Luke Maurits said:

Thanks for doing that little bit of analysis, rpulkrabek.  When you say you set the inside temperature to 1000 C, exactly what do you mean?

I set the inside of the aluminium combustion chamber to be 1000 C. I would have rather set the fuel to be at that temperature, but I would have needed to know the convection coefficient, which, I believe, is determined experimentally. In any case, I feel that the high temperature of combustion is too much for aluminium.

My gut feeling is that it is best to move forward with using phenolic insulation. I don't know much about it, but I think we will be able to put a thin layer between the HDPE and Aluminium. I think this thin layer of phenolic insulation will be much less massive and take up much less volume than planning to leave unburned HDPE. But, I don't know for sure. I would like to find out more about phenolic insulation first.

With that said, can anyone provide information about phenolic insulation? How is it used? How expensive is it? How efficient is it? How much heat can it withstand?

Thanks very much for your comment.  I just realised that I've hot-linked to your image above, sorry for being so rude, I hope it hasn't caused you any inconvenience.  I don't think CSTART forum posts pull an awful lot of traffic, but if you want us to re-host it locally please just say so.

It's good to know that both arrangements of the main and drogue chute are feasible, I was wondering if some diagrams I had seen around the place where incorrect.  Last we discussed this, I think we were starting to consider side-deploying chutes to eliminate the mechanical complexity of a rocket body made of separable sections.  I don't suppose you have any knowledge or experience with such arrangements?

Putting the drogue in the middle has the advantage of rocket separating into two piece of approximately same length and likely to go into a tumble, as opposed to a backend glide (more drift). Also, the shear pins preventing the main parachute compartment from opening don't have to carry the weight of the booster and thus the premature deployment of the main parachute (because of shear pins breaking before ejection) is less likely. 

Feel free to contact me (pavelp@dnastar.com) if you want more details regarding dual deployment.

It's worth noting that for the majority of a burn, there is some PE fuel between the actual combustion region and the inner wall of the combustion chamber and that will act as insulation.  Combustion will only actually happen directly adjacent to the inner wall at the very end of combustion when the PE grain has been almost completely burned out - and, in fact, we could design the grain geometry so that oxidiser runs out and combustion ends before this happens.  I have no idea how good an insulator HDPE is.  Could you model the transfer of heat from the combustion region through the PE and then through the Al and then to the ABS easily using some software?

I've also seen the phenolic insulation a lot and agree it can't hurt to look into that more.

My calculations went on to show that the temperature at the point where the Al connects with the abs is equal to 999.3 deg C. Not much changed from the 1000 deg inside temp. This would definitely melt the abs tube, which has a melting point of about 105 deg C! I suppose this should have been obvious in the first place for such a thin piece of metal. Well, as one of my old professors once said, "Common sense seems to come to us once we have already done something." 

Something clearly needs to be done, if the combustion temperature is to be this high. I have noticed recently that other rockets are using phenolic insulation on the inside of the combustion chamber. I think we should look more into this.

Did you have any particular plastics in mind?

I was thinking about this recently. If the combustion temperature and heat transfer from Al to ABS (or whatever plastic) is too high, we could simply implement some sort of insulation, such polystyrene foam (styrofoam).

There is one other worry that I have. Will there be heat generation from atmospheric reentry? Or, could it be so, that the drogue parachute will prevent this?

I am also happy to lock down the material choice for the oxidizer tank and combustion chamber. Maybe we can put a green check mark on this in the wiki and create the blog posts and such with a comment that describes our decision and to provide details if there are any disagreements. With regards to the diameters, I agree with you that we should focus to about 30cm diameter, but to try to relate this to what is readily available, if, for example, we can only get 29cm diameter.

I basically agree that  the first two could be finalised without much more discussion.  We definitely shouldn't vote on 6061 Al T6: this is obviously the correct choice on the basis of all our research and modelling.  I am happy to lock that one in right now.  As for choosing the outer diameter of the engine / the diameter of the aeroshell: I feel like this is the only point above that we would really need to make a hard choice on before we could move forward with meaningful extra work, so we should try to make a choice on it soonest.  I think maybe the main driver for this decision should be material availability.  Is there a measurement somewhere between 25 and 30 cm such that findings things like thin-walled aluminium tubing or tubing of strong engineering plastics in that diameter is relatively easy?  If so, we should just choose that.

