import warnings
import numpy as np
from rocketpy.control.controller import _Controller
from rocketpy.mathutils.function import Function
from rocketpy.motors.motor import EmptyMotor
from rocketpy.plots.rocket_plots import _RocketPlots
from rocketpy.prints.rocket_prints import _RocketPrints
from rocketpy.rocket.aero_surface import (
AirBrakes,
EllipticalFins,
Fins,
NoseCone,
RailButtons,
Tail,
TrapezoidalFins,
)
from rocketpy.rocket.components import Components
from rocketpy.rocket.parachute import Parachute
[docs]
class Rocket:
"""Keeps rocket information.
Attributes
----------
Rocket.radius : float
Rocket's largest radius in meters.
Rocket.area : float
Rocket's circular cross section largest frontal area in squared
meters.
Rocket.center_of_dry_mass_position : float
Position, in m, of the rocket's center of dry mass (i.e. center of
mass without propellant) relative to the rocket's coordinate system.
See :doc:`Positions and Coordinate Systems </user/positions>`
for more information
regarding the rocket's coordinate system.
Rocket.center_of_mass_without_motor : int, float
Position, in m, of the rocket's center of mass without motor
relative to the rocket's coordinate system. This does not include
the motor or propellant mass.
Rocket.motor_center_of_mass_position : Function
Position, in meters, of the motor's center of mass relative to the user
defined rocket coordinate system. This is a function of time since the
propellant mass decreases with time. For more information, see the
:doc:`Positions and Coordinate Systems </user/positions>`.
Rocket.motor_center_of_dry_mass_position : float
Position, in meters, of the motor's center of dry mass (i.e. center of
mass without propellant) relative to the user defined rocket coordinate
system. This is constant since the motor dry mass is constant.
Rocket.coordinate_system_orientation : string
String defining the orientation of the rocket's coordinate system.
The coordinate system is defined by the rocket's axis of symmetry.
The system's origin may be placed anywhere along such axis, such as
in the nozzle or in the nose cone, and must be kept the same for all
other positions specified. If "tail_to_nose", the coordinate system
is defined with the rocket's axis of symmetry pointing from the
rocket's tail to the rocket's nose cone. If "nose_to_tail", the
coordinate system is defined with the rocket's axis of symmetry
pointing from the rocket's nose cone to the rocket's tail.
Rocket.mass : float
Rocket's mass without motor and propellant, measured in kg.
Rocket.dry_mass : float
Rocket's mass without propellant, measured in kg. It does include the
motor mass.
Rocket.center_of_mass : Function
Position of the rocket's center of mass, including propellant, relative
to the user defined rocket reference system.
See :doc:`Positions and Coordinate Systems </user/positions>`
for more information
regarding the coordinate system.
Expressed in meters as a function of time.
Rocket.reduced_mass : Function
Function of time expressing the reduced mass of the rocket,
defined as the product of the propellant mass and the mass
of the rocket without propellant, divided by the sum of the
propellant mass and the rocket mass.
Rocket.total_mass : Function
Function of time expressing the total mass of the rocket,
defined as the sum of the propellant mass and the rocket
mass without propellant.
Rocket.thrust_to_weight : Function
Function of time expressing the motor thrust force divided by rocket
weight. The gravitational acceleration is assumed as 9.80665 m/s^2.
Rocket.cp_eccentricity_x : float
Center of pressure position relative to center of mass in the x
axis, perpendicular to axis of cylindrical symmetry, in meters.
Rocket.cp_eccentricity_y : float
Center of pressure position relative to center of mass in the y
axis, perpendicular to axis of cylindrical symmetry, in meters.
Rocket.thrust_eccentricity_y : float
Thrust vector position relative to center of mass in the y
axis, perpendicular to axis of cylindrical symmetry, in meters.
Rocket.thrust_eccentricity_x : float
Thrust vector position relative to center of mass in the x
axis, perpendicular to axis of cylindrical symmetry, in meters.
Rocket.aerodynamic_surfaces : list
Collection of aerodynamic surfaces of the rocket. Holds Nose cones,
Fin sets, and Tails.
Rocket.parachutes : list
Collection of parachutes of the rocket.
Rocket.air_brakes : list
Collection of air brakes of the rocket.
Rocket._controllers : list
Collection of controllers of the rocket.
Rocket.cp_position : Function
Function of Mach number expressing the rocket's center of pressure
position relative to user defined rocket reference system.
See :doc:`Positions and Coordinate Systems </user/positions>`
for more information.
Rocket.stability_margin : Function
Stability margin of the rocket, in calibers, as a function of mach
number and time. Stability margin is defined as the distance between
the center of pressure and the center of mass, divided by the
rocket's diameter.
Rocket.static_margin : Function
Static margin of the rocket, in calibers, as a function of time. Static
margin is defined as the distance between the center of pressure and the
center of mass, divided by the rocket's diameter.
Rocket.static_margin : float
Float value corresponding to rocket static margin when
loaded with propellant in units of rocket diameter or calibers.
Rocket.power_off_drag : Function
Rocket's drag coefficient as a function of Mach number when the
motor is off.
Rocket.power_on_drag : Function
Rocket's drag coefficient as a function of Mach number when the
motor is on.
Rocket.rail_buttons : RailButtons
RailButtons object containing the rail buttons information.
Rocket.motor : Motor
Rocket's motor. See Motor class for more details.
Rocket.motor_position : float
Position, in meters, of the motor's coordinate system origin
relative to the user defined rocket coordinate system.
See :doc:`Positions and Coordinate Systems </user/positions>`
for more information.
regarding the rocket's coordinate system.
Rocket.nozzle_position : float
Position, in meters, of the motor's nozzle exit relative to the user
defined rocket coordinate system.
See :doc:`Positions and Coordinate Systems </user/positions>`
for more information.
