def test(self):
from openaerostruct.geometry.utils import generate_mesh, write_FFD_file
from openaerostruct.geometry.geometry_group import Geometry
from openaerostruct.transfer.displacement_transfer import DisplacementTransfer
from openaerostruct.aerodynamics.aero_groups import AeroPoint
from openmdao.api import IndepVarComp, Problem, Group, NewtonSolver, ScipyIterativeSolver, LinearBlockGS, NonlinearBlockGS, DirectSolver, LinearBlockGS, PetscKSP, ScipyOptimizeDriver# TODO, SqliteRecorder, CaseReader, profile
from openmdao.devtools import iprofile
from openmdao.api import view_model
from six import iteritems
from pygeo import DVGeometry
# Create a dictionary to store options about the surface
mesh_dict = {'num_y' : 7,
'num_x' : 3,
'wing_type' : 'CRM',
'symmetry' : True,
'num_twist_cp' : 5,
'span_cos_spacing' : 0.}
mesh, _ = generate_mesh(mesh_dict)
surf_dict = {
# Wing definition
'name' : 'wing', # name of the surface
'symmetry' : True, # if true, model one half of wing
# reflected across the plane y = 0
'S_ref_type' : 'wetted', # how we compute the wing area,
# can be 'wetted' or 'projected'
'fem_model_type' : 'tube',
'DVGeo' : True,
'mesh' : mesh,
'mx' : 2,
'my' : 3,
# Aerodynamic performance of the lifting surface at
# an angle of attack of 0 (alpha=0).
# These CL0 and CD0 values are added to the CL and CD
# obtained from aerodynamic analysis of the surface to get
# the total CL and CD.
# These CL0 and CD0 values do not vary wrt alpha.
'CL0' : 0.0, # CL of the surface at alpha=0
'CD0' : 0.015, # CD of the surface at 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_cp' : np.array([0.15]), # thickness over chord ratio (NACA0015)
'c_max_t' : .303, # chordwise location of maximum (NACA0015)
# thickness
'with_viscous' : True, # if true, compute viscous drag
'with_wave' : False, # if true, compute wave drag
}
surfaces = [surf_dict]
# Create the problem and the model group
prob = Problem()
indep_var_comp = IndepVarComp()
indep_var_comp.add_output('v', val=248.136, units='m/s')
indep_var_comp.add_output('alpha', val=6.64, units='deg')
indep_var_comp.add_output('Mach_number', val=0.84)
indep_var_comp.add_output('re', val=1.e6, units='1/m')
indep_var_comp.add_output('rho', val=0.38, units='kg/m**3')
indep_var_comp.add_output('cg', val=np.zeros((3)), units='m')
prob.model.add_subsystem('prob_vars',
indep_var_comp,
promotes=['*'])
# Loop over each surface in the surfaces list
for surface in surfaces:
filename = write_FFD_file(surface, surface['mx'], surface['my'])
DVGeo = DVGeometry(filename)
geom_group = Geometry(surface=surface, DVGeo=DVGeo)
# Add tmp_group to the problem as the name of the surface.
# Note that is a group and performance group for each
# individual surface.
prob.model.add_subsystem(surface['name'], geom_group)
# Loop through and add a certain number of aero points
for i in range(1):
# Create the aero point group and add it to the model
aero_group = AeroPoint(surfaces=surfaces)
point_name = 'aero_point_{}'.format(i)
prob.model.add_subsystem(point_name, aero_group)
# Connect flow properties to the analysis point
prob.model.connect('v', point_name + '.v')
prob.model.connect('alpha', point_name + '.alpha')
prob.model.connect('Mach_number', point_name + '.Mach_number')
prob.model.connect('re', point_name + '.re')
prob.model.connect('rho', point_name + '.rho')
prob.model.connect('cg', point_name + '.cg')
# Connect the parameters within the model for each aero point
for surface in surfaces:
name = surface['name']
# Connect the mesh from the geometry component to the analysis point
prob.model.connect(name + '.mesh', point_name + '.' + name + '.def_mesh')
