NASASpaceFlight.com Forum
General Discussion => Q&A Section => Topic started by: strangequark on 09/04/2009 02:29 am
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Trying to get a clearer understanding on the challenges for staged combustion kerosene engines, and figured I'd toss this out there. I understand the reasons why oxidizer-rich cycles are desirable for a kerolox engine, but I'm curious why the preburner environment is seen as so challenging to choose materials for. I keep mentally comparing to jet engine combustors, which are also run at extremely oxidizer rich ratios, and at similar combustion temperatures. Is it that pure oxygen is that much more corrosive than air at high temp, and if so does anyone know how low the temperature would have to be to use comparable materials? Or is there some other factor at work that I'm not picking up on (likely)? I know this is kind of an obscure question, but I would greatly appreciate any insight that anyone can offer. Failing that, I would really like any pointers for good reading material on the subject.
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Is it that pure oxygen is that much more corrosive than air at high temp[...]?
Yes. For example: if the partial pressure of oxygen is high enough, aluminum will burn. Vigorously.
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Is it that pure oxygen is that much more corrosive than air at high temp[...]?
Yes. For example: if the partial pressure of oxygen is high enough, aluminum will burn. Vigorously.
But wouldn't the same be a problem for a hydrogen engine?
Danny Deger
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A hydrogen engine typically runs fuel-rich; this is not to say that hot hydrogen doesn't have materials issues. Just about any other chemical propellant in a stage combustion design has to run oxidizer-rich to prevent coking and sooting of internal components.
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Yes. For example: if the partial pressure of oxygen is high enough, aluminum will burn. Vigorously.
That does make a fair bit of sense, I guess. So, would the overall static pressure be a factor too? For instance, would air at 5000psi be as bad as oxygen at 1000psi (assuming air is about 20% ox)? Increased pressure increases reaction rate for gas-gas reactions, but I don't know if it does for reactions occuring on solid surfaces too.
A hydrogen engine typically runs fuel-rich; this is not to say that hot hydrogen doesn't have materials issues. Just about any other chemical propellant in a stage combustion design has to run oxidizer-rich to prevent coking and sooting of internal components.
Well, that and there's benefits to the power balances for the turbopumps when running oxidizer rich. AIAA has a great book called Liquid Rocket Thrust Chambers that walks through it, and gives pretty good clarity as to why kerosene should be ox-rich and hydrogen should be fuel-rich. However, it has just about jack regarding the material sciences issues, and I've been really curious.
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Failing that, I would really like any pointers for good reading material on the subject.
http://www.lpre.de/
In russian, but it seems for me readable through Google translator.
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To me the classic example of extreme is the Russian RD-253, a staged combustion engine that uses corrosive hydrazines and nitrogen tetroxide. I haven't found out much specific about the materials other than a long-standing mention of using zirconium as a lining in the hot sections. Other mention of oxidizer rich engines point out running with oxygen-compatible materials and limiting peak temperatures until it can run relatively fuel-rich in the combustion chamber. It all sounds like a very delicate and dynamic balancing act...
And that just a bit of foreign material getting into the hot sections and scratching the passivated surfaces can lead to a Bad Day At The Launch Pad.
Getting down to specifics seems to involve proprietary details the manufacturers are reluctant to talk about, but there must be a body of literature dealing with the general subject.
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Yes. For example: if the partial pressure of oxygen is high enough, aluminum will burn. Vigorously.
That does make a fair bit of sense, I guess. So, would the overall static pressure be a factor too? For instance, would air at 5000psi be as bad as oxygen at 1000psi (assuming air is about 20% ox)? Increased pressure increases reaction rate for gas-gas reactions, but I don't know if it does for reactions occuring on solid surfaces too.
When you use the word 'static pressure', one assumes it's in the context of supersonic flow, e.g., P-static/ P-total relationship, but I assume you're referring to the partial pressure of O2 in Air?
When compare pure oxygen with air (rocket engine vs. airbreathing engine), one must consider the effect of concentration of oxygen on materials. One way to understand this is to look at typical storage bottles for pure oxygen (such as in hospital/ transport trucks/ K-bottle, etc.) vs. those for air. The material of construction would be very different.
A hydrogen engine typically runs fuel-rich; this is not to say that hot hydrogen doesn't have materials issues. Just about any other chemical propellant in a stage combustion design has to run oxidizer-rich to prevent coking and sooting of internal components.
Well, that and there's benefits to the power balances for the turbopumps when running oxidizer rich. AIAA has a great book called Liquid Rocket Thrust Chambers that walks through it, and gives pretty good clarity as to why kerosene should be ox-rich and hydrogen should be fuel-rich. However, it has just about jack regarding the material sciences issues, and I've been really curious.
You may try to google "oxidizer-rich material compatibility" and gain some insights. I might add that the U.S. and the Russians have very different approaches on how to handle oxygen-rich environment.
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High pressure air is going to act somewhat differently than pure oxygen at the same partial pressure. This is because the N2 is going to act as both a chemical buffer and increase the thermal inertia of the system. Both retard reaction (heating) rates and maximum temperature.
I like Tom Flynn's book Cryogenic Engineering. It's generic to all cryo systems, but it's got lots of useful info and a chapter on safety.
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And that just a bit of foreign material getting into the hot sections and scratching the passivated surfaces can lead to a Bad Day At The Launch Pad.
