Saturday, July 23, 2022

A Reusable SLS? UPDATED: July 31, 2022.

 Copyright 2022 Robert Clark

 The SLS is now projected to cost $4 .1 billion per flight. Because of that severe cost it is projected to only fly once per year. This can not form the basis of a sustainable Moon colonization plan. But suppose we could make the SLS reusable? It’s already known the side boosters can be made reusable as with the shuttle program. The engines on the SLS core stage were derived from the shuttle engines which were intended to be reused up to 100 times. However, since the SLS was intended to be expendable the shuttle-derived engines on the core were designed cheaper to be expendable. However, any rocket engine even an expendable in reality is reusable at least 10 times or more. This is because they have to be certified for several firings for testing purposes. This is described by the well-regarded space expert Henry Spencer:


From: Henry Spencer <>


Subject: RLV engines (was Re: X-33 Concepts: Lockheed, Mac Dac, Rockwell)

Date: Wed, 19 Jun 1996 13:03:12 GMT

In article <4q6am4$> (Andy Haber) writes:

>I think this is an area where critics can speak the loudest.  Today's 

>existing engines all leave something be be desired as true, good SSTO engines.

>This is mostly due to history.  Most engines (other than SSME's) were 

>designed for ELV's, not SSTO's.

Actually, this does not have a lot of bearing on their suitability for

RLVs.  Most ELV engines are, despite their application, reusable, because

they have to be developed and tested.  The F-1 was specified for 20 starts

and 2250s of life, the J-2 for 30 and 3750s.  Six F-1s ran over 5000s each

as part of the service-life tests.  DC-X's RL10s looked "pristine" after

20 starts; the RL10 is nominally rated for 10 starts and 4000s of firing.

>...In terms of using SSME's, sure those can used,

>although doing something to reduce the required level on maintenance on

>the existing engines is quite desirable...

Unfortunately, it probably can't go far enough.  Rocketdyne's own estimate

was that, with a *lot* of work, you could probably get SSME maintenance

costs down to $750k/engine/flight, which is unsatisfactory if you're aiming

for really large cost reductions.


If we feared danger, mankind would never           |       Henry Spencer

go to space.                  --Ellison S. Onizuka |


 Then even reusing the vehicle 10 times could result in a factor of 10 reduction of launch cost, if the maintenance cost could be kept relatively low. That quote about $750, 000 maintenance cost after a lot of work may seem low but from memory I recall it being in the range of $1 million to $2 million per engine after several years into the shuttle program.


 But how to land the SLS core? Starting the SSME’s is a complex process. Modifying them to be air-startable would not be trivial. Instead, I suggest using the method proposed for making the Centaur a lunar lander, multiple pressure-fed side thrusters for a horizontal landing. 

Robust Lunar Exploration Using an Efficient Lunar Lander Derived from Existing Upper Stages. 

 Note then that for a stage reentering to Earth broad-side almost all the reentry velocity is burned off aerodynamically just by air drag so that the stage reaches terminal velocity at approx. 100 m/s. For a stage nearly empty of fuel, this low amount of velocity could be cancelled relatively easily by pressure-fed thrusters with the thrusters running on just the residual of propellant left in the tanks.

 About the landing, there would be additional development cost for the horizontal landing thrusters.  But pressure-fed thrusters are a relatively simple technology. Compare for example the time SpaceX spent developing the Draco thrusters on the Dragon to the time developing the Merlin engine. And from discussion of the thrusters on the Starship they seem more like an afterthought compared to the cost, time, and complexity put into the Raptor engines.

 How about giving the RS-25’s on the SLS core restart capability? Again I’ll refer to the redoubtable Henry Spencer:



From: (Henry Spencer)

Subject: Re: One part Oxygen, two parts Hydrogen and BOOM!

Date: Sat, 14 Oct 2000 03:37:23 GMT

In article <>,

Pete Zaitcev <> wrote:

>> The SSMEs use "torch" igniters, little oxygen/hydrogen burners firing into

>> the preburners and chambers.  The igniters themselves are ignited by,

>> essentially, high-tech spark plugs.