• The tank and chamber should be made of 6061 Al T6.
• The engine should have a maximum outer diameter between 25 and 30 cm (anything narrower is going to have an absurdly high fineness ratio).
• The total length of the engine should not exceed 5m (the 175 kg propellant engine I used to break the Karman line in OR is 4.3 m, so we know this is possible)
• The total propellant mass should be 180 kg (this is more than enough with our current design (which can probably be made lighter) and has the nice property that a 8:1 O:F ratio results in 160 kg of N2O and 20 kg of PE, i.e. we have nice round numbers for both propellants).
• The inner diameter of the chamber should be such that it is easy to buy PE rods of that diameter to use as grains.
• The average thrust or maximum burn time should be SOMETHING WE STILL NEED TO FIGURE OUT

How should these decisions be handled? A vote? In my opinion, the first 2 could be finalized, although, it's best to get everyone's opinion too. 

How does one go about creating the counter-sunk holes?  Are there drill-bits that can achieve this?

There are counter sunk cutters, that are similar to drill bits. I think it would be easy enough to use in something as simple as a drill press.

Luke Maurits said:

Also, with regards to attaching things to the ends of the combustion chamber (like the injector manifold and the nozzle), I suspect it will be necessary to use O-rings or something like them due to the relatively high pressures of hot gasses inside them. So when designing the manner in which all these things fit together (like in your final graphic above) we should probably take this into account.

I think some sort of gasket is needed for the injector manifold. I am still not entirely certain how this should be designed, but I am confident we can figure it out with a little bit of research. As for on the nozzle end, I am not sure that it's needed. Somebody, please, correct me if I am wrong, but, I don't think we need to worry about any fluid leaking from the nozzle. Is there another reason a gasket would be used on the nozzle? Do other rockets use gaskets near the nozzle?

One benefit of all the recent mass-modelling work from an avionics perspective has been that we can now get relatively decent upper bounds on the maximum acceleration (linear and angular) the avionics module will be subject to, which is an important design requirement - among other things, it will help us select appropriate accelerometers and gyroscopes.

Also, with regards to attaching things to the ends of the combustion chamber (like the injector manifold and the nozzle), I suspect it will be necessary to use O-rings or something like them due to the relatively high pressures of hot gasses inside them.  So when designing the manner in which all these things fit together (like in your final graphic above) we should probably take this into account.

With regards to moving forward, I think now is perhaps actually a good time to move toward a natural stopping point - we should make what decisions we want to/can from the basis of our modelling so far (like the requirements listed in this post) and write some reports and officiate some of the decisions we have made (e.g. use of 6061 Al T6).  This lets people following the blog etc. know what we've been doing and also puts all of this forum-based word somewhere more accessible and permanent.  Then we can move forward, making sure we keep everything consistent with the decisions/requirements in these reports.

That said, if we can agree on requirements quickly, I am happy to start work on the reports and let you go ahead with modelling stuff like heat transfer (maybe also, if you can, thermal expansion of the chamber?  We haven't spoken about this before, but I guess it's kind of important)?  I'm really starting to get out of my depth engineering-wise anyway, and want to start moving on to more organisational stuff, so doing the reports is a step in that direction.

Once the modelling is done, I can assign materials and then determine the complete mass as well as the center of gravity. This would then be useful, since OR only understands a rocket made entirely from one material (am I right?). 

I would also like to see the effects of heat from the combustion transferred to the aeroshell. This way we would be able to rule out a plastic for it's material or not.

The aeroshell then needs to have a counter sunk hole. As you can see in the picture below, the rocket looks much more streamlined.

Here is a close up of the nozzle section:

Finally, here is the cross section of where the components are assembled. Hopefully this allows you to see how each component is constrained and the role of the socket cap screw: