def main():
# Load spice kernels.
spice_interface.load_standard_spice_kernels()
# Set simulation start epoch.
simulation_start_epoch = 0.0
# Set numerical integration fixed step size.
fixed_step_size = 1.0
# Set simulation end epoch.
simulation_end_epoch = constants.JULIAN_DAY
###########################################################################
# CREATE ENVIRONMENT ######################################################
###########################################################################
bodies_to_create = ["Sun", "Earth", "Moon", "Mars", "Venus"]
env = Environment(bodies_to_create,
user_defined_bodies=None,
frame_origin="SSB",
frame_orientation="ECLIPJ2000",
custom_body_settings=None,
alternative_kernels=None,
start_epoch=simulation_start_epoch,
end_epoch=simulation_end_epoch,
fixed_dt=fixed_step_size,
epoch_margin=300.0)
###########################################################################
# CREATE VEHICLE ##########################################################
###########################################################################
env.vehicles = ["Delfi-C3"]
env.bodies["Delfi-C3"].set_constant_body_mass(400.0)
###########################################################################
# CREATE VEHICLE - ENVIRONMENT INTERFACE ##################################
###########################################################################
# Create aerodynamic interface settings for Delfi-C3.
env.set_aerodynamic_coefficient_interface(
target_body_name="Delfi-C3",
reference_area=4.0,
coefficients=1.2 * np.ones(3),
coefficients_in_aerodynamic_frame=True,
coefficients_in_negative_axis_direction=True,
model="constant_coefficients")
# Create radiation pressure interface settings for Delfi-C3.
env.set_radiation_pressure_interface(target_body_name="Delfi-C3",
reference_area=4.0,
coefficient=1.2,
occulting=["Earth"],
source="Sun",
model="cannon_ball")
###########################################################################
# FINALIZE BODIES #########################################################
###########################################################################
# Finalize body creation.
env.finalize_environment(frame_origin="SSB", frame_orientation="J2000")
###########################################################################
# CREATE ACCELERATIONS ####################################################
###########################################################################
# Define bodies that are propagated.
bodies_to_propagate = env.vehicles
# Define central bodies.
central_bodies = ["Earth"]
# Define unique (Sun, Earth) accelerations acting on Delfi-C3.
accelerations_of_delfi_c3 = dict(
Sun=[
simulation_setup.Acceleration.canon_ball_radiation_pressure()
# AccelerationSettings(AvailableAcceleration.cannon_ball_radiation_pressure) # LEGACY DESIGN.
],
Earth=[
simulation_setup.Acceleration.spherical_harmonic_gravity(5, 5),
# SphericalHarmonicAccelerationSettings(5, 5), # LEGACY DESIGN.
simulation_setup.Acceleration.aerodynamic()
# AccelerationSettings(AvailableAcceleration.aerodynamic) # LEGACY DESIGN.
])
# Define other point mass accelerations acting on Delfi-C3.
for other in set(bodies_to_create).difference({"Sun", "Earth"}):
accelerations_of_delfi_c3[other] = [
simulation_setup.Acceleration.point_mass_gravity()
]
# Create global accelerations dictionary.
acceleration_dict = {"Delfi-C3": accelerations_of_delfi_c3}
# Create acceleration models.
acceleration_models = simulation_setup.create_acceleration_models_dict(
bodies_to_create, acceleration_dict, bodies_to_propagate,
central_bodies)
###########################################################################
# CREATE PROPAGATION SETTINGS #############################################
###########################################################################
# Set initial conditions for the Asterix satellite that will be
# propagated in this simulation. The initial conditions are given in
# Keplerian elements and later on converted to Cartesian elements.
# Set Keplerian elements for Delfi-C3
earth_gravitational_parameter = env.bodies[
"Earth"].gravity_field_model.get_gravitational_parameter()
