def test_visible_above_horizon(self):
"""
visible_above_horizon test
"""
#equitorial orbit
tle_line1 = "1 44235U 19029A 20178.66667824 .02170155 00000-0 40488-1 0 9998"
tle_line2 = "2 44235 00.0000 163.9509 0005249 306.3756 83.0170 15.45172567 61683"
#high inclination orbit
tle2_line1 = "1 44235U 19029A 20178.66667824 .02170155 00000-0 40488-1 0 9998"
tle2_line2 = "2 44235 70.0000 163.9509 0005249 306.3756 83.0170 15.45172567 61683"
prop1 = orekit_utils.str_tle_propagator(tle_line1, tle_line2)
prop2 = orekit_utils.str_tle_propagator(tle2_line1, tle2_line2)
time = AbsoluteDate(2020, 6, 26, 1, 40, 00.000,
TimeScalesFactory.getUTC())
#verified with graphing overviews
self.assertFalse(orekit_utils.visible_above_horizon(
prop1, prop2, time))
self.assertTrue(
orekit_utils.visible_above_horizon(prop1, prop2,
time.shiftedBy(60. * 10.)))
self.assertTrue(
orekit_utils.visible_above_horizon(
prop1, prop2, time.shiftedBy(60. * 10. + 45. * 60.)))
time_period_visible = orekit_utils.visible_above_horizon(
prop1, prop2, time, 60 * 30)[0]
self.assertTrue(
time.shiftedBy(60. * 10.).isBetween(time_period_visible[0],
time_period_visible[1]))
def orekit_test_data(body,
filename,
satellite,
stations,
range_sigma=20.0,
range_rate_sigma=0.001,
range_base_weight=1.0,
range_rate_base_weight=1.0,
az_sigma=0.02,
el_sigma=0.02):
"""Load test data from W3B.aer"""
azels = []
ranges = []
rates = []
f = open(filename, 'r')
for line in f:
if line[0] == '#':
continue
elif len(line) < 25:
continue
fields = line.split()
date_components = DateTimeComponents.parseDateTime(fields[0])
date = AbsoluteDate(date_components, TimeScalesFactory.getUTC())
if fields[1] == 'ONEWAY':
two_way = False
fields.pop(1)
elif fields[1] == 'TWOWAY':
two_way = True
fields.pop(1)
else:
two_way = True
mode = fields[1]
station_name = fields[2]
station = stations[station_name]
if mode == 'RANGE':
rho = float(fields[3])
range_obj = Range(station, True, date, rho, range_sigma,
range_base_weight, satellite)
ranges.append(range_obj)
elif mode == 'AZ_EL':
az = float(fields[3]) * math.pi / 180.0
el = float(fields[4]) * math.pi / 180.0
azel_obj = AngularAzEl(station, date, [az, el],
[az_sigma, el_sigma],
[az_base_weight, el_base_weight], satellite)
azels.append(azel_obj)
elif mode in ('RATE', 'RRATE'):
rate_obj = RangeRate(station, date, float(fields[3]),
range_rate_sigma, range_rate_base_weight,
two_way, satellite)
rates.append(rate_obj)
else:
raise SyntaxError("unrecognized mode '{}'".format(mode))
return stations, ranges, rates, azels
def itrs_gcrs_single(itrs):
""" Transformation ITRS to GCRS.
