def test_read_write_bsp( plot = False ):
'''
Propagate a 30 degree inclination orbit (20 periods),
write BSP kernel, then read back the BSP kernel, ensuring
that latitudes are within +-30 degrees before and after
writing kernel
'''
spice.furnsh( sd.pck00010 )
sc = SC( {
'coes' : [ 8000.0, 0.01, 30.0, 0, 0, 0 ],
'tspan': '20'
} )
'''
Ensure latitudes are +-30 degrees
'''
sc.calc_latlons()
assert np.all( sc.latlons[ :, 2 ] <= 30.0 )
assert np.all( sc.latlons[ :, 2 ] >= -30.0 )
'''
spice.spkopn will error if a bsp with the requested
filename already exists
'''
filename = 'test_read_write_bsp.bsp'
if os.path.isfile( filename ):
os.remove( filename )
'''
Write bsp with all the default arguments except filename
'''
st.write_bsp( sc.ets, sc.states[ :, :6 ],
{ 'bsp_fn': filename } )
'''
Now read back the bsp and ensure that
latitudes are still within +-30 degrees
'''
spice.furnsh( filename )
states = st.calc_ephemeris( -999, sc.ets, 'IAU_EARTH', 399 )
latlons = nt.cart2lat( states[ :, :3 ] )
assert np.all( latlons[ :, 2 ] <= 30.0 )
assert np.all( latlons[ :, 2 ] >= -30.0 )
os.remove( filename )
def test_Spacecraft_minimum_altitude_stop_condition(plot=False):
sc = SC({
'coes': [pd.earth['radius'] + 1000.0, 0.5, 0.0, 90.0, 0.0, 0.0],
'tspan': '1',
'dt': 100.0,
'rtol': 1e-9,
'stop_conditions': {
'min_alt': 100.0
} # km
})
assert sc.cb == pd.earth
'''
Since inclination is 0, all z-axis components of Spacecraft
position and velocity should be 0
'''
assert np.all(sc.states[:, 2] == 0.0)
assert np.all(sc.states[:, 5] == 0.0)
'''
The orbital elements were set such that minimum altitude
crossing would occur, triggering the min_alt stop condition
'''
assert sc.ode_sol.success
assert sc.ode_sol.message == 'A termination event occurred.'
assert sc.ode_sol.status == 1
assert pytest.approx(
np.linalg.norm( sc.ode_sol.y_events[ 0 ][ :3 ] ) -\
sc.cb[ 'radius' ] - 100.0, abs = 6.0e-3 ) == 0.0
if plot:
sc.plot_altitudes({
'hlines': [{
'val': 100.0,
'color': 'c'
}],
'time_unit': 'hours',
'title': 'test_Spacecraft_minimum_altitude_stop_condition',
'show': True
})
def test_Spacecraft_basic_propagation(plot=False):
sc = SC({
'coes': [pd.earth['radius'] + 1000.0, 0.0, 0.0, 0.0, 0.0, 0.0],
'tspan': '1',
'dt': 100.0,
'rtol': 1e-8
})
assert sc.cb == pd.earth
'''
Since inclination is 0, all z-axis components of Spacecraft
position and velocity should be 0
'''
assert np.all(sc.states[:, 2] == 0.0)
assert np.all(sc.states[:, 5] == 0.0)
'''
Since there are no orbital perturbations, all COEs except
true anomaly should be close to constant
(not exactly constant due to numerical error).
