def test_orbit_wrong_dimensions_fails():
bad_r = [[1000, 0, 0]] * u.km
bad_v = [[[0, 10, 0]]] * u.km / u.s
with pytest.raises(ValueError) as excinfo:
Orbit.from_vectors(Earth, bad_r, bad_v)
assert "ValueError: Vectors must have dimension 1, got 2 and 3" in excinfo.exconly()
def test_from_horizons_raise_valueerror():
with pytest.raises(ValueError) as exep:
Orbit.from_horizons(name="Dummy")
assert (
"ValueError: Unknown target (Dummy). Maybe try different id_type?"
in exep.exconly()
)
def test_frozen_orbit_altitude():
with pytest.raises(ValueError) as excinfo:
Orbit.frozen(Earth, -1 * u.m)
assert excinfo.type == ValueError
assert (
str(excinfo.value)
== "The semimajor axis may not be smaller that Earth's radius"
)
def test_frozen_orbit_non_spherical_arguments():
with pytest.raises(AttributeError) as excinfo:
Orbit.frozen(Jupiter, 1 * u.m)
assert excinfo.type == AttributeError
assert (
str(excinfo.value)
== "Attractor Jupiter has not spherical harmonics implemented"
)
def test_geostationary_input(attractor):
with pytest.raises(ValueError) as excinfo:
Orbit.geostationary(attractor=attractor)
assert (
"ValueError: At least one among angular_velocity or period must be passed"
in excinfo.exconly()
)
def test_geostationary_non_existence_condition(attractor, period, hill_radius):
with pytest.raises(ValueError) as excinfo:
Orbit.geostationary(attractor=attractor, period=period, hill_radius=hill_radius)
assert (
"Geostationary orbit for the given parameters doesn't exist"
in excinfo.exconly()
)
def test_from_sbdb_raise_valueerror():
with pytest.raises(ValueError) as excinfo:
Orbit.from_sbdb(name="Halley")
assert (
str(excinfo.value)
== "2 different objects found: \n2688 Halley (1982 HG1)\n1P/Halley\n"
)
def test_from_ephem_raises_warning_if_time_is_not_tdb_with_proper_time(recwarn):
body = Earth
epoch = Time("2017-09-29 07:31:26", scale="utc")
expected_epoch_string = "2017-09-29 07:32:35.182" # epoch.tdb.value
Orbit.from_body_ephem(body, epoch)
w = recwarn.pop(TimeScaleWarning)
assert expected_epoch_string in str(w.message)
def test_bad_hyperbolic_raises_exception():
bad_a = 1.0 * u.AU
ecc = 1.5 * u.one
_inc = 100 * u.deg # Unused inclination
_a = 1.0 * u.deg # Unused angle
_body = Sun # Unused body
with pytest.raises(ValueError) as excinfo:
Orbit.from_classical(_body, bad_a, ecc, _inc, _a, _a, _a)
assert "Hyperbolic orbits have negative semimajor axis" in excinfo.exconly()
def test_state_raises_unitserror_if_rv_units_are_wrong():
_d = [1.0, 0.0, 0.0] * u.AU
wrong_v = [0.0, 1.0e-6, 0.0] * u.AU
with pytest.raises(u.UnitsError) as excinfo:
Orbit.from_vectors(Sun, _d, wrong_v)
assert (
"UnitsError: Argument 'v' to function 'from_vectors' must be in units convertible to 'm / s'."
in excinfo.exconly()
)
def test_plot_trajectory_sets_label():
op = OrbitPlotter()
earth = Orbit.from_body_ephem(Earth)
mars = Orbit.from_body_ephem(Mars)
trajectory = earth.sample()
op.plot(mars, label="Mars")
op.plot_trajectory(trajectory, label="Earth")
legend = plt.gca().get_legend()
assert legend.get_texts()[1].get_text() == "Earth"
def test_bad_inclination_raises_exception():
_d = 1.0 * u.AU # Unused distance
_ = 0.5 * u.one # Unused dimensionless value
_a = 1.0 * u.deg # Unused angle
_body = Sun # Unused body
bad_inc = 200 * u.deg
with pytest.raises(ValueError) as excinfo:
Orbit.from_classical(_body, _d, _, bad_inc, _a, _a, _a)
assert (
"ValueError: Inclination must be between 0 and 180 degrees" in excinfo.exconly()
)
def test_state_raises_unitserror_if_elements_units_are_wrong():
_d = 1.0 * u.AU # Unused distance
_ = 0.5 * u.one # Unused dimensionless value
_a = 1.0 * u.deg # Unused angle
wrong_angle = 1.0 * u.AU
with pytest.raises(u.UnitsError) as excinfo:
Orbit.from_classical(Sun, _d, _, _a, _a, _a, wrong_angle)
assert (
"UnitsError: Argument 'nu' to function 'from_classical' must be in units convertible to 'rad'."
