(bug) Are thrust values too large?

[runphilrun]

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; 

[runphilrun]

Thrust is seriously wrong, which throws off isp.
Mass flow rate seems right, ratios seem right...

[runphilrun]

Related to #28, after patching, outputs are

Exit pressure     9475.40195891 [Pa]
Thrust      11.60342097 [N]
Exit Mach        2.20004582 [-]
A/At         2.00505600 [-]
T/Tc         0.50811967 [-]
P/Pc         0.09351495 [-]
v/at         1.71792956 [-]

[runphilrun]

ignore thrust values until exit pressure and mass flow rate are verified