RS-68

[Aeneas]

Is someone out there actually ashamed for the RS-68?

TWR of ~ 50
vacuum Isp of ~412 s
and still ~ 14 m USD?

[MATTBLAK]

The RS-68 is a 'dead engine walking'. When Delta IV-Heavy goes, so will the engine. It's unlikely to be used on another booster ever again. Yes, it's relatively cheap, yes it's relatively powerful for a LH2 fueled engine. But it can't be air-started and it weighs more than 14,000 pounds (6.3 metric tons) so it would make a lousy upper stage engine. It weighs so much because of it's huge ablative nozzle.

The only real use I could see for it again would be if funding were allocated to turn that nozzle into a regeneratively cooled one like the Shuttle/SLS's RS-25 engines. The specific impulse would increase and then there would be no problem clustering them closely together as was first proposed in Constellation. The Isp increase on a big thrusting engine like RS-68 would give the SLS Corestage a modest payload increase to orbit. Five, maybe 7 metric tons.

But such an upgrade to RS-68 would not come cheap and the SLS would not have a high enough flight rate to justify the multi-billion dollar upgrade so...

[ugordan]

Ashamed? Not really. It was good enough for what it was meant to be at the time: a low(er) cost booster engine alternative to the SSME. As an engine, I think it's an indicator on where the US went wrong - focusing pretty much all of the development on either LH2 engines or solid rockets after Apollo, neglecting hydrocarbon engines.

Unfortunately, the choice of a LH2 booster engine and the low impulse density of that propellant made it uncompetitive with the more traditional approach of a hydrocarbon engine for boost stage, LH2 for upper stage. The large tankage and associated difficulties with cryogenic hydrogen were apparently enough to drown out any Isp advantage over hydrocarbon booster engines.* That's not to say a U.S. made staged combustion RP-1 engine would have been any cheaper, but the overall system cost probably would have been lower.

* Atlas V has entered the chat

[Aeneas]

The RS-58 is a 'dead engine walking'. When Delta IV-Heavy goes, so will the engine. It's unlikely to be used on another booster ever again. Yes, it's relatively cheap, yes it's relatively powerful for a LH2 fueled engine. But it can't be air-started and it weighs more than 14,000 pounds (6.3 metric tons) so it would make a lousy upper stage engine. It weighs so much because of it's huge ablative nozzle.

The only real use I could see for it again would be if funding were allocated to turn that nozzle into a regeneratively cooled one like the Shuttle/SLS's RS-25 engines. The specific impulse would increase and there would be no problem clustering them closely together as was first proposed in Constellation. The Isp increase on a big thrusting engine like RS-68 would give the SLS Corestage a modest payload increase to orbit. Five, maybe 7 metric tons.

But such an upgrade to RS-68 would not come cheap and the SLS would not have a high enough flight rate to justify the multi-billion dollar upgrade so...

True, true... But maybe it should not go through a re-design but new approach using Hydrolox for the 1. stage. RS-68 has about 120ish bar. Apparently there are alloys allowing for more than doubling this value. Going beyond 250 bar would drastically reduce size, probably increasing the TWR and adding FFSC probably would add some amazing SL Isp. I played around with RPA and adjusting SSME to 262 bar chamber pressure and an expansion ratio of ~35, it'll generate >404 s of SL Isp, going >430 s in vacuum. This would actually make the first stage much lighter than hydrocarbon riding rockets.

[baldusi]

So you are saying that if you use a different cycle, with 80s technology, for a different use, for a not needed requirement, hitting never achieved specs on an hydrolox engine, and without doing the numbers, you wish it to be amazing?

[dglow]

Can anyone recommend good articles or books examining the development of Delta IV – what led McDonnell to choose hydrolox for their boosters?

[GreenShrike]

This would actually make the first stage much lighter than hydrocarbon riding rockets.

...and why does this matter? Kerosene isn't exactly expensive, methane is even cheaper and, combined with lox (also very low cost), either would comprise the vast majority of the stage's mass.

You're effectively suggesting a very expensive redesign, to end up producing expensive* engines, just so you could save the cost of a bunch of cheap fuel and maybe a barrel section or two of tankage?

*When your competition is mass producing hydrocarbon engines for well under $1M (and has a target under $250K for its twice-as-powerful replacement), any hydrolox engine is more than likely going to be expensive.**

$14M  for an RS-68, you say? Consider that if Raptor only gets down to $500K, you could outfit a Superheavy with engines for about the same cost as powering a Delta IV Medium. If Raptors were $1M each, then a Superheavy would still cost less to power than a Delta IV Heavy.

While the disadvantage of an expensive first stage might be mitigated via reusability, that just ends up driving the cost per flight down to the cost of your propellant, and LH2 would remain expensive (and a pain to work with). And you'll less easily be able to expend an expensive booster stage if you happen to need more performance on a given mission.


Overall, I'd say admiring the RS-68 (and, even more so, the RS-25) is like admiring a fine Maserati, with its wonderful craftsmanship and marvelous engineering. Except if you just care about getting the job done, then a basic Ford F-150 pickup truck will do just as well. And if we want to truly expand into space, a hundred F-150 equivalents will serve us a hell of a lot better than a few showpieces, no matter how pretty and fast the showpieces might be.

So, no, there's nothing shameful about the RS-68 -- it just belongs in the past. The world has moved on, and before any redesign of old technology is warranted, there needs to be sufficient justification of the result against the current state-of-the-art -- and not in terms of ISP, but rather in terms of cost effectiveness.


** And if you really want to stick with hydrolox, I'd ignore the RS-68 and take a closer look at the LE-9. If JAXA and Mitsubishi can hit the H3's price targets, then its a hydrolox booster engine that might actually be cost effective, at least vis-a-vis the Merlin (absent F9 reusability, at any rate).

[Aeneas]

So you are saying that if you use a different cycle, with 80s technology, for a different use, for a not needed requirement, hitting never achieved specs on an hydrolox engine, and without doing the numbers, you wish it to be amazing?

Yes. Isn't it obvious? ^^

I'm well aware, that this is highly speculative, I just tend to dream a bit about possible rocket futures.

But regarding requirements, I wouldn't go that way. It's risky but having a technology in place offering cheap and large unit access to space will probably provoke the requirement once it's there.

[Aeneas]

This would actually make the first stage much lighter than hydrocarbon riding rockets.

...and why does this matter? Kerosene isn't exactly expensive, methane is even cheaper and, combined with lox (also very low cost), either would comprise the vast majority of the stage's mass.

You're effectively suggesting a very expensive redesign, to end up producing expensive* engines, just so you could save the cost of a bunch of cheap fuel and maybe a barrel section or two of tankage?

*When your competition is mass producing hydrocarbon engines for well under $1M (and has a target under $250K for its twice-as-powerful replacement), any hydrolox engine is more than likely going to be expensive.**

$14M  for an RS-68, you say? Consider that if Raptor only gets down to $500K, you could outfit a Superheavy with engines for about the same cost as powering a Delta IV Medium. If Raptors were $1M each, then a Superheavy would still cost less to power than a Delta IV Heavy.

While the disadvantage of an expensive first stage might be mitigated via reusability, that just ends up driving the cost per flight down to the cost of your propellant, and LH2 would remain expensive (and a pain to work with). And you'll less easily be able to expend an expensive booster stage if you happen to need more performance on a given mission.


Overall, I'd say admiring the RS-68 (and, even more so, the RS-25) is like admiring a fine Maserati, with its wonderful craftsmanship and marvelous engineering. Except if you just care about getting the job done, then a basic Ford F-150 pickup truck will do just as well. And if we want to truly expand into space, a hundred F-150 equivalents will serve us a hell of a lot better than a few showpieces, no matter how pretty and fast the showpieces might be.

So, no, there's nothing shameful about the RS-68 -- it just belongs in the past. The world has moved on, and before any redesign of old technology is warranted, there needs to be sufficient justification of the result against the current state-of-the-art -- and not in terms of ISP, but rather in terms of cost effectiveness.


** And if you really want to stick with hydrolox, I'd ignore the RS-68 and take a closer look at the LE-9. If JAXA and Mitsubishi can hit the H3's price targets, then its a hydrolox booster engine that might actually be cost effective, at least vis-a-vis the Merlin (absent F9 reusability, at any rate).

Fuel cost will likely not play a role for at least the next decade - though I'd be happy to be wrong (looking at you Starship!). And maybe, if there's something like a space trucking established, H2 production on site might reduce costs considerably. The one driver I have in mind, is the thrust needed at launch. If I want to go big, this is actually something that needs considering. Otherwise, it's a path into the Sea Dragon ocean which might be something yet it's a whole bunch of new considerations that would be needed to be taken into account.

If one can drive down costs with Raptor, one could do the same with a hydrolox equivalent. And yes, Delta IV is something of the past, which is said by Tory Bruno, too. I just open this thread to rant about the RS-68.

What's so good about LE-9 accept the possible price tag?

[envy887]

The RS-58 is a 'dead engine walking'. When Delta IV-Heavy goes, so will the engine. It's unlikely to be used on another booster ever again. Yes, it's relatively cheap, yes it's relatively powerful for a LH2 fueled engine. But it can't be air-started and it weighs more than 14,000 pounds (6.3 metric tons) so it would make a lousy upper stage engine. It weighs so much because of it's huge ablative nozzle.

The only real use I could see for it again would be if funding were allocated to turn that nozzle into a regeneratively cooled one like the Shuttle/SLS's RS-25 engines. The specific impulse would increase and there would be no problem clustering them closely together as was first proposed in Constellation. The Isp increase on a big thrusting engine like RS-68 would give the SLS Corestage a modest payload increase to orbit. Five, maybe 7 metric tons.

But such an upgrade to RS-68 would not come cheap and the SLS would not have a high enough flight rate to justify the multi-billion dollar upgrade so...

True, true... But maybe it should not go through a re-design but new approach using Hydrolox for the 1. stage. RS-68 has about 120ish bar. Apparently there are alloys allowing for more than doubling this value. Going beyond 250 bar would drastically reduce size, probably increasing the TWR and adding FFSC probably would add some amazing SL Isp. I played around with RPA and adjusting SSME to 262 bar chamber pressure and an expansion ratio of ~35, it'll generate >404 s of SL Isp, going >430 s in vacuum. This would actually make the first stage much lighter than hydrocarbon riding rockets.

SSME is 4 or 5 times as expensive, mostly because of the complexity required to get that high pressure. For a vertically launched main stage, it's more cost-effective to use hydrocarbons and pursue better mass fractions than to use hydrolox and chase ISP, because there is effectively no limit on liftoff mass.

The only place I can see LH2 main stages being optimal is for a horizontally launched or air-dropped stage, where gross liftoff mass is a hard constraint.

[Hog]

The RS-58 is a 'dead engine walking'. When Delta IV-Heavy goes, so will the engine. It's unlikely to be used on another booster ever again. Yes, it's relatively cheap, yes it's relatively powerful for a LH2 fueled engine. But it can't be air-started and it weighs more than 14,000 pounds (6.3 metric tons) so it would make a lousy upper stage engine. It weighs so much because of it's huge ablative nozzle.

