Skip to content

models.propulsion.motors

machwave.models.propulsion.motors

LiquidEngine

Bases: Motor[BiliquidPropellant, LiquidEngineThrustChamber]

Source code in machwave/models/propulsion/motors/liquid.py 
 13
 14
 15
 16
 17
 18
 19
 20
 21
 22
 23
 24
 25
 26
 27
 28
 29
 30
 31
 32
 33
 34
 35
 36
 37
 38
 39
 40
 41
 42
 43
 44
 45
 46
 47
 48
 49
 50
 51
 52
 53
 54
 55
 56
 57
 58
 59
 60
 61
 62
 63
 64
 65
 66
 67
 68
 69
 70
 71
 72
 73
 74
 75
 76
 77
 78
 79
 80
 81
 82
 83
 84
 85
 86
 87
 88
 89
 90
 91
 92
 93
 94
 95
 96
 97
 98
 99
100
101
102
103
104
105
106
107
108
109
110
111
112
113
114
115
116
117
118
119
120
121
122
123
124
125
126
127
128
129
130
131
132
133
134
135
136
137
138
139
140
141
142
143
144
145
146
147
148
149
150
151
class LiquidEngine(Motor[BiliquidPropellant, LiquidEngineThrustChamber]):
    def __init__(
        self,
        propellant: BiliquidPropellant,
        thrust_chamber: LiquidEngineThrustChamber,
        feed_system: FeedSystem,
        oxidizer_tank_cog: float | None = None,
        fuel_tank_cog: float | None = None,
        other_losses: float = 12.0,
    ) -> None:
        """
        Initialize a liquid rocket engine.

        Args:
            propellant: Bi-liquid propellant properties (oxidizer + fuel).
            thrust_chamber: Thrust chamber assembly (nozzle + combustion chamber + injector).
            feed_system: Propellant feed system (tanks, lines, pumps/pressurization).
            oxidizer_tank_cog: Axial position of the oxidizer tank center (where propellant CoG is),
                measured from the nozzle exit, in meters. If None, uses a default estimate.
            fuel_tank_cog: Axial position of the fuel tank center (where propellant CoG is),
                measured from the nozzle exit, in meters. If None, uses a default estimate.
            other_losses: Additional engine losses not accounted for by specific
                loss mechanisms, in percent. Defaults to 12%.
        """
        super().__init__(propellant, thrust_chamber, other_losses)
        self.feed_system = feed_system
        self.oxidizer_tank_cog = oxidizer_tank_cog
        self.fuel_tank_cog = fuel_tank_cog

    @property
    def initial_propellant_mass(self) -> float:
        """
        Returns the initial propellant mass in kg.
        """
        return self.feed_system.get_propellant_mass()

    def get_launch_mass(self) -> float:
        return self.thrust_chamber.dry_mass + self.initial_propellant_mass

    def get_dry_mass(self) -> float:
        return self.thrust_chamber.dry_mass

    def get_center_of_gravity(
        self, propellant_fraction: float = 0.0
    ) -> np.typing.NDArray[np.float64]:
        """
        Calculate the center of gravity of the liquid engine including structural
        dry mass, oxidizer, and fuel.

        The calculation uses a mass-weighted average of:
        1. Structural dry mass (thrust chamber, tanks structure, feed lines, etc.)
        2. Oxidizer mass (from oxidizer tank)
        3. Fuel mass (from fuel tank)

        Args:
            propellant_fraction: Fraction of propellant consumed (0.0 = full, 1.0 = empty).

        Returns:
            Center of gravity in 3D space [x, y, z], in meters.
            Origin is at the nozzle exit on the chamber axis.
            Positive x-direction points forward (toward bulkhead/away from nozzle exit).

        Raises:
            ValueError: If thrust_chamber.center_of_gravity_coordinate, oxidizer_tank_cog,
                or fuel_tank_cog is not defined.
        """
        # Get current propellant masses
        initial_ox_mass = self.feed_system.oxidizer_tank.initial_fluid_mass
        initial_fuel_mass = self.feed_system.fuel_tank.initial_fluid_mass

        ox_mass = initial_ox_mass * (1.0 - propellant_fraction)
        fuel_mass = initial_fuel_mass * (1.0 - propellant_fraction)

        # Structural dry mass (thrust chamber, tank structure, feed lines, etc.)
        dry_mass = self.thrust_chamber.dry_mass

        if self.thrust_chamber.center_of_gravity_coordinate is None:
            raise ValueError(
                "Thrust chamber center of gravity coordinate is not defined."
            )

        dry_cog = self.thrust_chamber.center_of_gravity_coordinate

        # Oxidizer tank CoG
        if self.oxidizer_tank_cog is None:
            raise ValueError("Oxidizer tank center of gravity is not defined.")

        ox_cog = np.array([self.oxidizer_tank_cog, 0.0, 0.0], dtype=np.float64)

