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 get_aero(alpha, taper_ratio):
airplane = asb.Airplane(wings=[
asb.Wing(symmetric=True,
xsecs=[
asb.WingXSec(
xyz_le=[-0.25, 0, 0],
chord=1,
),
asb.WingXSec(
xyz_le=[-0.25 * taper_ratio, 1, 0],
chord=taper_ratio,
)
])
])
op_point = asb.OperatingPoint(
velocity=1,
alpha=alpha,
)
vlm = asb.VortexLatticeMethod(
airplane,
op_point,
chordwise_resolution=6,
spanwise_resolution=6,
)
return vlm.run()
import aerosandbox as asb
import aerosandbox.numpy as np
airfoil = asb.Airfoil("naca0008")
airplane = asb.Airplane(name="Flat Plate",
xyz_ref=[0, 0, 0],
wings=[
asb.Wing(name="Wing",
symmetric=True,
xsecs=[
asb.WingXSec(
xyz_le=[0, 0, 0],
chord=1,
twist=0,
airfoil=airfoil,
),
asb.WingXSec(
xyz_le=[0, 5, 0],
chord=1,
twist=0,
airfoil=airfoil,
),
])
])
if __name__ == '__main__':
airplane.draw()
airplane = asb.Airplane(
name="Vanilla",
xyz_ref=[0.5, 0, 0],
s_ref=9,
c_ref=0.9,
b_ref=10,
wings=[
asb.Wing(name="Wing",
symmetric=True,
xsecs=[
asb.WingXSec(
xyz_le=[0, 0, 0],
chord=1,
twist=2,
airfoil=sd7037,
),
asb.WingXSec(
xyz_le=[0.2, 5, 1],
chord=0.6,
twist=2,
airfoil=sd7037,
)
]),
asb.Wing(name="H-stab",
symmetric=True,
xyz_le=[4, 0, 0],
xsecs=[
asb.WingXSec(
xyz_le=[0, 0, 0],
chord=0.7,
),
asb.WingXSec(xyz_le=[0.14, 1.25, 0], chord=0.42),
]),
asb.Wing(name="V-stab",
xyz_le=[4, 0, 0],
xsecs=[
asb.WingXSec(
xyz_le=[0, 0, 0],
chord=0.7,
),
asb.WingXSec(xyz_le=[0.14, 0, 1], chord=0.42)
])
],
fuselages=[
asb.Fuselage(name="Fuselage",
xyz_le=[0, 0, 0],
xsecs=[
asb.FuselageXSec(
xyz_c=[xi * 5 - 0.5, 0, 0],
radius=asb.Airfoil("naca0024").local_thickness(
x_over_c=xi)) for xi in np.cosspace(0, 1, 30)
])
])
airplane = asb.Airplane(
name="Peter's Glider",
xyz_ref=[0.04, 0, 0], # CG location
wings=[
asb.Wing(
name="Main Wing",
symmetric=True, # Should this wing be mirrored across the XZ plane?
xsecs=[ # The wing's cross ("X") sections
asb.WingXSec( # Root
xyz_le=[0, 0, 0], # Coordinates of the XSec's leading edge, relative to the wing's leading edge.
chord=0.18,
twist=0, # degrees
airfoil=wing_airfoil, # Airfoils are blended between a given XSec and the next one.
