def test_fuselage_aerodynamics_optimization():
opti = asb.Opti()
alpha = opti.variable(init_guess=0, lower_bound=0, upper_bound=30)
beta = opti.variable(init_guess=0)
fuselage = asb.Fuselage(xsecs=[
asb.FuselageXSec(xyz_c=[xi, 0, 0],
radius=asb.Airfoil("naca0010").local_thickness(
0.8 * xi)) for xi in np.cosspace(0, 1, 20)
], )
aero = asb.AeroBuildup(airplane=asb.Airplane(fuselages=[fuselage]),
op_point=asb.OperatingPoint(velocity=10,
alpha=alpha,
beta=beta)).run()
opti.minimize(-aero["L"] / aero["D"])
sol = opti.solve(verbose=False)
print(sol.value(alpha))
assert sol.value(alpha) > 10 and sol.value(alpha) < 20
assert sol.value(beta) == pytest.approx(0, abs=1e-3)
opti.minimize(aero["D"])
sol = opti.solve(verbose=False)
assert sol.value(alpha) == pytest.approx(0, abs=1e-2)
assert sol.value(beta) == pytest.approx(0, abs=1e-2)
def dynamics(t, y):
dyn = asb.FreeBodyDynamics(
*y,
g=1,
)
aero = asb.AeroBuildup(
airplane=airplane,
op_point=dyn.op_point
).run()
dyn.X, dyn.Y, dyn.Z = aero["F_b"]
dyn.L, dyn.M, dyn.N = aero["M_b"]
derivatives = dyn.state_derivatives()
derivatives["u"] = np.where(
dyn.u < 0,
0,
derivatives["u"]
)
return np.array(list(dyn.state_derivatives().values()))
import aerosandbox as asb
import aerosandbox.numpy as np
from aerosandbox.aerodynamics.aero_3D.test_aero_3D.conventional import airplane
analysis = asb.AeroBuildup(
airplane=airplane,
op_point=asb.OperatingPoint(
atmosphere=asb.Atmosphere(altitude=0, method="ISA"),
velocity=10, # m/s
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
),
)
import matplotlib.pyplot as plt
import aerosandbox.tools.pretty_plots as p
fuselage = asb.Fuselage(xsecs=[
asb.FuselageXSec(xyz_c=[xi, 0, 0],
radius=asb.Airfoil("naca0020").local_thickness(xi) / 2)
for xi in np.cosspace(0, 1, 20)
], )
fig, ax = plt.subplots(figsize=(7, 6))
V = np.linspace(10, 1000, 1001)
op_point = asb.OperatingPoint(velocity=V, )
aero = asb.AeroBuildup(
airplane=asb.Airplane(fuselages=[fuselage]),
op_point=op_point,
).run()
plt.plot(op_point.mach(), aero["CD"], label="Full Model")
aero = asb.AeroBuildup(airplane=asb.Airplane(fuselages=[fuselage]),
op_point=op_point,
include_wave_drag=False).run()
plt.plot(op_point.mach(),
aero["CD"],
zorder=1.9,
label="Model without Wave Drag")
p.show_plot("Transonic Fuselage Drag", "Mach [-]",
"Drag Area $C_D \cdot A$ [m$^2$]")
sol = integrate.solve_ivp(
fun=dynamics,
t_span=t_span,
y0=np.array(list(init_state.values())),
method="LSODA",
vectorized=True,
dense_output=True,
atol=np.array(list(atols.values())),
)
time = np.linspace(sol.t.min(), sol.t.max(), 300)
dyn = asb.FreeBodyDynamics(
time,
*sol.sol(time),
)
aero = asb.AeroBuildup(
airplane=airplane,
op_point=dyn.op_point
).run()
dyn.X, dyn.Y, dyn.Z = aero["F_b"]
dyn.L, dyn.M, dyn.N = aero["M_b"]
from aerosandbox.tools.pretty_plots import plt, show_plot, equal
vars_to_plot = {
**dyn.state,
"alpha" : dyn.alpha,
"beta" : dyn.beta,
"speed" : dyn.speed,
"altitude": dyn.altitude,
"CL" : aero["CL"],
"CY" : aero["CY"],
"CD" : aero["CD"],
af = xsec.airfoil
af.generate_polars(
cache_filename=f"cache/{af.name}.json",
xfoil_kwargs=dict(xfoil_repanel="naca" in af.name.lower()))
airplane.wings[i].xsecs[j].airfoil = af
op_point = asb.OperatingPoint(velocity=25, alpha=3)
vlm = asb.VortexLatticeMethod(
airplane,
op_point,
align_trailing_vortices_with_wind=True,
chordwise_resolution=12,
spanwise_resolution=12,
)
vlm_aero = vlm.run()
ab = asb.AeroBuildup(airplane, op_point)
ab_aero = ab.run()
avl = asb.AVL(
airplane,
op_point,
)
avl_aero = avl.run()
for k, v in {"VLM": vlm_aero, "AVL": avl_aero, "AB": ab_aero}.items():
print(f"{k}:")
for f in ["CL", "CD", "Cm"]:
print(f"\t{f} : {v[f]}")
print(f"\tL/D : {v['CL'] / v['CD']}")
def test_aero_buildup():
analysis = asb.AeroBuildup(
airplane=airplane,
op_point=asb.OperatingPoint(),
)
return analysis.run()
import aerosandbox.numpy as np
import matplotlib.pyplot as plt
import aerosandbox.tools.pretty_plots as p
fuselage = asb.Fuselage(xsecs=[
asb.FuselageXSec(xyz_c=[xi, 0, 0],
radius=asb.Airfoil("naca0010").local_thickness(xi))
for xi in np.cosspace(0, 1, 20)
], )
fig, ax = plt.subplots(figsize=(7, 6))
Beta, Alpha = np.meshgrid(np.linspace(-90, 90, 500), np.linspace(-90, 90, 500))
aero = asb.AeroBuildup(airplane=asb.Airplane(fuselages=[fuselage]),
op_point=asb.OperatingPoint(
velocity=10,
alpha=Alpha,
beta=Beta,
)).run()
from aerosandbox.tools.string_formatting import eng_string
p.contour(
Beta,
Alpha,
aero["L"],
colorbar_label="Lift $L$ [N]",
# levels=100,
linelabels_format=lambda s: eng_string(s, unit="N"),
cmap=plt.get_cmap("coolwarm"))
p.equal()
p.show_plot("3D Fuselage Lift", r"$\beta$ [deg]", r"$\alpha$ [deg]")
def get_aero(xyz_ref):
return asb.AeroBuildup(airplane=asb.Airplane(fuselages=[fuselage],
xyz_ref=xyz_ref),
op_point=asb.OperatingPoint(velocity=10,
alpha=5,
beta=5)).run()