# LEGACY DESIGN.
# KEI = orbital_element_conversions.KeplerianElementIndices
# asterix_initial_state_in_keplerian_elements = np.zeros(6)
# kep_state = asterix_initial_state_in_keplerian_elements
# kep_state[int(KEI.semi_major_axis_index)] = 7500.0E3
# kep_state[int(KEI.eccentricity_index)] = 0.1
# kep_state[int(KEI.inclination_index)] = np.deg2rad(85.3)
# kep_state[int(KEI.argument_of_periapsis_index)] = np.deg2rad(235.7)
# kep_state[int(KEI.longitude_of_ascending_node_index)] = np.deg2rad(23.4)
# kep_state[int(KEI.true_anomaly_index)] = np.deg2rad(139.87)
# system_initial_state = corbital_element_conversions.onvert_keplerian_to_cartesian_elements(
# kep_state, earth_gravitational_parameter)
# REVISED CONTEMPORARY DESIGN.
system_initial_state = elements.keplerian2cartesian(
mu=earth_gravitational_parameter,
a=7500.0E3,
ecc=0.1,
inc=np.deg2rad(85.3),
raan=np.deg2rad(23.4),
argp=np.deg2rad(235.7),
nu=np.deg2rad(139.87))
# Create propagation settings.
propagator_settings = propagators.TranslationalStatePropagatorSettings(
central_bodies, acceleration_models, bodies_to_propagate,
system_initial_state, simulation_end_epoch)
# Create numerical integrator settings.
integrator_settings = numerical_integrators.IntegratorSettings(
numerical_integrators.AvailableIntegrators.rungeKutta4,
simulation_start_epoch, fixed_step_size)
###########################################################################
# PROPAGATE ORBIT #########################################################
###########################################################################
# Create simulation object and propagate dynamics.
dynamics_simulator = propagators.SingleArcDynamicsSimulator(
bodies, integrator_settings, propagator_settings, True)
result = dynamics_simulator.get_equations_of_motion_numerical_solution()
###########################################################################
# PRINT INITIAL AND FINAL STATES ##########################################
###########################################################################
print(f"""
Single Earth-Orbiting Satellite Example.
The initial position vector of Delfi-C3 is [km]: \n{
result[simulation_start_epoch][:3] / 1E3}
The initial velocity vector of Delfi-C3 is [km]: \n{
result[simulation_start_epoch][3:] / 1E3}
After {simulation_end_epoch} seconds the position vector of Delfi-C3 is [km]: \n{
result[simulation_end_epoch][:3] / 1E3}
And the velocity vector of Delfi-C3 is [km]: \n{
result[simulation_start_epoch][3:] / 1E3}
""")
###########################################################################
# SAVE RESULTS ############################################################
###########################################################################
io.save2txt(
solution=result,
filename="singlePerturbedSatellitePropagationHistory.dat",
directory="./tutorial_2",
)
# Final statement (not required, though good practice in a __main__).
return 0
def single_satellite_propagator(
start_epoch=0.0,
fixed_step_size=10.0,
end_epoch=constants.JULIAN_DAY,
parent_body="Earth",
frame_orientation="ECLIPJ2000",
frame_origin="SSB",
satellite_name="Delfi-C3",
sat_sma=7500.0E3,
sat_ecc=0.1,
sat_inc=np.deg2rad(85.3),
sat_raan=np.deg2rad(23.4),
sat_argp=np.deg2rad(235.7),
sat_nu=np.deg2rad(139.87),
return_output=False,
print_output=True,
output_type='markdown' # [array, markdown, latex, string]
):
"""
Tutorial 1 of the tudatpy library in function form.
Warning
-------
This function is not suited to mass generation of output data, as it
is creating the environment and simulation from scratch for each call.