        # Fuel tank CoG
        if self.fuel_tank_cog is None:
            raise ValueError("Fuel tank center of gravity is not defined.")

        fuel_cog = np.array([self.fuel_tank_cog, 0.0, 0.0], dtype=np.float64)

        # Calculate total mass and weighted CoG
        total_mass = dry_mass + ox_mass + fuel_mass

        if total_mass <= 0:
            # Fallback if calculation fails
            return dry_cog

        weighted_cog = (
            dry_cog * dry_mass + ox_cog * ox_mass + fuel_cog * fuel_mass
        ) / total_mass

        return weighted_cog.astype(np.float64)

    def get_thrust_coefficient(
        self,
        chamber_pressure: float,
        exit_pressure: float,
        external_pressure: float,
        expansion_ratio: float,
        k_ex: float,
        n_cf: float,
    ) -> float:
        """Get thrust coefficient.

        Args:
            chamber_pressure: Chamber pressure [Pa].
            exit_pressure: Exit pressure [Pa].
            external_pressure: External pressure [Pa].
            expansion_ratio: Expansion ratio.
            k_ex: Two-phase isentropic coefficient.
            n_cf: Thrust coefficient correction factor.

        Returns:
            Instantaneous thrust coefficient.
        """
        cf_ideal = get_ideal_thrust_coefficient(
            chamber_pressure=chamber_pressure,
            exit_pressure=exit_pressure,
            external_pressure=external_pressure,
            expansion_ratio=expansion_ratio,
            k_ex=k_ex,
        )
        n_cf = self.get_thrust_coefficient_correction_factor()
        return cf_ideal * n_cf

initial_propellant_mass property

Returns the initial propellant mass in kg.

__init__(propellant, thrust_chamber, feed_system, oxidizer_tank_cog=None, fuel_tank_cog=None, other_losses=12.0)

Initialize a liquid rocket engine.

Parameters:

Name Type Description Default
propellant BiliquidPropellant

Bi-liquid propellant properties (oxidizer + fuel).

 required
thrust_chamber LiquidEngineThrustChamber

Thrust chamber assembly (nozzle + combustion chamber + injector).

 required
feed_system FeedSystem

Propellant feed system (tanks, lines, pumps/pressurization).

 required
oxidizer_tank_cog float | None

Axial position of the oxidizer tank center (where propellant CoG is), measured from the nozzle exit, in meters. If None, uses a default estimate.

 None
fuel_tank_cog float | None

Axial position of the fuel tank center (where propellant CoG is), measured from the nozzle exit, in meters. If None, uses a default estimate.

 None
other_losses float

Additional engine losses not accounted for by specific loss mechanisms, in percent. Defaults to 12%.

 12.0
Source code in machwave/models/propulsion/motors/liquid.py 
14
15
16
17
18
19
20
21
22
23
24
25
26
27
28
29
30
31
32
33
34
35
36
37
38
39
40
def __init__(
    self,
    propellant: BiliquidPropellant,
    thrust_chamber: LiquidEngineThrustChamber,
    feed_system: FeedSystem,
    oxidizer_tank_cog: float | None = None,
    fuel_tank_cog: float | None = None,
    other_losses: float = 12.0,
) -> None:
    """
    Initialize a liquid rocket engine.

    Args:
        propellant: Bi-liquid propellant properties (oxidizer + fuel).
        thrust_chamber: Thrust chamber assembly (nozzle + combustion chamber + injector).
        feed_system: Propellant feed system (tanks, lines, pumps/pressurization).
        oxidizer_tank_cog: Axial position of the oxidizer tank center (where propellant CoG is),
            measured from the nozzle exit, in meters. If None, uses a default estimate.
        fuel_tank_cog: Axial position of the fuel tank center (where propellant CoG is),
            measured from the nozzle exit, in meters. If None, uses a default estimate.
        other_losses: Additional engine losses not accounted for by specific
            loss mechanisms, in percent. Defaults to 12%.
    """
    super().__init__(propellant, thrust_chamber, other_losses)
    self.feed_system = feed_system
    self.oxidizer_tank_cog = oxidizer_tank_cog
    self.fuel_tank_cog = fuel_tank_cog

get_center_of_gravity(propellant_fraction=0.0)

Calculate the center of gravity of the liquid engine including structural dry mass, oxidizer, and fuel.

The calculation uses a mass-weighted average of: 1. Structural dry mass (thrust chamber, tanks structure, feed lines, etc.) 2. Oxidizer mass (from oxidizer tank) 3. Fuel mass (from fuel tank)

Parameters:

Name Type Description Default
propellant_fraction float

Fraction of propellant consumed (0.0 = full, 1.0 = empty).

 0.0

Returns:

Type Description
NDArray[float64]

Center of gravity in 3D space [x, y, z], in meters.
NDArray[float64]

Origin is at the nozzle exit on the chamber axis.
NDArray[float64]

Positive x-direction points forward (toward bulkhead/away from nozzle exit).

