num)
# --------------------------------------------------------------------
# Initialize variables needed for CL and CD calculations
# Use a seeded random order for values
# --------------------------------------------------------------------
random.seed(1)
Mc = np.linspace(0.05,0.9,test_num)
rho = np.linspace(0.3,1.3,test_num)
mu = np.linspace(5*10**-6,20*10**-6,test_num)
T = np.linspace(200,300,test_num)
pressure = np.linspace(10**5,10**6,test_num)
random.shuffle(Mc)
random.shuffle(rho)
random.shuffle(mu)
random.shuffle(T)
# Changed after to preserve seed for initial testing
Mc = Mc[:,None]
rho = rho[:,None]
mu = mu[:,None]
T = T[:,None]
pressure = pressure[:,None]
air = Air()
a = air.compute_speed_of_sound(T,pressure)
re = rho*a*Mc/mu
state.conditions.freestream.mach_number = Mc
state.conditions.freestream.density = rho
state.conditions.freestream.dynamic_viscosity = mu
state.conditions.freestream.temperature = T
state.conditions.freestream.pressure = pressure
state.conditions.freestream.reynolds_number = re
state.conditions.aerodynamics.angle_of_attack = angle_of_attacks
# --------------------------------------------------------------------
# Surrogate
# --------------------------------------------------------------------
#call the aero model
results = aerodynamics.evaluate(state)
#build a polar for the markup aero
CL = results.lift.total
CD = results.drag.total
# --------------------------------------------------------------------
# Test compute Lift
# --------------------------------------------------------------------
#compute_aircraft_lift(conditions, configuration, geometry)
lift = state.conditions.aerodynamics.lift_coefficient
lift_r = np.array([-2.84689226, -1.06501674, -0.63426096, -0.35809118, -0.04487569,
0.36343181, 0.61055156, 0.90742419, 1.43504496, 2.18401103, 1.81298486])[:,None]
print 'lift = ', lift
lift_test = np.abs((lift-lift_r)/lift)
print '\nCompute Lift Test Results\n'
#print lift_test
assert(np.max(lift_test)<1e-4), 'Aero regression failed at compute lift test'
def main():
vehicle = vehicle_setup()
weight = Tube_Wing.empty(vehicle)
# regression values
actual = Data()
actual.payload = 27349.9081525 #includes cargo #17349.9081525 #without cargo
actual.pax = 15036.587065500002
actual.bag = 2313.3210870000003
actual.fuel = 12977.803363592691 #includes cargo #22177.6377131 #without cargo
actual.empty = 38688.08848390731
actual.wing = 6649.709658738429
actual.fuselage = 6642.061164271899
actual.propulsion = 6838.185174956626
actual.landing_gear = 3160.632
actual.systems = 13479.10479056802
actual.wt_furnish = 6431.80372889
actual.horizontal_tail = 1037.7414196819743
actual.vertical_tail = 629.0387683502595
actual.rudder = 251.61550734010382
# error calculations
error = Data()
error.payload = (actual.payload - weight.payload) / actual.payload
error.pax = (actual.pax - weight.pax) / actual.pax
error.bag = (actual.bag - weight.bag) / actual.bag
error.fuel = (actual.fuel - weight.fuel) / actual.fuel
error.empty = (actual.empty - weight.empty) / actual.empty
error.wing = (actual.wing - weight.wing) / actual.wing
error.fuselage = (actual.fuselage - weight.fuselage) / actual.fuselage
error.propulsion = (actual.propulsion -
weight.propulsion) / actual.propulsion
error.landing_gear = (actual.landing_gear -
weight.landing_gear) / actual.landing_gear
error.systems = (actual.systems - weight.systems) / actual.systems
error.wt_furnish = (actual.wt_furnish -
weight.systems_breakdown.furnish) / actual.wt_furnish
error.horizontal_tail = (actual.horizontal_tail -
weight.horizontal_tail) / actual.horizontal_tail
error.vertical_tail = (actual.vertical_tail -
weight.vertical_tail) / actual.vertical_tail
error.rudder = (actual.rudder - weight.rudder) / actual.rudder
print('Results (kg)')
print(weight)
print('Relative Errors')
print(error)
for k, v in list(error.items()):
assert (np.abs(v) < 1E-6)
#General Aviation weights; note that values are taken from Raymer,
#but there is a huge spread among the GA designs, so individual components
#differ a good deal from the actual design
vehicle = vehicle_setup_general_aviation()
GTOW = vehicle.mass_properties.max_takeoff
weight = General_Aviation.empty(vehicle)
weight.fuel = vehicle.fuel.mass_properties.mass
actual = Data()
actual.bag = 0.
actual.empty = 618.485310343
actual.fuel = 144.69596603
actual.wing = 124.673093906
actual.fuselage = 119.522072873
actual.propulsion = 194.477769922 #includes power plant and propeller, does not include fuel system
actual.landing_gear = 44.8033840543 + 5.27975390045
actual.furnishing = 37.8341395817
actual.electrical = 36.7532226254
actual.control_systems = 14.8331955546
actual.fuel_systems = 15.6859717453
actual.systems = 108.096549345
error = Data()
error.fuel = (actual.fuel - weight.fuel) / actual.fuel
error.empty = (actual.empty - weight.empty) / actual.empty
error.wing = (actual.wing - weight.wing) / actual.wing
error.fuselage = (actual.fuselage - weight.fuselage) / actual.fuselage
error.propulsion = (actual.propulsion -
weight.propulsion) / actual.propulsion
error.landing_gear = (actual.landing_gear -
(weight.landing_gear_main +
weight.landing_gear_nose)) / actual.landing_gear
error.furnishing = (actual.furnishing -
weight.systems_breakdown.furnish) / actual.furnishing
error.electrical = (actual.electrical - weight.systems_breakdown.electrical
) / actual.electrical
error.control_systems = (
actual.control_systems -
weight.systems_breakdown.control_systems) / actual.control_systems
error.fuel_systems = (
actual.fuel_systems -
weight.systems_breakdown.fuel_system) / actual.fuel_systems
error.systems = (actual.systems - weight.systems) / actual.systems
print('actual.systems=', actual.systems)
print('General Aviation Results (kg)')
print(weight)
print('Relative Errors')
print(error)
for k, v in list(error.items()):
assert (np.abs(v) < 1e-6)
# BWB WEIGHTS
vehicle = bwb_setup()
weight = BWB.empty(vehicle)
# regression values
actual = Data()
actual.payload = 27349.9081525 #includes cargo #17349.9081525 #without cargo
actual.pax = 15036.587065500002
actual.bag = 2313.3210870000003
actual.fuel = 24860.343951919327
actual.empty = 26805.547895580676
actual.wing = 6576.679767012152
actual.fuselage = 1.0
actual.propulsion = 1413.8593105126783
actual.landing_gear = 3160.632
actual.systems = 15654.376818055844
actual.wt_furnish = 8205.349895589
# error calculations
error = Data()
error.payload = (actual.payload - weight.payload) / actual.payload
error.pax = (actual.pax - weight.pax) / actual.pax
error.bag = (actual.bag - weight.bag) / actual.bag
error.fuel = (actual.fuel - weight.fuel) / actual.fuel
error.empty = (actual.empty - weight.empty) / actual.empty
error.wing = (actual.wing - weight.wing) / actual.wing
error.fuselage = (actual.fuselage -
(weight.fuselage + 1.0)) / actual.fuselage
error.propulsion = (actual.propulsion -
weight.propulsion) / actual.propulsion
error.systems = (actual.systems - weight.systems) / actual.systems
error.wt_furnish = (actual.wt_furnish -
weight.systems_breakdown.furnish) / actual.wt_furnish
print('Results (kg)')
print(weight)
print('Relative Errors')
print(error)
for k, v in list(error.items()):
assert (np.abs(v) < 1E-6)
# Human Powered Aircraft
vehicle = hp_setup()
weight = HP.empty(vehicle)
# regression values
actual = Data()
actual.empty = 138.02737768459374
actual.wing = 89.86286881794777
actual.fuselage = 1.0
actual.horizontal_tail = 31.749272074174737
actual.vertical_tail = 16.415236792471237
# error calculations
error = Data()
error.empty = (actual.empty - weight.empty) / actual.empty
error.wing = (actual.wing - weight.wing) / actual.wing
error.fuselage = (actual.fuselage -
(weight.fuselage + 1.0)) / actual.fuselage
error.horizontal_tail = (actual.horizontal_tail -
weight.horizontal_tail) / actual.horizontal_tail
error.vertical_tail = (actual.vertical_tail -
weight.vertical_tail) / actual.vertical_tail
print('Results (kg)')
print(weight)
print('Relative Errors')
print(error)
for k, v in list(error.items()):
assert (np.abs(v) < 1E-6)
return
def main():
# initialize the vehicle
vehicle = vehicle_setup()
for wing in vehicle.wings:
wing.areas.wetted = 2.0 * wing.areas.reference
wing.areas.exposed = 0.8 * wing.areas.wetted
wing.areas.affected = 0.6 * wing.areas.wetted
# initalize the aero model
aerodynamics = SUAVE.Analyses.Aerodynamics.Fidelity_Zero()
aerodynamics.settings.number_spanwise_vortices = 5
aerodynamics.settings.number_chordwise_vortices = 2
aerodynamics.geometry = vehicle
aerodynamics.initialize()
#no of test points
test_num = 11
#specify the angle of attack
angle_of_attacks = np.linspace(-.174,.174,test_num)[:,None] #* Units.deg
# Cruise conditions (except Mach number)
state = SUAVE.Analyses.Mission.Segments.Conditions.State()
state.conditions = SUAVE.Analyses.Mission.Segments.Conditions.Aerodynamics()
state.expand_rows(test_num)
# --------------------------------------------------------------------
# Initialize variables needed for CL and CD calculations
# Use a pre-run random order for values
# --------------------------------------------------------------------
Mc = np.array([[0.9 ],
[0.475],
[0.05 ],
[0.39 ],
[0.815],
[0.645],
[0.305],
[0.22 ],
[0.56 ],
[0.73 ],
[0.135]])
rho = np.array([[0.8],
[1. ],
[0.5],
[1.1],
[0.4],
[1.3],
[0.6],
[0.3],
[0.9],
[0.7],
[1.2]])
mu = np.array([[1.85e-05],
[1.55e-05],
[1.40e-05],
[1.10e-05],
[2.00e-05],
[8.00e-06],
[6.50e-06],
[9.50e-06],
[1.70e-05],
[1.25e-05],
[5.00e-06]])
T = np.array([[270.],
[250.],
[280.],
[260.],
[240.],
[200.],
[290.],
[230.],
[210.],
[300.],
[220.]])
pressure = np.array([[ 100000.],
[ 190000.],
[ 280000.],
[ 370000.],
[ 460000.],
[ 550000.],
[ 640000.],
[ 730000.],
[ 820000.],
[ 910000.],
[1000000.]])
