def orbit_impulse(
orbit, vector
): #alter an obejcts orbit due an impulse (used for debris projectile collisions)
dv = [vector[0].value, vector[1].to(vector[0].unit).value, 0
] * vector[0].unit
man = Maneuver.impulse(dv)
return orbit.apply_maneuver(man)
def Earth_to_Jupiter(delta_v): """ Calculates the manuever and the launch windows from Earth to Jupiter, assuming tangential impulsive maneuver from Earth orbit around the Sun Earth_to_Jupiter(dv): 'dv' is given with velocity units (u.m/u.s) Returns: ss_transfer = trasfer orbit object t = time to the first next launch window T = period between possible launch windows t_transfer = time of transfer from launch to Jupiter arrival """ dv = delta_v * v_Earth_dir # Initial transfer orbit calculation (to obtain transfer time) man_tmp = Maneuver.impulse(dv) transfer = ss_Earth.apply_maneuver(man_tmp) # Angles calculation # nu_Jupiter: nu of transfer orbit where it crosses Jupiter nu_Jupiter = (np.math.acos((transfer.p/R_J - 1)/transfer.ecc)*u.rad) ang_vel_Earth = (norm(v_Earth).value / R_E.value)*u.rad/u.s ang_vel_Jupiter = (norm(v_Jupiter).value / R_J.value)*u.rad/u.s # Transfer time M = nu_to_M(nu_Jupiter, transfer.ecc) t_transfer = M/transfer.n # Delta_nu at launch (Final nu - nu that Jupiter walks during transfer) delta_nu_at_launch = nu_Jupiter - ang_vel_Jupiter * t_transfer # Angles in radians nu_init = angle(r_Earth, r_Jupiter)*u.rad # Next launch window in t: t = (nu_init - delta_nu_at_launch) / (ang_vel_Earth - ang_vel_Jupiter) # Time between windows T = (2*np.pi*u.rad) / (ang_vel_Earth - ang_vel_Jupiter) # If time to the windows is negative (in the past), we should wait until next window while t < 0: t = t + T # Earth at launch ss_Earth_launch = ss_Earth.propagate(t) launch_dir = ss_Earth_launch.v / norm(ss_Earth_launch.v) dv = delta_v * launch_dir # Maneuver calculation man = Maneuver((t, dv)) ss_transfer = ss_Earth.apply_maneuver(man) return ss_transfer, t, T, t_transfer
def generate_fitness_score(w1, w2, w3, w4, delt_v_size, test_orb, seq):
num_vs = 0
for man in seq:
if man == '1':
test_orb = test_orb.apply_maneuver(
Maneuver.impulse((delt_v_size, 0, 0) * u.m / u.s))
num_vs += 1
if man == '2':
test_orb = test_orb.apply_maneuver(
Maneuver.impulse((0, delt_v_size, 0) * u.m / u.s))
num_vs += 1
if man == '3':
test_orb = test_orb.apply_maneuver(
Maneuver.impulse((0, 0, delt_v_size) * u.m / u.s))
num_vs += 1
if man == '4':
test_orb = test_orb.apply_maneuver(
Maneuver.impulse((-delt_v_size, 0, 0) * u.m / u.s))
num_vs += 1
if man == '5':
test_orb = test_orb.apply_maneuver(
Maneuver.impulse((0, -delt_v_size, 0) * u.m / u.s))
num_vs += 1
if man == '6':
test_orb = test_orb.apply_maneuver(
Maneuver.impulse((0, 0, -delt_v_size) * u.m / u.s))
num_vs += 1
test_orb = test_orb.propagate(20 * u.min)
return np.abs(float(test_orb.inc / u.rad - 1.57)) * w1 + float(
test_orb.ecc) * w2 + np.abs(float(
test_orb.raan / u.rad)) * w3 + np.abs(
float(test_orb.a / u.km - 15000) / 15000) * w4
def test_maneuver_impulse():
dv = [1, 0, 0] * u.m / u.s
man = Maneuver.impulse(dv)
assert man.impulses[0] == (0 * u.s, dv)
def maneuver(self, dv): man = Maneuver.impulse(dv*u.km/u.s) self.orbit = self.orbit.apply_maneuver(man)
#This is a plot of all of the initial spacecraft orbits
op = OrbitPlotter3D()
op.plot(tests[0].orbit)
op.plot(tests[1].orbit)
op.plot(tests[2].orbit)
op.plot(tests[3].orbit)
#This is a plot of the final spacecraft orbits, after applying the fittest maneuver strings for each of them.
f_orbs = [test.orbit for test in tests]
delt_v_size = 500
for i in range(0,len(f_orbs)):
for man in fittest[i][0]:
if man == '1':
f_orbs[i] = f_orbs[i].apply_maneuver(Maneuver.impulse((delt_v_size,0,0)*u.m/u.s))
if man == '2':
f_orbs[i] = f_orbs[i].apply_maneuver(Maneuver.impulse((0,delt_v_size,0)*u.m/u.s))
if man == '3':
f_orbs[i] = f_orbs[i].apply_maneuver(Maneuver.impulse((0,0,delt_v_size)*u.m/u.s))
if man == '4':
f_orbs[i] = f_orbs[i].apply_maneuver(Maneuver.impulse((-delt_v_size,0,0)*u.m/u.s))
if man == '5':
f_orbs[i] = f_orbs[i].apply_maneuver(Maneuver.impulse((0,-delt_v_size,0)*u.m/u.s))
if man == '6':
f_orbs[i] = f_orbs[i].apply_maneuver(Maneuver.impulse((0,0,-delt_v_size)*u.m/u.s))
f_orbs[i] = f_orbs[i].propagate(20*u.min)
op = OrbitPlotter3D()
op.plot(f_orbs[0])
op.plot(f_orbs[1])
iss_30m = iss.propagate(30 * u.min) iss_30m.epoch iss_30m.nu.to(u.deg) #plot (iss_30m) earth = Orbit.from_body_ephem(Earth) #plot(earth) earth_30d = earth.propagate(30 * u.day) #plot(earth_30d) from poliastro.maneuver import Maneuver dv = [5, 0, 0] * u.m / u.s man = Maneuver.impulse(dv) man = Maneuver((0 * u.s, dv)) ss_i = Orbit.circular(Earth, alt=700 * u.km) ss_i #plot(ss_i) hoh = Maneuver.hohmann(ss_i, 36000 * u.km) hoh.get_total_cost() hoh.get_total_time() print(hoh.get_total_cost()) print(hoh.get_total_time()) hoh.impulses[0] hoh[0] tuple(val.decompose([u.km, u.s]) for val in hoh[1])
def test_maneuver_impulse():
dv = [1, 0, 0] * u.m / u.s
man = Maneuver.impulse(dv)
assert man.impulses[0] == (0 * u.s, dv)
def test_apply_manuever_correct_plane():
ceres = Orbit.from_sbdb("Ceres")
imp = Maneuver.impulse([0, 0, 0] * u.km / u.s)
new_ceres = ceres.apply_maneuver(imp)
assert ceres.plane == Planes.EARTH_ECLIPTIC
assert new_ceres.plane == ceres.plane