However, since this is a circular orbit, periapsis is loosely
defined, causing errors in true anomaly (not exactly linear),
argument of periapsis (bounces back and forth between
0 and 359.9 degrees), eccentricity (random noise), and
semi-major axis (drift / random noise)
'''
sc.calc_coes()
assert sc.coes_calculated
assert pytest.approx(sc.coes[:, 0],
abs=1e-3) == pd.earth['radius'] + 1000.0 # sma
assert pytest.approx(sc.coes[:, 1], abs=1e-6) == 0.0 # ecc
assert np.all(sc.coes[:, 2] == 0.0) # inc
assert np.all(sc.coes[:, 5] == 0.0) # raan
sc.calc_apoapses_periapses()
apse_diffs = sc.apoapses - sc.periapses
assert pytest.approx(apse_diffs, abs=1e-3) == 0.0
if plot:
sc.plot_coes()
def test_Spacecraft_inclination_latitude(plot=False):
spice.furnsh(sd.pck00010)
sc = SC({
'coes': [pd.earth['radius'] + 1000.0, 0.01, 50.0, 0.0, 0.0, 0.0],
'tspan': '3',
'dt': 10.0,
'rtol': 1e-8
})
assert sc.cb == pd.earth
'''
Given that this spacecraft has 50 degrees inclination,
the latitude coordinates should remain in between
-50 and 50 degrees, since inclination is defined as
the angle between the orbital plane and Earth's
equatorial plane
'''
sc.calc_latlons()
assert np.all( sc.latlons[ :, 2 ] <= 50.0 ) and\
np.all( sc.latlons[ :, 2 ] >= -50.0 )
if plot:
sc.plot_groundtracks()
def test_Spacecraft_maximum_altitude_stop_condition(plot=False):
stop_conditions = {'max_alt': pd.earth['SOI'] - pd.earth['radius']}
sc = SC({
'coes': [-(pd.earth['radius'] + 1000.0), 1.8, 0.0, 10.0, 0.0, 0.0],
'tspan': 5 * 24 * 3600.0,
'rtol': 1e-9,
'stop_conditions': stop_conditions
})
assert sc.cb == pd.earth
'''
Since inclination is 0, all z-axis components of Spacecraft
position and velocity should be 0
'''
assert np.all(sc.states[:, 2] == 0.0)
assert np.all(sc.states[:, 5] == 0.0)
'''
The orbital elements were set such that maximum altitude
crossing would occur, triggering the max_alt stop condition
'''
assert sc.ode_sol.success
assert sc.ode_sol.message == 'A termination event occurred.'
assert sc.ode_sol.status == 1
assert pytest.approx(
np.linalg.norm( sc.ode_sol.y_events[ 1 ][ :3 ] ) -\
sc.cb[ 'radius' ], abs = 1e-3 ) ==\
sc.cb[ 'SOI' ] - sc.cb[ 'radius' ]
if plot:
hline = {'val': pd.earth['SOI'] - pd.earth['radius'], 'color': 'c'}
sc.plot_altitudes({
'hlines': [hline],
'time_unit': 'hours',
'title': 'test_Spacecraft_maximum_altitude_stop_condition',
'show': True
})
'''
state0 = [
7.44304239e+03,
-5.12480267e+02,
2.37748995e+03, # km
5.99287925e+00,
1.19056063e+01,
5.17252832e+00
] # km / s
# propagate until exit Earth SOI
sc0 = SC({
'orbit_state': state0,
'et0': et0,
'frame': FRAME,
'tspan': 100000,
'dt': 10000,
'stop_conditions': {
'exit_SOI': True
}
})
'''
Calculate spacecraft state w.r.t solar system barycenter
at ephemeris time when spacecraft left Earth SOI
'''