re = np.array([[12819987.97468646],
[ 9713525.47464844],
[ 599012.59815633],
[12606549.94372309],
[ 5062187.10214493],
[29714816.00808047],
[ 9611290.40694227],
[ 2112171.68320523],
[ 8612638.72342302],
[14194381.78364854],
[ 9633881.90543247]])
air = Air()
a = air.compute_speed_of_sound(T,pressure)
re = rho*a*Mc/mu
state.conditions.freestream.mach_number = Mc
state.conditions.freestream.density = rho
state.conditions.freestream.dynamic_viscosity = mu
state.conditions.freestream.temperature = T
state.conditions.freestream.pressure = pressure
state.conditions.freestream.reynolds_number = re
state.conditions.aerodynamics.angle_of_attack = angle_of_attacks
# --------------------------------------------------------------------
# Surrogate
# --------------------------------------------------------------------
#call the aero model
results = aerodynamics.evaluate(state)
#build a polar for the markup aero
polar = Data()
CL = results.lift.total
CD = results.drag.total
polar.lift = CL
polar.drag = CD
# --------------------------------------------------------------------
# Test compute Lift
# --------------------------------------------------------------------
#compute_aircraft_lift(conditions, configuration, geometry)
lift = state.conditions.aerodynamics.lift_coefficient
print('lift = \n', lift)
print('\nCompute Lift Test Results\n')
#print lift_test
# --------------------------------------------------------------------
# Test compute drag
# --------------------------------------------------------------------
#compute_aircraft_drag(conditions, configuration, geometry)
# Pull calculated values
drag_breakdown = state.conditions.aerodynamics.drag_breakdown
# Only one wing is evaluated since they rely on the same function
cd_c = drag_breakdown.compressible['main_wing'].compressibility_drag
cd_i = drag_breakdown.induced.total
cd_m = drag_breakdown.miscellaneous.total
# cd_m_fuse_base = drag_breakdown.miscellaneous.fuselage_base
# cd_m_fuse_up = drag_breakdown.miscellaneous.fuselage_upsweep
# cd_m_nac_base = drag_breakdown.miscellaneous.nacelle_base['turbofan']
# cd_m_ctrl = drag_breakdown.miscellaneous.control_gaps
cd_p_fuse = drag_breakdown.parasite['fuselage'].parasite_drag_coefficient
cd_p_wing = drag_breakdown.parasite['main_wing'].parasite_drag_coefficient
cd_tot = drag_breakdown.total
print('cd_m =\n', cd_m)
#drag_tests = Data()
## Commented lines represent values not set by current drag functions, but to be recreated in the future
# Line below is not normalized since regression values are 0, insert commented line if this changes
# drag_tests.cd_m_fuse_base = np.abs((cd_m_fuse_base-cd_m_fuse_base_r)) # np.abs((cd_m_fuse_base-cd_m_fuse_base_r)/cd_m_fuse_base)
# drag_tests.cd_m_fuse_up = np.abs((cd_m_fuse_up - cd_m_fuse_up_r)/cd_m_fuse_up)
# drag_tests.cd_m_ctrl = np.abs((cd_m_ctrl - cd_m_ctrl_r)/cd_m_ctrl)
print('\nCompute Drag Test Results\n')
print('cd_tot=', cd_tot)
# --------------------------------------------------------------------
# Process All Results
# --------------------------------------------------------------------
results = make_results_object(lift, cd_c, cd_i, cd_m, cd_p_fuse, cd_p_wing, cd_tot)
save_results(results, SAVE=False)
results_regression = load_results()
test_results(results, results_regression)
#return conditions, configuration, geometry, test_num
return
def main():
# initialize the vehicle
vehicle = vehicle_setup()
for wing in vehicle.wings:
wing.areas.wetted = 2.0 * wing.areas.reference
wing.areas.exposed = 0.8 * wing.areas.wetted
wing.areas.affected = 0.6 * wing.areas.wetted
# initalize the aero model
aerodynamics = SUAVE.Analyses.Aerodynamics.Fidelity_Zero()
aerodynamics.geometry = vehicle
aerodynamics.initialize()
#no of test points
test_num = 11
#specify the angle of attack
angle_of_attacks = np.linspace(-.174,.174,test_num)[:,None] #* Units.deg
# Cruise conditions (except Mach number)
state = SUAVE.Analyses.Mission.Segments.Conditions.State()
state.conditions = SUAVE.Analyses.Mission.Segments.Conditions.Aerodynamics()
state.expand_rows(test_num)
# --------------------------------------------------------------------
# Initialize variables needed for CL and CD calculations
# Use a seeded random order for values
# --------------------------------------------------------------------
random.seed(1)
Mc = np.linspace(0.05,0.9,test_num)
random.shuffle(Mc)
rho = np.linspace(0.3,1.3,test_num)
random.shuffle(rho)
mu = np.linspace(5*10**-6,20*10**-6,test_num)
random.shuffle(mu)
T = np.linspace(200,300,test_num)
random.shuffle(T)
pressure = np.linspace(10**5,10**6,test_num)
# Changed after to preserve seed for initial testing
Mc = Mc[:,None]
rho = rho[:,None]
mu = mu[:,None]
T = T[:,None]
pressure = pressure[:,None]
air = Air()
a = air.compute_speed_of_sound(T,pressure)
re = rho*a*Mc/mu
state.conditions.freestream.mach_number = Mc
state.conditions.freestream.density = rho
state.conditions.freestream.dynamic_viscosity = mu
state.conditions.freestream.temperature = T
state.conditions.freestream.pressure = pressure
state.conditions.freestream.reynolds_number = re
state.conditions.aerodynamics.angle_of_attack = angle_of_attacks
# --------------------------------------------------------------------
# Surrogate
# --------------------------------------------------------------------
#call the aero model
results = aerodynamics.evaluate(state)
#build a polar for the markup aero
polar = Data()
CL = results.lift.total
CD = results.drag.total
polar.lift = CL
polar.drag = CD
# --------------------------------------------------------------------
# Test compute Lift
# --------------------------------------------------------------------
#compute_aircraft_lift(conditions, configuration, geometry)
lift = state.conditions.aerodynamics.lift_coefficient
lift_r = np.array( [-2.17277359, -0.77516232, -0.41769607, -0.16530511, 0.19456498, 0.49425496, \
0.67481247, 0.93041268, 1.41531217, 2.1033578, 1.71822138])[:,None]
print 'lift = ', lift
lift_test = np.abs((lift-lift_r)/lift)
print '\nCompute Lift Test Results\n'
#print lift_test
assert(np.max(lift_test)<1e-4), 'Aero regression failed at compute lift test'
# --------------------------------------------------------------------
# Test compute drag
# --------------------------------------------------------------------
#compute_aircraft_drag(conditions, configuration, geometry)
# Pull calculated values
drag_breakdown = state.conditions.aerodynamics.drag_breakdown
# Only one wing is evaluated since they rely on the same function
cd_c = drag_breakdown.compressible['main_wing'].compressibility_drag
cd_i = drag_breakdown.induced.total
cd_m = drag_breakdown.miscellaneous.total
# cd_m_fuse_base = drag_breakdown.miscellaneous.fuselage_base
# cd_m_fuse_up = drag_breakdown.miscellaneous.fuselage_upsweep
# cd_m_nac_base = drag_breakdown.miscellaneous.nacelle_base['turbofan']
# cd_m_ctrl = drag_breakdown.miscellaneous.control_gaps
cd_p_fuse = drag_breakdown.parasite['fuselage'].parasite_drag_coefficient
cd_p_wing = drag_breakdown.parasite['main_wing'].parasite_drag_coefficient
cd_tot = drag_breakdown.total
print 'cd_m =', cd_m
(cd_c_r, cd_i_r, cd_m_r, cd_m_fuse_base_r, cd_m_fuse_up_r, cd_m_nac_base_r, cd_m_ctrl_r, cd_p_fuse_r, cd_p_wing_r, cd_tot_r) = reg_values()
drag_tests = Data()
drag_tests.cd_c = np.abs((cd_c-cd_c_r)/cd_c)
for ii,cd in enumerate(drag_tests.cd_c):
if np.isnan(cd):
drag_tests.cd_c[ii] = np.abs((cd_c[ii]-cd_c_r[ii])/np.min(cd_c[cd_c!=0]))
drag_tests.cd_i = np.abs((cd_i-cd_i_r)/cd_i)
drag_tests.cd_m = np.abs((cd_m-cd_m_r)/cd_m)
## Commented lines represent values not set by current drag functions, but to be recreated in the future
# Line below is not normalized since regression values are 0, insert commented line if this changes
# drag_tests.cd_m_fuse_base = np.abs((cd_m_fuse_base-cd_m_fuse_base_r)) # np.abs((cd_m_fuse_base-cd_m_fuse_base_r)/cd_m_fuse_base)
# drag_tests.cd_m_fuse_up = np.abs((cd_m_fuse_up - cd_m_fuse_up_r)/cd_m_fuse_up)
# drag_tests.cd_m_ctrl = np.abs((cd_m_ctrl - cd_m_ctrl_r)/cd_m_ctrl)
drag_tests.cd_p_fuse = np.abs((cd_p_fuse - cd_p_fuse_r)/cd_p_fuse)
drag_tests.cd_p_wing = np.abs((cd_p_wing - cd_p_wing_r)/cd_p_wing)
drag_tests.cd_tot = np.abs((cd_tot - cd_tot_r)/cd_tot)
print '\nCompute Drag Test Results\n'
print 'cd_tot=', cd_tot
for i, tests in drag_tests.items():
assert(np.max(tests)<1e-4),'Aero regression test failed at ' + i
def main():
vehicle = vehicle_setup()
weight = Tube_Wing.empty(vehicle)
# regression values
actual = Data()
actual.payload = 27349.9081525 #includes cargo #17349.9081525 #without cargo
actual.pax = 15036.587065500002
actual.bag = 2313.3210870000003
actual.fuel = 12990.957450008464 #includes cargo #22177.6377131 #without cargo
actual.empty = 38674.934397491539
actual.wing = 6649.7096587384294
actual.fuselage = 6642.0611642718986
actual.propulsion = 6838.1851749566231
actual.landing_gear = 3160.632
actual.systems = 13479.10479056802
actual.wt_furnish = 6431.80372889
actual.horizontal_tail = 1024.5873332662029
actual.vertical_tail = 629.03876835025949
actual.rudder = 251.61550734010382
# error calculations
error = Data()
error.payload = (actual.payload - weight.payload)/actual.payload
error.pax = (actual.pax - weight.pax)/actual.pax
error.bag = (actual.bag - weight.bag)/actual.bag
error.fuel = (actual.fuel - weight.fuel)/actual.fuel
error.empty = (actual.empty - weight.empty)/actual.empty
error.wing = (actual.wing - weight.wing)/actual.wing
error.fuselage = (actual.fuselage - weight.fuselage)/actual.fuselage
error.propulsion = (actual.propulsion - weight.propulsion)/actual.propulsion
error.landing_gear = (actual.landing_gear - weight.landing_gear)/actual.landing_gear
error.systems = (actual.systems - weight.systems)/actual.systems
error.wt_furnish = (actual.wt_furnish - weight.systems_breakdown.furnish)/actual.wt_furnish
error.horizontal_tail = (actual.horizontal_tail - weight.horizontal_tail)/actual.horizontal_tail
error.vertical_tail = (actual.vertical_tail - weight.vertical_tail)/actual.vertical_tail
error.rudder = (actual.rudder - weight.rudder)/actual.rudder
print 'Results (kg)'
print weight
print 'Relative Errors'
print error
for k,v in error.items():
assert(np.abs(v)<1E-6)
#General Aviation weights; note that values are taken from Raymer,
#but there is a huge spread among the GA designs, so individual components
#differ a good deal from the actual design
vehicle = vehicle_setup_general_aviation()
GTOW = vehicle.mass_properties.max_takeoff
weight = General_Aviation.empty(vehicle)
weight.fuel = vehicle.fuel.mass_properties.mass
actual = Data()
actual.bag = 0.