Rocket.center_of_propellant_position : Function
Position of the propellant's center of mass relative to the user defined
rocket reference system. See
:doc:`Positions and Coordinate Systems </user/positions>` for more
information regarding the rocket's coordinate system. Expressed in
meters as a function of time.
Rocket.I_11_without_motor : float
Rocket's inertia tensor 11 component without any motors, in kg*m^2. This
is the same value that is passed in the Rocket.__init__() method.
Rocket.I_22_without_motor : float
Rocket's inertia tensor 22 component without any motors, in kg*m^2. This
is the same value that is passed in the Rocket.__init__() method.
Rocket.I_33_without_motor : float
Rocket's inertia tensor 33 component without any motors, in kg*m^2. This
is the same value that is passed in the Rocket.__init__() method.
Rocket.I_12_without_motor : float
Rocket's inertia tensor 12 component without any motors, in kg*m^2. This
is the same value that is passed in the Rocket.__init__() method.
Rocket.I_13_without_motor : float
Rocket's inertia tensor 13 component without any motors, in kg*m^2. This
is the same value that is passed in the Rocket.__init__() method.
Rocket.I_23_without_motor : float
Rocket's inertia tensor 23 component without any motors, in kg*m^2. This
is the same value that is passed in the Rocket.__init__() method.
Rocket.dry_I_11 : float
Rocket's inertia tensor 11 component with unloaded motor,in kg*m^2.
Rocket.dry_I_22 : float
Rocket's inertia tensor 22 component with unloaded motor,in kg*m^2.
Rocket.dry_I_33 : float
Rocket's inertia tensor 33 component with unloaded motor,in kg*m^2.
Rocket.dry_I_12 : float
Rocket's inertia tensor 12 component with unloaded motor,in kg*m^2.
Rocket.dry_I_13 : float
Rocket's inertia tensor 13 component with unloaded motor,in kg*m^2.
Rocket.dry_I_23 : float
Rocket's inertia tensor 23 component with unloaded motor,in kg*m^2.
"""
[docs]
def __init__(
self,
radius,
mass,
inertia,
power_off_drag,
power_on_drag,
center_of_mass_without_motor,
coordinate_system_orientation="tail_to_nose",
):
"""Initializes Rocket class, process inertial, geometrical and
aerodynamic parameters.
Parameters
----------
radius : int, float
Rocket largest outer radius in meters.
mass : int, float
Rocket total mass without motor in kg.
inertia : tuple, list
Tuple or list containing the rocket's inertia tensor components,
in kg*m^2. This should be measured without motor and propellant so
that the inertia reference point is the
`center_of_mass_without_motor`.
Assuming e_3 is the rocket's axis of symmetry, e_1 and e_2 are
orthogonal and form a plane perpendicular to e_3, the inertia tensor
components must be given in the following order: (I_11, I_22, I_33,
I_12, I_13, I_23), where I_ij is the component of the inertia tensor
in the direction of e_i x e_j. Alternatively, the inertia tensor can
be given as (I_11, I_22, I_33), where I_12 = I_13 = I_23 = 0. This
can also be called as "rocket dry inertia tensor".
power_off_drag : int, float, callable, string, array
Rocket's drag coefficient when the motor is off. Can be given as an
entry to the Function class. See help(Function) for more
information. If int or float is given, it is assumed constant. If
callable, string or array is given, it must be a function of Mach
number only.
power_on_drag : int, float, callable, string, array
Rocket's drag coefficient when the motor is on. Can be given as an
entry to the Function class. See help(Function) for more
information. If int or float is given, it is assumed constant. If
callable, string or array is given, it must be a function of Mach
number only.
center_of_mass_without_motor : int, float
Position, in m, of the rocket's center of mass without motor
relative to the rocket's coordinate system. Default is 0, which
means the center of dry mass is chosen as the origin, to comply
with the legacy behavior of versions 0.X.Y.
See :doc:`Positions and Coordinate Systems </user/positions>`
for more information
regarding the rocket's coordinate system.
coordinate_system_orientation : string, optional
String defining the orientation of the rocket's coordinate system.
The coordinate system is defined by the rocket's axis of symmetry.
The system's origin may be placed anywhere along such axis, such as
in the nozzle or in the nose cone, and must be kept the same for all
other positions specified. The two options available are:
"tail_to_nose" and "nose_to_tail". The first defines the coordinate
system with the rocket's axis of symmetry pointing from the rocket's
tail to the rocket's nose cone. The second option defines the
coordinate system with the rocket's axis of symmetry pointing from
the rocket's nose cone to the rocket's tail. Default is
"tail_to_nose".
Returns
-------
None
"""
# Define coordinate system orientation
self.coordinate_system_orientation = coordinate_system_orientation
if coordinate_system_orientation == "tail_to_nose":
self._csys = 1
elif coordinate_system_orientation == "nose_to_tail":
self._csys = -1
else:
raise TypeError(
"Invalid coordinate system orientation. Please choose between "
+ '"tail_to_nose" and "nose_to_tail".'
