def planet_asteroid(start_planet, target_name, tlaunch, tarrive, rev, N):
# Create PyKEP epoch objects and calculate flight time
t1 = epoch(tlaunch)
t2 = epoch(tarrive)
dt = (tarrive - tlaunch) * DAY2SEC
import py.AsteroidDB as asteroidDB
neo_db = asteroidDB.neo
target = (item for item in neo_db if item["name"] == target_name).next()
ep = epoch(target["epoch_mjd"], epoch.epoch_type.MJD)
a = target["a"] * AU
e = target["e"]
i = target["i"] * DEG2RAD
om = target["om"] * DEG2RAD
w = target["w"] * DEG2RAD
ma = target["ma"] * DEG2RAD
as_mu = 1E17 * 6.67384E-11 # maybe need to calculate actual mass from density and radius
r = (target["diameter"] / 2) * 1000
sr = r * 1.1
OBJ2 = planet(ep, (a, e, i, om, w, ma), MU_SUN, as_mu, r, sr)
OBJ1 = planet_ss(start_planet) # Calculate location of objects in flight path
r1, v1 = OBJ1.eph(t1)
r2, v2 = OBJ2.eph(t2)
# Find trajectory
l = lambert_problem(r1, r2, dt, MU_SUN)
#extract relevant information from solution
r = l.get_r1()
v = l.get_v1()[rev]
mu = l.get_mu()
#define the integration time
dtn = dt / (N - 1)
dtn_days = dtn * SEC2DAY
#alocate the cartesian components for r
t = np.array([0.0] * N)
x = np.array([0.0] * N)
y = np.array([0.0] * N)
z = np.array([0.0] * N)
#calculate the spacecraft position at each dt
for i in range(N):
t[i] = tlaunch + dtn_days * i
x[i] = r[0] / AU
y[i] = r[1] / AU
z[i] = r[2] / AU
r, v = propagate_lagrangian(r, v, dtn, mu)
#traj = [t, x, y, z]
vin = l.get_v1()[rev]
vout = l.get_v2()[rev]
#dV=fb_vel(vin,vout,planet_ss(arrive_planet))
#dV=np.sqrt( np.square(vin[0]/vout[0])+np.square(vin[1]/vout[1])+np.square(vin[2]/vout[2]))
#dV=np.sqrt( np.square(vin[0]-v1[0])+np.square(v1[1]-vin[1])+np.square(v1[2]-vin[2]))
#dV=np.sqrt( np.square(v2[0]-vout[0])+np.square(v2[1]-vout[1])+np.square(v2[2]-vout[2]))
#dV=np.sqrt( np.square(v1[0]/vin[0])+np.square(v1[1]/vin[1])+np.square(v1[2]/vin[2]))
C3_launch = (np.sqrt(np.square(vin[0] - v1[0]) + np.square(vin[1] - v1[1]) + np.square(vin[2] - v1[2]))) ** 2
C3_arrive = (np.sqrt(np.square(vout[0] - v2[0]) + np.square(vout[1] - v2[1]) + np.square(vout[2] - v2[2]))) ** 2
C3 = np.sqrt((C3_arrive ** 2) + (C3_launch ** 2))
return C3
def traj_planet_asteroid(source, dest, tlaunch, tarrive, rev, N):
t1 = epoch(tlaunch)
t2 = epoch(tarrive)
dt = (tarrive - tlaunch) * DAY2SEC
target = source['orbit']
ep = epoch(jd_to_mjd(tlaunch), epoch.epoch_type.MJD)
a = target["a"] * AU
e = target["e"]
i = target["i"] * DEG2RAD
om = target["om"] * DEG2RAD
w = target["w"] * DEG2RAD
ma = target["ma"] * DEG2RAD
as_mu = 1E17 * 6.67384E-11 # maybe need to calculate actual mass from density and radius
r = (10 / 2) * 1000
sr = r * 1.1
OBJ1 = planet(ep, (a, e, i, om, w, ma), MU_SUN, as_mu, r, sr)
target = dest['orbit']
ep = epoch(jd_to_mjd(tarrive), epoch.epoch_type.MJD)
a = target["a"] * AU
e = target["e"]
i = target["i"] * DEG2RAD
om = target["om"] * DEG2RAD
w = target["w"] * DEG2RAD
ma = target["ma"] * DEG2RAD
as_mu = 1E17 * 6.67384E-11 # maybe need to calculate actual mass from density and radius
r = (10 / 2) * 1000
sr = r * 1.1
OBJ2 = planet(ep, (a, e, i, om, w, ma), MU_SUN, as_mu, r, sr)
