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

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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.

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)

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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

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DELTA-V BUDGET MAP

Total Thrust

Result

N

TWR

Result

Thrust to Weight Ratio

∆V

Result

Impulse per unit mass

(space-faring "budget")

km/s

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© Michael Middleton 2020

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