Re: Mining the Moons of Mars

And yet our physically dark as coal Selene/moon is forever taboo/
nondisclosure rated, with more excluded evidence than our mutually
perpetrated cold-war, JFK and TWA flight 800 combined.

We can't even establish a platform of science instruments within
Selene L1. Is that pathetic, or what?

~ BG


On Jan 5, 11:44 am, willie.moo...@xxxxxxxxx wrote:
On Jan 5, 9:03 am, Ian Parker <ianpark...@xxxxxxxxx> wrote:

http://www.agu.org/pubs/crossref/1989/89GL00073.shtml

Should anyone want to go there in preferance to Mars there should be
plenty of water. You might even generate the hydrogen/oxygen required
for a Martian landing. Let's have an open mind! We could at a pinch
explore Mars by telepresence.

  - Ian Parker

Here is a photograph of the martian moon that's causing this
speculation;

http://upload.wikimedia.org/wikipedia/en/5/59/Stickney_mro.jpg

Definitely lots of interesting things that *might* be there.  You've
got something like 10 trillion metric tons of materials.  Loosely
compacted.  Easily accessible.  Bathed in sunlight 24/7 at about the
same intensity you get on Earth's surface.   So, that's very
interesting

You do need a means to mine the moon and process it for the materials
you want while handling the materials you don't want.

While it is true you may use aerobraking to enter orbit around Mars
and then make use of resources - if they can be found - to make
hydrogen and oxygen from water resources on one or both of the moons -
to land and return, and even refuel for departure back to Earth (as
well as supply crews with water and oxygen on orbit) - there isn't
much advantage from doing the Zubrin approach of using aerobraking to
land, and then doing all that on the Mars surface.

In fact, landing on the Mars surface gives you access to Mars'
atmosphere which allows you to process the gases with pumps and
filters and cryo coolers.  A far simpler and less risky process.

Aerobraking to the surface has been done;

http://sirius.bu.edu/aeronomy/withersmericarus2006.pdf

Mars entry speeds are about 5.7 km/sec with a descent angle of 11.5
degrees took 251 seconds.  That's 2.2 gees acceleration (v=at -->  a =
v/t) which is less than what a launch from Earth imposes.

To skip off the Mars atmosphere into Mars orbit requires that 3.6 km/
sec be lost, with a far shallower descent angle, at a lower gee
force.  So that's easy to do too.

So, the big difference is being able to process the moons of mars into
useful stuff easier and more safely and more efficiently than doing
the same thing with Mars atmosphere.

As Zubrin and others show in their works you can aerobrake and land on
Mars empty, or nearly so and then use a nuclear or solar energy source
to extract water vapor from the Mars atmosphere, and break that water
down into oxygen and hydrogen.  CO2 may be combined with H2 brought
from Earth, producing CH4 and H2O - and that water  broken back into 2
H2 and O2  from 2 H2O.

    4 H2 +  CO2 --->   CH4 + 2 H2O
    2 H2O + energy -->   2 H2 + O2

In the end, 2 H2 is converted to CH4 and O2 - which is far more
massive than the H2 brought along.
                    4 amu                  16 amu  32 amu
                                                   48 amu total

So, each ton of hydrogen is converted into 12 tons of propellant by
absorbing 11 tons of CO2 from the martian atmosphere.

http://www.space.com/businesstechnology/070507_methane_rocket.htmlhttp://www.space-travel.com/reports/ATK_Test_Fires_Liquid_Oxygen_Meth...

A 3.5 km/sec exhaust speed, combined with a 5.7 km/sec transfer
velocity from the Mars surface so, this is a velocity ratio of 5.7/3.5
which means a propellant fraction of 80.4%.  With an 8.6% structural
fraction - this leaves 11.0% payload fraction.     But 6.7% of that
must be hydrogen - leaving 4.3% payload - which must include the power
plant, and atmospheric processing - but part of that could be left
behind.

Using hydrogen and extracting both hydrogen and oxygen from water
vapor present in Mars' atmosphere - provides a 4.5 km/sec exhaust
speed, produces a propellant fraction of 71.9%.  With the same 8.6%
structural fraction - 19.5% payload fraction - a far simpler system,
but more power is needed.

Landing on Diemos or Phobos after aerobraking - reduces aerobraking
requirements slightly - lowering gee forces from 2.2 gees to 1.2
gees.  Adds complexity of processing solids in the vacuum of space.
Adds the benefit of solar energy available at higher intensity 24/7.
To deorbit from the moons requires 2.6 km/sec delta vee capability.
With a 4.5 km/sec exhaust speed this produces a propellant fraction of
43.9% - with the same 8.6% structural fraction - leaves 47.5% payload
fraction.

The propellant processing station - can be left on orbit during
landing - the lander can be use over and over again to visit several
sites, while the orbiting station can reconnoiter sites - to provide
planetary coverage in a single mission.

Finally, the 3.6 km/sec delta vee requirement to escape from Mars
orbit to Earth transfer - using hydrogen/oxygen is provided by a
propellant fraction of 55.1% - with the same 8.6% structural fraction
- leaves 36% payload fraction.

