def test_aero_analysis_flat_multiple(self):
OAS_prob = OASProblem({'type' : 'aero',
'optimize' : False})
OAS_prob.add_surface({'span_cos_spacing' : 0.})
OAS_prob.add_surface({'name' : 'tail',
'span_cos_spacing' : 0.,
'offset' : numpy.array([0., 0., 1000000.])})
OAS_prob.setup()
OAS_prob.run()
prob = OAS_prob.prob
self.assertAlmostEqual(prob['CL'], .46173591841167, places=5)
self.assertAlmostEqual(prob['tail_CL'], .46173591841167, places=5)
def test_aero_analysis_flat_symmetry(self):
OAS_prob = OASProblem({'type': 'aero', 'optimize': False})
surf_dict = {'symmetry': True}
OAS_prob.add_surface(surf_dict)
OAS_prob.setup()
OAS_prob.run()
prob = OAS_prob.prob
self.assertAlmostEqual(prob['wing_perf.CL'], .45655138, places=5)
def test_struct_analysis_symmetry(self):
OAS_prob = OASProblem({'type': 'struct', 'optimize': False})
surf_dict = {'symmetry': True}
OAS_prob.add_surface(surf_dict)
OAS_prob.setup()
OAS_prob.run()
prob = OAS_prob.prob
self.assertAlmostEqual(prob['wing.weight'], 83.2113646, places=3)
def test_aero_analysis_flat(self):
OAS_prob = OASProblem({'type': 'aero', 'optimize': False})
OAS_prob.add_surface({'span_cos_spacing': 0})
OAS_prob.setup()
OAS_prob.run()
prob = OAS_prob.prob
self.assertAlmostEqual(prob['wing_perf.CL'], .46173591841167, places=5)
self.assertAlmostEqual(prob['wing_perf.CD'], .005524603647, places=5)
def test_aero_analysis_flat_side_by_side(self):
OAS_prob = OASProblem({'type' : 'aero',
'optimize' : False})
OAS_prob.add_surface({'name' : 'wing',
'span' : 5.,
'num_y' : 3,
'span_cos_spacing' : 0.,
'offset' : numpy.array([0., -2.5, 0.])})
OAS_prob.add_surface({'name' : 'tail',
'span' : 5.,
'num_y' : 3,
'span_cos_spacing' : 0.,
'offset' : numpy.array([0., 2.5, 0.])})
OAS_prob.setup()
OAS_prob.run()
prob = OAS_prob.prob
self.assertAlmostEqual(prob['wing_CL'], .46173591841167, places=5)
self.assertAlmostEqual(prob['tail_CL'], .46173591841167, places=5)
self.assertAlmostEqual(prob['wing_CD'], .005524603647, places=5)
self.assertAlmostEqual(prob['tail_CD'], .005524603647, places=5)
def test_aerostruct_analysis(self):
OAS_prob = OASProblem({'type' : 'aerostruct',
'optimize' : False})
surf_dict = {'num_y' : 13,
'num_x' : 2,
'wing_type' : 'CRM',
'CL0' : 0.2,
'CD0' : 0.015}
OAS_prob.add_surface(surf_dict)
OAS_prob.setup()
OAS_prob.run()
prob = OAS_prob.prob
self.assertAlmostEqual(prob['CL'], .58245256)
self.assertAlmostEqual(prob['failure'], -.431801158, places=5)
self.assertAlmostEqual(prob['fuelburn'], 1400891.8033734, places=2)
def test_aerostruct_analysis_symmetry(self):
OAS_prob = OASProblem({'type': 'aerostruct', 'optimize': False})
surf_dict = {
'symmetry': True,
'num_y': 13,
'num_x': 2,
'wing_type': 'CRM',
'CL0': 0.2,
'CD0': 0.015
}
OAS_prob.add_surface(surf_dict)
OAS_prob.setup()
OAS_prob.run()
prob = OAS_prob.prob
self.assertAlmostEqual(prob['wing_perf.CL'], .58245256)
self.assertAlmostEqual(prob['wing_perf.failure'],
-.5011158763,
places=5)
self.assertAlmostEqual(prob['fuelburn'], 139970.19140120025, places=2)
def test_aero_optimization_flat(self):
OAS_prob = OASProblem({'type': 'aero', 'optimize': True})
OAS_prob.add_surface()
OAS_prob.setup()
OAS_prob.add_desvar('wing.twist_cp', lower=-10., upper=15.)
