May 12, 2011. Two
interesting articles about plasma propulsion thruster. One is autoreferat by
Sysoev Denis Vacheslavovich -
http://www.google.ca/url?sa=t&source=web&cd=11&ved=0CBYQFjAAOAo&url=http%3A%2F%2Fwww.mai.ru%2Fupload%2Fiblock%2F7c3%2F7c348b547a3efd385f6e6d8d1ea47e05.doc&ei=9mbNTeXMA4mosAOP0qSzCw&usg=AFQjCNGNiZmiiXXA-OKFyeAZ8Fk-tjaZDA
And another is article by Morozov from БЭС -
http://andriuha077.narod.ru/cad/plasmu.html . Text is kainda not readable
(black background) – to read copy to clipboard and then past to editor.
Looks like it is possible to make this device:
Liquid propellant enriched with lithium will be delivered by micro pump into a
micro chamber via capillary. At engine’s chamber, impulse of a laser beam will
ionize propellant, make it gaseous and via micro nozzle plasma will be expanded
and injected into a second chamber. At time of injection radial current from
cathode (steel road) to anode (aluminum tube) will create magnetic field which
will push plasma out of tube, making more acceleration then just regular
expansion of a propellant via nozzle.
It will be interesting to make experiments in a vacuum chamber. Needs to have –
quartz glass plate, etched steel plate, 5-6 laser diodes from DVD writes
(strongly – no experiments without eyes protection – permanent damage to retina
guaranteed – better to do all experiments inside optically sealed vacuum
chamber), lenses from same DVD writers, aluminum tube, capacitors 10-50 Farads,
electronics. Nozzles 0.2 mm (better smaller – no idea how to make). Yes – micro
pump required – again no idea what type – but with delivery of a 0.001g of
liquid per cycle.
3 experiments using torsion balance:
a) measure force of expansion of liquid propellant via nozzle – the basic
value;
b) measure force of expansion of liquid propellant with laser beam ionized
propellant in gaseous form.
c) measure force of impulse with plasma self pushing out of a can.
I believe all including torsion balance can be fitted into my favorite cooking
pod – sorry - vacuum chamber.
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Old:
From 2009 investigation was driven that low-thrust engine will be based on
graphite body, capable to heat by reflector from direct sunlight. Reflector will
be fixed on a craft’s frame, craft’s guidance system will orient a probe to
focus sunlight on a graphite engine’s body. Then probe will be orient low-thrust
engine to desired direction and small portion of a propellant will applied to
achieve presize impulse.
Here the materials from 2009 investigation.
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.
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.