Help needed for modelling a M20V Acclaim Ultra.

[chrismakhoul]

Hello,

I am trying to model a M20V Acclaim Ultra on python.
This is the code I have

The output is some sort of a flat surface that doesn't look like a plane at all !

Any ideas on how to help me ?

[cdhainaut]

Hi,
Your wings have a span equal to zero because of how you define xyz_le of each wing tip section.
Try this (just changed x coordinate to y coordinate):

from aerosandbox import *
airfoil_main = Airfoil(name="naca4412")
airfoil_vertical = Airfoil(name="naca0012")
m20v_acclaim_ultra = Airplane(
    name="Mooney M20V Acclaim Ultra",
    xyz_ref=[0, 0, 0],  # CG location
    wings=[
        Wing(
            name="Main Wing",
            xyz_le=[0, 0, 0],  # Position of the wing's leading edge
            symmetric=True,
            xsecs=[  # Wing cross-sections
                WingXSec(  # Root
                    xyz_le=[0, 0, 0],  # Position of the root section's leading edge
                    chord=1.9,
                    twist=1.7,
                    airfoil=airfoil_main,
                ),
                WingXSec(  # Tip
                    xyz_le=[0, 5, 0],  # Position of the tip section's leading edge
                    chord=0.6,
                    twist=0,
                    airfoil=airfoil_main,
                ),
            ],
        ),
        Wing(
            name="Horizontal Stabilizer",
            xyz_le=[8, 0, 0],  # Position of the stabilizer's leading edge
            symmetric=True,
            xsecs=[
                WingXSec(  # Root
                    xyz_le=[0, 0, 0],  # Position of the root section's leading edge
                    chord=0.6,
                    twist=0,
                    airfoil=airfoil_main,
                ),
                WingXSec(  # Tip
                    xyz_le=[0, 2, 0],  # Position of the tip section's leading edge
                    chord=0.3,
                    twist=0,
                    airfoil=airfoil_main,
                ),
            ],
        ),
        Wing(
            name="Vertical Stabilizer",
            xyz_le=[8, 0, 0.5],  # Position of the stabilizer's leading edge
            symmetric=False,
            xsecs=[
                WingXSec(
                    xyz_le=[0, 0, 0],  # Position of the root section's leading edge
                    chord=0.5,
                    twist=0,
                    airfoil=airfoil_vertical,
                ),
                WingXSec(
                    xyz_le=[0.2, 0, 0.5],  # Position of the tip section's leading edge
                    chord=0.25,
                    twist=0,
                    airfoil=airfoil_vertical,
                ),
            ],
        ),
    ],
)
m20v_acclaim_ultra.draw_three_view()

Gives this:

You may want to translate the stabilizers using the Wing.translate() function:

        Wing(
            name="Horizontal Stabilizer",
            xyz_le=[8, 0, 0],  # Position of the stabilizer's leading edge
            symmetric=True,
            xsecs=[
                WingXSec(  # Root
                    xyz_le=[0, 0, 0],  # Position of the root section's leading edge
                    chord=0.6,
                    twist=0,
                    airfoil=airfoil_main,
                ),
                WingXSec(  # Tip
                    xyz_le=[0, 2, 0],  # Position of the tip section's leading edge
                    chord=0.3,
                    twist=0,
                    airfoil=airfoil_main,
                ),
            ],
        ).translate([10.0, 0, 0.0]),
        Wing(
            name="Vertical Stabilizer",
            xyz_le=[8, 0, 0.5],  # Position of the stabilizer's leading edge
            symmetric=False,
            xsecs=[
                WingXSec(
                    xyz_le=[0, 0, 0],  # Position of the root section's leading edge
                    chord=0.5,
                    twist=0,
                    airfoil=airfoil_vertical,
                ),
                WingXSec(
                    xyz_le=[0.2, 0, 0.5],  # Position of the tip section's leading edge
                    chord=0.25,
                    twist=0,
                    airfoil=airfoil_vertical,
                ),
            ],
        ).translate([10.0, 0, 0.0]),

And this will give this:

[chrismakhoul]

Thank you so much for your answer ! I've been struggling with this since yesterday.
I also want to add the fuselage, any idea on why my code is not working ?