OAS_prob.add_desvar('wing.sweep', lower=10., upper=30.)
OAS_prob.add_desvar('wing.dihedral', lower=-10., upper=20.)
OAS_prob.add_desvar('wing.taper', lower=.5, upper=2.)
OAS_prob.add_constraint('wing_perf.CL', equals=0.5)
OAS_prob.add_objective('wing_perf.CD', scaler=1e4)
OAS_prob.run()
prob = OAS_prob.prob
self.assertAlmostEqual(prob['wing_perf.CD'],
.004048702908627036,
places=5)
def test_aerostruct_optimization_symmetry_multiple(self):
OAS_prob = OASProblem({'type': 'aerostruct', 'optimize': True})
surf_dict = {
'name': 'wing',
'symmetry': True,
'num_y': 13,
'num_x': 2,
'wing_type': 'CRM',
'CL0': 0.2,
'CD0': 0.015
}
OAS_prob.add_surface(surf_dict)
surf_dict.update({
'name': 'tail',
'offset': numpy.array([0., 0., 1.e7])
})
OAS_prob.add_surface(surf_dict)
OAS_prob.setup()
# Add design variables and constraints for both the wing and tail
OAS_prob.add_desvar('wing.twist_cp', lower=-15., upper=15.)
OAS_prob.add_desvar('wing.thickness_cp',
lower=0.01,
upper=0.25,
scaler=1e2)
OAS_prob.add_constraint('wing_perf.failure', upper=0.)
OAS_prob.add_desvar('tail.twist_cp', lower=-15., upper=15.)
OAS_prob.add_desvar('tail.thickness_cp',
lower=0.01,
upper=0.25,
scaler=1e2)
OAS_prob.add_constraint('tail_perf.failure', upper=0.)
OAS_prob.add_desvar('alpha', lower=-10., upper=10.)
OAS_prob.add_constraint('eq_con', equals=0.)
OAS_prob.add_objective('fuelburn', scaler=1e-5)
OAS_prob.run()
prob = OAS_prob.prob
self.assertAlmostEqual(prob['fuelburn'], 160496.22753448418, places=0)
self.assertAlmostEqual(prob['wing_perf.failure'], 1e-9, places=5)
self.assertAlmostEqual(numpy.linalg.norm(prob['wing.twist_cp']),
numpy.linalg.norm(prob['tail.twist_cp']),
places=2)
def test_aerostruct_optimization_symmetry(self):
OAS_prob = OASProblem({'type': 'aerostruct', 'optimize': True})
surf_dict = {
'symmetry': True,
'num_y': 13,
'num_x': 2,
'wing_type': 'CRM',
'CL0': 0.2,
'CD0': 0.015
}
OAS_prob.add_surface(surf_dict)
OAS_prob.setup()
OAS_prob.add_desvar('wing.twist_cp', lower=-15., upper=15.)
OAS_prob.add_desvar('wing.thickness_cp',
lower=0.01,
upper=0.25,
scaler=1e2)
OAS_prob.add_constraint('wing_perf.failure', upper=0.)
OAS_prob.add_desvar('alpha', lower=-10., upper=10.)
OAS_prob.add_constraint('eq_con', equals=0.)
OAS_prob.add_objective('fuelburn', scaler=1e-4)
OAS_prob.run()
prob = OAS_prob.prob
self.assertAlmostEqual(prob['fuelburn'], 80248.113842385414, places=0)
self.assertAlmostEqual(prob['wing_perf.failure'], 1e-9)
'\n +---------------------------------------------------------------+')
print(' | ERROR: Please supply a correct input argument to this script. |')
print(' | Possible options are ' + print_str[:-2] + ' |')
print(
' +---------------------------------------------------------------+\n')
raise
# Check input_arg and run prob1 if selected
if input_arg == 'prob1':
# Set problem type. Any option not set here will be set in the method
# `get_default_prob_dict` within `run_classes.py`
prob_dict = {'type': 'struct'}
# Instantiate OpenAeroStruct (OAS) problem
OAS_prob = OASProblem(prob_dict)