Actually, it must be bigger 0.16х0.16mm. (size of cell in filter before RD-170-derived turbopumps) and new ox-rich turbopumps have not only 'passivated surfaces' as NK-33, but thick inert-material inserts in the most dangerous places, so 'foreign particle' not a such big deal, as killing possible rotor-stator recontacts in transient start/stop modes.
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Just about any other chemical propellant in a stage combustion design has to run oxidizer-rich to prevent coking and sooting of internal components.
As an aside: I believe that is why the X-15 used ammonia.
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A hydrogen engine typically runs fuel-rich; this is not to say that hot hydrogen doesn't have materials issues. Just about any other chemical propellant in a stage combustion design has to run oxidizer-rich to prevent coking and sooting of internal components.
Well, actually LCH4/LOX could run with a 1:1 MR in the preburner (fuel-rich) without coking. For RP-1 (approximated as poly-CH2), the minimum is 8:7. In general, one must burn all carbon to CO.
This is based on equilibrium and on the assumption that CO forms before H2O. In some places in the preburner, soot might form.
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A hydrogen engine typically runs fuel-rich; this is not to say that hot hydrogen doesn't have materials issues. Just about any other chemical propellant in a stage combustion design has to run oxidizer-rich to prevent coking and sooting of internal components.
Well, actually LCH4/LOX could run with a 1:1 MR in the preburner (fuel-rich) without coking. For RP-1 (approximated as poly-CH2), the minimum is 8:7. In general, one must burn all carbon to CO.
This is based on equilibrium and on the assumption that CO forms before H2O. In some places in the preburner, soot might form.
Can ethane or propane be burned stoichiometric or slightly fuel-rich without excessive soot formation?
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Suggest looking for a program called Stanjan. It's a freeware combustion solver. You can do your own calculations with it.
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Suggest looking for a program called Stanjan. It's a freeware combustion solver. You can do your own calculations with it.
A *real* engineer does it by hand calculation ! ;-)
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Only to cleanse the mental palate or after an EMP 8) Hand-calcs are for math majors.
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Suggest looking for a program called Stanjan. It's a freeware combustion solver. You can do your own calculations with it.
I've found CEA more user friendly.
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I believe the Russians used a very high chromium content Stainless Steel (.33 Cr or so) in their oxygen rich cycle engines...if it is passivated first by pickling in pure oxygen in a furnace under pressure, a relatively thick shell of chromium oxide forms at the surfaces exposed to the oxygen.
The chromium oxide layer prevents further oxidation of the internal components exposed to high pressure oxygen-rich gasses.
I believe the environment is too hot and too corrosive for more traditional passivization techniques such as gold plating which is used in the hydrogen rich preburners and turbines of the SSME. Infact, some of the strange discolorations inside the engine bells of fired engines must be a small amount of that gold plating from the high pressure turbopumps. :)
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Has anyone ever built a full combustion rocket where the propellants are so fuel rich there is no oxidizer left after the 'pre'- burner? LOX/Methane might work well. Instead of trading mass fraction for delivery simplicity as in a pressure-fed, this would be trading isp for decent mass fraction, pump simplicity and pump robustness.
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Has anyone ever built a full combustion rocket where the propellants are so fuel rich there is no oxidizer left after the 'pre'- burner? LOX/Methane might work well. Instead of trading mass fraction for delivery simplicity as in a pressure-fed, this would be trading isp for decent mass fraction, pump simplicity and pump robustness.
I belive that "fuel rich" means exactly that - all the oxygen in the pre-burner is consumed, and the exhaust is a mix of combustion products & unburnt fuel. (I guess this means the input is richer than stoichiometric). I believe all US engines operate this way.
Oxidizer-rich would exhaust a mix of combustion products & unused oxidizer. RD-180 (Russian, obviously) operates this way.
cheers, Martin
Edit: "Russion"
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Has anyone ever built a full combustion rocket where the propellants are so fuel rich there is no oxidizer left after the 'pre'- burner? LOX/Methane might work well. Instead of trading mass fraction for delivery simplicity as in a pressure-fed, this would be trading isp for decent mass fraction, pump simplicity and pump robustness.
I belive that "fuel rich" means exactly that - all the oxygen in the pre-burner is consumed, and the exhaust is a mix of combustion products & unburnt fuel. (I guess this means the input is richer than stoichiometric). I believe all US engines operate this way.
Oxidizer-rich would exhaust a mix of combustion products & unused oxidizer. RD-180 (Russian, obviously) operates this way.
cheers, Martin
Edit: "Russion"
Sorry I didn't express that very clearly. I meant there is no combustion chamber after the preburner. I.E. the preburner IS the combustion chamber. Running so fuel rich that the drive turbine doesn't melt.
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Sorry I didn't express that very clearly. I meant there is no combustion chamber after the preburner. I.E. the preburner IS the combustion chamber. Running so fuel rich that the drive turbine doesn't melt.
Basically you're asking if there as ever been an engine that has the turbine downstream of the combustion chamber. I believe the answer is no just for heat flux reasons (how would you cool the turbine blades?). The closest thing that comes to mind is a tapoff cycle like in the J-2S, where the gas to power the turbine isn't generated in a separate gas generator but is tapped off from the combustion chamber.
If the engine output was low enough temperature due to highly non-stoichiometric fuel-rich combustion to power a turbine I don't think you would get anything near a normal ISP.
