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Clarification on Russian Engines
by
drbobguy
on 07 Feb, 2008 21:38
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Hello all,
My understanding is that many Russian engines (NK-33, RD-170, RD-171, RD-180) run oxidizer-rich preburners. But I'm not sure on the question if the exhaust is also oxidizer-rich.
And is the purpose of running oxidizer-rich preburners only to prevent carbon deposition, or is there some kind of other tradeoff? I assume running oxidizer-rich in the combustion chamber/nozzle is because O2 is lighter weight than most hydrocarbon/oxygen products so you get higher ISP from the lighter molecules (e.g., why H2-LOX engines run fuel-rich).
Does anyone know specifically what metallurgical skills are necessary for the plumbing for the hot oxidizer-rich gas in between the preburner and combustion chamber?
Why was the US so reluctant to follow this path?
Sorry if these are naive questions. If someone could point me to literature on this I'd be very thankful.
I'm also keenly interested on the initial research in the Soviet Union on staged combustion.
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#1
by
meiza
on 08 Feb, 2008 00:32
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Actually kerosene rockets exhaust mainly CO and H2O from what I remember offhand - both are lighter than O2 although CO marginally. And you don't want heavy CO2! So run fuel rich.
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#2
by
drbobguy
on 08 Feb, 2008 02:30
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Russian rockets do not run fuel rich. It appears I was wrong and that you actually get MUCH higher chamber pressures (around double) by running oxidizer-rich as opposed to fuel-rich in the staged-combustion preburner, for LOX-hydrocarbon (Note I am NOT talking about the combuster/nozzle). For LOX-hydrogen, you're better off running fuel-rich like the SSME.
Does anyone have any good advice to a detailed history as to why the Soviets went for LOX-hydrocarbon and never LOX-H2 in first stages? (Even Zenit/Energiya was LOX-hydrocarbon).
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#3
by
pm1823
on 08 Feb, 2008 02:55
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LH2 is a too costly fuel for the first stage.
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#4
by
drbobguy
on 08 Feb, 2008 03:00
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The Soviet Union didn't care about costs. LH2 is cheap if you can just order a couple thousand people to make it. I mean, of course economics does come into play, but I suspect that's not the whole issue.
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#5
by
Jim
on 08 Feb, 2008 03:04
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drbobguy - 7/2/2008 5:38 PM
Sorry if these are naive questions. If someone could point me to literature on this I'd be very thankful.
I'm also keenly interested on the initial research in the Soviet Union on staged combustion.
Good luck.
Russians hold this close to their chest, they have a version of ITAR
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#6
by
hop
on 08 Feb, 2008 03:30
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#7
by
simonbp
on 08 Feb, 2008 04:58
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The problem with (and beauty of) LH2 is that it burns so hot, but is stored so cold (20 K), requiring both high-temp metallurgy and cryogenic fuel pumps. The US had a step up because it spend a bunch of money to develop a hydrogen-powered P&W J-57 jet for the (canceled) successor to the A-12/SR-71. P&W took that technology/experience and applied to make first the RL-10, and then the J-2. The Russians decided instead to invest in improving the RP-1 rockets, and never really deployed LH2 until Energia...
Simon
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#8
by
pm1823
on 08 Feb, 2008 11:58
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Historically, from beginning Korolev planed to use LH2 on 2th and 3th stages of N-1, but had to adjust plans to KER-LOX with a lack of time and money to build engines for it and big LH2-plant on Baikonur, also he had a strong disagree with Glushko - which fight for UDMH+NT in all heavy rockets, because this type of engine\fuel is cost effective and also can be stored\used for ICBMs. When Korolev gone, POV of Glushko to 'stinky fuel' won. So, when I said "LH2 is a too costly fuel" this meant summary prise to have it on a first stage, not the prise of pound of LH2. Even now, in the new concept they have plans to use LCH4(liquid methane) on the first stages, but not LH2.
