def test_elevator_deflection():
control_input = ControlInput()
elevator_deflection = 0.2
control_input.elevator_deflection = elevator_deflection
assert control_input.elevator_deflection == pytest.approx(
elevator_deflection)
assert np.allclose(control_input.control_input,
[elevator_deflection, 0.0, 0.0, 0.0])
def test_dynamics_forces():
control_input = ControlInput()
prop = SimpleTestAircraftNoMoments(control_input)
t = 0
for i in range(-50, 101, 50):
control_input.throttle = 0.8
control_input.elevator_deflection = i
control_input.aileron_deflection = i
control_input.rudder_deflection = i
state = State()
state.vx = 20.0
state.vy = 1
state.vz = 0
params = {"prop": prop, "wind": no_wind()}
update = dynamics_kinetmatics_update(t=t,
x=state.state,
u=control_input.control_input,
params=params)
V_a = np.sqrt(np.sum(calc_airspeed(state, params['wind'].get(0.0))**2))
forces_aero_wind_frame = np.array([
-np.abs(control_input.elevator_deflection),
control_input.aileron_deflection, -control_input.rudder_deflection
])
forces_aero_body_frame = wind2body(forces_aero_wind_frame, state,
params['wind'].get(0))
force_propulsion = np.array([(2 * control_input.throttle)**2 - V_a**2,
0, 0])
force_gravity = inertial2body(
np.array([0, 0, prop.mass() * GRAVITY_CONST]), state)
forces_body = forces_aero_body_frame + force_propulsion + force_gravity
vx_update_expect = (1 / prop.mass()) * forces_body[0]
vy_update_expect = (1 / prop.mass()) * forces_body[1]
vz_update_expect = (1 / prop.mass()) * forces_body[2]
# No moments
ang_rate_x_update_expect = 0
ang_rate_y_update_expect = 0
ang_rate_z_update_expect = 0
assert np.allclose(vx_update_expect, update[6])
assert np.allclose(vy_update_expect, update[7])
assert np.allclose(vz_update_expect, update[8])
assert np.allclose(ang_rate_x_update_expect, update[9])
assert np.allclose(ang_rate_y_update_expect, update[10])
assert np.allclose(ang_rate_z_update_expect, update[11])
def test_dynamics_moments():
control_input = ControlInput()
t = 0
for i in range(-50, 51, 50):
control_input.throttle = i
control_input.elevator_deflection = i
control_input.aileron_deflection = i
control_input.rudder_deflection = i
prop = SimpleTestAircraftNoForces(control_input)
state = State()
state.vx = 20.0
state.vy = 1
state.vz = 0
state.ang_rate_x = 0.157079633
state.ang_rate_y = 0.157079633
state.ang_rate_z = 0.157079633
params = {"prop": prop, "wind": no_wind()}
update = dynamics_kinetmatics_update(t=t,
x=state.state,
u=control_input.control_input,
params=params)
moments_aero = np.array([
control_input.elevator_deflection,
control_input.aileron_deflection, control_input.rudder_deflection
])
omega = np.array(
[state.ang_rate_x, state.ang_rate_y, state.ang_rate_z])
coreolis_term = prop.inv_inertia_matrix().dot(
np.cross(omega,
prop.inertia_matrix().dot(omega)))
ang_rate_x_update_expect = (2 / 3) * moments_aero[0] - (
1 / 3) * moments_aero[2] - coreolis_term[0]
ang_rate_y_update_expect = (1 / 2) * moments_aero[1] - coreolis_term[1]
ang_rate_z_update_expect = (2 / 3) * moments_aero[2] - (
1 / 3) * moments_aero[0] - coreolis_term[2]
assert np.allclose(ang_rate_x_update_expect, update[9])
assert np.allclose(ang_rate_y_update_expect, update[10])
assert np.allclose(ang_rate_z_update_expect, update[11])
def drag_coeff_test_cases():
state1 = State()
state1.vx = 20.0
state1.vz = 0.0
wind1 = np.zeros(6)
wind2 = np.zeros(6)
wind2[0] = 1.0
wind2[2] = -19.0 * np.tan(6 * np.pi / 180.0)
zero_input = ControlInput()
elevator_input = ControlInput()
elevator_input.elevator_deflection = 2.0 * np.pi / 180.0
return [(state1, wind1, 0.0, zero_input, 0.015039166436721),
(state1, wind1, 1.0, zero_input, 0.043224285541117),
(state1, wind2, 0.0, zero_input, 0.027939336576642),
(state1, wind2, 1.0, zero_input, 0.082879203470842),
(state1, wind1, 0.0, elevator_input,
0.015039166436721 + 0.0633 * elevator_input.elevator_deflection),
(state1, wind1, 1.0, elevator_input,
0.043224285541117 + 0.0633 * elevator_input.elevator_deflection)]
def pitch_moment_coeff_test_cases():
state1 = State()
state1.vx = 20.0
state1.vz = 0.0
wind1 = np.zeros(6)
wind2 = np.zeros(6)
wind2[0] = 1.0
wind2[2] = -19.0 * np.tan(6 * np.pi / 180.0)
zero_input = ControlInput()
elevator_input = ControlInput()
elevator_input.elevator_deflection = 2.0 * np.pi / 180.0
return [(state1, wind1, 0.0, zero_input, 0.001161535578433),
(state1, wind1, 1.0, zero_input, -0.004297270367523),
(state1, wind2, 0.0, zero_input, -0.064477317568596),
(state1, wind2, 1.0, zero_input, -0.01808304889307),
(state1, wind1, 0.0, elevator_input,
0.001161535578433 - 0.206 * elevator_input.elevator_deflection),
(state1, wind1, 1.0, elevator_input,
-0.004297270367523 - 0.206 * elevator_input.elevator_deflection)]
def lift_coeff_test_cases():
constants = SkywalkerX8Constants()
c = constants.mean_chord
state1 = State()
state1.vx = 20.0
state1.vz = 0.0
state2 = State()
state2.vx = 20.0
state2.vz = 0.0
state2.ang_rate_y = 5 * np.pi / 180
wind1 = np.zeros(6)
wind2 = np.zeros(6)
wind3 = np.zeros(6)
wind2[0] = 1.0
wind2[2] = -19.0 * np.tan(8 * np.pi / 180.0)
wind3[0] = 1.0
wind3[2] = -19.0 * np.tan(8 * np.pi / 180.0)
wind3[4] = 3 * np.pi / 180
ang_rate_y2 = state2.ang_rate_y - wind3[4]
airspeed2 = np.sqrt(np.sum(calc_airspeed(state2, wind2)**2))
zero_input = ControlInput()
elevator_input = ControlInput()
elevator_input.elevator_deflection = 2.0 * np.pi / 180.0
return [(state1, wind1, 0.0, zero_input, 0.030075562375465),
(state1, wind1, 1.0, zero_input, 0.018798581619545),
(state1, wind2, 0.0, zero_input, 0.609296679062686),
(state1, wind2, 1.0, zero_input, 0.454153721254944),
(state1, wind1, 0.0, elevator_input,
0.030075562375465 + 0.278 * elevator_input.elevator_deflection),
(state1, wind1, 1.0, elevator_input,
0.018798581619545 + 0.278 * elevator_input.elevator_deflection),
(state2, wind3, 0.0, elevator_input,
0.609296679062686 + 4.60 * c / (2 * airspeed2) * ang_rate_y2 +
0.278 * elevator_input.elevator_deflection),
(state2, wind3, 1.0, elevator_input,
0.454153721254944 - 3.51 * c / (2 * airspeed2) * ang_rate_y2 +
0.278 * elevator_input.elevator_deflection)]