actual.empty = 618.485310343
actual.fuel = 144.69596603
actual.wing = 124.673093906
actual.fuselage = 119.522072873
actual.propulsion = 194.477769922 #includes power plant and propeller, does not include fuel system
actual.landing_gear = 44.8033840543+5.27975390045
actual.furnishing = 37.8341395817
actual.electrical = 36.7532226254
actual.control_systems = 14.8331955546
actual.fuel_systems = 15.6859717453
actual.systems = 108.096549345
error = Data()
error.fuel = (actual.fuel - weight.fuel)/actual.fuel
error.empty = (actual.empty - weight.empty)/actual.empty
error.wing = (actual.wing - weight.wing)/actual.wing
error.fuselage = (actual.fuselage - weight.fuselage)/actual.fuselage
error.propulsion = (actual.propulsion - weight.propulsion)/actual.propulsion
error.landing_gear = (actual.landing_gear - (weight.landing_gear_main+weight.landing_gear_nose))/actual.landing_gear
error.furnishing = (actual.furnishing-weight.systems_breakdown.furnish)/actual.furnishing
error.electrical = (actual.electrical-weight.systems_breakdown.electrical)/actual.electrical
error.control_systems = (actual.control_systems-weight.systems_breakdown.control_systems)/actual.control_systems
error.fuel_systems = (actual.fuel_systems-weight.systems_breakdown.fuel_system)/actual.fuel_systems
error.systems = (actual.systems - weight.systems)/actual.systems
print 'actual.systems=', actual.systems
print 'General Aviation Results (kg)'
print weight
print 'Relative Errors'
print error
for k,v in error.items():
assert(np.abs(v)<1e-6)
return
def main():
# initialize the vehicle
vehicle = vehicle_setup()
for wing in vehicle.wings:
wing.areas.wetted = 2.0 * wing.areas.reference
wing.areas.exposed = 0.8 * wing.areas.wetted
wing.areas.affected = 0.6 * wing.areas.wetted
# initalize the aero model
aerodynamics = SUAVE.Analyses.Aerodynamics.Fidelity_Zero()
aerodynamics.process.compute.lift.inviscid_wings = SUAVE.Analyses.Aerodynamics.Lifting_Line(
)
aerodynamics.geometry = vehicle
aerodynamics.initialize()
#no of test points
test_num = 11
#specify the angle of attack
angle_of_attacks = np.linspace(-.174, .174, test_num)[:,
None] #* Units.deg
# Cruise conditions (except Mach number)
state = SUAVE.Analyses.Mission.Segments.Conditions.State()
state.conditions = SUAVE.Analyses.Mission.Segments.Conditions.Aerodynamics(
)
state.expand_rows(test_num)
# --------------------------------------------------------------------
# Initialize variables needed for CL and CD calculations
# Use a seeded random order for values
# --------------------------------------------------------------------
random.seed(1)
Mc = np.linspace(0.05, 0.9, test_num)
rho = np.linspace(0.3, 1.3, test_num)
mu = np.linspace(5 * 10**-6, 20 * 10**-6, test_num)
T = np.linspace(200, 300, test_num)
pressure = np.linspace(10**5, 10**6, test_num)
random.shuffle(Mc)
random.shuffle(rho)
random.shuffle(mu)
random.shuffle(T)
# Changed after to preserve seed for initial testing
Mc = Mc[:, None]
rho = rho[:, None]
mu = mu[:, None]
T = T[:, None]
pressure = pressure[:, None]
air = Air()
a = air.compute_speed_of_sound(T, pressure)
re = rho * a * Mc / mu
state.conditions.freestream.mach_number = Mc
state.conditions.freestream.density = rho
state.conditions.freestream.dynamic_viscosity = mu
state.conditions.freestream.temperature = T
state.conditions.freestream.pressure = pressure
state.conditions.freestream.reynolds_number = re
state.conditions.aerodynamics.angle_of_attack = angle_of_attacks
# --------------------------------------------------------------------
# Surrogate
# --------------------------------------------------------------------
#call the aero model
results = aerodynamics.evaluate(state)
#build a polar for the markup aero
CL = results.lift.total
CD = results.drag.total
# --------------------------------------------------------------------
# Test compute Lift
# --------------------------------------------------------------------
#compute_aircraft_lift(conditions, configuration, geometry)
lift = state.conditions.aerodynamics.lift_coefficient
lift_r = np.array([
-2.84689226, -1.06501674, -0.63426096, -0.35809118, -0.04487569,
0.36343181, 0.61055156, 0.90742419, 1.43504496, 2.18401103, 1.81298486
])[:, None]
print 'lift = ', lift
lift_test = np.abs((lift - lift_r) / lift)
print '\nCompute Lift Test Results\n'
#print lift_test
assert (np.max(lift_test) <
1e-4), 'Aero regression failed at compute lift test'
def main():
vehicle = vehicle_setup() # Create the vehicle for testing
for wing in vehicle.wings:
wing.areas.wetted = 2.0 * wing.areas.reference
wing.areas.exposed = 0.8 * wing.areas.wetted
wing.areas.affected = 0.6 * wing.areas.wetted
aerodynamics = SUAVE.Analyses.Aerodynamics.Supersonic_Zero()
aerodynamics.geometry = vehicle
aerodynamics.settings.drag_coefficient_increment = 0.0000
vehicle.aerodynamics_model = aerodynamics
vehicle.aerodynamics_model.initialize()
test_num = 11 # Length of arrays used in this test
# --------------------------------------------------------------------
# Test Lift Surrogate
# --------------------------------------------------------------------
AoA = np.linspace(-.174,.174,test_num)[:,None] # +- 10 degrees
# Cruise conditions (except Mach number)
state = SUAVE.Analyses.Mission.Segments.Conditions.State()
state.conditions = SUAVE.Analyses.Mission.Segments.Conditions.Aerodynamics()
state.expand_rows(test_num)
# --------------------------------------------------------------------
# Initialize variables needed for CL and CD calculations
# Use a seeded random order for values
# --------------------------------------------------------------------
random.seed(1)
Mc = np.linspace(0.05,0.9,test_num)
random.shuffle(Mc)
AoA = AoA.reshape(test_num,1)
Mc = Mc.reshape(test_num,1)
rho = np.linspace(0.3,1.3,test_num)
random.shuffle(rho)
rho = rho.reshape(test_num,1)
mu = np.linspace(5*10**-6,20*10**-6,test_num)
random.shuffle(mu)
mu = mu.reshape(test_num,1)
T = np.linspace(200,300,test_num)
random.shuffle(T)
T = T.reshape(test_num,1)
pressure = np.linspace(10**5,10**6,test_num)
pressure = pressure.reshape(test_num,1)
state.conditions.freestream.mach_number = Mc
state.conditions.freestream.density = rho
state.conditions.freestream.dynamic_viscosity = mu
state.conditions.freestream.temperature = T
state.conditions.freestream.pressure = pressure
state.conditions.aerodynamics.angle_of_attack = AoA
# --------------------------------------------------------------------
# Surrogate
# --------------------------------------------------------------------
#call the aero model
results = aerodynamics.evaluate(state)
#build a polar for the markup aero
polar = Data()
CL = results.lift.total
CD = results.drag.total
polar.lift = CL
polar.drag = CD
# --------------------------------------------------------------------
# Test compute Lift
# --------------------------------------------------------------------
#compute_aircraft_lift(conditions, configuration, geometry)
lift = state.conditions.aerodynamics.lift_coefficient
# Truth value
lift_r = np.array([-2.42489437, -0.90696416, -0.53991953, -0.3044834 , -0.03710598,
0.31061936 , 0.52106899, 0.77407765, 1.22389024, 1.86240501,
1.54587835])[:,None]
lift_r = lift_r.reshape(test_num,1)
lift_test = np.abs((lift-lift_r)/lift)
print('\nCompute Lift Test Results\n')
print(lift_test)
assert(np.max(lift_test)<1e-4), 'Supersonic Aero regression failed at compute lift test'
# --------------------------------------------------------------------
# Test compute drag
# --------------------------------------------------------------------
#compute_aircraft_drag(conditions, configuration, geometry)
# Pull calculated values
drag_breakdown = state.conditions.aerodynamics.drag_breakdown
# Only one wing is evaluated since they rely on the same function
cd_c = drag_breakdown.compressible['main_wing'].compressibility_drag
cd_i = drag_breakdown.induced.total
cd_m = drag_breakdown.miscellaneous.total
cd_m_fuse_base = drag_breakdown.miscellaneous.fuselage_base
cd_m_fuse_up = drag_breakdown.miscellaneous.fuselage_upsweep
cd_m_nac_base = drag_breakdown.miscellaneous.nacelle_base['turbo_fan']
cd_m_ctrl = drag_breakdown.miscellaneous.control_gaps
cd_p_fuse = drag_breakdown.parasite['fuselage'].parasite_drag_coefficient
cd_p_wing = drag_breakdown.parasite['main_wing'].parasite_drag_coefficient
cd_tot = drag_breakdown.total
print(cd_c)
print(cd_i)
print(cd_m)
print(cd_m_fuse_base)
print(cd_m_fuse_up)
print(cd_m_nac_base)
print(cd_m_ctrl)
print(cd_p_fuse)
print(cd_p_wing)
print(cd_tot)
# Truth values
(cd_c_r, cd_i_r, cd_m_r, cd_m_fuse_base_r, cd_m_fuse_up_r, cd_m_nac_base_r, cd_m_ctrl_r, cd_p_fuse_r, cd_p_wing_r, cd_tot_r) = reg_values()
cd_c_r = cd_c_r.reshape(test_num,1)
cd_i_r = cd_i_r.reshape(test_num,1)
cd_m_r = cd_m_r.reshape(test_num,1)
cd_m_fuse_base_r = cd_m_fuse_base_r.reshape(test_num,1)
cd_m_fuse_up_r = cd_m_fuse_up_r.reshape(test_num,1)
cd_m_nac_base_r = cd_m_nac_base_r.reshape(test_num,1)
cd_m_ctrl_r = cd_m_ctrl_r.reshape(test_num,1)
cd_p_fuse_r = cd_p_fuse_r.reshape(test_num,1)
cd_p_wing_r = cd_p_wing_r.reshape(test_num,1)
cd_tot_r = cd_tot_r.reshape(test_num,1)
drag_tests = Data()
drag_tests.cd_c = np.abs((cd_c-cd_c_r)/cd_c)
drag_tests.cd_i = np.abs((cd_i-cd_i_r)/cd_i)
drag_tests.cd_m = np.abs((cd_m-cd_m_r)/cd_m)
# Line below is not normalized since regression values are 0, insert commented line if this changes
drag_tests.cd_m_fuse_base = np.abs((cd_m_fuse_base-cd_m_fuse_base_r)) # np.abs((cd_m_fuse_base-cd_m_fuse_base_r)/cd_m_fuse_base)
drag_tests.cd_m_fuse_up = np.abs((cd_m_fuse_up - cd_m_fuse_up_r)/cd_m_fuse_up)
drag_tests.cd_m_ctrl = np.abs((cd_m_ctrl - cd_m_ctrl_r)/cd_m_ctrl)
drag_tests.cd_p_fuse = np.abs((cd_p_fuse - cd_p_fuse_r)/cd_p_fuse)
drag_tests.cd_p_wing = np.abs((cd_p_wing - cd_p_wing_r)/cd_p_wing)
drag_tests.cd_tot = np.abs((cd_tot - cd_tot_r)/cd_tot)
print('\nCompute Drag Test Results\n')
#print drag_tests
for i, tests in list(drag_tests.items()):
assert(np.max(tests)<1e-4),'Supersonic Aero regression test failed at ' + i
if __name__ == '__main__':
#(conditions, configuration, geometry, test_num) = main()
main()
print('Supersonic Aero regression test passed!')
