Modest Proposal - Common Interplanetary Booster
Williamknowsbest
I want to talk first about the M1
Its a 1.5 million to 1.8 million lbf rocket engine developed the US
Army/Air Force back in the day, and turned over to NASA in 1960.
http://en.wikipedia.org/wiki/M-1_(rocket_engine)
Then, there was the J2 rocket engine with 200,000 lbf to 230,000 lbf
http://en.wikipedia.org/wiki/J-2_(rocket_engine)
This was used on the S-II and S-IVB stages of the Saturn V
moonrocket. The S-II was a 1,0060,000 lb mass system and the S-IVB
was a 253,000 lb mass system.
http://en.wikipedia.org/wiki/S-II
http://en.wikipedia.org/wiki/S-IVB
Special mention should be made of the S-IV's original configuration -
with 6 RL-10 engines. The RL-10 is a deeply throttable engine - and
restartable - perfect for a high performance lunar landing vehicle
http://en.wikipedia.org/wiki/S-IV
http://en.wikipedia.org/wiki/RL-10
Now I also want to discuss a little bit, the aerospike engine. This
is an inside out nozzle arrangement that allows any engine pumpset to
operate in a wide range of pressure conditions.
In fact aerospike engines have been produced using existing pumpsets
http://en.wikipedia.org/wiki/Image:Annular-Aerospike.jpg
Finally, there are innovations that were developed by legendary
aerospace engineering pioneer, Phillip Bono
http://en.wikipedia.org/wiki/Philip_Bono
http://www.google.com/patents?id=CpVzAAAAEBAJ&printsec=abstract&zoom=4
Here we have a spacecraft that launches vertically and re-enters tail
first, and lands vertically under rocket power using an aerospike
engine, a method very similar to the DC-X and Delta Clipper designs 35
years later (but without the altitude compensating nozzle)
http://en.wikipedia.org/wiki/Delta_Clipper
So, here's the deal,
Three elements, built around 7 M-1 pumpsets, into a single large
annular aerospike engine. Each element produces 12.6 million pounds
of thrust and masses 9.7 million pounds fully loaded and 1.2 million
pounds empty. These three elements are lashed together like a Delta
IV Heavy, but the two outboard elements are equipped to feed
propellant to the core stage, while the entire system lifts off.
Thus all engines fire at lift off, which is a good thing, and the two
outboard elements are drained forming in effect a first stage.
Furthermore, we get two stages for the price of one smaller stage,
because all three flight elements are nearly identical.
The entire system masses 31.3 million pounds at lift off, and
generates 37.8 million pounds of thrust. It burns 29.1 million pounds
of liquid hydrogen and liquid oxygen and accelerates to 3.5 km/sec -
not counting gravity and air drag losses during the ascent.
The two outboard elements fall away, and re-enter downrange. There
they deploy fold-away wings, and glides subsonically with GPS
assistance, to each meet up with their own B737 tow plane. The tow
plane snags the glider with a tow line, and each tow each stage back
to the launch center for release - and automatic landing.
Meanwhile, the core booster continues on its flight to orbit, pushing
two fully loaded S-IIs and a 280,000 lb payload.
When the core booster is emptied, it releases its stack, located on
the nose of the core booster, and descends toward the launch center
for a recovery very similar to that of the outboard boosters. All
three flight elements are returned to the launch center within 90
minutes of launch. Ideal delta vee is 9.08 km/sec not counting air
drag and gravity losses.
The first S-II in the stack, does a brief burn to circularize the
orbit. This S-II is capable of boosting the rest of the stack on any
of the following four missions;
Mission 1 - GEO - 600,000 pounds to GEO - power satellite deployment -
2 days
Mission 2 - Lunar Landing - 280,000 pounds on the lunar surface with
recovery of all components - 8 days to 30 days
Mission 3 - Mars Landing - 280,000 pounds in the mars system including
mars surface - with recovery of all components - 24 months
Mission 4 - NEA Landing - 280,000 pounds on any NEA with recovery of
all components - 36 months
The first S-II masses 1 million pounds and carries 875,000 pounds of
propellant. It imparts 2.2 km/sec to the remaining stack. This
allows recovery of this S-II in a manner similar to that of a ROMBUS
core booster, or Delta Clipper booster. The aerospike nozzle is
designed to withstand high speed re-entry, and the vehicle descends
vertically, and small pump sets fire up and brake the rocket in a soft
landing.
The second S-II has an integrated payload module atop its length,
which carries 280,000 pounds to 600,000 pounds. In the GEO
application this merely circularizes the orbit, releases the payload,
and then deorbits landing back at the launch center.
In the moon landing system, the S-II goes in for a direct ascent to
the moon, and lands vertically on the moon by rocket action alone. It
takes off the same way. In this application 280,000 pounds of
payload, 125,000 pounds of structure, and 875,000 pounds of propellant
operate on the stage to impart up to 5.4 km/sec to the stage. More
than sufficient to land on the moon and return to Earth. With 280,000
pounds of payload, 60 people could stay for up to a year on the moon.
One way 'cargo' flights could deliver more than double this payload,
if the vehicle returned nearly empty.
In the mars landing system, the upper S-II flies to the Mars, and uses
the aerospike/heat sheild arrangement to enter the Mars atmosphere,
and brake directly from an interplanetary trajectory, to either a Mars
landing, or Mars orbital capture. Reducing payload to 200,000 lbs
and increasing propellant mass 80,000 pounds in this system, allows a
delta vee of 6 km/sec - which is more than sufficient to launch off
the Mars surface to an Earth transfer orbit in one stage.
Of course, use of propellants and consumables in flight, lower mass
upon arrival and departure, so leaving 80,000 pounds or so on the Mars
surface, has the same impact as it does on the moon system - so it may
be possible to do more with an optimized system - these are just
preliminary figures based on preliminary analysis.
Obviously, operating stages for a year or more on the moon with 60
people on board, provide powerful assurance that such systems would
operate similarly on a multi-year Mars mission. Also a large vehicle,
provides adequate mass for radiation protection during an extended
voyage, While large crew size and large vehicle size provide a means
to address probable psychological difficulties associated with such a
mission.
Four vehicles launched simultaneously from four launch centers,
1) in USA
2) in Russia
3) in China
4) in EU (South America)
provide a means to send 120 people on expeditions to the moon, once a
year. Spreading the cost of the vehicle development over four groups
of nations, allow reduction of costs. Having two pairs of vehicles,
provide a means to create a bolo-style gravity system during transit.
Having four vehicles altogether, provide a back up capability similar
to that of Apollo 13 - using the lunar lander as a life boat.
A fleet of 3 vehicles from each group, 12 altogether, provide a means
to launch on a monthly basis, solar power satellites to GEO - while
launching 1 year expeditions to the moon, to four lunar outposts
operated by each agency, once a year - all four providing quarterly
launches. And then, the piece de resistance' - all four agencies
salvo launch four mars vehicles on a two year trip to mars every
synodic period.
Again, spreading the cost of the vehicle development, creating a
common mode system, provides a means to reduce costs of sustaining a
manned presence on the moon and mars. Periodically, journeys can also
take place to Venus, and Mercury as well as NEAs and Ceres and other
Asteroids.
This sort of thing makes more sense than NASA building an inferior
version of the Saturn I around Shuttle hardware.
12.6 thrust
1.3 gee
9.692307692 mass
1.211538462 structure
8.480769231 propellant
0.25 payload 2.25 S-II
1 S-II 0.875 propellant
1 S-II
29.07692308 S-0 0.388888889 u
31.32692308 GLOW 4.5 Ve
16.96153846 P1 2.216144183 Vf
0.541436464 u1 1.25 S-IV
4.5 Ve 0.875 propellant
3.508453906 Vf1
0.7 u
11.94230769 S-I 4.5 Ve
8.480769231 propellant 5.417877619 Vf
0.710144928 u2
4.5 Ve
5.57268404 Vf2
9.081137945 Vf1,2
BradGuth
> I want to talk about airframes and engines.
>
> I want to talk first about the M1
>
> Its a 1.5 million to 1.8 million lbf rocket engine developed the US
> Army/Air Force back in the day, and turned over to NASA in 1960.
>
> http://en.wikipedia.org/wiki/M-1_(rocket_engine)
>
> Then, there was the J2 rocket engine with 200,000 lbf to 230,000 lbf
>
> http://en.wikipedia.org/wiki/J-2_(rocket_engine)
>
> This was used on the S-II and S-IVB stages of the Saturn V
> moonrocket. The S-II was a 1,0060,000 lb mass system and the S-IVB
> was a 253,000 lb mass system.
>
>
> Special mention should be made of the S-IV's original configuration -
> with 6 RL-10 engines. The RL-10 is a deeply throttable engine - and
> restartable - perfect for a high performance lunar landing vehicle
>
>
> Now I also want to discuss a little bit, the aerospike engine. This
> is an inside out nozzle arrangement that allows any engine pumpset to
> operate in a wide range of pressure conditions.
>
> In fact aerospike engines have been produced using existing pumpsets
>
> http://en.wikipedia.org/wiki/Image:Annular-Aerospike.jpg
>
> Finally, there are innovations that were developed by legendary
> aerospace engineering pioneer, Phillip Bono
>
You call that a "Modest Proposal"?
What would you call a sophisticated or complex proposal?
Going extremely big seems worth doing, as I too could use a few
million pounds deployed to the Selene/moon L1, or that of my Venus L2
POOF City.
btw, since human DNA still isn't rad-hard, we'll need to deploy lots
and lots of shielding (namely water or perhaps better to deploy h2o2)
in order to protect our frail DNA.
~ Brad Guth Brad_Guth Brad.Guth BradGuth
Williamknowsbest
BradGuth
> I see you're still playing sci.space usenet's own Karl Rowe of
> disinformation.
We can see that if you were in charge of our DARPA and NASA, you'd be
it. Meaning that for other than clones of yourself and your computers
holding all the works of others, there would not be anyone else on
your staff or board of directors, and otherwise only yes boys and
girls would ever get hired.
What's more disinformation worthy than walking upon our physically
dark or darker than coal Selene/moon, that's otherwise loaded with
local substances of great value, plus cosmic deposits of nifty
minerals, crystal and gas elements. At least the crust of our moon
offers 260,000 ppb worth of h2o, and thus far that's 260,000 ppb more
than Mars has to offer.