)
# Define rocket inertia attributes in SI units
self.mass = mass
inertia = (*inertia, 0, 0, 0) if len(inertia) == 3 else inertia
self.I_11_without_motor = inertia[0]
self.I_22_without_motor = inertia[1]
self.I_33_without_motor = inertia[2]
self.I_12_without_motor = inertia[3]
self.I_13_without_motor = inertia[4]
self.I_23_without_motor = inertia[5]
# Define rocket geometrical parameters in SI units
self.center_of_mass_without_motor = center_of_mass_without_motor
self.radius = radius
self.area = np.pi * self.radius**2
# Eccentricity data initialization
self.cp_eccentricity_x = 0
self.cp_eccentricity_y = 0
self.thrust_eccentricity_y = 0
self.thrust_eccentricity_x = 0
# Parachute, Aerodynamic and Rail buttons data initialization
self.parachutes = []
# Controllers data initialization
self._controllers = []
# AirBrakes data initialization
self.air_brakes = []
# Aerodynamic data initialization
self.aerodynamic_surfaces = Components()
self.rail_buttons = Components()
self.cp_position = Function(
lambda mach: 0,
inputs="Mach Number",
outputs="Center of Pressure Position (m)",
)
self.total_lift_coeff_der = Function(
lambda mach: 0,
inputs="Mach Number",
outputs="Total Lift Coefficient Derivative",
)
self.static_margin = Function(
lambda time: 0, inputs="Time (s)", outputs="Static Margin (c)"
)
self.stability_margin = Function(
lambda mach, time: 0,
inputs=["Mach", "Time (s)"],
outputs="Stability Margin (c)",
)
# Define aerodynamic drag coefficients
self.power_off_drag = Function(
power_off_drag,
"Mach Number",
"Drag Coefficient with Power Off",
"linear",
"constant",
)
self.power_on_drag = Function(
power_on_drag,
"Mach Number",
"Drag Coefficient with Power On",
"linear",
"constant",
)
# Create a, possibly, temporary empty motor
# self.motors = Components() # currently unused, only 1 motor is supported
self.add_motor(motor=EmptyMotor(), position=0)
# Important dynamic inertial quantities
self.center_of_mass = None
self.reduced_mass = None
self.total_mass = None
self.dry_mass = None
# calculate dynamic inertial quantities
self.evaluate_dry_mass()
self.evaluate_total_mass()
self.evaluate_center_of_dry_mass()
self.evaluate_center_of_mass()
self.evaluate_reduced_mass()
self.evaluate_thrust_to_weight()
# Evaluate stability (even though no aerodynamic surfaces are present yet)
self.evaluate_center_of_pressure()
self.evaluate_stability_margin()
self.evaluate_static_margin()
# Initialize plots and prints object
self.prints = _RocketPrints(self)
self.plots = _RocketPlots(self)
@property
def nosecones(self):
"""A list containing all the nose cones currently added to the rocket."""
return self.aerodynamic_surfaces.get_by_type(NoseCone)
@property
def fins(self):
"""A list containing all the fins currently added to the rocket."""
return self.aerodynamic_surfaces.get_by_type(Fins)
@property
def tails(self):
"""A list with all the tails currently added to the rocket"""
return self.aerodynamic_surfaces.get_by_type(Tail)
[docs]
def evaluate_total_mass(self):
"""Calculates and returns the rocket's total mass. The total
mass is defined as the sum of the motor mass with propellant and the
rocket mass without propellant. The function returns an object
of the Function class and is defined as a function of time.
Returns
-------
self.total_mass : Function
Function of time expressing the total mass of the rocket,
defined as the sum of the propellant mass and the rocket
mass without propellant.
"""
# Make sure there is a motor associated with the rocket
if self.motor is None:
print("Please associate this rocket with a motor!")
return False
self.total_mass = self.mass + self.motor.total_mass
self.total_mass.set_outputs("Total Mass (Rocket + Motor + Propellant) (kg)")
self.total_mass.set_title("Total Mass (Rocket + Motor + Propellant)")
return self.total_mass
[docs]
def evaluate_dry_mass(self):
"""Calculates and returns the rocket's dry mass. The dry
mass is defined as the sum of the motor's dry mass and the
rocket mass without motor.
Returns
-------
self.dry_mass : float
Rocket's dry mass (Rocket + Motor) (kg)
"""
# Make sure there is a motor associated with the rocket
if self.motor is None:
print("Please associate this rocket with a motor!")
return False
self.dry_mass = self.mass + self.motor.dry_mass
return self.dry_mass
[docs]
def evaluate_center_of_mass(self):
"""Evaluates rocket center of mass position relative to user defined
rocket reference system.
Returns
-------
self.center_of_mass : Function
Function of time expressing the rocket's center of mass position
relative to user defined rocket reference system.
See :doc:`Positions and Coordinate Systems </user/positions>`
for more information.
"""
self.center_of_mass = (
self.center_of_mass_without_motor * self.mass
+ self.motor_center_of_mass_position * self.motor.total_mass
) / self.total_mass
self.center_of_mass.set_inputs("Time (s)")
self.center_of_mass.set_outputs("Center of Mass Position (m)")
self.center_of_mass.set_title(
"Center of Mass Position (Rocket + Motor + Propellant)"
)
return self.center_of_mass
[docs]
def evaluate_center_of_dry_mass(self):
"""Evaluates the rocket's center of dry mass (i.e. rocket with motor but
without propellant) position relative to user defined rocket reference
system.
Returns
-------
self.center_of_dry_mass_position : int, float
Rocket's center of dry mass position (with unloaded motor)
"""
self.center_of_dry_mass_position = (
self.center_of_mass_without_motor * self.mass
+ self.motor_center_of_dry_mass_position * self.motor.dry_mass
) / self.dry_mass
return self.center_of_dry_mass_position
[docs]
def evaluate_reduced_mass(self):
"""Calculates and returns the rocket's total reduced mass. The reduced
mass is defined as the product of the propellant mass and the rocket dry
mass (i.e. with unloaded motor), divided by the loaded rocket mass.
The function returns an object of the Function class and is defined as a
function of time.
Returns
-------
self.reduced_mass : Function
Function of time expressing the reduced mass of the rocket.
"""
# TODO: add tests for reduced_mass values
# Make sure there is a motor associated with the rocket
if self.motor is None:
print("Please associate this rocket with a motor!")
return False
# Get nicknames
prop_mass = self.motor.propellant_mass
dry_mass = self.dry_mass
# calculate reduced mass and return it
self.reduced_mass = prop_mass * dry_mass / (prop_mass + dry_mass)
self.reduced_mass.set_outputs("Reduced Mass (kg)")
self.reduced_mass.set_title("Reduced Mass")
return self.reduced_mass
[docs]
def evaluate_thrust_to_weight(self):
"""Evaluates thrust to weight as a Function of time. This is defined as
the motor thrust force divided by rocket weight. The gravitational
acceleration is assumed constant and equals to 9.80665 m/s^2.