# Perform the connections with the modified names within the
# 'aero_states' group.
prob.model.connect(name + '.mesh', point_name + '.aero_states.' + name + '_def_mesh')
prob.model.connect(name + '.t_over_c', point_name + '.' + name + '_perf.' + 't_over_c')
from openmdao.api import pyOptSparseDriver
prob.driver = pyOptSparseDriver()
prob.driver.options['optimizer'] = "SNOPT"
prob.driver.opt_settings = {'Major optimality tolerance': 1.0e-6,
'Major feasibility tolerance': 1.0e-6}
# Setup problem and add design variables, constraint, and objective
prob.model.add_design_var('alpha', lower=-15, upper=15)
prob.model.add_design_var('wing.shape', lower=-3, upper=2)
# prob.model.add_constraint('wing.shape', equals=0., indices=range(surf_dict['my'] * 2), linear=True)
prob.model.add_constraint(point_name + '.wing_perf.CL', equals=0.5)
prob.model.add_objective(point_name + '.wing_perf.CD', scaler=1e4)
# Set up the problem
prob.setup()
# view_model(prob, outfile='aero.html', show_browser=False)
# prob.run_model()
prob.run_driver()
assert_rel_error(self, prob['aero_point_0.wing_perf.CD'][0], 0.03398038, 1e-6)
assert_rel_error(self, prob['aero_point_0.wing_perf.CL'][0], 0.5, 1e-6)
assert_rel_error(self, prob['aero_point_0.CM'][1], -1.7736315914915437, 1e-5)
def test(self):
from openaerostruct.geometry.utils import generate_mesh, write_FFD_file
from openaerostruct.integration.aerostruct_groups import AerostructGeometry, AerostructPoint
import openmdao.api as om
from pygeo import DVGeometry
# Create a dictionary to store options about the surface
mesh_dict = {'num_y' : 5,
'num_x' : 2,
'wing_type' : 'CRM',
'symmetry' : True,
'num_twist_cp' : 5}
mesh, twist_cp = generate_mesh(mesh_dict)
surf_dict = {
# Wing definition
'name' : 'wing', # name of the surface
'symmetry' : True, # if true, model one half of wing
# reflected across the plane y = 0
'S_ref_type' : 'wetted', # how we compute the wing area,
# can be 'wetted' or 'projected'
'fem_model_type' : 'tube',
'thickness_cp' : np.array([.1, .2, .3]),
'mesh' : mesh,
'geom_manipulator' : 'FFD',
'mx' : 2,
'my' : 3,
# Aerodynamic performance of the lifting surface at
# an angle of attack of 0 (alpha=0).
# These CL0 and CD0 values are added to the CL and CD
# obtained from aerodynamic analysis of the surface to get
# the total CL and CD.
# These CL0 and CD0 values do not vary wrt alpha.
'CL0' : 0.0, # CL of the surface at alpha=0
'CD0' : 0.015, # CD of the surface at 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_cp' : np.array([0.15]), # thickness over chord ratio (NACA0015)
'c_max_t' : .303, # chordwise location of maximum (NACA0015)
# thickness
'with_viscous' : True,
'with_wave' : False, # if true, compute wave drag
# Structural values are based on aluminum 7075
'E' : 70.e9, # [Pa] Young's modulus of the spar
'G' : 30.e9, # [Pa] shear modulus of the spar
'yield' : 500.e6 / 2.5, # [Pa] yield stress divided by 2.5 for limiting case
'mrho' : 3.e3, # [kg/m^3] material density
'fem_origin' : 0.35, # normalized chordwise location of the spar
'wing_weight_ratio' : 2.,
'struct_weight_relief' : False, # True to add the weight of the structure to the loads on the structure
'distributed_fuel_weight' : False,
# Constraints
'exact_failure_constraint' : False, # if false, use KS function
}
surfaces = [surf_dict]
# Create the problem and assign the model group
prob = om.Problem()
# Add problem information as an independent variables component
indep_var_comp = om.IndepVarComp()
indep_var_comp.add_output('v', val=248.136, units='m/s')
indep_var_comp.add_output('alpha', val=5., units='deg')
indep_var_comp.add_output('Mach_number', val=0.84)
indep_var_comp.add_output('re', val=1.e6, units='1/m')
indep_var_comp.add_output('rho', val=0.38, units='kg/m**3')
indep_var_comp.add_output('CT', val=grav_constant * 17.e-6, units='1/s')
indep_var_comp.add_output('R', val=11.165e6, units='m')
indep_var_comp.add_output('W0', val=0.4 * 3e5, units='kg')
indep_var_comp.add_output('speed_of_sound', val=295.4, units='m/s')
indep_var_comp.add_output('load_factor', val=1.)
indep_var_comp.add_output('empty_cg', val=np.zeros((3)), units='m')
prob.model.add_subsystem('prob_vars',
indep_var_comp,
promotes=['*'])
# Loop over each surface in the surfaces list
for surface in surfaces:
# Get the surface name and create a group to contain components
# only for this surface
name = surface['name']
filename = write_FFD_file(surface, surface['mx'], surface['my'])
DVGeo = DVGeometry(filename)
aerostruct_group = AerostructGeometry(surface=surface, DVGeo=DVGeo)
