flight.py

# pylint: disable=too-many-lines
import math
import warnings
from copy import deepcopy
from functools import cached_property

import numpy as np
from scipy.integrate import BDF, DOP853, LSODA, RK23, RK45, OdeSolver, Radau

from rocketpy.simulation.flight_data_exporter import FlightDataExporter

from ..mathutils.function import Function, funcify_method
from ..mathutils.vector_matrix import Matrix, Vector
from ..motors.point_mass_motor import PointMassMotor
from ..plots.flight_plots import _FlightPlots
from ..prints.flight_prints import _FlightPrints
from ..rocket import PointMassRocket
from ..tools import (
    calculate_cubic_hermite_coefficients,
    deprecated,
    euler313_to_quaternions,
    find_closest,
    find_root_linear_interpolation,
    find_roots_cubic_function,
    quaternions_to_nutation,
    quaternions_to_precession,
    quaternions_to_spin,
)

ODE_SOLVER_MAP = {
    "RK23": RK23,
    "RK45": RK45,
    "DOP853": DOP853,
    "Radau": Radau,
    "BDF": BDF,
    "LSODA": LSODA,
}


# pylint: disable=too-many-public-methods
# pylint: disable=too-many-instance-attributes
class Flight:
    """Keeps all flight information and has a method to simulate flight.

    Attributes
    ----------
    Flight.env : Environment
        Environment object describing rail length, elevation, gravity and
        weather condition. See Environment class for more details.
    Flight.rocket : Rocket
        Rocket class describing rocket. See Rocket class for more
        details.
    Flight.parachutes : Parachute
        Direct link to parachutes of the Rocket. See Rocket class
        for more details.
    Flight.frontal_surface_wind : float
        Surface wind speed in m/s aligned with the launch rail.
    Flight.lateral_surface_wind : float
        Surface wind speed in m/s perpendicular to launch rail.
    Flight.FlightPhases : class
        Helper class to organize and manage different flight phases.
    Flight.TimeNodes : class
        Helper class to manage time discretization during simulation.
    Flight.time_iterator : function
        Helper iterator function to generate time discretization points.
    Flight.rail_length : float, int
        Launch rail length in meters.
    Flight.effective_1rl : float
        Original rail length minus the distance measured from nozzle exit
        to the upper rail button. It assumes the nozzle to be aligned with
        the beginning of the rail.
    Flight.effective_2rl : float
        Original rail length minus the distance measured from nozzle exit
        to the lower rail button. It assumes the nozzle to be aligned with
        the beginning of the rail.
    Flight.name: str
        Name of the flight.
    Flight._controllers : list
        List of controllers to be used during simulation.
    Flight.max_time : int, float
        Maximum simulation time allowed. Refers to physical time
        being simulated, not time taken to run simulation.
    Flight.max_time_step : int, float
        Maximum time step to use during numerical integration in seconds.
    Flight.min_time_step : int, float
        Minimum time step to use during numerical integration in seconds.
    Flight.rtol : int, float
        Maximum relative error tolerance to be tolerated in the
        numerical integration scheme.
    Flight.atol : int, float
        Maximum absolute error tolerance to be tolerated in the
        integration scheme.
    Flight.time_overshoot : bool, optional
        If True, decouples ODE time step from parachute and controller trigger
        functions sampling rate. The time steps can overshoot the necessary
        trigger function evaluation points and then interpolation is used to
        calculate them and feed the triggers. Can greatly improve run
        time in some cases.
    Flight.terminate_on_apogee : bool
        Whether to terminate simulation when rocket reaches apogee.
    Flight.solver : scipy.integrate.LSODA
        Scipy LSODA integration scheme.
    Flight.x : Function
        Rocket's X coordinate (positive east) as a function of time.
    Flight.y : Function
        Rocket's Y coordinate (positive north) as a function of time.
    Flight.z : Function
        Rocket's z coordinate (positive up) as a function of time.
    Flight.vx : Function
        Velocity of the rocket's center of dry mass in the X (East) direction of
        the inertial frame as a function of time.
    Flight.vy : Function
        Velocity of the rocket's center of dry mass in the Y (North) direction of
        the inertial frame as a function of time.
    Flight.vz : Function
        Velocity of the rocket's center of dry mass in the Z (Up) direction of
        the inertial frame as a function of time.
    Flight.e0 : Function
        Rocket's Euler parameter 0 as a function of time.
    Flight.e1 : Function
        Rocket's Euler parameter 1 as a function of time.
    Flight.e2 : Function
        Rocket's Euler parameter 2 as a function of time.
    Flight.e3 : Function
        Rocket's Euler parameter 3 as a function of time.
    Flight.w1 : Function
        Angular velocity of the rocket in the x direction of the rocket's
        body frame as a function of time, in rad/s. Sometimes referred to as
        pitch rate (q).
    Flight.w2 : Function
        Angular velocity of the rocket in the y direction of the rocket's
        body frame as a function of time, in rad/s. Sometimes referred to as
        yaw rate (r).
    Flight.w3 : Function
        Angular velocity of the rocket in the z direction of the rocket's
        body frame as a function of time, in rad/s. Sometimes referred to as
        roll rate (p).
    Flight.latitude: Function
        Rocket's latitude coordinates (positive North) as a function of time.
        The Equator has a latitude equal to 0, by convention.
    Flight.longitude: Function
        Rocket's longitude coordinates (positive East) as a function of time.
        Greenwich meridian has a longitude equal to 0, by convention.
    Flight.inclination : int, float
        Launch rail inclination angle relative to ground, given in degrees.
    Flight.heading : int, float
        Launch heading angle relative to north given in degrees.
    Flight.initial_solution : list
        List defines initial condition - [t_initial, x_init,
        y_init, z_init, vx_init, vy_init, vz_init, e0_init, e1_init,
        e2_init, e3_init, w1_init, w2_init, w3_init]
    Flight.t_initial : int, float
        Initial simulation time in seconds. Usually 0.
    Flight.solution : list
        Solution array which keeps results from each numerical
        integration.
    Flight.t : float
        Current integration time.
    Flight.y : list
        Current integration state vector u.
    Flight.out_of_rail_time : int, float
        Time, in seconds, in which the rocket completely leaves the
        rail.
    Flight.out_of_rail_state : list
        State vector u corresponding to state when the rocket
        completely leaves the rail.
    Flight.out_of_rail_velocity : int, float
        Velocity, in m/s, with which the rocket completely leaves the
        rail.
    Flight.apogee_state : array
        State vector u corresponding to state when the rocket's
        vertical velocity is zero in the apogee.
    Flight.apogee_time : int, float
        Time, in seconds, in which the rocket's vertical velocity
        reaches zero in the apogee.
    Flight.apogee_x : int, float
        X coordinate (positive east) of the center of mass of the
        rocket when it reaches apogee.
    Flight.apogee_y : int, float
        Y coordinate (positive north) of the center of mass of the
        rocket when it reaches apogee.
    Flight.apogee : int, float
        Z coordinate, or altitude, of the center of mass of the
        rocket when it reaches apogee.
    Flight.x_impact : int, float
        X coordinate (positive east) of the center of mass of the
        rocket when it impacts ground.
    Flight.y_impact : int, float
        Y coordinate (positive east) of the center of mass of the
        rocket when it impacts ground.
    Flight.impact_velocity : int, float
        Velocity magnitude of the center of mass of the rocket when
        it impacts ground.
    Flight.impact_state : array
        State vector u corresponding to state when the rocket
        impacts the ground.
    Flight.parachute_events : array
        List that stores parachute events triggered during flight.
    Flight.function_evaluations : array
        List that stores number of derivative function evaluations
        during numerical integration in cumulative manner.
    Flight.function_evaluations_per_time_step : list
        List that stores number of derivative function evaluations
        per time step during numerical integration.
    Flight.time_steps : array
        List of time steps taking during numerical integration in
        seconds.
    Flight.flight_phases : Flight.FlightPhases
        Stores and manages flight phases.
    Flight.wind_velocity_x : Function
        Wind velocity X (East) experienced by the rocket as a
        function of time.
    Flight.wind_velocity_y : Function
        Wind velocity Y (North) experienced by the rocket as a
        function of time.
    Flight.density : Function
        Air density experienced by the rocket as a function of
        time.
    Flight.pressure : Function
        Air pressure experienced by the rocket as a function of
        time.
    Flight.dynamic_viscosity : Function
        Air dynamic viscosity experienced by the rocket as a function of
        time.
    Flight.speed_of_sound : Function
        Speed of Sound in air experienced by the rocket as a
        function of time.
    Flight.ax : Function
        Acceleration of the rocket's center of dry mass along the X (East)
        axis in the inertial frame as a function of time.
    Flight.ay : Function
        Acceleration of the rocket's center of dry mass along the Y (North)
        axis in the inertial frame as a function of time.
    Flight.az : Function
        Acceleration of the rocket's center of dry mass along the Z (Up)
        axis in the inertial frame as a function of time.
    Flight.alpha1 : Function
        Angular acceleration of the rocket in the x direction of the rocket's
        body frame as a function of time, in rad/s. Sometimes referred to as
        yaw acceleration.
    Flight.alpha2 : Function
        Angular acceleration of the rocket in the y direction of the rocket's
        body frame as a function of time, in rad/s. Sometimes referred to as
        yaw acceleration.
    Flight.alpha3 : Function
        Angular acceleration of the rocket in the z direction of the rocket's
        body frame as a function of time, in rad/s. Sometimes referred to as
        roll acceleration.
    Flight.speed : Function
        Rocket velocity magnitude in m/s relative to ground as a
        function of time.
    Flight.max_speed : float
        Maximum velocity magnitude in m/s reached by the rocket
        relative to ground during flight.
    Flight.max_speed_time : float
        Time in seconds at which rocket reaches maximum velocity
        magnitude relative to ground.
    Flight.horizontal_speed : Function
        Rocket's velocity magnitude in the horizontal (North-East)
        plane in m/s as a function of time.
    Flight.acceleration : Function
        Rocket acceleration magnitude in m/s² relative to ground as a
        function of time.
    Flight.max_acceleration : float
        Maximum acceleration magnitude in m/s² reached by the rocket
        relative to ground during flight.
    Flight.max_acceleration_time : float
        Time in seconds at which rocket reaches maximum acceleration
        magnitude relative to ground.
    Flight.path_angle : Function
        Rocket's flight path angle, or the angle that the
        rocket's velocity makes with the horizontal (North-East)
        plane. Measured in degrees and expressed as a function
        of time.
    Flight.attitude_vector_x : Function
        Rocket's attitude vector, or the vector that points
        in the rocket's axis of symmetry, component in the X
        direction (East) as a function of time.
    Flight.attitude_vector_y : Function
        Rocket's attitude vector, or the vector that points
        in the rocket's axis of symmetry, component in the Y
        direction (East) as a function of time.
    Flight.attitude_vector_z : Function
        Rocket's attitude vector, or the vector that points
        in the rocket's axis of symmetry, component in the Z
        direction (East) as a function of time.
    Flight.attitude_angle : Function
        Rocket's attitude angle, or the angle that the
        rocket's axis of symmetry makes with the horizontal (North-East)
        plane. Measured in degrees and expressed as a function
        of time.
    Flight.lateral_attitude_angle : Function
        Rocket's lateral attitude angle, or the angle that the
        rocket's axis of symmetry makes with plane defined by
        the launch rail direction and the Z (up) axis.
        Measured in degrees and expressed as a function
        of time.
    Flight.phi : Function
        Rocket's Spin Euler Angle, φ, according to the 3-2-3 rotation
        system nomenclature (NASA Standard Aerospace). Measured in degrees and
        expressed as a function of time.
    Flight.theta : Function
        Rocket's Nutation Euler Angle, θ, according to the 3-2-3 rotation
        system nomenclature (NASA Standard Aerospace). Measured in degrees and
        expressed as a function of time.
    Flight.psi : Function
        Rocket's Precession Euler Angle, ψ, according to the 3-2-3 rotation
        system nomenclature (NASA Standard Aerospace). Measured in degrees and
        expressed as a function of time.
    Flight.R1 : Function
        Aerodynamic force acting along the x-axis of the rocket's body frame
        as a function of time. Expressed in Newtons (N).
    Flight.R2 : Function
        Aerodynamic force acting along the y-axis of the rocket's body frame
        as a function of time. Expressed in Newtons (N).
    Flight.R3 : Function
        Aerodynamic force acting along the z-axis of the rocket's body frame
        as a function of time. Expressed in Newtons (N).
    Flight.M1 : Function
        Aerodynamic moment acting along the x-axis of the rocket's body
        frame as a function of time. Expressed in Newtons (N).
    Flight.M2 : Function
        Aerodynamic moment acting along the y-axis of the rocket's body
        frame as a function of time. Expressed in Newtons (N).
    Flight.M3 : Function
        Aerodynamic moment acting along the z-axis of the rocket's body
        frame as a function of time. Expressed in Newtons (N).
    Flight.net_thrust : Function
        Rocket's engine net thrust as a function of time in Newton.
        This is the actual thrust force experienced by the rocket.
        It may be corrected with the atmospheric pressure if a reference
        pressure is defined.
    Flight.aerodynamic_lift : Function
        Resultant force perpendicular to rockets axis due to
        aerodynamic effects as a function of time. Units in N.
        Expressed as a function of time. Can be called or accessed
        as array.
    Flight.aerodynamic_drag : Function
        Resultant force aligned with the rockets axis due to
        aerodynamic effects as a function of time. Units in N.
        Expressed as a function of time. Can be called or accessed
        as array.
    Flight.aerodynamic_bending_moment : Function
        Resultant moment perpendicular to rocket's axis due to
        aerodynamic effects as a function of time. Units in N m.
        Expressed as a function of time. Can be called or accessed
        as array.
    Flight.aerodynamic_spin_moment : Function
        Resultant moment aligned with the rockets axis due to
        aerodynamic effects as a function of time. Units in N m.
        Expressed as a function of time. Can be called or accessed
        as array.
    Flight.rail_button1_normal_force : Function
        Upper rail button normal force in N as a function
        of time.
    Flight.max_rail_button1_normal_force : float
        Maximum upper rail button normal force experienced
        during rail flight phase in N.
    Flight.rail_button1_shear_force : Function
        Upper rail button shear force in N as a function
        of time.
    Flight.max_rail_button1_shear_force : float
        Maximum upper rail button shear force experienced
        during rail flight phase in N.
    Flight.rail_button2_normal_force : Function
        Lower rail button normal force in N as a function
        of time.
    Flight.max_rail_button2_normal_force : float
        Maximum lower rail button normal force experienced
        during rail flight phase in N.
    Flight.rail_button2_shear_force : Function
        Lower rail button shear force in N as a function
        of time.
    Flight.max_rail_button2_shear_force : float
        Maximum lower rail button shear force experienced
        during rail flight phase in N.
    Flight.rotational_energy : Function
        Rocket's rotational kinetic energy as a function of time.
        Units in J.
    Flight.translational_energy : Function
        Rocket's translational kinetic energy as a function of time.
        Units in J.
    Flight.kinetic_energy : Function
        Rocket's total kinetic energy as a function of time.
        Units in J.
    Flight.potential_energy : Function
        Rocket's gravitational potential energy as a function of
        time. Units in J.
    Flight.total_energy : Function
        Rocket's total mechanical energy as a function of time.
        Units in J.
    Flight.thrust_power : Function
        Rocket's engine thrust power output as a function
        of time in Watts.
    Flight.drag_power : Function
        Aerodynamic drag power output as a function
        of time in Watts.
    Flight.attitude_frequency_response : Function
        Fourier Frequency Analysis of the rocket's attitude angle.
        Expressed as the absolute value of the magnitude as a function
        of frequency in Hz.
    Flight.omega1_frequency_response : Function
        Fourier Frequency Analysis of the rocket's angular velocity omega 1.
        Expressed as the absolute value of the magnitude as a function
        of frequency in Hz.
    Flight.omega2_frequency_response : Function
        Fourier Frequency Analysis of the rocket's angular velocity omega 2.
        Expressed as the absolute value of the magnitude as a function
        of frequency in Hz.
    Flight.omega3_frequency_response : Function
        Fourier Frequency Analysis of the rocket's angular velocity omega 3.
        Expressed as the absolute value of the magnitude as a function
        of frequency in Hz.
    Flight.static_margin : Function
        Rocket's static margin during flight in calibers.
    Flight.stability_margin : Function
            Rocket's stability margin during flight, in calibers.
    Flight.initial_stability_margin : float
        Rocket's initial stability margin in calibers.
    Flight.out_of_rail_stability_margin : float
        Rocket's stability margin in calibers when it leaves the rail.
    Flight.stream_velocity_x : Function
        Freestream velocity x (East) component, in m/s, as a function of
        time.
    Flight.stream_velocity_y : Function
        Freestream velocity y (North) component, in m/s, as a function of
        time.
    Flight.stream_velocity_z : Function
        Freestream velocity z (up) component, in m/s, as a function of
        time.
    Flight.free_stream_speed : Function
        Freestream velocity magnitude, in m/s, as a function of time.
    Flight.apogee_freestream_speed : float
        Freestream speed of the rocket at apogee in m/s.
    Flight.mach_number : Function
        Rocket's Mach number defined as its freestream speed
        divided by the speed of sound at its altitude. Expressed
        as a function of time.
    Flight.max_mach_number : float
        Rocket's maximum Mach number experienced during flight.
    Flight.max_mach_number_time : float
        Time at which the rocket experiences the maximum Mach number.
    Flight.reynolds_number : Function
        Rocket's Reynolds number, using its diameter as reference
        length and free_stream_speed as reference velocity. Expressed
        as a function of time.
    Flight.max_reynolds_number : float
        Rocket's maximum Reynolds number experienced during flight.
    Flight.max_reynolds_number_time : float
        Time at which the rocket experiences the maximum Reynolds number.
    Flight.dynamic_pressure : Function
        Dynamic pressure experienced by the rocket in Pa as a function
        of time, defined by 0.5*rho*V^2, where rho is air density and V
        is the freestream speed.
    Flight.max_dynamic_pressure : float
        Maximum dynamic pressure, in Pa, experienced by the rocket.
    Flight.max_dynamic_pressure_time : float
        Time at which the rocket experiences maximum dynamic pressure.
    Flight.total_pressure : Function
        Total pressure experienced by the rocket in Pa as a function
        of time.
    Flight.max_total_pressure : float
        Maximum total pressure, in Pa, experienced by the rocket.
    Flight.max_total_pressure_time : float
        Time at which the rocket experiences maximum total pressure.
    Flight.angle_of_attack : Function
        Rocket's angle of attack in degrees as a function of time.
        Defined as the minimum angle between the attitude vector and
        the freestream velocity vector. Can be called or accessed as
        array.
    Flight.simulation_mode : str
        Simulation mode for the flight. Can be "6 DOF" or "3 DOF".
    Flight.rail_button1_bending_moment : Function
        Internal bending moment at upper rail button attachment point in N·m
        as a function of time. Calculated using beam theory during rail phase.
    Flight.max_rail_button1_bending_moment : float
        Maximum internal bending moment experienced at upper rail button
        attachment point during rail flight phase in N·m.
    Flight.rail_button2_bending_moment : Function
        Internal bending moment at lower rail button attachment point in N·m
        as a function of time. Calculated using beam theory during rail phase.
    Flight.max_rail_button2_bending_moment : float
        Maximum internal bending moment experienced at lower rail button
        attachment point during rail flight phase in N·m.
    """

    def __init__(  # pylint: disable=too-many-arguments,too-many-statements
        self,
        rocket,
        environment,
        rail_length,
        inclination=80.0,
        heading=90.0,
        initial_solution=None,
        terminate_on_apogee=False,
        max_time=600,
        max_time_step=np.inf,
        min_time_step=0,
        rtol=1e-6,
        atol=None,
        time_overshoot=True,
        verbose=False,
        name="Flight",
        equations_of_motion="standard",
        ode_solver="LSODA",
        simulation_mode="6 DOF",
        weathercock_coeff=0.0,
    ):
        """Run a trajectory simulation.

