The Dual-stage 4-grid Ion Thruster
Rémy MERCIER
This is a new esa' ion thruster, THE 'DS4G' THRUSTER, with isp = 19200s
nice!
http://www.esa.int/gsp/ACT/propulsion/safe_test_diaries_wk1.htm
--
Rémy MERCIER
meiza
> This is a new esa' ion thruster, THE 'DS4G' THRUSTER,
> with isp = 19200s
> nice!
> http://www.esa.int/gsp/ACT/propulsion/safe_test_diaries_wk1.htm
Huh, quite a great isp, but where do you need such a thing?
If a vehicle had just one third of it's mass fuel,
this engine could accelerate it to 80 km/s! (In space of
course, and it would take a lot of time.)
A cheat sheet* tells the trip to Mars from LEO needs less
than 5 km/s.
But the power consumption must be big compared to thrust and
total impulse too. The higher the exhaust velocity, the more
energy you actually give to the expelled fuel instead of the
vehicle. So the crafts with these motors would need to have
either big solar cells (makes them heavy, reduces mass
fraction and acceleration), or then really low thrust
requirements (i.e. these would be used on stationkeeping).
Has someone calculated optimal isp:s for certain trips
and ion engines / hall thrusters?
Thrust is proportional to exhaust velocity but power
is proportional to exhaust velocity squared.
There's no point in having a huge isp craft
if it reaches it's good velocity only after 10
years of thrusting, when the other one with
worse isp would already be at the destination.
I guess it boils down to the rocket equation.
Deltav is proportional to both exhaust velocity
and the logarithm of mass ratio.
With low mass ratios (only a bit above 1), the
derivative of the logarithm is close to 1, and
it's easier to grow that than isp.
Acceleration drops then with increasing fuel mass.
If instead isp (v_ex) was grown, and the thrust
kept at constant, the power would grow to the
power of two - the solar arrays balloon faster
than the fuel tank. So acceleration would
drop faster.
Only at high mass ratios does the logarithmic
behaviour begin to show and you hit diminishing
returns. Only then it's wiser to increase isp.
But with this high isp (20 000) and for example a
mass ratio of 2.7 (1 part rocket, 1.7 part fuel),
you'd get the 200 km/s delta v. Where do you need
that?
And how long would even a no-payload thingy
take to accelerate to that speed?
I guess for a trip to Jupiter or further you'd need
nuclear energy, which isn't very power-dense, so you're
better off just having a worse isp and more fuel. (I'd
guess this gives smaller total mass at same triptime.)
Not even trips to asteroids or comets and back, all close
to Sun...
So the main problem with ion engines to me seems to be
*too high* isp!
*) handy "cheat sheet":
http://www.pma.caltech.edu/~chirata/deltav.html
OM
>That's what THEY want you to think. The next thing you know, Ariane
>launches the secret EU deathray.
...Which, being from a French design, will turn us all into
emasculated socialist poofters with cravings for snails, grape juice
posing as fine wine, and a cable network dedicated to showing nothing
but Jerry Lewis films. The plans for this device will, no doubt, be
the contents of the Scroll that the Angel of the Lord currently has
wrapped with the Seventh Seal.
OM
--
]=====================================[
] OMBlog - http://www.io.com/~o_m/omworld [
] Let's face it: Sometimes you *need* [
] an obnoxious opinion in your day! [
]=====================================[
frédéric haessig
"meiza" <me...@inva.com> a écrit dans le message de news:
dnats5$pbb$1...@epityr.hut.fi...
> Rémy MERCIER <Rmy.MERCI...@spacebanter.com> wrote:
>
>> This is a new esa' ion thruster, THE 'DS4G' THRUSTER,
>> with isp = 19200s
>> nice!
>> http://www.esa.int/gsp/ACT/propulsion/safe_test_diaries_wk1.htm
>
> Huh, quite a great isp, but where do you need such a thing?
>
Several possible reasons.
1) lower consumable mass. This means higher useful payload ( for a given
launch vehicule )
2) Continuous thrust transfer. Hohman transfer ( aka minimum energy
transfer ) take a lot of time when you go to target far away. Continuous
thrusting reduces this by a big factor.