# Add tmp_group to the problem with the name of the surface.
prob.model.add_subsystem(name, aerostruct_group)
# Loop through and add a certain number of aero points
for i in range(1):
point_name = 'AS_point_{}'.format(i)
# Connect the parameters within the model for each aero point
# Create the aero point group and add it to the model
AS_point = AerostructPoint(surfaces=surfaces)
prob.model.add_subsystem(point_name, AS_point)
# Connect flow properties to the analysis point
prob.model.connect('v', point_name + '.v')
prob.model.connect('alpha', point_name + '.alpha')
prob.model.connect('Mach_number', point_name + '.Mach_number')
prob.model.connect('re', point_name + '.re')
prob.model.connect('rho', point_name + '.rho')
prob.model.connect('CT', point_name + '.CT')
prob.model.connect('R', point_name + '.R')
prob.model.connect('W0', point_name + '.W0')
prob.model.connect('speed_of_sound', point_name + '.speed_of_sound')
prob.model.connect('empty_cg', point_name + '.empty_cg')
prob.model.connect('load_factor', point_name + '.load_factor')
for surface in surfaces:
com_name = point_name + '.' + name + '_perf'
prob.model.connect(name + '.local_stiff_transformed', point_name + '.coupled.' + name + '.local_stiff_transformed')
prob.model.connect(name + '.nodes', point_name + '.coupled.' + name + '.nodes')
# Connect aerodyamic mesh to coupled group mesh
prob.model.connect(name + '.mesh', point_name + '.coupled.' + name + '.mesh')
# Connect performance calculation variables
prob.model.connect(name + '.radius', com_name + '.radius')
prob.model.connect(name + '.thickness', com_name + '.thickness')
prob.model.connect(name + '.nodes', com_name + '.nodes')
prob.model.connect(name + '.cg_location', point_name + '.' + 'total_perf.' + name + '_cg_location')
prob.model.connect(name + '.structural_mass', point_name + '.' + 'total_perf.' + name + '_structural_mass')
prob.model.connect(name + '.t_over_c', com_name + '.t_over_c')
# Import the Scipy Optimizer and set the driver of the problem to use
# it, which defaults to an SLSQP optimization method
prob.driver = om.ScipyOptimizeDriver()
recorder = om.SqliteRecorder("aerostruct_ffd.db")
prob.driver.add_recorder(recorder)
prob.driver.recording_options['record_derivatives'] = True
prob.driver.recording_options['includes'] = ['*']
# Setup problem and add design variables, constraint, and objective
prob.model.add_design_var('wing.shape', lower=-3, upper=2)
prob.model.add_design_var('wing.thickness_cp', lower=0.01, upper=0.5, scaler=1e2)
prob.model.add_constraint('AS_point_0.wing_perf.failure', upper=0.)
prob.model.add_constraint('AS_point_0.wing_perf.thickness_intersects', upper=0.)
# Add design variables, constraisnt, and objective on the problem
prob.model.add_design_var('alpha', lower=-10., upper=10.)
prob.model.add_constraint('AS_point_0.L_equals_W', equals=0.)
prob.model.add_objective('AS_point_0.fuelburn', scaler=1e-5)
# iprofile.setup()
# iprofile.start()
# Set up the problem
prob.setup()
# om.view_model(prob, outfile='aerostruct_ffd', show_browser=False)
# prob.run_model()
prob.run_driver()
# prob.check_partials(compact_print=True)
# print("\nWing CL:", prob['aero_point_0.wing_perf.CL'])
# print("Wing CD:", prob['aero_point_0.wing_perf.CD'])
# from helper import plot_3d_points
#
# mesh = prob['aero_point_0.wing.def_mesh']
# plot_3d_points(mesh)
#
# filename = mesh_dict['wing_type'] + '_' + str(mesh_dict['num_x']) + '_' + str(mesh_dict['num_y'])
# filename += '_' + str(surf_dict['mx']) + '_' + str(surf_dict['my']) + '.mesh'
# np.save(filename, mesh)
assert_rel_error(self, prob['AS_point_0.fuelburn'][0], 97680.8964568375, 1e-3)
def test(self):
from openaerostruct.geometry.utils import generate_mesh, write_FFD_file
from openaerostruct.geometry.geometry_group import Geometry
from openaerostruct.transfer.displacement_transfer import DisplacementTransfer
from openaerostruct.aerodynamics.aero_groups import AeroPoint
from openaerostruct.integration.multipoint_comps import MultiCD
from openmdao.api import IndepVarComp, Problem, Group, NewtonSolver, ScipyIterativeSolver, LinearBlockGS, NonlinearBlockGS, DirectSolver, LinearBlockGS, PetscKSP, ScipyOptimizeDriver, ExplicitComponent# TODO, SqliteRecorder, CaseReader, profile
from openmdao.devtools import iprofile
from openmdao.api import view_model
from openmdao.utils.assert_utils import assert_check_partials
from pygeo import DVGeometry
# Create a dictionary to store options about the surface
mesh_dict = {'num_y' : 5,
'num_x' : 3,
'wing_type' : 'CRM',
'symmetry' : True,
'num_twist_cp' : 5,
'span_cos_spacing' : 0.}
mesh, _ = generate_mesh(mesh_dict)
