Actuator#

class Actuator(name: str, spice_id: int | str = None)#

Bases: Spacecraft

Any spacecraft component that applies thrust, torque, or other forces.

Parameters:

  • name (str) – The name of the Actuator object.

  • spice_id (int or str, optional) – The NAIF ID code of the object. Defaults to None.

Notes

Actuator is a lightweight subclass of Spacecraft. It inherits all spacecraft state and SPICE-registration methods while adding no additional properties, leaving specialisation to concrete subclasses (e.g. thrusters).

See also

scarabaeus.Spacecraft

Parent class providing SPICE registration and mass tracking.

scarabaeus.Instrument

Sibling subclass for sensing instruments.

Attributes:
area

The cross-sectional area of the spacecraft.

body_fixed_frame_id

The SPICE ID assigned to the spacecraft’s body-fixed frame.

instrument_list

The list of instruments attached to the spacecraft.

mass_profile

Stored constant or polynomial mass profile.

n_plate_model

The N-Plate model representing the spacecraft’s geometry.

name

The name of the spacecraft.

ref_coeff

The reflectivity coefficient of the spacecraft.

spice_id

The SPICE ID of the Body object.

thruster

Thruster actuator attached to this spacecraft, or None if not set.

Methods

add_instrument(instruments)

Adds an instrument or instruments to the spacecraft.

fit_mass_profile_from_data(t_array, m_array, ...)

Fits a polynomial to a mass time history and sets it as the new mass profile.

get_attitude(from_frame, epoch[, tol])

Computes the Direction Cosine Matrix (DCM) to transform vectors from a given inertial frame to the spacecraft-fixed frame.

get_dependent_spice_ids()

Collects the SPICE ID's of all attached frames and instruments into a dictionary.

get_mass(t)

Evaluates the time-varying mass profile at specified epochs.

get_state(epoch_0[, reference_frame, origin])

Retrieves the state of the body relative to a given origin in a specified reference frame.

inertial_to_body_DCM_nadir_pointing_to_sun(...)

Computes the Direction Cosine Matrix (DCM) that points the body frame's -X axis toward the Sun (nadir-to-sun), using inertial quantities.

mass([t])

Returns the mass of the body.

set_mass_profile(poly, units)

Sets the mass profile of the body using a polynomial function and associated units.
add_instrument(instruments) → None#

Adds an instrument or instruments to the spacecraft.

Parameters:

instruments (Instrument or list of Instrument) – The instrument or instruments to be attached to the spacecraft.

Return type:

None

fit_mass_profile_from_data(t_array: ndarray, m_array: ndarray, domain: tuple, units: Units) → None#

Fits a polynomial to a mass time history and sets it as the new mass profile.

Parameters:

  • t_array (numpy.ndarray) – Time values in seconds (ephemeris time).

  • m_array (numpy.ndarray) – Corresponding mass values.

  • deg (int) – Degree of the polynomial.

  • domain (tuple) – Domain [t0, tf] for the polynomial conversion.

  • units (ArrayWUnits) – Units to assign to the mass profile. If None, uses existing mass profile units.

Return type:

None

get_attitude(from_frame: Frame, epoch: EpochArray, tol: int = 0) → ndarray#

Computes the Direction Cosine Matrix (DCM) to transform vectors from a given inertial frame to the spacecraft-fixed frame.

Parameters:

  • from_frame (Frame) – The intertial reference frame for which the spacecraft-fixed frame will be transformed from.

  • epoch (EpochArray) – The epoch at which the transformation will be performed. Given in ephemeris time (seconds past J2000).

  • tol (int, optional) – The time tolerance for retrieving attitude and angular velocity from the specified spacecraft clock time, given in sclock ticks. Defaults to 0.

Returns:

DCM – The DCM for transforming inertial vectors to spacecraft-fixed coordinates.

Return type:

numpy.ndarray

Notes

The epoch argument is converted from ephemeris time in seconds to continuous encoded spacecraft clock ticks.

get_dependent_spice_ids()#

Collects the SPICE ID’s of all attached frames and instruments into a dictionary.

