Re: Some proposals for low cost heavy lift launchers.
Robert Clark
launchers by using cross-feed parallel staging and using both weight
optimized structures and high efficiency engines at the same time.
Now all three of the companies offering liquid fueled alternatives to
NASA's shuttle-derived heavy lift vehicle, Boeing, Lockheed, and
SpaceX argue they can increase the payload for their heavy lift
vehicles by using cross-feed fueling:
DELTA IV LAUNCH VEHICLE GROWTH OPTIONS
TO SUPPORT NASA’S SPACE EXPLORATION VISION.
http://www.unitedlaunchalliance.com/site/docs/publications/DeltaIVLaunchVehicle%20GrowthOptionstoSupportNASA'sSpaceExplorationVision.pdf
Phase 2 EELV – An Old Configuration Option with
New Relevance to Future Heavy Lift Cargo.
http://www.ulalaunch.com/site/docs/publications/EELVPhase2_2010.pdf
NASA Studies Scaled-Up Falcon, Merlin.
Dec 02 , 2010
By Guy Norris, Madhu Unnikrishnan
Los Angeles, Los Angeles
http://web02.aviationweek.com/aw/mstory.do?id=news/awst/2010/11/29/AW_11_29_2010_p28-271784.xml&channel=space&headline=NASA%20Studies%20Scaled-Up%20Falcon,%20Merlin
Note that cross-feed fueling has been common on aircraft for decades.
Here's a description of it for the Concorde:
Concorde.
Balancing by Fuel-Pumping.
The Concorde Tank-Schematic:
"1 + 2 + 3 + 4 are the Collector-Tanks, feeding the engines directly.
Usually they feed there counterpart engines – but they can be cross-
switched to feed more and/or other engines at the same time.
5 + 7 and 8 + 6 are the Main-Transfer Tanks, feeding the 4 Collector-
Tanks. Initially 5 + 7 are active. If those are empty 6 + 8 take over
(or must be activated from the Engineering Panel!).
5a + 7a are Auxiliary-Tanks (to 5 and 7).
9 + 10 are the Trim-Tanks for balancing forward
11 is the Trim-Tank for balancing afterward"
http://wiki.flightgear.org/index.php/Concorde#Balancing_by_Fuel-Pumping
Now, if these companies with the heavy lift proposals would also use
weight optimized structures as well as high efficiency engines at the
same time on their individual stages, they would increase their
payload even further as well as getting SSTO-capable stages.
Bob Clark
=================================================================
Newsgroups: sci.space.policy, sci.astro, sci.physics
From: Robert Clark <rgregorycl...@yahoo.com>
Date: Mon, 2 Aug 2010 17:34:27 -0700 (PDT)
Local: Mon, Aug 2 2010 7:34 pm
Subject: Re: Some proposals for low cost heavy lift launchers.
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
===============================================================
http://groups.google.com/group/sci.space.policy/msg/206d19932302b043?hl=en
willia...@mokenergy.com
http://www.scribd.com/doc/30943696/ETDHLRLV
http://www.scribd.com/doc/31261680/Etdhlrlv-Addendum
http://www.scribd.com/doc/35439593/Solar-Power-Satellite-GEO
http://www.scribd.com/doc/45631474/Sea-Dragon-Derived-Launcher
http://www.scribd.com/doc/38432542/Multi-Element-Staging
Robert Clark
for a heavy lift launcher within its current budget limits:
NASA report favors SD HLV for SLS, complains Agency can’t afford 2016
target.
January 12th, 2011 by Chris Bergin
http://www.nasaspaceflight.com/2011/01/nasa-report-favors-sd-hlv-sls-complains-cant-afford-2016/
However, SpaceX has claimed they can deliver an HLV within 3 years
for a price of $2.5 billion:
NASA Studies Scaled-Up Falcon, Merlin.
Dec 02 , 2010
By Guy Norris, Madhu Unnikrishnan
Los Angeles, Los Angeles
"SpaceX will respond to NASA’s heavy-lift launch vehicle study with
concepts that can carry 150 tons to orbit and cost no more than $300
million per launch."
...
"Based on a roughly evenly split $10 billion budget for heavy lift,
with half for the boost stage and half for the upper stage, “we’re
confident we could get a fully operational vehicle to the pad for $2.5
billion—and not only that, I will personally guarantee it,” Musk says.
In addition, the final product would be a fully accounted cost per
flight of $300 million, he asserts. “I’ll also guarantee that,” he
adds, though he cautions this does not include a potential upper-stage
upgrade."
http://web02.aviationweek.com/aw/mstory.do?id=news/awst/2010/11/29/AW_11_29_2010_p28-271784.xml&channel=space&headline=NASA%20Studies%20Scaled-Up%20Falcon,%20Merlin
This would include the development of a 1.7 million lbs. sea level
thrust hydrocarbon engine at a $1 billion development cost for this
engine alone, the Merlin 2:
SpaceX Unveils Heavy-Lift Vehicle Plan.
Aug 5, 2010
By Guy Norris
http://www.aviationweek.com/aw/generic/story_channel.jsp?channel=space&id=news/asd/2010/08/05/07.xml&headline=SpaceX%20Unveils%20Heavy-Lift%20Vehicle%20Plan
Given SpaceX's success in delivering on their vehicle development
promises, their proposal deserves serious consideration.
However, I think NASA should consider a modification to their
proposal. SpaceX has shown they can produce a first stage booster at a
20 to 1 mass ratio in their Falcon 9 first stage. However, the larger
core stage that SpaceX is proposing for their heavy lift vehicle,
referred to as the Falcon X in their reports, won't have this high
mass ratio. On page 17 of this report the inert mass fraction, which
is mathematically equivalent to the reciprocal of the mass ratio, is
given as 0.06:
http://images.spaceref.com/news/2010/SpaceX_Propulsion.pdf
So the mass ratio for the stage would be 16.7 to 1. However, for the
larger stage the principles of scaling suggest the mass ratio should
get better for larger vehicles. Then my recommendation is for NASA to
make a formal request to SpaceX to investigate the reanalysis of the
structural design for their heavy lift core vehicle to see if it can
also be made to exceed a 20 to 1 mass ratio. This would have the
benefit of being able to lift more cargo to orbit but a potentially
more important benefit is that you would have a large core stage
capable of acting as an SSTO with large payload capability.
However, the benefits of a SSTO are most apparent when it is
reusable. Then I also suggest that NASA instead of paying SpaceX $1
billion for the development of the Merlin 2, that it should instead
restart the development of the RS-84 heavy thrust, reusable
hydrocarbon engine. As I argued here the RS-84 development could be
completed within perhaps 3 years at 2/3rds the cost of the SpaceX
Merlin 2:
Newsgroups: sci.space.policy, sci.astro, sci.physics
From: Robert Clark <rgregorycl...@yahoo.com>
Date: Fri, 18 Dec 2009 04:49:56 -0800 (PST)
Subject: Re: >>> so... the Ares-1 is DEAD >>>
http://groups.google.com/group/sci.space.policy/msg/ee460e256c24bbe6?hl=en
The RS-84 would have about 2/3rds the thrust of the Merlin 2 so you
would need more of them to do the thrust of the Merlin 2, but in being
reusable the cost would still be much less per launch than with the
expendable Merlin 2.
And what is the benefit to SpaceX under this plan? Since engines are
the lion share of the development cost for a space vehicle, SpaceX
makes the most money on their launches on the costs plus profit
percentage for their engines on their launches. SpaceX in developing a
billion dollar engine at a tens of millions of dollars cost per
engine, would make a quite a large profit in absolute terms per
launch. But the number of launches for such a vehicle would
necessarily be few.
I'm proposing instead a different business model for SpaceX. That
they forego the development of engines and instead focus on
reusability using the high efficiency (russian) engines already
available to start with, and with engines such as RS-84 designed to be
reusable with a high number of uses to be phased in soon. This would
cut launch costs by a factor at least of 10 and perhaps as much as by
a factor of 100.
With launch costs this low SpaceX would make far greater profit
simply on volume even with the greatly reduced cost per launch even
for the heavy lift launch vehicles.
Bob Clark
On Jan 1, 11:16 am, Robert Clark <rgregorycl...@yahoo.com> wrote:
> In the post copied below I suggested getting low cost heavy lift
> launchers by using cross-feed parallel staging and using both weight
> optimized structures and high efficiency engines at the same time.
> Now all three of the companies offering liquid fueled alternatives to
> NASA's shuttle-derived heavy lift vehicle, Boeing, Lockheed, and
> SpaceX argue they can increase the payload for their heavy lift
> vehicles by using cross-feed fueling:
>
> DELTA IV LAUNCH VEHICLE GROWTH OPTIONS
> TO SUPPORT NASA’S SPACE EXPLORATION VISION.http://www.unitedlaunchalliance.com/site/docs/publications/DeltaIVLau...
>
> Phase 2 EELV – An Old Configuration Option with
> New Relevance to Future Heavy Lift Cargo.http://www.ulalaunch.com/site/docs/publications/EELVPhase2_2010.pdf
>
> NASA Studies Scaled-Up Falcon, Merlin.
> Dec 02 , 2010
> By Guy Norris, Madhu Unnikrishnan
> Los Angeles, Los Angeleshttp://web02.aviationweek.com/aw/mstory.do?id=news/awst/2010/11/29/AW...
> by Ed Kyle, Updated 11/29/2009http://www.spacelaunchreport.com/liquidhllv.html
> April 25 - 27, 1996http://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 launcherhttp://i27.tinypic.com/2yxn2oz.jpg
>
> Bob Clark
> ===============================================================
http://groups.google.com/group/sci.space.policy/msg/206d19932302b043?...
Robert Clark
launcher and a manned launch system by 2016 or earlier:
NASA needs billions to get Lockheed Martin's Orion into space.
New Mexico Business Weekly
Date: Sunday, January 16, 2011, 7:03am MST
http://www.bizjournals.com/albuquerque/news/2011/01/16/NASA-lockheed-martin-orion-into-space.html
NASA estimates it will take an additional $6.6 billion to $7.1
billion to get Orion ready for first flight by 2015. On top of the
$4.9 billion already spent that is a total of $11 billion to $12
billion development cost for this capsule alone. The problem is that
Congress has only budgeted $3.92 billion over three years for a crewed
spacecraft, and maximum of $5.5 billion through 2017, still not
enough.
SpaceX has shown it can create a manned capsule for less than 1/10
the total cost of Orion even when you include the $300 million SpaceX
estimates will be required to develop an emergency abort system.
SpaceX claims even that their Dragon capsule can perform the beyond
low-Earth-orbit functions planned for the Orion spacecraft. However,
even if it can't, note that lunar missions would only take place at
best ten years from now. Then my suggestion, palatable even to the
supporters of Orion in Congress, is not to have to have Orion ready by
2016 to do the usual crew delivery functions to LEO such as to the
ISS, but rather stretch out its additional funding over ten years.
Then the cost might be as low as $660 million per year over the next
10 years developing Orion.
The usual crew to LEO functions would be left to the SpaceX Dragon or
to other companies developing manned spacecraft. For the Dragon, NASA
might only have to spend $100 million per year over 3 years to further
develop Dragon. This amount is so low NASA could also afford to
contribute to the development of the other commercial providers
developing manned spacecraft such as the DreamChaser by Sierra Nevada
and the winged lifting body design of Orbital Sciences.
This plan provides some leeway and options about which manned
spacecraft it is allowed to use. Note that some congressional critics
of SpaceX don't like relying on them for crewed launches. However,
we'll know within 3 years if their spacecraft will be reliable enough
for manned flights. During that time Orion will also continue its
development. If the final product of the Dragon, or perhaps the other
manned spacecraft contenders, is not satisfactory then we can re-ramp
up the development of the Orion, with a relatively low outlay having
been paid to the other proposed commercial spacecraft providers.
The key point is that you have significantly better leeway in your
options and choices with relatively low financial risk.
Bob Clark
On Jan 24, 3:16 pm, Robert Clark <rgregorycl...@yahoo.com> wrote:
> NASA has said it can't reach the Congress imposed deadline of 2016
> for a heavy lift launcher within its current budget limits:
>
> NASA report favors SD HLV for SLS, complains Agency can’t afford 2016
> target.
> January 12th, 2011 by Chris Berginhttp://www.nasaspaceflight.com/2011/01/nasa-report-favors-sd-hlv-sls-...
>
> However, SpaceX has claimed they can deliver an HLV within 3 years
> for a price of $2.5 billion:
>
> NASA Studies Scaled-Up Falcon, Merlin.
> Dec 02 , 2010
> By Guy Norris, Madhu Unnikrishnan
> Los Angeles, Los Angeles
> "SpaceX will respond to NASA’s heavy-lift launch vehicle study with
> concepts that can carry 150 tons to orbit and cost no more than $300
> million per launch."
> ...
> "Based on a roughly evenly split $10 billion budget for heavy lift,
> with half for the boost stage and half for the upper stage, “we’re
> confident we could get a fully operational vehicle to the pad for $2.5
> billion—and not only that, I will personally guarantee it,” Musk says.
> In addition, the final product would be a fully accounted cost per
> flight of $300 million, he asserts. “I’ll also guarantee that,” he
> adds, though he cautions this does not include a potential upper-stage
> upgrade."http://web02.aviationweek.com/aw/mstory.do?id=news/awst/2010/11/29/AW...
>
> This would include the development of a 1.7 million lbs. sea level
> thrust hydrocarbon engine at a $1 billion development cost for this
> engine alone, the Merlin 2:
>
> SpaceX Unveils Heavy-Lift Vehicle Plan.
> Aug 5, 2010
> By Guy Norrishttp://www.aviationweek.com/aw/generic/story_channel.jsp?channel=spac...
> Subject: Re: >>> so... the Ares-1 is DEAD >>>http://groups.google.com/group/sci.space.policy/msg/ee460e256c24bbe6?...
> ...
>
> read more »
Rick Jones
> http://www.bizjournals.com/albuquerque/news/2011/01/16/NASA-lockheed-martin-orion-into-space.html
> SpaceX has shown it can create a manned capsule for less than 1/10
> the total cost of Orion even when you include the $300 million SpaceX
> estimates will be required to develop an emergency abort system.
SpaceX have been doing great and wonderful things, but "shown?"
Figured, guessed, estimated, postulated even, but "shown" to me at
least means "demonstrated" and while sure, they have an unmanned
Dragon launch under their belt, I would be leery indeed of
ass-u-me-ing their estimates are spot-on. Is a SpaceX estimate better
than a LockMark estimate? Quite possibly (I don't recall SpaceX's
earliest estimates for things like when Falcon1 and Falcon9 would fly,
nor how close to their cost estimates they came). But it is still
just an estimate.
Manned Dragon may not be a "paper rocket" (a la "paper submarine") but
I'd call it closer to paper-board or perhaps corrugated cardboard than
bent metal.
rick jones
--
portable adj, code that compiles under more than one compiler
these opinions are mine, all mine; HP might not want them anyway... :)
feel free to post, OR email to rick.jones2 in hp.com but NOT BOTH...
Jeff Findley
says...
>
> In sci.space.history Robert Clark <rgrego...@yahoo.com> wrote:
>
> > http://www.bizjournals.com/albuquerque/news/2011/01/16/NASA-lockheed-martin-orion-into-space.html
>
> > SpaceX has shown it can create a manned capsule for less than 1/10
> > the total cost of Orion even when you include the $300 million SpaceX
> > estimates will be required to develop an emergency abort system.
>
> SpaceX have been doing great and wonderful things, but "shown?"
> Figured, guessed, estimated, postulated even, but "shown" to me at
> least means "demonstrated" and while sure, they have an unmanned
> Dragon launch under their belt, I would be leery indeed of
> ass-u-me-ing their estimates are spot-on. Is a SpaceX estimate better
> than a LockMark estimate? Quite possibly (I don't recall SpaceX's
> earliest estimates for things like when Falcon1 and Falcon9 would fly,
> nor how close to their cost estimates they came). But it is still
> just an estimate.
Agreed.
> Manned Dragon may not be a "paper rocket" (a la "paper submarine") but
> I'd call it closer to paper-board or perhaps corrugated cardboard than
> bent metal.
Let's be fair, they have bent metal and flown a successful unmanned
orbital test flight. That's something that NASA's Orion is still years
and billions of dollars away from doing. SpaceX has shown tremendous
progress, considering the amount of money they've spent so far on Falcon
9 and Dragon.
The Ares program has flown one "test flight" that consisted of a four
segment shuttle SRB with dummy fifth segment, dummy upper stage, and
dummy Orion capsule on top of that. That's a lot more dummies for a
supposed test flight!
Jeff
--
" Solids are a branch of fireworks, not rocketry. :-) :-) ", Henry
Spencer 1/28/2011
Rick Jones
> In article <ii7967$k7q$1...@usenet01.boi.hp.com>, rick....@hp.com
> > but I'd call it closer to paper-board or perhaps corrugated
> > cardboard than bent metal.
