Public Member Functions
float	R_specific (self)
float	A_throat (self)
float	A_exit (self)
float	expansion_ratio (self)
float	A_chamber (self)
float	moment_arm (self)

Public Attributes
int	A_throat = 0:

Static Public Attributes
float	throat_diameter = 0.0
float	exit_diameter = 0.0
float	bell_percent = 0.80
str	nozzle_method = "top"
float	inlet_pressure = 0.0
float	inlet_temperature = 293.0
float	gamma = CO2_GAMMA
float	molar_mass = CO2_MOLAR_MASS
float	nozzle_efficiency = 0.95
int	num_thrusters = 4
float	ambient_pressure = DEFAULT_AMBIENT_PA
float	rocket_length = 5.9182
float	rocket_radius = 0.1015
float	rocket_mass = 63.5029
float	cg_from_tail = 2.6162
float	chamber_diameter = 0.010
float	chamber_length = 0.015
float	volume_chamber = 0.0
float	Lstar = 0.0

Protected Member Functions
float	_effective_volume_chamber (self)
bool	_use_solved_plot (self)

Detailed Description

All inputs for one cold-gas thruster analysis.

Nozzle geometry
---------------
throat_diameter : float   — Throat diameter [m]
exit_diameter   : float   — Exit diameter [m]
bell_percent    : float   — Rao bell length fraction (default 0.80)
nozzle_method   : str     — 'garcia' or 'top' (default 'garcia')

Gas state
---------
inlet_pressure    : float — CO₂ stagnation (total) pressure [Pa]
inlet_temperature : float — CO₂ stagnation temperature [K]  (default 293)
gamma             : float — Specific heat ratio  (default 1.289 for CO₂)
molar_mass        : float — Molar mass [kg/mol]  (default 0.04401)
nozzle_efficiency : float — Velocity efficiency factor (default 0.95)

Thruster arrangement
--------------------
num_thrusters    : int   — Number of thrusters (default 4)
ambient_pressure : float — Back-pressure [Pa]  (default 12 psi = 82 737 Pa)

Rocket geometry (for torque calculation)
----------------------------------------
rocket_length : float — Total length tail→nose [m]  (default 7.5)
rocket_radius : float — Body tube radius [m]        (default 0.1015)
rocket_mass   : float — Total mass [kg]             (default 130.0)
cg_from_tail  : float — CG position from tail [m]   (default 3.0)

Definition at line 50 of file cold_gas_thruster.py.

Member Function Documentation

◆ _effective_volume_chamber()

float payload.cold_gas_thruster.ColdGasThrusterConfig._effective_volume_chamber ( self )

protected

Chamber volume [m³].  If not explicitly set, estimate from A_chamber × length.

Definition at line 142 of file cold_gas_thruster.py.

References A_chamber(), chamber_length, and volume_chamber.

◆ _use_solved_plot()

bool payload.cold_gas_thruster.ColdGasThrusterConfig._use_solved_plot ( self )

protected

True when the caller has provided enough geometry for solved_chamber_plot.

Definition at line 149 of file cold_gas_thruster.py.

References volume_chamber.

◆ A_chamber()

float payload.cold_gas_thruster.ColdGasThrusterConfig.A_chamber ( self )

Chamber area [m²].

Definition at line 137 of file cold_gas_thruster.py.

References engine.pipeline.config_schemas.ChamberGeometryConfig.chamber_diameter, and chamber_diameter.

Referenced by _effective_volume_chamber().

◆ A_exit()

float payload.cold_gas_thruster.ColdGasThrusterConfig.A_exit ( self )

Exit area [m²].

Definition at line 125 of file cold_gas_thruster.py.

References engine.pipeline.config_schemas.ChamberConfig.exit_diameter, engine.pipeline.config_schemas.ChamberGeometryConfig.exit_diameter, engine.pipeline.config_schemas.NozzleConfig.exit_diameter, and exit_diameter.

◆ A_throat()

float payload.cold_gas_thruster.ColdGasThrusterConfig.A_throat ( self )

Throat area [m²].

Definition at line 120 of file cold_gas_thruster.py.

References engine.pipeline.thermal.graphite_geometry.GraphiteInsertSizing.throat_diameter, and throat_diameter.

◆ expansion_ratio()

float payload.cold_gas_thruster.ColdGasThrusterConfig.expansion_ratio ( self )

Area ratio ε = A_exit / A_throat.

Definition at line 130 of file cold_gas_thruster.py.

◆ moment_arm()

float payload.cold_gas_thruster.ColdGasThrusterConfig.moment_arm ( self )

Distance from nosecone (thruster location) to CG [m].

Definition at line 154 of file cold_gas_thruster.py.

References cg_from_tail, backend.routers.flight.RocketConfig.rocket_length, engine.pipeline.config_schemas.RocketConfig.rocket_length, and rocket_length.

◆ R_specific()

float payload.cold_gas_thruster.ColdGasThrusterConfig.R_specific ( self )

Specific gas constant R = R_universal / molar_mass  [J/(kg·K)].

Definition at line 115 of file cold_gas_thruster.py.

References molar_mass.

Member Data Documentation

◆ A_throat

int payload.cold_gas_thruster.ColdGasThrusterConfig.A_throat = 0:

Definition at line 132 of file cold_gas_thruster.py.