in excinfo.exconly()
)
def test_parabolic_elements_fail_early():
attractor = Earth
ecc = 1.0 * u.one
_d = 1.0 * u.AU # Unused distance
_a = 1.0 * u.deg # Unused angle
with pytest.raises(ValueError) as excinfo:
Orbit.from_classical(attractor, _d, ecc, _a, _a, _a, _a)
assert (
"ValueError: For parabolic orbits use Orbit.parabolic instead"
in excinfo.exconly()
)
def test_from_coord_fails_if_no_time_differential():
pos = [30000, 0, 0] * u.km
cartrep = CartesianRepresentation(*pos)
# Method fails if coordinate instance doesn't contain a differential with
# respect to time
with pytest.raises(ValueError) as excinfo:
Orbit.from_coords(Earth, SkyCoord(cartrep))
assert (
"ValueError: Coordinate instance doesn't have a differential with respect to time"
in excinfo.exconly()
)
def test_from_coord_fails_for_multiple_positions(obstime):
cartdiff = CartesianDifferential(
[[0, 1, 0], [-0.1, 0.9, 0]] * u.km / u.s, xyz_axis=1
)
cartrep = CartesianRepresentation(
[[1, 0, 0], [0.9, 0.1, 0]] * u.km, differentials=cartdiff, xyz_axis=1
)
coords = GCRS(cartrep, representation_type=CartesianRepresentation, obstime=obstime)
with pytest.raises(ValueError) as excinfo:
Orbit.from_coords(Earth, coords)
assert (
"ValueError: Coordinate instance must represents exactly 1 position, found: 2"
in excinfo.exconly()
)
def orbit_from_record(record):
"""Return :py:class:`~poliastro.twobody.orbit.Orbit` given a record.
Retrieve info from JPL DASTCOM5 database.
Parameters
----------
record : int
Object record.
Returns
-------
orbit : ~poliastro.twobody.orbit.Orbit
NEO orbit.
"""
body_data = read_record(record)
a = body_data['A'].item() * u.au
ecc = body_data['EC'].item() * u.one
inc = body_data['IN'].item() * u.deg
raan = body_data['OM'].item() * u.deg
argp = body_data['W'].item() * u.deg
m = body_data['MA'].item() * u.deg
nu = M_to_nu(m, ecc)
epoch = Time(body_data['EPOCH'].item(), format='jd', scale='tdb')
orbit = Orbit.from_classical(Sun, a, ecc, inc, raan, argp, nu, epoch)
orbit._frame = HeliocentricEclipticJ2000(obstime=epoch)
return orbit
def test_inertial_body_centered_to_pqw():
molniya_r_peri, molniya_v_peri = coordinates.inertial_body_centered_to_pqw(molniya.r, molniya.v, bodies.Earth)
molniya_peri = Orbit.from_vectors(bodies.Earth, molniya_r_peri, molniya_v_peri, molniya.epoch)
assert_quantity_allclose(molniya_peri.e_vec[-2:], [0, 0], atol=1e-12)
assert_quantity_allclose(norm(molniya_peri.e_vec), norm(molniya.e_vec))
def test_perifocal_points_to_perigee():
_d = 1.0 * u.AU # Unused distance
_ = 0.5 * u.one # Unused dimensionless value
_a = 1.0 * u.deg # Unused angle
ss = Orbit.from_classical(Sun, _d, _, _a, _a, _a, _a)
p, _, _ = ss.pqw()
assert_allclose(p, ss.e_vec / ss.ecc)
def transform(orbit, frame_orig, frame_dest):
"""Transforms Orbit from one frame to another.
