def test_u1_read_in_from_file():
# Generate airplane from file
filename = "test/use_cases/u1_from_file/use_case_1.json"
myPlane = geom.Airplane(inputfile=filename)
# Generate lifting-line model for airplane
myLLModel = LLModel(myPlane)
# Solve the lifting-line model for the given condition
controls = {
"aileron": 10.,
"elevator": 5.,
}
aero_state = {"V_mag": 100., "alpha": 5., "rho": 1.}
results = myLLModel.solve(stype="linear",
control_state=controls,
aero_state=aero_state)
# compare results with expected values
if myLLModel._machup_compare:
test = np.array([
498.563823789387015, -355.559967409590001, -22164.650382783511304,
-15767.900012819352924, -6069.667491323640206, -485.172664364582431
])
else:
test = np.array([
488.73383601455544, -355.57437639926093, -22143.208875850567,
-15767.798773697272, -6079.3470120008151, -485.09007262607491
])
assert np.allclose(results["FX"], test[0], rtol=0., atol=1e-10) is True
assert np.allclose(results["FY"], test[1], rtol=0., atol=1e-10) is True
assert np.allclose(results["FZ"], test[2], rtol=0., atol=1e-10) is True
assert np.allclose(results["l"], test[3], rtol=0., atol=1e-10) is True
assert np.allclose(results["m"], test[4], rtol=0., atol=1e-10) is True
assert np.allclose(results["n"], test[5], rtol=0., atol=1e-10) is True
def test_u2_simple_monowing():
myPlane = geom.Airplane()
myPlane.cg_location(-0.29, 0., 0.25)
# add main wing and aileron control surface
mainWing = myPlane.add_wing("main_wing",
position=[0., 0., 0.],
semispan=4.,
root_chord=1.,
tip_chord=0.5,
dihedral=3.)
mainWing.airfoil("NACA2412",
alpha_L0=-0.036899751,
CL_alpha=6.283185307,
Cm_L0=-0.0527,
Cm_alpha=-0.08,
CD0=0.0055,
CD0_L=-0.0045,
CD0_L2=0.01,
CL_max=1.4)
mainWing.control_surface(percent_span=(0.4, 0.9),
percent_chord=0.25,
mix={"aileron": 1.})
# add horizontal tail and elevator control surface
horizontalTail = myPlane.add_wing("horizontal_tail",
position=[-4., 0., 0.],
semispan=1.,
root_chord=1.,
tip_chord=0.5,
mount_angle=1.,
sweep=20.)
horizontalTail.airfoil("NACA0012",
alpha_L0=0.,
CL_alpha=6.283185307,
Cm_L0=0.,
Cm_alpha=0.,
CD0=0.0055,
CD0_L=-0.0045,
CD0_L2=0.01,
CL_max=1.)
horizontalTail.control_surface(percent_span=(0., 1.),
percent_chord=(0.25, 0.5),
mix={"elevator": 1.})
# add vertical tail and rudder control surface
verticalTail = myPlane.add_wing("vertical_tail",
position=[-4., 0., 0.],
side="right",
semispan=0.8,
root_chord=1.,
tip_chord=0.5,
sweep=25.,
dihedral=90.)
verticalTail.airfoil("NACA0012",
alpha_L0=0.,
CL_alpha=6.283185307,
Cm_L0=0.,
Cm_alpha=0.,
CD0=0.0055,
CD0_L=-0.0045,
CD0_L2=0.01,
CL_max=1.)
verticalTail.control_surface(percent_span=(0., 1.),
percent_chord=(0.25, 0.5),
mix={"rudder": 1.})
# Generate lifting-line model for airplane
myLLModel = LLModel(myPlane)
# Solve the lifting-line model for the given condition
controls = {
"aileron": 10.,
"elevator": 5.,
}
aero_state = {"V_mag": 100., "alpha": 5., "rho": 1.}
results = myLLModel.solve(stype="linear",
control_state=controls,
aero_state=aero_state)
# compare results with expected values
if myLLModel._machup_compare:
test = np.array([
498.563823789387015, -355.559967409590001, -22164.650382783511304,
-15767.900012819352924, -6069.667491323640206, -485.172664364582431
])
else:
test = np.array([
488.73383601455544, -355.57437639926093, -22143.208875850567,
-15767.798773697272, -6079.3470120008151, -485.09007262607491
])
assert np.allclose(results["FX"], test[0], rtol=0., atol=1e-10) is True
assert np.allclose(results["FY"], test[1], rtol=0., atol=1e-10) is True
assert np.allclose(results["FZ"], test[2], rtol=0., atol=1e-10) is True
assert np.allclose(results["l"], test[3], rtol=0., atol=1e-10) is True
assert np.allclose(results["m"], test[4], rtol=0., atol=1e-10) is True
assert np.allclose(results["n"], test[5], rtol=0., atol=1e-10) is True
def test_u3_multisegment_wing():
my_plane = geom.Airplane()
# add multisegment wing with aileron and flap control surfaces
main_inboard = my_plane.add_wing("main_inboard",
location=[0., 0., 0.],
semispan=2.,
root_chord=1.,
dihedral=3.)
