def main():
for size in 10, 1000, 100000:
jd = np.linspace(2414992.5, 2471184.50, size)
kernel = SPK.open('de421.bsp')
ephem = Ephemeris(de421)
mars = kernel[0,4]
print(size)
print('-- old code (2 successive runs):')
t0 = time()
ephem.position('mars', jd)
print(time() - t0)
t0 = time()
ephem.position('mars', jd)
print(time() - t0)
print('-- new SPK-powered code (2 successive runs):')
t0 = time()
mars.compute(jd)
print(time() - t0)
t0 = time()
mars.compute(jd)
print(time() - t0)
print()
def test_scalar_input(self):
import de421
e = Ephemeris(de421)
self.check0(*e.compute("earthmoon", 2414994.0))
self.check1(*e.compute("earthmoon", 2415112.5))
def get_earth_to_sun(JD):
eph = Ephemeris(de421)
AU = 149597870.7 # km/AU
barycenter = eph.position('earthmoon', JD)
moonvector = eph.position('moon', JD)
earth = barycenter - moonvector * eph.earth_share
return -1.0 * vector(earth)/AU
def setUp(self):
try:
import de421
except ImportError:
raise SkipTest('the "de421" ephemeris package is not installed')
self.eph = Ephemeris(de421)
self.jalpha = self.eph.jalpha
self.jomega = self.eph.jomega
def test_ephemeris_end_date(self):
import de421
e = Ephemeris(de421)
x, y, z = e.position("earthmoon", e.jomega)
self.assertAlmostEqual(x, -2.81196460e07, delta=1.0)
self.assertAlmostEqual(y, 1.32000379e08, delta=1.0)
self.assertAlmostEqual(z, 5.72139011e07, delta=1.0)
def setUp(self):
try:
import de421
except ImportError:
raise SkipTest('the "de421" ephemeris package has not been'
' installed with "pip install de421"')
self.eph = Ephemeris(de421)
self.jalpha = self.eph.jalpha
self.jomega = self.eph.jomega
def test_array_input(self):
import de421
e = Ephemeris(de421)
v = e.compute("earthmoon", np.array([2414994.0, 2415112.5]))
v = np.array(v)
self.check0(*v[:, 0])
self.check1(*v[:, 1])
def find_residuals(r, rdot, inputJulian, dec_to_check, ra_to_check, timePass):
#Verlet Integrator
k = 0.01720209895 # Gaussian gravitational consgtant
dtau = 0.001 # This is Gaussian Days
dt = dtau / k # This is in mean solar days
tao = 0.0
deltao = 0.001
r = r2guess
rdot = r2Dot
r0Verlet = r
r1Verlet = r0Verlet + (rdot * deltao) + (0.5*acceleration(r0Verlet, Rho2hat*Rho2)*(deltao**2))
while tao <= timePass*k:
r2Update = (2*r1Verlet) - r0Verlet + acceleration(r1Verlet, Rho2hat*Rho2)*(deltao**2)
r0Verlet = r1Verlet
r1Verlet = r2Update
tao += deltao
#print r1Verlet
#Ephemeris Generation
AU = 149597870.7 # km/AU
eph = Ephemeris(de421)
barycenter = eph.position('earthmoon', inputJulian)
moonvector = eph.position('moon', inputJulian)
earth = barycenter - moonvector * eph.earth_share
R0 = vector(earth)/AU # This is the Sun to Earth center vector in Equatorial system
#print R0
R_geocentric = -1.0*R0
#gg = ?
rInput = r1Verlet
rho = (rInput + R_geocentric)
rhohat = norm(rho)
declination = arcsin(rhohat.z)
RA = arccos(rhohat.x/(cos(declination)))
if rhohat.y < 0:
RA = 2*math.pi - RA
RA_residuals = abs(RA - ra_to_check)/ra_to_check
Dec_residuals = abs(declination - dec_to_check)/dec_to_check
return RA_residuals, Dec_residuals
def EphemR (JulianDate):
#Import Ephem 421 to get vector from Earth to Sun (R)
AU = 149597870.7 # km/AU
eph = Ephemeris(de421)
barycenter = eph.position('earthmoon', JulianDate)
moonvector = eph.position('moon', JulianDate)
earth = barycenter - moonvector * eph.earth_share
R0 = vector(earth)/(AU) #This is the Sun to Earth center vector in Equatorial system
R_geocentric = -1.*R0 #Multiply by -1 to get Earth to Sun
#gg = #Our geocentric vector
#R_topocentric = R_geocentric - gg #Subtract our topocentric vector
return R_geocentric
def ephemeris_sun_lunar_ECRF(Julian_date):
# import Julian_date
# month = 6
# day = 1
# year = 1980
# hour = 12
# minute = 0
# second = 0
# jd = Julian_date.Julian_date(month, day, year, hour, minute, second)
# k = ephemeris('moon', jd)#2444391.5
# print('position=',k[0,:])
# print('velocity=', k[1, :])
barycenter_position = ephemeris_position_ICRF('earthmoon', Julian_date)
import de405
from jplephem import Ephemeris
eph = Ephemeris(de405)
Position_sun = ephemeris_position_ICRF('sun', Julian_date)
Position_moon_ECRF = ephemeris_position_ICRF('moon',
Julian_date) # 本身就是相对地球的
Position_earth = barycenter_position - Position_moon_ECRF * eph.earth_share
Position_sun_ECRF = Position_sun - Position_earth
return Position_sun_ECRF, Position_moon_ECRF
def planets_pos(self):
from jplephem import Ephemeris
import de421
from PyKEP import epoch
self.eph = Ephemeris(de421)
earthpos = []
marspos = []
venuspos = []
for ep in self.sc_state[3]:
posSun, __ = self.eph.position_and_velocity('sun', epoch(ep, epoch.epoch_type.MJD2000).jd)
positione, __ = self.eph.position_and_velocity('earthmoon', epoch(ep, epoch.epoch_type.MJD2000).jd)
positione = self.eq2eclipt(positione - posSun)
earthpos.append(positione)
positionm, __ = self.eph.position_and_velocity('mars', epoch(ep, epoch.epoch_type.MJD2000).jd)
positionm = self.eq2eclipt(positionm - posSun)
marspos.append(positionm)
positionv, __ = self.eph.position_and_velocity('venus', epoch(ep, epoch.epoch_type.MJD2000).jd)
positionv = self.eq2eclipt(positionv - posSun)
venuspos.append(positionv)
self.earthpos_km = [earthpos[i].reshape((1,3)).tolist()[0] for i in range(len(earthpos))]
self.marspos_km = [marspos[i].reshape((1,3)).tolist()[0] for i in range(len(earthpos))]
self.venuspos_km = [venuspos[i].reshape((1,3)).tolist()[0] for i in range(len(earthpos))]
def test_ephemeris():
# import Julian_date
# month = 6
# day = 1
# year = 1980
# hour = 12
# minute = 0
# second = 0
# jd = Julian_date.Julian_date(month, day, year, hour, minute, second)
# k = ephemeris('moon', jd)#2444391.5
# print('position=',k[0,:])
# print('velocity=', k[1, :])
Sun_state = ephemeris('sun', 2444391.5) # 太阳星历
barycenter_statement = ephemeris('earthmoon', 2444391.5)
import de405
from jplephem import Ephemeris
eph = Ephemeris(de405)
Position_sun = Sun_state[0, :]
Velocity_sun = Sun_state[1, :]
Moon_state = ephemeris('moon', 2444391.5) # 本身就是相对地球的
Position_moon = Moon_state[0, :]
Velocity_moon = Moon_state[1, :]
Earth_state = barycenter_statement - Moon_state * eph.earth_share
Position_earth = Earth_state[0, :]
Velocity_earth = Earth_state[1, :]
R_sun_earth = Position_sun - Position_earth # 地心ICRF参考系下太阳的位置矢量
V_sun_earth = Velocity_sun - Velocity_earth # 地心ICRF参考系下太阳的速度矢量
print(R_sun_earth, V_sun_earth)
class LegacyTests(_CommonTests, TestCase):
def setUp(self):
try:
import de421
except ImportError:
raise SkipTest('the "de421" ephemeris package has not been'
' installed with "pip install de421"')
self.eph = Ephemeris(de421)
self.jalpha = self.eph.jalpha
self.jomega = self.eph.jomega
def position(self, name, tdb, tdb2=0.0):
return self.eph.position(name, tdb, tdb2)
def position_and_velocity(self, name, tdb, tdb2=0.0):
return self.eph.position_and_velocity(name, tdb, tdb2)
def test_names(self):
self.assertEqual(self.eph.names, (
'earthmoon',
'jupiter',
'librations',
'mars',
'mercury',
'moon',
'neptune',
'nutations',
'pluto',
'saturn',
'sun',
'uranus',
'venus',
))
def test_legacy_compute_method(self):
pv = self.eph.compute('earthmoon', 2414994.0)
self.check0(pv[:3], pv[3:])
pv = self.eph.compute('earthmoon', np.array([2414994.0, 2415112.5]))
self.check0(pv[:3, 0], pv[3:, 0])
self.check1(pv[:3, 1], pv[3:, 1])
def test_ephemeris_end_date(self):
x, y, z = self.position('earthmoon', self.jomega)
self.assertAlmostEqual(x, -94189805.73967789, delta=epsilon_m)
self.assertAlmostEqual(y, 1.05103857e+08, delta=1.0)
self.assertAlmostEqual(z, 45550861.44383482, delta=epsilon_m)
def setUp(self):
try:
import de421
except ImportError:
raise SkipTest('the "de421" ephemeris package is not installed')
self.eph = Ephemeris(de421)
self.jalpha = self.eph.jalpha
self.jomega = self.eph.jomega
def planet_elements_and_sv(planet_id, year, month, day, hour, minute, second):
global mu
mu = 1.327124e11
#calculating julian time
j0 = sum(jdcal.gcal2jd(year, month, day))
ut = (hour+minute/60+second/3600)/24
jd = j0+ut
#determining state vectors
eph = Ephemeris(de421)
planet = month_planet_names.planet_name(planet_id)
if planet=="earth":
barycenter = eph.position_and_velocity('earthmoon', jd)
moonvector = eph.position_and_velocity('moon', jd)
r = barycenter[0] - eph.earth_share*moonvector[0]
v = barycenter[1] - eph.earth_share*moonvector[1]
elif planet=="moon":
barycenter = eph.position_and_velocity('earthmoon', jd)
moonvector = eph.position_and_velocity('moon', jd)
r = barycenter[0] - eph.moon_share*moonvector[0]
v = barycenter[1] - eph.moon_share*moonvector[1]
mu = 3.986e14
else:
r, v = eph.position_and_velocity(planet, jd)
r = r.flatten()
v = v.flatten()/(24*3600)
coe = coe_from_sv.coe_from_sv(r, v, mu)
return [coe, r, v, jd]
def setUp(self):
try:
import de421
except ImportError:
raise SkipTest('the "de421" ephemeris package has not been'
' installed with "pip install de421"')
self.eph = Ephemeris(de421)
self.jalpha = self.eph.jalpha
self.jomega = self.eph.jomega
def JPL_Eph_DE405(JD):
'''
:param JD:
:return: unit: m type: ndarray r_Mercury(ECRF), r_Venus(ECRF), r_Earth(ICRF), r_Mars(ECRF),
r_Jupiter(ECRF), r_Saturn(ECRF), r_Uranus(ECRF), \
r_Neptune(ECRF), r_Pluto(ECRF), r_Moon(ECRF), r_Sun(ECRF),r_SunSSB(ICRF)
'''
Sun_state = ephemeris('sun', JD) # 太阳星历
barycenter_statement = ephemeris('earthmoon', JD)
import de405
from jplephem import Ephemeris
eph = Ephemeris(de405)
Moon_To_Earth_state = ephemeris('moon', JD) # 本身就是相对地球的
Earth_state = barycenter_statement - Moon_To_Earth_state * eph.earth_share
r_Earth = Earth_state[0, :]
r_Sun = Sun_state[0, :] - r_Earth
Moon_state = Moon_To_Earth_state
r_Moon = Moon_state[0, :]
Mercury_state = ephemeris('mercury', JD)
r_Mercury = Mercury_state[0, :] - r_Earth
Venus_state = ephemeris('venus', JD)
r_Venus = Venus_state[0, :] - r_Earth
Mars_state = ephemeris('mars', JD)
r_Mars = Mars_state[0, :] - r_Earth
Jupiter_state = ephemeris('jupiter', JD)
r_Jupiter = Jupiter_state[0, :] - r_Earth
Saturn_state = ephemeris('saturn', JD)
r_Saturn = Saturn_state[0, :] - r_Earth
Uranus_state = ephemeris('uranus', JD)
r_Uranus = Uranus_state[0, :] - r_Earth
Neptune_state = ephemeris('neptune', JD)
r_Neptune = Neptune_state[0, :] - r_Earth
Pluto_state = ephemeris('pluto', JD)
r_Pluto = Pluto_state[0, :] - r_Earth
r_SunSSB = Sun_state[0, :]
#获取的数据是以km为单位，返回值要求为m为单位，需要乘以1e3转化
return r_Mercury*1e3,r_Venus*1e3,r_Earth*1e3,r_Mars*1e3,r_Jupiter*1e3,r_Saturn*1e3,r_Uranus*1e3, \
r_Neptune*1e3,r_Pluto*1e3,r_Moon*1e3,r_Sun*1e3,r_SunSSB*1e3
def ephemeris_position_ICRF(celestial_body, Julian_date):
'''
:param celestial_body: earthmoon mercury pluto venus
jupiter moon saturn
librations neptune sun
mars nutations uranus
:param Julian_date:儒略日
:return:天体的位置（km）,在ICRF下的，以太阳系质心为原点
'''
import de405
from jplephem import Ephemeris
eph = Ephemeris(de405)
position = eph.position(celestial_body, Julian_date) # 1980.06.01
position = np.transpose(position)
position = np.array(position).flatten()
return np.array(position)
def main():
for size in 10, 1000, 100000:
jd = np.linspace(2414992.5, 2471184.50, size)
kernel = SPK.open('de432.bsp')
ephem = Ephemeris(de421)
mars = kernel[0,4]
print(size)
print('-- old code (2 successive runs):')
t0 = time()
ephem.position('mars', jd)
print(time() - t0)
t0 = time()
ephem.position('mars', jd)
print(time() - t0)
print('-- new SPK-powered code (2 successive runs):')
t0 = time()
mars.compute(jd)
print(time() - t0)
t0 = time()
mars.compute(jd)
print(time() - t0)
print()
class LegacyTests(_CommonTests, TestCase):
def setUp(self):
try:
import de421
except ImportError:
raise SkipTest('the "de421" ephemeris package has not been'
' installed with "pip install de421"')
self.eph = Ephemeris(de421)
self.jalpha = self.eph.jalpha
self.jomega = self.eph.jomega
def position(self, name, tdb, tdb2=0.0):
return self.eph.position(name, tdb, tdb2)
def position_and_velocity(self, name, tdb, tdb2=0.0):
return self.eph.position_and_velocity(name, tdb, tdb2)
def test_names(self):
self.assertEqual(self.eph.names, (
'earthmoon', 'jupiter', 'librations', 'mars', 'mercury',
'moon', 'neptune', 'nutations', 'pluto', 'saturn', 'sun',
'uranus', 'venus',
))
def test_legacy_compute_method(self):
pv = self.eph.compute('earthmoon', 2414994.0)
self.check0(pv[:3], pv[3:])
pv = self.eph.compute('earthmoon', np.array([2414994.0, 2415112.5]))
self.check0(pv[:3,0], pv[3:,0])
self.check1(pv[:3,1], pv[3:,1])
def test_ephemeris_end_date(self):
x, y, z = self.position('earthmoon', self.jomega)
self.assertAlmostEqual(x, -94189805.73967789, delta=epsilon_m)
self.assertAlmostEqual(y, 1.05103857e+08, delta=1.0)
self.assertAlmostEqual(z, 45550861.44383482, delta=epsilon_m)
def ephemeris(celestial_body, Julian_date):
'''
:param celestial_body: earthmoon mercury pluto venus
jupiter moon saturn
librations neptune sun
mars nutations uranus
:param Julian_date:儒略日
:return:天体的位置（km）,速度（km/s）
'''
import de405
from jplephem import Ephemeris
eph = Ephemeris(de405)
position, velocity = eph.position_and_velocity(celestial_body,
Julian_date) # 1980.06.01
position = np.transpose(position)
position = np.array(position).flatten()
velocity = np.transpose(velocity)
velocity = np.array(velocity).flatten()
velocity = velocity / 86400
state = np.array([position, velocity])
return np.array(state)
def find_residuals(check_index, decs, ras, JDs, r, r_dot):
dec_res_total = 0
ra_res_total = 0
total_ra = 0
total_dec = 0
for i in range(len(decs)):
#print str(i + 1) + "/" + str(len(decs))
k = 0.01720209895
time_difference = JDs[i] - JDs[check_index]
tau = 0.0
d_tau = 0.0001 * is_negative(time_difference)
end_tau = time_difference * k
mu = 1.0
r_last = r
r_now = r_last + r_dot * d_tau + 0.5 * accel(r_last, mu) * (d_tau**2)
tau += d_tau
while abs(tau) <= abs(end_tau):
tau += d_tau
temp = verlet_position(r_now, r_last, d_tau, mu)
r_last = temp[1]
r_now = temp[0]
AU = 149597870.7 # km/AU
jd = JDs[i]
eph = Ephemeris(de421)
R_geocentric = get_earth_to_sun(JD)
sun_to_asteroid = r_now
earth_to_asteroid = norm(
R_geocentric +
sun_to_asteroid) # Gets unit vector from Earth to Asteroid
RA, Dec = location_to_angles(earth_to_asteroid)
RA_Residual = abs(ras[i] - RA)
Dec_Residual = abs(decs[i] - Dec)
ra_res_total += RA_Residual**2
dec_res_total += Dec_Residual**2
tau = 0
d_tau = 0.000001
k = 0.01720209895
mu = 1.0
sigma = sqrt(ra_res_total + dec_res_total)
return sigma
def find_residuals(check_index, decs, ras, JDs, r, r_dot):
dec = []
ra = []
for i in range(len(decs)):
k = 0.01720209895
time_difference = JDs[i] - JDs[check_index]
tau = 0.0
d_tau = 0.0001 * neg(time_difference)
end_tau = time_difference * k
mu = 1.0
r_last = r
r_now = r_last + r_dot * d_tau + 0.5 * accel(r_last, mu) * (d_tau ** 2)
tau += d_tau
while abs(tau) <= abs(end_tau):
tau += d_tau
temp = verlet_position(r_now, r_last, d_tau, mu)
r_last = temp[1]
r_now = temp[0]
AU = 149597870.7# km/AU
jd = JDs[i]
eph = Ephemeris(de421)
R_geocentric = get_earth_to_sun(jd)
sun_to_asteroid = r_now
earth_to_asteroid = norm(R_geocentric + sun_to_asteroid) # Gets unit vector from Earth to Asteroid
RA, Dec = location_to_angles(earth_to_asteroid)
ra.append(RA)
dec.append(Dec)
tau = 0
d_tau = 0.000001
k = 0.01720209895
mu = 1.0
return ra, dec
def find_endpoints(check_index, decs, ras, JDs, r, r_dot):
dec_res_total = 0
ra_res_total = 0
total_ra = 0
total_dec = 0
times = []
times.append(JDs[0])
times.append(JDs[len(JDs) - 1])
new_RAs = []
new_decs = []
for i in range(len(JDs)):
k = 0.01720209895
time_difference = JDs[i] - JDs[check_index]
tau = time_difference * k
f, g = best_f_and_g(tau, r, r_dot)
r_now = f * r + g * r_dot
AU = 149597870.7 #km/AU
jd = JDs[i]
eph = Ephemeris(de421)
R_geocentric = get_earth_to_sun(JD)
sun_to_asteroid = r_now
earth_to_asteroid = norm(
R_geocentric +
sun_to_asteroid) #Gets unit vector from Earth to Asteroid
RA, Dec = location_to_angles(earth_to_asteroid)
new_RAs.append(RA)
new_decs.append(Dec)
"""
tau = 0
d_tau = 0.000001
k = 0.01720209895
mu = 1.0
"""
return new_RAs, new_decs
def find_residuals(check_index, decs, ras, JDs, r, r_dot):
dec_res_total = 0
ra_res_total = 0
total_ra = 0
total_dec = 0
for i in range(len(decs)):
k = 0.01720209895
time_difference = JDs[i] - JDs[check_index]
tau = time_difference * k
mu = 1.0
f, g = better_f_and_g(tau, r, r_dot)
r_now = f * r + g * r_dot
AU = 149597870.7 # km/AU
jd = JDs[i]
eph = Ephemeris(de421)
R_geocentric = get_earth_to_sun(JD)
sun_to_asteroid = r_now
earth_to_asteroid = norm(
R_geocentric +
sun_to_asteroid) # Gets unit vector from Earth to Asteroid
RA, Dec = location_to_angles(earth_to_asteroid)
RA_Residual = abs(ras[i] - RA)
Dec_Residual = abs(decs[i] - Dec)
ra_res_total += RA_Residual**2
dec_res_total += Dec_Residual**2
tau = 0
d_tau = 0.000001
k = 0.01720209895
mu = 1.0
sigma = sqrt(ra_res_total + dec_res_total)
return sigma
class pos(object):
def __init__(self):
self.trajectory_positions = []
axtl = 23.43929*DEG2RAD # Earth axial tlit
# Macierz przejscia z ECI do helio
self.mactran = matrix([ [1, 0, 0],
[0, cos(axtl), sin(axtl)],
[0, -sin(axtl), cos(axtl)] ])
self.day = 86400.0
def positions_lambert(self, l, sol=0, units = 1.0, index = 0):
"""
Plots a particular solution to a Lambert's problem
USAGE: plot_lambert(ax,l, N=60, sol=0, units = 'PyKEP.AU', legend = 'False')
* l: PyKEP.lambert_problem object
* sol: solution to the Lambert's problem we want to plot (must be in 0..Nmax*2)
where Nmax is the maximum number of revolutions for which there exist a solution.
