Dedicated SLS Block 1B Thread

[MP99]

Well, as I understand it (which could be wrong) the J2X is really a 2nd stage engine powerful enough and designed to do a significant amount of the ascent through the atmosphere.

As I understand it, the vehicle is pretty much in vacuum by the time the J-2 kicks in, but it is still fighting gravity losses, so velocity rather than atmosphere seems to be the variable. Remember that the S-IV on Saturn I had 6 RL-10s with a total of 90k lb. thrust and ISP of 410. The Saturn I could put 19,800 lb. in LEO. The S-IV was replaced with the S-IVB which had a single J-2 @ 232K lb. thrust @ ISP of 421 and could place 41,000 lb. in LEO. On Saturn I-B the single J-2 on the S-IVB was able to place well more than twice the payload to LEO as 6 RL-10s on the S-IV of the Saturn I.

Even on Saturn V, the S-IVB still had to contribute a Delta Vee of over 3000 mph to reach temporary parking orbit before its restart as an EDS, so during that first burn, it was fighting more severe gravity losses than during its TLI burn.

I think the main issue on the Saturn I-B version of the S-IVB is that as an US, its  purpose was to reach LEO, not to be an EDS. The single J-2 was able to fight gravity loss much better than the combined 6 RL-10s. On lunar mission Saturn Vs, if the S-II had been able to place the payload into LEO, perhaps the third stage could have employed an RL-10 as that EDS would not have to fight gravity loss very much during TLI.

I agree (as above). Although I'd think SLS's core would burnout higher & faster than S-I (hydrolox 1.5 stage vs kerolox first stage), 1BUS is a higher payload with fewer RL-10s, so the same principles should apply.

It seems to me that what has not been well defined on SLS is whether a second stage would be used

A.) solely for reaching LEO.
B.) solely as an EDS.
C.) as a restartable stage which has a substantial first burn to reach LEO prior to a 2nd (Earth Departure) burn.
D.) as a restartable stage which has a short first burn for circularization prior to a 2nd (Earth Departure) burn.

If the objective is A or C, then the J-2X (ISP 448) would seem better in fighting gravity losses. If the objective is B or D, then the RL-10B2 (ISP 462) has a 14 sec ISP advantage.

Block 2 is very clearly option (A) - a core stage with advanced boosters and a 2x or 3x J-2X second stage makes a single burn, leaving the payload to perform it's own apogee/circ burn. Ironically, the second stage role would also be ideal for an air-start RS-25. Single burn so no restart required, and commonality with the core engine.

The in-space-only CPS with RL10s then becomes option (D) - a third stage performing circ burn & EDS.

ISTM the direction of the program comes down to some extent to an argument between chemical propulsion or SEP for moving stuff BLEO. If SEP is chosen, then they'd just want SLS to max it's payload to LEO, which is the block 2 config.

I have to say I prefer option (C) - block 1B, maximised by ballistic trajectory to EML. It avoids the need to develop separate upper stage & EDS, or upper stage & big SEP stage.

cheers, Martin

[USFdon]

This was to the detriment of the engine's use on the Ares V US. Again, if memory is correct, when CxP was cancelled, but J-2X kept alive, there was some discussion about reversing those parameters to that in its use primarily as an Earth Departure engine, J-2X could be pushed back to its earlier specs with a higher ISP. If anyone has details on this, please advise.

Is this the J-2x-285 variant that was mentioned in the RP-1 section in HEFT or was that just a hypothetical higher thrust variant?

[Hyperion5]

Elsewhere

Ed Kyle has made a compelling case for an SLS derived launcher with a single J-2X powering an appropriately sized middle stage topped by a DIVH derived iCPS. For a single launch that's going to be difficult to beat. The engine count is low (leading to high reliability) and the performance would be darn awful good.

I'll agree that the performance would be excellent.  However the reliability vs the 4-engine CPS currently envisioned on the SLS Bloc IB is likely to be lower. 

Reasons

1) Serial staging risk--adding another stage always ups the risk of something going wrong in stage separation, although for major US LVs recently staging failures have been a fairly rare occurrence. 

2) The real reason why the the Bloc IB as it stands now would be more reliable with 3 additional engines on the CPS is engine-out redundancy.  It's true that single engine stages can be fairly reliable because they eliminate a lot of risk that comes with extra engines.  However, ULA has actually done the math on this.  If you factor in losing an engine on the CPS, and I imagine the SLS engineers have, the risk of engine-related stage failure drops to almost 1/16th that of the single RL-10 stage.  That's better than an order of magnitude improvement in reliability, which has to look attractive to NASA. 

1 RL-10 engine stage: 79 stage failures per 100,000 flights
4 RL-10 engine stage: 5 stage failures per 100,000 flights

This is based on ULA's study of the Atlas V Phase 2 (see page 5): http://www.ulalaunch.com/site/docs/publications/EELVPhase2_2010.pdf

Well, as I understand it (which could be wrong) the J2X is really a 2nd stage engine powerful enough and designed to do a significant amount of the ascent through the atmosphere.

As I understand it, the vehicle is pretty much in vacuum by the time the J-2 kicks in, but it is still fighting gravity losses, so velocity rather than atmosphere seems to be the variable..............

Even on Saturn V, the S-IVB still had to contribute a Delta Vee of over 3000 mph to reach temporary parking orbit before its restart as an EDS, so during that first burn, it was fighting more severe gravity losses than during its TLI burn.

The SLS in contrast to the Saturn V is pretty much in vacuum and I might add that the core stage is nearly at orbital speed when stage separation would occur.  This means that unless you're lugging up a massive payload which would cause lower stage separation, there isn't going to be a huge benefit to adding J-2X engines. 

Of course any discussion of an upper stage must consider what booster is being used: 5-seg solid, ATK adv. SRB, AJ adv. LRB, Dynetics adv. LRB. ATK just said their advanced SRB could not meet the high profile Mars objectives unless the core uses 5 RS-25Es. The latest from NASA was an intent always to use 4 RS-25s on the core. The more powerful Dynetics booster may have the ability to make any US purely an EDS, thereby enabling a low number of RL-10B2s.

So basically the Bloc II SLS could simply be an SLS Bloc IB with Dynetics or Aerojet boosters and still be hugely capable.  That's a very nice evolution path for NASA given how much less expense it'd require versus developing an enlarged core or an all-new stage.  Perhaps ATK will have more difficulties selling NASA on advanced SRBs than I thought. 

[Lobo]

So basically the Bloc II SLS could simply be an SLS Bloc IB with Dynetics or Aerojet boosters and still be hugely capable.  That's a very nice evolution path for NASA given how much less expense it'd require versus developing an enlarged core or an all-new stage.  Perhaps ATK will have more difficulties selling NASA on advanced SRBs than I thought. 

Yea, to be honest, if or when there is a Block 2, this is what I'd expect it to look like.  I'm pretty sure NASA is designing at least Block 1A or 1B loads into the block 1 core so they don't need to upgrade the core itself for either of those upgrades.
It might be that they design all the way to block 2 loads into it too, or at least design it so that strengthening the core for the block 2 loads would be a relatively easy upgrade when/if that becomes a reality.

If Block 2 is just Block 1B with Advanced boosters, and doesn't have the additional LUS, then Block 2 really shouldn't have any additional bending loads due to a longer length.  So strengthening then, or designing it into Block 1 might not be too bad.  Just enough to handle the greater mass in the PLF, but the aerodynamic loads could be pretty similar if the core's not longer.

I would think/hope that this option is seriously looked at.  If the boosters have enough performance, then you don't need the J2X powered 2nd stage, or a 5th RS-25 on the core.  Since there is no existing advanced booster, SRB or LRB, then NASA could figure out what they need the booster to do to get 130mt to LEO with no or little help from the block 1B stage, and put that into the RFQ and let the companies meet that performance.   Sounds like Dynetics would be able to.  And I'd guess Aerojet could too because they could just design their offerings with enough AJ-1000 engines too get the necessary performance.
That's the advantage of LRB's, they can be designed as necessary.