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#9
by
pippin
on 08 Feb, 2008 13:42
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drbobguy - 8/2/2008 4:30 AM
1. Russian rockets do not run fuel rich. It appears I was wrong and that you actually get MUCH higher chamber pressures (around double) by running oxidizer-rich as opposed to fuel-rich in the staged-combustion preburner, for LOX-hydrocarbon (Note I am NOT talking about the combuster/nozzle). For LOX-hydrogen, you're better off running fuel-rich like the SSME.
2. Does anyone have any good advice to a detailed history as to why the Soviets went for LOX-hydrocarbon and never LOX-H2 in first stages? (Even Zenit/Energiya was LOX-hydrocarbon).
1. For staged combustion engines it's easy: if you run the preburner fuel rich you get gaseous kerosene which will cause massive coking in your turbines and ruin them right away, so you have to run oxy-rich. This in turn is much more challenging from a materials POV (hot oxygen will burn up almost everything), so you avoid it on LH2, where you don't have the coking issue.
2. Kerolox generally is not a bad idea for first stages (as you can see on Atlas and Saturn 5) because it has a much higher density than LH2 so your first stage can get much smaller, albeit heavier, but that doesn't matter much for a first stage. It does matter for upper stages so LH2 is the better choice here.
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#10
by
wingod
on 08 Feb, 2008 15:44
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drbobguy - 7/2/2008 9:30 PM
Russian rockets do not run fuel rich. It appears I was wrong and that you actually get MUCH higher chamber pressures (around double) by running oxidizer-rich as opposed to fuel-rich in the staged-combustion preburner, for LOX-hydrocarbon (Note I am NOT talking about the combuster/nozzle). For LOX-hydrogen, you're better off running fuel-rich like the SSME.
Does anyone have any good advice to a detailed history as to why the Soviets went for LOX-hydrocarbon and never LOX-H2 in first stages? (Even Zenit/Energiya was LOX-hydrocarbon).
It results in a more efficient launch vehicle to have a LOX/RP first stage. The booster is smaller in proportion to the overall system due to the density of the RP fuel. The russians and the Germans and the Lockheed people understand this.
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#11
by
drbobguy
on 08 Feb, 2008 18:39
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It results in a more efficient launch vehicle to have a LOX/RP first stage. The booster is smaller in proportion to the overall system due to the density of the RP fuel. The russians and the Germans and the Lockheed people understand this.
I understand there is a tradeoff in density of LH2 (and hence tank/structure volume and weight) and the higher ISP with LH2/LOX due to lighter molecules at higher speeds exiting the nozzle.
1. For staged combustion engines it's easy: if you run the preburner fuel rich you get gaseous kerosene which will cause massive coking in your turbines and ruin them right away, so you have to run oxy-rich. This in turn is much more challenging from a materials POV (hot oxygen will burn up almost everything), so you avoid it on LH2, where you don't have the coking issue.
Actually from what I read last night in
Liquid Rocket Thrust Chambers: Aspects of Modeling, Analysis, and Design this is only part of the reason (coking). The main reason is that with hydrocarbons, you can get much higher chamber pressure by running the precombustion chamber oxidizer-rich. You have to use fuel for cooling, so if you run the pre-burner fuel rich you have the problem of not being able to put much fuel through the preburner.
Maybe my understanding is wrong and someone can clarify.
From the textbook:
For oxygen and kerosene propellants, the energy release of the oxidizer-rich approach is over 400% greater than for the fuel-rich approach, within the same turbine temperature limits. The important fator, however, is the potential for this energy release to be converted into useful turbine work, which is discussed next. pg. 631
Therefore, considering equal system temperatures and peak system pressures, the oxygen-rich cycle provides a chamber pressure that is 87% higher. In practice, this ratio could be even higher because the selection of 20% of the kerosene to cool the fuel-rich cycle main chamber was very aggressive. pg. 632
But for hydrogen, you want to run fuel-rich:
The combined effect of improved energy release and improved turbine work potential more than offsets a moderate increase in required pumpwork, and results in a fuel-rich preference for staged combustion cycles if hydrogen is the fuel. pg. 632
These are completely theoretical thermodynamic and mechanical arguments that ignore materials effects (hydrogen embrittling metal or oxygen eating away at metal or coking).
