SLS for Return to the Moon by the 50th Anniversary of Apollo 11, page 4: further on lightweighting the SLS core.

                                             Copyright 2013 Robert Clark

 NASA has decided to revert to the original Al 2219 aluminum alloy that was first used on the shuttle external tank for the SLS core:

SLS takes on new buckling standards, drops Super Light alloy.
February 18, 2013 by Martin Payne 
http://www.nasaspaceflight.com/2013/02/sls-new-buckling-standards-drops-super-light-alloy/

 This is due to the greater brittleness of the lighter aluminum-lithium alloys used on the later super lightweight ET tank (SLWT). And because the later alloys were not available in the greater thickness needed for optimal lightweight performance. 
 However, NASA itself estimated the Al-li alloys could save 25% off the weight of a propellant tank over the Al 2219 alloy:

RELEASE : 09-096
NASA Uses Twin Processes to Develop New Tank Dome Technology
http://www.nasa.gov/centers/langley/news/releases/2009/09-096.htm

 Still NASA estimated in regards to the SLS tank, reverting back to the Al 2219 alloy would only cost 3,000 kg in lost payload, much smaller than 25%. Apparently, the reduced thickness of the plates available for the aluminum-lithium alloys used on the SLWT results in reduced weight efficiency. 
 However, a new aluminum-lithium alloy Al-Li 2050 has similar strength at lightweight to the SLWT alloys and is available in thicker plate sizes:

Shell Buckling Knockdown Factor (SBKF) Project Update.
http://www.nasa.gov/offices/nesc/home/Feature_ShellBuckling_Test.html

 Then we could recover the ca. 25% saving over using the Al 2219 alloy. This now is a significant increase in payload, beyond just 3,000 kg. The original ET tank using Al 2219 alloy weighed 35,000 kg. The new SLS tank is scaled up 33%, so under the same Al 2219 alloy would weigh in the range of 46,000 kg. Then the new Al-Li alloy saving 25% off this would be a saving of 11,500 kg. 
 NASA made an assessment of cost benefit analysis and decided on the older Al 2219 alloy. But this is Apollo era, 1960's, technology. This is going backwards not forwards in our technological development. 
 Further weight saving can be achieved by using composites for the intertank. NASA with Boeing is investigating large cryogenic composite tanks. This is still a research project. However the intertank is an unpressurized structure. Structures like this have been made of composites for decades. 
 To estimate the weight that can be saved, note the intertank in the al-li SLWT weighed 5,500 kg:

Single Stage To Orbit Mass Budgets Derived From Propellant Density and Specific Impulse.
John C. Whitehead
32nd AIAA/ASME/SAE/ASEE Joint Propulsion ConferenceLake Buena Vista, FL July 1-3, 1996

http://www.osti.gov/bridge/servlets/purl/379977-2LwFyZ/webviewable/379977.pdf

 Then the intertank of the SLS of 33% larger size may be estimated to weigh 7,300 kg. A new composite material known as an isotruss saves significantly on weight:

 It weighs less than 1/7th that of aluminum at the same strength. This would reduce the intertank mass to less than 1,000 kg. This would subtract off an additional 6,000 kg from the tank mass to bring it down to 28,500 kg. This is nearly 18,000 kg in total off from the original SLS tank weight, which could go to extra payload.
 As I mentioned in the blog post SLS for Return to the Moon by the 50th Anniversary of Apollo 11, page 3: lightweighting the SLS core, internal NASA estimates put the actual payload of the SLS as significantly above the 70 mT mark often cited by NASA. Then an additional 18,000 kg added to this payload capability would put the SLS payload to LEO at ca. 100 mT. This is important because it would mean the SLS would have the capability to do manned lunar lander missions, not just lunar flybys.
 NASA administrator Charles Bolden has said NASA, meaning the administrators, has no plans on a Moon mission, being more focused on a mission to an asteroid. However, the public in general, space advocates, industry, and even NASA's own ranks have shown no interest in the asteroid mission:

Back to the Moon? Not any time soon, says Bolden.