Raises:

Type Description
ValueError

If thrust_chamber.center_of_gravity_coordinate, oxidizer_tank_cog, or fuel_tank_cog is not defined.
Source code in machwave/models/propulsion/motors/liquid.py 
 55
 56
 57
 58
 59
 60
 61
 62
 63
 64
 65
 66
 67
 68
 69
 70
 71
 72
 73
 74
 75
 76
 77
 78
 79
 80
 81
 82
 83
 84
 85
 86
 87
 88
 89
 90
 91
 92
 93
 94
 95
 96
 97
 98
 99
100
101
102
103
104
105
106
107
108
109
110
111
112
113
114
115
116
117
118
119
def get_center_of_gravity(
    self, propellant_fraction: float = 0.0
) -> np.typing.NDArray[np.float64]:
    """
    Calculate the center of gravity of the liquid engine including structural
    dry mass, oxidizer, and fuel.

    The calculation uses a mass-weighted average of:
    1. Structural dry mass (thrust chamber, tanks structure, feed lines, etc.)
    2. Oxidizer mass (from oxidizer tank)
    3. Fuel mass (from fuel tank)

    Args:
        propellant_fraction: Fraction of propellant consumed (0.0 = full, 1.0 = empty).

    Returns:
        Center of gravity in 3D space [x, y, z], in meters.
        Origin is at the nozzle exit on the chamber axis.
        Positive x-direction points forward (toward bulkhead/away from nozzle exit).

    Raises:
        ValueError: If thrust_chamber.center_of_gravity_coordinate, oxidizer_tank_cog,
            or fuel_tank_cog is not defined.
    """
    # Get current propellant masses
    initial_ox_mass = self.feed_system.oxidizer_tank.initial_fluid_mass
    initial_fuel_mass = self.feed_system.fuel_tank.initial_fluid_mass

    ox_mass = initial_ox_mass * (1.0 - propellant_fraction)
    fuel_mass = initial_fuel_mass * (1.0 - propellant_fraction)

    # Structural dry mass (thrust chamber, tank structure, feed lines, etc.)
    dry_mass = self.thrust_chamber.dry_mass

    if self.thrust_chamber.center_of_gravity_coordinate is None:
        raise ValueError(
            "Thrust chamber center of gravity coordinate is not defined."
        )

    dry_cog = self.thrust_chamber.center_of_gravity_coordinate

    # Oxidizer tank CoG
    if self.oxidizer_tank_cog is None:
        raise ValueError("Oxidizer tank center of gravity is not defined.")

    ox_cog = np.array([self.oxidizer_tank_cog, 0.0, 0.0], dtype=np.float64)

    # Fuel tank CoG
    if self.fuel_tank_cog is None:
        raise ValueError("Fuel tank center of gravity is not defined.")

    fuel_cog = np.array([self.fuel_tank_cog, 0.0, 0.0], dtype=np.float64)

    # Calculate total mass and weighted CoG
    total_mass = dry_mass + ox_mass + fuel_mass

    if total_mass <= 0:
        # Fallback if calculation fails
        return dry_cog

    weighted_cog = (
        dry_cog * dry_mass + ox_cog * ox_mass + fuel_cog * fuel_mass
    ) / total_mass

    return weighted_cog.astype(np.float64)

get_thrust_coefficient(chamber_pressure, exit_pressure, external_pressure, expansion_ratio, k_ex, n_cf)

Get thrust coefficient.

Parameters:

Name Type Description Default
chamber_pressure float

Chamber pressure [Pa].

 required
exit_pressure float

Exit pressure [Pa].

 required
external_pressure float

External pressure [Pa].

 required
expansion_ratio float

Expansion ratio.

 required
k_ex float

Two-phase isentropic coefficient.

 required
n_cf float

Thrust coefficient correction factor.

 required

Returns:

Type Description
float

Instantaneous thrust coefficient.
Source code in machwave/models/propulsion/motors/liquid.py 
121
122
123
124
125
126
127
128
129
130
131
132
133
134
135
136
137
138
139
140
141
142
143
144
145
146
147
148
149
150
151
def get_thrust_coefficient(
    self,
    chamber_pressure: float,
    exit_pressure: float,
    external_pressure: float,
    expansion_ratio: float,
    k_ex: float,
    n_cf: float,
) -> float:
    """Get thrust coefficient.

    Args:
        chamber_pressure: Chamber pressure [Pa].
        exit_pressure: Exit pressure [Pa].
        external_pressure: External pressure [Pa].
        expansion_ratio: Expansion ratio.
        k_ex: Two-phase isentropic coefficient.
        n_cf: Thrust coefficient correction factor.