Parameters
----------
orbit : ~poliastro.bodies.Orbit
Orbit to transform
frame_orig : ~astropy.coordinates.BaseCoordinateFrame
Initial frame
frame_dest : ~astropy.coordinates.BaseCoordinateFrame
Final frame
Returns
-------
orbit: ~poliastro.bodies.Orbit
Orbit in the new frame
"""
orbit_orig = frame_orig(x=orbit.r[0], y=orbit.r[1], z=orbit.r[2],
v_x=orbit.v[0], v_y=orbit.v[1], v_z=orbit.v[2],
representation=CartesianRepresentation,
differential_type=CartesianDifferential)
orbit_dest = orbit_orig.transform_to(frame_dest(obstime=orbit.epoch))
orbit_dest.representation = CartesianRepresentation
return Orbit.from_vectors(orbit.attractor,
orbit_dest.data.xyz,
orbit_dest.data.differentials['s'].d_xyz,
epoch=orbit.epoch)
def plot_solar_system(outer=True, epoch=None):
"""
Plots the whole solar system in one single call.
.. versionadded:: 0.9.0
Parameters
------------
outer : bool, optional
Whether to print the outer Solar System, default to True.
epoch: ~astropy.time.Time, optional
Epoch value of the plot, default to J2000.
"""
bodies = [Mercury, Venus, Earth, Mars]
if outer:
bodies.extend([Jupiter, Saturn, Uranus, Neptune])
op = OrbitPlotter()
for body in bodies:
orb = Orbit.from_body_ephem(body, epoch)
op.plot(orb, label=str(body))
# Sets frame to the orbit of the Earth by default
# TODO: Wait until https://github.com/poliastro/poliastro/issues/316
# op.set_frame(*Orbit.from_body_ephem(Earth, epoch).pqw())
return op
def test_parabolic_has_proper_eccentricity():
attractor = Earth
_d = 1.0 * u.AU # Unused distance
_a = 1.0 * u.deg # Unused angle
expected_ecc = 1.0 * u.one
ss = Orbit.parabolic(attractor, _d, _a, _a, _a, _a)
assert_allclose(ss.ecc, expected_ecc)
def test_orbit_representation():
ss = Orbit.circular(
Earth, 600 * u.km, 20 * u.deg, epoch=Time("2018-09-08 09:04:00", scale="tdb")
)
expected_str = "6978 x 6978 km x 20.0 deg (GCRS) orbit around Earth (\u2641) at epoch 2018-09-08 09:04:00.000 (TDB)"
assert str(ss) == repr(ss) == expected_str
def test_orbit_accepts_ecliptic_plane():
r = [1e09, -4e09, -1e09] * u.km
v = [5e00, -1e01, -4e00] * u.km / u.s
ss = Orbit.from_vectors(Sun, r, v, plane=Planes.EARTH_ECLIPTIC)
assert ss.frame.is_equivalent_frame(HeliocentricEclipticJ2000(obstime=J2000))
def test_orbit_plot_static_3d():
# Data from Curtis, example 4.3
r = [-6045, -3490, 2500] * u.km
v = [-3.457, 6.618, 2.533] * u.km / u.s
ss = Orbit.from_vectors(Earth, r, v)
with pytest.raises(ValueError, match="static and use_3d cannot be true"):
ss.plot(static=True, use_3d=True)
def test_sample_numpoints():
_d = 1.0 * u.AU # Unused distance
_ = 0.5 * u.one # Unused dimensionless value
_a = 1.0 * u.deg # Unused angle
_body = Sun # Unused body
ss = Orbit.from_classical(_body, _d, _, _a, _a, _a, _a)
positions = ss.sample(values=50)
assert len(positions) == 50
def test_geostationary_creation_with_Hill_radius(
attractor, period, hill_radius, expected_a
):
ss = Orbit.geostationary(
attractor=attractor, period=period, hill_radius=hill_radius
)
assert_quantity_allclose(ss.a, expected_a, rtol=1.0e-7)