Damon Hill
(crickets)
--Damon
Williamknowsbest
Williamknowsbest
> (crickets)
>
> --Damon
There are crickets on usenet? lol.
BradGuth
> You can't stand being called out for the person you are.
I really don't terribly mind being associated with the bipolar likes
of William Mook, because at least we each care about our portions of
humanity, and of our frail environment that's badly in need of being
better understood and salvaged.
You have your all or nothing methods that clearly favor the upper most
0.1% of the sufficiently faith-based Americans (even if most of them
are having to be pretend-Atheists), as well as per sustaining their
trickle up status quo economy with intentions of never having to
revise history in order to reflect the truth, or having remorse about
one damn thing, and I favor the lower 99.9% of this entire world
(including all forms of life and of its environment) that's trying to
survive in spite of your 0.1% that you don't hold accountable for much
of anything that turns out bad and ugly, not to mention spendy and/or
inflationary as hell.
Ian Parker
reason. Once you are in space you can use the highly efficient ion
propusion motor.
No, I will correct myself LEO and high energy weight solar systems. If
an objective is SSP what will be needed is just that. Let us think in
terms of a squae kilometer of aluminium 1 micron thick. Weight 2.7T.
This can be used for reflectors. Potentially 2GW is falling on that
sqare kililometer. OK you will need silicon cells struts to give some
degree of mechanical stability. You will only get a limited efficiency
too.
If you could get 500MW for 10 tons you would be well placed not only
to have a good ion drive system, but also a stepping stone to SSP.
To get to LEO only rockets are really feasible. From LEO to wherever
there are a lot of other concepts that should be explored.
- Ian Parker
BradGuth
You do realize that you're speaking to our resident God, don't you?
Our resident lord Mook and substitute wizard of Oz is more than a wee
bit bipolar, and doesn't take kindly to folks that do not 100% accept
his proposal as is.
Imagine what a fully complex and maximum kind of proposal from lord
Mook is like. Just ask and you will receive tens of thousands of his
pirated words and plagiarized science as based almost entirely upon
the hard works of others that don't always get credit.
Technically most anything William Mook has to suggest is doable as
long as you believe everything published by those of of his DARPA/NASA
Old Testament, and that it's either 100% public funded as open-ended
to boot, and/or reverse tax funded is even better, and never mind the
next round of global inflation that'll be created.
Your basic 400~500 km LEO stuff that can manage to always avoid the
SAA contour while being assembled and/or maintained by us humans is
worth doing, although from the tether dipole element of my LSE-CM/ISS
should be a whole lot better.
Ian Parker
remarks I have made. He is the great fan of SSP. How can you want to
develop SSP and no apply the technology to space propulsion? I would
in fact have expected him to come back and say that what I had posted
was unduly pessimistic.
- Ian Parker
BradGuth
Usually he doesn't respond well to those he thinks are beneath his
Godly all-knowing expertise in everything under the sun.
It's not that many of Mook's notions are not without technological
merit, as long as the time required for their R&D plus public funding
is open-ended and without chance of remorse slipping in.
Mook only believes in the future, because the past as having been
scripted as history is forever unchangeable, no matter how skewed,
corrupted or dead wrong that history is. Therefore everything
officially recorded of our DARPA and NASA is absolute matter of fact,
or better than the word of God. This is what drives folks like
William Mook to believe that frail human DNA can easily cope with
whatever's within or outside of our protective magnetosphere, as well
as for easily surviving upon asteroids or that of our gamma and
secondary/recoil X-ray environment of our Selene/moon.
Mook is not a big supporter of rad-hard robotics, or in doing things
in the smallest and most efficient way possible. In the bipolar good
book of Mook, bigger is always better, and yet oddly he doesn't like
my 256e6 tonne LSE-CM/ISS, or forbid anything having to do with Venus.
Derek Lyons
>I feel that we should concentrate on low cost to LEO for the following
>reason. Once you are in space you can use the highly efficient ion
>propusion motor.
So long as you don't intend to actually go anywhere or do anything, an
ion motor suffices.
D.
--
Touch-twice life. Eat. Drink. Laugh.
http://derekl1963.livejournal.com/
-Resolved: To be more temperate in my postings.
Oct 5th, 2004 JDL
Ian Parker
> Ian Parker <ianpark...@gmail.com> wrote:
> >I feel that we should concentrate on low cost to LEO for the following
> >reason. Once you are in space you can use the highly efficient ion
> >propusion motor.
>
> So long as you don't intend to actually go anywhere or do anything, an
> ion motor suffices.
>
William Mook has made a big point about SSP and the amount of solar
power generated per ton of photovoltaics and mirrors. Ion drives
should be viewed in this context. Let us suppose tou have 10MW per
ton. This is in fact a very conservative estimate in terms of what we
are talking about for SSP. If our exhaust velocity is 50km/s we have a
thrust of 400N per ton of cells. This is going to take you quite a
way.
If you want the ultra high performance systems being talked about you
need to start somewhere. To me an ion system with this sort of level
of performance is the place to start. There seems to be little point
in carrying SSP to GEO in a rocket. You take it to LEO in a rocket and
use an ion drive to take it to GEO. What would we be talking about in
a high performance system? 4,000N/T (4N/kg) ? Something of that sort.
- Ian Parker
Pat Flannery
Derek Lyons wrote:
>> I feel that we should concentrate on low cost to LEO for the following
>> reason. Once you are in space you can use the highly efficient ion
>> propusion motor.
>>
>
> So long as you don't intend to actually go anywhere or do anything, an
> ion motor suffices.
>
Assuming you have the time to wait while it accelerates you, you can get
out of LEO with a ion engine.
You'd have to weigh (literally) the savings in conventional propellants
versus the food and water you'd have to add for the crew as they take
several weeks or months to get on their way to their destination.
Certainly this is something that favors a very small crew, or a unmanned
spacecraft.
At lower orbital altitudes air drag versus the ion engine's anemic
thrust could also be a real problem.
Pat
Jeff Findley
"Pat Flannery" <fla...@daktel.com> wrote in message
news:esudndtkZ-bcXSPV...@posted.northdakotatelephone...
> Assuming you have the time to wait while it accelerates you, you can get
> out of LEO with a ion engine.
> You'd have to weigh (literally) the savings in conventional propellants
> versus the food and water you'd have to add for the crew as they take
> several weeks or months to get on their way to their destination.
> Certainly this is something that favors a very small crew, or a unmanned
> spacecraft.
> At lower orbital altitudes air drag versus the ion engine's anemic thrust
> could also be a real problem.
You also have to look at the damage caused by moving slowly through the
van-Allen radiation belts. The radiation in those belts has a nasty
tendency to damage electronics, especially solar arrays.
You really want to start your ion engine journey *above* the van-Allen
belts. Say one of the earth-moon Lagrange points?
Jeff
--
A clever person solves a problem.
A wise person avoids it. -- Einstein
Willie...@gmail.com
is, the money you spend on building ion rockets, is really a
substitute for larger launcher sizes. Since launchers are too small
anyway, for powersat and factorysat and asteroidal capture, the first
step is clear. And once you have a lever, you use it - without
waiting around for efficiency improvements, though you do those too.
To go from LEO to GEO we can figure out from the vis-viva equation;
v^2 = mu*(2/a - 1/r)
To kick a payload from LEO to GEO requires adding 2.47 km/sec to the
payload. Then, when you're at altitude, you have to add another 1.47
km/sec. Then, to recover the booster, you have to subtract 1.5 km/sec
- to re-enter. Total delta vee is 5.44 km/sec.
Most of the payload is deposited at GEO.
Start with 2 million pounds at GEO and look at two different
conditions;
1) chemical kick stage with 4.5 km/sec exhaust speed
2) ion kick stage with 45 km/sec exhaust speed.
The structural fraction of the chemical kick stage is 12.5%. The
structural fraction of the ion kick stage is 37.5% - only 3x the
figure of the chemical rocket.
So, the chemical rocket needs;58.3% propellant fraction to accelerate
its payload to 3.94 km/sec. This leaves 29.2% for payload. Around
584,000 pounds the 600,000 pounds I mentioned.
Now you've got to recover the 250,000 lb stage - there's actually
two in this scenario, and one falls back immediately, while the other
has to be deorbited. Still, we have to subtract 1.5 km/sec - and with
a 4.5 km/sec rocket We need 28.4% the empty mass - which is 70,800
lbs for the larger mass,and 35,400 lbs for the smaller mass - this
reduces the payload at GEO from 584,000 pounds to 429,200 pounds in
the first instance, and 550,000 pounds in the second instance.
So, we have a 29 million pound launcher putting up 550,000 pounds into
GEO - with an all chemical booster -
Now, we have a 45 km/sec ion rocket achieving the same thing - with a
37.5% structural fraction. That's 650,000 pounds of structure.
We have the same 2 milion lbs on LEO. The same delta vees to carry
out. 3.94 km/sec - requires
u = 1 - 1/exp(3.94/45) = 8.4% propellant on the boost up.
This is 167,700 pounds of propellant on a 2 million pound starting
mass. Adding this to the structural fraction, we have 817,700 pounds
of stage and propellant, leaving 1,183,000 pounds of payload. About
double the payload. We have to figure out the deorbit propellant now.
The 650,000 pound stage has to deorbit so, it must go through a delta
vee of 1.5 km/sec. That means 22,000 pounds of propellant are
needed. This reduces the payload on GEO to 1,161,000 pounds.
CHEMICAL ION
4.5 km/sec 45.0 km/sec
459 sec Isp 4590 sec Is
2,000,000 stage 2,000,000 stage
250,000 structure 650,000 structure
1,150,000 propellant 189,700 propellant
550,000 payload 1,161,000 payload
We've more than doubled the payload FOR THIS LAUNCHER by adding a
higher performing upper stage. The question we mst always ask, is the
complexity and cost of adding this sort of technology to the upper
stage worth the improved performance? That is, if we take the dollars
and time to build a larger launcher, would we be ahead?
The answer I get is yes - using money at this juncture to build larger
launchers and launch them from adequately maintained launch centers at
appropriate locations at cost effective launch rates - is the quickest
easiest way to imrprove our capabilities in space. Once we've maxed
that out, we can start talking about improved propulsion - on existing
airframes and so forth.