surf_dict = {
# Wing definition
'name' : 'wing', # name of the surface
'type' : 'aero',
'symmetry' : True, # if true, model one half of wing
# reflected across the plane y = 0
'S_ref_type' : 'wetted', # how we compute the wing area,
# can be 'wetted' or 'projected'
'fem_model_type' : 'tube',
'mesh' : mesh,
'mx' : 2,
'my' : 3,
'geom_manipulator' : 'FFD',
# Aerodynamic performance of the lifting surface at
# an angle of attack of 0 (alpha=0).
# These CL0 and CD0 values are added to the CL and CD
# obtained from aerodynamic analysis of the surface to get
# the total CL and CD.
# These CL0 and CD0 values do not vary wrt alpha.
'CL0' : 0.0, # CL of the surface at alpha=0
'CD0' : 0.015, # CD of the surface at 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_cp' : np.array([0.15]), # thickness over chord ratio (NACA0015)
'c_max_t' : .303, # chordwise location of maximum (NACA0015)
# thickness
'with_viscous' : True, # if true, compute viscous drag
'with_wave' : False, # if true, compute wave drag
}
surfaces = [surf_dict]
n_points = 2
# Create the problem and the model group
prob = Problem()
indep_var_comp = IndepVarComp()
indep_var_comp.add_output('v', val=248.136, units='m/s')
indep_var_comp.add_output('alpha', val=np.ones(n_points)*6.64, units='deg')
indep_var_comp.add_output('M', val=0.84)
indep_var_comp.add_output('re', val=1.e6, units='1/m')
indep_var_comp.add_output('rho', val=0.38, units='kg/m**3')
indep_var_comp.add_output('cg', val=np.zeros((3)), units='m')
prob.model.add_subsystem('prob_vars',
indep_var_comp,
promotes=['*'])
# Loop through and add a certain number of aero points
for i in range(n_points):
# Create the aero point group and add it to the model
aero_group = AeroPoint(surfaces=surfaces)
point_name = 'aero_point_{}'.format(i)
prob.model.add_subsystem(point_name, aero_group)
# Connect flow properties to the analysis point
prob.model.connect('v', point_name + '.v')
prob.model.connect('alpha', point_name + '.alpha', src_indices=[i])
prob.model.connect('M', point_name + '.M')
prob.model.connect('re', point_name + '.re')
prob.model.connect('rho', point_name + '.rho')
prob.model.connect('cg', point_name + '.cg')
# Connect the parameters within the model for each aero point
for surface in surfaces:
filename = write_FFD_file(surface, surface['mx'], surface['my'])
DVGeo = DVGeometry(filename)
geom_group = Geometry(surface=surface, DVGeo=DVGeo)
# Add tmp_group to the problem as the name of the surface.
# Note that is a group and performance group for each
# individual surface.
aero_group.add_subsystem(surface['name'] + '_geom', geom_group)
name = surface['name']
prob.model.connect(point_name + '.CD', 'multi_CD.' + str(i) + '_CD')
# Connect the mesh from the geometry component to the analysis point
prob.model.connect(point_name + '.' + name + '_geom.mesh', point_name + '.' + name + '.def_mesh')
# Perform the connections with the modified names within the
# 'aero_states' group.
prob.model.connect(point_name + '.' + name + '_geom.mesh', point_name + '.aero_states.' + name + '_def_mesh')
prob.model.connect(point_name + '.' + name + '_geom.t_over_c', point_name + '.' + name + '_perf.' + 't_over_c')
prob.model.add_subsystem('multi_CD', MultiCD(n_points=n_points), promotes_outputs=['CD'])
from openmdao.api import pyOptSparseDriver
prob.driver = pyOptSparseDriver()
prob.driver.options['optimizer'] = "SNOPT"
prob.driver.opt_settings = {'Major optimality tolerance': 1.0e-5,
'Major feasibility tolerance': 1.0e-5}
# # Setup problem and add design variables, constraint, and objective
prob.model.add_design_var('alpha', lower=-15, upper=15)
prob.model.add_design_var('aero_point_0.wing_geom.shape', lower=-3, upper=2)
prob.model.add_constraint('aero_point_0.wing_perf.CL', equals=0.45)
prob.model.add_design_var('aero_point_1.wing_geom.shape', lower=-3, upper=2)
prob.model.add_constraint('aero_point_1.wing_perf.CL', equals=0.5)
prob.model.add_objective('CD', scaler=1e4)
# Set up the problem
prob.setup()
prob.run_model()
# Check the partials at this point in the design space
data = prob.check_partials(compact_print=True, out_stream=None, method='cs', step=1e-40)
assert_check_partials(data, atol=1e20, rtol=1e-6)
def test(self):
from openaerostruct.geometry.utils import generate_mesh, write_FFD_file
from openaerostruct.geometry.geometry_group import Geometry
from openaerostruct.aerodynamics.aero_groups import AeroPoint
from openaerostruct.integration.multipoint_comps import MultiCD
import openmdao.api as om
from openmdao.utils.assert_utils import assert_check_partials
from pygeo import DVGeometry
# Create a dictionary to store options about the surface
mesh_dict = {
'num_y': 5,
'num_x': 3,
'wing_type': 'CRM',
'symmetry': True,
'num_twist_cp': 5,
'span_cos_spacing': 0.0,
}
mesh, _ = generate_mesh(mesh_dict)