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You may try to google "oxidizer-rich material compatibility" and gain some insights. I might add that the U.S. and the Russians have very different approaches on how to handle oxygen-rich environment.
There are some good results from searching for that phrase. But does anyone have any more detailed references of exactly how these pickling/passivation processes work, particularly related to similarities in afterburning jet engines (where the gas passing through the turbine is also oxidizer-rich, although to a lesser extent).
Also any citations on how the US does this and how that differs from Russian technology would be much appreciated.
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Why are oxidiser rich preburners so much more difficult than peroxide gas generators? Both involve hot oxygen rich gas and the temperatures should be similar if the turbine is expected to survive in them. Is it that you would have much higher mole fractions of oxygen, given that peroxide decomposes into two thirds steam and one third oxygen?
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Close: it's partial pressure, not mole fraction. For a constant main chamber pressure, preburners almost certainly have higher pressure than GGs. Then, ORPB exhaust will be almost all oxidizer.
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Thanks. Reading back I see I more or less recapitulated strangequark's opening question. Can you give a few pointers as to why higher partial pressures mean higher reactivity?
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OK, let me have a stab at my own question. Partial oxygen pressure is proportional to the number of collisions between oxygen molecules and the preburner / piping wall, and the reaction rate increases with the number of collisions?
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Right, in many gaseous processes it's the partial pressure that determines the reaction chemistry. It's the same as if it were pure at the same (lower) pressure. Concentration is at best a second order influence on these processes.
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Right, in many gaseous processes it's the partial pressure that determines the reaction chemistry. It's the same as if it were pure at the same (lower) pressure. Concentration is at best a second order influence on these processes.
These reactions must be strongly nonlinear with respect to temperature and pressure (partial pressure of O2). I mean aircraft turbojets (esp. afterburning ones) also have hot oxidizer rich gas passing through the turbine, but the materials compatibility question doesn't seem to be nearly so demanding. So I assume there's some cutoff below which there are no major oxidation problems (at least when compared to similar systems like jet engines), but then above that all hell breaks lose.
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What's the partial pressure of GO2 in an airbreather? It's been a long time since I've thought about it.
Worse (or better), though, you've got a huge amount of GN2 sucking up energy of reaction.
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Air-breathing gas turbine engines have a lot of atmospheric nitrogen passing over their turbines. Does that help reduce the oxidative potential?
Another question I've been curious about: for a hydrolox *full-flow* staged combustion engine, is the oxidizer-rich turbine environment any less demanding than that of a kerolox staged combustion engine?
For the hydrolox engine, the oxidizer-rich turbine is only driving the oxidizer pump, and the stoichiometric O/F ratio for hydrolox is higher than for kerolox, so I assume that the hydrolox oxidizer turbine would be more oxidizer-rich than the kerolox turbine, but with a lower temperature since there is less preburner heat gain for approximately the same mass flow rate.
On a related note, does anyone know if the integrated powerhead demonstrator is headed toward any production engine development? Or is it being passed over in favor of a kerolox booster engine?
http://en.wikipedia.org/wiki/Integrated_powerhead_demonstrator
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What's the partial pressure of GO2 in an airbreather? It's been a long time since I've thought about it.
Worse (or better), though, you've got a huge amount of GN2 sucking up energy of reaction.
Low, in rocket terms. Modern engines have a pressure ratio of ~40 at altitude. Even assuming they could pull that at sea level, you're looking at ~150psi P,O2 in the combustor, which is substantially higher than what you'll get in the afterburner.
(Back in airbreathing engines now, with a bunch of jet engine types. Have to chuckle when they speak of "2500 fahrenheit" in tones of awe.)
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Worse (or better), though, you've got a huge amount of GN2 sucking up energy of reaction.
But the temperatures are similar and you have much less fuel for cooling. Isn't that the reason for strangequark's colleagues' awe of the 2500 F mark?
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Sorry I didn't express that very clearly. I meant there is no combustion chamber after the preburner. I.E. the preburner IS the combustion chamber. Running so fuel rich that the drive turbine doesn't melt.
Basically you're asking if there as ever been an engine that has the turbine downstream of the combustion chamber. I believe the answer is no just for heat flux reasons (how would you cool the turbine blades?). The closest thing that comes to mind is a tapoff cycle like in the J-2S, where the gas to power the turbine isn't generated in a separate gas generator but is tapped off from the combustion chamber.
If the engine output was low enough temperature due to highly non-stoichiometric fuel-rich combustion to power a turbine I don't think you would get anything near a normal ISP.
Don't cool the turbine! Make it out of tungsten. Tungsten has a higher melting point than the combustion temp, if I have a correct figure for the latter-at least in very fuel-rich hydrolox engines
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But the temperatures are similar and you have much less fuel for cooling. Isn't that the reason for strangequark's colleagues' awe of the 2500 F mark?
The temperature is similar, but the O2 partial pressure is phenomenally lower, 150psi, versus maybe 7000 psi. As for 2500F, I am kind of amused to hear those temperatures referred to as "high". It's impressive, given the design challenges for a jet engine turbine (moving parts, only having air for coolant, etc), but I'm used to rocket MCC temps, and thinking of 6000-7000F as "high".
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Don't cool the turbine! Make it out of tungsten. Tungsten has a higher melting point than the combustion temp, if I have a correct figure for the latter-at least in very fuel-rich hydrolox engines
Manufacturability. That which cannot be melted cannot be cast, and tungsten is virtually impossible to machine, much less in the shapes you want for a turbine. Also, the density will kill you.
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Can W be stamped? Doubtful.
Would powder metallurgy help?
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A cooled rotor is a much larger degree of complication than goes into current engines. I can't think of one, though they probably exist.
Powder metallurgy is the secret sauce.
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So the materials compatibility problem on the NK-33 was much easier to solve than on the RD-170 because of the much lower chamber pressure (and hence PPO2)?
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A cooled rotor is a much larger degree of complication than goes into current engines. I can't think of one, though they probably exist.
I thought modern compressors typically used internal air cooling.
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You'd have to ask a Russian about that.
Sure, aircraft engines are, but not rocket engines that I can think of.
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Ah, I thought you were talking about aircraft engines.
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Tungsten carbide can be sintered, so tungsten blades could be made by this process.
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The problem isn't the temperature at which a particular material melts per-se, the big question is what structural properties does the material have at a particular temperature (yes, I know that W is really hard at room temperature; I have no idea about really high temperatures)? A really good counter example is trying to make a turbine blade out of Niobium just because it has a high melting point.
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The problem isn't the temperature at which a particular material melts per-se, the big question is what structural properties does the material have at a particular temperature (yes, I know that W is really hard at room temperature; I have no idea about really high temperatures)? A really good counter example is trying to make a turbine blade out of Niobium just because it has a high melting point.
Apparently tungsten has a pretty large drop-off in strength at around 1200C, due to a crystal structure change, and creep can be a big issue for polycrystalline tungsten.
Google Books Search (http://books.google.com/books?id=foLRISkt9gcC&pg=PA28&lpg=PA28&dq=tungsten+creep&source=bl&ots=-rANG4o-z0&sig=c2kgP7lmW6VIpPCKVqFkrfQ4_6w&hl=en&ei=6s_gTL7bE4aglAfl3qC3Aw&sa=X&oi=book_result&ct=result&resnum=10&ved=0CF8Q6AEwCQ#v=onepage&q=tungsten%20creep&f=false).
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The problem isn't the temperature at which a particular material melts per-se, the big question is what structural properties does the material have at a particular temperature (yes, I know that W is really hard at room temperature; I have no idea about really high temperatures)? A really good counter example is trying to make a turbine blade out of Niobium just because it has a high melting point.
Apparently tungsten has a pretty large drop-off in strength at around 1200C, due to a crystal structure change, and creep can be a big issue for polycrystalline tungsten.
Google Books Search (http://books.google.com/books?id=foLRISkt9gcC&pg=PA28&lpg=PA28&dq=tungsten+creep&source=bl&ots=-rANG4o-z0&sig=c2kgP7lmW6VIpPCKVqFkrfQ4_6w&hl=en&ei=6s_gTL7bE4aglAfl3qC3Aw&sa=X&oi=book_result&ct=result&resnum=10&ved=0CF8Q6AEwCQ#v=onepage&q=tungsten%20creep&f=false).
According to the same book (no attack or disrespect intended-your statement is true for most alloys of W), non-sag (= creep resistant) W forms exist.
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The problem isn't the temperature at which a particular material melts per-se, the big question is what structural properties does the material have at a particular temperature (yes, I know that W is really hard at room temperature; I have no idea about really high temperatures)? A really good counter example is trying to make a turbine blade out of Niobium just because it has a high melting point.
Apparently tungsten has a pretty large drop-off in strength at around 1200C, due to a crystal structure change, and creep can be a big issue for polycrystalline tungsten.
Google Books Search (http://books.google.com/books?id=foLRISkt9gcC&pg=PA28&lpg=PA28&dq=tungsten+creep&source=bl&ots=-rANG4o-z0&sig=c2kgP7lmW6VIpPCKVqFkrfQ4_6w&hl=en&ei=6s_gTL7bE4aglAfl3qC3Aw&sa=X&oi=book_result&ct=result&resnum=10&ved=0CF8Q6AEwCQ#v=onepage&q=tungsten%20creep&f=false).
According to the same book (no attack or disrespect intended-your statement is true for most alloys of W), non-sag (= creep resistant) W forms exist.
None taken. But I also took it to imply that those are monocrystalline forms.
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They are long-grained forms.
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Actually, tungsten would be catastrophically oxidized by H2O at that heat. Oops!
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Exactly what is it about oxygen rich staged combustion engines that make the Russian engines so efficient?
I understand the advantages of staged combustion, but what is the advantage of the oxygen rich variety?
Is it taking advantage of the phase change of LOX to GOX?
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Exactly what is it about oxygen rich staged combustion engines that make the Russian engines so efficient?
I understand the advantages of staged combustion, but what is the advantage of the oxygen rich variety?
Is it taking advantage of the phase change of LOX to GOX?
Efficiency is due to being staged-combustion, which allows the entire mass of propellant to be used as high-velocity reaction mass.
Running oxygen rich is a necessity with kerosene fuel, to prevent coking of the turbopumps and ductwork/injectors in a staged combustion scheme. If anything, I believe the main combustion chamber still runs slightly fuel-rich, because the average molecular weight of the exhaust is a bit lower and hence specific impulse is higher.
Also, to control the peak temperature of the turbopump components; running the mixture ratio at stoichiometric would melt those components. The actively cooled main combustion chamber is better able to stand those temperatures.
Hydrogen simply approaches these issues from the opposite direction and has a big advantage by running fuel rich of lower molecular weight and thus higher exhaust velocity, equating to a higher specific impulse.
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Exactly what is it about oxygen rich staged combustion engines that make the Russian engines so efficient?
I understand the advantages of staged combustion, but what is the advantage of the oxygen rich variety?
Is it taking advantage of the phase change of LOX to GOX?
As I understand it, the rationale for the oxidizer-rich staged-combustion in the Russian kerolox engines is that a fuel-rich preburner would produce soot when using hydrocarbon fuels which may cause problems with downstream valves and injectors.
Fuel-rich preburners are generally preferable for hydrolox staged-combustion. Note that the hydrolox staged-combustion engines for Energia/Buran used fuel-rich preburners.
I don't think that anyone has produced an oxidizer-rich hydrolox powerhead, unless we count the American Integrated Powerhead Demonstrator, which used both fuel-rich and oxidizer-rich preburners in a hydrolox full-flow staged-combustion cycle.
What makes the Russian kerolox engines special by comparison is that the U.S. has never flown a staged-combustion kerolox engine of any kind. The closest America has ever come is RS-84, which was also an oxidizer-rich design.
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Isn't another reason for using an oxidiser-rich preburner that you have more power to drive your pumps, which allows higher chamber pressures and thus high thrust at lift-off without sacrificing too much Isp?
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Does anyone know why the rd-180 only has a thrust/weight ratio of 80? The nk-33 has T/W of 136, with half the chamber pressure, in comparison.
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Exactly what is it about oxygen rich staged combustion engines that make the Russian engines so efficient?
I understand the advantages of staged combustion, but what is the advantage of the oxygen rich variety?
Is it taking advantage of the phase change of LOX to GOX?
Efficiency is due to being staged-combustion, which allows the entire mass of propellant to be used as high-velocity reaction mass.
Running oxygen rich is a necessity with kerosene fuel, to prevent coking of the turbopumps and ductwork/injectors in a staged combustion scheme. If anything, I believe the main combustion chamber still runs slightly fuel-rich, because the average molecular weight of the exhaust is a bit lower and hence specific impulse is higher.
Also, to control the peak temperature of the turbopump components; running the mixture ratio at stoichiometric would melt those components. The actively cooled main combustion chamber is better able to stand those temperatures.
Hydrogen simply approaches these issues from the opposite direction and has a big advantage by running fuel rich of lower molecular weight and thus higher exhaust velocity, equating to a higher specific impulse.
A pet peeve: the reason fuel-rich is more efficient is because the smaller molecules release their energy more efficiently at the limited expansion ratios practical.
The decrease in mean molecular weight is more than canceled out by the decrease in temperature.
Also, running fuel-rich makes it FAR easier to avoid burning up the MCC-both by lowering the temp and by making the mixture reducing, thereby preventing oxidation of the walls.
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Would coking still be a problem with pure dodecane instead of RP-1, say produced through a Gas To Liquid process?
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Does anyone know why the rd-180 only has a thrust/weight ratio of 80? The nk-33 has T/W of 136, with half the chamber pressure, in comparison.
The RD-180 is a cut down version of a much larger engine, and still has many of the larger-engine components involved, in order to reduce the cost. This makes it heavier as a result.
In addition, the NK company normally makes jet turbines. Their turbopump system reflects this, single shaft. It came at a cost of rotational physics getting in the way, but if you counter-mount them, the engines cut each other out. Or if your control system knows how to compensate for the spin. The reduction in components enables further weight savings.
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ok, so could the nk-33 raise its chamber pressure to the rd-180's level and have even higher T/W?
A second unrelated question: is the chamber temperature lower for kerosene/lox engines than hydrolox engines assuming the same pressure?
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ok, so could the nk-33 raise its chamber pressure to the rd-180's level and have even higher T/W?
A second unrelated question: is the chamber temperature lower for kerosene/lox engines than hydrolox engines assuming the same pressure?
That is part of what Aerojet is doing. The AJ-26 has higher thrust than the NK-33 on the N-1, which means, yes, better T/W. But it cannot go as high as the RD-180 without failure. IIRC, it can handle a 40% boost in pressure at most, but it fatigues the system faster.
And I think the temperature for hydrogen is lower.
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IMEO, thrust to weight of an engine is a dumb metric. It has no bearing on the final goal of getting something to a position and velocity in space. Weight of the whole vehicle and thrust to weight of the whole vehicle are what matter. The parochial metrics of an engine guy aren't for systems thinkers.
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IMEO, thrust to weight of an engine is a dumb metric.
It would be stupid to compare engines by T/W without looking at the fuel, then yes. For example, all LH engines will look "worse" than RP-1 ones.
If you do consider the kind of fuel used, it is a completely valid metric, because weight of the engine eats into payload.
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IMEO, thrust to weight of an engine is a dumb metric.
It would be stupid to compare engines by T/W without looking at the fuel, then yes. For example, all LH engines will look "worse" than RP-1 ones.
If you do consider the kind of fuel used, it is a completely valid metric, because weight of the engine eats into payload.
If Isp & thrust (ie gravity losses) are the same, then that would be true for an upper stage engine. However, I'd suspect that the additional mass would often be associated with improved performance (Isp or Thrust).
