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#20
by
gin455res
on 11 Nov, 2010 14:58
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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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#21
by
drbobguy
on 12 Nov, 2010 01:23
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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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#22
by
drbobguy
on 12 Nov, 2010 01:26
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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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#23
by
mmeijeri
on 12 Nov, 2010 11:57
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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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#24
by
Antares
on 12 Nov, 2010 18:35
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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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#25
by
mmeijeri
on 12 Nov, 2010 18:40
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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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#26
by
mmeijeri
on 12 Nov, 2010 19:17
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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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#27
by
Antares
on 13 Nov, 2010 01:09
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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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#28
by
drbobguy
on 14 Nov, 2010 04:46
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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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#29
by
Antares
on 14 Nov, 2010 05:23
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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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#30
by
butters
on 14 Nov, 2010 05:25
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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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#31
by
strangequark
on 14 Nov, 2010 05:27
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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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#32
by
mmeijeri
on 14 Nov, 2010 14:43
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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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#33
by
tnphysics
on 14 Nov, 2010 19:44
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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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#34
by
strangequark
on 14 Nov, 2010 19:53
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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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#35
by
strangequark
on 14 Nov, 2010 19:54
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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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#36
by
tnphysics
on 14 Nov, 2010 20:13
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Can W be stamped? Doubtful.
Would powder metallurgy help?
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#37
by
Antares
on 14 Nov, 2010 21:33
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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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#38
by
drbobguy
on 14 Nov, 2010 21:35
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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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#39
by
mmeijeri
on 14 Nov, 2010 21:36
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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.