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, the types of propellants,
and the flow conditions upstream and downstream of the nozzle throat,
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 latest on-line student version of the program.
It 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.
is also available for more inexperienced users.
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 or Metric units for input and output and to
select the particular input panel.
Computed Thrust, Mass Flow,
Mass Flow , and
are 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 using
a slider, simply click on the slider button, hold down and drag to a
new position. You may also click on the arrows at either end of the
Details of the Input Variables
are given below.
- At the bottom right of the screen is the output from
the program displayed in output boxes. All of the values of the input
variables are also displayed on the output panel.
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 nozzle throat
is at "Throat-th."
The "Exit-ex" and "Free Stream-fs" conditions are
also noted. Free stream conditions exist far away from the nozzle.
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.
Depending on the input conditions, the exit pressure can be greater,
equal, or less than the free stream pressure. If the exit pressure is
greater than free stream, the nozzle is said to be Under Expanded
and the condition is noted on the schematic. If the exit pressure is
less than free stream, the nozzle is Over Expanded. If the exit
pressure is much less than free stream, a
may appear in the nozzle. This is a very undesirable design condition for the
nozzle because there are large entropy losses associated with the
normal shock. Further reduction in the nozzle pressure ratio causes the
shock to move upstream. In real nozzles, the shock interacts with the
wall boundary layer causes separation, and highly non-uniform flow. This
simple simulator can not calculate the details of these flow conditions,
so the performance variables are set to NA, not available.
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.
By convention, input boxes
have a white background and black numerals, output boxes have a
black background and yellow numerals. You can change the value of input
boxes by clicking on the box, backspace over the old value, type in
the new value, then hit the "Enter" button on your keyboard. You must
hit Enter to send the new value to the program.
- 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 by choice buttons at the
top of the panel. Propellants are selected as Fuels and
Oxidizers. Some of the fuels are mono-propellants, which
means that they do not need an oxidizer to produce combustion, only
a catalyst or a source of heat. For these fuels, the oxidizers are
specified as "NONE". All fuels can be combined with different amounts
of oxidizer to produce rocket exhaust. The weight ratio of oxidizer to
fuel is noted as the O/F ratio. For this simple simulator, we
only provide a couple of examples of the effects of O/F ratio which
you can select with the Option button. Again, for this simple
simulator, only selected fuel/oxidizer combinations are available. There
are more complex "combustion codes" which allow any combination of
fuel and oxidizer, but you need a more thorough understanding of
combustion chemistry to use such a code.
Selecting a propellant sets a value of
the molecular weight of the exhaust, 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
located next to the label.
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
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 , the computed
thrust of the rocket nozzle,
specific impulse Isp.
At the bottom, we show the selected fuel and
oxidizer and the oxidizer/fuel ratio.
The input combustion chamber
ratio of specific heats, Gamma, and the
molecular weight of the exhaust products are displayed on the second row.
Input combustion chamber total
and free stream pressure, Pfs, are displayed on the third row of outputs.
The nozzle exit pressure, Pex is computed from the
total pressure and the expansion ratio, Aex/Ath using
conditions from the throat to the exit.
The input throat area, Ath, and exit area, Aex, are
displayed on the fourth row with the
nozzle pressure ratio, NPR, which is the ratio of the
combustion chamber pressure to the free stream.
The throat Mach number Mth is set to 1.0 for the
The computed value of
exit velocity, Uex,
and exit Mach number Mex
are displayed on the fifth row.
compute all of the output variables is based on isentropic
flow through the nozzle.
When a non-isentropic shock wave appears in the nozzle, the analysis is
modified; isentropic flow is assumed up to the shock, then the
normal shock relations
are imposed, then a subsonic isentropic analysis from downstream of the
shock to the exit is performed.
Rocket Thrust Simulator:
Gas Density Activity: Grade 10-12
Mach Number Activity: Grade 9-12
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