This started as a short thought on nasaspaceflight and grew into something that resembled a blog post. So I copied and pasted it here. I screwed up the link but this is an ongoing thought from the turbine in chamber stuff from a decade and a half back. November 10 2008 Performance Monoprop being one of the series. I recently realized that the concept could serve as a pressure booster in a gas generator cycle to get performance almost matching that of staged combustion with a higher thrust to weight ratio and much faster and cheaper development.
I seem to be drifting back to an ability to do some simple prototyping and am interested in finding an engine company that might do a bit of business with an inventor.
The provenance of the regenerative cooled turbopump includes Rotary Rocket, ATREX, and a patented “single rotor turbine” from LANL. https://image-ppubs.uspto.gov/dirsearch-public/print/downloadPdf/6807802
Rotary Rocket introduced a concept to avoid turbopumps (kinda) by putting the thrust chambers on the ends of rotating arms. The canted thrust chambers spun the assembly which pressurized the propellants in the feed pipes with the centrifugal force. There was no central drive shaft as the assembly freewheeled. It eliminated the separate gas generator, turbine, drive shaft, impeller, and all of their housings and accessories. The engine was never completed beyond some subscale testing. Which makes sense when one thinks of the problems of getting 72 thrust chambers to operate properly in the high gee field at the ends of those arms. Just the weight of all those engines rotating at high speeds should make one nervous. Still, the idea of simplifying the engine process is appealing.
The ATREX was a Japanese hydrogen fueled air-turborocket with tip turbine blades mounted to the outside of the compressor. It was claimed to be lightweight and compact. Whether all the claimes made sense or not I don’t know. But putting tip turbine blades on the Roton structure seemed to be an easier path than the mass of whirling engines. It could be a much smaller structure delivering propellants into a thrust chamber bolted onto the vehicle in the normal way. Then the gas in the thrust chamber could drive the blades on the way through the throat. Seemed simpler, but exposes the turbine blades to more heat than even regeneratively cooled blades would be able to stand. Also, it leaves no path for cooling the nozzle and thrust chamber.
The Single Rotor Turbine https://image-ppubs.uspto.gov/dirsearch-public/print/downloadPdf/6807802 patented by LANL has the hollow turbine blades as the last stage of the compressor. The air flows through the inside of the blades as coolant while being accelerated and compressed. Then the air flows out of the blades into a volute for pressure recovery before entering the combustion chamber. In this way, all of the air is used as turbine coolant before all of it is burned in the combustion chamber and used to drive the turbine. The goal being to raise the allowable turbine inlet temperature for a more efficient engine. Primary direction seemed to be power generation. There seems to quite a bit of prior work with ideas of this nature as several previous patents are referenced.
Liquid propellants by nature have about three orders of magnitude more mass for a given volume than gas. As such, there is an enormous amount of coolant available inside the turbine blades limited by the requirement that the propellants exit the turbine blades as liquid. It also takes far less energy and velocity to pressurize a given amount of liquid than it does to pressurize gas. It seems apparent that using the LANL concept applied to liquid propellant would result in a relatively simple high pressure turbopump.
In the last post on the regenertive cooled turbine, I skipped over the pump characteristics suggesting using it in a fairly normal layout that happened to allow much higher turbine inlet temperatures. This actually is better as a stand alone turbopump as the driveshaft, housings, and torques involved would be just as heavy as the standard units for a fairly modest gain in capabilities at the expense of R and D on a new system, not to mention the uncertainty. Also, it would almost certainly be tasked to a staged combustion engine which is one of the most expensive and difficult engines to develop.
What I am going to suggest here is more modest in some ways and more radical in others. Develop a small turbopump and gas generator of this nature that free wheels on its’ shaft in the same manner as the Roton engine. Without having the stresses of torque through the turbine blades and disk, torsion in the shaft, and torque limitations in the impellers, the single rotor turbopump can operate at much higher speeds than any normal layout. The turbine inlet temperature can be much higher than any normal turbine even while the disk and blades are much cooler. The torque stresses are minimal with the drive gas on one side of the turbine blade driving against the liquid propellants on the other side of one thickness of metal. The centrifugal stresses will be the same for a given rpm and radius.
With the “brakes off” as compared to normal systems, the tip velocities of the turbine/impeller combo can reach speeds normally reserved for turbo compressors. Tip velocities creating velocity head, and velocity head going as the square of velocity, 40,000 feet of head pressure is theoretically achievable with tip velocities of 1,600 fps. LOX to over 19,000 psi and RP to over 14,000 psi is theoretically achievable. Cutting those in half with decent pressure recovery might give something close to reality. Using the RP at 7,000 psi as pressure in the gas generator and running as hot as the turbine blades allow should bring the turbine pressure drop to a very modest value allowing 5,000+ psi at turbine exit.
The interesting thing about the single rotor turbopump is that it doesn’t gain weight at the same rate as a conventional system. One that would run all the RP and much of the LOX, or vice versa, in a Merlin would be a unit you could pick up in one hand and not strain. But that would be several bridges too far. Especially as on of the weaknesses of this system is that it gives no reasonable path for regenerative cooling of the thrust chamber and nozzle.
I suggest it might be worthwhile to use such a unit in place of the gas generator on a gas generator cycle engine. Gas generators seem out of style at the moment with so many going for the staged combustion and full flow staged combustion. The reasons are the improved performance compared to the gas generator due to the gas generator exhaust being both lower temperature and lower velocity than the main flow. That percentage of propellant “not pulling its’ weight” is a potent argument. I suggest that replacing the passive gas generator (illustration 2-20 Huzel and Huang page 45) with a very small active gas generator with the single rotor turbopump could boost gas pressure to the standard turbine by 5,000 psi. This could allow either much less propellant to run the standard turbine which would boost system Isp. Or it could have and exhaust of 1,000 psi into a secondary combustion chamber with additional oxidizer creating a respectable Isp in itself. If the gas generator is using 10% of the total and getting 320 seconds in vacuum and the main chamber is getting 360, then the system would only be about 3-4 seconds below that obtained with staged combustion at similar pressures. Considering that the gas generator cycle has inherently better thrust to weight and is easier to develop, a new conservatively designed gas generator engine might be a contender against the state of the art stuff being developed now.
And then there is the option of using them as boost injectors. Initially one. Then later possibly 19 in the manner that the V2/A4 engine recycled the burner cups from a previous.
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