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I have been working on a rocket engine and I have cad models and everything and I am getting ready to build it. Then, I was running over the numbers one more time and I realized that the thrust I was getting was completely different from what I should have been getting.
The big problem is the NASA equations. It uses the variables Ttotal and Ptotal. NASA states, "pt is the total pressure in the combustion chamber, Tt is the total temperature in the combustion chamber." However, when I plug in 300 psi and 5742 degrees fahrenheit and the exhaust values for each variable as given by NASA's isentropic equations, I get a thrust of 735 pounds, whereas the first time I ran the numbers I got 32 pounds of thrust.
Here is my data, most of it was from a formula. The top ten rows are all givens. I got that information from https://risacher.org/rocket/eqns.html as I am building a gasoline-GOX engine.
What it is Value Unit
Combustion pressure (Pc) 300 PSI
Mixture Ratio 2.5 n/a
Flame Temp 5742 Fahrenheit
ISP 261 Seconds
GOX Density 0.083 lb/ft^3
Gas Density 44.5 lb/ft^3
Nozzle Throat Cross Sectional Area 0.00173611 ft^2
Gas constant (R) 65 ftlb/lb
Gamma (heat ratio) 1.2
gc (gravitation) 32 ft/second/second
Tt (temperature of gasses at nozzle throat) 5637.618 Fahrenheit
Tc (temperature of cumbustion flame) 6202 Rankine
Pt (gas pressure at nozzle throat) 169.2 PSI
Me (mach number of gasses at exit) 2.55
Area Exhaust/Area throat 3.65
Temperature Exhuast/Temperature combustion 0.606
Thrust 31.84104294 Pounds
Mass flow rate 0.0009229501076 pounds per second
Exhaust velocity 1380.669321 ft/second
Temperature Exhuast 3479.652 Fahrenheit
Exhaust pressure 24.4297 psiAs you might imagine, I would really appreciate someone going over and double checking my numbers. Thank you so much!
The big problem is the NASA equations. It uses the variables Ttotal and Ptotal. NASA states, "pt is the total pressure in the combustion chamber, Tt is the total temperature in the combustion chamber." However, when I plug in 300 psi and 5742 degrees fahrenheit and the exhaust values for each variable as given by NASA's isentropic equations, I get a thrust of 735 pounds, whereas the first time I ran the numbers I got 32 pounds of thrust.
Here is my data, most of it was from a formula. The top ten rows are all givens. I got that information from https://risacher.org/rocket/eqns.html as I am building a gasoline-GOX engine.
What it is Value Unit
Combustion pressure (Pc) 300 PSI
Mixture Ratio 2.5 n/a
Flame Temp 5742 Fahrenheit
ISP 261 Seconds
GOX Density 0.083 lb/ft^3
Gas Density 44.5 lb/ft^3
Nozzle Throat Cross Sectional Area 0.00173611 ft^2
Gas constant (R) 65 ftlb/lb
Gamma (heat ratio) 1.2
gc (gravitation) 32 ft/second/second
Tt (temperature of gasses at nozzle throat) 5637.618 Fahrenheit
Tc (temperature of cumbustion flame) 6202 Rankine
Pt (gas pressure at nozzle throat) 169.2 PSI
Me (mach number of gasses at exit) 2.55
Area Exhaust/Area throat 3.65
Temperature Exhuast/Temperature combustion 0.606
Thrust 31.84104294 Pounds
Mass flow rate 0.0009229501076 pounds per second
Exhaust velocity 1380.669321 ft/second
Temperature Exhuast 3479.652 Fahrenheit
Exhaust pressure 24.4297 psiAs you might imagine, I would really appreciate someone going over and double checking my numbers. Thank you so much!