        Parameters
        ----------
        rocket : Rocket
            Rocket to simulate.
        environment : Environment
            Environment to run simulation on.
        rail_length : int, float
            Length in which the rocket will be attached to the rail, only
            moving along a fixed direction, that is, the line parallel to the
            rail. Currently, if the an initial_solution is passed, the rail
            length is not used.
        inclination : int, float, optional
            Rail inclination angle relative to ground, given in degrees.
            Default is 80.
        heading : int, float, optional
            Heading angle relative to north given in degrees.
            Default is 90, which points in the x (east) direction.
        initial_solution : array, Flight, optional
            Initial solution array to be used. Format is:

            .. code-block:: python

                initial_solution = [
                    self.t_initial,
                    x_init, y_init, z_init,
                    vx_init, vy_init, vz_init,
                    e0_init, e1_init, e2_init, e3_init,
                    w1_init, w2_init, w3_init
                ]

            If a Flight object is used, the last state vector will be
            used as initial solution. If None, the initial solution will start
            with all null values, except for the euler parameters which will be
            calculated based on given values of inclination and heading.
            Default is None.
        terminate_on_apogee : boolean, optional
            Whether to terminate simulation when rocket reaches apogee.
            Default is False.
        max_time : int, float, optional
            Maximum time in which to simulate trajectory in seconds.
            Using this without setting a max_time_step may cause unexpected
            errors. Default is 600.
        max_time_step : int, float, optional
            Maximum time step to use during integration in seconds.
            Default is 0.01.
        min_time_step : int, float, optional
            Minimum time step to use during integration in seconds.
            Default is 0.01.
        rtol : float, array, optional
            Maximum relative error tolerance to be tolerated in the
            integration scheme. Can be given as array for each
            state space variable. Default is 1e-6.
        atol : float, optional
            Maximum absolute error tolerance to be tolerated in the
            integration scheme. Can be given as array for each
            state space variable. Default is 6*[1e-3] + 4*[1e-6] + 3*[1e-3].
        time_overshoot : bool, optional
            If True, decouples ODE time step from parachute and controller
            trigger functions sampling rate. The time steps can overshoot the
            necessary trigger function evaluation points and then interpolation
            is used to calculate them and feed the triggers. Can greatly improve
            run time in some cases. Default is True.
        verbose : bool, optional
            If true, verbose mode is activated. Default is False.
        name : str, optional
            Name of the flight. Default is "Flight".
        equations_of_motion : str, optional
            Type of equations of motion to use. Can be "standard" or
            "solid_propulsion". Default is "standard". Solid propulsion is a
            more restricted set of equations of motion that only works for
            solid propulsion rockets. Such equations were used in RocketPy v0
            and are kept here for backwards compatibility.
        ode_solver : str, ``scipy.integrate.OdeSolver``, optional
            Integration method to use to solve the equations of motion ODE.
            Available options are: 'RK23', 'RK45', 'DOP853', 'Radau', 'BDF',
            'LSODA' from ``scipy.integrate.solve_ivp``.
            Default is 'LSODA', which is recommended for most flights.
            A custom ``scipy.integrate.OdeSolver`` can be passed as well.
            For more information on the integration methods, see the scipy
            documentation [1]_.
        weathercock_coeff : float, optional
            Proportionality coefficient (rate coefficient) for the alignment rate of the rocket's body axis
            with the relative wind direction in 3-DOF simulations, in rad/s. The actual angular velocity
            applied to align the rocket is calculated as ``weathercock_coeff * sin(angle)``, where ``angle``
            is the angle between the rocket's axis and the wind direction. A higher value means faster alignment
            (quasi-static weathercocking). This parameter is only used when simulation_mode is '3 DOF'.
            Default is 0.0 to mimic a pure 3-DOF simulation without any weathercocking (fixed attitude).
            Set to a positive value to enable quasi-static weathercocking behaviour.


        Returns
        -------
        None

        References
        ----------
        .. [1] https://docs.scipy.org/doc/scipy/reference/generated/scipy.integrate.solve_ivp.html
        """
        # Save arguments
        self.env = environment
        self.rocket = rocket
        self.rail_length = rail_length
        if self.rail_length <= 0:  # pragma: no cover
            raise ValueError("Rail length must be a positive value.")
        self.parachutes = self.rocket.parachutes[:]
        self.inclination = inclination
        self.heading = heading
        self.max_time = max_time
        self.max_time_step = max_time_step
        self.min_time_step = min_time_step
        self.rtol = rtol
        self.atol = atol or 6 * [1e-3] + 4 * [1e-6] + 3 * [1e-3]
        self.initial_solution = initial_solution
        self.time_overshoot = time_overshoot
        self.terminate_on_apogee = terminate_on_apogee
        self.name = name
        self.equations_of_motion = equations_of_motion
        self.simulation_mode = simulation_mode
        self.ode_solver = ode_solver
        self.weathercock_coeff = weathercock_coeff

        # Controller initialization
        self.__init_controllers()

        # Flight initialization
        self.__init_solution_monitors()
        self.__init_equations_of_motion()
        self.__init_solver_monitors()

        # Create known flight phases
        self.flight_phases = self.FlightPhases()
        self.flight_phases.add_phase(
            self.t_initial, self.initial_derivative, clear=False
        )
        self.flight_phases.add_phase(self.max_time)

        # Simulate flight
        self.__simulate(verbose)

        # Initialize prints and plots objects
        self.prints = _FlightPrints(self)
        self.plots = _FlightPlots(self)

    def __repr__(self):
        return (
            f"<Flight(rocket= {self.rocket}, "
            f"environment= {self.env}, "
            f"rail_length= {self.rail_length}, "
            f"inclination= {self.inclination}, "
            f"heading = {self.heading},"
            f"name= {self.name})>"
        )

    # pylint: disable=too-many-nested-blocks, too-many-branches, too-many-locals,too-many-statements
    def __simulate(self, verbose):
        """Simulate the flight trajectory."""
        for phase_index, phase in self.time_iterator(self.flight_phases):
            # Determine maximum time for this flight phase
            phase.time_bound = self.flight_phases[phase_index + 1].t

            # Evaluate callbacks
            for callback in phase.callbacks:
                callback(self)

            # Create solver for this flight phase
            self.function_evaluations.append(0)

            phase.solver = self._solver(
                phase.derivative,
                t0=phase.t,
                y0=self.y_sol,
                t_bound=phase.time_bound,
                rtol=self.rtol,
                atol=self.atol,
                max_step=self.max_time_step,
                min_step=self.min_time_step,
            )

            # Initialize phase time nodes
            self.__setup_phase_time_nodes(phase)

            # Iterate through time nodes
            for node_index, node in self.time_iterator(phase.time_nodes):
                # Determine time bound for this time node
                node.time_bound = phase.time_nodes[node_index + 1].t
                phase.solver.t_bound = node.time_bound
                if self.__is_lsoda:
                    phase.solver._lsoda_solver._integrator.rwork[0] = (
                        phase.solver.t_bound
                    )
                    phase.solver._lsoda_solver._integrator.call_args[4] = (
                        phase.solver._lsoda_solver._integrator.rwork
                    )
                phase.solver.status = "running"

                # Feed required parachute and discrete controller triggers
                # TODO: parachutes should be moved to controllers
                for callback in node.callbacks:
                    callback(self)

                self.__process_sensors_and_controllers_at_current_node(node, phase)

                for controller in node._controllers:
                    controller(
                        self.t,
                        self.y_sol,
                        self.solution,
                        self.sensors,
                        self.env,
                    )

                for parachute in node.parachutes:
                    # Calculate and save pressure signal
                    (
                        noisy_pressure,
                        height_above_ground_level,
                    ) = self.__calculate_and_save_pressure_signals(
                        parachute, node.t, self.y_sol[2]
                    )
                    if parachute.triggerfunc(
                        noisy_pressure,
                        height_above_ground_level,
                        self.y_sol,
                        self.sensors,
                    ):
                        # Remove parachute from flight parachutes
                        self.parachutes.remove(parachute)
                        # Create phase for time after detection and before inflation
                        # Must only be created if parachute has any lag
                        i = 1
                        if parachute.lag != 0:
                            self.flight_phases.add_phase(
                                node.t,
                                phase.derivative,
                                clear=True,
                                index=phase_index + i,
                            )
                            i += 1
                        # Create flight phase for time after inflation
                        callbacks = [
                            lambda self, parachute_cd_s=parachute.cd_s: setattr(
                                self, "parachute_cd_s", parachute_cd_s
                            ),
                            lambda self, parachute_radius=parachute.radius: setattr(
                                self, "parachute_radius", parachute_radius
                            ),
                            lambda self, parachute_height=parachute.height: setattr(
                                self, "parachute_height", parachute_height
                            ),
                            lambda self, parachute_porosity=parachute.porosity: setattr(
                                self, "parachute_porosity", parachute_porosity
                            ),
                            lambda self,
                            added_mass_coefficient=parachute.added_mass_coefficient: setattr(
                                self,
                                "parachute_added_mass_coefficient",
                                added_mass_coefficient,
                            ),
                        ]
                        self.flight_phases.add_phase(
                            node.t + parachute.lag,
                            self.u_dot_parachute,
                            callbacks,
                            clear=False,
                            index=phase_index + i,
                        )
                        # Prepare to leave loops and start new flight phase
                        phase.time_nodes.flush_after(node_index)
                        phase.time_nodes.add_node(self.t, [], [], [])
                        phase.solver.status = "finished"
                        # Save parachute event
                        self.parachute_events.append([self.t, parachute])
                if self.__check_and_handle_parachute_triggers(
                    node, phase, phase_index, node_index
                ):
                    break  # Stop simulation if parachute is deployed

                # Step through simulation
                while phase.solver.status == "running":
                    # Execute solver step, log solution and function evaluations
                    phase.solver.step()
                    self.solution += [[phase.solver.t, *phase.solver.y]]
                    self.function_evaluations.append(phase.solver.nfev)

                    # Update time and state
                    self.t = phase.solver.t
                    self.y_sol = phase.solver.y
                    if verbose:
                        print(f"Current Simulation Time: {self.t:3.4f} s", end="\r")

                    if self.__check_simulation_events(phase, phase_index, node_index):
                        break  # Stop if simulation termination event occurred

                    # Process overshootable time nodes if enabled
                    if self.time_overshoot and self.__process_overshootable_nodes(
                        phase, phase_index, node_index
                    ):
                        break

                    # If controlled flight, post process must be done on sim time
                    # Post-process controllers if needed
                    if self._controllers:
                        phase.derivative(self.t, self.y_sol, post_processing=True)

        self.t_final = self.t
        self.__transform_pressure_signals_lists_to_functions()
        if self._controllers:
            # cache post process variables
            self.__evaluate_post_process = np.array(self.__post_processed_variables)
        if self.sensors:
            self.__cache_sensor_data()
        if verbose:
            print(f"\n>>> Simulation Completed at Time: {self.t:3.4f} s")

    def __setup_phase_time_nodes(self, phase):
        """Set up time nodes for the current phase.

        Parameters
        ----------
        phase : FlightPhase
            The current flight phase.
        """
        phase.time_nodes = self.TimeNodes()

        # Add first time node
        phase.time_nodes.add_node(phase.t, [], [], [])

        if self.time_overshoot is False:
            phase.time_nodes.add_parachutes(self.parachutes, phase.t, phase.time_bound)
            phase.time_nodes.add_sensors(self.rocket.sensors, phase.t, phase.time_bound)
            phase.time_nodes.add_controllers(
                self._controllers, phase.t, phase.time_bound
            )

        # Add last time node
        phase.time_nodes.add_node(phase.time_bound, [], [], [])

        # Organize time nodes
        phase.time_nodes.sort()
        phase.time_nodes.merge()

        # Clear triggers from first time node if necessary
        if phase.clear:
            phase.time_nodes[0].parachutes = []
            phase.time_nodes[0].callbacks = []

    def __process_sensors_and_controllers_at_current_node(self, node, phase):
        """Process sensors and controllers at the current node.

        Parameters
        ----------
        node : TimeNode
            The current time node.
        phase : FlightPhase
            The current flight phase.
        """
        if self.sensors:
            u_dot = phase.derivative(self.t, self.y_sol)
            self.__measure_sensors(node._component_sensors, u_dot)

        for controller in node._controllers:
            controller(
                self.t,
                self.y_sol,
                self.solution,
                self.sensors,
            )

    def __measure_sensors(self, component_sensors, u_dot, t=None, y_sol=None):
        """Measure sensors with the given state and derivative.

        Parameters
        ----------
        component_sensors : list
            List of (sensor, position) tuples.
        u_dot : array_like
            State derivative vector.
        t : float, optional
            Time for measurement. If None, uses self.t.
        y_sol : array_like, optional
            State vector. If None, uses self.y_sol.
        """
        if t is None:
            t = self.t
        if y_sol is None:
            y_sol = self.y_sol

        for sensor, position in component_sensors:
            relative_position = position - self.rocket._csys * Vector(
                [0, 0, self.rocket.center_of_dry_mass_position]
            )
            sensor.measure(
                t,
                u=y_sol,
                u_dot=u_dot,
                relative_position=relative_position,
                environment=self.env,
                gravity=self.env.gravity.get_value_opt(
                    y_sol[2] if len(y_sol) > 2 else self.solution[-1][3]
                ),
                pressure=self.env.pressure,
                earth_radius=self.env.earth_radius,
                initial_coordinates=(self.env.latitude, self.env.longitude),
            )

    def __check_and_handle_parachute_triggers(
        self, node, phase, phase_index, node_index
    ):
        """Check for parachute triggers and handle deployment.

        Parameters
        ----------
        node : TimeNode
            The current time node.
        phase : FlightPhase
            The current flight phase.
        phase_index : int
            The index of the current phase.
        node_index : int
            The index of the current node.

        Returns
        -------
        bool
            True if a parachute was triggered and the phase should break.
        """
        for parachute in node.parachutes:
            # Calculate and save pressure signal
            (
                noisy_pressure,
                height_above_ground_level,
            ) = self.__calculate_and_save_pressure_signals(
                parachute, node.t, self.y_sol[2]
            )
            if not parachute.triggerfunc(
                noisy_pressure,
                height_above_ground_level,
                self.y_sol,
                self.sensors,
            ):
                continue  # Check next parachute

            # Remove parachute from flight parachutes (if not already removed)
            if parachute in self.parachutes:
                self.parachutes.remove(parachute)
            else:
                continue  # Parachute already triggered, skip to next

            # Create phase for time after detection and before inflation
            # Must only be created if parachute has any lag
            i = 1
            if parachute.lag != 0:
                self.flight_phases.add_phase(
                    node.t,
                    phase.derivative,
                    clear=True,
                    index=phase_index + i,
                )
                i += 1

            # Create flight phase for time after inflation
            callbacks = [
                lambda self, parachute_cd_s=parachute.cd_s: setattr(
                    self, "parachute_cd_s", parachute_cd_s
                ),
                lambda self, parachute_radius=parachute.radius: setattr(
                    self, "parachute_radius", parachute_radius
                ),
                lambda self, parachute_height=parachute.height: setattr(
                    self, "parachute_height", parachute_height
                ),
                lambda self, parachute_porosity=parachute.porosity: setattr(
                    self, "parachute_porosity", parachute_porosity
                ),
                lambda self,
                added_mass_coefficient=parachute.added_mass_coefficient: setattr(
                    self,
                    "parachute_added_mass_coefficient",
                    added_mass_coefficient,
                ),
            ]
            self.flight_phases.add_phase(
                node.t + parachute.lag,
                self.u_dot_parachute,
                callbacks,
                clear=False,
                index=phase_index + i,
            )

            # Prepare to leave loops and start new flight phase
            phase.time_nodes.flush_after(node_index)
            phase.time_nodes.add_node(self.t, [], [], [])
            phase.solver.status = "finished"
            self.parachute_events.append([self.t, parachute])
            return True

        return False

    def __check_simulation_events(self, phase, phase_index, node_index):
        """Check for simulation events like out of rail, apogee, and impact.

        Parameters
        ----------
        phase : FlightPhase
            The current flight phase.
        phase_index : int
            The index of the current phase.
        node_index : int
            The index of the current node.

        Returns
        -------
        bool
            True if an event occurred and the simulation should break.
        """
        # Check for first out of rail event
        if len(self.out_of_rail_state) == 1 and (
            self.y_sol[0] ** 2
            + self.y_sol[1] ** 2
            + (self.y_sol[2] - self.env.elevation) ** 2
            >= self.effective_1rl**2
        ):
            return self.__handle_out_of_rail_event(phase, phase_index, node_index)

        # Check for apogee event
        # TODO: negative vz doesn't really mean apogee. Improve this.
        if len(self.apogee_state) == 1 and self.y_sol[5] < 0:
            return self.__handle_apogee_event(phase, phase_index, node_index)

        # Check for impact event
        if self.y_sol[2] < self.env.elevation:
            return self.__handle_impact_event(phase, phase_index, node_index)

        return False

    def __handle_out_of_rail_event(self, phase, phase_index, node_index):
        """Handle the out of rail event.

        Parameters
        ----------
        phase : FlightPhase
            The current flight phase.
        phase_index : int
            The index of the current phase.
        node_index : int
            The index of the current node.

        Returns
        -------
        bool
            True to indicate the simulation should break.
        """
        # Check exactly when it went out using root finding
        # Disconsider elevation
        self.solution[-2][3] -= self.env.elevation
        self.solution[-1][3] -= self.env.elevation
        # Get points
        y0 = sum(self.solution[-2][i] ** 2 for i in [1, 2, 3]) - self.effective_1rl**2
        yp0 = 2 * sum(
            self.solution[-2][i] * self.solution[-2][i + 3] for i in [1, 2, 3]
        )
        t1 = self.solution[-1][0] - self.solution[-2][0]
        y1 = sum(self.solution[-1][i] ** 2 for i in [1, 2, 3]) - self.effective_1rl**2
        yp1 = 2 * sum(
            self.solution[-1][i] * self.solution[-1][i + 3] for i in [1, 2, 3]
        )
        # Put elevation back
        self.solution[-2][3] += self.env.elevation
        self.solution[-1][3] += self.env.elevation
        # Cubic Hermite interpolation (ax**3 + bx**2 + cx + d)
        a, b, c, d = calculate_cubic_hermite_coefficients(
            0,
            float(phase.solver.step_size),
            y0,
            yp0,
            y1,
            yp1,
        )
        a += 1e-5  # TODO: why??
        # Find roots
        t_roots = find_roots_cubic_function(a, b, c, d)
        # Find correct root
        valid_t_root = [
            t_root.real
            for t_root in t_roots
            if 0 < t_root.real < t1 and abs(t_root.imag) < 0.001
        ]
        if len(valid_t_root) > 1:  # pragma: no cover
            raise ValueError("Multiple roots found when solving for rail exit time.")
        if len(valid_t_root) == 0:  # pragma: no cover
            raise ValueError("No valid roots found when solving for rail exit time.")
        # Determine final state when upper button is going out of rail
        self.t = valid_t_root[0] + self.solution[-2][0]
        interpolator = phase.solver.dense_output()
        self.y_sol = interpolator(self.t)
        self.solution[-1] = [self.t, *self.y_sol]
        self.out_of_rail_time = self.t
        self.out_of_rail_time_index = len(self.solution) - 1
        self.out_of_rail_state = self.y_sol
        # Create new flight phase
        self.flight_phases.add_phase(
            self.t,
            self.u_dot_generalized,
            index=phase_index + 1,
        )
        # Prepare to leave loops and start new flight phase
        phase.time_nodes.flush_after(node_index)
        phase.time_nodes.add_node(self.t, [], [], [])
        phase.solver.status = "finished"
        return True

    def __handle_apogee_event(self, phase, phase_index, node_index):
        """Handle the apogee event.

        Parameters
        ----------
        phase : FlightPhase
            The current flight phase.
        phase_index : int
            The index of the current phase.
        node_index : int
            The index of the current node.

        Returns
        -------
        bool
            True if simulation should break, False otherwise.
        """
        # Assume linear vz(t) to detect when vz = 0
        t0, vz0 = self.solution[-2][0], self.solution[-2][6]
        t1, vz1 = self.solution[-1][0], self.solution[-1][6]
        t_root = find_root_linear_interpolation(t0, t1, vz0, vz1, 0)
        # Fetch state at t_root
        interpolator = phase.solver.dense_output()
        self.apogee_state = interpolator(t_root)
        # Store apogee data
        self.apogee_time = t_root
        self.apogee_x = self.apogee_state[0]
        self.apogee_y = self.apogee_state[1]
        self.apogee = self.apogee_state[2]

        if self.terminate_on_apogee:
            self.t = self.t_final = t_root
            # Roll back solution
            self.solution[-1] = [self.t, *self.apogee_state]
            # Set last flight phase
            self.flight_phases.flush_after(phase_index)
            self.flight_phases.add_phase(self.t)
            # Prepare to leave loops and start new flight phase
            phase.time_nodes.flush_after(node_index)
            phase.time_nodes.add_node(self.t, [], [], [])
            phase.solver.status = "finished"
            return True
        elif len(self.solution) > 2:
            # adding the apogee state to solution increases accuracy
            # we can only do this if the apogee is not the first state
            self.solution.insert(-1, [t_root, *self.apogee_state])
        return False

    def __handle_impact_event(self, phase, phase_index, node_index):
        """Handle the impact event.

        Parameters
        ----------
        phase : FlightPhase
            The current flight phase.
        phase_index : int
            The index of the current phase.
        node_index : int
            The index of the current node.

