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CQ-UAV

Generative UAV models with CadQuery and SplineCloud

CadQuery

CQ-UAV

Installation

pip install git+https://github.com/nomad-vagabond/cq-uav.git@<version_tag>

Replace <version_tag> with the recent tag

Wing Model

Airfoil data

Airfoil geometry and aerodynamic coefficients are collected from the open SplineCloud repositories.

The airfoil collection can be loaded this way:

from cquav.wing.airfoil import load_airfoils_collection

airfoils_collection = load_airfoils_collection()
airfoil_data = airfoils_collection["NACA 6 series airfoils"]["NACA 64(3)-218 (naca643218-il)"]

Retangular Wing Console

The wing model with a constant chord.

Geometry

Airfoil shapes are approximated with smoothing B-Splines, while sharp tails are thickened to avoid malformed geometry.

Wing console consists of three parts:

  • three-chamber box compartment (made of construction material, such as carbon or fiberglass);
  • shaper (XPS foam, PLA, or other lightweight material);
  • shell (coating layer: fiberglass or other).

Geometry is generated based on the airfoil type and wing size. Currently, box thickness is selected automatically based on the profile height.

Instantiating model

from cquav.wing.airfoil import Airfoil
from cquav.wing.profile import AirfoilSection
from cquav.wing.rect_console import RectangularWingConsole

airfoil = Airfoil(airfoil_data)
airfoil_section = AirfoilSection(airfoil, chord=200)
wing_console = RectangularWingConsole(airfoil_section, length=800)

Displaying wing model in Jupyter Notebook (requires jupyter-cadquery ):

import cadquery as cq
from jupyter_cadquery import show

assy = cq.Assembly()
assy.add(wing_console.foam, name="foam", color=cq.Color("lightgray"))
assy.add(wing_console.front_box, name="left_box", color=cq.Color("yellow"))
assy.add(wing_console.central_box, name="central_box", color=cq.Color("yellow"))
assy.add(wing_console.rear_box, name="right_box", color=cq.Color("yellow"))
assy.add(wing_console.shell, name="shell", color=cq.Color("lightskyblue2"))
show(assy, angular_tolerance=0.1)

Assigning materials, calculating and displaying wing console properties

from cquav.materials import IsotropicMaterial, FluidProperties

air_density = 1.225 # [kg/m^3]
velocity = 35 # [m/s]
kinematic_viscosity = 1.460*1e-5 ## [m^2/s]

air_props = FluidProperties(air_density, velocity, kinematic_viscosity)
alpha = 5 # [degrees] - angle of attack

XPS_foam = IsotropicMaterial(30, 1e3, 25*1e6) # (density, tensile strength, tensile modulus)
Fiberglass_laminate = IsotropicMaterial(1800, 290*1e6, 12.4*1e9) # I know it is not isotropic - just another assumption )
materials = {"box": Fiberglass_laminate, "shell": Fiberglass_laminate, "foam": XPS_foam}

wing_console.assign_materials(materials)
wing_console.stats(alpha, air_props)
============================
Length: 1457.6691355004075, [mm]
Chord: 150, [mm]
Aspect ratio: 9.717794236669384
Area: 0.2186503703250611, [m^2]
----------
Mass: 1.5465428361878988, [kg] (box: 0.9566951728265329, foam: 0.027119495312970882, shell: 0.562728168048395)
Angle of attack: 0.3604633440024395, [degrees]
Excess lift force: 32.73014596820137, [N]
Console bend force: 35.77102468212604, [N]
Drag force: 1.3201885631129033, [N]
Lift to weight ratio: 2.7983907782328705
Center of aerodynamic pressure (lift): (37.5, 728.8345677502037), [mm, mm]
----------
Reinforcement box thickness: 1.427, [mm]
Shell thickness: 1, [mm]
Bend stress: 26.0240758429659, [MPa]
Shear stress: 0.6649548985267989, [MPa]
Von Mises stress: 26.049549296133048, [MPa]
Safety factor: 11.132630231074645
===========================

Building wing model with lattice shaper body

wing_console = RectangularWingConsole(airfoil_section, length=800, make_lattice=True)

Solving for wing properties

Finding wing span with the desired relative tip deflection (takes much time for lattice shaper)

solved_console_length = wing_console.fit_length_to_required_tip_deflection(
    alpha, air_props, tip_delta_max=0.1
)

Finding wing span with the desired relative tip deflection and desired excess lift force (unstable, work in progress)

lift_force = 100 # [N]

solved_console_chord = wing_console.fit_chord_to_required_lift_force(
    alpha, air_props, lift_force, tip_delta_max=0.1
)

Other Examples

More detailed examples are covered in the Jupyter notebooks in the 'examples' directory.

Assumptions

  • considered uniform distribution of pressure along the wing span (which is not true in reality);
  • UAV pitch angle = 0, meaning horizontal flight mode with arbitrary angle of attack;
  • only the wing box accepts aerodynamic load, which is applied along its central line;
  • pure box bending is considered to calculate tip deflection;
  • wing console is fixed on one of its ends, another end is free;
  • no torque is analyzed;
  • wing materials are considered to be isotropic;