Parameters
----------
start_epoch
fixed_step_size
end_epoch
parent_body
frame_orientation
frame_origin
satellite_name
sat_sma
sat_ecc
sat_inc
sat_raan
sat_argp
sat_nu
return_output
print_output
output_type
Returns
-------
"""
# Load spice kernels.
spice_interface.load_standard_kernels()
# Set simulation start epoch.
simulation_start_epoch = start_epoch
# Set numerical integration fixed step size.
fixed_step_size = fixed_step_size
# Set simulation end epoch.
simulation_end_epoch = end_epoch
###########################################################################
# CREATE ENVIRONMENT ######################################################
###########################################################################
# Create body objects.
bodies_to_create = [parent_body]
body_settings = simulation_setup.get_default_body_settings(
bodies_to_create)
body_settings[
parent_body].ephemeris_settings = simulation_setup.ConstantEphemerisSettings(
np.zeros(6))
body_settings[parent_body].rotation_model_settings.reset_original_frame(
frame_orientation)
# Create Earth Object.
bodies = simulation_setup.create_bodies(body_settings)
###########################################################################
# CREATE VEHICLE ##########################################################
###########################################################################
# Create vehicle objects.
bodies[satellite_name] = simulation_setup.Body()
###########################################################################
# FINALIZE BODIES #########################################################
###########################################################################
simulation_setup.set_global_frame_body_ephemerides(bodies, frame_origin,
frame_orientation)
###########################################################################
# CREATE ACCELERATIONS ####################################################
###########################################################################
# Define bodies that are propagated.
bodies_to_propagate = [satellite_name]
# Define central bodies.
central_bodies = [parent_body]
# Define accelerations acting on Delfi-C3.
accelerations_of_satellite = {
parent_body: [simulation_setup.Acceleration.point_mass_gravity()]
}
# Create global accelerations dictionary.
accelerations = {satellite_name: accelerations_of_satellite}
# Create acceleration models.
acceleration_models = simulation_setup.create_acceleration_models_dict(
bodies, accelerations, bodies_to_propagate, central_bodies)
###########################################################################
# CREATE PROPAGATION SETTINGS #############################################
###########################################################################
# Set initial conditions for the Asterix satellite that will be
# propagated in this simulation. The initial conditions are given in
# Keplerian elements and later on converted to Cartesian elements.
# Set Keplerian elements for satellite
parent_gravitational_parameter = bodies[
parent_body].gravity_field_model.get_gravitational_parameter()
# REVISED CONTEMPORARY DESIGN.
system_initial_state = elements.keplerian2cartesian(
mu=parent_gravitational_parameter,
a=sat_sma,
ecc=sat_ecc,
inc=sat_inc,
raan=sat_raan,
argp=sat_argp,
nu=sat_nu)
# Create propagation settings.
propagator_settings = propagators.TranslationalStatePropagatorSettings(
central_bodies, acceleration_models, bodies_to_propagate,
system_initial_state, simulation_end_epoch)
# Create numerical integrator settings.
integrator_settings = numerical_integrators.IntegratorSettings(
numerical_integrators.AvailableIntegrators.rungeKutta4,
simulation_start_epoch, fixed_step_size)
###########################################################################
# PROPAGATE ORBIT #########################################################
###########################################################################
# Create simulation object and propagate dynamics.
dynamics_simulator = propagators.SingleArcDynamicsSimulator(
bodies, integrator_settings, propagator_settings, True)
result = dynamics_simulator.get_equations_of_motion_numerical_solution()
###########################################################################
# OUTPUT HANDLING #########################################################
###########################################################################
if output_type != 'array':
df = pd.DataFrame(data=np.vstack((
result[simulation_start_epoch] / 1E3,
result[simulation_end_epoch] / 1E3,
)),
columns=[
'R_x [km]', 'R_y [km]', 'R_z [km]', 'V_x [km/s]',
'V_y [km/s]', 'V_z [km/s]'
])
df.index = ['t_0', 't_f']
# Output type
if output_type == 'markdown':
output = df.to_markdown()
elif output_type == 'latex':
output = df.to_latex(escape=False)
elif output_type == 'df':
output = df
else:
output = np.vstack((
result[simulation_start_epoch] / 1E3,
result[simulation_end_epoch] / 1E3,
))
# Print output
if print_output:
print(output)
# Return output
if return_output:
return output
else:
return None
def main():
# Load spice kernels.
spice_interface.load_standard_spice_kernels()
