Low Cost HLV, page 2: Comparison to the S-IC Stage.

Copyright 2012 Robert Clark 

The dry mass of the first stage of the vehicle described in the Low Cost HLV post was 1/20th the gross mass of the stage at 110,000 kg. Interestingly the propellant mass of the S-IC first stage of the Saturn V was about the same as in this HLV proposal, about 2,100,000 kg. Then it will be interesting to make a comparison to the dry mass of the S-IC stage. This page gives it as 130,000 kg:

Ground Ignition Weights.
http://history.nasa.gov/SP-4029/Apollo_18-19_Ground_Ignition_Weights.htm

 However, the Saturn V had a quite heavy first stage thrust structure:

SP-4206 Stages to Saturn.
7. The Lower Stages: S-IC and S-II.

Rosen apparently took the lead in pressing for the fifth engine, consistent with his obstinate push for a "big rocket." The MSFC contingent during the meetings included William Mrazek, Hans Maus, and James Bramlet. Rosen argued long and hard with Mrazek, until Mrazek bought the idea, carried the argument to his colleagues, and together they ultimately swayed von Braun. Adding the extra power plant really did not call for extensive design changes; this was Rosen's most convincing argument. Marshall engineers had drawn up the first stage to mount the original four engines at the ends of two heavy crossbeams at the base of the rocket. The innate conservatism of the von Braun design team was fortunate here, because the crossbeams were much heavier than required. Their inherent strength meant no real problems in mounting the fifth powerplant at the junction of the crossbeams, and the Saturn thus gained the added thrust to handle the increasingly heavy payloads of the later Apollo missions. "Conservative design," Rosen declared, "saved Apollo."²

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

Indeed it was heaviest single component of the S-IC stage, and so of the Saturn V:

S-IC.
2. Components.
http://en.wikipedia.org/wiki/S-IC#Components

 This online lecture of Dr. David Akin of the University of Maryland gives mass estimating relationships for various rocket components, taken from the reports NASA uses in designing rockets:


Mass Estimating Relations.
• Review of iterative design approach
• Mass Estimating Relations (MERs)
• Sample vehicle design analysis
http://spacecraft.ssl.umd.edu/academics/483F09/483F09L13.mass_est/483F09L13.MER.pdf

 On page 17 is given a relation between the thrust of the stage in Newtons and the mass of the thrust structure in kilograms:

M_{Thrust structure}(kg) = 2.55×10⁻⁴T(N)

 For 4 RD-171's at 7,900 kN each, this would be 8,000 kg. So we can subtract off 13,000 kg from that S-IC dry mass to get 117,000 kg. We're also using one less engine, so subtract off 8,350 kg for the one less engine. However, the RD-171 weighs about 1,000 kg more than the F-1, so add on 4,000 kg to get about 113,000 kg for the stage, quite close to the 20 to 1 mass ratio estimate. Note this is even without the weight saving alloys and composites now used.




  Bob Clark

Low Cost HLV.

Copyright 2012 Robert Clark

Credit: Modified from:
NASA's Space Launch System - Winners and Losers
by Ed Kyle, 06/17/2011
http://www.spacelaunchreport.com/sls4.html

 The announcement by two separate teams backed by highly regarded scientists and entrepreneurs for asteroidal or lunar mining means that quite likely there will be a significant market for super heavy lift. Note too that there were separate shuttle privatization plans with business models that involved privately investing perhaps $2 billion to produce a "shuttle 2.0". Quite key here though is a vehicle this size could serve as a super HLV without the ca. 80 mT shuttle orbiter. 

  I think at this point it is abundantly clear NASA can not be expected to make a cost effective launcher. An internal NASA estimate put the total development and launch cost of just four of the interim 70 mT SLS vehicles as $41 billion, which amounts to over $10 billion per launch.

 SpaceX has shown by using good cost-saving business practices to be able to produce a launcher at greatly reduced costs. They estimate their upcoming Falcon Heavy will break the $1,000 per pound barrier, or $2,000 per kg. Keep in mind then that increasing the size of your launcher is supposed to reduce your per kg costs. So likewise using good business practices, a super heavy lift launcher privately developed should be able to at least match this or exceed it. This would be in the range of $200 million per launch for a ca. 100 mT launcher, a radically reduced cost over that of the SLS. 