The only real use I could see for it again would be if funding were allocated to turn that nozzle into a regeneratively cooled one like the Shuttle/SLS's RS-25 engines. The specific impulse would increase and there would be no problem clustering them closely together as was first proposed in Constellation. The Isp increase on a big thrusting engine like RS-68 would give the SLS Corestage a modest payload increase to orbit. Five, maybe 7 metric tons.

But such an upgrade to RS-68 would not come cheap and the SLS would not have a high enough flight rate to justify the multi-billion dollar upgrade so...

True, true... But maybe it should not go through a re-design but new approach using Hydrolox for the 1. stage. RS-68 has about 120ish bar. Apparently there are alloys allowing for more than doubling this value. Going beyond 250 bar would drastically reduce size, probably increasing the TWR and adding FFSC probably would add some amazing SL Isp. I played around with RPA and adjusting SSME to 262 bar chamber pressure and an expansion ratio of ~35, it'll generate >404 s of SL Isp, going >430 s in vacuum. This would actually make the first stage much lighter than hydrocarbon riding rockets.

SSME is 4 or 5 times as expensive, mostly because of the complexity required to get that high pressure. For a vertically launched main stage, it's more cost-effective to use hydrocarbons and pursue better mass fractions than to use hydrolox and chase ISP, because there is effectively no limit on liftoff mass.

The only place I can see LH2 main stages being optimal is for a horizontally launched or air-dropped stage, where gross liftoff mass is a hard constraint.

Is Delta-IV H the only launcher with a pure HydroLOx booster/first stage?

[baldusi]

[...]
Is Delta-IV H the only launcher with a pure HydroLOx booster/first stage?

Only one I can think of. And personally think that's a good thing.

[envy887]

[...]
Is Delta-IV H the only launcher with a pure HydroLOx booster/first stage?

Only one I can think of. And personally think that's a good thing.

In addition to DIVH, there was also an all-hydrolox DIVM variant, with no SRBs. And there is a all-hydrolox variant of the H3, but it hasn't flown yet.

[edkyle99]

Is someone out there actually ashamed for the RS-68?

TWR of ~ 50
vacuum Isp of ~412 s
and still ~ 14 m USD?

Ashamed?  Not in the least.  Just three of these are needed to lift what is still the heaviest-hauling operational launch vehicle, carrying important DoD payloads.  The world's most powerful LH2/LOX engine.  Lots of thrust and decent ISP given the simple, reliable gas generator cycle.  No solids or kerosene to stain the atmosphere on ascent.  No failures in flight.  Good luck to those planning on replacing Delta 4 Heavy during NSSL.  It won't be trivial.

 - Ed Kyle

[Aeneas]

Is someone out there actually ashamed for the RS-68?

TWR of ~ 50
vacuum Isp of ~412 s
and still ~ 14 m USD?

Ashamed?  Not in the least.  Just three of these are needed to lift what is still the heaviest-hauling operational launch vehicle, carrying important DoD payloads.  The world's most powerful LH2/LOX engine.  Lots of thrust and decent ISP given the simple, reliable gas generator cycle.  No solids or kerosene to stain the atmosphere on ascent.  No failures in flight.  Good luck to those planning on replacing Delta 4 Heavy during NSSL.  It won't be trivial.

 - Ed Kyle

The Isp isn't decent, it's rather poor compared to RS-25. But thanks for the picture, it shows the thick ablative nozzle really well!
Delta IV is already replaceable by Falcon Heavy, isn't it?

[ugordan]

The Isp isn't decent, it's rather poor compared to RS-25.

The Isp by all means is very decent for a booster engine. You are setting the bar too high when you are comparing it to the RS-25, arguably one of if not the most complex and highest-performing engine, with a cost to accompany all that.

[Kryten]

Is someone out there actually ashamed for the RS-68?

TWR of ~ 50
vacuum Isp of ~412 s
and still ~ 14 m USD?

$14 million? Try $60 million. You're repeating stuff they were saying about the engine before it got into operation which they quickly found out they could not do; it's like asking why the space shuttle got retired when it could fly weekly.

[brickmack]

$14 million? Try $60 million. You're repeating stuff they were saying about the engine before it got into operation which they quickly found out they could not do; it's like asking why the space shuttle got retired when it could fly weekly.

Those projections were the worst case in an uncertain market. RS-68 settled around 28 million, RD-180 at around 23 million, RL10A-4-2 at 11-12 million, RL10B-2 at 6-7 million.

Anyway though, I think engine choice was the biggest problem for Delta. They spent half a billion dollars developing RS-68? At DIVs demonstrated flightrate, thats almost 10 million an engine just in amortized dev cost.

Two RS-25s in a reusable engine pod (ie, the original Boeing EELV bid) would have had a much higher ISP, more total thrust, and a slightly better TWR. And even at the most pessimistic estimates I've seen for RS-25s refurb cost, that'd still be cheaper than a single RS-68. Recent work on AR-22 (10 firings in 10 days with zero refurb needed, on an engine built from obsolete scraps) implies a much better cost. This added performance would've also made it a lot easier to achieve single-core heavy performance like Vulcan, which should drastically reduce costs (namely from fewer unique hardware designs). Plus the demonstrated reliability, plus the synergies with the still-flying Shuttle.

I'd really love to know what justified Boeing losing EELV, given from what information is available it sure looks like they had the cheapest to develop, cheapest to fly, most capable, and highest-heritage vehicle. Engine reuse could have been seen as a risk, but RS-25 reuse was already well-proven, their recovery system was pretty simple, and they did a demonstration drop test that went well

[edkyle99]

Delta IV is already replaceable by Falcon Heavy, isn't it?

It could be in theory, but not the way that SpaceX has flown it so far.  Delta 4 Heavy can lift almost 14 tonnes to GTO.  Falcon Heavy with recovery of all three boosters as flown to date can lift 8 tonnes.  Falcon Heavy could beat Delta 4 Heavy payload to GTO only by expending the center core, I think.  Falcon Heavy will also need a bigger fairing and SpaceX will need to add vertical payload integration to win NSSL, both steps the company has apparently proposed.

 - Ed Kyle

[Coastal Ron]

Delta IV is already replaceable by Falcon Heavy, isn't it?

It could be in theory, but not the way that SpaceX has flown it so far.  Delta 4 Heavy can lift almost 14 tonnes to GTO.  Falcon Heavy with recovery of all three boosters as flown to date can lift 8 tonnes.

You know quite well that it is a customer decision as to what version of Falcon Heavy they want to use, since the price is determined by the capability.

Falcon Heavy could beat Delta 4 Heavy payload to GTO only by expending the center core, I think.

SpaceX says that the Falcon Heavy has the capability to put 26.7mT to GTO, so it can certainly exceed the mass capabilities of Delta IV Heavy. It is only a matter of what price the customer is willing to pay, and Elon Musk has said a fully expendable Falcon Heavy would be about $150M.

As to the other requirements to take away payloads from Delta IV Heavy, I don't think SpaceX will want to modify Falcon Heavy very much when they plan to have Starship operational in a few years - and let's not debate whether that will happen...

[Aeneas]

Delta IV is already replaceable by Falcon Heavy, isn't it?

It could be in theory, but not the way that SpaceX has flown it so far.  Delta 4 Heavy can lift almost 14 tonnes to GTO.  Falcon Heavy with recovery of all three boosters as flown to date can lift 8 tonnes.  Falcon Heavy could beat Delta 4 Heavy payload to GTO only by expending the center core, I think.  Falcon Heavy will also need a bigger fairing and SpaceX will need to add vertical payload integration to win NSSL, both steps the company has apparently proposed.

 - Ed Kyle

Next step would be landing the side boosters on a drone ship, too. Would it be realistic to put and RL10-stage above 2. stage of the Falcon Heavy configuration?

[Welsh Dragon]

Delta IV is already replaceable by Falcon Heavy, isn't it?

It could be in theory, but not the way that SpaceX has flown it so far.  Delta 4 Heavy can lift almost 14 tonnes to GTO.  Falcon Heavy with recovery of all three boosters as flown to date can lift 8 tonnes.  Falcon Heavy could beat Delta 4 Heavy payload to GTO only by expending the center core, I think.  Falcon Heavy will also need a bigger fairing and SpaceX will need to add vertical payload integration to win NSSL, both steps the company has apparently proposed.

 - Ed Kyle

Next step would be landing the side boosters on a drone ship, too. Would it be realistic to put and RL10-stage above 2. stage of the Falcon Heavy configuration?

Permit me to do a Jim. No.*

*(SpaceX doesn't make RL10s, so would have to buy them. This goes against their philosophy and Aerojet wouldn't need to sell it to them. Also, they're pretty damn expensive. The real kicker is the lack of hydrogen infrastructure on the pad though. Lastly, if you're working on the assumption that Starship will be a success (as SpaceX will be), it would make no sense to upgrade Heavy.)

[baldusi]

I'd really love to know what justified Boeing losing EELV, given from what information is available it sure looks like they had the cheapest to develop, cheapest to fly, most capable, and highest-heritage vehicle.

They cheated... criminally so. They were left with some of the EELV because it was too late to swap for another bidder.

[brickmack]

I'd really love to know what justified Boeing losing EELV, given from what information is available it sure looks like they had the cheapest to develop, cheapest to fly, most capable, and highest-heritage vehicle.

They cheated... criminally so. They were left with some of the EELV because it was too late to swap for another bidder.

No, Boeing lost EELV before that. Then they bought McDonnell Douglas, who had won with their Delta IV bid. Then Boeing broke the law.

[edkyle99]

SpaceX says that the Falcon Heavy has the capability to put 26.7mT to GTO, so it can certainly exceed the mass capabilities of Delta IV Heavy. It is only a matter of what price the customer is willing to pay, and Elon Musk has said a fully expendable Falcon Heavy would be about $150M.

I have in my hands right now a copy of Launchspace magazine, September 2000 issue.  On page 41 it lists Delta IV Heavy prices in the $130-150 million per launch range, a fantastic reduction from the $365-435 million per launch for Titan IV listed on page 43.  I recall that when Titan IV development began, similar low-ball estimates were given.  Hopeful, but ultimately ridiculous price projections seem to always precede the actual implementation of these things, especially for big DoD missions.  But good luck to SpaceX, maybe they will break the trend.  It won't be easy on missions where they have to expend stages to meet customer performance requirements.

 - Ed Kyle

[dglow]

...
Hopeful, but ultimately ridiculous price projections seem to always precede the actual implementation of these things, especially for big DoD missions.  But good luck to SpaceX, maybe they will break the trend.  It won't be easy on missions where they have to expend stages to meet customer performance requirements.

IIRC SpaceX was already pricing their flights lower before they began recovering boosters.

[envy887]

SpaceX says that the Falcon Heavy has the capability to put 26.7mT to GTO, so it can certainly exceed the mass capabilities of Delta IV Heavy. It is only a matter of what price the customer is willing to pay, and Elon Musk has said a fully expendable Falcon Heavy would be about $150M.