"""
utc = TimeScalesFactory.getUTC()
date = itrs.index[0]
X = float(itrs['x[m]'][0])
Y = float(itrs['y[m]'][0])
Z = float(itrs['z[m]'][0])
Vx = float(itrs['vx[m/s]'][0])
Vy = float(itrs['vy[m/s]'][0])
Vz = float(itrs['vz[m/s]'][0])
ok_date = AbsoluteDate(date.year, date.month, date.day, date.hour, date.minute,
date.second + float(date.microsecond) / 1000000., utc)
PV_coordinates = PVCoordinates(Vector3D(X, Y, Z), Vector3D(Vx, Vy, Vz))
start_state = TimeStampedPVCoordinates(ok_date, PV_coordinates)
state_itrf = AbsolutePVCoordinates(itrf, start_state)
state_gcrf = AbsolutePVCoordinates(gcrf, state_itrf.getPVCoordinates(gcrf))
pos = state_gcrf.position
vel = state_gcrf.velocity
X = pos.getX()
Y = pos.getY()
Z = pos.getZ()
Vx = vel.getX()
Vy = vel.getY()
Vz = vel.getZ()
dframe = pd.DataFrame({'x[m]': [X], 'y[m]': [Y], 'z[m]': [Z], 'vx[m/s]': [Vx], 'vy[m/s]': [Vy], 'vz[m/s]': [Vz]}, index=itrs.index)
return dframe
def absolute_time_converter_utc_string(time_string):
"""
turn time_string into orekit absolute time object
Inputs: time scales in UTC
Output: absolute time object from orekit
"""
return AbsoluteDate(time_string, TimeScalesFactory.getUTC())
def test_get_ground_passes(self):
"""
Testing whether OreKit finds ground passess
"""
# test will fail if setup_orekit fails
utc = TimeScalesFactory.getUTC()
ra = 500 * 1000 # Apogee
rp = 400 * 1000 # Perigee
i = radians(55.0) # inclination
omega = radians(20.0) # perigee argument
raan = radians(10.0) # right ascension of ascending node
lv = radians(0.0) # True anomaly
epochDate = AbsoluteDate(2020, 1, 1, 0, 0, 00.000, utc)
initial_date = epochDate
a = (rp + ra + 2 * Constants.WGS84_EARTH_EQUATORIAL_RADIUS) / 2.0
e = 1.0 - (rp + Constants.WGS84_EARTH_EQUATORIAL_RADIUS) / a
## Inertial frame where the satellite is defined
inertialFrame = FramesFactory.getEME2000()
## Orbit construction as Keplerian
initialOrbit = KeplerianOrbit(a, e, i, omega, raan, lv,
PositionAngle.TRUE,
inertialFrame, epochDate, Constants.WGS84_EARTH_MU)
propagator = KeplerianPropagator(initialOrbit)
durand_lat = 37.4269
durand_lat = -122.1733
durand_alt = 10.0
output = get_ground_passes(propagator, durand_lat, durand_lat, durand_alt, initial_date, initial_date.shiftedBy(3600.0 * 24), ploting_param=False)
assert len(output) == 5
def test_field_of_view_detector(self):
"""
field_of_view_detector tests
"""
tle = TLE(
"1 44235U 19029A 20178.66667824 .02170155 00000-0 40488-1 0 9998",
"2 44235 00.0000 163.9509 0005249 306.3756 83.0170 15.45172567 61683"
)
parameters = {"frame": "TEME"}
attitudeProvider = orekit_utils.nadir_pointing_law(parameters)
propogator_fov = TLEPropagator.selectExtrapolator(
tle, attitudeProvider, 4.)
startDate = AbsoluteDate(2020, 6, 26, 1, 40, 00.000,
TimeScalesFactory.getUTC())
#should have one passes
self.assertTrue(
len(
orekit_utils.field_of_view_detector(propogator_fov, 0, 0, 0,
startDate, 20, 5400)) == 1)
#should have zero passes
self.assertTrue(
len(
orekit_utils.field_of_view_detector(propogator_fov, 0, 0, 0,
startDate, 20, 100)) == 0)
#test exact time input
self.assertFalse(
orekit_utils.field_of_view_detector(propogator_fov, 0, 0, 0,
startDate, 20))
def absolute_time_converter_utc_manual(year,
month,
day,
hour=0,
minute=0,
second=0.0):
"""
turn time into orekit absolute time object
Inputs: time scales in UTC
Output: absolute time object from orekit
"""
return AbsoluteDate(int(year), int(month), int(day), int(hour),
int(minute), float(second), TimeScalesFactory.getUTC())
def test_check_iot_in_range(self):
"""
check_iot_in_range test
"""
tle_line1 = "1 44235U 19029A 20178.66667824 .02170155 00000-0 40488-1 0 9998"
tle_line2 = "2 44235 00.0000 163.9509 0005249 306.3756 270.0170 15.45172567 61683"
prop1 = orekit_utils.str_tle_propagator(tle_line1, tle_line2)
lat = 0.
lon = 0.
alt = 10.
time = AbsoluteDate(2020, 6, 26, 1, 40, 00.000, TimeScalesFactory.getUTC())
self.assertTrue(check_iot_in_range(prop1, lat, lon, alt, time))
self.assertFalse(check_iot_in_range(prop1, lat, lon, alt, time.shiftedBy(60.*30.)))
def to_orekit_date(epoch):
"""Convert UTC simulation time from python's datetime object to Orekit's AbsoluteDate object.