# Calculate location of objects in flight path
r1, v1 = OBJ1.eph(t1)
r2, v2 = OBJ2.eph(t2)
#Find trajectory
l = lambert_problem(r1, r2, dt, MU_SUN)
#extract relevant information from solution
r = l.get_r1()
v = l.get_v1()[0]
mu = l.get_mu() #define the integration time
dtn = dt / (N - 1)
dtn_days = dtn * SEC2DAY
#alocate the cartesian components for r
t = np.array([0.0] * N)
x = np.array([0.0] * N)
y = np.array([0.0] * N)
z = np.array([0.0] * N)
#calculate the spacecraft position at each dt
for i in range(N):
t[i] = tlaunch + dtn_days * i
x[i] = r[0] / AU
y[i] = r[1] / AU
z[i] = r[2] / AU
r, v = propagate_lagrangian(r, v, dtn, mu)
traj = [t.tolist(), x.tolist(), y.tolist(), z.tolist()]
return traj
def planet_asteroid(start_planet, target_name, tlaunch, tarrive, rev, N):
# Create PyKEP epoch objects and calculate flight time
t1 = epoch(tlaunch)
t2 = epoch(tarrive)
dt = (tarrive - tlaunch) * DAY2SEC
import py.AsteroidDB as asteroidDB
neo_db = asteroidDB.neo
target = (item for item in neo_db if item["name"] == target_name).next()
ep = epoch(target["epoch_mjd"], epoch.epoch_type.MJD)
a = target["a"] * AU
e = target["e"]
i = target["i"] * DEG2RAD
om = target["om"] * DEG2RAD
w = target["w"] * DEG2RAD
ma = target["ma"] * DEG2RAD
as_mu = 1E17 * 6.67384E-11 # maybe need to calculate actual mass from density and radius
r = (target["diameter"] / 2) * 1000
sr = r * 1.1
OBJ2 = planet(ep, (a, e, i, om, w, ma), MU_SUN, as_mu, r, sr)
OBJ1 = planet_ss(
start_planet) # Calculate location of objects in flight path
r1, v1 = OBJ1.eph(t1)
r2, v2 = OBJ2.eph(t2)
# Find trajectory
l = lambert_problem(r1, r2, dt, MU_SUN)
#extract relevant information from solution
r = l.get_r1()
v = l.get_v1()[rev]
mu = l.get_mu()
#define the integration time
dtn = dt / (N - 1)
dtn_days = dtn * SEC2DAY
#alocate the cartesian components for r
t = np.array([0.0] * N)
x = np.array([0.0] * N)
y = np.array([0.0] * N)
z = np.array([0.0] * N)
#calculate the spacecraft position at each dt
for i in range(N):
t[i] = tlaunch + dtn_days * i
x[i] = r[0] / AU
y[i] = r[1] / AU
z[i] = r[2] / AU
r, v = propagate_lagrangian(r, v, dtn, mu)
#traj = [t, x, y, z]
vin = l.get_v1()[rev]
vout = l.get_v2()[rev]
#dV=fb_vel(vin,vout,planet_ss(arrive_planet))
#dV=np.sqrt( np.square(vin[0]/vout[0])+np.square(vin[1]/vout[1])+np.square(vin[2]/vout[2]))
#dV=np.sqrt( np.square(vin[0]-v1[0])+np.square(v1[1]-vin[1])+np.square(v1[2]-vin[2]))
#dV=np.sqrt( np.square(v2[0]-vout[0])+np.square(v2[1]-vout[1])+np.square(v2[2]-vout[2]))
#dV=np.sqrt( np.square(v1[0]/vin[0])+np.square(v1[1]/vin[1])+np.square(v1[2]/vin[2]))
C3_launch = (np.sqrt(
np.square(vin[0] - v1[0]) + np.square(vin[1] - v1[1]) +
np.square(vin[2] - v1[2])))**2
C3_arrive = (np.sqrt(
np.square(vout[0] - v2[0]) + np.square(vout[1] - v2[1]) +
np.square(vout[2] - v2[2])))**2
C3 = np.sqrt((C3_arrive**2) + (C3_launch**2))
return C3
def traj_planet_asteroid(source, dest, tlaunch, tarrive, rev, N):
t1 = epoch(tlaunch)
t2 = epoch(tarrive)
dt = (tarrive - tlaunch) * DAY2SEC
target = source['orbit']
ep = epoch(jd_to_mjd(tlaunch), epoch.epoch_type.MJD)
a = target["a"] * AU
e = target["e"]
i = target["i"] * DEG2RAD
om = target["om"] * DEG2RAD