I have in my hands right now a copy of Launchspace magazine, September 2000 issue.  On page 41 it lists Delta IV Heavy prices in the $130-150 million per launch range, a fantastic reduction from the $365-435 million per launch for Titan IV listed on page 43.  I recall that when Titan IV development began, similar low-ball estimates were given.  Hopeful, but ultimately ridiculous price projections seem to always precede the actual implementation of these things, especially for big DoD missions.  But good luck to SpaceX, maybe they will break the trend.  It won't be easy on missions where they have to expend stages to meet customer performance requirements.

 - Ed Kyle

SpaceX expended a stage on STP-2. How much was that contract?

[butters]

With all the cautionary tales available to guide design engineers, there's no excuse for developing a new hydrolox booster engine unless you've gone mad enough to attempt an SSTO. Maybe in the coming decades, additive manufacturing technology will reach a point where many of the cooling challenges and plumbing complexities associated with linear aerospike nozzles will become practical to solve, which could justify another stab at that white whale.

If the objective is a high-thrust upper stage engine for a heavy earth departure stage, then the most reasonable approach is to ditch the gas generator and go with some variant of the expander cycle or a tap-off cycle depending on how much thrust is required. This makes flight-start much easier and more reliable. So one could argue that BE-3U is the spiritual successor to RS-68.

[edkyle99]

So one could argue that BE-3U is the spiritual successor to RS-68.

Except for thrust.  BE-4 (and Raptor too) is closer to RS-68 in that regard than BE-3U.   I still don't get it myself.  BE-4 still uses cryogenic fuel like RS-68, but has more pump-cycle complexity for less thrust and ISP than RS-68.  Less cost maybe, but we'll have to see about that.

 - Ed Kyle

[edkyle99]

SpaceX expended a stage on STP-2. How much was that contract?

Not on purpose.  The core stage intended to land OCISLY, but missed.

 - Ed Kyle

[dglow]

What in your opinion might constitute a fair LV comparison such that we could compare prices?

Assume ULA and SpaceX win NSSL awards. Vulcan is ULA's future so let's set aside DIVH. A FH, expending its center core and recovering its side boosters, would best be compared to Vulcan flying in what configuration?

[joek]

Except for thrust. ...  Less cost maybe, but we'll have to see about that.

Re-usability and commonality--which relates to cost?  Obviously not applicable to BE-3U for second-stage use, but as a derivative of BE-3 and interchangeability-commonality with BE-3, a consideration.  For BE-4, expect re-usability is definitely a factor.  Likely changing to an ablative nozzle as RS-68 would require significant changes and reduce interchangeability-commonality--not to mention re usability, so costs would go up.  Operative question for BE-3U is whether BE-3 has a future, and thus whether maintaining commonality is worth it.  If Blue intends to continue to produce New Shepard, likely yes.

In short, too many differences in goals-history between RS-68, BE-3, BE-3U and B-4 to make comparisons--and IMO off-topic for this thread

[joek]

...
Assume ULA and SpaceX win NSSL awards. Vulcan is ULA's future so let's set aside DIVH. A FH, expending its center core and recovering its side boosters, would best be compared to Vulcan flying in what configuration?

Not comparable.  ULA-Vulcan is not competing on LV pricing alone.

[su27k]

SpaceX expended a stage on STP-2. How much was that contract?

https://spaceflightnow.com/2019/06/25/falcon-heavy-launches-on-military-led-rideshare-mission-boat-catches-fairing/

The Air Force’s launch contract with SpaceX for the STP-2 mission was previously valued at $185 million, according to Lt. Col. Ryan Rose, chief of the small launch and targets division at Kirtland Air Force Base.

The launch is now costing the Air Force around $160 million, and a “big factor” in the cost reduction was the military’s agreement to fly the STP-2 mission with reused rocket boosters, Bongiovi said.

But I expect NRO launches will be more expensive since it requires vertical integration, someone will have to pay for the new MST.

[dglow]

...
Assume ULA and SpaceX win NSSL awards. Vulcan is ULA's future so let's set aside DIVH. A FH, expending its center core and recovering its side boosters, would best be compared to Vulcan flying in what configuration?

Not comparable.  ULA-Vulcan is not competing on LV pricing alone.

That's true. And it's also off-topic. My apologies.

By way of explanation I was really asking that to Ed following the back & forth upthread. However, upon rereading his post, I think Ed's point wasn't about comparing prices but rather the general sentiment that historically launch providers find it difficult to keep initial costs estimates in check for DoD missions.

Still I do not see how recovering 'only' 58% of a FH will substantially impact SpaceX's position, especially given the other likely winners of Phase 2. But we'll see. And of course, the price of a good or service is not necessarily related to its cost.

[Coastal Ron]

SpaceX says that the Falcon Heavy has the capability to put 26.7mT to GTO, so it can certainly exceed the mass capabilities of Delta IV Heavy. It is only a matter of what price the customer is willing to pay, and Elon Musk has said a fully expendable Falcon Heavy would be about $150M.

I have in my hands right now a copy of Launchspace magazine, September 2000 issue.  On page 41 it lists Delta IV Heavy prices in the $130-150 million per launch range, a fantastic reduction from the $365-435 million per launch for Titan IV listed on page 43.  I recall that when Titan IV development began, similar low-ball estimates were given.  Hopeful, but ultimately ridiculous price projections seem to always precede the actual implementation of these things, especially for big DoD missions.  But good luck to SpaceX, maybe they will break the trend.  It won't be easy on missions where they have to expend stages to meet customer performance requirements.

Boeing and Lockheed Martin were bleeding cash and lobbying the U.S. Government to allow them to create a launch monopoly, so of course they were using every tool in their political belt to justify the merger. Once they merged it was easy to justify price increases and get away from it because the budgets were opaque.

SpaceX has not increased Falcon 9 prices for previously unblown 1st stages in 4 years, so they have very good price stability. That means Falcon Heavy is not likely to follow in the footsteps of Delta IV Heavy price-wise.

The challenge the RS-68 had, now that we speak of it in the past tense since we know the only launcher that uses it is End Of Life, is that it was created for a business environment where optimization was more important than price. As the commercial launch sector becomes more diverse, and as reusability increases, that business model is no longer supportable.

[Aeneas]

Except for thrust. ...  Less cost maybe, but we'll have to see about that.

In short, too many differences in goals-history between RS-68, BE-3, BE-3U and B-4 to make comparisons--and IMO off-topic for this thread

Don't worry: I opened the thread to rant about RS-68, so if it actually develops into a productive thread, keep the factual spirit going.

[Aeneas]

If the objective is a high-thrust upper stage engine for a heavy earth departure stage, then the most reasonable approach is to ditch the gas generator and go with some variant of the expander cycle or a tap-off cycle depending on how much thrust is required. This makes flight-start much easier and more reliable. So one could argue that BE-3U is the spiritual successor to RS-68.

Why not go crazy and use FFSC in a Raptor fashion of mass production and extreme reliability through heavy testing?

[Aeneas]

So one could argue that BE-3U is the spiritual successor to RS-68.

Except for thrust.  BE-4 (and Raptor too) is closer to RS-68 in that regard than BE-3U.   I still don't get it myself.  BE-4 still uses cryogenic fuel like RS-68, but has more pump-cycle complexity for less thrust and ISP than RS-68.  Less cost maybe, but we'll have to see about that.

 - Ed Kyle

That's true but BE-4 uses methane instead of hydrogen. So much less tank needed, smaller piping at the engine, and "only" ~90 instead of crazy low ~20 Kelvin to manage. Nonetheless, I love hydrogen because it burns so beautiful and has such amazing Isp-potential...

[butters]

If the objective is a high-thrust upper stage engine for a heavy earth departure stage, then the most reasonable approach is to ditch the gas generator and go with some variant of the expander cycle or a tap-off cycle depending on how much thrust is required. This makes flight-start much easier and more reliable. So one could argue that BE-3U is the spiritual successor to RS-68.

Why not go crazy and use FFSC in a Raptor fashion of mass production and extreme reliability through heavy testing?

That's the origin story of Raptor. It was a hydrolox FFSC for three years before they pivoted to methalox. Mars ISRU was one reason for that, but also, pumping LH2 to high pressures is very difficult. SpaceX knew they were gonna be up against it to develop the ox-rich LOX pump. They probably wanted to focus their attention on overcoming that challenge and avoid biting off more than they could chew on the fuel side on the engine. If it was too big a leap for SpaceX, it's too big a leap for most engine suppliers.

[baldusi]

If the objective is a high-thrust upper stage engine for a heavy earth departure stage, then the most reasonable approach is to ditch the gas generator and go with some variant of the expander cycle or a tap-off cycle depending on how much thrust is required. This makes flight-start much easier and more reliable. So one could argue that BE-3U is the spiritual successor to RS-68.

Why not go crazy and use FFSC in a Raptor fashion of mass production and extreme reliability through heavy testing?

That's the origin story of Raptor. It was a hydrolox FFSC for three years before they pivoted to methalox. Mars ISRU was one reason for that, but also, pumping LH2 to high pressures is very difficult. SpaceX knew they were gonna be up against it to develop the ox-rich LOX pump. They probably wanted to focus their attention on overcoming that challenge and avoid biting off more than they could chew on the fuel side on the engine. If it was too big a leap for SpaceX, it's too big a leap for most engine suppliers.

It was hydrolox FRSC, nos FF. The problem to do FFSC on hydrolox is that you get something like 80% of the power on the H2 turbine, where you need 30% of your power, and 20% of your power on the O2 turbine where you need 70% of your power. Methalox has an almost 30%/70% ration on both turbines and pumps. That's why going FF means using methalox.

[envy887]

So one could argue that BE-3U is the spiritual successor to RS-68.

Except for thrust.  BE-4 (and Raptor too) is closer to RS-68 in that regard than BE-3U.   I still don't get it myself.  BE-4 still uses cryogenic fuel like RS-68, but has more pump-cycle complexity for less thrust and ISP than RS-68.  Less cost maybe, but we'll have to see about that.

 - Ed Kyle

But Vulcan is both higher-performing and cheaper than any single-stick Delta IV. This suggest that the higher engine complexity and count actually makes for a better overall system, by simplifying other components and systems. RS-68's fuel adds a lot of tankage and dry mass that adds expense and decreases performance compared to BE-4.

[envy887]

SpaceX expended a stage on STP-2. How much was that contract?

Not on purpose.  The core stage intended to land OCISLY, but missed.

 - Ed Kyle

They said before the launch that it was going to be an extremely difficult landing with a low probability of success. I highly doubt they were pricing the mission below normal profit margin, with the intent to make it up on later flights of that core. Doubly so because FH has a very low flight rate so such a reflight would not happen for a long time, and no FH customers have accepted used cores so far as we know, and SpaceX has never recovered and reused a FH core, and the USAF isn't very sensitive to price, and I could go on.