Args:
epoch (datetime.datetime): UTC simulation time
Returns:
AbsoluteDate: simulation time in UTC
"""
seconds = float(epoch.second) + float(epoch.microsecond) / 1e6
orekit_date = AbsoluteDate(epoch.year,
epoch.month,
epoch.day,
epoch.hour,
epoch.minute,
seconds,
TimeScalesFactory.getUTC())
return orekit_date
def change_initial_conditions(self, initial_state, date, mass):
"""
Allows to change the initial conditions given to the propagator without initializing it again.
Args:
initial_state (Cartesian): New initial state of the satellite in cartesian coordinates.
date (datetime): New starting date of the propagator.
mass (float64): New satellite mass.
"""
# Redefine the start date
self.date = date
# Create position and velocity vectors as Vector3D
p = Vector3D(
float(initial_state.R[0]) * 1e3,
float(initial_state.R[1]) * 1e3,
float(initial_state.R[2]) * 1e3)
v = Vector3D(
float(initial_state.V[0]) * 1e3,
float(initial_state.V[1]) * 1e3,
float(initial_state.V[2]) * 1e3)
# Initialize orekit date
seconds = float(date.second) + float(date.microsecond) / 1e6
orekit_date = AbsoluteDate(date.year, date.month, date.day, date.hour,
date.minute, seconds,
TimeScalesFactory.getUTC())
# Extract frame
inertialFrame = FramesFactory.getEME2000()
# Evaluate new initial orbit
initialOrbit = CartesianOrbit(PVCoordinates(p, v), inertialFrame,
orekit_date, Cst.WGS84_EARTH_MU)
# Create new spacecraft state
newSpacecraftState = SpacecraftState(initialOrbit, mass)
# Rewrite propagator initial conditions
self.orekit_prop._propagator_num.setInitialState(newSpacecraftState)
def validate_event_detector(event_name):
# Unpack orekit and poliastro events
orekit_event, poliastro_event = DICT_OF_EVENTS[event_name]
# Time of fliht for propagating the orbit
tof = float(2 * 24 * 3600)
# Build the orekit propagator and add the event to it.
propagator = KeplerianPropagator(ss0_orekit)
propagator.addEventDetector(orekit_event)
# Propagate orekit's orbit
state = propagator.propagate(epoch0_orekit, epoch0_orekit.shiftedBy(tof))
orekit_event_epoch_raw = state.orbit.getPVCoordinates().getDate()
# Convert orekit epoch to astropy Time instance
orekit_event_epoch_str = orekit_event_epoch_raw.toString(
TimeScalesFactory.getUTC())
orekit_event_epoch = Time(orekit_event_epoch_str,
scale="utc",
format="isot")
orekit_event_epoch.format = "iso"
print(f"{orekit_event_epoch}")
# Propagate poliastro's orbit
_, _ = cowell(
Earth.k,
ss0_poliastro.r,
ss0_poliastro.v,
# Generate a set of tofs, each one for each propagation second
np.linspace(0, tof, 100) * u.s,
events=[poliastro_event],
)
poliastro_event_epoch = ss0_poliastro.epoch + poliastro_event.last_t
print(f"{poliastro_event_epoch}")
# Test both event epochs by checking the distance in seconds between them
dt = np.abs((orekit_event_epoch - poliastro_event_epoch).to(u.s))
assert_quantity_allclose(dt, 0 * u.s, atol=5 * u.s, rtol=1e-7)
def get_sun_max_elevation(self, year, month, day):
"""
Returns the maximum elevation angle (in radians) of the Sun (that is the zenith) a given day.
Parameters
----------
year : int
Year.
month : int
Month.
day : int
Day.
Returns
-------
current_ele0 : TYPE
DESCRIPTION.
currentutc_date : TYPE
DESCRIPTION.