> Let's be fair, they have bent metal and flown a successful unmanned
> orbital test flight.
Agreed, I was too harsh. I'll agree that in the context of Manned
Dragon is more than paper-board or corrugated cardboard. For some
reason though "bent metal" feels too strong to me - perhaps I just
have a warped sense of terminology, or perhaps I don't know just how
far SpaceX have actually gotten on the LES and environmental systems.
Might there be another term then? Or is my interpretation of "bent
metal" off?
> That's something that NASA's Orion is still years and billions of
> dollars away from doing.
Very true, and while I can see it being interpreted that way, I did
not mean my reality check on Manned Dragon to be interpreted as a
defense of anything else :)
> " Solids are a branch of fireworks, not rocketry. :-) :-) ", Henry
> Spencer 1/28/2011
Gee, I thought that solids were the foundation of rocketry?-)
(foundation, beginning that sort of thing)
rick jones
--
oxymoron n, commuter in a gas-guzzling luxury SUV with an American flag
Brad Guth
> http://www.scribd.com/doc/30877060/External-Tank-Derived-Heavy-Lift-L...http://www.scribd.com/doc/30943696/ETDHLRLVhttp://www.scribd.com/doc/31261680/Etdhlrlv-Addendumhttp://www.scribd.com/doc/35439593/Solar-Power-Satellite-GEOhttp://www.scribd.com/doc/45631474/Sea-Dragon-Derived-Launcherhttp://www.scribd.com/doc/38432542/Multi-Element-Staging
Do you really think anyone here gives a tinkers damn about your ideas,
or those of anyone else?
This Usenet/newsgroup is strictly a mainstream infomercial saturated
cesspool of insiders and special-interest cabals kicking others around
for sport. Remember that 3/4 of Americans are in some way obligated
to support their government agencies no matters what their crimes
against humanity involve, and 99.9% of all others are simply too
snookered, dumbfounded and/or afraid of their own shadow. Hitler
could be in charge and there'd be no difference in how these exact
same folks (especially the well-to-do Semites) would support that
little bastard.
http://translate.google.com/#
Brad Guth, Brad_Guth, Brad.Guth, BradGuth, BG / “Guth Usenet”
Sam Wormley
Brad--Give this a try: Aerobic exercise bulks up hippocampus,
improving memory in older adults.
New research shows that at least some parts of the brain can be
saved from age-related atrophy by relatively modest amounts of
activity late in life.
Brad Guth
> On 2/1/11 2:47 PM, Brad Guth wrote:
>
>
>
> > On Jan 6, 4:07 am, william.m...@mokenergy.com wrote:
> >>http://www.scribd.com/doc/30877060/External-Tank-Derived-Heavy-Lift-L...
>
> > or those of anyone else?
>
> > This Usenet/newsgroup is strictly a mainstream infomercial saturated
> > cesspool of insiders and special-interest cabals kicking others around
> > for sport. Remember that 3/4 of Americans are in some way obligated
> > to support their government agencies no matters what their crimes
> > against humanity involve, and 99.9% of all others are simply too
> > snookered, dumbfounded and/or afraid of their own shadow. Hitler
> > could be in charge and there'd be no difference in how these exact
> > same folks (especially the well-to-do Semites) would support that
> > little bastard.
>
> > http://translate.google.com/#
> > Brad Guth, Brad_Guth, Brad.Guth, BradGuth, BG / “Guth Usenet”
>
> Brad--Give this a try: Aerobic exercise bulks up hippocampus,
> improving memory in older adults.
>
> New research shows that at least some parts of the brain can be
> saved from age-related atrophy by relatively modest amounts of
> activity late in life.
>
Good one, now pretend that you and other old farts actually give a
damn about others, and the environment.
If I were in charge, problem solved because your NASA as you know it
would be terminated, including all retirement and benefits. At least
that's what JFK had intended to do before your friends resolved that
little threat.
Robert Clark
> In sci.space.history Robert Clark <rgregorycl...@yahoo.com> wrote:
>
> >http://www.bizjournals.com/albuquerque/news/2011/01/16/NASA-lockheed-...
> > SpaceX has shown it can create a manned capsule for less than 1/10
> > the total cost of Orion even when you include the $300 million SpaceX
> > estimates will be required to develop an emergency abort system.
>
> SpaceX have been doing great and wonderful things, but "shown?"
> Figured, guessed, estimated, postulated even, but "shown" to me at
> least means "demonstrated" and while sure, they have an unmanned
> Dragon launch under their belt, I would be leery indeed of
> ass-u-me-ing their estimates are spot-on. Is a SpaceX estimate better
> than a LockMark estimate? Quite possibly (I don't recall SpaceX's
> earliest estimates for things like when Falcon1 and Falcon9 would fly,
> nor how close to their cost estimates they came). But it is still
> just an estimate.
>
> Manned Dragon may not be a "paper rocket" (a la "paper submarine") but
> I'd call it closer to paper-board or perhaps corrugated cardboard than
> bent metal.
>
> rick jones
>
The $300 million number for manrating Dragon/Falcon 9 I got from
this article from 2009:
Is It Safe?
The first company with a plan—and a rocket—to send humans to orbit
answers the existential question.
* By Michael Milstein
* Air & Space Magazine, May 01, 2009
"For SpaceX, the only upgrades required for Dragon to carry people are
the Apollo-style abort-and-escape system, seats, and a full life
support system. It will cost about $300 million to go from
transporting cargo to transporting people, most of it for the escape
system and the test flights the human-rating rules require. SpaceX has
already negotiated the finances of this step with NASA."
http://www.airspacemag.com/space-exploration/Is-It-Safe.html?c=y&page=2
However, this is the latest estimate:
SpaceX proposes rocket-powered landing system.
BY STEPHEN CLARK
SPACEFLIGHT NOW
Posted: January 18, 2011
"LOS ANGELES -- SpaceX announced Monday it submitted a proposal to
NASA last month to start an estimated $1 billion process upgrading the
company's Dragon capsule, the first step in making the ship ready for
crew rotation flights to the International Space Station."
http://www.spaceflightnow.com/news/n1101/18spacex/
My least charitable interpretation for this price increase is that
SpaceX is playing the game. The game being you can be much more cost
conscious when spending your own money than when spending the
governments money - especially when that money you are getting from
the government covers not just cost but profit.
My more charitable interpretation is that SpaceX wanted Dragon to be
reusable. However, NASA has said it would prefer not to use the
reusable version. So perhaps the higher cost is coming from the fact
that a new Dragon has to be used for every test launch.
Bob Clark
Robert Clark
options and choices with relatively low financial risk.[/B]
Another option for a manned launcher. In this report Boeing proposes
heavy lift launchers using existing components:
Heavy Lift Launch Vehicles with Existing Propulsion Systems.
Benjamin Donahue, Lee Brady, Mike Farkas, Shelley LeRoy, Neal Graham
Boeing Phantom Works,Huntsville, AL 35824
Doug Blue
Boeing Space Exploration,Huntington Beach, CA 92605
http://www.launchcomplexmodels.com/Direct/documents/AIAA-2010-2370-650.pdf
One of the proposals is of a manned launcher for the Orion capsule
using a shuttle ET propellant tank and four RS-68 engines. This does
not use an upper stage but is not a single-stage-to-orbit vehicle
because the final push to orbit is made by the onboard thrusters on
the Orion spacecraft.
However, it is interesting in this report comparison is made to the S-
IVB upper stage on the Apollo rocket. I was reminded of a suggestion
of Gary Hudson that the S-IVB would be single-stage-to-orbit with
significant payload if it used the high efficiency SSME rather than
the J-2 engine:
A Single-Stage-to-Orbit Thought Experiment.
Gary C Hudson
http://www.spacefuture.com/archive/a_single_stage_to_orbit_thought_experiment.shtml
In Hudson's proposal the vehicle could lift 10,360 lbs, 4,710 kg.
This would be just enough to carry the crewed version of the
[URL="http://en.wikipedia.org/wiki/Dragon_(spacecraft)"]Dragon
spacecraft[/URL] without cargo. Boeing's proposal for a manned capsule
the CST-100 might be launchable by this also since it is of comparable
size and design to the Dragon:
Boeing space capsule could be operational by 2015.
BY STEPHEN CLARK
SPACEFLIGHT NOW
Posted: July 21, 2010
http://www.spaceflightnow.com/news/n1007/21boeing/
NASA has shown in their crewed spacecraft versions to want to hearken
back to Apollo in their use of capsules. This SSTO idea of Hudson
would have the advantage of using a proven Apollo component that is
already manrated. The SSME's are also already manrated rather than the
RS-68 of the Boeing proposal.
Because of its small small size compared to the shuttle ET propellant
tank it would also be relatively low cost as well as only needing one
SSME engine. In fact it would even be smaller than the Falcon 9, Delta
IV, and Atlas V expendable launchers. Note as well NASA is leaning now
to using SSME's or their expendable versions rather than the RS-68 for
their shuttle derived manned launchers.
Hudson in his article stated the S-IVB was designed by the Douglas
Aircraft Company, which merged with McDonnell Aircraft to form
McDonnell Douglas. McDonnell Douglas is now a division of Boeing, so
Boeing should have access to the design documents of the S-IVB.
NASA in their shuttle-derived launcher studies have focused on
getting a cheaper version of the SSME by making an expendable version.
However, the greatest advantage of a SSTO is in being reusable. Then I
suggest studies be made on the SSME going the opposite direction: how
can it be made to be reusable at much reduced maintenance cost? Now
the SSME's have to be rebuilt after every flight costing ten's of
millions of dollars. However, Henry Spencer a highly regarded expert
on the history of space flight has said Rocketdyne studies show that
with a lot of work maintenance could be reduced to $750K per flight:
Engine reusability (Henry Spencer)
http://yarchive.net/space/rocket/engine_reusability.html
Spencer here said this would not be satisfactory for really large
reductions in space costs. But this would be a reduction in SSME
maintenance costs by 1 to 2 orders of magnitude, a major reduction in
the costs for using the engine. The question is how much would it cost
to make the necessary upgrades to the engine?
Bob Clark
On Jan 31, 4:24 pm, Robert Clark <rgregorycl...@yahoo.com> wrote:
> Further on the budget shortfalls for fielding both a heavy lift
> launcher and a manned launch system by 2016 or earlier:
>
> NASA needs billions to get Lockheed Martin's Orion into space.
> New Mexico Business Weekly
> Date: Sunday, January 16, 2011, 7:03am MSThttp://www.bizjournals.com/albuquerque/news/2011/01/16/NASA-lockheed-...
hal...@aol.com
> [B]The key point is that you have significantly better leeway in your
> options and choices with relatively low financial risk.[/B]
>
> Another option for a manned launcher. In this report Boeing proposes
> heavy lift launchers using existing components:
>
> Heavy Lift Launch Vehicles with Existing Propulsion Systems.
> Benjamin Donahue, Lee Brady, Mike Farkas, Shelley LeRoy, Neal Graham
> Boeing Phantom Works,Huntsville, AL 35824
> Doug Blue
>
> One of the proposals is of a manned launcher for the Orion capsule
> using a shuttle ET propellant tank and four RS-68 engines. This does
> not use an upper stage but is not a single-stage-to-orbit vehicle
> because the final push to orbit is made by the onboard thrusters on
> the Orion spacecraft.
> However, it is interesting in this report comparison is made to the S-
> IVB upper stage on the Apollo rocket. I was reminded of a suggestion
> of Gary Hudson that the S-IVB would be single-stage-to-orbit with
> significant payload if it used the high efficiency SSME rather than
> the J-2 engine:
>
> A Single-Stage-to-Orbit Thought Experiment.
>
> In Hudson's proposal the vehicle could lift 10,360 lbs, 4,710 kg.
> This would be just enough to carry the crewed version of the
> [URL="http://en.wikipedia.org/wiki/Dragon_(spacecraft)"]Dragon
> spacecraft[/URL] without cargo. Boeing's proposal for a manned capsule
> the CST-100 might be launchable by this also since it is of comparable
> size and design to the Dragon:
>
> Boeing space capsule could be operational by 2015.
> BY STEPHEN CLARK
> SPACEFLIGHT NOW
>
> NASA has shown in their crewed spacecraft versions to want to hearken
> back to Apollo in their use of capsules. This SSTO idea of Hudson
> would have the advantage of using a proven Apollo component that is
> already manrated. The SSME's are also already manrated rather than the
> RS-68 of the Boeing proposal.
> Because of its small small size compared to the shuttle ET propellant
> tank it would also be relatively low cost as well as only needing one
> SSME engine. In fact it would even be smaller than the Falcon 9, Delta
> IV, and Atlas V expendable launchers. Note as well NASA is leaning now
> to using SSME's or their expendable versions rather than the RS-68 for
> their shuttle derived manned launchers.
> Hudson in his article stated the S-IVB was designed by the Douglas
> Aircraft Company, which merged with McDonnell Aircraft to form
> McDonnell Douglas. McDonnell Douglas is now a division of Boeing, so
> Boeing should have access to the design documents of the S-IVB.
> NASA in their shuttle-derived launcher studies have focused on
> getting a cheaper version of the SSME by making an expendable version.
> However, the greatest advantage of a SSTO is in being reusable. Then I
> suggest studies be made on the SSME going the opposite direction: how
> can it be made to be reusable at much reduced maintenance cost? Now
> the SSME's have to be rebuilt after every flight costing ten's of
> millions of dollars. However, Henry Spencer a highly regarded expert
> on the history of space flight has said Rocketdyne studies show that
> with a lot of work maintenance could be reduced to $750K per flight:
>
>
> - Show quoted text -
Anything from the apollo era MUST be redesigned and recertified to be
man rated.
suppliers, technology, everything will be at least slightly different.
basically a new start with a vague plan to duplicate apollo
components.
nasa was dumb they should of done something like this right after the
columbia loss
Jeff Findley
b666af...@l11g2000yqb.googlegroups.com>, rgrego...@yahoo.com
says...
>
> NASA in their shuttle-derived launcher studies have focused on
> getting a cheaper version of the SSME by making an expendable version.
Because all the engineers know the flight rate would be so low that it
wouldn't be worth the development costs to try to recover and reuse the
SSME's. Recovering the SSME's isn't a trivial thing to do.
It's kind of a sad state of affairs that the agency which should be
developing next generation technologies for space travel isn't itself
investing any significant amounts of money into developing reusable
hardware.
The first, and arguably easiest, step would be a liquid fueled reusable
first stage (to be used with a two stage to orbit launcher with a
conventional upper stage). Pretty much everything about reusability is
easier to do with a first stage rather than an upper stage due to the
lower altitude and velocity at burn-out, which makes recovery a lot
easier.
> However, the greatest advantage of a SSTO is in being reusable. Then I
> suggest studies be made on the SSME going the opposite direction: how
> can it be made to be reusable at much reduced maintenance cost? Now
> the SSME's have to be rebuilt after every flight costing ten's of
> millions of dollars. However, Henry Spencer a highly regarded expert
> on the history of space flight has said Rocketdyne studies show that
> with a lot of work maintenance could be reduced to $750K per flight:
>
> Engine reusability (Henry Spencer)
> http://yarchive.net/space/rocket/engine_reusability.html
>
> Spencer here said this would not be satisfactory for really large
> reductions in space costs. But this would be a reduction in SSME
> maintenance costs by 1 to 2 orders of magnitude, a major reduction in
> the costs for using the engine. The question is how much would it cost
> to make the necessary upgrades to the engine?
I recall (from discussions here) that the SSME's aren't torn down after
every flight like they used to be. They are, however, inspected after
every flight. What the exact procedures are (and current costs are), I
don't know for sure.
Jeff
--
Jeff Findley
@v7g2000yqh.googlegroups.com>, hal...@aol.com says...
>
> Anything from the apollo era MUST be redesigned and recertified to be
> man rated.
"Man rated" is a term which means anything NASA decides it means. They
write the rules, and they also write the wavers to the rules. Because
of this, the term "man rated" is essentially meaningless. It's just a
term which says that NASA is in control.
Brad Guth
> In article <32e2dccf-80ae-458d-9c23-
> b666af0d6...@l11g2000yqb.googlegroups.com>, rgregorycl...@yahoo.com
We'd first have to get rid of Fred and company of ZNR/GOPs that are
just out to screws us, then put Mook in charge. Are you game?
Brad Guth
> In article <687881c0-4a0c-40bd-81f7-4eedf406b846
>
> > Anything from the apollo era MUST be redesigned and recertified to be
> > man rated.
>
> "Man rated" is a term which means anything NASA decides it means. They
> write the rules, and they also write the wavers to the rules. Because
> of this, the term "man rated" is essentially meaningless. It's just a
> term which says that NASA is in control.