# We specify the necessary parameters here.
# Defaults are set in `get_default_surf_dict` within `run_classes.py` and
# then are overwritten with the settings here.
surface = {
'name': 'wing', # name of the surface
'num_x': 2, # number of chordwise points
'num_y': 15, # number of spanwise points
'span': 10., # full wingspan
'chord': 1., # root chord
'span_cos_spacing': 0, # 0 for uniform spanwise panels
# 1 for cosine-spaced panels
# any value between 0 and 1 for
# a mixed spacing
def setup(prob_dict={}, surfaces=[{}]):
''' Setup the aerostruct mesh
Default wing mesh (single lifting surface):
-------------------------------------------
name = 'wing' # name of the surface
num_x = 3 # number of chordwise points
num_y = 5 # number of spanwise points
root_chord = 1. # root chord
span_cos_spacing = 1 # 0 for uniform spanwise panels
# 1 for cosine-spaced panels
# any value between 0 and 1 for a mixed spacing
chord_cos_spacing = 0. # 0 for uniform chordwise panels
# 1 for cosine-spaced panels
# any value between 0 and 1 for a mixed spacing
wing_type = 'rect' # initial shape of the wing either 'CRM' or 'rect'
# 'CRM' can have different options after it, such as 'CRM:alpha_2.75' for the CRM shape at alpha=2.75
offset = np.array([0., 0., 0.]) # coordinates to offset the surface from its default location
symmetry = True # if true, model one half of wing reflected across the plane y = 0
S_ref_type = 'wetted' # 'wetted' or 'projected'
# Simple Geometric Variables
span = 10. # full wingspan
dihedral = 0. # wing dihedral angle in degrees positive is upward
sweep = 0. # wing sweep angle in degrees positive sweeps back
taper = 1. # taper ratio; 1. is uniform chord
# B-spline Geometric Variables. The number of control points for each of these variables can be specified in surf_dict
# by adding the prefix "num" to the variable (e.g. num_twist)
twist_cp = None
chord_cp = None
xshear_cp = None
zshear_cp = None
thickness_cp = None
Default wing parameters:
------------------------
Zero-lift aerodynamic performance
CL0 = 0.0 # CL value at AoA (alpha) = 0
CD0 = 0.0 # CD value at AoA (alpha) = 0
Airfoil properties for viscous drag calculation
k_lam = 0.05 # percentage of chord with laminar flow, used for viscous drag
t_over_c = 0.12 # thickness over chord ratio (NACA0012)
c_max_t = .303 # chordwise location of maximum (NACA0012) thickness
Structural values are based on aluminum
E = 70.e9 # [Pa] Young's modulus of the spar
G = 30.e9 # [Pa] shear modulus of the spar
yield = 20.e6 # [Pa] yield stress
mrho = 3.e3 # [kg/m^3] material density
fem_origin = 0.35 # chordwise location of the spar
Other
W0 = 0.4 * 3e5 # [kg] MTOW of B777 is 3e5 kg with fuel
Default problem parameters:
---------------------------
Re = 1e6 # Reynolds number
reynolds_length = 1.0 # characteristic Reynolds length
alpha = 5. # angle of attack
CT = 9.80665 * 17.e-6 # [1/s] (9.81 N/kg * 17e-6 kg/N/s)
R = 14.3e6 # [m] maximum range
M = 0.84 # Mach number at cruise
rho = 0.38 # [kg/m^3] air density at 35,000 ft
a = 295.4 # [m/s] speed of sound at 35,000 ft
with_viscous = False # if true, compute viscous drag
'''
# Use steps in run_aerostruct.py to add wing surface to problem
# Set problem type
prob_dict.update({'type' : 'aerostruct'}) # this doesn't really matter since we aren't calling OASProblem.setup()
# Instantiate problem
OAS_prob = OASProblem(prob_dict)
for surface in surfaces:
# Add SpatialBeamFEM size
FEMsize = 6 * surface['num_y'] + 6
surface.update({'FEMsize': FEMsize})