* units: the length unit to be used in the plot
"""
from PyKEP import propagate_lagrangian, AU
if sol > l.get_Nmax()*2:
raise ValueError("sol must be in 0 .. NMax*2 \n * Nmax is the maximum number of revolutions for which there exist a solution to the Lambert's problem \n * You can compute Nmax calling the get_Nmax() method of the lambert_problem object")
#We extract the relevant information from the Lambert's problem
r = l.get_r1()
v = l.get_v1()[sol]
T = l.get_tof()
mu = l.get_mu()
#We define the integration time ...
if T/86400. < 1:
N = int(T) #...compute number of points...
dt = T / (N-1.)
else:
N = int(2*T/86400.) #...compute number of points...
dt = T / (N-1.)
timetuple = [i*dt for i in range(N)]
#... and alocate the cartesian components for r
x = [0.0]*N
y = [0.0]*N
z = [0.0]*N
#We calculate the spacecraft position at each dt
for i in range(N):
x[i] = r[0]/units
y[i] = r[1]/units
z[i] = r[2]/units
r,v = propagate_lagrangian(r,v,dt,mu)
self.trajectory_positions.append([index, x, y, z, timetuple])
def positions_kepler(self, r,v,t,mu, units = 1, index = 0):
"""
Plots the result of a keplerian propagation
USAGE: plot_kepler(ax,r,v,t,mu, N=60, units = 1, color = 'b', legend = False):
* r: initial position (cartesian coordinates)
* v: initial velocity (cartesian coordinates)
* t: propagation time
* mu: gravitational parameter
* units: the length unit to be used in the plot
"""
from PyKEP import propagate_lagrangian
#We define the integration time ...
if t/86400. < 1:
N = int(t) #...compute number of points...
dt = t / (N-1.)
else:
N = int(2*t/86400.) #...compute number of points...
dt = t / (N-1.)
timetuple = [i*dt for i in range(N)]
#... and calcuate the cartesian components for r
x = [0.0]*N
y = [0.0]*N
z = [0.0]*N
#We calculate the spacecraft position at each dt
for i in range(N):
x[i] = r[0]/units
y[i] = r[1]/units
z[i] = r[2]/units
r,v = propagate_lagrangian(r,v,dt,mu)
self.trajectory_positions.append([index, x, y, z, timetuple])
def return_sc_positions(self):
return self.trajectory_positions
def set_launch_epoch(self, mjd2000_epoch):
self.Lepoch = mjd2000_epoch
def rework_time(self):
T = self.trajectory_positions
n = len(T)
ep = self.Lepoch
timespans = [T[i][4][-1]/86400. for i in range(n)]
beginings = [ep + sum(timespans[:i]) for i in range(n)]
# initialize lists...
t = []
x = []
y = []
z = []
# ...and join values, correcting for time
for g in range(n):
t = t + [beginings[g]+i/86400. for i in T[g][4]]
x = x + T[g][1]
y = y + T[g][2]
z = z + T[g][3]
# save spacecraft state
self.sc_state = [x, y, z, t]
def eq2eclipt(self, xyz):
macxyz = matrix(xyz)
return self.mactran.dot(macxyz)
def planets_pos(self):
from jplephem import Ephemeris
import de421
from PyKEP import epoch
self.eph = Ephemeris(de421)
earthpos = []
marspos = []
venuspos = []
for ep in self.sc_state[3]:
posSun, __ = self.eph.position_and_velocity('sun', epoch(ep, epoch.epoch_type.MJD2000).jd)
positione, __ = self.eph.position_and_velocity('earthmoon', epoch(ep, epoch.epoch_type.MJD2000).jd)
positione = self.eq2eclipt(positione - posSun)
earthpos.append(positione)
positionm, __ = self.eph.position_and_velocity('mars', epoch(ep, epoch.epoch_type.MJD2000).jd)
positionm = self.eq2eclipt(positionm - posSun)
marspos.append(positionm)
positionv, __ = self.eph.position_and_velocity('venus', epoch(ep, epoch.epoch_type.MJD2000).jd)
positionv = self.eq2eclipt(positionv - posSun)
venuspos.append(positionv)
self.earthpos_km = [earthpos[i].reshape((1,3)).tolist()[0] for i in range(len(earthpos))]
self.marspos_km = [marspos[i].reshape((1,3)).tolist()[0] for i in range(len(earthpos))]
self.venuspos_km = [venuspos[i].reshape((1,3)).tolist()[0] for i in range(len(earthpos))]
def distance_to_planets(self):
dre = []
drm = []
drv = []
for i in range(len(self.sc_state[3])):
ep = self.earthpos_km[i]
dist = [ep[0] - self.sc_state[0][i] / 1000.,
ep[1] - self.sc_state[1][i] / 1000.,
ep[2] - self.sc_state[2][i] / 1000.]
dre.append((sum([g**2 for g in dist]))**.5)
mp = self.marspos_km[i]
dist = [mp[0] - self.sc_state[0][i] / 1000.,
mp[1] - self.sc_state[1][i] / 1000.,
mp[2] - self.sc_state[2][i] / 1000.]
drm.append((sum([g**2 for g in dist]))**.5)
vp = self.venuspos_km[i]
dist = [vp[0] - self.sc_state[0][i] / 1000.,
vp[1] - self.sc_state[1][i] / 1000.,
vp[2] - self.sc_state[2][i] / 1000.]
drv.append((sum([g**2 for g in dist]))**.5)
return dre,drm, drv, self.sc_state[3]
def GenerateJourneyDataPlot(self, DataAxesLeft, DataAxesRight, PlotOptions, firstpass):
#generate a vector of dates
date_string_vector = []
for event in self.missionevents:
if event.EventType != 'match_point' and event.EventType != 'upwr_flyby' and event.EventType != 'pwr_flyby':
date_string_vector.append(event.GregorianDate)
date_vector = [datetime.datetime.strptime(d,'%m/%d/%Y').date() for d in date_string_vector]
#dummy line across the bottom so that neither axes object crashes
DataAxesLeft.plot(date_vector, np.zeros_like(date_vector), c='w', lw = 0.1)
DataAxesRight.plot(date_vector, np.zeros_like(date_vector), c='w', lw = 0.1)
#plot distance from central body
if PlotOptions.PlotR:
Rvector = []
for event in self.missionevents:
if event.EventType != 'match_point' and event.EventType != 'upwr_flyby' and event.EventType != 'pwr_flyby':
Rvector.append(math.sqrt(event.SpacecraftState[0]**2 + event.SpacecraftState[1]**2 + event.SpacecraftState[2]**2) / self.LU)
if firstpass:
unitstring = 'LU'
if self.central_body.lower() == 'sun' and math.fabs(self.LU - 149597870.69100001) < 1.0:
unitstring = 'AU'
labelstring = 'Distance from ' + self.central_body + ' (' + unitstring + ')'
DataAxesLeft.plot(date_vector, Rvector, c='k', lw=2, label=labelstring)
else:
DataAxesLeft.plot(date_vector, Rvector, c='k', lw=2)
#plot velocity
if PlotOptions.PlotV:
Vvector = []
for event in self.missionevents:
if event.EventType != 'match_point' and event.EventType != 'upwr_flyby' and event.EventType != 'pwr_flyby':
Vvector.append(math.sqrt(event.SpacecraftState[3]**2 + event.SpacecraftState[4]**2 + event.SpacecraftState[5]**2) / self.LU * self.TU)
if firstpass:
DataAxesLeft.plot(date_vector, Vvector, c='k', lw=2, ls='-.', label='Velocity magnitude (LU/TU)')
else:
DataAxesLeft.plot(date_vector, Vvector, c='k', lw=2, ls='-.')
#plot Thrust
if PlotOptions.PlotThrust:
Thrustvector = []
for event in self.missionevents:
if event.EventType != 'match_point' and event.EventType != 'upwr_flyby' and event.EventType != 'pwr_flyby':
Thrustvector.append(math.sqrt(event.Thrust[0]**2 + event.Thrust[1]**2 + event.Thrust[2]**2) * 10.0)
if firstpass:
DataAxesLeft.plot(date_vector, Thrustvector, c='r', lw=2, ls='-', label='Applied thrust (0.1 N)')
else:
DataAxesLeft.plot(date_vector, Thrustvector, c='r', lw=2, ls='-')
#plot Isp
if PlotOptions.PlotIsp:
Ispvector = []
for event in self.missionevents:
if event.EventType != 'match_point' and event.EventType != 'upwr_flyby' and event.EventType != 'pwr_flyby':
Ispvector.append(event.Isp / 1000.0)
if firstpass:
DataAxesLeft.plot(date_vector, Ispvector, c='c', lw=2, ls='-', label='Isp (1000 s)')
else:
DataAxesLeft.plot(date_vector, Ispvector, c='c', lw=2, ls='-')
#plot mass flow rate
if PlotOptions.PlotMdot:
Mdotvector = []
for event in self.missionevents:
if event.EventType != 'match_point' and event.EventType != 'upwr_flyby' and event.EventType != 'pwr_flyby':
Mdotvector.append(event.MassFlowRate * 1.0e6)
if firstpass:
DataAxesLeft.plot(date_vector, Mdotvector, c='brown', lw=2, ls='-', label='Mass flow rate (mg/s)')
else:
DataAxesLeft.plot(date_vector, Mdotvector, c='brown', lw=2, ls='-')
#plot Efficiency
if PlotOptions.PlotEfficiency:
Efficiencyvector = []
for event in self.missionevents:
if event.EventType == 'SFthrust' or event.EventType == 'FBLTthrust' or event.EventType == "PSBIthrust":
Efficiencyvector.append(event.AvailableThrust * event.Isp * 9.80665 / (2000 * event.ActivePower))
elif event.EventType != 'match_point' and event.EventType != 'upwr_flyby' and event.EventType != 'pwr_flyby':
Efficiencyvector.append(0.0)
if firstpass:
DataAxesLeft.plot(date_vector, Efficiencyvector, c='DarkGreen', lw=2, ls='-', label='Propulsion system efficiency')
else:
DataAxesLeft.plot(date_vector, Efficiencyvector, c='DarkGreen', lw=2, ls='-')
#plot Throttle
if PlotOptions.PlotThrottle:
Throttlevector = []
for event in self.missionevents:
if event.EventType == 'SFthrust' or event.EventType == 'FBLTthrust' or event.EventType == "PSBIthrust":
Throttlevector.append(math.sqrt(event.Thrust[0]**2 + event.Thrust[1]**2 + event.Thrust[2]**2) / (event.AvailableThrust * self.thruster_duty_cycle))
elif event.EventType != 'match_point' and event.EventType != 'upwr_flyby' and event.EventType != 'pwr_flyby':
Throttlevector.append(0.0)
if firstpass:
DataAxesLeft.plot(date_vector, Throttlevector, c='r', lw=2, ls='--', label='Throttle')
else:
DataAxesLeft.plot(date_vector, Throttlevector, c='r', lw=2, ls='--')
#plot power
if PlotOptions.PlotPower:
Powervector = []
for event in self.missionevents:
if event.EventType != 'match_point' and event.EventType != 'upwr_flyby' and event.EventType != 'pwr_flyby':
if event.AvailablePower > 0.0:
Powervector.append(event.AvailablePower)
else:
Powervector.append(0.0)
if firstpass:
DataAxesLeft.plot(date_vector, Powervector, c='Navy', lw=2, ls='-', label='Power produced by spacecraft (kW)')
else:
DataAxesLeft.plot(date_vector, Powervector, c='Navy', lw=2, ls='-')
#plot gamma
if PlotOptions.PlotGamma:
gammavector = []
for event in self.missionevents:
if event.EventType == 'SFthrust' or event.EventType == 'FBLTthrust' or event.EventType == "PSBIthrust":
gammavector.append(math.atan2(event.Thrust[1], event.Thrust[0]) * 180.0 / math.pi)
elif event.EventType != 'match_point' and event.EventType != 'upwr_flyby' and event.EventType != 'pwr_flyby':
gammavector.append(0.0)
if firstpass:
DataAxesRight.plot(date_vector, gammavector, c='DarkGreen', lw=2, ls='--', label=r'$\gamma$ (radians)')
else:
DataAxesRight.plot(date_vector, gammavector, c='DarkGreen', lw=2, ls='--')
#plot delta
if PlotOptions.PlotDelta:
deltavector = []
for event in self.missionevents:
if event.EventType == 'SFthrust' or event.EventType == 'FBLTthrust' or event.EventType == "PSBIthrust":
AppliedThrust = math.sqrt(event.Thrust[0]**2 + event.Thrust[1]**2 + event.Thrust[2]**2)
deltavector.append(math.asin(event.Thrust[2] / AppliedThrust) * 180.0 / math.pi)
elif event.EventType != 'match_point' and event.EventType != 'upwr_flyby' and event.EventType != 'pwr_flyby':
deltavector.append(0.0)
if firstpass:
DataAxesRight.plot(date_vector, deltavector, c='LightGreen', lw=2, ls='--', label=r'$\delta$ (radians)')
else:
DataAxesRight.plot(date_vector, deltavector, c='LightGreen', lw=2, ls='--')
#plot central body to thrust vector angle
if PlotOptions.PlotCB_thrust_angle:
CBthrustvector = []
for event in self.missionevents:
if event.EventType == 'SFthrust' or event.EventType == 'FBLTthrust' or event.EventType == "PSBIthrust":
r = math.sqrt(event.SpacecraftState[0]**2 + event.SpacecraftState[1]**2 + event.SpacecraftState[2]**2)
AppliedThrust = math.sqrt(event.Thrust[0]**2 + event.Thrust[1]**2 + event.Thrust[2]**2)
rdotT = event.SpacecraftState[0]*event.Thrust[0] + event.SpacecraftState[1]*event.Thrust[1] + event.SpacecraftState[2]*event.Thrust[2]
CBthrustvector.append( math.acos( rdotT / (r * AppliedThrust) ) * 180.0 / math.pi)
elif event.EventType != 'match_point' and event.EventType != 'upwr_flyby' and event.EventType != 'pwr_flyby':
CBthrustvector.append(0.0)
if firstpass:
DataAxesRight.plot(date_vector, CBthrustvector, c='Salmon', lw=2, ls='--', label='CB-thrust angle (radians)')
else:
DataAxesRight.plot(date_vector, CBthrustvector, c='Salmon', lw=2, ls='--')
#plot mass
if PlotOptions.PlotMass:
mass = []
for event in self.missionevents:
if event.EventType != 'match_point' and event.EventType != 'upwr_flyby' and event.EventType != 'pwr_flyby':
mass.append(event.Mass * 1.0e-3)
if firstpass:
DataAxesLeft.plot(date_vector, mass, c='DarkGrey', lw=2, ls='-', label='Mass (1000 kg)')
else:
DataAxesLeft.plot(date_vector, mass, c='DarkGrey', lw=2, ls='-')
#plot number of active thrusters
if PlotOptions.PlotNumberOfEngines:
numberofengines = []
for event in self.missionevents:
if event.EventType == 'SFthrust' or event.EventType == 'FBLTthrust' or event.EventType == "PSBIthrust":
numberofengines.append(event.Number_of_Active_Engines)
elif event.EventType != 'match_point' and event.EventType != 'upwr_flyby' and event.EventType != 'pwr_flyby':
numberofengines.append(0)
if firstpass:
DataAxesLeft.plot(date_vector, numberofengines, c='Orange', lw=2, ls='-', label='Number of active thrusters')
else:
DataAxesLeft.plot(date_vector, numberofengines, c='Orange', lw=2, ls='-')
#plot power actively used by the thrusters
if PlotOptions.PlotActivePower:
activepowervector = []
for event in self.missionevents:
if event.EventType == 'SFthrust' or event.EventType == 'FBLTthrust' or event.EventType == "PSBIthrust":
activepowervector.append(event.ActivePower)
elif event.EventType != 'match_point' and event.EventType != 'upwr_flyby' and event.EventType != 'pwr_flyby':
activepowervector.append(0.0)
if firstpass:
DataAxesLeft.plot(date_vector, activepowervector, c='Navy', lw=2, ls='--', label='Power used by the propulsion system (kW)')
else:
DataAxesLeft.plot(date_vector, activepowervector, c='Navy', lw=2, ls='--')
#plot waste heat from the propulsion system
if PlotOptions.PlotWasteHeat:
WasteHeatvector = []
for event in self.missionevents:
if event.EventType == 'SFthrust' or event.EventType == 'FBLTthrust' or event.EventType == "PSBIthrust":
WasteHeatvector.append( (1 - event.AvailableThrust * event.Isp * 9.80665 / (2000 * event.ActivePower)) * event.ActivePower )
elif event.EventType != 'match_point' and event.EventType != 'upwr_flyby' and event.EventType != 'pwr_flyby':
WasteHeatvector.append(0.0)
if firstpass:
DataAxesLeft.plot(date_vector, WasteHeatvector, c='Crimson', lw=2, ls='--', label='Waste heat from propulsion system (kW)')
else:
DataAxesLeft.plot(date_vector, WasteHeatvector, c='Crimson', lw=2, ls='--')
#plot Earth distance in LU and SPE angle
if PlotOptions.PlotEarthDistance or PlotOptions.PlotSunEarthSpacecraftAngle:
if self.central_body.lower() != 'sun':
print 'getting Earth distance and Sun-Spacecraft-Earth angle is only supported if the central body is the sun'
else:
import de423
from jplephem import Ephemeris
eph = Ephemeris(de423)
EarthDistanceVector = []
SunEarthSpacecraftAngle = []
for event in self.missionevents:
if event.EventType != 'match_point' and event.EventType != 'upwr_flyby' and event.EventType != 'pwr_flyby':
#get the position of the central body relative to the solar system barycenter
cb_relative_to_solar_system_barycenter_in_ICRC = eph.position(self.central_body.lower(), event.JulianDate)
#get the position of the Earth relative to the central body
embarycenter_in_ICRC = eph.position('earthmoon', event.JulianDate)
moonvector_in_ICRC = eph.position('moon', event.JulianDate)
Earth_relative_to_solar_system_barycenter_in_ICRC = embarycenter_in_ICRC - moonvector_in_ICRC * eph.earth_share
Earth_relative_to_cb_in_ICRC = Earth_relative_to_solar_system_barycenter_in_ICRC + cb_relative_to_solar_system_barycenter_in_ICRC
#rotate the spacecraft state relative to the central body into the ICRF frame
spacecraft_state_relative_to_central_body_in_ICRF = AstroFunctions.rotate_from_ecliptic_to_equatorial3(np.array(event.SpacecraftState[0:3]))
#get the vector from the Earth to the Spacecraft
Earth_Spacecraft_Vector = spacecraft_state_relative_to_central_body_in_ICRF - np.transpose(Earth_relative_to_cb_in_ICRC)[0]
#return the magnitude of the Earth-centered position vector
EarthDistanceVector.append(np.linalg.norm(Earth_Spacecraft_Vector) / self.LU)
#if applicable, compute sun-spacecraft-Earth angle
if PlotOptions.PlotSunEarthSpacecraftAngle:
cosAngle = np.dot(-Earth_Spacecraft_Vector, -spacecraft_state_relative_to_central_body_in_ICRF)/np.linalg.norm(spacecraft_state_relative_to_central_body_in_ICRF)/np.linalg.norm(Earth_Spacecraft_Vector)
SunEarthSpacecraftAngle.append(np.arccos(cosAngle) * 180.0/math.pi)
if PlotOptions.PlotEarthDistance:
if firstpass:
unitstring = 'LU'
if self.central_body.lower() == 'sun' and math.fabs(self.LU - 149597870.69100001) < 1.0:
unitstring = 'AU'
labelstring = 'Distance from Earth (' + unitstring + ')'
DataAxesLeft.plot(date_vector, EarthDistanceVector, c='g', lw=2, label=labelstring)
else:
DataAxesLeft.plot(date_vector, EarthDistanceVector, c='g', lw=2)
if PlotOptions.PlotSunEarthSpacecraftAngle:
if firstpass:
DataAxesRight.plot(date_vector, SunEarthSpacecraftAngle, c='orangered', lw=2, ls = '-.', label='Sun-Spacecraft-Earth Angle')
else:
DataAxesRight.plot(date_vector, SunEarthSpacecraftAngle, c='orangered', lw=2, ls = '-.')