ATK's SRB offering might be a little trickier though. 
I'm wondering if ATK has the ability to upsize their advanced solid if it doesn't have enough performance to meet a larger advanced booster requirement?  Would they be constrained by their existing infrastructure and tooling to handle solid segments? 

If that's the case, that would give LRB's an advantage if NASA were to require the extra booster performance.  And that could make "Block 2" not such a costly version after all.   If Block 1B is chosen, -something- will have to be done after the existing stock of steel SRB casings is flown out.  Either restart steel casing production, or upgrade to advanced solids or liquids.
Advanced solids would have the advantages of likely being cheaper to switch to from an upgrade to existing hardware standpoint.
But if they can't get 130mt to LEO, and NASA decides they want that performance from the booster, Advanced liquids might be a real boost, pardon the pun.
;-)

[sdsds]

the reliability vs the 4-engine CPS currently envisioned on the SLS Bloc IB is likely to be lower

Your reasoning is good, but there are some constraints imposed by the "3 of 4" approach. Versus J-2X, it looks like a difficult trade to conduct, particularly without something like a specific "mass through TLI" requirement.

the core stage is nearly at orbital speed when stage separation would occur. 

True for Block I. But for Block IB? The attached (crudely drawn) map shows an alternative. Launching to an orbit with a 28.5 degree inclination the Block I core might end up in the Pacific, as did the external tanks for Shuttle launches to Hubble. That's shown as the red marker with the square. I assert without proof that for Block IB, mass sent through TLI would be greater if the core were dropped in the Indian Ocean (plain red marker) with some CPS propellant used to reach orbit.

[Hyperion5]

the core stage is nearly at orbital speed when stage separation would occur. 

True for Block I. But for Block IB? The attached (crudely drawn) map shows an alternative. Launching to an orbit with a 28.5 degree inclination the Block I core might end up in the Pacific, as did the external tanks for Shuttle launches to Hubble. That's shown as the red marker with the square. I assert without proof that for Block IB, mass sent through TLI would be greater if the core were dropped in the Indian Ocean (plain red marker) with some CPS propellant used to reach orbit.

Well obviously Bloc IB is going to stage lower and slower than Bloc I due to having a bigger stage up top and probably a bigger payload.  I would think the Indian Ocean would be a quite reasonable place for the core to drop.  Add some LRBs onto the Bloc IB to make it the Bloc 2 and we may see the core drop into the Pacific once again unless it's a particularly heavy payload.

[HappyMartian]

the core stage is nearly at orbital speed when stage separation would occur. 

True for Block I. But for Block IB? The attached (crudely drawn) map shows an alternative. Launching to an orbit with a 28.5 degree inclination the Block I core might end up in the Pacific, as did the external tanks for Shuttle launches to Hubble. That's shown as the red marker with the square. I assert without proof that for Block IB, mass sent through TLI would be greater if the core were dropped in the Indian Ocean (plain red marker) with some CPS propellant used to reach orbit.

Well obviously Bloc IB is going to stage lower and slower than Bloc I due to having a bigger stage up top and probably a bigger payload.  I would think the Indian Ocean would be a quite reasonable place for the core to drop.  Add some LRBs onto the Bloc IB to make it the Bloc 2 and we may see the core drop into the Pacific once again unless it's a particularly heavy payload.

Having an engine out capability to reduce mission risks is reasonable, but what is being ignored is the eventual significant potential benefits of the LRBs and/or propellant tanker/depot to increase the size and therefore thrust requirements of the of the CPS.

An SLS Block 1B with LRBs coming and using propellant tankers/depot is going to become a different and much more capable launcher than the standard J-241SH or J-246SH. The CPS could use almost all of its propellant in getting into LEO.

A J-2X doesn't give you engine out capability and has a normal two starts and a maximum of only four engine starts - five starts would require a new ignition system.

The NGE isn't powerful enough unless you go to eight to ten engines and that would be a bit crowded.

