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)

Match Exhaust Pressure to Ambient

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.

Expansion Ratio

Result

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)

Show Exhaust Plume

1 Number of

Engines

1

9 I

Result

SP

(Specific Impulse)

361 Vehicle
Fuel Mass

1 3000

361 Vehicle
Payload Mass

1 3000

s

361 Engine
Unit Mass

1 1000

Total Thrust

Result

N

TWR

Result

Thrust to Weight Ratio

∆V

Result

km/s

Impulse per unit mass

(space-faring "budget")

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

Expansion Ratio

Result

Exhaust Mach

Result

Pressure Ratio

Result

Ex. Temperature

Result

°K

Mass Flow Rate

Result

kg/s

Thrust

Result

N

I

Vehicle
Fuel Mass

Vehicle
Payload Mass

Engine
Unit Mass