By Jeff Foust on 2013 April 5 at 1:05 pm ET

A week from Monday marks the third anniversary of President Obama’s speech at the Kennedy Space Center where he formally announced the goal of a human mission to an asteroid by 2025. While that is an official goal of NASA’s human space exploration program, there remains some opposition or, at the very least, lack of acceptance of the goal by many people, including some with NASA, as a report on NASA’s strategic direction concluded last December.
At a joint meeting of the Space Studies Board and the Aeronautics and Space Engineering Board in Washington on Thursday, the head of that study, Al Carnesale of UCLA, reiterated those concerns. “Since it was announced, there was less enthusiasm for it among the community broadly,” he said of the asteroid mission goal. “The more we learn about it, the more we hear about it, people seem less enthusiastic about it.”
Carnesale suggested that, in his opinion, it might be better to shelve the asteroid mission goal in favor of a human return to the Moon. “There’s a great deal of enthusiasm, almost everywhere, for the Moon,” he said. “I think there might be, if no one has to swallow their pride and swallow their words, and you can change the asteroid mission a little bit… it might be possible to move towards something that might be more of a consensus.”

http://www.spacepolitics.com/2013/04/05/back-to-the-moon-not-any-time-soon-says-bolden/

 The SLS even by its first mission in 2017 can do manned lunar landing missions by incorporating well known and relatively low cost weight saving methods to its core and upper stages.

 This would go a long way towards garnering support both among the public and those in  industry to know that a return to the Moon is in the offing and in the very near term.

  Bob Clark


Update, Sept. 28, 2013:

 Finally, NASA has acknowledged that the Block 1, first version of the SLS to launch in 2017 will have a 90+ mT payload capacity not the 70 mT always stated by NASA:

SLS Dual Use Upper Stage (DUUS).
http://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/20130013953_2013013757.pdf

 This is important since it means we will have the capability to do manned lunar landing missions by the 2017 first launch of the SLS:

SLS for Return to the Moon by the 50th Anniversary of Apollo 11, page 5: A 90+ metric ton first launch of the SLS.
http://exoscientist.blogspot.com/2013/09/sls-for-return-to-moon-by-50th.html

Budget Moon flights.

Copyright 2013 Robert Clark 

In the blog post SpaceX Dragon spacecraft for low cost trips to the Moon, I argued that manned flights to the Moon could be mounted for costs in the few hundred million dollars range. This is compared to the $100 billion cost estimated for the now defunct Constellation program.
 The key to the low cost was using the SpaceX Falcon Heavy booster, and using already existing components, such as the Centaur upper stages. But I wondered could we bring the cost down even further using even smaller upper stages? Perhaps to even a few 10's of millions of dollars??
 I'll use again a combination of hydrogen-fueled stages, two copies of the Ariane 4 third stage, version H-10-III this time. There are variations in the cited specifications for this stage in various sources:

Ariane 4.
http://www.b14643.de/Spacerockets_1/West_Europe/Ariane/Design/Ariane_2.htm 

Die Oberstufen H-8, H-10 und ESC-A.
http://www.bernd-leitenberger.de/h-10.shtml 

Ariane-44L H10-3.
http://space.skyrocket.de/doc_lau_det/ariane-44l_h10-3.htm 

Ariane H10-3.
http://www.astronautix.com/stages/arieh103.htm

ARIANE 4 SPECIFICATIONS.
http://www.braeunig.us/space/specs/ariane.htm 

 It may be some sources are including the weight of the Vehicle Equipment Bay (VEB), others not. The VEB carried the avionics and telemetry equipment for the Ariane 4. We may suppose these functions carried out by the crew capsule, and at lighter weight than that used on the Ariane 4, first launched in the mid-90's.
 I'll use the numbers on Braeunig's "ARIANE 4 SPECIFICATIONS" page. It gives the dry mass of the stage as 1,240 kg and the propellant mass as 11,860 kg. The HM-7B engine used on that stage has an Isp of ca. 445 s.