  I think consideration should also be given to an all liquid system. You would use the DIRECT team's Jupiter HLV hydrogen-fueled upper stage but instead of using the shuttle ET for the first stage to hold hydrolox, use the same sized tank for kerolox. This would give a super heavy lift vehicle without the SRB's.

 The DIRECT team wanted to use the same size tank to save on costs since you can use the same existing tooling in this case. However, a key fact is this will still be the case even if you switch to kerolox propellant. You would have to change the location of the divider between the fuel and oxidizer of course, but this is comparatively low cost compared to producing whole new tooling for a different size tank.

Now kerolox is a denser propellant so you are going to get a higher propellant load in this case. The density is about 3 times that of hydrolox, so lets say the propellant load of the first stage is now 2,100 mT. What about the dry mass? 

 At this point I think we should take note of the lightweight characteristics of the Falcon 9 that SpaceX was able to achieve. SpaceX has said the first stage of the Falcon 9 has a mass ratio better than 20 to 1. SpaceX did this by using well known techniques such as a common bulkhead design for the tanks. So we could follow this also to minimize first stage dry mass.

 Also, note that by scaling our propellant tanks up, we actually improve our mass ratio. So likely we can get an even better mass ratio than this for our large first stage. But using the 20 to 1 figure, or 19 to 1 for propellant to dry mass ratio, we get a dry mass of 110 mT.

 We need heavy thrust kerosene engines. I'll use the RD-171, with a sea level thrust of about 1,700,000 lbs, and vacuum Isp of 337 s. This will require 4 of the engines. This could be replaced later with the F-1 but using the RD-171 would allow you to start now on the vehicle development with better performance.

 For the specifications of the upper stage, the DIRECT team went through several versions of their Jupiter super heavy lift vehicle. I'll use the one they referred to as Jupiter-246 Heavy, LV 41.5004.08001. For whatever reason, the DIRECT team no longer has the specifications for their vehicles up on their web site. This version's specifications are within this post to the SpaceFellowship.com forum:


Re: An SSTO as "God and Robert Heinlein intended".
Posted on: Sat Mar 12, 2011 9:49 pm
http://spacefellowship.com/Forum/viewtopic.php?p=44979&sid=3d11bfaff22840cdcd239a4c452c48d6#p44979

though you'll have to register on that forum to view it.

 This version used a propellant mass of 190,849 kg, a dry mass of 11,825 kg and 6 of the RL-10B2 engines, with an Isp of 459 s. Other versions have used the new J-2X engine, but just 6 RL-10's are likely to be cheaper. 

 This version's interstage and payload fairing were at about 4,000 kg each. We'll round off the upper stage propellant mass to 190 mT and dry mass to 11 mT. Then using a 9,150 m/s delta-V to orbit we can estimate the payload to orbit as 145 mT:

337*9.81ln(1+2100/(110+201+4+4+145)) + 459*9.81ln(1+190/(11+4+145))=9,176 m/s

 Admittedly, this payload estimate seems high so I plugged some numbers into John Schilling's "Launch Vehicle Performance Calculator" and got:

Mission Performance:Launch Vehicle: User-Defined Launch Vehicle
Launch Site: Cape Canaveral / KSC
Destination Orbit: 185 x 185 km, 28 deg
Estimated Payload: 102573 kg
95% Confidence Interval: 86317 - 121993 kg

 So it's still, likely, ca. 100 mT.

 There are several variations on this theme. For example to save on development costs we could use the Ariane 5 core stage as the upper stage. Since the ESA was amenable to using it for an upper stage for a re-booted Ares I, i.e., ATK's "Liberty" rocket, they would likely be amenable to this as well. You could also make this be parallel staging with cross feed fueling to improve performance.

 Another possibility would be to make the upper stage also be kerosene-fueled. Say you used the same light-weight tooling and tank diameter for the hydrogen fueled upper stage but using now kerolox propellant. Again, you could improve performance by making it parallel staged with cross-feed fueling. But this has an additional advantage in that you could take one of the engines off the first stage to use it for the upper stage. This would result in a lower dry weight for the first stage. The upper stage though would then be somewhat overpowered using a RD-171, so it may suffice instead to use a RD-180 just for the upper stage.

 



   Bob Clark