        Returns
        -------
        bool
            True to indicate the simulation should break.
        """
        # Check exactly when it happened using root finding
        # Cubic Hermite interpolation (ax**3 + bx**2 + cx + d)
        a, b, c, d = calculate_cubic_hermite_coefficients(
            x0=0,  # t0
            x1=float(phase.solver.step_size),  # t1 - t0
            y0=float(self.solution[-2][3] - self.env.elevation),  # z0
            yp0=float(self.solution[-2][6]),  # vz0
            y1=float(self.solution[-1][3] - self.env.elevation),  # z1
            yp1=float(self.solution[-1][6]),  # vz1
        )
        # Find roots
        t_roots = find_roots_cubic_function(a, b, c, d)
        # Find correct root
        t1 = self.solution[-1][0] - self.solution[-2][0]
        valid_t_root = [
            t_root.real
            for t_root in t_roots
            if abs(t_root.imag) < 0.001 and 0 < t_root.real < t1
        ]
        if len(valid_t_root) > 1:  # pragma: no cover
            raise ValueError("Multiple roots found when solving for impact time.")
        # Determine impact state at t_root
        self.t = self.t_final = valid_t_root[0] + self.solution[-2][0]
        interpolator = phase.solver.dense_output()
        self.y_sol = self.impact_state = interpolator(self.t)
        # Roll back solution
        self.solution[-1] = [self.t, *self.y_sol]
        # Save impact state
        self.x_impact = self.impact_state[0]
        self.y_impact = self.impact_state[1]
        self.z_impact = self.impact_state[2]
        self.impact_velocity = self.impact_state[5]
        # Set last flight phase
        self.flight_phases.flush_after(phase_index)
        self.flight_phases.add_phase(self.t)
        # Prepare to leave loops and start new flight phase
        phase.time_nodes.flush_after(node_index)
        phase.time_nodes.add_node(self.t, [], [], [])
        phase.solver.status = "finished"
        return True

    def __process_overshootable_nodes(self, phase, phase_index, node_index):
        """Process overshootable time nodes for parachutes, controllers, and sensors.

        Parameters
        ----------
        phase : FlightPhase
            The current flight phase.
        phase_index : int
            The index of the current phase.
        node_index : int
            The index of the current node.

        Returns
        -------
        bool
            True if a parachute was triggered and the simulation should break.
        """
        overshootable_nodes = self.TimeNodes()

        overshootable_nodes.add_parachutes(
            self.parachutes, self.solution[-2][0], self.t
        )
        overshootable_nodes.add_controllers(
            self._controllers, self.solution[-2][0], self.t
        )
        overshootable_nodes.add_sensors(
            self.rocket.sensors, self.solution[-2][0], self.t
        )

        # Add last time node (always skipped)
        overshootable_nodes.add_node(self.t, [], [], [])

        if len(overshootable_nodes) < 1:
            return False  # Early exit

        overshootable_nodes.sort()
        overshootable_nodes.merge()

        # Clear if necessary
        if overshootable_nodes[0].t == phase.t and phase.clear:
            overshootable_nodes[0].parachutes = []
            overshootable_nodes[0].callbacks = []

        # Feed overshootable time nodes trigger
        interpolator = phase.solver.dense_output()
        for overshootable_index, overshootable_node in self.time_iterator(
            overshootable_nodes
        ):
            # Calculate state at node time
            overshootable_node.y_sol = interpolator(overshootable_node.t)

            # Check for parachute triggers
            if self.__check_overshootable_parachute_triggers(
                overshootable_node,
                overshootable_nodes,
                overshootable_index,
                phase,
                phase_index,
                node_index,
            ):
                return True

            # Process controllers at overshootable node
            for controller in overshootable_node._controllers:
                controller(
                    overshootable_node.t,
                    overshootable_node.y_sol,
                    self.solution,
                    self.sensors,
                )

            # Process sensors at overshootable node
            if overshootable_node._component_sensors:
                # Calculate u_dot for sensors at interpolated state
                u_dot = phase.derivative(overshootable_node.t, overshootable_node.y_sol)
                self.__measure_sensors(
                    overshootable_node._component_sensors,
                    u_dot,
                    overshootable_node.t,
                    overshootable_node.y_sol,
                )
        return False

    def __check_overshootable_parachute_triggers(
        self,
        overshootable_node,
        overshootable_nodes,
        overshootable_index,
        phase,
        phase_index,
        node_index,
    ):
        """Check for parachute triggers in overshootable nodes.

        Parameters
        ----------
        overshootable_node : TimeNode
            The current overshootable node.
        overshootable_nodes : TimeNodes
            The overshootable nodes collection.
        overshootable_index : int
            Index of the current overshootable node.
        phase : FlightPhase
            The current flight phase.
        phase_index : int
            The index of the current phase.
        node_index : int
            The index of the current node.

        Returns
        -------
        bool
            True if a parachute was triggered and the simulation should break.
        """
        for parachute in overshootable_node.parachutes:
            # Calculate and save pressure signal
            (
                noisy_pressure,
                height_above_ground_level,
            ) = self.__calculate_and_save_pressure_signals(
                parachute,
                overshootable_node.t,
                overshootable_node.y_sol[2],
            )

            # Check for parachute trigger
            if not parachute.triggerfunc(
                noisy_pressure,
                height_above_ground_level,
                overshootable_node.y_sol,
                self.sensors,
            ):
                continue  # Check next parachute

            # Remove parachute from flight parachutes
            self.parachutes.remove(parachute)

            # Create phase for time after detection and before inflation
            # Must only be created if parachute has any lag
            i = 1
            if parachute.lag != 0:
                self.flight_phases.add_phase(
                    overshootable_node.t,
                    phase.derivative,
                    clear=True,
                    index=phase_index + i,
                )
                i += 1

            # Create flight phase for time after inflation
            callbacks = [
                lambda self, parachute_cd_s=parachute.cd_s: setattr(
                    self, "parachute_cd_s", parachute_cd_s
                ),
                lambda self, parachute_radius=parachute.radius: setattr(
                    self, "parachute_radius", parachute_radius
                ),
                lambda self, parachute_height=parachute.height: setattr(
                    self, "parachute_height", parachute_height
                ),
                lambda self, parachute_porosity=parachute.porosity: setattr(
                    self, "parachute_porosity", parachute_porosity
                ),
                lambda self,
                added_mass_coefficient=parachute.added_mass_coefficient: setattr(
                    self,
                    "parachute_added_mass_coefficient",
                    added_mass_coefficient,
                ),
            ]
            self.flight_phases.add_phase(
                overshootable_node.t + parachute.lag,
                self.u_dot_parachute,
                callbacks,
                clear=False,
                index=phase_index + i,
            )

            # Rollback history
            self.t = overshootable_node.t
            self.y_sol = overshootable_node.y_sol
            self.solution[-1] = [overshootable_node.t, *overshootable_node.y_sol]

            # Prepare to leave loops and start new flight phase
            overshootable_nodes.flush_after(overshootable_index)
            phase.time_nodes.flush_after(node_index)
            phase.time_nodes.add_node(self.t, [], [], [])
            phase.solver.status = "finished"

            # Save parachute event
            self.parachute_events.append([self.t, parachute])
            return True

        return False

    def __calculate_and_save_pressure_signals(self, parachute, t, z):
        """Gets noise and pressure signals and saves them in the parachute
        object given the current time and altitude.

        Parameters
        ----------
        parachute : Parachute
            The parachute object to calculate signals for.
        t : float
            The current time in seconds.
        z : float
            The altitude above sea level in meters.

        Returns
        -------
        tuple[float, float]
            The noisy pressure and height above ground level.
        """
        # Calculate pressure and noise
        pressure = self.env.pressure.get_value_opt(z)
        noise = parachute.noise_function()
        noisy_pressure = pressure + noise

        # Stores in the parachute object
        parachute.clean_pressure_signal.append([t, pressure])
        parachute.noise_signal.append([t, noise])

        # Gets height above ground level considering noise
        height_above_ground_level = (
            self.env.barometric_height.get_value_opt(noisy_pressure)
            - self.env.elevation
        )

        return noisy_pressure, height_above_ground_level

    def __init_solution_monitors(self):
        # Initialize solution monitors
        self.out_of_rail_time = 0
        self.out_of_rail_time_index = 0
        self.out_of_rail_state = np.array([0])
        self.apogee_state = np.array([0])
        self.apogee = 0
        self.apogee_time = 0
        self.x_impact = 0
        self.y_impact = 0
        self.impact_velocity = 0
        self.impact_state = np.array([0])
        self.parachute_events = []
        self.__post_processed_variables = []

    def __init_flight_state(self):
        """Initialize flight state variables."""
        if self.initial_solution is None:
            # Initialize time and state variables
            self.t_initial = 0
            x_init, y_init, z_init = 0, 0, self.env.elevation
            vx_init, vy_init, vz_init = 0, 0, 0
            w1_init, w2_init, w3_init = 0, 0, 0
            # Initialize attitude
            # Precession / Heading Angle
            self.psi_init = np.radians(-self.heading)
            # Nutation / Attitude Angle
            self.theta_init = np.radians(self.inclination - 90)
            # Spin / Bank Angle
            self.phi_init = 0

            # Consider Rail Buttons position, if there is rail buttons
            try:
                self.phi_init += (
                    self.rocket.rail_buttons[0].component.angular_position_rad
                    if self.rocket._csys == 1
                    else 2 * np.pi
                    - self.rocket.rail_buttons[0].component.angular_position_rad
                )
            except IndexError:
                pass

            # 3-1-3 Euler Angles to Euler Parameters
            e0_init, e1_init, e2_init, e3_init = euler313_to_quaternions(
                self.phi_init, self.theta_init, self.psi_init
            )
            # Store initial conditions
            self.initial_solution = [
                self.t_initial,
                x_init,
                y_init,
                z_init,
                vx_init,
                vy_init,
                vz_init,
                e0_init,
                e1_init,
                e2_init,
                e3_init,
                w1_init,
                w2_init,
                w3_init,
            ]
            # Set initial derivative for rail phase
            self.initial_derivative = self.udot_rail1
        elif isinstance(self.initial_solution, Flight):
            # Initialize time and state variables based on last solution of
            # previous flight
            self.initial_solution = self.initial_solution.solution[-1]
            # Set unused monitors
            self.out_of_rail_state = self.initial_solution[1:]
            self.out_of_rail_time = self.initial_solution[0]
            self.out_of_rail_time_index = 0
            # Set initial derivative for 6-DOF flight phase
            self.initial_derivative = self.u_dot_generalized
        else:
            # Initial solution given, ignore rail phase
            # TODO: Check if rocket is actually out of rail. Otherwise, start at rail
            self.out_of_rail_state = self.initial_solution[1:]
            self.out_of_rail_time = self.initial_solution[0]
            self.out_of_rail_time_index = 0
            self.t_initial = self.initial_solution[0]
            self.initial_derivative = self.u_dot_generalized
        if self._controllers or self.sensors:
            # Handle post process during simulation, get initial accel/forces
            self.initial_derivative(
                self.t_initial, self.initial_solution[1:], post_processing=True
            )

    def __init_solver_monitors(self):
        # Initialize solver monitors
        self.function_evaluations = []
        # Initialize solution state
        self.solution = []
        self.__init_flight_state()

        self.t_initial = self.initial_solution[0]
        self.solution.append(self.initial_solution)
        self.t = self.solution[-1][0]
        self.y_sol = self.solution[-1][1:]

        self.__set_ode_solver(self.ode_solver)

    def __init_equations_of_motion(self):
        """Initialize equations of motion."""
        # Determine if a point-mass model is used.
        is_point_mass = isinstance(self.rocket, PointMassRocket) or (
            hasattr(self.rocket, "motor")
            and isinstance(self.rocket.motor, PointMassMotor)
        )
        # Set simulation mode based on model type.
        if is_point_mass:
            if self.simulation_mode != "3 DOF":
                warnings.warn(
                    "A point-mass model was detected. Simulation mode should be '3 DOF'.",
                    UserWarning,
                )
            self.simulation_mode = "3 DOF"

        # Set the equations of motion based on the final simulation mode.
        if self.simulation_mode == "3 DOF":
            self.u_dot_generalized = self.u_dot_generalized_3dof
        elif self.simulation_mode == "6 DOF":
            self.u_dot_generalized = (
                self.u_dot
                if self.equations_of_motion == "solid_propulsion"
                else self.u_dot_generalized
            )
        else:
            raise ValueError(
                f"Invalid simulation_mode: {self.simulation_mode}. "
                "Must be '3 DOF' or '6 DOF'."
            )

    def __init_controllers(self):
        """Initialize controllers and sensors"""
        self._controllers = self.rocket._controllers[:]
        self.sensors = self.rocket.sensors.get_components()

        # reset controllable object to initial state (only airbrakes for now)
        for air_brakes in self.rocket.air_brakes:
            air_brakes._reset()

        self.sensor_data = {}
        for sensor in self.sensors:
            sensor._reset(self.rocket)  # resets noise and measurement list
            self.sensor_data[sensor] = []

    def __cache_sensor_data(self):
        """Cache sensor data for simulations with sensors."""
        sensor_data = {}
        sensors = []
        for sensor in self.sensors:
            # skip sensors that are used more then once in the rocket
            if sensor not in sensors:
                sensors.append(sensor)
                sensor_data[sensor] = sensor.measured_data[:]
        self.sensor_data = sensor_data

    def __set_ode_solver(self, solver):
        """Sets the ODE solver to be used in the simulation.

        Parameters
        ----------
        solver : str, ``scipy.integrate.OdeSolver``
            Integration method to use to solve the equations of motion ODE,
            or a custom ``scipy.integrate.OdeSolver``.
        """
        if isinstance(solver, OdeSolver):
            self._solver = solver
        else:
            try:
                self._solver = ODE_SOLVER_MAP[solver]
            except KeyError as e:  # pragma: no cover
                raise ValueError(
                    f"Invalid ``ode_solver`` input: {solver}. "
                    f"Available options are: {', '.join(ODE_SOLVER_MAP.keys())}"
                ) from e

        self.__is_lsoda = issubclass(self._solver, LSODA)

    @cached_property
    def effective_1rl(self):
        """Original rail length minus the distance measured from nozzle exit
        to the upper rail button. It assumes the nozzle to be aligned with
        the beginning of the rail."""
        nozzle = self.rocket.nozzle_position
        try:
            rail_buttons = self.rocket.rail_buttons[0]
            upper_r_button = (
                rail_buttons.component.buttons_distance * self.rocket._csys
                + rail_buttons.position.z
            )
        except IndexError:  # No rail buttons defined
            upper_r_button = nozzle
        effective_1rl = self.rail_length - abs(nozzle - upper_r_button)
        return effective_1rl

    @cached_property
    def effective_2rl(self):
        """Original rail length minus the distance measured from nozzle exit
        to the lower rail button. It assumes the nozzle to be aligned with
        the beginning of the rail."""
        nozzle = self.rocket.nozzle_position
        try:
            rail_buttons = self.rocket.rail_buttons[0]
            lower_r_button = rail_buttons.position.z
        except IndexError:  # No rail buttons defined
            lower_r_button = nozzle
        effective_2rl = self.rail_length - abs(nozzle - lower_r_button)
        return effective_2rl

    @cached_property
    def frontal_surface_wind(self):
        """Frontal wind velocity at the surface level. The frontal wind is
        defined as the wind blowing in the direction of the rocket's heading.

        Returns
        -------
        float
            Wind velocity in the frontal direction at the surface level.
        """
        wind_u = self.env.wind_velocity_x.get_value_opt(self.env.elevation)
        wind_v = self.env.wind_velocity_y.get_value_opt(self.env.elevation)
        heading_rad = self.heading * np.pi / 180
        return wind_u * np.sin(heading_rad) + wind_v * np.cos(heading_rad)

    @cached_property
    def lateral_surface_wind(self):
        """Lateral wind velocity at the surface level. The lateral wind is
        defined as the wind blowing perpendicular to the rocket's heading.

        Returns
        -------
        float
            Wind velocity in the lateral direction at the surface level.
        """
        wind_u = self.env.wind_velocity_x.get_value_opt(self.env.elevation)
        wind_v = self.env.wind_velocity_y.get_value_opt(self.env.elevation)
        heading_rad = self.heading * np.pi / 180

        return -wind_u * np.cos(heading_rad) + wind_v * np.sin(heading_rad)

    def __get_drag_coefficient(self, drag_function, mach, z, freestream_velocity_body):
        """Calculate drag coefficient, handling both 1D and multi-dimensional functions.

        Parameters
        ----------
        drag_function : Function
            The drag coefficient function (power_on_drag or power_off_drag)
        mach : float
            Mach number
        z : float
            Altitude in meters
        freestream_velocity_body : Vector or array-like
            Freestream velocity in body frame [stream_vx_b, stream_vy_b, stream_vz_b]

        Returns
        -------
        float
            Drag coefficient value
        """
        # Early return for 1D drag functions (only mach number)
        if not isinstance(drag_function, Function) or not getattr(
            drag_function, "is_multidimensional", False
        ):
            return drag_function.get_value_opt(mach)

        # Multi-dimensional drag function - calculate additional parameters

        # Calculate Reynolds number: Re = rho * V * L / mu
        # where L is characteristic length (rocket diameter)
        rho = self.env.density.get_value_opt(z)
        mu = self.env.dynamic_viscosity.get_value_opt(z)
        freestream_speed = np.linalg.norm(freestream_velocity_body)
        characteristic_length = 2 * self.rocket.radius  # Diameter
        # Defensive: avoid division by zero or non-finite viscosity values.
        # Use a small epsilon fallback if `mu` is zero, negative, NaN or infinite.
        try:
            mu_val = float(mu)
        except (TypeError, ValueError, OverflowError):
            # Only catch errors related to invalid numeric conversion.
            # Avoid catching broad Exception to satisfy linters and
            # allow other unexpected errors to surface.
            mu_val = 0.0
        if not np.isfinite(mu_val) or mu_val <= 0.0:
            mu_safe = 1e-10
        else:
            mu_safe = mu_val

        reynolds = rho * freestream_speed * characteristic_length / mu_safe

        # Calculate angle of attack
        # Angle between freestream velocity and rocket axis (z-axis in body frame)
        # The z component of freestream velocity in body frame
        if hasattr(freestream_velocity_body, "z"):
            stream_vz_b = -freestream_velocity_body.z
        else:
            stream_vz_b = -freestream_velocity_body[2]

        # Normalize and calculate angle
        if freestream_speed > 1e-6:
            cos_alpha = stream_vz_b / freestream_speed
            # Clamp to [-1, 1] to avoid numerical issues
            cos_alpha = np.clip(cos_alpha, -1.0, 1.0)
            alpha_rad = np.arccos(cos_alpha)
            alpha_deg = np.rad2deg(alpha_rad)
        else:
            alpha_deg = 0.0

        # Determine which parameters to pass based on input names
        input_names = [name.lower() for name in drag_function.__inputs__]
        args = []

        for name in input_names:
            if "mach" in name or name == "m":
                args.append(mach)
            elif "reynolds" in name or name == "re":
                args.append(reynolds)
            elif "alpha" in name or name == "a" or "attack" in name:
                args.append(alpha_deg)
            else:
                # Unknown parameter, default to mach
                args.append(mach)

        return drag_function.get_value_opt(*args)

    def udot_rail1(self, t, u, post_processing=False):
        """Calculates derivative of u state vector with respect to time
        when rocket is flying in 1 DOF motion in the rail.

        Parameters
        ----------
        t : float
            Time in seconds
        u : list
            State vector defined by u = [x, y, z, vx, vy, vz, e0, e1,
            e2, e3, omega1, omega2, omega3].
        post_processing : bool, optional
            If True, adds flight data information directly to self
            variables such as self.angle_of_attack. Default is False.

        Return
        ------
        u_dot : list
            State vector defined by u_dot = [vx, vy, vz, ax, ay, az,
            e0dot, e1dot, e2dot, e3dot, alpha1, alpha2, alpha3].