>
> Jeff
> --
> " Solids are a branch of fireworks, not rocketry. :-) :-) ", Henry
> Spencer 1/28/2011
Just like the Skull and Bones is still in control of our government,
whereas pretty much anything goes as long as they don't get caught
with their pants down or red-handed (so to speak).
Anything that's fly-by-rocket capable of taking human butts into space
is man-rated, even if a good many end up paying the ultimate price for
a one-way ticket. A safe and ideally reusable reentry capability
seems to be a much greater complexity and safety factor than just
getting up there to begin with.
Brad Guth
> [B]The key point is that you have significantly better leeway in your
> options and choices with relatively low financial risk.[/B]
>
> Another option for a manned launcher. In this report Boeing proposes
> heavy lift launchers using existing components:
>
> Heavy Lift Launch Vehicles with Existing Propulsion Systems.
> Benjamin Donahue, Lee Brady, Mike Farkas, Shelley LeRoy, Neal Graham
> Boeing Phantom Works,Huntsville, AL 35824
> Doug Blue
>
> One of the proposals is of a manned launcher for the Orion capsule
> using a shuttle ET propellant tank and four RS-68 engines. This does
> not use an upper stage but is not a single-stage-to-orbit vehicle
> because the final push to orbit is made by the onboard thrusters on
> the Orion spacecraft.
> However, it is interesting in this report comparison is made to the S-
> IVB upper stage on the Apollo rocket. I was reminded of a suggestion
> of Gary Hudson that the S-IVB would be single-stage-to-orbit with
> significant payload if it used the high efficiency SSME rather than
> the J-2 engine:
>
> A Single-Stage-to-Orbit Thought Experiment.
>
> In Hudson's proposal the vehicle could lift 10,360 lbs, 4,710 kg.
> This would be just enough to carry the crewed version of the
> [URL="http://en.wikipedia.org/wiki/Dragon_(spacecraft)"]Dragon
> spacecraft[/URL] without cargo. Boeing's proposal for a manned capsule
> the CST-100 might be launchable by this also since it is of comparable
> size and design to the Dragon:
>
> Boeing space capsule could be operational by 2015.
> BY STEPHEN CLARK
> SPACEFLIGHT NOW
>
> NASA has shown in their crewed spacecraft versions to want to hearken
> back to Apollo in their use of capsules. This SSTO idea of Hudson
> would have the advantage of using a proven Apollo component that is
> already manrated. The SSME's are also already manrated rather than the
> RS-68 of the Boeing proposal.
> Because of its small small size compared to the shuttle ET propellant
> tank it would also be relatively low cost as well as only needing one
> SSME engine. In fact it would even be smaller than the Falcon 9, Delta
> IV, and Atlas V expendable launchers. Note as well NASA is leaning now
> to using SSME's or their expendable versions rather than the RS-68 for
> their shuttle derived manned launchers.
> Hudson in his article stated the S-IVB was designed by the Douglas
> Aircraft Company, which merged with McDonnell Aircraft to form
> McDonnell Douglas. McDonnell Douglas is now a division of Boeing, so
> Boeing should have access to the design documents of the S-IVB.
> NASA in their shuttle-derived launcher studies have focused on
> getting a cheaper version of the SSME by making an expendable version.
> However, the greatest advantage of a SSTO is in being reusable. Then I
> suggest studies be made on the SSME going the opposite direction: how
> can it be made to be reusable at much reduced maintenance cost? Now
> the SSME's have to be rebuilt after every flight costing ten's of
> millions of dollars. However, Henry Spencer a highly regarded expert
> on the history of space flight has said Rocketdyne studies show that
> with a lot of work maintenance could be reduced to $750K per flight:
>
Since our NASA is dysfunctional and worse than some agencies, spending
loot they really do not have (If I did such they'd toss my ass in
prison), then perhaps the private funded and strictly for-profit Mook
way is worth considering.
Robert Clark
> [B]The key point is that you have significantly better leeway in your
> options and choices with relatively low financial risk.[/B]
>
> Another option for a manned launcher. In this report Boeing proposes
> heavy lift launchers using existing components:
>
> Heavy Lift Launch Vehicles with Existing Propulsion Systems.
> Benjamin Donahue, Lee Brady, Mike Farkas, Shelley LeRoy, Neal Graham
> Boeing Phantom Works,Huntsville, AL 35824
> Doug Blue
>
> One of the proposals is of a manned launcher for the Orion capsule
> using a shuttle ET propellant tank and four RS-68 engines. This does
> not use an upper stage but is not a single-stage-to-orbit vehicle
> because the final push to orbit is made by the onboard thrusters on
> the Orion spacecraft.
> However, it is interesting in this report comparison is made to the S-
> IVB upper stage on the Apollo rocket. I was reminded of a suggestion
> of Gary Hudson that the S-IVB would be single-stage-to-orbit with
> significant payload if it used the high efficiency SSME rather than
> the J-2 engine:
>
> A Single-Stage-to-Orbit Thought Experiment.
>
> In Hudson's proposal the vehicle could lift 10,360 lbs, 4,710 kg.
> This would be just enough to carry the crewed version of the
> [URL="http://en.wikipedia.org/wiki/Dragon_(spacecraft)"]Dragon
> spacecraft[/URL] without cargo.
It is notable that the upper stage of the Ares I is based on this S-
IVB stage. Then this upper stage as well should be able to act as an
SSTO with an SSME engine. This is important because the Ares I upper
stage was originally planned to use the SSME, so this means much of
the technical and financial analysis of using the SSME for the upper
stage of the Ares I has already been done.
However, because of the cost of the SSME engine and technical risk in
making it airstartable, the decision was made to use the J-2X engine
instead. But for the SSTO purpose you don't have the problem of making
it airstartable, and as I discussed the reusability maintenance costs
can be reduced by an order of magnitude for the SSME.
This report contains some of the specifications on the Ares I upper
stage:
NASA’s Ares I Upper Stage.
http://www.nasa.gov/pdf/231430main_UpperStage_FS_final.pdf
The propellant mass is listed as 138 mT, the dry mass of the stage as
17.5 mT, and the interstage mass, as 4.1 mT. See the second attached
image below taken from page 2 of the report. The interstage supports
the weight of the upper stage on top of the lower stage so won't be
needed for the SSTO version. So we can take the dry mass now as 13.4
mT.
We need to add onto this now the extra weight of using the SSME over
the J-2X engine. The report lists the J-2X mass as 2.5 mT. The SSME
mass is 3.1 mT, .6 mT heavier. This brings the dry weight to 14 mT.
A puzzlingly high value of 2.5 mT however is given for the avionics.
You wouldn't think it would need to be this high if it consisted of
just electronics and computer systems with modern miniaturization.
Most of the avionics is included in the "instrument unit". As you can
see from the first attached image below, the instrument unit is
regarded as a separate element of the upper stage and is contained
within the forward skirt of the stage. The forward skirt serves to
support the weight of the Orion CEV, so needs to have significant
strength and mass to support the 20,000+ kg weight of the Orion
spacecraft. So I'm wondering if that 2.5 mT mass is including the mass
of this forward skirt.
The forward skirt mass can certainly be reduced if using a Dragon
spacecraft at only one quarter the mass of the Orion. So that part of
the dry mass will be reduced, though it's uncertain if the avionics
mass itself can be reduced. In any case using 14 mT dry mass of the
Ares I upper stage, the 138 mT propellant mass, the 425 s average Isp
of the SSME, and a 9,200 m/s required delta-V to orbit, we can
calculate the payload to orbit can be 3 mT:
425*9.8ln(1 + 138/(14 + 3)) = 9,205 m/s.
This payload mass would not be enough for the Dragon spacecraft but
might be enough for an innovative new spacecraft proposal from the
University of Maryland:
Phoenix: A Low-Cost Commercial Approach to the Crew Exploration
Vehicle.
http://www.nianet.org/rascal/forum2006/presentations/1010_umd_paper.pdf
This uses a cylindrical shape for the capsule so would have more space
for the crew/passengers. It also uses a new design for a thermal
protection system called a "parashield" that would save weight over
the traditional ablative design. The mass of the capsule in this study
is given as 3,268 kg, so would only have to be reduced by a small
proportion to fit within the payload mass constraints.
However, it might be possible to increase the payload capability of
the SSME-powered Ares I upper stage to be able to carry even the
Dragon spacecraft. First, more propellant can be carried in the same
size tanks by densifying the propellant by subcooling:
Liquid Oxygen Propellant Densification Unit Ground Tested With a Large-
Scale Flight-Weight Tank for the X-33 Reusable Launch Vehicle.
http://www.grc.nasa.gov/WWW/RT/RT2001/5000/5870tomsik.html
As much as 10% more propellant can be carried by subcooled
densification. This corresponds to 10% greater mass that can lofted to
orbit. So from a 17 mT total of (launch vehicle + payload), up to 18.7
mT. This extra mass can go to extra payload, so to 4.7 mT payload.
Secondly, recent research has shown that from 10% to 20% weight
savings can be made off the structural weight on launch vehicles:
NASA Recalculates To Save Weight On Launchers.
Jan 5, 2011
By Frank Morring, Jr.
http://www.aviationweek.com/aw/generic/story.jsp?id=news/awst/2011/01/03/AW_01_03_2011_p53-277413.xml&headline=NASA%20Recalculates%20To%20Save%20Weight%20On%20Launchers&channel=space
The structural mass sans engine is 11 mT. If 10% weight can be saved
off this then that can be transferred to extra payload, bringing the
payload capacity to 5.8 mT. This would then be within the payload
capacity to carry the Dragon spacecraft.
Bob Clark
Images:
Ares I elements.
http://oi56.tinypic.com/5voinb.jpg
Ares I upper stage.
http://oi55.tinypic.com/1bxvr.jpg
Robert Clark
> [B]The key point is that you have significantly better leeway in your
> options and choices with relatively low financial risk.[/B]
>
> Another option for a manned launcher. In this report Boeing proposes
> heavy lift launchers using existing components:
>
> Heavy Lift Launch Vehicles with Existing Propulsion Systems.
> Benjamin Donahue, Lee Brady, Mike Farkas, Shelley LeRoy, Neal Graham
> Boeing Phantom Works,Huntsville, AL 35824
> Doug Blue
>
> One of the proposals is of a manned launcher for the Orion capsule
> using a shuttle ET propellant tank and four RS-68 engines. This does
> not use an upper stage but is not a single-stage-to-orbit vehicle
> because the final push to orbit is made by the onboard thrusters on
> the Orion spacecraft.
> However, it is interesting in this report comparison is made to the S-
> IVB upper stage on the Apollo rocket. I was reminded of a suggestion
> of Gary Hudson that the S-IVB would be single-stage-to-orbit with
> significant payload if it used the high efficiency SSME rather than
> the J-2 engine:
>
> A Single-Stage-to-Orbit Thought Experiment.
>
> In Hudson's proposal the vehicle could lift 10,360 lbs, 4,710 kg.
> This would be just enough to carry the crewed version of the
> [URL="http://en.wikipedia.org/wiki/Dragon_(spacecraft)"]Dragon
> spacecraft[/URL] without cargo.
The point of the matter is that if you use highly weight optimized
structures and high efficiency engines at the same time then what you
wind up with will be a SSTO capable stage. The Ariane 5 core stage is
another weight optimized structure using common bulkhead design for
its propellant tanks. The Ariane 5 core stage will also become SSTO if
using high efficiency SSME's instead of the Vulcain engines.
The specifications of the Ariane 5 are given here:
Ariane 5 Data Sheet.
http://www.spacelaunchreport.com/ariane5.html
The Ariane 5 generic "G" version could be lofted by a single SSME.
It's gross mass is listed as 170 mT, and the propellant mass as 158
mT, for a dry mass of 12 mT. The Vulcain engine is listed on this page
as weighing 1,700 kg:
Vulcain - Specifications.
http://www.spaceandtech.com/spacedata/engines/vulcain_specs.shtml
Switching to a heavier SSME engine would add 1.4 mT to the vehicle dry
mass, so to 13.4 mT for the dry mass. Using a 425s average Isp again
for the SSME, this would allow a 6,000 kg payload:
425*9.8ln(1 + 158/(13.4+6)) = 9,218 m/s.
We wish to use this for a man-rated vehicle though. The Ariane 5 was
originally intended to be man-rated using the Hermes spaceplane to
carry crew. However, it's not certain the degree this was followed-
through when the Hermes was canceled.
As with the Ares I upper stage, there are means to increase the
payload capacity. Subcooled densification allows 10% greater
propellant to be carried, so then 10% greater mass can be lofted to
orbit. This brings the total lofted weight from 19.4 mT to 21.3 mT.
This extra weight can go to extra payload, so from 6 mT to about 8 mT
in payload.
The Ariane 5 uses an aluminum alloy, but not the aluminum-lithium
alloy being used now for the lightest weight designs. Switching to
aluminum-lithium allows approx. 10% weight saving over the previous
aluminum alloy. The structural mass sans the SSME engine is 10.3 mT,
so about 1 mT would be saved that could go to extra payload.
I also mentioned before the new research that suggests 10% to 20% can
be saved in structural mass because of overly conservative design now
used. This would be another 1 mT that could be saved off the dry
weight. These weight savings could go to extra payload, bringing the
payload capacity to 10 mT.
ESA appears to be amenable to adapting the Ariane 5 core stage for
other uses, considering its agreement with ATK to use it for an upper
stage. So NASA or a private company should be able to make an
agreement with the ESA to use it for this purpose, based on getting
sufficient financing. In this regard, to get a prototype done at low
cost I suggest using the RD-0120 russian analogue of the SSME's. These
are in mothballs and probably can be obtained at greatly reduced
price. As a point of comparison the NK-33 was mothballed by the
russians and Aerojet was able to buy 36 of them for only $1.1 million
each(!) Aerojets version of the NK-33 is now on track to be used by
Orbital Sciences on their Taurus II launcher.
Then the Ariane 5 core version of this SSTO has the advantage over the
Ares I upper stage and S-IVB versions in being already built and in
current use. It also has now the capability when powered by an SSME or
RD-0120 to launch a SpaceX Dragon sized spacecraft to orbit without
having to use special fuel densifying or lightweighting methods.
NASA has said they want to support commercial space. Support for this
launcher would allow for a small, relatively low cost launcher that
would permit independent private companies to launch their own manned,
or cargo flights to space.
Bob Clark
Robert Clark
In these examples of using the SSME engine on existing stages to turn
them into SSTO's, I was using the trajectory averaged Isp value for
the SSME that Gary Hudson uses here:
A Single-Stage-to-Orbit Thought Experiment.
Gary C Hudson
http://www.spacefuture.com/archive/a_single_stage_to_orbit_thought_experiment.shtml
Note that a trajectory averaged Isp is always higher than just the
midpoint value between the sea level and vacuum values because the
rocket spends most of the time at high altitude, where the Isp is
close to the vacuum value.
However, I myself have not seen this actually computed. I have not
even seen it stated anywhere else except in this calculation by
Hudson. It should not be hard to do this calculation. You would need
to know the value of the thrust over the flight of the shuttle. I'm
sure this exists somewhere, possibly in graphical form. For instance
it's presented here for the thrust of the SRB's:
Space Shuttle Solid Rocket Booster.
3.1 Ignition
http://en.wikipedia.org/wiki/Space_Shuttle_Solid_Rocket_Booster#Ignition
You could also get a fairly good approximation to this trajectory
averaged Isp by knowing the altitude over the time of the flight and
using the formula for how the thrust for a rocket varies with ambient
air pressure.
Anyone know where the thrust or altitude profile for the shuttle is
given over the flight of the vehicle?
Bob Clark
Brad Guth
>
> Note that a trajectory averaged Isp is always higher than just the
> midpoint value between the sea level and vacuum values because the
> rocket spends most of the time at high altitude, where the Isp is
> close to the vacuum value.
> However, I myself have not seen this actually computed. I have not
> even seen it stated anywhere else except in this calculation by
> Hudson. It should not be hard to do this calculation. You would need
> to know the value of the thrust over the flight of the shuttle. I'm
> sure this exists somewhere, possibly in graphical form. For instance
> it's presented here for the thrust of the SRB's:
>
> Space Shuttle Solid Rocket Booster.
>
> You could also get a fairly good approximation to this trajectory
> averaged Isp by knowing the altitude over the time of the flight and
> using the formula for how the thrust for a rocket varies with ambient
> air pressure.
> Anyone know where the thrust or altitude profile for the shuttle is
> given over the flight of the vehicle?
>
> Bob Clark
Using HTP and a little something of high-energy hydrocarbon density is
perhaps the one and only SSTO option that isn't going to be too
volumetric bulky to start with. Even reusable liquid boosters should
not be excluded.
Solidified pure hydrogen doesn't exist, and apparently fusion is not
an option.
