For the following inputs:
specific_heat_ratio = 1.4;
molecular_weight = .0289645; %kg/mol
chamber_temperature = 273; %K
chamber_pressure = 101325; %Pa
exit_radius = .014160; %m
throat_radius = .01; %m
conical_half_angle = 15; %deg
I'm getting these outputs:
Thrust 93.044764 [N]
Mass flow rate 0.077861 [kg/s]
Exhaust velocity 519.450384 [m/s]
Exit pressure 83503.927140 [Pa]
Exit area 0.000630 [m^2]
The equation for thrust is thrust = mass_flowrate*exit_velocity + (exit_pressure-ambient_pressure)*exit_area.
Ambient pressure is assumed to be 0 (vacuum). Exit pressure and exit velocity are verified to be correct with NASA 1135. Exit area is simple geometry and verified correct.
That leaves mass flow rate. Is this reasonable?
The derivation of mass flow rate is as follows:
throat_density = throat_pressure/(specific_gas_constant*throat_temperature);
throat_flowrate = sqrt(specific_heat_ratio*specific_gas_constant*throat_temperature);
mass_flowrate = throat_density*throat_flowrate*throat_area;
For the following inputs:
I'm getting these outputs:
The equation for thrust is
thrust = mass_flowrate*exit_velocity + (exit_pressure-ambient_pressure)*exit_area.Ambient pressure is assumed to be 0 (vacuum). Exit pressure and exit velocity are verified to be correct with NASA 1135. Exit area is simple geometry and verified correct.
That leaves mass flow rate. Is this reasonable?
The derivation of mass flow rate is as follows: