# Orbit¶

This is probably one of the main important modules of poliastro since it enables the user to create the poliastro.twobody.orbit.Orbit objects.

User can create or initialize an orbit in a very different ways but the most common ones are the following:

## ¶

Howeverm it is also possible to create ‘special’ orbits such us poliastro.twobody.orbit.Orbit.circular or even poliastro.twobody.orbit.Orbit.parabolic.

Furthermore, orbits can be sampled by poliastro.twobody.orbit.Orbit.sample or propagated in time by poliastro.twobody.orbit.Orbit.propagate. But even it is possible to apply a maneuver to an orbit with the poliastro.twobody.orbit.Orbit.apply_maneuver.

exception poliastro.twobody.orbit.TimeScaleWarning
exception poliastro.twobody.orbit.OrbitSamplingWarning
exception poliastro.twobody.orbit.PatchedConicsWarning
class poliastro.twobody.orbit.Orbit(state, epoch)

Position and velocity of a body with respect to an attractor at a given time (epoch).

Regardless of how the Orbit is created, the implicit reference system is an inertial one. For the specific case of the Solar System, this can be assumed to be the International Celestial Reference System or ICRS.

property attractor

Main attractor.

property epoch

Epoch of the orbit.

property plane

Fundamental plane of the frame.

r

Position vector.

v

Velocity vector.

a

Semimajor axis.

p

Semilatus rectum.

r_p

r_a

ecc

Eccentricity.

inc

Inclination.

raan

Right ascension of the ascending node.

argp

Argument of the perigee.

property nu

True anomaly.

f

Second modified equinoctial element.

g

Third modified equinoctial element.

h

Fourth modified equinoctial element.

k

Fifth modified equinoctial element.

L

True longitude.

period

Period of the orbit.

n

Mean motion.

energy

Specific energy.

e_vec

Eccentricity vector.

h_vec

Specific angular momentum vector.

h_mag

Specific angular momentum.

arglat

Argument of latitude.

M

Mean anomaly.

t_p

Elapsed time since latest perifocal passage.

classmethod from_vectors(attractor, r, v, epoch=<Time object: scale='tt' format='jyear_str' value=J2000.000>, plane=<Planes.EARTH_EQUATOR: 'Earth mean Equator and Equinox of epoch (J2000.0)'>)

Return Orbit from position and velocity vectors.

Parameters: attractor (Body) – Main attractor. r (Quantity) – Position vector wrt attractor center. v (Quantity) – Velocity vector. epoch (Time, optional) – Epoch, default to J2000. plane (Planes) – Fundamental plane of the frame.
classmethod from_coords(attractor, coord, plane=<Planes.EARTH_EQUATOR: 'Earth mean Equator and Equinox of epoch (J2000.0)'>)

Creates an Orbit from an attractor and astropy SkyCoord or BaseCoordinateFrame instance.

This method accepts position and velocity in any reference frame unlike Orbit.from_vector which can accept inputs in only inertial reference frame centred at attractor. Also note that the frame information is lost after creation of the orbit and only the inertial reference frame at body centre will be used for all purposes.

Parameters: attractor (Body) – Main attractor coord (astropy.coordinates.SkyCoord or BaseCoordinateFrame) – Position and velocity vectors in any reference frame. Note that coord must have a representation and its differential with respect to time. plane (Planes, optional) – Final orbit plane, default to Earth Equator.
classmethod from_classical(attractor, a, ecc, inc, raan, argp, nu, epoch=<Time object: scale='tt' format='jyear_str' value=J2000.000>, plane=<Planes.EARTH_EQUATOR: 'Earth mean Equator and Equinox of epoch (J2000.0)'>)

Return Orbit from classical orbital elements.

Parameters: attractor (Body) – Main attractor. a (Quantity) – Semi-major axis. ecc (Quantity) – Eccentricity. inc (Quantity) – Inclination raan (Quantity) – Right ascension of the ascending node. argp (Quantity) – Argument of the pericenter. nu (Quantity) – True anomaly. epoch (Time, optional) – Epoch, default to J2000. plane (Planes) – Fundamental plane of the frame.
classmethod from_equinoctial(attractor, p, f, g, h, k, L, epoch=<Time object: scale='tt' format='jyear_str' value=J2000.000>, plane=<Planes.EARTH_EQUATOR: 'Earth mean Equator and Equinox of epoch (J2000.0)'>)

Return Orbit from modified equinoctial elements.

