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  #11  
Old 04-09-2015, 02:46 PM
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On a somewhat related and serious note, the Smithsonian might possibly get the last production Delta II. There are three left, two are scheduled to launch, but there is no interest in the last one. Pads 17a and 17b have been decommissioned in Florida and the only active Delta II pad remaining is at Vandenberg, but nobody else wants a Delta II sized payload in polar orbit. Employees have volunteered to come in on their own time and build up the last Delta II for display if no one ends up buying it for a flight in the next year or two.

It has flown 98 flawless missions in a row and has only two failures out of 153 flights since the late 80's. That's darn impressive and deserving of a permanent display if it doesn't get used.
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Old 04-10-2015, 07:47 PM
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I wonder why we have to use Russian engines when those work so well?
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Old 04-10-2015, 10:34 PM
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Quote:
Originally Posted by bob jablonski
I wonder why we have to use Russian engines when those work so well?
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Because they're much smaller and for some reason they didn't want to use a cluster...

Later! OL JR
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Old 04-10-2015, 11:20 PM
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Quote:
Originally Posted by luke strawwalker
Because they're much smaller and for some reason they didn't want to use a cluster...

Later! OL JR

Yet the chosen Atlas RD-180 replacement is to be a pair of Blue Origin BE-4 fart burners, still under development. Once available, you may be able to purchase them on Amazon with free Prime shipping.
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Old 04-11-2015, 12:31 AM
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Quote:
Originally Posted by tbzep
Yet the chosen Atlas RD-180 replacement is to be a pair of Blue Origin BE-4 fart burners, still under development. Once available, you may be able to purchase them on Amazon with free Prime shipping.


LOL

How many RS-27's would it take to equal a single twin-chambered RD-180, or even the pair of Bezos "fart burners"?? I think that would answer the question, but I'm tired and need to hit the sack-- no time to look that up myself... let someone else do their own homework.

I think they like the Bezos engines because basically they'll fit in the "same holes" as the twin chambered RD-180... (more or less). Sure they'll have their own turbomachinery for each chamber, unlike the shared turbomachinery on the RD-180, but that's a minor quibble... I'm sure there'll be SOME modification to the thrust structure, plumbing, etc. for the "fart burners" but it's probably a lot less than would be required to add a big cluster of RS-27's to do the same job.

Remember, broadly speaking, that the RS-27 is an outgrowth of the old Saturn IB H-1 engine... where the RD-180 is more an updated, complicated combustion cycle competitor to the venerable old gas generator F-1...

Surely got to be easier to stick a pair of "Bezos fart burners" under the existing (but no doubt will be modified) Atlas V than building an all-new thrust structure to hold a half dozen or more RS-27's...

Later! OL JR
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Old 04-11-2015, 09:01 AM
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Quote:
Originally Posted by luke strawwalker
LOL

How many RS-27's would it take to equal a single twin-chambered RD-180, or even the pair of Bezos "fart burners"?? I think that would answer the question, but I'm tired and need to hit the sack-- no time to look that up myself... let someone else do their own homework.

I think they like the Bezos engines because basically they'll fit in the "same holes" as the twin chambered RD-180... (more or less). Sure they'll have their own turbomachinery for each chamber, unlike the shared turbomachinery on the RD-180, but that's a minor quibble... I'm sure there'll be SOME modification to the thrust structure, plumbing, etc. for the "fart burners" but it's probably a lot less than would be required to add a big cluster of RS-27's to do the same job.

Remember, broadly speaking, that the RS-27 is an outgrowth of the old Saturn IB H-1 engine... where the RD-180 is more an updated, complicated combustion cycle competitor to the venerable old gas generator F-1...

Surely got to be easier to stick a pair of "Bezos fart burners" under the existing (but no doubt will be modified) Atlas V than building an all-new thrust structure to hold a half dozen or more RS-27's...

Later! OL JR


To elaborate on this, you couldn't stuff enough Delta II engines in the Atlas V to replace the RD-180 without extreme modifications. It would take about 3 1/2 R-27's to equal the RD-180, so four engines would have to be stuffed into what basically would end up being a new airframe.