 We'll use again this table of Earth/Moon delta-V's:

Delta-V budget.
Earth–Moon space.

https://en.m.wikipedia.org/wiki/Delta-v_budget#Earth%E2%80%93Moon_space%E2%80%94high_thrust

 With aerobraking on the return to the Earth, the total round-trip delta-V is 8,650 m/s.  We'll  use the architecture that the landing stage is used to return all the way back to Earth, not just to link up with a stage waiting in lunar orbit. And just a single crew capsule will be used that carries the crew all the way from Earth to the Moon and back again, no separate command and lunar modules, as with Apollo. This is analogous to the Early Lunar Access proposal of the early 90's.
 Then we could bring a payload of 2.4 metric tons(mT) to the Moon and back, sufficient for a half-Dragon sized capsule:

445*9.81ln(1 + 11,860/(1,240 + 13,100 + 2,400)) + 445*9.81ln(1 + 11,860/(1,240 + 2,400)) = 8,660 m/s.

 Various weight saving techniques can be used to save further weight on the stages.  The propellant tanks are made of aluminum rather than the heavier steel of the Centaurs. But they can be made lighter in the rang of 15% to 25% by using the aluminum-lithium alloy used on the later versions of the shuttle external tank (ET). Further weight saving techniques would be to use the common bulkhead and "balloon tank", i.e., pressure-stabilized, design of the Centaurs. 
 The tanks can be made even lighter by using composites, perhaps by 30% over aluminum-lithium:

The Composite Cryotank Technologies and Demonstration Project.

http://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/20120002943_2012002501.pdf

 This means composites can save as much as 50% (!) off the weight of standard aluminum. Aside from tanks, composites can be used on the structural members to save weight. Composites could be used on the thrust structure, the intertank, the payload and interstage adapters, and the tank structural support members.
 A new composite structural technique is that of the isotruss:

Isotruss

 According to the manufacturer, it weighs only 1/12th as much as steel at the same strength:

isotruss strength comparison

 Boeing has already developed a 2.4 meter wide composite cryogenic tank: 

Boeing Develops Game Changing Composite Propellant Tank

 This would be sufficient for the H10-III stage. By next year they expect to have a 5.4 meter wide tank. This would allow a 20 mT standard sized Centaur to have the composite tanks. This size tank would also work for a scaled-up 40 mT Centaur-style stage. ULA has argued a stage this size following various weight saving techniques could get a 20 to 1 mass ratio. This would allow a stage that could make a manned round-trip lunar flight from LEO in a single stage.  

 ULA has discussed improving mass ratio as the best way to improve performance at the NewSpace 2012 conference. 

 From the graphs shown it is apparent propellant mass fraction, or mass ratio, is far and away the best way of increasing performance.  
Then to maximize the payload we can deliver to planetary targets, including on manned missions, it is important to implement such weight saving methods. For instance a single in-space stage for a lunar mission could be made reusable, thus cutting the cost of the in-space stage. 
 Conceivably just as important is that high mass ratio would also allow a single stage to orbit vehicle, SSTO. This has relevance to another ULA proposal of establishing propellant depots. It is known that a single stage capable of reaching LEO could also with orbital refueling make the round-trip from LEO to the lunar surface and back again as a single stage. So this one stage could make the entire trip from Earth to the Moon, with that one stop for refueling.
 The DC-X test VTVL test vehicle will be having its 20th anniversary this year. The event is to be celebrated by the participants in the program in August this year. It will also be marked by restoration of the DC-X for display in the New Mexico Space Museum.
 The DC-X itself was not intended to reach orbit. But interestingly the H-10-III stage, of similar size to the DC-X, may have this capability when lightweighted with composites. The single HM-7B engine it has though does not provide sufficient thrust for liftoff. You may need to add two to three additional engines. The increased thrust  however would add additional stress to the structure. It may require additional strengthening mass.   