control_surface_is_symmetric=True,
# Flap (ctrl. surfs. applied between this XSec and the next one.)
control_surface_deflection=0, # degrees
),
asb.WingXSec( # Mid
xyz_le=[0.01, 0.5, 0],
chord=0.16,
twist=0,
airfoil=wing_airfoil,
control_surface_is_symmetric=False, # Aileron
control_surface_deflection=0,
),
asb.WingXSec( # Tip
xyz_le=[0.08, 1, 0.1],
chord=0.08,
twist=0,
airfoil=wing_airfoil,
),
]
),
asb.Wing(
name="Horizontal Stabilizer",
symmetric=True,
xsecs=[
asb.WingXSec( # root
xyz_le=[0, 0, 0],
chord=0.1,
twist=-2,
airfoil=tail_airfoil,
control_surface_is_symmetric=True, # Elevator
control_surface_deflection=0,
),
asb.WingXSec( # tip
xyz_le=[0.02, 0.17, 0],
chord=0.08,
twist=-2,
airfoil=tail_airfoil
)
]
).translate([0.6, 0, 0.06]),
asb.Wing(
name="Vertical Stabilizer",
symmetric=False,
xsecs=[
asb.WingXSec(
xyz_le=[0, 0, 0],
chord=0.1,
twist=0,
airfoil=tail_airfoil,
control_surface_is_symmetric=True, # Rudder
control_surface_deflection=0,
),
asb.WingXSec(
xyz_le=[0.04, 0, 0.15],
chord=0.06,
twist=0,
airfoil=tail_airfoil
)
]
).translate([0.6, 0, 0.07])
],
fuselages=[
asb.Fuselage(
name="Fuselage",
xsecs=[
asb.FuselageXSec(
xyz_c=[0.8 * xi - 0.1, 0, 0.1 * xi - 0.03],
radius=0.6 * asb.Airfoil("dae51").local_thickness(x_over_c=xi)
)
for xi in np.cosspace(0, 1, 30)
]
)
]
)
right_fuse.xyz_le += cas.vertcat(0, wing_span / 2 * boom_location, 0)
fuses.extend([left_fuse, center_fuse, right_fuse])
left_hstab = copy.deepcopy(hstab)
center_hstab = copy.deepcopy(hstab)
right_hstab = copy.deepcopy(hstab)
left_hstab.xyz_le += cas.vertcat(0, -wing_span / 2 * boom_location, 0)
right_hstab.xyz_le += cas.vertcat(0, wing_span / 2 * boom_location, 0)
hstabs.extend([left_hstab, center_hstab, right_hstab])
left_vstab = copy.deepcopy(vstab)
center_vstab = copy.deepcopy(vstab)
right_vstab = copy.deepcopy(vstab)
left_vstab.xyz_le += cas.vertcat(0, -wing_span / 2 * boom_location, 0)
right_vstab.xyz_le += cas.vertcat(0, wing_span / 2 * boom_location, 0)
vstabs.extend([left_vstab, center_vstab, right_vstab])
else:
raise ValueError("Bad value of n_booms!")
airplane = asb.Airplane(
name="Solar1",
x_ref=0,
y_ref=0,
z_ref=0,
wings=[wing] + hstabs + vstabs,
fuselages=fuses,
)
# airplane.draw()
airplane.write_aswing("solar1.asw")
import aerosandbox as asb
import aerosandbox.numpy as np
import pytest
airplane = asb.Airplane(wings=[
asb.Wing(xsecs=[
asb.WingXSec(
xyz_le=[0, 0, 0], chord=1, airfoil=asb.Airfoil("naca0001")),
asb.WingXSec(
xyz_le=[0, 1, 0], chord=1, airfoil=asb.Airfoil("naca0001"))
])
])
def LD_from_alpha(alpha):
op_point = asb.OperatingPoint(
velocity=1,
alpha=alpha,
)
vlm = asb.VortexLatticeMethod(airplane,
op_point,
align_trailing_vortices_with_wind=True)
aero = vlm.run()
CD0 = 0.01
LD = aero["CL"] / (aero["CD"] + CD0)
return LD
airplane = asb.Airplane(
name="Peter's Glider",
xyz_ref=np.array([0, 0, 0]), # CG location
wings=[
asb.Wing(
name="Main Wing",
xyz_le=np.array([0, 0,
0]), # Coordinates of the wing's leading edge
symmetric=True,
xsecs=[ # The wing's cross ("X") sections
asb.WingXSec( # Root
xyz_le=np.array([0, 0, 0]),
# Coordinates of the XSec's leading edge, relative to the wing's leading edge.
chord=0.18,
twist_angle=2, # degrees
airfoil=
e216, # Airfoils are blended between a given XSec and the next one.
