aerosandbox.library.propulsion_small_solid_rocket#

Module Contents#

Functions#

burn_rate_coefficient(oxamide_fraction)

Burn rate vs oxamide content model.

c_star(oxamide_fraction)

Characteristic velocity vs. oxamide content model.

dubious_min_combustion_pressure(oxamide_fraction)

Minimum pressure for stable combustion vs. oxamide content model.

gamma(oxamide_fraction)

Ratio of specific heats vs. oxamide content model.

expansion_ratio_from_pressure(chamber_pressure, ...)

Find the nozzle expansion ratio from the chamber and exit pressures.

thrust_coefficient(chamber_pressure, exit_pressure, gamma)

Nozzle thrust coefficient, \(C_F\).

Attributes#

n

lamb

a_0

strand_reduction_factor

zeta_c_star

chamber_pressure_max

W_OM_VALID_RANGE

OUT_OF_RANGE_ERROR_STRING

oxamides

aerosandbox.library.propulsion_small_solid_rocket.n = 0.402[source]#
aerosandbox.library.propulsion_small_solid_rocket.lamb = 6.2[source]#
aerosandbox.library.propulsion_small_solid_rocket.a_0[source]#
aerosandbox.library.propulsion_small_solid_rocket.strand_reduction_factor[source]#
aerosandbox.library.propulsion_small_solid_rocket.zeta_c_star = 0.9[source]#
aerosandbox.library.propulsion_small_solid_rocket.chamber_pressure_max = 2000000.0[source]#
aerosandbox.library.propulsion_small_solid_rocket.W_OM_VALID_RANGE = (0, 0.22)[source]#
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

aerosandbox.library.propulsion_small_solid_rocket.oxamides[source]#