The objective of this project is to conceptualize, simulate, and build a functioning rocket with the following capabilities:

• Wireless igntion
• Controlled decent with the aide of a parachute
• Wireless real-time data logging of: Altitude, acceleration, and temperature
• Take and save multiple images onto an onboard SD card

All parts of this rocket have been included as .STL files and the original models have been designed using “AutoDesk Inventor Student Version” and can be provided to anyone interested via e-mail.

The .STL files are found here:

• launch_bracket.stl
• rocket_fins.stl
• nose_cone.stl
• The main tube can be made of 40mm OD PVC tubing.

The Part “Rocket Fins” has been designed with a wall an overall wall thickness of 2.4mm. In order to print a solid, perfectly fused part, this needs to be taken into account. If the 3D printer is equipt with a 0.4mm extruding nozzle, 3 shells should be selected for the printing parameter. If the 3D printer is equipt with a 0.8 extruding nozzle, this part MAY NOT PRINT WELL (his has not been tested).

Fabrication of the Rocket: (COMING SOON)

The fabrication of rocket motors is inherently dangerous and this tutorial has only been included for educational purposes. I would advise that any user purchase an Estes class F motor which has an outer diameter of 29mm.

…that said, I have used an artcile, "Potassium Nitrate based Rocket Propulsion" as starting point to design a reliable motor.

Within the article the authors designed a motor with the dimensions: 21cm long by 2.5 cm in diamanter and therefore a total volume of approximately 104cm cubed.

According to the author, the motor has the impulse rating of 80-160 Ns (Class G rocket)

However, this article stated that the Rocket Nozzle was exerimentally designed by Richard Nakka.

After doing some research online, I quickly determined that the physics of a rocket nozzle is not something that can be learned inside of a week. The previous mentioned page by Mr. Nakka offers a host of open-source simulation software to help in the design process of amateur rockets. After visiting a number of rocketry club websites and other tutorials, I have seen his website come up numerous times and have taken the personal choice to trust the integrity of his software.

The software required is called: ProPEP3

This software does not require GUIPEP to run the graphics interface. By following the basic installation instructions on the website you should have the program up and running very shortly.

• Open the program “ProPEP 3”, a shortcut should have been installed onto your desktop.

This program will allow us to input up to ten different chemical compounds and will out put the combustion temperature, molar mass, and chemical composition of a combustion reaction. I advise very heavily that you take the time to read the provided user manual as it will explain the program in much more detail than i will provide here. In the following tutorial, I will be identifing only the parameters and mixtures that will help us design our rocket nozzle.

• This is the main graphics interface of the ProPEP 3 work space.

The GUI provides 4 tabs:

1. Propellant Formulation
2. Grain Information
3. Test Burns
4. Compute A&N

• We are only to be concerned with the first tab “Propellant Formulation
• The other 3 are to be used with a force meter for engine testing and data acquisition

1. Under the menu “Ingredients”, give your recipe a 10 character name
2. Select a drop box and scroll down to fill out each given ingredient (Potassium Nitrate, Sorbitl, Iron Oxide)

• Fill out the weigh. This progam asks for input in grams, however the total weight can not excede 100grams and therefore can be treated as a percentage. This will not effect the combustion temperature in theory. The combustion temperature is of the primary parameter when determining a nozzles expansion ratio.
  • I have chosen: KNO3=64g ; Sor=32g ; FE03=2g

• Within the menu “Operating Conditions” we are presented with 3 Parameters:
• Temperature of ingredients(K) : The temperature of the motor before launch. In our case, room temperature, or 298K
  • Chamber Pressure (PSI): 1000PSI = 68ATM. The user manual advises to keep this at 1000PSI for most tests. However I would like to further investigate this as chamber pressure has an important role in burn rate/exhaust speed/burn temperature
  • Exhaust Pressure (PSI): This is the pressure in which the exhaust will be being dumped into. In other words, since the rocket is being launched at sea level, we must input the pressure at sea level (14.7PSI). However if the rocket is to be fired at elevation we can correct the exhaust pressure using the function:
    • p = 101325*(1-(2.2557E-5)*h)^5.2558 p=Pascals ; h=meters
    • 1Pa = 1.45E-4 PSI
    • 1Meter = 3.2 feet
• Select “Boost Velocity and Nozzle Design”
• Select “Calculate”

