From: Henry Spencer <firstname.lastname@example.org>
Subject: Re: new info on MacDac and Rockwell's X-33 designs: staged combustion
Date: Wed, 10 Jul 1996 15:21:00 GMT
In article <Pine.A32.3.92a.960708235014.123868Aemail@example.com> "'Larry' L Gales" <firstname.lastname@example.org> writes:
>> Probably. Proton has been flying with staged-combustion engines since
>> 1965. In fact, my impression is that all the big modern Soviet engines
>> use that cycle. They've been a little more cautious than the SSME's
>> designers, and have made it work pretty well.
>Could someone provide a good definition of staged combustion
Staged combustion burns part of the propellant in one or more preburners,
uses the resulting hot gas to spin the pump turbines, and then injects the
gas and the rest of the propellant into the chamber.
The advantage is that it gives you lots of pump power, which lets you run
the chamber at quite high pressure. The disadvantages are that high pump
power makes pump design very demanding, you have lots of high-pressure
hot-gas plumbing, and you've got a complex engine with a lot of feedback
loops to complicate debugging.
Life becomes much easier if you dump the turbine gas overboard instead of
feeding it into the chamber (the "gas generator" cycle). That does hurt
performance, however, both because the turbine gas contributes little
thrust and because the easier life means lower chamber pressure.
Another alternative is to drive the turbines with warm gas, obtained by
running the fuel (which must be cryogenic -- hydrogen, methane, or chilled
propane) through the chamber wall, driving the turbine(s) with that, and
then injecting it into the chamber. This is the "expander cycle" which you
hear a fair bit about in this group. The turbine-gas environment is much
more benign and easier to deal with, and the engine is somewhat simpler.
The downside is that pump power is limited because there's only so much
energy to be had.
If we feared danger, mankind would never | Henry Spencer
go to space. --Ellison S. Onizuka | email@example.com
From: firstname.lastname@example.org (Henry Spencer)
Subject: Re: Rocket Engine Questions and Considerations
Date: Tue, 2 May 2000 03:11:46 GMT
In article <email@example.com>, <firstname.lastname@example.org> wrote:
>...To put it another way, has it been
>determined for a particular vehicle type that one combination of number
>of preburners, shaft configuration, and preburner mixture richness lends
>itself to easier development, while another is most efficient, while
>another is more robust, and so forth?
In a word, no. Design experience is inadequate. When the number of
staged-combustion LOX/LH2 engines can be counted on the fingers of one
hand, there simply aren't enough data points for any general conclusions.
The Russians may, by now, be able to speak with some authority on the
development characteristics of staged-combustion engines using dense
propellants. But LH2 is a whole different can of worms, and conclusions
based on non-LH2 experience don't necessarily apply.
>...in addition to a longer development time, it can be argued that
>the RD-0120 benefited from a five year U.S. head start and so was able
>to both include materials or technologies not earlier available to
>NASA/Rocketdyne and avoid some incidental design drawbacks that the SSME
>might have been saddled with (though, specifically what, I wouldn't know...
Careful here: you're implicitly assuming that the Russians based the
RD-0120 on what they could read about the SSME, rather than on their own
far greater experience with staged-combustion cycles. It's much more
likely that they used their own materials and technologies and based their
design on their own past experience rather than on the woes of the SSME.
>...It can also be argued that the RD-0120, being
>non-reusable, can afford to push the operating envelope for an engine of
Not by very much. Practical liquid-fuel engines, except for types using
ablative cooling, *have* to be moderately reusable simply to get through
their typical test programs. There are considerations you don't have to
worry about *as much* if the engine life can be relatively short, like
cooling-tube hot-side wall fatigue, but it's not as much help as you might
>...separate preburners for the hydrogen
>and oxygen turbopumps (AW&ST, April 5, 1999, p. 57), albeit with the
>oxygen preburner running oxygen-rich. This is not the same scheme that
>they funded under the XLR-129 program.
Since some bad experiences on the SSME program, US engine designers (well,
engine concept designers) have strongly preferred schemes which don't
involve shaft seals between fuel-rich turbine gas and the oxidizer pump.
Whether this is a realistic view or sheer superstition, overgeneralized
from a single data point, is left to the student.
>1) Why did the Russians disbelieve the SSME design? Did they question
>the building of a staged combustion hydrogen engine, one that didn't use
>an oxygen-rich preburner, one with multiple preburners and shafts, or
>just the particular one which NASA and Rocketdyne pursued?
I don't have specific information on Russian views, but I would guess that
the sheer complexity of the design was what astonished them. The US had a
hard time making it work at all, as witness the complexity and elaborate
timing of the startup sequence.
>2) When it came to the SSME upgrade award in the 1980s, was it that NASA
>or P&W determined that the dual preburner design was fundamentally
>better, or did P&W's experience with the XLR-129 lead them to still
>maintain that a single preburner powerhead was better but were overruled
>by NASA, figuring that development costs would be too high?
Again, I don't know for sure, but my impression is that NASA wanted and
got re-implemented pumps, *not* a wholesale redesign of the engine. The
idea was to replace (what were seen as) marginal components, with minimum
"Be careful not to step | Henry Spencer email@example.com
in the Microsoft." -- John Denker | (aka firstname.lastname@example.org)