I have already mentioned elsewhere, on the very large launcher posts I
made a few weeks ago, that laser powered propulsion units are logical
next steps once the laser powersats are installed and excess power is
available at reasonable costs.
This is not the case today since we're suffering from high energy
prices a shortage of supply and increasing demand. Once this is
usefully addressed with the program described here, then it makes
sense to invest in some form of laser/ion propulsion - done at a power
level and at a structural fraction that beats the pants off of
conventional ion propulsion touted here.
Obviously, I'm looking at this as a business proposition.
Step 1: Create ultra-low-cost terrestrial solar panels.
I've done this.
and here's how you use them
http://www.ohiochamber.com/governmental/pdfs/William%20Mook_021308.pdf
Make hydrogen from solar DC and burn hydrogen in coal fired plants to
make AC on demand. Then take the coal not burned combine it with more
hydrogen to make liquid fuel products.
This supplies all our oil needs worldwide, and cuts our carbon use
more than half. This is sufficient to reverse the trend in carbon
build up since nature does have some capacity to absorb carbon in the
carbon cycle.
Step 2: Buy space launch assets from major aerospace firms.
Once this is in place, use the revenues to buy the space launch assets
of the major aerospace companies throughout the world. Those are
reorganize to build up space launch abilities. With this kind of
money I joint venture with other publicly owned business-like
entitites.
Step 3: Build subscale fully reusable commercial launcher.
Basically, I propose the Comon Interplanetary Booster and offer
contracts to help build and operate it - while reserving use for
powersat experimentatoin.
Take a small portion of the nearly $4 trillion earned in fuel and
electricity sales, and invest it in a large heavy lift launcher -
first a 500 ton to orbit. This is described here - and later, when
SSP technology is proven out - a larger 10,000 ton to orbit heavy lift
vehicle. Translating of course the ability to loft 10,000 tons into
25 million pounds of payload on Mars.
Step 4: Deploy a global wireless internet satellite constellation.
Orbiting 660 satellites in 33 sun-synch oribts of 20 satellites each -
each satellite massing 20 tons - provide 50 billion channels of
wireless broadband throughout the world, and capture $300 billion in
communications revenue and trillions of dollars per year in online
banking, financial services, and insurance revenues.
Step 5. Develop and deploy new powersat technology
Using revenues from space based assets, invest in developing new space
based assets, principally powersats. Do this in conjunction with
privately funded exploration along the lines described here, using the
same launcher set, with custom built flight elements to carry out Mars
expeditions, lunar development, and exploration, and asteroidal
exploration and development.
Step 6. Once powersat technology is proven, build larger launchers.
Using a portion of synfuels revenue, build larger launchers along the
lines described elsewhere, capable of putting up 10,000 tons (200
million pounds) into LEO with 12 million pounds (6,000 tons) into GEO
and 5,000,000 pounds (2,500 tons) to the surface of the moon and mars
and the Near Earth Asteroids.
Step 7. Once large powersats are operating on orbit, upgrade upper
stages to use high specific impulse laser propulsion and laser light
sail technology. Use this to harvest asteroids - and double payloads
from Earth to high orbit - and triple payloads to Mars and the Moon
and the asteroids from Earth.
Terrestrial solar power systems that are providing hydrogen for
massive synfuel production have their output increased 16x with the
addition of bandgap matched lasers on orbit - increased energy
translates directly to 16x the energy from hydrogen. As the
hydrocarbon fuels max out - additional demand is fulfilled with
hydrogen fuels.
Step 8. Develop MEMs based laser powered propulsive skin spacecraft
to implement personal ballistic transport on Earth and beyond Earth.
As the ability to absorb increasing amounts of power become bound by
our ability to ship and handle increasing amounts of hydrogen, direct
beaming of laser energy to end users begins. One of the central
consuming sectors is personal ballistic transport. Moving from a
pedestriatn socieety to an automotive society increases energy use
rate by 11x. Increasing from an automotible society to a personal jet
increases use rate by another factor of 9 - 100x more than
pedestrian. Increasing from aircraft to ballistic spacecraft
increases demand for energy another factor 30 - 3,000x pedestrian.
We have sufficient power on orbit if we beam energy directly to users
on demand.
As the cost of power and energy decreases, the cost of handling fuels
comes to dominate the cost - particularly if the fuels are high
pressure gases, or cryogenic fuels. So, when the handling costs
dominate, direct beaming will be preferred.
Instituting a Moore type curve in reucing the cost of energy and power
- from space - we can even predict when these sea changes come about.
When everyone can afford cars, airplanes, and spaceships - and when
they move from hydrocarbon,to hydrogen, to direct beaming.
Willie...@gmail.com
high specific impulses. That's the plus side.
What you have to ask yourself is does the increased cost, complexity
and so forth, pay sufficient dividends to be worth this? Why not
just double the size of the launcher? Would that be prefereable?
That is, I've proposed a 30 million pound vehicle here that puts
550,000 pounds on GEO. Putting some sort of nuclear electric system
together in operate nearly 5,000 sec Isp - doubles your payload to GEO
to 1,100,000 pounds. A large ion rocket that size is an expensive and
complex thing. What about going from a 3 element launcher to a 7
element launcher? That is, add 4 more booster elements to teh first
stage and have a 70 million pound vehicle at lift off. How does that
compare in complexity and cost to building a 650,000 pound ion rocket
engine?
I'm not saying we shouldn't do both. But every battle has a most
effective order to it. The question we have to ask, what's the best
way to proceed today?
I think we need to build heavier launchers and bigger payloads up
there.
This vehicle described here is bigger than anything ever seriously
contemplated before. It also has zero technical complexity (the three
element one) and it puts a crew of 60 on the moon for a year or two -
haha - and a similar crew on Mars for the same period - but only 90
days or so on mars - 2 years in transit.
This is HUGE - compared to what we've got so far.
This system could over a three year period launch a global wireless
hotspot with 50 billion channels - it could land hotels and labs and
big stuff on the moon and mars - launch serious power satellites to
test systems designs and make money doing it - before launching into
really big stuff - put people across the entire inner solar system out
to Ceres.
Once a few power satellites are up, I think beamed propulsion stages,
built around the existing stages would make sense. Laser thermal -
with 10 km/sec Ve (1,000 sec Isp) - laser sustained detonation - with
20 km/sec Ve (2,000 sec Isp) - laser photovoltaic ion rockets - with
50 km/sec Ve (5,000 sec Isp) - laser light sails (infinity Isp no
propellants at all) - These are natural research projects, and ion is
included. When you are power limited, lower isp like low gears give
you more force - at reduced speed.
But, there's a lot that can be done with plain vanilla stuff,and when
you're making money from royalties on the wireless web,and beamed
power sales, then you will increase the efficiency of already
operating upper stages -
Model: Saturn II.
Gross Mass: 490,778 kg (1,081,980 lb).
Empty Mass: 39,048 kg (86,086 lb).
Thrust (vac): 5,165.790 kN (1,161,316 lbf).
Isp: 421 sec.
Burn time: 390 sec.
Propellants: Lox/LH2.
Diameter: 10.06 m (33.00 ft).
Span: 10.06 m (33.00 ft).
Length: 24.84 m (81.49 ft).
Country: USA.
No Engines: 5.
Motor: J-2.
Cost $ : 290.000 million.
First Flight: 1967.
Last Flight: 1973.
No Launched: 24.
I'm proposing a reusable configuration, with thermal protection, and a
zero height annular aerospike engine configured for re-entry base
first and vertial powered touchdown. Landing on the moon and mars also
possible. Two of these guys stacked inline atop the central of three
flight elemets. TPS landing gear and so forth - increases mass to
125,000 pounds - using modern techniques.
The bottom S-II stage carries another S-II stage, that has a 63 ft
tall cone with a 33 ft base - atop the 82 ft tall cylinder - this is
a total length of 125 ft. Its a narrower taller version of this core
stage.
http://www.astronautix.com/lvs/rombus.htm
1/5th the structural mass and 1/10th the mass - though the size of
rombus is the same size as the three flight elements described
elsewhere.
A 45 ft diameter element - and 125 ft tall - and masses 10 million
pounds - is midway between an ET and rombus core booster.- ET is 28 ft
x 158 ft length and masses 1.68 million pounds.
BradGuth
wrote:
> "Pat Flannery" <flan...@daktel.com> wrote in message
I would tend to agree. However, the Selene/moon L1 is taboo/
nondisclosure rated, as I'd bet all other such Ls are either off-
limits or useless according to our resident wizard of Oz.
BradGuth
And that's supposedly modest?
Do you know of the all-inclusive and thus birth-to-grave accounting?
(apparently not)
~ BG
Pat Flannery
Jeff Findley wrote:
> You also have to look at the damage caused by moving slowly through the
> van-Allen radiation belts. The radiation in those belts has a nasty
> tendency to damage electronics, especially solar arrays.
>
Yeah, I hadn't thought of the VA belts; that really would screw things
up for a slow climb to escape velocity via ion propulsion.
So much for Ernst Stuhlinger's ion-driven parasol Mars ships.
Of course WvB had his space station orbiting at 1,000 miles altitude,
so that wasn't a good idea either.
> You really want to start your ion engine journey *above* the van-Allen
> belts. Say one of the earth-moon Lagrange points?
>
That would probably be as good a point as any to start from. Just make
sure you don't collide with the Moon as you spiral out.
Pat
Ian Parker
> "Pat Flannery" <flan...@daktel.com> wrote in message
take todays power weight ratio of solar cells you are absolutely
right. However viewing it in the context of SSP we get a different set
of answers. The thrust you get from ion propulsion (or variants on
same like a linear induction motor accelerating plasma) would not be
all that anaemic. You are talking about 4N/Kg when you are in the SSP
region. Moreover there is STILL quite a lot of radiation at GEO. You
have a point with Van Allen, but it would not be a show stopper for
SSP.
To me the main problem about looking at this as a business proposition
is the size of the technological leaps involved. This does not mean we
should not be discussing it, but I fear that it is before its time. To
me the big question of feasibility is the intermediate steps. Are
there intermediate steps which a hard headed business person would go
for now?
There is, of course, always terrestrial solar power. To me the main
driver for space feasibility is going to be the power weight ratio. I
am saying you need an ion drive giving 4N/Kg if you are going to be
taken seriously. Either that or something approaching a Von Neumann
machine with fabication and self repair, if not complete replication
on asteroids. I can see these two strands producing SSP at an
acceptable cost.
