Micro-Space >> Ultralight Manned Spaceflight

[Jim]

NASA made a 3 micro-satellite support structure for Pegasus as part of its ST-5 mission.
http://www.nasa.gov/mission_pages/st-5/main/index.html

No, NASA didn't make it, ST-5 project made it for themselves

[rpspeck]

GLXP Mass Still Growing Smaller

I have just added a near one gram HDTV Camera, capable of meeting the Google Prize requirements, to Micro-Space prototype parts.  Recognizing that the landed mass of a GLXP winner could be as small as One Kilogram, our systems may grow very small indeed!  We have also downsized our minimum Radio Link (for the HDTV transmission) to a few grams.  At this mass, the initial mass in LEO might be as little as 10 to 20 kg, with a launch cost of $100 Thousand (NOT Millions!).

The lightest system will have No Rover.  Since very accurate performance is required for the lander, this craft can easily perform the required displacements as subsequent “Hops”.   Assuming 300 sec ISP for good storable propellants (including Hydrogen Peroxide burned with a high energy fuel), a 500 meter Hop will require less than 1% of the landed mass in usable fuel, with an equal supply for the second landing.  Increasing these to 3.2% of the landed mass, gives a 5000 meter Hop.  Doing both hops, one after the other, totals 8.4% of the landed mass in usable fuel.  For a 1 kg landed mass, this comes to <84 grams of usable fuel, and eliminates the communication problems to a Rover!

[rpspeck]

Progress

My newly purchased HDTV camera actually weights in at 0.4 grams!  It has only a simple lens, not a Zoom model.  For the GLXP imaging it will be necessary to add a small “Converter” lens unit, particularly for telephoto operation (0.3 milliradian per pixel required, or about 3x telephoto effect.)  This small lens unit could be flipped over to produce a wide angle mode, but it is easier to slide in a  wide angle or “Fish Eye” adapter.  In fact, these cameras are so low mass – compared to any reliable mechanical mechanism – that it is better to provide three separate cameras, two with the special optics attached.  Providing three cameras also makes video capture possible – with reduced options – if one camera stops working – nice insurance when $$ Millions have been invested in the flight!

The telephoto optical system will need adjustable focus, for the depth of field at this magnification becomes quite limited.  No aperture adjustment is necessary as the gain control modes on the camera are sufficient.  I envision the Lander “Self Portrait” actually using the adjustable focus telephoto system, and a convex mirror.  A very light convex mirror, held away from the lander on a long arm, could reflect an image of the entire lander.  Viewed by the telephoto optics, this could provide a usable “Portrait”.  Computer image correction could remove distortions and yield a good photo.  Alternatively, an additional camera unit could be mounted on that “arm” and be lighter than a mirror.

The Lunar Navigation “Cameras” I proposed in one of our SBIR submissions, may be desirable to quantify the location of our lander, when it completes its 5km “Hop”.  Several companies are working on “HDR” (High Dynamic Range) video cameras for automotive and other uses.  The military has been working with high sensitivity versions of these for over a decade.  In low light, the “High Dynamic Range” keeps bright lights from masking important details in the glare.  In daylight, glare from the sun produces similar problems.  Photographers quickly learn to avoid these problems, but “full time” cameras – for continuous monitoring – encounter these conditions with some regularity. 

One recently advertised commercial unit promises 110 dB dynamic range: a brightness ratio of 300,000:1.  This makes it possible to image the naked sun, and sun lit Earth or Moon surface details in the same image.  (Roughly a 40,000 brightness ratio, Sun's surface brightness, to white target on the Earth.)  At night, it makes it possible to image both bright stars and the sun lit Earth in the sky.  These combinations are often required for Lunar navigation (if Lunar expeditions will continue after sundown  and on the “Far Side”.)   The worst case situation will still require use of a “Sun Shade”, but that is not a serious problem when accurate navigation is needed to correct for accumulated errors in “Dead Reckoning”.  This extreme occurs in daylight on the Far Side, when only the sun and stars are visible, and the sun – a single point reference – provides a good direction, but a second reference is necessary to compute an absolute position.

[rpspeck]

Refractory Motors

We have begun fabrication work with refractory (high melting temperature) C-103 alloy.  This Columbium (Niobium) alloy is the most common material used to make “Radiation Cooled” (otherwise uncooled) rocket motors.  It is used in all long life spacecraft motors, and was very visible in the webcasts of the SpaceX Falcon 1 launch vehicle.  The entire engine and nozzle of the Falcon 1 second stage are made from this material.  It glowed red during the launch.  This is, however, far below the service temperature limit of this alloy.  It retains usable strength to at least 3000 degrees F (1648 C), well above the “White Hot” melting point of Iron! 