# Set simulation start epoch.
simulation_start_epoch = 0.0
# Set numerical integration fixed step size.
fixed_step_size = 10.0
# Set simulation end epoch.
simulation_end_epoch = constants.JULIAN_DAY
###########################################################################
# CREATE ENVIRONMENT ######################################################
###########################################################################
# Create body objects.
bodies_to_create = ["Earth"]
body_settings = simulation_setup.get_default_body_settings(
bodies_to_create)
body_settings[
"Earth"].ephemeris_settings = simulation_setup.ConstantEphemerisSettings(
np.zeros(6))
body_settings["Earth"].rotation_model_settings.reset_original_frame(
"ECLIPJ2000")
# Create Earth Object.
bodies = simulation_setup.create_bodies(body_settings)
###########################################################################
# CREATE VEHICLE ##########################################################
###########################################################################
# Create vehicle objects.
bodies["Delfi-C3"] = simulation_setup.Body()
###########################################################################
# FINALIZE BODIES #########################################################
###########################################################################
simulation_setup.set_global_frame_body_ephemerides(bodies, "SSB",
"ECLIPJ2000")
###########################################################################
# CREATE ACCELERATIONS ####################################################
###########################################################################
# Define bodies that are propagated.
bodies_to_propagate = ["Delfi-C3"]
# Define central bodies.
central_bodies = ["Earth"]
# Define accelerations acting on Delfi-C3.
accelerations_of_delfi_c3 = dict(
Earth=[simulation_setup.Acceleration.point_mass_gravity()])
# Create global accelerations dictionary.
accelerations = {"Delfi-C3": accelerations_of_delfi_c3}
# Create acceleration models.
acceleration_models = simulation_setup.create_acceleration_models_dict(
bodies, accelerations, bodies_to_propagate, central_bodies)
###########################################################################
# CREATE PROPAGATION SETTINGS #############################################
###########################################################################
# Set initial conditions for the Asterix satellite that will be
# propagated in this simulation. The initial conditions are given in
# Keplerian elements and later on converted to Cartesian elements.
# Set Keplerian elements for Delfi-C3
earth_gravitational_parameter = bodies[
"Earth"].gravity_field_model.get_gravitational_parameter()
# LEGACY DESIGN.
# KEI = orbital_element_conversions.KeplerianElementIndices
# asterix_initial_state_in_keplerian_elements = np.zeros(6)
# kep_state = asterix_initial_state_in_keplerian_elements
# kep_state[int(KEI.semi_major_axis_index)] = 7500.0E3
# kep_state[int(KEI.eccentricity_index)] = 0.1
# kep_state[int(KEI.inclination_index)] = np.deg2rad(85.3)
# kep_state[int(KEI.argument_of_periapsis_index)] = np.deg2rad(235.7)
# kep_state[int(KEI.longitude_of_ascending_node_index)] = np.deg2rad(23.4)
# kep_state[int(KEI.true_anomaly_index)] = np.deg2rad(139.87)
# system_initial_state = corbital_element_conversions.onvert_keplerian_to_cartesian_elements(
# kep_state, earth_gravitational_parameter)
# REVISED CONTEMPORARY DESIGN.
system_initial_state = elements.keplerian2cartesian(
mu=earth_gravitational_parameter,
a=7500.0E3,
ecc=0.1,
inc=np.deg2rad(85.3),
raan=np.deg2rad(23.4),
argp=np.deg2rad(235.7),
nu=np.deg2rad(139.87))
# Create propagation settings.
propagator_settings = propagators.TranslationalStatePropagatorSettings(
central_bodies, acceleration_models, bodies_to_propagate,
system_initial_state, simulation_end_epoch)
# Create numerical integrator settings.
integrator_settings = numerical_integrators.IntegratorSettings(
numerical_integrators.AvailableIntegrators.rungeKutta4,
simulation_start_epoch, fixed_step_size)
###########################################################################
# PROPAGATE ORBIT #########################################################
###########################################################################
t0 = time.time()
# Create simulation object and propagate dynamics.
dynamics_simulator = propagators.SingleArcDynamicsSimulator(
bodies, integrator_settings, propagator_settings, True)
result = dynamics_simulator.get_equations_of_motion_numerical_solution()
###########################################################################
# PRINT INITIAL AND FINAL STATES ##########################################
###########################################################################
print(f"""
Single Earth-Orbiting Satellite Example.