def WriteDateReport(self, PlotOptions, reportfile):
#write header
reportfile.write('JD, MM-DD-YYYY, Event duration (days), Event type, Event location')
if PlotOptions.IncludeStateVectorInReport:
reportfile.write(',x (km), y (km), z (km), xdot (km/s), ydot (km/s), zdot (km/s)')
if PlotOptions.PlotR:
unitstring = 'LU'
if 'sun' in self.central_body.lower() and math.fabs(self.LU - 149597870.69100001) < 1.0:
unitstring = 'AU'
reportfile.write(', Distance from ' + self.central_body + ' (' + unitstring + ')')
if PlotOptions.PlotV:
reportfile.write(', Velocity magnitude (LU/TU)')
if PlotOptions.PlotThrust:
reportfile.write(', Applied thrust (N)')
if PlotOptions.PlotIsp:
reportfile.write(', Isp (s)')
if PlotOptions.PlotMdot:
reportfile.write(', Mass flow rate (kg/s)')
if PlotOptions.PlotEfficiency:
reportfile.write(', Propulsion system efficiency')
if PlotOptions.PlotThrottle:
reportfile.write(', Control magnitude')
if PlotOptions.PlotPower:
reportfile.write(', Power produced by spacecraft (kW)')
if PlotOptions.PlotGamma:
reportfile.write(', in-plane control angle (degrees)')
if PlotOptions.PlotDelta:
reportfile.write(', out-of-plane control angle (degrees)')
if PlotOptions.PlotArray_Thrust_Angle:
reportfile.write(', Array to thrust angle (degrees)')
if PlotOptions.PlotMass:
reportfile.write(', Mass (kg)')
if PlotOptions.PlotNumberOfEngines:
reportfile.write(', Number of active thrusters')
if PlotOptions.PlotActivePower:
reportfile.write(', Power used by the propulsion system (kW)')
if PlotOptions.PlotWasteHeat:
reportfile.write(', Waste heat from propulsion system (kW)')
if PlotOptions.PlotEarthDistance:
unitstring = 'LU'
if 'sun' in self.central_body.lower() and math.fabs(self.LU - 149597870.69100001) < 1.0:
unitstring = 'AU'
reportfile.write(', Distance from Earth (' + unitstring + ')')
if PlotOptions.PlotSunSpacecraftEarthAngle:
reportfile.write(', Sun-Spacecraft-Earth Angle')
if PlotOptions.PlotSpacecraftViewingAngle:
reportfile.write(', Latitude of Earth-Spacecraft Vector')
if PlotOptions.PlotThrottleLevel:
reportfile.write(', Throttle level')
if PlotOptions.PlotSunBoresightAngle:
reportfile.write(', Sun-Boresight Angle')
reportfile.write('\n')
#load stuff if necessary
if PlotOptions.PlotEarthDistance or PlotOptions.PlotSunSpacecraftEarthAngle or PlotOptions.PlotSpacecraftViewingAngle:
if not 'sun' in self.central_body.lower():
print('getting Earth distance and Sun-Spacecraft-Earth angle is only supported if the central body is the sun')
else:
import de423
from jplephem import Ephemeris
self.eph = Ephemeris(de423)
if PlotOptions.PlotThrottleLevel:
#instantiate a throttle table object
thruster_throttle_table = ThrottleTable.ThrottleTable(PlotOptions.throttletablefile)
self.high_Mdot_set, self.low_Mdot_set = thruster_throttle_table.create_non_dominated_sets()
#now print the report
for event in self.missionevents:
reportfile.write(str(event.JulianDate))
reportfile.write(', ' + event.GregorianDate)
reportfile.write(', ' + str(event.TimestepLength))
reportfile.write(', ' + event.EventType)
reportfile.write(', ' + event.Location)
if PlotOptions.IncludeStateVectorInReport:
for state in event.SpacecraftState:
reportfile.write(', ' + str(state))
if PlotOptions.PlotR:
reportfile.write(', ' + str(math.sqrt(event.SpacecraftState[0]**2 + event.SpacecraftState[1]**2 + event.SpacecraftState[2]**2) / self.LU))
if PlotOptions.PlotV:
reportfile.write(', ' + str(math.sqrt(event.SpacecraftState[3]**2 + event.SpacecraftState[4]**2 + event.SpacecraftState[5]**2) / self.LU * self.TU))
if PlotOptions.PlotThrust:
if event.EventType not in ['match_point','upwr_flyby','pwr_flyby','LT_rndzvs','LT_spiral']:
reportfile.write(', ' + str(math.sqrt(event.Thrust[0]**2 + event.Thrust[1]**2 + event.Thrust[2]**2)))
elif event.EventType == 'LT_spiral':
reportfile.write(', ' + str(event.AvailableThrust * self.thruster_duty_cycle))
else:
reportfile.write(', 0.0')
if PlotOptions.PlotIsp:
reportfile.write(', ' + str(event.Isp))
if PlotOptions.PlotMdot:
reportfile.write(', ' + str(event.MassFlowRate))
if PlotOptions.PlotEfficiency:
if event.EventType in ['SFthrust', 'SSFthrust','FBLTSthrust','FBLTthrust','PSBIthrust','PSFBthrust','LT_spiral']:
reportfile.write(', ' + str(event.AvailableThrust * event.Isp * 9.80665 / (2000 * event.ActivePower)))
else:
reportfile.write(', 0.0')
if PlotOptions.PlotThrottle:
if event.EventType in ['SFthrust', 'SSFthrust','FBLTSthrust','FBLTthrust','PSBIthrust','PSFBthrust']:
reportfile.write(', ' + str(math.sqrt(event.Thrust[0]**2 + event.Thrust[1]**2 + event.Thrust[2]**2) / (event.AvailableThrust * self.thruster_duty_cycle)))
elif event.EventType == 'LT_spiral':
reportfile.write(', ' + str(1.0))
elif event.EventType not in ['match_point','upwr_flyby','pwr_flyby','LT_rndzvs']:
reportfile.write(', ' + str(0.0))
if PlotOptions.PlotPower:
if event.EventType not in ['match_point','upwr_flyby','pwr_flyby','LT_rndzvs']:
if (event.AvailablePower > 0.0):
reportfile.write(', ' + str(event.AvailablePower))
else:
reportfile.write(', ' + str(0.0))
else:
reportfile.write(', ' + str(0.0))
if PlotOptions.PlotGamma:
if event.EventType in ['SFthrust', 'SSFthrust','FBLTSthrust','FBLTthrust','PSBIthrust','PSFBthrust','LT_spiral']:
reportfile.write(', ' + str(math.atan2(event.Thrust[1], event.Thrust[0]) * 180.0 / math.pi))
else:
reportfile.write(', ' + str(0.0))
if PlotOptions.PlotDelta:
if event.EventType in ['SFthrust', 'SSFthrust','FBLTSthrust','FBLTthrust','PSBIthrust','PSFBthrust','LT_spiral']:
AppliedThrust = math.sqrt(event.Thrust[0]**2 + event.Thrust[1]**2 + event.Thrust[2]**2)
reportfile.write(', ' + str(math.asin(event.Thrust[2] / AppliedThrust) * 180.0 / math.pi))
else:
reportfile.write(', ' + str(0.0))
if PlotOptions.PlotArray_Thrust_Angle:
if event.EventType in ['SFthrust', 'SSFthrust','FBLTSthrust','FBLTthrust','PSBIthrust','PSFBthrust','LT_spiral']:
r = math.sqrt(event.SpacecraftState[0]**2 + event.SpacecraftState[1]**2 + event.SpacecraftState[2]**2)
AppliedThrust = math.sqrt(event.Thrust[0]**2 + event.Thrust[1]**2 + event.Thrust[2]**2)
rdotT = event.SpacecraftState[0]*event.Thrust[0] + event.SpacecraftState[1]*event.Thrust[1] + event.SpacecraftState[2]*event.Thrust[2]
reportfile.write(', ' + str(math.acos( rdotT / (r * AppliedThrust) ) * 180.0 / math.pi))
else:
reportfile.write(', ' + str(0.0))
if PlotOptions.PlotMass:
reportfile.write(', ' + str(event.Mass))
if PlotOptions.PlotNumberOfEngines:
if event.EventType in ['SFthrust', 'SSFthrust','FBLTSthrust','FBLTthrust','PSBIthrust','PSFBthrust','LT_spiral']:
reportfile.write(', ' + str(event.Number_of_Active_Engines))
else:
reportfile.write(', ' + str(0))
if PlotOptions.PlotActivePower:
if event.EventType in ['SFthrust', 'SSFthrust','FBLTSthrust','FBLTthrust','PSBIthrust','PSFBthrust','LT_spiral']:
reportfile.write(', ' + str(event.ActivePower))
else:
reportfile.write(', ' + str(0.0))
if PlotOptions.PlotWasteHeat:
if event.EventType in ['SFthrust', 'SSFthrust','FBLTSthrust','FBLTthrust','PSBIthrust','PSFBthrust','LT_spiral']:
reportfile.write(', ' + str((1.0 - event.AvailableThrust * event.Isp * 9.80665 / (2000.0 * event.ActivePower)) * event.ActivePower))
else:
reportfile.write(', ' + str(0.0))
if PlotOptions.PlotEarthDistance or PlotOptions.PlotSunSpacecraftEarthAngle or PlotOptions.PlotSpacecraftViewingAngle:
if not 'sun' in self.central_body.lower():
print('getting Earth distance and Sun-Spacecraft-Earth angle is only supported if the central body is the sun')
else:
#get the position of the central body relative to the solar system barycenter
cb_relative_to_solar_system_barycenter_in_ICRF = self.eph.position('sun', event.JulianDate)
#get the position of the Earth relative to the central body
embarycenter_in_ICRF = self.eph.position('earthmoon', event.JulianDate)
moonvector_in_ICRF = self.eph.position('moon', event.JulianDate)
Earth_relative_to_solar_system_barycenter_in_ICRF = embarycenter_in_ICRF - moonvector_in_ICRF * self.eph.earth_share
Earth_relative_to_cb_in_ICRF = Earth_relative_to_solar_system_barycenter_in_ICRF + cb_relative_to_solar_system_barycenter_in_ICRF
#rotate the spacecraft state relative to the central body into the ICRF frame
spacecraft_state_relative_to_central_body_in_ICRF = AstroFunctions.rotate_from_ecliptic_to_equatorial3(np.array(event.SpacecraftState[0:3]))
#get the vector from the Earth to the Spacecraft
Earth_Spacecraft_Vector = spacecraft_state_relative_to_central_body_in_ICRF - np.transpose(Earth_relative_to_cb_in_ICRF)[0]
#return the magnitude of the Earth-centered position vector
EarthDistance = np.linalg.norm(Earth_Spacecraft_Vector) / self.LU
if PlotOptions.PlotEarthDistance:
reportfile.write(', ' + str(EarthDistance))
if PlotOptions.PlotSunSpacecraftEarthAngle:
cosAngle = np.dot(-Earth_Spacecraft_Vector, -spacecraft_state_relative_to_central_body_in_ICRF)/np.linalg.norm(spacecraft_state_relative_to_central_body_in_ICRF)/np.linalg.norm(Earth_Spacecraft_Vector)
reportfile.write(', ' + str(np.arccos(cosAngle) * 180.0/math.pi))
if PlotOptions.PlotSpacecraftViewingAngle:
tanAngle = Earth_Spacecraft_Vector[2]/math.sqrt(Earth_Spacecraft_Vector[0]*Earth_Spacecraft_Vector[0]+Earth_Spacecraft_Vector[1]*Earth_Spacecraft_Vector[1])
reportfile.write(', ' + str(np.arctan(tanAngle) * 180.0/math.pi))
if PlotOptions.PlotThrottleLevel:
if event.EventType in ['SFthrust', 'SSFthrust', 'FBLTSthrust', 'FBLTthrust', 'PSBIthrust','PSFBthrust', 'end_spiral']:
if PlotOptions.throttlesetmode == 0:
reportfile.write(', ' + str(int(self.high_Mdot_set.find_nearest_throttle_setting_1D(event.ActivePower / event.Number_of_Active_Engines).TL.strip('TL'))))
elif PlotOptions.throttlesetmode == 1:
reportfile.write(', ' + str(int(self.low_Mdot_set.find_nearest_throttle_setting_1D(event.ActivePower / event.Number_of_Active_Engines).TL.strip('TL'))))
elif PlotOptions.throttlesetmode == 2:
print('2D throttle matching not currently implemented')
else:
reportfile.write(', ' + str(0.0))
#plot sun-boresight angle
if PlotOptions.PlotSunBoresightAngle:
if event.EventType in ['SFthrust', 'SSFthrust', 'FBLTSthrust', 'FBLTthrust', "PSBIthrust",'PSFBthrust']:
r = math.sqrt(event.SpacecraftState[0]**2 + event.SpacecraftState[1]**2 + event.SpacecraftState[2]**2)
AppliedThrust = math.sqrt(event.Thrust[0]**2 + event.Thrust[1]**2 + event.Thrust[2]**2)
rdotT = event.SpacecraftState[0]*event.Thrust[0] + event.SpacecraftState[1]*event.Thrust[1] + event.SpacecraftState[2]*event.Thrust[2]
reportfile.write(', ' + str(math.acos( rdotT / (r * AppliedThrust) ) * 180.0 / math.pi))
else:
reportfile.write(', ' + str(0.0))
reportfile.write('\n')
mu = 1.0
r_dot /= k
r_now = r_last + r_dot * d_tau + 0.5 * accel(r_last, mu) * (d_tau ** 2)
tau += d_tau
while abs(tau) <= abs(end_tau):
tau += d_tau
temp = verlet_position(r_now, r_last, d_tau, mu)
r_last = temp[1]
r_now = temp[0]
return r_now
AU = 149597870.7 # km/AU
epsilon = 23.43687 # obliquity of the Ecliptic
eph = Ephemeris(de421)
def location_to_angles(location):
dec = asin(location.z)
#print dec
ra = acos(location.x/cos(dec))
if location.y < 0:
ra = 2 * pi - ra
return degrees(ra), degrees(dec)
def get_ra_dec(r, julian_date):
barycenter = eph.position('earthmoon', julian_date)
moonvector = eph.position('moon', julian_date)
earth = barycenter - moonvector * eph.earth_share
R0 = vector(earth)/AU # This is the Sun to Earth center vector in Equatorial system
class Journey(object):
def __init__(self, parent):
self.parent = parent
self.missionevents = []
self.journey_name = "AJourney"
self.journey_number = 0
self.central_body = "Sun"
self.central_body_radius = 4.379e+6
self.mu = 132712440017.99
self.LU = 1.49597870691e+8
self.TU = 5022642.890912973
self.boundary_states = []
self.flyby_periapse_states = [[0.0]*6]
self.thruster_duty_cycle = 1.0
self.journey_post_arrival_deltav = 0.0
self.journey_post_arrival_propellant_used = 0.0
self.journey_initial_mass_increment = 0.0
self.journey_final_mass_increment = 0.0
#default values for entry things
self.journey_entryinterface_velocity_with_respect_to_rotating_atmosphere = 0.0
self.journey_entry_landing_interface_point = [0.0]*6
self.entry_latitude = 0.0
self.entry_longitude = 0.0
self.entry_sun_angle = 0.0
self.journey_arrival_spacecraft_sun_Earth_angle = 180.0
self.state_frame = ''
self.alpha0 = 0.0
self.delta0 = 0.0
def PlotJourney(self, JourneyAxes, PlotOptions, CustomLabels = None):
#plot each mission event
BeforeMatchPoint = True
if PlotOptions.ShowMissionEvents:
for event in self.missionevents:
#do we have a custom label?