[Aeneas]

It was hydrolox FRSC, nos FF. The problem to do FFSC on hydrolox is that you get something like 80% of the power on the H2 turbine, where you need 30% of your power, and 20% of your power on the O2 turbine where you need 70% of your power. Methalox has an almost 30%/70% ration on both turbines and pumps. That's why going FF means using methalox.

I don't get it. What's the problem with having different sized turbo pumps on the hydrolox FFSC? You still have the advantage of gas in the combustion chamber, quite cool pumps and no complex bearings that need extra helium.

[edkyle99]

But Vulcan is both higher-performing and cheaper than any single-stick Delta IV. This suggest that the higher engine complexity and count actually makes for a better overall system, by simplifying other components and systems. RS-68's fuel adds a lot of tankage and dry mass that adds expense and decreases performance compared to BE-4.

Vulcan is called a "single core" design, different than the Delta 4 single core Medium and triple core Heavy alternatives, but those Vulcan "single cores" are different for the Medium and Heavy Vulcan versions.  The Heavy upper stage is stretched and may use more engines than the Medium upper stage.  Vulcan Heavy also leans heavily on those six big GEM63XL boosters, and requires two BE-4 engines per core rather than only one RS-68 for Delta 4 CBC.  Vulcan's first stage is fatter than Delta 4 CBC, though slightly shorter.  I think we'll see solids on most Vulcans, something not needed by Delta 4 Heavy. 

 - Ed Kyle

[baldusi]

It was hydrolox FRSC, nos FF. The problem to do FFSC on hydrolox is that you get something like 80% of the power on the H2 turbine, where you need 30% of your power, and 20% of your power on the O2 turbine where you need 70% of your power. Methalox has an almost 30%/70% ration on both turbines and pumps. That's why going FF means using methalox.

I don't get it. What's the problem with having different sized turbo pumps on the hydrolox FFSC? You still have the advantage of gas in the combustion chamber, quite cool pumps and no complex bearings that need extra helium.

  I thought you understood the cycle. Let me explain. In a full flow cycle, you take all the fuel, mix it with a tiny bit of oxidizer on the preburner to convert it to hot gas (actually it's critical fluid, but bear with me). The turbines are heat engines, so you run that hot gas through the turbine to power the fuel pump. Remember this: fuel preburner->fuel turbine->fuel pump. The oxidiezer is the same oxidizer preburner->oxidizer turbine->oxidizer pump.
  Now, the oxidizer to fuel ratio (O/F), is given by the main combustion chamber ratio, so that's a given. And how much heat is transfered to the turbines is a function of the amount of gas and it's elements. The formula is very correlated to:
 Power = Massflow x element specific heat.
 It just happens that for hydrolox engines, O/F ratio is somewhere between 5.5 and 6. And hydrogen has a specific heat around 15 times higher than oxygen. That means that you have 15 time more power (per unit of mass) on the hydrogen side, than on the oxygen side.
  So, you can not get higher pressure on the whole engine (and thus efficiency) than your oxygen side, which is really, really under powered. In other words, you are actually better off running the pumps with hydrogen, and still have a higher performing (and cooler turbines) than full flow. The SSME designers knew quite well what they were doing.
  Methalox, on the other side, have a very similar relation of power available to each turbine to power needed by each turbopump. So going full flow does allows you to extract a lot more work.

[Welsh Dragon]

But Vulcan is both higher-performing and cheaper than any single-stick Delta IV. This suggest that the higher engine complexity and count actually makes for a better overall system, by simplifying other components and systems. RS-68's fuel adds a lot of tankage and dry mass that adds expense and decreases performance compared to BE-4.

Vulcan is called a "single core" design, different than the Delta 4 single core Medium and triple core Heavy alternatives, but those Vulcan "single cores" are different for the Medium and Heavy Vulcan versions.  The Heavy upper stage is stretched and may use more engines than the Medium upper stage.  Vulcan Heavy also leans heavily on those six big GEM63XL boosters, and requires two BE-4 engines per core rather than only one RS-68 for Delta 4 CBC.  Vulcan's first stage is fatter than Delta 4 CBC, though slightly shorter.  I think we'll see solids on most Vulcans, something not needed by Delta 4 Heavy. 

 - Ed Kyle

But to be fair, every single Delta IV version also has different cores, every core of the Heavy is also different, so Vulcan is hardly at an advantage or disadvantage there. (Oh, if only we'd have had Atlas V Heavy!)

[su27k]

It was hydrolox FRSC, nos FF. The problem to do FFSC on hydrolox is that you get something like 80% of the power on the H2 turbine, where you need 30% of your power, and 20% of your power on the O2 turbine where you need 70% of your power. Methalox has an almost 30%/70% ration on both turbines and pumps. That's why going FF means using methalox.

There were proposals for hydrolox FFSC though, for example RS-2100. Also IPD is hydrolox.

[LouScheffer]

[...] hydrogen has a specific heat around 15 times higher than oxygen. That means that you have 15 time more power (per unit of mass) on the hydrogen side, than on the oxygen side.
  So, you can not get higher pressure on the whole engine (and thus efficiency) than your oxygen side, which is really, really under powered. In other words, you are actually better off running the pumps with hydrogen, and still have a higher performing (and cooler turbines) than full flow.

I believe that in theory you could just increase the temperature of the oxygen pre-burner until you get the power you need.  But in practice oxygen rich combustion is already a hideous materials problem.  Increasing the temperature enough to balance the pumps probably makes it a problem where no known materials suffice.

[Aeneas]

[...] hydrogen has a specific heat around 15 times higher than oxygen. That means that you have 15 time more power (per unit of mass) on the hydrogen side, than on the oxygen side.
  So, you can not get higher pressure on the whole engine (and thus efficiency) than your oxygen side, which is really, really under powered. In other words, you are actually better off running the pumps with hydrogen, and still have a higher performing (and cooler turbines) than full flow.

I believe that in theory you could just increase the temperature of the oxygen pre-burner until you get the power you need.  But in practice oxygen rich combustion is already a hideous materials problem.  Increasing the temperature enough to balance the pumps probably makes it a problem where no known materials suffice.

But with Raptor, it seems to work with a special alloy - whatever that may be. And RD-180 works with oxygen rich combustion anyway. So this given, why it's not possible to feed some hydrogen into the oxygen rich environment to enable full flow. The hydrogen side shouldn't be a problem anyway since RS-25 already prove it to work.

[Prettz]

It was hydrolox FRSC, nos FF. The problem to do FFSC on hydrolox is that you get something like 80% of the power on the H2 turbine, where you need 30% of your power, and 20% of your power on the O2 turbine where you need 70% of your power. Methalox has an almost 30%/70% ration on both turbines and pumps. That's why going FF means using methalox.

There were proposals for hydrolox FFSC though, for example RS-2100. Also IPD is hydrolox.

There's also more about converting SSME to FFSC here: https://eaglehill.us/spev-pdfs-articles/A005%20Knuth.pdf

[brickmack]

Vulcan is called a "single core" design, different than the Delta 4 single core Medium and triple core Heavy alternatives, but those Vulcan "single cores" are different for the Medium and Heavy Vulcan versions.  The Heavy upper stage is stretched and may use more engines than the Medium upper stage.  Vulcan Heavy also leans heavily on those six big GEM63XL boosters, and requires two BE-4 engines per core rather than only one RS-68 for Delta 4 CBC.  Vulcan's first stage is fatter than Delta 4 CBC, though slightly shorter.  I think we'll see solids on most Vulcans, something not needed by Delta 4 Heavy.

No, the core is identical for all Vulcan variants. Same tank size, and they always include the booster mounts whether they're used or not. Only the upper stage changes, and only the tank length (engines, thrust structure, etc are all identical). The 4 engine configuration was dropped a while ago, can get basically identical performance at drastically lower cost with 2 uprated RL10s (lower thrust, but also lower dry mass, and theres room for larger nozzles so higher ISP). I think the 4 engine option looked a lot more attractive when AR-1 was still on the table and the core stage would provide a lot less performance, needing the upper stage to have more thrust just to reach orbit

Not sure what point you're trying to make with the rest of that

[edkyle99]

Vulcan is called a "single core" design, different than the Delta 4 single core Medium and triple core Heavy alternatives, but those Vulcan "single cores" are different for the Medium and Heavy Vulcan versions.  The Heavy upper stage is stretched and may use more engines than the Medium upper stage.  Vulcan Heavy also leans heavily on those six big GEM63XL boosters, and requires two BE-4 engines per core rather than only one RS-68 for Delta 4 CBC.  Vulcan's first stage is fatter than Delta 4 CBC, though slightly shorter.  I think we'll see solids on most Vulcans, something not needed by Delta 4 Heavy.

No, the core is identical for all Vulcan variants. Same tank size, and they always include the booster mounts whether they're used or not. Only the upper stage changes, and only the tank length (engines, thrust structure, etc are all identical). The 4 engine configuration was dropped a while ago, can get basically identical performance at drastically lower cost with 2 uprated RL10s (lower thrust, but also lower dry mass, and theres room for larger nozzles so higher ISP). I think the 4 engine option looked a lot more attractive when AR-1 was still on the table and the core stage would provide a lot less performance, needing the upper stage to have more thrust just to reach orbit

Not sure what point you're trying to make with the rest of that

I was considering the second stage to be part of the "single core", so different designs for Medium versus Heavy though the first stage itself appears to be the same based on current information.

My point was that Delta 4 Heavy is not that easy to replace.  It requires for Vulcan, a stretched upper stage and full SRB complement, for Falcon Heavy at least a partly expendable flight and a new service tower, and for Omega bigger first and maybe third stages plus strap-on boosters, etc.  Thanks in large part to high-thrust LH2 propulsion of RS-68.

 - Ed Kyle   

[spacenut]

I thought Vulcan was the same diameter as Delta IV and used the same tooling.  5m.  Or did they go 5.5m diameter for Vulcan because of two engines? 

RS-68 if it didn't have ablative nozzle and had a cooled nozzle, may have been a really good engine for SLS core.  However, SLS would have required a decent sized upper stage.

[Aeneas]

I thought Vulcan was the same diameter as Delta IV and used the same tooling.  5m.  Or did they go 5.5m diameter for Vulcan because of two engines? 

It's 5.4 m diameter. Like the SLS Block 1 upper stage.

[baldusi]

[...] hydrogen has a specific heat around 15 times higher than oxygen. That means that you have 15 time more power (per unit of mass) on the hydrogen side, than on the oxygen side.
  So, you can not get higher pressure on the whole engine (and thus efficiency) than your oxygen side, which is really, really under powered. In other words, you are actually better off running the pumps with hydrogen, and still have a higher performing (and cooler turbines) than full flow.

I believe that in theory you could just increase the temperature of the oxygen pre-burner until you get the power you need.  But in practice oxygen rich combustion is already a hideous materials problem.  Increasing the temperature enough to balance the pumps probably makes it a problem where no known materials suffice.

Nobody like to leave free performance on the table. Staged engines are usually limited but the turbine blade materials. You can't really run them any hotter without sacrificing margin. And I'm not saying you can't get a bit better performance by going FFSC on H2 instead of FRSC. I'm saying that you don't get the astounding differences you get on methalox.