The initial position vector of Delfi-C3 is [km]: \n{
result[simulation_start_epoch][:3] / 1E3}
The initial velocity vector of Delfi-C3 is [km]: \n{
result[simulation_start_epoch][3:] / 1E3}
After {simulation_end_epoch} seconds the position vector of Delfi-C3 is [km]: \n{
result[simulation_end_epoch][:3] / 1E3}
And the velocity vector of Delfi-C3 is [km]: \n{
result[simulation_start_epoch][3:] / 1E3}
""")
###########################################################################
# SAVE RESULTS ############################################################
###########################################################################
#
# io.save2txt(
# solution=result,
# filename="singleSatellitePropagationHistory.dat",
# directory="./tutorial_1",
# )
# Final statement (not required, though good practice in a __main__).
return 0
def main():
# Load spice kernels.
spice_interface.load_standard_spice_kernels()
# Set simulation start epoch.
simulation_start_epoch = 0.0
# Set numerical integration fixed step size.
fixed_step_size = 1.0
# Set simulation end epoch.
simulation_end_epoch = constants.JULIAN_DAY
###########################################################################
# CREATE ENVIRONMENT ######################################################
###########################################################################
bodies_to_create = ["Sun", "Earth", "Moon", "Mars", "Venus"]
env = Environment(tudatpy_bodies=bodies_to_create,
custom_bodies=None,
frame_origin="SSB",
frame_orientation="ECLIPJ2000",
start_epoch=simulation_start_epoch,
end_epoch=simulation_end_epoch,
epoch_margin=300.0)
###########################################################################
# CREATE VEHICLE ##########################################################
###########################################################################
env["Delfi-C3"] = Body(env)
env["Delfi-C3"].set_constant_body_mass(400.0)
###########################################################################
# CREATE VEHICLE - ENVIRONMENT INTERFACE ##################################
###########################################################################
# Create aerodynamic coefficient interface settings
reference_area = 4.0
aerodynamic_coefficient = 1.2
aero_c_settings = simulation_setup.ConstantAerodynamicCoefficientSettings(
reference_area,
aerodynamic_coefficient * np.ones(3),
are_coefficients_in_aerodynamic_frame=True,
are_coefficients_in_negative_axis_direction=True)
# Create and set aerodynamic coefficients object
env._bodies["Delfi-C3"].set_aerodynamic_coefficient_interface(
simulation_setup.create_aerodynamic_coefficient_interface(
aero_c_settings, "Delfi-C3"))
# TODO: Simplify (post 1.0.0 work)
# Create radiation pressure settings
reference_area_radiation = 4.0
radiation_pressure_coefficient = 1.2
occulting_bodies = ["Earth"]
rad_press_settings = simulation_setup.CannonBallRadiationPressureInterfaceSettings(
"Sun", reference_area_radiation, radiation_pressure_coefficient,
occulting_bodies)
# Create and set radiation pressure settings
env._bodies["Delfi-C3"].set_radiation_pressure_interface(
"Sun",
simulation_setup.create_radiation_pressure_interface(
rad_press_settings, "Delfi-C3", env._bodies))