Parameters: attractor (Body) – Main attractor. p (Quantity) – Semilatus rectum. f (Quantity) – Second modified equinoctial element. g (Quantity) – Third modified equinoctial element. h (Quantity) – Fourth modified equinoctial element. k (Quantity) – Fifth modified equinoctial element. L (Quantity) – True longitude. epoch (Time, optional) – Epoch, default to J2000. plane (Planes) – Fundamental plane of the frame.
classmethod from_body_ephem(body, epoch=None)

Return osculating Orbit of a body at a given time.

get_frame()

Get equivalent reference frame of the orbit.

New in version 0.14.0.

change_attractor(new_attractor, force=False)

Changes orbit attractor.

Only changes from attractor to parent or the other way around are allowed.

Parameters: new_attractor (poliastro.bodies.Body) – Desired new attractor. force (boolean) – If True, changes attractor even if physically has no-sense. ss – Orbit with new attractor poliastro.twobody.orbit.Orbit
classmethod from_horizons(name, attractor, epoch=None, plane=<Planes.EARTH_EQUATOR: 'Earth mean Equator and Equinox of epoch (J2000.0)'>, id_type='smallbody')

Return osculating Orbit of a body using JPLHorizons module of Astroquery.

Parameters: name (string) – Name of the body to query for. epoch (Time, optional) – Epoch, default to None. plane (Planes) – Fundamental plane of the frame. id_type (string, optional) – Use “smallbody” for Asteroids and Comets, and “majorbody” for Planets and Satellites.
classmethod from_sbdb(name, **kargs)

Return osculating Orbit by using SBDB from Astroquery.

Parameters: body_name (string) – Name of the body to make the request. ss – Orbit corresponding to body_name poliastro.twobody.orbit.Orbit

Examples

>>> from poliastro.twobody.orbit import Orbit
>>> apophis_orbit = Orbit.from_sbdb('apophis')

classmethod circular(attractor, alt, inc=<Quantity 0. deg>, raan=<Quantity 0. deg>, arglat=<Quantity 0. deg>, epoch=<Time object: scale='tt' format='jyear_str' value=J2000.000>, plane=<Planes.EARTH_EQUATOR: 'Earth mean Equator and Equinox of epoch (J2000.0)'>)

Return circular Orbit.

Parameters: attractor (Body) – Main attractor. alt (Quantity) – Altitude over surface. inc (Quantity, optional) – Inclination, default to 0 deg (equatorial orbit). raan (Quantity, optional) – Right ascension of the ascending node, default to 0 deg. arglat (Quantity, optional) – Argument of latitude, default to 0 deg. epoch (Time, optional) – Epoch, default to J2000. plane (Planes) – Fundamental plane of the frame.
classmethod geostationary(attractor, angular_velocity=None, period=None, hill_radius=None)

Return the geostationary orbit for the given attractor and its rotational speed.

Parameters: attractor (Body) – Main attractor. angular_velocity (Quantity) – Rotational angular velocity of the attractor. period (Quantity) – Attractor’s rotational period, ignored if angular_velocity is passed. hill_radius (Quantity) – Radius of Hill sphere of the attractor (optional). Hill sphere radius(in contrast with Laplace’s SOI) is used here to validate the stability of the geostationary orbit, that is to make sure that the orbital radius required for the geostationary orbit is not outside of the gravitational sphere of influence of the attractor. Hill SOI of parent(if exists) of the attractor is ignored if hill_radius is not provided.
classmethod heliosynchronous(a=None, ecc=None, inc=None, ltan=<Quantity 10. hourangle>, argp=<Quantity 0. deg>, nu=<Quantity 0. deg>, epoch=<Time object: scale='tt' format='jyear_str' value=J2000.000>, plane=<Planes.EARTH_EQUATOR: 'Earth mean Equator and Equinox of epoch (J2000.0)'>)

Solves for a Sun-Synchronous orbit. These orbits make use of the J2 perturbation to precess in order to be always towards Sun. At least two parameters of the set {a, ecc, inc} are needed in order to solve for these kind of orbits. Relationships among them are given by:

