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

The Coming SSTO's: multi-Vulcain Ariane.

Copyright 2013 Robert Clark

 The option the ESA decided on for the planned Ariane 6 was the version using a solid propellant first stage:

CNES, ASI Favor Solid-Rocket Design For Ariane 6.

By Amy Svitak

Source: Aviation Week & Space Technology

October 15, 2010

https://web.archive.org/web/20130419155317/http://www.aviationweek.com/Article.aspx?id=/article-xml/AW_10_15_2012_p26-505016.xml&p=1

 However,  one of the other options discussed for the Ariane 6 would also allow manned European flight capability. This would be the two Vulcain core version.

CNES is evaluating these three launch vehicle concepts for a next-generation Ariane 6: two based on solid-rocket-motor technology plus an all-liquid-fueled launcher with optional solid-motor boosters. (Credit: CNES)

 To estimate the payload capability for the twin Vulcain core I'll use John Schillings launch performance calculator:

Launch Vehicle Performance Calculator.
http://www.silverbirdastronautics.com/LVperform.html

 In the calculations for this multi-Vulcain Ariane core stage, I used this page for the specifications on the Ariane:

Space Launch Report:  Ariane 5 Data Sheet.
http://www.spacelaunchreport.com/ariane5.html#config

 For the Vulcain 2 specifications, I've seen different numbers in different sources, though close to each other. I'll use this source:

Vulcain 2.
http://www.astronautix.com/engines/vulcain2.htm

 I'll also use the earlier Ariane 5 "G" version that is lighter than the current "E" version to be lofted by two Vulcains without side boosters. According to the SpaceLaunchReport page it had a 170 mT gross mass for the core at a 158 mT propellant load, giving a 12 mT dry mass.
 According to the Astronautix page, Vulcain 2 has a 434 s vacuum Isp and 1350 kN vacuum thrust. So two will have a 2700 kN vacuum thrust. The Vulcain's mass is listed as 1,800 kg. So adding another will bring the stage dry mass to 13,800 kg.
Now input this data into Schilling's calculator. Select again default residuals and select "No" for the "Restartable Upper Stage?" option. Select the Kourou launch site for this Ariane 5 core rocket. For the orbital inclination, I input 5.2 degrees. I gather Schilling uses this for Kourou's latitude since deviating from this decreases the payload. I chose also direct ascent for the trajectory.

Then the result I got was 7,456 kg(!) to orbit:

================================
Mission Performance:
Launch Vehicle:     User-Defined Launch Vehicle
Launch Site:     Guiana Space Center (Kourou)
Destination Orbit:      185 x 185 km, 5 deg
Estimated Payload:      7456 kg
95% Confidence Interval:      4528 - 10898 kg
================================

 We should be able to remove a component on the Ariane 5 core to lighten the weight for this application. Ed Kyle on his Spacelaunchreport.com page discusses the Liberty rocket that had been planned to use a SRB first stage and an Ariane 5 core second stage. For the Liberty application, a forward skirt on the core called the JAVE ("Jupe AVant Equipée") that transmits the forces of the two solid boosters to the core would be removed. This will also be removed for our application without solid boosters.
 The JAVE massed 1,700 kg. So our payload could be increased to 9,156 kg. However, Kyle also discusses on his page on the Liberty rocket that the increased thrust from the SRB first stage would require thicker walls on the Ariane core now used as an upper stage.
 The thicker walls on the Ariane 5 core for the Liberty rocket are indicated in this video:

 The 5-segment SRB to be used on the Liberty rocket has 12 times the thrust of the Vulcain engine. Yet as seen in the video the increased thickness to handle the increased axial load is only 50%. Then only doubling the thrust by adding a second Vulcain quite likely will require a much smaller increase in thickness. I'm informed that the 158 mT propellant mass tank has a dry weight in the range of 4,400 kg. So even increasing the thickness 50% increases the weight by ca. 2,200 kg, and the payload would still be approx. 7,000 kg.  A problem with this estimate though, aside from the unknown accuracy of the video, is that it is based on the larger Evolution "E" version of the Ariane 5 core, which might not require as much strengthening to handle the higher thrust loads as the smaller "G" version.  So we'll use a formula for calculating the thickness of a propellant tank based on the axial load as given on this lecture page: Launch Vehicle Design: Configurations and Structures. Space System Design, MAE 342, Princeton University Robert Stengel http://www.princeton.edu/~stengel/MAE342Lecture4.pdf  on page 9:

  From the first formula the critical buckling load without the pressurization effect is:   σ_{c,w/o pressure} = [9(t/R)^1.6 + 0.16(t/L)^1.3]*E.   Multiplying out the second formula for critical buckling with the pressurization effect you see it's:   σ_{c,w/ pressure} = σ_{c,w/o pressure} + 0.191p(R/t).   Now use the formula on p. 8 that relates the tensile strength of the material to the thickness required of a pressurized tank:

   _{You see that}  σ_hoop = p(R/t)  so that the formula above becomes:     σ_{c,w/ pressure} = σ_{c,w/o pressure} + 0.191σ_{hoop .   Now use values for the tensile strength of aluminum alloy. The aluminum alloy used on the Ariane 5 core tanks, Al 2219, happens to get stronger at cryogenic temperatures:}

 Table taken from Properties of Aluminum Alloys: Tensile, Creep, and Fatigue Data at High and Low Temperatures, page 86. The table gives the aluminum alloy strength at liquid hydrogen temperatures as 685 MPa and elasticity modulus, E, as 85 GPa.  For the Ariane 5 core "G" version, the hydrogen tank walls are only 1.3 mm thick, while the oxygen's, 4.7 mm. The diameter of the tanks is 5.4 m. Because of its extreme wall thinness it's the _{hydrogen tank whose stress has to be limited. It's length is about 18 m. Then the formula for the critical buckling load without pressurization gives:} σ_{c,w/o pressure} = [9(t/R)^1.6 + 0.16(t/L)^1.3]*E = [9(0.0013/2.7)^1.6 + 0.16(0.0013/18)^1.3]*85*10⁹ = 3,800,000 Pa.  And the additional buckling strength due to pressurization is 0.191σ_hoop = 0.191*685,000,000 = 130,800,000 Pa, for a total critical buckling load of 134,600,000 Pa.  The maximum thrust of two Vulcain 2's will be 2,700,000 N. The cross-sectional area of the hydrogen tank walls is 2*π*R*t = 2(3.14)(2.7)(0.0013) = 0.022 m² . Then the maximum axial pressure is 2,700,000/0.022 = 123,000,000 Pa.  This is indeed less than the critical buckling load of 134.6 MPa. However, for a manned launcher a safety factor of 1.4 is usually included. This will require the maximum axial pressure to be less than 96 MPa. This requires a wall thickness of 1.6 mm, about a 25% increase. This still only increases the tank weight by 1,000 kg, so the payload becomes now ca. 8,000 kg, still quite high for a SSTO. Remember also switching to aluminum-lithium alloy can save as much as 25% off the dry weight which would bring us again to the 9,000 kg payload range.


   Bob Clark