>I see... obviously there cannot be a spark in a vacuum.

Not entirely true, but irrelevant -- when the igniter fires up, there's an

oxygen/hydrogen gas mixture there for the spark to travel through.

>Is the plug the reason engines cannot be restarted in orbit or

>there is more to the story?

There's nothing *fundamental* in the SSME which makes an in-space restart

impossible -- no one-shot parts or anything like that -- but it's a

complicated engine which has to be set up exactly right for a successful

start, and ground equipment (and gravity!) helps out with that.  It would

not be difficult to develop a variant which could start itself in space,

but there has been no reason to do that.


Microsoft shouldn't be broken up.       |  Henry Spencer

It should be shut down.  -- Phil Agre   |      (aka


 So likely it could be done by Aerojet, but I have no confidence they could do it in an affordable manner. Or more precisely, I have no confidence they would do it at an affordable price charged to NASA. For instance the RS-25 engine used on the SLS is derived from the SSME. It was expected to be cheaper than the SSME as it it used a lower parts counts and was not required to have the 100 times reusability of the SSME. But instead Aerojet charged more for this engine than the SSME even when accounting for inflation:

NASA will pay a staggering $146 million for each SLS rocket engine. The rocket needs four engines, and it is expendable. ERIC BERGER - 5/1/2020, 6:55 PM

 About the payload lost on reusability, a stage that goes to LEO can remain in orbit for a few orbits to come back over the landing site so minimal propellant is burned to return to launch site. 

 If we do use a large upper stage, then the SLS would not go to orbit and as SpaceX showed you would need minimal fuel burned if landed down range, and so minimal payload lost, rather than returning to launch site. However, there is then the cost of the upper stage. If it were the Ariane 5/6, the cost of the Ariane 6 being as low as $77 million, it should be even lower than that without the Ariane side boosters or upper stage. 

ARIANE 6 VS. SPACEX: HOW THE ROCKETS STACK UP The European Space Agency is planning to use the Ariane 6 for a variety of missions. ESA MIKE BROWN 1.24.2022 2:00 PM In January 2021, Politico reported that the Ariane 6 could launch for as little as $77 million. That’s a steep discount from the $177 million price tag for the Ariane 5.

  About the landing thrusters, I wouldn’t give a contract for it to any of the usual aerospace companies under NASA’s cost-plus contracts. Instead I would prefer doing it “in house”, so to speak. I was quite impressed by a team at Johnson Space Center led by chief NASA engineer Stephen Altemus developing an unmanned lunar lander for only $14 million development cost:

The Morpheus lunar lander as a manned lander for the Moon.

  The approach the NASA team used on saving costs was likely analogous to that used by commercial space in cutting costs.  No doubt also the pressure-fed engines being used rather than complex turbo-pump engines contributed to the low development cost.


 This report estimates the launch market as approx. $48 Billion per year by 2030:

Global Space Launch Services Market is projected to reach at a market value of US$ 47.6 Billion by 2030: Visiongain Research Inc October 05, 2021 09:33 ET | Source: Visiongain Ltd

 At a going rate of approx. $10,000 per kilo to LEO that would amount to 4,800 tons to orbit.  For a SLS lofting nearly 100 tons to orbit even in SLS 1 form, that’s quite a lot of launches it could take part in per year IF it could do it at a competitive price. If  it could do 10 reuses, that could bring the price down to $400 million per flight, or $4,000 per kilo, about the price of the reusable F9 when new, or a bit more than $3,000 per kilo of the used F9. But IF it could do 20 reuses, within the capabilities of some expendable engines, it would be $2,000 per kilo which would beat even the F9 used reusable price.

 In point of fact though the first four SLS vehicles would all use original Space Shuttle engines. Then likely each has dozens of uses left in their operational lifetimes:

Apr 6, 2021
RS-25 Rocket Engines Return to Launch NASA’s Artemis Moon Missions.