    Returns:
        Instantaneous thrust coefficient.
    """
    cf_ideal = get_ideal_thrust_coefficient(
        chamber_pressure=chamber_pressure,
        exit_pressure=exit_pressure,
        external_pressure=external_pressure,
        expansion_ratio=expansion_ratio,
        k_ex=k_ex,
    )
    n_cf = self.get_thrust_coefficient_correction_factor()
    return cf_ideal * n_cf

Motor

Bases: Generic[P, T], ABC

Abstract rocket motor/engine class. Can be used to model any chemical rocket propulsion system, such as Solid, Hybrid and Liquid.

Source code in machwave/models/propulsion/motors/base.py 
 18
 19
 20
 21
 22
 23
 24
 25
 26
 27
 28
 29
 30
 31
 32
 33
 34
 35
 36
 37
 38
 39
 40
 41
 42
 43
 44
 45
 46
 47
 48
 49
 50
 51
 52
 53
 54
 55
 56
 57
 58
 59
 60
 61
 62
 63
 64
 65
 66
 67
 68
 69
 70
 71
 72
 73
 74
 75
 76
 77
 78
 79
 80
 81
 82
 83
 84
 85
 86
 87
 88
 89
 90
 91
 92
 93
 94
 95
 96
 97
 98
 99
100
101
102
103
104
105
106
107
108
109
110
111
112
113
114
115
116
117
118
119
120
121
122
123
124
125
126
127
128
class Motor(Generic[P, T], ABC):
    """
    Abstract rocket motor/engine class. Can be used to model any chemical
    rocket propulsion system, such as Solid, Hybrid and Liquid.
    """

    def __init__(
        self,
        propellant: P,
        thrust_chamber: T,
        other_losses: float = DEFAULT_OTHER_MOTOR_LOSSES,
    ) -> None:
        """
        Instantiates object attributes common to any motor/engine (Solid,
        Hybrid or Liquid).

        Args:
            propellant: Object representing the propellant used in the motor.
            thrust_chamber: Object representing the thrust chamber of the motor.
            other_losses: Other motor losses, in percent.
        """
        self.propellant = propellant
        self.thrust_chamber = thrust_chamber

        self.other_losses = other_losses

    @abstractmethod
    def get_launch_mass(self) -> float:
        """
        Calculates the total mass of the motor before launch.

        Returns:
            Total mass of the motor before launch, in kg
        """
        pass

    @abstractmethod
    def get_dry_mass(self) -> float:
        """
        Calculates the dry mass of the rocket at any time.

        Returns:
            Dry mass of the rocket, in kg
        """
        pass

    @abstractmethod
    def get_center_of_gravity(self, *args, **kwargs) -> np.typing.NDArray[np.float64]:
        """
        Calculate the center of gravity of the propulsion system.

        The coordinate system origin corresponds to the combustion chamber axis
        at the nozzle exit plane, with positive x pointing toward the bulkhead.

        Returns:
            A 1D array of shape (3,) representing the [x, y, z] coordinates of
            the center of gravity, in meters.
        """
        pass

    def get_thrust_coefficient_correction_factor(self, *args, **kwargs) -> float:
        """
        Calculates the thrust coefficient correction factor. This factor is
        adimensional and should be applied to the ideal thrust coefficient to
        get the real thrust coefficient.

        Returns:
            Thrust coefficient correction factor
        """
        return (
            (100.0 - self.other_losses) / 100.0
        ) * self.propellant.combustion_efficiency

    @abstractmethod
    def get_thrust_coefficient(self, *args, **kwargs) -> float:
        """
        Calculates the thrust coefficient at a particular instant.

        Returns:
            Thrust coefficient
        """
        pass

    @property
    @abstractmethod
    def initial_propellant_mass(self) -> float:
        """
        Returns:
            Initial propellant mass, in kg
        """
        pass

    def get_thrust(self, cf: float, chamber_pressure: float) -> float:
        """
        Calculates the thrust based on instantaneous thrust coefficient and
        chamber pressure.

        Utilized nozzle throat area from the structure and nozzle classes.

        Args:
            cf: Instantaneous thrust coefficient, adimensional
            chamber_pressure: Instantaneous chamber pressure, in Pa

        Returns:
            Instantaneous thrust, in Newtons
        """
        return get_thrust_from_thrust_coefficient(
            cf,
            chamber_pressure,
            self.thrust_chamber.nozzle.get_throat_area(),
        )

initial_propellant_mass abstractmethod property

Returns:

Type Description
float

Initial propellant mass, in kg

__init__(propellant, thrust_chamber, other_losses=DEFAULT_OTHER_MOTOR_LOSSES)

Instantiates object attributes common to any motor/engine (Solid, Hybrid or Liquid).

Parameters:

Name Type Description Default
propellant P

Object representing the propellant used in the motor.

 required
thrust_chamber T

Object representing the thrust chamber of the motor.

 required
other_losses float

Other motor losses, in percent.