3) In-space maneuvering. Either attitude correction or orbit changes. When
you want to visit several celestial objects with the same mission ( like a
planet and several of its moons ) or change orbit around a single object.
...... etc
The usual problem with ION thruster is not Too high an Isp ( as if there was
such a thing ) but too low a thrust level level for some applications. It
seems technology is slowly solving this.
meiza
> "meiza" <me...@inva.com> a écrit dans le message de news:
> dnats5$pbb$1...@epityr.hut.fi...
>> Rémy MERCIER <Rmy.MERCI...@spacebanter.com> wrote:
>>
>>> with isp = 19200s
>> Huh, quite a great isp, but where do you need such a thing?
> Several possible reasons.
> 1) lower consumable mass. This means higher useful payload ( for a given
> launch vehicule )
As I explained, propellant mass ratios get ridiculously low but
the *power source mass* grows by the specific impulse.
(if you keep same thrust.)
Even with isp as low as 3100-3500 s (deep space 1), and a mass ratio
of 1.2, you get over 5 km/s of delta v. That is for example
a 1000 kg dry mass probe with 200 kg of propellant.
Deep Space 1 had total mass of 500 kg, of which used
fuel was 74 kg. (ratio 1.15) It gained 4.3 km/s of
speed when thrusting for over 400 days. The engine
used 2.1 kW of power.
Let's take that probe and do nothing but increase isp
5-fold to 17500.
Use basic Newtonian equations:
P = 0.5*(-dm/dt)*v_ex^2 , power
F = (-dm/dt)*v_ex , thrust
Since power stays the same, we can solve
mass flow:
-dm/dt=2*P/v_ex^2
and thus the thrust is:
F = 2*P/v_ex
So the thrust drops to one fifth.
When the original deep space 1 had thrusted for 400
days and gained 4.3 km/s of speed, our new probe
would have gained less than 1 km/s.
Only after 2000 days (5.5 years), we would be going
that speed. Sure, we would have used fuel only about 10 kg,
but whether it's 10 or 70 kg, that is only a vanishingly small
mass of the whole 500 kg probe.
Our new probe would eventually reach a 20 km/s delta v,
but it would take still 4 times more time, so it would
last in total close to 22 years to reach full speed.
So, one can say, let us increase power so that the high-isp
design can accelerate as fast as the original. But this
asks for 5X the solar cells. The original had two 5m long
"wings". Even if we saved 60 kg in
the form of fuel, it is not enough to make
four more of those for 8 kilowatts. It's hard to get
good figures for achievable solar cell power density, but
one quoted number is 10 kg/kW. This would make the
new cells weigh 80 kg. I don't know about the power
electronics.
> 2) Continuous thrust transfer. Hohman transfer ( aka minimum energy
> transfer ) take a lot of time when you go to target far away. Continuous
> thrusting reduces this by a big factor.
I don't know about this but are you sure you can get to a target
faster by thrusting low for one year compared to getting the same delta-v
in one hour? Or one month? Seems counter-intuitive.
> 3) In-space maneuvering. Either attitude correction or orbit changes. When
> you want to visit several celestial objects with the same mission ( like a
> planet and several of its moons ) or change orbit around a single object.
Deep Space 1 had hydrazine for attitude control (actually running out of that
forced mission end), as did Smart-1.
Ion engines are probably not very good for that (at least directly)... the
turning rate gets very slow, and if your solar arrays point in the wrong
direction and you'd need an attitude correction, you can't use the ion engine
because, uh, your solar arrays are pointing in the wrong direction. :)
If you want to visit several moons in one mission, you're talking Jupiter or
beyond, and then you don't do much with solar cells and have to use lower
power per kg nuclear energy. If you need more delta v, it makes less mass to
put in more fuel than to put more nuclear power and increase isp. Europe
so far hasn't invested in nuclear space power sources.
> The usual problem with ION thruster is not Too high an Isp ( as if there was
> such a thing ) but too low a thrust level level for some applications. It
> seems technology is slowly solving this.
The low thrust is precisely because of high power needed which is precisely
because of the high isp. This is Newtonian physics and it can't be solved any
other way than by higher mass efficiency power sources.
Higher isp in ion engines automatically means less thrust for same power.
There is no way around that.