        """
        # Retrieve integration data
        _, _, z, vx, vy, vz, e0, e1, e2, e3, _, _, _ = u

        # Retrieve important quantities
        # Mass
        total_mass_at_t = self.rocket.total_mass.get_value_opt(t)

        # Get freestream speed
        free_stream_speed = (
            (self.env.wind_velocity_x.get_value_opt(z) - vx) ** 2
            + (self.env.wind_velocity_y.get_value_opt(z) - vy) ** 2
            + (vz) ** 2
        ) ** 0.5
        free_stream_mach = free_stream_speed / self.env.speed_of_sound.get_value_opt(z)

        # For rail motion, rocket is constrained - velocity mostly along z-axis in body frame
        # Calculate velocity in body frame (simplified for rail)
        a11 = 1 - 2 * (e2**2 + e3**2)
        a12 = 2 * (e1 * e2 - e0 * e3)
        a13 = 2 * (e1 * e3 + e0 * e2)
        a21 = 2 * (e1 * e2 + e0 * e3)
        a22 = 1 - 2 * (e1**2 + e3**2)
        a23 = 2 * (e2 * e3 - e0 * e1)
        a31 = 2 * (e1 * e3 - e0 * e2)
        a32 = 2 * (e2 * e3 + e0 * e1)
        a33 = 1 - 2 * (e1**2 + e2**2)

        # Freestream velocity in body frame
        wind_vx = self.env.wind_velocity_x.get_value_opt(z)
        wind_vy = self.env.wind_velocity_y.get_value_opt(z)
        stream_vx_b = a11 * (wind_vx - vx) + a21 * (wind_vy - vy) + a31 * (-vz)
        stream_vy_b = a12 * (wind_vx - vx) + a22 * (wind_vy - vy) + a32 * (-vz)
        stream_vz_b = a13 * (wind_vx - vx) + a23 * (wind_vy - vy) + a33 * (-vz)

        drag_coeff = self.__get_drag_coefficient(
            self.rocket.power_on_drag,
            free_stream_mach,
            z,
            [stream_vx_b, stream_vy_b, stream_vz_b],
        )

        # Calculate Forces
        pressure = self.env.pressure.get_value_opt(z)
        net_thrust = max(
            self.rocket.motor.thrust.get_value_opt(t)
            + self.rocket.motor.pressure_thrust(pressure),
            0,
        )
        rho = self.env.density.get_value_opt(z)
        R3 = -0.5 * rho * (free_stream_speed**2) * self.rocket.area * (drag_coeff)

        # Calculate Linear acceleration
        a3 = (R3 + net_thrust) / total_mass_at_t - (
            e0**2 - e1**2 - e2**2 + e3**2
        ) * self.env.gravity.get_value_opt(z)
        if a3 > 0:
            ax = 2 * (e1 * e3 + e0 * e2) * a3
            ay = 2 * (e2 * e3 - e0 * e1) * a3
            az = (1 - 2 * (e1**2 + e2**2)) * a3
        else:
            ax, ay, az = 0, 0, 0

        if post_processing:
            # Use u_dot post processing code for forces, moments and env data
            self.u_dot_generalized(t, u, post_processing=True)
            # Save feasible accelerations
            self.__post_processed_variables[-1][1:7] = [ax, ay, az, 0, 0, 0]

        return [vx, vy, vz, ax, ay, az, 0, 0, 0, 0, 0, 0, 0]

    def udot_rail2(self, t, u, post_processing=False):  # pragma: no cover
        """[Still not implemented] Calculates derivative of u state vector with
        respect to time when rocket is flying in 3 DOF motion in the rail.

        Parameters
        ----------
        t : float
            Time in seconds
        u : list
            State vector defined by u = [x, y, z, vx, vy, vz, e0, e1,
            e2, e3, omega1, omega2, omega3].
        post_processing : bool, optional
            If True, adds flight data information directly to self
            variables such as self.angle_of_attack, by default False

        Returns
        -------
        u_dot : list
            State vector defined by u_dot = [vx, vy, vz, ax, ay, az,
            e0dot, e1dot, e2dot, e3dot, alpha1, alpha2, alpha3].
        """
        # Hey! We will finish this function later, now we just can use u_dot
        return self.u_dot_generalized(t, u, post_processing=post_processing)

    def u_dot(self, t, u, post_processing=False):  # pylint: disable=too-many-locals,too-many-statements
        """Calculates derivative of u state vector with respect to time
        when rocket is flying in 6 DOF motion during ascent out of rail
        and descent without parachute.

        Parameters
        ----------
        t : float
            Time in seconds
        u : list
            State vector defined by u = [x, y, z, vx, vy, vz, e0, e1,
            e2, e3, omega1, omega2, omega3].
        post_processing : bool, optional
            If True, adds flight data information directly to self
            variables such as self.angle_of_attack, by default False

        Returns
        -------
        u_dot : list
            State vector defined by u_dot = [vx, vy, vz, ax, ay, az,
            e0dot, e1dot, e2dot, e3dot, alpha1, alpha2, alpha3].
        """

        # Retrieve integration data
        _, _, z, vx, vy, vz, e0, e1, e2, e3, omega1, omega2, omega3 = u
        # Determine lift force and moment
        omega1, omega2, omega3 = 0, 0, 0
        R1, R2, M1, M2, M3 = 0, 0, 0, 0, 0
        # Thrust correction parameters
        pressure = self.env.pressure.get_value_opt(z)
        # Determine current behavior
        if self.rocket.motor.burn_start_time < t < self.rocket.motor.burn_out_time:
            # Motor burning
            # Retrieve important motor quantities
            # Inertias
            motor_I_33_at_t = self.rocket.motor.I_33.get_value_opt(t)
            motor_I_11_at_t = self.rocket.motor.I_11.get_value_opt(t)
            motor_I_33_derivative_at_t = self.rocket.motor.I_33.differentiate(
                t, dx=1e-6
            )
            motor_I_11_derivative_at_t = self.rocket.motor.I_11.differentiate(
                t, dx=1e-6
            )
            # Mass
            mass_flow_rate_at_t = self.rocket.motor.mass_flow_rate.get_value_opt(t)
            propellant_mass_at_t = self.rocket.motor.propellant_mass.get_value_opt(t)
            # Thrust

            net_thrust = max(
                self.rocket.motor.thrust.get_value_opt(t)
                + self.rocket.motor.pressure_thrust(pressure),
                0,
            )
            # Off center moment
            M1 += self.rocket.thrust_eccentricity_y * net_thrust
            M2 -= self.rocket.thrust_eccentricity_x * net_thrust
        else:
            # Motor stopped
            # Inertias
            (
                motor_I_33_at_t,
                motor_I_11_at_t,
                motor_I_33_derivative_at_t,
                motor_I_11_derivative_at_t,
            ) = (0, 0, 0, 0)
            # Mass
            mass_flow_rate_at_t, propellant_mass_at_t = 0, 0
            # thrust
            net_thrust = 0

        # Retrieve important quantities
        # Inertias
        rocket_dry_I_33 = self.rocket.dry_I_33
        rocket_dry_I_11 = self.rocket.dry_I_11
        # Mass
        rocket_dry_mass = self.rocket.dry_mass  # already with motor's dry mass
        total_mass_at_t = propellant_mass_at_t + rocket_dry_mass
        mu = (propellant_mass_at_t * rocket_dry_mass) / (
            propellant_mass_at_t + rocket_dry_mass
        )
        # Geometry
        # b = -self.rocket.distance_rocket_propellant
        b = (
            -(
                self.rocket.center_of_propellant_position.get_value_opt(0)
                - self.rocket.center_of_dry_mass_position
            )
            * self.rocket._csys
        )
        c = self.rocket.nozzle_to_cdm
        nozzle_radius = self.rocket.motor.nozzle_radius
        # Prepare transformation matrix
        a11 = 1 - 2 * (e2**2 + e3**2)
        a12 = 2 * (e1 * e2 - e0 * e3)
        a13 = 2 * (e1 * e3 + e0 * e2)
        a21 = 2 * (e1 * e2 + e0 * e3)
        a22 = 1 - 2 * (e1**2 + e3**2)
        a23 = 2 * (e2 * e3 - e0 * e1)
        a31 = 2 * (e1 * e3 - e0 * e2)
        a32 = 2 * (e2 * e3 + e0 * e1)
        a33 = 1 - 2 * (e1**2 + e2**2)
        # Transformation matrix: (123) -> (XYZ)
        K = Matrix([[a11, a12, a13], [a21, a22, a23], [a31, a32, a33]])
        Kt = K.transpose

        # Calculate Forces and Moments
        # Get freestream speed
        wind_velocity_x = self.env.wind_velocity_x.get_value_opt(z)
        wind_velocity_y = self.env.wind_velocity_y.get_value_opt(z)
        speed_of_sound = self.env.speed_of_sound.get_value_opt(z)
        free_stream_speed = (
            (wind_velocity_x - vx) ** 2 + (wind_velocity_y - vy) ** 2 + (vz) ** 2
        ) ** 0.5
        free_stream_mach = free_stream_speed / speed_of_sound

        # Get rocket velocity in body frame (needed for drag calculation)
        vx_b = a11 * vx + a21 * vy + a31 * vz
        vy_b = a12 * vx + a22 * vy + a32 * vz
        vz_b = a13 * vx + a23 * vy + a33 * vz

        # Calculate freestream velocity in body frame
        stream_vx_b = (
            a11 * (wind_velocity_x - vx) + a21 * (wind_velocity_y - vy) + a31 * (-vz)
        )
        stream_vy_b = (
            a12 * (wind_velocity_x - vx) + a22 * (wind_velocity_y - vy) + a32 * (-vz)
        )
        stream_vz_b = (
            a13 * (wind_velocity_x - vx) + a23 * (wind_velocity_y - vy) + a33 * (-vz)
        )

        # Determine aerodynamics forces
        # Determine Drag Force
        if t < self.rocket.motor.burn_out_time:
            drag_coeff = self.__get_drag_coefficient(
                self.rocket.power_on_drag,
                free_stream_mach,
                z,
                [stream_vx_b, stream_vy_b, stream_vz_b],
            )
        else:
            drag_coeff = self.__get_drag_coefficient(
                self.rocket.power_off_drag,
                free_stream_mach,
                z,
                [stream_vx_b, stream_vy_b, stream_vz_b],
            )
        rho = self.env.density.get_value_opt(z)
        R3 = -0.5 * rho * (free_stream_speed**2) * self.rocket.area * drag_coeff
        for air_brakes in self.rocket.air_brakes:
            if air_brakes.deployment_level > 0:
                air_brakes_cd = air_brakes.drag_coefficient.get_value_opt(
                    air_brakes.deployment_level, free_stream_mach
                )
                air_brakes_force = (
                    -0.5
                    * rho
                    * (free_stream_speed**2)
                    * air_brakes.reference_area
                    * air_brakes_cd
                )
                if air_brakes.override_rocket_drag:
                    R3 = air_brakes_force  # Substitutes rocket drag coefficient
                else:
                    R3 += air_brakes_force
        # Off center moment
        M1 += self.rocket.cp_eccentricity_y * R3
        M2 -= self.rocket.cp_eccentricity_x * R3
        # Calculate lift and moment for each component of the rocket
        velocity_in_body_frame = Vector([vx_b, vy_b, vz_b])
        w = Vector([omega1, omega2, omega3])
        for aero_surface, _ in self.rocket.aerodynamic_surfaces:
            # Component cp relative to CDM in body frame
            comp_cp = self.rocket.surfaces_cp_to_cdm[aero_surface]
            # Component absolute velocity in body frame
            comp_vb = velocity_in_body_frame + (w ^ comp_cp)
            # Wind velocity at component altitude
            comp_z = z + (K @ comp_cp).z
            comp_wind_vx = self.env.wind_velocity_x.get_value_opt(comp_z)
            comp_wind_vy = self.env.wind_velocity_y.get_value_opt(comp_z)
            # Component freestream velocity in body frame
            comp_wind_vb = Kt @ Vector([comp_wind_vx, comp_wind_vy, 0])
            comp_stream_velocity = comp_wind_vb - comp_vb
            comp_stream_speed = abs(comp_stream_velocity)
            comp_stream_mach = comp_stream_speed / speed_of_sound
            # Reynolds at component altitude
            # TODO: Reynolds is only used in generic surfaces. This calculation
            # should be moved to the surface class for efficiency
            comp_reynolds = (
                self.env.density.get_value_opt(comp_z)
                * comp_stream_speed
                * aero_surface.reference_length
                / self.env.dynamic_viscosity.get_value_opt(comp_z)
            )
            # Forces and moments
            X, Y, Z, M, N, L = aero_surface.compute_forces_and_moments(
                comp_stream_velocity,
                comp_stream_speed,
                comp_stream_mach,
                rho,
                comp_cp,
                w,
                comp_reynolds,
            )
            R1 += X
            R2 += Y
            R3 += Z
            M1 += M
            M2 += N
            M3 += L
        # Off center moment
        M3 += self.rocket.cp_eccentricity_x * R2 - self.rocket.cp_eccentricity_y * R1

        # Calculate derivatives
        # Angular acceleration
        alpha1 = (
            M1
            - (
                omega2
                * omega3
                * (
                    rocket_dry_I_33
                    + motor_I_33_at_t
                    - rocket_dry_I_11
                    - motor_I_11_at_t
                    - mu * b**2
                )
                + omega1
                * (
                    (
                        motor_I_11_derivative_at_t
                        + mass_flow_rate_at_t
                        * (rocket_dry_mass - 1)
                        * (b / total_mass_at_t) ** 2
                    )
                    - mass_flow_rate_at_t
                    * ((nozzle_radius / 2) ** 2 + (c - b * mu / rocket_dry_mass) ** 2)
                )
            )
        ) / (rocket_dry_I_11 + motor_I_11_at_t + mu * b**2)
        alpha2 = (
            M2
            - (
                omega1
                * omega3
                * (
                    rocket_dry_I_11
                    + motor_I_11_at_t
                    + mu * b**2
                    - rocket_dry_I_33
                    - motor_I_33_at_t
                )
                + omega2
                * (
                    (
                        motor_I_11_derivative_at_t
                        + mass_flow_rate_at_t
                        * (rocket_dry_mass - 1)
                        * (b / total_mass_at_t) ** 2
                    )
                    - mass_flow_rate_at_t
                    * ((nozzle_radius / 2) ** 2 + (c - b * mu / rocket_dry_mass) ** 2)
                )
            )
        ) / (rocket_dry_I_11 + motor_I_11_at_t + mu * b**2)
        alpha3 = (
            M3
            - omega3
            * (
                motor_I_33_derivative_at_t
                - mass_flow_rate_at_t * (nozzle_radius**2) / 2
            )
        ) / (rocket_dry_I_33 + motor_I_33_at_t)
        # Euler parameters derivative
        e0dot = 0.5 * (-omega1 * e1 - omega2 * e2 - omega3 * e3)
        e1dot = 0.5 * (omega1 * e0 + omega3 * e2 - omega2 * e3)
        e2dot = 0.5 * (omega2 * e0 - omega3 * e1 + omega1 * e3)
        e3dot = 0.5 * (omega3 * e0 + omega2 * e1 - omega1 * e2)
        # Linear acceleration
        L = [
            (
                R1
                - b * propellant_mass_at_t * (omega2**2 + omega3**2)
                - 2 * c * mass_flow_rate_at_t * omega2
            )
            / total_mass_at_t,
            (
                R2
                + b * propellant_mass_at_t * (alpha3 + omega1 * omega2)
                + 2 * c * mass_flow_rate_at_t * omega1
            )
            / total_mass_at_t,
            (R3 - b * propellant_mass_at_t * (alpha2 - omega1 * omega3) + net_thrust)
            / total_mass_at_t,
        ]
        ax, ay, az = K @ Vector(L)
        az -= self.env.gravity.get_value_opt(z)  # Include gravity

        # Coriolis acceleration
        _, w_earth_y, w_earth_z = self.env.earth_rotation_vector
        ax -= 2 * (vz * w_earth_y - vy * w_earth_z)
        ay -= 2 * (vx * w_earth_z)
        az -= 2 * (-vx * w_earth_y)

        # Create u_dot
        u_dot = [
            vx,
            vy,
            vz,
            ax,
            ay,
            az,
            e0dot,
            e1dot,
            e2dot,
            e3dot,
            alpha1,
            alpha2,
            alpha3,
        ]

        if post_processing:
            self.__post_processed_variables.append(
                [
                    t,
                    ax,
                    ay,
                    az,
                    alpha1,
                    alpha2,
                    alpha3,
                    R1,
                    R2,
                    R3,
                    M1,
                    M2,
                    M3,
                    net_thrust,
                ]
            )
        return u_dot

    def u_dot_generalized_3dof(self, t, u, post_processing=False):
        """Calculates derivative of u state vector with respect to time when the
        rocket is flying in 3 DOF motion in space and significant mass variation
        effects exist. Includes a weathercocking model that evolves the body axis
        direction toward the relative wind direction.

        Parameters
        ----------
        t : float
            Time in seconds.
        u : list
            State vector: [x, y, z, vx, vy, vz, e0, e1, e2, e3, omega1, omega2, omega3].
        post_processing : bool, optional
            If True, adds flight data to self variables like self.angle_of_attack.

        Returns
        -------
        list
            Derivative state vector: [vx, vy, vz, ax, ay, az,
            e0_dot, e1_dot, e2_dot, e3_dot, alpha1, alpha2, alpha3].
        """
        # Unpack state
        _, _, z, vx, vy, vz, e0, e1, e2, e3, omega1, omega2, omega3 = u

        # Define vectors
        v = Vector([vx, vy, vz])
        e = [e0, e1, e2, e3]
        w = Vector([omega1, omega2, omega3])

        # Mass and transformation
        total_mass = self.rocket.total_mass.get_value_opt(t)
        K = Matrix.transformation(e)
        Kt = K.transpose

        # Atmospheric and wind data
        rho = self.env.density.get_value_opt(z)
        wind_vx = self.env.wind_velocity_x.get_value_opt(z)
        wind_vy = self.env.wind_velocity_y.get_value_opt(z)
        wind_velocity = Vector([wind_vx, wind_vy, 0])

        free_stream_velocity = wind_velocity - v
        free_stream_speed = abs(free_stream_velocity)
        speed_of_sound = self.env.speed_of_sound.get_value_opt(z)
        mach = free_stream_speed / speed_of_sound

        # Drag computation
        if t < self.rocket.motor.burn_out_time:
            cd = self.rocket.power_on_drag.get_value_opt(mach)
        else:
            cd = self.rocket.power_off_drag.get_value_opt(mach)

        R1, R2 = 0, 0
        R3 = -0.5 * rho * free_stream_speed**2 * self.rocket.area * cd

        for air_brake in self.rocket.air_brakes:
            if air_brake.deployment_level > 0:
                ab_cd = air_brake.drag_coefficient.get_value_opt(
                    air_brake.deployment_level, mach
                )
                ab_force = (
                    -0.5 * rho * free_stream_speed**2 * air_brake.reference_area * ab_cd
                )
                if air_brake.override_rocket_drag:
                    R3 = ab_force
                else:
                    R3 += ab_force

        # Velocity in body frame
        vb_body = Kt @ v

        for surface, _ in self.rocket.aerodynamic_surfaces:
            cp = self.rocket.surfaces_cp_to_cdm[surface]
            vb_component = vb_body + (w ^ cp)

            comp_z = z + (K @ cp).z
            wind_cx = self.env.wind_velocity_x.get_value_opt(comp_z)
            wind_cy = self.env.wind_velocity_y.get_value_opt(comp_z)
            wind_body = Kt @ Vector([wind_cx, wind_cy, 0])

            rel_velocity = wind_body - vb_component
            rel_speed = abs(rel_velocity)
            rel_mach = rel_speed / speed_of_sound

            reynolds = (
                self.env.density.get_value_opt(comp_z)
                * rel_speed
                * surface.reference_length
                / self.env.dynamic_viscosity.get_value_opt(comp_z)
            )

            fx, fy, fz, *_ = surface.compute_forces_and_moments(
                rel_velocity, rel_speed, rel_mach, rho, cp, w, reynolds
            )
            R1 += fx
            R2 += fy
            R3 += fz

        # Thrust and weight
        # Calculate net thrust including pressure thrust correction if motor is burning
        if self.rocket.motor.burn_start_time < t < self.rocket.motor.burn_out_time:
            pressure = self.env.pressure.get_value_opt(z)
            net_thrust = max(
                self.rocket.motor.thrust.get_value_opt(t)
                + self.rocket.motor.pressure_thrust(pressure),
                0,
            )
        else:
            net_thrust = 0
        gravity = self.env.gravity.get_value_opt(z)
        weight_body = Kt @ Vector([0, 0, -total_mass * gravity])

        total_force = Vector([0, 0, net_thrust]) + weight_body + Vector([R1, R2, R3])

        # Dynamics
        v_dot = K @ (total_force / total_mass)
        r_dot = [vx, vy, vz]
        # Weathercocking: evolve body axis direction toward relative wind
        # The body z-axis (attitude vector) should align with -freestream_velocity
        if self.weathercock_coeff > 0 and free_stream_speed > 1e-6:
            # Current body z-axis in inertial frame (attitude vector)
            # From rotation matrix: column 3 gives the body z-axis in inertial frame
            body_z_inertial = Vector(
                [
                    2 * (e1 * e3 + e0 * e2),
                    2 * (e2 * e3 - e0 * e1),
                    1 - 2 * (e1**2 + e2**2),
                ]
            )

            # Desired direction: opposite of freestream velocity (into the wind)
            # This is the direction the rocket nose should point
            # Division by free_stream_speed ensures the result is a unit vector
            desired_direction = -free_stream_velocity / free_stream_speed