   Robert Clark


Could the SpaceX SuperHeavy be a SSTO?

 Copyright 2022 Robert Clark

 Running some  numbers for the SuperHeavy+Starship launcher, I was surprised to get that an expendable SuperHeavy alone could be SSTO with quite high payload.  Wikipedia gives the propellant mass of the SuperHeavy as 3,400 tons, but does not give the dry mass. We can do an estimate of that based on information Elon provided in a tweet:

 This is for a stripped down Starship, no reusability systems, no passenger quarters, and reduced number of engines.  But this could not lift-off from ground because of the reduced thrust with only 3 engines plus being vacuum optimized these could not operate at sea level. So up the number of engines to 9 using sea level Raptors. According to wiki the Raptors have a mass of 1,500 kg. So adding 6 more brings the dry mass to 49 tons, call it 50 tons, for a mass ratio of 25 to 1.

By the way, there have been many estimates of the capabilities of the Starship for a use other than that with the many passengers, say 50 to 100 , to LEO or as colonists to Mars, for instance, such as the tanker use or only as the lander vehicle transporting a capsule for astronauts for lunar missions.  But surprisingly they all use the ca. 100 ton dry mass of the passenger Starship. But without this large passenger compartment it should be a much smaller dry mass used in the calculations. For instance, the Dragon 2 crew capsule dry mass without the trunk is in the range of 7 to 8 tons for up to 7 astronauts. So imagine a scaled up passenger compartment for 50 passengers or more. That passenger compartment itself could well mass over 60 tons.

 So the dry mass estimate of a stripped down, expendable, reduced engine Starship of 40 tons offered by Elon does make sense. 

Based on this, an expendable Starship with sufficient engines for ground launch could be SSTO:

the ISP of the Raptors for both sea level and vacuum-optimized versions have been given various numbers. I’ll use 358 s as the vacuum ISP of the sea level Raptor. For calculating payload using the rocket equation, the vacuum Isp is commonly used even for the ground stage, since the diminution in Isp at sea level can be regarded as a loss just like air drag and gravity loss for which you compensate by adding additional amount to required delta-v to orbit just like the other losses.

 Then 3580ln(1 +1200/(50 + 50)) = 9,180 m/s sufficient for LEO.      

 But as of now, SpaceX has no plans of making the Starship a ground-launched vehicle. So we’ll look instead at the SuperHeavy. For an expendable version with no reusability systems, we’ll estimate the dry mass using a mass ratio of 25 to 1, same as for a ground-launched expendable Starship. Actually, likely the Superheavy mass ratio will be even better than this since it is known scaling a rocket up improves the mass ratio. So this gives a dry mass of 136 tons. Then the expendable SuperHeavy could get 150 tons to LEO as an expendable SSTO: 3580ln(1 + 3,400/(136 + 150)) = 9,150 m/s, sufficient for LEO.

 But what about a reusable version? Reusability systems added to a stage should add less than 10% to the dry mass:


From: (Henry Spencer)


Subject: Re: The cost (in weight) for Reusable SSTO

Date: Sun, 28 Mar 1999 22:37:10 GMT

In article <kemJ2.876$>,

Larry Gales <> wrote:

>An SSTO with a useful payload using Kero/LOX is easy to do -- provided that

>it is *expendable*.  All of the difficulty lies in making it reusable...

There are people who are sufficiently anti-SSTO that they will dispute the

feasibility of even expendable SSTOs (apparently not having read the specs

for the Titan II first stage carefully).

>   (1) De-orbit fuel: I understand that it takes about 100 m/s to de-orbit.

That's roughly right.  Of course, in favorable circumstances you could play

tricks like using a tether to simultaneously boost a payload higher and

de-orbit your vehicle.  (As NASA's Ivan Bekey pointed out, this is one case

where the extra dry mass of a reusable vehicle is an *advantage*, because

the heavier the vehicle, the greater the boost given to the payload.)

