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

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_orbit_from_horizons_has_expected_elements():
    epoch = Time("2018-07-23", scale="tdb")
    # Orbit Parameters of Ceres
    # Taken from https://ssd.jpl.nasa.gov/horizons.cgi
    ss = Orbit.from_classical(
        Sun,
        2.76710759221651 * u.au,
        0.07554803091400027 * u.one,
        27.18502494739172 * u.deg,
        23.36913218336299 * u.deg,
        132.2919809219236 * u.deg,
        21.28957916690369 * u.deg,
        epoch,
    )
    ss1 = Orbit.from_horizons(name="Ceres", attractor=Sun, epoch=epoch)
    assert ss.pqw()[0].value.all() == ss1.pqw()[0].value.all()
    assert_quantity_allclose(ss.r_a, ss1.r_a, rtol=1.0e-4)
    assert_quantity_allclose(ss.a, ss1.a, rtol=1.0e-4)

def test_convert_from_rv_to_coe():
    # Data from Vallado, example 2.6
    attractor = Earth
    p = 11_067.790 * 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 = [6_525.344, 6_861.535, 6_449.125] * 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_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, a, ecc, _a, _a, _a, nu)
    orbit_t_p = orbit.t_p

    assert_quantity_allclose(orbit_t_p, expected_t_p, rtol=1e-2)

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_apply_maneuver_changes_epoch():
    _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)
    dt = 1 * u.h
    dv = [0, 0, 0] * u.km / u.s
    orbit_new = ss.apply_maneuver([(dt, dv)])
    assert orbit_new.epoch == ss.epoch + dt