So, a 100 ton payload arriving on the mars surface requires the
following departing mars, which gives the scale of teh ship and power
supply needed;

Approach                   Payload  Vehicle     Energy
Zubrin  direct descent     4.3% 2,326 tons     577 MWh
HyZ     direct descent   11.0%    909 tons  5,656 MWh
orbit     phobos/diemos 36.0%    278 tons  1,729 MWh
surf      phobos diemos 47.5%    106 tons*    660 MWh**

*NOTE:  The surface vehicle is assumed to be 50 tons empty, while the
entire payload is 100 tons
*NOTE:  The energy for the surface vehicle is for landing and take off
only.

A primary energy system - either nuclear or solar - large enough for
the HyZ approach would allow 6 landings and take offs at six different
sites on Mars while the return stage was being refilled.

So there is are some things to recommend this approach.

Of course a bimodal nuclear rocket engine - improves things greatly!

http://www.grc.nasa.gov/WWW/RT/2004/PB/PBM-mcguire.html

Here was have a self contained power supply for transit, as well as
for use on orbit around Mars.  We also have a nuclear thermal rocket
capable of generating 8.5 km/sec to 9.5 km/sec exhaust speeds.  That
means far less propellant is needed to depart mars orbit.    Using 8.5
km/sec exhaust speed and 3.6 km/sec delta vee to leave mars orbit,
requires a propellant fraction of 34.6%   Increasing structural
fraction to 15.4% - leaves 50% payload fraction throughout.  Here we
assume half the payload - 50 tons is the same chemically powered
lander.

Using the bimodal nuclear engine to land on mars creates a nuclear
radiation hazard that presumably is taken care of in space by
approaching and departing the nuclear stage along shadow lines of the
gamma shield used in transit.

Here, despite large increase in structural fraction due to the nuclear
thermal engine, and shield, we only need 200 tons on departure and 100
tons of propellant - 622 MWh of energy which means that at the same
power level as the HyZ mission above, 7 to 8 landings may be made with
a chemical stage.

Aerobraking at Earth, with this system at mission end is also
possible.  The nuclear component enters high orbit after aerobraking -
as it does in Mars.  The lander then descends to land on Earth with
crew and retrieved materials.

To reuse the system, the lander is refilled and relaunched, along with
164 tons of liquid hydrogen.  A total of 288 tons of payload.  This is
about half the capacity of a super-heavy lift launcher massing 5,000
tons at lift off - I have designed to support the deployment of solar-
power satellites.  Obviously, such a launcher would easily support the
return of two vehicles every synodic period.

The heavy launcher would also deploy a single launch to mars at the
outset - six launches deploying a fleet of six vehicles.  This gives
sufficient backup to assure crew survival in the event of mechanical
failure of any one two or three critical systems.  Then after the
first mission, the six vehicles are reprovisioned with three launches
before the next launch window opens.

This approach provides a low cost means - especially if power
satellite revenues are taxed to support space operations beyond Earth
-  to maintain a continuous presence on Mars.

The bimodal engine is really a redesign of the 1950s era Rover/Nerva
program.

http://www.fas.org/nuke/space/c04rover.htm

With full support of military and intelligence agencies this could be
redeployed for about $6 billion today.  The full-scale heavy lift
launcher, would cost another $7 billion.  The Mars lander, and deep
space transfer technologies, related to Mars and deep space operations
would cost another $7 billion - a total of $30 billion - if approached
efficiently (not using current method of contractor management now
used by NASA).

I could be completed within 5 years (the next two synodic periods) and
result in a heavy lift launcher capable of supporting power satellites
(the powersats themselves are developed and paid for by commercial
interests, the launchers are built owned and operated by NASA in this
example, and rented to the powersat builders to cover costs
of deep space exploration stages and their operation by NASA and NSF
and universities)

http://sci.tech-archive.net/Archive/sci.space.policy/2008-06/msg00005...http://www.astronautix.com/lvs/searagon.htm

Return to the moon and expansion of the space station are a natural
consequence of this as well - paid for ultimately from power sales on
orbit.

Commercial and exploration success, leads naturally to even LARGER
launchers, and larger payloads, and more capable systems

http://sci.tech-archive.net/pdf/Archive/sci.space.policy/2008-08/msg0...

* * *

Here's more general information on the martian moons;

http://en.wikipedia.org/wiki/Phobos_(moon)http://en.wikipedia.org/wiki/Deimos_(moon)

PHOBOS (Fear)
Dimensions      26.8 × 22.4 × 18.4 km
Mean radius     11.1 km  (0.002 1 Earths)
Surface area    ~6 100 km²  (11.9 µEarths)
Volume  5 680 km³  (5.0 nEarths)
Mass    1.072×10^16 kg (1.8 nEarths)
Mean density    1.887 g/cm³
Equatorial surface gravity      0.008 4–0.001 9 m/s²
...

read more »

.

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