state_earth = spice.spkgeo(399, sc0.ets[-1], FRAME, 0)[0]
state1 = sc0.states[-1, :6] + state_earth
'''
Now model the spacecraft as a heliocentric elliptical orbit
and propagate until enter Jupiter SOI
'''
sc1 = SC({
from Spacecraft import Spacecraft as SC
import orbit_calculations as oc
import plotting_tools as pt
import planetary_data as pd
if __name__ == '__main__':
periapsis = pd.earth[ 'radius' ] + 4000.0 # km
coes_circular = [ periapsis, 0, 0, 0, 0, 0 ]
coes_elliptical = [ periapsis / 0.3, 0.7, 0, 0, 0, 0 ]
state_parabolic = spice.conics(
[ periapsis, 1.0, 0, 0, 0, -10.0, 0, pd.earth[ 'mu' ] ], 0 )
state_hyperbolic = spice.conics(
[ periapsis, 2.5, 0, 0, 0, -10.0, 0, pd.earth[ 'mu' ] ], 0 )
sc_circular = SC( { 'coes': coes_circular, 'tspan': '1' } )
sc_elliptical = SC( { 'coes': coes_elliptical, 'tspan': '2' } )
sc_parabolic = SC( { 'orbit_state': state_parabolic, 'tspan': 70000.0 } )
sc_hyperbolic = SC( { 'orbit_state': state_hyperbolic, 'tspan': 30000.0 } )
rs = [ sc_circular.states [ :, :3 ], sc_elliptical.states[ :, :3 ],
sc_parabolic.states[ :, :3 ], sc_hyperbolic.states[ :, :3 ] ]
labels = [ 'Circular $(ecc=0.0)$', 'Elliptical $(ecc=0.7)$',
'Parabolic $(ecc=1.0)$', 'Hyperbolic $(ecc=3.0)$' ]
sc_elliptical.plot_states( {
'time_unit': 'hours',
'show' : True
} )
pt.plot_orbits( rs,
from Spacecraft import Spacecraft as SC
import orbit_calculations as oc
import numpy as np
import matplotlib.pyplot as plt
plt.style.use('dark_background')
# Molniya orbital elements
a = 6778.0
coes = [a, 0.0, 0.0, 0.0, 0.0, 0.0]
sc_config = {'coes': coes, 'tspan': '2.5', 'dt': 100.0}
if __name__ == '__main__':
sc = SC(sc_config)
accels = np.zeros((sc.states.shape[0], 3))
for n in range(sc.states.shape[0]):
accels[n] = oc.two_body_ode(sc.ets[n], sc.states[n])[3:]
rnorms = np.linalg.norm(sc.states[:, :3], axis=1)
vnorms = np.linalg.norm(sc.states[:, 3:6], axis=1)
anorms = np.linalg.norm(accels, axis=1)
fig, (ax0, ax1, ax2) = plt.subplots(3, 1, figsize=(20, 10))
ets = (sc.ets - sc.ets[0]) / 3600.0
ax0.plot(ets, accels[:, 0], 'r', label=r'$a_x$')
ax0.plot(ets, accels[:, 1], 'g', label=r'$a_y$')
import ode_tools as ot
import plotting_tools as pt
import planetary_data as pd
if __name__ == '__main__':
r0 = pd.earth['radius'] + 1000.0
v0_circ = (pd.earth['mu'] / r0)**0.5
state0_sc0 = [r0, 0, 0, 0, v0_circ, 0.0]
state0_sc1 = [r0, 0, 0, 2.0, 9.0, 0.0]
state0_sc2 = [r0, 0, 0, 1.0, 0.0, 8.0]
state0_sc3 = [r0, 0, 0, 1.0, 6.0, 6.0]
states0 = [state0_sc0, state0_sc1, state0_sc2, state0_sc3]
rs = []
sc_config = {'tspan': '1'}
for state0 in states0:
sc_config['orbit_state'] = state0
rs.append(SC(sc_config).states[:, :3])
pt.plot_orbits(
rs, {
'labels': range(4),
'colors': ['C3', 'm', 'c', 'lime'],
'traj_lws': 2,
'azimuth': -32,
'elevation': 70,
'show': True
})
'''