I do have one last question for someone knowledgable.... in the book cited above, the calculation assumes that fuel sent to cool the chamber can't be used for the preburner. Is there a reason this is true? Why can't fuel be sent to cool the chamber/nozzle and then run through the preburner?
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#12
by
pippin
on 08 Feb, 2008 18:48
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Well, they are not completely ignoring them since they assume the same turbine temperature. But you're right, that's a strong point.
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#13
by
drbobguy
on 08 Feb, 2008 18:54
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pippin - 8/2/2008 10:48 PM
Well, they are not completely ignoring them since they assume the same turbine temperature. But you're right, that's a strong point.
I don't think you are wrong in saying that coking is a problem if run fuel-rich. I assume the combustion products of RP-1 and LOX in a fuel-rich equilibrates with much more carbon soot (since it's not clean burning). It's just that in this case, you want to run oxidizer-rich both for the coking reason, and because theoretical efficiency is much higher running oxidizer-rich.
Although like I said, this seems to depend on the assumption that fuel sent for reactive cooling cannot be used in the preburner. I don't know why that necessarily the case, but I am quite ignorant on these matters.
To me it is just very interesting that the rocket design approach in the Soviet Union was completely different than in the US, and that this problem of oxidizer-rich hydrocarbon staged combustion engines was deemed to be intractable in the US, whereas the Soviets made it work.
One more question: does staged-combustion become impractical past a certain point? I mean, the pressures are so high that there are limits to tubing diameters, etc. I'd imagine an engine the size of the F-1 could not be made staged as opposed to GG, is that correct?
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#14
by
edkyle99
on 08 Feb, 2008 18:59
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wingod - 8/2/2008 10:44 AM
It results in a more efficient launch vehicle to have a LOX/RP first stage. The booster is smaller in proportion to the overall system due to the density of the RP fuel. The russians and the Germans and the Lockheed people understand this.
NASA had this all figured out and successfully demonstrated in 1963-64, 45 years ago. The next-closest thing to a high thrust kerosene first stage is a solid propellant first stage. Thus Ares I.
BTW, the checkerboard tracking pattern on the SA-5 S-IV interstage is the inspiration for the ID image that accompanies my messages. The pattern also appeared on Corporal and Redstone missiles, among others.
- Ed Kyle
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#15
by
drbobguy
on 08 Feb, 2008 19:47
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Ed, did NASA have a staged LOX/RP-1 technology demonstrator? Or are you talking about the F-1?
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#16
by
meiza
on 08 Feb, 2008 23:33
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edkyle99 - 8/2/2008 7:59 PM
wingod - 8/2/2008 10:44 AM
It results in a more efficient launch vehicle to have a LOX/RP first stage. The booster is smaller in proportion to the overall system due to the density of the RP fuel. The russians and the Germans and the Lockheed people understand this.
NASA had this all figured out and successfully demonstrated in 1963-64, 45 years ago.
I think Saturn I fits under the umbrella of "the Germans" in wingod's comment.
The next-closest thing to a high thrust kerosene first stage is a solid propellant first stage. Thus Ares I.
BTW, the checkerboard tracking pattern on the SA-5 S-IV interstage is the inspiration for the ID image that accompanies my messages. The pattern also appeared on Corporal and Redstone missiles, among others.
- Ed Kyle
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#17
by
meiza
on 08 Feb, 2008 23:34
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drbobguy - 8/2/2008 8:47 PM
Ed, did NASA have a staged LOX/RP-1 technology demonstrator? Or are you talking about the F-1?
He meant kerosene for first stage and hydrogen for upper stages. Which the Saturn rockets had...
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#18
by
edkyle99
on 09 Feb, 2008 00:11
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meiza - 8/2/2008 6:34 PM
drbobguy - 8/2/2008 8:47 PM
Ed, did NASA have a staged LOX/RP-1 technology demonstrator? Or are you talking about the F-1?