◆ ambient_pressure

float payload.cold_gas_thruster.ColdGasThrusterConfig.ambient_pressure = DEFAULT_AMBIENT_PA

static

Definition at line 96 of file cold_gas_thruster.py.

Referenced by engine.layerx.card._ExpansionScaling.evaluate(), and feedtwin.engine.card.EngineCard.install().

◆ bell_percent

float payload.cold_gas_thruster.ColdGasThrusterConfig.bell_percent = 0.80

static

Definition at line 84 of file cold_gas_thruster.py.

◆ cg_from_tail

float payload.cold_gas_thruster.ColdGasThrusterConfig.cg_from_tail = 2.6162

static

Definition at line 102 of file cold_gas_thruster.py.

Referenced by moment_arm().

◆ chamber_diameter

float payload.cold_gas_thruster.ColdGasThrusterConfig.chamber_diameter = 0.010

static

Definition at line 105 of file cold_gas_thruster.py.

Referenced by A_chamber().

◆ chamber_length

float payload.cold_gas_thruster.ColdGasThrusterConfig.chamber_length = 0.015

static

Definition at line 106 of file cold_gas_thruster.py.

Referenced by _effective_volume_chamber().

◆ exit_diameter

float payload.cold_gas_thruster.ColdGasThrusterConfig.exit_diameter = 0.0

static

Definition at line 83 of file cold_gas_thruster.py.

Referenced by A_exit().

◆ gamma

float payload.cold_gas_thruster.ColdGasThrusterConfig.gamma = CO2_GAMMA

static

Definition at line 90 of file cold_gas_thruster.py.

Referenced by test_combustion_stability_audit.TestNyquistAllCrossings.case(), test_combustion_stability_audit.TestNyquistAllCrossings.L_indep(), and engine.pipeline.stability.chug.ChugChamber.theta_c().

◆ inlet_pressure

float payload.cold_gas_thruster.ColdGasThrusterConfig.inlet_pressure = 0.0

static

Definition at line 88 of file cold_gas_thruster.py.

◆ inlet_temperature

float payload.cold_gas_thruster.ColdGasThrusterConfig.inlet_temperature = 293.0

static

Definition at line 89 of file cold_gas_thruster.py.

◆ Lstar

float payload.cold_gas_thruster.ColdGasThrusterConfig.Lstar = 0.0

static

Definition at line 111 of file cold_gas_thruster.py.

Referenced by test_combustion_stability_audit.TestNyquistAllCrossings.case(), test_combustion_stability_audit.TestNyquistAllCrossings.L_indep(), and engine.pipeline.stability.chug.ChugChamber.theta_c().

◆ molar_mass

float payload.cold_gas_thruster.ColdGasThrusterConfig.molar_mass = CO2_MOLAR_MASS

static

Definition at line 91 of file cold_gas_thruster.py.

Referenced by R_specific().

◆ nozzle_efficiency

float payload.cold_gas_thruster.ColdGasThrusterConfig.nozzle_efficiency = 0.95

static

Definition at line 92 of file cold_gas_thruster.py.

◆ nozzle_method

str payload.cold_gas_thruster.ColdGasThrusterConfig.nozzle_method = "top"

static

Definition at line 85 of file cold_gas_thruster.py.

◆ num_thrusters

int payload.cold_gas_thruster.ColdGasThrusterConfig.num_thrusters = 4

static

Definition at line 95 of file cold_gas_thruster.py.

◆ rocket_length

float payload.cold_gas_thruster.ColdGasThrusterConfig.rocket_length = 5.9182

static

Definition at line 99 of file cold_gas_thruster.py.

Referenced by moment_arm().

◆ rocket_mass

float payload.cold_gas_thruster.ColdGasThrusterConfig.rocket_mass = 63.5029

static

Definition at line 101 of file cold_gas_thruster.py.

◆ rocket_radius

float payload.cold_gas_thruster.ColdGasThrusterConfig.rocket_radius = 0.1015

static

Definition at line 100 of file cold_gas_thruster.py.

◆ throat_diameter

float payload.cold_gas_thruster.ColdGasThrusterConfig.throat_diameter = 0.0

static

Definition at line 82 of file cold_gas_thruster.py.

Referenced by A_throat().

◆ volume_chamber

float payload.cold_gas_thruster.ColdGasThrusterConfig.volume_chamber = 0.0

static

Definition at line 110 of file cold_gas_thruster.py.

Referenced by _effective_volume_chamber(), and _use_solved_plot().

The documentation for this class was generated from the following file:

EngineDesign/payload/cold_gas_thruster.py

Public Member Functions

Public Attributes

Static Public Attributes

Protected Member Functions

Detailed Description

Member Function Documentation

◆ _effective_volume_chamber()

◆ _use_solved_plot()

◆ A_chamber()

◆ A_exit()

◆ A_throat()

◆ expansion_ratio()

◆ moment_arm()

◆ R_specific()

Member Data Documentation

◆ A_throat

◆ ambient_pressure

◆ bell_percent

◆ cg_from_tail

◆ chamber_diameter

◆ chamber_length

◆ exit_diameter

◆ gamma

◆ inlet_pressure

◆ inlet_temperature

◆ Lstar

◆ molar_mass

◆ nozzle_efficiency

◆ nozzle_method

◆ num_thrusters

◆ rocket_length

◆ rocket_mass

◆ rocket_radius

◆ throat_diameter

◆ volume_chamber