surf_dict = {
# Wing definition
'name': 'wing', # name of the surface
'symmetry': True, # if true, model one half of wing
# reflected across the plane y = 0
'S_ref_type': 'wetted', # how we compute the wing area,
# can be 'wetted' or 'projected'
'fem_model_type': 'tube',
'mesh': mesh,
'mx': 2,
'my': 3,
'geom_manipulator': 'FFD',
# Aerodynamic performance of the lifting surface at
# an angle of attack of 0 (alpha=0).
# These CL0 and CD0 values are added to the CL and CD
# obtained from aerodynamic analysis of the surface to get
# the total CL and CD.
# These CL0 and CD0 values do not vary wrt alpha.
'CL0': 0.0, # CL of the surface at alpha=0
'CD0': 0.015, # CD of the surface at 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_cp':
np.array([0.15]), # thickness over chord ratio (NACA0015)
'c_max_t': 0.303, # chordwise location of maximum (NACA0015)
# thickness
'with_viscous': True, # if true, compute viscous drag
'with_wave': False, # if true, compute wave drag
}
surfaces = [surf_dict]
n_points = 2
# Create the problem and the model group
prob = om.Problem()
indep_var_comp = om.IndepVarComp()
indep_var_comp.add_output('v', val=248.136, units='m/s')
indep_var_comp.add_output('alpha',
val=np.ones(n_points) * 6.64,
units='deg')
indep_var_comp.add_output('Mach_number', val=0.84)
indep_var_comp.add_output('re', val=1.0e6, units='1/m')
indep_var_comp.add_output('rho', val=0.38, units='kg/m**3')
indep_var_comp.add_output('cg', val=np.zeros((3)), units='m')
prob.model.add_subsystem('prob_vars', indep_var_comp, promotes=['*'])
# Loop over each surface in the surfaces list
for surface in surfaces:
# Get the surface name and create a group to contain components
# only for this surface
name = surface['name']
# FFD setup
filename = write_FFD_file(surface, surface['mx'], surface['my'])
DVGeo = DVGeometry(filename)
geom_group = Geometry(surface=surface, DVGeo=DVGeo)
# Add tmp_group to the problem with the name of the surface.
prob.model.add_subsystem(name + '_geom', geom_group)
# Loop through and add a certain number of aero points
for i in range(n_points):
# Create the aero point group and add it to the model
aero_group = AeroPoint(surfaces=surfaces)
point_name = 'aero_point_{}'.format(i)
prob.model.add_subsystem(point_name, aero_group)
# Connect flow properties to the analysis point
prob.model.connect('v', point_name + '.v')
prob.model.connect('alpha', point_name + '.alpha', src_indices=[i])
prob.model.connect('Mach_number', point_name + '.Mach_number')
prob.model.connect('re', point_name + '.re')
prob.model.connect('rho', point_name + '.rho')
prob.model.connect('cg', point_name + '.cg')
# Connect the parameters within the model for each aero point
for surface in surfaces:
name = surface['name']
# Connect the drag coeff at this point to the multi_CD component, which does the summation.
prob.model.connect(point_name + '.CD',
'multi_CD.' + str(i) + '_CD')
# Connect the mesh from the geometry component to the analysis point
prob.model.connect(name + '_geom.mesh',
point_name + '.' + name + '.def_mesh')
# Perform the connections with the modified names within the
# 'aero_states' group.
prob.model.connect(
name + '_geom.mesh',
point_name + '.aero_states.' + name + '_def_mesh')
prob.model.connect(
name + '_geom.t_over_c',
point_name + '.' + name + '_perf.' + 't_over_c')
prob.model.add_subsystem('multi_CD',
MultiCD(n_points=n_points),
promotes_outputs=['CD'])
prob.driver = om.ScipyOptimizeDriver()
# Setup problem and add design variables, constraint, and objective
# design variable is the wing shape, and angle-of-attack at each point.
prob.model.add_design_var('alpha', lower=-15, upper=15)
prob.model.add_design_var('wing_geom.shape', lower=-3, upper=2)
# set different target CL value at each point.
prob.model.add_constraint('aero_point_0.wing_perf.CL', equals=0.45)
prob.model.add_constraint('aero_point_1.wing_perf.CL', equals=0.5)