"""
currentutc_date = AbsoluteDate(year, month, day, 0, 0, 0.0,
TimeScalesFactory.getUTC())
is_max_elevation = False
time_step = 10.0
current_ele0 = self.get_sun_elevation(currentutc_date)
current_ele1 = self.get_sun_elevation(
currentutc_date.shiftedBy(time_step))
if current_ele0 > current_ele1:
is_max_elevation = True
while not is_max_elevation:
current_ele0 = current_ele1
current_ele1 = self.get_sun_elevation(
currentutc_date.shiftedBy(time_step))
if current_ele0 > current_ele1:
is_max_elevation = True
currentutc_date.shiftedBy(-time_step)
currentutc_date = currentutc_date.shiftedBy(time_step)
return current_ele0, currentutc_date
def osc_elems_transformation_ore(self, other, dir):
self._orekit_lock.acquire()
self.load_orekit()
KepOrbElem.vm.attachCurrentThread()
utc = TimeScalesFactory.getUTC()
orekit_date = AbsoluteDate(2017, 1, 1, 12, 1, 1.0, utc)
inertialFrame = FramesFactory.getEME2000()
a = float(other.a)
e = float(other.e)
i = float(other.i)
w = float(other.w)
O = float(other.O)
v = float(other.v)
initialOrbit = KeplerianOrbit(a * 1000.0, e, i, w, O, v,
PositionAngle.TRUE, inertialFrame,
orekit_date, Constants.mu_earth * 1e9)
initialState = SpacecraftState(initialOrbit, 1.0)
#zonal_forces= DSSTZonal(provider,2,1,5)
zonal_forces = DSSTZonal(KepOrbElem.provider, 6, 4, 6)
forces = ArrayList()
forces.add(zonal_forces)
try:
equinoctial = None
if dir:
equinoctial = DSSTPropagator.computeMeanState(
initialState, None, forces)
else:
equinoctial = DSSTPropagator.computeOsculatingState(
initialState, None, forces)
newOrbit = KeplerianOrbit(equinoctial.getOrbit())
self.a = newOrbit.getA() / 1000.0
self.e = newOrbit.getE()
self.i = newOrbit.getI()
self.w = newOrbit.getPerigeeArgument()
self.O = newOrbit.getRightAscensionOfAscendingNode()
self.v = newOrbit.getAnomaly(PositionAngle.TRUE)
# correct ranges
if self.i < 0:
self.i += 2 * np.pi
if self.w < 0:
self.w += 2 * np.pi
if self.O < 0:
self.O += 2 * np.pi
if self.v < 0:
self.v += 2 * np.pi
finally:
self._orekit_lock.release()
def dlAndParseCpfData(username_edc, password_edc, url, datasetIdList,
startDate, endDate):
"""
Downloads and parses CPF prediction data. A CPF file usually contains one week of data. Using both startDate and
endDate parameters, it is possible to truncate this data.
:param username_edc: str, username for the EDC API
:param password_edc: str, password for the EDC API
:param url: str, URL for the EDC API
:param datasetIdList: list of dataset ids to download.