# --------------------------------------------------------------------
# Drag Polar
# --------------------------------------------------------------------
# --------------------------------------------------------------------
# Drag Polar
# --------------------------------------------------------------------
# initialize the vehicle
vehicle = vehicle_setup()
for wing in vehicle.wings:
wing.areas.wetted = 2.0 * wing.areas.reference
wing.areas.exposed = 0.8 * wing.areas.wetted
wing.areas.affected = 0.6 * wing.areas.wetted
# initalize the aero model
aerodynamics = SUAVE.Analyses.Aerodynamics.Supersonic_Zero()
aerodynamics.geometry = vehicle
## modify inviscid wings - linear model
#inviscid_wings = SUAVE.Analyses.Aerodynamics.Linear_Lift()
#inviscid_wings.settings.slope_correction_coefficient = 1.04
#inviscid_wings.settings.zero_lift_coefficient = 2.*np.pi* 3.1 * Units.deg
#aerodynamics.process.compute.lift.inviscid_wings = inviscid_wings
def setup():
nexus = Nexus()
problem = Data()
nexus.optimization_problem = problem
# -------------------------------------------------------------------
# Inputs
# -------------------------------------------------------------------
# [ tag , initial, [lb,ub], scaling, units ]
problem.inputs = np.array([
[ 'wing_area' , 124.8 , ( 70. , 200. ) , 124.8 , Units.meter**2],
[ 'wing_aspect_ratio' , 10.18, ( 5. , 20. ) , 10.18, Units.less],
[ 'wing_sweep' , 25. , ( 0. , 35. ) , 25. , Units.degrees],
[ 'wing_thickness' , 0.105 , ( 0.07 , 0.20 ) , 0.105, Units.less],
[ 'design_thrust' , 52700. , ( 10000. , 70000. ) , 52700. , Units.N],
[ 'MTOW' , 79090. , ( 20000. ,100000. ) , 79090. , Units.kg],
[ 'MZFW_ratio' , 0.77 , ( 0.6 , 0.99 ) , 0.77 , Units.less],
[ 'flap_takeoff_angle' , 10. , ( 0. , 20. ) , 10. , Units.degrees],
[ 'flap_landing_angle' , 40. , ( 0. , 50. ) , 40. , Units.degrees],
[ 'short_field_TOW' , 64030. , ( 20000. ,100000. ) , 64030. , Units.kg],
[ 'design_TOW' , 68520. , ( 20000. ,100000. ) , 68520. , Units.kg],
[ 'noise_takeoff_speed_increase' , 10.0 , ( 10. , 20. ) , 10.0 , Units.knots],
[ 'noise_cutback_altitude' , 304.8 , ( 240. , 400. ) , 304.8 , Units.meter],
])
# -------------------------------------------------------------------
# Objective
# -------------------------------------------------------------------
problem.objective = np.array([
[ 'noise_cumulative_margin', 17, Units.less ],
])
# -------------------------------------------------------------------
# Constraints
# -------------------------------------------------------------------
# [ tag, sense, edge, scaling, units ]
problem.constraints = np.array([
[ 'MZFW consistency' , '>' , 0. , 10 , Units.less],
[ 'design_range_fuel_margin' , '>', 0., 10, Units.less],
[ 'short_field_fuel_margin' , '>' , 0. , 10, Units.less],
[ 'max_range_fuel_margin' , '>' , 0. , 10, Units.less],
[ 'wing_span' , '<', 35.9664, 35.9664, Units.less],
[ 'noise_flyover_margin' , '>', 0. , 10., Units.less],
[ 'noise_sideline_margin' , '>', 0. , 10. , Units.less],
[ 'noise_approach_margin' , '>', 0., 10., Units.less],
[ 'takeoff_field_length' , '<', 1985., 1985., Units.meters],
[ 'landing_field_length' , '<', 1385., 1385., Units.meters],
[ '2nd_segment_climb_max_range' , '>', 0.024, 0.024, Units.less],
[ '2nd_segment_climb_short_field' , '>', 0.024, 0.024, Units.less],
[ 'max_throttle' , '<', 1., 1., Units.less],
[ 'short_takeoff_field_length' , '<', 1330., 1330., Units.meters],
[ 'noise_cumulative_margin' , '>', 10., 10., Units.less],
])
# -------------------------------------------------------------------
# Aliases
# -------------------------------------------------------------------
problem.aliases = [
[ 'wing_area' , ['vehicle_configurations.*.wings.main_wing.areas.reference',
'vehicle_configurations.*.reference_area' ]],
[ 'wing_aspect_ratio' , 'vehicle_configurations.*.wings.main_wing.aspect_ratio' ],
[ 'wing_incidence' , 'vehicle_configurations.*.wings.main_wing.twists.root' ],
[ 'wing_tip_twist' , 'vehicle_configurations.*.wings.main_wing.twists.tip' ],
[ 'wing_sweep' , 'vehicle_configurations.*.wings.main_wing.sweeps.quarter_chord' ],
[ 'wing_thickness' , 'vehicle_configurations.*.wings.main_wing.thickness_to_chord' ],
[ 'wing_taper' , 'vehicle_configurations.*.wings.main_wing.taper' ],
[ 'wing_location' , 'vehicle_configurations.*.wings.main_wing.origin[0]' ],
[ 'horizontal_tail_area' , 'vehicle_configurations.*.wings.horizontal_stabilizer.areas.reference'],
[ 'horizontal_tail_aspect_ratio' , 'vehicle_configurations.*.wings.horizontal_stabilizer.aspect_ratio' ],
[ 'vertical_tail_area' , 'vehicle_configurations.*.wings.vertical_stabilizer.areas.reference' ],
[ 'vertical_tail_aspect_ratio' , 'vehicle_configurations.*.wings.vertical_stabilizer.aspect_ratio' ],
[ 'design_thrust' , 'vehicle_configurations.*.propulsors.turbofan.thrust.total_design' ],
[ 'MTOW' , ['vehicle_configurations.*.mass_properties.takeoff' ,
'vehicle_configurations.*.mass_properties.max_takeoff' ]],
[ 'design_TOW' , 'vehicle_configurations.base.mass_properties.takeoff' ],
[ 'short_field_TOW' , 'vehicle_configurations.short_field_takeoff.mass_properties.takeoff' ],
[ 'flap_takeoff_angle' , ['vehicle_configurations.takeoff.wings.main_wing.flaps.angle',
'vehicle_configurations.short_field_takeoff.wings.main_wing.flaps.angle']],
[ 'flap_landing_angle' , 'vehicle_configurations.landing.wings.main_wing.flaps.angle' ],
[ 'slat_takeoff_angle' , ['vehicle_configurations.takeoff.wings.main_wing.slats.angle',
'vehicle_configurations.short_field_takeoff.wings.main_wing.slats.angle']],
[ 'slat_landing_angle' , 'vehicle_configurations.landing.wings.main_wing.slats.angle' ],
[ 'wing_span' , 'vehicle_configurations.base.wings.main_wing.spans.projected' ],
[ 'noise_approach_margin' , 'summary.noise_approach_margin' ],
[ 'noise_sideline_margin' , 'summary.noise_sideline_margin' ],
[ 'noise_flyover_margin' , 'summary.noise_flyover_margin' ],
[ 'static_stability' , 'summary.static_stability' ],
[ 'vertical_tail_volume_coefficient' , 'summary.vertical_tail_volume_coefficient' ],
[ 'horizontal_tail_volume_coefficient', 'summary.horizontal_tail_volume_coefficient' ],
[ 'wing_max_cl_norm' , 'summary.maximum_cl_norm' ],
[ 'design_range_fuel_margin' , 'summary.design_range_fuel_margin' ],
[ 'takeoff_field_length' , 'summary.takeoff_field_length' ],
[ 'landing_field_length' , 'summary.landing_field_length' ],
[ 'short_takeoff_field_length' , 'summary.short_takeoff_field_length' ],
[ '2nd_segment_climb_max_range' , 'summary.second_segment_climb_gradient_takeoff' ],
[ '2nd_segment_climb_short_field' , 'summary.second_segment_climb_gradient_short_field' ],
[ 'max_throttle' , 'summary.max_throttle' ],
[ 'short_field_fuel_margin' , 'summary.short_field_fuel_margin' ],
[ 'max_range_fuel_margin' , 'summary.max_range_fuel_margin' ],
[ 'max_range' , 'missions.max_range_distance' ],
[ 'MZFW consistency' , 'summary.MZFW_consistency' ],
[ 'MZFW_ratio' , 'MZFW_ratio' ],
[ 'noise_takeoff_speed_increase' , 'noise_V2_increase' ],
[ 'noise_cutback_altitude' , 'missions.takeoff.segments.climb.altitude_end' ],
[ 'noise_cumulative_margin' , 'summary.noise_margin' ],
[ 'weighted_sum_objective' , 'summary.weighted_sum_objective' ],
]
# -------------------------------------------------------------------
# Vehicles
# -------------------------------------------------------------------
vehicle = vehicle_setup()
nexus.vehicle_configurations = configs_setup(vehicle)
# -------------------------------------------------------------------
# Analyses
# -------------------------------------------------------------------
nexus.analyses = Analyses.setup(nexus.vehicle_configurations)
# -------------------------------------------------------------------
# Missions
# -------------------------------------------------------------------
nexus.missions = Missions.setup(nexus.analyses)
# -------------------------------------------------------------------
# Procedure
# -------------------------------------------------------------------
nexus.procedure = Procedure.setup()
# -------------------------------------------------------------------
# Summary
# -------------------------------------------------------------------
nexus.summary = Data()
return nexus
def main():
# vehicle data
vehicle = vehicle_setup()
configs = configs_setup(vehicle)
# vehicle analyses
configs_analyses = analyses_setup(configs)
# append AVL aerodynamic analysis
aerodynamics = SUAVE.Analyses.Aerodynamics.AVL()
aerodynamics.process.compute.lift.inviscid.regression_flag = True
aerodynamics.process.compute.lift.inviscid.keep_files = True
aerodynamics.geometry = copy.deepcopy(configs.cruise)
aerodynamics.process.compute.lift.inviscid.training_file = 'cruise_data_aerodynamics.txt'
configs_analyses.cruise.append(aerodynamics)
# append AVL stability analysis
stability = SUAVE.Analyses.Stability.AVL()