The most suitable engine seems to be the MB-60. There doesn't seem to be any logical reason to not use four to five MB-60 Cryogenic Upper Stage Engines. Japan does need to make a real and significant contribution to SLS Lunar and other missions and it has the most suitable engine.


"Assume a refuelable CPS in LEO. This will mitigate the impact of propellant losses for missions with long wait times before Earth departure" Page 20

"Propulsion"  "Next Generation Engine (NGE)" "MB-60 Engine" Page 25

"CPS Gamma" with "5x MB-60" Page 29

From: A Study of CPS Stages for Missions beyond LEO   By Mark Schaffer May 16, 2012
At: http://spirit.as.utexas.edu/~fiso/telecon/Schaffer_5-16-12/Schaffer_5-16-12%20Rev%20A.pdf

Also at: http://www.sei.aero/eng/papers/uploads/archive/SpaceWorks%20CPS%20Study%20Final%20Distribution.pdf


I added the bold.



"MHI and PWR successfully assembled and tested the first
MB-XX demonstrator engine in the summer of 2005. Testing
was performed at the Tashiro Field Laboratory located in
northern Japan."

"These data confirm that the demonstrator engine
operated much as expected and confirm that the design can be
operated as predicted at the 267 kN (60 Klbf) thrust condition."

From: The MB-60 Cryogenic Upper Stage Engine - A World Class Propulsion System  By William Sack, Kenji Kishimoto, Akira Ogawara, Kimito Yoshikawa, and Masahiro Atsumi    2009
At: http://archive.ists.or.jp/upload_pdf/2009-a-03.pdf 


Edited. 

[sdsds]

Well obviously Bloc IB is going to stage lower and slower than Bloc I due to having a bigger stage up top and probably a bigger payload.

"CPS Gamma" with "5x MB-60" Page 29

From: A Study of CPS Stages for Missions beyond LEO   By Mark Schaffer May 16, 2012

Schaffer's work is great, but his CPS Gamma is going to be a long way out on the timeline. As you suggest, he assumes something else put a bunch of propellant into orbit! Back as 2011 others at NASA suggested an iCPS-2. It would carry 57.8 tonnes of usable propellant, compared with iCPS-1 which is reported as carrying 27.2 tonnes.

If I understand that plan correctly, the core would deliver iCPS-1 and payload into a -87 km x 241 km target orbit. After coasting to apogee, iCPS-1 would provide the 100 m/s orbit circularization delta-v.

In contrast, I'm estimating the core could deliver iCPS-2 and payload into an orbit that requires a circularization delta-v of about 700 m/s.

I'm sensitive to the "garbage in, garbage out" criticism, but my estimates are that iCPS-1 would deliver 36 tonnes through TLI and iCPS-2 would deliver 41 tonnes.

In both cases I assumed the iCPS Isp was 462 s, like RL10. J2X is never going to have an Isp like that. So I used what seemed like a worst case: 438 s. But I also assumed it had enough thrust to contribute more than just a circularization burn; indeed I assumed it provided 1250 m/s of ascent delta-v. The propellant mass of the stage would grow to 89 tonnes, but even with its poor Isp, it still looks like it delivers 42 tonnes through TLI.

These results suggest that for an SLS second stage, having enough engine thrust to materially participate in ascent can overcome even a fairly severe Isp penalty.

[HappyMartian]

Well obviously Bloc IB is going to stage lower and slower than Bloc I due to having a bigger stage up top and probably a bigger payload.

"CPS Gamma" with "5x MB-60" Page 29

From: A Study of CPS Stages for Missions beyond LEO   By Mark Schaffer May 16, 2012

Schaffer's work is great, but his CPS Gamma is going to be a long way out on the timeline. As you suggest, he assumes something else put a bunch of propellant into orbit! Back as 2011 others at NASA suggested an iCPS-2. It would carry 57.8 tonnes of usable propellant, compared with iCPS-1 which is reported as carrying 27.2 tonnes.

