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Project, Ideas, Flight, Competition, and Next Steps.

Click for main mission description.In Russian.

Click for test mission description.

Click for concept for mission support.

Click for Metallurgy on Lunar surface.

Click for 3D printing on the moon.

Click for (Old) 2010 Proposed Flight schema.

Click for (Old) 2010 Attude control design and requirements.

Click for (Old) 2010 Communication system design and requirements.

 

 

Main mission description

The primary mission will be done by a direct flight to the Moon from a trans- geostationary orbit. Sample of such orbit is the Astra satellite. That transfer orbit characterized as elliptic with the perigee close to the Earth and the apogee touching the geostationary orbit's circle. Depend on the orientation of the ellipse will needs to wait around 1 month to a correct position of the Sun and the Moon to make the impulse to reach the Moon by a direct flight. After separation from the tag platform, an automatic system check will take place, resulting in an “READY” message being sent over a backup communications link to the mission control, to acknowledg that the craft is functional.

picture of the craft (without payload adapter).

Next, the craft will start an orientation sequence. The sequence is using a solar sensor for a search of the Sun. The vector (calculated direction) to the Sun from the gyro platform data will be stored for a future use. As a backup the orientation to Sun will be determined by the maximum current of one of the solar panels (the craft turns its solar panel to the Sun, to maximize the power harvest). The craft uses an “adaptive attitude” system for all maneuvers. The principle for a “adaptive attitude" control system is based on calculations of 125 possible movements, performed by combination of stepper motor's steps, and a reaction of resulting quaternions rotations recodred and stored for a future use. Those values allow to precisely controlling the craft's rotation (direction), and to optimize movements to reach orientation during flight.

The process to stop undesirable rotations after separation from a tag-platform, or after "sleep" mode, will be performed by a calculating the period of the function of the solar sensor's current. Algorithmically to stop rotation, (which is equal to the maximization of a measured period of the luminosity function), will be attempted for movements with biggest possible rotation angles from 125 possible quaternions.

In uncontrollable rotation case the communication will be possible on the perigee part of the orbit via the backup communication. At that moment recorded GPS/GALELEO raw data will be transferred back to the mission control. The continuous recording of the raw signal is not necessary, and will be need to get 5-8 records from any navigation satellite for 30-60 minutes interval. Analyzing that data in the mission control will allow determining the trajectory, by the software running in the distributed calculations mode. If the GPS system's raw data could be not be available, then a raw stream of digitized data on L1 frequency will be stored, with a later extraction of a raw navigational satellites coordinates / velocities from stored recording. That step can be done by the signal's extraction software on board. After determining the orbit parameters, mission control initiates the command for the attitude control to find a vector of the direction to the Earth.

The craft then starts rotating itself with some constant speed, and infrared sensors, detects crossing the edge of Earth. Each moment of the crossing of the edge will be a signal to record the quaternion of the orientation of the craft (provided by gyro-platform). The direction to the center of the Earth is determined via onboard calculation.

The adaptive attitude system is equipped with a special mechanism for tracking and performing rotations without processing the data from the gyro-platform. Rotation can be initiated from any still position. Reversed set of the commands repeated in backward order from a previous sequence of rotation movements..

To confirm the period of the orbit, the craft perform rotation to conform direction to Earth edge. It is not desirable to keep a constant track of the direction to the center of Earth. After detection the center of the Earth and the detection the direction to the Sun, mission control can send the sequence of orientation maneuvers for the communication session's orientation (on 2.4GHz frequency). The session’s orientation commands include vector's values, stored in onboard system, and time marks for each vectors. The attitude control linearly extrapolates vectors in the time between sequential time marks. It is the task of the mission control to split vectors of orientation to make possible linear interpolation between time marks. At the session will be transferred telemetry, and data from the imagining system.