# Add the specified wing surface to the problem.
OAS_prob.add_surface(surface)
# Add materials properties for the wing surface to the surface dict in OAS_prob
for idx, surface in enumerate(OAS_prob.surfaces):
A, Iy, Iz, J = materials_tube(surface['radius'], surface['thickness'], surface)
OAS_prob.surfaces[idx].update({
'A': A,
'Iy': Iy,
'Iz': Iz,
'J': J
})
# Get total panels and save in prob_dict
tot_panels = 0
for surface in OAS_prob.surfaces:
ny = surface['num_y']
nx = surface['num_x']
tot_panels += (nx - 1) * (ny - 1)
OAS_prob.prob_dict.update({'tot_panels': tot_panels})
# Assume we are only using a single lifting surface for now
surface = OAS_prob.surfaces[0]
# Initialize the OpenAeroStruct components and save them in a component dictionary
comp_dict = {}
comp_dict['MaterialsTube'] = MaterialsTube(surface)
comp_dict['GeometryMesh'] = GeometryMesh(surface)
comp_dict['TransferDisplacements'] = TransferDisplacements(surface)
comp_dict['VLMGeometry'] = VLMGeometry(surface)
comp_dict['AssembleAIC'] = AssembleAIC([surface])
comp_dict['AeroCirculations'] = AeroCirculations(OAS_prob.prob_dict['tot_panels'])
comp_dict['VLMForces'] = VLMForces([surface])
comp_dict['TransferLoads'] = TransferLoads(surface)
comp_dict['ComputeNodes'] = ComputeNodes(surface)
comp_dict['AssembleK'] = AssembleK(surface)
comp_dict['SpatialBeamFEM'] = SpatialBeamFEM(surface['FEMsize'])
comp_dict['SpatialBeamDisp'] = SpatialBeamDisp(surface)
OAS_prob.comp_dict = comp_dict
return OAS_prob
import numpy
from run_classes import OASProblem
if __name__ == "__main__":
# Set problem type
prob_dict = {'type' : 'aerostruct'}
if sys.argv[1].startswith('0'): # run analysis once
prob_dict.update({'optimize' : False})
else: # perform optimization
prob_dict.update({'optimize' : True})
# Instantiate problem and add default surface
OAS_prob = OASProblem(prob_dict)
# Create a dictionary to store options about the surface
surf_dict = {'name' : 'wing',
'symmetry' : False,
'num_y' : 13,
'num_x' : 2,
'wing_type' : 'CRM',
'CL0' : 0.2,
'CD0' : 0.015}
# Add the specified wing surface to the problem
OAS_prob.add_surface(surf_dict)
# Single lifting surface
if not sys.argv[1].endswith('m'):
def setup(prob_dict={}, surfaces=[{}]):
''' Setup the aerostruct mesh
Default wing mesh (single lifting surface):
-------------------------------------------
name = 'wing' # name of the surface
num_x = 3 # number of chordwise points
num_y = 5 # number of spanwise points
root_chord = 1. # root chord
span_cos_spacing = 1 # 0 for uniform spanwise panels
# 1 for cosine-spaced panels
# any value between 0 and 1 for a mixed spacing
chord_cos_spacing = 0. # 0 for uniform chordwise panels
# 1 for cosine-spaced panels
# any value between 0 and 1 for a mixed spacing
wing_type = 'rect' # initial shape of the wing either 'CRM' or 'rect'
# 'CRM' can have different options after it, such as 'CRM:alpha_2.75' for the CRM shape at alpha=2.75
offset = np.array([0., 0., 0.]) # coordinates to offset the surface from its default location
symmetry = True # if true, model one half of wing reflected across the plane y = 0
S_ref_type = 'wetted' # 'wetted' or 'projected'
# Simple Geometric Variables
span = 10. # full wingspan
dihedral = 0. # wing dihedral angle in degrees positive is upward
sweep = 0. # wing sweep angle in degrees positive sweeps back
taper = 1. # taper ratio; 1. is uniform chord