If I understand that plan correctly, the core would deliver iCPS-1 and payload into a -87 km x 241 km target orbit. After coasting to apogee, iCPS-1 would provide the 100 m/s orbit circularization delta-v.

In contrast, I'm estimating the core could deliver iCPS-2 and payload into an orbit that requires a circularization delta-v of about 700 m/s.

I'm sensitive to the "garbage in, garbage out" criticism, but my estimates are that iCPS-1 would deliver 36 tonnes through TLI and iCPS-2 would deliver 41 tonnes.

In both cases I assumed the iCPS Isp was 462 s, like RL10. J2X is never going to have an Isp like that. So I used what seemed like a worst case: 438 s. But I also assumed it had enough thrust to contribute more than just a circularization burn; indeed I assumed it provided 1250 m/s of ascent delta-v. The propellant mass of the stage would grow to 89 tonnes, but even with its poor Isp, it still looks like it delivers 42 tonnes through TLI.

These results suggest that for an SLS second stage, having enough engine thrust to materially participate in ascent can overcome even a fairly severe Isp penalty.

A J-2X doesn't provide more than four starts. Two starts would be the norm.

A single J-2X doesn't provide engine out capability.

A J-2X isn't capable of small and precise correction burns to adjust TLI or other mission delta-v corrections. This implies the need for a smaller rocket engine.

A J-2X Isp (vac.) 448 seconds is less than the MB-60 Isp (vac.) 465, or 
RL10-A4-2 Isp (vac.) 451 sec.

The RL10-A4-2 isn't powerful enough unless we go to twenty engines and that would not fit under the CPS.

The NGE isn't powerful enough unless we go to ten engines and that would be difficult to fit under the CPS.

Our most likely choice becomes the MB-60 which is a modern, compact, and efficient rocket engine and has a useful thrust of 60,000 lbf.

Five MB-60s give 300,000 lbs of thrust which is slightly more than the 294,000 pounds of thrust of the J-2X. 

Five MB-60s, with multiple restarts, also allow for engine out capability, and the ability to make precise small delta-v corrections for TLI, TMI, or other missions.

The CPS as arrives in LEO with a payload and almost no propellant.

Hopefully all the claims about hydrolox transfers and LEO propellant depots or buddy tankers are accurate and we will be able to fill the CPS with propellant. 

Note that the largest CPS without a payload might also serve as a tanker once it is in Lunar orbit and enable multiple flights of a hydrolox reusable Lander. The largest CPS perhaps could also ferry itself back to a high Earth orbit or LEO to be resupplied with propellant for another mission. 

The main disadvantage of the five MB-60s versus the J-2X is additional engine weight.

A J-2X is about: 5,450 pounds (2,470 kg)

A MB-60 is about: 1,302 pounds (591 kg )



For various versions of the MB-60 CPS see Page 33 of:

From: A Study of CPS Stages for Missions beyond LEO   By Mark Schaffer May 16, 2012

At: http://spirit.as.utexas.edu/~fiso/telecon/Schaffer_5-16-12/Schaffer_5-16-12%20Rev%20A.pdf

Also at: http://www.sei.aero/eng/papers/uploads/archive/SpaceWorks%20CPS%20Study%20Final%20Distribution.pdf

[sdsds]

Forgive my brevity; I don't want to overload other readers' eyeballs with long responses.

A J-2X doesn't provide more than four starts. Two starts would be the norm.

To power the CxP EDS, J-2X was designed for exactly this application.

A single J-2X doesn't provide engine out capability.

True but over-rated. It's a complex trade.

A J-2X isn't capable of small and precise correction burns

Not needed.

A J-2X Isp (vac.) 448 seconds is less

Yet its net performance through TLI in this application is better.

Our most likely choice becomes the MB-60

It does not exist.

[Lobo]

MB-60 seems like a very interesting little engine.

sdsds is right, it doesn't exist, but how long would it take to exist?  Do we know how far along it got?  Or how long it would take to finish?