if CustomLabels != None:
for label in CustomLabels: #this is a list of lists, [eventNumber, labelstring]
if label[0] == event.EventNumber:
BeforeMatchPoint = event.PlotEvent(JourneyAxes, self.LU, self.TU, self.mu, PlotOptions, BeforeMatchPoint, label[1])
else:
BeforeMatchPoint = event.PlotEvent(JourneyAxes, self.LU, self.TU, self.mu, PlotOptions, BeforeMatchPoint)
else:
BeforeMatchPoint = event.PlotEvent(JourneyAxes, self.LU, self.TU, self.mu, PlotOptions, BeforeMatchPoint)
def PlotJourneyBoundaryOrbits(self, JourneyAxes, PlotOptions):
from scipy.integrate import ode
for boundaryStateIndex in range(0, len(self.boundary_states)):
if not ((boundaryStateIndex == 0 and self.missionevents[0].Location == 'free point') \
or (boundaryStateIndex == (len(self.boundary_states) - 1) and self.missionevents[-1].Location == 'free point')) \
or PlotOptions.ShowFreePointBoundaryOrbits:
boundarystate = self.boundary_states[boundaryStateIndex]
BoundaryStateScaled = np.array(boundarystate) / self.LU
BoundaryStateScaled[3:6] *= self.TU
r = np.linalg.norm(BoundaryStateScaled[0:3])
v = np.linalg.norm(BoundaryStateScaled[3:6])
if r == 0.0:
continue
a = r / (2.0 - r*v*v)
if a > 0.0:
T = 2*math.pi*math.sqrt(a**3)
else:
T = 2*math.pi*self.LU / self.TU
StateIntegrateObject = ode(EOM.EOM_inertial_2body, jac=EOM.EOM_jacobian_intertial_2body).set_integrator('dop853', atol=1.0e-8, rtol=1.0e-8)
StateIntegrateObject.set_initial_value(BoundaryStateScaled).set_f_params(1.0).set_jac_params(1.0)
dt = T / 10000
StateHistory = []
epoch = (self.missionevents[0].JulianDate - 2451544.5) * 86400.0
while StateIntegrateObject.successful() and StateIntegrateObject.t < T * 1.01:
epoch += dt * self.TU
stateArray = StateIntegrateObject.integrate(StateIntegrateObject.t + dt)
state = [stateArray[0] * self.LU, stateArray[1] * self.LU, stateArray[2] * self.LU, epoch]
StateHistory.append(state)
X = []
Y = []
Z = []
for StateLine in StateHistory:
if PlotOptions.PlotCentralBody != 'Journey central body':
import spiceypy
journey_central_body = self.central_body
if journey_central_body in ['Mars','Jupiter','Uranus','Neptune','Pluto']:
journey_central_body = journey_central_body + ' Barycenter'
if journey_central_body != PlotOptions.PlotCentralBody:
body_state, LT = spiceypy.spkezr(journey_central_body, StateLine[-1], 'J2000', "None", PlotOptions.PlotCentralBody)
X.append(StateLine[0] + body_state[0])
Y.append(StateLine[1] + body_state[1])
Z.append(StateLine[2] + body_state[2])
else:
X.append(StateLine[0])
Y.append(StateLine[1])
Z.append(StateLine[2])
else:
X.append(StateLine[0])
Y.append(StateLine[1])
Z.append(StateLine[2])
JourneyAxes.plot(X, Y, Z, lw=1, c='0.75')
def PlotJourneyCentralBodyOrbits(self, JourneyAxes, PlotOptions):
if PlotOptions.PlotCentralBody != 'Journey central body':
import spiceypy
journey_central_body = self.central_body
if journey_central_body in ['Mars','Jupiter','Uranus','Neptune','Pluto']:
journey_central_body = journey_central_body + ' Barycenter'
if journey_central_body != PlotOptions.PlotCentralBody:
body_state, LT = spiceypy.spkezr(journey_central_body, (self.missionevents[0].JulianDate - 2451545.0) * 86400.0, 'J2000', "None", PlotOptions.PlotCentralBody)
r = np.linalg.norm(body_state[0:3])
v = np.linalg.norm(body_state[3:6])
mu = {'Mercury': 22031.780000,
'Venus': 324858.592000,
'Earth': 398600.435436,
'Mars': 42828.375214,
'Jupiter Barycenter': 126712764.800000,
'Saturn Barycenter': 37940585.200000,
'Uranus Barycenter': 5794548.600000,
'Neptune Barycenter': 6836527.100580,
'Pluto Barycenter': 977.000000,
'Sun': 132712440041.939400,
'Moon': 4902.800066,
'Earth-Moon Barycenter': 403503.235502}
elements = spiceypy.oscltx(body_state, (self.missionevents[0].JulianDate - 2451545.0) * 86400.0, mu[PlotOptions.PlotCentralBody])
a = elements[-2]
#a = r / (2.0 - r*v*v)
if a > 0.0:
T = 2*math.pi *math.sqrt(a**3 / mu[PlotOptions.PlotCentralBody])
else:
T = 2*math.pi * self.TU
t0 = (self.missionevents[0].JulianDate - 2451545.0) * 86400.0
tf = t0 + T
X = []
Y = []
Z = []
for epoch in np.linspace(t0, tf, num=100):
body_state, LT = spiceypy.spkezr(journey_central_body, epoch, 'J2000', "None", PlotOptions.PlotCentralBody)
X.append(body_state[0])
Y.append(body_state[1])
Z.append(body_state[2])
JourneyAxes.plot(X, Y, Z, lw=1, c='0.75')
def UpdateLabelPosition(self, Figure, Axes):
for event in self.missionevents:
event.UpdateLabelPosition(Figure, Axes)
def PlotPhaseBoundariesOnDataPlot(self, DataAxes, PlotOptions, firstpass, Ybounds):
#create vertical lines at important events
date_string_vector = []
boundarylegendflag = True
burnlegendflag = True
for event in self.missionevents:
if event.EventType in ['upwr_flyby', 'pwr_flyby', 'LT_rndzvs', 'rendezvous', 'intercept', 'insertion', 'match-vinf', 'launch', 'departure', "begin_spiral", "end_spiral"]:
event_epoch = datetime.datetime.strptime(event.GregorianDate,'%m/%d/%Y').date()
if firstpass and boundarylegendflag:
DataAxes.plot([event_epoch]*2, Ybounds, c='k', marker='+', ls = '-.', lw=3)#, label='Phase boundary')
boundarylegendflag = False
else:
DataAxes.plot([event_epoch]*2, Ybounds, c='k', marker='+', ls = '-.', lw=3)
if event.EventType == 'chem_burn':
event_epoch = datetime.datetime.strptime(event.GregorianDate,'%m/%d/%Y').date()
if firstpass and burnlegendflag:
DataAxes.plot([event_epoch]*2, Ybounds, c='r', marker='+', ls = '-.', lw=3, label='Deep-Space Maneuver')
burnlegendflag = False
else:
DataAxes.plot([event_epoch]*2, Ybounds, c='r', marker='+', ls = '-.', lw=3)
def GenerateJourneyDataPlot(self, DataAxesLeft, DataAxesRight, PlotOptions, firstpass):
import math
import astropy
#generate a vector of dates
date_string_vector = []
for event in self.missionevents:
if event.EventType != 'match_point' and event.EventType != 'upwr_flyby' and event.EventType != 'pwr_flyby' and event.EventType != 'LT_rndzvs':
date_string_vector.append(event.GregorianDate)
date_vector = [datetime.datetime.strptime(d,'%m/%d/%Y').date() for d in date_string_vector]
#dummy line across the bottom so that neither axes object crashes
DataAxesLeft.plot(date_vector, np.zeros_like(date_vector), c='w', lw = 0.1)
DataAxesRight.plot(date_vector, np.zeros_like(date_vector), c='w', lw = 0.1)
#plot distance from central body
if PlotOptions.PlotR:
Rvector = []
for event in self.missionevents:
if event.EventType != 'match_point' and event.EventType != 'upwr_flyby' and event.EventType != 'pwr_flyby' and event.EventType != 'LT_rndzvs':
Rvector.append(math.sqrt(event.SpacecraftState[0]**2 + event.SpacecraftState[1]**2 + event.SpacecraftState[2]**2) / self.LU)
if firstpass:
unitstring = 'LU'
if 'sun' in self.central_body.lower() and math.fabs(self.LU - 149597870.69100001) < 1.0:
unitstring = 'AU'
labelstring = 'Distance from ' + self.central_body + ' (' + unitstring + ')'
DataAxesLeft.plot(date_vector, Rvector, c='k', lw=2, label=labelstring)
else:
DataAxesLeft.plot(date_vector, Rvector, c='k', lw=2)
#plot velocity
if PlotOptions.PlotV:
Vvector = []
for event in self.missionevents:
if event.EventType != 'match_point' and event.EventType != 'upwr_flyby' and event.EventType != 'pwr_flyby' and event.EventType != 'LT_rndzvs':
Vvector.append(math.sqrt(event.SpacecraftState[3]**2 + event.SpacecraftState[4]**2 + event.SpacecraftState[5]**2) / self.LU * self.TU)
if firstpass:
DataAxesLeft.plot(date_vector, Vvector, c='k', lw=2, ls='-.', label='Velocity magnitude (LU/TU)')
else:
DataAxesLeft.plot(date_vector, Vvector, c='k', lw=2, ls='-.')
#plot Thrust
if PlotOptions.PlotThrust:
Thrustvector = []
shortDateVector = []
for event in self.missionevents:
epoch = astropy.time.Time(event.JulianDate, format='jd', scale='tdb')
if event.EventType in ['SFthrust', 'SSFthrust', 'FBLTSthrust', 'FBLTthrust', 'PSBIthrust', 'PSFBthrust']:
Thrustvector.append(math.sqrt(event.Thrust[0]**2 + event.Thrust[1]**2 + event.Thrust[2]**2) * 10.0)
shortDateVector.append((epoch - event.TimestepLength / 2).datetime)
Thrustvector.append(math.sqrt(event.Thrust[0]**2 + event.Thrust[1]**2 + event.Thrust[2]**2) * 10.0)
shortDateVector.append((epoch + event.TimestepLength / 2).datetime)
elif event.EventType == 'LT_spiral':
Thrustvector.append(event.AvailableThrust * self.thruster_duty_cycle*10.0)
shortDateVector.append((epoch - event.TimestepLength / 2).datetime)
Thrustvector.append(event.AvailableThrust * self.thruster_duty_cycle*10.0)
shortDateVector.append((epoch + event.TimestepLength / 2).datetime)
elif event.EventType != 'match_point':
Thrustvector.append(0.0)
shortDateVector.append((epoch - event.TimestepLength / 2).datetime)
Thrustvector.append(0.0)
shortDateVector.append((epoch + event.TimestepLength / 2).datetime)
if firstpass:
DataAxesLeft.plot(shortDateVector, Thrustvector, c='r', lw=2, ls='--', label='Applied thrust (0.1 N)')
else:
DataAxesLeft.plot(shortDateVector, Thrustvector, c='r', lw=2, ls='--')
#plot Isp
if PlotOptions.PlotIsp:
Ispvector = []
shortDateVector = []
for event in self.missionevents:
epoch = astropy.time.Time(event.JulianDate, format='jd', scale='tdb')
if event.EventType in ['SFthrust', 'SSFthrust', 'FBLTSthrust', 'FBLTthrust', 'PSBIthrust', 'PSFBthrust', 'LT_spiral']:
Ispvector.append(event.Isp / 1000.0)
shortDateVector.append((epoch - event.TimestepLength / 2).datetime)
Ispvector.append(event.Isp / 1000.0)
shortDateVector.append((epoch + event.TimestepLength / 2).datetime)
elif event.EventType != 'match_point':
Ispvector.append(0.0)
shortDateVector.append((epoch - event.TimestepLength / 2).datetime)
Ispvector.append(0.0)
shortDateVector.append((epoch + event.TimestepLength / 2).datetime)
if firstpass:
DataAxesLeft.plot(shortDateVector, Ispvector, 'dodgerblue', lw=4, ls='-.', label='Isp (1000 s)')
else:
DataAxesLeft.plot(shortDateVector, Ispvector, 'dodgerblue', lw=4, ls='-.')
#plot mass flow rate
if PlotOptions.PlotMdot:
Mdotvector = []
shortDateVector = []
for event in self.missionevents:
epoch = astropy.time.Time(event.JulianDate, format='jd', scale='tdb')
if event.EventType in ['SFthrust', 'SSFthrust', 'FBLTSthrust', 'FBLTthrust', 'PSBIthrust', 'PSFBthrust', 'LT_spiral']:
Mdotvector.append(event.MassFlowRate * 1.0e6)
shortDateVector.append((epoch - event.TimestepLength / 2).datetime)
Mdotvector.append(event.MassFlowRate * 1.0e6)
shortDateVector.append((epoch + event.TimestepLength / 2).datetime)
elif event.EventType != 'match_point':
Mdotvector.append(0.0)
shortDateVector.append((epoch - event.TimestepLength / 2).datetime)
Mdotvector.append(0.0)
shortDateVector.append((epoch + event.TimestepLength / 2).datetime)
if firstpass:
DataAxesLeft.plot(shortDateVector, Mdotvector, c='brown', lw=2, ls='-', label='Mass flow rate (mg/s)')
else:
DataAxesLeft.plot(shortDateVector, Mdotvector, c='brown', lw=2, ls='-')
#plot Efficiency
if PlotOptions.PlotEfficiency:
Efficiencyvector = []
shortDateVector = []
for event in self.missionevents:
epoch = astropy.time.Time(event.JulianDate, format='jd', scale='tdb')
if event.EventType in ['SFthrust', 'SSFthrust', 'FBLTSthrust', 'FBLTthrust', 'PSBIthrust', 'PSFBthrust', 'LT_spiral']:
Efficiencyvector.append(event.AvailableThrust * event.Isp * 9.80665 / (2000 * event.ActivePower))
shortDateVector.append((epoch - event.TimestepLength / 2).datetime)
Efficiencyvector.append(event.AvailableThrust * event.Isp * 9.80665 / (2000 * event.ActivePower))
shortDateVector.append((epoch + event.TimestepLength / 2).datetime)
elif event.EventType != 'match_point':
Efficiencyvector.append(0.0)
shortDateVector.append((epoch - event.TimestepLength / 2).datetime)
Efficiencyvector.append(0.0)
shortDateVector.append((epoch + event.TimestepLength / 2).datetime)
if firstpass:
DataAxesLeft.plot(shortDateVector, Efficiencyvector, c='DarkGreen', lw=2, ls='-', label='Propulsion system efficiency')
else:
DataAxesLeft.plot(shortDateVector, Efficiencyvector, c='DarkGreen', lw=2, ls='-')
#plot control magnitude
if PlotOptions.PlotThrottle:
Throttlevector = []
shortDateVector = []
for event in self.missionevents:
epoch = astropy.time.Time(event.JulianDate, format='jd', scale='tdb')
if event.EventType in ['SFthrust', 'SSFthrust', 'PSBIthrust']:
Throttlevector.append(math.sqrt(event.Thrust[0]**2 + event.Thrust[1]**2 + event.Thrust[2]**2) / (event.AvailableThrust))
shortDateVector.append((epoch - event.TimestepLength / 2).datetime)
Throttlevector.append(math.sqrt(event.Thrust[0]**2 + event.Thrust[1]**2 + event.Thrust[2]**2) / (event.AvailableThrust))
shortDateVector.append((epoch + event.TimestepLength / 2).datetime)
elif event.EventType in ['FBLTthrust','FBLTSthrust','PSFBthrust']:
Throttlevector.append(event.DVmagorThrottle)
shortDateVector.append((epoch - event.TimestepLength / 2).datetime)
Throttlevector.append(event.DVmagorThrottle)
shortDateVector.append((epoch + event.TimestepLength / 2).datetime)
elif event.EventType == 'LT_spiral':
Throttlevector.append(1.0)
shortDateVector.append((epoch - event.TimestepLength / 2).datetime)
Throttlevector.append(1.0)
shortDateVector.append((epoch + event.TimestepLength / 2).datetime)
elif event.EventType != 'match_point':
Throttlevector.append(0.0)
shortDateVector.append((epoch - event.TimestepLength / 2).datetime)
Throttlevector.append(0.0)
shortDateVector.append((epoch + event.TimestepLength / 2).datetime)
if firstpass:
DataAxesLeft.plot(shortDateVector, Throttlevector, c='r', lw=2, ls='--', label='Control magnitude')
else:
DataAxesLeft.plot(shortDateVector, Throttlevector, c='r', lw=2, ls='--')
#plot power
if PlotOptions.PlotPower:
Powervector = []
for event in self.missionevents:
if event.EventType != 'match_point' and event.EventType != 'upwr_flyby' and event.EventType != 'pwr_flyby' and event.EventType != 'LT_rndzvs':
if event.AvailablePower > 0.0:
Powervector.append(event.AvailablePower)
else:
Powervector.append(0.0)
if firstpass:
DataAxesLeft.plot(date_vector, Powervector, c='Navy', lw=2, ls='-.', label='Power available for propulsion (kW)')
else:
DataAxesLeft.plot(date_vector, Powervector, c='Navy', lw=2, ls='-.')