In the communication session (on 2.4 GHz), additional measurements data will be collected. It is the traveling time for a RF signal from the ground station to the craft and the signal traveling time from the craft to the ground station. Two sequential “ping" measurements can be used in the orbit determination.

From that moment, all efforts will be concentrated on finding proper trajectory from trans-geostationary trajectory (see sample) to the target on the lunar surface. Preliminary calculations about possible orbits will be performed before a launch.

As much as possible data from GPS (main) and from GPS/GALILEO front-end RF system needs to be recorded. GPS satellites are flying at distances of 5 Earth diameters and their transmitted signal is beaming toward the Earth direction. Ideal will be records of the last (before leaving beams) data. That, far from the Earth, recordings will give a max accuracy of the orbit determination. Because the craft will separates from the tag platform, the orbit of a platform will be tracked by independent measurements, and the conformation of the orbit’s parameters can be cross verified.

It is expected 2-3 weeks delay before main trans-lunar impulse. Another method of an orbit determination will be performed to measure 3 directions, each in different time - to the Sun, to the center of the Earth (it will be good visibility on trans-geostationary trajectory), and to the center of the Moon.

Also is considered to get measurements (and times), of 3 points of the orbit. Repeating the process can give another 3 coordinates. On the way to the Moon the communication session become longer and ground station locates around the globe. The window of session with 2 ground stations simultaneously can be important in the orbit’s determination. Such events are planned with the pairs in Donetsk and Kazakhstan, as well in Hawaii and Cook Island. These two ground stations pairs have near perpendicular connecting lines. Which will allows the measure RF signal’s time to estimate the orbit with better precision.

The word "orbit" means “flying around another celestial body without collision”, and it assumes period, inclination, and 5 others values. In simulation software, the orbit can be calculated from the position and the velocity (2 vectors), those vectors are the prime source for all calculation. On the main burn accelerometer will record a real performed impulse. The error in impulse's calculations will be less than error in position measurement by RF signal travel time. In trajectory calculation / orbit's determination software the velocity will be less volatile. “Distributed mode” of calculations will speed up processing of series of measurements.

The main impulse (sends craft to the Moon) defines a landing point and imperfections in impulse can influence on the landing target.

   Picture of a main fixed impulse engine.

For the prime lending point was chosen coordinates= 2S15E on the lunar surface. That is highland area with steady inclination. Landing in a radius of 1km from the point will give steady sloops to travel 500m in any directions. 30km on east from the exact S2E15 there are two interesting sites - Theon Junior, and Theon Senior. Both landmarks are 18-19km in the diameter, with the rim to the floor 3km difference and with the length of the slope - 5km. Shifting landing point to the rim of Theon Junior (S2.1312 E15.7745) mission will have ability to travel distance 4-5 km in short period of time (8 minutes) and HD video can give detailed pictures of a geological structures on all way from the rim to the floor.

Picture of a desired landing agrea S2.1312 E15.7745

The calculated error is around 600m on latitude and 300m on longitude. In the case of the flight with help of a gravitation of the sun and the moon the landing time will be at the lunar night. The best precision (minimum error) at targeting the landing point is depends on a “lunar” daytime. If a daytime is close to sunrise, then the error will be bigger, and if a daytime is close to "lunar midnight" then the precision is better. The ideal will be landing at the Sunrise line (terminator).

On a trans-lunar trajectory the orientation of the craft become the challenge: the direction to the Sun can be determined correctly. Two maneuvers will be performed every half an hour: to take a picture of the Moon and the Earth. An imagining data then will be transferred to the Earth for a trans-lunar trajectory’s conformation. The picture (the Sun at this moment will perfectly eliminate the Earth) of the Earth with the orientation vector. At mission control adjustment for vector can be calculated based on images and that correction will be delivered to a craft for delta correction. Gyro-platform to perform such operation needs to be capable to track “drift of zero” with minimum precision - 0.5 degree per 30 min.