assert_quantity_allclose(ss.period, period, rtol=1.0e-7)
def test_default_time_for_new_state():
_d = 1.0 * u.AU # Unused distance
_ = 0.5 * u.one # Unused dimensionless value
_a = 1.0 * u.deg # Unused angle
_body = Sun # Unused body
expected_epoch = J2000
ss = Orbit.from_classical(_body, _d, _, _a, _a, _a, _a)
assert ss.epoch == expected_epoch
def test_show_calls_prepare_plot(mock_prepare_plot, mock_iplot):
m = OrbitPlotter2D()
earth = Orbit.from_body_ephem(Earth)
m.plot(orbit=earth, label="Obj")
m.show()
assert mock_iplot.call_count == 1
mock_prepare_plot.assert_called_once_with()
def test_orbit_no_frame_representation():
date_launch = Time("2011-11-26 15:02", scale="utc")
r = [61445.76498656, 24827.93010168, 0.0] * u.km
v = [-0.42581645, -0.18867869, 0.0] * u.km / u.s
ss = Orbit.from_vectors(Moon, r, v, date_launch)
expected_str = "106 x -142299 km x 180.0 deg orbit around Moon (\u263E) at epoch 2011-11-26 15:02:00.000 (UTC)"
assert str(ss) == repr(ss) == expected_str
def orbit_from_orbit(init_orbit, part_of_period, alfa_init): # alfa in degrees
period_init = init_orbit.period
t_maneuver = part_of_period * period_init
mean_motion_init = init_orbit.n
anomaly_init = (alfa_init / 180 * np.pi +
mean_motion_init * t_maneuver) * u.rad
orbit_dprt = init_orbit.propagate_to_anomaly(anomaly_init)
r = orbit_dprt.r
v = orbit_dprt.v
date_arrival = time.Time((DATA_LAUNCH.unix + t_maneuver), format='unix')
orbit = Orbit.from_vectors(Earth, r, v, epoch=date_arrival)
return orbit
def test_orbit_represent_as_produces_correct_data():
r = [1e09, -4e09, -1e09] * u.km
v = [5e00, -1e01, -4e00] * u.km / u.s
ss = Orbit.from_vectors(Sun, r, v)
expected_result = CartesianRepresentation(
*r, differentials=CartesianDifferential(*v))
result = ss.represent_as(CartesianRepresentation)
# We can't directly compare the objects, see
# https://github.com/astropy/astropy/issues/7793
assert (result.xyz == expected_result.xyz).all()
assert (result.differentials["s"].d_xyz ==
expected_result.differentials["s"].d_xyz).all()
def test_expected_angular_momentum():
# Example from Curtis
expected_ang_mag = 72472 * u.km**2
attractor = Earth
_a = 1.0 * u.deg # Unused angle
a = 15_300 * u.km
ecc = 0.37255 * u.one
nu = 120 * u.deg
orbit = Orbit.from_classical(attractor, a, ecc, _a, _a, _a, nu)
orbit_h_mag = orbit.h_mag
assert_quantity_allclose(orbit_h_mag.value,
expected_ang_mag.value,
rtol=1e-2)
def test_convert_from_rv_to_coe():
# Data from Vallado, example 2.6
attractor = Earth
p = 11067.790 * u.km
ecc = 0.83285 * u.one
inc = 87.87 * u.deg
raan = 227.89 * u.deg
argp = 53.38 * u.deg
nu = 92.335 * u.deg
expected_r = [6525.344, 6861.535, 6449.125] * u.km
expected_v = [4.902276, 5.533124, -1.975709] * u.km / u.s
r, v = Orbit.from_classical(attractor, p / (1 - ecc**2), ecc, inc, raan,
argp, nu).rv()
assert_quantity_allclose(r, expected_r, rtol=1e-5)
assert_quantity_allclose(v, expected_v, rtol=1e-5)
def test_expected_mean_anomaly():
# Example from Curtis
expected_mean_anomaly = 77.93 * u.deg
attractor = Earth
_a = 1.0 * u.deg # Unused angle
a = 15_300 * u.km
ecc = 0.37255 * u.one