Even if Isp & thrust are the same for a 1st stage engine, mass of the 1st stage (and therefore of the engine) is certainly less than 1:1 with payload to orbit. I suspect a small additional engine mass balanced by a small improvement to Isp or thrust may deliver better performance overall. Especially with a 1st stage that stages relatively early (eg Kerolox 1st stage with high thrust Hydrolox 2nd stage).
cheers, Martin
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Dry mass is dry mass. Within a stage I don't care if it's tank, engine or computer. It all counts the same.
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Dry mass is dry mass. Within a stage I don't care if it's tank, engine or computer. It all counts the same.
OH SACRILEGE !!
Now listen here <pulling up pants>, don't you realize that engine is KING in the launch vehicle business? The rest of vehicle is just a flying engine test stand, and payload is just an excuse so someone will pay us to fly engines.
How dare you to relate dead structures & electrons to a LIVING, BREATHING MACHINE ! ;D
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I'll note that if you have an Oxidizer and a Fuel rich pre-burner each can drive a matching turbopump with *no* concerns about a seal failure let the drive gases mix with the pump contents.
This is not the case with the SSME where there are 3 seals between the turbine and the HPOTP and 2 chambers with separate venting and monitoring.
During testing it was found the planned volume of purge helium in these inter seal cavities had to be raised nearly 5x to ensure no hazardous mixing.
Going to dual fuel rich/ox rich pre-burners would eliminate a *major* failure mode.
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I'll note that if you have an Oxidizer and a Fuel rich pre-burner each can drive a matching turbopump with *no* concerns about a seal failure let the drive gases mix with the pump contents.
This is not the case with the SSME where there are 3 seals between the turbine and the HPOTP and 2 chambers with separate venting and monitoring.
During testing it was found the planned volume of purge helium in these inter seal cavities had to be raised nearly 5x to ensure no hazardous mixing.
Going to dual fuel rich/ox rich pre-burners would eliminate a *major* failure mode.
And double some other risks. It's all trade offs by trying to estimate probabilities.
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Going to dual fuel rich/ox rich pre-burners would eliminate a *major* failure mode.
And then you have the entire ox-rich problem to deal with. Engineering rule no. 1: TANSTAAFL.
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Going to dual fuel rich/ox rich pre-burners would eliminate a *major* failure mode.
And then you have the entire ox-rich problem to deal with. Engineering rule no. 1: TANSTAAFL.
True. Such is life.
But note this.
The seal issue is a *dynamic* problem. Excessive leakage through the seals, insufficient purge pressure are 2 outcomes (with *many* possible root causes leading to them). It all has to work on the vehicle, on the day (and on the SSME *every* day) in a heavy and complex way.
Going to an Ox rich pre-burner driven LOX turbo pump designs out these issues but demands addressing material science issues that have not been properly addressed in the US for *decades*, despite having been successfully solved in the FSU decades ago.
Regarding T/W I'll note H2/O2 engines have the worst T/W of *any* rocket engine not explicitly designed for launch vehicle use.
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I understand that staged combustion give better thermal efficiency to an engine. But I was wondering, if full stage combustion is so clear cut. Or is just an issue of easier turbine sealing and simplified injector design? I mean the injector because I'm assuming it would be an only gas injector against a mixed gas/liquid (SC) or liquid/liquid(GG) injector.
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Full staged as in both the fuel side AND the oxidizer side? I thought oxidizer-rich hot gas was a bad thing... leads to burned engine components, fires, etc.
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Isn't the Integrated Powerhead Demonstrator and the RD-270 that?
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Isn't the Integrated Powerhead Demonstrator and the RD-270 that?
Isn't the RD-270 late sixties-early seventies technology, and I can't find anything less than a couple of years old on the IPD. Were either successful?
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The RD-170 and NK-15 families are both oxidizer rich preburners. The ducts just have to be made out of metals that can take supercritical GOx at that temperature. Western technology has only gotten there in the last 15 years and still about TRL 6.
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I understand that staged combustion give better thermal efficiency to an engine. But I was wondering, if full stage combustion is so clear cut. Or is just an issue of easier turbine sealing and simplified injector design? I mean the injector because I'm assuming it would be an only gas injector against a mixed gas/liquid (SC) or liquid/liquid(GG) injector.
I know I'm kind of resurrecting an old thread (and one I started in my younger days), but I found this question while doing a google search on IPD from another question you asked (http://forum.nasaspaceflight.com/index.php?topic=15457.msg886637#msg886637), baldusi.
Full flow staged combustion is very clear cut from a performance standpoint. The power available to your turbines is linearly proportional to the mass flow you can put through them. Then, the pressure rise your pumps give to your propellant is linearly proportional to power.
So, full flow-->lots of mass flow through turbines-->lots of pumping power-->higher pump output pressure-->higher chamber pressure
For a first stage engine, higher chamber pressure means you can have a larger expansion ratio, while still avoiding flow separation from the nozzle. It also means a smaller throat for the thrust level, which means your nozzle (and engine as a whole) can be smaller for the same expansion ratio.
Lots and lots of performance benefits aside from the turbine seals and injection. However, you are also talking about an engine that may eat you alive in development costs. I think I remember Jon Goff saying that anything over 1500 psi just isn't worth it :).
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I understand that staged combustion give better thermal efficiency to an engine. But I was wondering, if full stage combustion is so clear cut. Or is just an issue of easier turbine sealing and simplified injector design? I mean the injector because I'm assuming it would be an only gas injector against a mixed gas/liquid (SC) or liquid/liquid(GG) injector.
