def initialize(self,vehicle):
# unpack
conditions_table = self.conditions_table
geometry = self.geometry
configuration = self.configuration
#
AoA = conditions_table.angle_of_attack
n_conditions = len(AoA)
# copy geometry
for k in ['fuselages','wings','propulsors']:
geometry[k] = deepcopy(vehicle[k])
# reference area
geometry.reference_area = vehicle.reference_area
# arrays
CL = np.zeros_like(AoA)
# condition input, local, do not keep
konditions = Conditions()
konditions.aerodynamics = Data()
# calculate aerodynamics for table
for i in xrange(n_conditions):
# overriding conditions, thus the name mangling
konditions.aerodynamics.angle_of_attack = AoA[i]
# these functions are inherited from Aerodynamics() or overridden
CL[i] = calculate_lift_vortex_lattice(konditions, configuration, geometry)
# store table
conditions_table.lift_coefficient = CL
# build surrogate
self.build_surrogate_sub()
# calculate aerodynamics for table
for i in xrange(n_conditions):
# overriding conditions, thus the name mangling
konditions.aerodynamics.angle_of_attack = AoA[i]
# these functions are inherited from Aerodynamics() or overridden
CL[i] = calculate_lift_linear_supersonic(konditions, configuration, geometry)
# store table
conditions_table.lift_coefficient = CL
# build surrogate
self.build_surrogate_sup()
return
def __defaults__(self):
self.unknowns = Unknowns()
self.conditions = Conditions()
self.residuals = Residuals()
self.numerics = Numerics()
self.initials = Conditions()
def initialize(self):
# build the conditions table
self.build_conditions_table()
# unpack
conditions_table = self.conditions_table
geometry = self.geometry
configuration = self.configuration
#
AoA = conditions_table.angle_of_attack
#Ma = conditions_table.mach_number
#Re = conditions_table.reynolds_number
# check
#assert len(AoA) == len(Ma) == len(Re) , 'condition length mismatch'
n_conditions = len(AoA)
# arrays
CL = np.zeros_like(AoA)
CD = np.zeros_like(AoA)
# condition input, local, do not keep
konditions = Conditions()
# calculate aerodynamics for table
for i in xrange(n_conditions):
# overriding conditions, thus the name mangling
konditions.angle_of_attack = AoA[i]
#konditions.mach_number = Ma[i]
#konditions.reynolds_number = Re[i]
# these functions are inherited from Aerodynamics() or overridden
CL[i] = self.calculate_lift(konditions, configuration, geometry)
CD[i] = self.calculate_drag(konditions, configuration, geometry)
# store table
conditions_table.lift_coefficient = CL
conditions_table.drag_coefficient = CD
# build surrogate
self.build_surrogate()
return
def initialize(self):
# build the conditions table
self.build_conditions_table()
# unpack
conditions_table = self.conditions_table
geometry = self.geometry
configuration = self.configuration
#
AoA = conditions_table.angle_of_attack
Ma = conditions_table.mach_number
Re = conditions_table.reynolds_number
# check
assert len(AoA) == len(Ma) == len(Re), 'condition length mismatch'
n_conditions = len(AoA)
# arrays
CL = np.zeros_like(AoA)
CD = np.zeros_like(AoA)
# condition input, local, do not keep
konditions = Conditions()
# calculate aerodynamics for table
for i in xrange(n_conditions):
# overriding conditions, thus the name mangling
konditions.angle_of_attack = AoA[i]
konditions.mach_number = Ma[i]
konditions.reynolds_number = Re[i]
# these functions are inherited from Aerodynamics() or overridden
CL[i] = self.calculate_lift(konditions, configuration, geometry)
CD[i] = self.calculate_drag(konditions, configuration, geometry)
# store table
conditions_table.lift_coefficient = CL
conditions_table.drag_coefficient = CD
# build surrogate
self.build_surrogate()
return
def __defaults__(self):
self.tag = 'numerics'
self.number_control_points = 16
self.discretization_method = chebyshev_data
self.solver_jacobian = "none"
self.tolerance_solution = 1e-8
self.tolerance_boundary_conditions = 1e-8
self.dimensionless = Conditions()
self.dimensionless.control_points = np.empty([0, 0])
self.dimensionless.differentiate = np.empty([0, 0])
self.dimensionless.integrate = np.empty([0, 0])