w = target["w"] * DEG2RAD
ma = target["ma"] * DEG2RAD
as_mu = 1E17 * 6.67384E-11 # maybe need to calculate actual mass from density and radius
r = (10 / 2) * 1000
sr = r * 1.1
OBJ1 = planet(ep, (a, e, i, om, w, ma), MU_SUN, as_mu, r, sr)
target = dest['orbit']
ep = epoch(jd_to_mjd(tarrive), epoch.epoch_type.MJD)
a = target["a"] * AU
e = target["e"]
i = target["i"] * DEG2RAD
om = target["om"] * DEG2RAD
w = target["w"] * DEG2RAD
ma = target["ma"] * DEG2RAD
as_mu = 1E17 * 6.67384E-11 # maybe need to calculate actual mass from density and radius
r = (10 / 2) * 1000
sr = r * 1.1
OBJ2 = planet(ep, (a, e, i, om, w, ma), MU_SUN, as_mu, r, sr)
# Calculate location of objects in flight path
r1, v1 = OBJ1.eph(t1)
r2, v2 = OBJ2.eph(t2)
#Find trajectory
l = lambert_problem(r1, r2, dt, MU_SUN)
#extract relevant information from solution
r = l.get_r1()
v = l.get_v1()[0]
mu = l.get_mu() #define the integration time
dtn = dt / (N - 1)
dtn_days = dtn * SEC2DAY
#alocate the cartesian components for r
t = np.array([0.0] * N)
x = np.array([0.0] * N)
y = np.array([0.0] * N)
z = np.array([0.0] * N)
#calculate the spacecraft position at each dt
for i in range(N):
t[i] = tlaunch + dtn_days * i
x[i] = r[0] / AU
y[i] = r[1] / AU
z[i] = r[2] / AU
r, v = propagate_lagrangian(r, v, dtn, mu)
traj = [t.tolist(), x.tolist(), y.tolist(), z.tolist()]
return traj
def getTraj(K1, K2, tlaunch, tarrive, N):
'''
Finds a trajectory between two objects orbiting the Sun
USAGE: traj = getTraj(K1, K2, tlaunch, tarrive)
K: array of object parameters.
epoch: epoch of Keplerian orbital elements (JD)
a: semimajor axis (AU)
e: eccentricity (none)
i: inclination (deg)
om: longitude of the ascending node (deg)
w: argument of perihelion (deg)
ma: mean anomaly at epoch (deg)
mass: mass of object (kg)
r: radius of object (m)
sr: safe radius to approach object (m)
K1: [epoch1,a1,e1,i1,om1,w1,ma1,mass1,r1,sr1]
K2: [epoch2,a2,e2,i2,om2,w2,ma2,mass2,r2,sr2]
tlaunch: launch time (JD)
tarrive: arrival time (JD)
N: number of points in calculated trajectory
'''
import numpy as np
from PyKEP import epoch, DAY2SEC, SEC2DAY, AU, DEG2RAD, MU_SUN, planet, lambert_problem, propagate_lagrangian
#Create PyKEP epoch objects and calculate flight time
t1 = epoch(tlaunch, epoch.epoch_type.JD)
t2 = epoch(tarrive, epoch.epoch_type.JD)
dt = (t2.mjd2000 - t1.mjd2000) * DAY2SEC
#First object
K1[0] = epoch(K1[0],
epoch.epoch_type.JD) #convert epoch to PyKEP epoch object
K1[1] = K1[1] * AU #convert AU to meters
K1[3] = K1[3] * DEG2RAD #convert angles from degrees to radians
K1[4] = K1[4] * DEG2RAD
K1[5] = K1[5] * DEG2RAD
K1[6] = K1[6] * DEG2RAD
K1[7] = K1[7] * 6.67384E-11 #convert mass to gravitational parameter mu
OBJ1 = planet(K1[0], K1[1:7], MU_SUN, K1[7], K1[8], K1[9])
#Second object
K2[0] = epoch(K2[0],
epoch.epoch_type.JD) #convert epoch to PyKEP epoch object
K2[1] = K2[1] * AU #convert AU to meters
K2[3] = K2[3] * DEG2RAD #convert angles from degrees to radians
K2[4] = K2[4] * DEG2RAD
K2[5] = K2[5] * DEG2RAD
K2[6] = K2[6] * DEG2RAD
K2[7] = K2[7] * 6.67384E-11 #convert mass to gravitational parameter mu
OBJ2 = planet(K2[0], K2[1:7], MU_SUN, K2[7], K2[8], K2[9])