#plot gamma
if PlotOptions.PlotGamma:
gammavector = []
shortDateVector = []
for event in self.missionevents:
if event.EventType in ['SFthrust', 'SSFthrust', 'FBLTSthrust', 'FBLTthrust', 'PSBIthrust', 'PSFBthrust', 'LT_spiral']:
gammavector.append(math.atan2(event.Thrust[1], event.Thrust[0]) * 180.0 / math.pi)
shortDateVector.append(datetime.datetime.strptime(event.GregorianDate,'%m/%d/%Y').date())
if firstpass:
DataAxesRight.plot(shortDateVector, gammavector, c='DarkGreen', lw=2, ls='--', label=r'$\gamma$ (degrees)')
else:
DataAxesRight.plot(shortDateVector, gammavector, c='DarkGreen', lw=2, ls='--')
#plot delta
if PlotOptions.PlotDelta:
deltavector = []
shortDateVector = []
for event in self.missionevents:
if event.EventType in ['SFthrust', 'SSFthrust', 'FBLTSthrust', 'FBLTthrust', 'PSBIthrust', 'PSFBthrust', 'LT_spiral']:
AppliedThrust = math.sqrt(event.Thrust[0]**2 + event.Thrust[1]**2 + event.Thrust[2]**2)
deltavector.append(math.asin(event.Thrust[2] / AppliedThrust) * 180.0 / math.pi)
shortDateVector.append(datetime.datetime.strptime(event.GregorianDate,'%m/%d/%Y').date())
if firstpass:
DataAxesRight.plot(shortDateVector, deltavector, c='LightGreen', lw=2, ls='--', label=r'$\delta$ (degrees)')
else:
DataAxesRight.plot(shortDateVector, deltavector, c='LightGreen', lw=2, ls='--')
#plot central body to thrust vector angle
if PlotOptions.PlotArray_Thrust_Angle:
CBthrustvector = []
shortDateVector = []
for event in self.missionevents:
if event.EventType in ['SFthrust', 'SSFthrust', 'FBLTSthrust', 'FBLTthrust', 'PSBIthrust', 'PSFBthrust', 'LT_spiral']:
r = math.sqrt(event.SpacecraftState[0]**2 + event.SpacecraftState[1]**2 + event.SpacecraftState[2]**2)
AppliedThrust = math.sqrt(event.Thrust[0]**2 + event.Thrust[1]**2 + event.Thrust[2]**2)
rdotT = event.SpacecraftState[0]*event.Thrust[0] + event.SpacecraftState[1]*event.Thrust[1] + event.SpacecraftState[2]*event.Thrust[2]
CBthrustvector.append( 90.0 - math.acos( rdotT / (r * AppliedThrust) ) * 180.0 / math.pi)
shortDateVector.append(datetime.datetime.strptime(event.GregorianDate,'%m/%d/%Y').date())
if firstpass:
DataAxesRight.plot(shortDateVector, CBthrustvector, c='Salmon', marker='o', ls='None', label='Array to thrust angle (degrees)')
else:
DataAxesRight.plot(shortDateVector, CBthrustvector, c='Salmon', marker='o', ls='None')
#plot mass
if PlotOptions.PlotMass:
mass = []
for event in self.missionevents:
if event.EventType != 'match_point' and event.EventType != 'upwr_flyby' and event.EventType != 'pwr_flyby' and event.EventType != 'LT_rndzvs':
mass.append(event.Mass * 1.0e-3)
if firstpass:
DataAxesLeft.plot(date_vector, mass, c='DarkGrey', lw=2, ls='-', label='Mass (1000 kg)')
else:
DataAxesLeft.plot(date_vector, mass, c='DarkGrey', lw=2, ls='-')
#plot number of active thrusters
if PlotOptions.PlotNumberOfEngines:
numberofengines = []
shortDateVector = []
for event in self.missionevents:
if event.EventType in ['SFthrust', 'SSFthrust', 'FBLTSthrust', 'FBLTthrust', 'PSBIthrust', 'PSFBthrust', 'LT_spiral']:
numberofengines.append(event.ActivePower)
shortDateVector.append(datetime.datetime.strptime(event.GregorianDate,'%m/%d/%Y').date())
if firstpass:
DataAxesLeft.plot(shortDateVector, numberofengines, 'k*', mew=2, markersize=7, label='Number of active thrusters')
else:
DataAxesLeft.plot(shortDateVector, numberofengines, 'k*', mew=2, markersize=7)
#plot power actively used by the thrusters
if PlotOptions.PlotActivePower:
activepowervector = []
shortDateVector = []
for event in self.missionevents:
if event.EventType in ['SFthrust', 'SSFthrust', 'FBLTSthrust', 'FBLTthrust', 'PSBIthrust', 'PSFBthrust', 'LT_spiral']:
activepowervector.append(event.ActivePower)
shortDateVector.append(datetime.datetime.strptime(event.GregorianDate,'%m/%d/%Y').date())
if firstpass:
DataAxesLeft.plot(shortDateVector, activepowervector, c='Navy', lw=2, ls='-', label='Power used by the propulsion system (kW)')
else:
DataAxesLeft.plot(shortDateVector, activepowervector, c='Navy', lw=2, ls='-')
#plot waste heat from the propulsion system
if PlotOptions.PlotWasteHeat:
WasteHeatvector = []
shortDateVector = []
for event in self.missionevents:
if event.EventType in ['SFthrust', 'SSFthrust', 'FBLTSthrust', 'FBLTthrust', 'PSBIthrust', 'PSFBthrust', 'LT_spiral']:
WasteHeatvector.append( (1 - event.AvailableThrust * event.Isp * 9.80665 / (2000 * event.ActivePower)) * event.ActivePower )
shortDateVector.append(datetime.datetime.strptime(event.GregorianDate,'%m/%d/%Y').date())
if firstpass:
DataAxesLeft.plot(shortDateVector, WasteHeatvector, c='Crimson', lw=2, ls='--', label='Waste heat from propulsion system (kW)')
else:
DataAxesLeft.plot(shortDateVector, WasteHeatvector, c='Crimson', lw=2, ls='--')
#plot Earth distance in LU and SPE angle
if PlotOptions.PlotEarthDistance or PlotOptions.PlotSunSpacecraftEarthAngle or PlotOptions.PlotSpacecraftViewingAngle:
if not 'sun' in self.central_body.lower():
print('getting Earth distance and Sun-Spacecraft-Earth angle is only supported if the central body is the sun')
else:
import de423
from jplephem import Ephemeris
eph = Ephemeris(de423)
EarthDistanceVector = []
SunSpacecraftEarthAngle = []
SpacecraftViewingAngle = []
for event in self.missionevents:
if event.EventType != 'match_point' and event.EventType != 'upwr_flyby' and event.EventType != 'pwr_flyby' and event.EventType != 'LT_rndzvs':
#get the position of the central body relative to the solar system barycenter
cb_relative_to_solar_system_barycenter_in_ICRF = eph.position('sun', event.JulianDate)
#get the position of the Earth relative to the central body
embarycenter_in_ICRF = eph.position('earthmoon', event.JulianDate)
moonvector_in_ICRF = eph.position('moon', event.JulianDate)
Earth_relative_to_solar_system_barycenter_in_ICRF = embarycenter_in_ICRF - moonvector_in_ICRF * eph.earth_share
Earth_relative_to_cb_in_ICRF = Earth_relative_to_solar_system_barycenter_in_ICRF + cb_relative_to_solar_system_barycenter_in_ICRF
#rotate the spacecraft state relative to the central body into the ICRF frame
spacecraft_state_relative_to_central_body_in_ICRF = AstroFunctions.rotate_from_ecliptic_to_equatorial3(np.array(event.SpacecraftState[0:3]))
#get the vector from the Earth to the Spacecraft
Earth_Spacecraft_Vector = spacecraft_state_relative_to_central_body_in_ICRF - np.transpose(Earth_relative_to_cb_in_ICRF)[0]
#return the magnitude of the Earth-centered position vector
EarthDistanceVector.append(np.linalg.norm(Earth_Spacecraft_Vector) / self.LU)
#if applicable, compute sun-spacecraft-Earth angle
if PlotOptions.PlotSunSpacecraftEarthAngle:
cosAngle = np.dot(-Earth_Spacecraft_Vector, -spacecraft_state_relative_to_central_body_in_ICRF)/np.linalg.norm(spacecraft_state_relative_to_central_body_in_ICRF)/np.linalg.norm(Earth_Spacecraft_Vector)
SunSpacecraftEarthAngle.append(np.arccos(cosAngle) * 180.0/math.pi)
if PlotOptions.PlotSpacecraftViewingAngle:
tanAngle = Earth_Spacecraft_Vector[2]/math.sqrt(Earth_Spacecraft_Vector[0]*Earth_Spacecraft_Vector[0]+Earth_Spacecraft_Vector[1]*Earth_Spacecraft_Vector[1])
SpacecraftViewingAngle.append(np.arctan(tanAngle) * 180.0/math.pi)
if PlotOptions.PlotEarthDistance:
if firstpass:
unitstring = 'LU'
if 'sun' in self.central_body.lower() and math.fabs(self.LU - 149597870.69100001) < 1.0:
unitstring = 'AU'
labelstring = 'Distance from Earth (' + unitstring + ')'
DataAxesLeft.plot(date_vector, EarthDistanceVector, c='g', lw=2, label=labelstring)
else:
DataAxesLeft.plot(date_vector, EarthDistanceVector, c='g', lw=2)
if PlotOptions.PlotSunSpacecraftEarthAngle:
if firstpass:
DataAxesRight.plot(date_vector, SunSpacecraftEarthAngle, c='orangered', lw=2, ls = '-.', label='Sun-Spacecraft-Earth Angle')
else:
DataAxesRight.plot(date_vector, SunSpacecraftEarthAngle, c='orangered', lw=2, ls = '-.')
if PlotOptions.PlotSpacecraftViewingAngle:
if firstpass:
DataAxesRight.plot(date_vector, SpacecraftViewingAngle, c='orangered', lw=2, ls = '-.', label='Latitude of Earth-Spacecraft Vector')
else:
DataAxesRight.plot(date_vector, SpacecraftViewingAngle, c='orangered', lw=2, ls = '-.')
#plot throttle level
if PlotOptions.PlotThrottleLevel:
ThrottleLevelVector = []
shortDateVector = []
for event in self.missionevents:
if event.EventType in ['SFthrust', 'SSFthrust', 'FBLTSthrust', 'FBLTthrust', 'PSBIthrust', 'PSFBthrust', 'LT_spiral']:
ThrottleLevelVector.append(event.ThrottleLevel)
shortDateVector.append(datetime.datetime.strptime(event.GregorianDate,'%m/%d/%Y').date())
if firstpass:
DataAxesLeft.scatter(shortDateVector, ThrottleLevelVector, c='midnightblue', marker='h' , label='Throttle level')
else:
DataAxesLeft.scatter(shortDateVector, ThrottleLevelVector, c='midnightblue', marker='h' )
#plot sun-boresight angle
if PlotOptions.PlotSunBoresightAngle:
SunBoresightAngle = []
for event in self.missionevents:
if event.EventType in ['SFthrust', 'SSFthrust', 'FBLTSthrust', 'FBLTthrust', 'PSFBthrust', "PSBIthrust"]:
r = math.sqrt(event.SpacecraftState[0]**2 + event.SpacecraftState[1]**2 + event.SpacecraftState[2]**2)
AppliedThrust = math.sqrt(event.Thrust[0]**2 + event.Thrust[1]**2 + event.Thrust[2]**2)
rdotT = -event.SpacecraftState[0]*event.Thrust[0] - event.SpacecraftState[1]*event.Thrust[1] - event.SpacecraftState[2]*event.Thrust[2]
SunBoresightAngle.append(math.acos( rdotT / (r * AppliedThrust) ) * 180.0 / math.pi)
elif event.EventType != 'match_point' and event.EventType != 'upwr_flyby' and event.EventType != 'pwr_flyby' and event.EventType != 'LT_rndzvs':
SunBoresightAngle.append(0.0)
if firstpass:
DataAxesRight.plot(date_vector, SunBoresightAngle, c='purple', marker='o', ls='None', label='Sun-to-Boresight angle (degrees)')
else:
DataAxesRight.plot(date_vector, SunBoresightAngle, c='purple', marker='o', ls='None')
#DataAxesLeft.set_ylim(bottom=-0.7)
#DataAxesRight.set_ylim(bottom=-0.7)
#DataAxesLeft.set_ylim([-0.2,2.5])
#DataAxesLeft.set_ylim([-0.2,2.5])
def WriteDateReport(self, PlotOptions, reportfile):
#write header
reportfile.write('JD, MM-DD-YYYY, Event duration (days), Event type, Event location')
if PlotOptions.IncludeStateVectorInReport:
reportfile.write(',x (km), y (km), z (km), xdot (km/s), ydot (km/s), zdot (km/s)')
if PlotOptions.PlotR:
unitstring = 'LU'
if 'sun' in self.central_body.lower() and math.fabs(self.LU - 149597870.69100001) < 1.0:
unitstring = 'AU'
reportfile.write(', Distance from ' + self.central_body + ' (' + unitstring + ')')
if PlotOptions.PlotV:
reportfile.write(', Velocity magnitude (LU/TU)')
if PlotOptions.PlotThrust:
reportfile.write(', Applied thrust (N)')
if PlotOptions.PlotIsp:
reportfile.write(', Isp (s)')
if PlotOptions.PlotMdot:
reportfile.write(', Mass flow rate (kg/s)')
if PlotOptions.PlotEfficiency:
reportfile.write(', Propulsion system efficiency')
if PlotOptions.PlotThrottle:
reportfile.write(', Control magnitude')
if PlotOptions.PlotPower:
reportfile.write(', Power produced by spacecraft (kW)')
if PlotOptions.PlotGamma:
reportfile.write(', in-plane control angle (degrees)')
if PlotOptions.PlotDelta:
reportfile.write(', out-of-plane control angle (degrees)')
if PlotOptions.PlotArray_Thrust_Angle:
reportfile.write(', Array to thrust angle (degrees)')
if PlotOptions.PlotMass:
reportfile.write(', Mass (kg)')
if PlotOptions.PlotNumberOfEngines:
reportfile.write(', Number of active thrusters')
if PlotOptions.PlotActivePower:
reportfile.write(', Power used by the propulsion system (kW)')
if PlotOptions.PlotWasteHeat:
reportfile.write(', Waste heat from propulsion system (kW)')
if PlotOptions.PlotEarthDistance:
unitstring = 'LU'
if 'sun' in self.central_body.lower() and math.fabs(self.LU - 149597870.69100001) < 1.0:
unitstring = 'AU'
reportfile.write(', Distance from Earth (' + unitstring + ')')
if PlotOptions.PlotSunSpacecraftEarthAngle:
reportfile.write(', Sun-Spacecraft-Earth Angle')
if PlotOptions.PlotSpacecraftViewingAngle:
reportfile.write(', Latitude of Earth-Spacecraft Vector')
if PlotOptions.PlotThrottleLevel:
reportfile.write(', Throttle level')
if PlotOptions.PlotSunBoresightAngle:
reportfile.write(', Sun-Boresight Angle')
reportfile.write('\n')
#load stuff if necessary
if PlotOptions.PlotEarthDistance or PlotOptions.PlotSunSpacecraftEarthAngle or PlotOptions.PlotSpacecraftViewingAngle:
if not 'sun' in self.central_body.lower():
print('getting Earth distance and Sun-Spacecraft-Earth angle is only supported if the central body is the sun')
else:
import de423
from jplephem import Ephemeris
self.eph = Ephemeris(de423)
if PlotOptions.PlotThrottleLevel:
#instantiate a throttle table object
thruster_throttle_table = ThrottleTable.ThrottleTable(PlotOptions.throttletablefile)
self.high_Mdot_set, self.low_Mdot_set = thruster_throttle_table.create_non_dominated_sets()
#now print the report
for event in self.missionevents:
reportfile.write(str(event.JulianDate))
reportfile.write(', ' + event.GregorianDate)
reportfile.write(', ' + str(event.TimestepLength))
reportfile.write(', ' + event.EventType)
reportfile.write(', ' + event.Location)
if PlotOptions.IncludeStateVectorInReport:
for state in event.SpacecraftState:
reportfile.write(', ' + str(state))
if PlotOptions.PlotR:
reportfile.write(', ' + str(math.sqrt(event.SpacecraftState[0]**2 + event.SpacecraftState[1]**2 + event.SpacecraftState[2]**2) / self.LU))
if PlotOptions.PlotV:
reportfile.write(', ' + str(math.sqrt(event.SpacecraftState[3]**2 + event.SpacecraftState[4]**2 + event.SpacecraftState[5]**2) / self.LU * self.TU))
if PlotOptions.PlotThrust:
if event.EventType not in ['match_point','upwr_flyby','pwr_flyby','LT_rndzvs','LT_spiral']:
reportfile.write(', ' + str(math.sqrt(event.Thrust[0]**2 + event.Thrust[1]**2 + event.Thrust[2]**2)))
elif event.EventType == 'LT_spiral':
reportfile.write(', ' + str(event.AvailableThrust * self.thruster_duty_cycle))
else:
reportfile.write(', 0.0')
if PlotOptions.PlotIsp:
reportfile.write(', ' + str(event.Isp))
if PlotOptions.PlotMdot:
reportfile.write(', ' + str(event.MassFlowRate))
if PlotOptions.PlotEfficiency:
if event.EventType in ['SFthrust', 'SSFthrust','FBLTSthrust','FBLTthrust','PSBIthrust','PSFBthrust','LT_spiral']:
reportfile.write(', ' + str(event.AvailableThrust * event.Isp * 9.80665 / (2000 * event.ActivePower)))
else:
reportfile.write(', 0.0')
if PlotOptions.PlotThrottle:
if event.EventType in ['SFthrust', 'SSFthrust','FBLTSthrust','FBLTthrust','PSBIthrust','PSFBthrust']:
reportfile.write(', ' + str(math.sqrt(event.Thrust[0]**2 + event.Thrust[1]**2 + event.Thrust[2]**2) / (event.AvailableThrust * self.thruster_duty_cycle)))
elif event.EventType == 'LT_spiral':
reportfile.write(', ' + str(1.0))
elif event.EventType not in ['match_point','upwr_flyby','pwr_flyby','LT_rndzvs']:
reportfile.write(', ' + str(0.0))
if PlotOptions.PlotPower:
if event.EventType not in ['match_point','upwr_flyby','pwr_flyby','LT_rndzvs']:
if (event.AvailablePower > 0.0):
reportfile.write(', ' + str(event.AvailablePower))
else:
reportfile.write(', ' + str(0.0))
else:
reportfile.write(', ' + str(0.0))
if PlotOptions.PlotGamma:
if event.EventType in ['SFthrust', 'SSFthrust','FBLTSthrust','FBLTthrust','PSBIthrust','PSFBthrust','LT_spiral']:
reportfile.write(', ' + str(math.atan2(event.Thrust[1], event.Thrust[0]) * 180.0 / math.pi))
else:
reportfile.write(', ' + str(0.0))
if PlotOptions.PlotDelta:
if event.EventType in ['SFthrust', 'SSFthrust','FBLTSthrust','FBLTthrust','PSBIthrust','PSFBthrust','LT_spiral']:
AppliedThrust = math.sqrt(event.Thrust[0]**2 + event.Thrust[1]**2 + event.Thrust[2]**2)
reportfile.write(', ' + str(math.asin(event.Thrust[2] / AppliedThrust) * 180.0 / math.pi))
else:
reportfile.write(', ' + str(0.0))
if PlotOptions.PlotArray_Thrust_Angle:
if event.EventType in ['SFthrust', 'SSFthrust','FBLTSthrust','FBLTthrust','PSBIthrust','PSFBthrust','LT_spiral']:
r = math.sqrt(event.SpacecraftState[0]**2 + event.SpacecraftState[1]**2 + event.SpacecraftState[2]**2)
AppliedThrust = math.sqrt(event.Thrust[0]**2 + event.Thrust[1]**2 + event.Thrust[2]**2)
rdotT = event.SpacecraftState[0]*event.Thrust[0] + event.SpacecraftState[1]*event.Thrust[1] + event.SpacecraftState[2]*event.Thrust[2]
reportfile.write(', ' + str(math.acos( rdotT / (r * AppliedThrust) ) * 180.0 / math.pi))
else:
reportfile.write(', ' + str(0.0))
if PlotOptions.PlotMass:
reportfile.write(', ' + str(event.Mass))
if PlotOptions.PlotNumberOfEngines:
if event.EventType in ['SFthrust', 'SSFthrust','FBLTSthrust','FBLTthrust','PSBIthrust','PSFBthrust','LT_spiral']:
reportfile.write(', ' + str(event.Number_of_Active_Engines))
else:
reportfile.write(', ' + str(0))
if PlotOptions.PlotActivePower:
if event.EventType in ['SFthrust', 'SSFthrust','FBLTSthrust','FBLTthrust','PSBIthrust','PSFBthrust','LT_spiral']:
reportfile.write(', ' + str(event.ActivePower))
else:
reportfile.write(', ' + str(0.0))
if PlotOptions.PlotWasteHeat:
if event.EventType in ['SFthrust', 'SSFthrust','FBLTSthrust','FBLTthrust','PSBIthrust','PSFBthrust','LT_spiral']:
reportfile.write(', ' + str((1.0 - event.AvailableThrust * event.Isp * 9.80665 / (2000.0 * event.ActivePower)) * event.ActivePower))
else:
reportfile.write(', ' + str(0.0))
if PlotOptions.PlotEarthDistance or PlotOptions.PlotSunSpacecraftEarthAngle or PlotOptions.PlotSpacecraftViewingAngle:
if not 'sun' in self.central_body.lower():
print('getting Earth distance and Sun-Spacecraft-Earth angle is only supported if the central body is the sun')
else:
#get the position of the central body relative to the solar system barycenter
cb_relative_to_solar_system_barycenter_in_ICRF = self.eph.position('sun', event.JulianDate)
#get the position of the Earth relative to the central body
embarycenter_in_ICRF = self.eph.position('earthmoon', event.JulianDate)
moonvector_in_ICRF = self.eph.position('moon', event.JulianDate)
Earth_relative_to_solar_system_barycenter_in_ICRF = embarycenter_in_ICRF - moonvector_in_ICRF * self.eph.earth_share
Earth_relative_to_cb_in_ICRF = Earth_relative_to_solar_system_barycenter_in_ICRF + cb_relative_to_solar_system_barycenter_in_ICRF
#rotate the spacecraft state relative to the central body into the ICRF frame
spacecraft_state_relative_to_central_body_in_ICRF = AstroFunctions.rotate_from_ecliptic_to_equatorial3(np.array(event.SpacecraftState[0:3]))
#get the vector from the Earth to the Spacecraft
Earth_Spacecraft_Vector = spacecraft_state_relative_to_central_body_in_ICRF - np.transpose(Earth_relative_to_cb_in_ICRF)[0]
#return the magnitude of the Earth-centered position vector
EarthDistance = np.linalg.norm(Earth_Spacecraft_Vector) / self.LU
if PlotOptions.PlotEarthDistance:
reportfile.write(', ' + str(EarthDistance))
if PlotOptions.PlotSunSpacecraftEarthAngle:
cosAngle = np.dot(-Earth_Spacecraft_Vector, -spacecraft_state_relative_to_central_body_in_ICRF)/np.linalg.norm(spacecraft_state_relative_to_central_body_in_ICRF)/np.linalg.norm(Earth_Spacecraft_Vector)
reportfile.write(', ' + str(np.arccos(cosAngle) * 180.0/math.pi))
if PlotOptions.PlotSpacecraftViewingAngle:
tanAngle = Earth_Spacecraft_Vector[2]/math.sqrt(Earth_Spacecraft_Vector[0]*Earth_Spacecraft_Vector[0]+Earth_Spacecraft_Vector[1]*Earth_Spacecraft_Vector[1])
reportfile.write(', ' + str(np.arctan(tanAngle) * 180.0/math.pi))
if PlotOptions.PlotThrottleLevel:
if event.EventType in ['SFthrust', 'SSFthrust', 'FBLTSthrust', 'FBLTthrust', 'PSBIthrust','PSFBthrust', 'end_spiral']:
if PlotOptions.throttlesetmode == 0:
reportfile.write(', ' + str(int(self.high_Mdot_set.find_nearest_throttle_setting_1D(event.ActivePower / event.Number_of_Active_Engines).TL.strip('TL'))))
elif PlotOptions.throttlesetmode == 1:
reportfile.write(', ' + str(int(self.low_Mdot_set.find_nearest_throttle_setting_1D(event.ActivePower / event.Number_of_Active_Engines).TL.strip('TL'))))
elif PlotOptions.throttlesetmode == 2:
print('2D throttle matching not currently implemented')
else:
reportfile.write(', ' + str(0.0))
#plot sun-boresight angle
if PlotOptions.PlotSunBoresightAngle:
if event.EventType in ['SFthrust', 'SSFthrust', 'FBLTSthrust', 'FBLTthrust', "PSBIthrust",'PSFBthrust']:
r = math.sqrt(event.SpacecraftState[0]**2 + event.SpacecraftState[1]**2 + event.SpacecraftState[2]**2)
AppliedThrust = math.sqrt(event.Thrust[0]**2 + event.Thrust[1]**2 + event.Thrust[2]**2)
rdotT = event.SpacecraftState[0]*event.Thrust[0] + event.SpacecraftState[1]*event.Thrust[1] + event.SpacecraftState[2]*event.Thrust[2]
reportfile.write(', ' + str(math.acos( rdotT / (r * AppliedThrust) ) * 180.0 / math.pi))
else:
reportfile.write(', ' + str(0.0))
reportfile.write('\n')
def AutoTableJourney(self, TableFile, PlotOptions, skipNext):
for event in self.missionevents:
skipNext = event.AutoTableLine(TableFile, PlotOptions, skipNext)
return skipNext
def getManeuver(self, phaseIndex = 0, maneuverIndex = 0):
#function to get the nth DSM/TCM in a phase
#if that maneuver does not exist, return None
#Step 1: search through the journey to find the right phase. We do this by counting the number of boundary events until we get to the end of the journey or the phase that we want
#since phase boundaries are always upwr_flyby or pwr_flyby, this isn't too hard!
