def setup(self):
nn = self.options['num_nodes']
flight_phase = self.options['flight_phase']
# a propulsion system needs to be defined in order to provide thrust
# information for the mission analysis code
# propulsion_promotes_outputs = ['fuel_flow', 'thrust']
propulsion_promotes_inputs = ["fltcond|*", "throttle"]
self.add_subsystem('propmodel', CFM56(num_nodes=nn, plot=False),
promotes_inputs=propulsion_promotes_inputs)
doubler = om.ExecComp(['thrust=2*thrust_in', 'fuel_flow=2*fuel_flow_in'],
thrust_in={'val': 1.0*np.ones((nn,)),
'units': 'kN'},
thrust={'val': 1.0*np.ones((nn,)),
'units': 'kN'},
fuel_flow={'val': 1.0*np.ones((nn,)),
'units': 'kg/s',
'tags': ['integrate', 'state_name:fuel_used', 'state_units:kg', 'state_val:1.0', 'state_promotes:True']},
fuel_flow_in={'val': 1.0*np.ones((nn,)),
'units': 'kg/s'})
self.add_subsystem('doubler', doubler, promotes_outputs=['*'])
self.connect('propmodel.thrust', 'doubler.thrust_in')
self.connect('propmodel.fuel_flow', 'doubler.fuel_flow_in')
# use a different drag coefficient for takeoff versus cruise
if flight_phase not in ['v0v1', 'v1v0', 'v1vr', 'rotate']:
cd0_source = 'ac|aero|polar|CD0_cruise'
else:
cd0_source = 'ac|aero|polar|CD0_TO'
self.add_subsystem('drag', PolarDrag(num_nodes=nn),
promotes_inputs=['fltcond|CL', 'ac|geom|*', ('CD0', cd0_source),
'fltcond|q', ('e', 'ac|aero|polar|e')],
promotes_outputs=['drag'])
# generally the weights module will be custom to each airplane
passthru = om.ExecComp('OEW=x',
x={'val': 1.0,
'units': 'kg'},
OEW={'val': 1.0,
'units': 'kg'})
self.add_subsystem('OEW', passthru,
promotes_inputs=[('x', 'ac|weights|OEW')],
promotes_outputs=['OEW'])
self.add_subsystem('weight', oc.AddSubtractComp(output_name='weight',
input_names=['ac|weights|MTOW', 'fuel_used'],
units='kg', vec_size=[1, nn],
scaling_factors=[1, -1]),
promotes_inputs=['*'],
promotes_outputs=['weight'])
def test_defaults(self):
p = Problem()
p.model = CFM56()
p.setup(force_alloc_complex=True)
p.set_val('throttle', 0.5)
p.set_val('fltcond|h', 10e3, units='ft')
p.set_val('fltcond|M', 0.5)
p.run_model()
assert_near_equal(p.get_val('thrust', units='lbf'), 7050.73840869*np.ones(1), tolerance=1e-6)
assert_near_equal(p.get_val('fuel_flow', units='kg/s'), 0.50273824*np.ones(1), tolerance=1e-6)
assert_near_equal(p.get_val('T4', units='degK'), 1432.06813946*np.ones(1), tolerance=1e-6)
partials = p.check_partials(method='cs',compact_print=True)
assert_check_partials(partials)
def test_vectorized(self):
nn = 5
p = Problem()
p.model = CFM56(num_nodes=nn)
p.setup(force_alloc_complex=True)
p.set_val('throttle', np.linspace(0.0001, 1., nn))
p.set_val('fltcond|h', np.linspace(0, 40e3, nn), units='ft')
p.set_val('fltcond|M', np.linspace(0.1, 0.9, nn))
p.run_model()
assert_near_equal(p.get_val('thrust', units='lbf'), np.array([1445.41349482, 3961.46624224,
5278.43191982, 5441.44404298, 6479.00525867]), tolerance=5e-3)
assert_near_equal(p.get_val('fuel_flow', units='kg/s'), np.array([0.17032429, 0.25496437,
0.35745638, 0.40572545, 0.4924194]), tolerance=5e-3)
assert_near_equal(p.get_val('T4', units='degK'), np.array([1005.38911171, 1207.57548728,
1381.94820904, 1508.07901676, 1665.37063872]), tolerance=5e-3)
partials = p.check_partials(method='cs',compact_print=True)
assert_check_partials(partials)
def setup(self):
nn = self.options['num_nodes']
flight_phase = self.options['flight_phase']
# a propulsion system needs to be defined in order to provide thrust
# information for the mission analysis code
propulsion_promotes_inputs = ["fltcond|*", "throttle"]
self.add_subsystem('propmodel',
CFM56(num_nodes=nn, plot=False),
promotes_inputs=propulsion_promotes_inputs)
doubler = om.ExecComp(
['thrust=2*thrust_in', 'fuel_flow=2*fuel_flow_in'],
thrust_in={
'val': 1.0 * np.ones((nn, )),
'units': 'kN'
},
thrust={
'val': 1.0 * np.ones((nn, )),
'units': 'kN'
},
fuel_flow={
'val':
1.0 * np.ones((nn, )),
'units':
'kg/s',
'tags': [
'integrate', 'state_name:fuel_used', 'state_units:kg',
'state_val:1.0', 'state_promotes:True'
]
},
fuel_flow_in={
'val': 1.0 * np.ones((nn, )),
'units': 'kg/s'
})
self.add_subsystem('doubler', doubler, promotes_outputs=['*'])
self.connect('propmodel.thrust', 'doubler.thrust_in')