I agree entirely that lasers are the only feasible low cost means of
propulsion. But, you have to get there and demonstrate how you can get
there. Small scale demonstrator projects?
Suppose we can get 0.5N/Kg. That would be a start. We could start our
ion drive with conventional rocket power being used to reach escape
velocity.
There is just one point here concerning Georgia, the gap and the ISS.
Nobody seems to be saying that the ISS would be a good place to try
out new ideas. I tihnk this is telling.
- Ian Parker
Willie...@gmail.com
There's also the cycle time.
Operating commercially to GEO there is a cost penalty in tarrying too
long at any one stage of the journey.
While 2 days versus 70 days may not seem like such a big deal with low
launch rates, especially in the present anti-space environment, that
changes in a commercial operation.
You don't make money waiting.
Putting up a powersat a week or heaven forfend, a powersat a day,
which is what's needed to supply the power needs of a growing planet,
means that one or two of the chemical boosters will be needed while
10 to 70 ion boosters are needed for the same mass flow rate to
GEO..
Now, the savings achieved sending up twice the payload to GEO per
ground launch which is what ion gives you, pales in comparison to the
costs of maintaining 10 to 70 of them in transit to maintain your mass
flow rate to GEO.
As costs drop, demand increases. Withi increasing demand, there is an
increasing mass flow rate from Earth to GEO in this case. A similar
analysis is done for luna and mars and NEAs and Ceres etc. Each
system has an order of battle dictated by the astrogation
requirements, the technical requirements, and the business
requirements.
At present we have intelligence and communications assets at GEO and
MEOI and LEO. I am contemplating global network assets, energy assets
as well as material resource assets, and a human presence at MEO, GEO
and beyond. This changes things.
As I said previously, at some point increasing specific impulse makes
sense. With a successful powersat constellation on orbit it makes
sense to use available spare laser capacity to implement a variety of
laser propulsion technologies. These include;
1) laser thermal 1,000 sec isp 44.5 MW/tonne
2) laser detonation 2,000 sec isp 89.0 MW/tonne
3) laser electric ion 5,000 sec isp 222.5 MW/tonne
4) laser light sail (infinity isp) 1.47 GW/kgf
We have increasing power levels for a given thrust at higher specific
impulses. This is analogous to a gear in a transmission. So, a
gigawatt at 1,000 sec Isp produces about 50,000 lbf engine. That same
gigawatt at 2,000 sec Isp produces only 25,000 lbf but better gas
mileage. At 5,000 sec Isp, we have 10,000 lbf engine and with a laser
light sail the same GW produces 680 grams of force.
With the energy source removed from the rocket, and power delivered by
laser beam, we have the ability to increase thrust to weight - which
shortens boost times and mission times - which maintains flight rate
and cost efficiency - providing balance of system costs are kept under
control.
The launcher therefore benetis most from laser thermal, the upper
stage, laser detonatoin, and kick stage laser-electric while
interplanetary and interstellar stages use laser light sail.
This only makes sense economically when the cost of laser photons from
solar pumped space lasers drop below a certain price.- throughout
their entire cycle of use.
This along with the other considerations determine the 'order of
battle' in introducing the technology. Also open issues in each
system described determine R&D efforts and maturity of the technology
determines the level of R&D effort.
To get an idea of cost - consider that a barrel of crude oil contains
6.1 GJ of energy. It costs these days in excess of $122 - that's $20
per GJ. A 300hp engine costs about $2,350 - that's $10,000 per MW. At
these prices we have built the automotive age - though during its peak
costs were more in the $1 per GJ range.
For laser detonation engines to operate at the same cost as
automobiles requires that each tonne of thrust cost less than $890,000
and each second of illumination cost less than $1.78
At these prices we can begin to consider the use of laser detonation
in our rockets, both on the ground and in deep space.
Now a short range ballistic flier is possible with this system. To
travel further requires more energy than is typically used in
automotive travel. So to get reasonable costs, for daily driver
rockets prices have to drop. Which is another way of saying that
prices have to drop for demand to increase. This is an unremarkable
statement. However, what should the prices be? The answer is around
1% the cost of what we pay for energy and power in automobiles. That
is a tonne of thrust with a laser detonation engine has to cost around
$8,900 and each second of illumination by a space laser to drive it
has got to cost about $0.02 At these prices, we can begin to
realistically consider personal ballistic transport off world, and to
any point on world. The age of the daily driver rocket will have
arrived.
The way to achieve these prices is to invest in technologies that
lower the cost of generation and increase power to weight. I have
engaged in a 12 year program to cut the costs in terrestrial solar
power
and I have developed a business model to implement this technology for
profit
http://www.ohiochamber.com/governmental/pdfs/William%20Mook_021308.pdf
my intent is to continue my research into creating low cost laser
powersats and low cost launchers to support them.
Terrestrial solar at 7 cents per peak wtt produces energy at 1/3 to
1/5 cent per kWh - , that's $0.92 to $0.56 per GJ - which is cheap
enough to electrolyze water into hydrogen and oxygen, and use those
chemicals in a variety of ways to make synthetic fuels of
exceptionally high quality - and do so at a profit! And restore the
price points for energy that prevailed in the 1950s and 60s -
restoring our economic vigor.
I am not content to rest there. Lower costs are possible! The same
terrestrial arrays that will resolve our energy problems today, will
be the basis upon which we will grow into the future! Adding a power
satellite on orbit, the same area as the terrestrial array allows the
terrestrial array to produce 16x more energy in a year. Do this with
only twice the cost, and the cost of power drops to 1/8th the figures
above - doing in the 2010s what we should have done in the 1970s -
made power too cheap to meter!! Of course the costs of meters have
dropped since the 1950s, and the uses we'll put this energy to is
unimagined in the 1950s yet, at $0.12 to $0.07 per GJ - we are at a
price point that makes personal ballistic transport a reality. Also,
resolving the cost issues associated with the powersat in the first
place, helps solve some key issues to make laser detonation -
especially when done in MEMs arrays - I like to call propulsive skin -
gets the price for an engine to where it needs to be as well.
BradGuth
And the trillions upon trillions of public loot in order to pay for
all of this off-world and terrestrial infrastructure is materializing
exactly how?
~ BG
Willie...@gmail.com
> I feel that we should concentrate on low cost to LEO for the following
> reason. Once you are in space you can use the highly efficient ion
> propusion motor.
Recall, that high specific impulse on orbit, translates to more
payload on high orbit for a given launcher size.
Given our current capabilities in space, apogee kick motors and
perigess kick motors are nearly all solid propellant. They are also
all expendable. This means they're low performing and high cost.
I am proposing here that they be replaced with a reusable hydrogen/
oxygen system that is more than twice as efficient when compare to
solids, and less than 1/50th the cost when compared to expendables..
These two improvements increase payloads to deep space and lowers
their cost already.
Now, increasing specific impulse beyond that of hydrogen/oxygen and
lowering costs below reusable chemical propellants will occur along
the lines I've described elsewhere. It will be done as part of a
continuing chain of commercial development.
Low cost to LEO is achieved by building fully reusable systems,
operating from adequately maintained infrastructure, having sufficient
operating rates and process automation to end the need for standing
armies of technicians to fly them.
Anyone who has paid a premium for a swiss watch knows that smaller can
mean more expensive - depending on the details. An appropriately
sized vehicle can do much to reduce costs as well.
Yet, there is something to be said about smaller systems built around
existing engine sets with modern airframes and avionics.
The smallest vehicle I have analyzed involves building an annular
aerospike nozzle around an Pratt&Whitney RL-10 pumpset. Last time I
looked this it cost $5 million. A flight element, when you include
avionics and airframe, cost $12 million for the flight article.
The engine is capable of 4.5 km/sec exhaust speed and 20,000 lbs
thrust. It is attached to a flight element that has 15,384 pounds
mass, carrying 14,000 pounds of propellant - massing 1,384 pounds
emtpy - using SSTO technology
This vehicle with 750 pounds payload,is capable of attaining orbit.
More importantly it is capable of being tested through all flight
regimes before being put into service.
Ganging together 3 flight elements, the two outboard elements
operating as first stage, feeding propellant to the central element,
which operates as second stage, along the lines already described for
the larger vehicle, puts 4,500 pounds into orbit.
The vehicle masses 50,000 pounds a lift off and produces 60,000 pounds
of thrust. When the two outboard elements separate, the system is
flying at 5.47 km/sec - less air and gravity drag losses. The upper
stage adds another 3.63 km/sec - again without air and gravity drag
losses.
This system costs $36 to build, another $14 million for non-recurring
engineering charges. Add another $50 million for launch center and
operations center. This is $100 million - and it the minimum you can
do.
A 4,500 pound payload is about the size of a Gemini capsule or Soyuz
orbital module. With modern avionics and structures, and best
engineering practices, it should be possible to put 3 to 5 astronauts
aboard - if that's the direction you want to go.
The piloted stage - is likely to add another $100 million to the
overall costs - and add another $100 million to build a fleet of 3
launchers for a reasonable flight rate of once per month trending with
learning curve effects to 2 per month.
So, we're talking $300 million minimum cost - in this sort of program
- to prove out the various ideas I'm putting forth here.
Putting in a couple of expendable kick stages like TE-473 at perigee
and TE-416 at apogee - both from Thiokol - gives you a capacity to
take 500 pounds or so to GEO - at several million dollars added cost
per launch.
In short, for $300 million gets you into the launch business - at the
same scale most others are in these days.
But with the exception of expendable upper stages, you've dramatically
lowered the launch costs. A flight rate of one launch per month is
maintained, and about $60 million per year is earned - $5 million per
flight - trending to $120 million per year for your $300 million
invested.
Insurance costs are extra.
Increasing the size of the vehicle is easily achieved by increasing
the number of pumpsets per aeropike engine. Six to eight seems to be
the practical limit - with 160,000 pounds thrust topping out the
system using these pumps. This is about 1/70th the size of the M1
based system I've described elsewhere and miniscule compared to the
super-heavies needed to develop the moon and mars and asteroids
industrially.
Since payload to orbit scales with launch weight, a 8 pump system
launches 31,500 pounds into LEO - and vehicle elements are twice the
linear dimension of the 'starter rocket'.