:param startDate: datetime object. Data prior to this date will be removed
:param endDate: datetime object. Data after this date will be removed
:return: pandas DataFrame containing:
- index: datetime object of the data point, in UTC locale
- columns 'x', 'y', and 'z': float, satellite position in ITRF frame in meters
"""
import requests
import json
from org.orekit.time import AbsoluteDate
from org.orekit.time import TimeScalesFactory
from orekit.pyhelpers import absolutedate_to_datetime
utc = TimeScalesFactory.getUTC()
dl_args = {}
dl_args['username'] = username_edc
dl_args['password'] = password_edc
dl_args['action'] = 'data-download'
dl_args['data_type'] = 'CPF'
import pandas as pd
cpfDataFrame = pd.DataFrame(columns=['x', 'y', 'z'])
for datasetId in datasetIdList:
dl_args['id'] = str(datasetId)
dl_response = requests.post(url, data=dl_args)
if dl_response.status_code == 200:
""" convert json string in python list """
data = json.loads(dl_response.text)
currentLine = ''
i = 0
n = len(data)
while (not currentLine.startswith('10')
) and i < n: # Reading lines until the H4 header
currentLine = data[i]
i += 1
while currentLine.startswith('10') and i < n:
lineData = currentLine.split()
mjd_day = int(lineData[2])
secondOfDay = float(lineData[3])
position_ecef = [
float(lineData[5]),
float(lineData[6]),
float(lineData[7])
]
absolutedate = AbsoluteDate.createMJDDate(
mjd_day, secondOfDay, utc)
currentdatetime = absolutedate_to_datetime(absolutedate)
if (currentdatetime >= startDate) and (currentdatetime <=
endDate):
cpfDataFrame.loc[currentdatetime] = position_ecef
currentLine = data[i]
i += 1
else:
print(dl_response.status_code)
print(dl_response.text)
return cpfDataFrame
Constants.WGS84_EARTH_FLATTENING,
ITRF_Frame)
# CREATE THE GROUND STATION
# First create the Topocentric Frame
station_coord = GeodeticPoint(radians(40), radians(-110), 2000.0)
station_frame = TopocentricFrame(earth, station_coord, 'MyGndStation')
# Now create the Ground station itself and the satellite object
GndStation = GroundStation(station_frame)
Sat = ObservableSatellite(1) # create the observable satellite object, name it 1 as default
# SET THE DAY OF THE OBSERVATIONS
yr_mo_day = (2012, 8, 20)
utc = TimeScalesFactory.getUTC()
# OPEN THE CSV FILE CONTAINING THE OBSERVATIONS
data = open('Observations.csv')
csv_data = csv.reader(data)
data_lines = list(csv_data)
# LOOP OVER THE FILE AND CREATE THE OBSERVATION OBJECTS
# AngularRaDec objects hold Ra and Dec in this order
dates = []
obs = []
# I use the list method here to append rows and then transpose it, the next for
class Propagator:
def __init__(self, initial_orbit, initial_date, shifted_date):
self.shifted_date = shifted_date
self.initial_orbit = initial_orbit
self.initial_date = initial_date
def __str__(self):
return str(self.__dict__)
def propagate_keplerian_elements(self):
propagator = KeplerianPropagator(self.initial_orbit)
return propagator.propagate(self.initial_date, self.shifted_date)
if __name__ == "__main__":
instance = KeplerianElements(
24464560.0, 0.7311, 0.122138, 3.10686, 1.00681, 0.048363,
PositionAngle.MEAN, FramesFactory.getEME2000(),
AbsoluteDate(2020, 1, 1, 0, 0, 00.000,
TimeScalesFactory.getUTC()), Constants.GRS80_EARTH_MU)
k = instance.init_keplerian_elements()
init_date = AbsoluteDate(2020, 1, 1, 0, 0, 00.000,
TimeScalesFactory.getUTC())
shift = init_date.shiftedBy(3600.0 * 48)
kp = Propagator(k, init_date, shift)
kp.propagate_keplerian_elements()
print(k)
print(kp)
def __init__(self, orekit_data, settings=None, **kwargs):
super(Orekit, self).__init__(settings=settings, **kwargs)
if self.logger is not None:
self.logger.debug(f'sorts.propagator.Orekit:init')
if self.profiler is not None:
self.profiler.start('Orekit:init')
setup_orekit_curdir(filename=orekit_data)
if self.logger is not None:
self.logger.debug(f'Orekit:init:orekit-data = {orekit_data}')
self.utc = TimeScalesFactory.getUTC()
self.__settings = dict()
self.__settings['SolarStrengthLevel'] = getattr(
org.orekit.forces.drag.atmosphere.data.