# objective is the sum of CDs at each point.
prob.model.add_objective('CD', scaler=1e4)
# Set up the problem
prob.setup()
prob.run_model()
# Check the partials at this point in the design space
data = prob.check_partials(compact_print=True,
out_stream=None,
method='fd',
step=1e-5)
assert_check_partials(data, atol=1e20, rtol=1e-3)
def test(self):
from openaerostruct.geometry.utils import generate_mesh, write_FFD_file
from openaerostruct.integration.aerostruct_groups import AerostructGeometry, AerostructPoint
from openmdao.api import IndepVarComp, Problem, Group, SqliteRecorder
from pygeo import DVGeometry
# Create a dictionary to store options about the surface
mesh_dict = {'num_y' : 5,
'num_x' : 2,
'wing_type' : 'CRM',
'symmetry' : True,
'num_twist_cp' : 5}
mesh, twist_cp = generate_mesh(mesh_dict)
surf_dict = {
# Wing definition
'name' : 'wing', # name of the surface
'type' : 'aerostruct',
'symmetry' : True, # if true, model one half of wing
# reflected across the plane y = 0
'S_ref_type' : 'wetted', # how we compute the wing area,
# can be 'wetted' or 'projected'
'fem_model_type' : 'tube',
'thickness_cp' : np.array([.1, .2, .3]),
'mesh' : mesh,
'geom_manipulator' : 'FFD',
'mx' : 2,
'my' : 3,
# Aerodynamic performance of the lifting surface at
# an angle of attack of 0 (alpha=0).
# These CL0 and CD0 values are added to the CL and CD
# obtained from aerodynamic analysis of the surface to get
# the total CL and CD.
# These CL0 and CD0 values do not vary wrt alpha.
'CL0' : 0.0, # CL of the surface at alpha=0
'CD0' : 0.015, # CD of the surface at 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_cp' : np.array([0.15]), # thickness over chord ratio (NACA0015)
'c_max_t' : .303, # chordwise location of maximum (NACA0015)
# thickness
'with_viscous' : True,
'with_wave' : False, # if true, compute wave drag
# Structural values are based on aluminum 7075
'E' : 70.e9, # [Pa] Young's modulus of the spar
'G' : 30.e9, # [Pa] shear modulus of the spar
'yield' : 500.e6 / 2.5, # [Pa] yield stress divided by 2.5 for limiting case
'mrho' : 3.e3, # [kg/m^3] material density
'fem_origin' : 0.35, # normalized chordwise location of the spar
'wing_weight_ratio' : 2.,
'struct_weight_relief' : False, # True to add the weight of the structure to the loads on the structure
'distributed_fuel_weight' : False,
# Constraints
'exact_failure_constraint' : False, # if false, use KS function
}
surfaces = [surf_dict]
# Create the problem and assign the model group
prob = Problem()
# Add problem information as an independent variables component
indep_var_comp = IndepVarComp()
indep_var_comp.add_output('v', val=248.136, units='m/s')
indep_var_comp.add_output('alpha', val=5., units='deg')
indep_var_comp.add_output('M', val=0.84)
indep_var_comp.add_output('re', val=1.e6, units='1/m')
indep_var_comp.add_output('rho', val=0.38, units='kg/m**3')
indep_var_comp.add_output('CT', val=9.80665 * 17.e-6, units='1/s')
indep_var_comp.add_output('R', val=11.165e6, units='m')
indep_var_comp.add_output('W0', val=0.4 * 3e5, units='kg')
indep_var_comp.add_output('a', val=295.4, units='m/s')
indep_var_comp.add_output('load_factor', val=1.)
indep_var_comp.add_output('empty_cg', val=np.zeros((3)), units='m')
prob.model.add_subsystem('prob_vars',
indep_var_comp,
promotes=['*'])
# Loop over each surface in the surfaces list
for surface in surfaces:
# Get the surface name and create a group to contain components
# only for this surface
name = surface['name']
filename = write_FFD_file(surface, surface['mx'], surface['my'])
DVGeo = DVGeometry(filename)
aerostruct_group = AerostructGeometry(surface=surface, DVGeo=DVGeo)
