That page is about the flight itself and what will be after the end of the
competition.
Posts in backward chronological order.
ON MARCH 14, 2013 05:13 PM Aluminum, again.
With full interaction 27g Al (1 mol) with H2O to form the amorphous Al(OH)3 and H2 by reaction produces 418 kJ of heat. According to calculations by the stoichiometric equation of the chemical reaction for the complete oxidation of 27g (1 mol) of aluminum requires 54g (3 mol) of H2O, with proportion mass of water to the mass of aluminum in 2 times: H2O: Al = 54:27 = 2 : 1: reaction:
2Al+6H2O3 = 2Al(OH)3 + 3H2
H2
or
2Al + 3H2O = Al2O3 + 3H2
Released at the same time the amount of heat, in the absence of cooling, it is enough to heat the reaction product to a temperature of 2300°C. (based on in
http://www.nanometer.ru/2008/04/23/nanoporoshok_aluminia_48221.html).
On http://altinfoyg.ru/index.php/rashot/rachotidei/pva.html for comparison - 1kg of aluminum can give 0.11kr of hydrogen with volume 1.24m3 and it is equivalent of a (to compare with gasoline 46mJ/kg) 0.296kg of gasoline. All this articles was about methods how to produce hydrogen using fine aluminum and the water.
Interesting table on page 103 of
http://www.duskyrobin.com/tpu/2007-01-00025.pdf Introducing more water makes process more controllable.
Victor can you check (just curious!). If, instead of H2O, it is mixture of 50%H2O and 50%H2O2 (or may be in another proportion), than: 2H2O2 = 2H2O + O2, and 2Al + 3H2O = Al2O3 + 3H2, and combustion of hydrogen and oxygen also adds released energy. Under normal temperature all components are stable (well, some sort of for H2O2), but if cylinder shaped piece of aluminum will be ignited and mixture will be delivered to a place of burning via cooling channels(inside aluminum cylinder). Then process of burning can be controllable (ignore Ignition problem for now). Where I am wrong?
JANUARY 04, 2013 03:52 PM Al2O3, SiO2.
Thanks Boris, in vacuum under temperature Ti2O3 on C
surface will be perfectly react with outcome of Ti and CO. Layer of Ti can be
formed on top of layer of a C (carbon particles, nanotubes, graphene, just name
it). Question how to recycle C – as a first task to catch CO, and second
question is to separate CO (as I understand that can be done by chlorella –
living creature does it perfectly).
And thanks for tip how to make experiments – Clay is Al2O3 + SiO2 — needs to grind to smallest as
possible particles. Dried as much as possible. Place dust into a vacuum chamber
and try to separate by electrostatics. Low grade of vacuum (1mm) actually will
help to charge dust particle. 6 time bigger gravitation force can be
compensated. If separation can be achieved – well, it will be good.
DECEMBER 28, 2012 05:24 PM Aluminum. 3D printing.
Thanks, V. Existing, formulas about aluminum extraction process in vacuum and
1/6 earth gravity.
- Conductive aluminum surface as forming layer
for 3D printing part.
- Al2O3 distributed by electrostatic on
top of forming layer.
- K3ALF6 as a catalyst, distributed by
PVD (physical vapor deposition) precisely on a place of a next slice of a
printed part.
- Beam of electrons on the same square area: Al3+ + 3 e− → 2Al ; O2− - 2 e−→ O ; K3ALF6evaporates
and can be reused.
In vacuum do no need for a carbon electrode – oxygen will be separated without
chemical reaction. Also electrical current in vacuum better be done by emitted
electrons. Surface of particles are big for unit mass. In 1/6 gravity different
fractions does not separates properly, but it will be layer of Al2O3 , isn’t it? Is it possible to
reuse K3ALF6? How to collect O2 effectively?
Well, bicycle as it is. Legend from “Historia naturalis” about Tiberius and
goldsmith, brought imperator “silver” looking, light, metal’s plate. On
imperator’s question: “How did you make it?” Inventor replied: “From clay.” On
next question: “Who else does know the process?” Inventor proudly said: “Only Me
and Gods”. It was a mistake, to protect his recently investments in silver,
imperator ordered to cut of inventor’s head. Nice, unconfirmed story was, until
couple years ago some students got aluminum by technology and chemical
components available 2000 years ago only. Again, bicycle’s invention is a
bicycle invention. Needs to do this on The Moon, not on the earth! Anyway we do
not know investments portfolio of other people.