Robert Clark
> The point of the matter is that if you use highly weight optimized
> structures and high efficiency engines at the same time then what you
> wind up with will be a SSTO capable stage. The Ariane 5 core stage is
> another weight optimized structure using common bulkhead design for
> its propellant tanks. The Ariane 5 core stage will also become SSTO if
> using high efficiency SSME's instead of the Vulcain engines.
> The specifications of the Ariane 5 are given here:
>
>
> The Ariane 5 generic "G" version could be lofted by a single SSME.
> It's gross mass is listed as 170 mT, and the propellant mass as 158
> mT, for a dry mass of 12 mT. The Vulcain engine is listed on this page
> as weighing 1,700 kg:
>
>
> Switching to a heavier SSME engine would add 1.4 mT to the vehicle dry
> mass, so to 13.4 mT for the dry mass. Using a 425s average Isp again
> for the SSME, this would allow a 6,000 kg payload:
>
> 425*9.8ln(1 + 158/(13.4+6)) = 9,218 m/s.
>
> We wish to use this for a man-rated vehicle though. The Ariane 5 was
> originally intended to be man-rated using the Hermes spaceplane to
> carry crew. However, it's not certain the degree this was followed-
> through when the Hermes was canceled.
> As with the Ares I upper stage, there are means to increase the
> payload capacity. Subcooled densification allows 10% greater
> propellant to be carried, so then 10% greater mass can be lofted to
> orbit. This brings the total lofted weight from 19.4 mT to 21.3 mT.
> This extra weight can go to extra payload, so from 6 mT to about 8 mT
> in payload.
> The Ariane 5 uses an aluminum alloy, but not the aluminum-lithium
> alloy being used now for the lightest weight designs. Switching to
> aluminum-lithium allows approx. 10% weight saving over the previous
> aluminum alloy. The structural mass sans the SSME engine is 10.3 mT,
> so about 1 mT would be saved that could go to extra payload.
> I also mentioned before the new research that suggests 10% to 20% can
> be saved in structural mass because of overly conservative design now
> used. This would be another 1 mT that could be saved off the dry
> weight. These weight savings could go to extra payload, bringing the
> payload capacity to 10 mT.
The advantages of a SSTO are best utilized as a reusable vehicle.
Then you would have to subtract from this estimated payload mass the
mass needed for reentry and landing systems.
However, this SSTO could still be useful as an expendable vehicle.
Then you could have up to a 9,000 kg payload without the reentry and
landing systems. This is close to the 10,000 kg payload capacity of
the Falcon 9.
I saw this article that had an estimate for the price of an
expendable version of the SSME's:
PWR Offers Shuttle Engine Alternative.
Jul 15, 2009
By Joseph C. Anselmo
"The company also would manufacture additional engines using the
existing SSME design while beginning work on a modified design that
incorporates advances in the construction of nozzles and combustion
chambers. That would be ready to go into production within 3-4 years.
Maser estimates the modified SSME would cost two-thirds to four-fifths
of the original model - depending on the number ordered - and would be
'a little more expensive' than the company's RS-68 engine 'but in that
ballpark.'"
http://www.aviationweek.com/aw/generic/story_channel.jsp?channel=space&id=news/Engine071509.xml&headline=PWR%20Offers%20Shuttle%20Engine%20Alternative
Using a price of $40 million for the current SSME's this would
correspond to a price of from $26.7 to $32 million for the expendable
versions. Considering the fact the engines make up the bulk of the
cost of an expendable launcher, this expendable SSTO launcher very
well could be comparable in cost to the Falcon 9 at $50 million.
Bob Clark
willia...@mokenergy.com
make in the future to fund the development of a 200,000 barrel per day
production plant that used the Bergius Process to make oil from US
coal and hydrogen made from water with very low cost solar panels I
had developed. After fulfilling the contracts I would have been left
with an $85 billion asset and a supply chain to build more hydrogen
producing solar panels at very low cost, and more coal to liquid
Bergius reactors - gaining $85 billion each time one was finished.
There were some shenannigans in New York with some of the banks
involved and when that didn't work, the rules were changed about
testing procedures. As it currently stands, there is no approved test
for coal derived liquids that any futures market will accept. ASTM
hopes to have one in 2016, but that's just for Fischer-Tropsch.
Bergius isn't even considered.
In any case, until then, none of the coal-to-liquid deals like I
described can go forward. They will.
Now, had I completed the deal in 2005 I would be producing 200,000
barrels per day today, and be worth $85 billion. I would also be
churning out one of these systems every month. The world presently
uses enough oil today to need 395 of these systems. So, it would take
30 years to make a big dent - against growing demand. Oh yeah, and
oil would be trading in the $30 per barrel range.
This last piece was the killer. Because economically recoverable oil
is based on what it costs to extract versus what it brings in the
market. So, if oil is trading at $140 per barrel, there is more you
can get since you are willing to spend more. If oil is $30 per
barrel, oil wells will be shut down, and the economically recoverable
reserves of the oil companies is far smaller at $30 selling price than
at $120 selling price. So, their market capitalization, which is
based on their economically recoverable reserves, is a quadratic
function of oil price. That's why they like to see gradual oil price
rises over time - not a reversal of that trend - which a new non-oil
source brings.
There is also the issue of very low cost hydrogen made from water and
sunlight.
That makes things worse if you're an oil company with depleting
reserves.
Why post all of this in a thread about heavy lift launchers? lol.
Because of what I had planned to do with my SECOND $85 billion.
The first $85 billion went to pay for the over-sized supply chain I
was planning. It also went to acquire strategic assets, like coal
companies and rail roads - stuff like that.
The second $85 billion went to pay for GROWTH - and to stay ahead of
the competition.
This meant that I would use the second $85 billion to acquire LMT and
BA - restructure the two companies into three or four smaller
companies, and sell controlling interest in three of those companies
and keep the fourth. These companies would be to aircraft
manufacturers, one weapons manufacturer and a spaceship manufacturer.
Since the spaceship component was a drain on the companies before
acquisition, the pieces without that drain would be worth about $8
billion to $12 billion more structured this way. This money would pay
for the transaction, and get a small return on the $85 billion you
started with.
With this money - which is more money than these companies have
received from NASA over the past 20 years -- would be used to build a
rather large heavy lift launcher built around some existing engine and
airframe technologies - to place 632 metric tons into LEO at an
incredibly low cost per flight - with highly reusable launcher
technology.
http://www.scribd.com/doc/31261680/Etdhlrlv-Addendum
What would I do with this launcher?
Launch solar power satellites that beamed IR laser energy to the
terrestrial solar units that made hydrogen for the coal to liquid
processes described above.
http://www.scribd.com/doc/35439593/Solar-Power-Satellite-GEO
Why would I do that?
Because doing that would increase the amount of hydrogen produced by
these systems SEVEN (7) times! So, I need only 57 instead of 395
units to meet today's energy needs - which means I get things done in
5 years instead of 35 years.
Like I said, to remain competitive in the energy business and keep the
oil companies that produce conventional oil out of the market I have
come to dominate with my technology.
This is an exciting development! It is something that still can
happen, once we sort through the saleability of coal derived
liquids.
Even more exciting is what this means to space travel enthusiasts. A
fleet of highly reusable heavy lift launchers putting 632 tons into
LEO every week - with one unallocated launch every month - which can
be donated to the national or global space effort if not needed for
the commercial program - along with a sizeable charitable contribution
to develop payloads for it (which would be larger than the total of
all space budgets world wide) - would be a very positive development
indeed.
Of course this heavy lift launcher would also put up other commercial
systems. For example, a global wireless internet would be deployed in
very short order. Money from that asset would largely be the source
of donated dollars. The donations also are done partly in self-
interest. I have an $80 billion + asset that makes spacecraft and
rockets. It benefits me to have as many people as possible thinking
about uses for that asset and how to make it more valuable to the
human community.
Had the oil been floated in 2005 without mishap - we would be buying
up LMT/BA today. Oil would be trading at $30 per barrel range. And
likely the huge transfer of wealth out of the US banking system would
not have occurred, or been only a minor blip as wealth shifted from
the Middle East who is unhappy about US policies to the US.
By 2015 we would be back on the Moon and on our way to Mars. We would
be experimenting with solar power satellites beaming energy from the
vicinity of Sol. By 2020 we would have an outpost near Sedna and be
experimenting with sending useful energy across the solar system. At
the same time we would be using the gravity lens of the Sun as a
telescope objective and have detailed information of our cosmos on a
scale unimagined today. By 2025 we would be sending probes to nearby
stars even as we began shipping more material from the asteroids than
is mined on Earth today. By 2030 we would have remotely operated
robots operating throughout the galaxy - through a negative time delay
signal shunted through Sgr A* - as MEMS based spacecraft - powered by
laser beams from space - filled every garage on the planet - giving
first ballistic transport to everyone on Earth - and later allowing
people to live on orbit in their own space homes - and commute to
Earth. By 2040 with sufficient energy collected from the Sun, many of
those space homes would travel first acorss the solar system, and
later as technology developed, from star to star. By the 100th
anniversary of Sputnik, we would have the first cities around other
star systems.
Brad Guth
You lack focus, not to mention team members. Even the likes of
Bigelow Aerospace has more focus and his own team of expertise besides
himself.
When is Mokenergy going to deliver its first tonne of hydrogen for
$100?
hal...@aol.com
east war that shuts down exports for 6 months is probably enough
foreigners can buy us our assets, hope the chinese treat us decently
Brad Guth
At least it'll eliminate those spendy and time consuming reelections,
that we also can't afford. Perhaps India can run Texas and Florida,
or we can always sell Florida to Cuba and the southern half of
California to Mexico.
hal...@aol.com
nope mexico the US and canada except for the french part will all be
rolled into one country.
this will end illegal immigration from mexico, the will be citizens of
the North American Union
hopefully this will decrease a bit of the overhead of 3 seperate
countries.
just one president one legislature and one of everything
Brad Guth
It should cut the collective overhead by 90%, because so much is
either dysfunctional or over-lapping that it's currently far worse
than silly.
By all means, include Cuba plus a few other island nations that need
to become part of this multinational union.
How many bloody, nasty wars do you think this merger is going to take?
willia...@mokenergy.com
Its not the Chinese, its the bankers who backed Mao and Lenin in the
first place to create a bogus enemy for the USA to fight with so both
would run up massive debts. Its part of their plan to end the age of
nation states and private property - which began with their defeat at
the hands of Napoleon with the signing of the Treaty of Westphalia..
The central banks who work for themselves at each central bank
throughout the world, not the nation they're operating in and
chartered to, decide to make money easy or scarce based on whether
they think they can make the most money by growing an economy or
looting it. The business cycle for example, isn't natural, its
engineered by the banks and financiers to harvest wealth that has been
created by others. Its just like feeding and breeding sheep to shear
them every 8 months or so. You have a growth cycle and then a harvest
cycle where you get the profits for essentially nothing as people
become bankrupt and you pick up assets for nothing during bad times.
You then loan money to new owners at inflated prices and interest at
the beginning of the next cycle. Then, with everyone properly
chastised, you make money easy and build up the economy again.
The central problem for today is that these ruling oligarchs were
content to earn money by growing the economy when they felt it was
possible for the economy to grow. Today they believe universally that
we've reached peak oil output, and they're universally looting areas
that they feel are not 'sustainable' against the 'inevitable die off'
coming.
Those who use the most energy per person are the ones most
vulnerable.
They've taken care to develop a cover story and their shock troops to
enforce hard to accept changes in those regions. These are generally
known as green parties and movements.
If the shearing is particularly painful, and these days there are very
sophisticated econometric models to guide the financiers in making
these decisions, they arrange to have some sort of civil unrest, or
even limited war between those who are being sheared ruthlessly.
To this end there is always international tension maintained, which
can be exploited to explain this or that economic problem to those who
really don't understand or care to understand what's going on.
Who knows what the plans for the USA are? I don't know. I do know
that the USA has the highest per capita energy use on the planet. So
it likely won't be pleasant. I've taken up tax residency in New
Zealand - while still maintaining my US Citizenship. This is
permitted under new UN Rules passed in 2005 which essentially
dissolved the nation state's ability to collect taxes, and puts it in
the hands of the ruling oligarchy.
I do know its not the Chinese that are the ones in charge.
The Chinese owe the same bankers more percentage of their future
earnings than America does. That is changing. What they're likely
doing to America in my opinion is bringing it in line with the rest of
the world - making us more dependent on the international money system
they're creating and less dependent on ourselves. To achieve this end
they've got to get rid of what Marx would call the American
Bourgeoisie who know the score and will take effective action to break
their stranglehold if unopposed. Their numbers are small, so they
have to be rather ruthless about this.
That's the point of the current economic condition.
In the end, the new global dollar will be strongly based on the
productive capacity of 1 billion Chinese laborers along with the raw
material resources still remaining in Oceania which is untapped.
That's why according to some who claim to know that Canberra will be
the new capital of the new global governing bodies and why US built
Pine Gap near Alice Springs several decades ago and buried all the
world's computers there (and the basis of the fabulous econometric
models - soon to be joined with personality modeling based on internet
presence) and connected it to the world's most advanced global
surveillance and communications systems, and why Rupert Murdoch was
given the go ahead by these banks to buy out all the world's media
following US loss in Vietnam (which spawned infowar to counter the
perceived threat of a media out of control).
All these arguments were shaped by the bankers who had a long term
plan to defeat the nation states and re-establish their stranglehold
on the productive capacity of the bulk of their fellow men and women.
willia...@mokenergy.com
The North American Union and other changes to our ability to collect
taxes which the UN imposed on the nations of North America in 2005 -
signed into law by George Bush - mean the borders between nations are
non existent. This will allow shock troops to be brought in from
Mexico and South America to quell riots in the USA just as US troops
are sent to South America and Asia to quell riots there. The Russians
proved you couldn't trust local troops to fire on their own citizens
during the Moscow riots. The Chinese too at Tianemen Square proved
you couldn't trust local troops to kill their own citizens. So, the
present machinations were set into place.
The real surprise will come when ALL troops everywhere decide not to
fire on unarmed civilians. What you can do with a few fantatical
'expert' marksmen in a covert op is quite different with a lot of
troops marching around the world. My bet is that humanity will
prevail and they will ultimately turn on the small group of
controllers at the top and re-establish our humanity and freedoms.
Not just in America, everywhere.
Brad Guth
My guess is that they'll make a large number of mistakes that'll cost
humanity a billion or so lives as well as the rest of us having to
lose yet another decade, but in the end it'll be as you say. Perhaps
open technology and the best available science on a global scale will
prevail after all.
It seems that our government and their many agencies that seldom
bother to communicate with one another in order to supposedly save us
from the bad guys (I’m thinking perhaps 9/11 was just a dry run for
what’s to come), or to simply constructively interact for the
efficiency and greater common sense good, instead wants all the rest
of us village idiots as loyal taxpayers and dumbfounded consumers that
are essentially paying for absolutely everything (including Big
Energy, banking, investment, health care and insurance screw-ups), to
always fully trust the authority of our government and all of their
agencies plus special-interest operatives and each and every one of
their money changers, as to only believing that the extremely nearby
planet Venus is simply too freaking greenhouse hot and nasty for even
the smartest of intelligent life or whatever technology can possibly
deal with. Well guess what folks, they is worse than dead wrong about
that, because even a dysfunctional Goldilocks knows better.
http://nssdc.gsfc.nasa.gov/imgcat/hires/mgn_c115s095_1.gif
Guth Venus, at ten times resample/enlargement of the area in
question:
https://docs.google.com/File?id=ddsdxhv_4fdgd46df_b
http://translate.google.com/#
Brad Guth / Guth Venus / Blog and my Google document pages:
http://groups.google.com/group/guth-usenet?hl=en
http://bradguth.blogspot.com/
http://docs.google.com/View?id=ddsdxhv_0hrm5bdfj
Robert Clark
> heavy lift launchers using existing components:
>
> Heavy Lift Launch Vehicles with Existing Propulsion Systems.
> Benjamin Donahue, Lee Brady, Mike Farkas, Shelley LeRoy, Neal Graham
> Boeing Phantom Works,Huntsville, AL 35824
> Doug Blue
>
> One of the proposals is of a manned launcher for the Orion capsule
> using a shuttle ET propellant tank and four RS-68 engines. This does
> not use an upper stage but is not a single-stage-to-orbit vehicle
> because the final push to orbit is made by the onboard thrusters on
> the Orion spacecraft.
> However, it is interesting in this report comparison is made to the S-
> IVB upper stage on the Apollo rocket. I was reminded of a suggestion
> of Gary Hudson that the S-IVB would be single-stage-to-orbit with
> significant payload if it used the high efficiency SSME rather than
> the J-2 engine:
>
> A Single-Stage-to-Orbit Thought Experiment.
> Gary C Hudson
> http://www.spacefuture.com/archive/a_single_stage_to_orbit_thought_experiment.shtml
>
> In Hudson's proposal the vehicle could lift 10,360 lbs, 4,710 kg.