Although it doesn't exist, there could be a case to be made if it would be adopted accorss the EELV line too, as it was originally intended for an upgraded Delta IV upper stage.  IF ULA were to adopt it as a new common replacement for RL-10, as well as SLS, that could allow for some cost savings.  Likewise if it were used for a lunar lander crasher stage, and/or descent stage, a deep throttling version like CECE. 

But that could all be done with RL-10 too, and that engine exists and will be man-rated for commercial Atlas V and iCPS, so it might make sense to stick with that engine until there's a need accross the SLS and EELV lines to evolve or upgrade it.

[sdsds]

I'm sensitive to the "garbage in, garbage out" criticism, but my estimates are that iCPS-1 would deliver 36 tonnes through TLI and iCPS-2 would deliver 41 tonnes.

Oops! I somehow allowed "iCPS-1" to hold way too much propellant. (The 5m Delta stage holds only 27 tonnes.) But I'm pleased to see the 41 tonne through TLI estimate for "iCPS-2" is more or less validated by Jim Chilton's October 18, 2012 presentation. (Thanks to Steven Pietrobon for the link: https://info.aiaa.org/Regions/SE/HSV_AIAA/Downloadable%20Items/AIAA-Chilton_18Oct2012_Final2.pdf)

Chilton shows 43 tonnes through TLI for this configuration.

[Hyperion5]

I'm sensitive to the "garbage in, garbage out" criticism, but my estimates are that iCPS-1 would deliver 36 tonnes through TLI and iCPS-2 would deliver 41 tonnes.

Oops! I somehow allowed "iCPS-1" to hold way too much propellant. (The 5m Delta stage holds only 27 tonnes.) But I'm pleased to see the 41 tonne through TLI estimate for "iCPS-2" is more or less validated by Jim Chilton's October 18, 2012 presentation. (Thanks to Steven Pietrobon for the link: https://info.aiaa.org/Regions/SE/HSV_AIAA/Downloadable%20Items/AIAA-Chilton_18Oct2012_Final2.pdf)

Chilton shows 43 tonnes through TLI for this configuration.

Actually I think "Beyond Earth Orbit" means beyond the earth's gravity entirely.  If, as I suspect, they mean that's the maximum escape velocity payload, then the TLI number should be around 45 mt according to Lobo, a near exact match for the Saturn V.  Only the later Saturn Vs, which were optimized for greater TLI mass in order to carry the lunar rover (Apollo 15-17), surpassed 45 mt through TLI.  They pushed 47 mt through TLI.   

While it doesn't surpass the Saturn V, clongton was right that the IB could single-launch Apollo missions, or at least the early ones.  I know it's bigger than what he wants, but I say, heck, if we're going to have a big launcher, we might as well get more than just 70 mt to LEO & 19 mt BEO out of it. 

[Lobo]

Actually I think "Beyond Earth Orbit" means beyond the earth's gravity entirely.  If, as I suspect, they mean that's the maximum escape velocity payload, then the TLI number should be around 45 mt according to Lobo, a near exact match for the Saturn V.  Only the later Saturn Vs, which were optimized for greater TLI mass in order to carry the lunar rover (Apollo 15-17), surpassed 45 mt through TLI.  They pushed 47 mt through TLI.   

While it doesn't surpass the Saturn V, clongton was right that the IB could single-launch Apollo missions, or at least the early ones.  I know it's bigger than what he wants, but I say, heck, if we're going to have a big launcher, we might as well get more than just 70 mt to LEO & 19 mt BEO out of it. 

I would think even if Block 1 was only on par with early Satrun V's, the mission would probably still be even more capable than later Apollo missions, if the lander is hyrolox or methalox.  Lighter and better ISP.

I think if there's a hydrolox crasher stage, and reusable methalox lander, with even less mass actually landed, that would probably get about the best landad mass of any configuration. (although I still like landing the descender on the surface like ACES, so it can be used, and packed with cargo on Earth...but jettisoning the descender prior to terminal landing I think gets a bit more landed mass)

So even a bit less than Saturn V's best TLI capacity could yield somewhat more landed mass.

[HappyMartian]

Well obviously Bloc IB is going to stage lower and slower than Bloc I due to having a bigger stage up top and probably a bigger payload.

