for i in range(0, number_RAAN):
bodies.create_empty_body(SC_NAMES[i])
bodies.get_body(SC_NAMES[i]).set_constant_mass(4.874)
environment_setup.add_radiation_pressure_interface(
bodies, SC_NAMES[i], radiation_pressure_settings)
central_bodies.append("Earth")
acceleration_settings[SC_NAMES[i]] = accelerations_settings_chess
initial_state_temp = conversion.keplerian_to_cartesian(
gravitational_parameter=earth_gravitational_parameter,
semi_major_axis=7078136,
eccentricity=0.0424,
inclination=np.deg2rad(97.5926),
argument_of_periapsis=np.deg2rad(RAAN[i]),
longitude_of_ascending_node=np.deg2rad(100.0),
true_anomaly=np.deg2rad(0.0))
if i == 0:
initial_state = initial_state_temp
else:
initial_state = np.hstack([initial_state, initial_state_temp])
dependent_variables_to_save.append(
propagation_setup.dependent_variable.single_acceleration_norm(
propagation_setup.acceleration.cannonball_radiation_pressure_type,
SC_NAMES[i], "Sun"))
acceleration_models = propagation_setup.create_acceleration_models(
bodies, acceleration_settings, bodies_to_propagate, central_bodies)
Saturn=[propagation_setup.acceleration.point_mass_gravity()],
Uranus=[propagation_setup.acceleration.point_mass_gravity()],
Neptune=[propagation_setup.acceleration.point_mass_gravity()],
Pluto=[propagation_setup.acceleration.point_mass_gravity()])
acceleration_settings = {"CubeSat": accelerations_settings_3u_cubesat}
acceleration_models = propagation_setup.create_acceleration_models(
bodies, acceleration_settings, bodies_to_propagate, central_bodies)
# # Create propagation settings
# define initial conditions
sun_gravitational_parameter = bodies.get_body("Sun").gravitational_parameter
initial_state = conversion.keplerian_to_cartesian(
gravitational_parameter=sun_gravitational_parameter,
semi_major_axis=1.3 * 149597870700.0,
eccentricity=0.1,
inclination=np.deg2rad(1.0),
argument_of_periapsis=np.deg2rad(235.7),
longitude_of_ascending_node=np.deg2rad(23.4),
true_anomaly=np.deg2rad(139.87))
# define the variables we need
dependent_variables_to_save = [
propagation_setup.dependent_variable.total_acceleration("CubeSat"),
propagation_setup.dependent_variable.keplerian_state("CubeSat", "Sun"),
]
# create propagation settings
propagator_settings = propagation_setup.propagator.translational(
central_bodies,
acceleration_models,
bodies_to_propagate,
initial_state,
simulation_end_epoch,
def fitness(self,
orbit_parameters: List[float]) -> List[float]:
"""
Computes the fitness value for the problem.
Parameters
----------
orbit_parameters : List[float]
Vector of decision variables of size n (for this problem, n = 4).
Returns
-------
List[float]
List of size p with the values for each objective (for this multi-objective optimization problem, p=2).
"""
# Retrieves system of bodies
current_bodies = self.bodies_function()
# Retrieves Itokawa gravitational parameter
itokawa_gravitational_parameter = current_bodies.get_body("Itokawa").gravitational_parameter
# Reset the initial state from the decision variable vector
new_initial_state = conversion.keplerian_to_cartesian(
gravitational_parameter=itokawa_gravitational_parameter,
semi_major_axis=orbit_parameters[0],
eccentricity=orbit_parameters[1],
inclination=np.deg2rad(orbit_parameters[2]),
argument_of_periapsis=np.deg2rad(235.7),
longitude_of_ascending_node=np.deg2rad(orbit_parameters[3]),
true_anomaly=np.deg2rad(139.87))
# Retrieves propagator settings object
propagator_settings = self.propagator_settings_function()
# Retrieves integrator settings object
integrator_settings = self.integrator_settings_function()
# Reset the initial state
propagator_settings.reset_initial_states(new_initial_state)
# Propagate orbit
dynamics_simulator = propagation_setup.SingleArcDynamicsSimulator(current_bodies,
integrator_settings,
propagator_settings,
print_dependent_variable_data=False)
# Update dynamics simulator function
self.dynamics_simulator_function = lambda: dynamics_simulator
# Retrieve dependent variable history
dependent_variables = dynamics_simulator.dependent_variable_history
dependent_variables_list = np.vstack(list(dependent_variables.values()))
# Retrieve distance
distance = dependent_variables_list[:, 0]
# Retrieve latitude
latitudes = dependent_variables_list[:, 1]
# Compute mean latitude
mean_latitude = np.mean(np.absolute(latitudes))
# Computes fitness as mean latitude
current_fitness = 1.0 / mean_latitude
# Exaggerate fitness value if the spacecraft has broken out of the selected distance range
current_penalty = 0.0
if (max(dynamics_simulator.dependent_variable_history.keys()) < self.mission_final_time):
current_penalty += 1.0E4
return [current_fitness + current_penalty, np.mean(distance) + current_penalty * 1.0E3]
accelerations_settings_chess = dict(
Earth=[propagation_setup.acceleration.spherical_harmonic_gravity(4, 4)])
acceleration_settings = {"Spacecraft": accelerations_settings_chess}
acceleration_models = propagation_setup.create_acceleration_models(
bodies, acceleration_settings, bodies_to_propagate, central_bodies)
# Initial System State
earth_gravitational_parameter = bodies.get_body(
"Earth").gravitational_parameter
initial_state = conversion.keplerian_to_cartesian(
gravitational_parameter=earth_gravitational_parameter,
semi_major_axis=6928136,
eccentricity=0.0,
inclination=np.deg2rad(97.5919),
argument_of_periapsis=np.deg2rad(0.0),
longitude_of_ascending_node=np.deg2rad(90.0),
true_anomaly=np.deg2rad(0.0))
# Variables to Export
dependent_variables_to_save = [
propagation_setup.dependent_variable.relative_position(
"Spacecraft", "Earth")
]
# Propagator settings
propagator_settings = propagation_setup.propagator.translational(
central_bodies,
acceleration_models,
bodies_to_propagate,