 DEFAULT_OTHER_MOTOR_LOSSES
Source code in machwave/models/propulsion/motors/base.py 
24
25
26
27
28
29
30
31
32
33
34
35
36
37
38
39
40
41
42
def __init__(
    self,
    propellant: P,
    thrust_chamber: T,
    other_losses: float = DEFAULT_OTHER_MOTOR_LOSSES,
) -> None:
    """
    Instantiates object attributes common to any motor/engine (Solid,
    Hybrid or Liquid).

    Args:
        propellant: Object representing the propellant used in the motor.
        thrust_chamber: Object representing the thrust chamber of the motor.
        other_losses: Other motor losses, in percent.
    """
    self.propellant = propellant
    self.thrust_chamber = thrust_chamber

    self.other_losses = other_losses

get_center_of_gravity(*args, **kwargs) abstractmethod

Calculate the center of gravity of the propulsion system.

The coordinate system origin corresponds to the combustion chamber axis at the nozzle exit plane, with positive x pointing toward the bulkhead.

Returns:

Type Description
NDArray[float64]

A 1D array of shape (3,) representing the [x, y, z] coordinates of
NDArray[float64]

the center of gravity, in meters.
Source code in machwave/models/propulsion/motors/base.py 
64
65
66
67
68
69
70
71
72
73
74
75
76
@abstractmethod
def get_center_of_gravity(self, *args, **kwargs) -> np.typing.NDArray[np.float64]:
    """
    Calculate the center of gravity of the propulsion system.

    The coordinate system origin corresponds to the combustion chamber axis
    at the nozzle exit plane, with positive x pointing toward the bulkhead.

    Returns:
        A 1D array of shape (3,) representing the [x, y, z] coordinates of
        the center of gravity, in meters.
    """
    pass

get_dry_mass() abstractmethod

Calculates the dry mass of the rocket at any time.

Returns:

Type Description
float

Dry mass of the rocket, in kg
Source code in machwave/models/propulsion/motors/base.py 
54
55
56
57
58
59
60
61
62
@abstractmethod
def get_dry_mass(self) -> float:
    """
    Calculates the dry mass of the rocket at any time.

    Returns:
        Dry mass of the rocket, in kg
    """
    pass

get_launch_mass() abstractmethod

Calculates the total mass of the motor before launch.

Returns:

Type Description
float

Total mass of the motor before launch, in kg
Source code in machwave/models/propulsion/motors/base.py 
44
45
46
47
48
49
50
51
52
@abstractmethod
def get_launch_mass(self) -> float:
    """
    Calculates the total mass of the motor before launch.

    Returns:
        Total mass of the motor before launch, in kg
    """
    pass

get_thrust(cf, chamber_pressure)

Calculates the thrust based on instantaneous thrust coefficient and chamber pressure.

Utilized nozzle throat area from the structure and nozzle classes.

Parameters:

Name Type Description Default
cf float

Instantaneous thrust coefficient, adimensional

 required
chamber_pressure float

Instantaneous chamber pressure, in Pa

 required

Returns:

Type Description
float

Instantaneous thrust, in Newtons
Source code in machwave/models/propulsion/motors/base.py 
110
111
112
113
114
115
116
117
118
119
120
121
122
123
124
125
126
127
128
def get_thrust(self, cf: float, chamber_pressure: float) -> float:
    """
    Calculates the thrust based on instantaneous thrust coefficient and
    chamber pressure.

    Utilized nozzle throat area from the structure and nozzle classes.

    Args:
        cf: Instantaneous thrust coefficient, adimensional
        chamber_pressure: Instantaneous chamber pressure, in Pa

    Returns:
        Instantaneous thrust, in Newtons
    """
    return get_thrust_from_thrust_coefficient(
        cf,
        chamber_pressure,
        self.thrust_chamber.nozzle.get_throat_area(),
    )

get_thrust_coefficient(*args, **kwargs) abstractmethod

Calculates the thrust coefficient at a particular instant.

Returns:

Type Description
float

Thrust coefficient
Source code in machwave/models/propulsion/motors/base.py 
91
92
93
94
95
96
97
98
99
@abstractmethod
def get_thrust_coefficient(self, *args, **kwargs) -> float:
    """
    Calculates the thrust coefficient at a particular instant.

    Returns:
        Thrust coefficient
    """
    pass

get_thrust_coefficient_correction_factor(*args, **kwargs)

Calculates the thrust coefficient correction factor. This factor is adimensional and should be applied to the ideal thrust coefficient to get the real thrust coefficient.

Returns:

Type Description
float

Thrust coefficient correction factor
Source code in machwave/models/propulsion/motors/base.py 
78
79
80
81
82
83
84
85
86
87
88
89
def get_thrust_coefficient_correction_factor(self, *args, **kwargs) -> float:
    """
    Calculates the thrust coefficient correction factor. This factor is
    adimensional and should be applied to the ideal thrust coefficient to
    get the real thrust coefficient.