Risk assessment at this stage of the mission – if main impulse will not bring craft to the lunar surface (the missing trajectory) then the brake impulse will be used to correct the error and send the craft into a collision with the Moon. That will be done to reduce space debris on the earth orbit. That scenario is not a perfect mission ending, but at least hockey puck will be delivered to the moon.

Brake impulse is performed at the last moments before the landing on the Moon. The Moon will be close to the craft and the command’s sequence will be: the detection of the edge of the Moon (by 2 infrared sensors - values gives the direction to the center of the moon), the calculation of the direction to the Sun (by solar sensor). After two vectors measured executed command to orient the craft to the direction of the (solid state) engine’s firing. The craft will be rotated (by attitude control with the hockey puck embeded) to stabilize in the time of firing. Exact ignition time will be determined by a laser range finder. In a study provided in 2011, ignition point was around 24km from the lunar surface at 4 sec before impact. The precision of the ignition (in study) was 10ms, and the ignition error brought 300m error above the lunar surface. Burn of the last engine will take around 20 seconds, and at that time (engine still burns), it ignite the pyro-bolts to separate the engine shell from the landing craft and impact adsorbdion shield (6 kg in mass). Rotation of the rover and the impact shield will be compensated (stopping the rotation) by the carbon fiber spring, released on separation.

picture of a brake fixed impulse engine.

Calculations performed in 2011 showed that from a random chosen, low earth orbit, landing point on the moon can have the error in longitude - 600 m, in latitude - 300 m. In study the expected landing time was 6-24 hours before Sun rise.

Before the ignition of the brake engine, the HD camera starts to record a video. 3 previous (pre-landing) recordings (length 1 min each) will be done to observe lunar approach. Recording will stop after 5 minutes, and that will cover the entire process of the landing, 4 minutes before impact and 1 minute after. In the case of "before sunrise" landing, to record the performance of the impact shield, the illumination by LED will be turned on.

From the moment of the landing, the mission will be moved to a mobility phase.

 

 

[translation to Russian - it is a pain to express ideas in time strustual language - I'll better right information convient for me way - anyway - nobody will read it:]

 Концепция главного полета к Луне в рамках Гугл Лунар ИКСПРАЙЗ.

Главная попытка достичь Луны будет осуществлена по таектории прямого прелета к Луне с орбиты переходной к геостационарной. Кредит у такой схеме прелета принадлежит Александру Михалову. Пример такой орбиты - вывод на геостационарную орбиту коммуникационного спутника Astra в ноябре 2013 года. Запуски спутников на геостацинарные орбиты проводятся регулярно, с установишимя маркетом, и как следствие цены за килограмм нагрузки более стабильны, чем цены при прелете с низкой околоземной орбиты. Преходная геостационарная орбита характерезуется элипсом с точкой касания в апогее окружности геостацинарных орбит. Перигей орбиты находится близко к земле ниже спутников глобального позиционированияю. Так-как  основной нагрузкой в полете является геостационарный спутник, то отделение луного модуля возможно после отработки двигателя, выводящего общюю нагрузку на преходную к геостационарной орбиту. С этого момента лунный модуль может находиться в свободном полете. Ориентация элипса орбиты не гарантирует немедленную возможность достич Луны. Необходимо ждать относительной позиции Луны и Солнца, чтобы выбрать момент для достижения цели. Использования фиксированного импульса позволяет упростить и удешевить лунный модуль, но с другой стороны требует точного расчета имульса и точной ориетировки лунного модуля в момент прожигания двигателя. Такая схема по предварительной прикидке требует от 30 до 45 кг общей массы лунного модуля при доставке к луне 4 килограммового лунохода с 2 килограммовым посадочным демпфером. Коммуникация с лунным модулем на участке перигея возможна на всенаправленную антену (планируется использовать существующий спутниковый коммуникационный модем). А на участке апогея коммуникация будет осуществляться в 2.4ГГц диапазоне на хеликовскую или уменьшенного диаметра поляризованную хеликовскую антенну.

