def from_sbdb(cls, name, **kargs):
"""Return osculating `Orbit` by using `SBDB` from Astroquery.
Parameters
----------
body_name: string
Name of the body to make the request.
Returns
-------
ss: poliastro.twobody.orbit.Orbit
Orbit corresponding to body_name
Examples
--------
>>> from poliastro.twobody.orbit import Orbit
>>> apophis_orbit = Orbit.from_sbdb('apophis')
"""
obj = SBDB.query(name, full_precision=True, **kargs)
if "count" in obj:
# no error till now ---> more than one object has been found
# contains all the name of the objects
objects_name = obj["list"]["name"]
objects_name_in_str = (
""
) # used to store them in string form each in new line
for i in objects_name:
objects_name_in_str += i + "\n"
raise ValueError(
str(obj["count"]) + " different objects found: \n" + objects_name_in_str
)
a = obj["orbit"]["elements"]["a"].to(u.AU) * u.AU
ecc = float(obj["orbit"]["elements"]["e"]) * u.one
inc = obj["orbit"]["elements"]["i"].to(u.deg) * u.deg
raan = obj["orbit"]["elements"]["om"].to(u.deg) * u.deg
argp = obj["orbit"]["elements"]["w"].to(u.deg) * u.deg
# Since JPL provides Mean Anomaly (M) we need to make
# the conversion to the true anomaly (\nu)
nu = M_to_nu(obj["orbit"]["elements"]["ma"].to(u.deg) * u.deg, ecc)
epoch = time.Time(obj["orbit"]["epoch"].to(u.d), format="jd")
ss = cls.from_classical(
Sun,
a,
ecc,
inc,
raan,
argp,
nu,
epoch=epoch.tdb,
plane=Planes.EARTH_ECLIPTIC,
)
return ss
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')
orbit = Orbit.from_classical(Sun, a, ecc, inc, raan, argp, nu, epoch)
return orbit
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 mean_motion(orbit, tof, **kwargs):
r"""Propagates orbit using mean motion
.. versionadded:: 0.9.0
Parameters
----------
orbit : ~poliastro.twobody.orbit.Orbit
the Orbit object to propagate.
tof : float
Time of flight (s).
Notes
-----
This method takes initial :math:`\vec{r}, \vec{v}`, calculates classical orbit parameters,
increases mean anomaly and performs inverse transformation to get final :math:`\vec{r}, \vec{v}`
The logic is based on formulae (4), (6) and (7) from http://dx.doi.org/10.1007/s10569-013-9476-9
"""
k = orbit.attractor.k.to(u.km**3 / u.s**2).value
r0 = orbit.r.to(u.km).value
v0 = orbit.v.to(u.km / u.s).value
# get the initial true anomaly and orbit parameters that are constant over time
p, ecc, inc, raan, argp, nu0 = rv2coe(k, r0, v0)
# get the initial mean anomaly
M0 = nu_to_M(nu0, ecc)
# strong elliptic or strong hyperbolic orbits
if np.abs(ecc - 1.0) > 1e-2:
a = p / (1.0 - ecc**2)
# given the initial mean anomaly, calculate mean anomaly
# at the end, mean motion (n) equals sqrt(mu / |a^3|)
with u.set_enabled_equivalencies(u.dimensionless_angles()):
M = M0 + tof * np.sqrt(k / np.abs(a**3)) * u.rad
nu = M_to_nu(M, ecc)
# near-parabolic orbit
else:
q = p * np.abs(1.0 - ecc) / np.abs(1.0 - ecc**2)
# mean motion n = sqrt(mu / 2 q^3) for parabolic orbit
with u.set_enabled_equivalencies(u.dimensionless_angles()):
M = M0 + tof * np.sqrt(k / 2.0 / (q**3))
nu = M_to_nu(M, ecc)
with u.set_enabled_equivalencies(u.dimensionless_angles()):
return coe2rv(k, p, ecc, inc, raan, argp, nu)
def mean_motion(k, r0, v0, tof, **kwargs):
r"""Propagates orbit using mean motion
Parameters
----------
k : float
Gravitational constant of main attractor (km^3 / s^2).
r0 : array
Initial position (km).
v0 : array
Initial velocity (km).
ad : function(t0, u, k), optional
Non Keplerian acceleration (km/s2), default to None.
tof : float
Time of flight (s).
