Latest News - 8/4/08

The detailed design of the new 250 lbf chamber has been completed. Designing in the O-ring seals took a bit more effort than I anticipated. I also ended up using a 12-pass cooling tube instead of the 10-pass tube I was planning so the pressure drop is also higher than I had hoped. My estimate of the pressure drop through the cooling jacket is about 300 psi which will keep the maximum fuel temperature at around 430 degF. I'll post the thermal analysis estimates when I get a chance.

In the hopes that the rocket powered by this engine will get some altitude, I designed the engine to have optimum expansion at 15000 feet. If I did the calculations correctly, it should provide about 270 lbf at sea level (assuming no separation), 250 lbf at 15000 ft, and 238 lbf at 30000 feet. The total engine weight not counting fasteners should be around 5.2 lbm but I should be able to make some additional cuts and easily get it down to about 4.5 lbm.

I picked up the 5 inch diameter 6061-T6 bar stock last week and I'm just waiting on a new boring bar before I start making chips. I'm trying some new cutting inserts this time (TCGT with AK chipbreaker) which are supposed to work great with aluminum. I actually was considering a slightly larger flange than 5 inches on the top of the chamber when I realized that 5 inches was the max diameter I can turn on my lathe using the inside jaws. The outside jaws can turn much larger but the chamber is fairly long so I want to have a good grip on it. Instead of using NPT fittings, this time I'm using SAE O-ring ports. The only problem is that the port reamers are very expensive (~$200 each!) but this should allow me to avoid stripping out the aluminum threads with the stainless steel fasteners.

The test stand will need some upgrades to support a 30-40 second run time on this engine in addition to the higher thrust. The load cell has a maximum capacity of 500 lbf but I need to add some linear slides or something similar to keep the engine from being cantilevered out so much. I'm also considering adding a cascaded purge system to blow out the extra propellants after each run.

6/1/08

I've been working on the design for the next engine the past few months. It will be a 200-250 lbf regen engine with LOX/Kerosene as the propellants, essentially the same design as the 100 lbf engine I've been testing but with some changes (O-rings for seals instead of Grafoil, etc.). One of the parameters I wanted to determine from the 100 lbf regen tests was the gas side heat transfer coefficient so I could use it to correct my design equations. I've done that and used an hg correction value (x1.45) for the design of the new one but I'm wondering whether I'm using the proper methodology to determine it. I'll go over the method I used when I get a chance to write it up.

Since I plan to use this new engine to fly in a rocket, I'm trying to decide on the exact thrust level I need. I don't have a lot of empirical data yet for empty weights of the various parts in a vehicle but I suspect the empty weight will be heavier than I expect. The vehicle will be designed for a 30-40 second burn time so I'm hoping the new amateur rocketry regulations get approved in the next year or so which lift the 15-second burn time limit.

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