Costs are 6x larger as well. We're talking a $1.8 billion program
now. But, we can use the smaller vehicles as high energy reusable
upper stages!
Two of the flight elements described above can carried aloft the
larger system, fired in parallel to put 9,000 pounds into GEO with
recovery of ALL stages.
Alternatively, the two flight elements can be fired in series, and
place 4,500 pounds on the moon and return it to Earth. Also, 4,500
pounds can be landed on Mars and returned to Earth.
This achieved with RL-10 pumps configured as described.
This replicates what we're doing now in space, and doesn't
substantially advance the art - though it does apply some technology
that has been developed over the past 20 years.
Once I have the money to treat rockets like sportscars or race planes,
I'd build the following $30 million sportscar;
Two RL-10 pumpsets feeding an annular aerospike engine can be tweaked
to produce 44,000 pounds of thrust. This is 1/500th the thrust of the
ROMBUS
http://www.astronautix.com/lvs/rombus.htm
Cutting all the numbers back by 500 - and the dimensions back by the
cube root of 500 - you have micro-m-bus haha..
Mirco-moon bus.
LEO Payload: 1,980 lb to: 185 km Orbit. at: 28.00 degrees.
Liftoff Thrust: 36,000 lbf
Total Mass: 28,000 lbf
Core Diameter: 9.82 ft
Total Length: 12 ft
Eight tanks,
Gross Mass 575 lbs
Empty Mass 80 lbs
Length: 12.61 ft
Diameter: 3.14 ft
Propellants: Lox/LH2.
Stage1: 1 x Rombus.
Gross Mass: 22,500 lbs
Empty Mass: 1,350 lbs
Motor: 1 x Plug-Nozzle Rombus.
Thrust (vac): 45,800 lbf.
Isp: 455 sec.
Burn time: 215 sec.
Length: 12 ft.
Diameter: 6.55 ft
Propellants: Lox/LH2.
This puts 1,980 pounds into orbit - this is a 1 person capsule -
integrated atop the micro-Rombus vehicle.
The interesting thing is that any launcher that can put 22,000 pounds
of propellant on orbit next to the micro-rombus, can cause it to carry
out the mission profiles described in the associated missions
Project Selena,
Project Deimos
- all done on a micro scale.
http://www.astronautix.com/craft/proelena.htm
http://www.astronautix.com/craft/proeimos.htm
The 3.14 ft diameter 12.61 ft long hydrogen tanks are a bit small for
habitats - perhaps individual sleeping compartments on the moon or
mars.
The vehicle that is 8x larger than the minimum vehicle - the one that
puts up 31,200 pounds - is fully capable of supporting the micro-
rombus.
So, you either put it as the upper stage aboard the vehicle, or you
launch both separately and refuel on orbit. The latter is useful if
the launcher is not man-rated.
These smaller systems as you can see do not lower the cost to LEO on a
per kg basis, as far as the larger systems. Nor do they increase the
mass flow rate to orbit to absorb the increase in demand caused by
lower costs.
> No, I will correct myself LEO and high energy weight solar systems.
As I said, putting a reusable hydrogen oxygen kick stage to work is a
vast improvement over current art and the addition of laser
propulsion, (thermal, detonation, electric, light sail) is premature
until laser costs are in line with other rocket costs.
> If
> an objective is SSP what will be needed is just that. Let us think in
> terms of a squae kilometer of aluminium 1 micron thick. Weight 2.7T.
> This can be used for reflectors. Potentially 2GW is falling on that
> sqare kililometer. OK you will need silicon cells struts to give some
> degree of mechanical stability. You will only get a limited efficiency
> too.
Use gas pressure - see Echo -
http://www.astronautix.com/craft/echo.htm
a transparent top sheet and a reflective back sheet - heating is an
issue - this is starting to sound like my low mass powersat.
This will be developed after laser power sats are operating - when it
makes sense to do so. The effect will be to raise the payload to
high orbit by a factor of 2 to 3.
The problem is the thrust to weight versus power and mission times.
> If you could get 500MW for 10 tons you would be well placed not only
> to have a good ion drive system, but also a stepping stone to SSP.
Making a lightweight laser powersat along the lines I've described
elsewhere, and launching it aboard a large launcher - provides a means
to directly enter a powersat program where the launcher and satellite
and everything else is developed in one program and paid for by power
sales.
> To get to LEO only rockets are really feasible.
No, once you have substantial power on orbit, available by laser or
maser beam - you have a variety of rocket propulsion elements to
consider;
laser thermal - 10 km/sec 44.5 MW/tonne
laser detonation - 20 km/sec 89.0 MW/tonn
laser electric - 50 km/sec 222.5 MW/tonne.
laser light sale - 300,000 km/sec 1,470 MW/kg
Reducing oxygen aboard to cover only early ascent, and then relying on
laser thermal rockets to be used at the tail end of launch, and then
switching to laser detonation on the ascent stage, and then, to laser
electric on the orbital stage, provides adequate thrust throughout the
flight cycle, and optimizes vehicle performance so that payload on GEO
per launch is maximized.
This requires the development of a few dozen technologies and the
successful resolution of nearly 100 open issues - AFTER the powersats
are operating.
Putting a powersat program at the tail end of resolving all this - is
just a way of saying you're not going to be doing it anytime soon.
> From LEO to wherever
> there are a lot of other concepts that should be explored.
I have looked at nearly everything out there, what I propose here is
the first step in the direction I need to go following the success of
my early coal-to-liquid projects. I have sponsored 8 projects around
the world - and intend to do 50 to 60 projects over the next 10
years. Upgrading each of these terrestrial solar arrays with bandgap
matched light sources - provides a means to provide ALL the world's
PRESENT energy supplies from space. This generates $4 trillion per
year - with high margins. Given the EBITDA I've projected for the
system, it will have a market capitalization - assuming no further
fuel price increases - of about $80 trillion. 1/3 of that will be
owned by me. This will provide sufficient capital to acquire all the
space launch assets of the planet and organize them to do something
interesting in space. This will include;
1) global wireless internet - telerobotics
2) global beamed power from space
3) industrial development of the moon and mars
4) capture of NEAs and industrial development on Earth orbit
5) telerobotic factory satellites
6) personal ballistic transport
7) personal spaceship, space home.
Then I will retire.
> ย - Ian Parker
Willie...@gmail.com
>
> - Show quoted text -
I have not proposed spending any public money on any of my projects.
The present day market for fuels on this planet is $4 trillion per
year. These fuels support a $70 trillion economy. That economy and
the demand for fuels, increases at 4% per year globally, and as much
as 12% per year in economic hotspots like China, India and
Indonesia.
Buying coal reserves, and stripped out surface mines, and putting
solar panels on the stripped out land, and making hydrogen and oxygen
gases from water using sunlight, and using the hydrogen and oxygen
with coal to make AC electricity (through co-gen) and gasoline, diesel
fuel and jet fuel - provides a way to increase the value of coal
acquired for less than $2 per tonne, to something worth on the order
of $1,000 per tonne.
This creates private funds. I pay taxes, I do not take them.
Like Howard Hughes back in the day I do not sell my technology. I
build own and operate facilities for a piece of the action. I own 35%
of the output of all my facilities. The mine owners and financiers
own the other parts of the output.
A 200,000 b/d facility takes 30,000 tons of coal per day and 6,000
tons of hydrogen to make $25 million in free cash flow per day.
Each facility is therefore worth $210 billion at today's fuel
prices. That makes my piece of each facility worth $70 billion -
that's $3 billion free cash flow per year from each facility.
The money comes from the sale of gasoline diesel fuel and jet fuel
from each facility.
I have 8 facilities under construction right now.
I will build 50 to 60 facilities over the next 10 years.
This will net me $3.5 trillion in assets, and $150 billion per year in
free cash flow.
Somewhere along this development arc I will acquire the space launch
assets of all the major aerospace companies. Either by buying the
companies outright, and splitting off their non-space launch
components - or by buying those portions from their parent. This is
likely to cost something on the order of $50 billion.
Another $65 billion will reorganize those assets and result in the
sorts of systems we're describing here - the first step is to stop the
bleeding and get off the government teat - this means finding a
revenue source.
The low hanging fruit is telecom. Teledesic had the right idea, but
they didn't grab their balls and start building the launchers they
needed. This killed them, as it killed Motorola's Iridium program. I
will launch 660 satellites into sun synch polar orbits using a fully
reusable launcher of the type described here. This will provide 50
billion wireless broadband channels with a very simple chipset that
can be installed in any sort of digital device. This will generate
about $300 billion per year in revenues, and make the $110 billion
acquisition and reorg costs, well worth it. The $300 billion per year
will generate another $100 billion per year to maintain the
infrastructure - which is about 3x what the world spends on space
development today. The rest will accumulate to a kitty that will
allow me to make bets on space related stuff. Also I will invest
heavily in tele-robotics for use on Earth - and telepresence.
After the communications constellation is producing a profit, I will
then build test satellites to do powersat experiments in space, and
telerobotic assembly and manufacturing in space - as well as
telepresence to implement space travel for the masses.
That is, while you will be able to buy a flight to orbit for $5
million along this development arc, you will also be able to hire a
live circuit to a telerobot on orbit for $5 per minute - this will be
popular at State Fairs and so forth - providing a full environment.
And for a few dollars per month - you'll be able to do telepresence
online - across the solar system - and there will of course be
computer modelled environments created from live data fromacross the
solar system.
This will get us to the moon, mars, and NEAs - allow us to develop
each, and when the cost of energy drops to an appropriate level, usher
in personal ballistic travel and personal spaceships, and finally
personal space homes the size of los angeles county for the cost of a
10 acre plot in Wyoming.
Willie...@gmail.com
Sunpower back junction cells) at most, and in space solar energy has a
density of 0.137 watts per sq cm and produces 0.03 watts electrical
per cm2, a dichroic mirror reflects away all wavelengths longer than
bandgap energy, with the cell radiating away 0.068 watts per sq cm as
heat.
0.068 W/cm2 = 5.67e-12 * T^4 = 331K
That's about 58C ... 136F which is warm, but operable. Cutting off
shorter wavelengths cools temps further, but at a cost of cutting
power output.
Now, consider that solar cell thicknesses can be as little at 275 um
thick - massing 64 milligrams per cm2.
