def test_mean_to_true():
for row in ELLIPTIC_ANGLES_DATA:
ecc, M, expected_nu = row
ecc = ecc * u.one
M = M * u.deg
expected_nu = expected_nu * u.deg
nu = angles.E_to_nu(angles.M_to_E(M, ecc), ecc)
assert_quantity_allclose(nu, expected_nu, rtol=1e-4)
def test_eccentric_to_true_range2pi(E, ecc):
nu = angles.E_to_nu(E, ecc)
print(E.value, ecc.value, nu.value)
E1 = angles.nu_to_E(nu, ecc)
assert_quantity_allclose(E1, E, rtol=1e-8)
def test_eccentric_to_true_range(E, ecc):
nu = angles.E_to_nu(E, ecc)
E1 = angles.nu_to_E(nu, ecc)
assert_quantity_allclose(E1, E, rtol=1e-8)
def _propagate(self):
''' Propagate orbits of all satellites for each timestep in self.timesteps_arr'''
if self.constellation_type == 'walker':
sats_per_plane = self.walker_params[
'num_sats'] / self.walker_params['num_planes']
a = self.walker_params['a_km'] * u.km
ecc = 0 * u.one # walks are all circular
inc = self.walker_params['i_deg'] * u.deg
argp = 0 * u.deg
# raan increment
raan_inc = 360 / self.walker_params['num_planes'] * u.deg
# nu increment
nu_inc = 360 / sats_per_plane * u.deg
raan = 0 * u.deg # use this to space out based on num_planes
sat_id = 0
for _ in range(0, self.walker_params['num_planes']):
nu = 0 * u.deg # use this to space out based on num_sats/num_planes
for _ in range(0, int(sats_per_plane)):
# create baseline orbit
curOrb = Orbit.from_classical(Earth, a, ecc, inc, raan,
argp, nu)
print('Propagating sat number %d of %d over %d timesteps' %
(sat_id + 1, self.num_sats, len(self.timesteps_arr)))
(r_list, v_list) = self._propagate_curOrb(curOrb)
self.time_s_pos_eci_km[sat_id, :, :] = np.hstack(
(self.timesteps_arr, np.array(r_list)))
self.time_s_vel_eci_km[sat_id, :, :] = np.hstack(
(self.timesteps_arr, np.array(v_list)))
sat_id += 1
nu += nu_inc
raan += raan_inc
elif self.constellation_type == 'kepler_meananom':
# since each satellite has it's own orbit specified, we propagate one satellite at a time
for sat_id in range(self.num_sats):
orbit_params = self.all_orbit_params[sat_id][
self.constellation_type]
a = orbit_params['a_km'] * u.km
ecc = orbit_params['e'] * u.one
inc = orbit_params['i_deg'] * u.deg
raan = orbit_params['RAAN_deg'] * u.deg
# ASSUMES ARGUMENT OF PERIGEE IS ZERO
argp = 0 * u.deg
# convert mean anomaly to true anomaly
M = orbit_params['M_deg'] * u.deg
E = twobody_angles.M_to_E(M, ecc) # eccentric anomaly
nu = twobody_angles.E_to_nu(E, ecc) # true anomaly
curOrb = Orbit.from_classical(Earth, a, ecc, inc, raan, argp,
nu)
print('Propagating Sat %d of %d over %d timesteps' %
(sat_id + 1, self.num_sats, len(self.timesteps_arr)))
(r_list, v_list) = self._propagate_curOrb(curOrb)
self.time_s_pos_eci_km[sat_id, :, :] = np.hstack(
(self.timesteps_arr, np.array(r_list)))
self.time_s_vel_eci_km[sat_id, :, :] = np.hstack(
(self.timesteps_arr, np.array(v_list)))
else:
raise NotImplementedError(
'Only Walker constellation and Kelper Mean Anomoly specifications implemented'
)
self.propagation_complete = True