> This would be just enough to carry the crewed version of the
> [URL="http://en.wikipedia.org/wiki/Dragon_(spacecraft)"]Dragon
> spacecraft[/URL] without cargo.
In doing some background web searches, I found that the upper stage
of the Direct team's Jupiter-246 vehicle also would become SSTO when
switched to a SSME engine. I guess I should not have been surprised by
this. The thesis I have been arguing repeatedly via email with
individuals in NASA and the industry and on space oriented forums such
as this one is that if you use BOTH the most weight optimized designs
AND the highest efficiency engines available, then what you will wind
up with will be SSTO capable whether you intend it to or not.
By highest efficiency engines I don't mean just an engine optimized
to have a high vacuum Isp only. I mean an engine of highest efficiency
over the entire flight range to orbit. For hydrogen engines that is
the SSME, and the Russian analogue RD-0120. However, the point of the
matter is that the same is true of kerosene-fueled vehicles, when
using both highly weight optimized structures and highest efficiency
engines, such as the NK-33 or RD-180.
That a SSME-powered Jupiter-246 upper stage would be SSTO capable is
important since the Direct team is more amenable to thinking outside
the box. So they would be more amenable to the idea you could have a
SSTO vehicle. And in fact at least the expendable version for this
SSTO would be no more difficult than their proposal for the upper
stage on the Jupiter-246.
Here is a diagram showing the specifications of the Direct team's
Jupiter-246:
It uses 6 RL-10B-2 engines. According to the specifications here
these weigh about 300 kg each:
RL10B-2
Propulsion System
http://www.pratt-whitney.com/StaticFiles/Pratt%20&%20Whitney%20New/Media%20Center/Assets/1%20Static%20Files/Docs/pwr_RL10B-2.pdf
So exchanging these for a SSME will add about 1,300 kg to the upper
stage weight. The dry mass will be increased then to 13,150 kg, and
the gross mass to 204,000 kg. However, by the Space Shuttle main
engine thrust specifications, even at 109% thrust this comes to only
417,300 lbs, or 189,700 kgf:
Space Shuttle main engine.
10 Thrust specifications.
http://en.wikipedia.org/wiki/Space_Shuttle_main_engine#Thrust_specifications
So we'll reduce the propellant load to be lifted by the SSME.
We'll take the liftoff thrust/weight ratio to be 1.2. This will bring
the gross mass down to 170,000 kg. Then the propellant mass has to be
reduced by 34,000 kg. This brings the propellant mass down to 156,850
kg. Note this results in a mass ratio close to 13, well sufficient for
SSTO with a hydrogen-fueled engine.
This mass ratio for a hydrogen-fueled stage of 13 is high, but the
original number for the Jupiter-246 upper stage is even higher at
above 17. These high values for the Direct teams launcher led to some
doubts about their calculations, but an analysis by Dr. Steven
Pietrobon showed it was in keeping with historical trends for upper
stages:
Analysis of Propellant Tank Masses.
http://www.nasa.gov/pdf/382034main_018%20-%2020090706.05.Analysis_of_Propellant_Tank_Masses.pdf
Then using Gary Hudson's 425s average Isp for the SSME from his "A
Single-Stage-to-Orbit Thought Experiment" article and the 9,200 m/s
required delta-V value for orbit, this stage as an SSTO could loft
6,200 kg to orbit:
425*9.8ln(1 + 170,000/(13,150 + 6,200) = 9,200 m/s.
Again, we might be able to loft 10% greater total mass to orbit with
propellant densification by subcooling and also shave 10% off the
structural mass of the stage with the recent weight saving research.
This will bring the payload mass up to about 9,000 kg.
In this calculation I kept the same size tanks and only used them
partially filled. This might be useful if for instance the Jupiter-246
upper stage was built to the original Direct teams specifications and
you wanted to use the same size stage, though switched to a SSME
engine, for the SSTO application to save on costs.
However, you could save additional weight off the stage if you used
smaller propellant tanks for the SSTO application. I estimate about
900 kg could be saved with the smaller tanks that could go to
additional payload.
You could also get the SSTO without reducing propellant by using two
SSME engines. For a manned launcher this would be preferred to have
engine out capability. The dry mass with one SSME I calculated to be
13,150 kg. Adding on a second SSME would bring the dry mass to about
16,300 kg.
The propellant load is 190,850 kg. Then you could loft 7,200 kg
payload:
425*9.8ln(1 + 190,850/(16,300 + 7,200)) = 9,210 kg.
Again with propellant densification and recent lightweighting
techniques this payload might be raised to about 10,000 kg.
Bob Clark
Robert Clark
> On Feb 3, 2:46 am, Robert Clark <rgregorycl...@yahoo.com> wrote:
> ...
> In doing some background web searches, I found that the upper stage
> of the Direct team's Jupiter-246 vehicle also would become SSTO when
> switched to a SSME engine. I guess I should not have been surprised by
> Then using Gary Hudson's 425s average Isp for the SSME from his "A
> Single-Stage-to-Orbit Thought Experiment" article and the 9,200 m/s
> required delta-V value for orbit, this stage as an SSTO could loft
> 6,200 kg to orbit:
>
> 425*9.8ln(1 + 170,000/(13,150 + 6,200) = 9,200 m/s.
>
> Again, we might be able to loft 10% greater total mass to orbit with
> propellant densification by subcooling and also shave 10% off the
> structural mass of the stage with the recent weight saving research.
> This will bring the payload mass up to about 9,000 kg.
> In this calculation I kept the same size tanks and only used them
> partially filled. This might be useful if for instance the Jupiter-246
> upper stage was built to the original Direct teams specifications and
> you wanted to use the same size stage, though switched to a SSME
> engine, for the SSTO application to save on costs.
>
That 170,000kg number should be the propellant mass which is
156,850kg. So that calculation should read:
425*9.8ln(1 + 156,850/(13,150 + 6,200)) = 9,200 m/s.
Bob Clark
Robert Clark
> ... In doing some background web searches, I found that the upper stage
> of the Direct team's Jupiter-246 vehicle also would become SSTO when
> switched to a SSME engine. I guess I should not have been surprised by
> this. The thesis I have been arguing repeatedly via email with
> individuals in NASA and the industry and on space oriented forums such
> as this one is that if you use BOTH the most weight optimized designs
> AND the highest efficiency engines available, then what you will wind
> up with will be SSTO capable whether you intend it to or not.
> By highest efficiency engines I don't mean just an engine optimized
> to have a high vacuum Isp only. I mean an engine of highest efficiency
> over the entire flight range to orbit. For hydrogen engines that is
> the SSME, and the Russian analogue RD-0120.
There might still be some resistance to using the upper stage as a
SSTO. However, the point of the matter is even if you use these upper
stages as part of a multistage system you are still better off using
*both* highly weight optimized structures *and* engines of highest
surface-to-orbit-efficiency (not just vacuum optimized engines) *at
the same time*.
We'll use in this case parallel staging of the same sized stages, a
bimese version, but using cross-feed fueling. This is a fueling method
that has both stages firing, as with parallel staging, but all the
propellant is coming from only a single stage at a time. Then when
that stage exhausts its propellant, it is jettisoned, and the
remaining stage proceeds on with its own full tank of propellant still
remaining.
Let's see how much payload we can carry in this case. Again assume the
425s trajectory averaged Isp of Hudson, and the 9,200 m/s required
delta-V for orbit.
Estimate the possible payload as 29,000 kg. For the first segment of
the flight the achieved delta-V would be: 425*9.8ln(1+156,850/
(2*13,150 + 156,850 +29,000)) = 2,305 m/s.
For the second segment, use the 455s vacuum Isp of the SSME's:
455*9.8ln(1 + 156,850/(13,150 + 29,000)) =6,921. And the total delta-V
is 9,226 m/s, sufficient for orbit with a 29,000 kg payload.
Note that a 29,000 kg payload is sufficient to even carry a Orion
capsule, at least in an expendable version of the staged vehicle
without reentry and landing systems.
Then you have different options for the vehicle. As a single stage it
could carry a small capsule such as the SpaceX Dragon, or the Boeing
CST-100. But using twinned copies of it, it would be able to loft the
heavier Orion spacecraft.
Bob Clark
Robert Clark
> ... In doing some background web searches, I found that the upper stage
> of the Direct team's Jupiter-246 vehicle also would become SSTO when
> switched to a SSME engine. I guess I should not have been surprised by
> this. The thesis I have been arguing repeatedly via email with
> individuals in NASA and the industry and on space oriented forums such
> as this one is that if you use BOTH the most weight optimized designs
> AND the highest efficiency engines available, then what you will wind
> up with will be SSTO capable whether you intend it to or not.
> By highest efficiency engines I don't mean just an engine optimized
> to have a high vacuum Isp only. I mean an engine of highest efficiency
> over the entire flight range to orbit. For hydrogen engines that is
> the SSME, and the Russian analogue RD-0120.
>...
Another highly weight optimized stage was the S-II second stage on
the Saturn V. According to this Wikipedia page it was even better
optimized than the S-IVB stage :
Saturn V.
S-II second stage.
http://en.wikipedia.org/wiki/Saturn_V#S-II_second_stage
The 5 J-2 engines used had a mass of 1,580 kg each, for a total mass
of 7,900 kg. You'll need 3 of the SSME's operating at 109% thrust to
lift the mass. So the 7,900 kg mass of the engines is replaced with
9,300 kg. And the 36,000 kg S-II dry mass is raised to 37,400 kg and
the gross mass is raised to 481,400 kg.
Now using Gary Hudson's 425s trajectory averaged Isp for the SSME
engines, and the 9,200 m/s required delta-V to orbit. We get a 17,000
kg payload:
425*9.8ln((481400 + 17000)/(37400 + 17000)) = 9,225 m/s
However, again we can get 10% greater total mass to orbit by
propellant densification. This brings the payload to 22,440 kg. Also
perhaps 10% off the structural mass can be saved by using aluminum-
lithium alloy. And an additional 10% mass can be saved by the new
weight saving methods. These weight savings can go to extra payload to
bring the payload mass up to 28,000 kg. Note this is sufficient now to
carry the Orion spacecraft as a SSTO.
Bob Clark
willia...@mokenergy.com
hydrogen, with J2 pumpsets, feeding an aerospike nozzle achieves the
highest specific impulse for a chemical propellant ever test-fired in
a rocket engine was lithium, fluorine, and hydrogen (a tripropellant):
542 seconds (5,320 m/s).
http://www.aiaa.org/content.cfm?pageid=406&gTable=mtgpaper&gID=40999
http://upload.wikimedia.org/wikipedia/en/1/19/Annular-Aerospike.jpg
While this combination is reportedly impractical, those
impracticalities are easily sorted out by colloidally suspending the
lithium within the hydrogen after coating with lithium hydride. The
propellant combination is a tank of hydrogen/lithium mix, which is the
fuel consisting of 20% lithium by weight, and 80% hydrogen by weight
and fluorine which is the oxidizer. The fuel suspension averages 162
kg per cubic meter and the oxidizer suspension averages 1,520 kg per
per cubic meter and they mix in a 9:1 ratio - oxidizer to fuel
producing an average density of 1,366 kg per cubic meter.
Isp=542 seconds
Ve = 5,320 m/s
rho=1,366 kg per cubic meter.
http://en.wikipedia.org/wiki/S-II
http://www.astronautix.com/stages/saturnii.htm
http://upload.wikimedia.org/wikipedia/commons/2/22/Saturn_V_second_stage.jpg
structure mass: 40 metric tons
Volume: 1,937.7 cubic meters
rho = 1.366 metric tons per cubic meter
Propellant mass: 2,660 metric tons
Take off weight: 2700 metric tons
Ve = 5320 (vacuum)
Ve = 5000 (sea level)
Ve = 5150 (average)
mu = 1 - 1/exp(Vf/Ve) = 1 - 1/exp(9.2/5.15) = 0.83244
Take off weight = Propellant weight / mu = 2,660 / 0.83244 = 3,195
metric tons
Final weight = take off weight - propellant weight = 3,195 - 2,660 =
535 metric tons
Payload weight = final weight - structure = 535 - 40 = 495 metric tons
Take off thrust = 1.4x 3,195 metric tons = 4,473 metric tons = 43.844
Mega Newtons.
Mass flow rate = Thrust/ Ve = 43.844e+6 / 5,150 = 8,513.47 kg/sec
Fluorine 7,662.1 kg/sec --> 5,040.9 liters/sec
Lithium/H2 851.5 kg/sec --> 7,339.2 liters/sec
The J2 engine pumps 504.5 liters/sec of H2 and 179.6 liters/sec of LOX
- so, five modified LOX pumps and three modified H2 sets are needed
for each aerospike.
Jeff Findley
17g2000prr.googlegroups.com>, willia...@mokenergy.com says...
>
> Using fluorine and hydrogen with lithium colloidally suspended in the
> hydrogen, with J2 pumpsets, feeding an aerospike nozzle achieves the
> highest specific impulse for a chemical propellant ever test-fired in
> a rocket engine was lithium, fluorine, and hydrogen (a tripropellant):
> 542 seconds (5,320 m/s).
Extremely toxic propellants don't exactly make for cheap launches. To
say nothing of the environmental impact.
What's wrong with LOX and LH2? You seem to be getting all worked up
over the maximum theoretical ISP. This is the "performance uber alles"
mindset that Henry Spencer repeatedly warned about in these newsgroups.
As Henry always said, it's price per lb to LEO you want to optimize, not
ISP. Do you just not care about the economics, or are you deluded by
the "conventional wisdom" that you need to maximize ISP, minimize the
dry mass fraction, and the like? Note that "conventional wisdom"
brought us Delta IV and Atlas V, which aren't exactly shining examples
of low cost.
Jeff Findley
fjmc...@gmail.com says...
>
> willia...@mokenergy.com wrote:
>
> >Using fluorine and hydrogen with lithium colloidally suspended in the
>
> Is a science experiment that will consume years of effort.
Not to mention the fact that it's pointless if your goal is minimizing
launch costs ($ per lb to LEO). LOX and LH2 may be a p.i.t.a. to work
with since they're cryogenic, but adding fluorine and lithium to the mix
is literally a toxic recipe for high costs.
I think Mookie is succumbing to what Henry Spencer called the
"performance uber alles" mindset. Damn the complexity, damn the cost,
we're going to optimize our launch vehicles for maximum performance.
John Park
> In article <0b61c1af-7981-4bfc-96ae-69d1dc3ac08f@
> 17g2000prr.googlegroups.com>, willia...@mokenergy.com says...
>>
>> Using fluorine and hydrogen with lithium colloidally suspended in the
>> hydrogen, with J2 pumpsets, feeding an aerospike nozzle achieves the
>> highest specific impulse for a chemical propellant ever test-fired in
>> a rocket engine was lithium, fluorine, and hydrogen (a tripropellant):
>> 542 seconds (5,320 m/s).
>
> Extremely toxic propellants don't exactly make for cheap launches. To
> say nothing of the environmental impact.
Would such a launch ever be permmitted? The product, HF, isn't that much nicer
than F2 itself. (I suppose you might get away with it for a while somewhere in
the Sahara...)
--John Park
Pat Flannery
> Would such a launch ever be permmitted? The product, HF, isn't that much nicer
> than F2 itself. (I suppose you might get away with it for a while somewhere in
> the Sahara...)
Well, fluorine engines got ground tested by both the US and USSR. Here's
the AMPS-1 fluorine/LH2 engine that was going to power the Lockheed
FDL-5 Air Force spaceplane under test:
http://www.picturetrail.com/sfx/album/view/8379229
I don't even like of the looks of the exhaust on that thing.
In that case the spacecraft would probably have been launched from a
carrier aircraft out over the ocean rather than off of a launchpad:
http://www.picturetrail.com/sfx/album/view/8379229
The Soviet fluorine engine was never flown either:
http://www.astronautix.com/engines/rd301.htm
Pat
John Park
Thanks. Ick. Though I notice that the Soviet vehicle was intended as an
upper stage, so presumably it wouldn't have been used in the atmosphere.
(But the model builder for the Lockheed spaceplane seems to imagine
using LF2 as a fuel raises no particular diificulties. Ick, again.)
--John Park
willia...@mokenergy.com
> Jeff Findley (jeff.find...@ugs.nojunk.com) writes:
> > In article <0b61c1af-7981-4bfc-96ae-69d1dc3ac08f@
>
> >> Using fluorine and hydrogen with lithium colloidally suspended in the
> >> hydrogen, with J2 pumpsets, feeding an aerospike nozzle achieves the
> >> highest specific impulse for a chemical propellant ever test-fired in
> >> a rocket engine was lithium, fluorine, and hydrogen (a tripropellant):
> >> 542 seconds (5,320 m/s).
>
> > Extremely toxic propellants don't exactly make for cheap launches. To
> > say nothing of the environmental impact.
>
> Would such a launch ever be permmitted? The product, HF, isn't that much nicer
> than F2 itself. (I suppose you might get away with it for a while somewhere in
> the Sahara...)