What equipment will be needed to make such experiments?
DECEMBER 18, 2012 01:56 AM Possible scientific experiment?
Well, Victor – you right, absolutely - two visible way today – in 2000 was
experiments to make Al from Al2O3 in hydrogen plasma (in some blog reference to
UDK 541.14http://hbar.phys.msu.ru/gorm/forum/index.php?t=msg&th=2433&goto=50817&S=8dc2fbb39a516315e8538348f84b2e64#msg_50817 experiments was done in Krasnoiarsk) ,
and in 2006 was theoretical study on formula2Al2O3+3C=4AL+3CO2 (http://www.itp.nsc.ru/Laboratory/LAB_2_1/papers/5.pdf ) again in plasma, was done estimates:
12-14 kWt for 1kg of aluminum at 2400K temperature. Both methods require ether
hydrogen or carbon. Carbon is preferable to bring from earth. In second method
some amount of a carbon react with aluminum and creates Al4C3, that will require
carbon extraction from carbide. Second way requires less energy. Recycle CO2-> C
+ O2 also will be needed. What do you think ?– can we formulate the real task to
investigate instead of just bla-bla - “How to restore aluminum from 2Al2O3 → 4Al
+3O2 under vacuum conditions and lunar environment”, or to forget all this for a
first stage and to try separation of Al metal from a lunar dust (some amount
already present in moon rock's samples).
Experiment
in that case can involve extraction, by static, Al particles, and distribute it
to collector to form a wire, then evaporation of the formed wire can create on a
concave (or flat, or prism formed) surface some reflector mirror. Simple, light
and effective experiment (instead of bringing laser reflector or optical device
from the earth will be possible to print it on the moon surface).
Dec 7, 2012.
Bla-Bla-Bla continues. Victor – ...With 3D printing on the moon
first step will be to separate fractions of a lunar dust. For such porpoise can
be used low gravity, vacuum conditions and old vacuum tubes technology.
Cathode's material under the heat (as it was in tubes) emits electrons. Anode on
some distance from cathode, over regolith (I believe long time ago was an idea
to use cathode-anode as a solar battery and theoretical efficiency was pretty
high, and heavy) accelerates electrons and part of electrons beamed to a surface
of regolith, after some period of charging, charging process stopped, and
another plate (charges negatively) placed over regolith, acting as a capacitor’s
plate (all of this is from 10 grade physics’ practice book, I believe). Depend
of a voltage on a plate, smaller fractions of duct will levitates from a
surface, and distributes by weight vertically. Last step - another positive
charge applying by tore shaped collector will discharge particles of regolith.
Lowing or elevating collector (or by changing voltage of the “capacitor” plate -
preferably) will select different fractions. Actually it does not matter how to
harvest fraction - static charges, and known formulas can apply. If some
fraction will be suitable to make practical object (glass will be first to try)
the same principle can apply to a way of delivery dust particle to a surface of
printing layers (it will be less mechanical parts in 3D printer).
Surprisingly some of old “school’s” physics bring modern interests – for example
charged, same sign, 0.1-100 micron particles, in some conditions can attract
each other instead of commonly believed action.
http://journals.ioffe.ru/jtf/2010/05/p75-79.pdf
Compact version of useful formulas are in:
http://genphys1.phys.spbu.ru/People/Karasev/docs/meth2.pdf
Effects on charged dust particles on Moon -
http://nuclphys.sinp.msu.ru/school/s10/10_01.pdf
Effects on charged dust particles on equipment:
http://144.206.159.178/ft/7938/739975/13859267.pdf
Plasma with injected 5-60 mkm particles:
http://www.ebiblioteka.lt/resursai/Uzsienio%20leidiniai/ioffe/ztf/2004/11/ztf_t74v11_24.pdf
Despite of your “contra” about complication of metallurgy process on the moon –
and you correct point that this is not our priority today - it is known that
aluminum is like “accumulator”, it accumulates energy. Circle – (a) day time
producing Al from Al2O3 and (b) night time producing from Fe2O3 +2Al -> heat +
2Fe0 + Al2O3 solves a problem - technological process can work all time without
nuclear source.
