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Merged
jkirk5 merged 16 commits into from
Sep 8, 2026
Merged

Convert Aircraft.Wing.LOAD_DISTRIBUTION_CONTROL to variable #1273

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6b22105
convert Aircraft.Wing.LOAD_DISTRIBUTION_CONTROL from option to variable
xjjiang Aug 20, 2026
e2ca8eb
Merge branch 'main' into Var_LOAD_DISTRIBUTION_CONTROL
xjjiang Aug 20, 2026
082a2aa
added a note
xjjiang Aug 20, 2026
826bc2f
added two cases: LOAD_DISTRIBUTION_CONTROL = 1.25 and 2.75
xjjiang Aug 21, 2026
24b9608
Merge branch 'OpenMDAO:main' into Var_LOAD_DISTRIBUTION_CONTROL
xjjiang Aug 24, 2026
4bdce0a
Merge branch 'OpenMDAO:main' into Var_LOAD_DISTRIBUTION_CONTROL
xjjiang Aug 25, 2026
63f05e2
Merge branch 'OpenMDAO:main' into Var_LOAD_DISTRIBUTION_CONTROL
xjjiang Aug 26, 2026
9e6e52e
Merge branch 'OpenMDAO:main' into Var_LOAD_DISTRIBUTION_CONTROL
xjjiang Aug 27, 2026
59f97e0
Merge branch 'OpenMDAO:main' into Var_LOAD_DISTRIBUTION_CONTROL
xjjiang Aug 27, 2026
495af6e
Merge branch 'OpenMDAO:main' into Var_LOAD_DISTRIBUTION_CONTROL
xjjiang Aug 27, 2026
370b299
Merge branch 'OpenMDAO:main' into Var_LOAD_DISTRIBUTION_CONTROL
xjjiang Aug 31, 2026
15979fb
Merge branch 'OpenMDAO:main' into Var_LOAD_DISTRIBUTION_CONTROL
xjjiang Sep 1, 2026
002d581
modified a note in code
xjjiang Sep 2, 2026
9fffbc1
Merge branch 'Var_LOAD_DISTRIBUTION_CONTROL' of github.com:xjjiang/om…
xjjiang Sep 2, 2026
2ffe5dd
Merge branch 'OpenMDAO:main' into Var_LOAD_DISTRIBUTION_CONTROL
xjjiang Sep 4, 2026
167ab78
Update aviary/subsystems/mass/flops_based/test/test_wing_detailed.py
jkirk5 Sep 8, 2026
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81 changes: 55 additions & 26 deletions aviary/subsystems/mass/flops_based/test/test_wing_detailed.py
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Open in desktop
Original file line number Diff line number Diff line change
Expand Up @@ -53,9 +53,6 @@ def test_case(self, case_name):
Aircraft.Wing.INPUT_STATION_DISTRIBUTION: inputs.get_val(
Aircraft.Wing.INPUT_STATION_DISTRIBUTION
),
Aircraft.Wing.LOAD_DISTRIBUTION_CONTROL: inputs.get_val(
Aircraft.Wing.LOAD_DISTRIBUTION_CONTROL
),
Aircraft.Wing.NUM_INTEGRATION_STATIONS: inputs.get_val(
Aircraft.Wing.NUM_INTEGRATION_STATIONS
),
Expand Down Expand Up @@ -86,6 +83,7 @@ def test_case(self, case_name):
Aircraft.Engine.WING_LOCATIONS,
Aircraft.Wing.THICKNESS_TO_CHORD,
Aircraft.Wing.THICKNESS_TO_CHORD_REFERENCE,
Aircraft.Wing.LOAD_DISTRIBUTION_CONTROL,
],
output_keys=[
Aircraft.Wing.BENDING_MATERIAL_FACTOR,
Expand Down Expand Up @@ -125,9 +123,6 @@ def test_case_multiengine(self):
Aircraft.Wing.INPUT_STATION_DISTRIBUTION: aviary_options.get_val(
Aircraft.Wing.INPUT_STATION_DISTRIBUTION
),
Aircraft.Wing.LOAD_DISTRIBUTION_CONTROL: aviary_options.get_val(
Aircraft.Wing.LOAD_DISTRIBUTION_CONTROL
),
Aircraft.Wing.NUM_INTEGRATION_STATIONS: aviary_options.get_val(
Aircraft.Wing.NUM_INTEGRATION_STATIONS
),
Expand All @@ -150,6 +145,7 @@ def test_case_multiengine(self):
Aircraft.Wing.AEROELASTIC_TAILORING_FACTOR,
Aircraft.Wing.THICKNESS_TO_CHORD,
Aircraft.Wing.THICKNESS_TO_CHORD_REFERENCE,
Aircraft.Wing.LOAD_DISTRIBUTION_CONTROL,
]