AWP | Astrodynamics with Python by Alfonso Gonzalez
https://github.com/alfonsogonzalez/AWP
https://www.youtube.com/c/AlfonsoGonzalezSpaceEngineering
Hello world of Spacecraft class
Two-body propagation with J2 perturbation for 100 periods
'''
# Python standard libraries
# AWP libraries
from Spacecraft import Spacecraft as SC
from planetary_data import earth
if __name__ == '__main__':
coes = [earth['radius'] + 1000, 0.05, 30.0, 0.0, 0.0, 0.0]
sc = SC({
'coes': coes,
'tspan': '100',
'dt': 100.0,
'orbit_perts': {
'J2': True
}
})
sc.plot_3d()
# AWP library
from Spacecraft import Spacecraft as SC
import spice_data as sd
import spice_tools as st
# 3rd party libraries
import spiceypy as spice
if __name__ == '__main__':
spice.furnsh(sd.leapseconds_kernel)
spice.furnsh(sd.de432)
spice.furnsh(sd.pck00010)
sc = SC({
'date0': '2021-03-03 22:10:35 TDB',
'coes': [7480.0, 0.09, 5.5, 6.26, 5.95, 0.2],
'tspan': '40',
})
st.write_bsp(sc.ets, sc.states[:, :6], {'bsp_fn': 'leo.bsp'})
spice.furnsh('leo.bsp')
et0 = spice.str2et('2021-03-03 22:10:40 TDB')
etf = spice.str2et('2021-03-04 TDB')
timecell = spice.utils.support_types.SPICEDOUBLE_CELL(2)
spice.appndd(et0, timecell)
spice.appndd(etf, timecell)
cell = spice.gfoclt('ANY', '399', 'ELLIPSOID', 'IAU_EARTH', '10',
'ELLIPSOID', 'IAU_SUN', 'LT', '-999', 120.0, timecell)
def test_Spacecraft_enter_SOI_voyager2_jupiter(plot=False):
spice.furnsh(sd.leapseconds_kernel)
spice.furnsh(sd.de432)
frame = 'ECLIPJ2000'
# beginning of coverage for Voyager 2 SPK kernel
et0 = spice.str2et('1977 AUG 20 15:32:32.182')
'''
Voyager 2 initial state w.r.t Earth at et0
'''
state0 = [
7.44304239e+03,
-5.12480267e+02,
2.37748995e+03, # km
5.99287925e+00,
1.19056063e+01,
5.17252832e+00
] # km / s
'''
Propagate Voyager 2 trajectory until it exits Earth's sphere of influence,
where maximum altitude stop condition will be triggered
'''
stop_conditions = {'max_alt': pd.earth['SOI'] - pd.earth['radius']}
sc0 = SC({
'orbit_state': state0,
'et0': et0,
'frame': frame,
'tspan': 100000,
'stop_conditions': stop_conditions
})
assert sc0.cb == pd.earth
assert sc0.ode_sol.success
assert sc0.ode_sol.message == 'A termination event occurred.'
assert sc0.ode_sol.status == 1
'''
Calculate spacecraft state w.r.t solar system barycenter
at ephemeris time when spacecraft left Earth SOI
'''
state_earth = spice.spkgeo(399, sc0.ets[-1], frame, 0)[0]
state1 = sc0.states[-1, :6] + state_earth
'''
Now model the spacecraft as a heliocentric elliptical orbit
and propagate until enter Jupiter SOI
'''
sc1 = SC({
'orbit_state': state1,
'et0': sc0.ets[-1],
'frame': frame,
'tspan': 5 * 365 * 24 * 3600.0,
'dt': 30000,
'stop_conditions': {
'enter_SOI': pd.jupiter
},
'cb': pd.sun
})
assert sc1.cb == pd.sun
assert sc1.ode_sol.success
assert sc1.ode_sol.message == 'A termination event occurred.'