nu = 120 * u.deg
orbit = Orbit.from_classical(attractor, a, ecc, _a, _a, _a, nu)
orbit_M = orbit.M
assert_quantity_allclose(orbit_M.value,
expected_mean_anomaly.value,
rtol=1e-2)
def test_orbit_creation_using_skycoord(attractor):
vel = [0, 2, 0] * u.km / u.s
cartdiff = CartesianDifferential(*vel)
pos = [30_000, 0, 0] * u.km
cartrep = CartesianRepresentation(*pos, differentials=cartdiff)
coord = SkyCoord(cartrep, frame="icrs")
o = Orbit.from_coords(attractor, coord)
inertial_frame_at_body_centre = get_frame(
attractor, Planes.EARTH_EQUATOR, obstime=coord.obstime
)
coord_transformed_to_irf = coord.transform_to(inertial_frame_at_body_centre)
pos_transformed_to_irf = coord_transformed_to_irf.cartesian.xyz
vel_transformed_to_irf = coord_transformed_to_irf.cartesian.differentials["s"].d_xyz
assert (o.r == pos_transformed_to_irf).all()
assert (o.v == vel_transformed_to_irf).all()
def test_from_sbdb():
# Dictionary with structure: 'Object': [a, e, i, raan, argp, nu, epoch]
# Notice JPL provides Mean anomaly, a conversion is needed to obtain nu
SBDB_DATA = {
"Ceres": (
2.76916515450648 * u.AU,
0.07600902910070946 * u.one,
10.59406704424526 * u.deg,
80.30553156826473 * u.deg,
73.597694115971 * u.deg,
86.01851107780747 * u.deg,
)
}
for target_name in SBDB_DATA.keys():
ss_target = Orbit.from_sbdb(target_name)
ss_classical = ss_target.classical()
assert ss_classical == SBDB_DATA[target_name]
def from_vectors(cls, r, v):
"""Return EarthSatellite with the `Orbit` from position and velocity vectors.
Parameters
----------
r : ~astropy.units.Quantity
Position vector wrt attractor center.
v : ~astropy.units.Quantity
Velocity vector.
Returns
-------
EarthSatellite
New EarthSatellite with the'Orbit' position and velocity vectors.
"""
orbit = Orbit.from_vectors(Earth, r, v)
return cls(orbit)
def test_apply_maneuver_returns_intermediate_states_if_true():
_d = 1.0 * u.AU # Unused distance
_ = 0.5 * u.one # Unused dimensionless value
_a = 1.0 * u.deg # Unused angle
ss = Orbit.from_classical(attractor=Sun,
a=_d,
ecc=_,
inc=_a,
raan=_a,
argp=_a,
nu=_a)
dt1 = 0.5 * u.h
dv1 = [5, 0, 10] * u.km / u.s
dt2 = 0.5 * u.h
dv2 = [0, 5, 10] * u.km / u.s
states = ss.apply_maneuver([(dt1, dv1), (dt2, dv2)], intermediate=True)
assert len(states) == 2
assert states[-1].epoch == ss.epoch + dt1 + dt2
def test_orbit_creation_using_frame_obj(attractor, frame, obstime):
vel = [0, 2, 0] * u.km / u.s
cartdiff = CartesianDifferential(*vel)
pos = [30_000, 0, 0] * u.km
cartrep = CartesianRepresentation(*pos, differentials=cartdiff)
coord = frame(cartrep, obstime=obstime)
o = Orbit.from_coords(attractor, coord)
inertial_frame_at_body_centre = get_frame(
attractor, Planes.EARTH_EQUATOR, obstime=coord.obstime
)
coord_transformed_to_irf = coord.transform_to(inertial_frame_at_body_centre)
pos_transformed_to_irf = coord_transformed_to_irf.cartesian.xyz
vel_transformed_to_irf = coord_transformed_to_irf.cartesian.differentials["s"].d_xyz
assert_quantity_allclose(o.r, pos_transformed_to_irf, atol=1e-5 * u.km)
assert_quantity_allclose(o.v, vel_transformed_to_irf, atol=1e-5 * u.km / u.s)