Why bother optimising for fuel efficiency if your already oversize motor
is weighing more than the fuel tank. Only if you're going for a *really*
long trip so that the fuel tank actually starts to matter.
So that's why I was asking. High (>3000) isp is not useful unless you start
talking about much more than 5 km/s deltavees or alternatively really
low mass power sources.
I think Beppi-Colombo to Mercury will use Hall effect thrusters with
about 1500 isp, like Smart-1.
Does anyone have links or references to ion engine power system masses?
(Solar cells, electronics, thrusters.) It's really hard to find anything
in the net.
--
-meiza
Henry Spencer
>Has someone calculated optimal isp:s for certain trips
>and ion engines / hall thrusters?
Yes. :-) With any system that's energy-limited rather than mass-limited
(and ion thrusters almost invariably qualify), there is an optimum Isp due
to the mass of the power system. Unless the mission is very ambitious,
the optimum Isp typically isn't all that high.
The dominant problem of ion-thruster design, usually, is to get the Isp
*down* and the thrust and efficiency up.
--
spsystems.net is temporarily off the air; | Henry Spencer
mail to henry at zoo.utoronto.ca instead. | he...@spsystems.net
Rémy MERCIER
Here, you can find a new interresting example of the use of ion
thrusters:
http://www.orbitalrecovery.com/cxolev.htm
With the future two stages Hall thruster it will be possible to reduce
(or increase) the isp and simultaneously to increase (or reduce) the
thrust.
In my mind it is quite simple.
My ideal rocket:
first stage: isp=300s
second stage: isp=360s
third stage: isp=450s
fourth stage: isp=500 to 5000 (geo, moon and Mars tug)
fith stage: isp > 10000s (beyond Mars)
--
Rémy MERCIER
Rémy MERCIER
> start
> talking about much more than 5 km/s deltavees or alternatively really
> low mass power sources.
>
> I think Beppi-Colombo to Mercury will use Hall effect thrusters with
> about 1500 isp, like Smart-1.
>
> Does anyone have links or references to ion engine power system
> masses?
> (Solar cells, electronics, thrusters.) It's really hard to find
> anything
> in the net.
>
> --
> -meiza
hydrazine? Why? You can use a lithium battery when your solar arrays
are pointing in the wrong direction.
About Beppi-Colombo I wonder if the next generation Hall thruster will
be ready (dual stage, 6kw and higher isp, near 2000s).
And more interresting, since a few months there is a strong interest
about the HDLT: (Helicon Double Layer Thruster):
http://tinyurl.com/bss28
http://tinyurl.com/a3ga6
http://www.esa.int/gsp/ACT/propulsion/helicon_double_layer.htm
http://www.abc.net.au/catalyst/stories/s1185537.htm
"""Europe so far hasn't invested in nuclear space power sources."""
Yes but there is a CEA team (french) working on this questions for ESA
and the russian could help with their thermoionic technology.
To go far and fast (Mars) we need a great tug: 70% mass = solair cells
and structure, propellant=20% and cargo=10%
--
Rémy MERCIER
Michael Gallagher
<Rmy.MERCI...@spacebanter.com> wrote:
>
>This is a new esa' ion thruster, THE 'DS4G' THRUSTER, with isp = 19200s
>nice!
>http://www.esa.int/gsp/ACT/propulsion/safe_test_diaries_wk1.htm
Classic case of most of the package being taken up by packaging. :)
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meiza
> meiza Wrote:
>> Rémy MERCIER Rmy.MERCI...@spacebanter.com wrote:
>> with isp = 19200s
>>
>> Huh, quite a great isp, but where do you need such a thing?
>...
> To go far and fast (Mars) we need a great tug: 70% mass = solair cells
> and structure, propellant=20% and cargo=10%
I actually did some calculations.
Let's assume we go to Mars and get back and can do that with 20 km/s of
delta v. If you assume a certain solar cell mass efficiency k(I assumed
100 W/kg), and required initial acceleration a_i, you can solve the mass
of the solar cells and the mass of propellant analytically as a function
of exhaust velocity. The rest is then left for the non-propulsion-related
mass which is proportional to payload. I've marked this as m_u. I made
some pictures:
http://users.tkk.fi/~tmaja/ion_optimality/20km.png
With very low acceleration, 36 days for a kilometre a second (going
10 km/s would take thus a whole year, not smart on a mars trip)
demonstrated by the black curve, the optimal exhaust speed is about
10^5 m/s, or in other words, specific impulse is around 10000 s.