Surpluses oxygen with aluminum by itself is perfect solid (Al) + liquid (O2)
rocket’s fuel+oxidizer.
Can you answer me how that aluminothermic reaction will work not on a tank’s
plates welding but on a kainda micro-level – in vacuum, 25 -50 micron size Al
and similar size Fe2O3 intersect each other at specific point where controlled
heat can apply? It will be nice to make such experiment, today on earth, in big
vacuum chamber – but no time, – can you answer what it will be at least
theoretically?
And with 3d printing - look what did Markus Kauser :
http://www.thisiscolossal.com/2011/06/markus-kayser-builds-a-solar-powered-3d-printer-that-prints-glass-from-sand-and-a-sun-powered-laser-cutter/
Dec 6, 2012.
for 3D printing. OK, Victor - in Lunar soil samples:
Plagioclase 30%-35% compositions NaAlSi3O8 to CaAl2Si2O8 , melting point
1100—1550 °C.
Pyroxenes 54%-60% usually for earth that is formula XY(Si,Al)2O6 (X = calcium,
sodium, iron+2 , magnesium, rarely zinc, lithium; Y = in smaller size chromium,
aluminium, manganese, scandium, titanium, vanadium, iron+2), with melting point
1300°C.
Olivine 3%-8% formula (Mg,Fe)2SiO4 melting point Mg2SiO4 = 1897°C.
Ilmenite (absent in Lunar highland areas) 18%-2% formula (FeTiO3). Melting point
1400°C.
I understand that to get Ti, Fe, or Al from oxides needs to do all chemical
reaction. And standard way does not work, no CO, no H2, no C, no H2O. To get
Iron from Fe2O3 it is possible to use Aluminum: Fe+32O3 +2Al -> 2Fe0 + Al2O3,
and not sure about Ti. Looks like way to start will be Aluminum, with molten
oxide electrolysis (it is 1200C) - will requare a lot of electrisity - but side
effect – it will be 40% of the oxygen!
http://en.wikipedia.org/wiki/File:Moon_vs_earth_composition.svg . I am
not talking about reaction like Fe+32O3 +2Al it surplus heat at night time!
With 3D printing on the moon first step will be to separate fractions of a lunar
dust.
Nov 16 2012.
3D printed CPUs on the moon? Well, first processors for IBM
S/360 computers was done by interesting technology, CPU by itself was a BOOK(!),
with conductors laminated into a plastic pages. Current introduced by input
inductors around BOOK (sorry CPU) creates tiny current on output conductors on
pages, and that current does all processor/logics job.
Yes, cooling was major problem (efficiency) to be resolved.
Yes, to change microcode (operations performed by CPU) needs to open CPU’s BOOK,
remove old pages, and insert new.
Yes, it was difficult to design such computer (USSR for example just copycat
existing pages from S/360), and etc. problems.
But today it is possible to 3D print such BOOK (CPU) as one device from
titanium, debug it, tested it (CPU for example can support system of command of
a specific, exsisting processor) and then simply re-print CPU on the moon.
Yep – speed can be slow. Well, size will be big, but old technology can help to
make self-suntanned base on the Moon. All computers required for the mission can
be built on the moon except initial amount delivered from the earth.
The same can be done for amplifiers devices required for a
communications/input/output units – vacuum and dust on the moon is free,
emissions of electrons works same way under solar radiation (instead of heat)
and old century’s triodes, pentodes are perfect to replicate by current 3D
printing technology. Why tubes? – Well solar radiation can kill electronics, but
tubes will be OK! Story of Kemurdjian's team (Lunokxod-1,2 designer) show
interesting twist about tubes – after Chernobil explosions, on roof of the
reactor to clean radioactive debris was tried different rovers. All advanced
robotics failed in first minutes – and only one was working – on tubes – it was
a rover from Kemurdjian's team.
Below are components from
two year old design.
Proposed Flight schema.
1. Launch on a “low earth” orbit can be done via commercial vehicle or as additional
payload form regular government’s funded space launch. Sample is amateur’s radio
satellites, which has low weight and was launched as by-product of commercial /
government satellite’s launches. Payment for launch must be done as a requirements
for participation in competition. Orbit can be pre-calculated but likely be unpredictable.