Using the 0.03 watts/electrical per cm2 this is 469 kW/tonne.
Now there are added things, like the film mentioned, foils to conduct
electricity, mounts structure etc., etc., etc., - but the solar cells
can be thinned by chemical and mechanical polishing - and so forth -
this is a good estimate of what's possible.
Now consider a bandgap matched laser illuminating the solar panel at
6.75 watts/cm2 being converte with 99% efficiency. With a optical
bandpass filter that admits only the laser wavelength, the same 0.068
W/cm2 is radiated away, so the same operating temperature. The same
64 milligrams per cm2 - but now there's a whopping 6.68 watts/cm2 of
electrical power!!! That's 104.4 MW/tonne
A resistojet rocket with 1,000 sec Isp requires 44.5 MW/tonne thrust
A plasma arc rocket with 2,000 sec Isp requires 89.0 MW/tonne thrust.
An ion rocket with 5,000 sec Isp requires 222.5 MW/tonne thrust.
Now consider that in order to create an ionized stream you need to
make a plasma, and to make a plasma you need to make a hot gas.
Consider that with 104.4 MW/tonne you and create a laser electric
rocket that has three modes of operation;
A resistojet mode that heats cryogenic hydrogen to high temperatures
and produces 2.34 thrust to weight.
A plasma arc mode that heats a high temperature gas and produces 1.17
thrust to weight.
An ion rocket that takes a plasma and accelerates it and produces 0.47
thrust to weight.
A chemical rocket has a 70:1 thrust to weight typically, and to
produce 1.3 gees at lift off it comprises 1.85% total lift off
weight. With a 8% structural fraction - the engine is 25% of the
empty vehicle weight.
Limiting our engine mass to 35% of the total vehicle weight - we have
the following gee forces possible with this improved system;
resistojet mode - 0.819 gees - sufficient for an upper stage during
ascent.
plasmajet mode - 0.409 gees - sufficient for an upper stage late
ascent
ion mode -0.164 gees - again very late upper stage and into orbit.
With 1/6th gee you can almost do a lunar landing at 5,000 sec Isp.
Which is pretty amazing. Transfer times aren't going to be a problem
with this system.
Modifying the central stage with this sort of system means deploying a
wing - how large does the wing have to be? Well if the stage mass is
9.7 million pounds and gee force is 0.819 - then we have 3608 metric
tons of thrust
That's 160.5 GW.
This is a little less than the 220 GW powersat I've spoken of
elsewhere.
At 6.68 watts/cm2 - 66.8 kW/m2 - 66.8 GW/km2 - A square on each side
1.1 km!!
Quite a sizeable wing!
and not doable.
Increasing intensities 1 million times reduces wing sizes to 1 meter
area - which is far more reasonable size. The problem now is
temperature.
For the laser, 1,000,000x is possible, and for the PV too if it can be
kept cool. Another advantage is that the mass is reuced by a factor
of 1,000,000 - 104.4 TW/tonne so, 160.5 GW engine has a power plant
that masses only 1.53 kg!!!!
The power plant is now way way down there - and we can reduce engine
size to that of a chemical engine, and launch from the surface!! In
fact increasing the power levels to 802.5 GW - increases the mass of
the power plant to only 7.65 kg.
The problem is heating the PV cell.
One way to do that is to pass the propellant through the PV wing - to
keep it cool.
But, when we do that some argue we are making the PV a heat exchanger,
and we might as well do that directly. because the first two methods,
the resistojet and arcjet, can be done by DIRECT LASER ACTION.
So, we can see that going 1 million times the intensity above - which
is perfectly doable - provides a means to reduce the wings to meter in
size this way and up our thrust to weight while increasing thrust and
power levels and whatnot.
The ion rocket can also be done this way, using the cooling action of
the propellant moving through the PV wing to keep it cool during
operation, and then use the tremendous electrical output to operate
really intense arcjet and ion rocket - using laser action itself to do
direct heating - with maybe some resistojet assistance.
At this point we dispense with stages and turn EACH of our launch
elements into super massive cruisers.
With a 50 km/sec exhaust velocity launched with a 1.2 TW beam from the
surface of Earth - we have to increase our thrust to produce 1.3 gees
at lift off - and now our 9.7 million pound vehicle requires only 2.4
million pounds of propellant, and 1.3 million pounds of structure -
providing a lift capacity of 6.0 million pounds to GEO and return.
Lifting 5 million pounds to the moon and returning it. Lifting 5
million to Mars and returning.
ALL WITH THE SAME AIRFRAME AND AVIONICS
Merely replacing the Aerospike engine, with a laser/thermoelectric
engine.
The 'power pads' are so small and lightweight, the base of the vehicle
is replete with them. They attach to the skin of the vehicle, and on
its base, and propellant is passed automatically through the hot spot,
regardless of how the vehicle is oriented. The intense laser beam is
also redirected from the ground during launch.
So, for every single launcher described here, that puts 600,000 pounds
into GEO -and 280,000 pounds to the moon or mars, with the erection of
the powersat network, and the perfection of the thermoelectric laser
engine, we now have THREE vehicles capable of putting up 6,000,000
pounds to GEO and 5,000,000 pounds to Mars!!!
I mentioned these elswhere in other posts, but perhaps it wasn't
clear.
A fleet of a three dozen vehicles lifting 15 million pounds across the
solar system each year becomes a fleet of 108 vehicles lifting 780
million pounds across the solar system each year.
This by only continuing power plant construction and solving a few
open issues with cooling PV cells operating under intense
conditions.
Clearly the most efficient growth path in performance with the least
R&D and investment dollars.
The next step is to take this technology, and put it on smaller
vehicles using highly reliable MEMs rocket arrays
http://www.nsti.org/procs/Nanotech2006v3/3/W51.03
http://www.me.berkeley.edu/mrcl/rockets.html
using the same approach...
A 1/2 ton vehicle carrying four adults - requires no more than 200 MW
when pulling a little less than 2 gees - and requires 84 kg of
propellant - a little more than 1 cubic meter of liquid hydrogen - to
attain orbit.
In this application, it may be possible to use superheated steam that
decomposes by thermolysis - and then ionize both the oxygen and the
hydrogen - but there are open issues. Even so, being able to use
water - reduces propellant volume to 84 liters - about the size of an
automobile gas tank today!!!
The age of the daily driver rocket will certainly have arrived.
I have related here the thought process that went into my decisions to
proceed as I have - I believe these are the best decisions by far.
BradGuth
>
> read more ยป
Unless you're and ET with advanced anti-aging genetic biology, you're
not going to live long enough to accomplish 1%, if that much.
Of course, as long as you manage to keep the cost of energy inflated
and/or De Beers and ExxonMobil like cartel or cabal regulated, you'll
get enough of our public loot, and then some. Clearly you have no
problems affording $5/gallon or even $10/gallon of refined fossil or
synfuel.
BradGuth
>
> using the same approach...
>
> A 1/2 ton vehicle carrying four adults - requires no more than 200 MW
> when pulling a little less than 2 gees - and requires 84 kg of
> propellant - a little more than 1 cubic meter of liquid hydrogen - to
> attain orbit.
>
> In this application, it may be possible to use superheated steam that
> decomposes by thermolysis - and then ionize both the oxygen and the
> hydrogen - but there are open issues. Even so, being able to use
> water - reduces propellant volume to 84 liters - about the size of an
> automobile gas tank today!!!
>
> The age of the daily driver rocket will certainly have arrived.
>
> I have related here the thought process that went into my decisions to
> proceed as I have - I believe these are the best decisions by far.
There's no question that LEO or GSO space and that of our extremely
nearby Selene/moon offers nearly unlimited and relatively clean
energy. It'll be just ten times as spendy and decades if not
centuries down the road.
In the mean time, there's ten times as much renewable and clean
terrestrial accessible solar, wind, geothermal, tidal and thorium
based energy for accommodating humanity as is. We need only to
develop 100 TW of such clean and affordable energy, plus a
distribution grid that's mostly terrestrial.
Although, if your private space based alternatives can be competitive,
then so be it.
Willie...@gmail.com
> > [snip]
>
> Unless you're and ET with advanced anti-aging genetic biology, you're
> not going to live long enough to accomplish 1%, if that much.
>
> Of course, as long as you manage to keep the cost of energy inflated
> and/or De Beers and ExxonMobil like cartel or cabal regulated, you'll
> get enough of our public loot, and then some. ย Clearly you have no
> problems affording $5/gallon or even $10/gallon of refined fossil or
> synfuel.
>
> ย ~ Brad Guth Brad_Guth Brad.Guth BradGuth
Last time I filled up by Alpha 8C Spider in Geneva I paid $8.06 per
gallon. Cost of living in Geneva is double that of New York City.
When I filled up my Maserati in Sydney last December I paid $6.40 per
gallon. Cost ofliving in Sydney is 50% higher than living in LA.
USA pays the least for all its commodities of all the nations on
Earth, 4% of the people consume 30% of everything at deeply discounted
prices and then complain the most about what they pay.
As usual, your commentary is so off-the-mark I am left speechless by
its total lack of sensibility. Commenting on what you meant by what
you said would be akin to commenting on what the surf outside my
Malibu beach house meant by this are that crashing of waves.
Willie...@gmail.com
>
> > using the same approach...
>
> > A 1/2 ton vehicle carrying four adults - requires no more than 200 MW
> > when pulling a little less than 2 gees - and requires 84 kg of
> > propellant - a little more than 1 cubic meter of liquid hydrogen - to
> > attain orbit.
>
> > In this application, it may be possible to use superheated steam that
> > decomposes by thermolysis - and then ionize both the oxygen and the
> > hydrogen - but there are open issues. ย Even so, being able to use
> > water - reduces propellant volume to 84 liters - about the size of an
> > automobile gas tank today!!!
>
> > The age of the daily driver rocket will certainly have arrived.
>
> > I have related here the thought process that went into my decisions to
> > proceed as I have - I believe these are the best decisions by far.
>
> There's no question that LEO or GSO space and that of our extremely
> nearby Selene/moon offers nearly unlimited and relatively clean
> energy. ย It'll be just ten times as spendy and decades if not
> centuries down the road.
>
> In the mean time, there's ten times as much renewable and clean
> terrestrial accessible solar, wind, geothermal, tidal and thorium
> based energy for accommodating humanity as is. ย We need only to
> develop 100 TW of such clean and affordable energy, plus a
> distribution grid that's mostly terrestrial.