[Aeneas]

Nobody like to leave free performance on the table. Staged engines are usually limited but the turbine blade materials. You can't really run them any hotter without sacrificing margin. And I'm not saying you can't get a bit better performance by going FFSC on H2 instead of FRSC. I'm saying that you don't get the astounding differences you get on methalox.

Then I didn't understand it yet. Why is the difference between methalox FRSC vs methalox FFSC larger than hydrolox FRSC vs hydrolox FFSC? Yes, the LH2 pump is much bigger than the LCH4 pump but that's the case for both FRSC and FFSC.
So the only but major difference is spraying a little methane into the LOX stream behaves vastly different than spraying a little hydrogen in the LOX stream?

[Prettz]

[...] hydrogen has a specific heat around 15 times higher than oxygen. That means that you have 15 time more power (per unit of mass) on the hydrogen side, than on the oxygen side.
  So, you can not get higher pressure on the whole engine (and thus efficiency) than your oxygen side, which is really, really under powered. In other words, you are actually better off running the pumps with hydrogen, and still have a higher performing (and cooler turbines) than full flow.

I believe that in theory you could just increase the temperature of the oxygen pre-burner until you get the power you need.  But in practice oxygen rich combustion is already a hideous materials problem.  Increasing the temperature enough to balance the pumps probably makes it a problem where no known materials suffice.

Nobody like to leave free performance on the table. Staged engines are usually limited but the turbine blade materials. You can't really run them any hotter without sacrificing margin. And I'm not saying you can't get a bit better performance by going FFSC on H2 instead of FRSC. I'm saying that you don't get the astounding differences you get on methalox.

Isn't the main point of doing it with hydrolox improving reliability and reusability? (over what the SSME was ever capable of)

[LouScheffer]

So, you can not get higher pressure on the whole engine (and thus efficiency) than your oxygen side, which is really, really under powered. In other words, you are actually better off running the pumps with hydrogen, and still have a higher performing (and cooler turbines) than full flow. The SSME designers knew quite well what they were doing.

I don't see how this can be right.  Take the RD-180 oxygen side.  This uses only the oxygen flow to pump all oxygen up to record-breaking pressures.  Since it's using oxygen flow to pump oxygen, this is (almost) independent of the fuel used.

But the RD-180 is not only using the oxygen flow to pump oxygen, it's also using it to pump fuel up to even higher pressure (because of cooling losses).  Take away the need to pump fuel, and oxygen flow has the power to pump oxygen to even more eye-popping pressures.   So it's really hard for me to see how the oxygen side could limit the attainable pressure of the engine.

[baldusi]

Nobody like to leave free performance on the table. Staged engines are usually limited but the turbine blade materials. You can't really run them any hotter without sacrificing margin. And I'm not saying you can't get a bit better performance by going FFSC on H2 instead of FRSC. I'm saying that you don't get the astounding differences you get on methalox.

Then I didn't understand it yet. Why is the difference between methalox FRSC vs methalox FFSC larger than hydrolox FRSC vs hydrolox FFSC? Yes, the LH2 pump is much bigger than the LCH4 pump but that's the case for both FRSC and FFSC.
So the only but major difference is spraying a little methane into the LOX stream behaves vastly different than spraying a little hydrogen in the LOX stream?

I don't think I was clear enough. You have two sides for each propellant:
1) How much power they can PROVIDE to the turbines (massflow*specific heat).
2) How much power the pumps NEED to increase the pressure to the desired level.

Now, material science is what limits turbines (usually to a max inlet temperature of 750K to 850K). So, you can not get more power than what your masswflow*specific heat provides.
Hydrogen provides so much power, that you might end with actually less power on the LOX side (because you can only use O2 to run the LOX turbine) than splitting the Hydrogen flow (in a Fuel Rich SC). Not really a loss, but your gain is relatively low. Your H2 side will run really cool. But the LOX side will be working well to the material limit to keep up.
In the methalox case, the power GENERATED by the CH4 side is almost what the CH4 pumps need for matching the LOX side pressure (all at the same turbine temperature). So you can go with low performance and super life (low temperature turbines) or dial up the performance as much as you dare, that both sides will be evenly matched.
Again, your LOX side will limit you very fast while the H2 side will basically idle in a FFSC. Was I a bit clearer?

[baldusi]

[...] hydrogen has a specific heat around 15 times higher than oxygen. That means that you have 15 time more power (per unit of mass) on the hydrogen side, than on the oxygen side.
  So, you can not get higher pressure on the whole engine (and thus efficiency) than your oxygen side, which is really, really under powered. In other words, you are actually better off running the pumps with hydrogen, and still have a higher performing (and cooler turbines) than full flow.

I believe that in theory you could just increase the temperature of the oxygen pre-burner until you get the power you need.  But in practice oxygen rich combustion is already a hideous materials problem.  Increasing the temperature enough to balance the pumps probably makes it a problem where no known materials suffice.

Nobody like to leave free performance on the table. Staged engines are usually limited but the turbine blade materials. You can't really run them any hotter without sacrificing margin. And I'm not saying you can't get a bit better performance by going FFSC on H2 instead of FRSC. I'm saying that you don't get the astounding differences you get on methalox.

Isn't the main point of doing it with hydrolox improving reliability and reusability? (over what the SSME was ever capable of)

Hydrolox was used because of performance reasons, not reliability. H2 is horrible to insulate, leaks as crazy and is nasty anywhere where it finds atmosphere. And the engines and tanks are really heavy for the propellant mass. I would argue that reusability was also affected because seals and tanks have to undergo extra stress. Yes, kerosene soots. Yet, Falcon 9 does not seems too concerned with that.
USA was afraid of ORSC, and for upper stages it does offer excellent performance. But look what private industry has embraced for first stages once the sacred mantra of H2 was actually lifted and people who wanted to post a profit on a competitive market had to make the calls.

[Lars-J]

Almost everyone (in the U.S.) at the time had convinced themselves that it was "hydrolox or bust", Isp was all that mattered. Another factor was a loss of cutting edge hydrocarbon engine experience. So with the SSME in production, it turned into a "if all you have is a hammer, every problem looks like a nail" scenario. We have the best hydrolox engines, so naturally any new engine should be hydrolox.

It took SpaceX to finally challenge and break that mindset.

[Prettz]

[...] hydrogen has a specific heat around 15 times higher than oxygen. That means that you have 15 time more power (per unit of mass) on the hydrogen side, than on the oxygen side.
  So, you can not get higher pressure on the whole engine (and thus efficiency) than your oxygen side, which is really, really under powered. In other words, you are actually better off running the pumps with hydrogen, and still have a higher performing (and cooler turbines) than full flow.

I believe that in theory you could just increase the temperature of the oxygen pre-burner until you get the power you need.  But in practice oxygen rich combustion is already a hideous materials problem.  Increasing the temperature enough to balance the pumps probably makes it a problem where no known materials suffice.

Nobody like to leave free performance on the table. Staged engines are usually limited but the turbine blade materials. You can't really run them any hotter without sacrificing margin. And I'm not saying you can't get a bit better performance by going FFSC on H2 instead of FRSC. I'm saying that you don't get the astounding differences you get on methalox.

Isn't the main point of doing it with hydrolox improving reliability and reusability? (over what the SSME was ever capable of)

Hydrolox was used because of performance reasons, not reliability. H2 is horrible to insulate, leaks as crazy and is nasty anywhere where it finds atmosphere. And the engines and tanks are really heavy for the propellant mass. I would argue that reusability was also affected because seals and tanks have to undergo extra stress. Yes, kerosene soots. Yet, Falcon 9 does not seems too concerned with that.
USA was afraid of ORSC, and for upper stages it does offer excellent performance. But look what private industry has embraced for first stages once the sacred mantra of H2 was actually lifted and people who wanted to post a profit on a competitive market had to make the calls.

That's not at all what I was getting at. It's that the main point of moving from a hydrolox FRSC (the SSME) to full-flow is to make the engine more practical for reuse, and possibly easier to build. Not the performance gain, although that's nice to have. It was laid out in that paper I linked to earlier.

[spacenut]

Until SpaceX came along, I don't think any American made kerolox engines were developed since the 1960's.  Atlas I, II, and Delta II used the H1 engine which was used for Saturn I.  The Saturn V F1 engine was the last until Merlin came along.  Atlas V used the Russian RD-180. 

Now SpaceX and Blue Origin have developed metholox engines. 

Blue developed new BE-3's which are far cheaper than any of the other American made hydrolox engines.  It also fills a nich in hydrolox power range, between RL-10 and J2X.  I'm surprised NASA hasn't considered the BE-3U as an upper stage engine for SLS, or just use the New Glenn upper stage on SLS, because of development costs. 

RS-68 is probably the last hydrolox booster engine to ever be developed, as metholox seems to be the best compromise between kerolox reusable boosters and possible hydrolox reusable boosters.  Metholox is non-sooting and clean burning and uses smaller tankage per unit of thrust than hydrolox. 

[baldusi]

[...]
That's not at all what I was getting at. It's that the main point of moving from a hydrolox FRSC (the SSME) to full-flow is to make the engine more practical for reuse, and possibly easier to build. Not the performance gain, although that's nice to have. It was laid out in that paper I linked to earlier.

That paper was that... a paper. Among other things, they reduced Pc from 21.3MPa to 19.3MPa. So they were going to a lower performance engine. There were going for a cheaper and simpler engine, but didn't had any experience in ORSC.

[baldusi]

So, you can not get higher pressure on the whole engine (and thus efficiency) than your oxygen side, which is really, really under powered. In other words, you are actually better off running the pumps with hydrogen, and still have a higher performing (and cooler turbines) than full flow. The SSME designers knew quite well what they were doing.

I don't see how this can be right.  Take the RD-180 oxygen side.  This uses only the oxygen flow to pump all oxygen up to record-breaking pressures.  Since it's using oxygen flow to pump oxygen, this is (almost) independent of the fuel used.

But the RD-180 is not only using the oxygen flow to pump oxygen, it's also using it to pump fuel up to even higher pressure (because of cooling losses).  Take away the need to pump fuel, and oxygen flow has the power to pump oxygen to even more eye-popping pressures.   So it's really hard for me to see how the oxygen side could limit the attainable pressure of the engine.

I didn't said you will get lower performance. I said you will not get as much performance gain when going from FRSC to FFSC with hydrolox as you will with methalox (and you would actually go ORSC to FFSC, but that's a small detail).
The RD-170 family has an oxidizer outlet pressure (output of the main pump) of 60MPa (8,700psi) vs SSME 47.85(6,939psi). So they pushed they pressure a 25% above SSME, while still providing 30% of power to the fuel side. So, you should be able to increase pressure a 60% above that of SSME, or about 34MPa (say 5,000psi). But I remember someone doing some back of the envelop calculation and stating that 50MPa (7,200psi) was possible with methalox under the same assumptions.