            # Compute rotation axis (cross product of current and desired)
            rotation_axis = body_z_inertial ^ desired_direction
            rotation_axis_mag = abs(rotation_axis)

            # Determine omega_body based on alignment state
            omega_body = None

            if rotation_axis_mag > 1e-8:
                # Normal case: compute angular velocity from misalignment
                rotation_axis = rotation_axis / rotation_axis_mag

                # The magnitude of the cross product of two unit vectors equals
                # the sine of the angle between them
                sin_angle = min(1.0, max(-1.0, rotation_axis_mag))

                # Angular velocity magnitude proportional to misalignment angle
                omega_mag = self.weathercock_coeff * sin_angle

                # Angular velocity in inertial frame, then transform to body frame
                omega_body = Kt @ (rotation_axis * omega_mag)
            else:
                # Check if aligned or anti-aligned using dot product
                dot = body_z_inertial @ desired_direction
                if dot < -0.999:  # Anti-aligned
                    # Choose an arbitrary perpendicular axis
                    x_axis = Vector([1.0, 0.0, 0.0])
                    perp_axis = body_z_inertial ^ x_axis
                    if abs(perp_axis) < 1e-6:
                        y_axis = Vector([0.0, 1.0, 0.0])
                        perp_axis = body_z_inertial ^ y_axis
                        if abs(perp_axis) < 1e-6:
                            raise ValueError(
                                "Cannot determine a valid rotation axis: "
                                "body_z_inertial is parallel to both x and y axes."
                            )
                    rotation_axis = perp_axis.unit_vector
                    # 180 degree rotation: sin(angle) = 1
                    omega_mag = self.weathercock_coeff * 1.0
                    omega_body = Kt @ (rotation_axis * omega_mag)
                # else: aligned (dot > 0.999) - no rotation needed, omega_body stays None

            # Compute quaternion derivatives from omega_body
            if omega_body is not None:
                omega1_wc, omega2_wc, omega3_wc = (
                    omega_body.x,
                    omega_body.y,
                    omega_body.z,
                )
                e0_dot = 0.5 * (-omega1_wc * e1 - omega2_wc * e2 - omega3_wc * e3)
                e1_dot = 0.5 * (omega1_wc * e0 + omega3_wc * e2 - omega2_wc * e3)
                e2_dot = 0.5 * (omega2_wc * e0 - omega3_wc * e1 + omega1_wc * e3)
                e3_dot = 0.5 * (omega3_wc * e0 + omega2_wc * e1 - omega1_wc * e2)
                e_dot = [e0_dot, e1_dot, e2_dot, e3_dot]
            else:
                e_dot = [0, 0, 0, 0]
            w_dot = [0, 0, 0]  # No angular acceleration in 3DOF model
        else:
            # No weathercocking or negligible freestream speed
            e_dot = [0, 0, 0, 0]
            w_dot = [0, 0, 0]

        u_dot = [*r_dot, *v_dot, *e_dot, *w_dot]

        if post_processing:
            self.__post_processed_variables.append(
                [t, *v_dot, *w_dot, R1, R2, R3, 0, 0, 0, net_thrust]
            )

        return u_dot

    def u_dot_generalized(self, t, u, post_processing=False):  # pylint: disable=too-many-locals,too-many-statements
        """Calculates derivative of u state vector with respect to time when the
        rocket is flying in 6 DOF motion in space and significant mass variation
        effects exist. Typical flight phases include powered ascent after launch
        rail.

        Parameters
        ----------
        t : float
            Time in seconds
        u : list
            State vector defined by u = [x, y, z, vx, vy, vz, q0, q1,
            q2, q3, omega1, omega2, omega3].
        post_processing : bool, optional
            If True, adds flight data information directly to self variables
            such as self.angle_of_attack, by default False.

        Returns
        -------
        u_dot : list
            State vector defined by u_dot = [vx, vy, vz, ax, ay, az,
            e0_dot, e1_dot, e2_dot, e3_dot, alpha1, alpha2, alpha3].
        """
        # Retrieve integration data
        _, _, z, vx, vy, vz, e0, e1, e2, e3, omega1, omega2, omega3 = u

        # Create necessary vectors
        # r = Vector([x, y, z])  # CDM position vector
        v = Vector([vx, vy, vz])  # CDM velocity vector
        e = [e0, e1, e2, e3]  # Euler parameters/quaternions
        w = Vector([omega1, omega2, omega3])  # Angular velocity vector

        # Retrieve necessary quantities
        ## Rocket mass
        total_mass = self.rocket.total_mass.get_value_opt(t)
        total_mass_dot = self.rocket.total_mass_flow_rate.get_value_opt(t)
        total_mass_ddot = self.rocket.total_mass_flow_rate.differentiate_complex_step(t)
        ## CM position vector and time derivatives relative to CDM in body frame
        r_CM_z = self.rocket.com_to_cdm_function
        r_CM_t = r_CM_z.get_value_opt(t)
        r_CM = Vector([0, 0, r_CM_t])
        r_CM_dot = Vector([0, 0, r_CM_z.differentiate_complex_step(t)])
        r_CM_ddot = Vector([0, 0, r_CM_z.differentiate(t, order=2)])
        ## Nozzle position vector
        r_NOZ = Vector([0, 0, self.rocket.nozzle_to_cdm])
        ## Nozzle gyration tensor
        S_nozzle = self.rocket.nozzle_gyration_tensor
        ## Inertia tensor
        inertia_tensor = self.rocket.get_inertia_tensor_at_time(t)
        ## Inertia tensor time derivative in the body frame
        I_dot = self.rocket.get_inertia_tensor_derivative_at_time(t)

        # Calculate the Inertia tensor relative to CM
        H = (r_CM.cross_matrix @ -r_CM.cross_matrix) * total_mass
        I_CM = inertia_tensor - H

        # Prepare transformation matrices
        K = Matrix.transformation(e)
        Kt = K.transpose

        # Compute aerodynamic forces and moments
        R1, R2, R3, M1, M2, M3 = 0, 0, 0, 0, 0, 0

        ## Drag force
        rho = self.env.density.get_value_opt(z)
        wind_velocity_x = self.env.wind_velocity_x.get_value_opt(z)
        wind_velocity_y = self.env.wind_velocity_y.get_value_opt(z)
        wind_velocity = Vector([wind_velocity_x, wind_velocity_y, 0])
        free_stream_speed = abs((wind_velocity - Vector(v)))
        speed_of_sound = self.env.speed_of_sound.get_value_opt(z)
        free_stream_mach = free_stream_speed / speed_of_sound

        # Get rocket velocity in body frame (needed for drag calculation)
        velocity_in_body_frame = Kt @ v
        # Calculate freestream velocity in body frame
        freestream_velocity = wind_velocity - v
        freestream_velocity_body = Kt @ freestream_velocity

        if self.rocket.motor.burn_start_time < t < self.rocket.motor.burn_out_time:
            pressure = self.env.pressure.get_value_opt(z)
            net_thrust = max(
                self.rocket.motor.thrust.get_value_opt(t)
                + self.rocket.motor.pressure_thrust(pressure),
                0,
            )
            drag_coeff = self.__get_drag_coefficient(
                self.rocket.power_on_drag,
                free_stream_mach,
                z,
                freestream_velocity_body,
            )
        else:
            net_thrust = 0
            drag_coeff = self.__get_drag_coefficient(
                self.rocket.power_off_drag,
                free_stream_mach,
                z,
                freestream_velocity_body,
            )
        R3 += -0.5 * rho * (free_stream_speed**2) * self.rocket.area * drag_coeff
        for air_brakes in self.rocket.air_brakes:
            if air_brakes.deployment_level > 0:
                air_brakes_cd = air_brakes.drag_coefficient.get_value_opt(
                    air_brakes.deployment_level, free_stream_mach
                )
                air_brakes_force = (
                    -0.5
                    * rho
                    * (free_stream_speed**2)
                    * air_brakes.reference_area
                    * air_brakes_cd
                )
                if air_brakes.override_rocket_drag:
                    R3 = air_brakes_force  # Substitutes rocket drag coefficient
                else:
                    R3 += air_brakes_force
        # Calculate lift and moment for each component of the rocket
        for aero_surface, _ in self.rocket.aerodynamic_surfaces:
            # Component cp relative to CDM in body frame
            comp_cp = self.rocket.surfaces_cp_to_cdm[aero_surface]
            # Component absolute velocity in body frame
            comp_vb = velocity_in_body_frame + (w ^ comp_cp)
            # Wind velocity at component altitude
            comp_z = z + (K @ comp_cp).z
            comp_wind_vx = self.env.wind_velocity_x.get_value_opt(comp_z)
            comp_wind_vy = self.env.wind_velocity_y.get_value_opt(comp_z)
            # Component freestream velocity in body frame
            comp_wind_vb = Kt @ Vector([comp_wind_vx, comp_wind_vy, 0])
            comp_stream_velocity = comp_wind_vb - comp_vb
            comp_stream_speed = abs(comp_stream_velocity)
            comp_stream_mach = comp_stream_speed / speed_of_sound
            # Reynolds at component altitude
            # TODO: Reynolds is only used in generic surfaces. This calculation
            # should be moved to the surface class for efficiency
            comp_reynolds = (
                self.env.density.get_value_opt(comp_z)
                * comp_stream_speed
                * aero_surface.reference_length
                / self.env.dynamic_viscosity.get_value_opt(comp_z)
            )
            # Forces and moments
            X, Y, Z, M, N, L = aero_surface.compute_forces_and_moments(
                comp_stream_velocity,
                comp_stream_speed,
                comp_stream_mach,
                rho,
                comp_cp,
                w,
                comp_reynolds,
            )
            R1 += X
            R2 += Y
            R3 += Z
            M1 += M
            M2 += N
            M3 += L

        # Off center moment
        M1 += (
            self.rocket.cp_eccentricity_y * R3
            + self.rocket.thrust_eccentricity_y * net_thrust
        )
        M2 -= (
            self.rocket.cp_eccentricity_x * R3
            + self.rocket.thrust_eccentricity_x * net_thrust
        )
        M3 += self.rocket.cp_eccentricity_x * R2 - self.rocket.cp_eccentricity_y * R1

        weight_in_body_frame = Kt @ Vector(
            [0, 0, -total_mass * self.env.gravity.get_value_opt(z)]
        )

        T00 = total_mass * r_CM
        T03 = 2 * total_mass_dot * (r_NOZ - r_CM) - 2 * total_mass * r_CM_dot
        T04 = (
            Vector([0, 0, net_thrust])
            - total_mass * r_CM_ddot
            - 2 * total_mass_dot * r_CM_dot
            + total_mass_ddot * (r_NOZ - r_CM)
        )
        T05 = total_mass_dot * S_nozzle - I_dot

        T20 = (
            ((w ^ T00) ^ w)
            + (w ^ T03)
            + T04
            + weight_in_body_frame
            + Vector([R1, R2, R3])
        )

        T21 = (
            ((inertia_tensor @ w) ^ w)
            + T05 @ w
            - (weight_in_body_frame ^ r_CM)
            + Vector([M1, M2, M3])
        )

        # Angular velocity derivative
        w_dot = I_CM.inverse @ (T21 + (T20 ^ r_CM))

        # Euler parameters derivative
        e_dot = [
            0.5 * (-omega1 * e1 - omega2 * e2 - omega3 * e3),
            0.5 * (omega1 * e0 + omega3 * e2 - omega2 * e3),
            0.5 * (omega2 * e0 - omega3 * e1 + omega1 * e3),
            0.5 * (omega3 * e0 + omega2 * e1 - omega1 * e2),
        ]

        # Velocity vector derivative + Coriolis acceleration
        w_earth = Vector(self.env.earth_rotation_vector)
        v_dot = K @ (T20 / total_mass - (r_CM ^ w_dot)) - 2 * (w_earth ^ v)

        # Position vector derivative
        r_dot = [vx, vy, vz]

        # Create u_dot
        u_dot = [*r_dot, *v_dot, *e_dot, *w_dot]

        if post_processing:
            self.__post_processed_variables.append(
                [t, *v_dot, *w_dot, R1, R2, R3, M1, M2, M3, net_thrust]
            )

        return u_dot

    def u_dot_parachute(self, t, u, post_processing=False):
        """Calculates derivative of u state vector with respect to time
        when rocket is flying under parachute. A 3 DOF approximation is
        used.

        Parameters
        ----------
        t : float
            Time in seconds
        u : list
            State vector defined by u = [x, y, z, vx, vy, vz, e0, e1,
            e2, e3, omega1, omega2, omega3].
        post_processing : bool, optional
            If True, adds flight data information directly to self
            variables such as self.angle_of_attack. Default is False.

        Return
        ------
        u_dot : list
            State vector defined by u_dot = [vx, vy, vz, ax, ay, az,
            e0dot, e1dot, e2dot, e3dot, alpha1, alpha2, alpha3].

        """
        # Get relevant state data
        z, vx, vy, vz = u[2:6]

        # Get atmospheric data
        rho = self.env.density.get_value_opt(z)
        wind_velocity_x = self.env.wind_velocity_x.get_value_opt(z)
        wind_velocity_y = self.env.wind_velocity_y.get_value_opt(z)

        # Get the mass of the rocket
        mp = self.rocket.dry_mass

        # to = 1.2
        # eta = 1
        # Rdot = (6 * R * (1 - eta) / (1.2**6)) * (
        #     (1 - eta) * t**5 + eta * (to**3) * (t**2)
        # )
        # Rdot = 0

        # tf = 8 * nominal diameter / velocity at line stretch

        # Calculate added mass
        ma = (
            self.parachute_added_mass_coefficient
            * rho
            * (2 / 3)
            * np.pi
            * self.parachute_radius**2
            * self.parachute_height
        )

        # Calculate freestream speed
        freestream_x = vx - wind_velocity_x
        freestream_y = vy - wind_velocity_y
        freestream_z = vz
        free_stream_speed = (freestream_x**2 + freestream_y**2 + freestream_z**2) ** 0.5

        # Determine drag force
        pseudo_drag = -0.5 * rho * self.parachute_cd_s * free_stream_speed
        # pseudo_drag = pseudo_drag - ka * rho * 4 * np.pi * (R**2) * Rdot
        Dx = pseudo_drag * freestream_x  # add eta efficiency for wake
        Dy = pseudo_drag * freestream_y
        Dz = pseudo_drag * freestream_z
        ax = Dx / (mp + ma)
        ay = Dy / (mp + ma)
        az = (Dz - mp * self.env.gravity.get_value_opt(z)) / (mp + ma)

        # Add coriolis acceleration
        _, w_earth_y, w_earth_z = self.env.earth_rotation_vector
        ax -= 2 * (vz * w_earth_y - vy * w_earth_z)
        ay -= 2 * (vx * w_earth_z)
        az -= 2 * (-vx * w_earth_y)

        if post_processing:
            self.__post_processed_variables.append(
                [t, ax, ay, az, 0, 0, 0, Dx, Dy, Dz, 0, 0, 0, 0]
            )

        return [vx, vy, vz, ax, ay, az, 0, 0, 0, 0, 0, 0, 0]

    @cached_property
    def solution_array(self):
        """Returns solution array of the rocket flight."""
        return np.array(self.solution)

    @property
    def function_evaluations_per_time_step(self):
        """Get the number of function evaluations per time step. This method
        calculates the difference between consecutive function evaluations
        during numerical integration and returns it as a list.

        Returns
        -------
        list
            The list of differences in function evaluations per time step.
        """
        return np.diff(self.function_evaluations).tolist()

    @cached_property
    def time(self):
        """Returns time array from solution."""
        return self.solution_array[:, 0]

    @cached_property
    def time_steps(self):
        """Returns time step array."""
        return np.diff(self.time)

    def get_solution_at_time(self, t, atol=1e-3):
        """Returns the solution state vector at a given time. If the time is
        not found in the solution, the closest time is used and a warning is
        raised.

        Parameters
        ----------
        t : float
            Time in seconds.
        atol : float, optional
            Absolute tolerance for time comparison. Default is 1e-3. If the
            difference between the time and the closest time in the solution
            is greater than this value, a warning is raised.

        Returns
        -------
        solution : np.array
            Solution state at time t. The array follows the format of the
            solution array, with the first column being time like this:
            [t, x, y, z, vx, vy, vz, e0, e1, e2, e3, omega1, omega2, omega3].

        """
        time_index = find_closest(self.time, t)
        if abs(self.time[time_index] - t) > atol:
            warnings.warn(
                f"Time {t} not found in solution. Closest time is "
                f"{self.time[time_index]}. Using closest time.",
                UserWarning,
            )
        return self.solution_array[time_index, :]

    # Process first type of outputs - state vector
    # Transform solution array into Functions
    @funcify_method("Time (s)", "X (m)", "spline", "constant")
    def x(self):
        """Rocket x position relative to the launch pad as a Function of
        time."""
        return self.solution_array[:, [0, 1]]

    @funcify_method("Time (s)", "Y (m)", "spline", "constant")
    def y(self):
        """Rocket y position relative to the launch pad as a Function of
        time."""
        return self.solution_array[:, [0, 2]]

    @funcify_method("Time (s)", "Z (m)", "spline", "constant")
    def z(self):
        """Rocket z position relative to the launch pad as a Function of
        time."""
        return self.solution_array[:, [0, 3]]

    @funcify_method("Time (s)", "Altitude AGL (m)", "spline", "constant")
    def altitude(self):
        """Rocket altitude above ground level as a Function of time. Ground
        level is defined by the environment elevation."""
        return self.z - self.env.elevation

    @funcify_method("Time (s)", "Vx (m/s)", "spline", "zero")
    def vx(self):
        """Velocity of the rocket's center of dry mass in the X (East) direction
        of the inertial frame as a function of time."""
        return self.solution_array[:, [0, 4]]

    @funcify_method("Time (s)", "Vy (m/s)", "spline", "zero")
    def vy(self):
        """Velocity of the rocket's center of dry mass in the Y (North)
        direction of the inertial frame as a function of time."""
        return self.solution_array[:, [0, 5]]

    @funcify_method("Time (s)", "Vz (m/s)", "spline", "zero")
    def vz(self):
        """Velocity of the rocket's center of dry mass in the Z (Up) direction of
        the inertial frame as a function of time."""
        return self.solution_array[:, [0, 6]]

    @funcify_method("Time (s)", "e0", "spline", "constant")
    def e0(self):
        """Rocket quaternion e0 as a Function of time."""
        return self.solution_array[:, [0, 7]]

    @funcify_method("Time (s)", "e1", "spline", "constant")
    def e1(self):
        """Rocket quaternion e1 as a Function of time."""
        return self.solution_array[:, [0, 8]]

    @funcify_method("Time (s)", "e2", "spline", "constant")
    def e2(self):
        """Rocket quaternion e2 as a Function of time."""
        return self.solution_array[:, [0, 9]]

    @funcify_method("Time (s)", "e3", "spline", "constant")
    def e3(self):
        """Rocket quaternion e3 as a Function of time."""
        return self.solution_array[:, [0, 10]]

    @funcify_method("Time (s)", "ω1 (rad/s)", "spline", "zero")
    def w1(self):
        """Angular velocity of the rocket in the x direction of the rocket's
        body frame as a function of time, in rad/s. Sometimes referred to as
        pitch rate (q)."""
        return self.solution_array[:, [0, 11]]

    @funcify_method("Time (s)", "ω2 (rad/s)", "spline", "zero")
    def w2(self):
        """Angular velocity of the rocket in the y direction of the rocket's
        body frame as a function of time, in rad/s. Sometimes referred to as
        yaw rate (r)."""
        return self.solution_array[:, [0, 12]]

    @funcify_method("Time (s)", "ω3 (rad/s)", "spline", "zero")
    def w3(self):
        """Angular velocity of the rocket in the z direction of the rocket's
        body frame as a function of time, in rad/s. Sometimes referred to as
        roll rate (p)."""
        return self.solution_array[:, [0, 13]]

    # Process second type of outputs - accelerations components
    @funcify_method("Time (s)", "Ax (m/s²)", "spline", "zero")
    def ax(self):
        """Acceleration of the rocket's center of dry mass along the X (East)
        axis in the inertial frame as a function of time."""
        return self.__evaluate_post_process[:, [0, 1]]

    @funcify_method("Time (s)", "Ay (m/s²)", "spline", "zero")
    def ay(self):
        """Acceleration of the rocket's center of dry mass along the Y (North)
        axis in the inertial frame as a function of time."""
        return self.__evaluate_post_process[:, [0, 2]]

    @funcify_method("Time (s)", "Az (m/s²)", "spline", "zero")
    def az(self):
        """Acceleration of the rocket's center of dry mass along the Z (Up)
        axis in the inertial frame as a function of time."""
        return self.__evaluate_post_process[:, [0, 3]]