На рисунке приведен пакет лунного модуля с луноходом, многослойным демпфером, двумя твердотопливными двигателями. В сопле тормозного двигателя смонтированный лазерный дальномер и маховое колесо (специально для Канады выбрана хокейная шайба закрепленная на шаговом двигателе. Все 3х мерные модели лунного модуля доступны на - https://github.com/alexdobrianski/ROVER6.

Ориентация лунного модуля осуществляется на базе гиро-платформы и активных вращающихся маховиков. Главный маховик установлен на оси ценрта тяжести всего модуля в сопле тормозного двигателя. Два других маховика - это колеса лунохода. Требемые вращения для ориентации вычисляеются на основе кватернионной матрицы получаемой экспериментальным путем в полете. Матрица заполняется на основе возможных 5 х 5 х 5 =125 комбинация моментов создаваемых тремя двигателями. У каждого двигателя 5 возможный движений - (а) по часовой с фиксированным моментом вращения, (б) - по часовой с 1/2 фиксированного момента (в) - неподвижно (г) - против часовой с 1/2 фиксированного момента (д) против часовой с фиксированным моментом.

Гироплатформа использует попарное количество твердотельных гироскопов. Оси каждой пары взаимно компенсированны. Направлене оси +Z одного совпадает с -Z направлением оси другого, ось +X первого соответсвует направлению оси -Y второго. Пара из двух гироскопов позволяет обнаружить вращения 0.004 градуса в секунду, а компенсация по осям позволила в 6 часовом прогоне правильно определить напрваление на полярную звезду с точностью до 30 гарадусов (что составляет ошибке 5 градусов в час). Потребляемая мощность одной пары составляла 0.1 ватта. Схема позволяет увеличивать точность гиро-платформы за счет увеличения числа пар гироскопов.

Гироплатформа также содержит 3-х мерный акселерометр, достаточный для оценки направления к центру Луны (луноход на поверхности), чувствительностью 0.032g. Акселерометер позоляет замерять импульс работающего твердотопливного двигателя.

В оринтации лунного модуля используется три ифракрасных сенсора для вычисления направления на солнце и детектирование земного или лунного горизонта. Детектирование горизонта двумя сенсорами позволяет найти направление на центр земли или луны. В качестве запасного способа солнечной ориентации используется солнечные батареи.

Для остановки вращения лунного модуля, после отделения или после слепого полета, анализируется на периодичность сила тока генерируемая солнечными элементами. Период соотвествует периоду вращения модуля. В то же время вращение, зафиксорованное гироплатформой, может быть вычеслено с большой точность с помощью процедуры усреднения, т.к. вращение постоянно вокруг одной оси. Данное значение вращения является калибровочным значением для всей гироплатформы (по тому же принципу как для вычисления дрейфа нуля в на земле может быть использовано вращение земли). Главная и две вспомагательные солнечные батареи расположенны перпендикулярно друг другу так чтобы получить одновременно информацию о направлении на солнце по трем осям. После калибровки гироплатформы на маховики системы ориентции подаются мерные вращения (всего 125) и реакции по изменению вращения записываьюся  в таблицу для последующего использования. Этой информации достаточно чтобы осуществлить остановка вращения с ориентацией в направлении максимальной ЭДС вырабатываемого солнечными элементами. Оригинальная ось вращения является главным моментом инерции лунного модулм записывается для дальнейшего использования.