phaseCount = 0
maneuverCount = 0
for event in self.missionevents:
if event.EventType in ['pwr_flyby','upwr_flyby']:
phaseCount += 1
maneuverCount = 0
if event.EventType == 'chem_burn':
if maneuverCount == maneuverIndex:
#this is the maneuver that we want, so return it
return event
#if we didn't just return, increment the maneuver count
maneuverCount += 1
#if we got this far then the event that the user asked for does not exist, so return "none"
return None
def getPeriapseDistance(self, scale = None):
r_min = 1.0e+10
if scale == None:
scale = self.LU
for event in self.missionevents:
r = (event.SpacecraftState[0]**2 + event.SpacecraftState[1]**2 + event.SpacecraftState[2]**2 ) ** 0.5
if r < r_min:
r_min = r
return r_min / scale
def GenerateJourneyDataPlot(self, DataAxesLeft, DataAxesRight, PlotOptions, firstpass):
import math
import astropy
#generate a vector of dates
date_string_vector = []
for event in self.missionevents:
if event.EventType != 'match_point' and event.EventType != 'upwr_flyby' and event.EventType != 'pwr_flyby' and event.EventType != 'LT_rndzvs':
date_string_vector.append(event.GregorianDate)
date_vector = [datetime.datetime.strptime(d,'%m/%d/%Y').date() for d in date_string_vector]
#dummy line across the bottom so that neither axes object crashes
DataAxesLeft.plot(date_vector, np.zeros_like(date_vector), c='w', lw = 0.1)
DataAxesRight.plot(date_vector, np.zeros_like(date_vector), c='w', lw = 0.1)
#plot distance from central body
if PlotOptions.PlotR:
Rvector = []
for event in self.missionevents:
if event.EventType != 'match_point' and event.EventType != 'upwr_flyby' and event.EventType != 'pwr_flyby' and event.EventType != 'LT_rndzvs':
Rvector.append(math.sqrt(event.SpacecraftState[0]**2 + event.SpacecraftState[1]**2 + event.SpacecraftState[2]**2) / self.LU)
if firstpass:
unitstring = 'LU'
if 'sun' in self.central_body.lower() and math.fabs(self.LU - 149597870.69100001) < 1.0:
unitstring = 'AU'
labelstring = 'Distance from ' + self.central_body + ' (' + unitstring + ')'
DataAxesLeft.plot(date_vector, Rvector, c='k', lw=2, label=labelstring)
else:
DataAxesLeft.plot(date_vector, Rvector, c='k', lw=2)
#plot velocity
if PlotOptions.PlotV:
Vvector = []
for event in self.missionevents:
if event.EventType != 'match_point' and event.EventType != 'upwr_flyby' and event.EventType != 'pwr_flyby' and event.EventType != 'LT_rndzvs':
Vvector.append(math.sqrt(event.SpacecraftState[3]**2 + event.SpacecraftState[4]**2 + event.SpacecraftState[5]**2) / self.LU * self.TU)
if firstpass:
DataAxesLeft.plot(date_vector, Vvector, c='k', lw=2, ls='-.', label='Velocity magnitude (LU/TU)')
else:
DataAxesLeft.plot(date_vector, Vvector, c='k', lw=2, ls='-.')
#plot Thrust
if PlotOptions.PlotThrust:
Thrustvector = []
shortDateVector = []
for event in self.missionevents:
epoch = astropy.time.Time(event.JulianDate, format='jd', scale='tdb')
if event.EventType in ['SFthrust', 'SSFthrust', 'FBLTSthrust', 'FBLTthrust', 'PSBIthrust', 'PSFBthrust']:
Thrustvector.append(math.sqrt(event.Thrust[0]**2 + event.Thrust[1]**2 + event.Thrust[2]**2) * 10.0)
shortDateVector.append((epoch - event.TimestepLength / 2).datetime)
Thrustvector.append(math.sqrt(event.Thrust[0]**2 + event.Thrust[1]**2 + event.Thrust[2]**2) * 10.0)
shortDateVector.append((epoch + event.TimestepLength / 2).datetime)
elif event.EventType == 'LT_spiral':
Thrustvector.append(event.AvailableThrust * self.thruster_duty_cycle*10.0)
shortDateVector.append((epoch - event.TimestepLength / 2).datetime)
Thrustvector.append(event.AvailableThrust * self.thruster_duty_cycle*10.0)
shortDateVector.append((epoch + event.TimestepLength / 2).datetime)
elif event.EventType != 'match_point':
Thrustvector.append(0.0)
shortDateVector.append((epoch - event.TimestepLength / 2).datetime)
Thrustvector.append(0.0)
shortDateVector.append((epoch + event.TimestepLength / 2).datetime)
if firstpass:
DataAxesLeft.plot(shortDateVector, Thrustvector, c='r', lw=2, ls='--', label='Applied thrust (0.1 N)')
else:
DataAxesLeft.plot(shortDateVector, Thrustvector, c='r', lw=2, ls='--')
#plot Isp
if PlotOptions.PlotIsp:
Ispvector = []
shortDateVector = []
for event in self.missionevents:
epoch = astropy.time.Time(event.JulianDate, format='jd', scale='tdb')
if event.EventType in ['SFthrust', 'SSFthrust', 'FBLTSthrust', 'FBLTthrust', 'PSBIthrust', 'PSFBthrust', 'LT_spiral']:
Ispvector.append(event.Isp / 1000.0)
shortDateVector.append((epoch - event.TimestepLength / 2).datetime)
Ispvector.append(event.Isp / 1000.0)
shortDateVector.append((epoch + event.TimestepLength / 2).datetime)
elif event.EventType != 'match_point':
Ispvector.append(0.0)
shortDateVector.append((epoch - event.TimestepLength / 2).datetime)
Ispvector.append(0.0)
shortDateVector.append((epoch + event.TimestepLength / 2).datetime)
if firstpass:
DataAxesLeft.plot(shortDateVector, Ispvector, 'dodgerblue', lw=4, ls='-.', label='Isp (1000 s)')
else:
DataAxesLeft.plot(shortDateVector, Ispvector, 'dodgerblue', lw=4, ls='-.')
#plot mass flow rate
if PlotOptions.PlotMdot:
Mdotvector = []
shortDateVector = []
for event in self.missionevents:
epoch = astropy.time.Time(event.JulianDate, format='jd', scale='tdb')
if event.EventType in ['SFthrust', 'SSFthrust', 'FBLTSthrust', 'FBLTthrust', 'PSBIthrust', 'PSFBthrust', 'LT_spiral']:
Mdotvector.append(event.MassFlowRate * 1.0e6)
shortDateVector.append((epoch - event.TimestepLength / 2).datetime)
Mdotvector.append(event.MassFlowRate * 1.0e6)
shortDateVector.append((epoch + event.TimestepLength / 2).datetime)
elif event.EventType != 'match_point':
Mdotvector.append(0.0)
shortDateVector.append((epoch - event.TimestepLength / 2).datetime)
Mdotvector.append(0.0)
shortDateVector.append((epoch + event.TimestepLength / 2).datetime)
if firstpass:
DataAxesLeft.plot(shortDateVector, Mdotvector, c='brown', lw=2, ls='-', label='Mass flow rate (mg/s)')
else:
DataAxesLeft.plot(shortDateVector, Mdotvector, c='brown', lw=2, ls='-')
#plot Efficiency
if PlotOptions.PlotEfficiency:
Efficiencyvector = []
shortDateVector = []
for event in self.missionevents:
epoch = astropy.time.Time(event.JulianDate, format='jd', scale='tdb')
if event.EventType in ['SFthrust', 'SSFthrust', 'FBLTSthrust', 'FBLTthrust', 'PSBIthrust', 'PSFBthrust', 'LT_spiral']:
Efficiencyvector.append(event.AvailableThrust * event.Isp * 9.80665 / (2000 * event.ActivePower))
shortDateVector.append((epoch - event.TimestepLength / 2).datetime)
Efficiencyvector.append(event.AvailableThrust * event.Isp * 9.80665 / (2000 * event.ActivePower))
shortDateVector.append((epoch + event.TimestepLength / 2).datetime)
elif event.EventType != 'match_point':
Efficiencyvector.append(0.0)
shortDateVector.append((epoch - event.TimestepLength / 2).datetime)
Efficiencyvector.append(0.0)
shortDateVector.append((epoch + event.TimestepLength / 2).datetime)
if firstpass:
DataAxesLeft.plot(shortDateVector, Efficiencyvector, c='DarkGreen', lw=2, ls='-', label='Propulsion system efficiency')
else:
DataAxesLeft.plot(shortDateVector, Efficiencyvector, c='DarkGreen', lw=2, ls='-')
#plot control magnitude
if PlotOptions.PlotThrottle:
Throttlevector = []
shortDateVector = []
for event in self.missionevents:
epoch = astropy.time.Time(event.JulianDate, format='jd', scale='tdb')
if event.EventType in ['SFthrust', 'SSFthrust', 'PSBIthrust']:
Throttlevector.append(math.sqrt(event.Thrust[0]**2 + event.Thrust[1]**2 + event.Thrust[2]**2) / (event.AvailableThrust))
shortDateVector.append((epoch - event.TimestepLength / 2).datetime)
Throttlevector.append(math.sqrt(event.Thrust[0]**2 + event.Thrust[1]**2 + event.Thrust[2]**2) / (event.AvailableThrust))
shortDateVector.append((epoch + event.TimestepLength / 2).datetime)
elif event.EventType in ['FBLTthrust','FBLTSthrust','PSFBthrust']:
Throttlevector.append(event.DVmagorThrottle)
shortDateVector.append((epoch - event.TimestepLength / 2).datetime)
Throttlevector.append(event.DVmagorThrottle)
shortDateVector.append((epoch + event.TimestepLength / 2).datetime)
elif event.EventType == 'LT_spiral':
Throttlevector.append(1.0)
shortDateVector.append((epoch - event.TimestepLength / 2).datetime)
Throttlevector.append(1.0)
shortDateVector.append((epoch + event.TimestepLength / 2).datetime)
elif event.EventType != 'match_point':
Throttlevector.append(0.0)
shortDateVector.append((epoch - event.TimestepLength / 2).datetime)
Throttlevector.append(0.0)
shortDateVector.append((epoch + event.TimestepLength / 2).datetime)
if firstpass:
DataAxesLeft.plot(shortDateVector, Throttlevector, c='r', lw=2, ls='--', label='Control magnitude')
else:
DataAxesLeft.plot(shortDateVector, Throttlevector, c='r', lw=2, ls='--')
#plot power
if PlotOptions.PlotPower:
Powervector = []
for event in self.missionevents:
if event.EventType != 'match_point' and event.EventType != 'upwr_flyby' and event.EventType != 'pwr_flyby' and event.EventType != 'LT_rndzvs':
if event.AvailablePower > 0.0:
Powervector.append(event.AvailablePower)
else:
Powervector.append(0.0)
if firstpass:
DataAxesLeft.plot(date_vector, Powervector, c='Navy', lw=2, ls='-.', label='Power available for propulsion (kW)')
else:
DataAxesLeft.plot(date_vector, Powervector, c='Navy', lw=2, ls='-.')
#plot gamma
if PlotOptions.PlotGamma:
gammavector = []
shortDateVector = []
for event in self.missionevents:
if event.EventType in ['SFthrust', 'SSFthrust', 'FBLTSthrust', 'FBLTthrust', 'PSBIthrust', 'PSFBthrust', 'LT_spiral']:
gammavector.append(math.atan2(event.Thrust[1], event.Thrust[0]) * 180.0 / math.pi)
shortDateVector.append(datetime.datetime.strptime(event.GregorianDate,'%m/%d/%Y').date())
if firstpass:
DataAxesRight.plot(shortDateVector, gammavector, c='DarkGreen', lw=2, ls='--', label=r'$\gamma$ (degrees)')
else:
DataAxesRight.plot(shortDateVector, gammavector, c='DarkGreen', lw=2, ls='--')
#plot delta
if PlotOptions.PlotDelta:
deltavector = []
shortDateVector = []
for event in self.missionevents:
if event.EventType in ['SFthrust', 'SSFthrust', 'FBLTSthrust', 'FBLTthrust', 'PSBIthrust', 'PSFBthrust', 'LT_spiral']:
AppliedThrust = math.sqrt(event.Thrust[0]**2 + event.Thrust[1]**2 + event.Thrust[2]**2)
deltavector.append(math.asin(event.Thrust[2] / AppliedThrust) * 180.0 / math.pi)
shortDateVector.append(datetime.datetime.strptime(event.GregorianDate,'%m/%d/%Y').date())
if firstpass:
DataAxesRight.plot(shortDateVector, deltavector, c='LightGreen', lw=2, ls='--', label=r'$\delta$ (degrees)')
else:
DataAxesRight.plot(shortDateVector, deltavector, c='LightGreen', lw=2, ls='--')
#plot central body to thrust vector angle
if PlotOptions.PlotArray_Thrust_Angle:
CBthrustvector = []
shortDateVector = []
for event in self.missionevents:
if event.EventType in ['SFthrust', 'SSFthrust', 'FBLTSthrust', 'FBLTthrust', 'PSBIthrust', 'PSFBthrust', 'LT_spiral']:
r = math.sqrt(event.SpacecraftState[0]**2 + event.SpacecraftState[1]**2 + event.SpacecraftState[2]**2)
AppliedThrust = math.sqrt(event.Thrust[0]**2 + event.Thrust[1]**2 + event.Thrust[2]**2)
rdotT = event.SpacecraftState[0]*event.Thrust[0] + event.SpacecraftState[1]*event.Thrust[1] + event.SpacecraftState[2]*event.Thrust[2]
CBthrustvector.append( 90.0 - math.acos( rdotT / (r * AppliedThrust) ) * 180.0 / math.pi)
shortDateVector.append(datetime.datetime.strptime(event.GregorianDate,'%m/%d/%Y').date())
if firstpass:
DataAxesRight.plot(shortDateVector, CBthrustvector, c='Salmon', marker='o', ls='None', label='Array to thrust angle (degrees)')
else:
DataAxesRight.plot(shortDateVector, CBthrustvector, c='Salmon', marker='o', ls='None')
#plot mass
if PlotOptions.PlotMass:
mass = []
for event in self.missionevents:
if event.EventType != 'match_point' and event.EventType != 'upwr_flyby' and event.EventType != 'pwr_flyby' and event.EventType != 'LT_rndzvs':
mass.append(event.Mass * 1.0e-3)
if firstpass:
DataAxesLeft.plot(date_vector, mass, c='DarkGrey', lw=2, ls='-', label='Mass (1000 kg)')
else:
DataAxesLeft.plot(date_vector, mass, c='DarkGrey', lw=2, ls='-')
#plot number of active thrusters
if PlotOptions.PlotNumberOfEngines:
numberofengines = []
shortDateVector = []
for event in self.missionevents:
if event.EventType in ['SFthrust', 'SSFthrust', 'FBLTSthrust', 'FBLTthrust', 'PSBIthrust', 'PSFBthrust', 'LT_spiral']:
numberofengines.append(event.ActivePower)
shortDateVector.append(datetime.datetime.strptime(event.GregorianDate,'%m/%d/%Y').date())
if firstpass:
DataAxesLeft.plot(shortDateVector, numberofengines, 'k*', mew=2, markersize=7, label='Number of active thrusters')
else:
DataAxesLeft.plot(shortDateVector, numberofengines, 'k*', mew=2, markersize=7)
#plot power actively used by the thrusters
if PlotOptions.PlotActivePower:
activepowervector = []
shortDateVector = []
for event in self.missionevents:
if event.EventType in ['SFthrust', 'SSFthrust', 'FBLTSthrust', 'FBLTthrust', 'PSBIthrust', 'PSFBthrust', 'LT_spiral']:
activepowervector.append(event.ActivePower)
shortDateVector.append(datetime.datetime.strptime(event.GregorianDate,'%m/%d/%Y').date())
if firstpass:
DataAxesLeft.plot(shortDateVector, activepowervector, c='Navy', lw=2, ls='-', label='Power used by the propulsion system (kW)')
else:
DataAxesLeft.plot(shortDateVector, activepowervector, c='Navy', lw=2, ls='-')
#plot waste heat from the propulsion system
if PlotOptions.PlotWasteHeat:
WasteHeatvector = []
shortDateVector = []
for event in self.missionevents:
if event.EventType in ['SFthrust', 'SSFthrust', 'FBLTSthrust', 'FBLTthrust', 'PSBIthrust', 'PSFBthrust', 'LT_spiral']:
WasteHeatvector.append( (1 - event.AvailableThrust * event.Isp * 9.80665 / (2000 * event.ActivePower)) * event.ActivePower )
shortDateVector.append(datetime.datetime.strptime(event.GregorianDate,'%m/%d/%Y').date())
if firstpass:
DataAxesLeft.plot(shortDateVector, WasteHeatvector, c='Crimson', lw=2, ls='--', label='Waste heat from propulsion system (kW)')
else:
DataAxesLeft.plot(shortDateVector, WasteHeatvector, c='Crimson', lw=2, ls='--')
#plot Earth distance in LU and SPE angle
if PlotOptions.PlotEarthDistance or PlotOptions.PlotSunSpacecraftEarthAngle or PlotOptions.PlotSpacecraftViewingAngle:
if not 'sun' in self.central_body.lower():
print('getting Earth distance and Sun-Spacecraft-Earth angle is only supported if the central body is the sun')
else:
import de423
from jplephem import Ephemeris
eph = Ephemeris(de423)
EarthDistanceVector = []
SunSpacecraftEarthAngle = []
SpacecraftViewingAngle = []
for event in self.missionevents:
if event.EventType != 'match_point' and event.EventType != 'upwr_flyby' and event.EventType != 'pwr_flyby' and event.EventType != 'LT_rndzvs':
#get the position of the central body relative to the solar system barycenter
cb_relative_to_solar_system_barycenter_in_ICRF = eph.position('sun', event.JulianDate)
#get the position of the Earth relative to the central body
embarycenter_in_ICRF = eph.position('earthmoon', event.JulianDate)
moonvector_in_ICRF = eph.position('moon', event.JulianDate)
Earth_relative_to_solar_system_barycenter_in_ICRF = embarycenter_in_ICRF - moonvector_in_ICRF * eph.earth_share
Earth_relative_to_cb_in_ICRF = Earth_relative_to_solar_system_barycenter_in_ICRF + cb_relative_to_solar_system_barycenter_in_ICRF
#rotate the spacecraft state relative to the central body into the ICRF frame
spacecraft_state_relative_to_central_body_in_ICRF = AstroFunctions.rotate_from_ecliptic_to_equatorial3(np.array(event.SpacecraftState[0:3]))
#get the vector from the Earth to the Spacecraft
Earth_Spacecraft_Vector = spacecraft_state_relative_to_central_body_in_ICRF - np.transpose(Earth_relative_to_cb_in_ICRF)[0]
#return the magnitude of the Earth-centered position vector
EarthDistanceVector.append(np.linalg.norm(Earth_Spacecraft_Vector) / self.LU)
#if applicable, compute sun-spacecraft-Earth angle
if PlotOptions.PlotSunSpacecraftEarthAngle:
cosAngle = np.dot(-Earth_Spacecraft_Vector, -spacecraft_state_relative_to_central_body_in_ICRF)/np.linalg.norm(spacecraft_state_relative_to_central_body_in_ICRF)/np.linalg.norm(Earth_Spacecraft_Vector)
SunSpacecraftEarthAngle.append(np.arccos(cosAngle) * 180.0/math.pi)
if PlotOptions.PlotSpacecraftViewingAngle:
tanAngle = Earth_Spacecraft_Vector[2]/math.sqrt(Earth_Spacecraft_Vector[0]*Earth_Spacecraft_Vector[0]+Earth_Spacecraft_Vector[1]*Earth_Spacecraft_Vector[1])
SpacecraftViewingAngle.append(np.arctan(tanAngle) * 180.0/math.pi)
if PlotOptions.PlotEarthDistance:
if firstpass:
unitstring = 'LU'
if 'sun' in self.central_body.lower() and math.fabs(self.LU - 149597870.69100001) < 1.0:
unitstring = 'AU'
labelstring = 'Distance from Earth (' + unitstring + ')'
DataAxesLeft.plot(date_vector, EarthDistanceVector, c='g', lw=2, label=labelstring)
else:
DataAxesLeft.plot(date_vector, EarthDistanceVector, c='g', lw=2)
if PlotOptions.PlotSunSpacecraftEarthAngle:
if firstpass:
DataAxesRight.plot(date_vector, SunSpacecraftEarthAngle, c='orangered', lw=2, ls = '-.', label='Sun-Spacecraft-Earth Angle')
else:
DataAxesRight.plot(date_vector, SunSpacecraftEarthAngle, c='orangered', lw=2, ls = '-.')