for key in input_keys:
Expand Down Expand Up @@ -213,9 +209,6 @@ def test_case_fuselage_engines(self):
Aircraft.Wing.INPUT_STATION_DISTRIBUTION: aviary_options.get_val(
Aircraft.Wing.INPUT_STATION_DISTRIBUTION
),
Aircraft.Wing.LOAD_DISTRIBUTION_CONTROL: aviary_options.get_val(
Aircraft.Wing.LOAD_DISTRIBUTION_CONTROL
),
Aircraft.Wing.NUM_INTEGRATION_STATIONS: aviary_options.get_val(
Aircraft.Wing.NUM_INTEGRATION_STATIONS
),
Expand All @@ -240,6 +233,7 @@ def test_case_fuselage_engines(self):
Aircraft.Wing.AEROELASTIC_TAILORING_FACTOR,
Aircraft.Wing.THICKNESS_TO_CHORD,
Aircraft.Wing.THICKNESS_TO_CHORD_REFERENCE,
Aircraft.Wing.LOAD_DISTRIBUTION_CONTROL,
]

for key in input_keys:
Expand Down Expand Up @@ -296,9 +290,6 @@ def test_case_fuselage_multiengine(self):
Aircraft.Wing.INPUT_STATION_DISTRIBUTION: aviary_options.get_val(
Aircraft.Wing.INPUT_STATION_DISTRIBUTION
),
Aircraft.Wing.LOAD_DISTRIBUTION_CONTROL: aviary_options.get_val(
Aircraft.Wing.LOAD_DISTRIBUTION_CONTROL
),
Aircraft.Wing.NUM_INTEGRATION_STATIONS: aviary_options.get_val(
Aircraft.Wing.NUM_INTEGRATION_STATIONS
),
Expand All @@ -323,6 +314,7 @@ def test_case_fuselage_multiengine(self):
Aircraft.Wing.AEROELASTIC_TAILORING_FACTOR,
Aircraft.Wing.THICKNESS_TO_CHORD,
Aircraft.Wing.THICKNESS_TO_CHORD_REFERENCE,
Aircraft.Wing.LOAD_DISTRIBUTION_CONTROL,
]

for key in input_keys:
Expand Down Expand Up @@ -374,9 +366,6 @@ def test_extreme_engine_loc(self):
Aircraft.Wing.INPUT_STATION_DISTRIBUTION: aviary_options.get_val(
Aircraft.Wing.INPUT_STATION_DISTRIBUTION
),
Aircraft.Wing.LOAD_DISTRIBUTION_CONTROL: aviary_options.get_val(
Aircraft.Wing.LOAD_DISTRIBUTION_CONTROL
),
Aircraft.Wing.NUM_INTEGRATION_STATIONS: aviary_options.get_val(
Aircraft.Wing.NUM_INTEGRATION_STATIONS
),
Expand All @@ -401,6 +390,7 @@ def test_extreme_engine_loc(self):
Aircraft.Wing.AEROELASTIC_TAILORING_FACTOR,
Aircraft.Wing.THICKNESS_TO_CHORD,
Aircraft.Wing.THICKNESS_TO_CHORD_REFERENCE,
Aircraft.Wing.LOAD_DISTRIBUTION_CONTROL,
]

for key in input_keys:
Expand All @@ -426,7 +416,7 @@ def test_extreme_engine_loc(self):
def test_intensity_factor(self):
"""
data taken from high_wing_single_aisle.csv
Aircraft.Wing.LOAD_DISTRIBUTION_CONTROL = 3, 1, 1.5, 2.5
Aircraft.Wing.LOAD_DISTRIBUTION_CONTROL = 3, 1, 1.5, 2.5, 1.25, 2.75
"""
options = get_option_defaults()
options.set_val(Aircraft.Engine.NUM_ENGINES, val=[2], units='unitless')
Expand All @@ -435,7 +425,6 @@ def test_intensity_factor(self):
options.set_val(
Aircraft.Wing.INPUT_STATION_DISTRIBUTION, val=[0.0, 0.5, 1.0], units='unitless'
)
options.set_val(Aircraft.Wing.LOAD_DISTRIBUTION_CONTROL, val=3, units='unitless')
options.set_val(Aircraft.Wing.NUM_INTEGRATION_STATIONS, val=46, units='unitless')

