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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