This alloy is an essential replacement for the Stainless Steel in our current Hydrogen Peroxide – Methanol motors.  Switching to high concentration Peroxide will push the combustion temperature much higher than the Steel can handle.  The combustion temperature of the Peroxide fuel is relatively modest (compared to LOX reactions), but the performance is good for a “Storable” propellant.  Peroxide gives Specific Impulse virtually identical to “Nitrogen Tetroxide” combustion with the same fuel component, with low toxicity and lower combustion temperature.

Using this allow will allow us to continue to use our simple, reliable, no moving part motor designs with higher performance fuel.  The high concentration peroxide will allow acceleration to near escape velocity, orbital insertion at the Moon, and landing with an overall ten to one mass ratio:  10 kg landed on the Moon starting with 100 kg in Earth Orbit.  The very low mass of our fuel tanks and control systems make these numbers quite achievable.  I had not initially anticipated that we would be able to win the Google Lunar X PRIZE starting with only 1/10 of the Falcon 1 Payload mass in orbit, but that now seems to be a HIGH estimate of the requirement.       

The C-103 alloy is surprising easy to form and work with.  It resembles a sheet of Stainless Steel, but is as soft as a sheet of annealed Iron.  Reportedly, it can be reliably TIG welded in shielding gas, and we will attempt that soon.

Our development strategy remains “Last Stage First!”.  A deeply embedded “Mantra” in the professional Space Community is that a “Second Stage ALWAYS Costs More than the First Stage.”  There are reasons why this is a tolerable - and possibly optimum situation. However, it is not automatically true, and certainly was not true for the “Juno 1”, used to launch the US “Explorer 1” in 1958.  Given that “Mantra”, only success with the upper stages and guidance systems of our GLXP system will make “Professional” evaluation of our plans meaningful, with proper emphasis on the dominant, “Off the Shelf” launch cost to orbit.  At some point, even visionary investors usually get input from “Professionals”.

[Lampyridae]

Have you finalised your design yet? Is it going to be a pogo lander then?

[rpspeck]

Lunar Light Show

In the process of developing the Laser Altimeter needed in our lander, we realized that equipping our Lander (even our lightest version) with a Laser visible from the Earth is quite feasible!  A fraction of a Watt output red Diode Laser equals many “candle power”.  When this light is formed  into a one milliradian  beam, it equals millions of “Beam Candlepower”, and will be visible from the 400,000 km distance of the Moon.  This beam would illuminate a spot 400 km in diameter on the Earth.  The optimum shape will probably be elongated, 400 km north to south, and 100 km east to west.  Viewers in the right latitude range (selected by laser beam positioning commands sent to the lander) would see the red Laser light for about 3 minutes as the Earth's rotation carried them through the beam.  Flashing of the Laser, as well as its color, would aid in detecting the light visually. 

Our plans call for landing just beyond the “sunrise” line on the waxing Moon.  Only 2 to 3 days of operation will be possible before the initially cold temperature changes to excessively hot, with the sun high above the lander.  Since we do not plan to land in the dark, extremely cold environment preceding sunrise,  viewers will have to spot the Laser against a sunlit portion of the Moon.  This will be easier just after sunset on the Earth, since the Earth's dark sky will produce less veiling glare.  Viewing should be fairly easy with common sizes of amateur telescopes.

The purpose of this demonstration (beyond showing the possibility of optical communication) would be to draw thousands of people into direct participation – seeing the lander's position on the Moon for themselves - and avoiding the “Apollo Controversy” (“They didn't really land on the Moon, because ...”).  Many of those who see a “low budget” lander blinking at them from the Moon, will begin thinking about the possibility of “low budget” Human landings.

[A_M_Swallow]

The media will have to be told where to place their cameramen to film the lander's light.  They will need more than one observation point to allow for cloud cover.

Spinoffs - distress signal for broken down rovers.  If we can see the rover from Earth it is easier to route a rescue vehicle to the location.

In theory the same applies to injured hikers and broken down vehicles on Earth.  We will also need a satellite that can spot the warning signal within a couple of hours.

[rpspeck]

GLXP – Indirect Progresses

Our efforts to develop a Lunar Lander for the Google Lunar X PRIZE have led us into interesting technologies with much wider applications.  For example, our successful efforts to breadboard an adequate Laser Altimeter for our lander have led to Laser Interferometer capabilities with good potential for robotic and machine tool applications, as well as for monitoring precision satellite formations.  Consideration of “Sample Return” options led us to upgrade our hardware for “Autonomous Rendezvous and Docking” including upgrading the demo hardware for our “Electro Magnetic Tractor Beam”.  The later can be used to assemble and maintain satellite formations in orbit with no propellant consumption, as well as to effect precision docking.  We have in fact developed a cluster of interesting technologies for innovative applications in space, all of which can be “Proven in Space” in CubeSat demonstrations.  I will discuss this more in the next few days.