В сосстоянии неконтролируемого вращения коммуникационная сессия возможна только на перигейном участке арбиты, под коммуникационными спутниками. До сеанса связи (окно каждые 12 часов) модуль собирает данные о своей орбите. Выделенный GPS сигнал, содержащий вектора позиции и скорости навигационного спутника, вместе с точным временем его (сигнала) приема подготавливается к отправке при приближении к перигею. Обычно модем (запасной системы коммуникации)  детектирует сигнал спутника связи и сигнализирует внешнему устройству о возможности начать сеанс. Для отправки подготавливаются 5-6 записей навигацинных спутников, отстоящих по времени одна от другой, с как можно большим интервалом между записями. Это позволяет точнее определить орбиту. Сигнал навигационных спутников предстваляет собой 3 сигнала наложенных один на другой - первый и главный из них это псевдо-случайная, похожая на шум, индивидульная для каждого навигационного спутника последовательность. На него модулируется данные о позиции и скорости спутника. Прием сообщения занимет несколько секунд. Начальный таймстамп вычисляется после приема всего сообщения. Обычно навигационные приемники используют несколько принятых сообщений, чтобы расчитать позицию. Скорость и позиция навигационного спутника представленна в геостационарной ситеме координат. Если приемник движется относительно земли, то система защиты премника (игнорирование сигнала с дорлеровским смещением при пороговой скорости более 500 м с) отказывается обрабатывать такой сигнал. Поэтому на "хороших" спутниках применяют специальные приемнику со снятой защитой (разрешение на такою разработку  или покупку необходимо получать у бог весть кого на юге). Но даже "неправильный" (c системой защиты) навигационный приемник на разных участках орбиты  оказывается в позиции когда доплеровское смещение небольшое. Эта запись не позволяет вычислить координаты, но приемник принятую запись рапортует через сериальный порт по специальному протоколу (ID 29/30). Сколлеционированные в течении нескольких часов такие записи и будут теми данными, передаваемыми по запасному каналу связи на перигейном участке орбиты. Кроме навигационных примеников с записями ID 29/30, на лунном модуле размещена альтернативная ситема приема и обработки сигнала навигационных спутников. Это так называемые RF front-end приемники . Фактически это радиоусилитель с оцифровкой. Частота оцифровки 40 мегагерц, оцифровка 2 битами (что дает хорошую компрессию принятого сигнала). Принятый такой raw (сырой) сигнал записявается со скоростью 10 мегабайт в секунду в оперативную память. Участка записи в 300мегабайт достаточно для последующего выделения возможных сигналов навигационных спутников. Схематика такого устройства достаточно проста, вся сложность в обработке софтвером.  Передача выделенных из сырого сигнала координат навигационных спутников осуществляется на том же перигейном участке орбиты. По такой схеме снимается ограничение на секретность разработки оборудования для приема сигналов навигационных спутников.

 After determining the orbit parameters, mission control initiates the command for the attitude control to find a vector of the direction to the Earth.

The craft then starts rotating itself with some constant speed, and infrared sensors, detects crossing the edge of Earth. Each moment of the crossing of the edge will be a signal to record the quaternion of the orientation of the craft (provided by gyro-platform). The direction to the center of the Earth is determined via onboard calculation.

The adaptive attitude system is equipped with a special mechanism for tracking and performing rotations without processing the data from the gyro-platform. Rotation can be initiated from any still position. Reversed set of the commands repeated in backward order from a previous sequence of rotation movements..

To confirm the period of the orbit, the craft perform rotation to conform direction to Earth edge. It is not desirable to keep a constant track of the direction to the center of Earth. After detection the center of the Earth and the detection the direction to the Sun, mission control can send the sequence of orientation maneuvers for the communication session's orientation (on 2.4GHz frequency). The session’s orientation commands include vector's values, stored in onboard system, and time marks for each vectors. The attitude control linearly extrapolates vectors in the time between sequential time marks. It is the task of the mission control to split vectors of orientation to make possible linear interpolation between time marks. At the session will be transferred telemetry, and data from the imagining system.

In the communication session (on 2.4 GHz), additional measurements data will be collected. It is the traveling time for a RF signal from the ground station to the craft and the signal traveling time from the craft to the ground station. Two sequential “ping" measurements can be used in the orbit determination.

From that moment, all efforts will be concentrated on finding proper trajectory from trans-geostationary trajectory (see sample) to the target on the lunar surface. Preliminary calculations about possible orbits will be performed before a launch.