\begin{split}\begin{align} a &= \left (\frac{-3R_{\bigoplus}J_{2}\sqrt{\mu}\cos(i)}{2\dot{\Omega}(1-e^2)^2} \right ) ^ {\frac{2}{7}}\\ e &= \sqrt{1 - \sqrt{\frac{-3R_{\bigoplus}J_{2}\sqrt{\mu}cos(i)}{2a^{\frac{7}{2}}\dot{\Omega}}}}\\ i &= \arccos{\left ( \frac{-2a^{\frac{7}{2}}\dot{\Omega}(1-e^2)^2}{3R_{\bigoplus}J_{2}\sqrt{\mu}} \right )}\\ \end{align}\end{split}
Parameters: a (Quantity) – Semi-major axis. ecc (Quantity) – Eccentricity. inc (Quantity) – Inclination. ltan (Quantity) – Local time of the ascending node which will be translated to the Right ascension of the ascending node. argp (Quantity) – Argument of the pericenter. nu (Quantity) – True anomaly. epoch (Time, optional) – Epoch, default to J2000. plane (Planes) – Fundamental plane of the frame.
classmethod parabolic(attractor, p, inc, raan, argp, nu, epoch=<Time object: scale='tt' format='jyear_str' value=J2000.000>, plane=<Planes.EARTH_EQUATOR: 'Earth mean Equator and Equinox of epoch (J2000.0)'>)

Return parabolic Orbit.

Parameters: attractor (Body) – Main attractor. p (Quantity) – Semilatus rectum or parameter. inc (Quantity, optional) – Inclination. raan (Quantity) – Right ascension of the ascending node. argp (Quantity) – Argument of the pericenter. nu (Quantity) – True anomaly. epoch (Time, optional) – Epoch, default to J2000. plane (Planes) – Fundamental plane of the frame.
classmethod frozen(attractor, alt, inc=None, argp=None, raan=<Quantity 0. deg>, arglat=<Quantity 0. deg>, ecc=None, epoch=<Time object: scale='tt' format='jyear_str' value=J2000.000>, plane=<Planes.EARTH_EQUATOR: 'Earth mean Equator and Equinox of epoch (J2000.0)'>)

Return frozen Orbit. If any of the given arguments results in an impossibility, some values will be overwritten

To achieve frozen orbit these two equations have to be set to zero.

$\dfrac {d\overline {e}}{dt}=\dfrac {-3\overline {n}J_{3}R^{3}_{E}\sin \left( \overline {i}\right) }{2a^{3}\left( 1-\overline {e}^{2}\right) ^{2}}\left( 1-\dfrac {5}{4}\sin ^{2}\overline {i}\right) \cos \overline {w}$
$\dfrac {d\overline {\omega }}{dt}=\dfrac {3\overline {n}J_{2}R^{2}_{E}}{a^{2}\left( 1-\overline {e}^{2}\right) ^{2}}\left( 1-\dfrac {5}{4}\sin ^{2}\overline {i}\right) \left[ 1+\dfrac {J_{3}R_{E}}{2J_{2}\overline {a}\left( 1-\overline {e}^{2}\right) }\left( \dfrac {\sin ^{2}\overline {i}-\overline {e}\cos ^{2}\overline {i}}{\sin \overline {i}}\right) \dfrac {\sin \overline {w}}{\overline {e}}\right]$

The first approach would be to nullify following term to zero:

$( 1-\dfrac {5}{4}\sin ^{2})$

For which one obtains the so-called critical inclinations: i = 63.4349 or 116.5651 degrees. To escape the inclination requirement, the argument of periapsis can be set to w = 90 or 270 degrees to nullify the second equation. Then, one should nullify the right-hand side of the first equation, which yields an expression that correlates the inclination of the object and the eccentricity of the orbit:

$\overline {e}=-\dfrac {J_{3}R_{E}}{2J_{2}\overline {a}\left( 1-\overline {e}^{2}\right) }\left( \dfrac {\sin ^{2}\overline {i}-\overline {e}\cos ^{2} \overline {i}}{\sin \overline {i}}\right)$

Assuming that e is negligible compared to J2, it can be shown that:

$\overline {e}\approx -\dfrac {J_{3}R_{E}}{2J_{2}\overline {a}}\sin \overline {i}$

The implementation is divided in the following cases:

1. When the user gives a negative altitude, the method will raise a ValueError

2. When the attractor has not defined J2 or J3, the method will raise an AttributeError

3. When the attractor has J2/J3 outside of range 1 to 10 , the method will raise an NotImplementedError. Special case for Venus.See “Extension of the critical inclination” by Xiaodong Liu, Hexi Baoyin, and Xingrui Ma

4. If argp is not given or the given argp is a critical value:

• if eccentricity is none and inclination is none, the inclination is set with a critical value and the eccentricity is obtained from the last formula mentioned
• if only eccentricity is none, we calculate this value with the last formula mentioned
• if only inclination is none ,we calculate this value with the formula for eccentricity with critical argp.
5. If inc is not given or the given inc is critical:

• if the argp and the eccentricity is given we keep these values to create the orbit
• if the eccentricity is given we keep this value, if not, default to the eccentricity of the Moon’s orbit around the Earth
6. if it’s not possible to create an orbit with the the argp and the inclination given, both of them are set to the critical values and the eccentricity is calculate with the last formula