Parameters of orbit after successful launch are to be made independently (from a
launcher company) calculated and verified by GPS module(s).
2. Up to 5-20 circulation will required the check of all on-board equipment/systems
and for preparation for low-orbit high-orbit manoeuvre. Crucial systems functionality
has to be confirmed: astro-orientation, orbit parameter's calculation, communication,
image delivery, data transfer.
3. “Low-orbit” to “high-orbit” transfer has to be done via two impulses. High orbit
has to be achieved because of unknowing parameters of low-orbit, resulting in an
unknowing waiting time (up to one month) before flight to the moon. Impulses can
be roughly calculated and preformed by less precision high-thrust engines. High
orbit has to be achieved to avoid atmosphere influence on waiting orbit, manipulators.
4. “High-orbit” to “waiting-orbit” transfer must be done via low-thrust precision
engine. The engine needs to be fired up a couple of times to delivered impulses
for orbit correction. Without rushing all system functionality needs to be tested
at “waiting-orbit”.
5. “Waiting-orbit” to “flight-orbit” transfer must be done via one impulse less
precision high-thrust engine. After impulse used engine’s parts are to be disconnected
or ejected. All command for orbit correction has to be verified before this manoeuvre.
6. At “Earth-moon orbit” intensive orbit corrections must be performed. All correction
must be done by low-thrust engine.
7. Brake impulse can be preset and constant at design stage, plus-minus variation
verified at testing stage. “Earth-moon orbit” correction together with intensive
calculations on mission control system, and on-board computers to delivery probe
in specific point, with specific velocity in sun-earth-moon celestial point. Astro-orientation
system must set impulse vector based on desired landing place. Brake impulse should
slow probe enough to allow for soft-landing (air-bag based) system to adsorb impact
of a probe in the lunar surface. Before soft lading all unused parts and engines
need to be disconnected /ejected. Probe manipulators / components need to be placed
into a landing position.
8. Soft-landing system (air-bag(s)) will somehow adsorb impact and deliver probe
without some / serious damage to the moon. Air-bags will deflate and probe will
be oriented on surface.
9. Check of all probe system. Broken probe’s components have to be detected. Communication
with mission control has to be performed. Backup systems activated. Decision made
on travel ability on terrain.
10. Travel / filming / data transferring according X PRIZE rules. If possible travel
500M and attempt to survive a night.
Astro-orientation and gyro-orientation system design and requirements.
It is impossible to do a mission without the ability to orient probe in flight,
to calculate position, orientation, center of gravity and speed of a probe. All
this is a task for orientation system. Modern technologies allow the use of compact
gyroscopes to detect probe's X-Y-Z axes orientation, cameras to make pictures of
starts, infrared sensible sensors (cameras) to calculate celestial’s bodies horizons,
computers to make calculation and predict location of a probe. Antenna's manipulator's
controls (motors / rotation motion) can be used to change orientation of a probe.
Different type gyroscopes are to be investigate by parameters and tested on different
testing platforms. Results of testing should give requirements for gyro-platform
design, which include temperature restriction, functionality, location, precision,
data bandwidth requirements. Precision of a collected data should give requirements
for high-thrust engine and low-thrust engine. For example: delays in gyroscopes
sensor’s calculation and delays in orientation’s controls motors can create unstable
thrust’s vectors, and compensation by reducing thrust must be apply.
Performance in vacuum, in different temperature conditions, under stress of
vibration, at landing impact have to be tested. Performance of different lenses
(plastic, glass, diaphragms) for different wavelength has to be performed. Power
consumptions, assembling capabilities, supporting electronics have to investigate
and consider in requirements for probe’s imaging system. Data capabilities / compression
capabilities of supporting electronics has to give requirements for Communication
system and Data Transfer system.
Position, orientation, speed, centre of gravity of a probe must be stored in 3 separated
(considered prime) location on a probe. This should include computer's memory and
non voltage memory. Verification, calculation and correction of position + orientation
can be update by astro-orientation, gyro-orientation system and from mission control.
History data of position and orientation enough for on board calculation should
be stored in on board computers and can be update/corrected by mission control.
