peterdsharpe
commented
4 years ago Whoops! Seems I missed a little syntax error in the example file! I've just committed and pushed a fix to the master branch.
You can either pull the latest changes, or I've copied the fixed example script here for your convenience:
import copy
from aerosandbox import *
opti = cas.Opti() # Initialize an analysis/optimization environment
# Here's how you can make design variables:
def variable(init_val, lb=None, ub=None):
"""
Initialize a scalar design variable.
:param init_val: Initial guess
:param lb: Optional lower bound
:param ub: Optional upper bound
:return: The variable.
"""
var = opti.variable()
opti.set_initial(var, init_val)
if lb is not None:
opti.subject_to(var >= lb)
if ub is not None:
opti.subject_to(var <= ub)
return var
def quasi_variable(val):
"""
Initialize a scalar variable that is fixed at the value you specify. Useful, as the optimization environment will start
creating a computational graph for all downstream quantities of this, so you can later take gradients w.r.t. this.
:param val: Value to set.
:return: The quasi-variable.
"""
var = opti.variable()
opti.set_initial(var, val)
opti.subject_to(var == val)
return var
# Define the 2D properties of all airfoils you want to use
generic_cambered_airfoil = Airfoil(
CL_function=lambda alpha, Re, mach, deflection,: ( # Lift coefficient function
(alpha * np.pi / 180) * (2 * np.pi) + 0.4550
),
CDp_function=lambda alpha, Re, mach, deflection: ( # Profile drag coefficient function
(1 + (alpha / 5) ** 2) * 2 * (0.074 / Re ** 0.2)
),
Cm_function=lambda alpha, Re, mach, deflection: ( # Moment coefficient function
0
)
)
generic_airfoil = Airfoil(
CL_function=lambda alpha, Re, mach, deflection,: ( # Lift coefficient function
(alpha * np.pi / 180) * (2 * np.pi)
),
CDp_function=lambda alpha, Re, mach, deflection: ( # Profile drag coefficient function
(1 + (alpha / 5) ** 2) * 2 * (0.074 / Re ** 0.2)
),
Cm_function=lambda alpha, Re, mach, deflection: ( # Moment coefficient function
0
)
)
# Define the 3D geometry you want to analyze/optimize.
# Here, all distances are in meters and all angles are in degrees.
airplane = Airplane(
name="Peter's Glider",
x_ref=0, # CG location
y_ref=0, # CG location
z_ref=0, # CG location
fuselages=[
Fuselage(
name="Fuselage",
x_le=-0.25,
y_le=0,
z_le=-0.1,
symmetric=False,
xsecs=[
FuselageXSec(
x_c = 0,
y_c=0,
z_c=0,
radius=0
),
FuselageXSec(
x_c = 0.2,
y_c=0,
z_c=0,
radius=0.1
),
FuselageXSec(
x_c = 0.5,
y_c=0,
z_c=0,
radius=0.12
),
FuselageXSec(
x_c = 0.8,
y_c=0,
z_c=0,
radius=0.1
),
FuselageXSec(
x_c = 1,
y_c=0,
z_c=0.08,
radius=0
),
]
)
],
wings=[
Wing(
name="Main Wing",
x_le=0, # Coordinates of the wing's leading edge
y_le=0, # Coordinates of the wing's leading edge
z_le=0, # Coordinates of the wing's leading edge
symmetric=True,
xsecs=[ # The wing's cross ("X") sections
WingXSec( # Root
x_le=0, # Coordinates of the XSec's leading edge, relative to the wing's leading edge.
y_le=0, # Coordinates of the XSec's leading edge, relative to the wing's leading edge.
z_le=0, # Coordinates of the XSec's leading edge, relative to the wing's leading edge.
chord=0.18,
twist=2, # degrees
airfoil=generic_cambered_airfoil, # Airfoils are blended between a given XSec and the next one.
control_surface_type='symmetric',
# Flap # Control surfaces are applied between a given XSec and the next one.
control_surface_deflection=0, # degrees
),
WingXSec( # Mid
x_le=0.01,
y_le=0.5,
z_le=0,
chord=0.16,
twist=0,
airfoil=generic_cambered_airfoil,
control_surface_type='asymmetric', # Aileron
control_surface_deflection=0,
),
WingXSec( # Tip
x_le=0.08,
y_le=1,
z_le=0.1,
chord=0.08,
twist=-2,
airfoil=generic_cambered_airfoil,
),
]
),
]
)
# airplane.draw()
airplane.set_spanwise_paneling_everywhere(8) # Set the resolution of your analysis
ap = Casll1( # Set up the AeroProblem
airplane=airplane,
op_point=OperatingPoint(
density=1.225, # kg/m^3
viscosity=1.81e-5, # kg/m-s
velocity=10, # m/s
mach=0, # Freestream mach number
alpha=5, # In degrees
beta=0, # In degrees
p=0, # About the body x-axis, in rad/sec
q=0, # About the body y-axis, in rad/sec
r=0, # About the body z-axis, in rad/sec
),
opti=opti # Pass it an optimization environment to work in
)
# Solver options
p_opts = {}
s_opts = {}
s_opts["mu_strategy"] = "adaptive"
opti.solver('ipopt', p_opts, s_opts)
# Solve
try:
sol = opti.solve()
except RuntimeError:
sol = opti.debug
# Postprocess
# Create solved object
ap_sol = copy.deepcopy(ap)
ap_sol.substitute_solution(sol)
ap_sol.draw(show=True) # Generates a pretty picture!
print("CL:", ap_sol.CL)
print("CD:", ap_sol.CD)
print("CY:", ap_sol.CY)
print("Cl:", ap_sol.Cl)
print("Cm:", ap_sol.Cm)
print("Cn:", ap_sol.Cn)
I had this error running this example Traceback (most recent call last): File "C:/Users/Gyaase Elvis/PycharmProjects/Aerosoft_2/aircraft_analysis/temp.py", line 154, in
ap = Casll1Fuse( # Set up the AeroProblem
NameError: name 'Casll1Fuse' is not defined
Any help would be appreciated