COST. 
 For the lunar flight scenario, the total weight for the two H-10-III stages and the 2.4 mT capsule would be 28.4 mT. Using the estimated $2,000 per kilo cost for the Falcon Heavy launcher this would be a launch cost of $56.8 million. There is the cost also of the H-10-III stages. According to the "Ariane H10-3" Astronautix page, it's listed as $12 million. So the total cost of the launch would be $80.8 million.
 There is also however the cost of the capsule. We may suppose though the capsule is reusable so its cost per use might be only in the few million dollars range. Actually at a weight penalty of a few hundred pounds of propellant kept in reserve, which would subtract a proportionally smaller amount from the payload, the first H-100-III stage might be returnable to Earth to also be reusable. The second stage, used as the lander, could already be reusable since under the Early Lunar Access architecture it would also serve as the propulsive stage to return the capsule all the way to Earth. This would reduce the per use cost for the upper stages as well.
You would not be limited to using the Falcon Heavy, though this would be the lowest cost. The Delta IV Heavy with upgraded RS-68a engine has a 28 mT payload capacity to LEO, slightly less than the 28.4 mT needed at around a $300 million launch cost. However, the Delta IV Heavy is not expected to be man-rated anyway so you might as well launch the capsule on another rocket and link up with the propulsive stages in orbit.

ESA VERSION.
 This may be an architecture that could be implemented by the ESA. The Ariane 5 ME  is to have a 20% increase in payload to GTO to 12 mT. If the increase in payload to LEO is also 20% then that would bring the payload capacity to 24 mT. This would be enough to carry the propellant of the two H-10-III stages. Neither the current Ariane 5 nor the ME version will be man-rated however. We will need a separate man-rated rocket to carry the crew to orbit. This would have to lift the two H-10-III dry masses plus the capsule mass for 2*1.24 mT + 2.4 mT = 4.88 mT. I'll show in an accompanying blog post this is well within the capability of a vehicle made from a just a single stage of the Ariane 5 core with a second Vulcain 2 engine added. 
 The cost of the Ariane 5 is about $200 million. If the cost is also increased by 20% on the ME this would bring it to $240 million. And the cost for the manned launcher? I'm informed by a member of ESA that the estimated cost for a modified Ariane 5 core with a second Vulcain 2 engine would be ca. $60 milion. Plus the $24 million cost of the two upper stages brings the cost to $324 million.



  Bob Clark

SLS for Return to the Moon by the 50th Anniversary of Apollo 11, page 3: lightweighting the SLS core.

Copyright 2013 Robert Clark

 In the blog post SLS for Return to the Moon by the 50th Anniversary of Apollo 11, page 2: Orion + SEV design, I made the argument that the SLS Block I, scheduled to launch in 2017, should have significantly more payload than the 70 mT cited by NASA, which is no more than that of the smaller, less powerful Block 0 SLS. Recent news reports also indicated that a Freedom Of Information Act (FOIA) request to NASA on the SLS specifications was denied:

NASA MSFC Says That SLS Performance Specs Fall Under ITAR.
http://spaceref.com/news/viewnews.html?id=1697

Report: NASA in Huntsville won't release performance specifications for new rocket.
By Lee Roop | ****@al.com
on January 25, 2013 at 3:23 PM, updated January 25, 2013 at 3:51 PM
blog.al.com/breaking/2013/01/report_nasa_in_huntsville_wont.html

 Rand Simberg had suggested to me that the reason why NASA did not want to release the actual capabilities of the Block I SLS is that it would negate the need for even proceeding with the expensive Block II SLS. Thus it was an attempt, he argued, to maintain the "pork" of the expensive Block II development.

 However, I have been informed by those in the know that the Block I will indeed likely have a greater payload capacity than the 70 mT of the Block 0 version. However, a problem with providing such specifications for a new rocket is there is always weight growth beyond that which was originally expected.

 I had argued that scaling up a rocket should result in increased payload. But an additional factor to consider is that the new SLS core will not be scaled up in all dimensions. It is to be kept at the same width of the shuttle external tank (ET) while its length is stretched 33%. The same diameter is maintained to use the same tooling as that used to build the ET. However, stretching the length while maintaining the same diameter means additional strengthening members have to be attached to maintain its strength against bending and buckling loads. So it's not just a straight-forward matter of scaling up the mass of the core stage to estimate the payload capacity of the Block I. So I fully believe as the SLS core stage comes closer to completion then more accurate values for the payload capacity will be released.

