示例#1
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 def _propagate_ecef(self, when_utc=None):
     """Return position and velocity in the given date using ECEF coordinate system."""
     position_eci, velocity_eci = self._propagate_eci(when_utc)
     gmst = gstime_from_datetime(when_utc)
     position_ecef = coordinate_systems.eci_to_ecef(position_eci, gmst)
     velocity_ecef = coordinate_systems.eci_to_ecef(velocity_eci, gmst)
     return position_ecef, velocity_ecef
示例#2
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    def from_tle(cls, sate_id, source, date=None):
        """Returns approximate keplerian elements from TLE.

        The conversion between mean elements in the TEME reference
        frame to osculating elements in any standard reference frame
        is not well defined in literature (see Vallado 3rd edition, pp 236 to 240)

        """
        # Get latest TLE, or the one corresponding to a specified date
        if date is None:
            date = datetime.datetime.utcnow()

        tle = source.get_tle(sate_id, date)

        # Retrieve TLE epoch and corresponding position
        epoch = twoline2rv(tle.lines[0], tle.lines[1], wgs84).epoch
        pos = TLEPredictor(sate_id, source).get_position(epoch)

        # Convert position from ECEF to ECI
        gmst = gstime_from_datetime(epoch)
        position_eci = coordinate_systems.ecef_to_eci(pos.position_ecef, gmst)
        velocity_eci = coordinate_systems.ecef_to_eci(pos.velocity_ecef, gmst)

        # Convert position to Keplerian osculating elements
        p, ecc, inc, raan, argp, ta = rv2coe(wgs84.mu, np.array(position_eci),
                                             np.array(velocity_eci))
        sma = p / (1 - ecc**2)

        return cls(sma, ecc, degrees(inc), degrees(raan), degrees(argp),
                   degrees(ta), epoch)
示例#3
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    def osculating_elements(self):
        """Osculating Keplerian orbital elements.

        Semimajor axis (km), eccentricity, inclination (deg),
        right ascension of the ascending node or RAAN (deg),
        argument of perigee (deg), true anomaly (deg).

        """
        gmst = gstime_from_datetime(self.when_utc)
        position_eci = coordinate_systems.ecef_to_eci(self.position_ecef, gmst)
        velocity_eci = coordinate_systems.ecef_to_eci(self.velocity_ecef, gmst)

        # Convert position to Keplerian osculating elements
        p, ecc, inc, raan, argp, ta = rv2coe(MU_E, np.array(position_eci),
                                             np.array(velocity_eci))
        # Transform to more familiar semimajor axis
        sma = p / (1 - ecc**2)

        return sma, ecc, degrees(inc), degrees(raan), degrees(argp), degrees(
            ta)
示例#4
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    def osculating_elements(self):
        """Osculating Keplerian orbital elements.

        Semimajor axis (km), eccentricity, inclination (deg),
        right ascension of the ascending node or RAAN (deg),
        argument of perigee (deg), true anomaly (deg).

        """
        gmst = gstime_from_datetime(self.when_utc)
        position_eci = coordinate_systems.ecef_to_eci(self.position_ecef, gmst)
        velocity_eci = coordinate_systems.ecef_to_eci(self.velocity_ecef, gmst)

        # Convert position to Keplerian osculating elements
        p, ecc, inc, raan, argp, ta = rv2coe(MU_E, np.array(position_eci),
                                             np.array(velocity_eci))
        # Transform to more familiar semimajor axis
        sma = p / (1 - ecc**2)

        # NOTE: rv2coe already does % (2 * np.pi)
        # but under some circumstances this might require another pass,
        # see https://github.com/satellogic/orbit-predictor/pull/106#issuecomment-730177598
        return sma, ecc, degrees(inc), degrees(raan), degrees(
            argp), degrees(ta) % 360