def side_force_coeff_test_cases():
constants = SkywalkerX8Constants()
b = constants.wing_span
state1 = State()
state1.vx = 20.0
state1.vy = 0.0
state1.vz = 0.0
state2 = State()
state2.vx = 28.6362532829
state2.vy = 1.0
state2.vz = 0.0
state2.ang_rate_x = 5 * np.pi / 180
state2.ang_rate_z = 5 * np.pi / 180
wind = np.zeros(6)
airspeed = np.sqrt(np.sum(calc_airspeed(state2, wind)**2))
zero_input = ControlInput()
aileron_input = ControlInput()
aileron_input.aileron_deflection = 2.0 * np.pi / 180.0
return [
(state1, wind, 0.0, zero_input, 2.99968641720902E-08),
(state1, wind, 1.0, zero_input, -7.94977329537982E-06),
(state2, wind, 0.0, aileron_input,
-0.008604744865183 + -0.085 * b / (2 * airspeed) * state2.ang_rate_x +
0.005 * b / (2 * airspeed) * state2.ang_rate_z +
0.0433 * aileron_input.aileron_deflection),
(state2, wind, 1.0, aileron_input,
-0.007089388672593 + -0.133 * b / (2 * airspeed) * state2.ang_rate_x +
0.002 * b / (2 * airspeed) * state2.ang_rate_z +
0.0433 * aileron_input.aileron_deflection)
]
def yaw_moment_coeff_test_cases():
constants = SkywalkerX8Constants()
b = constants.wing_span
state1 = State()
state1.vx = 20.0
state1.vy = 0.0
state1.vz = 0.0
state2 = State()
state2.vx = 28.6362532829
state2.vy = 1.0
state2.vz = 0.0
state2.ang_rate_x = 5 * np.pi / 180
state2.ang_rate_z = 5 * np.pi / 180
wind = np.zeros(6)
airspeed = np.sqrt(np.sum(calc_airspeed(state2, wind)**2))
zero_input = ControlInput()
aileron_input = ControlInput()
aileron_input.aileron_deflection = 2.0 * np.pi / 180.0
return [
(state1, wind, 0.0, zero_input, 4.9176697574439E-06),
(state1, wind, 1.0, zero_input, 1.96093394589053E-05),
(state2, wind, 0.0, aileron_input,
0.000825947539055 + 0.027 * b / (2 * airspeed) * state2.ang_rate_x +
-0.022 * b / (2 * airspeed) * state2.ang_rate_z -
0.00339 * aileron_input.aileron_deflection),
(state2, wind, 1.0, aileron_input,
0.001052911121301 + 0.017 * b / (2 * airspeed) * state2.ang_rate_x +
-0.049 * b / (2 * airspeed) * state2.ang_rate_z -
0.00339 * aileron_input.aileron_deflection)
]
def roll_moment_coeff_test_cases():
constants = SkywalkerX8Constants()
b = constants.wing_span
state1 = State()
state1.vx = 20.0
state1.vy = 0.0
state1.vz = 0.0
state2 = State()
state2.vx = 28.6362532829
state2.vy = 1.0
state2.vz = 0.0
state2.ang_rate_x = 5 * np.pi / 180
state2.ang_rate_z = 5 * np.pi / 180
wind = np.zeros(6)
airspeed = np.sqrt(np.sum(calc_airspeed(state2, wind)**2))
zero_input = ControlInput()
aileron_input = ControlInput()
aileron_input.aileron_deflection = 2.0 * np.pi / 180.0
return [
(state1, wind, 0.0, zero_input, -8.40821757613653E-05),
(state1, wind, 1.0, zero_input, -7.34515369827804E-05),
(state2, wind, 0.0, aileron_input,
-0.00380800071177 + -0.409 * b / (2 * airspeed) * state2.ang_rate_x +
0.039 * b / (2 * airspeed) * state2.ang_rate_z +
0.12 * aileron_input.aileron_deflection),
(state2, wind, 1.0, aileron_input,
-0.003067251004494 + -0.407 * b / (2 * airspeed) * state2.ang_rate_x +
0.158 * b / (2 * airspeed) * state2.ang_rate_z +
0.12 * aileron_input.aileron_deflection)
]
def body2wind_test_cases():
state1 = State()
vec1 = np.array([1.0, 0.0, 0.0])
wind1 = np.zeros(6)
expect1 = np.array([1.0, 0.0, 0.0])
state2 = State()
state2.vx = 10.0
vec2 = np.array([1.0, 0.0, 0.0])
wind2 = np.array([0.0, 10.0, 0.0])
expect2 = np.array([np.sqrt(2.0) / 2.0, -np.sqrt(2.0) / 2.0, 0.0])
state3 = State()
state3.vx = 10.0
vec3 = np.array([0.0, 1.0, 0.0])
wind3 = np.array([0.0, 10.0, 0.0])
expect3 = np.array([np.sqrt(2.0) / 2.0, np.sqrt(2.0) / 2.0, 0.0])
state4 = State()
state4.vx = 10.0
state4.vz = 10.0
vec4 = np.array([1.0, 0.0, 0.0])
wind4 = np.array([0.0, 0.0, 0.0])
expect4 = np.array([np.sqrt(2.0) / 2.0, 0.0, -np.sqrt(2.0) / 2.0])
return [(state1, vec1, wind1, expect1), (state2, vec2, wind2, expect2),
(state3, vec3, wind3, expect3), (state4, vec4, wind4, expect4)]
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)]
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)]