    Returns:
        Thrust coefficient correction factor
    """
    return (
        (100.0 - self.other_losses) / 100.0
    ) * self.propellant.combustion_efficiency

SolidMotor

Bases: Motor[SolidPropellant, SolidMotorThrustChamber]

Source code in machwave/models/propulsion/motors/solid.py 
 12
 13
 14
 15
 16
 17
 18
 19
 20
 21
 22
 23
 24
 25
 26
 27
 28
 29
 30
 31
 32
 33
 34
 35
 36
 37
 38
 39
 40
 41
 42
 43
 44
 45
 46
 47
 48
 49
 50
 51
 52
 53
 54
 55
 56
 57
 58
 59
 60
 61
 62
 63
 64
 65
 66
 67
 68
 69
 70
 71
 72
 73
 74
 75
 76
 77
 78
 79
 80
 81
 82
 83
 84
 85
 86
 87
 88
 89
 90
 91
 92
 93
 94
 95
 96
 97
 98
 99
100
101
102
103
104
105
106
107
108
109
110
111
112
113
114
115
116
117
118
119
120
121
122
123
124
125
126
127
128
129
130
131
132
133
134
135
136
137
138
139
140
141
142
143
144
145
146
147
148
149
150
151
152
153
154
155
156
157
158
159
160
161
162
163
164
165
166
167
168
169
170
171
172
class SolidMotor(
    motor_base.Motor[
        propellants.SolidPropellant, thrust_chamber.SolidMotorThrustChamber
    ]
):
    def __init__(
        self,
        grain: grain.Grain,
        propellant: propellants.SolidPropellant,
        thrust_chamber: thrust_chamber.SolidMotorThrustChamber,
        other_losses: float = motor_base.DEFAULT_OTHER_MOTOR_LOSSES,
    ) -> None:
        """
        Initialize a solid rocket motor.

        Args:
            grain: Grain geometry configuration.
            propellant: Solid propellant properties.
            thrust_chamber: Thrust chamber model.
            other_losses: Additional motor losses not accounted for by specific
                loss mechanisms (0-1), defaults to 0.12 (12%).
        """
        super().__init__(propellant, thrust_chamber, other_losses)

        self.grain = grain
        self.propellant: propellants.SolidPropellant = propellant
        self.cf_ideal = None  # ideal thrust coefficient
        self.cf_real = None  # real thrust coefficient

    def get_free_chamber_volume(self, propellant_volume: float) -> float:
        """
        Calculates the chamber volume without any propellant.

        Args:
            propellant_volume: Propellant volume, in m^3

        Returns:
            Free chamber volume, in m^3
        """
        return (
            self.thrust_chamber.combustion_chamber.internal_volume - propellant_volume
        )

    @property
    def initial_propellant_mass(self) -> float:
        """
        Returns:
            Initial propellant mass, in kg
        """
        return self.grain.get_propellant_mass(
            web_distance=0, ideal_density=self.propellant.ideal_density
        )

    def get_thrust_coefficient_correction_factor(
        self, n_kin: float, n_bl: float, n_tp: float
    ) -> float:
        """
        Calculates the thrust coefficient correction factor including all
        losses.

        Args:
            n_kin: Kinematic correction factor, adimensional, in percent
            n_bl: Boundary layer correction factor, adimensional, in percent
            n_tp: Two-phase correction factor, adimensional, in percent

        Returns:
            Thrust coefficient correction factor, adimensional
        """
        return (
            (100 - (n_kin + n_bl + n_tp + self.other_losses))
            * losses.get_nozzle_divergent_percentage_loss(
                self.thrust_chamber.nozzle.throat_diameter
            )
            / 100
            * self.propellant.combustion_efficiency
        )

    def get_thrust_coefficient(
        self,
        chamber_pressure: float,
        exit_pressure: float,
        external_pressure: float,
        expansion_ratio: float,
        k_ex: float,
        n_cf: float,
    ) -> float:
        """
        Args:
            chamber_pressure: Chamber pressure, in Pa
            exit_pressure: Exit pressure, in Pa
            external_pressure: External pressure, in Pa
            expansion_ratio: Expansion ratio, adimensional
            k_ex: Two-phase isentropic coefficient, adimensional
            n_cf: Thrust coefficient correction factor, adimensional

        Returns:
            Instanteneous thrust coefficient, adimensional
        """
        self.cf_ideal = nozzle.get_ideal_thrust_coefficient(
            chamber_pressure,
            exit_pressure,
            external_pressure,
            expansion_ratio,
            k_ex,
        )
        self.cf_real = nozzle.apply_thrust_coefficient_correction(self.cf_ideal, n_cf)
        return self.cf_real

    def get_launch_mass(self) -> float:
        return self.thrust_chamber.dry_mass + self.initial_propellant_mass

    def get_dry_mass(self) -> float:
        return self.thrust_chamber.dry_mass

    def get_center_of_gravity(
        self, web_distance: float = 0.0
    ) -> np.typing.NDArray[np.float64]:
        """
        Calculates the center of gravity of the solid motor including
        propellant grain (wet mass) and dry mass.