Parameters:

None

Returns:

ids – A dictionary containing the keys:

  • 'instruments' : list The SPICE ID’s of all attached Instrument objects.

  • 'plates' : list The SPICE ID’s of all C-frames for CK type panels defined in the nPlateModel if one has been assigned to the Spacecraft.

Return type:

dict

get_mass(t: EpochArray) → ArrayWUnits#

Evaluates the time-varying mass profile at specified epochs.

Parameters:

t (EpochArray) – Time(s) in ephemeris seconds (TDB).

Returns:

mass – Evaluated mass at each input epoch.

Return type:

ArrayWUnits

get_state(epoch_0, reference_frame: str = 'J2000', origin: str = 'EARTH') → ArrayWFrame#

Retrieves the state of the body relative to a given origin in a specified reference frame.

Parameters:

  • epoch_0 (EpochArray) – The epoch times for which the state is to be computed.

  • reference_frame (str) – The reference frame in which the state is desired. Defaults to ‘J2000’.

  • origin (str) – The origin body relative to which the state is computed. Defaults to 'EARTH'.

Returns:

state_vector – The state vector of the body relative to the origin.

Return type:

ArrayWFrame

inertial_to_body_DCM_nadir_pointing_to_sun(origin_name, frame, sun2body_pos, current_state: dict, epoch, body)#

Computes the Direction Cosine Matrix (DCM) that points the body frame’s -X axis toward the Sun (nadir-to-sun), using inertial quantities.

Parameters:

  • origin_name (str) – Name of the origin at which current_state is defined (e.g., ‘EARTH’, ‘SUN’).

  • frame (Frame or str) – Frame in which current_state is expressed (e.g., J2000).

  • sun2body_pos (ArrayWUnits) – Position of the body w.r.t. SUN (km).

  • current_state (dict) –

    State of the body w.r.t. the origin. Accepts either: - {(‘velocity’, spice_id): Array/ArrayWUnits or ‘velocity’: Array/ArrayWUnits}

    (and similarly for ‘position’ if ever needed)

  • epoch (float) – Epoch (TDB seconds past J2000) at which the state is computed.

  • body (CelestialBody or Spacecraft) – The body whose state we are extracting from current_state (used for spice_id lookup).

Returns:

DCM_inertial_to_body – DCM that rotates inertial vectors into the body-fixed frame realizing the nadir-to-sun pointing.

Return type:

(3,3) ndarray

mass(t: EpochArray | float | ndarray = None) → ArrayWUnits#

Returns the mass of the body. If the mass is time-varying, evaluates it at time t. If the mass is constant, returns the same value regardless of t.

Parameters:

t (EpochArray, float, or np.ndarray, optional) – Time(s) in ephemeris seconds (TDB) at which to evaluate the mass. Or None for constant mass.

Returns:

Mass at time t or constant mass.

Return type:

ArrayWUnits

set_mass_profile(poly: SCBPolynomial, units: Units) → None#

Sets the mass profile of the body using a polynomial function and associated units.

Parameters:

  • poly (SCBPolynomial) – A polynomial object representing the mass profile.

  • units (Units) – The units associated with the mass profile.

Return type:

None

property area: ArrayWUnits#

The cross-sectional area of the spacecraft. Expressed in units of square meters.

property body_fixed_frame_id: int#

The SPICE ID assigned to the spacecraft’s body-fixed frame.

property instrument_list: list#

The list of instruments attached to the spacecraft.

property mass_profile: ArrayWUnits | SCBPolynomial#

Stored constant or polynomial mass profile.

property n_plate_model: nPlateModel#

The N-Plate model representing the spacecraft’s geometry.

property name: str#

The name of the spacecraft.

property ref_coeff: float#

The reflectivity coefficient of the spacecraft.

property spice_id: int | str#

The SPICE ID of the Body object.

property thruster#

Thruster actuator attached to this spacecraft, or None if not set.