As much as possible data from GPS (main) and from GPS/GALILEO front-end RF system needs to be recorded. GPS satellites are flying at distances of 5 Earth diameters and their transmitted signal is beaming toward the Earth direction. Ideal will be records of the last (before leaving beams) data. That, far from the Earth, recordings will give a max accuracy of the orbit determination. Because the craft will separates from the tag platform, the orbit of a platform will be tracked by independent measurements, and the conformation of the orbit’s parameters can be cross verified.

It is expected 2-3 weeks delay before main trans-lunar impulse. Another method of an orbit determination will be performed to measure 3 directions, each in different time - to the Sun, to the center of the Earth (it will be good visibility on trans-geostationary trajectory), and to the center of the Moon.

Also is considered to get measurements (and times), of 3 points of the orbit. Repeating the process can give another 3 coordinates. On the way to the Moon the communication session become longer and ground station locates around the globe. The window of session with 2 ground stations simultaneously can be important in the orbit’s determination. Such events are planned with the pairs in Donetsk and Kazakhstan, as well in Hawaii and Cook Island. These two ground stations pairs have near perpendicular connecting lines. Which will allows the measure RF signal’s time to estimate the orbit with better precision.

The word "orbit" means “flying around another celestial body without collision”, and it assumes period, inclination, and 5 others values. In simulation software, the orbit can be calculated from the position and the velocity (2 vectors), those vectors are the prime source for all calculation. On the main burn accelerometer will record a real performed impulse. The error in impulse's calculations will be less than error in position measurement by RF signal travel time. In trajectory calculation / orbit's determination software the velocity will be less volatile. “Distributed mode” of calculations will speed up processing of series of measurements.

The main impulse (sends craft to the Moon) defines a landing point and imperfections in impulse can influence on the landing target.

 

Test mission description

On 2011 team summit one of Google X Prize teams presented partner developing a launch vehicle, capable of delivering a tube-sat payload to a low earth orbit. Even in the case of a failure of such project, spending 14 grand is not the same as to spend millions on a development of a satellite for a testing mission. Before the main mission flight, the decision was made to use this opportunity. For such the test mission, was outlined following key goals:

1. The orientation of a craft/satellite needs to be calculated with the precision of 0.3 degrees. (With an error of 0.3 degrees, the Moon will be missed). If the nano-satellite will be capable to achieve such precision, then it will be a “green light” for a craft to develop.

2. The communication over main and backup communication channels needs to be capable of transferring data to and from ground station(s).

3. The requirements for the satellites testing needs to be verified and passed, prior to placing the craft inside any launch vehicle. That has to be done long before signing launch agreement. Passing ground tests is a good step, especially on an unknown territory.

4. The ground station communication has to be developed and tested. Ideally it will communicate with this nano-satellite on low Earth orbit. And ground station in compact form will be the same rover (to be sent to the Moon).

5. Gyro platform and attitude control must be in developed.

6. Independent ways to determine satellite’s orbit must be verified.

All system on the nano-satellite, on the ground station and on the rover will be the same: a camera unit (2 units looking in opposite directions), the gyro platform unit, the main computer with data storage unit, the Communication unit, the backup communication unit, the power plant unit, the attitude control/orientation unit. If the development of such units and software for it will be successful, then the software can be re-used on a rover/craft for main mission.

For the backup communication was the decision to use existing satellite communication. Modems with such function appeared on the market over the last years. To overcome restrictions (for modem) are temperature requirements needs to place the modem into a sealed epoxy box / compartment. Power requirements on transmitter for main communication system will be 1.5 Watt instead of 10 Watts for the main mission.