Returns
-------
None
"""
self.thrust_to_weight = self.motor.thrust / (9.80665 * self.total_mass)
self.thrust_to_weight.set_inputs("Time (s)")
self.thrust_to_weight.set_outputs("Thrust/Weight")
self.thrust_to_weight.set_title("Thrust to Weight ratio")
[docs]
def evaluate_center_of_pressure(self):
"""Evaluates rocket center of pressure position relative to user defined
rocket reference system. It can be called as many times as needed, as it
will update the center of pressure function every time it is called. The
code will iterate through all aerodynamic surfaces and consider each of
their center of pressure position and derivative of the coefficient of
lift as a function of Mach number.
Returns
-------
self.cp_position : Function
Function of Mach number expressing the rocket's center of pressure
position relative to user defined rocket reference system.
See :doc:`Positions and Coordinate Systems </user/positions>`
for more information.
"""
# Re-Initialize total lift coefficient derivative and center of pressure position
self.total_lift_coeff_der.set_source(lambda mach: 0)
self.cp_position.set_source(lambda mach: 0)
# Calculate total lift coefficient derivative and center of pressure
if len(self.aerodynamic_surfaces) > 0:
for aero_surface, position in self.aerodynamic_surfaces:
# ref_factor corrects lift for different reference areas
ref_factor = (aero_surface.rocket_radius / self.radius) ** 2
self.total_lift_coeff_der += ref_factor * aero_surface.clalpha
self.cp_position += (
ref_factor
* aero_surface.clalpha
* (position - self._csys * aero_surface.cpz)
)
self.cp_position /= self.total_lift_coeff_der
return self.cp_position
[docs]
def evaluate_stability_margin(self):
"""Calculates the stability margin of the rocket as a function of mach
number and time.
Returns
-------
stability_margin : Function
Stability margin of the rocket, in calibers, as a function of mach
number and time. Stability margin is defined as the distance between
the center of pressure and the center of mass, divided by the
rocket's diameter.
"""
self.stability_margin.set_source(
lambda mach, time: (
(self.center_of_mass(time) - self.cp_position(mach)) / (2 * self.radius)
)
* self._csys
)
return self.stability_margin
[docs]
def evaluate_static_margin(self):
"""Calculates the static margin of the rocket as a function of time.
Returns
-------
static_margin : Function
Static margin of the rocket, in calibers, as a function of time.
Static margin is defined as the distance between the center of
pressure and the center of mass, divided by the rocket's diameter.
"""
# Calculate static margin
self.static_margin.set_source(
lambda time: (self.center_of_mass(time) - self.cp_position(0))
/ (2 * self.radius)
)
# Change sign if coordinate system is upside down
self.static_margin *= self._csys
self.static_margin.set_inputs("Time (s)")
self.static_margin.set_outputs("Static Margin (c)")
self.static_margin.set_title("Static Margin")
self.static_margin.set_discrete(
lower=0, upper=self.motor.burn_out_time, samples=200
)
return self.static_margin
[docs]
def evaluate_dry_inertias(self):
"""Calculates and returns the rocket's dry inertias relative to
the rocket's center of mass. The inertias are saved and returned
in units of kg*m². This does not consider propellant mass but does take
into account the motor dry mass.
Returns
-------
self.dry_I_11 : float
Float value corresponding to rocket inertia tensor 11
component, which corresponds to the inertia relative to the
e_1 axis, centered at the instantaneous center of mass.
self.dry_I_22 : float
Float value corresponding to rocket inertia tensor 22
component, which corresponds to the inertia relative to the
e_2 axis, centered at the instantaneous center of mass.
self.dry_I_33 : float
Float value corresponding to rocket inertia tensor 33
component, which corresponds to the inertia relative to the
e_3 axis, centered at the instantaneous center of mass.
self.dry_I_12 : float
Float value corresponding to rocket inertia tensor 12
component, which corresponds to the inertia relative to the
e_1 and e_2 axes, centered at the instantaneous center of mass.
self.dry_I_13 : float
Float value corresponding to rocket inertia tensor 13
component, which corresponds to the inertia relative to the
e_1 and e_3 axes, centered at the instantaneous center of mass.
self.dry_I_23 : float
Float value corresponding to rocket inertia tensor 23
component, which corresponds to the inertia relative to the
e_2 and e_3 axes, centered at the instantaneous center of mass.
Notes
-----
The e_1 and e_2 directions are assumed to be the directions
perpendicular to the rocket axial direction.
The e_3 direction is assumed to be the direction parallel to the axis
of symmetry of the rocket.
RocketPy follows the definition of the inertia tensor as in [1], which
includes the minus sign for all products of inertia.
References
----------
.. [1] https://en.wikipedia.org/wiki/Moment_of_inertia#Inertia_tensor
"""
# Compute axes distances
noMCM_to_CDM = (
self.center_of_mass_without_motor - self.center_of_dry_mass_position
)
motorCDM_to_CDM = (
self.motor_center_of_dry_mass_position - self.center_of_dry_mass_position
)
# Compute dry inertias
self.dry_I_11 = (
self.I_11_without_motor
+ self.mass * noMCM_to_CDM**2
+ self.motor.dry_I_11
+ self.motor.dry_mass * motorCDM_to_CDM**2
)
self.dry_I_22 = (
self.I_22_without_motor
+ self.mass * noMCM_to_CDM**2
+ self.motor.dry_I_22
+ self.motor.dry_mass * motorCDM_to_CDM**2
)
self.dry_I_33 = self.I_33_without_motor + self.motor.dry_I_33
self.dry_I_12 = self.I_12_without_motor + self.motor.dry_I_12
self.dry_I_13 = self.I_13_without_motor + self.motor.dry_I_13
self.dry_I_23 = self.I_23_without_motor + self.motor.dry_I_23
return (
self.dry_I_11,
self.dry_I_22,
self.dry_I_33,
self.dry_I_12,
self.dry_I_13,
self.dry_I_23,
)
[docs]
def evaluate_inertias(self):
"""Calculates and returns the rocket's inertias relative to
the rocket's center of mass. The inertias are saved and returned
in units of kg*m².