# B-spline Geometric Variables. The number of control points for each of these variables can be specified in surf_dict
# by adding the prefix "num" to the variable (e.g. num_twist)
twist_cp = None
chord_cp = None
xshear_cp = None
zshear_cp = None
thickness_cp = None
Default wing parameters:
------------------------
Zero-lift aerodynamic performance
CL0 = 0.0 # CL value at AoA (alpha) = 0
CD0 = 0.0 # CD value at AoA (alpha) = 0
Airfoil properties for viscous drag calculation
k_lam = 0.05 # percentage of chord with laminar flow, used for viscous drag
t_over_c = 0.12 # thickness over chord ratio (NACA0012)
c_max_t = .303 # chordwise location of maximum (NACA0012) thickness
Structural values are based on aluminum
E = 70.e9 # [Pa] Young's modulus of the spar
G = 30.e9 # [Pa] shear modulus of the spar
stress = 20.e6 # [Pa] yield stress
mrho = 3.e3 # [kg/m^3] material density
fem_origin = 0.35 # chordwise location of the spar
Other
W0 = 0.4 * 3e5 # [kg] MTOW of B777 is 3e5 kg with fuel
Default problem parameters:
---------------------------
Re = 1e6 # Reynolds number
reynolds_length = 1.0 # characteristic Reynolds length
alpha = 5. # angle of attack
CT = 9.80665 * 17.e-6 # [1/s] (9.81 N/kg * 17e-6 kg/N/s)
R = 14.3e6 # [m] maximum range
M = 0.84 # Mach number at cruise
rho = 0.38 # [kg/m^3] air density at 35,000 ft
a = 295.4 # [m/s] speed of sound at 35,000 ft
with_viscous = False # if true, compute viscous drag
'''
# Use steps in run_aerostruct.py to add wing surface to problem
# Set problem type
prob_dict.update({
'type': 'aerostruct'
}) # this doesn't really matter since we aren't calling OASProblem.setup()
# Instantiate problem
OAS_prob = OASProblem(prob_dict)
for surface in surfaces:
# Add SpatialBeamFEM size
FEMsize = 6 * surface['num_y'] + 6
surface.update({'FEMsize': FEMsize})
# Add the specified wing surface to the problem.
OAS_prob.add_surface(surface)
# Add materials properties for the wing surface to the surface dict in OAS_prob
for idx, surface in enumerate(OAS_prob.surfaces):
A, Iy, Iz, J = materials_tube(surface['radius'], surface['thickness'],
surface)
OAS_prob.surfaces[idx].update({'A': A, 'Iy': Iy, 'Iz': Iz, 'J': J})
# Get total panels and save in prob_dict
tot_panels = 0
for surface in OAS_prob.surfaces:
ny = surface['num_y']
nx = surface['num_x']
tot_panels += (nx - 1) * (ny - 1)
OAS_prob.prob_dict.update({'tot_panels': tot_panels})
# Assume we are only using a single lifting surface for now
surface = OAS_prob.surfaces[0]
# Initialize the OpenAeroStruct components and save them in a component dictionary
comp_dict = {}
comp_dict['MaterialsTube'] = MaterialsTube(surface)
comp_dict['GeometryMesh'] = GeometryMesh(surface)
comp_dict['TransferDisplacements'] = TransferDisplacements(surface)
comp_dict['VLMGeometry'] = VLMGeometry(surface)
comp_dict['AssembleAIC'] = AssembleAIC([surface])
comp_dict['AeroCirculations'] = AeroCirculations(
OAS_prob.prob_dict['tot_panels'])
comp_dict['VLMForces'] = VLMForces([surface])
comp_dict['TransferLoads'] = TransferLoads(surface)
comp_dict['ComputeNodes'] = ComputeNodes(surface)
comp_dict['AssembleK'] = AssembleK(surface)
comp_dict['SpatialBeamFEM'] = SpatialBeamFEM(surface['FEMsize'])
comp_dict['SpatialBeamDisp'] = SpatialBeamDisp(surface)
OAS_prob.comp_dict = comp_dict
return OAS_prob
def setup(prob_dict={}, surfaces=[{}]):
''' Setup the aerostruct mesh
Default wing mesh (single lifting surface):
-------------------------------------------