• After the program has finished running (A couple seconds) you will see :
  • Isp* = Something
  • C* = Something
  • Ect…
• Select the button “Show Results”, if all has gone well the above screen will appear.
• This is the main Data display page. We will br primarily interested in the “Chamber Results Follow” Section

• Within this section it is interesting to note:
  • T(K) = Ignition temperature
  • Cp/Cv = Gamma
  • The molecular Weight of the mixture
• These 3 outputs are interesting as it will allow us to use yet another simulation in the future. The other simulation will allow us to test the perfomance (Thrust) of a rocket nozzle in respects to these principale parameters.
• I advise that you read the provided instruction manual for the software to better understand the rest of this page as it is very interesting, albeit beyond the scope of this tutorial.
• You may now close the Data menu and return to the main “Propellant Formulation” menu.
• Click “DISPLAY NOZZLE GRAPHS”

It should be noted that this simulation CAN NOT account for burn rates of a given solution as it is highly dependant on the volume/surface ratio of the final “fuel cell”. Beyond that, the propogation of the ignition, humidity, environment… all play a role in burn rates and therefore the overall debit of gas.

It is for this reason that, this simulation has given us a “rough idea” of the performance of our rocket motor. It is NOT to be treated as a perfect simulation by any means.

• This is a graphical output of the last section of the DATA menu (not discussed within this tutorial). What we have is the specific impulse in function to the expansion ratio of the Nozzle.

• Where the Expansion ratio is : the ratio : DE/DT
• Therefore, when we model our nozzle we should have a ratio of somewhere between 5-15 to increase the specific impulse (perfomance) of our designed rocket.

When designing the rocket & motor casing we must take consider the following:

1. The motor casing must be as light as possible
2. It must withstand 68ATM (Determined via Simulation)
3. It must sustain 1600K (Determined via Simulation)
4. Have the overall dimensions 2.98cm OD x 9cm (To maintain an equal volume with respects to the article
5. A “de Laval” nozzle is required to boost specific impulse of the rocket

The solution decided upon was that of a compound motor casing, made from a plaster mold and an ABS outter shell. A plaster mold was chosen as a potentionally good casing material as well as a nozzle. A 4 piece mold was designed on “Autodesk Inventor Student Version” and the .STLs are available here (Autodesk files are available upon request):

• base1.stl (Large Blue with the Set Studs)
• base2.stl (Large Blue with the Set Holes)
• internal_motor.stl (Long Grey)
• internal_nozzle.stl (Short Grey)

Base 1 & 2 are each 20cm long and may require access to a large format 3D printer. Our pieces were printed using an Ultimaker with 0.4mm extruders and were created using PLA.

1. We have our four individual mold parts
2. The grey “internal parts” fills the space where the future propellant and nozzle will be placed. The blue “bases” form the outter cylindrical shell of our future motor case. The space created between the grey/blue surfaces is 2mm wide and will be the cavity that we will be filling with plaster.
3. Upon closer inspection as to where the 2 “internal parts” one can see that the cavity will be much thicker in this area and the profile does in fact create our “de Laval” nozzle. The expansion ratio between the throat/exit is 3.3:1 with a 7mm throat and a simulated Specific Impulse of 130seconds.
4. The second Base closes on top of the other 3 pieces and a 3D cavity is created for the cylindrical motor casing.

A crude simulation was created using FEMM 4.2 to demonstrate the propagation of heat from the center of the motor to the outside of a 40mm OD Plaster insulator. FEMM To do this with FEMM 4.2 a few “liberties” were taken to try and approximate the burning of the fuel.

In this tutorial we will see how to very easily create a time dependant simulation using Euler's Method.

• What this means is that we will generate an initial condtion; the rocket is sitting outside on the launch pad with all components the same temperature (300K).
• Then at T(+0), we will assume that the rocket will burn from it's core outward. (this is our first “best case scenario assumption)

We assume also that the rocket will burn with peak thrust within the order of seconds. This has been based upon the article sited in the begining of this tutorial. We can see below that their rocket burned for near 45 seconds with a pressurized chamber for only 2.5-3.0 seconds. The temperature is assumed to be at the previously simulated 1600K. The time leading up to this intense burn can not be assumed to be at 1600K as it is under ambient pressure.

• This tells us that during our simulation the propagation of the “combustion surface” should be fast enough to burn 20mm outwards in under 5 seconds.

The final result of our simulation can be seen here in a GIF: This is a 13 second simulation of the heat transfered from the buring motor to the outside surface of the plaster casing.