Returns
-------
self.I_11 : float
Float value corresponding to rocket inertia tensor 11
component, which corresponds to the inertia relative to the
e_1 axis, centered at the instantaneous center of mass.
self.I_22 : float
Float value corresponding to rocket inertia tensor 22
component, which corresponds to the inertia relative to the
e_2 axis, centered at the instantaneous center of mass.
self.I_33 : float
Float value corresponding to rocket inertia tensor 33
component, which corresponds to the inertia relative to the
e_3 axis, centered at the instantaneous center of mass.
Notes
-----
The e_1 and e_2 directions are assumed to be the directions
perpendicular to the rocket axial direction.
The e_3 direction is assumed to be the direction parallel to the axis
of symmetry of the rocket.
RocketPy follows the definition of the inertia tensor as in [1], which
includes the minus sign for all products of inertia.
References
----------
.. [1] https://en.wikipedia.org/wiki/Moment_of_inertia#Inertia_tensor
"""
# Get masses
prop_mass = self.motor.propellant_mass # Propellant mass as a function of time
dry_mass = self.dry_mass # Constant rocket mass with motor, without propellant
# Compute axes distances
CM_to_CDM = self.center_of_mass - self.center_of_dry_mass_position
CM_to_CPM = self.center_of_mass - self.center_of_propellant_position
# Compute inertias
self.I_11 = (
self.dry_I_11
+ self.motor.I_11
+ dry_mass * CM_to_CDM**2
+ prop_mass * CM_to_CPM**2
)
self.I_22 = (
self.dry_I_22
+ self.motor.I_22
+ dry_mass * CM_to_CDM**2
+ prop_mass * CM_to_CPM**2
)
self.I_33 = self.dry_I_33 + self.motor.I_33
self.I_12 = self.dry_I_12 + self.motor.I_12
self.I_13 = self.dry_I_13 + self.motor.I_13
self.I_23 = self.dry_I_23 + self.motor.I_23
# Return inertias
return (
self.I_11,
self.I_22,
self.I_33,
self.I_12,
self.I_13,
self.I_23,
)
def evaluate_nozzle_gyration_tensor(self):
pass
[docs]
def add_motor(self, motor, position):
"""Adds a motor to the rocket.
Parameters
----------
motor : Motor, SolidMotor, HybridMotor, LiquidMotor, GenericMotor
Motor to be added to the rocket.
position : int, float
Position, in meters, of the motor's coordinate system origin
relative to the user defined rocket coordinate system.
See Also
--------
:ref:`add_surfaces`
Returns
-------
None
"""
if hasattr(self, "motor") and not isinstance(self.motor, EmptyMotor):
print(
"Only one motor per rocket is currently supported. "
+ "Overwriting previous motor."
)
self.motor = motor
self.motor_position = position
_ = self._csys * self.motor._csys
self.center_of_propellant_position = (
self.motor.center_of_propellant_mass * _ + self.motor_position
)
self.motor_center_of_mass_position = (
self.motor.center_of_mass * _ + self.motor_position
)
self.motor_center_of_dry_mass_position = (
self.motor.center_of_dry_mass_position * _ + self.motor_position
)
self.nozzle_position = self.motor.nozzle_position * _ + self.motor_position
self.evaluate_dry_mass()
self.evaluate_total_mass()
self.evaluate_center_of_dry_mass()
self.evaluate_center_of_mass()
self.evaluate_dry_inertias()
self.evaluate_inertias()
self.evaluate_reduced_mass()
self.evaluate_thrust_to_weight()
self.evaluate_center_of_pressure()
self.evaluate_stability_margin()
self.evaluate_static_margin()
[docs]
def add_surfaces(self, surfaces, positions):
"""Adds one or more aerodynamic surfaces to the rocket. The aerodynamic
surface must be an instance of a class that inherits from the
AeroSurface (e.g. NoseCone, TrapezoidalFins, etc.)
Parameters
----------
surfaces : list, AeroSurface, NoseCone, TrapezoidalFins, EllipticalFins, Tail
Aerodynamic surface to be added to the rocket. Can be a list of
AeroSurface if more than one surface is to be added.
positions : int, float, list
Position, in m, of the aerodynamic surface's center of pressure
relative to the user defined rocket coordinate system.
If a list is passed, it will correspond to the position of each item
in the surfaces list.
For NoseCone type, position is relative to the nose cone tip.
For Fins type, position is relative to the point belonging to
the root chord which is highest in the rocket coordinate system.
For Tail type, position is relative to the point belonging to the
tail which is highest in the rocket coordinate system.
See Also
--------
:ref:`add_surfaces`
Returns
-------
None
"""
try:
for surface, position in zip(surfaces, positions):
self.aerodynamic_surfaces.add(surface, position)
except TypeError:
self.aerodynamic_surfaces.add(surfaces, positions)
self.evaluate_center_of_pressure()
self.evaluate_stability_margin()
self.evaluate_static_margin()
[docs]
def _add_controllers(self, controllers):
"""Adds a controller to the rocket.
Parameters
----------
controllers : list of Controller objects
List of controllers to be added to the rocket. If a single
Controller object is passed, outside of a list, a try/except block
will be used to try to append the controller to the list.
Returns
-------
None
"""
try:
self._controllers.extend(controllers)
except TypeError:
self._controllers.append(controllers)
[docs]
def add_tail(
self, top_radius, bottom_radius, length, position, radius=None, name="Tail"
):
"""Create a new tail or rocket diameter change, storing its
parameters as part of the aerodynamic_surfaces list. Its
parameters are the axial position along the rocket and its
derivative of the coefficient of lift in respect to angle of
attack.