    @funcify_method("Time (s)", "α1 (rad/s²)", "spline", "zero")
    def alpha1(self):
        """Angular acceleration of the rocket in the x direction of the rocket's
        body frame as a function of time, in rad/s. Sometimes referred to as
        pitch acceleration."""
        return self.__evaluate_post_process[:, [0, 4]]

    @funcify_method("Time (s)", "α2 (rad/s²)", "spline", "zero")
    def alpha2(self):
        """Angular acceleration of the rocket in the y direction of the rocket's
        body frame as a function of time, in rad/s. Sometimes referred to as
        yaw acceleration."""
        return self.__evaluate_post_process[:, [0, 5]]

    @funcify_method("Time (s)", "α3 (rad/s²)", "spline", "zero")
    def alpha3(self):
        """Angular acceleration of the rocket in the z direction of the rocket's
        body frame as a function of time, in rad/s. Sometimes referred to as
        roll acceleration."""
        return self.__evaluate_post_process[:, [0, 6]]

    # Process third type of outputs - Temporary values
    @funcify_method("Time (s)", "R1 (N)", "spline", "zero")
    def R1(self):
        """Aerodynamic force acting along the x-axis of the rocket's body frame
        as a function of time. Expressed in Newtons (N)."""
        return self.__evaluate_post_process[:, [0, 7]]

    @funcify_method("Time (s)", "R2 (N)", "spline", "zero")
    def R2(self):
        """Aerodynamic force acting along the y-axis of the rocket's body frame
        as a function of time. Expressed in Newtons (N)."""
        return self.__evaluate_post_process[:, [0, 8]]

    @funcify_method("Time (s)", "R3 (N)", "spline", "zero")
    def R3(self):
        """Aerodynamic force acting along the z-axis of the rocket's body frame
        as a function of time. Expressed in Newtons (N)."""
        return self.__evaluate_post_process[:, [0, 9]]

    @funcify_method("Time (s)", "M1 (Nm)", "linear", "zero")
    def M1(self):
        """Aerodynamic moment acting along the x-axis of the rocket's body
        frame as a function of time. Expressed in Newtons (N)."""
        return self.__evaluate_post_process[:, [0, 10]]

    @funcify_method("Time (s)", "M2 (Nm)", "linear", "zero")
    def M2(self):
        """Aerodynamic moment acting along the y-axis of the rocket's body
        frame as a function of time. Expressed in Newtons (N)."""
        return self.__evaluate_post_process[:, [0, 11]]

    @funcify_method("Time (s)", "M3 (Nm)", "linear", "zero")
    def M3(self):
        """Aerodynamic moment acting along the z-axis of the rocket's body
        frame as a function of time. Expressed in Newtons (N)."""
        return self.__evaluate_post_process[:, [0, 12]]

    @funcify_method("Time (s)", "Net Thrust (N)", "linear", "zero")
    def net_thrust(self):
        """Net thrust of the rocket as a Function of time. This is the
        actual thrust force experienced by the rocket. It may be corrected
        with the atmospheric pressure if a reference pressure is defined."""
        return self.__evaluate_post_process[:, [0, 13]]

    @funcify_method("Time (s)", "Pressure (Pa)", "spline", "constant")
    def pressure(self):
        """Air pressure felt by the rocket as a Function of time."""
        return [(t, self.env.pressure.get_value_opt(z)) for t, z in self.z]

    @funcify_method("Time (s)", "Density (kg/m³)", "spline", "constant")
    def density(self):
        """Air density felt by the rocket as a Function of time."""
        return [(t, self.env.density.get_value_opt(z)) for t, z in self.z]

    @funcify_method("Time (s)", "Dynamic Viscosity (Pa s)", "spline", "constant")
    def dynamic_viscosity(self):
        """Air dynamic viscosity felt by the rocket as a Function of
        time."""
        return [(t, self.env.dynamic_viscosity.get_value_opt(z)) for t, z in self.z]

    @funcify_method("Time (s)", "Speed of Sound (m/s)", "spline", "constant")
    def speed_of_sound(self):
        """Speed of sound in the air felt by the rocket as a Function of time."""
        return [(t, self.env.speed_of_sound.get_value_opt(z)) for t, z in self.z]

    @funcify_method("Time (s)", "Wind Velocity X (East) (m/s)", "spline", "constant")
    def wind_velocity_x(self):
        """Wind velocity in the X direction (east) as a Function of time."""
        return [(t, self.env.wind_velocity_x.get_value_opt(z)) for t, z in self.z]

    @funcify_method("Time (s)", "Wind Velocity Y (North) (m/s)", "spline", "constant")
    def wind_velocity_y(self):
        """Wind velocity in the Y direction (north) as a Function of time."""
        return [(t, self.env.wind_velocity_y.get_value_opt(z)) for t, z in self.z]

    # Process fourth type of output - values calculated from previous outputs

    # Frame conversions

    @cached_property
    def _stacked_velocity_body_frame(self):
        """Stacked velocity array at the center of dry mass in the body frame
        at each time step."""
        Kt = self.direction_cosine_matrixes
        v = np.array(
            [
                self.vx.y_array,
                self.vy.y_array,
                self.vz.y_array,
            ]
        ).transpose()
        stacked_velocity_body_frame = np.squeeze(np.matmul(Kt, v[:, :, np.newaxis]))
        return stacked_velocity_body_frame

    @funcify_method("Time (s)", "Vx Body Frame (m/s)", "spline", "zero")
    def vx_body_frame(self):
        """Velocity of the rocket's center of dry mass along the x axis of the
        body frame as a function of time."""
        return np.array(
            [self.time, self._stacked_velocity_body_frame[:, 0]]
        ).transpose()

    @funcify_method("Time (s)", "Vy Body Frame (m/s)", "spline", "zero")
    def vy_body_frame(self):
        """Velocity of the rocket's center of dry mass along the y axis of the
        body frame as a function of time."""
        return np.array(
            [self.time, self._stacked_velocity_body_frame[:, 1]]
        ).transpose()

    @funcify_method("Time (s)", "Vz Body Frame (m/s)", "spline", "zero")
    def vz_body_frame(self):
        """Velocity of the rocket's center of dry mass along the z axis of the
        body frame as a function of time."""
        return np.array(
            [self.time, self._stacked_velocity_body_frame[:, 2]]
        ).transpose()

    @cached_property
    def _stacked_acceleration_body_frame(self):
        """Stacked acceleration array at the center of dry mass in the body
        frame at each time step."""
        Kt = self.direction_cosine_matrixes
        a = np.array(
            [
                self.ax.y_array,
                self.ay.y_array,
                self.az.y_array,
            ]
        ).transpose()
        stacked_acceleration_body_frame = np.squeeze(np.matmul(Kt, a[:, :, np.newaxis]))
        return stacked_acceleration_body_frame

    @funcify_method("Time (s)", "Ax Body Frame (m/s²)", "spline", "zero")
    def ax_body_frame(self):
        """Acceleration of the rocket's center of dry mass along the x axis of the
        body frame as a function of time."""
        return np.array(
            [self.time, self._stacked_acceleration_body_frame[:, 0]]
        ).transpose()

    @funcify_method("Time (s)", "Ay Body Frame (m/s²)", "spline", "zero")
    def ay_body_frame(self):
        """Acceleration of the rocket's center of dry mass along the y axis of the
        body frame as a function of time."""
        return np.array(
            [self.time, self._stacked_acceleration_body_frame[:, 1]]
        ).transpose()

    @funcify_method("Time (s)", "Az Body Frame (m/s²)", "spline", "zero")
    def az_body_frame(self):
        """Acceleration of the rocket's center of dry mass along the z axis of the
        body frame as a function of time."""
        return np.array(
            [self.time, self._stacked_acceleration_body_frame[:, 2]]
        ).transpose()

    # Kinematics functions and values
    # Velocity Magnitude
    @funcify_method("Time (s)", "Speed - Velocity Magnitude (m/s)")
    def speed(self):
        """Rocket speed, or velocity magnitude, as a Function of time."""
        return (self.vx**2 + self.vy**2 + self.vz**2) ** 0.5

    @property
    def out_of_rail_velocity(self):
        """Velocity at which the rocket leaves the launch rail."""
        return self.speed.get_value_opt(self.out_of_rail_time)

    @cached_property
    def max_speed_time(self):
        """Time at which the rocket reaches its maximum speed."""
        max_speed_time_index = np.argmax(self.speed.get_source()[:, 1])
        return self.speed[max_speed_time_index, 0]

    @cached_property
    def max_speed(self):
        """Maximum speed reached by the rocket."""
        return self.speed.get_value_opt(self.max_speed_time)

    # Accelerations
    @funcify_method("Time (s)", "acceleration Magnitude (m/s²)")
    def acceleration(self):
        """Rocket acceleration magnitude as a Function of time."""
        return (self.ax**2 + self.ay**2 + self.az**2) ** 0.5

    @funcify_method("Time (s)", "Axial Acceleration (m/s²)", "spline", "zero")
    def axial_acceleration(self):
        """Axial acceleration magnitude as a Function of time."""
        return (
            self.ax * self.attitude_vector_x
            + self.ay * self.attitude_vector_y
            + self.az * self.attitude_vector_z
        )

    @cached_property
    def max_acceleration_power_on_time(self):
        """Time at which the rocket reaches its maximum acceleration during
        motor burn."""
        burn_out_time_index = find_closest(
            self.acceleration.source[:, 0], self.rocket.motor.burn_out_time
        )
        if burn_out_time_index == 0:
            return 0  # the burn out time is before the first time step

        max_acceleration_time_index = np.argmax(
            self.acceleration[:burn_out_time_index, 1]
        )
        return self.acceleration[max_acceleration_time_index, 0]

    @cached_property
    def max_acceleration_power_on(self):
        """Maximum acceleration reached by the rocket during motor burn."""
        return self.acceleration.get_value_opt(self.max_acceleration_power_on_time)

    @cached_property
    def max_acceleration_power_off_time(self):
        """Time at which the rocket reaches its maximum acceleration after
        motor burn."""
        burn_out_time_index = find_closest(
            self.acceleration.source[:, 0], self.rocket.motor.burn_out_time
        )
        max_acceleration_time_index = np.argmax(
            self.acceleration[burn_out_time_index:, 1]
        )
        return self.acceleration[max_acceleration_time_index, 0]

    @cached_property
    def max_acceleration_power_off(self):
        """Maximum acceleration reached by the rocket after motor burn."""
        return self.acceleration.get_value_opt(self.max_acceleration_power_off_time)

    @cached_property
    def max_acceleration_time(self):
        """Time at which the rocket reaches its maximum acceleration."""
        max_acceleration_time_index = np.argmax(self.acceleration[:, 1])
        return self.acceleration[max_acceleration_time_index, 0]

    @cached_property
    def max_acceleration(self):
        """Maximum acceleration reached by the rocket."""
        max_acceleration_time_index = np.argmax(self.acceleration[:, 1])
        return self.acceleration[max_acceleration_time_index, 1]

    @funcify_method("Time (s)", "Horizontal Speed (m/s)")
    def horizontal_speed(self):
        """Rocket horizontal speed as a Function of time."""
        return (self.vx**2 + self.vy**2) ** 0.5

    # Path Angle
    @funcify_method("Time (s)", "Path Angle (°)", "spline", "constant")
    def path_angle(self):
        """Rocket path angle as a Function of time."""
        path_angle = (180 / np.pi) * np.arctan2(
            self.vz[:, 1], self.horizontal_speed[:, 1]
        )
        return np.column_stack([self.time, path_angle])

    # Attitude Angle
    @funcify_method("Time (s)", "Attitude Vector X Component")
    def attitude_vector_x(self):
        """Rocket attitude vector X component as a Function of time.
        Same as row 1, column 3 of the rotation matrix that defines
        the conversion from the body frame to the inertial frame
        at each time step."""
        return 2 * (self.e1 * self.e3 + self.e0 * self.e2)  # a13

    @funcify_method("Time (s)", "Attitude Vector Y Component")
    def attitude_vector_y(self):
        """Rocket attitude vector Y component as a Function of time.
        Same as row 2, column 3 of the rotation matrix that defines
        the conversion from the body frame to the inertial frame
        at each time step."""
        return 2 * (self.e2 * self.e3 - self.e0 * self.e1)  # a23

    @funcify_method("Time (s)", "Attitude Vector Z Component")
    def attitude_vector_z(self):
        """Rocket attitude vector Z component as a Function of time.
        Same as row 3, column 3 of the rotation matrix that defines
        the conversion from the body frame to the inertial frame
        at each time step."""
        return 1 - 2 * (self.e1**2 + self.e2**2)  # a33

    @funcify_method("Time (s)", "Attitude Angle (°)")
    def attitude_angle(self):
        """Rocket attitude angle as a Function of time."""
        horizontal_attitude_proj = (
            self.attitude_vector_x**2 + self.attitude_vector_y**2
        ) ** 0.5
        attitude_angle = (180 / np.pi) * np.arctan2(
            self.attitude_vector_z[:, 1], horizontal_attitude_proj[:, 1]
        )
        return np.column_stack([self.time, attitude_angle])

    # Lateral Attitude Angle
    @funcify_method("Time (s)", "Lateral Attitude Angle (°)")
    def lateral_attitude_angle(self):
        """Rocket lateral attitude angle as a Function of time."""
        # TODO: complex method, it should be defined elsewhere.
        lateral_vector_angle = (np.pi / 180) * (self.heading - 90)
        lateral_vector_x = np.sin(lateral_vector_angle)
        lateral_vector_y = np.cos(lateral_vector_angle)
        attitude_lateral_proj = (
            lateral_vector_x * self.attitude_vector_x[:, 1]
            + lateral_vector_y * self.attitude_vector_y[:, 1]
        )
        attitude_lateral_proj_x = attitude_lateral_proj * lateral_vector_x
        attitude_lateral_proj_y = attitude_lateral_proj * lateral_vector_y
        attitude_lateral_plane_proj_x = (
            self.attitude_vector_x[:, 1] - attitude_lateral_proj_x
        )
        attitude_lateral_plane_proj_y = (
            self.attitude_vector_y[:, 1] - attitude_lateral_proj_y
        )
        attitude_lateral_plane_proj_z = self.attitude_vector_z[:, 1]
        attitude_lateral_plane_proj = (
            attitude_lateral_plane_proj_x**2
            + attitude_lateral_plane_proj_y**2
            + attitude_lateral_plane_proj_z**2
        ) ** 0.5
        lateral_attitude_angle = (180 / np.pi) * np.arctan2(
            attitude_lateral_proj, attitude_lateral_plane_proj
        )
        lateral_attitude_angle = np.column_stack([self.time, lateral_attitude_angle])
        return lateral_attitude_angle

    # Euler Angles
    @funcify_method("Time (s)", "Precession Angle - ψ (°)", "spline", "constant")
    def psi(self):
        """Precession angle as a Function of time."""
        psi = quaternions_to_precession(
            self.e0.y_array, self.e1.y_array, self.e2.y_array, self.e3.y_array
        )
        return np.column_stack([self.time, psi])

    @funcify_method("Time (s)", "Spin Angle - φ (°)", "spline", "constant")
    def phi(self):
        """Spin angle as a Function of time."""
        phi = quaternions_to_spin(
            self.e0.y_array, self.e1.y_array, self.e2.y_array, self.e3.y_array
        )
        return np.column_stack([self.time, phi])

    @funcify_method("Time (s)", "Nutation Angle - θ (°)", "spline", "constant")
    def theta(self):
        """Nutation angle as a Function of time."""
        theta = quaternions_to_nutation(self.e1.y_array, self.e2.y_array)
        return np.column_stack([self.time, theta])

    # Fluid Mechanics variables
    # Freestream Velocity
    @funcify_method("Time (s)", "Freestream Velocity X (m/s)", "spline", "constant")
    def stream_velocity_x(self):
        """Freestream velocity X component as a Function of time."""
        return np.column_stack((self.time, self.wind_velocity_x[:, 1] - self.vx[:, 1]))

    @funcify_method("Time (s)", "Freestream Velocity Y (m/s)", "spline", "constant")
    def stream_velocity_y(self):
        """Freestream velocity Y component as a Function of time."""
        return np.column_stack((self.time, self.wind_velocity_y[:, 1] - self.vy[:, 1]))

    @funcify_method("Time (s)", "Freestream Velocity Z (m/s)", "spline", "constant")
    def stream_velocity_z(self):
        """Freestream velocity Z component as a Function of time."""
        return np.column_stack((self.time, -self.vz[:, 1]))

    @funcify_method("Time (s)", "Freestream Speed (m/s)", "spline", "constant")
    def free_stream_speed(self):
        """Freestream speed as a Function of time."""
        free_stream_speed = (
            self.stream_velocity_x**2
            + self.stream_velocity_y**2
            + self.stream_velocity_z**2
        ) ** 0.5
        return free_stream_speed.get_source()

    # Apogee Freestream speed
    @cached_property
    def apogee_freestream_speed(self):
        """Freestream speed at apogee in m/s."""
        return self.free_stream_speed.get_value_opt(self.apogee_time)

    # Mach Number
    @funcify_method("Time (s)", "Mach Number", "spline", "zero")
    def mach_number(self):
        """Mach number as a Function of time."""
        return self.free_stream_speed / self.speed_of_sound

    @cached_property
    def max_mach_number_time(self):
        """Time of maximum Mach number."""
        max_mach_number_time_index = np.argmax(self.mach_number[:, 1])
        return self.mach_number[max_mach_number_time_index, 0]

    @cached_property
    def max_mach_number(self):
        """Maximum Mach number."""
        return self.mach_number.get_value_opt(self.max_mach_number_time)

    # Stability Margin
    @cached_property
    def max_stability_margin_time(self):
        """Time of maximum stability margin."""
        max_stability_margin_time_index = np.argmax(self.stability_margin[:, 1])
        return self.stability_margin[max_stability_margin_time_index, 0]

    @cached_property
    def max_stability_margin(self):
        """Maximum stability margin."""
        return self.stability_margin.get_value_opt(self.max_stability_margin_time)

    @cached_property
    def min_stability_margin_time(self):
        """Time of minimum stability margin."""
        min_stability_margin_time_index = np.argmin(self.stability_margin[:, 1])
        return self.stability_margin[min_stability_margin_time_index, 0]

    @cached_property
    def min_stability_margin(self):
        """Minimum stability margin."""
        return self.stability_margin.get_value_opt(self.min_stability_margin_time)

    @property
    def initial_stability_margin(self):
        """Stability margin at time 0.

        Returns
        -------
        float
        """
        return self.stability_margin.get_value_opt(self.time[0])

    @property
    def out_of_rail_stability_margin(self):
        """Stability margin at the time the rocket leaves the rail.