def test_expected_last_perifocal_passage():
# Example from Curtis
expected_t_p = 4077 * u.s
attractor = Earth
_a = 1.0 * u.deg # Unused angle
a = 15_300 * u.km
ecc = 0.37255 * u.one
nu = 120 * u.deg
orbit = Orbit.from_classical(attractor=attractor,
a=a,
ecc=ecc,
inc=_a,
raan=_a,
argp=_a,
nu=nu)
orbit_t_p = orbit.t_p
assert_quantity_allclose(orbit_t_p, expected_t_p, rtol=1e-2)
def test_expected_mean_anomaly():
# Example from Curtis
expected_mean_anomaly = 77.93 * u.deg
attractor = Earth
_a = 1.0 * u.deg # Unused angle
a = 15_300 * u.km
ecc = 0.37255 * u.one
nu = 120 * u.deg
orbit = Orbit.from_classical(attractor=attractor,
a=a,
ecc=ecc,
inc=_a,
raan=_a,
argp=_a,
nu=nu)
orbit_M = E_to_M(nu_to_E(orbit.nu, orbit.ecc), orbit.ecc)
assert_quantity_allclose(orbit_M, expected_mean_anomaly, rtol=1e-2)
def test_from_orbit_has_desired_properties(method, rtol):
epochs = Time(
[
"2020-02-01 12:00:00",
"2020-02-13 12:00:00",
"2020-03-04 12:00:00",
"2020-03-17 12:00:00",
]
)
expected_coordinates = CartesianRepresentation(
[
(336.77109079, -1447.38211842, -743.72094119),
(-1133.43957703, 449.41297342, 3129.10416554),
(201.42480053, -1978.64139325, -287.25776291),
(-1084.94556884, -1713.5357774, 3298.72284309),
]
* u.km,
xyz_axis=1,
differentials=CartesianDifferential(
[
(-2.68502706, -14.85798508, 9.66683585),
(1.36841306, 7.30080155, -4.88822441),
(-3.46999908, -10.04899184, 11.19715233),
(-1.46069855, 5.88696886, 3.28112281),
]
* (u.km / u.s),
xyz_axis=1,
),
)
r = [-1000, -2000, 3100] * u.km
v = [-1.836, 5.218, 4.433] * u.km / u.s
orb = Orbit.from_vectors(Earth, r, v)
unused_plane = Planes.EARTH_EQUATOR
ephem = Ephem.from_orbit(orbit=orb, epochs=epochs, plane=unused_plane)
coordinates = ephem.sample()
assert ephem.epochs is epochs
assert_coordinates_allclose(coordinates, expected_coordinates, rtol=rtol)
def orbit_from_spk_id(spk_id, api_key='DEMO_KEY'):
"""Return :py:class:`~poliastro.twobody.orbit.Orbit` given a SPK-ID.
Retrieve info from NASA NeoWS API, and therefore
it only works with NEAs (Near Earth Asteroids).
Parameters
----------
spk_id : str
SPK-ID number, which is given to each body by JPL.
api_key : str
NASA OPEN APIs key (default: `DEMO_KEY`)
Returns
-------
orbit : ~poliastro.twobody.orbit.Orbit
NEA orbit.
"""
payload = {'api_key': api_key}
response = requests.get(NEOWS_URL + spk_id, params=payload)
response.raise_for_status()
orbital_data = response.json()['orbital_data']
attractor = Sun
a = float(orbital_data['semi_major_axis']) * u.AU
ecc = float(orbital_data['eccentricity']) * u.one
inc = float(orbital_data['inclination']) * u.deg
raan = float(orbital_data['ascending_node_longitude']) * u.deg
argp = float(orbital_data['perihelion_argument']) * u.deg
m = float(orbital_data['mean_anomaly']) * u.deg
nu = M_to_nu(m.to(u.rad), ecc)
epoch = Time(float(orbital_data['epoch_osculation']),
format='jd',
scale='tdb')
return Orbit.from_classical(attractor, a, ecc, inc, raan, argp, nu, epoch)
def transform(orbit, frame_orig, frame_dest):
"""Transforms Orbit from one frame to another.