name = 'wing' # name of the surface
num_x = 3 # number of chordwise points
num_y = 5 # number of spanwise points
root_chord = 1. # root chord
span_cos_spacing = 1 # 0 for uniform spanwise panels
# 1 for cosine-spaced panels
# any value between 0 and 1 for a mixed spacing
chord_cos_spacing = 0. # 0 for uniform chordwise panels
# 1 for cosine-spaced panels
# any value between 0 and 1 for a mixed spacing
wing_type = 'rect' # initial shape of the wing either 'CRM' or 'rect'
# 'CRM' can have different options after it, such as 'CRM:alpha_2.75' for the CRM shape at alpha=2.75
offset = np.array([0., 0., 0.]) # coordinates to offset the surface from its default location
symmetry = True # if true, model one half of wing reflected across the plane y = 0
S_ref_type = 'wetted' # 'wetted' or 'projected'
# Simple Geometric Variables
span = 10. # full wingspan
dihedral = 0. # wing dihedral angle in degrees positive is upward
sweep = 0. # wing sweep angle in degrees positive sweeps back
taper = 1. # taper ratio; 1. is uniform chord
# B-spline Geometric Variables. The number of control points for each of these variables can be specified in surf_dict
# by adding the prefix "num" to the variable (e.g. num_twist)
twist_cp = None
chord_cp = None
xshear_cp = None
zshear_cp = None
thickness_cp = None
Default wing parameters:
------------------------
Zero-lift aerodynamic performance
CL0 = 0.0 # CL value at AoA (alpha) = 0
CD0 = 0.0 # CD value at AoA (alpha) = 0
Airfoil properties for viscous drag calculation
k_lam = 0.05 # percentage of chord with laminar flow, used for viscous drag
t_over_c = 0.12 # thickness over chord ratio (NACA0012)
c_max_t = .303 # chordwise location of maximum (NACA0012) thickness
Structural values are based on aluminum
E = 70.e9 # [Pa] Young's modulus of the spar
G = 30.e9 # [Pa] shear modulus of the spar
<<<<<<< HEAD
stress = 20.e6 # [Pa] yield stress
=======
yield = 20.e6 # [Pa] yield stress
>>>>>>> 7eefd15e6c26c95cbbd9f303d23ecb4716c2fea3
mrho = 3.e3 # [kg/m^3] material density
fem_origin = 0.35 # chordwise location of the spar
Other
W0 = 0.4 * 3e5 # [kg] MTOW of B777 is 3e5 kg with fuel
Default problem parameters:
---------------------------
Re = 1e6 # Reynolds number
reynolds_length = 1.0 # characteristic Reynolds length
alpha = 5. # angle of attack
CT = 9.80665 * 17.e-6 # [1/s] (9.81 N/kg * 17e-6 kg/N/s)
R = 14.3e6 # [m] maximum range
M = 0.84 # Mach number at cruise
rho = 0.38 # [kg/m^3] air density at 35,000 ft
a = 295.4 # [m/s] speed of sound at 35,000 ft
with_viscous = False # if true, compute viscous drag
'''
# Use steps in run_aerostruct.py to add wing surface to problem
# Set problem type
prob_dict.update({'type' : 'aerostruct'}) # this doesn't really matter since we aren't calling OASProblem.setup()
# Instantiate problem
OAS_prob = OASProblem(prob_dict)
for surface in surfaces:
# Add SpatialBeamFEM size
FEMsize = 6 * surface['num_y'] + 6
surface.update({'FEMsize': FEMsize})
# Add the specified wing surface to the problem.
OAS_prob.add_surface(surface)
# Add materials properties for the wing surface to the surface dict in OAS_prob
for idx, surface in enumerate(OAS_prob.surfaces):
def test_aero_multiple_opt(self):
OAS_prob = OASProblem({'type': 'aero', 'optimize': True})
surf_dict = {
'name': 'wing',
'span': 5.,
'num_y': 3,
'span_cos_spacing': 0.
}
OAS_prob.add_surface(surf_dict)
surf_dict.update({
'name': 'tail',
'offset': numpy.array([0., 0., 10.])