He meant kerosene for first stage and hydrogen for upper stages. Which the Saturn rockets had...
That's right.
Going further, here is the reason that NASA was flying hydrogen upper stages as early as 1963.
http://en.wikipedia.org/wiki/Abe_Silversteinhttp://en.wikipedia.org/wiki/Silverstein_Committeehttp://www.hq.nasa.gov/office/pao/History/SP-4404/ch10-4.htmThis is also the reason that the Soviets did not fly hydrogen stages until the 1980s. The USSR did not have an Abe Silverstein. Even von Braun was averse to the idea of rapid development of hydrogen upper stages. He initially supported a Saturn B, rather than Silverstein's Saturn C. Saturn B would have used a big kerosene second stage. Silverstein's argument for rapid hydrogen development convinced von Braun.
Meanwhile, in the USSR, Glushko resisted hydrogen development to protect his storable propellant methods. Korolev focused on kerosene. Etc. No Abe Silverstein, no hydrogen.
- Ed Kyle
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#19
by
edkyle99
on 09 Feb, 2008 01:08
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Upon further consideration, I think that it would have been more correct to say that Glushko opposed hydrogen engine development based on his engineering beliefs.
- Ed Kyle
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#20
by
CFE
on 09 Feb, 2008 05:14
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simonbp - 7/2/2008 10:58 PM
The problem with (and beauty of) LH2 is that it burns so hot, but is stored so cold (20 K), requiring both high-temp metallurgy and cryogenic fuel pumps. The US had a step up because it spend a bunch of money to develop a hydrogen-powered P&W J-57 jet for the (canceled) successor to the A-12/SR-71. P&W took that technology/experience and applied to make first the RL-10, and then the J-2. The Russians decided instead to invest in improving the RP-1 rockets, and never really deployed LH2 until Energia...
Simon
Just a minor nitpick, but the hydrogen-fueled Lockheed "Suntan" actually preceded the A-12. It was the technical challenges of the CL-1200 Suntan that convinced the CIA and Skunk Works to pursue a more conventional successor to the U-2, which emerged as the Blackbird family of aircraft.
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#21
by
pippin
on 09 Feb, 2008 07:49
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drbobguy - 8/2/2008 8:54 PM
I don't think you are wrong in saying that coking is a problem if run fuel-rich. I assume the combustion products of RP-1 and LOX in a fuel-rich equilibrates with much more carbon soot (since it's not clean burning). It's just that in this case, you want to run oxidizer-rich both for the coking reason, and because theoretical efficiency is much higher running oxidizer-rich.
Although like I said, this seems to depend on the assumption that fuel sent for reactive cooling cannot be used in the preburner. I don't know why that necessarily the case, but I am quite ignorant on these matters.
I should have been more specific.
As I read your quoted article, they see an upper limit for the turbine temperature as the limiting factor. Since the efficiency of a turbo pump depends on the pressure differential, this means going lox rich (which has a way lower temp) gives you more potential (that's the 400%).
The issue with the kerosene is more or less the same. Cooling the nozzle means recovering heat energy from the nozzle and transferring it to the cooling liquid, expanding it and generate additional pressure (in an expander cycle engine it's the only power used to do that).
However, since your kerosene is not superchilled and since you don't want it to go gaseous there is quite a limit on how much energy it can absorb.
To me it is just very interesting that the rocket design approach in the Soviet Union was completely different than in the US, and that this problem of oxidizer-rich hydrocarbon staged combustion engines was deemed to be intractable in the US, whereas the Soviets made it work.
One more question: does staged-combustion become impractical past a certain point? I mean, the pressures are so high that there are limits to tubing diameters, etc. I'd imagine an engine the size of the F-1 could not be made staged as opposed to GG, is that correct?
The RD-171 is actually "bigger" (ie more thrust) than the F1. Even though it has four nozzels, it has just one set of turbomachinery, which is what counts here.