if PlotOptions.PlotSpacecraftViewingAngle:
if firstpass:
DataAxesRight.plot(date_vector, SpacecraftViewingAngle, c='orangered', lw=2, ls = '-.', label='Latitude of Earth-Spacecraft Vector')
else:
DataAxesRight.plot(date_vector, SpacecraftViewingAngle, c='orangered', lw=2, ls = '-.')
#plot throttle level
if PlotOptions.PlotThrottleLevel:
ThrottleLevelVector = []
shortDateVector = []
for event in self.missionevents:
if event.EventType in ['SFthrust', 'SSFthrust', 'FBLTSthrust', 'FBLTthrust', 'PSBIthrust', 'PSFBthrust', 'LT_spiral']:
ThrottleLevelVector.append(event.ThrottleLevel)
shortDateVector.append(datetime.datetime.strptime(event.GregorianDate,'%m/%d/%Y').date())
if firstpass:
DataAxesLeft.scatter(shortDateVector, ThrottleLevelVector, c='midnightblue', marker='h' , label='Throttle level')
else:
DataAxesLeft.scatter(shortDateVector, ThrottleLevelVector, c='midnightblue', marker='h' )
#plot sun-boresight angle
if PlotOptions.PlotSunBoresightAngle:
SunBoresightAngle = []
for event in self.missionevents:
if event.EventType in ['SFthrust', 'SSFthrust', 'FBLTSthrust', 'FBLTthrust', 'PSFBthrust', "PSBIthrust"]:
r = math.sqrt(event.SpacecraftState[0]**2 + event.SpacecraftState[1]**2 + event.SpacecraftState[2]**2)
AppliedThrust = math.sqrt(event.Thrust[0]**2 + event.Thrust[1]**2 + event.Thrust[2]**2)
rdotT = -event.SpacecraftState[0]*event.Thrust[0] - event.SpacecraftState[1]*event.Thrust[1] - event.SpacecraftState[2]*event.Thrust[2]
SunBoresightAngle.append(math.acos( rdotT / (r * AppliedThrust) ) * 180.0 / math.pi)
elif event.EventType != 'match_point' and event.EventType != 'upwr_flyby' and event.EventType != 'pwr_flyby' and event.EventType != 'LT_rndzvs':
SunBoresightAngle.append(0.0)
if firstpass:
DataAxesRight.plot(date_vector, SunBoresightAngle, c='purple', marker='o', ls='None', label='Sun-to-Boresight angle (degrees)')
else:
DataAxesRight.plot(date_vector, SunBoresightAngle, c='purple', marker='o', ls='None')
class Tests(TestCase):
def setUp(self):
import de421
self.e = Ephemeris(de421)
def check0(self, xyz, xyzdot=None):
eq = partial(self.assertAlmostEqual, delta=1.0)
x, y, z = xyz
eq(x, 39705023.28)
eq(y, 131195345.65)
eq(z, 56898495.41)
if xyzdot is None:
return
dx, dy, dz = xyzdot
eq(dx, -2524248.19)
eq(dy, 619970.11)
eq(dz, 268928.26)
def check1(self, xyz, xyzdot=None):
eq = partial(self.assertAlmostEqual, delta=1.0)
x, y, z = xyz
eq(x, -144692624.00)
eq(y, -32707965.14)
eq(z, -14207167.26)
if xyzdot is None:
return
dx, dy, dz = xyzdot
eq(dx, 587334.38)
eq(dy, -2297419.36)
eq(dz, -996628.74)
def test_names(self):
self.assertEqual(self.e.names, (
'earthmoon', 'jupiter', 'librations', 'mars', 'mercury',
'moon', 'neptune', 'nutations', 'pluto', 'saturn', 'sun',
'uranus', 'venus',
))
def test_scalar_tdb(self):
self.check0(self.e.position('earthmoon', 2414994.0))
self.check1(self.e.position('earthmoon', 2415112.5))
def test_scalar_tdb2(self):
self.check0(self.e.position('earthmoon', 2414990.0, 4.0))
self.check1(self.e.position('earthmoon', 2415110.0, 2.5))
def test_scalar_tdb_keyword(self):
self.check0(self.e.position('earthmoon', tdb=2414994.0))
self.check1(self.e.position('earthmoon', tdb=2415112.5))
def test_scalar_tdb2_keyword(self):
self.check0(self.e.position('earthmoon', tdb=2414990.0, tdb2=4.0))
self.check1(self.e.position('earthmoon', tdb=2415110.0, tdb2=2.5))
def check_2d_result(self, name, tdb, tdb2):
p = self.e.position(name, tdb + tdb2)
self.check0(p[:,0])
self.check1(p[:,1])
p = self.e.position(name, tdb, tdb2)
self.check0(p[:,0])
self.check1(p[:,1])
p, v = self.e.position_and_velocity(name, tdb + tdb2)
self.check0(p[:,0], v[:,0])
self.check1(p[:,1], v[:,1])
p, v = self.e.position_and_velocity(name, tdb, tdb2)
self.check0(p[:,0], v[:,0])
self.check1(p[:,1], v[:,1])
def test_array_tdb(self):
tdb = np.array([2414994.0, 2415112.5])
tdb2 = 0.0
self.check_2d_result('earthmoon', tdb, tdb2)
def test_array_tdb_scalar_tdb2(self):
tdb = np.array([2414991.5, 2415110.0])
tdb2 = 2.5
self.check_2d_result('earthmoon', tdb, tdb2)
def test_scalar_tdb_array_tdb2(self):
tdb = 2414990.0
d = 2415112.5 - tdb
tdb2 = np.array([4.0, d])
self.check_2d_result('earthmoon', tdb, tdb2)
def test_array_tdb_array_tdb2(self):
tdb = np.array([2414990.0, 2415110.0])
tdb2 = np.array([4.0, 2.5])
self.check_2d_result('earthmoon', tdb, tdb2)
def test_legacy_compute_method(self):
pv = self.e.compute('earthmoon', 2414994.0)
self.check0(pv[:3], pv[3:])
pv = self.e.compute('earthmoon', np.array([2414994.0, 2415112.5]))
self.check0(pv[:3,0], pv[3:,0])
self.check1(pv[:3,1], pv[3:,1])
def test_ephemeris_end_date(self):
x, y, z = self.e.position('earthmoon', self.e.jomega)
self.assertAlmostEqual(x, -94189805.73967789, delta=1.0)
self.assertAlmostEqual(y, 1.05103857e+08, delta=1.0)
self.assertAlmostEqual(z, 45550861.44383482, delta=1.0)
def test_too_early_date(self):
tdb = self.e.jalpha - 0.01
self.assertRaises(DateError, self.e.position, 'earthmoon', tdb)
def test_too_late_date(self):
tdb = self.e.jomega + 16.01
self.assertRaises(DateError, self.e.position, 'earthmoon', tdb)
def GenerateJourneyDataPlot(self, DataAxesLeft, DataAxesRight, PlotOptions,
firstpass):
#generate a vector of dates
date_string_vector = []
for event in self.missionevents:
if event.EventType != 'match_point' and event.EventType != 'upwr_flyby' and event.EventType != 'pwr_flyby':
date_string_vector.append(event.GregorianDate)
date_vector = [
datetime.datetime.strptime(d, '%m/%d/%Y').date()
for d in date_string_vector
]
#dummy line across the bottom so that neither axes object crashes
DataAxesLeft.plot(date_vector,
np.zeros_like(date_vector),
c='w',
lw=0.1)
DataAxesRight.plot(date_vector,
np.zeros_like(date_vector),
c='w',
lw=0.1)
#plot distance from central body
if PlotOptions.PlotR:
Rvector = []
for event in self.missionevents:
if event.EventType != 'match_point' and event.EventType != 'upwr_flyby' and event.EventType != 'pwr_flyby':
Rvector.append(
math.sqrt(event.SpacecraftState[0]**2 +
event.SpacecraftState[1]**2 +
event.SpacecraftState[2]**2) / self.LU)
if firstpass:
unitstring = 'LU'
if self.central_body.lower(
) == 'sun' and math.fabs(self.LU - 149597870.69100001) < 1.0:
unitstring = 'AU'
labelstring = 'Distance from ' + self.central_body + ' (' + unitstring + ')'
DataAxesLeft.plot(date_vector,
Rvector,
c='k',
lw=2,
label=labelstring)
else:
DataAxesLeft.plot(date_vector, Rvector, c='k', lw=2)
#plot velocity
if PlotOptions.PlotV:
Vvector = []
for event in self.missionevents:
if event.EventType != 'match_point' and event.EventType != 'upwr_flyby' and event.EventType != 'pwr_flyby':
Vvector.append(
math.sqrt(event.SpacecraftState[3]**2 +
event.SpacecraftState[4]**2 +
event.SpacecraftState[5]**2) / self.LU *
self.TU)
if firstpass:
DataAxesLeft.plot(date_vector,
Vvector,
c='k',
lw=2,
ls='-.',
label='Velocity magnitude (LU/TU)')
else:
DataAxesLeft.plot(date_vector, Vvector, c='k', lw=2, ls='-.')
#plot Thrust
if PlotOptions.PlotThrust:
Thrustvector = []
for event in self.missionevents:
if event.EventType != 'match_point' and event.EventType != 'upwr_flyby' and event.EventType != 'pwr_flyby':
Thrustvector.append(
math.sqrt(event.Thrust[0]**2 + event.Thrust[1]**2 +
event.Thrust[2]**2) * 10.0)
if firstpass:
DataAxesLeft.plot(date_vector,
Thrustvector,
c='r',
lw=2,
ls='-',
label='Applied thrust (0.1 N)')
else:
DataAxesLeft.plot(date_vector,
Thrustvector,
c='r',
lw=2,
ls='-')
#plot Isp
if PlotOptions.PlotIsp:
Ispvector = []
for event in self.missionevents:
if event.EventType != 'match_point' and event.EventType != 'upwr_flyby' and event.EventType != 'pwr_flyby':
Ispvector.append(event.Isp / 1000.0)
if firstpass:
DataAxesLeft.plot(date_vector,
Ispvector,
c='c',
lw=2,
ls='-',
label='Isp (1000 s)')
else:
DataAxesLeft.plot(date_vector, Ispvector, c='c', lw=2, ls='-')
#plot mass flow rate
if PlotOptions.PlotMdot:
Mdotvector = []
for event in self.missionevents:
if event.EventType != 'match_point' and event.EventType != 'upwr_flyby' and event.EventType != 'pwr_flyby':
Mdotvector.append(event.MassFlowRate * 1.0e6)
if firstpass:
DataAxesLeft.plot(date_vector,
Mdotvector,
c='brown',
lw=2,
ls='-',
label='Mass flow rate (mg/s)')
else:
DataAxesLeft.plot(date_vector,
Mdotvector,
c='brown',
lw=2,
ls='-')
#plot Efficiency
if PlotOptions.PlotEfficiency:
Efficiencyvector = []
for event in self.missionevents:
if event.EventType == 'SFthrust' or event.EventType == 'FBLTthrust' or event.EventType == "PSBIthrust":
Efficiencyvector.append(event.AvailableThrust * event.Isp *
9.80665 /
(2000 * event.ActivePower))
elif event.EventType != 'match_point' and event.EventType != 'upwr_flyby' and event.EventType != 'pwr_flyby':
Efficiencyvector.append(0.0)
if firstpass:
DataAxesLeft.plot(date_vector,
Efficiencyvector,
c='DarkGreen',
lw=2,
ls='-',
label='Propulsion system efficiency')
else:
DataAxesLeft.plot(date_vector,
Efficiencyvector,
c='DarkGreen',
lw=2,
ls='-')
#plot Throttle
if PlotOptions.PlotThrottle:
Throttlevector = []
for event in self.missionevents:
if event.EventType == 'SFthrust' or event.EventType == 'FBLTthrust' or event.EventType == "PSBIthrust":
Throttlevector.append(
math.sqrt(event.Thrust[0]**2 + event.Thrust[1]**2 +
event.Thrust[2]**2) /
(event.AvailableThrust * self.thruster_duty_cycle))
elif event.EventType != 'match_point' and event.EventType != 'upwr_flyby' and event.EventType != 'pwr_flyby':
Throttlevector.append(0.0)
if firstpass:
DataAxesLeft.plot(date_vector,
Throttlevector,
c='r',
lw=2,
ls='--',
label='Throttle')
else:
DataAxesLeft.plot(date_vector,
Throttlevector,
c='r',
lw=2,
ls='--')
#plot power
if PlotOptions.PlotPower:
Powervector = []
for event in self.missionevents:
if event.EventType != 'match_point' and event.EventType != 'upwr_flyby' and event.EventType != 'pwr_flyby':
if event.AvailablePower > 0.0:
Powervector.append(event.AvailablePower)
else:
Powervector.append(0.0)
if firstpass:
DataAxesLeft.plot(date_vector,
Powervector,
c='Navy',
lw=2,
ls='-',
label='Power produced by spacecraft (kW)')
else:
DataAxesLeft.plot(date_vector,
Powervector,
c='Navy',
lw=2,
ls='-')
#plot gamma
if PlotOptions.PlotGamma:
gammavector = []
for event in self.missionevents:
if event.EventType == 'SFthrust' or event.EventType == 'FBLTthrust' or event.EventType == "PSBIthrust":
gammavector.append(
math.atan2(event.Thrust[1], event.Thrust[0]) * 180.0 /
math.pi)
elif event.EventType != 'match_point' and event.EventType != 'upwr_flyby' and event.EventType != 'pwr_flyby':
gammavector.append(0.0)
if firstpass:
DataAxesRight.plot(date_vector,
gammavector,
c='DarkGreen',
lw=2,
ls='--',
label=r'$\gamma$ (radians)')
else:
DataAxesRight.plot(date_vector,
gammavector,
c='DarkGreen',
lw=2,
ls='--')
#plot delta
if PlotOptions.PlotDelta:
deltavector = []
for event in self.missionevents:
if event.EventType == 'SFthrust' or event.EventType == 'FBLTthrust' or event.EventType == "PSBIthrust":
AppliedThrust = math.sqrt(event.Thrust[0]**2 +
event.Thrust[1]**2 +
event.Thrust[2]**2)
deltavector.append(
math.asin(event.Thrust[2] / AppliedThrust) * 180.0 /
math.pi)
elif event.EventType != 'match_point' and event.EventType != 'upwr_flyby' and event.EventType != 'pwr_flyby':
deltavector.append(0.0)
if firstpass:
DataAxesRight.plot(date_vector,
deltavector,
c='LightGreen',
lw=2,
ls='--',
label=r'$\delta$ (radians)')
else:
DataAxesRight.plot(date_vector,
deltavector,
c='LightGreen',
lw=2,
ls='--')
#plot central body to thrust vector angle
if PlotOptions.PlotCB_thrust_angle:
CBthrustvector = []
for event in self.missionevents:
if event.EventType == 'SFthrust' or event.EventType == 'FBLTthrust' or event.EventType == "PSBIthrust":
r = math.sqrt(event.SpacecraftState[0]**2 +
event.SpacecraftState[1]**2 +
event.SpacecraftState[2]**2)
AppliedThrust = math.sqrt(event.Thrust[0]**2 +
event.Thrust[1]**2 +
event.Thrust[2]**2)
rdotT = event.SpacecraftState[0] * event.Thrust[
0] + event.SpacecraftState[1] * event.Thrust[
1] + event.SpacecraftState[2] * event.Thrust[2]
CBthrustvector.append(
math.acos(rdotT / (r * AppliedThrust)) * 180.0 /
math.pi)
elif event.EventType != 'match_point' and event.EventType != 'upwr_flyby' and event.EventType != 'pwr_flyby':
CBthrustvector.append(0.0)
if firstpass:
DataAxesRight.plot(date_vector,
CBthrustvector,
c='Salmon',
lw=2,
ls='--',
label='CB-thrust angle (radians)')
else:
DataAxesRight.plot(date_vector,
CBthrustvector,
c='Salmon',
lw=2,
ls='--')
#plot mass
if PlotOptions.PlotMass:
mass = []
for event in self.missionevents:
if event.EventType != 'match_point' and event.EventType != 'upwr_flyby' and event.EventType != 'pwr_flyby':
mass.append(event.Mass * 1.0e-3)
if firstpass:
DataAxesLeft.plot(date_vector,
mass,
c='DarkGrey',
lw=2,
ls='-',
label='Mass (1000 kg)')
else:
DataAxesLeft.plot(date_vector,
mass,
c='DarkGrey',
lw=2,
ls='-')
#plot number of active thrusters
if PlotOptions.PlotNumberOfEngines:
numberofengines = []
for event in self.missionevents:
if event.EventType == 'SFthrust' or event.EventType == 'FBLTthrust' or event.EventType == "PSBIthrust":
numberofengines.append(event.Number_of_Active_Engines)
elif event.EventType != 'match_point' and event.EventType != 'upwr_flyby' and event.EventType != 'pwr_flyby':
numberofengines.append(0)
if firstpass:
DataAxesLeft.plot(date_vector,
numberofengines,
c='Orange',
lw=2,
ls='-',
label='Number of active thrusters')
else:
DataAxesLeft.plot(date_vector,
numberofengines,
c='Orange',
lw=2,
ls='-')
#plot power actively used by the thrusters
if PlotOptions.PlotActivePower:
activepowervector = []
for event in self.missionevents:
if event.EventType == 'SFthrust' or event.EventType == 'FBLTthrust' or event.EventType == "PSBIthrust":
activepowervector.append(event.ActivePower)
elif event.EventType != 'match_point' and event.EventType != 'upwr_flyby' and event.EventType != 'pwr_flyby':
activepowervector.append(0.0)
if firstpass:
DataAxesLeft.plot(
date_vector,
activepowervector,
c='Navy',
lw=2,
ls='--',
label='Power used by the propulsion system (kW)')
else:
DataAxesLeft.plot(date_vector,
activepowervector,
c='Navy',
lw=2,
ls='--')
#plot waste heat from the propulsion system
if PlotOptions.PlotWasteHeat:
WasteHeatvector = []
for event in self.missionevents:
if event.EventType == 'SFthrust' or event.EventType == 'FBLTthrust' or event.EventType == "PSBIthrust":
WasteHeatvector.append(
(1 - event.AvailableThrust * event.Isp * 9.80665 /
(2000 * event.ActivePower)) * event.ActivePower)
elif event.EventType != 'match_point' and event.EventType != 'upwr_flyby' and event.EventType != 'pwr_flyby':
WasteHeatvector.append(0.0)
if firstpass:
DataAxesLeft.plot(
date_vector,
WasteHeatvector,
c='Crimson',
lw=2,
ls='--',
label='Waste heat from propulsion system (kW)')
else:
DataAxesLeft.plot(date_vector,
WasteHeatvector,
c='Crimson',
lw=2,
ls='--')
#plot Earth distance in LU and SPE angle
if PlotOptions.PlotEarthDistance or PlotOptions.PlotSunEarthSpacecraftAngle:
if self.central_body.lower() != 'sun':
print 'getting Earth distance and Sun-Spacecraft-Earth angle is only supported if the central body is the sun'
else:
import de423
from jplephem import Ephemeris
eph = Ephemeris(de423)
EarthDistanceVector = []
SunEarthSpacecraftAngle = []
for event in self.missionevents:
if event.EventType != 'match_point' and event.EventType != 'upwr_flyby' and event.EventType != 'pwr_flyby':
#get the position of the central body relative to the solar system barycenter
cb_relative_to_solar_system_barycenter_in_ICRC = eph.position(
self.central_body.lower(), event.JulianDate)
#get the position of the Earth relative to the central body
embarycenter_in_ICRC = eph.position(
'earthmoon', event.JulianDate)
moonvector_in_ICRC = eph.position(
'moon', event.JulianDate)
Earth_relative_to_solar_system_barycenter_in_ICRC = embarycenter_in_ICRC - moonvector_in_ICRC * eph.earth_share
Earth_relative_to_cb_in_ICRC = Earth_relative_to_solar_system_barycenter_in_ICRC + cb_relative_to_solar_system_barycenter_in_ICRC
#rotate the spacecraft state relative to the central body into the ICRF frame
spacecraft_state_relative_to_central_body_in_ICRF = AstroFunctions.rotate_from_ecliptic_to_equatorial3(
np.array(event.SpacecraftState[0:3]))
#get the vector from the Earth to the Spacecraft
Earth_Spacecraft_Vector = spacecraft_state_relative_to_central_body_in_ICRF - np.transpose(
Earth_relative_to_cb_in_ICRC)[0]
#return the magnitude of the Earth-centered position vector
EarthDistanceVector.append(
np.linalg.norm(Earth_Spacecraft_Vector) / self.LU)
#if applicable, compute sun-spacecraft-Earth angle
if PlotOptions.PlotSunEarthSpacecraftAngle:
cosAngle = np.dot(
-Earth_Spacecraft_Vector,
-spacecraft_state_relative_to_central_body_in_ICRF
) / np.linalg.norm(
spacecraft_state_relative_to_central_body_in_ICRF
) / np.linalg.norm(Earth_Spacecraft_Vector)
SunEarthSpacecraftAngle.append(
np.arccos(cosAngle) * 180.0 / math.pi)
if PlotOptions.PlotEarthDistance:
if firstpass:
unitstring = 'LU'
if self.central_body.lower() == 'sun' and math.fabs(
self.LU - 149597870.69100001) < 1.0:
unitstring = 'AU'
labelstring = 'Distance from Earth (' + unitstring + ')'
DataAxesLeft.plot(date_vector,
EarthDistanceVector,
c='g',
lw=2,
label=labelstring)
else:
DataAxesLeft.plot(date_vector,
EarthDistanceVector,
c='g',
lw=2)
if PlotOptions.PlotSunEarthSpacecraftAngle:
if firstpass:
DataAxesRight.plot(date_vector,
SunEarthSpacecraftAngle,
c='orangered',
lw=2,
ls='-.',
label='Sun-Spacecraft-Earth Angle')
else:
DataAxesRight.plot(date_vector,
SunEarthSpacecraftAngle,
c='orangered',
lw=2,
ls='-.')