stability.regression_flag = True
stability.keep_files = True
stability.geometry = copy.deepcopy(configs.cruise)
stability.training_file = 'cruise_data_stability.txt'
configs_analyses.cruise.append(stability)
# mission analyses
mission = mission_setup(configs_analyses)
missions_analyses = missions_setup(mission)
analyses = SUAVE.Analyses.Analysis.Container()
analyses.configs = configs_analyses
analyses.missions = missions_analyses
simple_sizing(configs, analyses)
configs.finalize()
analyses.finalize()
# mission analysis
mission = analyses.missions.base
results = mission.evaluate()
# lift coefficient check
lift_coefficient = results.conditions.cruise.aerodynamics.lift_coefficient[0]
lift_coefficient_true = 0.59495841
print lift_coefficient
diff_CL = np.abs(lift_coefficient - lift_coefficient_true)
print 'CL difference'
print diff_CL
assert np.abs((lift_coefficient - lift_coefficient_true)/lift_coefficient_true) < 1e-3
# moment coefficient check
moment_coefficient = results.conditions.cruise.stability.static.CM[0][0]
moment_coefficient_true = -0.620326644
print moment_coefficient
diff_CM = np.abs(moment_coefficient - moment_coefficient_true)
print 'CM difference'
print diff_CM
assert np.abs((moment_coefficient - moment_coefficient_true)/moment_coefficient_true) < 1e-3
return
def main():
vehicle = vehicle_setup()
weight = Tube_Wing.empty(vehicle)
# regression values
actual = Data()
actual.payload = 27349.9081525 #includes cargo #17349.9081525 #without cargo
actual.pax = 15036.587065500002
actual.bag = 2313.3210870000003
actual.fuel = 12977.803363592691 #includes cargo #22177.6377131 #without cargo
actual.empty = 38688.08848390731
actual.wing = 6649.709658738429
actual.fuselage = 6642.061164271899
actual.propulsion = 6838.185174956626
actual.landing_gear = 3160.632
actual.systems = 13479.10479056802
actual.wt_furnish = 6431.80372889
actual.horizontal_tail = 1037.7414196819743
actual.vertical_tail = 629.0387683502595
actual.rudder = 251.61550734010382
# error calculations
error = Data()
error.payload = (actual.payload - weight.payload)/actual.payload
error.pax = (actual.pax - weight.pax)/actual.pax
error.bag = (actual.bag - weight.bag)/actual.bag
error.fuel = (actual.fuel - weight.fuel)/actual.fuel
error.empty = (actual.empty - weight.empty)/actual.empty
error.wing = (actual.wing - weight.wing)/actual.wing
error.fuselage = (actual.fuselage - weight.fuselage)/actual.fuselage
error.propulsion = (actual.propulsion - weight.propulsion)/actual.propulsion
error.landing_gear = (actual.landing_gear - weight.landing_gear)/actual.landing_gear
error.systems = (actual.systems - weight.systems)/actual.systems
error.wt_furnish = (actual.wt_furnish - weight.systems_breakdown.furnish)/actual.wt_furnish
error.horizontal_tail = (actual.horizontal_tail - weight.horizontal_tail)/actual.horizontal_tail
error.vertical_tail = (actual.vertical_tail - weight.vertical_tail)/actual.vertical_tail
error.rudder = (actual.rudder - weight.rudder)/actual.rudder
print('Results (kg)')
print(weight)
print('Relative Errors')
print(error)
for k,v in list(error.items()):
assert(np.abs(v)<1E-6)
#General Aviation weights; note that values are taken from Raymer,
#but there is a huge spread among the GA designs, so individual components
#differ a good deal from the actual design
vehicle = vehicle_setup_general_aviation()
GTOW = vehicle.mass_properties.max_takeoff
weight = General_Aviation.empty(vehicle)
weight.fuel = vehicle.fuel.mass_properties.mass
actual = Data()
actual.bag = 0.
actual.empty = 618.485310343
actual.fuel = 144.69596603
actual.wing = 124.673093906
actual.fuselage = 119.522072873
actual.propulsion = 194.477769922 #includes power plant and propeller, does not include fuel system
actual.landing_gear = 44.8033840543+5.27975390045
actual.furnishing = 37.8341395817
actual.electrical = 36.7532226254
actual.control_systems = 14.8331955546
actual.fuel_systems = 15.6859717453
actual.systems = 108.096549345
error = Data()
error.fuel = (actual.fuel - weight.fuel)/actual.fuel
error.empty = (actual.empty - weight.empty)/actual.empty
error.wing = (actual.wing - weight.wing)/actual.wing
error.fuselage = (actual.fuselage - weight.fuselage)/actual.fuselage
error.propulsion = (actual.propulsion - weight.propulsion)/actual.propulsion
error.landing_gear = (actual.landing_gear - (weight.landing_gear_main+weight.landing_gear_nose))/actual.landing_gear
error.furnishing = (actual.furnishing-weight.systems_breakdown.furnish)/actual.furnishing
error.electrical = (actual.electrical-weight.systems_breakdown.electrical)/actual.electrical
error.control_systems = (actual.control_systems-weight.systems_breakdown.control_systems)/actual.control_systems
error.fuel_systems = (actual.fuel_systems-weight.systems_breakdown.fuel_system)/actual.fuel_systems
error.systems = (actual.systems - weight.systems)/actual.systems
print('actual.systems=', actual.systems)
print('General Aviation Results (kg)')
print(weight)
print('Relative Errors')
print(error)
for k,v in list(error.items()):
assert(np.abs(v)<1e-6)
# BWB WEIGHTS
vehicle = bwb_setup()
weight = BWB.empty(vehicle)
# regression values
actual = Data()
actual.payload = 27349.9081525 #includes cargo #17349.9081525 #without cargo
actual.pax = 15036.587065500002
actual.bag = 2313.3210870000003
actual.fuel = 24860.343951919327
actual.empty = 26805.547895580676
actual.wing = 6576.679767012152
actual.fuselage = 1.0
actual.propulsion = 1413.8593105126783
actual.landing_gear = 3160.632
actual.systems = 15654.376818055844
actual.wt_furnish = 8205.349895589
# error calculations
error = Data()
error.payload = (actual.payload - weight.payload)/actual.payload
error.pax = (actual.pax - weight.pax)/actual.pax
error.bag = (actual.bag - weight.bag)/actual.bag
error.fuel = (actual.fuel - weight.fuel)/actual.fuel
error.empty = (actual.empty - weight.empty)/actual.empty
error.wing = (actual.wing - weight.wing)/actual.wing
error.fuselage = (actual.fuselage - (weight.fuselage+1.0))/actual.fuselage
error.propulsion = (actual.propulsion - weight.propulsion)/actual.propulsion
error.systems = (actual.systems - weight.systems)/actual.systems
error.wt_furnish = (actual.wt_furnish - weight.systems_breakdown.furnish)/actual.wt_furnish
print('Results (kg)')
print(weight)
print('Relative Errors')
print(error)
for k,v in list(error.items()):
assert(np.abs(v)<1E-6)
# Human Powered Aircraft
vehicle = hp_setup()
weight = HP.empty(vehicle)
# regression values
actual = Data()
actual.empty = 138.02737768459374
actual.wing = 89.86286881794777
actual.fuselage = 1.0
actual.horizontal_tail = 31.749272074174737
actual.vertical_tail = 16.415236792471237
# error calculations
error = Data()
error.empty = (actual.empty - weight.empty)/actual.empty
error.wing = (actual.wing - weight.wing)/actual.wing
error.fuselage = (actual.fuselage - (weight.fuselage+1.0))/actual.fuselage
error.horizontal_tail = (actual.horizontal_tail - weight.horizontal_tail)/actual.horizontal_tail
error.vertical_tail = (actual.vertical_tail - weight.vertical_tail)/actual.vertical_tail
print('Results (kg)')
print(weight)
print('Relative Errors')
print(error)
for k,v in list(error.items()):
assert(np.abs(v)<1E-6)
return
def main():
# initialize the vehicle
vehicle = vehicle_setup()
for wing in vehicle.wings:
wing.areas.wetted = 2.0 * wing.areas.reference
wing.areas.exposed = 0.8 * wing.areas.wetted
wing.areas.affected = 0.6 * wing.areas.wetted
# initalize the aero model
aerodynamics = SUAVE.Analyses.Aerodynamics.Fidelity_Zero()
aerodynamics.process.compute.lift.inviscid_wings = SUAVE.Analyses.Aerodynamics.Lifting_Line()
aerodynamics.geometry = vehicle
aerodynamics.initialize()
#no of test points
test_num = 11
#specify the angle of attack
angle_of_attacks = np.linspace(-.174,.174,test_num)[:,None] #* Units.deg
# Cruise conditions (except Mach number)
state = SUAVE.Analyses.Mission.Segments.Conditions.State()
state.conditions = SUAVE.Analyses.Mission.Segments.Conditions.Aerodynamics()
state.expand_rows(test_num)
# --------------------------------------------------------------------
# Initialize variables needed for CL and CD calculations
# Use a pre-run random order for values
# --------------------------------------------------------------------
Mc = np.array([[0.9 ],
[0.475],
[0.05 ],
[0.39 ],
[0.815],
[0.645],
[0.305],
[0.22 ],
[0.56 ],
[0.73 ],
[0.135]])
rho = np.array([[0.8],
[1. ],
[0.5],
[1.1],
[0.4],
[1.3],
[0.6],
[0.3],
[0.9],
[0.7],
[1.2]])
mu = np.array([[1.85e-05],
[1.55e-05],
[1.40e-05],
[1.10e-05],
[2.00e-05],
[8.00e-06],
[6.50e-06],
[9.50e-06],
[1.70e-05],
[1.25e-05],
[5.00e-06]])
T = np.array([[270.],
[250.],
[280.],
[260.],
[240.],
[200.],
[290.],
[230.],
[210.],
[300.],
[220.]])
pressure = np.array([[ 100000.],
[ 190000.],
[ 280000.],
[ 370000.],
[ 460000.],
[ 550000.],
[ 640000.],
[ 730000.],
[ 820000.],
[ 910000.],
[1000000.]])
