With this software you can investigate how a rocket
nozzle produces thrust
by changing the values of different factors that affect thrust. By
changing the shape of the nozzle and the flow conditions upstream and
downstream, you can control both the amount of
gas that passes through the nozzle and the exit
There are several different versions of RocketThrust which
require different levels of experience with the package,
knowledge of thermodynamics, and computer technology.
This web page contains the on-line Version 1.4 of the program.
is also available.
Version 1.4b includes an on-line user's manual which describes the
various options available in the program and includes hyperlinks to
pages in the
Beginner's Guide to Rockets
describing the math and science of rockets.
More experienced users can select a
version of the program which does not include
these instructions and loads faster on your computer.
You can download these versions of the program to your computer
by clicking on this yellow button:
NOTE: If you experience
difficulties when using the sliders to change variables, simply click
away from the slider and then back to
If the arrows on the end of the
sliders disappear, click in the areas where the left and right arrow
Images should appear, and they should reappear.
If you see only a grey box at the top of this page, be sure that Java is
enabled in your browser. If Java is enabled, and you are using the Windows XP
operating system, you need to get a newer version of Java. Go to this link:
try the "Download It Now" button, and then select "Yes" when the download box from Sun
The program screen is divided into four main parts:
- On the left of the screen is a graphics window in which
you can display a drawing of the nozzle you are designing. You can
control the appearance of the graphics by using your mouse
and the slider
located in the graphics window. Details are given in
- On the top right of the screen are choice buttons to select
English and metric units for input and output and to
select the particular input panel. Computed Thrust and Flow through
the nozzle are also displayed here. The red "Reset" button
is used to return the program to its default conditions.
- On the middle right of the screen are the interactive
inputs to the program. Inputs to the program can be made using
sliders or input boxes. To change the value of an input variable,
simply move the slider. Or click on the input box, select and
replace the old value, and hit Enter to send the new value
to the program.
Details of the Input Variables
are given below.
- At the bottom right of the screen is additional output from
the program displayed in output boxes. By convention, input boxes
have a white background and black numerals, output boxes have a
black background and yellow numerals.
Details of the
Output Variables are given
On the left is a schematic drawing of the nozzle you are
designing. Flow is from
top to bottom for the rocket nozzle. The combustion chamber (or
plenum) conditions are noted by the "Plenum-0," and the throat
is at "Throat-th." The "Exit-ex" and "Free Stream-fs" conditions
You can move the schematic in the graphics window by clicking on the figure,
holding the left mouse button down and drag the schematic to a new location.
You can change the size of the schematic by using the Zoom slider at the
left of the graphics window. Click on the bar and move it along the line.
If you lose the schematic, click on the word "Find" to restore the schematic
to its default location.
You can change the length of the nozzle in the schematic by
using the "Length" slider on the "Geometry" input panel.
In real nozzles, the length to throat area ratio
is important for keeping the flow attached. In this simulator,
viscous effects are ignored, and the length is used only for "nice"
graphics--it does not affect the calculation of thrust.
The input variables are located at the middle right on three panels;
geometry, flow, and propellant.
You select the type of input panel by using the choice button above the panel.
- If you select the Geometry input,
you must specify the
throat area Ath. For a rocket or
convergent-divergent nozzle, you must also
specify the exit area ratio Aex/Ath.
The plenum area ratio Ao/Ath and the Length of
the nozzle are given for pleasing graphics, but are not used in
the calculation of performance.
If you select the Flow input, you can change the
chamber total pressure Pto, total temperature
Tto, and free stream static pressure Pfs.
The pressure and temperature are used in the calculation of the
mass flow through the nozzle.
For rocket calculations, if you change the propellants, the
plenum chamber temperature is re-set to the average combustion
temperature of the propellants. You may then change the chamber temperature
to see its effect on thrust by using the sliders and input box on the
Flow input panel.
If you select the Propellant input, you can change the
gas which passes through the nozzle.
The names of several propellants are given on the choice button next
to the "Mol. Wt." label. Selecting a propellant re-sets the value of
the molecular weight, the ratio of
specific heats gamma
combustion temperature .
The change in molecular weight changes the
gas constant used in the calculation of the
mass flow through the nozzle.
You can select to use a typical value for the molecular weight of the products
of combustion, or you can input your own value by using the choice button.
The value of the ratio of specific heats depends on the temperature of the
flow, and you can use a typical curve for the variation of gamma, or input
your own value by using the choice button next to the "Gamma" label.
Finally, the combustion of the propellants generates a typical combustion
temperature. You can use the typical value, or input your own value on the Propellant
input panel by using the choice button.
Output variables are located at the top and bottom on the right.
At the top of the output group are the
weight flow and the computed
thrust of the rocket nozzle.
At the bottom are the computed values of
specific impulse Isp and
exit velocity Uex.
The exit velocity depends on the exit area Aex and the exit
Mach number Mex, which are also
displayed. The exit Mach number is determined by the throat Mach
number Mth and the area ratio, which is input. The thrust
depends also on the pressure at the exit Pex and on the
overall nozzle pressure ratio NPR. The
compute these variables is based on isentropic
flow through the nozzle.