>   (2) TPS (heat shield): the figures I hear for this are around 15% of the

>orbital mass

Could be... but one should be very suspicious of this sort of parametric

estimate.  It's often possible to beat such numbers, often by quite a large

margin, by being clever and exploiting favorable conditions.  Any single

number for TPS in particular has a *lot* of assumptions in it.

>   (4) Landing gear: about 3%

Gary Hudson pointed out a couple of years ago that, while 3% is common

wisdom, the B-58 landing gear was 1.5%... and that was a very tall and

mechanically complex gear designed in the 1950s.  See comment above

about cleverness.

I would be very suspicious of any parametric number for landing gear which

doesn't at least distinguish between vertical and horizontal landing.

>   (5) Additional structure to meet loads from differnet directions (e.g.,


>        takeoff, semi-horizontal re-enttry, horizontal landing).  This is


>        guesswork on my part, but I assume about 8%

Of course, here the assumptions are up front:  you're assuming a flight

profile that many of us would say is simply inferior -- overly complex,

difficult to test incrementally, and hard on the structure.

>I would appreciate it if anyone could supply more accurate figures.

More accurate figures either have to be for a specific vehicle design,

or are so hedged about with assumptions that they are nearly meaningless.


The good old days                   |  Henry Spencer

weren't.                            |      (aka

 The 15% mentioned for thermal protecton(TPS) is for Apollo-era heat shields. But the PICA-X developed by SpaceX is 50% lighter so call it 7.5% for TPS.  And for the landing gear ca. 3%, but with carbon composites say half of that at 1.5%.

  But this would put the reusable payload at ca. 136 tons which is in the range of 100 to 150 tons of the full two stage reusable vehicle!

 How is that possible? A reusable multistage vehicle has a severe disadvantage. The fuel that needs to be kept on reserve for the first stage to slow down and boost back to the launch site subtracts greatly from the payload possible.  But for a reusable SSTO it can remain in orbit until the Earth rotates below until the landing site is once again below the vehicle.

   Robert Clark

Saturday, July 24, 2021

A Plea to Add EARLY Treatment to the Arsenal to Combat COVID-19.

 Copyright 2021 Robert Clark

Did WHO wipe from its web site a positive ivermectin report for safety?

 It was discussed on some online forums that WHO on it’s web site lauded ivermectin for its safety and effectiveness for treating parasitic disease:

Mass treatment with ivermectin: an underutilized public health strategy

 However, when I went to the link on the WHO site recently, that page is no longer there. Note the link I gave above is the one from the site, a site that archives previous web pages.

 The original link on the WHO site was:

 If you go to that link now though you’ll get an error message that WHO has been revamping its site since 2020 and some pages have been moved. However, I used the search function on the WHO site and that page no longer turns up.

 So WHO acknowledges IVM’s safety for wide spread use. Then for a global epidemic it makes no logical sense to not allow its use under a doctors care, at least under an emergency use authorization.

 Here’s a stunning fact when you think about it: 

For COVID, IF your case is going to require hospitalization it will happen on average within 7 days of symptoms appearing.

Then IF a proposed medication really is successful for EARLY treatment of COVID-19, we will have evidence of that within days, indicated by the reduction of hospitalizations.

 To me that is a stunning fact, if any of these proposed medications really is effective for EARLY treatment, then if they had been used wide-spread for EARLY treatment in the U.S. or other countries we would have known within days that they had indeed been successful. We would not have had to suffer through the pandemic for 18 months. It’s extent could have been radically cut within just days.

 This shows why the single-minded focus of relying on RCT’s is severely misguided in the midst of a pandemic. RCT’s take 3 months or more to complete. That’s too long during a pandemic undergoing exponential growth.

 WHO and other public health agencies are not considering this fact in their risk-benefit analysis, especially for repurposed drugs with well-known side-effect profiles: the risk is small but the benefits are profoundly important and could be confirmed within just days.

 Here’s another way of seeing this: suppose we just now instituted a national treatment guideline that EVERYONE be giving IVM or some other proposed drug for EARLY treatment on first signs of symptoms. And suppose then within days we saw a drop in hospitalizations by 80%, i.e., by 1/5th, across the board, for every city and state, and for every race and socio-economic group we saw this drop.