main_inboard.airfoil("NACA2412",
alpha_L0=-0.036899751,
CL_alpha=6.283185307,
Cm_L0=-0.0527,
Cm_alpha=-0.08,
CD0=0.0055,
CD0_L=-0.0045,
CD0_L2=0.01,
CL_max=1.4)
main_inboard.control_surface(percent_span=[0.2, 0.9],
percent_chord=0.25,
mix={"flap": 1.})
main_outboard = my_plane.add_wing("main_outboard",
connect_to=("main_inboard", "tip"),
semispan=2.,
root_chord=1.,
tip_chord=0.6,
sweep=-3.)
main_outboard.airfoil("NACA2412",
alpha_L0=-0.036899751,
CL_alpha=6.283185307,
Cm_L0=-0.0527,
Cm_alpha=-0.08,
CD0=0.0055,
CD0_L=-0.0045,
CD0_L2=0.01,
CL_max=1.4)
main_outboard.control_surface(percent_span=[0., 0.8],
percent_chord=0.25,
mix={"aileron": 1.})
# Generate lifting-line model for airplane
my_llmodel = LLModel(my_plane)
# Solve the lifting-line model for the given condition
controls = {
"aileron": 4.,
"flap": 7.,
}
aero_state = {
"V_mag": 100.,
"alpha": 1.,
"rho": 1.
}
results = my_llmodel.solve(stype="linear",
control_state=controls,
aero_state=aero_state)
# compare results with expected values
if my_llmodel._machup_compare:
test = np.array([-501.548045209023599,
-19.359345013074183,
-14896.07014164796783,
-5942.396955478406198,
-2711.801354010903196,
55.816799463192389])
else:
test = np.array([-501.76214132311253,
-19.339235392302943,
-14895.830345852655,
-5942.3558470931075,
-2719.492748965411,
54.85153478517509])
assert np.allclose(results["FX"], test[0], rtol=0., atol=1e-10) is True
assert np.allclose(results["FY"], test[1], rtol=0., atol=1e-10) is True
assert np.allclose(results["FZ"], test[2], rtol=0., atol=1e-10) is True
assert np.allclose(results["l"], test[3], rtol=0., atol=1e-10) is True
assert np.allclose(results["m"], test[4], rtol=0., atol=1e-10) is True
assert np.allclose(results["n"], test[5], rtol=0., atol=1e-10) is True
def tapered_control_surface_model():
"""Get a LLModel from the tapered controls example."""
filename = PLANE_DIR + "tapered_control_5sect.json"
plane = geom.Airplane(inputfile=filename)
model = LLModel(plane)
return model
def yoffset_wing_model():
"""Get a LLModel from the yoffset wing example."""
filename = PLANE_DIR + "yoffset_wing_5sect.json"
plane = geom.Airplane(inputfile=filename)
model = LLModel(plane)
return model
def aero_twist_wing_model():
"""Get a LLModel from the aero twist example."""
filename = PLANE_DIR + "aerodynamic_twist_wing_5sect.json"
plane = geom.Airplane(inputfile=filename)
model = LLModel(plane)
return model
def washout_wing_model():
"""Get a LLModel from the washout_wing_5sect.json example."""
filename = PLANE_DIR + "washout_wing_5sect.json"
plane = geom.Airplane(inputfile=filename)
model = LLModel(plane)
return model
def swept_wing_model():
"""Get a LLModel from the swept_wing.json example."""
filename = PLANE_DIR + "swept_wing.json"
plane = geom.Airplane(inputfile=filename)
model = LLModel(plane)
return model
def small_plane_model():
"""Get a LLModel from the straight_simple_plane.json example."""
filename = PLANE_DIR + "straight_simple_plane.json"
plane = geom.Airplane(inputfile=filename)
model = LLModel(plane)
return model
def v2_plane_model():
"""Get a LLModel from the test plane v2 example."""
filename = PLANE_DIR + "test_plane_v2.json"
plane = geom.Airplane(inputfile=filename)
model = LLModel(plane)
return model