self.time = Conditions()
self.time.control_points = np.empty([0, 0])
self.time.differentiate = np.empty([0, 0])
self.time.integrate = np.empty([0, 0])
def __defaults__(self):
self.tag = 'Aerodynamics'
self.geometry = Geometry()
self.configuration = Configuration()
self.conditions_table = Conditions(
angle_of_attack=np.linspace(-10., 10., 5) * Units.deg, )
self.model = Data()
def __defaults__(self):
self.tag = 'basic_conditions'
# start default row vectors
ones_1col = self.ones_row(1)
ones_2col = self.ones_row(2)
ones_3col = self.ones_row(3)
# top level conditions
self.frames = Conditions()
self.weights = Conditions()
self.energies = Conditions()
# inertial conditions
self.frames.inertial = Conditions()
self.frames.inertial.position_vector = ones_3col * 0
self.frames.inertial.velocity_vector = ones_3col * 0
self.frames.inertial.acceleration_vector = ones_3col * 0
self.frames.inertial.gravity_force_vector = ones_3col * 0
self.frames.inertial.total_force_vector = ones_3col * 0
self.frames.inertial.time = ones_1col * 0
# body conditions
self.frames.body = Conditions()
self.frames.body.inertial_rotations = ones_3col * 0
self.frames.body.transform_to_inertial = np.empty([0, 0, 0])
# weights conditions
self.weights.total_mass = ones_1col * 0
self.weights.weight_breakdown = Conditions()
# energy conditions
self.energies.total_energy = ones_1col * 0
self.energies.total_efficiency = ones_1col * 0
def __defaults__(self):
self.tag = 'Aerodynamics'
self.geometry = Geometry()
self.configuration = Configuration()
self.conditions_table = Conditions(
angle_of_attack=np.linspace(-10., 10., 5) * Units.deg,
mach_number=np.linspace(0.0, 1.0, 5),
reynolds_number=np.linspace(1.e4, 1.e10, 3),
)
self.model = Data()
def __defaults__(self):
"""This sets the default values.
Assumptions:
None
Source:
N/A
Inputs:
None
Outputs:
None
Properties Used:
None
"""
self.tag = 'numerics'
self.number_control_points = 16
self.discretization_method = chebyshev_data
self.solver_jacobian = "none"
self.tolerance_solution = 1e-8
self.tolerance_boundary_conditions = 1e-8
self.converged = None
self.dimensionless = Conditions()
self.dimensionless.control_points = np.empty([0, 0])
self.dimensionless.differentiate = np.empty([0, 0])
self.dimensionless.integrate = np.empty([0, 0])
self.time = Conditions()
self.time.control_points = np.empty([0, 0])
self.time.differentiate = np.empty([0, 0])
self.time.integrate = np.empty([0, 0])
def __defaults__(self):
""" This sets the default values.
Assumptions:
None
Source:
N/A
Inputs:
None
Outputs:
None
Properties Used:
None
"""
self.unknowns = Unknowns()
self.conditions = Conditions()
self.residuals = Residuals()
self.numerics = Numerics()
self.initials = Conditions()
class Container(State):
def __defaults__(self):
self.segments = Conditions()
def merged(self):
state_out = State()
for i,(tag,sub_state) in enumerate(self.segments.items()):
for key in ['unknowns','conditions','residuals']:
if i == 0:
state_out[key].update(sub_state[key])
else:
state_out[key] = state_out[key].do_recursive(append_array,sub_state[key])
return state_out
def getCurrent(self):
cmd = b'LOOP 1\n'
current = None
try:
self.port.write(cmd)
self.lastActv = datetime.datetime.now()
data = self.read(100)
if len(data) != 100:
return None
data = bytearray(data)
if data[0] == 0x06:
data.pop(0)
if VPCRC.validate(data):
current = Conditions(data)
except Exception as e:
Logger.error(e)
return current
def __defaults__(self):
self.tag = 'Fidelity_Zero'
self.geometry = Geometry()
self.configuration = Configuration()
# correction factors
self.configuration.fuselage_lift_correction = 1.14
self.configuration.trim_drag_correction_factor = 1.02
self.configuration.wing_parasite_drag_form_factor = 1.1
self.configuration.fuselage_parasite_drag_form_factor = 2.3
self.configuration.aircraft_span_efficiency_factor = 0.78
# vortex lattice configurations
self.configuration.number_panels_spanwise = 5
self.configuration.number_panels_chordwise = 1
self.conditions_table = Conditions(
angle_of_attack=np.array([-10., -5., 0., 5., 10.]) * Units.deg, )