>
> --John Park
Interesting quote you're responding to. The quote was extracted from
another HF booster design I put out a few months ago, and totally
ignored the fact that I was posting something new here.
Here I'm proposing a particle beam propulsion system would be
responsible for operations near Earth, while HF would be used near the
lunar surface only during the first three years of operation. After a
particle beam propulsion based launcher were erected on the moon, HF
would be used in course correction.
But let's address your legitimate concern about HF emissions.
Please understand I would not build any of these things until I've
made considerable money in the energy business. I would then use that
money to acquire major aerospace contractors. I would then reorganize
those contractors removing their money losing space faring
capabilities. I would then sell the pieces back to the market,
getting more than I paid. I would then use the money to combine all
the space faring components to build a series of space vehicles to
carry out economically important missions for Earth. These include;
(1) Global hotspot; (including telepresence)
(2) Space power
(3) Asteroid retrieval
(4) Space factory (via telepresence)
(5) Mining retrieved asteroids on Earth Orbit
(6) Personal spacehip
(7) Personal space home
Along the way I might build Li-H-F spacecraft that emit HF.
HF is relatively low cost compared to spacecraft and payload costs.
Lithium costs $6,600 per tonne. Fluorospar costs $400 per tonne. Its
CaF2. So, its 49% by weight Fluorine. So, counting the energy costs
involved, its cost is $1,200 per tonne in a dedicated facility that
makes hundreds of thousands of tons per year for space vehicle
application. Another $700 per tonne for NG derived hydrogen, though
solar derived hydrogen is available at less than $100 per tonne at
this point.
Propellant Ratio Cost Cost
Hydrogen 7.14% $ 700 $ 50.00
Lithium 25.00% $6600 $1,650.00
Fluorine 67.86% $1200 $ 814.29
TOTAL: $2,514.29
Space hardware runs $10,000,000 to $20,000,000 per tonne. So, these
costs are small compared to that. By reorganizing the space faring
assets of the planet for efficient operation, I plan to reduce costs
on average to $2,000,000 per tonne - first pass. And then implement
an investment program to maintain cost reductions in a very similar
way computing costs have been reduced over the years. This all part
of achieving my business goals in space.
Now, please understand, I posted a particle beam propulsion system
immediately above. But, since some fool randomly chose to respond to
a hydrogen-fluorine booster post months before - let's look at this
objection about HF emissions.
The booster that I proposed before the current particle beam system,
consisted of 7 components each carrying 600 tonnes of propellant iirc
and so would emit about 4,200 tonnes of reaction product at each
launch. A total of 31% or 1,302 tonnes of these emissions are HF.
The world produces 165,000 tons of hydrogen fluoride per year counting
only its aluminum production along with emissions from its use of
fossil fuels. By replacing the energy infrastructure on the planet
with low cost hydrogen including the use of hydrogen instead of carbon
or fluorine as a reducing agent, I would gain a large, number of
credits each year under the current system of GHG trading.
The total amount of credits is subject to negotiation with the folks
in Geneva who administer it. The argument is usually, that these
emissions go up 4% per year, and they agree to give you the current
year times the life span of the equipment, minus the contributions of
the equipment itself. So, lets say 33 years life span and 3 years
production is counted against it, this is 30 years net savings. So
that's a total of 4,950,000 tonnes of HF credits along with the
dollars to buy up the space faring assets and operate them in an
efficient manner.
Now, dividing 1,302 tonnes per launch by 4,950,000 tonnes of credits
obtains 3,801 launches. Over 360 months - 30 years - this is over 10
launches per month. One every three days. With no net emissions on
the planet. In fact, we would be migrating out of the chemical
booster in about 15 years - or 180 months. So, we would have a basis
to increase launch rates above that over this period, combined with
the solemn promise we'd stop terrestrial operations within that
period. We wouldn't be able to spend credits we didn't have anyway.
Finally, we're taking a global emission, and turning into a point
emission by this process. By selecting the launch location - say
somewhere in Libya as you suggest - this is also an advantage. So, we
would likely have some negotiating room around this.
Now, is this worth the advantage? Well, let's consider a 5.2 km/sec
exhaust speed with an average propellant density of 1.3x that of water
vs a 4.5 km/sec exhaust speed with an average propellant density 0.35x
that of water.
Oops - its about 4 AM here in Christchurch and I just had a wee
shake. I've got to go check a few things on the farm. I will be
back. Basically higher density and higher performance radically
improve lift to orbit for a given vehicle size, and despite increased
propellant cost over H2 and O2 - you do get early adopter advantage.
As the program to develop lower cost vehicles and payloads progresses,
there will be pressure to use less expensive propellants, and as
volumes to orbit increase, you will over-run your HF allotments. So,
this isn't a stationary thing.
Still, 1,000 tonnes to orbit every other day represents a great
advantage.
oh - just had a power fluctuation. Got to go check the hydro-power.
Ciao.
willia...@mokenergy.com
H-O: rho=0.35 MT/m3, Isp=4.5 sec.
Seven Elements:
600 tonnes propellant (Li-H-F)
161.5 tonnes propellant (H-O)
45 tonnes structure
Li-F-H: 750 tonnes payload to Low Earth Orbit (LEO)
H-O: 72 tonnes payload to LEO
(see calculations below)
STAGE 1 - TAKE OFF
750 72 Payload
5,265.0 1,517.8 TOW
2,400.0 646.2 S1-propellant
5.2 4.5 Ve
0.4558 0.4257 u-S1-prop frac
3.2 2.5 dV - S1
STAGE 2 -
2,685.0 691.6 S2-TOW
1,200.0 323.1 S2-propellant
5.2 4.5 Ve
0.4469 0.4671 u-S2-prop frac
3.1 2.8 dV - S2
STAGE 3 -
1,395.0 278.5 S3-TOW
600.0 161.5 S3-propellant
5.2 4.5 Ve
0.4301 0.5800 u-S3-prop frac
2.9 3.9 dV - S3
TOTAL IDEAL DELTA VEE
9.2 9.2 Total dV
(same as Ideal Shuttle Delta Vee)
willia...@mokenergy.com
Imagine 8 billion people living in their own self-sustaining space
homes the size of Los Angeles county, orbiting Earth, tended by highly
automated teleoperated robots, possessing a garage of interplanetary
spacecraft - powered by the Sun, and feeding from raw materials
imported from the asteroid belt.
We could have this today if appropriate investments were made.
We could have this within 30 years with the right investment program
were started today.
Where do we start?
http://www.scribd.com/doc/50130613/EU-Critical-Minerals-Report-2010-Annex
With the most valued materials.
The world production of things like Gallium and Germanium is in the
100 metric ton per year range. Indium 500 metric tons per year.
Palladium and Platinum group metals more than 100 metric tons, less
than 500 metric tons.
So, despite the fact that humanity consumes 30 billion metric tons of
raw material per year, and despite the fact that if eight billion of
us consumed at the rate of the average American millionaire we'd
consume 400 billion metric tons of stuff - we can begin to recover ALL
of this planet's needs for rarer materials from the asteroids with
quite modest infrastructure.
The ability to recover 2,500 metric tons per year of selected
materials from the asteroid belt, more cheaply than it can be
recovered on Earth allows us dominate a market that returns more money
than all the space programs in the world combined spent on space
travel at their peak!
The ability to recover 33,000 metric tons per year of these same
materials from the asteroid belt, in quantities unavailable on Earth,
allows us to make the point, that we are not constrained to this
planet's limited resources, or the restrictions imposed by the
biosphere.
Expanding on these core abilities once established and funded, will in
very short times lead to the end game where 8 billion of us are living
in space homes orbiting above the planet.
A laser powered rocket capable of accelerating asteroid mass as plasma
to 50 km/sec in a desired direction, requires 1.25 GJ per kg of
ejected material. To move an object from Ceres to Earth orbit along
a Hohman transfer orbit requires a delta vee of 16 km/sec. This means
that 27.4% of a rocket's mass must be ejected to carry out this
maneuver in two impulses. That means 377 kg of material must be
ejected at 50 km/sec to transport 1 metric ton into Earth orbit for
processing. So, each metric ton requires 471.4 GJ of energy expended
to return to Earth orbit. So, each metric ton returned per year
requires the installation of 14,938 Watts of laser propulsion
capability.
I have developed terrestrial solar installations that cost $0.07 to
$0.04 per peak watt. I have developed systems that on Earth orbit
will reduce these costs to $0.01 per peak watt. I have developed sun
orbiting systems that reduce these costs to less than 1/20 cent per
peak watt.
http://www.youtube.com/watch?v=dbWNnVsBhOg
http://www.youtube.com/watch?v=QvE-bkc0Uxo&playnext=1&list=PL838BBE4ECC513CCB&index=1
http://www.scribd.com/doc/20024019/White-Paper-to-Mok-FINAL-1
So, at these prices 15,000 watts have the following fixed costs
associated with return one ton per year;
CAPITAL COST OF 15,000 WATTS OF LASER ENERGY USING MOK TECH
Terrestrial: $1,050 - $600
Earth Orbit: $150
Sun Orbit $8
These fixed costs are amortized over 30 years. So, the cost for each
ton then becomes;
COST CHARGED PER METRIC TON RETURNED
Pct Terrest. Terrest. GEO Sun
Early Mature
$1,050 $600 $150 $8.00 Capital Cost
0% $35.00 $20.00 $5.00 $0.27 National Program
3% $53.57 $30.61 $7.65 $0.41 Subsidized
9% $102.20 $58.40 $14.60 $0.78 Commercial
27% $283.72 $162.12 $40.53 $2.16 Early-Stage Venture
So, even early-stage ventures that demand high discount rates using
terrestrial solar energy to power terrestrial laser systems, have the
potential to return anything found in the asteroid belt at less than
$300 per metric ton. As technology develops and matures, we can
expect this to grow to the point where we import anything that costs
more than $2 per metric ton to mine on Earth, and with the backing of
nation-states who print their own money, this could drop even further
still - to under $0.30 per metric ton!
A liter of water imported from Ceres then, would cost no more than
$0.30 and could drop below 1/3 cent. This means a liter of water
could be dropped from orbit to anyone who wanted a liter of water
anywhere on Earth.
2,500 metric tons per year require a modest 37.5 MW laser system.
We've already built 50 MW laser systems. At $1 million per ton and
higher costs, the recovery of rare materials from the asteroids could
get this started.
The Atacama desert is a good place to start.
willia...@mokenergy.com
Propulsive effects have been created on unprepared surfaces by laser
action. This technique will be developed to mine the asteroid belt
using entirely terrestrial based systems.
Low cost terrestrial solar panels creating hydrogen and oxygen from a
water store, being fed into a super-critical boiler that drives a
1,000 MW steam turbine system, which then drives a 300 MW high-
efficiency laser, operating in the Atacama desert. This region is
chosen for its high level of solar insolation, its broad view of the
sky, and its very low dispersion due to atmospheric effects. With
adequate optics a 300 MW laser will beam energy across the solar
system and produce propulsive effects on unconfined surfaces of
objects as far away as Ceres;
http://www.springerlink.com/content/v182427785713146/
http://oai.dtic.mil/oai/oai?verb=getRecord&metadataPrefix=html&identifier=ADA344774
http://adsabs.harvard.edu/abs/2009AIPC.1087..170T
We can also work with optical and IR telescopes that operate in
conjunction with laser pulses that cause asteroids to eject plasma
that are then characterized spectroscopically.
In this way a terrestrial solar powered center can survey all the
small bodies of the solar system, first identifying them with optical,
microwave, radar and lidar system - which then intensify the
characterization by creating plasma plumes on them to read their
composition. Then, on selected bodies more powerful structured pulses
modify their rotation and movement through space, giving us detailed
information on their density, moment of inertia and so forth.
Finally, for those rich bodies that interest us, we carve off pieces
if needed, and bring them back to Earth by directing structured plasma
plumes to impart the desired thrust effects. For 10 tonne to 100
tonne per day return rate, we can land the materials in a return zone
in the Atacama desert for recovery.
The revenue from this operation will then be used to expand the system
with space based lasers to enlarge the return rate, resulting in
ultimately, the entire industrial infrastructure of humanity being
placed on orbit powered by sunlight and fed from raw materials
recovered from the asteroids.
Brad Guth
> What good is space travel?
>
> Imagine 8 billion people living in their own self-sustaining space
> homes the size of Los Angeles county, orbiting Earth, tended by highly
> automated teleoperated robots, possessing a garage of interplanetary
> spacecraft - powered by the Sun, and feeding from raw materials
> imported from the asteroid belt.
>
> We could have this today if appropriate investments were made.
>
> We could have this within 30 years with the right investment program
> were started today.
>
> Where do we start?
>
Problem is, you'll be long dead before 10% of what you've suggested
takes place, if anything. Are you still planning on being immortal?
Brad Guth
> TERRESTRIAL LASER SUSTAINED DETONATION RETURN OF FREE FLYING ASTEROIDS
>
> Propulsive effects have been created on unprepared surfaces by laser
> action. This technique will be developed to mine the asteroid belt
> using entirely terrestrial based systems.
>
> Low cost terrestrial solar panels creating hydrogen and oxygen from a
> water store, being fed into a super-critical boiler that drives a
> 1,000 MW steam turbine system, which then drives a 300 MW high-
> efficiency laser, operating in the Atacama desert. This region is
> chosen for its high level of solar insolation, its broad view of the
> sky, and its very low dispersion due to atmospheric effects. With
> adequate optics a 300 MW laser will beam energy across the solar
> system and produce propulsive effects on unconfined surfaces of
> objects as far away as Ceres;
>
>
> We can also work with optical and IR telescopes that operate in
> conjunction with laser pulses that cause asteroids to eject plasma
> that are then characterized spectroscopically.
>
> In this way a terrestrial solar powered center can survey all the
> small bodies of the solar system, first identifying them with optical,
> microwave, radar and lidar system - which then intensify the
> characterization by creating plasma plumes on them to read their
> composition. Then, on selected bodies more powerful structured pulses
> modify their rotation and movement through space, giving us detailed
> information on their density, moment of inertia and so forth.
> Finally, for those rich bodies that interest us, we carve off pieces
> if needed, and bring them back to Earth by directing structured plasma
> plumes to impart the desired thrust effects. For 10 tonne to 100
> tonne per day return rate, we can land the materials in a return zone
> in the Atacama desert for recovery.
>
> The revenue from this operation will then be used to expand the system
> with space based lasers to enlarge the return rate, resulting in
> ultimately, the entire industrial infrastructure of humanity being
> placed on orbit powered by sunlight and fed from raw materials
> recovered from the asteroids.
How is that 10,000 page book "Mook, the new wizard of Oz" coming
along?
willia...@mokenergy.com
I'm staying in New Zealand until the the 160 bombs go off in America.
willia...@mokenergy.com
> http://www.youtube.com/watch?v=EUBf31VjfGE
>
> I'm staying in New Zealand until the the 160 bombs go off in America.
willia...@mokenergy.com
>
> I'm staying in New Zealand until the the 160 bombs go off in America.
If the Springfield, Cincinatti area is still surviving, I will then
set up productive capabilities in these areas
http://www.scribd.com/doc/43547498/Mok-Capabilities
And produce MEMS based tracking systems;
http://www.scribd.com/doc/43547079/Pv-Hte-Sketch
Which lead to MEMS based self-replicating machine systems
http://en.wikipedia.org/wiki/Self-replicating_machine
Which will create a global infrastructure...
willia...@mokenergy.com
>
> I'm staying in New Zealand until the the 160 bombs go off in America.
De-minimus. 1% of the world's land area, cultivated processed and
transported by smart smoke.