Parameters
----------
top_radius : int, float
Tail top radius in meters, considering positive direction
from center of mass to nose cone.
bottom_radius : int, float
Tail bottom radius in meters, considering positive direction
from center of mass to nose cone.
length : int, float
Tail length or height in meters. Must be a positive value.
position : int, float
Tail position relative to the rocket's coordinate system.
By tail position, understand the point belonging to the tail which
is highest in the rocket coordinate system (i.e. the point
closest to the nose cone).
See Also
--------
:ref:`add_surfaces`
Returns
-------
tail : Tail
Tail object created.
"""
# Modify reference radius if not provided
radius = self.radius if radius is None else radius
# Create tail, adds it to the rocket and returns it
tail = Tail(top_radius, bottom_radius, length, radius, name)
self.add_surfaces(tail, position)
return tail
[docs]
def add_nose(
self, length, kind, position, bluffness=0, name="Nose Cone", base_radius=None
):
"""Creates a nose cone, storing its parameters as part of the
aerodynamic_surfaces list. Its parameters are the axial position
along the rocket and its derivative of the coefficient of lift
in respect to angle of attack.
Parameters
----------
length : int, float
Nose cone length or height in meters. Must be a positive
value.
kind : string
Nose cone type. Von Karman, conical, ogive, and lvhaack are
supported.
position : int, float
Nose cone tip coordinate relative to the rocket's coordinate system.
See `Rocket.coordinate_system_orientation` for more information.
bluffness : float, optional
Ratio between the radius of the circle on the tip of the ogive and
the radius of the base of the ogive.
name : string
Nose cone name. Default is "Nose Cone".
base_radius : int, float, optional
Nose cone base radius in meters. If not given, the rocket radius
will be used.
See Also
--------
:ref:`add_surfaces`
Returns
-------
nose : Nose
Nose cone object created.
"""
nose = NoseCone(
length=length,
kind=kind,
base_radius=base_radius or self.radius,
rocket_radius=base_radius or self.radius,
bluffness=bluffness,
name=name,
)
self.add_surfaces(nose, position)
return nose
[docs]
def add_fins(self, *args, **kwargs):
"""See Rocket.add_trapezoidal_fins for documentation.
This method is set to be deprecated in version 1.0.0 and fully removed
by version 2.0.0. Use Rocket.add_trapezoidal_fins instead. It keeps the
same arguments and signature."""
warnings.warn(
"This method is set to be deprecated in version 1.0.0 and fully "
"removed by version 2.0.0. Use Rocket.add_trapezoidal_fins instead",
PendingDeprecationWarning,
)
return self.add_trapezoidal_fins(*args, **kwargs)
[docs]
def add_trapezoidal_fins(
self,
n,
root_chord,
tip_chord,
span,
position,
cant_angle=0,
sweep_length=None,
sweep_angle=None,
radius=None,
airfoil=None,
name="Fins",
):
"""Create a trapezoidal fin set, storing its parameters as part of the
aerodynamic_surfaces list. Its parameters are the axial position along
the rocket and its derivative of the coefficient of lift in respect to
angle of attack.
Parameters
----------
n : int
Number of fins, from 2 to infinity.
span : int, float
Fin span in meters.
root_chord : int, float
Fin root chord in meters.
tip_chord : int, float
Fin tip chord in meters.
position : int, float
Fin set position relative to the rocket's coordinate system.
By fin set position, understand the point belonging to the root
chord which is highest in the rocket coordinate system (i.e.
the point closest to the nose cone tip).
See Also
--------
:ref:`add_surfaces`
cant_angle : int, float, optional
Fins cant angle with respect to the rocket centerline. Must
be given in degrees.
sweep_length : int, float, optional
Fins sweep length in meters. By sweep length, understand the axial
distance between the fin root leading edge and the fin tip leading
edge measured parallel to the rocket centerline. If not given, the
sweep length is assumed to be equal the root chord minus the tip
chord, in which case the fin is a right trapezoid with its base
perpendicular to the rocket's axis. Cannot be used in conjunction
with sweep_angle.
sweep_angle : int, float, optional
Fins sweep angle with respect to the rocket centerline. Must be
given in degrees. If not given, the sweep angle is automatically
calculated, in which case the fin is assumed to be a right trapezoid
with its base perpendicular to the rocket's axis. Cannot be used in
conjunction with sweep_length.
radius : int, float, optional
Reference fuselage radius where the fins are located. This is used
to calculate lift coefficient and to draw the rocket. If None,
which is default, the rocket radius will be used.
airfoil : tuple, optional
Default is null, in which case fins will be treated as flat plates.
Otherwise, if tuple, fins will be considered as airfoils. The
tuple's first item specifies the airfoil's lift coefficient
by angle of attack and must be either a .csv, .txt, ndarray
or callable. The .csv and .txt files can contain a single line
header and the first column must specify the angle of attack, while
the second column must specify the lift coefficient. The
ndarray should be as [(x0, y0), (x1, y1), (x2, y2), ...]
where x0 is the angle of attack and y0 is the lift coefficient.
If callable, it should take an angle of attack as input and
return the lift coefficient at that angle of attack.
The tuple's second item is the unit of the angle of attack,
accepting either "radians" or "degrees".
Returns
-------
fin_set : TrapezoidalFins
Fin set object created.
"""
# Modify radius if not given, use rocket radius, otherwise use given.
radius = radius if radius is not None else self.radius
# Create a fin set as an object of TrapezoidalFins class
fin_set = TrapezoidalFins(
n,
root_chord,
tip_chord,
span,
radius,
cant_angle,
sweep_length,
sweep_angle,
airfoil,
name,
)
# Add fin set to the list of aerodynamic surfaces
self.add_surfaces(fin_set, position)
return fin_set
[docs]
def add_elliptical_fins(
self,
n,
root_chord,
span,
position,
cant_angle=0,
radius=None,
airfoil=None,
name="Fins",
):
"""Create an elliptical fin set, storing its parameters as part of the
aerodynamic_surfaces list. Its parameters are the axial position along
the rocket and its derivative of the coefficient of lift in respect to
angle of attack.