})
OAS_prob.add_surface(surf_dict)
OAS_prob.setup()
OAS_prob.add_desvar('tail.twist_cp', lower=-10., upper=15.)
OAS_prob.add_desvar('tail.sweep', lower=10., upper=30.)
OAS_prob.add_desvar('tail.dihedral', lower=-10., upper=20.)
OAS_prob.add_desvar('tail.taper', lower=.5, upper=2.)
OAS_prob.add_constraint('tail_perf.CL', equals=0.5)
OAS_prob.add_objective('tail_perf.CD', scaler=1e4)
OAS_prob.run()
prob = OAS_prob.prob
self.assertAlmostEqual(prob['wing_perf.CL'],
.41543435621928004,
places=5)
self.assertAlmostEqual(prob['tail_perf.CL'], .5, places=5)
self.assertAlmostEqual(prob['wing_perf.CD'],
.0075400306289957033,
places=5)
self.assertAlmostEqual(prob['tail_perf.CD'],
.00791118243006308,
places=5)
print(' | ERROR: Please supply a correct input argument to this script. |')
print(' | Possible options are ' + print_str[:-2] + ' |')
print(' | See the docstring at the top of this file for more info. |')
print(' +---------------------------------------------------------------+\n')
raise
# Set problem type
prob_dict = {'type' : 'struct'}
if sys.argv[1].startswith('0'): # run analysis once
prob_dict.update({'optimize' : False})
else: # perform optimization
prob_dict.update({'optimize' : True})
# Instantiate problem and add default surface
OAS_prob = OASProblem(prob_dict)
OAS_prob.add_surface({'name' : 'wing',
'num_y' : 5})
# Single lifting surface
if not sys.argv[1].endswith('m'):
# Setup problem and add design variables, constraint, and objective
OAS_prob.setup()
OAS_prob.add_desvar('wing.thickness_cp', lower=0.001, upper=0.25, scaler=1e2)
OAS_prob.add_constraint('wing.failure', upper=0.)
OAS_prob.add_objective('wing.weight', scaler=1e-3)
# Multiple lifting surfaces
else:
def test_struct_optimization(self):
OAS_prob = OASProblem({'type': 'struct', 'optimize': True})
OAS_prob.add_surface()
OAS_prob.setup()
OAS_prob.add_desvar('wing.thickness_cp',
lower=0.001,
upper=0.25,
scaler=1e2)
OAS_prob.add_constraint('wing.failure', upper=0.)
OAS_prob.add_objective('wing.weight', scaler=1e-3)
OAS_prob.run()
prob = OAS_prob.prob
self.assertAlmostEqual(prob['wing.weight'], 2010.4792274, places=2)
)
print(
' +---------------------------------------------------------------+\n'
)
raise
# Set problem type
prob_dict = {'type': 'aero'}
if sys.argv[1].startswith('0'): # run analysis once
prob_dict.update({'optimize': False})
else: # perform optimization
prob_dict.update({'optimize': True})
# Instantiate problem and add default surface
OAS_prob = OASProblem(prob_dict)
OAS_prob.add_surface({
'name': 'wing',
'symmetry': True,
'num_y': 5,
'num_x': 3
})
# Single lifting surface
if not sys.argv[1].endswith('m'):
# Setup problem and add design variables, constraint, and objective
OAS_prob.setup()
OAS_prob.add_desvar('wing.twist_cp', lower=-10., upper=15.)
OAS_prob.add_desvar('wing.sweep', lower=10., upper=30.)
OAS_prob.add_desvar('wing.dihedral', lower=-10., upper=20.)
def test_struct_optimization_symmetry(self):
OAS_prob = OASProblem({'type': 'struct', 'optimize': True})
surf_dict = {'symmetry': True}
OAS_prob.add_surface(surf_dict)
OAS_prob.setup()
OAS_prob.add_desvar('wing.thickness_cp',
lower=0.001,
upper=0.25,
scaler=1e2)
OAS_prob.add_constraint('wing.failure', upper=0.)
OAS_prob.add_objective('wing.weight', scaler=1e-3)
OAS_prob.run()
prob = OAS_prob.prob
self.assertAlmostEqual(prob['wing.weight'],
1908.6362044761127,
places=2)