 I was interested to note that, according to what I have been informed, that the current internal NASA estimates of the SLS payload to LEO would still allow my Orion+SEV lunar landing proposal, though not with as much leeway. Then to the end of increasing the payload and increasing the mass growth margins, I have some suggestions. Though having the propellant tanks being composite might be a bridge too far for a 2017 launch some of the structural strengthening members in the tanks might be. I was struck by this image while researching composite structures:

Isotruss

 It shows a new carbon composite structure called an isotruss. It just looks like it would be lightweight for the strength, doesn't it?

 SpaceX also uses composites to reduce weight in the interstage between the first and second stage of the Falcon 9. They could be used for the interstage for the SLS as well. An even more weight saving application of composites might be in the intertank though. This is the component of the external tank that supports the oxygen tank above the hydrogen tank. It's actually a heavy component of the ET, weighing more even than the oxygen tank.
  
 The tank mass of the ET and other rocket stages is discussed in the report:

Single Stage To Orbit Mass Budgets Derived From Propellant Density and Specific Impulse.
John C. Whitehead
32nd AIAA/ASME/SAE/ASEE Joint Propulsion ConferenceLake Buena Vista, FLJuly 1-3, 1996
http://www.osti.gov/bridge/servlets/purl/379977-2LwFyZ/webviewable/379977.pdf

 I consider this by the way to be the best aerospace engineering paper never published because of the importance of its conclusions. If the validity of its arguments had been recognized then, then we would already have by now routine manned space access.

 On page 8 in Fig. 6 is shown a diagram giving the mass of the intertank in the ET. It's mass is listed as 5.5 tons:

 For the tank stretched 33% in the SLS this might be 7.3 tons. Estimates put the weight savings in the structural mass in the 40% range by using composites. 

 A greater weight saving strategy would be to eliminate the intertank entirely. This is known as common bulkhead design. It also eliminates the weight of one of the bulkheads. This was used very successfully on the Saturn V cryogenic upper stages. Currently it is used on the Centaur upper stage, the Falcon 9 first stage, and the Ariane 5 core. The plan now is to use separated tanks with an interstage to maintain commonality with the current ET design. Still it might be advantageous to do a trade study on the increased payload in comparison to the increased cost of the common bulkhead design.

 Another possibility to save weight might be inflatable payload fairings. For such a large rocket, intended to carry such large payloads, the payload fairings would be quite heavy. There was a NASA RFI for innovative proposals on fairings and adapters, which has already expired. However, perhaps NASA can do a trade study of the weight saving possible under this method. Bigelow has been in the news for his inflatable habitats so this would not be so unusual for aerospace applications.

 Also, I want to argue again as I did in my Orion+SEV post for funding the ULA suggestions for lightweighting the Centaurs, to the extent of getting a 20 to 1(!) mass ratio. With a propellant size of 40 metric tons(mT) this would allow a round trip lunar landing mission with a single in-space stage. In fact such a lunar lander could be reusable, and it would be so small it could even be launched with a 70 mT sized launcher.

 It needs to be mentioned that many knowledgeable industry insiders do not believe the final Block II version of the SLS will ever fly. This is because of the long time frame, 20 years from now so over several presidential administrations, and because of its high cost. The SLS Block I scheduled for 2017 on the other hand very likely will.

 Wouldn't it be great for the public to learn as we more closely approach the completion date for this SLS in 2017 that it already will have the capacity for lunar landing flights and moreover we already have, by then, the in-space stages to accomplish it?

 NASA considering the very legitimate possibility that the Block II will not be funded should have as a contingency some plans of accomplishing the BEO explorations with the Block I SLS. This can be done with no new technology and not really very much extra cost, but just by using known methods of lightweighting the core and in-space stages.


