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"""SGP4 propagator. This is a wrapper around the PyPI SGP4 propagator.
However, this does not generate an artificial TLE. So there is no
string manipulation involved. Hence this is faster than sgp4_prop_string."""
from datetime import datetime
import numpy as np
from sgp4.model import Satellite
from sgp4.earth_gravity import wgs72
from sgp4.propagation import sgp4init
from orbitdeterminator.util.state_kep import state_kep
def __true_to_mean(T,e):
"""Converts true anomaly to mean anomaly.
Args:
T(float): true anomaly in degrees
e(float): eccentricity
Returns:
float: the mean anomaly in degrees
"""
T = np.radians(T)
E = np.arctan2((1-e**2)*np.sin(T),e+np.cos(T))
M = E - e*np.sin(E)
M = np.degrees(M)
M = M%360
return M
# Parts of this method have been copied from:
# https://github.com/brandon-rhodes/python-sgp4/blob/master/sgp4/io.py
def kep_to_sat(kep,epoch,bstar=0.21109E-4,whichconst=wgs72,afspc_mode=False):
"""kep_to_sat(kep,epoch,bstar=0.21109E-4,whichconst=wgs72,afspc_mode=False)
Converts a set of keplerian elements into a Satellite object.
Args:
kep(1x6 numpy array): the osculating keplerian elements at epoch
epoch(float): the epoch
bstar(float): bstar drag coefficient
whichconst(float): gravity model. refer pypi sgp4 documentation
afspc_mode(boolean): refer pypi sgp4 documentation
Returns:
Satellite object: an sgp4 satellite object encapsulating the arguments
"""
deg2rad = np.pi / 180.0; # 0.0174532925199433
xpdotp = 1440.0 / (2.0 * np.pi); # 229.1831180523293
tumin = whichconst.tumin
satrec = Satellite()
satrec.error = 0;
satrec.whichconst = whichconst # Python extension: remembers its consts
satrec.satnum = 0
dt_obj = datetime.utcfromtimestamp(epoch)
t_obj = dt_obj.timetuple()
satrec.epochdays = (t_obj.tm_yday +
t_obj.tm_hour/24 +
t_obj.tm_min/1440 +
t_obj.tm_sec/86400)
satrec.ndot = 0
satrec.nddot = 0
satrec.bstar = bstar
satrec.inclo = kep[2]
satrec.nodeo = kep[4]
satrec.ecco = kep[1]
satrec.argpo = kep[3]
satrec.mo = __true_to_mean(kep[5],kep[1])
satrec.no = 86400/(2*np.pi*(kep[0]**3/398600.4405)**0.5)
satrec.no = satrec.no / xpdotp; # rad/min
satrec.a = pow( satrec.no*tumin , (-2.0/3.0) );
# ---- find standard orbital elements ----
satrec.inclo = satrec.inclo * deg2rad;
satrec.nodeo = satrec.nodeo * deg2rad;
satrec.argpo = satrec.argpo * deg2rad;
satrec.mo = satrec.mo * deg2rad;
satrec.alta = satrec.a*(1.0 + satrec.ecco) - 1.0;
satrec.altp = satrec.a*(1.0 - satrec.ecco) - 1.0;
satrec.epochyr = dt_obj.year
satrec.jdsatepoch = epoch/86400.0 + 2440587.5
satrec.epoch = dt_obj
# ---------------- initialize the orbit at sgp4epoch -------------------
sgp4init(whichconst, afspc_mode, satrec.satnum, satrec.jdsatepoch-2433281.5, satrec.bstar,
satrec.ecco, satrec.argpo, satrec.inclo, satrec.mo, satrec.no,
satrec.nodeo, satrec)
return satrec
def propagate_kep(kep,t0,tf,bstar=0.21109E-4):
"""Propagates a set of keplerian elements.
Args:
kep(1x6 numpy array): osculating keplerian elements at epoch
t0(float): initial time (epoch)
tf(float): final time
Returns:
pos(1x3 numpy array): the position at tf
vel(1x3 numpy array): the velocity at tf
"""
sat = kep_to_sat(kep,t0,bstar=bstar)
tf = datetime.utcfromtimestamp(tf).timetuple()
pos, vel = sat.propagate(
tf.tm_year, tf.tm_mon, tf.tm_mday, tf.tm_hour, tf.tm_min, tf.tm_sec)
return np.array(list(pos)),np.array(list(vel))
def propagate_state(r,v,t0,tf,bstar=0.21109E-4):
"""Propagates a state vector
Args:
r(1x3 numpy array): the position vector at epoch
v(1x3 numpy array): the velocity vector at epoch
t0(float): initial time (epoch)
tf(float): final time
Returns:
pos(1x3 numpy array): the position at tf
vel(1x3 numpy array): the velocity at tf
"""
kep = state_kep(r,v)
return propagate_kep(kep,t0,tf,bstar)
if __name__ == "__main__":
t0 = 1526927274
tf = 1526932833
#kep = np.array([6782.96, 0.0004084, 51.6402, 108.2140, 150.4026, 238.0528])
r = np.array([-5.23684633e+03, 4.12417773e+03, -1.26294137e+03])
v = np.array([-3.86204515e+00, -3.12048032e+00, 5.83839029e+00])
#pos,vel = propagate_kep(kep,t0,tf)
pos,vel = propagate_state(r,v,t0,tf)
print(pos,vel)
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