self.models = Data()
def __defaults__(self):
self.tag = 'aerodynamic_conditions'
# start default row vectors
ones_1col = self.ones_row(1)
ones_2col = self.ones_row(2)
ones_3col = self.ones_row(3)
# wind frame conditions
self.frames.wind = Conditions()
self.frames.wind.body_rotations = ones_3col * 0 # rotations in [X,Y,Z] -> [phi,theta,psi]
self.frames.wind.velocity_vector = ones_3col * 0
self.frames.wind.lift_force_vector = ones_3col * 0
self.frames.wind.drag_force_vector = ones_3col * 0
self.frames.wind.transform_to_inertial = np.empty([0,0,0])
# body frame conditions
self.frames.body.thrust_force_vector = ones_3col * 0
# planet frame conditions
self.frames.planet = Conditions()
self.frames.planet.start_time = None
self.frames.planet.latitude = ones_1col * 0
self.frames.planet.longitude = ones_1col * 0
# freestream conditions
self.freestream = Conditions()
self.freestream.velocity = ones_1col * 0
self.freestream.mach_number = ones_1col * 0
self.freestream.pressure = ones_1col * 0
self.freestream.temperature = ones_1col * 0
self.freestream.density = ones_1col * 0
self.freestream.speed_of_sound = ones_1col * 0
self.freestream.dynamic_viscosity = ones_1col * 0
self.freestream.altitude = ones_1col * 0
self.freestream.gravity = ones_1col * 0
self.freestream.reynolds_number = ones_1col * 0
self.freestream.dynamic_pressure = ones_1col * 0
# aerodynamics conditions
self.aerodynamics = Conditions()
self.aerodynamics.angle_of_attack = ones_1col * 0
self.aerodynamics.side_slip_angle = ones_1col * 0
self.aerodynamics.roll_angle = ones_1col * 0
self.aerodynamics.lift_coefficient = ones_1col * 0
self.aerodynamics.drag_coefficient = ones_1col * 0
self.aerodynamics.lift_breakdown = Conditions()
self.aerodynamics.drag_breakdown = Conditions()
self.aerodynamics.drag_breakdown.parasite = Conditions()
self.aerodynamics.drag_breakdown.compressible = Conditions()
# stability conditions
self.stability = Conditions()
self.stability.static = Conditions()
self.stability.dynamic = Conditions()
# propulsion conditions
self.propulsion = Conditions()
self.propulsion.throttle = ones_1col * 0
self.propulsion.battery_energy = ones_1col * 0
self.propulsion.thrust_breakdown = Conditions()
# energy conditions
self.energies.gravity_energy = ones_1col * 0
self.energies.propulsion_power = ones_1col * 0
# weights conditions
self.weights.vehicle_mass_rate = ones_1col * 0
def __defaults__(self):
self.segments = Conditions()
aerodynamics = Aerodynamics_Surrogate()
def calculate_drag(conditions, configuration=None, geometry=None):
return (conditions.angle_of_attack * 2. * np.pi)**2.
def calculate_lift(conditions, configuration=None, geometry=None):
return conditions.angle_of_attack * 2. * np.pi
# monkey patch the models
aerodynamics.calculate_drag = calculate_drag
aerodynamics.calculate_lift = calculate_lift
aerodynamics.initialize()
# test conditions
conditions = Conditions(
angle_of_attack=8.0 * Units.deg,
mach_number=0.5,
reynolds_number=1.e6,
)
# call the surrogate
CL_I, CD_I = aerodynamics.__call__(conditions)
# truth data
CL_T = aerodynamics.calculate_lift(conditions)
CD_T = aerodynamics.calculate_drag(conditions)
print 'CL = %.4f = %.4f' % (CL_I[0], CL_T)
print 'CD = %.4f = %.4f' % (CD_I[0], CD_T)
shape = (0, 0, SCREEN_SHAPE[0], SCREEN_SHAPE[1])
Population = 1000
conditions = Conditions(
gene_weight_probability=0.8,
gene_random_probability=0.1,
genome_disable_probability=0.75,
genome_node_probability=0.03,
genome_connection_probability=0.05,
species_asexual_probability=0.25,
species_interspecies_reproduction_probability=0.001,
species_keep_ratio=.5,
gene_max_weight=1.0,
gene_min_weight=-1.0,
gene_weight_shift=.01,
genome_weight_coefficient=0.4,
genome_disjoint_coefficient=1.0,
genome_excess_coefficient=1.0,
genome_min_divide=20,
species_age_fertility_limit=15,
species_threshold=3.0,
species_keep_champion=True,
species_champion_limit=5,
species_niche_divide_min=0,
population_age_limit=20,
population_size=Population,
app_start_node_depth=0,
app_end_node_depth=100)
sim = DodgingSimulation(9,
5,