control_surface_is_symmetric=True, # Flap
control_surface_deflection=0, # degrees
),
asb.WingXSec( # Mid
xyz_le=np.array([0.01, 0.5, 0]),
chord=0.16,
twist_angle=0,
airfoil=e216,
control_surface_is_symmetric=False, # Aileron
control_surface_deflection=0,
),
asb.WingXSec( # Tip
xyz_le=np.array([0.08, 1, 0.1]),
chord=0.08,
twist_angle=-2,
airfoil=e216,
),
]),
asb.Wing(
name="Horizontal Stabilizer",
xyz_le=np.array([0.6, 0, 0.1]),
symmetric=True,
xsecs=[
asb.WingXSec( # root
xyz_le=np.array([0, 0, 0]),
chord=0.1,
twist_angle=-10,
airfoil=naca0008,
control_surface_is_symmetric=True, # Elevator
control_surface_deflection=0,
),
asb.WingXSec( # tip
xyz_le=np.array([0.02, 0.17, 0]),
chord=0.08,
twist_angle=-10,
airfoil=naca0008)
]),
asb.Wing(name="Vertical Stabilizer",
xyz_le=np.array([0.6, 0, 0.15]),
symmetric=False,
xsecs=[
asb.WingXSec(
xyz_le=np.array([0, 0, 0]),
chord=0.1,
twist_angle=0,
airfoil=naca0008,
control_surface_deflection=0,
),
asb.WingXSec(xyz_le=np.array([0.04, 0, 0.15]),
chord=0.06,
twist_angle=0,
airfoil=naca0008)
])
])
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$^2$]")
testPlane = ae.Airplane(
name='ju87',
xyz_ref=[0, 0, 0], # CG location
wings=[
ae.Wing(
name="Main Wing",
xyz_le=[0, 0, 0], # Coordinates of the wing's leading edge
symmetric=True,
xsecs=[ # The wing's cross ("X") sections
ae.WingXSec( # Root
xyz_le=[
0, 0, x_trans[0]
], # Coordinates of the XSec's leading edge, relative to the wing's leading edge.
chord=CHORD,
twist=0, # degrees
airfoil=ae.Airfoil(name="naca4412"),
control_surface_type='symmetric',
# Flap # Control surfaces are applied between a given XSec and the next one.
control_surface_deflection=0, # degrees
control_surface_hinge_point=0.75 # as chord fraction
),
ae.WingXSec( # Mid
xyz_le=[0, x_trans[1], 0],
chord=CHORD,
twist=0,
airfoil=ae.Airfoil(name="naca4412"),
control_surface_type='asymmetric', # Aileron
control_surface_deflection=0,
control_surface_hinge_point=0.75),
ae.WingXSec( # Tip
xyz_le=[0, 1, x_trans[2]],
chord=CHORD,
twist=0,
airfoil=ae.Airfoil(name="naca4412"),
)
])
])
airplane = asb.Airplane(
name="Peter's Glider",
xyz_ref=[0, 0, 0], # CG location
wings=[
asb.Wing(
name="Main Wing",
xyz_le=[0, 0, 0], # Coordinates of the wing's leading edge
symmetric=True, # Should this wing be mirrored across the XZ plane?
xsecs=[ # The wing's cross ("X") sections
asb.WingXSec( # Root
xyz_le=[
0, 0, 0
], # Coordinates of the XSec's leading edge, relative to the wing's leading edge.
chord=0.18,
twist_angle=2, # degrees
airfoil=
airfoil, # Airfoils are blended between a given XSec and the next one.
control_surface_is_symmetric=
True, # Flap (ctrl. surfs. applied between this XSec and the next one.)