Full flow staged combustion is very clear cut from a performance standpoint. The power available to your turbines is linearly proportional to the mass flow you can put through them. Then, the pressure rise your pumps give to your propellant is linearly proportional to power.
So, full flow-->lots of mass flow through turbines-->lots of pumping power-->higher pump output pressure-->higher chamber pressure
For a first stage engine, higher chamber pressure means you can have a larger expansion ratio, while still avoiding flow separation from the nozzle. It also means a smaller throat for the thrust level, which means your nozzle (and engine as a whole) can be smaller for the same expansion ratio.
Most of the benefit is that in a staged combustion engine all your propellant is exhausted at the full chamber pressure. In a gas generator cycle, you can't just dump the hot exhaust into the chamber because its pressure has dropped over the turbine, and you can't dump it upstream of the pumps or you'll flash-boil your propellants in the feed lines. You have to dump it overboard at some small fraction of main chamber pressure.
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I understand that staged combustion give better thermal efficiency to an engine. But I was wondering, if full stage combustion is so clear cut. Or is just an issue of easier turbine sealing and simplified injector design? I mean the injector because I'm assuming it would be an only gas injector against a mixed gas/liquid (SC) or liquid/liquid(GG) injector.
I know I'm kind of resurrecting an old thread (and one I started in my younger days), but I found this question while doing a google search on IPD from another question you asked (http://forum.nasaspaceflight.com/index.php?topic=15457.msg886637#msg886637), baldusi.
Your "younger days"? Strangequark, you're too young to be talking like that. It's like when my 33-year old English professor started going on a rant about "when I was your age" back when I was in college. ;) ;D
Full flow staged combustion is very clear cut from a performance standpoint. The power available to your turbines is linearly proportional to the mass flow you can put through them. Then, the pressure rise your pumps give to your propellant is linearly proportional to power.
So, full flow-->lots of mass flow through turbines-->lots of pumping power-->higher pump output pressure-->higher chamber pressure
For a first stage engine, higher chamber pressure means you can have a larger expansion ratio, while still avoiding flow separation from the nozzle. It also means a smaller throat for the thrust level, which means your nozzle (and engine as a whole) can be smaller for the same expansion ratio.
Lots and lots of performance benefits aside from the turbine seals and injection. However, you are also talking about an engine that may eat you alive in development costs. I think I remember Jon Goff saying that anything over 1500 psi just isn't worth it :).
I'm curious, but does anyone know how many different forms of staged combustion exist or have been thought up? I can think of at least three off the top of my head.
Twin-preburners (one oxygen rich, one fuel-rich)
Examples: RD-170 family
Single pre-burner & turbopump (runs oxygen-rich or fuel-rich)
NK-33, RD-162, RD-0120
Full-flow staged combustion
Examples??
Does the SSME count as a unique staged combustion cycle due to its resonance chambers?
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SSME is two fuel rich preburners
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SSME is two fuel rich preburners
Four?
cheers, Martin
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SSME is two fuel rich preburners
Four?
cheers, Martin
Two fuel rich + two LOX rich? ;)
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SSME is two fuel rich preburners
Four?
cheers, Martin
Two fuel rich + two LOX rich? ;)
No, now I look again Jim was right. ::)
Two fuel rich burners, with the LP turbo-pumps driven from the pumped output of the associated HP pumps (so two fuel-rich burners, each driving two pumps).
cheers, Martin
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Does anyone know why the rd-180 only has a thrust/weight ratio of 80? The nk-33 has T/W of 136, with half the chamber pressure, in comparison.
One reason is that NK-33 does not have swivel mechanism in it's nozzle, RD-180 has 2-way swivel
N-1 was steered by throttling the engines,
Sojuz-2-1v will be steered by RD-0110R vernier thruster.
AJ-26 weights much more than NK-33 because the swivel mechanism is added, but it seems the weight of AJ-26 is reported nowhere, all links to AJ-26 just point to NK-33 even though the engines have considerable differences.
But AJ26 still propably has slightly better T/W ratio than RD-180.
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I'm curious, but does anyone know how many different forms of staged combustion exist or have been thought up? I can think of at least three off the top of my head.
Twin-preburners (one oxygen rich, one fuel-rich)
Examples: RD-170 family
Single pre-burner & turbopump (runs oxygen-rich or fuel-rich)
NK-33, RD-162, RD-0120
Full-flow staged combustion
Examples??
Does the SSME count as a unique staged combustion cycle due to its resonance chambers?
You're confusing a few things.
1) About staged combustion the best way to think about it is the Russian nomenclature. They speak about how fuel and oxidizer arrive at the chamber's injectors. Thus, a gas generator is liquid/liquid, a staged combustion is gas/liquid and a full flow is gas gas. The only gas/gas engine that I know about, was the failed RD-270 engine. And there was the Integrated Powerhead Demonstrator. But I don't think they even had a thrust chamber for it.
2) RD-170/1/1M family has two identical oxidizer rich gas generators. That's why making the RD-180 required so little changes. When they developed it they didn't felt confident on making a single gas generator of that side. But both gas generators are oxidizer rich and only the oxidizer arrives as a gas to the chamber. The fuel is injected in liquid form.
3) The H2 and CH4 engines might extract some extra energy by doing something like an expander, but it's usually to avoid cavitation or such. In the RS-25D case the H2 LP turbine I think it does use expander power, but I'm not sure. I know that the only use of the LP turbines is to avoid cavitation.