def setUp(self):
import de421
self.e = Ephemeris(de421)
def find_residuals(check_index, decs, ras, JDs, r, r_dot):
dec_res_total = 0
ra_res_total = 0
total_ra = 0
total_dec = 0
date = "2018-7-16 01:52"
for i in range(len(decs)):
#print str(i + 1) + "/" + str(len(decs))
k = 0.01720209895
time_difference = JDs[i] - JDs[check_index]
tau = 0
d_tau = 0.001 * is_negative(time_difference)
end_tau = time_difference * k
#TEMP INTEGRATOR THINGY HERE (THE NEW ONE BRO)
sim = rebound.Simulation()
date = "2018-7-16 01:52"
#TEST INFO#
#SUN
sim.add(m=1.0)
#ASTEROID
sim.add(m=0.0, x=r.x, y=r.y, z=r.z, vx=r_dot.x, vy=r_dot.y, vz=r_dot.z)
#EARTH
sim.add(m=1.0 / 330000.0,
x=4.189675889417898E-01,
y=-8.496104448760934E-01,
z=-3.683119473905782E-01,
vx=1.539665060981763E-02 / k,
vy=6.454046892673179E-03 / k,
vz=2.797224517462366E-03 / k)
#sim.add('Earth')
#JUPITER
sim.add(m=1.0 / 1000.0,
x=-3.213494013743870E+00,
y=-4.009890793133180E+00,
z=-1.640521133996669E+00,
vx=5.974697377415830E-03 / k,
vy=-3.756443410778172E-03 / k,
vz=-1.755593829452442E-03 / k)
#sim.add('Jupiter')
#SATURN
sim.add(m=1.0 / 3300.0,
x=1.084878715668408E+00,
y=-9.231862109791130E+00,
z=-3.860012135382278E+00,
vx=5.246951399009119E-03 / k,
vy=6.191309101028240E-04,
vz=3.018304002418789E-05 / k)
#sim.add('Saturn')
sim.move_to_com()
ps = sim.particles
sim.integrate(end_tau)
r_now = vector(ps[1].x, ps[1].y, ps[1].z)
AU = 149597870.7 # km/AU
jd = JDs[i]
eph = Ephemeris(de421)
R_geocentric = get_earth_to_sun(JD)
sun_to_asteroid = r_now
earth_to_asteroid = norm(
R_geocentric +
sun_to_asteroid) # Gets unit vector from Earth to Asteroid
RA, Dec = location_to_angles(earth_to_asteroid)
RA_Residual = abs(ras[i] - RA)
Dec_Residual = abs(decs[i] - Dec)
ra_res_total += RA_Residual**2
dec_res_total += Dec_Residual**2
tau = 0
d_tau = 0.000001
k = 0.01720209895
mu = 1.0
sigma = sqrt(ra_res_total + dec_res_total)
return sigma
def pertubation(t, orbit_element):
'''
:param Orbit_Elements: 轨道六根数,单位为弧度
:param t: 时间
:return: 微分方程右边的表达式
'''
'''
d_earth_nonsphericfigure:地球非球型摄动影响
'''
t_start_jd = 2458454.0
R_earth = 6378 #km
mu = 398600 #km^3/s^(-2)
J2 = -1.08264 * 10**(-3)
semi_major_axis = orbit_element[0]
Eccentricity = orbit_element[1]
Inclination = orbit_element[2]
RAAN = orbit_element[3]
Perigee = orbit_element[4]
True_Anomaly = orbit_element[5]
p = semi_major_axis**(1 - Eccentricity**2)
n = sqrt(mu / (semi_major_axis)**3)
d_RAAN_earth_nonsphericfigure = (-2 / 3) * (
R_earth / p)**2 * n * J2 * cos(Inclination)
d_Perigee_earth_nonsphericfigure = -3 / 4 * (R_earth / p)**2 * n * J2 * (
1 - 5 * (cos(Inclination)**2))
d_semi_major_axis_earth_nonsphericfigure = 0
d_Eccentricity_earth_nonsphericfigure = 0
d_Inclination_earth_nonsphericfigure = 0
d_Mean_anomaly_earth_nonsphericfigure = 0
'''
d_lunar_solar日,月引力摄动
'''
Sun_state = ephemeris('sun', t_start_jd + t / 3600 / 24) #太阳星历
barycenter_statement = ephemeris('earthmoon', t_start_jd + t / 3600 / 24)
import de405
from jplephem import Ephemeris
eph = Ephemeris(de405)
Position_sun = Sun_state[0, :]
Velocity_sun = Sun_state[1, :]
Moon_state = ephemeris('moon', t_start_jd + t / 3600 / 24) #本身就是相对地球的
Position_moon = Moon_state[0, :]
Velocity_moon = Moon_state[1, :]
Earth_state = barycenter_statement - Moon_state * eph.earth_share
Position_earth = Earth_state[0, :]
Velocity_earth = Earth_state[1, :]
# 太阳引力摄动
G = 6.67259 * 1e-11 #N·m²/kg²
M_sun = 1.9885 * 1e30 #kg
R_sun_earth = Position_sun - Position_earth #地心ICRF参考系下太阳的位置矢量
V_sun_earth = Velocity_sun - Velocity_earth #地心ICRF参考系下太阳的速度矢量
r_sun = np.sqrt(R_sun_earth[0]**2 + R_sun_earth[1]**2 + R_sun_earth[2]**2)
K_sun = G * M_sun / (r_sun**3) #r_sun为日地距
#计算太阳在地心ICRF坐标系下的轨道根数
import RV_To_Orbit_Elements
mu_sun_earth = 398600 + G * M_sun / (1e9) # km^3/s^2
Sun_Orbit_Elements_ECRF = RV_To_Orbit_Elements.rv_to_orbit_element(
R_sun_earth, V_sun_earth, mu_sun_earth)
Inclination_sun = Sun_Orbit_Elements_ECRF[2]
RAAN_sun = Sun_Orbit_Elements_ECRF[3]
Perigee_sun = Sun_Orbit_Elements_ECRF[4]
True_anomaly_sun = Sun_Orbit_Elements_ECRF[5]
u_sun = True_anomaly_sun + Perigee_sun
#计算系数
A_sun = cos(RAAN - RAAN_sun) * cos(u_sun) + sin(RAAN - RAAN_sun) * sin(
u_sun) * cos(Inclination_sun)
B1_sun = cos(Inclination)
B2_sun = B1_sun * (-sin(RAAN - RAAN_sun) * cos(u_sun) + sin(u_sun) *
cos(Inclination_sun) * cos(RAAN - RAAN_sun))
B_sun = B2_sun + sin(Inclination) * sin(Inclination_sun) * sin(u_sun)
C1_sun = sin(Inclination)
C2_sun = C1_sun * (sin(RAAN - RAAN_sun) * cos(u_sun) - sin(u_sun) *
cos(Inclination_sun) * cos(RAAN - RAAN_sun))
C_sun = C2_sun + cos(Inclination) * sin(Inclination_sun) * sin(u_sun)
d_semi_major_axis_solar = 0
d_Eccentricity_solar = -(15 * K_sun / 2 / n) * Eccentricity * sqrt(
1 - Eccentricity**2) * (A_sun * B_sun * cos(2 * Perigee) - (1 / 2) *
(A_sun**2 - B_sun**2) * sin(2 * Perigee))
d_Inclination_solar=3*K_sun*C_sun/4/n/sqrt(1-Eccentricity**2)*(A_sun*(2+3*Eccentricity**2+\
5*Eccentricity**2*cos(2*Perigee))+5*(B_sun*Eccentricity**2)*sin(2*Perigee))
d_RAAN_solar = 3*K_sun*C_sun/4/n/sqrt(1-Eccentricity**2)/sin(Inclination)*(B_sun*(2+3*Eccentricity**2+\
5*Eccentricity**2*cos(2*Perigee))+5*A_sun*Eccentricity**2*sin(2*Perigee))
d_Perigee_solar = 3*K_sun/2/n*sqrt(1-Eccentricity**2)*((5*A_sun*B_sun*sin(Perigee*2)+\
1/2*(A_sun**2-B_sun**2)*cos(2*Perigee))-1+3/2*(A_sun**2+B_sun**2))\
+5*semi_major_axis/2/Eccentricity/r_sun*(1-5/4*(A_sun**2-B_sun**2))*(A_sun*cos(Perigee)\
+B_sun*sin(Perigee))-cos(Inclination)*d_RAAN_solar
d_Mean_anomaly_solar = 0
# 月球引力摄动
M_moon = 7.349 * 1e22 #kg
R_moon_earth = Position_moon #地心ICRF参考系下太阳的位置矢量
V_moon_earth = Velocity_moon #地心ICRF参考系下太阳的速度矢量
r_moon = np.sqrt(R_moon_earth.dot(R_moon_earth))
K_moon = G * M_moon / (r_moon**3) #r_moon为月地距
mu_moon_earth = 398600 + G * M_moon / (1e9)
#计算月球在地心ICRF坐标系下的轨道根数
Moon_Orbit_Elements_ECRF = RV_To_Orbit_Elements.rv_to_orbit_element(
R_moon_earth, V_moon_earth, mu_moon_earth)
Inclination_moon = Moon_Orbit_Elements_ECRF[2]
RAAN_moon = Moon_Orbit_Elements_ECRF[3]
Perigee_moon = Moon_Orbit_Elements_ECRF[4]
True_anomaly_moon = Moon_Orbit_Elements_ECRF[5]
u_moon = True_anomaly_moon + Perigee_moon
#计算系数
A_moon = cos(RAAN - RAAN_moon) * cos(u_moon) + sin(RAAN - RAAN_moon) * sin(
u_moon) * cos(Inclination_moon)
B1_moon = cos(Inclination)
B2_moon = B1_moon * (-sin(RAAN - RAAN_moon) * cos(u_moon) + sin(u_moon) *
cos(Inclination_moon) * cos(RAAN - RAAN_moon))
B_moon = B2_moon + sin(Inclination) * sin(Inclination_moon) * sin(u_moon)
C1_moon = sin(Inclination)
C2_moon = C1_moon * (sin(RAAN - RAAN_moon) * cos(u_moon) - sin(u_moon) *
cos(Inclination_moon) * cos(RAAN - RAAN_moon))
C_moon = C2_moon + cos(Inclination) * sin(Inclination_moon) * sin(u_moon)
d_semi_major_axis_lunar = 0
d_Eccentricity_lunar=-(15*K_moon/2/n)*Eccentricity*sqrt(1-Eccentricity**2)*\
(A_moon*B_moon*cos(2*Perigee)-(1/2)*(A_moon**2-B_moon**2)*sin(2*Perigee))
d_Inclination_lunar=3*K_moon*C_moon/4/n/sqrt(1-Eccentricity**2)*(A_moon*(2+3*Eccentricity**2+\
5*Eccentricity**2*cos(2*Perigee))+5*B_moon*Eccentricity**2*sin(2*Perigee))
d_RAAN_lunar = 3*K_moon*C_moon/4/n/sqrt(1-Eccentricity**2)/sin(Inclination)*(B_moon*(2+3*Eccentricity**2+\
5*Eccentricity**2*cos(2*Perigee))+5*A_moon*Eccentricity**2*sin(2*Perigee))
d_Perigee_lunar = 3*K_moon/2/n*sqrt(1-Eccentricity**2)*((5*A_moon*B_moon*sin(Perigee*2)+\
1/2*(A_moon**2-B_moon**2)*cos(2*Perigee))-1+3/2*(A_moon**2+B_moon**2))\
+5*semi_major_axis/2/Eccentricity/r_moon*(1-5/4*(A_moon**2-B_moon**2))*(A_moon*cos(Perigee)\
+B_moon*sin(Perigee))-cos(Inclination)*d_RAAN_lunar
d_Mean_anomaly_lunar = 0
d_semi_major_axis = d_semi_major_axis_earth_nonsphericfigure + d_semi_major_axis_lunar + d_semi_major_axis_solar
d_Eccentricity = d_Eccentricity_earth_nonsphericfigure + d_Eccentricity_lunar + d_Eccentricity_solar
d_Inclination = d_Inclination_earth_nonsphericfigure + d_Inclination_solar + d_Inclination_lunar
d_RAAN = d_RAAN_earth_nonsphericfigure + d_Inclination_lunar + d_RAAN_solar
d_Perigee = d_Perigee_earth_nonsphericfigure + d_Perigee_lunar + d_Perigee_solar
d_Mean_anomaly = n
semi_major_axis_earth_nonsphericfigure = 1
semi_major_axis_lunar_solar = 0
Eccentricity_earth_nonsphericfigure = 1
Eccentricity_lunar_solar = 0
Inclination_earth_nonsphericfigure = 1
Inclination_solar_solar = 0
RAAN_earth_nonsphericfigure = 1
RAAN_lunar_solar = 0
Perigee_earth_nonsphericfigure = 1
Perigee_lunar_solar = 0
return np.array([
d_semi_major_axis, d_Eccentricity, d_Inclination, d_RAAN, d_Perigee,
d_Mean_anomaly
])
RA_measured = array([[20, 56, 08.762], [20, 59, 20.410], [21, 8, 43.156]])
Dec_measured = array([[16, 29, 32.52], [18, 14, 10.95], [22, 28, 36.27]])
t1 = JD[0]
t2 = JD[1]
t3 = JD[2]
list_test = [JD[0], JD[2]]
tau2 = 0
tau1 = k * (t1 - t2)
tau3 = k * (t3 - t2)
R = []
for i in JD:
eph = Ephemeris(de421)
barycenter = eph.position('earthmoon', i)
moonvector = eph.position('moon', i)
earth = barycenter - moonvector * eph.earth_share
R0 = vector(
earth
) / AU # This is the Sun to Earth center vector in Equatorial system
R_geocentric = -1. * R0
R.append(array(R_geocentric))
print 'R_geocentric', R
print
def RAconverter(hrs, minutes, seconds):
return (hrs / 24. + minutes / (24. * 60) + seconds / (24. * 3600)) * 360.
# import os
from skyfield.api import *
years = range(-9999, 9999, 10)
north_pole = earth.topos('90 N', '0 W')
print(north_pole)
x, y, z = north_pole.gcrs(utc=(years, 1, 1)).position.km
print(x,y,z)
plot(x, y)
exit()
# from skyfield.api import *
import de421
from jplephem import Ephemeris
eph = Ephemeris(de421)
# x, y, z = eph.position('saturn', 2456656.501)
# x, y, z = eph.position('mars', 2456658.500000000);
# x, y, z = eph.position('mercury', 2414992.5);
x,y,z = eph.position('mercury', 2456876) - eph.position('jupiter', 2456876)
print x,y,z
for t in range (2456876,2456876+1000):
mx,my,mz = eph.position('mercury', t)
jx,jy,jz = eph.position('jupiter', t)
eq = -0.0145987218963993*(jx-mx)-0.430326550289791*(jy-my)+0.902555226805917*(jz-mz)
print "ALPHA",t,eq
exit()