Then proponents of RCT’s would still say it wasn’t real because it wasn’t conducted under the guise of an RCT.

 Note this is not just idle speculation. What we have seen in India may be an indication of such a rapid drop in cases and fatalities when IVM is put in wide spread use. 


 Note now the Indian variant that was spreading rapidly then IS the delta variant predominant in the UK, Israel and building now in the U.S.

 There was skepticism expressed by some that the effect in India could not be due to IVM because the drop occurred so rapidly after IVM being put in wide-spread use.

 But that’s the ENTIRE point of the matter: since IF for a particular case COVID is to progress to the hospitalization stage, it will happen within days. Then if a medication can greatly cut this risk by EARLY treatment, its effectiveness will also be seen within just days.

 Note too that such medications in their antiviral role if they really are effective as EARLY treatments quite likely will also be effective for prophylaxis. And that has indeed been seen to be the case for IVM. See for example this interview with a doctor actually treating patients with IVM in India:

Choosing and using ivermectin for covid-19 in India.

 So imagine such a drug also being used wide-scale in a country for prophylaxis, and it cutting transmission also by, say, 80%, i.e., by 1/5th.

 Then both these effects together prophylaxis and treatment will cut hospitalizations and fatalities by a factor of 1/25. This is now getting into the range of what vaccines can do, cutting severe cases and fatalities by double digits.

 Since the effect can be seen so rapidly, to test it this doesn’t have to be put in effect across an entire country. Just put it in effect in some city of some size, for both treatment and prophylaxis.

 I said WHO had earlier acknowledged ivermectins safety for mass treatment. But part of the reason it won’t grant its use for COVID treatment is the claim that not enough evidence supports its effectiveness. I have made this challenge to people I’ve corresponded with to perform their own judgement on this issue. The PubMed site has a listing of peer-reviewed published papers. Searching on there turns up literally dozens of published reports positive on ivermectins effectiveness:

 As of now there are 210 reports listed under this search. That’s quite a lot of papers to review. So I’ve challenged anyone to pick any 10 at random. Here’s a random number generator page: Set the max range of numbers to pick from at 210. Use the generator 10 times to come up with 10 numbers at random from 1 to 210 and select those papers to review. Count now how many papers are positive, negative, and neutral. You’ll find the reports are overwhelmingly positive for ivermectins usefulness.

 It is simply unreasonable to suppose all those reports are finding positive effects just by coincidence.

  Robert Clark

Monday, June 21, 2021

SSME based SSTO’s. UPDATED, 6/28/2021 - Extension to the Delta IV Heavy

 Copyright 2021 Robert Clark

 The SSME makes possible *practical* SSTO’s, that is, with significant payloads. Need to use known lightweight materials, such as carbon fiber. As discussed in the blog post,

DARPA's Spaceplane: an X-33 version, Page 2 

some ultra high strength metals would also work:

 The standard aluminum the Delta IV used for its propellant tanks had about a 25 to 1 propellant to tank mass ratio. At 200 ton propellant mass, this gives a tank mass of 8 tons. So using light weight materials we can get this down to 4 tons.

 To estimate payload possible we need the required delta-v to low Earth orbit(LEO). Various estimates are given for this depending on altitude and orbital inclination. I’ll use the estimate give in this report of 9,000 m/s:

Towards Reusable Launchers - A Widening Perspective.

 The Delta IV launch vehicle used the low cost, RS-68, an engine designed for low cost, but only at mid level performance. But for an SSTO you need both high performance engines and weight optimized structures. So we’ll replace the RS-68 with two SSME’s. The two SSME’s will have about the same mass as the RS-68 with slightly more thrust. But most importantly they have a much better vacuum Isp at 452.3 s compared to 412 s.

