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)
This code calculates ideal rocket nozzle geometry
for input parameters describing the combustion chamber environment and desired exhaust characteristics under isentropic conditions.
Vehicle structures and Delta-V
budgets are configured
and mapped below.
Engine Preset
Nozzle Length
Result
m
Exit Diameter
Result
m
Expansion Ratio
Result
NOZZLE GEOMETRY
Exhaust Mach
Result
Pressure Ratio
Result
Ex. Temperature
Result
°K
Mass Flow Rate
Result
kg/s
Thrust
Result
N
I
Result
SP
s
(Specific Impulse)
1
Number of
Engines
1
9
I
Result
SP
(Specific Impulse)
361
Vehicle
Fuel Mass
1
3000
361
Vehicle
Payload Mass
1
3000
s
VEHICLE STRUCTURE
361
Engine
Unit Mass
1
1000
DELTA-V BUDGET MAP
Click a starting point on the Delta-V map to predict the range of paths achievable by the above rocket engine and vehicle configuration, travelling between some of the nearer parts of the Solar System.
Default start: Earth.
Note: This Delta-V budget map does not take rocket-staging into account, nor does it facilitate boosters (hence why some of the presets fail to lift off from earth when in reality they are able to).
The calculation also freezes the above engine ISP at its design ambient pressure/altitude, so atmospheres of destinations/origins don't alter the engine performance like they would in real life.
It also assumes that at each subsequent waypoint after the original launch point, the vehicle has a sufficient thrust to weight ratio (TWR) to lift off again.
progress toward
waypoint
one-way trip between
waypoints achievable
round trip between
waypoints achievable