[Aeneas]

Blue developed new BE-3's which are far cheaper than any of the other American made hydrolox engines.  It also fills a nich in hydrolox power range, between RL-10 and J2X.  I'm surprised NASA hasn't considered the BE-3U as an upper stage engine for SLS, or just use the New Glenn upper stage on SLS, because of development costs. 

Do you know the Isp for BE-3 and BE-3U?

[Stan-1967]

  I'm surprised NASA hasn't considered the BE-3U as an upper stage engine for SLS, or just use the New Glenn upper stage on SLS, because of development costs.  g and uses smaller tankage per unit of thrust than hydrolox.

NASA did send out an RFI for proposals on a "drop in" upper stage replacement for SLS back in 2017.  It was probably a government formality to open the contract for competitive bids, but the way the contract was written to consider a "drop in" replacement made it pretty much impossible for any other rocket engine to meet the constraints.  Basically it was written around the specifications that 4 x RL-10's can deliver.

Where is this thread about the RS-68 going?  RS-68's fate & legacy were summed up in the first page of posts.  Maybe the thread should be modified to discuss what place hydrolox engine technology will have in new generation rockets?  Hydrolox in general, & the RS-68 in particular look to be relics as far as preferred booster technology goes.  Upper stages & in space applications combined with ISRU look more favorable for hydrolox.

[ZachS09]

Could the RS-68B designed for Ares V have been used on Delta IV?

It has higher specific impulse and thrust.

[envy887]

Vulcan is called a "single core" design, different than the Delta 4 single core Medium and triple core Heavy alternatives, but those Vulcan "single cores" are different for the Medium and Heavy Vulcan versions.  The Heavy upper stage is stretched and may use more engines than the Medium upper stage.  Vulcan Heavy also leans heavily on those six big GEM63XL boosters, and requires two BE-4 engines per core rather than only one RS-68 for Delta 4 CBC.  Vulcan's first stage is fatter than Delta 4 CBC, though slightly shorter.  I think we'll see solids on most Vulcans, something not needed by Delta 4 Heavy.

No, the core is identical for all Vulcan variants. Same tank size, and they always include the booster mounts whether they're used or not. Only the upper stage changes, and only the tank length (engines, thrust structure, etc are all identical). The 4 engine configuration was dropped a while ago, can get basically identical performance at drastically lower cost with 2 uprated RL10s (lower thrust, but also lower dry mass, and theres room for larger nozzles so higher ISP). I think the 4 engine option looked a lot more attractive when AR-1 was still on the table and the core stage would provide a lot less performance, needing the upper stage to have more thrust just to reach orbit

Not sure what point you're trying to make with the rest of that

I was considering the second stage to be part of the "single core", so different designs for Medium versus Heavy though the first stage itself appears to be the same based on current information.

My point was that Delta 4 Heavy is not that easy to replace.  It requires for Vulcan, a stretched upper stage and full SRB complement, for Falcon Heavy at least a partly expendable flight and a new service tower, and for Omega bigger first and maybe third stages plus strap-on boosters, etc.  Thanks in large part to high-thrust LH2 propulsion of RS-68.

 - Ed Kyle

I would primarily measure "hard to replace" in terms of cost for the same performance. Vulcan and FH are both ~1/2 the cost of DIVH, and both outperform DIVH by a substantial margin in most or all cases... which suggests the high thrust hydrogen boost is actually really easy to replace.

[brickmack]

Could the RS-68B designed for Ares V have been used on Delta IV?

It has higher specific impulse and thrust.

Yes, but theres other upgrade paths that would have offered equal or better performance at similar or lower dev cost and similar or lower per-flight cost. Depending on exactly what mission set was being targeted and what commonality with other rockets was deemed economically worthwhile, DIVH with 6 GEM-60s, or a larger upper stage like ACES or maybe even the Ares I upper stage, all seem more attractive. If RS-68B was used on Delta, it probably would be for crewrating purposes, since it would eliminate the fireball and have a bunch of other safety improvements. Higher performance would just be a happy accident

[JEF_300]

** And if you really want to stick with hydrolox, I'd ignore the RS-68 and take a closer look at the LE-9. If JAXA and Mitsubishi can hit the H3's price targets, then its a hydrolox booster engine that might actually be cost effective, at least vis-a-vis the Merlin (absent F9 reusability, at any rate).

What's so good about LE-9 accept the possible price tag?

I know it's off topic for this post, but the question was asked, and I really feel that the LE-9 doesn't get the credit it's due.

Summary: the LE-9 is an Expander-Bleed Cycle Hydrolox sea-level engine developed by Mitsubishi Heavy Industries. It's kinda like the lovechild of an RL-10 and an NK-33, and it's great, and yet its also apparently inexpensive.

Stats grabbed from the LE-9 Wikipedia page: https://en.wikipedia.org/wiki/LE-9

Dry weight                        2.4 t (5,300 lb)
Thrust (vac.)                     1,471 kN (331,000 lbf)
Isp (vac.)                         426 s (4.18 km/s)
Chamber pressure             10.0 MPa (1,450 psi)
Thrust-to-weight ratio        62.50
Mixture ratio                     5.9

And keep in mind that MHI wants to price the H3, which has either 2 or 3 of these engines on it's first stage, on-par with the Falcon 9.

It's an Expander-Bleed Cycle, which is the same cycle as the RL-10, but instead of dumping the expanded gasses used to run the pumps into the combustion chamber, they're dumped overboard (or more specifically, into the engine bell, which is the same thing as overboard to the people who named rocket engine cycles).

The RL-10 has to maintain a relatively low chamber pressure for it to be possible to dump the pump gas into the chamber. The LE-9, and all other Expander-Bleed Cycle engines, don't have this limitation, which makes them much easier to scale and more applicable to high-thrust designs, at the cost of some isp.

Expander Cylce designs are also a very simple and reliable, which is part of why the RL-10 has been so successful.

[baldusi]

It's an Expander-Bleed Cycle, which is the same cycle as the RL-10, but instead of dumping the expanded gasses used to run the pumps into the combustion chamber, they're dumped overboard (or more specifically, into the engine bell, which is the same thing as overboard to the people who named rocket engine cycles).

The RL-10 has to maintain a relatively low chamber pressure for it to be possible to dump the pump gas into the chamber. The LE-9, and all other Expander-Bleed Cycle engines, don't have this limitation, which makes them much easier to scale and more applicable to high-thrust designs, at the cost of some isp.

Expander Cylce designs are also a very simple and reliable, which is part of why the RL-10 has been so successful.

No it's not the same cycle. Expander Bleed Cycle has a lot more in common with Gas Generator (like the YF-77). In the Expander Cycle all of fuel is expanded through the cooling channels and used to drive the turbines. Then, it is injected as a gas into the combustion chamber. The problem is that your heating source growths to the square of your scaling of the engine, but your propellant mass growths to the cube. So you can not scale that much.
The Expander Bleed is a Gas Generator implementation on the expander cycle: you take a very small amount of propellant (say, 2%), get it as hot as your turbine's blades can take, and after passing them through the turbines you dump it. So, you still inject liquid fuel into the main combustion chamber.
So, overall, the RL-10 is a closed cycle, gas-liquid engine with turbines designed to take the whole mass of the fuel while the LE-9 is an open cycle, liquid-liquid engine with turbines sized for just 2% of the fuel mass. Anywhere you look, unless you talk about start-up, LE-9 looks just like the YF-77. Even have same YF-77 chamber pressure.

[MATTBLAK]

Could the RS-68B designed for Ares V have been used on Delta IV?

It has higher specific impulse and thrust.

Certainly. If the Delta IV-Heavy had used the RS-68B engine and had improvements such as aluminum-lithium structures, propellant densification, the 5 meter upper stage stretched and a second RL-10 engine added - that booster would get well over 30 metric tons into Low Earth Orbit and send about 14 tons on Earth Escape trajectories. Add 6x GEM-60 solid boosters on one side of the trio of boosters and the LEO payload would have been increased about another 10 metric tons and the Earth Escape payload might have approached 17 tons - roughly the equivalent of the Falcon Heavy all round. Still wouldn't have been as cheap, though.

[ZachS09]

Could the RS-68B designed for Ares V have been used on Delta IV?

It has higher specific impulse and thrust.

Certainly. If the Delta IV-Heavy had used the RS-68B engine and had improvements such as aluminum-lithium structures, propellant densification, the 5 meter upper stage stretched and a second RL-10 engine added - that booster would get well over 30 metric tons into Low Earth Orbit and send about 14 tons on Earth Escape trajectories. Add 6x GEM-60 solid boosters on one side of the trio of boosters and the LEO payload would have been increased about another 10 metric tons and the Earth Escape payload might have approached 17 tons - roughly the equivalent of the Falcon Heavy all round. Still wouldn't have been as cheap, though.

What currently flying upper stage would the stretched DCSS-5 with two RL-10 engines resemble?

[baldusi]

Could the RS-68B designed for Ares V have been used on Delta IV?

It has higher specific impulse and thrust.

Certainly. If the Delta IV-Heavy had used the RS-68B engine and had improvements such as aluminum-lithium structures, propellant densification, the 5 meter upper stage stretched and a second RL-10 engine added - that booster would get well over 30 metric tons into Low Earth Orbit and send about 14 tons on Earth Escape trajectories. Add 6x GEM-60 solid boosters on one side of the trio of boosters and the LEO payload would have been increased about another 10 metric tons and the Earth Escape payload might have approached 17 tons - roughly the equivalent of the Falcon Heavy all round. Still wouldn't have been as cheap, though.

What currently flying upper stage would the stretched DCSS-5 with two RL-10 engines resemble?

While Boeing's growth plan included a stretched DIVUS, once on ULA, the ACES was proposed as a common US. But then they decided Vulcan was a cheaper and simpler solution.

[ulm_atms]

DCSS could be stretched/enhanced quite a lot just by swapping out to a single BE3U.  I still don't understand why anyone would want multiple RL-10s over a single BE3U.  Size the US to the BE3U's specs...and you get a MUCH cheaper, MUCH more payload mass throwing US....but that has nothing to do with the RS-68...

As far as the RS-68.  Hydrolox boosters are just not mass efficient is the main problem as a SLT boosters.  RP-1 seems about the best except for the sooting issue if wanting to reuse.  Metholox does seem to be the best compromise between them IMO.

Biggest thing against the RS-68 is how HEAVY they are compared to every other hydrolox engine out there...

[Welsh Dragon]

DCSS could be stretched/enhanced quite a lot just by swapping out to a single BE3U.  I still don't understand why anyone would want multiple RL-10s over a single BE3U.  Size the US to the BE3U's specs...and you get a MUCH cheaper, MUCH more payload mass throwing US....but that has nothing to do with the RS-68...

As far as the RS-68.  Hydrolox boosters are just not mass efficient is the main problem as a SLT boosters.  RP-1 seems about the best except for the sooting issue if wanting to reuse.  Metholox does seem to be the best compromise between them IMO.

Biggest thing against the RS-68 is how HEAVY they are compared to every other hydrolox engine out there...