#Calculate location of objects in flight path
r1, v1 = OBJ1.eph(t1)
r2, v2 = OBJ2.eph(t2)
#Find trajectory
l = lambert_problem(r1, r2, dt, MU_SUN)
#extract relevant information from solution
r = l.get_r1()
v = l.get_v1()[0]
mu = l.get_mu()
#define the integration time
dtn = dt / (N - 1)
dtn_days = dtn * SEC2DAY
#alocate the cartesian components for r
t = np.array([0.0] * N)
x = np.array([0.0] * N)
y = np.array([0.0] * N)
z = np.array([0.0] * N)
#calculate the spacecraft position at each dt
for i in range(N):
t[i] = tlaunch + dtn_days * i
x[i] = r[0] / AU
y[i] = r[1] / AU
z[i] = r[2] / AU
r, v = propagate_lagrangian(r, v, dtn, mu)
traj = [t, x, y, z]
return traj
def getTraj(K1, K2, tlaunch, tarrive, N): ''' Finds a trajectory between two objects orbiting the Sun USAGE: traj = getTraj(K1, K2, tlaunch, tarrive) K: array of object parameters. epoch: epoch of Keplerian orbital elements (JD) a: semimajor axis (AU) e: eccentricity (none) i: inclination (deg) om: longitude of the ascending node (deg) w: argument of perihelion (deg) ma: mean anomaly at epoch (deg) mass: mass of object (kg) r: radius of object (m) sr: safe radius to approach object (m) K1: [epoch1,a1,e1,i1,om1,w1,ma1,mass1,r1,sr1] K2: [epoch2,a2,e2,i2,om2,w2,ma2,mass2,r2,sr2] tlaunch: launch time (JD) tarrive: arrival time (JD) N: number of points in calculated trajectory ''' import numpy as np from PyKEP import epoch, DAY2SEC, SEC2DAY, AU, DEG2RAD, MU_SUN, planet, lambert_problem, propagate_lagrangian #Create PyKEP epoch objects and calculate flight time t1 = epoch(tlaunch,epoch.epoch_type.JD) t2 = epoch(tarrive,epoch.epoch_type.JD) dt = (t2.mjd2000 - t1.mjd2000) * DAY2SEC #First object K1[0] = epoch(K1[0],epoch.epoch_type.JD) #convert epoch to PyKEP epoch object K1[1] = K1[1] * AU #convert AU to meters K1[3] = K1[3] * DEG2RAD #convert angles from degrees to radians K1[4] = K1[4] * DEG2RAD K1[5] = K1[5] * DEG2RAD K1[6] = K1[6] * DEG2RAD K1[7] = K1[7] * 6.67384E-11 #convert mass to gravitational parameter mu OBJ1 = planet(K1[0], K1[1:7], MU_SUN, K1[7], K1[8], K1[9]) #Second object K2[0] = epoch(K2[0],epoch.epoch_type.JD) #convert epoch to PyKEP epoch object K2[1] = K2[1] * AU #convert AU to meters K2[3] = K2[3] * DEG2RAD #convert angles from degrees to radians K2[4] = K2[4] * DEG2RAD K2[5] = K2[5] * DEG2RAD K2[6] = K2[6] * DEG2RAD K2[7] = K2[7] * 6.67384E-11 #convert mass to gravitational parameter mu OBJ2 = planet(K2[0], K2[1:7], MU_SUN, K2[7], K2[8], K2[9]) #Calculate location of objects in flight path r1, v1 = OBJ1.eph(t1) r2, v2 = OBJ2.eph(t2) #Find trajectory l = lambert_problem(r1, r2, dt, MU_SUN) #extract relevant information from solution r = l.get_r1() v = l.get_v1()[0] mu = l.get_mu() #define the integration time dtn = dt / (N - 1) dtn_days = dtn * SEC2DAY #alocate the cartesian components for r t = np.array([0.0] * N) x = np.array([0.0] * N) y = np.array([0.0] * N) z = np.array([0.0] * N) #calculate the spacecraft position at each dt for i in range(N): t[i] = tlaunch + dtn_days * i x[i] = r[0] / AU y[i] = r[1] / AU z[i] = r[2] / AU r, v = propagate_lagrangian(r, v, dtn, mu) traj = [t, x, y, z] return traj