Parameters: attractor (Body) – Main attractor. alt (Quantity) – Altitude over surface. inc (Quantity, optional) – Inclination, default to critical value. argp (Quantity, optional) – Argument of the pericenter, default to critical value. raan (Quantity, optional) – Right ascension of the ascending node, default to 0 deg. arglat (Quantity, optional) – Argument of latitude, default to 0 deg. ecc (Quantity) – Eccentricity, default to the eccentricity of the Moon’s orbit around the Earth epoch (Time, optional) – Epoch, default to J2000. plane (Planes) – Fundamental plane of the frame.
represent_as(representation, differential_class=None)

Converts the orbit to a specific representation.

New in version 0.11.0.

Parameters: representation (BaseRepresentation) – Representation object to use. It must be a class, not an instance. differential_class (BaseDifferential, optional) – Class in which the differential should be represented, default to None.

Examples

>>> from poliastro.examples import iss
>>> from astropy.coordinates import SphericalRepresentation
>>> iss.represent_as(CartesianRepresentation)
<CartesianRepresentation (x, y, z) in km
(859.07256, -4137.20368, 5295.56871)>
>>> iss.represent_as(CartesianRepresentation).xyz
<Quantity [  859.07256, -4137.20368,  5295.56871] km>
>>> iss.represent_as(CartesianRepresentation, CartesianDifferential).differentials['s']
<CartesianDifferential (d_x, d_y, d_z) in km / s
(7.37289205, 2.08223573, 0.43999979)>
>>> iss.represent_as(CartesianRepresentation, CartesianDifferential).differentials['s'].d_xyz
<Quantity [7.37289205, 2.08223573, 0.43999979] km / s>
>>> iss.represent_as(SphericalRepresentation, CartesianDifferential)
(4.91712525, 0.89732339, 6774.76995296)
(has differentials w.r.t.: 's')>

rv()

Position and velocity vectors.

classical()

Classical orbital elements.

pqw()

Perifocal frame (PQW) vectors.

propagate(value, method=<function mean_motion>, rtol=1e-10, **kwargs)

Propagates an orbit a specified time.

If value is true anomaly, propagate orbit to this anomaly and return the result. Otherwise, if time is provided, propagate this Orbit some time and return the result.

Parameters: value (Quantity, Time, TimeDelta) – Scalar time to propagate. rtol (float, optional) – Relative tolerance for the propagation algorithm, default to 1e-10. method (function, optional) – Method used for propagation **kwargs – parameters used in perturbation models New orbit after propagation. Orbit
time_to_anomaly(value)

Returns time required to be in a specific true anomaly.

Parameters: value (Quantity) – tof – Time of flight required. Quantity
propagate_to_anomaly(value)

Propagates an orbit to a specific true anomaly.

Parameters: value (Quantity) – Resulting orbit after propagation. Orbit
sample(values=100, *, min_anomaly=None, max_anomaly=None)

Samples an orbit to some specified time values.

New in version 0.8.0.

Parameters: values (int) – Number of interval points (default to 100). max_anomaly (min_anomaly,) – Anomaly limits to sample the orbit. For elliptic orbits the default will be $$E \in \left[0, 2 \pi \right]$$, and for hyperbolic orbits it will be $$\nu \in \left[-\nu_c, \nu_c \right]$$, where $$\nu_c$$ is either the current true anomaly or a value that corresponds to $$r = 3p$$. positions – Array of x, y, z positions, with proper times as the frame attributes if supported. BaseCoordinateFrame

Notes

When specifying a number of points, the initial and final position is present twice inside the result (first and last row). This is more useful for plotting.

Examples

>>> from astropy import units as u
>>> from poliastro.examples import iss
>>> iss.sample()
<CartesianRepresentation (x, y, z) in km ...
>>> iss.sample(10)
<CartesianRepresentation (x, y, z) in km ...

apply_maneuver(maneuver, intermediate=False)

Returns resulting Orbit after applying maneuver to self.

Optionally return intermediate states (default to False).

Parameters: maneuver (Maneuver) – Maneuver to apply. intermediate (bool, optional) – Return intermediate states, default to False.
plot(label=None, use_3d=False, interactive=False)

Plots the orbit as an interactive widget.

Parameters: label (str, optional) – Label for the orbit, defaults to empty. use_3d (bool, optional) – Produce a 3D plot, default to False. interactive (bool, optional) – Produce an interactive (rather than static) image of the orbit, default to False. This option requires Plotly properly installed and configured for your environment.