Parameters
----------
orbit : ~poliastro.bodies.Orbit
Orbit to transform
frame_orig : ~astropy.coordinates.BaseCoordinateFrame
Initial frame
frame_dest : ~astropy.coordinates.BaseCoordinateFrame
Final frame
Returns
-------
orbit: ~poliastro.bodies.Orbit
Orbit in the new frame
"""
orbit_orig = frame_orig(
x=orbit.r[0],
y=orbit.r[1],
z=orbit.r[2],
v_x=orbit.v[0],
v_y=orbit.v[1],
v_z=orbit.v[2],
representation=CartesianRepresentation,
differential_type=CartesianDifferential,
)
orbit_dest = orbit_orig.transform_to(frame_dest(obstime=orbit.epoch))
orbit_dest.representation = CartesianRepresentation
return Orbit.from_vectors(
orbit.attractor,
orbit_dest.data.xyz,
orbit_dest.data.differentials["s"].d_xyz,
epoch=orbit.epoch,
)
def test_from_sbdb():
# Dictionary with structure: 'Object': [a, e, i, raan, argp, nu, epoch]
# Notice JPL provides Mean anomaly, a conversion is needed to obtain nu
SBDB_DATA = {
"Ceres": (
2.769165146349478 * u.AU,
0.07600902762923671 * u.one,
10.59406732590292 * u.deg,
80.30553084093981 * u.deg,
73.59769486239257 * u.deg,
M_to_nu(77.37209773768207 * u.deg, 0.07600902762923671 * u.one),
)
}
for target_name in SBDB_DATA.keys():
ss_target = Orbit.from_sbdb(target_name)
ss_classical = ss_target.classical()
assert ss_classical == SBDB_DATA[target_name]
def orbit_from_record(record):
"""Return :py:class:`~poliastro.twobody.orbit.Orbit` given a record.
Retrieve info from JPL DASTCOM5 database.
Parameters
----------
record : int
Object record.
Returns
-------
orbit : ~poliastro.twobody.orbit.Orbit
NEO orbit.
"""
body_data = read_record(record)
a = body_data["A"].item() * u.au
ecc = body_data["EC"].item() * u.one
inc = body_data["IN"].item() * u.deg
raan = body_data["OM"].item() * u.deg
argp = body_data["W"].item() * u.deg
M = body_data["MA"].item() * u.deg
epoch = Time(body_data["EPOCH"].item(), format="jd", scale="tdb")
# NOTE: It is unclear how this conversion should happen,
# see https://ssd-api.jpl.nasa.gov/doc/sbdb.html
if ecc < 1:
M = (M + np.pi * u.rad) % (2 * np.pi * u.rad) - np.pi * u.rad
nu = E_to_nu(M_to_E(M, ecc), ecc)
elif ecc == 1:
nu = D_to_nu(M_to_D(M))
else:
nu = F_to_nu(M_to_F(M, ecc), ecc)
orbit = Orbit.from_classical(Sun, a, ecc, inc, raan, argp, nu, epoch)
orbit._frame = HeliocentricEclipticJ2000(obstime=epoch)
return orbit
def test_propagate_instance():
tof = 1.0 * u.min
ss0 = Orbit.from_classical(
Earth,
1000 * u.km,
0.75 * u.one,
63.4 * u.deg,
0 * u.deg,
270 * u.deg,
80 * u.deg,
)
C_D = 2.2 * u.one # dimentionless (any value would do)
A = ((np.pi / 4.0) * (u.m**2)).to(u.km**2)
m = 100 * u.kg
spacecraft = Spacecraft(A, C_D, m)
earth_satellite = EarthSatellite(ss0, spacecraft)
orbit_with_j2 = earth_satellite.propagate(tof=tof, gravity=EarthGravity.J2)
orbit_without_perturbation = earth_satellite.propagate(tof)
orbit_with_atmosphere_and_j2 = earth_satellite.propagate(
tof=tof, gravity=EarthGravity.J2, atmosphere=COESA76())
assert isinstance(orbit_with_j2, EarthSatellite)
assert isinstance(orbit_with_atmosphere_and_j2, EarthSatellite)
assert isinstance(orbit_without_perturbation, EarthSatellite)
def test_convert_from_coe_to_rv():
# Data from Vallado, example 2.5
attractor = Earth
r = [6524.384, 6862.875, 6448.296] * u.km
v = [4.901327, 5.533756, -1.976341] * u.km / u.s
expected_p = 11067.79 * u.km
expected_ecc = 0.832853 * u.one
expected_inc = 87.870 * u.deg
expected_raan = 227.89 * u.deg