        The calculation uses a mass-weighted average of:
        1. Propellant grain CoG;
        2. Thrust chamber dry mass CoG, considered constant.

        Args:
            web_distance: Web distance traveled [m].
                Defaults to ignition state.

        Returns:
            Center of gravity in 3D space (x, y, z) [m].

        Raises:
            ValueError: If thrust chamber dry mass CoG is not defined or if
                total mass is less than or equal to zero.
        """
        grain_cog_port = self.grain.get_center_of_gravity(web_distance=web_distance)
        propellant_mass = self.grain.get_propellant_mass(
            web_distance=web_distance, ideal_density=self.propellant.ideal_density
        )

        dry_mass = self.thrust_chamber.dry_mass
        dry_mass_cog = self.thrust_chamber.center_of_gravity_coordinate
        nozzle_exit_to_port = self.thrust_chamber.nozzle_exit_to_grain_port_distance

        # Transform grain CoG from port origin to nozzle exit origin
        grain_cog = grain_cog_port.copy()
        grain_cog[0] = nozzle_exit_to_port + grain_cog_port[0]

        if dry_mass_cog is None:
            raise ValueError("Dry mass center of gravity coordinate is not defined.")

        total_mass = propellant_mass + dry_mass
        if total_mass <= 0:
            raise ValueError("Total mass must be greater than zero to calculate CoG.")

        weighted_cog = (
            grain_cog * propellant_mass + dry_mass_cog * dry_mass
        ) / total_mass

        return weighted_cog.astype(np.float64)

initial_propellant_mass property

Returns:

Type Description
float

Initial propellant mass, in kg

__init__(grain, propellant, thrust_chamber, other_losses=motor_base.DEFAULT_OTHER_MOTOR_LOSSES)

Initialize a solid rocket motor.

Parameters:

Name Type Description Default
grain Grain

Grain geometry configuration.

 required
propellant SolidPropellant

Solid propellant properties.

 required
thrust_chamber SolidMotorThrustChamber

Thrust chamber model.

 required
other_losses float

Additional motor losses not accounted for by specific loss mechanisms (0-1), defaults to 0.12 (12%).

 DEFAULT_OTHER_MOTOR_LOSSES
Source code in machwave/models/propulsion/motors/solid.py 
17
18
19
20
21
22
23
24
25
26
27
28
29
30
31
32
33
34
35
36
37
38
39
def __init__(
    self,
    grain: grain.Grain,
    propellant: propellants.SolidPropellant,
    thrust_chamber: thrust_chamber.SolidMotorThrustChamber,
    other_losses: float = motor_base.DEFAULT_OTHER_MOTOR_LOSSES,
) -> None:
    """
    Initialize a solid rocket motor.

    Args:
        grain: Grain geometry configuration.
        propellant: Solid propellant properties.
        thrust_chamber: Thrust chamber model.
        other_losses: Additional motor losses not accounted for by specific
            loss mechanisms (0-1), defaults to 0.12 (12%).
    """
    super().__init__(propellant, thrust_chamber, other_losses)

    self.grain = grain
    self.propellant: propellants.SolidPropellant = propellant
    self.cf_ideal = None  # ideal thrust coefficient
    self.cf_real = None  # real thrust coefficient

get_center_of_gravity(web_distance=0.0)

Calculates the center of gravity of the solid motor including propellant grain (wet mass) and dry mass.

The calculation uses a mass-weighted average of: 1. Propellant grain CoG; 2. Thrust chamber dry mass CoG, considered constant.

Parameters:

Name Type Description Default
web_distance float

Web distance traveled [m]. Defaults to ignition state.

 0.0

Returns:

Type Description
NDArray[float64]

Center of gravity in 3D space (x, y, z) [m].

Raises:

Type Description
ValueError

If thrust chamber dry mass CoG is not defined or if total mass is less than or equal to zero.
Source code in machwave/models/propulsion/motors/solid.py 
126
127
128
129
130
131
132
133
134
135
136
137
138
139
140
141
142
143
144
145
146
147
148
149
150
151
152
153
154
155
156
157
158
159
160
161
162
163
164
165
166
167
168
169
170
171
172
def get_center_of_gravity(
    self, web_distance: float = 0.0
) -> np.typing.NDArray[np.float64]:
    """
    Calculates the center of gravity of the solid motor including
    propellant grain (wet mass) and dry mass.

    The calculation uses a mass-weighted average of:
    1. Propellant grain CoG;
    2. Thrust chamber dry mass CoG, considered constant.

    Args:
        web_distance: Web distance traveled [m].
            Defaults to ignition state.

    Returns:
        Center of gravity in 3D space (x, y, z) [m].