Because the RL-10 exists now and is a known quantity. That remains to be seen for the BE3U*. And no, because the BE3 has flown doesn't mean the BE3U has been validated. It's even a different engine cycle.

*That doesn't mean I think it won't fly, but right now it hasn't.

[ulm_atms]

DCSS could be stretched/enhanced quite a lot just by swapping out to a single BE3U.  I still don't understand why anyone would want multiple RL-10s over a single BE3U.  Size the US to the BE3U's specs...and you get a MUCH cheaper, MUCH more payload mass throwing US....but that has nothing to do with the RS-68...

As far as the RS-68.  Hydrolox boosters are just not mass efficient is the main problem as a SLT boosters.  RP-1 seems about the best except for the sooting issue if wanting to reuse.  Metholox does seem to be the best compromise between them IMO.

Biggest thing against the RS-68 is how HEAVY they are compared to every other hydrolox engine out there...

Because the RL-10 exists now and is a known quantity. That remains to be seen for the BE3U*. And no, because the BE3 has flown doesn't mean the BE3U has been validated. It's even a different engine cycle.

*That doesn't mean I think it won't fly, but right now it hasn't.

Well...there two rockets being built with BE4s...and they haven't flown either...but I get your point.  The BE3U has got to be about as far along as BE4...or close to it for NG to be built however.

EDIT TO ADD:

On the RS-68...I think I would take 6(up to 9?) BE3s then 1 RS-68.  I think you still come out lighter with 6 which is a direct thrust-to-thrust match...but I can't seem to find a dry mass number for BE3 to verify.

[Aeneas]

DCSS could be stretched/enhanced quite a lot just by swapping out to a single BE3U.  I still don't understand why anyone would want multiple RL-10s over a single BE3U.  Size the US to the BE3U's specs...and you get a MUCH cheaper, MUCH more payload mass throwing US....but that has nothing to do with the RS-68...

As far as the RS-68.  Hydrolox boosters are just not mass efficient is the main problem as a SLT boosters.  RP-1 seems about the best except for the sooting issue if wanting to reuse.  Metholox does seem to be the best compromise between them IMO.

Biggest thing against the RS-68 is how HEAVY they are compared to every other hydrolox engine out there...

Which ends still with the open question: What are the stats of the BE-3U?

[edkyle99]

Which ends still with the open question: What are the stats of the BE-3U?

Blue Origin, at:
https://www.blueorigin.com/engines/be-3
only tells us that it will produce 710 kN (160,000 lbf) thrust in vacuum, but there is a nice image of the engine on the page that someone might be able to use to reverse-engineer it a bit.  Since it is expander bleed cycle, it should have lower ISP than RL-10 and Vinci (which are 460-465 sec), but more thrust.  On the other hand it should or could have higher ISP than J-2X (448 sec).  Then again, I'm not a combustion engineer, so I could be wrong!  There is a chance that BE-3U could end up being a more important engine in future decades than BE-4.

 - Ed Kyle

[john smith 19]

The Expander Bleed is a Gas Generator implementation on the expander cycle: you take a very small amount of propellant (say, 2%), get it as hot as your turbine's blades can take, and after passing them through the turbines you dump it. So, you still inject liquid fuel into the main combustion chamber.

So no actual gas generator involved.

That means that the combustion chamber pressure is not limited to a level that the drive turbine can generate from the heat extracted from the CC.

It also means no GG to generate, start or re-start. The lightest, cheapest parts are the ones that aren't there.

And that T/W ratio for an LH2 engine is phenomenal

[JEF_300]

Which ends still with the open question: What are the stats of the BE-3U?

Blue Origin, at:
https://www.blueorigin.com/engines/be-3
only tells us that it will produce 710 kN (160,000 lbf) thrust in vacuum, but there is a nice image of the engine on the page that someone might be able to use to reverse-engineer it a bit.  Since it is expander bleed cycle, it should have lower ISP than RL-10 and Vinci (which are 460-465 sec), but more thrust.  On the other hand it should or could have higher ISP than J-2X (448 sec).  Then again, I'm not a combustion engineer, so I could be wrong!  There is a chance that BE-3U could end up being a more important engine in future decades than BE-4.

 - Ed Kyle

To my knowledge (maybe one of you more read up on China or India will correct me here), the only expander bleed upper stage engine in the world is the LE-5(A, B, B-2, and soon B-3) used on the H-II and soon the H3.

Rather than drop a list of specifications here, i'm just gonna add a screenshot of the chart on the LE-5 wiki page.

NOTE: The original LE-5 was gas generator, but all variants beyond it are expander-bleed.

What I find interesting, as a side note, is that between the A and B, they intentionally made changes that cost them 5 seconds of isp in order to increase thrust, make the engine more reliable, and cheaper to manufacture. That's very practical of MHI.

[baldusi]

The Expander Bleed is a Gas Generator implementation on the expander cycle: you take a very small amount of propellant (say, 2%), get it as hot as your turbine's blades can take, and after passing them through the turbines you dump it. So, you still inject liquid fuel into the main combustion chamber.

So no actual gas generator involved.

That means that the combustion chamber pressure is not limited to a level that the drive turbine can generate from the heat extracted from the CC.

It also means no GG to generate, start or re-start. The lightest, cheapest parts are the ones that aren't there.

And that T/W ratio for an LH2 engine is phenomenal

Yes it still is limited. In the closed-expander all the fuel is heated and passes through the turbine. In the bleed-expander, something like a 2% of fuel goes through the turbine. You can extract more work since you can extract more heat and have higher pressure losses. And you don't need the pump to push the pressure of all the fuel high enough that even after passing through the cooling channels and the turbines it still has main combustion chamber pressure. So you need less pump work. But you also use a lot less propellant.
As you go up in trust, you need more and more proportion of fuel (because it's heated less). Only that that relationship is a lot flatter than the closed case.

[sdsds]

World’s Most Powerful Hydrogen-Fueled Rocket Engine Completes Final Acceptance Test for ULA Delta IV Heavy Launch Vehicle

https://www.rocket.com/article/world%E2%80%99s-most-powerful-hydrogen-fueled-rocket-engine-completes-final-acceptance-test-ula

“We’ve continued to improve the RS-68 engine, which today remains the most powerful hydrogen-fueled rocket engine in the world,” said Jim Maser, Aerojet Rocketdyne senior vice president of Space. “This engine was developed entirely with company funds to be a very cost competitive and extremely reliable booster engine.”

This assertion from Maser (my emboldening above) caught me off guard.

Weren't the RS-68A improvements funded by the USAF?

[spacenut]

IF, big IF, they replaced the ablative nozzle on RS-68 and used 4 of them on SLS, wouldn't they have to have a larger upper stage?  This would be because the RS-68's would use more fuel than the RS-25's, right?  And IF they used a pair of BE-3U's on say a larger upper stage, wider anyway to leave room for cargo, what would be the SLS payload capability?

[Alberto-Girardi]

IF, big IF, they replaced the ablative nozzle on RS-68 and used 4 of them on SLS, wouldn't they have to have a larger upper stage?  This would be because the RS-68's would use more fuel than the RS-25's, right?  And IF they used a pair of BE-3U's on say a larger upper stage, wider anyway to leave room for cargo, what would be the SLS payload capability?

Maybe, but changing to a regeneratively cooled nozzle will be redesigning the engine, and SLS tank are built for RS-25, that uses staged combustion, unlicke the rs.68 that uses simpler and less efficent gas generator. Maybe even the proportion of fuel that the two engine need is different, given the different cycles. At the end you would need to re design and rebuilt the SLS.

[Hog]

IF, big IF, they replaced the ablative nozzle on RS-68 and used 4 of them on SLS, wouldn't they have to have a larger upper stage?  This would be because the RS-68's would use more fuel than the RS-25's, right?  And IF they used a pair of BE-3U's on say a larger upper stage, wider anyway to leave room for cargo, what would be the SLS payload capability?

Remember the original 10 meter rocket, Ares-V?  6 RS-68 engines and dual 5.5 segment SRBs?

Core Stage
"The 6 RS-68 engines powering the Core will fly at 108 percent power levels (6 percent higher than used on Delta-IV currently) and will each produce 702,055 lbs of thrust and have an Isp of 365 seconds at sea level and will have 797,000 lb of thrust and will have an Isp of 414 seconds in a vacuum."  NSF David Harris
 4,212,330 million pounds thrust for the 6 core stage engines

5.5 segment SRBs
"Each new reusable 5.5 segment SRB, will contain over 1.5 million pounds of propellant which will produce a peak of 3,774,000 million lbs of thrust and will have a vacuum Isp of 275.5 seconds. The 38 percent larger SRB’s will burn for 116 seconds – a full 8 seconds shorter burn time than Space Shuttle – before being jettisoned."  NSF David Harris

7,548,000 pounds thrust for the booster stage

Core Stage + Booster Stage= 4,212,330 + 7,548,000=EDIT make that 11,760,330 pounds of thrust (~52,313 kilonewtons) pounds thrust off the pad.  Would have been a monster.  The N-1 moon rocket was 10,200,000 lbf/45,400 kN off the pad.

[ZachS09]

IF, big IF, they replaced the ablative nozzle on RS-68 and used 4 of them on SLS, wouldn't they have to have a larger upper stage?  This would be because the RS-68's would use more fuel than the RS-25's, right?  And IF they used a pair of BE-3U's on say a larger upper stage, wider anyway to leave room for cargo, what would be the SLS payload capability?

Remember the original 10 meter rocket, Ares-V?  6 RS-68 engines and dual 5.5 segment SRBs?

Core Stage
"The 6 RS-68 engines powering the Core will fly at 108 percent power levels (6 percent higher than used on Delta-IV currently) and will each produce 702,055 lbs of thrust and have an Isp of 365 seconds at sea level and will have 797,000 lb of thrust and will have an Isp of 414 seconds in a vacuum."  NSF David Harris
 4,212,330 million pounds thrust for the 6 core stage engines

5.5 segment SRBs
"Each new reusable 5.5 segment SRB, will contain over 1.5 million pounds of propellant which will produce a peak of 3,774,000 million lbs of thrust and will have a vacuum Isp of 275.5 seconds. The 38 percent larger SRB’s will burn for 116 seconds – a full 8 seconds shorter burn time than Space Shuttle – before being jettisoned."  NSF David Harris

7,548,000 pounds thrust for the booster stage

Core Stage + Booster Stage= 4,212,330 + 7,548,000=8,424,660 pounds thrust off the pad.  Would have been a monster.  The N-1 moon rocket was 10,200,000 lbf/45,400 kN off the pad.

Actually, it's 11,760,330 pounds of thrust (~52,313 kilonewtons).

[Hog]

IF, big IF, they replaced the ablative nozzle on RS-68 and used 4 of them on SLS, wouldn't they have to have a larger upper stage?  This would be because the RS-68's would use more fuel than the RS-25's, right?  And IF they used a pair of BE-3U's on say a larger upper stage, wider anyway to leave room for cargo, what would be the SLS payload capability?