4) You have to consider the expander cycle, for example. You have both closed cycle (like the RL-10/Vinci/RD-0146), that inject gas fuel and liquid oxidizer. And you also have open cycle expander (like the LE-5A/B and LE-X), where the expanded fuel is dumped and the fuel that does arrives to the injectors is in liquid form.
5) strangequark did teased about a new cycle for the RL10 replacement program.
6) You also have to consider the pressure fed and piston pump fed engines, you might consider them different cycles.
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Does anyone know why the rd-180 only has a thrust/weight ratio of 80? The nk-33 has T/W of 136, with half the chamber pressure, in comparison.
One reason is that NK-33 does not have swivel mechanism in it's nozzle, RD-180 has 2-way swivel
N-1 was steered by throttling the engines,
Sojuz-2-1v will be steered by RD-0110R vernier thruster.
AJ-26 weights much more than NK-33 because the swivel mechanism is added, but it seems the weight of AJ-26 is reported nowhere, all links to AJ-26 just point to NK-33 even though the engines have considerable differences.
But AJ26 still propably has slightly better T/W ratio than RD-180.
I believe the t/w difference has a lot to do with the RD-170 family's combustion cycle. If you compared an single-chamber RD-191 with a single-chamber NK-33, the NK-33 is a nice, compact engine with a single turbopump and an oxygen-rich preburner. An RD-191 is a much larger engine with only modestly more thrust. If I'm reading this RD-180 schematic correctly (http://www.nap.edu/books/0309102472/xhtml/images/p20010c31g259001.jpg), it appears there are not one but two turbopumps feeding into a single pre-burner, which ignites some of what passes through the turbine that feeds the combustion chambers, or combustion chamber in the case of the RD-191. Looking at the NK-33 schematic, it presents a much simpler form of staged combustion: http://www.bing.com/images/search?q=nk-33+engine+schematic&qs=n&form=QBIR&pq=nk-33+engine+schematic&sc=0-23&sp=-1&sk=#view=detail&id=6B49A6F2969218BFD58C2EB34E09FED0C9561D0B&selectedIndex=1. Given an RD-191 has a respectable t/w ratio of 89 with two-plane swiveling up to 8 degrees, I'd guess it's the combustion cycle needed to achieve those incredible pressures that requires so much mass. Although the RD-193, essentially a lightened RD-191 lugging around 300 fewer kilograms, has a t/w ratio of 103.
I would guess that the AJ-26 probably still has a better t/w ratio than an RD-193, though the difference will be a lot less than between an RD-180 and an NK-33. Anyone know why an RD-191 has such a superior t/w ratio to its RD-180 cousin?
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After reading a few things, on this site and elsewhere, about this type of engine I've got a few questions:
Is it possible to change the fuel/oxidizer ratio in the preburners? I mean while the engine is running. Could this be used to change thrust (isp) and pressure in the combustion chamber?
In order to simplify plumbing, can a single set of pre-burners/turbopumps be used to feed several combustion chambers? Sort of using a single assembly for a small cluster of engines. Less mass per chamber/nozzle would improve T/W, right?
Actually, I'm thinking that 2-4 engines might be clustered as a single propulsion unit, combining also actuators for gimballing (i.e. the whole thing would gimball/throttle as a unit).
Thnx
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After reading a few things, on this site and elsewhere, about this type of engine I've got a few questions:
Is it possible to change the fuel/oxidizer ratio in the preburners? I mean while the engine is running. Could this be used to change thrust (isp) and pressure in the combustion chamber?
In order to simplify plumbing, can a single set of pre-burners/turbopumps be used to feed several combustion chambers? Sort of using a single assembly for a small cluster of engines. Less mass per chamber/nozzle would improve T/W, right?
Actually, I'm thinking that 2-4 engines might be clustered as a single propulsion unit, combining also actuators for gimballing (i.e. the whole thing would gimball/throttle as a unit).
Thnx
Yes on both accounts and you are describing what RD-170/171/180 engines are doing. There's a throttle for fuel going to the preburner and another throttle for rest of the fuel going to the nozzles. Oxidizer is free to flow to the preburner as fast the turbopump can pump it.
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BTW, the RD-0124 has a single TP and preburner that ffeds the 4 main combustion chambers AND the four vernier engines. But the single nozzle version (the RD-0125), appears to be lighter (but longer).
The RD-170/1 family, uses two preburners to run a single turbopump. They had instability problems at the preburner level (and the MCC, too). That's why they use two preburners and four MCC. But generally speaking, changing the O/F ratio is usually done to handle propellant management.
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I believe the t/w difference has a lot to do with the RD-170 family's combustion cycle. If you compared an single-chamber RD-191 with a single-chamber NK-33, the NK-33 is a nice, compact engine with a single turbopump and an oxygen-rich preburner. An RD-191 is a much larger engine with only modestly more thrust. If I'm reading this RD-180 schematic correctly (http://www.nap.edu/books/0309102472/xhtml/images/p20010c31g259001.jpg),
I think the combustion cycle is very much the same.
The schematic for the NK-33 shows less details.
The NK-33 has the 2 boost pumps integrated into the same housing as the TP. That might be the main reason, why a RD-180 looks more complicated.
That is one NK-33 boost pump, the center shaft is the main shaft for fuel/oxidizer, and this boostpump is driven by the flow of the TP:
http://www.lpre.de/sntk/NK-33/img/tna_o.gif