Parameters
----------
n : int
Number of fins, from 2 to infinity.
root_chord : int, float
Fin root chord in meters.
span : int, float
Fin span in meters.
position : int, float
Fin set position relative to the rocket's coordinate system. By fin
set position, understand the point belonging to the root chord which
is highest in the rocket coordinate system (i.e. the point
closest to the nose cone tip).
See Also
--------
:ref:`add_surfaces`
cant_angle : int, float, optional
Fins cant angle with respect to the rocket centerline. Must be given
in degrees.
radius : int, float, optional
Reference fuselage radius where the fins are located. This is used
to calculate lift coefficient and to draw the rocket. If None,
which is default, the rocket radius will be used.
airfoil : tuple, optional
Default is null, in which case fins will be treated as flat plates.
Otherwise, if tuple, fins will be considered as airfoils. The
tuple's first item specifies the airfoil's lift coefficient
by angle of attack and must be either a .csv, .txt, ndarray
or callable. The .csv and .txt files can contain a single line
header and the first column must specify the angle of attack, while
the second column must specify the lift coefficient. The
ndarray should be as [(x0, y0), (x1, y1), (x2, y2), ...]
where x0 is the angle of attack and y0 is the lift coefficient.
If callable, it should take an angle of attack as input and
return the lift coefficient at that angle of attack.
The tuple's second item is the unit of the angle of attack,
accepting either "radians" or "degrees".
Returns
-------
fin_set : EllipticalFins
Fin set object created.
"""
radius = radius if radius is not None else self.radius
fin_set = EllipticalFins(n, root_chord, span, radius, cant_angle, airfoil, name)
self.add_surfaces(fin_set, position)
return fin_set
[docs]
def add_parachute(
self, name, cd_s, trigger, sampling_rate=100, lag=0, noise=(0, 0, 0)
):
"""Creates a new parachute, storing its parameters such as
opening delay, drag coefficients and trigger function.
Parameters
----------
name : string
Parachute name, such as drogue and main. Has no impact in
simulation, as it is only used to display data in a more
organized matter.
cd_s : float
Drag coefficient times reference area for parachute. It is
used to compute the drag force exerted on the parachute by
the equation F = ((1/2)*rho*V^2)*cd_s, that is, the drag
force is the dynamic pressure computed on the parachute
times its cd_s coefficient. Has units of area and must be
given in squared meters.
trigger : function, float, string
This parameter defines the trigger condition for the parachute
ejection system. It can be one of the following:
- A callable function that takes three arguments:
1. Freestream pressure in pascals.
2. Height in meters above ground level.
3. The state vector of the simulation, which is defined as:
[x, y, z, vx, vy, vz, e0, e1, e2, e3, wx, wy, wz].
The function should return True if the parachute ejection system should
be triggered and False otherwise.
- A float value, representing an absolute height in meters. In this
case, the parachute will be ejected when the rocket reaches this height
above ground level.
- The string "apogee" which triggers the parachute at apogee, i.e.,
when the rocket reaches its highest point and starts descending.
Note: The function will be called according to the sampling rate
specified next.
sampling_rate : float, optional
Sampling rate in which the trigger function works. It is used to
simulate the refresh rate of onboard sensors such as barometers.
Default value is 100. Value must be given in hertz.
lag : float, optional
Time between the parachute ejection system is triggered and the
parachute is fully opened. During this time, the simulation will
consider the rocket as flying without a parachute. Default value
is 0. Must be given in seconds.
noise : tuple, list, optional
List in the format (mean, standard deviation, time-correlation).
The values are used to add noise to the pressure signal which is
passed to the trigger function. Default value is (0, 0, 0). Units
are in pascal.
Returns
-------
parachute : Parachute
Parachute containing trigger, sampling_rate, lag, cd_s, noise
and name. Furthermore, it stores clean_pressure_signal,
noise_signal and noisyPressureSignal which are filled in during
Flight simulation.
"""
parachute = Parachute(name, cd_s, trigger, sampling_rate, lag, noise)
self.parachutes.append(parachute)
return self.parachutes[-1]
[docs]
def add_air_brakes(
self,
drag_coefficient_curve,
controller_function,
sampling_rate,
clamp=True,
reference_area=None,
initial_observed_variables=None,
override_rocket_drag=False,
return_controller=False,
name="AirBrakes",
controller_name="AirBrakes Controller",
):
"""Creates a new air brakes system, storing its parameters such as
drag coefficient curve, controller function, sampling rate, and
reference area.
Parameters
----------
drag_coefficient_curve : int, float, callable, array, string, Function
This parameter represents the drag coefficient associated with the
air brakes and/or the entire rocket, depending on the value of
``override_rocket_drag``.
- If a constant, it should be an integer or a float representing a
fixed drag coefficient value.
- If a function, it must take two parameters: deployment level and
Mach number, and return the drag coefficient. This function allows
for dynamic computation based on deployment and Mach number.
- If an array, it should be a 2D array with three columns: the first
column for deployment level, the second for Mach number, and the
third for the corresponding drag coefficient.
- If a string, it should be the path to a .csv or .txt file. The
file must contain three columns: the first for deployment level,
the second for Mach number, and the third for the drag
coefficient.
- If a Function, it must take two parameters: deployment level and
Mach number, and return the drag coefficient.
.. note:: For ``override_rocket_drag = False``, at
deployment level 0, the drag coefficient is assumed to be 0,
independent of the input drag coefficient curve. This means that
the simulation always considers that at a deployment level of 0,
the air brakes are completely retracted and do not contribute to
the drag of the rocket.
controller_function : function, callable
An user-defined function responsible for controlling the simulation.