Notes
-----
This method takes initial :math:`\vec{r}, \vec{v}`, calculates classical orbit parameters,
increases mean anomaly and performs inverse transformation to get final :math:`\vec{r}, \vec{v}`
The logic is based on formulae (4), (6) and (7) from http://dx.doi.org/10.1007/s10569-013-9476-9
"""
# get the initial true anomaly and orbit parameters that are constant over time
p, ecc, inc, raan, argp, nu0 = rv2coe(k, r0, v0)
# get the initial mean anomaly
M0 = nu_to_M(nu0, ecc)
# elliptic or hyperbolic orbits
if not np.isclose(ecc, 1.0, rtol=1e-06):
a = p / (1.0 - ecc**2)
# given the initial mean anomaly, calculate mean anomaly
# at the end, mean motion (n) equals sqrt(mu / |a^3|)
with u.set_enabled_equivalencies(u.dimensionless_angles()):
M = M0 + tof * np.sqrt(k / np.abs(a**3)) * u.rad
nu = M_to_nu(M, ecc)
# parabolic orbit
else:
q = p / 2.0
# mean motion n = sqrt(mu / 2 q^3) for parabolic orbit
with u.set_enabled_equivalencies(u.dimensionless_angles()):
M = M0 + tof * np.sqrt(k / (2.0 * q**3))
# using Barker's equation, which is solved analytically
# for parabolic orbit, get true anomaly
B = 3.0 * M / 2.0
A = (B + np.sqrt(1.0 + B**2))**(2.0 / 3.0)
D = 2.0 * A * B / (1.0 + A + A**2)
nu = 2.0 * np.arctan(D)
with u.set_enabled_equivalencies(u.dimensionless_angles()):
return coe2rv(k, p, ecc, inc, raan, argp, nu)
def orbit_from_spk_id(spk_id, api_key=None):
"""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 or DEFAULT_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")
ss = Orbit.from_classical(attractor,
a,
ecc,
inc,
raan,
argp,
nu,
epoch,
plane=Planes.EARTH_ECLIPTIC)
return ss
def from_sbdb(cls, name, **kargs):
"""Return osculating `Orbit` by using `SBDB` from Astroquery.
Parameters
----------
body_name: string
Name of the body to make the request.
Returns
-------
ss: poliastro.twobody.orbit.Orbit
Orbit corresponding to body_name
Examples
--------
>>> from poliastro.twobody.orbit import Orbit
>>> apophis_orbit = Orbit.from_sbdb('apophis')
"""
obj = SBDB.query(name, full_precision=True, **kargs)
a = obj["orbit"]["elements"]["a"].to(u.AU) * u.AU
ecc = float(obj["orbit"]["elements"]["e"]) * u.one
inc = obj["orbit"]["elements"]["i"].to(u.deg) * u.deg
raan = obj["orbit"]["elements"]["om"].to(u.deg) * u.deg
argp = obj["orbit"]["elements"]["w"].to(u.deg) * u.deg
# Since JPL provides Mean Anomaly (M) we need to make
# the conversion to the true anomaly (\nu)
nu = M_to_nu(obj["orbit"]["elements"]["ma"].to(u.deg) * u.deg, ecc)
epoch = time.Time(obj["orbit"]["epoch"].to(u.d), format="jd")
ss = cls.from_classical(
Sun,
a,
ecc,
inc,
raan,
argp,
nu,
epoch=epoch.tdb,
plane=Planes.EARTH_ECLIPTIC,
)
return ss
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_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 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,
M_to_nu(77.37209588584763 * u.deg, 0.07600902910070946 * 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]
import numpy as np
from astropy import units as u
from poliastro.twobody import Orbit
from poliastro.twobody.angles import M_to_nu
from poliastro.bodies import Sun
MU69_ecc = 0.05046617 * u.one # Eccentricity
MU69_i = 2.44996247 * u.deg # Inclination
MU69_a = 6676495970 * u.km # Semimajor axisº
MU69_raan = 159.04395317345 * u.deg # Right ascension of the ascending node
MU69_argp = 181.14813924539 * u.deg # Argument of perihelion
MU69_M = 310.91947579009 * u.deg # Mean anomaly
MU69_nu = M_to_nu(MU69_M, MU69_ecc) # True anomaly
orbit = Orbit.from_classical(Sun, MU69_a, MU69_ecc, MU69_i, MU69_raan,
MU69_argp, MU69_nu)