Below are components from one 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:
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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.
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Design vacuum operated remote controllable valves, propellant’s hydraulics / injectors
components, and propellant’s tank.
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Design propellant tank’s temperature stabilization, or use liquids like alcogol
with low cristalizationtemperature.
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Testing prototype’s hydraulics / components in vacuum chamber, thermal chambers
to confirm its functionality.
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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.
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Testing prototype in air with reflector to approximate vacuum’s performance.
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Based on test’s results change design or make decision for total net thrust tests
and usability of engine.