        Returns
        -------
        float
        """
        return self.stability_margin.get_value_opt(self.out_of_rail_time)

    # Reynolds Number
    @funcify_method("Time (s)", "Reynolds Number", "spline", "zero")
    def reynolds_number(self):
        """Reynolds number as a Function of time."""
        return (self.density * self.free_stream_speed / self.dynamic_viscosity) * (
            2 * self.rocket.radius
        )

    @cached_property
    def max_reynolds_number_time(self):
        """Time of maximum Reynolds number."""
        max_reynolds_number_time_index = np.argmax(self.reynolds_number[:, 1])
        return self.reynolds_number[max_reynolds_number_time_index, 0]

    @cached_property
    def max_reynolds_number(self):
        """Maximum Reynolds number."""
        return self.reynolds_number.get_value_opt(self.max_reynolds_number_time)

    # Dynamic Pressure
    @funcify_method("Time (s)", "Dynamic Pressure (Pa)", "spline", "zero")
    def dynamic_pressure(self):
        """Dynamic pressure as a Function of time."""
        return 0.5 * self.density * self.free_stream_speed**2

    @cached_property
    def max_dynamic_pressure_time(self):
        """Time of maximum dynamic pressure."""
        max_dynamic_pressure_time_index = np.argmax(self.dynamic_pressure[:, 1])
        return self.dynamic_pressure[max_dynamic_pressure_time_index, 0]

    @cached_property
    def max_dynamic_pressure(self):
        """Maximum dynamic pressure."""
        return self.dynamic_pressure.get_value_opt(self.max_dynamic_pressure_time)

    # Total Pressure
    @funcify_method("Time (s)", "Total Pressure (Pa)", "spline", "zero")
    def total_pressure(self):
        """Total pressure as a Function of time."""
        return self.pressure * (1 + 0.2 * self.mach_number**2) ** (3.5)

    @cached_property
    def max_total_pressure_time(self):
        """Time of maximum total pressure."""
        max_total_pressure_time_index = np.argmax(self.total_pressure[:, 1])
        return self.total_pressure[max_total_pressure_time_index, 0]

    @cached_property
    def max_total_pressure(self):
        """Maximum total pressure."""
        return self.total_pressure.get_value_opt(self.max_total_pressure_time)

    # Dynamics functions and variables

    #  Aerodynamic Lift and Drag
    # TODO: These are not lift and drag, they are the aerodynamic forces in
    # the rocket frame, meaning they are normal and axial forces. They should
    # be renamed.
    @funcify_method("Time (s)", "Aerodynamic Lift Force (N)", "spline", "zero")
    def aerodynamic_lift(self):
        """Aerodynamic lift force as a Function of time."""
        return (self.R1**2 + self.R2**2) ** 0.5

    @funcify_method("Time (s)", "Aerodynamic Drag Force (N)", "spline", "zero")
    def aerodynamic_drag(self):
        """Aerodynamic drag force as a Function of time."""
        return -1 * self.R3

    @funcify_method("Time (s)", "Aerodynamic Bending Moment (Nm)", "spline", "zero")
    def aerodynamic_bending_moment(self):
        """Aerodynamic bending moment as a Function of time."""
        return (self.M1**2 + self.M2**2) ** 0.5

    @funcify_method("Time (s)", "Aerodynamic Spin Moment (Nm)", "spline", "zero")
    def aerodynamic_spin_moment(self):
        """Aerodynamic spin moment as a Function of time."""
        return self.M3

    # Energy
    # Kinetic Energy
    @funcify_method("Time (s)", "Rotational Kinetic Energy (J)")
    def rotational_energy(self):
        """Rotational kinetic energy as a Function of time."""
        rotational_energy = 0.5 * (
            self.rocket.I_11 * self.w1**2
            + self.rocket.I_22 * self.w2**2
            + self.rocket.I_33 * self.w3**2
        )
        rotational_energy.set_discrete_based_on_model(self.w1)
        return rotational_energy

    @funcify_method("Time (s)", "Translational Kinetic Energy (J)", "spline", "zero")
    def translational_energy(self):
        """Translational kinetic energy as a Function of time."""
        # Redefine total_mass time grid to allow for efficient Function algebra
        total_mass = deepcopy(self.rocket.total_mass)
        total_mass.set_discrete_based_on_model(self.vz)
        translational_energy = 0.5 * total_mass * (self.speed**2)
        return translational_energy

    @funcify_method("Time (s)", "Kinetic Energy (J)", "spline", "zero")
    def kinetic_energy(self):
        """Total kinetic energy as a Function of time."""
        return self.rotational_energy + self.translational_energy

    # Potential Energy
    @funcify_method("Time (s)", "Potential Energy (J)", "spline", "constant")
    def potential_energy(self):
        """Potential energy as a Function of time in relation to sea
        level."""
        # Constants
        # TODO: this constant should come from Environment.
        standard_gravitational_parameter = 3.986004418e14
        # Redefine total_mass time grid to allow for efficient Function algebra
        total_mass = deepcopy(self.rocket.total_mass)
        total_mass.set_discrete_based_on_model(self.z)
        return (
            standard_gravitational_parameter
            * total_mass
            * (1 / (self.z + self.env.earth_radius) - 1 / self.env.earth_radius)
        )

    # Total Mechanical Energy
    @funcify_method("Time (s)", "Mechanical Energy (J)", "spline", "constant")
    def total_energy(self):
        """Total mechanical energy as a Function of time."""
        return self.kinetic_energy + self.potential_energy

    # thrust Power
    @funcify_method("Time (s)", "Thrust Power (W)", "spline", "zero")
    def thrust_power(self):
        """Thrust power as a Function of time."""
        return self.net_thrust * self.speed

    # Drag Power
    @funcify_method("Time (s)", "Drag Power (W)", "spline", "zero")
    def drag_power(self):
        """Drag power as a Function of time."""
        drag_power = self.R3 * self.speed
        drag_power.set_outputs("Drag Power (W)")
        return drag_power

    # Angle of Attack
    @cached_property
    def direction_cosine_matrixes(self):
        """Direction cosine matrix representing the attitude of the body frame,
        relative to the inertial frame, at each time step."""
        # Stack the y_arrays from e0, e1, e2, and e3 along a new axis
        stacked_arrays = np.stack(
            [self.e0.y_array, self.e1.y_array, self.e2.y_array, self.e3.y_array],
            axis=-1,
        )

        # Apply the transformation to the stacked arrays along the last axis
        Kt = np.array([Matrix.transformation(row).transpose for row in stacked_arrays])

        return Kt

    @cached_property
    def stream_velocity_body_frame(self):
        """Stream velocity array at the center of dry mass in the body frame at
        each time step."""
        Kt = self.direction_cosine_matrixes
        stream_velocity = np.array(
            [
                self.stream_velocity_x.y_array,
                self.stream_velocity_y.y_array,
                self.stream_velocity_z.y_array,
            ]
        ).transpose()
        stream_velocity_body = np.squeeze(
            np.matmul(Kt, stream_velocity[:, :, np.newaxis])
        )
        return stream_velocity_body

    @funcify_method("Time (s)", "Angle of Attack (°)", "spline", "constant")
    def angle_of_attack(self):
        """Angle of attack of the rocket with respect to the freestream
        velocity vector. Sometimes called total angle of attack. Defined as the
        angle between the freestream velocity vector and the rocket's z-axis.
        All in the Body Axes Coordinate System."""
        # Define stream velocity z component in body frame
        stream_vz_body = (
            -self.attitude_vector_x.y_array * self.stream_velocity_x.y_array
            - self.attitude_vector_y.y_array * self.stream_velocity_y.y_array
            - self.attitude_vector_z.y_array * self.stream_velocity_z.y_array
        )
        # Define freestream speed list
        free_stream_speed = self.free_stream_speed.y_array

        stream_vz_body_normalized = np.divide(
            stream_vz_body,
            free_stream_speed,
            out=np.zeros_like(stream_vz_body),
            where=free_stream_speed > 1e-6,
        )
        stream_vz_body_normalized = np.clip(stream_vz_body_normalized, -1, 1)

        # Calculate angle of attack and convert to degrees
        angle_of_attack = np.rad2deg(np.arccos(stream_vz_body_normalized))
        angle_of_attack = np.nan_to_num(angle_of_attack)

        return np.column_stack([self.time, angle_of_attack])

    @funcify_method("Time (s)", "Partial Angle of Attack (°)", "spline", "constant")
    def partial_angle_of_attack(self):
        """Partial angle of attack of the rocket with respect to the stream
        velocity vector. By partial angle of attack, it is meant the angle
        between the stream velocity vector in the y-z plane and the rocket's
        z-axis. All in the Body Axes Coordinate System."""
        # Stream velocity in standard aerodynamic frame
        stream_velocity = -self.stream_velocity_body_frame
        alpha = np.arctan2(
            stream_velocity[:, 1],
            stream_velocity[:, 2],
        )  # y-z plane
        return np.column_stack([self.time, np.rad2deg(alpha)])

    @funcify_method("Time (s)", "Beta (°)", "spline", "constant")
    def angle_of_sideslip(self):
        """Angle of sideslip of the rocket with respect to the stream
        velocity vector. Defined as the angle between the stream velocity
        vector in the x-z plane and the rocket's z-axis. All in the Body
        Axes Coordinate System."""
        # Stream velocity in standard aerodynamic frame
        stream_velocity = -self.stream_velocity_body_frame
        beta = np.arctan2(
            stream_velocity[:, 0],
            stream_velocity[:, 2],
        )  # x-z plane
        return np.column_stack([self.time, np.rad2deg(beta)])

    # Frequency response and stability variables
    @funcify_method("Frequency (Hz)", "ω1 Fourier Amplitude", "spline", "zero")
    def omega1_frequency_response(self):
        """Angular velocity 1 frequency response as a Function of
        frequency, as the rocket leaves the launch rail for 5 seconds of flight.
        """
        return self.w1.to_frequency_domain(
            self.out_of_rail_time, self.out_of_rail_time + 5, 100
        )

    @funcify_method("Frequency (Hz)", "ω2 Fourier Amplitude", "spline", "zero")
    def omega2_frequency_response(self):
        """Angular velocity 2 frequency response as a Function of
        frequency, as the rocket leaves the launch rail for 5 seconds of flight.
        """
        return self.w2.to_frequency_domain(
            self.out_of_rail_time, self.out_of_rail_time + 5, 100
        )

    @funcify_method("Frequency (Hz)", "ω3 Fourier Amplitude", "spline", "zero")
    def omega3_frequency_response(self):
        """Angular velocity 3 frequency response as a Function of
        frequency, as the rocket leaves the launch rail for 5 seconds of flight.
        """
        return self.w3.to_frequency_domain(
            self.out_of_rail_time, self.out_of_rail_time + 5, 100
        )

    @funcify_method(
        "Frequency (Hz)", "Attitude Angle Fourier Amplitude", "spline", "zero"
    )
    def attitude_frequency_response(self):
        """Attitude frequency response as a Function of frequency, as
        the rocket leaves the launch rail for 5 seconds of flight."""
        return self.attitude_angle.to_frequency_domain(
            lower=self.out_of_rail_time,
            upper=self.out_of_rail_time + 5,
            sampling_frequency=100,
        )

    @cached_property
    def static_margin(self):
        """Static margin of the rocket."""
        return self.rocket.static_margin

    @funcify_method("Time (s)", "Stability Margin (c)", "linear", "zero")
    def stability_margin(self):
        """Stability margin of the rocket along the flight, it considers the
        variation of the center of pressure position according to the mach
        number, as well as the variation of the center of gravity position
        according to the propellant mass evolution.

        Parameters
        ----------
        None

        Returns
        -------
        stability : rocketpy.Function
            Stability margin as a rocketpy.Function of time. The stability margin
            is defined as the distance between the center of pressure and the
            center of gravity, divided by the rocket diameter.
        """
        return [(t, self.rocket.stability_margin(m, t)) for t, m in self.mach_number]

    # Rail Button Forces

    @funcify_method("Time (s)", "Upper Rail Button Normal Force (N)", "spline", "zero")
    def rail_button1_normal_force(self):
        """Upper rail button normal force as a Function of time. If
        there's no rail button defined, the function returns a null Function."""
        return self.__calculate_rail_button_forces[0]

    @funcify_method("Time (s)", "Upper Rail Button Shear Force (N)", "spline", "zero")
    def rail_button1_shear_force(self):
        """Upper rail button shear force as a Function of time. If
        there's no rail button defined, the function returns a null Function."""
        return self.__calculate_rail_button_forces[1]

    @funcify_method("Time (s)", "Lower Rail Button Normal Force (N)", "spline", "zero")
    def rail_button2_normal_force(self):
        """Lower rail button normal force as a Function of time. If
        there's no rail button defined, the function returns a null Function."""
        return self.__calculate_rail_button_forces[2]

    @funcify_method("Time (s)", "Lower Rail Button Shear Force (N)", "spline", "zero")
    def rail_button2_shear_force(self):
        """Lower rail button shear force as a Function of time. If
        there's no rail button defined, the function returns a null Function."""
        return self.__calculate_rail_button_forces[3]

    @property
    def max_rail_button1_normal_force(self):
        """Maximum upper rail button normal force, in Newtons."""
        return np.abs(self.rail_button1_normal_force.y_array).max()

    @property
    def max_rail_button1_shear_force(self):
        """Maximum upper rail button shear force, in Newtons."""
        return np.abs(self.rail_button1_shear_force.y_array).max()

    @property
    def max_rail_button2_normal_force(self):
        """Maximum lower rail button normal force, in Newtons."""
        return np.abs(self.rail_button2_normal_force.y_array).max()

    @property
    def max_rail_button2_shear_force(self):
        """Maximum lower rail button shear force, in Newtons."""
        return np.abs(self.rail_button2_shear_force.y_array).max()

    @cached_property
    def calculate_rail_button_bending_moments(self):
        """Calculate internal bending moments at rail button attachment points.

        Uses beam theory to determine the internal structural moments for
        stress analysis of the rail button attachments (fasteners and airframe).

        The bending moment at each button attachment consists of:

        1. Normal force moment: $M = N \\times d$, where $N$ is the normal
           reaction force and $d$ is the distance from button to center of
           dry mass.
        2. Shear force cantilever moment: $M = S \\times h$, where $S$ is the
           shear (tangential) force and $h$ is the button standoff height.

        Returns
        -------
        tuple
            rail_button1_bending_moment : Function
                Bending moment at upper rail button as a function of time (N·m).
            max_rail_button1_bending_moment : float
                Maximum upper rail button bending moment (N·m).
            rail_button2_bending_moment : Function
                Bending moment at lower rail button as a function of time (N·m).
            max_rail_button2_bending_moment : float
                Maximum lower rail button bending moment (N·m).

        Notes
        -----
        - Calculated only during the rail phase of flight
        - Maximum values use absolute values for worst-case stress analysis
        - The bending moments represent internal stresses in the rocket
          airframe at the rail button attachment points

        **Assumptions:**

        - Rail buttons act as simple supports: provide reaction forces (normal
          and shear) but no moment reaction at the rail contact point
        - The rocket acts as a beam supported at two points (rail buttons)
        - Bending moments arise from the lever arm effect of reaction forces
          and the cantilever moment from button standoff height
        """
        # Check if rail buttons exist
        null_moment = Function(0)
        if len(self.rocket.rail_buttons) == 0:
            warnings.warn(
                "Trying to calculate rail button bending moments without "
                "rail buttons defined. Setting moments to zero.",
                UserWarning,
            )
            return (null_moment, 0.0, null_moment, 0.0)

        # Get rail button geometry
        rail_buttons_tuple = self.rocket.rail_buttons[0]
        # Rail button standoff height
        h_button = rail_buttons_tuple.component.button_height
        if h_button is None:
            warnings.warn(
                "Rail button height not defined. Bending moments cannot be "
                "calculated. Setting moments to zero.",
                UserWarning,
            )
            return (null_moment, 0.0, null_moment, 0.0)
        upper_button_position = (
            rail_buttons_tuple.component.buttons_distance
            + rail_buttons_tuple.position.z
        )
        lower_button_position = rail_buttons_tuple.position.z

        # Get center of dry mass (handle both callable and property)
        if callable(self.rocket.center_of_dry_mass_position):
            cdm = self.rocket.center_of_dry_mass_position(self.rocket._csys)
        else:
            cdm = self.rocket.center_of_dry_mass_position

        # Distances from buttons to center of dry mass
        d1 = abs(upper_button_position - cdm)
        d2 = abs(lower_button_position - cdm)

        # forces
        N1 = self.rail_button1_normal_force
        N2 = self.rail_button2_normal_force
        S1 = self.rail_button1_shear_force
        S2 = self.rail_button2_shear_force
        t = N1.source[:, 0]

        # Calculate bending moments at attachment points
        # Primary contribution from shear force acting at button height
        # Secondary contribution from normal force creating moment about attachment
        m1_values = N2.source[:, 1] * d2 + S1.source[:, 1] * h_button
        m2_values = N1.source[:, 1] * d1 + S2.source[:, 1] * h_button

        rail_button1_bending_moment = Function(
            np.column_stack([t, m1_values]),
            inputs="Time (s)",
            outputs="Bending Moment (N·m)",
            interpolation="linear",
        )
        rail_button2_bending_moment = Function(
            np.column_stack([t, m2_values]),
            inputs="Time (s)",
            outputs="Bending Moment (N·m)",
            interpolation="linear",
        )

        # Maximum bending moments (absolute value for stress calculations)
        max_rail_button1_bending_moment = float(np.max(np.abs(m1_values)))
        max_rail_button2_bending_moment = float(np.max(np.abs(m2_values)))

        return (
            rail_button1_bending_moment,
            max_rail_button1_bending_moment,
            rail_button2_bending_moment,
            max_rail_button2_bending_moment,
        )

    @property
    def rail_button1_bending_moment(self):
        """Upper rail button bending moment as a Function of time."""
        return self.calculate_rail_button_bending_moments[0]

    @property
    def max_rail_button1_bending_moment(self):
        """Maximum upper rail button bending moment, in N·m."""
        return self.calculate_rail_button_bending_moments[1]

    @property
    def rail_button2_bending_moment(self):
        """Lower rail button bending moment as a Function of time."""
        return self.calculate_rail_button_bending_moments[2]

    @property
    def max_rail_button2_bending_moment(self):
        """Maximum lower rail button bending moment, in N·m."""
        return self.calculate_rail_button_bending_moments[3]

    @funcify_method(
        "Time (s)", "Horizontal Distance to Launch Point (m)", "spline", "constant"
    )
    def drift(self):
        """Rocket horizontal distance to the launch point, in meters, as a
        Function of time."""
        return np.column_stack(
            (self.time, (self.x[:, 1] ** 2 + self.y[:, 1] ** 2) ** 0.5)
        )

    @funcify_method("Time (s)", "Bearing (°)", "spline", "constant")
    def bearing(self):
        """Rocket bearing compass, in degrees, as a Function of time."""
        x, y = self.x[:, 1], self.y[:, 1]
        bearing = (2 * np.pi - np.arctan2(-x, y)) * (180 / np.pi)
        return np.column_stack((self.time, bearing))

    @funcify_method("Time (s)", "Latitude (°)", "linear", "constant")
    def latitude(self):
        """Rocket latitude coordinate, in degrees, as a Function of
        time.
        """
        lat1 = np.deg2rad(self.env.latitude)  # Launch lat point converted to radians

        # Applies the haversine equation to find final lat/lon coordinates
        latitude = np.rad2deg(
            np.arcsin(
                np.sin(lat1) * np.cos(self.drift[:, 1] / self.env.earth_radius)
                + np.cos(lat1)
                * np.sin(self.drift[:, 1] / self.env.earth_radius)
                * np.cos(np.deg2rad(self.bearing[:, 1]))
            )
        )
        return np.column_stack((self.time, latitude))

    # TODO: haversine should be defined in tools.py so we just invoke it in here.
    @funcify_method("Time (s)", "Longitude (°)", "linear", "constant")
    def longitude(self):
        """Rocket longitude coordinate, in degrees, as a Function of
        time.
        """
        lat1 = np.deg2rad(self.env.latitude)  # Launch lat point converted to radians
        lon1 = np.deg2rad(self.env.longitude)  # Launch lon point converted to radians

        # Applies the haversine equation to find final lat/lon coordinates
        longitude = np.rad2deg(
            lon1
            + np.arctan2(
                np.sin(np.deg2rad(self.bearing[:, 1]))
                * np.sin(self.drift[:, 1] / self.env.earth_radius)
                * np.cos(lat1),
                np.cos(self.drift[:, 1] / self.env.earth_radius)
                - np.sin(lat1) * np.sin(np.deg2rad(self.latitude[:, 1])),
            )
        )

        return np.column_stack((self.time, longitude))

    def get_controller_observed_variables(self):
        """Retrieve the observed variables related to air brakes from the
        controllers. If there is only one set of observed variables, it is
        returned as a list. If there are multiple sets, the list containing
        all sets is returned."""
        observed_variables = [
            controller.observed_variables for controller in self._controllers
        ]
        return (
            observed_variables[0]
            if len(observed_variables) == 1
            else observed_variables
        )

    @cached_property
    def __calculate_rail_button_forces(self):  # TODO: complex method.
        """Calculate the forces applied to the rail buttons while rocket is
        still on the launch rail. It will return 0 if no rail buttons are
        defined.