    Raises:
        ValueError: If thrust chamber dry mass CoG is not defined or if
            total mass is less than or equal to zero.
    """
    grain_cog_port = self.grain.get_center_of_gravity(web_distance=web_distance)
    propellant_mass = self.grain.get_propellant_mass(
        web_distance=web_distance, ideal_density=self.propellant.ideal_density
    )

    dry_mass = self.thrust_chamber.dry_mass
    dry_mass_cog = self.thrust_chamber.center_of_gravity_coordinate
    nozzle_exit_to_port = self.thrust_chamber.nozzle_exit_to_grain_port_distance

    # Transform grain CoG from port origin to nozzle exit origin
    grain_cog = grain_cog_port.copy()
    grain_cog[0] = nozzle_exit_to_port + grain_cog_port[0]

    if dry_mass_cog is None:
        raise ValueError("Dry mass center of gravity coordinate is not defined.")

    total_mass = propellant_mass + dry_mass
    if total_mass <= 0:
        raise ValueError("Total mass must be greater than zero to calculate CoG.")

    weighted_cog = (
        grain_cog * propellant_mass + dry_mass_cog * dry_mass
    ) / total_mass

    return weighted_cog.astype(np.float64)

get_free_chamber_volume(propellant_volume)

Calculates the chamber volume without any propellant.

Parameters:

Name Type Description Default
propellant_volume float

Propellant volume, in m^3

 required

Returns:

Type Description
float

Free chamber volume, in m^3
Source code in machwave/models/propulsion/motors/solid.py 
41
42
43
44
45
46
47
48
49
50
51
52
53
def get_free_chamber_volume(self, propellant_volume: float) -> float:
    """
    Calculates the chamber volume without any propellant.

    Args:
        propellant_volume: Propellant volume, in m^3

    Returns:
        Free chamber volume, in m^3
    """
    return (
        self.thrust_chamber.combustion_chamber.internal_volume - propellant_volume
    )

get_thrust_coefficient(chamber_pressure, exit_pressure, external_pressure, expansion_ratio, k_ex, n_cf)

Parameters:

Name Type Description Default
chamber_pressure float

Chamber pressure, in Pa

 required
exit_pressure float

Exit pressure, in Pa

 required
external_pressure float

External pressure, in Pa

 required
expansion_ratio float

Expansion ratio, adimensional

 required
k_ex float

Two-phase isentropic coefficient, adimensional

 required
n_cf float

Thrust coefficient correction factor, adimensional

 required

Returns:

Type Description
float

Instanteneous thrust coefficient, adimensional
Source code in machwave/models/propulsion/motors/solid.py 
 89
 90
 91
 92
 93
 94
 95
 96
 97
 98
 99
100
101
102
103
104
105
106
107
108
109
110
111
112
113
114
115
116
117
118
def get_thrust_coefficient(
    self,
    chamber_pressure: float,
    exit_pressure: float,
    external_pressure: float,
    expansion_ratio: float,
    k_ex: float,
    n_cf: float,
) -> float:
    """
    Args:
        chamber_pressure: Chamber pressure, in Pa
        exit_pressure: Exit pressure, in Pa
        external_pressure: External pressure, in Pa
        expansion_ratio: Expansion ratio, adimensional
        k_ex: Two-phase isentropic coefficient, adimensional
        n_cf: Thrust coefficient correction factor, adimensional

    Returns:
        Instanteneous thrust coefficient, adimensional
    """
    self.cf_ideal = nozzle.get_ideal_thrust_coefficient(
        chamber_pressure,
        exit_pressure,
        external_pressure,
        expansion_ratio,
        k_ex,
    )
    self.cf_real = nozzle.apply_thrust_coefficient_correction(self.cf_ideal, n_cf)
    return self.cf_real

get_thrust_coefficient_correction_factor(n_kin, n_bl, n_tp)

Calculates the thrust coefficient correction factor including all losses.

Parameters:

Name Type Description Default
n_kin float

Kinematic correction factor, adimensional, in percent

 required
n_bl float

Boundary layer correction factor, adimensional, in percent

 required
n_tp float

Two-phase correction factor, adimensional, in percent

 required

Returns:

Type Description
float

Thrust coefficient correction factor, adimensional
Source code in machwave/models/propulsion/motors/solid.py 
65
66
67
68
69
70
71
72
73
74
75
76
77
78
79
80
81
82
83
84
85
86
87
def get_thrust_coefficient_correction_factor(
    self, n_kin: float, n_bl: float, n_tp: float
) -> float:
    """
    Calculates the thrust coefficient correction factor including all
    losses.

    Args:
        n_kin: Kinematic correction factor, adimensional, in percent
        n_bl: Boundary layer correction factor, adimensional, in percent
        n_tp: Two-phase correction factor, adimensional, in percent

    Returns:
        Thrust coefficient correction factor, adimensional
    """
    return (
        (100 - (n_kin + n_bl + n_tp + self.other_losses))
        * losses.get_nozzle_divergent_percentage_loss(
            self.thrust_chamber.nozzle.throat_diameter
        )
        / 100
        * self.propellant.combustion_efficiency
    )