Useful mass is a whopping 65% of the whole mass!
Fuel mass fraction would be exp(2*10^(4-5))=exp(0.2)= 1.22 ->
18 percent of total mass would be propellant.
Solar cell mass percentage would be 100% - 18% - 65% = 17%.
With higher accelerations, the isp optimizes to a lower value, as can be
seen from the red and blue curve. So it's not useful to increase solar
cell mass to 70% and put a high isp engine, at least not in this delta v
and with these assumptions.
Isp=2500 is about optimal for a craft which is a quicker accelerater,
only 40 days to 10 km/s. Useful mass would be about ten percent of the
craft.
The next picture shows values for a lesser, 12 km/s deltav.
http://users.tkk.fi/~tmaja/ion_optimality/12km.png
Isp:s optimize even lower.
And so on.
So I actually then maximized useful mass as a function of exhaust
velocity and deltav. I solved for the zero point of the derivative and
used some positivity constraints. Exhaust velocity is hard to get out of
this thing though, so I used a shortcut, using a variable
s = v_ex/deltav.
The optimality points lie on this curve:
s^2 * e^s = 2k*deltav / a_i
Resulting picture is here:
http://users.tkk.fi/~tmaja/ion_optimality/max_r_u.png
Here we can see that for a deltav of 10 km/s (10^4 m/s), an optimal
exhaust velocity would be about 1.3 - 2.5 times the deltav, depending
on required initial acceleration. (higher acceleration, lower isp).
That would mean isp of 1300-2500.
The part where the curves go below 1 is when any isp isn't sufficient to
give that deltav with that acceleration and power source density so the
payload goes negative.
So the original question goes: where is the isp 20000 s or exhaust
velocity of 2*10^5 m/s useful? The answer is here:
http://users.tkk.fi/~tmaja/ion_optimality/100km.png
Only in very low acceleration very high speed probes you get
least total mass with high isp:
Useful mass: 30%
Fuel mass: 40%
Solar cells / nuclear reactor 100W/kg mass: 30%
Final speed: 100 km/s
Isp: 20000 s
Initial acceleration: 1 km/s / 36d
Approximate time to full speed: 10 years.
Is ESA planning something like this? I didn't know at least.
Oort cloud explorer? Or have they invented better solar cells?
The power density assumptions are the least sound here, and
I'm happy to add to my knowledge.
All the Matlab source code is in the directory if you want to
look. I might have done some mistakes.
Rémy MERCIER
Henry Spencer Wrote:
The SMART-1 Hall thruster isp = 1600s and ESA team is working on a
dual-stage Hall thruster to achieve a greater isp. If we want to go
far... unavoidably that will take a long time... and we’ll need a high
isp (the best solution). If we want to go very far... unavoidably that
will take a very long time... and we’ll need a very high isp (the best
solution). The "optimum" isp depend on the mission (how far: distance
and time). But, in fact, they hope to achieve a fully control of a
variable isp and thrust (from 1200 to 3000s).
perfect in that case:
http://tinyurl.com/cfcgm
______________________
wine is better with cheese
--
Rémy MERCIER
Rémy MERCIER
> Here we can see that for a deltav of 10 km/s (10^4 m/s), an optimal
> exhaust velocity would be about 1.3 - 2.5 times the deltav, depending
> on required initial acceleration. (higher acceleration, lower isp).
> That would mean isp of 1300-2500.
>
> The part where the curves go below 1 is when any isp isn't sufficient
> to
> give that deltav with that acceleration and power source density so
> the
> payload goes negative.
>
>
> So the original question goes: where is the isp 20000 s or exhaust
> velocity of 2*10^5 m/s useful? The answer is he
>
> Only in very low acceleration very high speed probes you get
> least total mass with high isp:
> Useful mass: 30%
> Fuel mass: 40%
> Solar cells / nuclear reactor 100W/kg mass: 30%
> Final speed: 100 km/s
> Isp: 20000 s
> Initial acceleration: 1 km/s / 36d
> Approximate time to full speed: 10 years.