# Add tmp_group to the problem with the name of the surface.
prob.model.add_subsystem(name, aerostruct_group)
# Loop through and add a certain number of aero points
for i in range(1):
point_name = 'AS_point_{}'.format(i)
# Connect the parameters within the model for each aero point
# Create the aero point group and add it to the model
AS_point = AerostructPoint(surfaces=surfaces)
prob.model.add_subsystem(point_name, AS_point)
# Connect flow properties to the analysis point
prob.model.connect('v', point_name + '.v')
prob.model.connect('alpha', point_name + '.alpha')
prob.model.connect('M', point_name + '.M')
prob.model.connect('re', point_name + '.re')
prob.model.connect('rho', point_name + '.rho')
prob.model.connect('CT', point_name + '.CT')
prob.model.connect('R', point_name + '.R')
prob.model.connect('W0', point_name + '.W0')
prob.model.connect('a', point_name + '.a')
prob.model.connect('empty_cg', point_name + '.empty_cg')
prob.model.connect('load_factor', point_name + '.load_factor')
for surface in surfaces:
prob.model.connect('load_factor', name + '.load_factor')
com_name = point_name + '.' + name + '_perf'
prob.model.connect(name + '.K', point_name + '.coupled.' + name + '.K')
prob.model.connect(name + '.nodes', point_name + '.coupled.' + name + '.nodes')
# Connect aerodyamic mesh to coupled group mesh
prob.model.connect(name + '.mesh', point_name + '.coupled.' + name + '.mesh')
# Connect performance calculation variables
prob.model.connect(name + '.radius', com_name + '.radius')
prob.model.connect(name + '.thickness', com_name + '.thickness')
prob.model.connect(name + '.nodes', com_name + '.nodes')
prob.model.connect(name + '.cg_location', point_name + '.' + 'total_perf.' + name + '_cg_location')
prob.model.connect(name + '.structural_weight', point_name + '.' + 'total_perf.' + name + '_structural_weight')
prob.model.connect(name + '.t_over_c', com_name + '.t_over_c')
# Import the Scipy Optimizer and set the driver of the problem to use
# it, which defaults to an SLSQP optimization method
from openmdao.api import ScipyOptimizeDriver
prob.driver = ScipyOptimizeDriver()
recorder = SqliteRecorder("aerostruct_ffd.db")
prob.driver.add_recorder(recorder)
prob.driver.recording_options['record_derivatives'] = True
prob.driver.recording_options['includes'] = ['*']
# Setup problem and add design variables, constraint, and objective
prob.model.add_design_var('wing.shape', lower=-3, upper=2)
prob.model.add_design_var('wing.thickness_cp', lower=0.01, upper=0.5, scaler=1e2)
prob.model.add_constraint('AS_point_0.wing_perf.failure', upper=0.)
prob.model.add_constraint('AS_point_0.wing_perf.thickness_intersects', upper=0.)
# Add design variables, constraisnt, and objective on the problem
prob.model.add_design_var('alpha', lower=-10., upper=10.)
prob.model.add_constraint('AS_point_0.L_equals_W', equals=0.)
prob.model.add_objective('AS_point_0.fuelburn', scaler=1e-5)
# iprofile.setup()
# iprofile.start()
# Set up the problem
prob.setup()
# from openmdao.api import view_model
# view_model(prob, outfile='aerostruct_ffd', show_browser=False)
# prob.run_model()
prob.run_driver()
# prob.check_partials(compact_print=True)
# print("\nWing CL:", prob['aero_point_0.wing_perf.CL'])
# print("Wing CD:", prob['aero_point_0.wing_perf.CD'])
# from helper import plot_3d_points
#
# mesh = prob['aero_point_0.wing.def_mesh']
# plot_3d_points(mesh)
#
# filename = mesh_dict['wing_type'] + '_' + str(mesh_dict['num_x']) + '_' + str(mesh_dict['num_y'])
# filename += '_' + str(surf_dict['mx']) + '_' + str(surf_dict['my']) + '.mesh'
# np.save(filename, mesh)
assert_rel_error(self, prob['AS_point_0.fuelburn'][0], 104675.0989232741, 1e-3)
def test(self):
from openaerostruct.geometry.utils import generate_mesh, write_FFD_file
from openaerostruct.geometry.geometry_group import Geometry
from openaerostruct.transfer.displacement_transfer import DisplacementTransfer
from openaerostruct.aerodynamics.aero_groups import AeroPoint
from openaerostruct.integration.multipoint_comps import MultiCD, GeomMatch
from openmdao.api import IndepVarComp, Problem, Group, NewtonSolver, ScipyIterativeSolver, LinearBlockGS, NonlinearBlockGS, DirectSolver, LinearBlockGS, PetscKSP, ScipyOptimizeDriver, ExplicitComponent# TODO, SqliteRecorder, CaseReader, profile
from openmdao.devtools import iprofile
from openmdao.api import view_model
from pygeo import DVGeometry
# Create a dictionary to store options about the surface
mesh_dict = {'num_y' : 5,
'num_x' : 3,
'wing_type' : 'CRM',
'symmetry' : True,
'num_twist_cp' : 5,
'span_cos_spacing' : 0.}
mesh, _ = generate_mesh(mesh_dict)