Food Pounds/Acre lbs/Per Ppl/Acre Acres Total Sq Miles
Spinach 11,000 2.2 5,000 1,400,000 2,188
Carrot 19,400 18.2 1,066 6,567,010 10,261
Onion 19,800 20.0 990 7,070,707 11,048
Squash 17,000 4.8 3,512 1,992,941 3,114
Potato 15,200 52.5 290 24,177,632 37,778
Celery 32,000 10.0 3,204 2,184,875 3,414
Cabbage 13,700 12.4 1,104 6,339,854 9,906
Tomato 11,000 74.4 148 47,345,455 73,977
bean 4,600 5.4 853 8,202,174 12,816
Lettuce 9,100 24.3 374 18,700,000 29,219
Turnip 12,000 6.2 1,941 3,606,167 5,635
Broccoli 7,300 7.6 965 7,256,986 11,339
Cauliflow 10,800 6.5 1,664 4,206,481 6,573
Bell Pep 6,900 7.1 971 7,208,986 11,264
Potato 6,000 1.3 4,471 1,565,667 2,446
Corn 6,200 9.5 649 10,780,000 16,844
Squash 9,700 5.0 1,951 3,588,041 5,606
Beet 10,800 0.8 13,636 513,333 802
Cantalou 9,800 6.2 1,574 4,447,143 6,949
Peas 2,200 5.3 413 16,940,000 26,469
Asparag 4,400 0.9 4,762 1,470,000 2,297
Cucumbe 8,400 9.3 901 7,773,333 12,146
Radish 12,000 1.3 9,569 731,500 1,143
Watermel 10,300 9.7 1,064 6,578,641 10,279
Bean-Lim 1,400 2.0 707 9,900,000 15,469
Almonds, 1,150 1.0 1,150 6,086,957 9,511
Pecans, 962 0.5 2,004 3,492,723 5,457
Pistachi 1,490 0.2 6,478 1,080,537 1,688
Walnuts 2,760 0.5 5,520 1,268,116 1,981
Apples 25,400 30.6 829 8,445,748 13,196
Apricots 11,260 0.4 26,938 259,858 406
Avocados 3,680 0.9 4,289 1,632,065 2,550
Cherries 6,660 1.0 6,441 1,086,787 1,698
Dates 9,420 3.2 2,913 2,403,185 3,755
Figs 4,620 0.9 5,000 1,400,000 2,188
Kiwi Fruit 7,860 1.3 6,160 1,136,387 1,776
Nectarine 16,080 2.5 6,411 1,091,791 1,706
Olives 4,320 0.1 65,455 106,944 167
Peaches 31,400 4.2 7,395 946,561 1,479
Pears 28,200 5.4 5,232 1,337,943 2,091
Pears Ba 31,600 7.0 4,489 1,559,494 2,437
Plums 10,400 2.6 4,040 1,732,500 2,707
Grapes 14,770 13.4 1,102 6,349,763 9,922
Oats 2,240 7.3 309 22,687,500 35,449
Barley 2,833 0.3 10,731 652,312 1,019
Win Whe 4,800 73.0 66 106,458,333 166,341
D Wheat 6,300 73.0 86 81,111,111 126,736
Rice 20,485 17.6 1,164 6,014,157 9,397
TOTALS 10,827 549.9 9,555 468,887,697 732,637
Brad Guth
But you can't even make good on your $100/tonne hydrogen that's
created entirely off-grid.
Why don't you just accomplish the Mokenergy hydrogen thing?
It seems by just not having your cheap and renewable hydrogen is by
itself a tragedy for humanity and the environment. So what's the big
holdup this time?
Ever heard of DARPA(E)?
Robert Clark
> On Mar 7, 1:52 am, Robert Clark <rgregorycl...@yahoo.com> wrote:
>
> > of the Direct team's Jupiter-246 vehicle also would become SSTO when
> > switched to a SSME engine. I guess I should not have been surprised by
> > this. The thesis I have been arguing repeatedly via email with
> > individuals in NASA and the industry and on space oriented forums such
> > as this one is that if you use BOTH the most weight optimized designs
> > AND the highest efficiency engines available, then what you will wind
> > up with will be SSTO capable whether you intend it to or not.
> > By highest efficiency engines I don't mean just an engine optimized
> > to have a high vacuum Isp only. I mean an engine of highest efficiency
> > over the entire flight range to orbit. For hydrogen engines that is
> > the SSME, and the Russian analogue RD-0120.
> >...
>
The previous posts in this thread were about using liquid-fueled only
stages. However, the shuttle-derived heavy lift vehicles proposed that
use SRB's can also reduce costs by being made fully reusable.
Some have soured on the idea of reusability because of the case of the
space shuttle. But the problem with the shuttle was that that
spacecraft that had to be carried to orbit was so heavy. It was nearly
four times the weight of the payload you could carry. I quoted before
a statement by Robert Zubrin in one of his books that emphasized this
point, which he argued contributed to the shuttle being a fiscal
disaster:
Newsgroups: sci.space.policy, sci.astro, sci.physics,
sci.space.history
From: Robert Clark <rgregorycl...@yahoo.com>
Date: Wed, 10 Feb 2010 22:04:01 -0800 (PST)
Subject: Re: A kerosene-fueled X-33 as a single stage to orbit
vehicle.
http://groups.google.com/group/sci.space.policy/msg/55d37c1a87c77c1e?hl=en
However actually for most launch vehicles the upper stage that
actually makes it to orbit along with the payload is usually
comparatively small, in fact, frequently smaller than the payload. And
this gets even better the larger the launcher gets. Note then for a
large launcher since the dry mass of the upper stage will be a
fraction of the payload mass and the reentry/landing systems will be a
fraction of the upper stage dry mass, the extra weight to make the
upper stage reusable is actually a small fraction of the payload mass,
so will only subtract a small amount from the payload.
This point is well illustrated by the DIRECT Jupiter-246 heavy lift
launcher. See the specifications in the diagram linked to below.
Note that the dry mass of the upper stage is less than 12,000 kg. But
the payload mass is in the range of 105,000 kg to 117,000 kg. However,
the extra mass for reentry/landing systems for an orbital stage is
commonly estimated to total about 28% of the dry mass of the stage
(see below), which is about 3,400 kg for this upper stage. This would
subtract off a comparatively small amount from the payload mass.
However, the biggest cost saving would not be in making the upper
stage reusable but in the reusability of the expensive core stage with
its 4 SSME engines and shuttle ET-derived propellant tank. This has a
dry mass of 66,895 kg. So if we used the 28% estimate for reentry/
landing systems this would be an extra 18,730 kg added to this stage
weight. So it would conceivable subtract this amount from the payload
weight.
But there are two key reasons why it will likely not have to be this
high an amount that has to be subtracted off from the payload weight.
First, another point Zubrin makes in that passage I quoted from his
book _Entering Space_ is that for a first stage every extra kilo added
to the first stage weight generally will only subtract about .1 of a
kilo from the payload weight. However, this core stage is not quite a
first stage; it's closer to being a second stage. The amount of
payload that has to be subtracted off will be somewhat more than
1/10th though not the full amount of this extra weight, depending on
how much delta-V this stage makes up.
The second key reason is that this core stage will not have to reach
all the way to orbit so its reentry regime will not be as severe as
for an orbital stage, so the reentry systems not as heavy. To see why,
notice that unlike the shuttle ET, this ET-syle propellant tank will
be carrying the 200,000 kg gross weight of the upper stage plus the
ca. 100,000 kg payload, much more mass to loft before staging than for
the shuttle. So it will reach significantly lower velocity.
For the 28% of the landing mass for reentry/landing systems, first
Robert Zubrin gives an estimate of about 15% for reentry thermal
protection:
Reentry heat shields.
http://en.wikipedia.org/wiki/Reusable_launch_system#Reentry_heat_shields
Secondly, in a discussion between Henry Spencer and Mitchell Burnside
Clapp on the relative benefits of horizontal vs. vertical landing, the
extra mass for winged landing or a powered descent is about 10%:
Horizontal vs. vertical landing (Henry Spencer; Mitchell Burnside
Clapp)
http://yarchive.net/space/launchers/horizontal_vs_vertical_landing.html
Finally, in another discussion on Yarchive.net/space, the landing gear
weight is given as about 3%:
Landing gear weight (Gary Hudson; George Herbert; Henry Spencer).
http://yarchive.net/space/launchers/landing_gear_weight.html
However, note with modern materials quite likely this 28% estimate for
the reentry/landing systems can be cut in half.
Bob Clark
Jupiter-246 Heavy - Lunar EDS Launch Vehicle Configuration.
http://www.launchcomplexmodels.com/Direct/documents/Baseball_Cards/J246H-41.5004.08001_EDS_090608.jpg
Robert Clark
launcher as an interim solution to developing a heavy lift vehicle.
This would use two 4-segment SRM's as does the shuttle and an ET. But
it would not have a shuttle orbiter, nor would this Phase I vehicle
have an upper stage:
SLS planning focuses on dual phase approach opening with SD HLV.
April 25th, 2011 by Chris Bergin
http://www.nasaspaceflight.com/2011/04/sls-planning-dual-phase-approach-opening-sd-hlv/
However, built into this plan is that at most 4 flights of this
vehicle will be made before it is discontinued in favor of a more
expensive, 130 mt payload upgrade. These 4 flights are to regarded as
"test flights" according to the Bergin article. They will use 3 SSME's
at a time and only 12 of those will be available including those taken
from the retired space shuttles, thus allowing only 4 flights.
Presumably after that either the production of new SSME's will be
started or their expendable versions will be, or NASA will choose
instead to use kerosene fueled engines for the core stage.
However, a better plan in my view would be to explore methods in
which this Phase I vehicle could be reusable. Then this low cost HLV
could have many more missions as well as cutting costs in being
reusable. This would give you more options as to when and if the more
expensive vehicle needed to be developed.
Many at NASA are not are favorably inclined towards reusable systems
because of the experience of the shuttle. However, as I mentioned
before in the post below a key reason for why the shuttle was not
economical would not hold in this case: it would not have to carry the
80 mt orbiter that took out most of the vehicles payload capacity.
Another reason why a reusable vehicle could be done better now is
because of the research that has already been done to address the
failings of the shuttle system. For instance, for the X-33/VentureStar
program the metallic shingles to be used for thermal protection have
confirmed in testing they would require less maintenance than the
ceramic tiles of the shuttle.
The advanced ceramics used on the Air Force's X-37B were also
expected to cut maintenance on thermal protection. It would be useful
to find out if they have been successful in that regard.
The X-37B may also serve as a good model to use for the reentry
system for the ET tank to be used on the Phase I vehicle. Note that
the X-37B's short stubby wings are much smaller in proportion to the
size of the vehicle than those of the space shuttle. That and the
composite materials used for the wings would result in much reduced
mass used for the wings for the ET tank.
Other lightweight reentry systems would be the German IXV program
which does not use wings:
IXV Program Aims to Put ESA at Cutting Edge of Re-entry Technology
Posted by Doug Messier on September 18, 2010, at 4:08 am in ESA.
http://www.parabolicarc.com/2010/09/18/ixv-program-aims-put-esa-cutting-edge-reentry-technology/
and the inflatable one NASA is investigating:
NASA Successfully Tests Vacuum-Packed Inflatable Heat Shield.
A vacuum-packed inflatable shroud could enable future spacecraft
reentry on both Earth and Mars.
By Jeremy HsuPosted 08.17.2009 at 3:00 pm
http://www.popsci.com/military-aviation-amp-space/article/2009-08/nasa-puts-inflatable-heat-shield-flight-test
Bob Clark
On Apr 4, 2:20 pm, Robert Clark <rgregorycl...@yahoo.com> wrote:
> On Mar 12, 12:34 pm, Robert Clark <rgregorycl...@yahoo.com> wrote:
> ...
> vehicle.http://groups.google.com/group/sci.space.policy/msg/55d37c1a87c77c1e?...
> Reentry heat shields.http://en.wikipedia.org/wiki/Reusable_launch_system#Reentry_heat_shields
>
> Secondly, in a discussion between Henry Spencer and Mitchell Burnside
> Clapp on the relative benefits of horizontal vs. vertical landing, the
> extra mass for winged landing or a powered descent is about 10%:
>
> Horizontal vs. vertical landing (Henry Spencer; Mitchell Burnside
> Clapp)http://yarchive.net/space/launchers/horizontal_vs_vertical_landing.html
>
> Finally, in another discussion on Yarchive.net/space, the landing gear
> weight is given as about 3%:
>
> Landing gear weight (Gary Hudson; George Herbert; Henry Spencer).http://yarchive.net/space/launchers/landing_gear_weight.html
>
> However, note with modern materials quite likely this 28% estimate for
> the reentry/landing systems can be cut in half.
>
> Bob Clark
>
> Jupiter-246 Heavy - Lunar EDS Launch Vehicle Configuration.http://www.launchcomplexmodels.com/Direct/documents/Baseball_Cards/J2...
Robert Clark
> NASA appears to be leaning to a 70 mt payload shuttle-derived
> launcher as an interim solution to developing a heavy lift vehicle.
> This would use two 4-segment SRM's as does the shuttle and an ET. But
> it would not have a shuttle orbiter, nor would this Phase I vehicle
> have an upper stage:
>
> SLS planning focuses on dual phase approach opening with SD HLV.
>
> and the inflatable one NASA is investigating:
>
> NASA Successfully Tests Vacuum-Packed Inflatable Heat Shield.
> A vacuum-packed inflatable shroud could enable future spacecraft
> reentry on both Earth and Mars.
>
The above discusses that maintenance costs for thermal protection
should be significantly less than for the space shuttle. But another
significant recurring cost for the shuttle program was for maintenance
on the engines. Now, the SSME's have to be overhauled after every
flight, costing ten's of millions of dollars. However, Henry Spencer a
highly regarded expert on the history of space flight has said
Rocketdyne studies show that with a lot of work to upgrade it,
maintenance could be reduced to $750K per flight per engine:
Engine reusability (Henry Spencer)
http://yarchive.net/space/rocket/engine_reusability.html
Spencer here said this would not be satisfactory for really large
reductions in space costs. But this would be a reduction in SSME
maintenance costs by 1 to 2 orders of magnitude, a major reduction in
the costs for using the engine. A key question though is how much
would be the cost to make the necessary upgrades to the engine.
I also did not estimate the extra mass of the reentry/landing
systems. Here's a diagram showing the specifications for the DIRECT
team's version of this Phase I ca. 70 mt launcher:
DIRECTv3 Jupiter-130 - LEO Cargo Launch Vehicle Configuration.
http://www.launchcomplexmodels.com/Direct/documents/Baseball_Cards/J130-41.4000.10051_CaLV_30x130nmi_29.0deg_090606.jpg
The dry mass of the core stage is given as 63.7 mt. This mass can be
reduced by going to a common bulkhead design for the propellant tanks
that for instance SpaceX was able to use to reduce dry mass for its
Falcon vehicles. The intertank on the shuttle ET actually weighs more
than the oxygen tank:
External Tank.
http://science.ksc.nasa.gov/shuttle/technology/sts-newsref/et.html
Going to a common bulkhead design would eliminate this mass, reducing
the dry mass by about 5 mt. Also recent research has shown that dry
mass of rocket vehicles in general can be reduced by 10% to 20%. This
would take off about another 5 mt to 10 mt.
I gave an estimate before in this thread of about 28% of the dry mass
for reentry/landing systems. However, as I said probably with modern
materials we can cut this in half. Then with all these reductions
together the extra mass for reentry/landing systems might only be in
the range of 7,000 kg. So we would still maintain 90% of the payload
mass while gaining reusability and a longer useful life for this low
cost heavy lift launcher.
Bob Clark
Robert Clark
> NASA appears to be leaning to a 70 mt payload shuttle-derived
> launcher as an interim solution to developing a heavy lift vehicle.
> This would use two 4-segment SRM's as does the shuttle and an ET. But
> it would not have a shuttle orbiter, nor would this Phase I vehicle
> have an upper stage:
>
> SLS planning focuses on dual phase approach opening with SD HLV.
>
> However, built into this plan is that at most 4 flights of this
> vehicle will be made before it is discontinued in favor of a more
> expensive, 130 mt payload upgrade. These 4 flights are to regarded as
> "test flights" according to the Bergin article. They will use 3 SSME's
> at a time and only 12 of those will be available including those taken
> from the retired space shuttles, thus allowing only 4 flights.
> Presumably after that either the production of new SSME's will be
> started or their expendable versions will be, or NASA will choose
> instead to use kerosene fueled engines for the core stage.
> However, a better plan in my view would be to explore methods in
> which this Phase I vehicle could be reusable. Then this low cost HLV
> could have many more missions as well as cutting costs in being
> reusable. This would give you more options as to when and if the more
> expensive vehicle needed to be developed.
This is for the interim, Phase I, 70 mt launcher. This is to use two
SRB's and an external tank as with the shuttle system, but no orbiter
and no upper stage. However, the possibilities become especially
interesting when we look at the case of making the Phase II, 100+ mt
payload launcher reusable. This vehicle will have an additional upper
stage. This has a significant advantage for the lightness of the
reentry/landing systems in that only the upper stage at a small dry
mass would have to have the full reentry systems of an orbiting
vehicle. The upper stage of the DIRECT teams's Jupiter-246 for
instance weighs less than 12,000 kg. Also, for this case the ET would
reach a much reduced velocity and would not reach orbit so its reentry
systems would be much simpler and lighter.[1]
To get the full benefits of reusability we'll switch out the RL-10's
or J-2X engines used on the upper stage for SSME(s). This does have a
problem though in that the SSME would have to be made air startable.
The benefits for reusability are so significant that costs estimates
for this upgrade should be made.
However, a different potential solution would also reap additional
benefits. If instead of placing this stage atop the ET tank, we put it
in parallel with it, then the stage could also be started on the
ground.
This has a benefit because now we could use cross-feed fueling
between the ET and upper stage tanks. Cross-feed fueling with parallel
staging is known to be able to increase your payload. For instance by
using it for their Falcon Heavy vehicle SpaceX was able to increase
its payload by 50%.