>
> Is ESA planning something like this? I didn't know at least.
> Oort cloud explorer? Or have they invented better solar cells?
> The power density assumptions are the least sound here, and
> I'm happy to add to my knowledge.
>
> All the Matlab source code is in the directory if you want to
> look. I might have done some mistakes.
wow!
very interesting. My calculations were wrong. Thank you.
--
Rémy MERCIER
klreed
How about if the solar arrays were made of our new CP1/a-Si:H
Ultra-lightweight solar cells at a power density of 4300 W/kg? Do you
think a 24 kilogram, 814 m², 100 kW space solar array could solve some
of the energy requirement problems for the DS4G?
klreed
Rémy MERCIER
klreed Wrote:
> Hi,
> How about if the solar arrays were made of our new CP1/a-Si:H
> Ultra-lightweight solar cells at a power density of 4300 W/kg? Do you
> think a 24 kilogram, 814 m², 100 kW space solar array could solve some
> of the energy requirement problems for the DS4G?
> klreed
>
hi
do you have some information about "CP1/a-Si:H" ???
--
Rémy MERCIER
klreed
Email me and I will send you pdf file. Pretty much made on 6 micron
thick CP1 polyimide 123 W/m2, 35 m2/kg, 4300 W/kg. Super flexible,
rollable. Needs no ffront coverslide or added protective layers. Needs
to backside reinofrcement for high G-force at launch. Fairly in
expensive to make.
100 kW of CP1/a-Si:H ultra-lightweight solar cells weigh about 24 kg.
klreed
Rémy MERCIER
Page 18: "low cost solar arrays"
http://emits.esa.int/emits-doc/emitsdata/booklets/C2002-11-25.pdf
Maybe you could try to contact Mr Sandberg
--
Rémy MERCIER
klreed
Correction CP1/a-Si:H post:
Needs NO (to) backside reinforcement for high G-force at launch. Fairly
inexpensive to make.
The idea here was that rigid panel solar array, like the Galaxy 3C
wings or Dutch Space rigid arrays, have mass add-ons like front/back
coverslides and backside reinforcement for high G-force space launch.
These add-ons are usually about 4 kg to 5 kg per m2. This is why using
high efficiency solar cells (+ 1000 W/kg) still have a result of 65
W/kg to 100 W/kg in a finished space array.
Even if CP1/a-Si:H were not record power density, and it is a record,
we would save about 1000 W/kg by leaving out these mass add-ons we do
not need for a space solar array.
I had suggested a 100 kW array for something like DS4G because they
have already made and tested a 20-meter thin film deployment boom for
space sails using CP1 with a reflective coating. The same booms could
deploy two 20-meter CP1/a-Si:H solar arrays as two wings and give the
required m2 to generate 100 kW of solar power.
Pictures of the boom tests are here:
http://www.inspacepropulsion.com/news_sail.html
Thanks for the good idea and pdf reference. I sent some information to
Jorgen Sandberg at ESA-ESTEC.
Rémy MERCIER
klreed Wrote:
> Hello Rémy,
>
> Correction CP1/a-Si:H post:
> Needs NO (to) backside reinforcement for high G-force at launch.
> Fairly
> inexpensive to make.
>
> The idea here was that rigid panel solar array, like the Galaxy 3C
> wings or Dutch Space rigid arrays, have mass add-ons like front/back
> coverslides and backside reinforcement for high G-force space launch.
> These add-ons are usually about 4 kg to 5 kg per m2. This is why using
> high efficiency solar cells (+ 1000 W/kg) still have a result of 65
> W/kg to 100 W/kg in a finished space array.
>
> Even if CP1/a-Si:H were not record power density, and it is a record,
> we would save about 1000 W/kg by leaving out these mass add-ons we do
> not need for a space solar array.
>
> I had suggested a 100 kW array for something like DS4G because they
> have already made and tested a 20-meter thin film deployment boom for
> space sails using CP1 with a reflective coating. The same booms could
> deploy two 20-meter CP1/a-Si:H solar arrays as two wings and give the
> required m2 to generate 100 kW of solar power.
>
> Pictures of the boom tests are he
Hi klreed,
intersting questions!