control_surface_deflection=0, # degrees
),
asb.WingXSec( # Mid
xyz_le=[0.01, 0.5, 0],
chord=0.16,
twist_angle=0,
airfoil=airfoil,
control_surface_is_symmetric=False, # Aileron
control_surface_deflection=0,
),
asb.WingXSec( # Tip
xyz_le=[0.08, 1, 0.1],
chord=0.08,
twist_angle=-2,
airfoil=airfoil,
),
]),
asb.Wing(
name="Horizontal Stabilizer",
xyz_le=[0.6, 0, 0.1],
symmetric=True,
xsecs=[
asb.WingXSec( # root
xyz_le=[0, 0, 0],
chord=0.1,
twist_angle=-10,
airfoil=airfoil,
control_surface_is_symmetric=True, # Elevator
control_surface_deflection=0,
),
asb.WingXSec( # tip
xyz_le=[0.02, 0.17, 0],
chord=0.08,
twist_angle=-10,
airfoil=airfoil)
]),
asb.Wing(
name="Vertical Stabilizer",
xyz_le=[0.6, 0, 0.15],
symmetric=False,
xsecs=[
asb.WingXSec(
xyz_le=[0, 0, 0],
chord=0.1,
twist_angle=0,
airfoil=airfoil,
control_surface_is_symmetric=True, # Rudder
control_surface_deflection=0,
),
asb.WingXSec(xyz_le=[0.04, 0, 0.15],
chord=0.06,
twist_angle=0,
airfoil=airfoil)
])
])
def make_airplane(
n_booms,
wing_span,
):
# n_booms = 3
# wing
# wing_span = 37.126
wing_root_chord = 2.316
wing_x_quarter_chord = -0.1
# hstab
hstab_span = 2.867
hstab_chord = 1.085
hstab_twist_angle = -7
# vstab
vstab_span = 2.397
vstab_chord = 1.134
# fuselage
boom_length = 6.181
nose_length = 1.5
fuse_diameter = 0.6
boom_diameter = 0.2
wing = asb.Wing(
name="Main Wing",
# x_le=-0.05 * wing_root_chord, # Coordinates of the wing's leading edge # TODO make this a free parameter?
x_le=
wing_x_quarter_chord, # Coordinates of the wing's leading edge # TODO make this a free parameter?
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
asb.WingXSec( # Root
x_le=-wing_root_chord / 4,
# 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=wing_root_chord,
twist=0, # degrees
airfoil=
e216, # 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
spanwise_panels=30,
),
asb.WingXSec( # Tip
x_le=-wing_root_chord * 0.5 / 4,
y_le=wing_span / 2,
z_le=0, # wing_span / 2 * cas.pi / 180 * 5,
chord=wing_root_chord * 0.5,
twist=0,
airfoil=e216,
),
],
)
hstab = asb.Wing(
name="Horizontal Stabilizer",
x_le=boom_length - vstab_chord * 0.75 -
hstab_chord, # Coordinates of the wing's leading edge
y_le=0, # Coordinates of the wing's leading edge
z_le=0.1, # Coordinates of the wing's leading edge
symmetric=True,
xsecs=[ # The wing's cross ("X") sections
asb.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=hstab_chord,
twist=-3, # degrees # TODO fix
airfoil=
naca0008, # 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
spanwise_panels=8,
),
asb.WingXSec( # Tip
x_le=0,
y_le=hstab_span / 2,
z_le=0,
chord=hstab_chord,
twist=-3, # TODO fix
airfoil=naca0008,
),
],
)
vstab = asb.Wing(
name="Vertical Stabilizer",
x_le=boom_length -
vstab_chord * 0.75, # Coordinates of the wing's leading edge
y_le=0, # Coordinates of the wing's leading edge
z_le=-vstab_span / 2 +
vstab_span * 0.15, # Coordinates of the wing's leading edge
symmetric=False,
xsecs=[ # The wing's cross ("X") sections
asb.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=vstab_chord,
twist=0, # degrees
airfoil=
naca0008, # 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
spanwise_panels=8,
),
asb.WingXSec( # Tip
x_le=0,
y_le=0,
z_le=vstab_span,
chord=vstab_chord,
twist=0,
airfoil=naca0008,
),
],
)
### Build the fuselage geometry
blend = lambda x: (1 - np.cos(np.pi * x)) / 2
fuse_x_c = []
fuse_z_c = []
fuse_radius = []
fuse_resolution = 10
# Nose geometry
fuse_nose_theta = np.linspace(0, np.pi / 2, fuse_resolution)