>
> Although, if your private space based alternatives can be competitive,
> then so be it.
>
>
> - Show quoted text -
Of course what you say has nothing at all to do with what Ian and I
are talking about.
A common fully reusable booster built around M1 and J2 engine sets -
using standardized components among the for major space agencies -
spreading the cost of operations and research and so forth - would not
only spend less money than todays agencies over all, but achieve far
more in space.
Three 9.7 million pound elements joined in parallel, operating as two
stages, lofting two 1 million pound stages in line with the central
element - is capable of sending 600,000 lbs to GEO and returning, or
sending 280,000 lbs to Lunar Surface or Mars system and returning with
the payload. 500,000 lbs can be sent one way with return of the
nearly empty vehicle.
280,000 pounds is enough to supply 60 people for 24 months - providing
an instant lunar base or mars base - and in cargo mode, 60 to 120
people per landed 500,000 lb module - particularly when using local
water supplies.
600,000 lbs to GEO is enough to do solar power satellites testing.
1,000,000 lbs to MEO is enough to put up large constellations of
50,000 lb communications and sensing satellites that provide a global
telecom and surveillance service.
By boosting a higher performing ion rocket into orbit, payloads to
high orbit can be doubled - though with limited power, thrusts are so
low, that mission times are adversely affected - particularly in
cislunar space.
Once adequate solar power satellites are operational, and laser
beaming is done cheaply enough, we might consider a laser/ion
combination or a very high performing laser/thermal/ion combination -
this last innovation once carried out successfully will put over 5
million pounds on the moon and mars, and 6 million pounds at GEO -
enough for very large solar power satellites.
The intense laser/pv device is a component for a near solar surface
solar laser system that is 1000s times less expensive per peak watt
than earth orbiting solar power satellites - and provides a means for
large scale interplanetary and even interstellar travel to be common
place.
BradGuth
And as per Mook usual, the lower 99.9% of humanity gets to pay for
everything, including paying the ultimate price or your perpetrated
wars.
We need a viable plan of action that'll offer 100 TW of clean/
renewable energy in order to safely and efficiently accommodate 1e10
humans.
Your ENRON form of off-world derived energy, at $10/kwhr by the time
it gets fully operational, isn't such a good idea. Otherwise I like
working with powerful laser cannons, especially those capable of
pumping out fast moving Rn ions.
~ BG
Rand Simberg
Flannery <fla...@daktel.com> made the phosphor on my monitor glow in
such a way as to indicate that:
>
>
>Derek Lyons wrote:
>>> I feel that we should concentrate on low cost to LEO for the following
>>> reason. Once you are in space you can use the highly efficient ion
>>> propusion motor.
>>>
>>
>> So long as you don't intend to actually go anywhere or do anything, an
>> ion motor suffices.
>>
>
>Assuming you have the time to wait while it accelerates you, you can get
>out of LEO with a ion engine.
>You'd have to weigh (literally) the savings in conventional propellants
>versus the food and water you'd have to add for the crew as they take
>several weeks or months to get on their way to their destination.
>Certainly this is something that favors a very small crew, or a unmanned
>spacecraft.
Or sending the vast bulk of the mission with an ion engine, and
sending the crew to catch up with it and rendezvous in an Orion or
similar after it's reached escape.
Ian Parker
> On Wed, 03 Sep 2008 11:50:02 -0500, in a place far, far away, Pat
>
This is (in effect) the quadrature idea. The ion motor would in fact
have to set up a quadruture between Mars and Earth. The first trip to
Mars would be supplies and be unmanned, while the second would be
manned. If you were just sending people off with supplies to last just
a few days you would not need that large a rocket. It would be a lot
smaller than Apollo. Your reentry spacecraft would make the initial
(unmanned) quadrature trip.
I believe that would be the cheapest way to get a manned expedition to
Mars. There is someone who has advocated using Martian CO2 + sunlight
to produce rocket fuel. Beauty of quadrature. This could be done and
tested BEFORE the manned expedition starts. They would know whether
the refueling would work or not.
- Ian Parker
BradGuth
Mars is simply a faith-based kind of inert failsafe planet, and that's
the only reason its getting our hard earned loot invested.
There's 260,000 ppb more h2o within the crust of our Selene/moon than
Mars has to offer. Go figure why there's not multiple underground
lunar habitats as of decades ago.
~ Brad Guth Brad_Guth Brad.Guth BradGuth BG
BradGuth
Honestly, are you bed-ridden or otherwise confined?
It's perfectly OK if you are, as it would make a great difference in
the way folks treat your topics and feedback, because it's not of
bigotry or much less racism if we're being intentionally misinformed.
~ Brad Guth Brad_Guth Brad.Guth BradGuth BG
Ian Parker
from it. I am saying though that if you do, the way NASA is going
about it is completely wrong.
- Ian Parker
BradGuth
I concur entirely, as well as of going back as of decades ago is when
I believe our DARPA / NASA started going South on us, so to speak.
Currently our NASA is still a very born-again kind of faith-based
cabal, that's taking us to the cleaners.
Williamknowsbest
>
> - Show quoted text -
As usual your commentary seems to a well-crafted disinformation sound-
bite in response to my commentary. That is, it took you several days
to come up with a response, and that response is targeted at an
emotional level, and makes absolutely no sense whatever when viewed
logically. I'm driving cars in Geneva and Sydney and LA, and you aske
me if I'm bed-ridden! lol. after a 4 day hiatus (here, elsewhere
you've made multiple posts to everyone of my comments) - so I gotta
believe this one stumped you in your continuing efforts to ratfuck
everthing I say. hahaha. You're a trip Guthball, and it will be a
very pleasant day when I depose you in court to find out what you're
really about! My guess is you will have a difficult time avoiding
jail.
Williamknowsbest
>
> - Show quoted text -
You want to do it as cheaply as possible, and as quickly as possible.
This means leveraging it off of another program that actually makes
money for the launch provider and vehicle developer, and leverages
common mode design, which is what my original post was all about - and
spreading it around to other folks who are already spending money on
space launch, is a good idea too. Kennedy had that idea, but died
before he could implement it.
BradGuth
Whatever makes you a happy bipolar camper.
~ BG
Willie...@gmail.com
offensive things regardless of what is said. I guess those things
must've tested well in audience testing hunh? lol.
Fact is, the idea of using Mars water and a solar or nuclear power
source to process water into hydrogen and oxygen, provides a means to
send an additional 600,000 lbs to Mars - nearly the entire stage
weight! Especially if oxygen is derived from water or CO2 for
breathing. Pressurizing the mars atmosphere to grow crops would also
be interesting.
A person takes 2 lbs of oxygen per day on average. That's about a
liter of water per day. That's 15.8 MJ per day. That's about 184
watts continuous. Four 300 Watt solar panels (power rating at Earth)
and intertie, and a laboratory electrolyzer - to generate enough power
at Mars to provide the oxygen needed by an astronaut! The astronaut
could breathe indefinitely with a water supply!
Another possibility is to pressurize the Mars atmosphere using a
molecular sieve or a cryogenic separator to separate out the 0.13%
oxygen and then compress it.
This requires that the 0.60 kpa atmosphere be pressurized 27,000x Mars
pressure with everything but oxygen filtered out.
Nitrogen would only need to be be pressurized 4,500x to have the same
ratio of oxygen and nitrogen as on Earth;
The energy needed to compress is not less than
W = nRT*ln(Pa/Pb)
Pa/Pb = 27,000 --> ln(Pa/Pb) = 10.2
2 pounds of O2 gas molecules 28.4 moles so n 28.4
The temperature of the Mars atmosphere ranges from 133K to 293K - lets
say an average of 220K
R is the rydberg constant and that's 8.314 J/(K*mol)
W = 28.4 * 8.314 * 220 * 10.2 = 530 kJ for the
oxygen/day
That's 6.13 watts!!!! A hampster in a cage could generate that!
lol.
Pa/Pb = 4500 --> ln(Pa/Pb) = 8.4
The nitrogen n=74.6 moles
T and R the same
W = 74.6 * 8.314 * 220 * 8.4 = 1,146 kJ for the
nitrogen
Since the nitrogen is not consumed, you could take two days to
pressurize it and it would take only 6.6 watts - that is the same
system with a different filter would pressurize nitrogen in the same
atmosphere to Earth normal pressure, in two days, and keep it there.
Water vapor in Mars' atmosphere is 0.03% in Earth's atmosphere its 1%
- this means that to attain the same level of water vapor in a
synthetic atmosphere as seen on Earth the water vapor must be
separated and compressed 5,700x
Pa/Pb = 5700 --> ln(Pa/Pb) = 8.6
n=0.02 - for the combined Nitrogen and Oxygen
T and R the same
W = 0.02 * 8.314 * 220 * 8.6 = 0.3 kJ for the water
vapor (at 1%)
This is trivial = 3.63 milliwatts!!!
This is telling us we can mine water from Mars' atmosphere very
easily!! At 6 watts we can get 2000X at much water as needed to make
the atmosphere as moist as Earth's this means liquid water is easily
obtained.
A single 300 watt solar panel with a molecular sieve and compression
pump, supplies 50 people air and water directly from Mars' atmosphere.
Is this right? Even if the molecular sieve introduces a 50% loss,
this is a remarkable calculation if I haven't made a mistake.
A simple pump would pressurize a PET film to pressures where crops and
people would easily survive.
Plants may not need much. Earth's atmosphere is 101.3 kpa and Mars'
atmosphere is 0.60 kpa on average. That's 0.59% partial pressure..
Mars' atmosphere is 96.5% CO2 - so the partial pressure of CO2 on Mars
is 0.57% of Earth's atmospheric pressure. Earth's atmosphere on the
other hand is 0.0384% of teh total, which makes Mars' atmosphere
nearly 15x more abundant in CO2 than Earth's even at the low pressure
found naturally. So, why pressurize for plants? To get liquid
water! Water cannot occur as a liquid at the pressures found on
Mars. Triple the atmospheric pressure, without filtering and you have
45x more CO2 than found on Earth and sufficient pressure to liquefy
water - so plants can transpire it.
