Some proposals for low cost heavy lift launchers.

I showed in this post:

Newsgroups: sci.space.policy, sci.astro, sci.physics,
sci.space.history
From: Robert Clark
Date: Tue, 4 May 2010 10:49:50 -0700 (PDT)
Subject: A kerosene-fueled X-33 as a single stage to orbit
vehicle.
http://groups.google.com/group/sci.s...aaf61151?hl=en

that two reconfigured X-33's mated bimese fashion and using a cross-
feed fueling system could reduce the costs to orbit by *two orders* of
magnitude. This shows there really is no logical objection to
developing an SSTO. Because even if it is argued multistaged systems
can carry more payload, you can carry *even* more payload by making
those stages be separately SSTO capable. *Multiple times* more.
I want to emphasize again the only reason why I used the Lockheed
version of the X-33 was because it was already largely built. The
other two proposed versions of a suborbital X-33 demonstrator by
Rockwell and McDonnell-Douglas would also become fully orbital when
switched from hydrogen to kerosene-fueled at comparable costs.
These would be easier to make because you wouldn't have the problem
that led to the
X-33's downfall of lightweighting the tanks. Then the only thing
keeping us from $100/lbs. launch costs is the acceptance that SSTO is
indeed possible.
That is why it is so imperative that the Falcon 1 first stage derived
SSTO I discussed before be done because it would be so easy and CHEAP
to achieve:

Newsgroups: sci.space.policy, sci.astro, sci.physics,
sci.space.history
From: Robert Clark
Date: Sun, 14 Mar 2010 18:24:37 -0700 (PDT)
Subject: A kerosene-fueled X-33 as a single stage to orbit
vehicle.
http://groups.google.com/group/sci.s...833c4470?hl=en

Then finally the light bulb would come on.

However, the bimese X-33 would involve some technical risk in that it
would require the building of a second hydrocarbon-fueled X-33 and the
low payload cost, due to the high payload capacity, would only obtain
if the untested tank lightweighting methods really did bring the
tankage ratio of the conformal tanks to be more in line with that of
cylindrical tanks.
Therefore I'll show here that an (expendable) heavy lift system can be
produced with a payload capacity in the range of 40,000 kg to 60,000
kg at a minimal cost compared to the other heavy lift systems being
proposed, and while using already existing components and at minimal
technical risk.
Previously I had argued that both the Falcon 1 and Falcon 9 first
stages had a 20 to 1 mass ratio, and that this was important because
this was the mass ratio often cited for a kerosene-fueled rocket to
have SSTO capability. But that was based on the data on the
SpaceLaunchReport.com site.
The numbers on this site though are estimates and can be inaccurate.
For instance from numbers actually released by SpaceX, the Falcon 1
first stage mass ratio is actually about 16.8 to 1.
However, I was surprised to see in this recent news release from
SpaceX that the Falcon 9 first stage mass ratio is actually better
than 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

Undoubtedly it is able to achieve this high mass ratio because it also
uses common bulkhead design for the propellant tanks as does Falcon 1.
Note that the original Atlas and the Saturn V upper stages nearly had
SSTO mass ratios because they used common bulkheads.
From this news release, we can also estimate the dry mass of the first
stage:

UPDATES: JULY 2009 - DECEMBER 2009.
DRAGON/FALCON 9 UPDATE.
Wednesday, September 23rd, 2009
"Weighing in at over 7,700 kg (17,000 lbs), the thrust assembly and
nine
Merlin engines represents over half the dry mass of the Falcon 9 first
stage."
http://www.spacex.com/updates_archive.php?page=2009_2

So I'll estimate the dry mass of the first stage as 15,000 kg, and the
first stage total mass as 300,000 kg, and so the propellant mass as
285,000 kg.
I'll again use three NK-33's as the engines, replacing the nine
Merlin's. Using 660 kg as an estimate of the Merlin 1C mass, and 1,222
as the NK-33 mass, the dry mass becomes 15,000 - 9*660 + 3*1,222 =
12,726 kg.
Again let's calculate what payload we can get using two of these
Falcon 9's mated bimese fashion using cross-feed propellant transfer.
This time I'll use a little more conservative average Isp of 335 s for
the first portion of the trip where they are still mated together, but
still assume some altitude compensation method is being used such as
an aerospike. Then I'll still take the vacuum Isp as 360 s.
Let's estimate the payload as 40,000 kg. Then we get a delta-V of:

335*9.8ln(1+285,000/(2*12,726+285,000+40,000)) = 1,954 m/s, for the
first
mated-together portion of the flight, and then:
360*9.8ln(1+285,000/(12,726+40,000)) = 6,552 m/s, for the upper stage
portion, giving a total of about 8,500 m/s.

Note again that by using more energetic hydrocarbon fuels, perhaps
also densified by subcooling, you can get perhaps 50% higher payload
to orbit than the 40,000 kg, so to perhaps 60,000 kg.
This certainly qualifies as heavy lift if not super heavy lift. And
could satisfy the requirements of a lunar mission at least for the
launch system by using two launches.


Bob Clark