It is interesting comparing the two best known first stages in the US that use kerosene and LOX. The Atlas 5 and the Falcon 9 use a similar fuel in their first stages and then diverge in the technical aspects. The Atlas 5 with the RD 180 engine has about 10% higher Isp at sea level while the Falcon 9 Merlin has nearly twice the thrust/weight ratio. The over all Falcon 9 first stage seems to have a much lower dry mass ratio which makes up the difference in engine performance and then some.
There are going to be new vehicles designed by the various companies eventually that would like to benefit from the competitive advantages of both vehicles. A high thrust/weight ratio engine with high Isp that also has low dry mass is a desirable target. The more these features can be designed in, the more mass is available for payload, reusability, or both.
One of the engine cycles that is discussed from time to time is the dual chamber concept. It is more or less a gas generator cycle with an exhaust pressure high enough to inject into a lower pressure thrust chamber to burn with fuel or oxidizer to get useful thrust. I suggest it might be possible to get very near RD 180 Isp with very near Merlin thrust/weight with a variation of the concept. A low stage dry mass being part of the goal, I add in a few features that may be unique.
In the cartoon I have two high pressure chambers on the outside with a lower pressure chamber in the middle with an altitude compensating nozzle.
The black boxes in the tanks are the electric inducer pumps from the previous post.  They are to keep the propellants at high enough pressure to the main pumps to suppress cavitation as well as keeping required tank pressurization to a minimum.
The small blue tanks in the inter tank area are for the liquid hydrogen that serves multiple purposes. First the hydrogen feed hits a heat exchanger in the LOX Â tank to keep it cold enough to stay liquid and suppress cavitation even as tank pressure drops. Then it hits a heat exchanger in the RP tank for the same purpose. Then it is used to cool the turbine blades the same way that jet engines use air cooling. Finally it burns with the excess LOX from the gas generator to produce thrust.
With the pumps providing pressures to the main engine similar to that of the RD 180, the Isp of them should be similar. About 10% of the propellant goes to the gas generator driving the pumps with a residual pressure of 300 psi after the turbine. If the 300 psi engine was a normal kerosene engine, one would expect an Isp in the 250s from that portion of the thrust system. With the lean (LOX rich) gas generator driving a hydrogen cooled turbine at much higher than normal turbine inlet temperatures, the warm hydrogen mixes with the hot oxygen as it is used for film cooling of the blades and burns in the secondary chamber above the throat. The hydrogen/kerosene/LOX engine at 300 psi could approach the ISP of the main engines due to the higher performance of hydrogen. Hydrogen usage will be a fraction of a percent of the total propellant load.
The compensating nozzle of the low pressure engine in the center would allow reasonable Isp of that portion at sea level, especially with the hydrogen component. The higher expansion ratio made possible would allow much higher Isp at altitude, which, with the hydrogen component, could give vacuum Isp higher than the RD 180. I think the potential result is low hardware mass combined with high first stage performance.
johnhare
Latest posts by johnhare (see all)
- SPS in the Van Allens - October 21, 2023
- Regenerative cooled turbopump - March 16, 2023
- Second Guessing Starship - September 19, 2022
Aren’t you also adding more tanks? You posted this very early in the morning my time. Get some sleep John.
Adding liquid hydrogen tanks in place of gaseous helium tanks. And it was late morning my time, 5:00-5:30 or so
Will the hydrogen tanks be as small as they are drawn in the cartoon? The hydrogen will be a small fraction of the propellant by mass, but not as small of a fraction by volume. Have you done a back of the envelope calculation for how much hydrogen will be used?
LH2 tanks would be bigger than the cartoon. 1/2% by mass would be 7%. by Volume.
It’s a nice comparison and a tempting sketch. But if you are prepared to take LH2 along and even run it through your turbine blades, you’re deep into tri-propellant territory and, in my view at least, not too far (in terms of technical complexity) from LH2/LOX/Li or even F engines where the Isp would be vastly superior, the sort Rocketdyne tested around 1963.
With the LH2 as a minor component and only used on the low pressure areas I don’t see it as nearly the technical complexity of the tri-props you mention. But then perhaps either I don’t get your point or explained mine poorly.
I certainly don’t want to discourage you, but LH2 is a tricky substance to work with. Check out the SSME for all the hoops they had to jump in order to use it for turbine blade cooling. Plus you get a super cold zone in your launch vehicle – one that requires He for precooling, adding to your launch table cost – and it sort of messes up the thermal design of the whole thing. It’s perfectly doable, no doubt, but it just adds a layer of complexity that, in my view at least, would require a very major performance boost to justify. If you can reach that, and specially for quite a large LV, it might be worth it.
The low-pressure constraint helps of course, although it also introduces a design constraint on your gas generator. Not sure that’s an issue though.
Max,
If I couldn’t deal with criticism, I shouldn’t post weird ideas. iPhone comment from work. More tonight.
John,
Good to know. I admire your ideas, btw.
I post ideas I think are interesting, at least to me. I also normally thing they may be useful to somebody even though that is a low probability. This particular one I think might be worth some thought if someone is going clean sheet vehicle, and it happens to fit their particular capabilities. I would say that this concept would be worth somewhere between a break time discussion, and a serious trade study to an interested party. It should almost go without saying that the particular layout I described is not likely to ever fly. If it serves as an idea generator for even a mild improvement to one entities vehicle though, it will have been well worth my time.
Pointing out problems with a concept at the amateur blog level is also massively cheaper than finding them during design, or worse, in hardware. LH2 is indeed a pain on several levels. The problem is that it is so useful on other levels that looking for ways around its’ problems will be a constant.
If I personally reach the point of having enough $olid fuel to investigate some ideas, the sequence will be more or less like this.
1. Do mock ups of the pumps, injectors, or nozzles I find interesting.
2. Possibly do some cold flow to get past the laugh test.
3. Involve someone with experience before building components for separate testing.
4. Test individual components somewhere that is already set up to get honest numbers and do it safely.
5. Avoid creating a situation that will bankrupt me. (Actually #0, but I’m normally optimistic)
6. If promising results are obtained by any of the components, decide how to move forward.