 The dry mass of the Delta IV first stage is about 26 tons. By light weighting the tank, we cut 4 tons from the tank mass to bring the dry mass to 22 tons. Then we can get a payload of 8 tons to orbit. By the rocket equation:

452.3*9.81ln(1 + 200/(22 + 8.2)) = 9,012 m/s.

 The Delta-IV TSTO, with no side boosters, has about the same payload to LEO using a Centaur upper stage:

The higher Isp of the SSME’s explains why the SSTO can match the performance of the TSTO to LEO.

 It is true that most launches now are for satellites actually to go to GEO, geosynchronous orbit. For this you can use the current Centaur upper stage. But firstly the use of the Centaur would allow higher payload to LEO than the current TSTO. The Centaur has about a 21 ton propellant load and 2 ton dry mass for a 23 ton gross mass with an Isp of 462 s. Then the TSTO could now get 21 tons to LEO:

452.3*9.81ln(1 + 200/(22 + 23 + 21.5)) + 462*9.81ln(1 + 21/(2 + 21.5)) = 9,050 m/s. 

 This is well above the 8 tons or so of the current TSTO version and is close to the payload of the Falcon 9 full thrust version.

 This TSTO does get more than the SSTO, but in point of fact you don’t need the maximal payload for most launches, so why use the upper stage if it is unneeded? 

 The additional delta-v required to GEO is about 3,800 m/s. Then a total of 12,800 m/s would be required to get the payload to GEO. This TSTO version could get 6 tons to GEO:

452.3*9.81ln(1 + 200/(22 + 23 + 6)) + 462*9.81ln(1 + 21/(2 + 6)) = 12,900 m/s. 

 This is multiple times higher than the current version:

However, the SSME’s are expensive engines. To get the full usefulness of an SSTO you really need reusability. In point of fact by using lightweight thermal protection and landing legs, and lightweight wings or propellant storage for final approach, the additional mass for reusability can be less than 10% of the landed mass. So the SSTO would still get significant mass to LEO.

 Reusables do get their most usefulness at high launch rates. Then to serve various markets we would also want a smaller vehicle than this SSTO. We could get this by cutting down the tank size and using a single SSME: 

 We’ll cut down the propellant size to 150 tons, 3/4ths the usual size, and again we’ll use light-weighting on the tank of the Delta IV first stage. Subtracting off the 8 tons for the standard tank the dry mass is 18 tons. The tank mass at 3/4ths size would be 6 tons. But we’re lightweighting it so it’ll be 3 tons. That brings the dry mass to 21 tons.

 But the SSME at 3.2 tons is also a lighter engine than the RS-68, at 6.6 tons. So we’ll subtract off an additional 3.4 tons from the dry mass to bring it to 17.6 tons. There will be additional reductions in the dry mass due to smaller thrust structure for the lower thrust of the single SSME, and smaller insulation and wiring for the smaller vehicle, but not as large as the other reductions. 

 So for the first estimate take the dry mass as 17.6 tons. Then we can get 5 tons to LEO with this smaller SSTO:

452.3*9.81ln(1 + 150/(17.6 + 5)) = 9,020 m/s.

 Since less than 10% would be needed for reusability we can still get ca. 3 tons to LEO as a reusable SSTO.


   Robert Clark

UPDATED, 6/28/2021.

 The advantage of using high performance engines and lightweight tanks is not just for getting a SSTO. As mentioned above using this, the two-stage-to-orbit(TSTO) Delta IV can double its LEO payload to ca. 20 tons to match that of the Falcon 9 Full Thrust. Its payload capacity to the lucrative GEO satellite market would also match that of the Falcon 9 FT.

 ULA could then be competitive with SpaceX for the largest current market for commercial launches. Note that ULA could also match SpaceX in it's partially reusable approach of returning only the first stage. ULA head Tory Bruno has said the SpaceX approach of boosting back the first stage to the launch site requires too much fuel, and loses too much payload.