It will take about 1.5 BE-4 fart burners to equal the thrust of the RD-180, , assuming their commercial production form meets their projections of 550,000 lbs. On the surface it looks like the BE-4 pair would give better performance (1.1 million lbs compared to the RD-180's 860,500 lbs), but I haven't looked at the comparison of fuel density and other factors that make the kero/lox RD-180 such a good performer on the Atlas first stage. It's possible that the Atlas would have to be stretched to carry enough fuel to equal the current Atlas' lifting capabilities.

It doesn't make much difference right now, as the BE-4 is still several years away. We still have a few RD-180's. By the time we run out, our issues with Russia may be resolved, or we may have achieved the concept of M.A.D. With the planned phase out of the Delta IV (except the Heavy), we'd better speed things up a little. The Delta IV could be used instead of the Atlas V on some payloads, but it's more violent acoustic and thrust profiles don't promote it as a complete replacement for the Atlas.
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Old 04-13-2015, 01:47 AM
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Quote:
Originally Posted by tbzep
To elaborate on this, you couldn't stuff enough Delta II engines in the Atlas V to replace the RD-180 without extreme modifications. It would take about 3 1/2 R-27's to equal the RD-180, so four engines would have to be stuffed into what basically would end up being a new airframe.

It will take about 1.5 BE-4 fart burners to equal the thrust of the RD-180, , assuming their commercial production form meets their projections of 550,000 lbs. On the surface it looks like the BE-4 pair would give better performance (1.1 million lbs compared to the RD-180's 860,500 lbs), but I haven't looked at the comparison of fuel density and other factors that make the kero/lox RD-180 such a good performer on the Atlas first stage. It's possible that the Atlas would have to be stretched to carry enough fuel to equal the current Atlas' lifting capabilities.

It doesn't make much difference right now, as the BE-4 is still several years away. We still have a few RD-180's. By the time we run out, our issues with Russia may be resolved, or we may have achieved the concept of M.A.D. With the planned phase out of the Delta IV (except the Heavy), we'd better speed things up a little. The Delta IV could be used instead of the Atlas V on some payloads, but it's more violent acoustic and thrust profiles don't promote it as a complete replacement for the Atlas.


Some quick comparisons...

Property..................................Kerosene .........................................Liquid Methane
Density....................................0.81-1.02 gm./CC...........................0.4226 gm/CC
Heat Content...........................43.1-46.2 MJ/kg..............................55.5 MJ/kg
Rocket engine max ISP...........~353 seconds..................................~363 seconds
Typical engine ISP (vac)........383 seconds (RD-180)......................363 sec. (Raptor)
........................(sea level).......311 seconds (RD-180)......................321 sec. (Raptor)
Fuel/oxy ratio (optimum).......2.72:1 (RD-180)................................~2.8:1

So, all things being equal, the thing that immediately jumps out is that LNG is about HALF the density of RP-1 kerosene... but the density of liquid hydrogen, by comparison, is 0.0708 gm/CC, or about 1/6 that of LNG... the energy content is somewhat higher, so technically speaking, that means slightly less fuel would be necessary to deliver the same amount of thermal energy, roughly speaking, and the specific impulse is also slightly higher (I couldn't get exact figures on the BE-4 engine, so I took the numbers I COULD find-- for SpaceX's Raptor engine-- which is a 1,000,000 lb thrust engine, versus the BE-4 which is designed for 550,000 lbs of thrust (roughly equivalent to SSME). The ISP for methane, at least on the drawing board, is slightly higher than for RP-1 at sea level, and slightly lower than kerosene in vacuum. For sake of argument here, we could probably call it a wash. The mixture ratios are also broadly similar as well, so the oxidizer tank size should be unaffected...

SO, we could expect to see the Atlas V get fuel tanks for LNG about 2X the size of the current RP-1 tanks... roughly speaking. Performance should be about the same, depending on the engine. The BE-4 sales literature (PR brochures) don't mention the expected specific impulse of the engine, but do state that it's going to be an advanced high pressure staged-combustion cycle engine like the RD-180... what that means in performance terms versus RD-180, which is probably the most highly advanced kerosene engine flying today, is a bit conjectural... but I'd say they'd be broadly similar.