self.connect('propmodel.fuel_flow', 'doubler.fuel_flow_in')
oas_surf_dict = {} # options for OpenAeroStruct
# Grid size and number of spline control points (must be same as B738AnalysisGroup)
global NUM_X, NUM_Y, NUM_TWIST, NUM_TOVERC, NUM_SKIN, NUM_SPAR, USE_SURROGATE
if USE_SURROGATE:
self.add_subsystem(
'drag',
OASAerostructDragPolar(num_nodes=nn,
num_x=NUM_X,
num_y=NUM_Y,
num_twist=NUM_TWIST,
num_toverc=NUM_TOVERC,
num_skin=NUM_SKIN,
num_spar=NUM_SPAR,
surf_options=oas_surf_dict),
promotes_inputs=[
'fltcond|CL', 'fltcond|M', 'fltcond|h', 'fltcond|q',
'ac|geom|wing|S_ref', 'ac|geom|wing|AR',
'ac|geom|wing|taper', 'ac|geom|wing|c4sweep',
'ac|geom|wing|twist', 'ac|geom|wing|toverc',
'ac|geom|wing|skin_thickness',
'ac|geom|wing|spar_thickness', 'ac|aero|CD_nonwing'
],
promotes_outputs=[
'drag', 'ac|weights|W_wing',
('failure', 'ac|struct|failure')
])
else:
self.add_subsystem(
'drag',
OASAerostructDragPolarExact(num_nodes=nn,
num_x=NUM_X,
num_y=NUM_Y,
num_twist=NUM_TWIST,
num_toverc=NUM_TOVERC,
num_skin=NUM_SKIN,
num_spar=NUM_SPAR,
surf_options=oas_surf_dict),
promotes_inputs=[
'fltcond|CL', 'fltcond|M', 'fltcond|h', 'fltcond|q',
'ac|geom|wing|S_ref', 'ac|geom|wing|AR',
'ac|geom|wing|taper', 'ac|geom|wing|c4sweep',
'ac|geom|wing|twist', 'ac|geom|wing|toverc',
'ac|geom|wing|skin_thickness',
'ac|geom|wing|spar_thickness', 'ac|aero|CD_nonwing'
],
promotes_outputs=[
'drag', 'ac|weights|W_wing',
('failure', 'ac|struct|failure')
])
# generally the weights module will be custom to each airplane
passthru = om.ExecComp('OEW=x',
x={
'val': 1.0,
'units': 'kg'
},
OEW={
'val': 1.0,
'units': 'kg'
})
self.add_subsystem('OEW',
passthru,
promotes_inputs=[('x', 'ac|weights|OEW')],
promotes_outputs=['OEW'])
# Use Raymer as estimate for 737 original wing weight, subtract it
# out, then add in OpenAeroStruct wing weight estimate
self.add_subsystem('weight',
oc.AddSubtractComp(output_name='weight',
input_names=[
'ac|weights|MTOW',
'fuel_used',
'ac|weights|orig_W_wing',
'ac|weights|W_wing'
],
units='kg',
vec_size=[1, nn, 1, 1],
scaling_factors=[1, -1, -1, 1]),
promotes_inputs=['*'],
promotes_outputs=['weight'])
def setup(self):
nn = self.options['num_nodes']
flight_phase = self.options['flight_phase']
# a propulsion system needs to be defined in order to provide thrust
# information for the mission analysis code
propulsion_promotes_inputs = ["fltcond|*", "throttle"]
self.add_subsystem('propmodel',
CFM56(num_nodes=nn, plot=False),
promotes_inputs=propulsion_promotes_inputs)
doubler = om.ExecComp(
['thrust=2*thrust_in', 'fuel_flow=2*fuel_flow_in'],
thrust_in={
'val': 1.0 * np.ones((nn, )),
'units': 'kN'
},
thrust={
'val': 1.0 * np.ones((nn, )),
'units': 'kN'
},
fuel_flow={
'val':
1.0 * np.ones((nn, )),
'units':
'kg/s',
'tags': [
'integrate', 'state_name:fuel_used', 'state_units:kg',
'state_val:1.0', 'state_promotes:True'
]
},
fuel_flow_in={
'val': 1.0 * np.ones((nn, )),
'units': 'kg/s'
})
self.add_subsystem('doubler', doubler, promotes_outputs=['*'])
self.connect('propmodel.thrust', 'doubler.thrust_in')
self.connect('propmodel.fuel_flow', 'doubler.fuel_flow_in')
# use a different drag coefficient for takeoff versus cruise
oas_surf_dict = {} # options for OpenAeroStruct
oas_surf_dict['t_over_c'] = acdata['ac']['geom']['wing']['toverc'][
'value']
self.add_subsystem('drag',
OASDragPolar(num_nodes=nn,
num_x=3,
num_y=7,
num_twist=3,
surf_options=oas_surf_dict),
promotes_inputs=[
'fltcond|CL', 'fltcond|M', 'fltcond|h',
'fltcond|q', 'ac|geom|wing|S_ref',
'ac|geom|wing|AR', 'ac|geom|wing|taper',
'ac|geom|wing|c4sweep', 'ac|geom|wing|twist',
'ac|aero|CD_nonwing'
],
promotes_outputs=['drag'])
self.set_input_defaults(
'ac|aero|CD_nonwing',
0.0145) # based on matching fuel burn of B738.py example
# generally the weights module will be custom to each airplane
passthru = om.ExecComp('OEW=x',
x={
'val': 1.0,
'units': 'kg'
},
OEW={
'val': 1.0,
'units': 'kg'
})
self.add_subsystem('OEW',
passthru,
promotes_inputs=[('x', 'ac|weights|OEW')],
promotes_outputs=['OEW'])
self.add_subsystem('weight',
oc.AddSubtractComp(
output_name='weight',
input_names=['ac|weights|MTOW', 'fuel_used'],
units='kg',
vec_size=[1, nn],
scaling_factors=[1, -1]),
promotes_inputs=['*'],
promotes_outputs=['weight'])