Note also that we wouldn't have the development cost for a new 4 SSME
engine core stage, as is currently planned for the Phase II vehicle.
We would use the same 3-engine core stage as used for the Phase I
vehicle. The extra thrust for the Phase II vehicle would come from the
upper stage now firing in parallel from the start.
However, another potentially game changing effect of doing this is
that if you look at the mass ratio of this reconfigured upper stage
with SSME(s) you see it has SSTO capability. This is because it has
the weight optimization of an upper stage and now using an engine
optimized to be most efficient during the entire flight to orbit it
can reach orbit in a single stage with significant payload.[2]
In fact not just the Jupiter-246 upper stage would have this
capability, but in fact the Ariane 5's upper stage, the Apollo's S-II
and S-IVB, and the planned Ares I upper stage would as well if
switched out to use SSME(s).
This is important because we will have already existing stages as
well as the engines to make at least an initial version of this upper
stage. This means we could have a significant cost reduction on an
initial version of the upper stage.
Another very key fact is because this upper stage can be used as a
separate launcher and even manned launcher, thus with its own market,
we could initiate it's development and production in parallel to the
low cost Phase I vehicle. So we would get in fact not only a 70+ mt
vehicle, but we would get the 100+ mt launcher and a manned launcher
in just the same short time frame of the Phase I launcher and at a
smaller cost than now planned for the Phase II, 100+ mt launcher.
Indeed because there would be such a significant market for this
manned SSTO vehicle, NASA might not have to pay for its development at
all.
Bob Clark
1.)Newsgroups: sci.space.policy, sci.astro, sci.physics,
sci.space.history
From: Robert Clark <rgregorycl...@yahoo.com>
Date: Mon, 4 Apr 2011 11:20:23 -0700 (PDT)
Local: Mon, Apr 4 2011 2:20 pm
Subject: Re: Some proposals for low cost heavy lift launchers.
http://groups.google.com/group/sci.space.policy/msg/126d7a70c08c32e3?hl=en
2.)Newsgroups: sci.space.policy, sci.astro, sci.physics,
sci.space.history
From: Robert Clark <rgregorycl...@yahoo.com>
Date: Sun, 6 Mar 2011 22:52:19 -0800 (PST)
Local: Mon, Mar 7 2011 2:52 am
Subject: Re: Some proposals for low cost heavy lift launchers.
http://groups.google.com/group/sci.space.policy/tree/browse_frm/thread/449ba3d5c3a04e59/444441f02bd4dd60?rnum=31&_done=%2Fgroup%2Fsci.space.policy%2Fbrowse_frm%2Fthread%2F449ba3d5c3a04e59%3Fscoring%3Dd%26&scoring=d#doc_97639eaf0b8ac44d
Robert Clark
> its payload by 50%.
> Note also that we wouldn't have the developmentcostfor a new 4 SSME
> engine core stage, as is currently planned for the Phase II vehicle.
> We would use the same 3-engine core stage as used for the Phase I
> vehicle. The extra thrust for the Phase II vehicle would come from the
> upper stage now firing in parallel from the start.
> However, another potentially game changing effect of doing this is
> that if you look at the mass ratio of this reconfigured upper stage
> with SSME(s) you see it has SSTO capability. This is because it has
> the weight optimization of an upper stage and now using an engine
> optimized to be most efficient during the entire flight to orbit it
> can reach orbit in a single stage with significant payload.[2]
> In fact not just the Jupiter-246 upper stage would have this
> capability, but in fact the Ariane 5's upper stage, the Apollo's S-II
> and S-IVB, and the planned Ares I upper stage would as well if
> switched out to use SSME(s).
> This is important because we will have already existing stages as
> well as the engines to make at least an initial version of this upper
> initial version of the upper stage.
> Another very key fact is because this upper stage can be used as a
> separate launcher and even manned launcher, thus with its own market,
> vehicle, but we would get the 100+ mt launcher and a manned launcher
> in just the same short time frame of the Phase I launcher and at a
> Indeed because there would be such a significant market for this
> manned SSTO vehicle, NASA might not have to pay for its development at
> all.
Nice article here arguing in favor of NASA promoting small manned
commercial vehicles:
Human spaceflight for less: the case for smaller launch vehicles,
revisited.
by Grant Bonin
Monday, June 6, 2011
http://www.thespacereview.com/article/1861/1
In the comments section, I commented that the capability to produce
such small, low cost, manned vehicles exists now. I estimated the cost
for a such a reusable vehicle in the range of a few tens of millions
of dollars unit cost, comparable to a medium sized business jet.
Bob Clark
Robert Clark
> On May 8, 9:20 pm, Robert Clark <rgregorycl...@yahoo.com> wrote:
>...
> commercial vehicles:
>
> Human spaceflight for less: the case for smaller launch vehicles,
> revisited.
> by Grant Bonin
> Monday, June 6, 2011
> http://www.thespacereview.com/article/1861/1
>
> In the comments section, I commented that the capability to produce
> such small, low cost, manned vehicles exists now. I estimated the cost
> for a such a reusable vehicle in the range of a few tens of millions
> of dollars unit cost, comparable to a medium sized business jet.
>
> Bob Clark
Just saw this on Hobbyspace.com:
Boeing proposes SSTO system for AF RBS program.
"The new issue of Aviation Week has a brief blurb about a Boeing
proposal for the Air Force's Reusable Booster System (RBS) program:
Boeing Offers AFRL Reusable Booster Proposal - AvWeek - June.13.11
(subscription required).
Darryl Davis, who leads Boeing's Phantom Works, tells AvWeek that they
are proposing a 3-4 year technology readiness assessment that would
lead up to a demonstration of a X-37B type of system but would be
smaller. Wind tunnel tests have been completed. Davis says the system
would be a single stage capable of reaching low Earth orbit and, with
a booster, higher orbits. The system would return to Earth as a
glider.
Davis says "that advances in lightweight composites warrant another
look" at single-stage-to-orbit launchers."
http://www.hobbyspace.com/nucleus/index.php?itemid=30110
I don't have a subscription to AV Week. If anyone does perhaps they
could look this up.
I'm curious about the statement it would be "smaller" than the X-37B.
I did some preliminary calculations that if you switched to kerosene
fuel and a high efficiency engine such as the NK-33, and filled every
scrap of internal volume with propellant, then a vehicle twice the
size of the X-37B could be SSTO. I would be surprised they are able to
get it to work with a smaller vehicle than the X-37B.
Perhaps they mean it would be smaller than the booster, Atlas V, and
X-37B system, as the Atlas V weighs upwards of 300,000 kg.
Bob Clark
Brad Guth
So what don't you like about Mokareospace?
Robert Clark
> On Jun 13, 4:53 am, Robert Clark <rgregorycl...@yahoo.com> wrote:
Any of these proposals discussed on these forums would have to be
initiated by NASA or the "NewSpace" or "Old Space" aerospace
companies. They would involve multi-million dollar investments. In
order for the proposals to be taken seriously they would have to be
either presented in a peer reviewed journal or presented at one of the
aerospace conferences.
This second might be the better way to go since you can defend your
ideas in person with people in the industry who would be making the
decisions on the feasibility of an idea, and also you wouldn't have to
depend on the vagaries of the journal publication process. Another
benefit is that you would have it at least published in the conference
proceedings.
Another avenue rather surprisingly is increasing in influence. That's
in writing your own blog. That might be especially useful for Mook
because of the amount of detail in his postings.
Blogs have been criticized in the news media in a variety of realms
most notably in the realm of politics because anyone can create their
own blog. Still these blogs do have their influence in politics and we
have seen also in the field of space policy.
Bob Clark
Brad Guth
I can 100% agree with that. However, our William Mook (aka
Mokaerospace and Mokenergy) seems to be a rogue wizard that can't
quite put it all together without looking a bit overly eccentric and
kind of a cranky wizard of Oz demanding. Mook certainly means well
and isn't the least bit dumb or unqualified, but he has been burned so
often that his battery of loose cannons only manage to shoot blanks at
those trying to terminate everything he has to offer.
Mook needs a lot more fire power, as well as bipolar disorder
medication from time to time, because his research and subsequent
ideas are sufficiently terrific, not that I agree with everything
that's served on his table.
Robert Clark
I have been advised to open up a blog, not always positively, so I have.
My first post is on super heavy lift vehicles:
Low Cost HLV.
http://exoscientist.blogspot.com/2012/05/low-cost-hlv.html
Suggestions on the post and on improving the blog are invited.
Bob Clark
Robert Clark
> I have been advised to open up a blog, not always positively, so I have.
> My first post is on super heavy lift vehicles:
>
> Suggestions on the post and on improving the blog are invited.
>
> Bob Clark
will make space flight routine. Then we will need low cost heavy lift
and low cost manned flight. This other blog post shows how you can get
a manned flight to the Moon at a cost in the few hundred million
dollar range, in contrast to the $100 billion proposed by NASA:
SpaceX Dragon spacecraft for low cost trips to the Moon.
http://exoscientist.blogspot.com/2012/05/spacex-dragon-spacecraft-for-low-cost.html
Comments on the blog post and on improving the blog are invited.
Bob Clark
Robert Clark
> On May 2, 2:11 am, Robert Clark <rgregorycl...@yahoo.com> wrote:
>
> > I have been advised to open up a blog, not always positively, so I have.
> > My first post is on super heavy lift vehicles:
>
> >LowCostHLV.
> >http://exoscientist.blogspot.com/2012/05/low-cost-hlv.html
>
> > Suggestions on the post and on improving the blog are invited.
>
> > Bob Clark
>
> I'm of the opinion that off-world mining is the "killer app" that
> will make space flight routine. Then we will needlowcostheavy lift
> a manned flight to the Moon at acostin the few hundred million
>
> Comments on the blog post and on improving the blog are invited.
>
> Bob Clark
HLV" post
is feasible by making a comparison to the Saturn S-IC stage:
MONDAY, MAY 7, 2012
Low Cost HLV, page 2: Comparison to the S-IC Stage.
http://exoscientist.blogspot.com/2012/05/low-cost-hlv-page-2.html
This blog post shows you can get an even higher mass ratio, close to
29 to 1,
by using modern materials and common bulkhead design. A mass ratio
this high
makes possible a SSTO:
THURSDAY, JUNE 21, 2012
Low Cost HLV, page 3: Lightweighting the S-IC Stage.
http://exoscientist.blogspot.com/2012/06/low-cost-hlv-page-3-lightweighting-s-ic.html
Bob Clark
mokme...@gmail.com
C + 2 H2O2 --> CO2 + 2 H2O
12 + 2 x 34 = 44 + 2 x 18
So, for each 150 kg of carbon you have 850 kg of H2O2 for every tonne of propellant. The graphite occupies 66.7 liters whilst the hydrogen-peroxide occupies 586.2 liters. 652.9 liters per metric ton of propellant. A specific gravity of 1.53 kg/liter.
The carbon is The exhaust velocity of this system is 3.3 km/sec. RocketLab has perfected a colloidal silver liquid that acts as a simple method of detonation for this mixture much more efficiently than silver mesh.
http://www.rocketlab.co.nz/propulsion/high-density-monopropellant/
At these densities, and temperatures, and pressures, 3% structural fractions are possible. A 7 element system with each element 42 feet long and 7 feet in diameter and equipped with cross-feeding with 4 elements draining at lift off, 2 elements draining for second stage, and 1 element for third stage - puts up 10 tons.
Brad Guth
obfuscation. What you propose here as using H2O2 and pure carbon
sounds like a perfectly good package deal, though I doubt most will
pay any attention as long as it was your idea or that of anything I
could support.
mokme...@gmail.com
RocketLab has already mixed up this fuel and fired it and gotten good results.
The thick liquid doesn't slosh and is highly storable.
The eight foot diameter and forty-two foot long fiberglass tank coated with foam, looks a lot like a miniature space shuttle external tank. The fuel is pressurized with internal bladder and refrigerated. The bladder keeps things flowing in the right direction. Eight auger-like pumps produce the right mass flow and pressure for the simple annular nozzle array feeding an aerospike engine with eight throttable elements which controls flight direction.
The 5,700 lb system carries 163,000 lbs of propellant.
With a 9,185 ft/sec exhaust speed at lift off rising to 10,825 ft/sec at 60,000 ft and above, the single stage has a terminal velocity of 30,000 ft/sec - which is sufficient to put it into orbit without payload.
A 20,000 lb payload is capable of being put into orbit with this system with three similar launch elements. Two outboard elements feed all three engines operating at the base of each of the three flight elements. Each outboard element is dropped when empty, and the central element continues to orbit with its payload.
A dozen satellites launched into a sun-synchronous polar orbit from New Zealand provide global internet/GPS services with recovery of the launch elements in the Southern Sea off the coast of Antarctica.
A launcher consisting of seven elements just described, with four elements forming the first stage, two the second stage, one the third stage, puts up 50,000 lbs into the same orbit.
This same propellant powering stages of this mass are capable of putting 5,950 lbs of the 20,000 lb satellite or 14,890 lbs of the 50,000 lb satellite, into a trans-lunar trajectory.
If we imagine these 5,950 lb and 14,890 lb payloads were a stage equipped to land on the moon and take off and return to Earth, we have 1,025 lbs and 2,560 lbs respectively for each of these stages.
Now a long-duration mechanical counter-pressure suit
http://dspace.mit.edu/handle/1721.1/63033
http://www.space.com/728-high-tech-spacesuits-eyed-extreme-exploration.html
is very lightweight, and dispenses with the need for cabins and things of that nature, especially with heads up displays and data gloves.
Also, tiny mechanisms (fuel cells, life support systems) made in quantity are highly reliable and very lightweight and capable.
http://www.youtube.com/watch?v=VxSs1kGZQqc
The spacesuit is also equipped for re-entry.
http://www.youtube.com/watch?v=V6xqgWVgAok
Basically, three people wearing this gear and massing on average 341 lbs each are strapped to the smaller lander and eleven people similarly equipped are strapped to the larger lander.
To build the flight system requires that $100 million to get to the first flight system and $200 million to get to the second flight system. So, this is $33.3 million and $18.2 million per person for the first flight.
I offer the first 33 flights to the moon for $330 million to a limited number of resellers ($66 million minimum investment).
I offer thirty flights to the moon for tourists using this system at $25 million each after funding is achieved. I will accept deposits which will be placed on deposit with a partner bank as Tier One Capital and earn interest during the wait period until the service is delivered - which will take up to three years for delivery.
After each flight tourists get to keep their suits and any of up to 20 lbs of materials that they can stow in their carrier bag. Professionally produced videos of their flights as well as all training prior to flight is included. Finally, for those who do something historic, special events will be arranged. (First woman on the moon, first X-national on the moon, etc.)
mokme...@gmail.com
20,000 lbs payload provides 5 layers of 11 passengers each, plus 8 crew. So, resellers take note. In addition to the eleven lunar flights, there are 44 orbital flights - all with re-entry - which is the first 'manned' three unit flight. When these seats are sold for $2.5 million each, the cost of the lunar flight is covered - for the reseller.
Of course a reseller can take a trip form themselves and their loved ones, and resell extra seats to offset costs or eliminate them entirely.
Of course resellers take greater risk than tourists.
We will also need pilots and crew as well.
Brad Guth
messages seems to be working, but the actual context of a given reply
isn't displayed (as though it has been intentionally moderated as
something nondisclosure rated so that we can't read it)
Brad Guth
> is very lightweight, and dispenses with the need for cabins and things of that nature, especially with heads up displays and data gloves.
>
> Also, tiny mechanisms (fuel cells, life support systems) made in quantity are highly reliable and very lightweight and capable.
>
> http://www.youtube.com/watch?v=VxSs1kGZQqc
>
> The spacesuit is also equipped for re-entry.
>
> http://www.youtube.com/watch?v=V6xqgWVgAok
>
> Basically, three people wearing this gear and massing on average 341 lbs each are strapped to the smaller lander and eleven people similarly equipped are strapped to the larger lander.
>
> To build the flight system requires that $100 million to get to the first flight system and $200 million to get to the second flight system. So, this is $33.3 million and $18.2 million per person for the first flight.
>
> I offer the first 33 flights to the moon for $330 million to a limited number of resellers ($66 million minimum investment).
>
> I offer thirty flights to the moon for tourists using this system at $25 million each after funding is achieved. I will accept deposits which will be placed on deposit with a partner bank as Tier One Capital and earn interest during the wait period until the service is delivered - which will take up to three years for delivery.
>
> After each flight tourists get to keep their suits and any of up to 20 lbs of materials that they can stow in their carrier bag. Professionally produced videos of their flights as well as all training prior to flight is included. Finally, for those who do something historic, special events will be arranged. (First woman on the moon, first X-national on the moon, etc.)
form of H2O2 that even Mokenergy should be able to provide at less
cost than most any other alternative.