Here are some possibilities for lower cost super heavy lift launchers,
in the 100,000+ kg payload range. As described in this article the
proposals for the heavy lift launchers using kerosene-fueled lower
stages are focusing on using diameters for the tanks of that of the
large size Delta IV, at 5.1 meters wide or the even larger shuttle ET,
at 8.4 meters wide:
All-Liquid: A Super Heavy Lift Alternative?
by Ed Kyle, Updated 11/29/2009
http://www.spacelaunchreport.com/liquidhllv.html
The reason for this is that it is cheaper to create new tanks of the
same diameter as already produced ones by using the same tooling as
those previous ones. This is true even if switching from hydrogen to
kerosene in the new tanks.
However, I will argue that you can get super heavy lift launchers
without using the expensive upper stages of the other proposals by
using the very high mass ratios proven possible by SpaceX with the
Falcon 9 lower stage, at above 20 to 1:
SPACEX ACHIEVES ORBITAL BULLSEYE WITH INAUGURAL FLIGHT OF FALCON 9
ROCKET.
Cape Canaveral, Florida – June 7, 2010
"The Merlin engine is one of only two orbit class rocket engines
developed in the United States in the last decade (SpaceX’s Kestrel is
the other), and is the highest efficiency American hydrocarbon engine
ever built.
"The Falcon 9 first stage, with a fully fueled to dry weight ratio of
over 20, has the world's best structural efficiency, despite being
designed to higher human rated factors of safety."
http://www.spacex.com/press.php?page=20100607
We will use tanks of the same size as these other proposals but will
use parallel, "bimese", staging with cross-feed fueling. This method
uses two copies of lower stages mated together in parallel with the
fueling for all the engines coming sequentially from only a single
stage, and with that stage being jettisoned when it's expended its
fuel. See the linked image below for how parallel staging with cross-
feed fueling works.
Do the calculation first for the large 8.4 meter wide tank version. At
the bottom of Kyle's "All-Liquid: A Super Heavy Lift Alternative?"
article is given the estimated mass values for the gross mass and
propellant mass of the 8.4 meter wide core first stage. The gross mass
of this single stage is given as 1,423 metric tons and the propellant
mass as 1,323 metric tons, so the empty mass of the stage would be
approx. 100 metric tons (a proportionally small amount is also taken
up by the residual propellant at the end of the flight.) Then the mass
ratio is 14 to 1. However, the much smaller Falcon 9 first stage has
already demonstrated a mass ratio of over 20 to 1.
A key fact about scaling is that you can increase your payload to
orbit more than the proportional amount indicated by scaling the
rocket up. Said another way, by scaling your rocket larger your mass
ratio in fact gets better. The reason is the volume and mass of your
propellant increases by cube of the increase and key weight components
such as the engines and tanks do also, but some components such as
fairings, avionics, wiring, etc. increase at a much smaller rate. That
savings in dry weight translates to a better mass ratio, and so a
payload even better than the proportional increase in mass.
This is the reason for example that proponents of the "big dumb
booster" concept say you reduce your costs to orbit just by making
very large rockets. It's also the reason that for all three of the
reusable launch vehicle (RLV's) proposals that had been made to NASA
in the 90's, for each them their half-scale demonstrators could only
be suborbital.
Then we would get an even better mass ratio for this "super Evolved
Atlas" core than the 20 to 1 of the Falcon 9 first stage, if we used
the weight saving methods of the Falcon 9 first stage, which used
aluminum-lithium tanks with common bulkhead design. It would also work
to get a comparable high mass ratio if instead the balloon tanks of
the earlier Atlas versions prior to the Atlas V were used.
So I'll use the mass ratio 20 to 1 to get a dry mass of 71.15 mT, call
it 70,000 kg, though we should be able to do better than this. We'll
calculate the case where we use the standard performance parameters of
the RD-180 first, i.e., without altitude compensation methods. I'll
use the average Isp of 329 s given in the Kyle article for the first
leg of the trip, and 338 s for the standard vacuum Isp of the RD-180.
For the required delta-V I'll use the 8,900 m/s often given for
kerosene fueled vehicles when you take into account the reduction of
the gravity drag using dense propellants. Estimate the payload as 115
mT. Then the delta-V for the first leg is 329*9.8ln(1 + 1,323/(2*70 +
1*1,323 + 115)) = 1,960 m/s. For the second leg the delta-V is
338*9.8ln(1 + 1,323/(70 + 115)) = 6,950 m/s. So the total delta-V is
8,910 m/s, sufficient for LEO with the 115 mT payload, by the 8,900 m/
s value I'm taking here as required for a dense propellant vehicle.
Now let's estimate it assuming we can use altitude compensation
methods. We'll use performance numbers given in this report:
Alternate Propellants for SSTO Launchers.
Dr. Bruce Dunn
Adapted from a Presentation at:
Space Access 96
Phoenix Arizona
April 25 - 27, 1996
http://www.dunnspace.com/alternate_ssto_propellants.htm
In table 2 is given the estimated average Isp for a high performance
kerolox engine with altitude compensation as 338.3 s. We'll take the
vacuum Isp as that reached by high performance vacuum optimized
kerolox engines as 360 s. Estimate payload as 145,000 kg. For the
first leg, the delta-V is 338.3*9.8ln(1 + 1,323/(2*70 + 1*1,323 +
145)) = 1,990 m/s. For the second leg the delta-V is 360*9.8ln(1 +
1,323/(70 + 145)) = 6,940 m/s, for a total delta-V of 8,930 m/s,
sufficient for orbit with the 145,000 kg payload.
Now we'll estimate the payload using the higher energy fuel
methylacetylene. The average Isp is given as 352 s in Dunn's report.
The theoretical vacuum Isp is given as 391 s. High performance engines
can get quite close to the theoretical value, at 97% and above. So
I'll take the vacuum Isp as 380 s. Estimate the payload as 175,000 kg.
Then the delta-V over the first leg is 352*9.8ln(1 + 1,323/(2*70 +
1*1,323 + 175)) = 2,040 s. For the second leg the delta-V will be
380*9.8ln(1 + 1,323/(70 + 175)) = 6,910 s, for a total delta-V of
8,950 m/s, sufficient for orbit with the 175,000 kg payload.
bimese Falcon 9 launcher
http://i27.tinypic.com/2yxn2oz.jpg
Bob Clark