def _create_winglet(self):
"""Create winglet.
Returns
-------
aerosandbox.Wing
"""
# Extract winglet configuration
parameters = self.winglet_parameters
dimensions = self.winglet_dimensions
location_tip = self.__get_winglet_vector__(
length=dimensions["length"],
sweep=parameters[W_ANGLE_SWEEP],
cant=parameters[W_ANGLE_CANT],
)
# Slightly separete from the wing
_epsilon_winglet_wing = Point([0, 0, 0.01])
# Get wing tip section
_section = self.__wingtip_section__
_coordinates = list(_section[LE_LOCATION])
coordinates_weld = _coordinates + _epsilon_winglet_wing
# Match TE of wing tip and winglet root chord
chord_root = dimensions["chord_root"]
chord_tip = dimensions["chord_tip"]
coordinates_weld.x += _section[CHORD] - chord_root
winglet_airfoil = sbx.Airfoil(name=parameters[W_AIRFOIL])
twist_root = parameters[W_ANGLE_TWIST_ROOT]
twist_tip = parameters[W_ANGLE_TWIST_TIP]
winglet = sbx.Wing(
name="Winglet",
xyz_le=list(coordinates_weld),
symmetric=True,
xsecs=[
sbx.WingXSec(
xyz_le=[0, 0, 0],
chord=chord_root,
twist=twist_root,
airfoil=winglet_airfoil,
),
sbx.WingXSec(
xyz_le=list(location_tip),
chord=chord_tip,
twist=twist_tip,
airfoil=winglet_airfoil,
),
],
)
self.winglet = [winglet]
return winglet
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()
def create_wing_planform(self):
"""Create planform object.
Returns
-------
aerosanbox.Wing
"""
# Create sections
sections = []
for _section in self.sections:
_coordinates = list(_section[LE_LOCATION])
_airfoil = sbx.Airfoil(name=_section[AIRFOIL])
_sbx_section = sbx.WingXSec(
xyz_le=
_coordinates, # Coordinates of the XSec's leading edge, **relative** to the wing's leading edge.
chord=_section[CHORD],
twist=_section[TWIST], # degrees
airfoil=_airfoil,
)
sections.append(_sbx_section)
# Create wing
planform = sbx.Wing(
name=self.NAME,
xyz_le=[0.0, 0.0, 0.0], # Coordinates of the wing's leading edge
symmetric=True,
xsecs=sections,
)
self.planform = planform
return planform
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()
sd7037 = asb.Airfoil("sd7037")
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],
### Define the 3D geometry you want to analyze/optimize.
# Here, all distances are in meters and all angles are in degrees.
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,
wing = asb.Wing(
name="Main Wing",
# x_le=-0.05 * wing_root_chord, # Coordinates of the wing's leading edge
x_le=wing_x_quarter_chord, # 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
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
),
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,
),
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
# Here, all distances are in meters and all angles are in degrees.
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,
#
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]],
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 ft_to_m(feet, inches=0): # Converts feet (and inches) to meters
return feet * u.foot + inches * u.inch
naca2412 = asb.Airfoil("naca2412")
naca0012 = asb.Airfoil("naca0012")
naca2412.generate_polars(cache_filename="assets/naca2412.json")
naca0012.generate_polars(cache_filename="assets/naca0012.json")
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),
### Define the 3D geometry you want to analyze/optimize.
# Here, all distances are in meters and all angles are in degrees.
airplane = asb.Airplane(
name="Peter's Glider",
xyz_ref=[0, 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=2, # degrees
airfoil=
wing_airfoil, # Airfoils are blended between a given XSec and the next one.
),
asb.WingXSec( # Mid
xyz_le=[0.01, 0.5, 0],
chord=0.16,
twist=0,
airfoil=wing_airfoil,
),
asb.WingXSec( # Tip
xyz_le=[0.08, 1, 0.1],
chord=0.08,
twist=-2,
airfoil=wing_airfoil,