Probe’s orientation on all stages must be able to execute 4 different tasks:
- centre of gravity positioning, for orientation at main impulses engine’s firing;
- centre of gravity positioning, for changing vector of thrust for low-thrust engine;
- probe’s rotation in axes X-Y-Z before engines firing;
- probe’s rotation in axes X-Y-Z of a probe's manipulator(s), antenna(s), tools,
engines components positioning.
All orientation commands have to be done by (a) main on board computer; (b) backup
on board computer; (c) mission control. In an absence of communication with mission
control all pre-programmed orientations targets for all flight + on moon mission
will be executed. Mission control can specify source for control devices either
with on-board computers, or mission control itself, all settings for all control
devices must time out to execute, first-in-last-out fixed size command stack with
default (top-of-stack) command's values.
Low-engine's system requirements and design.
It will require two type's of engines for the project. First type is high-thrust
engine with preset impulse. Second type is low-thrust orbit correction engine.
Design and development of any rocket engines is a tricky part of the project. There
are no theoretical way to design engine, calculations can done only to justify prototype's
development. Testing prototypes can prove intuition's decisions and calculations,
and engines (especially experimental prototypes) can / will have the tendency to
explode.
It is possible that safety rules will prohibit design of some engine's components
and engine itself. Security rules, government regulations, permits can be obstacle
for development / design also.
When dealing with high-thrust engines it is preferable to use something already
developed and available on the market. Solid state rocket motors, are used in modern
fireworks can satisfied all parameters and consideration. All task for solid state
rocket engine's use in this case will be to select suppliers, order rocket's motors,
confirm impulses, confirm precision of impulses, design containers, design ejection
/ disconnection system.
Concentration to develop low-thrust engine for orbit correction can be main priority
and task. It is unknown now, on what principles, and what parameters will be required
for those engine. It is unknown, whether there will be time and resources to design
and develop a low-thrust engine itself. The main goal is to “be safety” when designing,
“safety” in developing, and “safety” in production. Success in this task will be
crucial for project, scientific research can bring inventions, and commercial use
can be beneficiary for future business. Using vacuum chamber can help in this development.
All another testing platforms can help.
Based on simple calculation of sun, earth, and moon gravity on distances from 0.2
to 0.95 earth-to-moon range forces will be in range 8N to 0.34N for a probe with
mass 100Kg. Control flight will be enough for engine’s thrust to be in that range.
Communication system.
For communication system frequencies to communicate with probe on a distances ranged
low earth orbit (200-300km), high orbit (>500km), earth-moon flight (< 400,000km),
modulation type, error corrections methods, duplicated frequency, duplicated method
of communication.
Main and backup systems can be 10:1 transfer rate ratio and 5:1 weight ratio. Main
and backup antenna (after one year we still have no idea what will be the
backup antenna). The designed should be able to change its 3D orientation for better
transmitting and receiving data from mission control. Capability for changing orientation
of antenna can be used to change orientation of probe’s on-flight's stages, and
the same capability must be used to allow travel (crawling/creeping) on a moon surface.
Engineering task to develop this functionality will be great to achieve.
All communication for on board devices electronics have to be tested on vacuum,
thermal, vibration platforms. Range testing must perform by reducing transmitting
signal for on board equipment, and reducing transmitting signal for communication
testing platform.
Frequencies of a probe's transmitter can drift because of temperature. Testing for
temperature characteristics of a transmitter must be done and internal temperature
stabilization, temperature measurement of probe's transmitter can be used for reception's
turning. The same must be done for the on board receiver, it internal temperature
are to be recorded and used to turn frequencies of earth located transmitters.
As on Google Lunar X PRIZE web site announcement SETI radio telescopes can be used
for receiving and may transmit data to and from probe. Communication with SETI radio
array system is to be designed, developed and tested. Needs (a) convert data from
SETI radio array into internal data; (b) convert data for transmitting to frequencies
of signals acceptable by SETI radio array system.
Separate independent earth's located transmitter / receiver system have to be developed.
If it will be impossible to transmit data to probe from some distance then flight
by itself will continue in automatic mode.
Actually there is no guaranty that from first launch will be achieved long range
communication, In this case all failures in communication system has to be analyzed,
system can be redesigned and a second launch should be used to achieve mission goal.