This function is expected to take the following arguments, in order:
1. `time` (float): The current simulation time in seconds.
2. `sampling_rate` (float): The rate at which the controller
function is called, measured in Hertz (Hz).
3. `state` (list): The state vector of the simulation, structured as
`[x, y, z, vx, vy, vz, e0, e1, e2, e3, wx, wy, wz]`.
4. `state_history` (list): A record of the rocket's state at each
step throughout the simulation. The state_history is organized as a
list of lists, with each sublist containing a state vector. The last
item in the list always corresponds to the previous state vector,
providing a chronological sequence of the rocket's evolving states.
5. `observed_variables` (list): A list containing the variables that
the controller function returns. The initial value in the first
step of the simulation of this list is provided by the
`initial_observed_variables` argument.
6. `interactive_objects` (list): A list containing the objects that
the controller function can interact with. The objects are
listed in the same order as they are provided in the
`interactive_objects`
This function will be called during the simulation at the specified
sampling rate. The function should evaluate and change the observed
objects as needed. The function should return None.
.. note:: The function will be called according to the sampling rate
specified.
sampling_rate : float
The sampling rate of the controller function in Hertz (Hz). This
means that the controller function will be called every
`1/sampling_rate` seconds.
clamp : bool, optional
If True, the simulation will clamp the deployment level to 0 or 1 if
the deployment level is out of bounds. If False, the simulation will
not clamp the deployment level and will instead raise a warning if
the deployment level is out of bounds. Default is True.
reference_area : float, optional
Reference area used to calculate the drag force of the air brakes
from the drag coefficient curve. If None, which is default, use
rocket section area. Must be given in squared meters.
initial_observed_variables : list, optional
A list of the initial values of the variables that the controller
function returns. This list is used to initialize the
`observed_variables` argument of the controller function. The
default value is None, which initializes the list as an empty list.
override_rocket_drag : bool, optional
If False, the air brakes drag coefficient will be added to the
rocket's power off drag coefficient curve. If True, during the
simulation, the rocket's power off drag will be ignored and the air
brakes drag coefficient will be used for the entire rocket instead.
Default is False.
return_controller : bool, optional
If True, the function will return the controller object created.
Default is False.
name : string, optional
AirBrakes name, such as drogue and main. Has no impact in
simulation, as it is only used to display data in a more
organized matter.
controller_name : string, optional
Controller name. Has no impact in simulation, as it is only used to
display data in a more organized matter.
Returns
-------
air_brakes : AirBrakes
AirBrakes object created.
controller : Controller
Controller object created.
"""
reference_area = reference_area if reference_area is not None else self.area
air_brakes = AirBrakes(
drag_coefficient_curve=drag_coefficient_curve,
reference_area=reference_area,
clamp=clamp,
override_rocket_drag=override_rocket_drag,
deployment_level=0,
name=name,
)
_controller = _Controller(
interactive_objects=air_brakes,
controller_function=controller_function,
sampling_rate=sampling_rate,
initial_observed_variables=initial_observed_variables,
name=controller_name,
)
self.air_brakes.append(air_brakes)
self._add_controllers(_controller)
if return_controller:
return air_brakes, _controller
else:
return air_brakes
[docs]
def add_cm_eccentricity(self, x, y):
"""Moves line of action of aerodynamic and thrust forces by
equal translation amount to simulate an eccentricity in the
position of the center of mass of the rocket relative to its
geometrical center line.
Parameters
----------
x : float
Distance in meters by which the CM is to be translated in
the x direction relative to geometrical center line.
y : float
Distance in meters by which the CM is to be translated in
the y direction relative to geometrical center line.
Returns
-------
self : Rocket
Object of the Rocket class.
Notes
-----
Should not be used together with add_cp_eccentricity and
add_thrust_eccentricity.
"""
self.cp_eccentricity_x = -x
self.cp_eccentricity_y = -y
self.thrust_eccentricity_y = -x
self.thrust_eccentricity_x = -y
return self
[docs]
def add_cp_eccentricity(self, x, y):
"""Moves line of action of aerodynamic forces to simulate an
eccentricity in the position of the center of pressure relative
to the center of mass of the rocket.
Parameters
----------
x : float
Distance in meters by which the CP is to be translated in
the x direction relative to the center of mass axial line.
y : float
Distance in meters by which the CP is to be translated in
the y direction relative to the center of mass axial line.
Returns
-------
self : Rocket
Object of the Rocket class.
"""
self.cp_eccentricity_x = x
self.cp_eccentricity_y = y
return self
[docs]
def add_thrust_eccentricity(self, x, y):
"""Moves line of action of thrust forces to simulate a
misalignment of the thrust vector and the center of mass.
Parameters
----------
x : float
Distance in meters by which the line of action of the
thrust force is to be translated in the x direction
relative to the center of mass axial line.
y : float
Distance in meters by which the line of action of the
thrust force is to be translated in the x direction
relative to the center of mass axial line.
Returns
-------
self : Rocket
Object of the Rocket class.
"""
self.thrust_eccentricity_y = x
self.thrust_eccentricity_x = y
return self
[docs]
def draw(self, vis_args=None):
"""Draws the rocket in a matplotlib figure.
Parameters
----------
vis_args : dict, optional
Determines the visual aspects when drawing the rocket. If None,
default values are used. Default values are:
{
"background": "#EEEEEE",
"tail": "black",
"nose": "black",
"body": "dimgrey",
"fins": "black",
"motor": "black",
"buttons": "black",
"line_width": 2.0,
}
A full list of color names can be found at:
https://matplotlib.org/stable/gallery/color/named_colors
"""
self.plots.draw(vis_args)
return None
[docs]
def info(self):
"""Prints out a summary of the data and graphs available about
the Rocket.
Returns
-------
None
"""
self.prints.all()
[docs]
def all_info(self):
"""Prints out all data and graphs available about the Rocket.
Returns
-------
None
"""
self.info()
self.plots.all()