Fuel flow is the combined flow of fuel, oxidizer and propellant typically expressed in kilograms per second. A rocket going to low earth orbit will usually burn for about two hundred and thirty seconds and since most of the gross mass of the rocket is wet, the fuel flow is roughly equal to the mass divided by two hundred. A Saturn V class craft will have a fuel flow of about ten thousand kilograms per second at liftoff.

In a rocket, the fuel flow will often be throttled back towards the end of the burn to limit the peak acceleration; but, in these scripts for the sake of simplicity, the fuel flow is assumed to be constant. Given exhaust velocity, liftoff acceleration and mass the fuel flow can be calculated which is in turn used to calculate exit area and pump power.

fuel flow = mass * liftoff acceleration / exhaust velocity

k = 1.21
throat temperature = combustion temperature / ( 0.5 + 0.5 * k )
throat pressure = chamber pressure * pow( 0.5 + 0.5 * k, - k / ( k - 1.0 ) )
throat area = sqrt( 8,314 J * K / kmol * throat temperature / exhaust molecular / k ) / throat pressure
exit area = fuel flow * throat area * area ratio

propellant energy = propellant usage * pump efficiency * working temperature * 8,314 J * K / kmol / 0.2
combustion exhaust = combustion ratio / exhaust molecular
uncombustion = 1.0 - combustion ratio
pump energy = propellant energy * propellant mass ratio / propellant molecular + propellant energy * reactant propellant usage * ( fuel mass ratio * ( uncombustion / fuel molecular + combustion exhaust ) + oxidizer mass ratio * ( uncombustion / oxidizer molecular + combustion exhaust ) )
pump power = pump energy * fuel flow
 
 

This is used in pumped rocket and rocket cost.
 
 

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