MarshallSolarActivityFutureEstimation.StrengthLevel,
self.settings['solar_activity_strength'])
self._tolerances = None
if self.settings['constants_source'] == 'JPL-IAU':
self.mu = Constants.JPL_SSD_EARTH_GM
self.R_earth = Constants.IAU_2015_NOMINAL_EARTH_EQUATORIAL_RADIUS
self.f_earth = (Constants.IAU_2015_NOMINAL_EARTH_EQUATORIAL_RADIUS
- Constants.IAU_2015_NOMINAL_EARTH_POLAR_RADIUS
) / Constants.IAU_2015_NOMINAL_EARTH_POLAR_RADIUS
else:
self.mu = Constants.WGS84_EARTH_MU
self.R_earth = Constants.WGS84_EARTH_EQUATORIAL_RADIUS
self.f_earth = Constants.WGS84_EARTH_FLATTENING
self.M_earth = self.mu / scipy.constants.G
self.__params = None
self._forces = {}
if self.settings['radiation_pressure']:
self._forces['radiation_pressure'] = None
if self.settings['drag_force']:
self._forces['drag_force'] = None
if self.settings['earth_gravity'] == 'HolmesFeatherstone':
provider = org.orekit.forces.gravity.potential.GravityFieldFactory.getNormalizedProvider(
self.settings['gravity_order'][0],
self.settings['gravity_order'][1])
holmesFeatherstone = org.orekit.forces.gravity.HolmesFeatherstoneAttractionModel(
FramesFactory.getITRF(IERSConventions.IERS_2010, True),
provider,
)
self._forces['earth_gravity'] = holmesFeatherstone
elif self.settings['earth_gravity'] == 'Newtonian':
Newtonian = org.orekit.forces.gravity.NewtonianAttraction(self.mu)
self._forces['earth_gravity'] = Newtonian
else:
raise Exception(
'Supplied Earth gravity model "{}" not recognized'.format(
self.settings['earth_gravity']))
if self.settings['solarsystem_perturbers'] is not None:
for body in self.settings['solarsystem_perturbers']:
body_template = getattr(CelestialBodyFactory,
'get{}'.format(body))
body_instance = body_template()
perturbation = org.orekit.forces.gravity.ThirdBodyAttraction(
body_instance)
self._forces['perturbation_{}'.format(body)] = perturbation
if self.logger is not None:
for key in self._forces:
if self._forces[key] is not None:
self.logger.debug(
f'Orekit:init:_forces:{key} = {type(self._forces[key])}'
)
else:
self.logger.debug(f'Orekit:init:_forces:{key} = None')
if self.profiler is not None:
self.profiler.stop('Orekit:init')
def create_data_validity_checklist():
"""Get files loader by DataProvider and create dict() with valid dates for loaded data.
Creates a list with valid start and ending date for data loaded by the
DataProvider during building. The method looks for follwing folders
holding the correspoding files:
Earth-Orientation-Parameters: EOP files using IERS2010 convetions
MSAFE: NASA Marshall Solar Activity Future Estimation files
Magnetic-Field-Models: magentic field model data files
The list should be used before every propagation step, to check if data
is loaded/still exists for current simulation time. Otherwise
simulation could return results with coarse accuracy. If e.g. no
EOP data is available, null correction is used, which could worsen the
propagators accuracy.
"""
checklist = dict()
start_dates = []
end_dates = []
EOP_file = _get_name_of_loaded_files('Earth-Orientation-Parameters')
if EOP_file:
EOPHist = FramesFactory.getEOPHistory(IERS.IERS_2010, False)
EOP_start_date = EOPHist.getStartDate()
EOP_end_date = EOPHist.getEndDate()
checklist['EOP_dates'] = [EOP_start_date, EOP_end_date]
start_dates.append(EOP_start_date)
end_dates.append(EOP_end_date)
CelesTrack_file = _get_name_of_loaded_files('CELESTRACK')
if CelesTrack_file:
ctw = CelesTrackWeather("(?:sw|SW)\\p{Digit}+\\.(?:txt|TXT)")
ctw_start_date = ctw.getMinDate()
ctw_end_date = ctw.getMaxDate()
checklist['CTW_dates'] = [ctw_start_date, ctw_end_date]
start_dates.append(ctw_start_date)
end_dates.append(ctw_end_date)
MSAFE_file = _get_name_of_loaded_files('MSAFE')
if MSAFE_file:
msafe = MarshallSolarActivityFutureEstimation(
"(?:Jan|Feb|Mar|Apr|May|Jun|Jul|Aug|Sep|Oct|Nov|Dec)" +