surf_dict = {
# Wing definition
'name' : 'wing', # name of the surface
'type' : 'aero',
'symmetry' : True, # if true, model one half of wing
# reflected across the plane y = 0
'S_ref_type' : 'wetted', # how we compute the wing area,
# can be 'wetted' or 'projected'
'fem_model_type' : 'tube',
'mesh' : mesh,
'mx' : 2,
'my' : 3,
'geom_manipulator' : 'FFD',
# Aerodynamic performance of the lifting surface at
# an angle of attack of 0 (alpha=0).
# These CL0 and CD0 values are added to the CL and CD
# obtained from aerodynamic analysis of the surface to get
# the total CL and CD.
# These CL0 and CD0 values do not vary wrt alpha.
'CL0' : 0.0, # CL of the surface at alpha=0
'CD0' : 0.015, # CD of the surface at 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' : 0.15, # thickness over chord ratio (NACA0015)
'c_max_t' : .303, # chordwise location of maximum (NACA0015)
# thickness
'with_viscous' : True, # if true, compute viscous drag
}
surf_dict['num_x'], surf_dict['num_y'] = surf_dict['mesh'].shape[:2]
surfaces = [surf_dict]
n_points = 2
# Create the problem and the model group
prob = Problem()
indep_var_comp = IndepVarComp()
indep_var_comp.add_output('v', val=248.136, units='m/s')
indep_var_comp.add_output('alpha', val=np.ones(n_points)*6.64)
indep_var_comp.add_output('M', val=0.84)
indep_var_comp.add_output('re', val=1.e6, units='1/m')
indep_var_comp.add_output('rho', val=0.38, units='kg/m**3')
indep_var_comp.add_output('cg', val=np.zeros((3)), units='m')
prob.model.add_subsystem('prob_vars',
indep_var_comp,
promotes=['*'])
# Loop through and add a certain number of aero points
for i in range(n_points):
# Create the aero point group and add it to the model
aero_group = AeroPoint(surfaces=surfaces)
point_name = 'aero_point_{}'.format(i)
prob.model.add_subsystem(point_name, aero_group)
# Connect flow properties to the analysis point
prob.model.connect('v', point_name + '.v')
prob.model.connect('alpha', point_name + '.alpha', src_indices=[i])
prob.model.connect('M', point_name + '.M')
prob.model.connect('re', point_name + '.re')
prob.model.connect('rho', point_name + '.rho')
prob.model.connect('cg', point_name + '.cg')
# Connect the parameters within the model for each aero point
for surface in surfaces:
filename = write_FFD_file(surface, surface['mx'], surface['my'])
DVGeo = DVGeometry(filename)
geom_group = Geometry(surface=surface, DVGeo=DVGeo)
# Add tmp_group to the problem as the name of the surface.
# Note that is a group and performance group for each
# individual surface.
aero_group.add_subsystem(surface['name'] + '_geom', geom_group)
name = surface['name']
prob.model.connect(point_name + '.CD', 'multi_CD.' + str(i) + '_CD')
prob.model.connect(point_name + '.wing_geom.shape', 'geom_match.' + str(i) + '_shape')
# Connect the mesh from the geometry component to the analysis point
prob.model.connect(point_name + '.' + name + '_geom.mesh', point_name + '.' + name + '.def_mesh')
# Perform the connections with the modified names within the
# 'aero_states' group.
prob.model.connect(point_name + '.' + name + '_geom.mesh', point_name + '.aero_states.' + name + '_def_mesh')
prob.model.add_subsystem('multi_CD', MultiCD(n_points=n_points), promotes_outputs=['CD'])
prob.model.add_subsystem('geom_match', GeomMatch(n_points=n_points, mx=surf_dict['mx'], my=surf_dict['my']), promotes_outputs=['shape_diff'])
from openmdao.api import pyOptSparseDriver
prob.driver = pyOptSparseDriver()
prob.driver.options['optimizer'] = "SNOPT"
prob.driver.opt_settings = {'Major optimality tolerance': 1.0e-5,
'Major feasibility tolerance': 1.0e-5}
# # Setup problem and add design variables, constraint, and objective
prob.model.add_design_var('alpha', lower=-15, upper=15)
prob.model.add_design_var('aero_point_0.wing_geom.shape', lower=-3, upper=2)
prob.model.add_constraint('aero_point_0.wing_perf.CL', equals=0.45)
prob.model.add_design_var('aero_point_1.wing_geom.shape', lower=-3, upper=2)
prob.model.add_constraint('aero_point_1.wing_perf.CL', equals=0.5)
prob.model.add_constraint('shape_diff', equals=0., indices=range(surf_dict['my'] * (n_points - 1) * 1), linear=True)
# prob.model.add_constraint('shape_diff', equals=0., linear=True)
prob.model.add_objective('CD', scaler=1e4)
# Set up the problem
prob.setup()
# prob.run_model()
prob.run_driver()
# prob.check_partials(compact_print=True)
assert_rel_error(self, prob['aero_point_0.wing_perf.CL'][0], 0.45, 1e-6)
assert_rel_error(self, prob['aero_point_0.wing_perf.CD'][0], 0.03229527496681374, 1e-6)
assert_rel_error(self, prob['aero_point_1.wing_perf.CL'][0], 0.5, 1e-6)
assert_rel_error(self, prob['aero_point_1.wing_perf.CD'][0], 0.033767398588947985, 1e-6)
