def main():
vehicle = mission_B737.vehicle_setup()
configs = mission_B737.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 = mission.target_landing_weight - results.conditions.weights.total_mass[-1,0]
print 'landing weight error' , error
assert error < 1.
return
def main():
vehicle = mission_B737.vehicle_setup()
configs = mission_B737.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 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]
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 = 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([-2.42489437, -0.90696416, -0.53991953, -0.3044834 , -0.03710598,
0.31061936 , 0.52106899, 0.77407765, 1.22389024, 1.86240501,
1.54587835])[:,None]
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['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
(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 drag_tests
for i, tests in drag_tests.items():
assert(np.max(tests)<1e-4),'Aero regression test failed at ' + i
if __name__ == '__main__':
#(conditions, configuration, geometry, test_num) = main()
main()
print '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.Fidelity_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():
# 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]
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 = 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([
-2.42489437, -0.90696416, -0.53991953, -0.3044834, -0.03710598,
0.31061936, 0.52106899, 0.77407765, 1.22389024, 1.86240501, 1.54587835
])[:, None]
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['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
(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 drag_tests
for i, tests in drag_tests.items():
assert (np.max(tests) < 1e-4), 'Aero regression test failed at ' + i
if __name__ == '__main__':
#(conditions, configuration, geometry, test_num) = main()
main()
print '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.Fidelity_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