re = np.array([[12819987.97468646],
[ 9713525.47464844],
[ 599012.59815633],
[12606549.94372309],
[ 5062187.10214493],
[29714816.00808047],
[ 9611290.40694227],
[ 2112171.68320523],
[ 8612638.72342302],
[14194381.78364854],
[ 9633881.90543247]])
air = Air()
a = air.compute_speed_of_sound(T,pressure)
re = rho*a*Mc/mu
state.conditions.freestream.mach_number = Mc
state.conditions.freestream.density = rho
state.conditions.freestream.dynamic_viscosity = mu
state.conditions.freestream.temperature = T
state.conditions.freestream.pressure = pressure
state.conditions.freestream.reynolds_number = re
state.conditions.aerodynamics.angle_of_attack = angle_of_attacks
# --------------------------------------------------------------------
# Surrogate
# --------------------------------------------------------------------
#call the aero model
results = aerodynamics.evaluate(state)
#build a polar for the markup aero
CL = results.lift.total
CD = results.drag.total
# --------------------------------------------------------------------
# Test compute Lift
# --------------------------------------------------------------------
#compute_aircraft_lift(conditions, configuration, geometry)
lift = state.conditions.aerodynamics.lift_coefficient
lift_r = np.array([-2.84689226, -1.06501674, -0.63426096, -0.35809118, -0.04487569,
0.36343181, 0.61055156, 0.90742419, 1.43504496, 2.18401103, 1.81298486])[:,None]
print('lift = ', lift)
lift_test = np.abs((lift-lift_r)/lift)
print('\nCompute Lift Test Results\n')
#print lift_test
assert(np.max(lift_test)<1e-4), 'Aero regression failed at compute lift test'
def main():
# vehicle data
vehicle = vehicle_setup()
configs = configs_setup(vehicle)
# vehicle analyses
configs_analyses = analyses_setup(configs)
run_new_regression = False
# append AVL aerodynamic analysis
aerodynamics = SUAVE.Analyses.Aerodynamics.AVL()
aerodynamics.settings.number_spanwise_vortices = 30
aerodynamics.settings.keep_files = True
aerodynamics.geometry = copy.deepcopy(configs.cruise)
configs_analyses.cruise.append(aerodynamics)
# append AVL stability analysis
stability = SUAVE.Analyses.Stability.AVL()
stability.settings.number_spanwise_vortices = 30
stability.settings.keep_files = True
stability.geometry = copy.deepcopy(configs.cruise)
if run_new_regression:
# append AVL aerodynamic analysis
aerodynamics.settings.regression_flag = False
aerodynamics.process.compute.lift.inviscid.settings.filenames.avl_bin_name = 'CHANGE/TO/AVL/PATH'
aerodynamics.settings.save_regression_results = True
stability.settings.regression_flag = False
stability.settings.save_regression_results = True
stability.settings.filenames.avl_bin_name = 'CHANGE/TO/AVL/PATH'
else:
aerodynamics.settings.regression_flag = True
aerodynamics.settings.save_regression_results = False
aerodynamics.settings.training_file = 'cruise_aero_data.txt'
stability.settings.regression_flag = True
stability.settings.save_regression_results = False
stability.training_file = 'cruise_stability_data.txt'
configs_analyses.cruise.append(aerodynamics)
configs_analyses.cruise.append(stability)
# ------------------------------------------------------------------
# Initialize the Mission
# ------------------------------------------------------------------
mission = SUAVE.Analyses.Mission.Sequential_Segments()
mission.tag = 'the_mission'
#airport
airport = SUAVE.Attributes.Airports.Airport()
airport.altitude = 0.0 * Units.ft
airport.delta_isa = 0.0
airport.atmosphere = SUAVE.Attributes.Atmospheres.Earth.US_Standard_1976()
mission.airport = airport
# unpack Segments module
Segments = SUAVE.Analyses.Mission.Segments
# base segment
base_segment = Segments.Segment()
# ------------------------------------------------------------------
# Cruise Segment: constant speed, constant altitude
# ------------------------------------------------------------------
segment = Segments.Cruise.Constant_Speed_Constant_Altitude(base_segment)
segment.tag = "cruise"
segment.analyses.extend(configs_analyses.cruise)
segment.air_speed = 230. * Units['m/s']
segment.distance = 4000. * Units.km
segment.altitude = 10.668 * Units.km
segment.state.numerics.number_control_points = 4
# add to mission
mission.append_segment(segment)
missions_analyses = missions_setup(mission)
analyses = SUAVE.Analyses.Analysis.Container()
analyses.configs = configs_analyses
analyses.missions = missions_analyses
simple_sizing(configs, analyses)
configs.finalize()
analyses.finalize()
# mission analysis
mission = analyses.missions.base
results = mission.evaluate()
# lift coefficient check
lift_coefficient = results.segments.cruise.conditions.aerodynamics.lift_coefficient[
0][0]
lift_coefficient_true = 0.6124936427552575
print(lift_coefficient)
diff_CL = np.abs(lift_coefficient - lift_coefficient_true)
print('CL difference')
print(diff_CL)
assert np.abs((lift_coefficient - lift_coefficient_true) /
lift_coefficient_true) < 1e-6
# moment coefficient check
moment_coefficient = results.segments.cruise.conditions.stability.static.CM[
0][0]
moment_coefficient_true = -0.5764235338199974
print(moment_coefficient)
diff_CM = np.abs(moment_coefficient - moment_coefficient_true)
print('CM difference')
print(diff_CM)
assert np.abs((moment_coefficient - moment_coefficient_true) /
moment_coefficient_true) < 1e-6
return
def main():
# initialize the vehicle
vehicle = vehicle_setup()
for wing in vehicle.wings:
wing.areas.wetted = 2.0 * wing.areas.reference
wing.areas.exposed = 0.8 * wing.areas.wetted
wing.areas.affected = 0.6 * wing.areas.wetted
# initalize the aero model
aerodynamics = SUAVE.Analyses.Aerodynamics.Fidelity_Zero()
aerodynamics.process.compute.lift.inviscid_wings = SUAVE.Analyses.Aerodynamics.Lifting_Line()
aerodynamics.geometry = vehicle
aerodynamics.initialize()
#no of test points
test_num = 11
#specify the angle of attack
angle_of_attacks = np.linspace(-.174,.174,test_num)[:,None] #* Units.deg
# Cruise conditions (except Mach number)
state = SUAVE.Analyses.Mission.Segments.Conditions.State()
state.conditions = SUAVE.Analyses.Mission.Segments.Conditions.Aerodynamics()
state.expand_rows(test_num)
# --------------------------------------------------------------------
# Initialize variables needed for CL and CD calculations
# Use a pre-run random order for values
# --------------------------------------------------------------------
Mc = np.array([[0.9 ],
[0.475],
[0.05 ],
[0.39 ],
[0.815],
[0.645],
[0.305],
[0.22 ],
[0.56 ],
[0.73 ],
[0.135]])
rho = np.array([[0.8],
[1. ],
[0.5],
[1.1],
[0.4],
[1.3],
[0.6],
[0.3],
[0.9],
[0.7],
[1.2]])
mu = np.array([[1.85e-05],
[1.55e-05],
[1.40e-05],
[1.10e-05],
[2.00e-05],
[8.00e-06],
[6.50e-06],
[9.50e-06],
[1.70e-05],
[1.25e-05],
[5.00e-06]])
T = np.array([[270.],
[250.],
[280.],
[260.],
[240.],
[200.],
[290.],
[230.],
[210.],
[300.],
[220.]])
pressure = np.array([[ 100000.],
[ 190000.],
[ 280000.],
[ 370000.],
[ 460000.],
[ 550000.],
[ 640000.],
[ 730000.],
[ 820000.],
[ 910000.],
[1000000.]])
re = np.array([[12819987.97468646],
[ 9713525.47464844],
[ 599012.59815633],
[12606549.94372309],
[ 5062187.10214493],
[29714816.00808047],
[ 9611290.40694227],
[ 2112171.68320523],
[ 8612638.72342302],
[14194381.78364854],
[ 9633881.90543247]])
air = Air()
a = air.compute_speed_of_sound(T,pressure)
re = rho*a*Mc/mu
state.conditions.freestream.mach_number = Mc
state.conditions.freestream.density = rho
state.conditions.freestream.dynamic_viscosity = mu
state.conditions.freestream.temperature = T
state.conditions.freestream.pressure = pressure
state.conditions.freestream.reynolds_number = re
state.conditions.aerodynamics.angle_of_attack = angle_of_attacks
# --------------------------------------------------------------------
# Surrogate
# --------------------------------------------------------------------
#call the aero model
results = aerodynamics.evaluate(state)
#build a polar for the markup aero
CL = results.lift.total
CD = results.drag.total
# --------------------------------------------------------------------
# Test compute Lift
# --------------------------------------------------------------------
#compute_aircraft_lift(conditions, configuration, geometry)
lift = state.conditions.aerodynamics.lift_coefficient
lift_r = np.array([-2.92026278,-1.13251873,-0.72118981,-0.48215461,
-0.28934217, 0.14217826, 0.40422306, 0.67788365,
1.13167945, 1.77614421, 1.50398138])[:,None]
print('lift = ', lift)
lift_test = np.abs((lift-lift_r)/lift)
print('\nCompute Lift Test Results\n')
#print lift_test
assert(np.max(lift_test)<1e-4), 'Aero regression failed at compute lift test'
if __name__ == '__main__':
#(conditions, configuration, geometry, test_num) = main()
main()
print 'Supersonic Aero regression test passed!'
# --------------------------------------------------------------------
# Drag Polar
# --------------------------------------------------------------------
# --------------------------------------------------------------------
# Drag Polar
# --------------------------------------------------------------------
# initialize the vehicle
vehicle = vehicle_setup()
for wing in vehicle.wings:
wing.areas.wetted = 2.0 * wing.areas.reference
wing.areas.exposed = 0.8 * wing.areas.wetted
wing.areas.affected = 0.6 * wing.areas.wetted
# initalize the aero model
aerodynamics = SUAVE.Analyses.Aerodynamics.Supersonic_Zero()
aerodynamics.geometry = vehicle
## modify inviscid wings - linear model
#inviscid_wings = SUAVE.Analyses.Aerodynamics.Linear_Lift()
#inviscid_wings.settings.slope_correction_coefficient = 1.04
#inviscid_wings.settings.zero_lift_coefficient = 2.*np.pi* 3.1 * Units.deg
#aerodynamics.process.compute.lift.inviscid_wings = inviscid_wings
def main():
vehicle = vehicle_setup()
weight = Transport.empty(vehicle)
# regression values
actual = Data()
actual.payload = 27349.9081525 # includes cargo #17349.9081525 #without cargo
actual.pax = 15036.587065500002
actual.bag = 2313.3210870000003
actual.fuel = 16504.32206450931 # includes cargo #22177.6377131 #without cargo
actual.empty = 35161.56978299069
actual.wing = 3461.869204335895
actual.fuselage = 6700.709511002648
actual.propulsion = 6838.185174956626
actual.landing_gear = 3160.632
actual.systems = 13390.723085494214
actual.wt_furnish = 6431.803728889001
actual.horizontal_tail = 728.7965315109458
actual.vertical_tail = 880.6542756903633
actual.rudder = 251.61550734010382
actual.nose_gear = 316.06320000000005
actual.main_gear = 2844.5688
# error calculations
error = Data()
error.payload = (actual.payload - weight.payload) / actual.payload
error.pax = (actual.pax - weight.pax) / actual.pax
error.bag = (actual.bag - weight.bag) / actual.bag
error.fuel = (actual.fuel - weight.fuel) / actual.fuel
error.empty = (actual.empty - weight.empty) / actual.empty
error.wing = (actual.wing - weight.wing) / actual.wing
error.fuselage = (actual.fuselage - weight.fuselage) / actual.fuselage
error.propulsion = (actual.propulsion -
weight.propulsion) / actual.propulsion
error.nose_gear = (actual.nose_gear - weight.nose_gear) / actual.nose_gear
error.main_gear = (actual.main_gear - weight.main_gear) / actual.main_gear
error.systems = (actual.systems - weight.systems) / actual.systems
error.wt_furnish = (actual.wt_furnish -
weight.systems_breakdown.furnish) / actual.wt_furnish
error.horizontal_tail = (actual.horizontal_tail -
weight.horizontal_tail) / actual.horizontal_tail
error.vertical_tail = (actual.vertical_tail -
weight.vertical_tail) / actual.vertical_tail
error.rudder = (actual.rudder - weight.rudder) / actual.rudder
print('Results (kg)')
print(weight)
print('Relative Errors')
print(error)
for k, v in list(error.items()):
assert (np.abs(v) < 1E-6)
#General Aviation weights; note that values are taken from Raymer,
#but there is a huge spread among the GA designs, so individual components
#differ a good deal from the actual design
vehicle = vehicle_setup_general_aviation()
GTOW = vehicle.mass_properties.max_takeoff
weight = General_Aviation.empty(vehicle)
weight.fuel = vehicle.fuel.mass_properties.mass
actual = Data()
actual.bag = 0.