assert sc1.ode_sol.status == 1
if plot:
ets = np.concatenate((sc0.ets, sc1.ets))
rs_earth = st.calc_ephemeris(3, ets, frame, 10)[:, :3]
rs_jupiter = st.calc_ephemeris(5, ets, frame, 10)[:, :3]
rs0 = sc0.states[ :, :3 ] +\
st.calc_ephemeris( 3, sc0.ets, frame, 10 )[ :, :3 ]
labels = [
'$Voyager2_{EarthSOI}$', '$Voyager2_{SunSOI}$', 'Earth', 'Jupiter'
]
pt.plot_orbits([rs0, sc1.states[:, :3], rs_earth, rs_jupiter], {
'labels': labels,
'colors': ['m', 'c', 'b', 'C3'],
'show': True
})
# AWP libraries
from Spacecraft import Spacecraft as SC
from planetary_data import earth
import plotting_tools as pt
# 3rd party libraries
import numpy as np
aops = np.arange(0, 360, 90)
incs = np.arange(0, 90, 20)
tas = [0, 180]
coes = [earth['radius'] + 10000, 0.05, 0.0, 0.0, 0.0, 0.0]
scs = []
config = {'tspan': '1', 'dt': 100.0}
print(len(aops) * len(incs) * len(tas))
if __name__ == '__main__':
for inc in incs:
for aop in aops:
for ta in tas:
coes[2] = inc
coes[4] = ta
coes[5] = aop
config['coes'] = coes
sc = SC(config)
scs.append(sc)
rs = [sc.states[:, :3] for sc in scs]
pt.plot_orbits(rs, {'traj_lws': 1, 'show': True})
'''
AWP | Astrodynamics with Python by Alfonso Gonzalez
https://github.com/alfonsogonzalez/AWP
https://www.youtube.com/c/AlfonsoGonzalezSpaceEngineering
Sun-synchronous orbit spacecraft eclipse calculations
'''
# AWP library
from Spacecraft import Spacecraft as SC
import spice_data as sd
# 3rd party libraries
import spiceypy as spice
if __name__ == '__main__':
spice.furnsh( sd.leapseconds_kernel )
spice.furnsh( sd.de432 )
sc = SC( {
'date0': '2021-01-01',
'coes' : [ 6378 + 890.0, 0.0, 99.0, 0, 0, 80 ],
'tspan': '4',
} )
sc.calc_eclipses( vv = True )
sc.plot_eclipse_array( {
'time_unit': 'days',
'show' : True
} )
# 3rd party libraries
import numpy as np
# AWP library
from Spacecraft import Spacecraft as SC
import orbit_calculations as oc
import numerical_tools as nt
import plotting_tools as pt
import planetary_data as pd
if __name__ == '__main__':
periapsis = pd.earth[ 'radius' ] + 4000.0 # km
coes = [ periapsis / 0.3, 0.7, 0, 30.0, 0, 0 ]
state = oc.coes2state( coes )
v_rot = np.dot( nt.Cz( 40.0 * nt.d2r ), state[ 3: ] )
state_rot = np.concatenate( ( state[ :3 ], v_rot ) )
sc0 = SC( { 'orbit_state': state, 'tspan': '1' } )
sc_rot = SC( { 'orbit_state': state_rot, 'tspan': '1' } )
pt.plot_orbits( [ sc0.states[ :, :3 ], sc_rot.states[ :, :3 ] ],
{
'labels' : [ 'Before', 'After' ],
'colors' : [ 'c', 'm' ],
'azimuth' : -90.0,
'elevation' : 90.0,
'axes_custom': 36000.0,
'traj_lws' : 2.5,
'show' : True
} )
'''
AWP | Astrodynamics with Python by Alfonso Gonzalez
https://github.com/alfonsogonzalez/AWP
https://www.youtube.com/c/AlfonsoGonzalezSpaceEngineering
Geostationary orbit spacecraft eclipse calculations
'''
from Spacecraft import Spacecraft as SC
import spice_data as sd
import spiceypy as spice
if __name__ == '__main__':
spice.furnsh( sd.leapseconds_kernel )
spice.furnsh( sd.de432 )
sc = SC( {
'date0': '2021-01-01',
'coes' : [ 42164, 0, 0, 0, 0, 0 ],
'tspan': '365',
} )
sc.plot_3d()
sc.calc_eclipses( vv = True )
sc.plot_eclipse_array( {
'time_unit': 'days',
'show' : True,
} )