Assuming we've got plants growing at this lower pressure, producing
oxygen at the same rate its being consumed by humans,
Pa/Pb = 3 --> ln(Pa/Pb) = 1.1
n = 39.1
T and R same
W = 39.1 * 8.314 * 220 * 1.1 = 78.7 kW
Power is 918 milliwatts!!!!
So, compressing Mars' air 3x under a dome of thin film of PET to grow
plants, and compressing THAT air (enriched in O2 and water) as
described to produce Earth normal pressure and copious water for human
use under a separate dome - is easily achieved with very little power
indeed.
To heat the air is simple. Compress it and use a heat exchanger to
warm the discharged air after separation. It is only needed to raise
the temperature to 300K which gives the ratio of buffer gases needed
for a given insulating factor for the PET film. A layer of stagnant
CO2 trapped in inflatable layers, with an appropriate IR reflector
should keep the temperature within a comfortable zone with very little
external power input.
If there isn't a major error in calculation, it seems with a very
modest setup we could recharge our air and water supplies for the
return journey home using quite modest power supplies - and with
seeds, perhaps even FOOD supplies as well!!
Obviously we'd want to try this all out a few times before shipping
crews out lightly provisioned, but it seems doable! We could cut our
supplies in half, or alternatively, increase our crew by 3x, or even
5x if we left emigres there a few seasons.
The fuel is something else. The S-II stage I described earlier has
875,000 pounds of liquid oxygen liquid hydrogen. This requires 9,000
GJ of energy. The stay time on Mars might be 2.5 years if we wait a
full synodic period before returning. Especially if we can 'live of
the land' so to speak. This means we need a power supply of 115 kW
and 400 kiloliters of water. This water is most easily extracted
from the atmosphere.
With 500 W/m2 and 2 kWh/m2/da - we need 1,380 sq m of solar collectors
to provide this power reliably. That's 460 of my 4ft x 8ft panels. A
'half' string of panels at 550 - fit in an 8ft x 12ft x 24ft volume
and mass 11,000 lbs. Replace the water optics with mirror optics, and
minimal water for cooling in the lower concentration and lower
temperature mars atmosphere - further, replace the copper foil with
MEMs based klystron emitters, and a simple phase delay system - and
the panel delivers microwave power to a 'power tower' very very
reliably - and the mass is reduced to about 2,200 lbs.
The whole setup for 60 people, aluminum coated PET film, solar panels,
seeds, compressors, molecular sieves, etc, etc, etc. masses less than
22,000 lbs - and you have over 280,000 lbs ranging up to 800,000 lbs
capacity!!
So, day 1 you deploy the solar panels and the PET domes along with
compressors and air lines and so forth. Then, you charge up the
atmosphere within the domes in two weeks - by running the compressors
and whatnot at 7x the rate they're normally run to maintain oxygen and
water levels. Even so, water is electrolyzed at 80% the normal rate.
Then, air processing is cut back after pressure is attained - and fuel
processing is increased. Still there are huge quantities of reserves
in case there's a leak for examplei in one of the domes - increasing
loss rates 7x above normal, or if there's a cold snap and more heat
has to be generated by overpumping the air, and so forth.
It seems to me if the power requirements are so minimal, only a few
watts. A kg of hydrogen contains 143 MJ of energy. Combined with 8
kg of oxygen in a fuel cell it produces 114 MJ of electrical energy.
That's enough to supply 215 days worth of oxygen in a suit, along with
the 8 kg of oxygen needed to burn the hydrogen. As a side benefit,
you also get drinking water for the duration.
So, in a suit application, you have something very interesting without
screwing around with solar panels - for early voyages before the whole
living off the land is fully developed.
With a 100:1 buffer gas - you double the hydrogen consumption to build
up a houseful of air and then recharge it for the same period. It
seems that 5 kg of hydrogen gas burned in 40 kg of extracted
oxygenshould be enough to provide Earth normal air and drinking water
from the Mars atmosphere for a 2.5 year stay on Mars. .
Willie...@gmail.com
any reference to meaningful numbers.
Check this out;
Chemical Rocket
Isp = 465 sec ---> Ve = 4.558 km/sec
Ion Rocket
Isp = 5,000 sec ---> 49.011 km/sec
dV(Earth/Mars) = 3 km/sec (once on orbit)
dV(Mars Capture) = aerobrake to orbit
dV(Mars/Earth) = 1 km/sec (in Mars orbit)
dV(Earth Capture) = aerobrake to orbit
dV)Total) = 4 km/sec
Propellant fraction
u = 1 - 1/exp(Vf/Ve)
u(chemical) = 58.43%
u(ion) = 7.83%
Thrust to power
F = 2 * P / Ve
Specific thrust to power
So, each kiloNewton of thrust is
2.28 MW/kN chemical
22.55 MW/kN ion
Thrust to weight
Chemicall = 70:1
Ion = depends on power supply specific weight
22 MW/kg ---> 100:1
10 MW/kg ---> 45:1 <-- competitive with chemical
1 MW/kg --> 4.5:1
100 kW/kg --> 1:2.2
10 kW/kg --> 1:22 <--VASIMIR hi
1 kW/kg --> 1:220 <--VSIMIR lo
220 W/kg ---> 1:1000 <---typical hi
100 W/kg --> 1:2200
22 W/kg --> 1:10,000 <--typical lo
10 W/kg ---> 1:22,000
ENGINE FRACTION
Say we had a nuclear power supply that generate 10 kilowatts per kg
(including radiator) - and that we operated at 1/100th gee
throughout. Well, the engine by itself produces 1/22nd gee, with no
added mass, so this means the engine is 28.21% of the total vehicle
mass.
So the stage weight using ion breaks deown as;
28.21% ion engine
7.83% propellant
63.96% balance of system
With 3.96% - structure - this leaves 60% useful payload.
For the chemical rocket we have
1.43% - to lift entire vehicle in 1 gee
0.48% - to lift entire vehicle in Mars gravity
0.0143% - 1/100th gee (same as Ion rocket)
A low gee force means lightweight tanks and whatnot. The SIVB
accelerated at 1 gee and had a structural fraction of 5% or so
(including engine). SSTO program sought to achieve for a launcher
less than 5% structural fracxtion. In orbit, we can say that 4% is
easily achievable - even with aerobraking to orbit.
4.00% - engine + structure
58.43% - propellant
37.57% - useful payload
This is 62.62% of the ion rocket The ion rocket provides a 59.71%
improvement in payload to Mars. At $10,000 per kg to orbit this is a
savings, but is it the best way to go?
VEHICLE COST:
4% structure+ engine - chemical @ $100,000/kg
40% structure + engine - ion @ $100,000/kg
The ion rocket, upper stage is likely to cost AT LEAST 10x as much as
the chemical upper stage. It is also going to take 10x as long, and
very likely be 10x more risky. IT costs more to develop a vehicle
than to fly it. So, to obtain a real program savings requires 10
flights to Mars which likely will take at least 22 years to realize!
To achieve any real savings requires nearly 50 years and the savings
are modest.
HIGHEST BEST USE OF RESOURCES
A 37% reduction in the cost of LAUNCH - provides the SAME SAVINGS as
operating an ion engine at extremely high efficiency. WHAT WOULD THIS
COST? The payload on orbit is about 4% the GLOW of a chemical
rocket. That is, the launcher masses 25x as much as the payload. In
a two stage system this is the first stage is 20x and the upper stage
is 4x the mass of the payload.
This means that spending 4x the money spent on the chemical payload
(versus 10x the money for Ion rocket) if spent on the upper stage,
could provide a totally redsigned upper stage.
That is, for less than half the cost of a ion rocket upper stage, we
can redsign an upper stage of a conventional booster, for total
reusability - and if we cut costs by 37% - we are ahead on a cost
basis.
It is less costly less risky, and provides greater advantage to build
a fully reusable launcher and reduce costs MORE than 37% - which is
the point of the common interplanetary booster!
That is, if your goal is to get to MArs and back as quickly and as
inexpensively and as safely as possible - then reusable chemical upper
stages of adequate size, launched by reusable chemical booster stages
of adequate size - is the answer.
If we want to minimize even those costs, we can take existing stages,
like the EPC H173 cryogenic main stage and replace the two SRB (EAP
P241) with LRBs built around a modified core booster- each with 5
Vulcain engines - and cross - feeding to the core booster- and re-
design them all for recovery. Enhanced TPS and fold away wings for
both elements. Downrange recovery for the Cryogenic Main Stage. A
recoverable ESC A H 14.4 stage. - built along the lines of the DC-X -
and a reusable stage connection hardware. so the stages go together
like freight train cars.
This is likely to cost less than development of an ION rocket stages,
and likely to reduce costs to about 10% that of a throw away Ariane.
So, for the same money we get 10X as much payload on orbit (10
launches at the same cost as 1 expendable- i.e. 210,000 kg )- rather
than a 4,000 kg increase to Mars (8,000 kg with a chemical stage, vs.
12,000 kg with an ion stage) at the same cost.
That is, the development cost of the ion stage and flying it once will
cost as much as the improvements in the booster and flying it 10x at a
reduced cost!
So, a straight chemical stage 8,000 kg to mars and back
Ion stage program - 12,000 kg to mars and back
improved booster program - chemical upper stage - 80,000 kg
Obviously at this stage you get more bang for your buck spending your
dollars on launcher improvements and sticking with a chemical stage.
Later, when powersats are operating, then we can consider more
advanced propulsion.
BradGuth
With 99.9% of humanity broke and/or starving as can be, you'll also
have access to having as many servitude minions as you'd like. It
seems your New World Order will be the good life, at least for the
upper most 0.1%.
~ BG
BradGuth
Is our perpetual William Mook, aka Williamknowsbest, willie.moo,
Willie.Moo, William.M, tomcat, wizard of Oz and so forth, dead?
Hopefully he's simply off making himself a whole lot richer on behalf
of creating all that green and cheap hydrogen plus better synfuels
from coal without further polluting our environment.
~ BG
> >http://www.astronautix.com/craft/proelena.htmhttp://www.astronautix.c...
> ...
>
> read more ยป
BradGuth
JFK did not just go and die on us, but instead was systematically
whacked because he was getting ready to pull those ARPA(DARPA) and
NASA plugs before it was too late. We're talking about 10,000 plus
high level and/or cushy civil service jobs and nearly countless others
with nifty benefits that were put at risk as long as JFK was alive and
intending to pull their plugs.
~ BG