Rocket Nozzle Design
1.2
γ
(Ratio of Specific Heats)
1
2
361
J/kg.K
R
100
500
(Specific Gas Constant)
_
1.268
MPa
Chamber Pressure
1
25
2103
°K
Chamber Temperature
1500
4000
0.002
m
Nozzle Throat
Diameter
0.001
1.0
90
Combustion
85
90
%
Nozzle
98
Efficiency
1.01325
kPa
0
Exhaust Pressure
& Equivalent Altitude
0km
(sea lvl)
100km
(edge of space)
1.01325
kPa
0
Ambient Pressure
& Equivalent Altitude
0km
(sea lvl)
100km
(edge of space)
Engine Preset
Nozzle Length
Result
m
Exit Diameter
Result
m
This design code calculates ideal rocket nozzle geometry for given input parameters describing
the combustion chamber environment and
desired exhaust pressure under
isentropic conditions.