"\\p{Digit}\\p{Digit}\\p{Digit}\\p{Digit}F10\\" +
".(?:txt|TXT)",
MarshallSolarActivityFutureEstimation.StrengthLevel.AVERAGE)
MS_start_date = msafe.getMinDate()
MS_end_date = msafe.getMaxDate()
checklist['MSAFE_dates'] = [MS_start_date, MS_end_date]
start_dates.append(MS_start_date)
end_dates.append(MS_end_date)
if FileDataHandler._mag_field_coll:
coll_iterator = FileDataHandler._mag_field_coll.iterator()
first = coll_iterator.next()
GM_start_date = first.validFrom()
GM_end_date = first.validTo()
for GM in coll_iterator:
if GM_start_date > GM.validFrom():
GM_start_date = GM.validFrom()
if GM_end_date < GM.validTo():
GM_end_date = GM.validTo()
# convert to absolute date for later comparison
dec_year = GM_start_date
base = datetime(int(dec_year), 1, 1)
dec = timedelta(seconds=(base.replace(year=base.year + 1) -
base).total_seconds() * (dec_year-int(dec_year)))
GM_start_date = to_orekit_date(base + dec)
dec_year = GM_end_date
base = datetime(int(dec_year), 1, 1)
dec = timedelta(seconds=(base.replace(year=base.year + 1) -
base).total_seconds() * (dec_year-int(dec_year)))
GM_end_date = to_orekit_date(base + dec)
checklist['MagField_dates'] = [GM_start_date, GM_end_date]
start_dates.append(GM_start_date)
end_dates.append(GM_end_date)
if checklist: # if any data loaded define first and last date
# using as date zero 01.01.1850. Assuming no data before.
absolute_zero = AbsoluteDate(1850, 1, 1, TimeScalesFactory.getUTC())
first_date = min(start_dates, key=lambda p: p.durationFrom(absolute_zero))
last_date = min(end_dates, key=lambda p: p.durationFrom(absolute_zero))
checklist['MinMax_dates'] = [first_date, last_date]
mesg = "[INFO]: Simulation can run between epochs: " + \
str(first_date) + " & " + str(last_date) + \
" (based on loaded files)."
print mesg
FileDataHandler._data_checklist = checklist
def main():
a = 24396159.0 # semi major axis (m)
e = 0.720 # eccentricity
i = radians(10.0) # inclination
omega = radians(50.0) # perigee argument
raan = radians(150) # right ascension of ascending node
lM = 0.0 # mean anomaly
# Set inertial frame
inertialFrame = FramesFactory.getEME2000()
# Initial date in UTC time scale
utc = TimeScalesFactory.getUTC()
initial_date = AbsoluteDate(2004, 1, 1, 23, 30, 00.000, utc)
# Setup orbit propagator
# gravitation coefficient
mu = 3.986004415e+14
# Orbit construction as Keplerian
initialOrbit = KeplerianOrbit(a, e, i, omega, raan, lM, PositionAngle.MEAN,
inertialFrame, initial_date, mu)
initial_state = SpacecraftState(initialOrbit)
# use a numerical propogator
min_step = 0.001
max_step = 1000.0
position_tolerance = 10.0
propagation_type = OrbitType.KEPLERIAN
tolerances = NumericalPropagator.tolerances(position_tolerance,
initialOrbit, propagation_type)
integrator = DormandPrince853Integrator(min_step, max_step, 1e-5, 1e-10)
propagator = NumericalPropagator(integrator)
propagator.setOrbitType(propagation_type)
# force model gravity field
provider = GravityFieldFactory.getNormalizedProvider(10, 10)
holmesFeatherstone = HolmesFeatherstoneAttractionModel(
FramesFactory.getITRF(IERSConventions.IERS_2010, True), provider)
# SRP
# ssc = IsotropicRadiationSingleCoefficient(100.0, 0.8) # Spacecraft surface area (m^2), C_r absorbtion
# srp = SolarRadiationPressure(CelestialBodyFactory.getSun(), a, ssc) # sun, semi-major Earth, spacecraft sensitivity
propagator.addForceModel(holmesFeatherstone)
# propagator.addForceModel(ThirdBodyAttraction(CelestialBodyFactory.getSun()))
# propagator.addForceModel(ThirdBodyAttraction(CelestialBodyFactory.getMoon()))
# propagator.addForceModel(srp)
propagator.setMasterMode(60.0, TutorialStepHandler())
propagator.setInitialState(initial_state)
# propagator.setEphemerisMode()
finalstate = propagator.propagate(initial_date.shiftedBy(
10. * 24 * 60**2)) # TIme shift in seconds