 Instead what ULA has proposed doing is returning the engine package only, catching it in midair on return by parachute. The rest of the first stage would be thrown away. This is a kludge, an inelegant solution just to get by. A kludge never provides a long lasting solution. The Space Shuttle was a kludge, which explains why it never reached its goal of fast and low cost reusability. However, by increasing the Delta IV's payload capacity in the described manner, ULA can also boost back to the launch site and match SpaceX in reusable payload capacity.

 The startling increase in payload would also apply to the Delta IV Heavy, with the addition of cross-feed fueling. This is where for a parallel staged vehicle the fuel for the central core is taking from the side boosters during the parallel burn portion of the flight. This allows the central core to be fully fueled when the side boosters are jettisoned.

55th International Astronautical Congress 2004 - Vancouver, Canada 

1 IAC-04-V.4.03 


 We'll use the rocket equation to again estimate the payload possible for the Delta IV Heavy. Because of how cross-feed fueling works, the calculation works as if the it were a three stage vehicle with the fuel only being from the side-boosters during the "first stage":

452.3*9.81ln(1 + 400/(44 + 220 + 23 + 68)) + 452.3*9.81ln(1 + 200/(22 + 23 + 68)) +462*9.81ln(1 + 21/(2 + 68)) = 9,060 m/s.

 This payload of 68 tons nearly triples the payload of the current version of the Delta IV Heavy, and now matches the payload of the Falcon Heavy.

 This is still using the expensive SSME's however. But the point of the matter is this is doable for the current engine RS-68 used on the Delta IV as well if given altitude compensating nozzles.

 There are various means of adding altitude compensating nozzles to existing engines. They could be like the RL-10B with extendible nozzles, or flexible, expandable nozzles, or mechanically moving "petal" arrangements as with variable area nozzles on jet engines, or several other ways. 

 This last is interesting in that this method of achieving variable area nozzles has been in existence since the 70's. As a total WAG it might improve the performance for jets by, say, 10%. But for rockets using variable nozzles could improve performance by 100% or more and it still has not been used for rockets. 

 In such a way, a midlevel performance engine such as the RS-68 could be upgraded to achieve comparable performance of the high performance SSME's. I won't offer a calculation here using the rocket equation for this case. The reason is while such estimates frequently take the vacuum Isp for a fixed nozzle engine, this may or not be accurate for the case of an altitude compensating nozzle. On the one hand you can give the RS-68 a higher vacuum Isp than SSME, even at 470+ s, but the higher chamber pressure of the SSME, suggests it should have better performance at ground level than the RS-68 given alt.comp.  

 The calculation of the delta-v possible through altitude compensation is one that should be made by the launch companies.

Regular Manned Lunar Flights.

 The use of the higher performance engines gives us the capability of SSTO's with the Delta IV and also heavy lift with the Delta IV Heavy, with payload comparable to the first planned version of the SLS at ca. 70 tons. With this we have the possibility of routine manned flights to the Moon and beyond. Robert Zubrin has written a proposal for a low cost architecture for setting up a Moon base with regular flights to the Moon. Remarkable all it would need beyond the current capability is a reusable lunar lander.

 In the Zubrin proposal it would take only 3 flights of the Falcon Heavy and one flight of a manned Falcon 9 to set up the manned lunar base:

Op-ed | Moon Direct: How to build a moonbase in four years.

by Robert Zubrin — March 30, 2018

 But with the extra capabilities of the Delta IV Heavy, the three Falcon Heavy launches could be replaced by three launches of the Delta IV Heavy.

 Recall I said Zubrins plan only needed a lunar lander. ULA is planning the upper stage of its upcoming Vulcan lander, the Centaur V, to have in-space reusability. Then this upper stage could also be used as a reusable lunar lander. See discussion here:

Robust Lunar Exploration Using an Efficient Lunar

Lander Derived from Existing Upper Stages

AIAA 2009-6566

Bernard F. Kutter et al.

 In regards to adapting the Centaur V or a currently in use Centaur to a horizontal landing lunar lander, I advise the landing rockets be just pressure-fed thrusters fueled from the regular tanks of the hydrolox tanks of the Centaur.