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#22
by
drbobguy
on 09 Feb, 2008 19:02
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One last question then. My understanding is the Russian designs with multiple nozzles for one turbopump (like the RD-170) are done that to avoid combustion instability issues in large chambers/nozzles. Materials and volute flow don't scale (e.g. a house fly can't be the size of a car and still fly), so a RD-170 engine with one nozzle 4x the power of an individual nozzle might not work.
Are there other motivations for splitting up one chamber/nozzle into 4 smaller ones?
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#23
by
pippin
on 09 Feb, 2008 21:42
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drbobguy - 9/2/2008 9:02 PM
One last question then. My understanding is the Russian designs with multiple nozzles for one turbopump (like the RD-170) are done that to avoid combustion instability issues in large chambers/nozzles. Materials and volute flow don't scale (e.g. a house fly can't be the size of a car and still fly), so a RD-170 engine with one nozzle 4x the power of an individual nozzle might not work.
Are there other motivations for splitting up one chamber/nozzle into 4 smaller ones?
The main problem seems to be the mixing of propellant and oxydizer in a large chamber.
"Combustion Instabilities" sounds cute but typically leads to RUD-events (Rapid, Unscheduled Disassembly, which sounds cute, too), so it's quite an issue worth avoiding. It was the main issue during F1 development.
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#24
by
hop
on 09 Feb, 2008 23:32
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drbobguy - 9/2/2008 12:02 PM
Are there other motivations for splitting up one chamber/nozzle into 4 smaller ones?
I'm not sure if this was the original intent, but it also makes building a range of engines a lot easier. The fact that the RD-170, RD-180 and RD-191 share a significant number of components (astronautix says 80% between the 170 and 180 but it's not clear what is being measured) greatly simplifies both development and production. Not only is your chamber and nozzle already proven, but you can use the same tooling as well.
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#25
by
meiza
on 10 Feb, 2008 00:08
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Back to the original thing, why does running the preburner (assuming there is only one) oxidizer rich mean higher chamber pressure?
And I think the RD-180 is propellant rich in the *chamber*. You get higher ISP that way than running stoichiometric.
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#26
by
yinzer
on 10 Feb, 2008 03:29
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drbobguy - 9/2/2008 12:02 PM
One last question then. My understanding is the Russian designs with multiple nozzles for one turbopump (like the RD-170) are done that to avoid combustion instability issues in large chambers/nozzles. Materials and volute flow don't scale (e.g. a house fly can't be the size of a car and still fly), so a RD-170 engine with one nozzle 4x the power of an individual nozzle might not work.
Are there other motivations for splitting up one chamber/nozzle into 4 smaller ones?
It makes the overall engine shorter.
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#27
by
rumble
on 10 Feb, 2008 03:43
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Has the US ever considered making engines the same way...multiple chambers for a single set of pumps? I assume not, noticing that the RS-84 was supposed to have been a single-chamber 1M lbf engine (more powerful than the RD-180)...even though the RS-84 was supposed to be our first (?) attempt at bringing an oxygen-rich staged combustion LOX/RP-1 engine to production.
In other words, even though we seemed to be following the Russians in combustion cycle, we weren't using their "more of, but smaller" chamber design.
Did we even seriously consider a multi-chamber design? What were the decision points/issues?
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#28
by
Patchouli
on 10 Feb, 2008 04:39
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yinzer - 9/2/2008 10:29 PM
drbobguy - 9/2/2008 12:02 PM
One last question then. My understanding is the Russian designs with multiple nozzles for one turbopump (like the RD-170) are done that to avoid combustion instability issues in large chambers/nozzles. Materials and volute flow don't scale (e.g. a house fly can't be the size of a car and still fly), so a RD-170 engine with one nozzle 4x the power of an individual nozzle might not work.
Are there other motivations for splitting up one chamber/nozzle into 4 smaller ones?
It makes the overall engine shorter.
Also the Russians historically did not have as powerful super computers as the Americans or as large of budgets to allow lots of ground testing to iron out these issues.
One fast way to save time is just use several smaller nozzles also the tooling is cheaper.
