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)
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 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()
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 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 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 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 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 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 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 = []
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 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)):
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
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
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 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')
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
])
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 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='-.')