prob = self.prob
Expand Down Expand Up @@ -470,6 +459,9 @@ def test_intensity_factor(self):
prob.model.set_input_defaults(
Aircraft.Wing.THICKNESS_TO_CHORD_REFERENCE, val=0.01, units='unitless'
)
prob.model.set_input_defaults(
Aircraft.Wing.LOAD_DISTRIBUTION_CONTROL, val=3, units='unitless'
)

setup_model_options(prob, options)
prob.setup(check=False, force_alloc_complex=True)
Expand All @@ -485,7 +477,9 @@ def test_intensity_factor(self):
pod_inertia_factor = prob.get_val(Aircraft.Wing.ENG_POD_INERTIA_FACTOR)
assert_near_equal(pod_inertia_factor, 0.9772998541, tolerance=1e-9)

options.set_val(Aircraft.Wing.LOAD_DISTRIBUTION_CONTROL, val=1, units='unitless')
prob.model.set_input_defaults(
Aircraft.Wing.LOAD_DISTRIBUTION_CONTROL, val=1, units='unitless'
)
setup_model_options(prob, options)
prob.setup(check=False, force_alloc_complex=True)
prob.run_model()
Expand All @@ -496,7 +490,9 @@ def test_intensity_factor(self):
pod_inertia_factor = prob.get_val(Aircraft.Wing.ENG_POD_INERTIA_FACTOR)
assert_near_equal(pod_inertia_factor, 1.0, tolerance=1e-9)

options.set_val(Aircraft.Wing.LOAD_DISTRIBUTION_CONTROL, val=1.5, units='unitless')
prob.model.set_input_defaults(
Aircraft.Wing.LOAD_DISTRIBUTION_CONTROL, val=1.5, units='unitless'
)
setup_model_options(prob, options)
prob.setup(check=False, force_alloc_complex=True)
prob.run_model()
Expand All @@ -507,7 +503,9 @@ def test_intensity_factor(self):
pod_inertia_factor = prob.get_val(Aircraft.Wing.ENG_POD_INERTIA_FACTOR)
assert_near_equal(pod_inertia_factor, 1.0, tolerance=1e-9)

options.set_val(Aircraft.Wing.LOAD_DISTRIBUTION_CONTROL, val=2.5, units='unitless')
prob.model.set_input_defaults(
Aircraft.Wing.LOAD_DISTRIBUTION_CONTROL, val=2.5, units='unitless'
)
setup_model_options(prob, options)
prob.setup(check=False, force_alloc_complex=True)
prob.run_model()
Expand All @@ -518,6 +516,34 @@ def test_intensity_factor(self):
pod_inertia_factor = prob.get_val(Aircraft.Wing.ENG_POD_INERTIA_FACTOR)
assert_near_equal(pod_inertia_factor, 1.0, tolerance=1e-9)

prob.model.set_input_defaults(
Aircraft.Wing.LOAD_DISTRIBUTION_CONTROL, val=1.25, units='unitless'
)
setup_model_options(prob, options)
prob.setup(check=False, force_alloc_complex=True)
prob.run_model()

bending_mat_factor = prob.get_val(Aircraft.Wing.BENDING_MATERIAL_FACTOR)
# without smoothing, it is 1.7507575953
assert_near_equal(bending_mat_factor, 1.78032845, tolerance=1e-9)

pod_inertia_factor = prob.get_val(Aircraft.Wing.ENG_POD_INERTIA_FACTOR)
assert_near_equal(pod_inertia_factor, 1.0, tolerance=1e-9)

prob.model.set_input_defaults(
Aircraft.Wing.LOAD_DISTRIBUTION_CONTROL, val=2.75, units='unitless'
)
setup_model_options(prob, options)
prob.setup(check=False, force_alloc_complex=True)
prob.run_model()

bending_mat_factor = prob.get_val(Aircraft.Wing.BENDING_MATERIAL_FACTOR)
# without smoothing, it is 2.7838124896
assert_near_equal(bending_mat_factor, 2.8107811534, tolerance=1e-9)

pod_inertia_factor = prob.get_val(Aircraft.Wing.ENG_POD_INERTIA_FACTOR)
assert_near_equal(pod_inertia_factor, 1.0, tolerance=1e-9)

def test_IO(self):
assert_match_varnames(self.prob.model)