But there are disadvantages too such as increased plumbing complexity and making sure all combustion chambers light
For some reason they choose a very western design route with the LH2 RD120 engines on the energia core though this might have been due to a complete lack of experience with large lH2 engines on their part.
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#29
by
Patchouli
on 10 Feb, 2008 04:49
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rumble - 9/2/2008 10:43 PM
Has the US ever considered making engines the same way...multiple chambers for a single set of pumps? I assume not, noticing that the RS-84 was supposed to have been a single-chamber 1M lbf engine (more powerful than the RD-180)...even though the RS-84 was supposed to be our first (?) attempt at bringing an oxygen-rich staged combustion LOX/RP-1 engine to production.
In other words, even though we seemed to be following the Russians in combustion cycle, we weren't using their "more of, but smaller" chamber design.
Did we even seriously consider a multi-chamber design? What were the decision points/issues?
I think the American designers just figure a single large nozzle and combustion chamber would give better performance mostly in thrust to weight and ISP.
BTW I'd like the see the RS-84 program restarted as new f1 class liquid fueled engine esp a reusable one would be a very nice asset to have in our space program and would allow us to get away from large SRBs on crewed vehicles.
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#30
by
rumble
on 10 Feb, 2008 15:06
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Patchouli - 9/2/2008 11:49 PM
BTW I'd like the see the RS-84 program restarted as new f1 class liquid fueled engine esp a reusable one would be a very nice asset to have in our space program and would allow us to get away from large SRBs on crewed vehicles.
While I don't disagree, I'm afraid we're decade(s) away from such an eventuality. Nice thought, but ain't gonna happen. For a more full explanation, search this site to get pages of discussion on precisely what you mention.
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#31
by
Jim
on 10 Feb, 2008 16:50
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Patchouli - 10/2/2008 12:49 AM
BTW I'd like the see the RS-84 program restarted as new f1 class liquid fueled engine esp a reusable one would be a very nice asset to have in our space program and would allow us to get away from large SRBs on crewed vehicles.
wishing isn't going to make it happen, it is going to take big $'s, which NASA doesn't have
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#32
by
edkyle99
on 10 Feb, 2008 16:54
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yinzer - 9/2/2008 10:29 PM
drbobguy - 9/2/2008 12:02 PM
One last question then. My understanding is the Russian designs with multiple nozzles for one turbopump (like the RD-170) are done that to avoid combustion instability issues in large chambers/nozzles. Materials and volute flow don't scale (e.g. a house fly can't be the size of a car and still fly), so a RD-170 engine with one nozzle 4x the power of an individual nozzle might not work.
Are there other motivations for splitting up one chamber/nozzle into 4 smaller ones?
It makes the overall engine shorter.
In addition, use of multiple nozzles provides roll control authority, which one nozzle cannot provide.
- Ed Kyle
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#33
by
edkyle99
on 10 Feb, 2008 17:03
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rumble - 9/2/2008 10:43 PM
Has the US ever considered making engines the same way...multiple chambers for a single set of pumps?
Some of the Atlas MA-5(A) propulsion systems had twin booster engines that shared a turbopump.
- Ed Kyle
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#34
by
rumble
on 10 Feb, 2008 17:11
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edkyle99 - 10/2/2008 12:03 PM
rumble - 9/2/2008 10:43 PM
Has the US ever considered making engines the same way...multiple chambers for a single set of pumps?
Some of the Atlas MA-5(A) propulsion systems had twin booster engines that shared a turbopump.
- Ed Kyle
Oh. Whoops! I even looked this up on astronautix before asking, but I missed this text on the XLR-89-5 engine (emphasis added):
Designed for booster applications. Gas generator, pump-fed. Shared turbopumps for booster engines.
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#35
by
Damon Hill
on 10 Feb, 2008 18:32
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To go right to the source for many Russian engines, see Khimavtomatiki (Chemical Automation?):
http://www.kbkha.ru/?lang=en (top page)
http://www.kbkha.ru/?p=8&cat=3 (engines developed for space flight, sidebar for other types)
They have extensively redesigned their web page, which can be a small gold mine of information
and history.
