aerosandbox.library.propulsion_small_solid_rocket
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Module Contents#
Functions#
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Burn rate vs oxamide content model. |
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Characteristic velocity vs. oxamide content model. |
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Minimum pressure for stable combustion vs. oxamide content model. |
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Ratio of specific heats vs. oxamide content model. |
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Find the nozzle expansion ratio from the chamber and exit pressures. |
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Nozzle thrust coefficient, \(C_F\). |
Attributes#
- aerosandbox.library.propulsion_small_solid_rocket.OUT_OF_RANGE_ERROR_STRING = '{:.3f} is outside the model valid range of {:.3f} <= w_om <= {:.3f}'[source]#
- aerosandbox.library.propulsion_small_solid_rocket.burn_rate_coefficient(oxamide_fraction)[source]#
Burn rate vs oxamide content model. Valid from 0% to 15% oxamide. # TODO IMPLEMENT THIS
- Returns:
- propellant burn rate coefficient
[units: pascal**(-n) meter second**-1].
- Return type:
a
- aerosandbox.library.propulsion_small_solid_rocket.c_star(oxamide_fraction)[source]#
Characteristic velocity vs. oxamide content model. Valid from 0% to 20% oxamide. # TODO IMPLEMENT THIS
- Returns:
ideal characteristic velocity [units: meter second**-1].
- Return type:
c_star
- aerosandbox.library.propulsion_small_solid_rocket.dubious_min_combustion_pressure(oxamide_fraction)[source]#
Minimum pressure for stable combustion vs. oxamide content model.
Note: this model is of DUBIOUS accuracy. Don’t trust it.
- aerosandbox.library.propulsion_small_solid_rocket.gamma(oxamide_fraction)[source]#
Ratio of specific heats vs. oxamide content model.
- Returns:
Exhaust gas ratio of specific heats [units: dimensionless].
- Return type:
gamma
- aerosandbox.library.propulsion_small_solid_rocket.expansion_ratio_from_pressure(chamber_pressure, exit_pressure, gamma, oxamide_fraction)[source]#
Find the nozzle expansion ratio from the chamber and exit pressures.
See expansion-ratio-tutorial-label for a physical description of the expansion ratio.
Reference: Rocket Propulsion Elements, 8th Edition, Equation 3-25
- Parameters:
chamber_pressure (scalar) – Nozzle stagnation chamber pressure [units: pascal].
exit_pressure (scalar) – Nozzle exit static pressure [units: pascal].
gamma (scalar) – Exhaust gas ratio of specific heats [units: dimensionless].
- Returns:
Expansion ratio \(\epsilon = A_e / A_t\) [units: dimensionless]
- Return type:
scalar
- aerosandbox.library.propulsion_small_solid_rocket.thrust_coefficient(chamber_pressure, exit_pressure, gamma, p_a=None, er=None)[source]#
Nozzle thrust coefficient, \(C_F\).
The thrust coefficient is a figure of merit for the nozzle expansion process. See thrust-coefficient-label for a description of the physical meaning of the thrust coefficient.
Reference: Equation 1-33a in Huzel and Huang.
- Parameters:
chamber_pressure (scalar) – Nozzle stagnation chamber pressure [units: pascal].
exit_pressure (scalar) – Nozzle exit static pressure [units: pascal].
gamma (scalar) – Exhaust gas ratio of specific heats [units: dimensionless].
p_a (scalar, optional) – Ambient pressure [units: pascal]. If None, then p_a = exit_pressure.
er (scalar, optional) – Nozzle area expansion ratio [units: dimensionless]. If None, then p_a = exit_pressure.
- Returns:
The thrust coefficient, \(C_F\) [units: dimensionless].
- Return type:
scalar