Expand All @@ -542,7 +568,6 @@ def test_case1(self):
aviary_options.set_val(
Aircraft.Wing.INPUT_STATION_DISTRIBUTION, [0.5, 1.0], units='unitless'
)
aviary_options.set_val(Aircraft.Wing.LOAD_DISTRIBUTION_CONTROL, 2.0, units='unitless')
aviary_options.set_val(Aircraft.Wing.NUM_INTEGRATION_STATIONS, 50, units='unitless')
aviary_options.set_val(Aircraft.BWB.DETAILED_WING_PROVIDED, False, units='unitless')

Expand All @@ -564,6 +589,9 @@ def test_case1(self):
prob.model.set_input_defaults(
Aircraft.Wing.THICKNESS_TO_CHORD_REFERENCE, 0.11, units='unitless'
)
prob.model.set_input_defaults(
Aircraft.Wing.LOAD_DISTRIBUTION_CONTROL, 2.0, units='unitless'
)

setup_model_options(self.prob, aviary_options)
prob.setup(check=False, force_alloc_complex=True)
Expand Down Expand Up @@ -614,7 +642,6 @@ def test_case1(self):
[0.35, 0.4, 0.45, 0.5, 0.55, 0.6, 0.6499, 0.7, 0.75, 0.8, 0.85, 0.8999, 0.95, 1],
units='unitless',
)
aviary_options.set_val(Aircraft.Wing.LOAD_DISTRIBUTION_CONTROL, 2.0, units='unitless')
aviary_options.set_val(Aircraft.Wing.NUM_INTEGRATION_STATIONS, 100, units='unitless')
aviary_options.set_val(Aircraft.BWB.DETAILED_WING_PROVIDED, True, units='unitless')

Expand All @@ -636,6 +663,9 @@ def test_case1(self):
prob.model.set_input_defaults(
Aircraft.Wing.THICKNESS_TO_CHORD_REFERENCE, 0.11, units='unitless'
)
prob.model.set_input_defaults(
Aircraft.Wing.LOAD_DISTRIBUTION_CONTROL, 2.0, units='unitless'
)

setup_model_options(self.prob, aviary_options)
prob.setup(check=False, force_alloc_complex=True)
Expand Down Expand Up @@ -734,7 +764,6 @@ def test_case2(self):
[0.0, 0.2075, 0.415, 0.6927, 0.928, 1.0],
units='unitless',
)
aviary_options.set_val(Aircraft.Wing.LOAD_DISTRIBUTION_CONTROL, 2.0, units='unitless')
aviary_options.set_val(Aircraft.Wing.NUM_INTEGRATION_STATIONS, 10, units='unitless')
aviary_options.set_val(Aircraft.BWB.DETAILED_WING_PROVIDED, True, units='unitless')

Expand All @@ -755,6 +784,9 @@ def test_case2(self):
prob.model.set_input_defaults(
Aircraft.Wing.THICKNESS_TO_CHORD_REFERENCE, 0.11, units='unitless'
)
prob.model.set_input_defaults(
Aircraft.Wing.LOAD_DISTRIBUTION_CONTROL, 2.0, units='unitless'
)

setup_model_options(self.prob, aviary_options)
prob.setup(check=False, force_alloc_complex=True)
Expand Down Expand Up @@ -797,6 +829,3 @@ def test_case2(self):

if __name__ == '__main__':
unittest.main()
# test = DetailedWingBendingTest()
# test.setUp()
# test.test_intensity_factor()
38 changes: 32 additions & 6 deletions aviary/subsystems/mass/flops_based/wing_detailed.py
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Original file line number Diff line number Diff line change
Expand Up @@ -46,6 +46,32 @@ def load_intensity_by_factor(load_dist_factor, intn_stations):
return load_intensity


def smooth_load_intensity_by_factor(load_dist_factor, intn_stations):
"""
Calculate load intensity at wing integration stations for the given load_dist_factor.