Main communication will be on 2.4GHz. That band is designated for a “hopping frequency” of unrestricted communication’s band. As long as the transmitter does not stay on the same frequency channel, no license required to operate, plus no restriction on the antenna construction. The max powers for such transmitters are - in Canada 4Wt or in US - 1Wt. Recently developed OEM amplifiers modules has transition power up to 1-2Watts. The core module for a 2.4 GHz transmit and receive will be regular a Bluetooth RF device RF. The software requires error corrections for long distances. “Standard" methods with repeating of packets, send/receive ACK/NACKs, is not a practical way on lunar distances, as a result it has to be suppressed in hardware implementation.

The power plant designed to harvests as much as possible from the solar panel, and stores energy in 6 capacitors for the test mission. High volume capacitors were chosen instead of batteries because of their better performance in temperature ranges. Temperature range -50C +125C makes high volume capacitors be suitable for the flight. The power plant will check which solar panel performs best, and which unit needs to power most, and will distribute power source to the required unit. Performance of the power plant will be the main “source” of data for a rover's redesign. It was chosen to postpone the development of a power plant to as much as possible. Developments in power harvesting is booming, and waiting period for the flight (just for one year) can allow using new technologies for the mission. There as a risk for such postponing – common practice in the space industry dictates that the design of system has to be frozen for 1-2 years before scheduled flight. Engineers responsible for the integration satellite/craft into a flight vehicle can be reluctant to accept such approach.

The camera module will be a prime element for orientation maneuvers – it will not only deliver pictures, but the software in the main computer can detect the horizon and that detection will calculate the direction to the center of the Earth/Moon. The camera will be stripped from of original optics, and a pin hole will be the main optical element. The quartz glass on the sensor will be drilled to make a hole to avoid distraction in the vacuum.

The orientation and attitude control module on a nano-satellite will have 3 small stepper motors (instead of 4 on the rover, or 2 on a ground station). The task for the modules will be to perform rotation to keep orientation calculated by gyro platform unit. The gyro - platform was the main development in 2011 and is capable to desired presision.

The frame of the satellite will be built within specification. Team Plan B decided to build our own frame: the antenna for a 2.4 Ghz communication has to be deployed. One of our partners, Jetasonic Technologies Inc, a company, based in Coquitlam BC, Canada, designed and manufactured such frame.

For the ground station, one of the tasks will to orient the antenna to the flying satellite (to compare with task for rover to orient antenna to the Earth. The goal is to have the rover the same as the ground station.

The test mission is considered “the must” before attempts to reach the Moon. It will be premature and not serious to talk, to plan, or to build anything before the test mission will be successful.

 

Concept for support:

 

AUGUST 13, 2013 03:14 AM Lunaro Sterling coin. Space money that propel space exploration.


