guest blogger john hare
Original ideas on my part have slowed down a bit, so I stole borrowed some from other people. Jon, Ben, Paul, Thomas, and Eric all share in this hare-brained one. Shifting the blame, of course.
Jon introduced the Thrust Augmented Nozzle concept here a couple of years ago. By introducing one or more compact (very short L*) combustion chambers into the expansion region of the nozzle, it is possible to increase thrust at lift off, have a much larger expansion ratio in vacuum, and have a serious throttle down between so that the efficiencies and thrusts match launch requirements better than existing engines. It allows all the above to be done at much lower chamber pressures than currently associated with high performance rocket engines.
On arocket, Thomas suggested moving my in chamber turbine to an aerospike tip. Ben suggested that a regeneratively cooled turbine closer to the throat that also contributed thrust augmentation would be more useful. Moving parts inside the combustion chamber strike Eric as a Very Bad Idea, so Bens’ suggestion might address that issue as well. Thomas also mentioned using a center hybrid grain as drive shaft protection.Â
Paul was quite disturbed by my suggestion of injecting oxydizer and propellant in constant contact with a chamber wall. He feels that the close proximity of the mixed propellants  to the very hot ignition source of the thrust chamber interior would resemble an explosion more than a controlled rocket burn. By using the very short L* TAN chambers, the explosive burn can be directed down the nozzle almost unconfined.  Â
 
At lift off, this engine would be generating thrust from the main chamber with a bipropellant chamber with some hybrid augmentation from a very slow regression grain. The hybrid grain would be more of an ablative drive shaft protection than normal hybrid grain, except that it is selected to provide useful fuel. The TAN chambers would be fed from the turbine tips operating just below the throat in a similar manner to the discussions of several months ago. The turbine drives the pump impellers feeding the main chamber with the drive shaft protected by the hybrid grain. The turbine in this case resembles an aircraft propeller more than a tiny rocket turbine disk. The combined thrust of the main chamber with the TAN augmentation should get thrust/weight ratios in the 150-200 range at lift off.
With improved acceleration off the pad, the need for throttle down will occur much earlier in the launch profile than current practice. At thrust reduction 1, the turbine core flows are shut off so that only the cooling channels feed the TAN chambers. A thrust reduction of 50% or more is possible at this stage. The reduction in turbine flow to the TAN chambers has the odd effect of increasing available pressure in the main chamber. When the turbine is no longer extracting work to drive the main TAN flow, rpms will increase as the main chamber will still be providing as much turbine drive power as before with half or less of the fluid mass to pump. Â
At thrust reduction 2, the turbine is blown off and expended, leaving the stage to reach orbit with pressure feed to allow deep throttling at much reduced pressures. In vacuum, with the very high expansion ratios possible with TAN, good performance is possible down to very low pressures including the pressurant gasses in a VAPAC propellant selection.
johnhare
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Does the turbine have to be expended?
Depending upon the turbine design (rpm, solidity, blade number, etc) you will have a pretty challenging pulsed flow in and out of your TAN chamber(s). BTW, is the TAN chamber annular, or a ring of cans?
The turbine doesn’t have to be expended, but during very deep throttling there will be no coolant flow to save it from melting. Also the system could overspeed when pumping very low quantities.
I was picturing a semi pulsed flow with numerous chambers, somewhat similar to the earlier pulse detonation aerospike, except with deflagration instead of detonation for a somewhat smoother operation.
Not sure if this has been mentioned before, but wouldn’t a turbine shaft that long develop some pretty serious whip?
I’ve been thinking about using solid grain as ablative cooling too. I’d wondered about using it downstream of the throat (I think most nozzle extensions use ablative cooling?) to provide the inevitable TAN. One thing that occurs to me though is that the recession of the grain will alter the geometry of the combustion chamber. Have you considered the effect of the change in L* as the solid grain erodes?
The turbine shaft would require good support to avoid whip, as well as investigation of actual length desired. I don’t see it as a show stopper if the rest of the concept was desirable.
The ablative grains downstream of the throat would be operating in a fuel rich environment and wouldn’t provide any combustion energy. Even if running lean with excess oxygen, it seems unlikely that the burn time available in the supersonic stream would produce useful thrust, especially without some throat geometry to focus the energy.