        Returns
        -------
        F11: Function
            Rail Button 1 force in the 1 direction
        F12: Function
            Rail Button 1 force in the 2 direction
        F21: Function
            Rail Button 2 force in the 1 direction
        F22: Function
            Rail Button 2 force in the 2 direction
        """
        # First check for no rail phase or rail buttons
        null_force = Function(0)
        if self.out_of_rail_time_index == 0:  # No rail phase, no rail button forces
            warnings.warn(
                "Trying to calculate rail button forces without a rail phase defined. "
                + "The rail button forces will be set to zero.",
                UserWarning,
            )
            return null_force, null_force, null_force, null_force
        if len(self.rocket.rail_buttons) == 0:
            warnings.warn(
                "Trying to calculate rail button forces without rail buttons defined. "
                + "The rail button forces will be set to zero.",
                UserWarning,
            )
            return null_force, null_force, null_force, null_force

        # Distance from Rail Button 1 (upper) to CM
        rail_buttons_tuple = self.rocket.rail_buttons[0]
        upper_button_position = (
            rail_buttons_tuple.component.buttons_distance
            + rail_buttons_tuple.position.z
        )
        lower_button_position = rail_buttons_tuple.position.z
        angular_position_rad = rail_buttons_tuple.component.angular_position_rad
        D1 = (
            upper_button_position - self.rocket.center_of_dry_mass_position
        ) * self.rocket._csys
        # Distance from Rail Button 2 (lower) to CM
        D2 = (
            lower_button_position - self.rocket.center_of_dry_mass_position
        ) * self.rocket._csys
        F11 = (self.R1 * D2 - self.M2) / (D1 + D2)
        F11.set_outputs("Upper button force direction 1 (m)")
        F12 = (self.R2 * D2 + self.M1) / (D1 + D2)
        F12.set_outputs("Upper button force direction 2 (m)")
        F21 = (self.R1 * D1 + self.M2) / (D1 + D2)
        F21.set_outputs("Lower button force direction 1 (m)")
        F22 = (self.R2 * D1 - self.M1) / (D1 + D2)
        F22.set_outputs("Lower button force direction 2 (m)")

        model = Function(
            F11.get_source()[: self.out_of_rail_time_index + 1, :],
            interpolation=F11.__interpolation__,
        )

        # Limit force calculation to when rocket is in rail
        F11.set_discrete_based_on_model(model)
        F12.set_discrete_based_on_model(model)
        F21.set_discrete_based_on_model(model)
        F22.set_discrete_based_on_model(model)

        rail_button1_normal_force = F11 * np.cos(angular_position_rad) + F12 * np.sin(
            angular_position_rad
        )
        rail_button1_shear_force = F11 * -np.sin(angular_position_rad) + F12 * np.cos(
            angular_position_rad
        )
        rail_button2_normal_force = F21 * np.cos(angular_position_rad) + F22 * np.sin(
            angular_position_rad
        )
        rail_button2_shear_force = F21 * -np.sin(angular_position_rad) + F22 * np.cos(
            angular_position_rad
        )

        return (
            rail_button1_normal_force,
            rail_button1_shear_force,
            rail_button2_normal_force,
            rail_button2_shear_force,
        )

    def __transform_pressure_signals_lists_to_functions(self):
        """Calculate the pressure signal from the pressure sensor.
        It creates a signal_function attribute in the parachute object.
        Parachute works as a subclass of Rocket class.

        Returns
        -------
        None
        """
        # Transform parachute sensor feed into functions
        for parachute in self.rocket.parachutes:
            # TODO: these Functions do not need input validation
            parachute.clean_pressure_signal_function = Function(
                parachute.clean_pressure_signal,
                "Time (s)",
                "Pressure - Without Noise (Pa)",
                "linear",
            )
            parachute.noise_signal_function = Function(
                parachute.noise_signal, "Time (s)", "Pressure Noise (Pa)", "linear"
            )
            parachute.noisy_pressure_signal_function = (
                parachute.clean_pressure_signal_function
                + parachute.noise_signal_function
            )

    @cached_property
    def __evaluate_post_process(self):
        """Evaluate all post-processing variables by running the simulation
        again but with the post-processing flag set to True.

        Returns
        -------
        np.array
            An array containing all post-processed variables evaluated at each
            time step. Each element of the array is a list containing:
            [t, ax, ay, az, alpha1, alpha2, alpha3, R1, R2, R3, M1, M2, M3, net_thrust]
        """
        self.__post_processed_variables = []
        for phase_index, phase in self.time_iterator(self.flight_phases):
            init_time = phase.t
            final_time = self.flight_phases[phase_index + 1].t
            current_derivative = phase.derivative
            for callback in phase.callbacks:
                callback(self)
            for step in self.solution:
                if init_time < step[0] <= final_time or (
                    init_time == self.t_initial and step[0] == self.t_initial
                ):
                    current_derivative(step[0], step[1:], post_processing=True)

        return np.array(self.__post_processed_variables)

    def calculate_stall_wind_velocity(self, stall_angle):  # TODO: move to utilities
        """Function to calculate the maximum wind velocity before the angle of
        attack exceeds a desired angle, at the instant of departing rail launch.
        Can be helpful if you know the exact stall angle of all aerodynamics
        surfaces.

        Parameters
        ----------
        stall_angle : float
            Angle, in degrees, for which you would like to know the maximum wind
            speed before the angle of attack exceeds it

        Return
        ------
        None
        """
        v_f = self.out_of_rail_velocity

        theta = np.radians(self.inclination)
        stall_angle = np.radians(stall_angle)

        c = (math.cos(stall_angle) ** 2 - math.cos(theta) ** 2) / math.sin(
            stall_angle
        ) ** 2
        w_v = (
            2 * v_f * math.cos(theta) / c
            + (
                4 * v_f * v_f * math.cos(theta) * math.cos(theta) / (c**2)
                + 4 * 1 * v_f * v_f / c
            )
            ** 0.5
        ) / 2

        stall_angle = np.degrees(stall_angle)
        print(
            "Maximum wind velocity at Rail Departure time before angle"
            + f" of attack exceeds {stall_angle:.3f}°: {w_v:.3f} m/s"
        )

    @deprecated(
        reason="Moved to FlightDataExporter.export_pressures()",
        version="v1.12.0",
        alternative="rocketpy.simulation.flight_data_exporter.FlightDataExporter.export_pressures",
    )
    def export_pressures(self, file_name, time_step):
        """
        .. deprecated:: 1.11
           Use :class:`rocketpy.simulation.flight_data_exporter.FlightDataExporter`
           and call ``.export_pressures(...)``.
        """
        return FlightDataExporter(self).export_pressures(file_name, time_step)

    @deprecated(
        reason="Moved to FlightDataExporter.export_data()",
        version="v1.12.0",
        alternative="rocketpy.simulation.flight_data_exporter.FlightDataExporter.export_data",
    )
    def export_data(self, file_name, *variables, time_step=None):
        """
        .. deprecated:: 1.11
           Use :class:`rocketpy.simulation.flight_data_exporter.FlightDataExporter`
           and call ``.export_data(...)``.
        """
        return FlightDataExporter(self).export_data(
            file_name, *variables, time_step=time_step
        )

    @deprecated(
        reason="Moved to FlightDataExporter.export_sensor_data()",
        version="v1.12.0",
        alternative="rocketpy.simulation.flight_data_exporter.FlightDataExporter.export_sensor_data",
    )
    def export_sensor_data(self, file_name, sensor=None):
        """
        .. deprecated:: 1.11
           Use :class:`rocketpy.simulation.flight_data_exporter.FlightDataExporter`
           and call ``.export_sensor_data(...)``.
        """
        return FlightDataExporter(self).export_sensor_data(file_name, sensor=sensor)

    @deprecated(
        reason="Moved to FlightDataExporter.export_kml()",
        version="v1.12.0",
        alternative="rocketpy.simulation.flight_data_exporter.FlightDataExporter.export_kml",
    )
    def export_kml(
        self,
        file_name="trajectory.kml",
        time_step=None,
        extrude=True,
        color="641400F0",
        altitude_mode="absolute",
    ):
        """
        .. deprecated:: 1.11
           Use :class:`rocketpy.simulation.flight_data_exporter.FlightDataExporter`
           and call ``.export_kml(...)``.
        """
        return FlightDataExporter(self).export_kml(
            file_name=file_name,
            time_step=time_step,
            extrude=extrude,
            color=color,
            altitude_mode=altitude_mode,
        )

    def info(self):
        """Prints out a summary of the data available about the Flight."""
        self.prints.all()

    def all_info(self):
        """Prints out all data and graphs available about the Flight."""
        self.info()
        self.plots.all()

    def time_iterator(self, node_list):
        i = 0
        while i < len(node_list) - 1:
            yield i, node_list[i]
            i += 1

    def to_dict(self, **kwargs):
        data = {
            "rocket": self.rocket,
            "env": self.env,
            "rail_length": self.rail_length,
            "inclination": self.inclination,
            "heading": self.heading,
            "initial_solution": self.initial_solution,
            "terminate_on_apogee": self.terminate_on_apogee,
            "max_time": self.max_time,
            "max_time_step": self.max_time_step,
            "min_time_step": self.min_time_step,
            "rtol": self.rtol,
            "atol": self.atol,
            "time_overshoot": self.time_overshoot,
            "name": self.name,
            "equations_of_motion": self.equations_of_motion,
            # The following outputs are essential to run all_info method
            "solution": self.solution,
            "out_of_rail_time": self.out_of_rail_time,
            "out_of_rail_time_index": self.out_of_rail_time_index,
            "apogee_time": self.apogee_time,
            "apogee": self.apogee,
            "parachute_events": self.parachute_events,
            "impact_state": self.impact_state,
            "impact_velocity": self.impact_velocity,
            "x_impact": self.x_impact,
            "y_impact": self.y_impact,
            "t_final": self.t_final,
            "function_evaluations": self.function_evaluations,
            "ax": self.ax,
            "ay": self.ay,
            "az": self.az,
            "alpha1": self.alpha1,
            "alpha2": self.alpha2,
            "alpha3": self.alpha3,
            "R1": self.R1,
            "R2": self.R2,
            "R3": self.R3,
            "M1": self.M1,
            "M2": self.M2,
            "M3": self.M3,
            "net_thrust": self.net_thrust,
        }

        if kwargs.get("include_outputs", False):
            data.update(
                {
                    "time": self.time,
                    "out_of_rail_velocity": self.out_of_rail_velocity,
                    "out_of_rail_state": self.out_of_rail_state,
                    "apogee_x": self.apogee_x,
                    "apogee_y": self.apogee_y,
                    "apogee_state": self.apogee_state,
                    "x": self.x,
                    "y": self.y,
                    "z": self.z,
                    "vx": self.vx,
                    "vy": self.vy,
                    "vz": self.vz,
                    "e0": self.e0,
                    "e1": self.e1,
                    "e2": self.e2,
                    "e3": self.e3,
                    "w1": self.w1,
                    "w2": self.w2,
                    "w3": self.w3,
                    "altitude": self.altitude,
                    "mach_number": self.mach_number,
                    "stream_velocity_x": self.stream_velocity_x,
                    "stream_velocity_y": self.stream_velocity_y,
                    "stream_velocity_z": self.stream_velocity_z,
                    "free_stream_speed": self.free_stream_speed,
                    "angle_of_attack": self.angle_of_attack,
                    "static_margin": self.static_margin,
                    "stability_margin": self.stability_margin,
                    "latitude": self.latitude,
                    "longitude": self.longitude,
                }
            )

        return data

    @classmethod
    def from_dict(cls, data):
        return cls(
            rocket=data["rocket"],
            environment=data["env"],
            rail_length=data["rail_length"],
            inclination=data["inclination"],
            heading=data["heading"],
            initial_solution=None,
            terminate_on_apogee=data["terminate_on_apogee"],
            max_time=data["max_time"],
            max_time_step=data["max_time_step"],
            min_time_step=data["min_time_step"],
            rtol=data["rtol"],
            atol=data["atol"],
            time_overshoot=data["time_overshoot"],
            name=data["name"],
            equations_of_motion=data["equations_of_motion"],
        )

    class FlightPhases:
        """Class to handle flight phases. It is used to store the derivatives
        and callbacks for each flight phase. It is also used to handle the
        insertion of flight phases in the correct order, according to their
        initial time.

        Attributes
        ----------
        list : list
            A list of FlightPhase objects. See FlightPhase subclass.
        """

        def __init__(self, init_list=None):
            init_list = [] if init_list is None else init_list
            self.list = init_list[:]

        def __getitem__(self, index):
            return self.list[index]

        def __len__(self):
            return len(self.list)

        def __repr__(self):
            return str(self.list)

        def display_warning(self, *messages):  # pragma: no cover
            """A simple function to print a warning message."""
            print("WARNING:", *messages)

        def add(self, flight_phase, index=None):  # TODO: quite complex method
            """Add a flight phase to the list. It will be inserted in the
            correct position, according to its initial time. If no index is
            provided, it will be appended to the end of the list. If by any
            reason the flight phase cannot be inserted in the correct position,
            a warning will be displayed and the flight phase will be inserted
            in the closest position possible.

            Parameters
            ----------
            flight_phase : FlightPhase
                The flight phase object to be added. See FlightPhase class.
            index : int, optional
                The index of the flight phase to be added. If no index is
                provided, the flight phase will be appended to the end of the
                list. Default is None.

            Returns
            -------
            None
            """
            # Handle first phase
            if len(self.list) == 0:
                self.list.append(flight_phase)
                return None

            # Handle appending to last position
            if index is None:
                previous_phase = self.list[-1]
                if flight_phase.t > previous_phase.t:
                    self.list.append(flight_phase)
                    return None
                warning_msg = (
                    (
                        "Trying to add flight phase starting together with the "
                        "one preceding it. This may be caused by multiple "
                        "parachutes being triggered simultaneously."
                    )
                    if flight_phase.t == previous_phase.t
                    else (
                        "Trying to add flight phase starting *before* the one "
                        "*preceding* it. This may be caused by multiple "
                        "parachutes being triggered simultaneously "
                        "or by having a negative parachute lag.",
                    )
                )
                self.display_warning(*warning_msg)
                flight_phase.t += 1e-7 if flight_phase.t == previous_phase.t else 0
                self.add(
                    flight_phase, -2 if flight_phase.t < previous_phase.t else None
                )
                return None

            # Handle inserting into intermediary position.
            # Check if new flight phase respects time
            next_phase = self.list[index]
            previous_phase = self.list[index - 1]
            if previous_phase.t < flight_phase.t < next_phase.t:
                self.list.insert(index, flight_phase)
                return None
            warning_msg = (
                (
                    "Trying to add flight phase starting *together* with the one *preceding* it. ",
                    "This may be caused by multiple parachutes being triggered simultaneously.",
                )
                if flight_phase.t == previous_phase.t
                else (
                    (
                        "Trying to add flight phase starting *together* with the one *proceeding* it. ",
                        "This may be caused by multiple parachutes being triggered simultaneously.",
                    )
                    if flight_phase.t == next_phase.t
                    else (
                        (
                            "Trying to add flight phase starting *before* the one *preceding* it. ",
                            "This may be caused by multiple parachutes being triggered simultaneously",
                            " or by having a negative parachute lag.",
                        )
                        if flight_phase.t < previous_phase.t
                        else (
                            "Trying to add flight phase starting *after* the one *proceeding* it.",
                            "This may be caused by multiple parachutes being triggered simultaneously.",
                        )
                    )
                )
            )
            self.display_warning(*warning_msg)
            adjust = 1e-7 if flight_phase.t in {previous_phase.t, next_phase.t} else 0
            new_index = (
                index - 1
                if flight_phase.t < previous_phase.t
                else index + 1
                if flight_phase.t > next_phase.t
                else index
            )
            flight_phase.t += adjust
            self.add(flight_phase, new_index)

        def add_phase(self, t, derivatives=None, callback=None, clear=True, index=None):
            """Add a new flight phase to the list, with the specified
            characteristics. This method creates a new FlightPhase instance and
            adds it to the flight phases list, either at the specified index
            position or appended to the end. See FlightPhases.add() for more
            information.

            Parameters
            ----------
            t : float
                The initial time of the new flight phase.
            derivatives : function, optional
                A function representing the derivatives of the flight phase.
                Default is None.
            callback : list of functions, optional
                A list of callback functions to be executed during the flight
                phase. Default is None. You can also pass an empty list.
            clear : bool, optional
                A flag indicating whether to clear the solution after the phase.
                Default is True.
            index : int, optional
                The index at which the new flight phase should be inserted.
                If not provided, the flight phase will be appended
                to the end of the list. Default is None.

            Returns
            -------
            None
            """
            self.add(self.FlightPhase(t, derivatives, callback, clear), index)

        def flush_after(self, index):
            """This function deletes all flight phases after a given index.

            Parameters
            ----------
            index : int
                The index of the last flight phase to be kept.

            Returns
            -------
            None
            """
            del self.list[index + 1 :]

        class FlightPhase:
            """Class to store a flight phase. It stores the initial time, the
            derivative function, the callback functions and a flag to clear
            the solution after the phase.

            Attributes
            ----------
            t : float
                The initial time of the flight phase.
            derivative : function
                A function representing the derivatives of the flight phase.
            callbacks : list of functions
                A list of callback functions to be executed during the flight
                phase.
            clear : bool
                A flag indicating whether to clear the solution after the phase.
            """

            # TODO: add a "name" optional argument to the FlightPhase. Really helps.

            def __init__(self, t, derivative=None, callbacks=None, clear=True):
                self.t = t
                self.derivative = derivative
                self.callbacks = callbacks[:] if callbacks is not None else []
                self.clear = clear

            def __repr__(self):
                name = "None" if self.derivative is None else self.derivative.__name__
                return (
                    f"<FlightPhase(t= {self.t}, derivative= {name}, "
                    f"callbacks= {self.callbacks}, clear= {self.clear})>"
                )

    class TimeNodes:
        """TimeNodes is a class that stores all the time nodes of a simulation.
        It is meant to work like a python list, but it has some additional
        methods that are useful for the simulation. Them items stored in are
        TimeNodes object are instances of the TimeNode class.
        """

        def __init__(self, init_list=None):
            if not init_list:
                init_list = []
            self.list = init_list[:]

        def __getitem__(self, index):
            return self.list[index]

        def __len__(self):
            return len(self.list)

        def __repr__(self):
            return str(self.list)

        def add(self, time_node):
            self.list.append(time_node)

        def add_node(self, t, parachutes, controllers, sensors):
            self.list.append(self.TimeNode(t, parachutes, controllers, sensors))

        def add_parachutes(self, parachutes, t_init, t_end):
            for parachute in parachutes:
                # Calculate start of sampling time nodes
                sampling_interval = 1 / parachute.sampling_rate
                parachute_node_list = [
                    self.TimeNode(i * sampling_interval, [parachute], [], [])
                    for i in range(
                        math.ceil(t_init / sampling_interval),
                        math.floor(t_end / sampling_interval) + 1,
                    )
                ]
                self.list += parachute_node_list

        def add_controllers(self, controllers, t_init, t_end):
            for controller in controllers:
                # Calculate start of sampling time nodes
                controller_time_step = 1 / controller.sampling_rate
                controller_node_list = [
                    self.TimeNode(i * controller_time_step, [], [controller], [])
                    for i in range(
                        math.ceil(t_init / controller_time_step),
                        math.floor(t_end / controller_time_step) + 1,
                    )
                ]
                self.list += controller_node_list

        def add_sensors(self, sensors, t_init, t_end):
            # Iterate over sensors
            for sensor_component_tuple in sensors:
                # Calculate start of sampling time nodes
                sensor_time_step = 1 / sensor_component_tuple.component.sampling_rate
                sensor_node_list = [
                    self.TimeNode(
                        i * sensor_time_step, [], [], [sensor_component_tuple]
                    )
                    for i in range(
                        math.ceil(t_init / sensor_time_step),
                        math.floor(t_end / sensor_time_step) + 1,
                    )
                ]
                self.list += sensor_node_list

        def sort(self):
            self.list.sort()

        def merge(self):
            """Merge all the time nodes that have the same time. This is made to
            avoid multiple evaluations of the same time node. This method does
            not guarantee the order of the nodes in the list, so it is
            recommended to sort the list before or after using this method.
            """
            tmp_dict = {}
            for node in self.list:
                time = round(node.t, 7)
                try:
                    # Try to access the node and merge if it exists
                    tmp_dict[time].parachutes += node.parachutes
                    tmp_dict[time]._controllers += node._controllers
                    tmp_dict[time].callbacks += node.callbacks
                    tmp_dict[time]._component_sensors += node._component_sensors
                    tmp_dict[time]._controllers += node._controllers
                except KeyError:
                    # If the node does not exist, add it to the dictionary
                    tmp_dict[time] = node
            self.list = list(tmp_dict.values())

        def flush_after(self, index):
            del self.list[index + 1 :]

        class TimeNode:
            """TimeNode is a class that represents a time node in the time
            nodes list. It stores the time, the parachutes and the controllers
            that are active at that time. This class is supposed to work
            exclusively within the TimeNodes class.
            """

            def __init__(self, t, parachutes, controllers, sensors):
                """Create a TimeNode object.

                Parameters
                ----------
                t : float
                    Initial time of the time node.
                parachutes : list[Parachute]
                    List containing all the parachutes that should be evaluated
                    at this time node.
                controllers : list[_Controller]
                    List containing all the controllers that should be evaluated
                    at this time node.
                sensors : list[ComponentSensor]
                    List containing all the sensors that should be evaluated
                    at this time node.
                """
                self.t = t
                self.parachutes = parachutes
                self.callbacks = []
                self._controllers = controllers
                self._component_sensors = sensors

            def __repr__(self):
                return (
                    f"<TimeNode("
                    f"t: {self.t}, "
                    f"parachutes: {len(self.parachutes)}, "
                    f"controllers: {len(self._controllers)}, "
                    f"sensors: {len(self._component_sensors)})>"
                )

            def __lt__(self, other):
                """Allows the comparison of two TimeNode objects based on their
                initial time. This is particularly useful for sorting a list of
                TimeNode objects.

                Parameters
                ----------
                other : TimeNode
                    Another TimeNode object to compare with.

                Returns
                -------
                bool
                    True if the initial time of the current TimeNode is less
                    than the initial time of the other TimeNode, False
                    otherwise.
                """
                return self.t < other.t