"CPS Gamma" with "5x MB-60" Page 29

From: A Study of CPS Stages for Missions beyond LEO   By Mark Schaffer May 16, 2012

Schaffer's work is great, but his CPS Gamma is going to be a long way out on the timeline. As you suggest, he assumes something else put a bunch of propellant into orbit! Back as 2011 others at NASA suggested an iCPS-2. It would carry 57.8 tonnes of usable propellant, compared with iCPS-1 which is reported as carrying 27.2 tonnes.

If I understand that plan correctly, the core would deliver iCPS-1 and payload into a -87 km x 241 km target orbit. After coasting to apogee, iCPS-1 would provide the 100 m/s orbit circularization delta-v.

In contrast, I'm estimating the core could deliver iCPS-2 and payload into an orbit that requires a circularization delta-v of about 700 m/s.

I'm sensitive to the "garbage in, garbage out" criticism, but my estimates are that iCPS-1 would deliver 36 tonnes through TLI and iCPS-2 would deliver 41 tonnes.

In both cases I assumed the iCPS Isp was 462 s, like RL10. J2X is never going to have an Isp like that. So I used what seemed like a worst case: 438 s. But I also assumed it had enough thrust to contribute more than just a circularization burn; indeed I assumed it provided 1250 m/s of ascent delta-v. The propellant mass of the stage would grow to 89 tonnes, but even with its poor Isp, it still looks like it delivers 42 tonnes through TLI.

These results suggest that for an SLS second stage, having enough engine thrust to materially participate in ascent can overcome even a fairly severe Isp penalty.

A J-2X doesn't provide more than four starts. Two starts would be the norm.

A single J-2X doesn't provide engine out capability.

A J-2X isn't capable of small and precise correction burns to adjust TLI or other mission delta-v corrections. This implies the need for a smaller rocket engine.

A J-2X Isp (vac.) 448 seconds is less than the MB-60 Isp (vac.) 465, or 
RL10-A4-2 Isp (vac.) 451 sec.

The RL10-A4-2 isn't powerful enough unless we go to twenty engines and that would not fit under the CPS.

The NGE isn't powerful enough unless we go to ten engines and that would be difficult to fit under the CPS.

Our most likely choice becomes the MB-60 which is a modern, compact, and efficient rocket engine and has a useful thrust of 60,000 lbf.

Five MB-60s give 300,000 lbs of thrust which is slightly more than the 294,000 pounds of thrust of the J-2X. 

Five MB-60s, with multiple restarts, also allow for engine out capability, and the ability to make precise small delta-v corrections for TLI, TMI, or other missions.

The CPS as arrives in LEO with a payload and almost no propellant.

Hopefully all the claims about hydrolox transfers and LEO propellant depots or buddy tankers are accurate and we will be able to fill the CPS with propellant. 

Note that the largest CPS without a payload might also serve as a tanker once it is in Lunar orbit and enable multiple flights of a hydrolox reusable Lander. The largest CPS perhaps could also ferry itself back to a high Earth orbit or LEO to be resupplied with propellant for another mission. 

The main disadvantage of the five MB-60s versus the J-2X is additional engine weight.

A J-2X is about: 5,450 pounds (2,470 kg)

A MB-60 is about: 1,302 pounds (591 kg )



For various versions of the MB-60 CPS see Page 33 of:

From: A Study of CPS Stages for Missions beyond LEO   By Mark Schaffer May 16, 2012

At: http://spirit.as.utexas.edu/~fiso/telecon/Schaffer_5-16-12/Schaffer_5-16-12%20Rev%20A.pdf

Also at: http://www.sei.aero/eng/papers/uploads/archive/SpaceWorks%20CPS%20Study%20Final%20Distribution.pdf

Perhaps what is really needed is:

Ogawara et al, “Dual Hydrogen/Methane Propellant Capability of the
Expander-Bleed Cycle Engine” presented at 26th ISTS Conference,
Hamamatsu, Japan, June 2008, ISTS 2008-a-07.