So larger propellant tanks would seem necessary for an LNG powered Atlas V versus a kerosene powered one, based strictly on the difference in density alone... There was the Atlas V Phase 2 plans about 10-12 years ago, which would have phased out the current 3.81 meter (12.5 feet) diameter tankage in favor of building the Atlas V with reproportioned tanks built using the Delta IV's 5 meter (16.4 feet) diameter tanks... considering the density differences (and thus tank volume required) for kerosene propellant versus liquid hydrogen, the Atlas V Phase 2 would have been the same diameter at Delta IV, but much shorter. Using the same tooling as the LH2 powered Delta IV, an LNG powered Atlas V Phase 2 successor would be taller than the kerosene version, but shorter than the Delta IV. This would make both rockets on common tooling, cutting costs. The plan for Atlas V Phase 2, however, wasn't strictly to make the same rocket as Atlas V using the Delta IV tooling, but to add a second RD-180 (for four total thrust chambers (nozzles) thus increasing the capability of the rocket to almost Delta IV-Heavy class performance. Presumably the same thing could be achieved (broadly similar anyway) using a cluster of four BE-4 LNG engines.

Now, if they strictly want to keep the Atlas V as-is, all they really need to do is a tank stretch and probably tweak the propellant line sizing, etc. for LNG versus kerosene. LNG is kept at -260 degrees at 3.6 psi, versus kerosene at room temperature and atmospheric pressure. Of course this isn't really TOO much different from liquid oxygen, which is -297 degrees... this is of course much easier to insulate for than the difference between LO2 and kerosene at room temperature (to prevent excess LO2 boiloff and freezing of the kerosene into a solid) or LH2 at -423 degrees F... (to prevent excess LH2 boiloff from heat leaking in from the LO2 tank). One big benefit to LNG over hydrogen, other than the obvious density (and thus tank volume) differences, is that the higher liquid temperature of LNG vs. LH2 means that simple nitrogen gas pressurization and purge systems can be used versus requiring helium systems for liquid hydrogen propellant systems. Nitrogen systems are of course usually used for kerosene rockets. Also, LNG is a larger molecule and readily handled in industry-- much easier to store and eliminate leaks than the much smaller and much colder LH2 molecules which tend to leak like a sieve and are rather exotic to handle (problems with liquified air, seals, embrittlement, etc.) LNG should be no harder to handle in bulk than liquid oxygen, and much easier than liquid hydrogen.

Well, that's just off the top of my head...

Later! OL JR
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  #18  
Old 04-13-2015, 01:57 AM
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Found this on the net... might be of interest... it's about the SpaceX Raptor methane engine versus the Merlin D kerosene engine, but the same arguments apply between the BE-4 methane burner versus the RD-180 kerosene engine...


"Methane (CH4) and RP-1 are roughly equivalent in realizable performance. As previously mentioned by other posters, CH4 has slightly higher impulse – about 370 s in vacuum vs the 360 s – at the same chamber pressure of 7 MPa. But, this is counterbalanced by its lower bulk density of about 830 kg/m3 vs about 1030 kg/m3. Bulk Density is the density of the combined Fuel and Oxidizer load in their appropriate ratios. Even though Methane is "only" 430 kg/m3 it is burned with 3.5 parts oxygen compared to 2.1 parts for RP-1, hence a CH4 rocket will be carrying more oxygen and less fuel by weight. Oxygen is pretty dense at a little over 1140 kg/m3 which is denser in fact than RP-1 (about 810 kg/m3). If we assume that chamber pressures and engine cycle efficiency will be equal, RP-1 outperforms CH4 simply because a 20% larger tank will impose weigh penalties that slightly outweigh the 3% increase in specific impulse. However, the RP-1 advantage is contingent upon operating at an equal chamber pressure which may not be the case. And, Methane (CH4) has additional advantages that are applicable in specific scenarios.

The reasons CH4 is a front runner for SpaceX's Raptor can probably be attributed to four factors:

Methane does not coke (polymerize) at the operating temperatures of a rocket engine – it's coking point is roughly twice as high. This makes it easier to make an engine reusable and re-usability is a key SpaceX objective.

Because Methane does not coke, it is also easier to implement a full-flow stage combustion (FFSC) cycle where all the fuel and oxidizer flow goes through the pre-burner. Compared to contemporary Russian partial flow stage combustion engines higher chamber pressures are attainable resulting in a total impulse advantage of about 30 seconds, or 9%. This eliminates the performance deficiency of CH4 compared to RP-1.