[transcript]
Let's talk today about Lunaro Sterling Coin,which we think can lift Team PlanB, and not only us, to the moon.
If you look closely on "Space progress" in past 60-70 years, what you can see is milestones and breakthroughs were achieved by space enthusiasts like Goddard, Tsiolkovsky, Von Braun, Korolev.
Even today not many people will assign their life, and devote efforts for goals like making Mars habitable for human kind, or to build "space villages", capable to fly to other stars.
Theoretical calculations and stipulations, can be met with low resources. But practical implementation like building controllable explosions to propel payloads into the orbit, requires high resources.
At present and as it was in the last 60 years, nobody will vote -- "pro", to send chunk of money to the sky. The same chunk of money that is equivalent to building a small city.
But each action creates an opposite reaction in this universe, and such practical obstacle, as resources, century ago enthusiasts by-passed them with a smart move.
They promised for men with chevrons on shoulders to deliver dangerous payloads to any place on the globe by simple press of a button.
This decoy in goal worked for enthusiasts. And because of that they were employed, and each minute, every moment, when men in uniform was busy with nasty toys, enthusiasts use this opportunity to reach different targets - to build bigger rocket engine, launch satellite, that left their mark in space, on the moon.
Then, enthusiasm died. "Space progress" has killed the enthusiasm as reaction to massive "action".
Imagine: if the data from "Curiosity rover" would be available to Von Braun and Korolev. Such knowledge will deter any enthusiastic curiosity, because they would see human life on mars an impossible goal.
What choice do enthusiasts have today?,At best, probably, to make a movie. (As we do it now ) It can be done in color, in 3D, better be by a famous film director. And the movie can be watched in Dolby sound, together with popcorn, or on a smart phone on the way to work. Making it perfect sustainable approach.
What else? Crowd-sourcing? On Kikstarter, only one Google Lunar XPRIZE project was funded: "to write the article about the Lunar Google XPRIZE competition", and all other 3 technology-related projects did not reach funding level.
The National Space Agency and military will not support space enthusiasts. In the National Space agency, where enthusiasts are greeted with smile all and every time? Military is satisfied with their "status quo" - they are happy with their offset contracts and support of last century weapons technology.
Skip the business. Business usually interested in cheap "leftover" from space technologies. Even the mining industry with a 10-20 years return of investments period, hesitates to be involved.
It is common sense - if one wants to do something in The Space - it is just, their own problem.
It is essential to invent some independent mechanism to fund "Space devotees", to fund what people can create outside of the Earth surface.
In Team PlanB, we have made our attempt - it is Lunaro Sterling Coin.
First, it is merchandise. We can sell it, the same way as we sell t-shirt.
Second, as a coin, it is valued. It is made from silver. After each 2000 prints, a new design will be made to satisfy criteria: less mintage - more numismatic value.
Third, each coin has a certificate of authenticity provided online. The person owning the coin can set anonymous authentication by himself/herself. The owner of the coin can be changed only by providing such authentication.
More importantly the coin can be used as a barter payments, for service and equipment, outside the Earth surface.
It is similar to the crowd-sourcing approach, but it also keeps doors open for other funding possibilities. Instead of shares investor can purchase coins.
Truly speaking, IT IS "Space money".
In all regulations and laws, it is treated as merchandise. Merchandise on which taxes are paid. Merchandise that cannot can not be prohibited, in any part of the globe, to act as barter's payments for anything not belonging to this world. It is intended to be more than just circulating money, or payment method in Space.
It is designed to lift-off the technology from the earth surface.
"Space money" that does not have national boundaries, and takes advantage on the fact that space does not belong to any national jurisdiction.
The owners of the coins have the power to decide, how, and at what price to sell "Lunaro Sterling"
Team PlanB is ready to share "Lunaro Sterling Coin" design and authentication process with "Space devotees / Space enthusiasts" all over the world.
We do this because we think this is logical, innovative, effective, and “cool” approach to fund space enthusiasts.
What will happen if "Space devotees/ enthusiasts" fail to achieve their objectives and not reach their stars? Well, it will be nothing new in this world. However the coins, will tell the story.
Other updates for last week – communication protocol was ongoing project, (you can visit Github to see and download updates).
RF noise pattern recognition was on hold.
RF Front end prototype for GPS and Galileo system was obtained to develop and to debug positioning software.
Also on Github you can find 3d mold for wheel central holder.
It should be combined technology some elements of the mold (precision) will be made from alumini plastic, some (expendable) from PVA.
We look forward to seeing you again.

 


JUNE 27, 2013 11:11 PM     Modem, Lunaro, RF noise pattern detection.

Lunaro sterling coins database and web interface mock-up. Small program to detect noise patterns in RF.

Выделение паттерна шума из радио сигнала.

Metallurgy on Lunar surface:


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.

  1. Conductive aluminum surface as forming layer for 3D printing part.
  2. Al2O3 distributed by electrostatic on top of forming layer.
  3. K3ALF6 as a catalyst, distributed by PVD (physical vapor deposition) precisely on a place of a next slice of a printed part.
  4. 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:

2010 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.

2010 Astro-orientation and gyro-orientation system (attude control) 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.



2010. Communication system design and requarments.

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.