Parameters:
load_dist_factor (float): 0 or 1 <= load_dist_factor <= 3.
1.0 : triangular distribution
2.0 : elliptical distribution (default)
3.0 : rectangular distribution
1.0-2.0 : blend of triangular and elliptical by quartic polynomial
2.0-3.0 : blend of elliptical and rectangular by quartic polynomial
intn_stations (array): integration stations

Note: In FLOPS, load_dist_factor can be 0 in which case load intensities are computed
based on input pressure distribution. This is NOT currently implemented in Aviary.
"""
a = 1.0 - intn_stations
b = np.sqrt(1.0 - intn_stations**2)
c = np.ones(len(intn_stations))
t = load_dist_factor - 2.0
load_intensity = (
b + (c - a) / 2 * t + 7 * (a - 2 * b + c) / 6 * t**2 + 2 * (-a + 2 * b - c) / 3 * t**4
)
return load_intensity


class DetailedWingBendingFact(om.ExplicitComponent):
"""
Computation of wing bending factor and engine inertia relief factor
Expand All @@ -62,7 +88,6 @@ def initialize(self):
add_aviary_option(self, Aircraft.Engine.NUM_WING_ENGINES)
add_aviary_option(self, Aircraft.Propulsion.TOTAL_NUM_WING_ENGINES)
add_aviary_option(self, Aircraft.Wing.INPUT_STATION_DISTRIBUTION)
add_aviary_option(self, Aircraft.Wing.LOAD_DISTRIBUTION_CONTROL)
add_aviary_option(self, Aircraft.Wing.NUM_INTEGRATION_STATIONS)

def setup(self):
Expand Down Expand Up @@ -95,6 +120,7 @@ def setup(self):
add_aviary_input(self, Aircraft.Wing.ASPECT_RATIO_REFERENCE, units='unitless')
add_aviary_input(self, Aircraft.Wing.STRUT_BRACING_FACTOR, units='unitless')
add_aviary_input(self, Aircraft.Wing.AEROELASTIC_TAILORING_FACTOR, units='unitless')
add_aviary_input(self, Aircraft.Wing.LOAD_DISTRIBUTION_CONTROL)

if total_num_wing_engines > 1:
add_aviary_input(
Expand Down Expand Up @@ -167,8 +193,8 @@ def compute(self, inputs, outputs):
(dy[1:] + 2.0 * integration_stations[1:-1]) * dy[1:] * sweep_int_stations[1:-1]
)

load_distribution_factor = self.options[Aircraft.Wing.LOAD_DISTRIBUTION_CONTROL]
load_intensity = load_intensity_by_factor(load_distribution_factor, integration_stations)
load_distrib_factor = inputs[Aircraft.Wing.LOAD_DISTRIBUTION_CONTROL]
load_intensity = smooth_load_intensity_by_factor(load_distrib_factor, integration_stations)

chord_interp = InterpND(
method='slinear', points=(inp_stations), x_interp=integration_stations
Expand Down Expand Up @@ -308,7 +334,6 @@ def initialize(self):
add_aviary_option(self, Aircraft.Engine.NUM_WING_ENGINES)
add_aviary_option(self, Aircraft.Propulsion.TOTAL_NUM_WING_ENGINES)
add_aviary_option(self, Aircraft.Wing.INPUT_STATION_DISTRIBUTION)
add_aviary_option(self, Aircraft.Wing.LOAD_DISTRIBUTION_CONTROL)
add_aviary_option(self, Aircraft.Wing.NUM_INTEGRATION_STATIONS)
add_aviary_option(self, Aircraft.BWB.DETAILED_WING_PROVIDED)
add_aviary_option(self, Settings.VERBOSITY)
Expand Down Expand Up @@ -340,6 +365,7 @@ def setup(self):
add_aviary_input(self, Aircraft.Wing.AEROELASTIC_TAILORING_FACTOR, units='unitless')
add_aviary_input(self, Aircraft.Wing.SPAN, units='ft')
add_aviary_input(self, Aircraft.Fuselage.MAX_WIDTH, units='ft')
add_aviary_input(self, Aircraft.Wing.LOAD_DISTRIBUTION_CONTROL)

if total_num_wing_engines > 1:
add_aviary_input(
Expand Down Expand Up @@ -460,8 +486,8 @@ def compute(self, inputs, outputs):
(dy[0:] + 2.0 * integration_stations[0:-1]) * dy[0:] * sweep_int_stations[0:-1]
)

load_distribution_factor = self.options[Aircraft.Wing.LOAD_DISTRIBUTION_CONTROL]
load_intensity = load_intensity_by_factor(load_distribution_factor, integration_stations)
load_distrib_factor = inputs[Aircraft.Wing.LOAD_DISTRIBUTION_CONTROL]
load_intensity = smooth_load_intensity_by_factor(load_distrib_factor, integration_stations)

chord_interp = InterpND(
method='slinear', points=(inp_stations_mod), x_interp=integration_stations
Expand Down
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