Remember the original 10 meter rocket, Ares-V?  6 RS-68 engines and dual 5.5 segment SRBs?

Core Stage
"The 6 RS-68 engines powering the Core will fly at 108 percent power levels (6 percent higher than used on Delta-IV currently) and will each produce 702,055 lbs of thrust and have an Isp of 365 seconds at sea level and will have 797,000 lb of thrust and will have an Isp of 414 seconds in a vacuum."  NSF David Harris
 4,212,330 million pounds thrust for the 6 core stage engines

5.5 segment SRBs
"Each new reusable 5.5 segment SRB, will contain over 1.5 million pounds of propellant which will produce a peak of 3,774,000 million lbs of thrust and will have a vacuum Isp of 275.5 seconds. The 38 percent larger SRB’s will burn for 116 seconds – a full 8 seconds shorter burn time than Space Shuttle – before being jettisoned."  NSF David Harris

7,548,000 pounds thrust for the booster stage

Core Stage + Booster Stage= 4,212,330 + 7,548,000=8,424,660 pounds thrust off the pad.  Would have been a monster.  The N-1 moon rocket was 10,200,000 lbf/45,400 kN off the pad.

Actually, it's 11,760,330 pounds of thrust (~52,313 kilonewtons).

Excellent, do you happen to have a quick breakdown of the core vs booster thrust?   I can't seem to find an off the pad/sea level thrust figure for the 5.5 segment boosters.

This read agrees with you stating 5306 tonnes/11,697,727 pounds/thrust/52,034kN.

""One Later Design - Six RS-68s and Two 5.5 SRBs

On June 23, 2008, NASA's Constellation Program Manager Jeff Hanley announced that the Ares V baseline design had grown larger and more powerful than previous designs.  The trans-lunar insertion (TLI) capability had risen to 71 tonnes, a 7 tonne increase.   The theoretical Low Earth Orbit (LEO) payload had grown to 145 tonnes or more from the previous 130-ish tonnes (see Table 1 for Details).

Planners added a sixth RS-68 engine to the core stage and specified a pair of five-and-a-half segment solid rocket boosters.  The previous design used five RS-68s and two five-segment boosters.  The core stage was lengthened.  The Earth Departure Stage (EDS) diameter had already been increased to 10 meters to match the core diameter during recent design iterations.  The payload fairing diameter also grew to 10 meters.  Overall height jumped to 116.16 meters, 5.56 meters taller than Saturn V.  Liftoff weight increased to 3,699.23 tonnes and liftoff thrust to an unprecedented 5,306 tonnes. """
https://www.spacelaunchreport.com/ares5.html

[ZachS09]

IF, big IF, they replaced the ablative nozzle on RS-68 and used 4 of them on SLS, wouldn't they have to have a larger upper stage?  This would be because the RS-68's would use more fuel than the RS-25's, right?  And IF they used a pair of BE-3U's on say a larger upper stage, wider anyway to leave room for cargo, what would be the SLS payload capability?

Remember the original 10 meter rocket, Ares-V?  6 RS-68 engines and dual 5.5 segment SRBs?

Core Stage
"The 6 RS-68 engines powering the Core will fly at 108 percent power levels (6 percent higher than used on Delta-IV currently) and will each produce 702,055 lbs of thrust and have an Isp of 365 seconds at sea level and will have 797,000 lb of thrust and will have an Isp of 414 seconds in a vacuum."  NSF David Harris
 4,212,330 million pounds thrust for the 6 core stage engines

5.5 segment SRBs
"Each new reusable 5.5 segment SRB, will contain over 1.5 million pounds of propellant which will produce a peak of 3,774,000 million lbs of thrust and will have a vacuum Isp of 275.5 seconds. The 38 percent larger SRB’s will burn for 116 seconds – a full 8 seconds shorter burn time than Space Shuttle – before being jettisoned."  NSF David Harris

7,548,000 pounds thrust for the booster stage

Core Stage + Booster Stage= 4,212,330 + 7,548,000=8,424,660 pounds thrust off the pad.  Would have been a monster.  The N-1 moon rocket was 10,200,000 lbf/45,400 kN off the pad.

Actually, it's 11,760,330 pounds of thrust (~52,313 kilonewtons).

Excellent, do you happen to have a quick breakdown of the core vs booster thrust?   I can't seem to find an off the pad/sea level thrust figure for the 5.5 segment boosters.

This read agrees with you stating 5306 tonnes/11,697,727 pounds/thrust/52,034kN.

""One Later Design - Six RS-68s and Two 5.5 SRBs

On June 23, 2008, NASA's Constellation Program Manager Jeff Hanley announced that the Ares V baseline design had grown larger and more powerful than previous designs.  The trans-lunar insertion (TLI) capability had risen to 71 tonnes, a 7 tonne increase.   The theoretical Low Earth Orbit (LEO) payload had grown to 145 tonnes or more from the previous 130-ish tonnes (see Table 1 for Details).

Planners added a sixth RS-68 engine to the core stage and specified a pair of five-and-a-half segment solid rocket boosters.  The previous design used five RS-68s and two five-segment boosters.  The core stage was lengthened.  The Earth Departure Stage (EDS) diameter had already been increased to 10 meters to match the core diameter during recent design iterations.  The payload fairing diameter also grew to 10 meters.  Overall height jumped to 116.16 meters, 5.56 meters taller than Saturn V.  Liftoff weight increased to 3,699.23 tonnes and liftoff thrust to an unprecedented 5,306 tonnes. """
https://www.spacelaunchreport.com/ares5.html

That, I never thought of yet. And I honestly don’t know what the “off the pad/sea level thrust” thing means.

[Hog]

IF, big IF, they replaced the ablative nozzle on RS-68 and used 4 of them on SLS, wouldn't they have to have a larger upper stage?  This would be because the RS-68's would use more fuel than the RS-25's, right?  And IF they used a pair of BE-3U's on say a larger upper stage, wider anyway to leave room for cargo, what would be the SLS payload capability?

Remember the original 10 meter rocket, Ares-V?  6 RS-68 engines and dual 5.5 segment SRBs?

Core Stage
"The 6 RS-68 engines powering the Core will fly at 108 percent power levels (6 percent higher than used on Delta-IV currently) and will each produce 702,055 lbs of thrust and have an Isp of 365 seconds at sea level and will have 797,000 lb of thrust and will have an Isp of 414 seconds in a vacuum."  NSF David Harris
 4,212,330 million pounds thrust for the 6 core stage engines

5.5 segment SRBs
"Each new reusable 5.5 segment SRB, will contain over 1.5 million pounds of propellant which will produce a peak of 3,774,000 million lbs of thrust and will have a vacuum Isp of 275.5 seconds. The 38 percent larger SRB’s will burn for 116 seconds – a full 8 seconds shorter burn time than Space Shuttle – before being jettisoned."  NSF David Harris

7,548,000 pounds thrust for the booster stage

Core Stage + Booster Stage= 4,212,330 + 7,548,000=8,424,660 pounds thrust off the pad.  Would have been a monster.  The N-1 moon rocket was 10,200,000 lbf/45,400 kN off the pad.

Actually, it's 11,760,330 pounds of thrust (~52,313 kilonewtons).

Excellent, do you happen to have a quick breakdown of the core vs booster thrust?   I can't seem to find an off the pad/sea level thrust figure for the 5.5 segment boosters.

This read agrees with you stating 5306 tonnes/11,697,727 pounds/thrust/52,034kN.

""One Later Design - Six RS-68s and Two 5.5 SRBs

On June 23, 2008, NASA's Constellation Program Manager Jeff Hanley announced that the Ares V baseline design had grown larger and more powerful than previous designs.  The trans-lunar insertion (TLI) capability had risen to 71 tonnes, a 7 tonne increase.   The theoretical Low Earth Orbit (LEO) payload had grown to 145 tonnes or more from the previous 130-ish tonnes (see Table 1 for Details).

Planners added a sixth RS-68 engine to the core stage and specified a pair of five-and-a-half segment solid rocket boosters.  The previous design used five RS-68s and two five-segment boosters.  The core stage was lengthened.  The Earth Departure Stage (EDS) diameter had already been increased to 10 meters to match the core diameter during recent design iterations.  The payload fairing diameter also grew to 10 meters.  Overall height jumped to 116.16 meters, 5.56 meters taller than Saturn V.  Liftoff weight increased to 3,699.23 tonnes and liftoff thrust to an unprecedented 5,306 tonnes. """
https://www.spacelaunchreport.com/ares5.html

That, I never thought of yet. And I honestly don’t know what the “off the pad/sea level thrust” thing means.

"Off the pad/sea level thrust" meaning the sea level rating of the 5.5 segment boosters vs their vacuum thrust rating.rating

[ZachS09]

Okay. Using http://www.b14643.de/ as a source, here's the sea level/vacuum thrust rating of the Ares V 5.5 RSRMs.

Sea level average: 12,689.6 kN (~2,852,736 lbf)

Sea level maximum: 16,651 kN (~3,743,294 lbf)

Vacuum average: 14,000.5 kN (~3,147,438 lbf)

Vacuum maximum: 18,371.5 kN (~4,130,078 lbf)

[Hog]

IF, big IF, they replaced the ablative nozzle on RS-68 and used 4 of them on SLS, wouldn't they have to have a larger upper stage?  This would be because the RS-68's would use more fuel than the RS-25's, right?  And IF they used a pair of BE-3U's on say a larger upper stage, wider anyway to leave room for cargo, what would be the SLS payload capability?

Remember the original 10 meter rocket, Ares-V?  6 RS-68 engines and dual 5.5 segment SRBs?

Core Stage
"The 6 RS-68 engines powering the Core will fly at 108 percent power levels (6 percent higher than used on Delta-IV currently) and will each produce 702,055 lbs of thrust and have an Isp of 365 seconds at sea level and will have 797,000 lb of thrust and will have an Isp of 414 seconds in a vacuum."  NSF David Harris
 4,212,330 million pounds thrust for the 6 core stage engines

5.5 segment SRBs
"Each new reusable 5.5 segment SRB, will contain over 1.5 million pounds of propellant which will produce a peak of 3,774,000 million lbs of thrust and will have a vacuum Isp of 275.5 seconds. The 38 percent larger SRB’s will burn for 116 seconds – a full 8 seconds shorter burn time than Space Shuttle – before being jettisoned."  NSF David Harris

7,548,000 pounds thrust for the booster stage

Core Stage + Booster Stage= 4,212,330 + 7,548,000=8,424,660 pounds thrust off the pad.  Would have been a monster.  The N-1 moon rocket was 10,200,000 lbf/45,400 kN off the pad.

Actually, it's 11,760,330 pounds of thrust (~52,313 kilonewtons).

Jeez, I thought you were implying my input numbers were incorrect, when in fact it was my addition.

 It seems weird that the 5.5 segment SRBs don't perform better as when compared to the SLS 5 segment RSRMV, but I must remember that the  6 RS-68/5.5 segment iteration of ARES-V used SRBs that were reusable and were laden with parachutes and all the other recovery equipment.