actual.empty = 720.1834370409678
actual.fuel = 144.69596603
actual.wing = 152.25407206578896
actual.fuselage = 126.7421108234472
actual.propulsion = 224.40728553408732
actual.landing_gear = 67.81320006645151
actual.furnishing = 37.8341395817
actual.electrical = 41.28649399649684
actual.control_systems = 20.51671046011007
actual.fuel_systems = 20.173688786768366
actual.systems = 122.8010526627288
error = Data()
error.fuel = (actual.fuel - weight.fuel) / actual.fuel
error.empty = (actual.empty - weight.empty) / actual.empty
error.wing = (actual.wing - weight.wing) / actual.wing
error.fuselage = (actual.fuselage - weight.fuselage) / actual.fuselage
error.propulsion = (actual.propulsion -
weight.propulsion) / actual.propulsion
error.landing_gear = (actual.landing_gear -
(weight.landing_gear_main +
weight.landing_gear_nose)) / actual.landing_gear
error.furnishing = (actual.furnishing -
weight.systems_breakdown.furnish) / actual.furnishing
error.electrical = (actual.electrical - weight.systems_breakdown.electrical
) / actual.electrical
error.control_systems = (
actual.control_systems -
weight.systems_breakdown.control_systems) / actual.control_systems
error.fuel_systems = (
actual.fuel_systems -
weight.systems_breakdown.fuel_system) / actual.fuel_systems
error.systems = (actual.systems - weight.systems) / actual.systems
print('actual.systems=', actual.systems)
print('General Aviation Results (kg)')
print(weight)
print('Relative Errors')
print(error)
for k, v in list(error.items()):
assert (np.abs(v) < 1e-6)
# BWB WEIGHTS
vehicle = bwb_setup()
weight = BWB.empty(vehicle)
# regression values
actual = Data()
actual.payload = 27349.9081525 #includes cargo #17349.9081525 #without cargo
actual.pax = 15036.587065500002
actual.bag = 2313.3210870000003
actual.fuel = 26119.117465169547
actual.empty = 25546.774382330455
actual.wing = 5317.906253761935
actual.fuselage = 1.0
actual.propulsion = 1413.8593105126783
actual.landing_gear = 3160.632
actual.systems = 15654.376818055844
actual.wt_furnish = 8205.349895589
# error calculations
error = Data()
error.payload = (actual.payload - weight.payload) / actual.payload
error.pax = (actual.pax - weight.pax) / actual.pax
error.bag = (actual.bag - weight.bag) / actual.bag
error.fuel = (actual.fuel - weight.fuel) / actual.fuel
error.empty = (actual.empty - weight.empty) / actual.empty
error.wing = (actual.wing - weight.wing) / actual.wing
error.fuselage = (actual.fuselage -
(weight.fuselage + 1.0)) / actual.fuselage
error.propulsion = (actual.propulsion -
weight.propulsion) / actual.propulsion
error.systems = (actual.systems - weight.systems) / actual.systems
error.wt_furnish = (actual.wt_furnish -
weight.systems_breakdown.furnish) / actual.wt_furnish
print('Results (kg)')
print(weight)
print('Relative Errors')
print(error)
for k, v in list(error.items()):
assert (np.abs(v) < 1E-6)
# Human Powered Aircraft
vehicle = hp_setup()
weight = HP.empty(vehicle)
# regression values
actual = Data()
actual.empty = 143.59737768459374
actual.wing = 95.43286881794776
actual.fuselage = 1.0
actual.horizontal_tail = 31.749272074174737
actual.vertical_tail = 16.415236792471237
# error calculations
error = Data()
error.empty = (actual.empty - weight.empty) / actual.empty
error.wing = (actual.wing - weight.wing) / actual.wing
error.fuselage = (actual.fuselage -
(weight.fuselage + 1.0)) / actual.fuselage
error.horizontal_tail = (actual.horizontal_tail -
weight.horizontal_tail) / actual.horizontal_tail
error.vertical_tail = (actual.vertical_tail -
weight.vertical_tail) / actual.vertical_tail
print('Results (kg)')
print(weight)
print('Relative Errors')
print(error)
for k, v in list(error.items()):
assert (np.abs(v) < 1E-6)
return
def main():
# Transport Weights
vehicle = vehicle_setup()
method_types = [
'SUAVE', 'New SUAVE', 'FLOPS Simple', 'FLOPS Complex', 'Raymer'
]
for method_type in method_types:
print('Testing Method: ' + method_type)
if 'FLOPS' in method_type:
settings = Data()
settings.FLOPS = Data()
settings.FLOPS.aeroelastic_tailoring_factor = 0.
settings.FLOPS.strut_braced_wing_factor = 0.
settings.FLOPS.composite_utilization_factor = 0.5
settings.FLOPS.variable_sweep_factor = 1.
elif 'Raymer' in method_type:
settings = Data()
settings.Raymer = Data()
settings.Raymer.fuselage_mounted_landing_gear_factor = 1.
else:
settings = None
weight = Common.empty_weight(vehicle,
settings=settings,
method_type=method_type)
#save_results(weight, 'weights_'+method_type.replace(' ','_')+'.res')
old_weight = load_results('weights_' + method_type.replace(' ', '_') +
'.res')
check_list = [
'payload_breakdown.total', 'payload_breakdown.passengers',
'payload_breakdown.baggage', 'structures.wing',
'structures.fuselage', 'propulsion_breakdown.total',
'structures.nose_landing_gear', 'structures.main_landing_gear',
'systems_breakdown.total', 'systems_breakdown.furnish',
'structures.horizontal_tail', 'structures.vertical_tail', 'empty',
'fuel'
]
# do the check
for k in check_list:
print(k)
old_val = old_weight.deep_get(k)
new_val = weight.deep_get(k)
err = (new_val - old_val) / old_val
print('Error:', err)
assert np.abs(err) < 1e-6, 'Check Failed : %s' % k
print('')
#General Aviation weights; note that values are taken from Raymer,
#but there is a huge spread among the GA designs, so individual components
#differ a good deal from the actual design
vehicle = vehicle_setup_general_aviation()
weight = General_Aviation.empty(vehicle)
weight.fuel = vehicle.fuel.mass_properties.mass
actual = Data()
actual.bag = 0.