And I've found this:
"Thin film Module Pilot Line and Pre-qualification" Page 60, Mr PEROL
and Pages 61/215/246/247/313
http://emits.esa.int/emits-doc/emitsdata/booklets/C2005-11-9.pdf
@+
--
Rémy MERCIER
klreed
That is a very good find. I will contact Mr. Perol.
What type of business are you in Remy?
Kevin Reed
Brad Guth
http://groups.google.com/group/sci.space.history/browse_frm/thread/83d6304568528eac/7461dcce65a08519#7461dcce65a08519
http://www.esa.int/gsp/ACT/propulsion/safe_test_diaries_wk1.htm
The supposed wizardly topic contributions of:
>Huh, quite a great isp, but where do you need such a thing?
and
>But with this high isp (20,000) and for example a
>mass ratio of 2.7 (1 part rocket, 1.7 part fuel),
>you'd get the 200 km/s delta v. Where do you need
>that?
Such tells us exactly which naysay space-toilet of disinformation these
sorts of folks have to offer. Instead of think-tank we have MOS
wag-the-dog worth of their anti-everything under their sun which orbits
their flat-Earth of naysayers running us amuck.
What if instead of limited to wossy and entirely passive xenon-->ion,
whereas how about their using Rn222-->Ion, plus obviously taking
advantage of any number of other nifty Ra226 and/or Ra228 decay
elements that are clearly already highly reactive to start with, as
proceeding along their way to becoming the likes of lead?
http://www.ead.anl.gov/pub/doc/radium.pdf
That's a great amount of waste-not, want-not that'll keep going to
waste unless perhaps it's put to good Ion thrust usage.
If you wanted a great deal of applied ion thrust at liftoff, try
thinking LRn-->Rn-->Ion
-
Brad Guth
Rémy MERCIER
klreed Wrote:
> Hello Remy,
> That is a very good find. I will contact Mr. Perol.
> What type of business are you in Remy?
> Kevin Reed
>
Hi Kevin, I was inventor (heat pump, new processes, centrifugal heat
pump).
Rémy
--
Rémy MERCIER
Brad Guth
Ra-->LRn-->Rn-->Ion thruster capability, even though it's entirely
capable of becoming a truly impressive isp as well as the most ion
thrust potential/gram, especially considering that Rn is so active and
massive to start with. Perhaps my multitasking VL2-TRACE
station-keeping platform can be the first Radium-->Radon breeder
replenished Rn-->Ion thruster application that'll maintain the
necessary halo-orbit within the Venus-L2 zone for nearly 1600 years
worth, instead of 8 years via xenon. By this time next year, the
Ariane-5 shouldn't have any problems with deploying as much as 10
tonnes worth of this auto-parking package that'll include the necessary
tonnage of Radium and otherwise solar powered VL2-TRACE, as established
into the station-keeping sweet-spot that's continually above their
nighttime season.
BTW; I'm still looking for the best little 2-stage rocket in town
that'll deploy an individual or possibly a few microsatellites around
our moon. All that these have to accomplish is manage to get their
payloads a few km or so past the LL1/ME-L1 point of no-return, and
that's actually not very far away (upon average 316,500 km off our
deck, with good timing it's a whole lot closer yet), whereas from that
point on a very small (meaning slight) directional boost will do quite
nicely at getting my microsatellites into their relatively low orbits.
I'd like to think of having these starting off at not much greater than
25 km off the lunar deck, whereas atmospheric drag (which could be
artificially induced) should bring each of them down to the nasty
surface within a few weeks. At least I can't hardly imagine such
initial low altitude orbits lasting more than a couple of months
(that's at least 960 orbits before dust-breaking or otherwise slamming
into the lunar basalt at something less than 2.4 km/s). Actually, their
final approach and thus orbit termination could be sufficiently remote
controlled with deploying a large area (<100 m2) mylar or whatever
synthetic parachute as part of the final plan of action, that should
get the impact velocity down to a somewhat testy but if all goes well
survivable crash landing of perhaps 240 m/s (I'm also thinking the
larger the volume and thus composite shell area of the 10 kg
microsatellite the better).