fuse_x_c.extend([(wing_x_quarter_chord - wing_root_chord / 4) -
nose_length * np.cos(theta) for theta in fuse_nose_theta])
fuse_z_c.extend([-fuse_diameter / 2] * fuse_resolution)
fuse_radius.extend(
[fuse_diameter / 2 * np.sin(theta) for theta in fuse_nose_theta])
# Taper
fuse_taper_x_nondim = np.linspace(0, 1, fuse_resolution)
fuse_x_c.extend([
0.0 * boom_length + (0.6 - 0.0) * boom_length * x_nd
for x_nd in fuse_taper_x_nondim
])
fuse_z_c.extend([
-fuse_diameter / 2 * blend(1 - x_nd) - boom_diameter / 2 * blend(x_nd)
for x_nd in fuse_taper_x_nondim
])
fuse_radius.extend([
fuse_diameter / 2 * blend(1 - x_nd) + boom_diameter / 2 * blend(x_nd)
for x_nd in fuse_taper_x_nondim
])
# Tail
# fuse_tail_x_nondim = np.linspace(0, 1, fuse_resolution)[1:]
# fuse_x_c.extend([
# 0.9 * boom_length + (1 - 0.9) * boom_length * x_nd for x_nd in fuse_taper_x_nondim
# ])
# fuse_z_c.extend([
# -boom_diameter / 2 * blend(1 - x_nd) for x_nd in fuse_taper_x_nondim
# ])
# fuse_radius.extend([
# boom_diameter / 2 * blend(1 - x_nd) for x_nd in fuse_taper_x_nondim
# ])
fuse_straight_resolution = 4
fuse_x_c.extend([
0.6 * boom_length + (1 - 0.6) * boom_length * x_nd
for x_nd in np.linspace(0, 1, fuse_straight_resolution)[1:]
])
fuse_z_c.extend([-boom_diameter / 2] * (fuse_straight_resolution - 1))
fuse_radius.extend([boom_diameter / 2] * (fuse_straight_resolution - 1))
fuse = asb.Fuselage(
name="Fuselage",
x_le=0,
y_le=0,
z_le=0,
xsecs=[
asb.FuselageXSec(x_c=fuse_x_c[i],
z_c=fuse_z_c[i],
radius=fuse_radius[i])
for i in range(len(fuse_x_c))
],
)
# Assemble the airplane
fuses = []
hstabs = []
vstabs = []
if n_booms == 1:
fuses.append(fuse)
hstabs.append(hstab)
vstabs.append(vstab)
elif n_booms == 2:
boom_location = 0.40 # as a fraction of the half-span
left_fuse = copy.deepcopy(fuse)
right_fuse = copy.deepcopy(fuse)
left_fuse.xyz_le += cas.vertcat(0, -wing_span / 2 * boom_location, 0)
right_fuse.xyz_le += cas.vertcat(0, wing_span / 2 * boom_location, 0)
fuses.extend([left_fuse, right_fuse])
left_hstab = copy.deepcopy(hstab)
right_hstab = copy.deepcopy(hstab)
left_hstab.xyz_le += cas.vertcat(0, -wing_span / 2 * boom_location, 0)
right_hstab.xyz_le += cas.vertcat(0, wing_span / 2 * boom_location, 0)
hstabs.extend([left_hstab, right_hstab])
left_vstab = copy.deepcopy(vstab)
right_vstab = copy.deepcopy(vstab)
left_vstab.xyz_le += cas.vertcat(0, -wing_span / 2 * boom_location, 0)
right_vstab.xyz_le += cas.vertcat(0, wing_span / 2 * boom_location, 0)
vstabs.extend([left_vstab, right_vstab])
elif n_booms == 3:
boom_location = 0.57 # as a fraction of the half-span
left_fuse = copy.deepcopy(fuse)
center_fuse = copy.deepcopy(fuse)
right_fuse = copy.deepcopy(fuse)
left_fuse.xyz_le += cas.vertcat(0, -wing_span / 2 * boom_location, 0)
right_fuse.xyz_le += cas.vertcat(0, wing_span / 2 * boom_location, 0)
fuses.extend([left_fuse, center_fuse, right_fuse])
left_hstab = copy.deepcopy(hstab)
center_hstab = copy.deepcopy(hstab)
right_hstab = copy.deepcopy(hstab)
left_hstab.xyz_le += cas.vertcat(0, -wing_span / 2 * boom_location, 0)
right_hstab.xyz_le += cas.vertcat(0, wing_span / 2 * boom_location, 0)
hstabs.extend([left_hstab, center_hstab, right_hstab])
left_vstab = copy.deepcopy(vstab)
center_vstab = copy.deepcopy(vstab)
right_vstab = copy.deepcopy(vstab)
left_vstab.xyz_le += cas.vertcat(0, -wing_span / 2 * boom_location, 0)
right_vstab.xyz_le += cas.vertcat(0, wing_span / 2 * boom_location, 0)
vstabs.extend([left_vstab, center_vstab, right_vstab])
else:
raise ValueError("Bad value of n_booms!")