actual.empty = 700.0097482541994
actual.fuel = 48.417662245800784
actual.wing = 152.25407206578896
actual.fuselage = 126.7421108234472
actual.propulsion = 224.40728553408732
actual.landing_gear = 67.81320006645151
actual.furnishing = 37.8341395817
actual.electrical = 41.28649399649684
actual.control_systems = 20.51671046011007
actual.fuel_systems = 20.173688786768366
actual.systems = 102.62736387596043
error = Data()
error.fuel = (actual.fuel - weight.fuel) / actual.fuel
error.empty = (actual.empty - weight.empty) / actual.empty
error.wing = (actual.wing - weight.structures.wing) / actual.wing
error.fuselage = (actual.fuselage -
weight.structures.fuselage) / actual.fuselage
error.propulsion = (actual.propulsion -
weight.propulsion_breakdown.total) / actual.propulsion
error.landing_gear = (
actual.landing_gear -
(weight.structures.main_landing_gear +
weight.structures.nose_landing_gear)) / actual.landing_gear
error.furnishing = (actual.furnishing -
weight.systems_breakdown.furnish) / actual.furnishing
error.electrical = (actual.electrical - weight.systems_breakdown.electrical
) / actual.electrical
error.control_systems = (
actual.control_systems -
weight.systems_breakdown.control_systems) / actual.control_systems
error.fuel_systems = (
actual.fuel_systems -
weight.propulsion_breakdown.fuel_system) / actual.fuel_systems
error.systems = (actual.systems -
weight.systems_breakdown.total) / actual.systems
print('Results (kg)')
print(weight)
print('Relative Errors')
print(error)
for k, v in error.items():
assert (np.abs(v) < 1E-6)
# BWB WEIGHTS
vehicle = bwb_setup()
weight = BWB.empty(vehicle)
# regression values
actual = Data()
actual.payload = 27349.9081525 #includes cargo #17349.9081525 #without cargo
actual.pax = 15036.587065500002
actual.bag = 2313.3210870000003
actual.fuel = 23361.42500371662
actual.empty = 24417.180232883387
actual.wing = 7272.740220314861
actual.fuselage = 1.0
actual.propulsion = 1413.8593105126783
actual.landing_gear = 3160.632
actual.systems = 12569.948702055846
actual.wt_furnish = 8205.349895589
# error calculations
error = Data()
error.payload = (actual.payload -
weight.payload_breakdown.total) / actual.payload
error.pax = (actual.pax - weight.payload_breakdown.passengers) / actual.pax
error.bag = (actual.bag - weight.payload_breakdown.baggage) / actual.bag
error.fuel = (actual.fuel - weight.fuel) / actual.fuel
error.empty = (actual.empty - weight.empty) / actual.empty
error.wing = (actual.wing - weight.structures.wing) / actual.wing
error.fuselage = (actual.fuselage -
(weight.structures.fuselage + 1.0)) / actual.fuselage
error.propulsion = (actual.propulsion -
weight.propulsion_breakdown.total) / actual.propulsion
error.systems = (actual.systems -
weight.systems_breakdown.total) / actual.systems
error.wt_furnish = (actual.wt_furnish -
weight.systems_breakdown.furnish) / actual.wt_furnish
print('Results (kg)')
print(weight)
print('Relative Errors')
print(error)
for k, v in error.items():
assert (np.abs(v) < 1E-6)
# Human Powered Aircraft
vehicle = hp_setup()
weight = HP.empty(vehicle)
# regression values
actual = Data()
actual.empty = 143.59737768459374
actual.wing = 95.43286881794776
actual.fuselage = 1.0
actual.horizontal_tail = 31.749272074174737
actual.vertical_tail = 16.415236792471237
# error calculations
error = Data()
error.empty = (actual.empty - weight.empty) / actual.empty
error.wing = (actual.wing - weight.wing) / actual.wing
error.fuselage = (actual.fuselage -
(weight.fuselage + 1.0)) / actual.fuselage
error.horizontal_tail = (actual.horizontal_tail -
weight.horizontal_tail) / actual.horizontal_tail
error.vertical_tail = (actual.vertical_tail -
weight.vertical_tail) / actual.vertical_tail
print('Results (kg)')
print(weight)
print('Relative Errors')
print(error)
for k, v in error.items():
assert (np.abs(v) < 1E-6)
return
def main():
# initialize the vehicle
vehicle = vehicle_setup()
for wing in vehicle.wings:
wing.areas.wetted = 2.0 * wing.areas.reference
wing.areas.exposed = 0.8 * wing.areas.wetted
wing.areas.affected = 0.6 * wing.areas.wetted
# initalize the aero model
aerodynamics = SUAVE.Analyses.Aerodynamics.Fidelity_Zero()
aerodynamics.geometry = vehicle
aerodynamics.initialize()
#no of test points
test_num = 11
#specify the angle of attack
angle_of_attacks = np.linspace(-.174, .174, test_num)[:,
None] #* Units.deg
# Cruise conditions (except Mach number)
state = SUAVE.Analyses.Mission.Segments.Conditions.State()
state.conditions = SUAVE.Analyses.Mission.Segments.Conditions.Aerodynamics(
)
state.expand_rows(test_num)
# --------------------------------------------------------------------
# Initialize variables needed for CL and CD calculations
# Use a seeded random order for values
# --------------------------------------------------------------------
random.seed(1)
Mc = np.linspace(0.05, 0.9, test_num)
random.shuffle(Mc)
rho = np.linspace(0.3, 1.3, test_num)
random.shuffle(rho)
mu = np.linspace(5 * 10**-6, 20 * 10**-6, test_num)
random.shuffle(mu)
T = np.linspace(200, 300, test_num)
random.shuffle(T)
pressure = np.linspace(10**5, 10**6, test_num)
# Changed after to preserve seed for initial testing
Mc = Mc[:, None]
rho = rho[:, None]
mu = mu[:, None]
T = T[:, None]
pressure = pressure[:, None]
air = Air()
a = air.compute_speed_of_sound(T, pressure)
re = rho * a * Mc / mu
state.conditions.freestream.mach_number = Mc
state.conditions.freestream.density = rho
state.conditions.freestream.dynamic_viscosity = mu
state.conditions.freestream.temperature = T
state.conditions.freestream.pressure = pressure
state.conditions.freestream.reynolds_number = re
state.conditions.aerodynamics.angle_of_attack = angle_of_attacks
# --------------------------------------------------------------------
# Surrogate
# --------------------------------------------------------------------
#call the aero model
results = aerodynamics.evaluate(state)
#build a polar for the markup aero
polar = Data()
CL = results.lift.total
CD = results.drag.total
polar.lift = CL
polar.drag = CD
# --------------------------------------------------------------------
# Test compute Lift
# --------------------------------------------------------------------
#compute_aircraft_lift(conditions, configuration, geometry)
lift = state.conditions.aerodynamics.lift_coefficient
lift_r = np.array([
-2.17753919,
-0.7768714,
-0.41862788,
-0.16569318,
0.1949377,
0.49528782,
0.67624325,
0.93239723,
1.41834084,
2.1078681,
1.72191103,
])[:, None]
print 'lift = ', lift
lift_test = np.abs((lift - lift_r) / lift)
print '\nCompute Lift Test Results\n'
#print lift_test
assert (np.max(lift_test) <
1e-4), 'Aero regression failed at compute lift test'
# --------------------------------------------------------------------
# Test compute drag
# --------------------------------------------------------------------
#compute_aircraft_drag(conditions, configuration, geometry)
# Pull calculated values
drag_breakdown = state.conditions.aerodynamics.drag_breakdown
# Only one wing is evaluated since they rely on the same function
cd_c = drag_breakdown.compressible['main_wing'].compressibility_drag
cd_i = drag_breakdown.induced.total
cd_m = drag_breakdown.miscellaneous.total
# cd_m_fuse_base = drag_breakdown.miscellaneous.fuselage_base
# cd_m_fuse_up = drag_breakdown.miscellaneous.fuselage_upsweep
# cd_m_nac_base = drag_breakdown.miscellaneous.nacelle_base['turbofan']
# cd_m_ctrl = drag_breakdown.miscellaneous.control_gaps
cd_p_fuse = drag_breakdown.parasite['fuselage'].parasite_drag_coefficient
cd_p_wing = drag_breakdown.parasite['main_wing'].parasite_drag_coefficient
cd_tot = drag_breakdown.total
print 'cd_m =', cd_m
(cd_c_r, cd_i_r, cd_m_r, cd_m_fuse_base_r, cd_m_fuse_up_r, cd_m_nac_base_r,
cd_m_ctrl_r, cd_p_fuse_r, cd_p_wing_r, cd_tot_r) = reg_values()
drag_tests = Data()
drag_tests.cd_c = np.abs((cd_c - cd_c_r) / cd_c)
drag_tests.cd_i = np.abs((cd_i - cd_i_r) / cd_i)
drag_tests.cd_m = np.abs((cd_m - cd_m_r) / cd_m)
## Commented lines represent values not set by current drag functions, but to be recreated in the future
# Line below is not normalized since regression values are 0, insert commented line if this changes
# drag_tests.cd_m_fuse_base = np.abs((cd_m_fuse_base-cd_m_fuse_base_r)) # np.abs((cd_m_fuse_base-cd_m_fuse_base_r)/cd_m_fuse_base)
# drag_tests.cd_m_fuse_up = np.abs((cd_m_fuse_up - cd_m_fuse_up_r)/cd_m_fuse_up)
# drag_tests.cd_m_ctrl = np.abs((cd_m_ctrl - cd_m_ctrl_r)/cd_m_ctrl)
drag_tests.cd_p_fuse = np.abs((cd_p_fuse - cd_p_fuse_r) / cd_p_fuse)
drag_tests.cd_p_wing = np.abs((cd_p_wing - cd_p_wing_r) / cd_p_wing)
drag_tests.cd_tot = np.abs((cd_tot - cd_tot_r) / cd_tot)
print '\nCompute Drag Test Results\n'
print 'cd_tot=', cd_tot
for i, tests in drag_tests.items():
assert (np.max(tests) < 1e-4), 'Aero regression test failed at ' + i
def main():
vehicle = vehicle_setup()
weight = Tube_Wing.empty(vehicle)
# regression values
actual = Data()
actual.payload = 27349.9081525 #includes cargo #17349.9081525 #without cargo
actual.pax = 15036.587065500002
actual.bag = 2313.3210870000003
actual.fuel = 12990.957450008464 #includes cargo #22177.6377131 #without cargo
actual.empty = 38674.934397491539
actual.wing = 6649.7096587384294
actual.fuselage = 6642.0611642718986
actual.propulsion = 6838.1851749566231
actual.landing_gear = 3160.632
actual.systems = 13479.10479056802
actual.wt_furnish = 6431.80372889
actual.horizontal_tail = 1024.5873332662029
actual.vertical_tail = 629.03876835025949
actual.rudder = 251.61550734010382
# error calculations
error = Data()
error.payload = (actual.payload - weight.payload) / actual.payload
error.pax = (actual.pax - weight.pax) / actual.pax
error.bag = (actual.bag - weight.bag) / actual.bag
error.fuel = (actual.fuel - weight.fuel) / actual.fuel
error.empty = (actual.empty - weight.empty) / actual.empty
error.wing = (actual.wing - weight.wing) / actual.wing
error.fuselage = (actual.fuselage - weight.fuselage) / actual.fuselage
error.propulsion = (actual.propulsion -
weight.propulsion) / actual.propulsion
error.landing_gear = (actual.landing_gear -
weight.landing_gear) / actual.landing_gear
error.systems = (actual.systems - weight.systems) / actual.systems
error.wt_furnish = (actual.wt_furnish -
weight.systems_breakdown.furnish) / actual.wt_furnish
error.horizontal_tail = (actual.horizontal_tail -
weight.horizontal_tail) / actual.horizontal_tail
error.vertical_tail = (actual.vertical_tail -
weight.vertical_tail) / actual.vertical_tail
error.rudder = (actual.rudder - weight.rudder) / actual.rudder
print 'Results (kg)'
print weight
print 'Relative Errors'
print error
for k, v in error.items():
assert (np.abs(v) < 1E-6)
#General Aviation weights; note that values are taken from Raymer,
#but there is a huge spread among the GA designs, so individual components
#differ a good deal from the actual design
vehicle = vehicle_setup_general_aviation()
GTOW = vehicle.mass_properties.max_takeoff
weight = General_Aviation.empty(vehicle)
weight.fuel = vehicle.fuel.mass_properties.mass
actual = Data()
actual.bag = 0.
actual.empty = 618.485310343
actual.fuel = 144.69596603
actual.wing = 124.673093906
actual.fuselage = 119.522072873
actual.propulsion = 194.477769922 #includes power plant and propeller, does not include fuel system
actual.landing_gear = 44.8033840543 + 5.27975390045
actual.furnishing = 37.8341395817
actual.electrical = 36.7532226254
actual.control_systems = 14.8331955546
actual.fuel_systems = 15.6859717453
actual.systems = 108.096549345
error = Data()
error.fuel = (actual.fuel - weight.fuel) / actual.fuel
error.empty = (actual.empty - weight.empty) / actual.empty
error.wing = (actual.wing - weight.wing) / actual.wing
error.fuselage = (actual.fuselage - weight.fuselage) / actual.fuselage
error.propulsion = (actual.propulsion -
weight.propulsion) / actual.propulsion
error.landing_gear = (actual.landing_gear -
(weight.landing_gear_main +
weight.landing_gear_nose)) / actual.landing_gear
error.furnishing = (actual.furnishing -
weight.systems_breakdown.furnish) / actual.furnishing
error.electrical = (actual.electrical - weight.systems_breakdown.electrical
) / actual.electrical
error.control_systems = (
actual.control_systems -
weight.systems_breakdown.control_systems) / actual.control_systems
error.fuel_systems = (
actual.fuel_systems -
weight.systems_breakdown.fuel_system) / actual.fuel_systems
error.systems = (actual.systems - weight.systems) / actual.systems
print 'actual.systems=', actual.systems
print 'General Aviation Results (kg)'
print weight
print 'Relative Errors'
print error
for k, v in error.items():
assert (np.abs(v) < 1e-6)
return