-
Brad Guth
Rémy MERCIER
Brad Guth Wrote:
> With regard to the: The Dual-stage 4-grid Ion Thruster
> http://www.esa.int/gsp/ACT/propulsion/safe_test_diaries_wk1.htm
>
> The supposed wizardly topic contributions of:
> Huh, quite a great isp, but where do you need such a thing?
> and
> But with this high isp (20,000) and for example a
> mass ratio of 2.7 (1 part rocket, 1.7 part fuel),
> you'd get the 200 km/s delta v. Where do you need
> that?
> Such tells us exactly which naysay space-toilet of disinformation
> these
> sorts of folks have to offer. Instead of think-tank we have MOS
> wag-the-dog worth of their anti-everything under their sun which
> orbits
> their flat-Earth of naysayers running us amuck.
>
> What if instead of limited to wossy and entirely passive xenon--ion,
> advantage of any number of other nifty Ra226 and/or Ra228 decay
> elements that are clearly already highly reactive to start with, as
> proceeding along their way to becoming the likes of lead?
> http://www.ead.anl.gov/pub/doc/radium.pdf
> That's a great amount of waste-not, want-not that'll keep going to
> waste unless perhaps it's put to good Ion thrust usage.
>
> If you wanted a great deal of applied ion thrust at liftoff, try
> thinking LRn--Rn--Ion
> -
> Brad Guth
Hello Brad,
I have found this very interesting summarize about the future electric
propulsion:
http://www.esa.int/gsp/ACT/propulsion/ultra_ion.htm
Rémy
--
Rémy MERCIER
Brad Guth
Thanks so much. Looks like absolutely terrific Ion thrust information
we could all use.
ADVANCED PROPULSION SYSTEMS
http://www.esa.int/gsp/ACT/propulsion/ultra_ion.htm
I'll do my reading and share upon whatever I can interpret from those
doing what our pathetic NASA should have been accomplishing as of
decades ago.
I'm somewhat wondering as to how much Rn222 we can power-breed out of a
given kg or tonne of Radium?
-
Brad Guth
klreed
as well as the solar cells. Heat pumps might be useful to utilize waste
heat energy from H2 fuel cell components. I have been looking mostly at
carbon aerogel pressurized heat-exchangers,
Kevin
Brad Guth
various nuclear alternatives, having these few words to share upon Ion
thrusters.
>The 1 GW thermal space reactor can be used to place very large payloads
>around the inner solar system. It can also be adapted to produce 200
>MW electrical - to power either a very large space base, OR a
>moderately powerful 20,000 lb thrust ion rocket engine with 5,000 sec
>Isp. (my power levels may be a little off I'm typing this from memory
>and I don't want to recalc them)
Not that my math is always worth a hoot, though I believe that works
out to: 1.0 w = 0.04536 gf (45.36 mgf/joule)
Of course, this is supposedly having been based upon passive/inert
Xenon ions that are quite dead to start with, whereas the likes of
Rn222 ions are already seriously on the go to start with.
What if instead of limited to a finite cash of wossy and entirely
passive Xenon-->Ion, whereas how about using a Radium-->Radon breeder
that's capable of providing nearly an endless supply of Rn222-->Ion,
plus obviously taking advantage of any number of other nifty Ra226
and/or Ra228 decay elements that are clearly already highly reactive to
start with, as proceeding along their way to becoming the likes of lead
that's briefly worth 330 trillion Ci/g, or perhaps an extremely brief
Po212 element that's worthy of 180,000 trillion Ci/g.
http://www.ead.anl.gov/pub/doc/radium.pdf
Seems as though, that's a great amount of waste-not, want-not decay
worthy potential of ions that'll keep going to waste no matters what,
unless perhaps it's put to some good Ion thrust usage along the way.
Therefore, why the hell not employ such a Radium to Radon breeder (aka
Ra-->Rn-->ion), as that's only good for a half life of 1600 years.
How's that for an impressive Isp?
If you wanted a great deal of applied Ion thrust at liftoff, try
thinking LRn-->Rn-->Ion, whereas an unlimited cash of sub-frozen LRn222
would provide all the tonnage the launch phase could possibly get rid
of.
