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Isentropic compressible flow relations.

• [1] - Heister et al., Rocket Propulsion (https://doi.org/10.1017/9781108381376)

Functions

A_At(M, gamma)	Ratio of local flow area to the throat area, for a given Mach number [1].
M_from_A_subsonic(A, A_t, gamma)	Get the Mach number from the local flow area, assuming subsonic flow.
M_from_A_supersonic(A, A_t, gamma)	Get the Mach number from the local flow area, assuming supersonic flow.
M_from_p(p, p0, gamma)	Mach number from static pressure and stagnation pressure.
T(T0, M, gamma)	Get local temperature from the Mach number and stagnation temperature.
T0(T, M, gamma)	Get the stangation temperature from the static temperature and Mach number [1]
Tr(T, M, gamma, r)	Get the recovery temperature, which is the temperature you should use for convective heat transfer (i.e.
get_choked_mdot(perfect_gas, …)	Get the mass flow rate through a choked nozzle.
get_exit_area(perfect_gas, …)	Get the nozzle exit area, given the gas properties and combustion chamber conditions.
get_throat_area(perfect_gas, …)	Get the nozzle throat area, given the gas properties and combustion chamber conditions.
m_bar(M, gamma)	Non-dimensional mass flow rate, defined as m_bar = mdot * sqrt(cp*T0)/(A*p0).
p(p0, M, gamma)	Get local pressure from the Mach number and stagnation pressure.
p0(p, M, gamma)	Get stagnation pressure from static pressure and Mach number [1]