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#36
by
Damon Hill
on 10 Feb, 2008 19:04
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#37
by
CFE
on 10 Feb, 2008 19:08
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Jim - 10/2/2008 10:50 AM
Patchouli - 10/2/2008 12:49 AM
BTW I'd like the see the RS-84 program restarted as new f1 class liquid fueled engine esp a reusable one would be a very nice asset to have in our space program and would allow us to get away from large SRBs on crewed vehicles.
wishing isn't going to make it happen, it is going to take big $'s, which NASA doesn't have
If we see a new LOX-Kerosene engine in the same ballpark as the F1, it will probably be built with Elon Musk's money. If Falcon IX goes operational, it would probably be worthwhile to replace the nine-Merlin cluster with a Big "Falcon" Engine.
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#38
by
meiza
on 11 Feb, 2008 00:01
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Thanks Damon, didn't know there was a storable propellant "Russian F-1". Although staged combustion, with high pressure and ISP.
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#39
by
edkyle99
on 11 Feb, 2008 02:05
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CFE - 10/2/2008 2:08 PM
Jim - 10/2/2008 10:50 AM
Patchouli - 10/2/2008 12:49 AM
BTW I'd like the see the RS-84 program restarted as new f1 class liquid fueled engine esp a reusable one would be a very nice asset to have in our space program and would allow us to get away from large SRBs on crewed vehicles.
wishing isn't going to make it happen, it is going to take big $'s, which NASA doesn't have
If we see a new LOX-Kerosene engine in the same ballpark as the F1, it will probably be built with Elon Musk's money. If Falcon IX goes operational, it would probably be worthwhile to replace the nine-Merlin cluster with a Big "Falcon" Engine.
It is a shame that NASA can't buy RD-171 engines. Just one of those attached to a kerosene/LOX first stage topped by a smaller version of the Ares I upper stage could easily lift Orion to ISS or for lunar missions. A rocket of this type would weigh less and be a lot smaller than either Atlas V Heavy or Delta IV Heavy at liftoff. I suppose the improved Atlas variants looked a bit like this, with twin RD-180 engines but without a J-2X powered upper stage.
A gas generator kerosene first stage would need more thrust, perhaps 800 tonnes force (17-18 Merlin 1C engines or 9 RS-27As!) but would still produce a rocket weighing more than 300 tonnes less than Ares I at liftoff. An ideal configuration IMO would be four chambers/nozzles producing 200 tonnes thrust each at liftoff. Rocketdyne worked on an engine much like this during the late 1950s. E-1, I think it was called.
Bucks, I suppose. The SRB infrastructure exists. The U.S. kerosene propulsion would have to be created from scratch.
- Ed Kyle
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#40
by
meiza
on 11 Feb, 2008 11:49
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I don't think developing a decent performance RD-180 sized gas generator kerosene engine would take *that long*. Maybe 5 years. But it would be a relatively sure thing to produce something, it's known territory. Of course, its performance would suck compared to the RD-180. So it would have to have some other advantages to be able to exist...
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#41
by
William Barton
on 11 Feb, 2008 12:15
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There's been considerable discussion of how single-engine stages provide better LOM stats for LVs. is there a comparison between single-chamber vs. multichamber engines in that regard?
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#42
by
Jim
on 11 Feb, 2008 12:31
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William Barton - 11/2/2008 8:15 AM
There's been considerable discussion of how single-engine stages provide better LOM stats for LVs. is there a comparison between single-chamber vs. multi chamber engines in that regard?
multi chamber would allow for roll control The differences between an RD-180 and RS-68 as far as nozzles would be 1 less actuator (RD-180 has 2 each pitch and yaw, RS-68, has 1 each pitch, yaw and roll). this would be mean an RS-68 nozzle setup would be more reliable. Now take an SSME type engine: single nozzle and no generator output available for roll control. This means an additional roll control system would be needed. All these comparisons assume single engine installations.
Bottom line, it has to be evaluated at propulsion system level and not at the engine level