    Bob Clark

Update, Sept. 28, 2013:

 Finally, NASA has acknowledged that the Block 1, first version of the SLS to launch in 2017 will have a 90+ mT payload capacity not the 70 mT always stated by NASA:

SLS Dual Use Upper Stage (DUUS).
http://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/20130013953_2013013757.pdf

 This is important since it means we will have the capability to do manned lunar landing missions by the 2017 first launch of the SLS:

SLS for Return to the Moon by the 50th Anniversary of Apollo 11, page 5: A 90+ metric ton first launch of the SLS.
http://exoscientist.blogspot.com/2013/09/sls-for-return-to-moon-by-50th.html

Low Cost HLV, page 3: Lightweighting the S-IC Stage.

Copyright 2012 Robert Clark

                                                    SASSTO 
                                                    SASSTO - Saturn-derived SSTO Launch Vehicle 
                                                    Credit: © Mark Wade


 I showed in the post Low Cost HLV, page 2: Comparison to the S-IC Stage that the S-IC first stage of the Saturn V could give a nearly 20-to-1 mass ratio using a lighter thrust structure and using four RD-171 engines instead of five F-1 engines. But in fact we can do better than this. The S-IC of the 1960's did not have available the aluminum-lithium alloy used for example on the Falcon 9 and shuttle ET. Here I calculate a lighter structure using this lighter alloy and some mass reducing structural changes.

 The tank mass of the S-IC stage and of some other rocket stages is discussed in this key report:

Single Stage To Orbit Mass Budgets Derived From Propellant Density and Specific Impulse.
John C. Whitehead
32nd AIAA/ASME/SAE/ASEE Joint Propulsion ConferenceLake Buena Vista, FLJuly 1-3, 1996
http://www.osti.gov/bridge/servlets/purl/379977-2LwFyZ/webviewable/379977.pdf

 On page 8 in Fig. 6 is given a comparison of the tank weights of the Saturn S-IC, Atlas II, and shuttle ET:

 We've already reduced the mass of the stage down to 113,000 kg by using a lighter thrust structure in the prior post. Fins are not really needed for large rockets with computer guidance and control, so remove these to reduce the mass again to 112,000 kg. The  four 4 meter diameter RD-171 engines can fit under the 10 meter diameter S-IC tank, so remove the engine fairings to bring the mass down to 108,000 kg.
 Now use common bulkhead design to reduce the tank mass further. The question of using common bulkhead design for the S-IC arose during the Apollo design period:

SP-4206 Stages to Saturn.
7. The Lower Stages: S-IC and S-II.
THE S-IC AND THE HUNTSVILLE CONNECTION.

The main configuration of the S-IC had already been established by MSFC, including the decision to use RP-1, as opposed to the LH2 fuel used in the upper stages. Although LH2 promised greater power, some quick figuring indicated that it would not work for the first stage booster.Liquid hydrogen was only one half as dense as kerosene. This density ratio indicated that, for the necessary propellant, an LH2 tank design would require a far larger tank volume than required for RP-1. The size would create unacceptable penalties in tank weight and aerodynamic design. So, RP-1 became the fuel. In addition, because both the fuel and oxidant were relatively dense, engineers chose a separate, rather than integral, container configuration with a common bulkhead. The leading issue prior to the contract awards related to the number of engines the first stage would mount.

http://history.nasa.gov/SP-4206/ch7.htm#197

   This could be interpreted to mean the density of the propellant made it unfeasible, but I think the relatively smaller tankage mass using the dense propellants made the more difficult common bulkhead design unnecessary. For instance as you see in that Fig. 6 from Whitehead's report, in the shuttle ET tank the intertank weighs more than the entire oxygen tank. The relative weight of the intertank is not as bad for the kerosene S-IC. Still, common bulkhead design is used for the large kerosene first stage on the Falcon 9 to help save weight.

 So to minimize stage weight we will remove the interstage and one of the bulkheads. Assuming top and bottom bulkheads weigh the same for each of the LOX and kerosene tanks on the S-IC, then from the information in Fig. 6, the LOX bulkheads weigh 4 mT each and the kerosene, 3.3 mT. Conservatively, let's say we remove one of the kerosene bulkheads instead of a LOX bulkhead since we may need the larger LOX bulkhead for strength. Then also removing the 6 mT intertank, we bring the dry mass down to 99 mT.