If SpaceX intends to use the same fuel in all the stages, CH4 can be considered a better upper stage fuel and a worse lift-off fuel, even without enabling higher working pressures. This is because upper stages are typically 1/8th to 1/10th the size of the 1st stage, and here impulse is more important than density. Using Methane with the aforementioned FFSC cycle means that SpaceX can potentially get equivalent 1st stage performance and better upper stage performance.

Even though it is, IMHO, somewhat dubious that early Mars mission will use in-situ fuel production. If that ever becomes an applicable practice, Methane can be produced from water and CO2 while RP-1 cannot.

Other than that, there is the non-factor that somewhat favor Methane, such as regular grade Natural Gas being good enough and not having to highly refine the fuel from regular kerosene to RP-1 to achieve low coking characteristics and consistent densities. I say it is a non-factor, because fuel cost is such a negligible part of launch costs that it really doesn't matter if fuel cost a few times more or less. Fuel is typically only about 0.3% of the cost of flying a rocket to orbit, so fuel cost really doesn't matter – Not even when you compare highly expensive propellant combos like Hydrazine/Tetroxide to the relatively cheap Kerosene/Oxygen.
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edited Jun 30 '14 at 2:25


answered Jun 27 '14 at 8:23
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  #19  
Old 04-13-2015, 02:01 AM
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Found this too... more of the same...




"A good question. In pre-EELV studies, NASA and the U.S. Air Force looked at LOX/methane. EELV resulted in the LOX/kerosene Atlas V and the LOX/hydrogen Delta IV.

At the 4th International Conference on Launcher Technology in 2002, Burkhardt et al. compared a reusable LOX/kerosene launch vehicle using the RD-180 type engine from the Atlas V with a LOX/methane vehicle using a possible engine of the same efficient staged combustion cycle:

The LOX/methane engine had about 3% higher specific impulse but that advantage was outweighed by the lower density of the liquid methane compared to kerosene.

LOX/kerosene was slightly better performing overall in terms of payload and expected to be lower cost to build and operate, the same result as the pre-EELV studies.

The reason LOX/hydrogen is comparable to or better than LOX/kerosene is that the specific impulse is much higher overcoming the even lower density problem. For the Space Shuttle the main engines operated from ground to orbit so the higher specific impulse of hydrogen at higher altitude was the reason for its use.

For a first stage that only operates to low altitud followed by a LOX/hydrogen second stage as in the Atlas and Delta, kerosene has comparable payload performance and may be lower cost because of vehicle size. For the Delta IV another advantage is commonality with the upper stage propellants.

Methane is not currently supplied at launch sites so a major facility investment would be needed.

Lack of long experience with operation is another negative for methane.

If the Raptor were to be used in space as in a Mars mission then the fact that both LOX and liquid methane are relatively easy to store in space compared to hydrogen or kerosene would be an advantage.
References:

Comparative Study of Kerosene and Methane Propellant Engines for Reusable Liquid Booster Stages. Holger Burkhardt, Martin Sippel, Armin Herbertz, Josef Klevanski. 4th International Conference on Launcher Technology "Space Launcher Liquid Propulsion" 3-6 December 2002 – Liège (Belgium).

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edited May 29 '14 at 8:00
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answered May 29 '14 at 6:19
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Old 04-13-2015, 03:27 AM
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Firefly Space Systems (see: http://www.fireflyspace.com/ ), which was founded by a SpaceX alumnus, is also using LOX/methane in their two-stage vehicles, which will utilize autogenous (self-pressurizing) pressure-fed engines. The first stage engine (as well as the outboard booster engines of the larger of their two designs) is an altitude-compensating aerospike engine, which suggests that they too have reusability in mind for future variants. Also:

You beat me to it: Methane does not coke (unlike kerosene), which makes it very attractive for reusable engines. As with ion engines (which were an attractive option that no one wanted to try for spacecraft main propulsion because no one had done it before, until NASA's Deep Space 1 mission finally "broke its maiden"), methane engines have enough advantages that it's past time to "retire the risk" by trying them on flight vehicles. One of the reasons why the X-15's XLR-99 reusable "big engine" used anhydrous ammonia (which is an atrocious chemical to work with) was because it, too, does not coke; methane is a much friendlier alternative to ammonia.
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