The recession of the grain inside the chamber will increase L* for probably minor combustion efficiency improvement.
It seems to me that most ablative surfaces are designed so as to minimize the rate at which they burn through. I don’t think solid rocket fuel qualifies in this regard. Given the immense size of the SRB’s on the shuttle, and the fact that it only takes them a couple of minutes to completely burn out, I think you’d be hard pressed to find a fuel grain that would burn slowly enough to remain useful throughout the first stage of your flight profile, especially considering that you’re talking about a much smaller volume of fuel. And if you did find something that would burn through that slowly, I’d have to wonder just how much it would actually be contributing to the engine thrust.
Granted, I’m not much of an expert on solid fuels, but I was under the impression that, in most engines, there are usually chunks of solid fuel which flake off and continue to burn while still in the flow. If so, then there is probably a good chance that they could remain large enough to cause damage to the turbine at the throat.
I do like that almost all of the moving parts are now outside of the main combustion chamber, but I still have similar reservations about the TAN chambers and the tips of the blades.
Hybrid grains are supposed to have much lower regression rates than solid fuel grains.
I can’t believe I didn’t think about solid chunks hitting the turbine blade. SS1 had a problem with chunks even on the prize flights, so that would definately be a show stopper.
TAN chambers and blade tips would be the worlds first muzzle loading rocket. That has to count for something in the museum of dumb rocket tricks, even if not in the real world.
Hmmm, hadn’t considered the burn time, with that solid TAN thing, I figured the weight/thickness of the required coating would be what made it impractical. Out of interest, by “throat geometry to focus the energy”, do you mean something like a plug nozzle?
According to this site
http://www.braeunig.us/space/propuls.htm#solid
some fuels have a threshold flow velocity below which they don’t undergo erosive burning, so it might be possible to avoid chunks in the flow.
Not necessarily a plug nozzle, just a ring chamber geometry with throat that takes the gas to supersonic before it enters the expansion nozzle and merges with the primary stream.
It does seem likely that gas velocity in contact with the fuel grain could be low enough to discourage the chunks from erosive burning. While I had always thought in terms of regen cooling of parts in chamber, Thomas had a pretty good argument for just throwing a fuel grain around it instead and not trying to over engineer the concept.
Another fascinating subject to which I can add no technical commentary. I just like the TAN concept, and would be happy to see an operational engine of the basic vanilla version developed.
Has anyone figured out if the Aerojet HC Boost is/was a TAN concept?
What if you replace the whole turbin and place solid fuel in the TAN chambers and use them for the start and the first flight phase. In the second phase the nozzle had much better efficiency.
Jörg:
It sounds like you’re describing an “integrated rocket-rocket.” The most basic integrated rocket-ramjet works this way, with a solid fuel charge taking up a ramjet’s combustion chamber for the initial boost.
But the idea of the TAN nozzle is to work like the afterburner of a turbojet engine, meaning that you add extra propellant and oxidizer to the exhaust stream already coming from the actual combustion chamber.
Roderick
I know what TAN means, I wanted only deleting the turbin blades and shaft and keep the combustion chamber. And the solid fuel would be in separat chambers just like the picture above.
Hi John. I got to thinking about ways to minimise the velocity of the flow over the fuel grain so as to minimise errosive burning, and it struck me that since the shaft is rotating, the velocity would be minimised by a “rotating” flow, such as the “swirl” in Orbitec’s vortex engine. Could you bring the oxidiser injection from the top of the chamber to the bottom to create a vortex engine?
13Tim
Hi John. I got to thinking about ways to minimise the velocity of the flow over the fuel grain so as to minimise errosive burning, and it struck me that since the shaft is rotating, the velocity would be minimised by a “rotating†flow, such as the “swirl†in Orbitec’s vortex engine. Could you bring the oxidiser injection from the top of the chamber to the bottom to create a vortex engine?
I didn’t draw the shaft accurately. The shaft must be enclosed in a bearing housing all the way down to the turbine with multiple bearings. I didn’t realize the misdirection until one of my friends in the business pointed that out to me off line. We went a couple of rounds before I realized that I had implied an exposed shaft.
I don’t see any reason the swirl you suggest couldn’t be done with the injectors. I get real nervous now thinking about any solid fuel material immediately above the turbine.