-
Brad Guth
Brad Guth
times)
It seems that a Radon ion is in addition to being initially a whole lot
more massive than Xe to start with, is simply worthy of being
diverted/directed as well as expedited along it's natural path of decay
of having trillions of fold greater reactive inert energy to expend.
This is somewhat like a massive rocket consuming it's inert self along
the way, thus making whatever inert mass somewhat of a non issue until
you run out of rocket.
The Radium element itself is entirely expendable (all 1620 years worth
of it's half-life). Of course timing is everything. Too early and
it's still Radium, too late is going to lose out on whatever's
transpiring too fast for ion thrusters to take advantage of.
Collecting an initial sub-frozen cash of LRn certainly isn't the least
bit of a terrestrial problem. In fact, the collection process and good
riddance of providing an alternative usage for the likes of a somewhat
nasty Uranium-238 byproduct such as Radium, as well as the likes of the
subsequent Rn222 should be thought of as a good thing for having
excluded this from the environment of Earth, that which already has far
more than it's far share of lead poisoning to deal with. In fact, in
some locations we're being gradually killed off and/or having our DNA
mutated by the surplus of the Rn222 element itself. For some reason
New York and the northern half of New Jersey have more than their far
share of Radium-->Radon taking place, therefore the environmental
dosage of lead has got to be measurably greater to boot (no wonder
folks from our NE seem rather unusually dumbfounded, especially those
having lived down wind from all of those local reactors, where such
surplus radon gas has been nearly continually released directly into
the local environment), with oddly Sweden and Finland seemingly being
the most infected indoor lot at < 37,000 pCi/L (they must be the one's
hording Radium). Of course, our Nuclear Regulatory Agency(NRA) as well
as our EPA that's nearly always functioning a good century behind
whatever's needed, as well as local state wizards are usually going to
past the buck by way of offering their usual cover thy butt
infomercials in order to apease their local campers, as per doing
whatever it takes for their own job security or otherwise for getting
reelected (nothing new there).
Certainly a nearly continuous Isp worthy supply of Rn222 as a viable
ion producing cash of rocket fuel is capable of becoming a good thing
for extended space travels, that is as long as the necessary tonnage of
whatever the Radium-->Radon breeder reactor isn't overlooked. Our
somewhat naked moon should be an absolute windfall of radioactive
elements, plus just good old secondary/recoil dosage that's worthy of
having created considerable tonnage of Rn222 on the fly, sort of speak.
Since radioactive atoms in general are considered as unstable, whereas
spontaneous atomic decay produces of all things "Ionizing Radiation",
which subsequently breaks chemical bonds and thereby strips electrons,
as in no kidding folks. Without hardly applying a joule worth of
having to artificially produce ions, whereas it seems you've already
got a natural supply of Rn ions that are going to waste whether you
want them or not. Once Rn222 becomes Po214 is when it's near the end of
it's value for ion thrusting, upon eventually reaching a final decay of
Pb206 is about as useless as elements tend to get.
There are nearly countless studies and reports with redard to such
elements. A search for Radiun/Radon or similar topics will easily
overload your PC as well as your brain in no time at all.
Unfortunately, it seems the really important and/or most interesting
information isn't situated on top of the stack, that is unless you
already know of specifics within a given study/report.
http://en.wikipedia.org/wiki/Radon
http://www.ead.anl.gov/pub/doc/radium.pdf
http://www.wadsworth.org/radon/
http://www.epa.gov/radon/zonemap.html
http://www.csbsju.edu/MNradon/rnmaps.html
http://www.nj.gov/dep/rpp/radon/download/jgainv5.pdf
http://www.epa.gov/radon/zonemap/newjersey.htm
http://www.radonseal.com/radon-facts.htm
http://energy.cr.usgs.gov/radon/georadon/2.html
http://www.doh.wa.gov/ehp/rp/air/factsheets-htm/FactSht3.htm
http://www1.umn.edu/eoh/hazards/hazardssite/radon/radonforskeptics.html
BTW; wherever there's Rn222 there's Ra226 and/or Ra228, which implies
that U-238 is or has been nearby.
What I'm wondering and thus asking a somewhat loaded question upon, is
what's keeping our rocket-scientist from going with Rn222-->ion
thrusting?
What's all of the taboo/nondisclosure or need-to-know fuss all about?
-
Brad Guth