 Now estimate the weight saving using the lighter aluminum-lithium alloy. From the Wikipedia page on the shuttle ET, the tank weight reduced from 35,000 kg using aluminum alloy 2219, the same alloy used for the S-IC tanks, to 26,500 kg using aluminum-lithium alloy, a reduction of 24%.

 After the structural changes, the tanks now weigh 25.5 mT. Subtracting off 24% from this is a reduction in mass by 6 mT. This brings the stage mass down to 93 mT.

 Keep in mind though, the plan is to use a shuttle ET size tank to save cost on tooling. The ca. 720 mT hydrolox of the shuttle ET becomes ca. 2,100 mT with the 3 times denser kerolox. This turns out to be about the same kerolox carried by the S-IC. So the purpose here was just to get an idea of a lightweight stage you can get using modern materials.

 Now notice you get significant payload as a SSTO using the RD-171 engines at 338 s vacuum Isp. Taking the required delta-v to orbit as 9,150 m/s for kerolox, you can get 48 mT to orbit:

338*9.81ln(1 + 2,100/(93 + 48)) = 9,170 m/s.

 Note though that if we are to use the shuttle ET as a stage then the pointed end of the LOX tank would need to be removed. We could take the equivalent cylindrical LOX tank of the same volume. It would have the same dry weight, so the stage dry mass stays the same.

 However, if you take the full length of a cylindrical tank now as 46.9 m and the diameter as 8.4 m, per the specifications of the SLWT version of the ET, and the density of kerolox as about 1,030 kg/m^3, then we get about 2,600 mT kerolox. The tank weight would increase somewhat without the pointed end, but not by much compared to the entire stage weight. Then you could loft 82 mT to orbit:

338*9.81ln(1 + 2,600/(93 + 82)) = 9,160 m/s.

 A propellant load of 2,600 mT at dry mass of 93 mT corresponds to a mass ratio close to 29 to 1, rather high. But SpaceX has said with their side boosters on the Falcon Heavy they expect to achieve a mass ratio of 30 to 1, and mass ratio does get better as you scale up a stage,with this shuttle ET size stage being much larger.

 This payload of 82 mT is better than the 70 mT to be carried by the interim SLS. Remember our HLV is to be developed using the SpaceX-style commercial approach. Then based on a $2,000/kg price of the Falcon Heavy, the full two stage version of our HLV as comparably priced might only cost ca. $200 million per launch at a 100 mT payload to orbit.

 So the SSTO version would even cost less than this, perhaps only ca. $100 million per launch for the 82 mT payload to orbit.

 The dry weight could be lightened further by using composites. Estimates put the weight savings in the structural mass in the 40% range for a fully composite structure. In that case the payload could exceed 100 mT for this SSTO.

 An increase of the Isp could be possible by using an aerospike or plug nozzle, up to the range of 360 s. The multi-nozzle format of the RD-171 engines makes this feasible. The four nozzles of each engine would be shortened and arranged around a central aerospike. This was the idea behind the aerospikes planned for the X-33 and VentureStar. It was also used earlier in the planned Beta SSTO of Dietrich Koelle and the SASSTO SSTO of Phillip Bono.

 An argument against this was that the aerospike nozzle would make the propulsion system too heavy. For instance for the aerospike on the X-33 the thrust/weight ratio was only 40 to 1, compared to a 70 to 1 ratio for the SSME's for example. However, the lightweight, high temperature ceramics and composites available now should make the T/W comparable to bell nozzle engines:

Ceramic Materials for Reusable Liquid Fueled Rocket Engine Combustion Devices.
http://ammtiac.alionscience.com/pdf/AMPQ8_1ART06.pdf


  Bob Clark

Note: The SASSTO SSTO shown at the top and discussed near the end was derived from the Saturn S-IVB stage, not the S-IC, and was hydrogen fueled. The Beta SSTO discussed was also hydrogen fueled. However, a key result of the cited report of John C. Whitehead is that it is actually easier to make a kerosene-fueled SSTO. This is because the large and heavy hydrogen fuel tanks swamps out the advantage of its higher Isp.  - B.C., 6/22/2012.