airplane = asb.Airplane(
name="Solar1",
x_ref=0,
y_ref=0,
z_ref=0,
wings=[wing] + hstabs + vstabs,
fuselages=fuses,
)
return airplane
def CD_function(alpha, Re, mach, deflection):
return 0 * Re**0
def CM_function(alpha, Re, mach, deflection):
return 0 * alpha
ideal_airfoil = asb.Airfoil(name="Ideal Airfoil",
coordinates=get_NACA_coordinates('naca0012'),
CL_function=CL_function,
CD_function=CD_function,
CM_function=CM_function)
wing = asb.Wing(xsecs=[
asb.WingXSec(xyz_le=[0, y_i, 0], chord=1, airfoil=ideal_airfoil)
for y_i in [0, 10]
],
symmetric=True)
airplane = asb.Airplane(wings=[wing])
op_point = asb.OperatingPoint(velocity=340, alpha=0)
aero = asb.AeroBuildup(airplane=airplane, op_point=op_point).run()
from pprint import pprint
pprint(aero)
def airplane(self):
return sbx.Airplane(name=self.NAME,
xyz_ref=[0, 0, 0],
wings=self.wings)
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()
airplane = asb.Airplane(
name="Cessna 152",
wings=[
asb.Wing(name="Wing",
xsecs=[
asb.WingXSec(xyz_le=[0, 0, 0],
chord=ft_to_m(5, 4),
airfoil=naca2412),
asb.WingXSec(
xyz_le=[0, ft_to_m(7),
ft_to_m(7) * np.sind(1)],
chord=ft_to_m(5, 4),
airfoil=naca2412),
asb.WingXSec(xyz_le=[
ft_to_m(4, 3 / 4) - ft_to_m(3, 8 + 1 / 2),
ft_to_m(33, 4) / 2,
ft_to_m(33, 4) / 2 * np.sind(1)
],
chord=ft_to_m(3, 8 + 1 / 2),
airfoil=naca0012)
],
symmetric=True),
asb.Wing(name="Horizontal Stabilizer",
xsecs=[
asb.WingXSec(xyz_le=[0, 0, 0],
chord=ft_to_m(3, 8),
airfoil=naca0012,
twist=-2),
asb.WingXSec(xyz_le=[ft_to_m(1),
ft_to_m(10) / 2, 0],
chord=ft_to_m(2, 4 + 3 / 8),
airfoil=naca0012,
twist=-2)
],
symmetric=True).translate([ft_to_m(13, 3), 0,
ft_to_m(-2)]),
asb.Wing(name="Vertical Stabilizer",
xsecs=[
asb.WingXSec(
xyz_le=[ft_to_m(-5), 0, 0],
chord=ft_to_m(8, 8),
airfoil=naca0012,
),
asb.WingXSec(
xyz_le=[ft_to_m(0), 0, ft_to_m(1)],
chord=ft_to_m(3, 8),
airfoil=naca0012,
),
asb.WingXSec(
xyz_le=[ft_to_m(0, 8), 0,
ft_to_m(5)],
chord=ft_to_m(2, 8),
airfoil=naca0012,
),
]).translate(
[ft_to_m(16, 11) - ft_to_m(3, 8), 0,
ft_to_m(-2)])
],
fuselages=[
asb.Fuselage(xsecs=[
asb.FuselageXSec(
xyz_c=[0, 0, ft_to_m(-1)],
radius=0,
),
asb.FuselageXSec(xyz_c=[0, 0, ft_to_m(-1)], radius=ft_to_m(1.5)),
asb.FuselageXSec(
xyz_c=[ft_to_m(3), 0, ft_to_m(-0.85)], radius=ft_to_m(1.7)),
asb.FuselageXSec(xyz_c=[ft_to_m(5), 0, ft_to_m(0)],
radius=ft_to_m(2.7)),
asb.FuselageXSec(
xyz_c=[ft_to_m(10, 4), 0, ft_to_m(0.3)], radius=ft_to_m(2.3)),
asb.FuselageXSec(
xyz_c=[ft_to_m(21, 11), 0, ft_to_m(0.8)], radius=ft_to_m(0.3)),
]).translate([ft_to_m(-5), 0, ft_to_m(-3)])
])