Note also:

Excellence of the Japanese Expander-Bleed Cycle Rocket Engine and Enhancements for Future Engine Applications    By William Sack, Koichi Okita, Akihide Kurosu, Akira Ogawara, Kimito Yoshikawa, Masahiro Atsumi, Kenji Kishimoto, Kevin Lunde    2008
At: http://archive.ists.or.jp/upload_pdf/2008-a-03.pdf

"JAXA and MHI have successfully developed an excellent expander bleed power cycle rocket engine designated the LE-5B. This LOX-hydrogen engine currently provides upper stage propulsion for the H-2A launch vehicle and has proven to be reliable and robust in 15 flights. The engine was developed as an upgrade to the original gas generator powered LE-5 engine. Elimination of the gas generator simplified the design, required ignition of only a single combustion device, precluded a potential chug instability at throttled conditions, and eliminated moisture from the turbine drive gases preventing “freeze locking” of the turbine on restart. Future Japanese engine designs can benefit and build on this valuable experience. To further enhance this proven cycle, higher combustion pressure is desirable to optimize performance for potential booster engine application. In addition, higher injector and turbopump performance is required to further maximize the cycle’s specific impulse. These improvements have the potential to greatly benefit Japan’s new LE-X booster engine design. Another interesting feature of a LOX/H2 expander-bleed cycle engine is that, based on a recent study, the cycle can potentially be designed to function as a dual fuel engine utilizing LCH4 or LH2. This capability can facilitate the use of in-situ propellant systems for space exploration applications. The purpose of this paper is to discuss progress and background towards the development of these expander-bleed cycle engine improvements and capabilities."


I added the bold to the article's abstract.

Food for thought about what options the Dual Hydrogen/Methane Propellant Capability of the Expander-Bleed Cycle Engine might offer for the upper stage and Lunar Lander.

[newpylong]

Do we think they are even going to go with the 105 ton/1B configuration and not skip right to 2?

[PahTo]

I think the availability of SRB motor casings and RS-25 engines will determine the fate of 1B, and likely the fate of the US HLLV program for many years to come.
I hope the powers that be are really looking to Block 1 as a bridge to the preferred LRB + RS-68 powered 8.4 meter core, but in doing this, we run the risk of scuttling the whole deal...
That is, would it have been better to take longer to do it right (and risk the prolonged downtime killing the whole effort), or to keep elements flying, however infrequently, to keep the HLLV dream alive (and risk the costs of the interim program killing the whole effort).  ATK lobbying clearly has played a big role, and will continue to.

[RotoSequence]

I think the availability of SRB motor casings and RS-25 engines will determine the fate of 1B, and likely the fate of the US HLLV program for many years to come.
I hope the powers that be are really looking to Block 1 as a bridge to the preferred LRB + RS-68 powered 8.4 meter core, but in doing this, we run the risk of scuttling the whole deal...
That is, would it have been better to take longer to do it right (and risk the prolonged downtime killing the whole effort), or to keep elements flying, however infrequently, to keep the HLLV dream alive (and risk the costs of the interim program killing the whole effort).  ATK lobbying clearly has played a big role, and will continue to.

In detail, how does the 4x RS-68 8.4m stage compare to the 3x J2-X 8.4m stage?

[newpylong]

Well, they wouldn't use the J2-X's for the core. I believe he means whether they keep using the RS-25 or move to 68's when the SSME's run out instead of build RS-25E's.

The big problem with moving to the 68's is Man Rating them (the amount of modifications required for Ares V was going to be astronomical) and lower thrust. They were quite a bit cheaper to produce though.

[sdsds]

It seems that the "ICPS-2" moniker came from somewhere outside this thread, but it does not correctly describe the proposed 4xRL10 stage, which would not use ICPS tank tooling (it would have 8.4 m ET diameter and 5.5 m Ares I diameter tanks) or engines (or avionics).

I agree with your description, but think it would nonetheless still be an ICPS unless it had the loiter capability that is an essential part of the CPS {description, definition, requirements}.