[cdhainaut]

Could you please detail what is not working? Or maybe share your code using the method described here instead (using backticks):
https://docs.github.com/en/get-started/writing-on-github/working-with-advanced-formatting/creating-and-highlighting-code-blocks

Thanks

[chrismakhoul]

Here's my code :

from aerosandbox import *
airfoil_main = Airfoil(name="naca4412")
airfoil_vertical = Airfoil(name="naca0012")
m20v_acclaim_ultra = Airplane(
    name="Mooney M20V Acclaim Ultra",
    xyz_ref=[0, 0, 0],  # CG location
    wings=[
        Wing(
            name="Main Wing",
            xyz_le=[0, 0, 0],  # Position of the wing's leading edge
            symmetric=True,
            xsecs=[  # Wing cross-sections
                WingXSec(  # Root
                    xyz_le=[0, 0, 0],  # Position of the root section's leading edge
                    chord=1.9,
                    twist=1.7,
                    airfoil=airfoil_main,
                ),
                WingXSec(  # Tip
                    xyz_le=[0, 5, 0],  # Position of the tip section's leading edge
                    chord=0.6,
                    twist=0,
                    airfoil=airfoil_main,
                ),
            ],
        ),
        Wing(
            name="Horizontal Stabilizer",
            xyz_le=[8, 0, 0],  # Position of the stabilizer's leading edge
            symmetric=True,
            xsecs=[
                WingXSec(  # Root
                    xyz_le=[0, 0, 0],  # Position of the root section's leading edge
                    chord=0.6,
                    twist=0,
                    airfoil=airfoil_main,
                ),
                WingXSec(  # Tip
                    xyz_le=[0, 2, 0],  # Position of the tip section's leading edge
                    chord=0.3,
                    twist=0,
                    airfoil=airfoil_main,
                ),
            ],
        ),
        Wing(
            name="Vertical Stabilizer",
            xyz_le=[8, 0, 0.5],  # Position of the stabilizer's leading edge
            symmetric=False,
            xsecs=[
                WingXSec(
                    xyz_le=[0, 0, 0],  # Position of the root section's leading edge
                    chord=0.5,
                    twist=0,
                    airfoil=airfoil_vertical,
                ),
                WingXSec(
                    xyz_le=[0.2, 0, 0.5],  # Position of the tip section's leading edge
                    chord=0.25,
                    twist=0,
                    airfoil=airfoil_vertical,
                ),
            ],
        ),
    ],
    fuselages=[  # Including fuselage definition
        Fuselage(
            name="Fuselage",  # Name of the fuselage
            xyz_le=[-0.05, 0, 0],  # Location of the fuselage's reference point
            symmetric=True,  # Assuming symmetry
            xsecs=[  # Fuselage cross-sections
                FuselageXSec(  # First cross-section
                    xyz_c=[-0.05, 0, 0],  # Location of the cross-section's center, relative to the leading edge
                    radius=0.05,  # Radius
                ),
                FuselageXSec(  # Second cross-section
                    xyz_c=[1, 0, 0],  # Location of the cross-section's center, relative to the leading edge
                    radius=0.04,  # Radius
                ),
                FuselageXSec(  # Third cross-section
                    xyz_c=[2, 0, 0],  # Location of the cross-section's center, relative to the leading edge
                    radius=0.03,  # Radius
                ),
         ]
        )
    ]
)
m20v_acclaim_ultra.draw_three_view()

# Operating conditions
op_point = OperatingPoint(
    velocity=128.6,  # m/s
    alpha=5,  # Angle of attack in degrees
    beta=0,  # Sideslip angle in degrees
    p=0,  # Roll rate
    q=0,  # Pitch rate
    r=0,  # Yaw rate
)

# Setup the Vortex Lattice Method (VLM) analysis
aero_problem = VortexLatticeMethod(
    airplane=m20v_acclaim_ultra,
    op_point=op_point,
)

Again, thank you for your help.