# # Create aerodynamic interface settings for Delfi-C3.
# env["Delfi-C3"]._set_aerodynamic_coefficient_interface(
# reference_area=4.0,
# coefficients=1.2 * np.ones(3),
# coefficients_in_aerodynamic_frame=True,
# coefficients_in_negative_axis_direction=True,
# model="constant_coefficients")
#
# # Create radiation pressure interface settings for Delfi-C3.
# env["Delfi-C3"]._set_radiation_pressure_interface(
# reference_area=4.0,
# coefficient=1.2,
# occulting=["Earth"],
# source="Sun",
# model="cannon_ball")
###########################################################################
# FINALIZE BODIES #########################################################
###########################################################################
# Finalize body creation.
env.finalize_environment()
###########################################################################
# CREATE ACCELERATIONS ####################################################
###########################################################################
# Define bodies that are propagated.
bodies_to_propagate = ["Delfi-C3"]
# Define central bodies.
central_bodies = ["Earth"]
# Define unique (Sun, Earth) accelerations acting on Delfi-C3.
accelerations_of_delfi_c3 = dict(
Sun=[
simulation_setup.Acceleration.canon_ball_radiation_pressure()
# AccelerationSettings(AvailableAcceleration.cannon_ball_radiation_pressure) # LEGACY DESIGN.
],
Earth=[
simulation_setup.Acceleration.spherical_harmonic_gravity(5, 5),
# SphericalHarmonicAccelerationSettings(5, 5), # LEGACY DESIGN.
simulation_setup.Acceleration.aerodynamic()
# AccelerationSettings(AvailableAcceleration.aerodynamic) # LEGACY DESIGN.
])
# Define other point mass accelerations acting on Delfi-C3.
for other in set(bodies_to_create).difference({"Sun", "Earth"}):
accelerations_of_delfi_c3[other] = [
simulation_setup.Acceleration.point_mass_gravity()
]
# Create global accelerations dictionary.
acceleration_dict = {"Delfi-C3": accelerations_of_delfi_c3}
# Create acceleration models.
acceleration_models = simulation_setup.create_acceleration_models_dict(
env._bodies, acceleration_dict, bodies_to_propagate, central_bodies)
###########################################################################
# CREATE PROPAGATION SETTINGS #############################################
###########################################################################
# Set initial conditions for the Asterix satellite that will be
# propagated in this simulation. The initial conditions are given in
# Keplerian elements and later on converted to Cartesian elements.
# Set Keplerian elements for Delfi-C3
earth_gravitational_parameter = env[
"Earth"].gravity_field_model.get_gravitational_parameter()
# REVISED CONTEMPORARY DESIGN.
system_initial_state = elements.keplerian2cartesian(
mu=earth_gravitational_parameter,
a=7500.0E3,
ecc=0.1,
inc=np.deg2rad(85.3),
raan=np.deg2rad(23.4),
argp=np.deg2rad(235.7),
nu=np.deg2rad(139.87))
# Create propagation settings.
propagator_settings = propagators.TranslationalStatePropagatorSettings(
central_bodies, acceleration_models, bodies_to_propagate,
system_initial_state, simulation_end_epoch)
# Create numerical integrator settings.
integrator_settings = numerical_integrators.IntegratorSettings(
numerical_integrators.AvailableIntegrators.rungeKutta4,
simulation_start_epoch, fixed_step_size)
###########################################################################
# PROPAGATE ORBIT #########################################################
###########################################################################
t0 = time.time()
# Create simulation object and propagate dynamics.
print(env._bodies)
dynamics_simulator = propagators.SingleArcDynamicsSimulator(
env._bodies, integrator_settings, propagator_settings, True)
print("dynamics_simulator: ", time.time() - t0)
result = dynamics_simulator.get_equations_of_motion_numerical_solution()
###########################################################################
# PRINT INITIAL AND FINAL STATES ##########################################
###########################################################################
print(f"""
Single Earth-Orbiting Satellite Example.
The initial position vector of Delfi-C3 is [km]: \n{
result[simulation_start_epoch][:3] / 1E3}
The initial velocity vector of Delfi-C3 is [km/s]: \n{
result[simulation_start_epoch][3:] / 1E3}
After {simulation_end_epoch} seconds the position vector of Delfi-C3 is [km]: \n{
result[simulation_end_epoch][:3] / 1E3}
And the velocity vector of Delfi-C3 is [km/s]: \n{
result[simulation_start_epoch][3:] / 1E3}
""")
###########################################################################
# SAVE RESULTS ############################################################
###########################################################################
io.save2txt(
solution=result,
filename="singlePerturbedSatellitePropagationHistory.dat",
directory="./tutorial_2_prototype",
)
# Final statement (not required, though good practice in a __main__).
return 0