References for low-thrust engine calculations:
http://ru.wikipedia.org/wiki/Солнечная_энергетика
On earth orbit solar energy is 1367 Bt/m². Parabolic reflector with area 1m² (Radius
= 0.314m) and focusing area 0.01m² will give energy 136.7 KBt/m².
http://en.wikipedia.org/wiki/Thermolysis
http://en.wikipedia.org/wiki/High-temperature_electrolysis
Theoretically it is require 141.86 M joules to produce 1 Kg of Hydrogen in high
temperature electrolysis. Water by itself will separate to oxygen and hydrogen at
the temperature over 2000C.
http://en.wikipedia.org/wiki/Rocket_engine
http://en.wikipedia.org/wiki/Bipropellant_rocket
For H2+LOX liquid rocket engine’s parameters in vacuum with optimum expansion from
68.05 Atm to 0 Atm and Nozzle Area = 40:1 are:
r = 4.83 Mixture ratio: mass oxidizer / mass hydrogen.
Ve = 4462 Average exhaust velocity, m/s. The same measure as specific impulse in
different units, numerically equal to specific impulse in N·s/kg.
C* = 2386 Characteristic velocity, m/s. Equal to chamber pressure multiplied by
throat area, divided by mass flow rate. Used to check experimental rocket's combustion
efficiency.
Tc = 2978 Chamber temperature, °C.
d = 0.32 Bulk density of fuel and oxidizer, g/cm³.
Basically high temp electrolysis is reverse process for burning two water’s components.
1367 Bt energy enough to produce around 0.01gm of Hydrogen and 0.0483 gm of Oxygen.
That mean apply Solar energy in vacuum via reflector on area 0.01m² and heat 0.0583
gm water will be equivalent of an Hydrogen + Oxygen engine. For 0.06gm/sec propellant
use this will give a (theoretically) trust = 4462m/s*0.00006Kg = 0.26N. To compare
existing ion thrusters can give 0.018-0.024N (ref:
http://ru.wikipedia.org/wiki/Электрический_ракетный_двигатель).
In reality 68.05 Atm pressure have to be achieved in low-engine – lowing pressure
will lowing trust.
http://en.wikipedia.org/wiki/Graphite
http://en.wikipedia.org/wiki/Carbon
and
http://ru.wikipedia.org/wiki/Углерод
http://ru.wikipedia.org/wiki/Теплопроводность
Engine chamber’s temperature = 2978 also hard to achieve. It is possible to use
graphite can help to build engine’s parts. Melting point of a carbon is 3820K. It’s
thermal conductivity (for variety forms of carbons) is better then water.
Achieving at least 0.1N thrust (38% from theoretical value) with probe’s propellant
mass =20kg and 0.06 gm/s mass flow will give ability to correct “waiting-orbit”
and “earth-moon” orbit it total time 20kg / 0.00006kg/s = 5555 min = 92 hours =
3.8 days. Assuming propellant = water 20% of probe’s mass this will give net (theoretical)
trust 4462m/s*20Kg = 98240N. With 38% of a theoretical value it will be 37331N.
Design and development of this kind of engine can be done this way:
-
Design and build reflectors: light wait, probably other side of solar panel coated
with reflective material, sizes in proportions solar energy outsize of atmosphere
to solar energy at average sunny day in Vancouver. In-flight reflector will be smaller
then required for testing. Different size reflectors can used to achieve different
temperature in combustion chamber.
-
Design and build different heat adsorption engines with different way of heating
propellant.
-
Design vacuum operated remote controllable valves, propellant’s hydraulics / injectors
components, and propellant’s tank.
-
Design propellant tank’s temperature stabilization, or use liquids like alcogol
with low cristalizationtemperature.
-
Testing prototype’s hydraulics / components in vacuum chamber, thermal chambers
to confirm its functionality.
-
Testing prototype’s engines for 2-3 different temperature’s values in vacuum and
air to get data for characteristic (Pressure, mass flow, temperature in combustion
chamber) approximation.
-
Testing prototype in air with reflector to approximate vacuum’s performance.
-
Based on test’s results change design or make decision for total net thrust tests
and usability of engine.