Multiple Lifting Surfaces

It’s easily possible to simulate multiple lifting surfaces simultaneously in OpenAeroStruct. The most straightforward example is a wing and a tail for a conventional airplane, as shown below, though OpenAeroStruct can handle any arbitrary collection of lifting surfaces.

import numpy as np

import openmdao.api as om

from openaerostruct.geometry.utils import generate_mesh
from openaerostruct.geometry.geometry_group import Geometry
from openaerostruct.aerodynamics.aero_groups import AeroPoint

# Create a dictionary to store options about the surface
mesh_dict = {"num_y": 7, "num_x": 2, "wing_type": "CRM", "symmetry": True, "num_twist_cp": 5}

mesh, twist_cp = generate_mesh(mesh_dict)

surf_dict = {
    # Wing definition
    "name": "wing",  # name of the surface
    "symmetry": True,  # if true, model one half of wing
    # reflected across the plane y = 0
    "S_ref_type": "wetted",  # how we compute the wing area,
    # can be 'wetted' or 'projected'
    "fem_model_type": "tube",
    "twist_cp": twist_cp,
    "mesh": mesh,
    # Aerodynamic performance of the lifting surface at
    # an angle of attack of 0 (alpha=0).
    # These CL0 and CD0 values are added to the CL and CD
    # obtained from aerodynamic analysis of the surface to get
    # the total CL and CD.
    # These CL0 and CD0 values do not vary wrt alpha.
    "CL0": 0.0,  # CL of the surface at alpha=0
    "CD0": 0.015,  # CD of the surface at alpha=0
    # Airfoil properties for viscous drag calculation
    "k_lam": 0.05,  # percentage of chord with laminar
    # flow, used for viscous drag
    "t_over_c_cp": np.array([0.15]),  # thickness over chord ratio (NACA0015)
    "c_max_t": 0.303,  # chordwise location of maximum (NACA0015)
    # thickness
    "with_viscous": True,  # if true, compute viscous drag
    "with_wave": False,  # if true, compute wave drag
}

# Create a dictionary to store options about the surface
mesh_dict = {"num_y": 7, "num_x": 2, "wing_type": "rect", "symmetry": True, "offset": np.array([50, 0.0, 0.0])}

mesh = generate_mesh(mesh_dict)

surf_dict2 = {
    # Wing definition
    "name": "tail",  # name of the surface
    "symmetry": True,  # if true, model one half of wing
    # reflected across the plane y = 0
    "S_ref_type": "wetted",  # how we compute the wing area,
    # can be 'wetted' or 'projected'
    "twist_cp": twist_cp,
    "mesh": mesh,
    # Aerodynamic performance of the lifting surface at
    # an angle of attack of 0 (alpha=0).
    # These CL0 and CD0 values are added to the CL and CD
    # obtained from aerodynamic analysis of the surface to get
    # the total CL and CD.
    # These CL0 and CD0 values do not vary wrt alpha.
    "CL0": 0.0,  # CL of the surface at alpha=0
    "CD0": 0.0,  # CD of the surface at alpha=0
    "fem_origin": 0.35,
    # Airfoil properties for viscous drag calculation
    "k_lam": 0.05,  # percentage of chord with laminar
    # flow, used for viscous drag
    "t_over_c_cp": np.array([0.15]),  # thickness over chord ratio (NACA0015)
    "c_max_t": 0.303,  # chordwise location of maximum (NACA0015)
    # thickness
    "with_viscous": True,  # if true, compute viscous drag
    "with_wave": False,  # if true, compute wave drag
}

surfaces = [surf_dict, surf_dict2]

# Create the problem and the model group
prob = om.Problem()

indep_var_comp = om.IndepVarComp()
indep_var_comp.add_output("v", val=248.136, units="m/s")
indep_var_comp.add_output("alpha", val=5.0, units="deg")
indep_var_comp.add_output("Mach_number", val=0.84)
indep_var_comp.add_output("re", val=1.0e6, units="1/m")
indep_var_comp.add_output("rho", val=0.38, units="kg/m**3")
indep_var_comp.add_output("cg", val=np.zeros((3)), units="m")

prob.model.add_subsystem("prob_vars", indep_var_comp, promotes=["*"])

# Loop over each surface in the surfaces list
for surface in surfaces:
    geom_group = Geometry(surface=surface)

    # Add tmp_group to the problem as the name of the surface.
    # Note that is a group and performance group for each
    # individual surface.
    prob.model.add_subsystem(surface["name"], geom_group)

# Loop through and add a certain number of aero points
for i in range(1):
    # Create the aero point group and add it to the model
    aero_group = AeroPoint(surfaces=surfaces)
    point_name = "aero_point_{}".format(i)
    prob.model.add_subsystem(point_name, aero_group)

    # Connect flow properties to the analysis point
    prob.model.connect("v", point_name + ".v")
    prob.model.connect("alpha", point_name + ".alpha")
    prob.model.connect("Mach_number", point_name + ".Mach_number")
    prob.model.connect("re", point_name + ".re")
    prob.model.connect("rho", point_name + ".rho")
    prob.model.connect("cg", point_name + ".cg")

    # Connect the parameters within the model for each aero point
    for surface in surfaces:
        name = surface["name"]

        # Connect the mesh from the geometry component to the analysis point
        prob.model.connect(name + ".mesh", point_name + "." + name + ".def_mesh")

        # Perform the connections with the modified names within the
        # 'aero_states' group.
        prob.model.connect(name + ".mesh", point_name + ".aero_states." + name + "_def_mesh")

        prob.model.connect(name + ".t_over_c", point_name + "." + name + "_perf." + "t_over_c")

# Set up the problem
prob.setup()

prob.run_model()

print(prob["aero_point_0.wing_perf.CD"][0])

0.03539775778518085

print(prob["aero_point_0.wing_perf.CL"][0])

0.512473693224805

print(prob["aero_point_0.CM"][1])

-2.016825327289286