How to use the poliastro.twobody.orbit.Orbit.from_vectors function in poliastro

def test_earth_satellite_orbit():
    r = [3_539.08827417, 5_310.19903462, 3_066.31301457] * u.km
    v = [-6.49780849, 3.24910291, 1.87521413] * u.km / u.s
    ss = Orbit.from_vectors(Earth, r, v)
    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(ss, spacecraft)
    assert isinstance(earth_satellite.orbit, Orbit)

def near_parabolic():
    r = [8.0e3, 1.0e3, 0.0] * u.km
    v = [-0.5, -0.5, 0.0] * u.km / u.s

    return Orbit.from_vectors(Earth, r, v)

def test_orbit_change_attractor_open():
    r = [-6_045, -3_490, 2_500] * u.km
    v = [-15.457, 6.618, 2.533] * u.km / u.s
    ss = Orbit.from_vectors(Earth, r, v)

    with pytest.warns(PatchedConicsWarning) as record:
        ss.change_attractor(Sun)

    w = record.pop(PatchedConicsWarning)
    assert "Leaving the SOI of the current attractor" in w.message.args[0]

def test_orbit_no_frame_representation():
    date_launch = Time("2011-11-26 15:02", scale="utc")
    r = [61_445.76498656, 24_827.93010168, 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 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_orbit_from_custom_body_raises_error_when_asked_frame():
    attractor = Body(Sun, 1 * u.km ** 3 / u.s ** 2, "_DummyPlanet")

    r = [1e09, -4e09, -1e09] * u.km
    v = [5e00, -1e01, -4e00] * u.km / u.s

    ss = Orbit.from_vectors(attractor, r, v)

    with pytest.raises(NotImplementedError) as excinfo:
        ss.get_frame()
    assert (
        "Frames for orbits around custom bodies are not yet supported"
        in excinfo.exconly()
    )

def test_sample_big_orbits():
    # See https://github.com/poliastro/poliastro/issues/265
    ss = Orbit.from_vectors(
        Sun,
        [-9_018_878.6, -94_116_055, 22_619_059] * u.km,
        [-49.950923, -12.948431, -4.2925158] * u.km / u.s,
    )
    positions = ss.sample(15)
    assert len(positions) == 15

def _targetting(departure_body, target_body, t_launch, t_arrival):
    """This function returns the increment in departure and arrival velocities."""

    # Get position and velocities for departure and arrival
    rr_dpt_body, vv_dpt_body = _get_state(departure_body, t_launch)
    rr_arr_body, vv_arr_body = _get_state(target_body, t_arrival)

    # Transform into Orbit objects
    attractor = departure_body.parent
    ss_dpt = Orbit.from_vectors(attractor, rr_dpt_body, vv_dpt_body, epoch=t_launch)
    ss_arr = Orbit.from_vectors(attractor, rr_arr_body, vv_arr_body, epoch=t_arrival)

    # Define time of flight
    tof = ss_arr.epoch - ss_dpt.epoch

    if tof <= 0:
        return None, None, None, None, None

    try:
        # Lambert is now a Maneuver object
        man_lambert = Maneuver.lambert(ss_dpt, ss_arr)

        # Get norm delta velocities
        dv_dpt = norm(man_lambert.impulses[0][1])
        dv_arr = norm(man_lambert.impulses[1][1])

        # Compute all the output variables

def _targetting(departure_body, target_body, t_launch, t_arrival):
    """This function returns the increment in departure and arrival velocities."""

    # Get position and velocities for departure and arrival
    rr_dpt_body, vv_dpt_body = _get_state(departure_body, t_launch)
    rr_arr_body, vv_arr_body = _get_state(target_body, t_arrival)

    # Transform into Orbit objects
    attractor = departure_body.parent
    ss_dpt = Orbit.from_vectors(attractor, rr_dpt_body, vv_dpt_body, epoch=t_launch)
    ss_arr = Orbit.from_vectors(attractor, rr_arr_body, vv_arr_body, epoch=t_arrival)

    # Define time of flight
    tof = ss_arr.epoch - ss_dpt.epoch

    if tof <= 0:
        return None, None, None, None, None

    try:
        # Lambert is now a Maneuver object
        man_lambert = Maneuver.lambert(ss_dpt, ss_arr)

        # Get norm delta velocities
        dv_dpt = norm(man_lambert.impulses[0][1])
        dv_arr = norm(man_lambert.impulses[1][1])