expected_argp = 53.38 * u.deg
expected_nu = 92.335 * u.deg
ss = Orbit.from_vectors(attractor, r, v)
_, ecc, inc, raan, argp, nu = ss.classical()
p = ss.p
assert_quantity_allclose(p, expected_p, rtol=1e-4)
assert_quantity_allclose(ecc, expected_ecc, rtol=1e-4)
assert_quantity_allclose(inc, expected_inc, rtol=1e-4)
assert_quantity_allclose(raan, expected_raan, rtol=1e-4)
assert_quantity_allclose(argp, expected_argp, rtol=1e-4)
assert_quantity_allclose(nu, expected_nu, rtol=1e-4)
def test_propagate_instance():
tof = 1.0 * u.min
ss0 = Orbit.from_classical(
attractor=Earth,
a=1000 * u.km,
ecc=0.75 * u.one,
inc=63.4 * u.deg,
raan=0 * u.deg,
argp=270 * u.deg,
nu=80 * u.deg,
)
C_D = 2.2 * u.one # Dimensionless (any value would do)
A = ((np.pi / 4.0) * (u.m**2)).to(u.km**2)
m = 100 * u.kg
spacecraft = Spacecraft(A, C_D, m)
earth_satellite = EarthSatellite(ss0, spacecraft)
orbit_with_j2 = earth_satellite.propagate(tof=tof, gravity=EarthGravity.J2)
orbit_without_perturbation = earth_satellite.propagate(tof)
# orbit_with_atmosphere_and_j2 = earth_satellite.propagate(
# tof=tof, gravity=EarthGravity.J2, atmosphere=COESA76()
# )
assert isinstance(orbit_with_j2, EarthSatellite)
# assert isinstance(orbit_with_atmosphere_and_j2, EarthSatellite)
assert isinstance(orbit_without_perturbation, EarthSatellite)
def test_from_sbdb_raise_valueerror():
with pytest.raises(ValueError) as excinfo:
Orbit.from_sbdb(name="Halley")
assert (str(excinfo.value) ==
"2 different objects found: \n2688 Halley (1982 HG1)\n1P/Halley\n")
def test_arglat_within_range():
r = [3539.08827417, 5310.19903462, 3066.31301457] * u.km
v = [-6.49780849, 3.24910291, 1.87521413] * u.km / u.s
ss = Orbit.from_vectors(Earth, r, v)
assert 0 * u.deg <= ss.arglat <= 360 * u.deg
def test_geostationary_creation_from_period(attractor, period, expected_a):
ss = Orbit.geostationary(attractor=attractor, period=period)
assert_quantity_allclose(ss.a, expected_a, rtol=1.0e-7)
assert_quantity_allclose(ss.period, period, rtol=1.0e-7)
def test_plane_is_set_in_horizons():
plane = Planes.EARTH_ECLIPTIC
ss = Orbit.from_horizons(name="Ceres", plane=plane)
assert ss.plane == plane
def test_from_horizons_raise_valueerror():
with pytest.raises(ValueError) as exep:
Orbit.from_horizons(name="Dummy")
assert ("ValueError: Unknown target (Dummy). Maybe try different id_type?"
in exep.exconly())
def test_orbit_from_ephem_is_in_icrs_frame(body):
ss = Orbit.from_body_ephem(body)
assert ss.frame.is_equivalent_frame(ICRS())
def test_frozen_orbit_altitude():
with pytest.raises(ValueError) as excinfo:
Orbit.frozen(Earth, -1 * u.m)
assert excinfo.type == ValueError
assert (str(excinfo.value) ==
"The semimajor axis may not be smaller that Earth's radius")
def test_frozen_orbit_non_spherical_arguments():
with pytest.raises(AttributeError) as excinfo:
Orbit.frozen(Jupiter, 1 * u.m)
assert excinfo.type == AttributeError
assert (str(excinfo.value) ==
"Attractor Jupiter has not spherical harmonics implemented")
def test_frozen_orbit_venus_special_case():
with pytest.raises(NotImplementedError) as excinfo:
Orbit.frozen(Venus, 1 * u.m)
assert excinfo.type == NotImplementedError
assert str(excinfo.value) == "This has not been implemented for Venus"
def test_frozen_orbit_non_critical_inclination():
orbit = Orbit.frozen(Earth, 1e3 * u.km,
inc=0 * u.deg) # Non-critical value
assert orbit.argp in [np.pi / 2, 3 * np.pi / 2] * u.rad