Our ability to deliver advanced hardware in low mass packages was also advanced by our GLXP efforts, and our commitment to pursue that prize with the lowest possible systems mass.  The X PRIZE Foundation's recent provision of access to “Venture Capital” and “Angel Investor” experts was welcome and useful.  These experts made it clear that FUNDING a GLXP flight is probably more difficult than ENGINEERING a successful system.  This input only underscores the Micro-Space focus on pushing the mission cost as low as possible, and soliciting funding for other uses of the technology we are developing.  More later – I have Proposal Deadlines in two days!

[rpspeck]

LTO Systems Progress

Development of the LTO (Lunar Transfer Orbit) System – necessary for Micro-Space GLXP efforts -  is progressing.  This system, offering flight of small, experimental payloads ABOVE  LEO – has attracted attention from NASA and DARPA, since they can envision operational space hardware which will require or benefit from High Orbits.  By radically reducing the cost to place experimental hardware in such orbits, the Micro-Space system can greatly accelerate experiments and operational tests of technologies to be used in Molniya, GEO, Lagrangian and other special orbits.

Our present focus is on adding the control and navigational subsystems necessary for “Upper Stage” operations to our production, “Propulsion Module”, while retaining its high fuel mass fraction. With high density fuel, this module  starts with 40 pounds of propellant, and will have a 4 pound dry mass with all necessary Com, Nav, Thrust Vectoring and RCS components.  With a projected 300 sec ISP, and storable propellants, it can accelerate 16 pounds of payload  from LEO to escape velocity.  A reduced payload will allow interplanetary trajectories. 

Starting with a cluster of three “Propulsion Modules” in LEO (a capability always planned for these units) – and using our “Deep Throttling Parallel Stage” strategy -  at least 50 pounds of mass can be accelerated from LEO to escape velocity, and up to 15 pounds landed on the Moon.  This is now the “High Limit” for the mass we expect to need to win the Google Lunar X PRIZE.  A 150 pound mass in LEO will suffice. 

We have an alternate fuel planned which will reduce the efficiency but slash the risks so that these “High Orbit” flight systems will be accepted as commercial “Secondary Payload” for flight to LEO.


A derivative of the “Thrust Vectoring” systems we have flown successfully in the past promises to come in at <30 grams (1 ounce) and integrate well into our production Propulsion Module.  An  Inertial Navigation System  sufficient to stabilize all the powered operations should be under 10 grams.  And the 3 DOF RCS is now shaping up at 25 grams mass.  All of these are needed in our LTO stage, and these low masses make the projected performance achievable.

For experimental use, we expect to couple the RCS, NAV and COM systems to the Propulsion Module with our Wireless Data Link, and retain these systems in a detached NanoSat after acceleration is completed.  Adding a second RCS “Plate” will provide 6 DOF control to the satellite (25 grams added) and another  40 grams will provide a reaction wheel set for continuous attitude adjustment.  Thus, 100 grams (10% of the minimum CubeSat mass) will provide for both coarse and fine ADAC stabilization. 

This assembly can easily include both Sun Sensors and the ½ gram “Star Sensor” previously discussed.  We will be conducting sensitivity tests on one of these units very soon, although previous tests have indicated that the 19 stars and planets brighter than 1.0 magnitude will certainly be viewable, and the next 31, brighter than 2.0 magnitude, almost certainly will be also.  These fifty viewable targets will be more than enough for precision navigation.


An important use of our star camera will be to monitor the relative motion of select stars near the Moon's Limb as that body is approached.  Although this is less demanding than the Mar's Aerobraking application we studied for a University Paper, that relative motion is the best indicator of Lunar periapse during mid course corrections and setting up for Lunar Orbit Insertion.

Progress is being made, and the pieces are coming together.  We should be running the lightweight RCS systems next week, and plan to install a set – with gyro sensing – in a CubeSat demo soon thereafter.

[rpspeck]

40 Gram RCS + EMTB Sensing + Laser Interferometers

Progress is good in several areas.  Our low mass Reaction Control System prototypes are going together well, with actuators, electronics and machined hardware in the prototypes.  The 3.5 DOF unit will probably come in below the 40 grams mass expected, although additional mass reductions have been identified.  These RCS units work with full pressure CO2 to eliminate the need for a pressure regulator.  3.5 DOF means the unit will provide bidirectional control of the “Angular” Three Degrees of Freedom involved in setting of the attitude of our small spacecraft, or Lunar Lander, before main engine ignition.  The extra ½ Degree of Freedom represents the availability of a unidirectional linear acceleration – used to settle the liquid fuel before main engine ignition. 

Two of these units will provide full 6 DOF, all direction linear and angular adjustments (Attitude and “Divert”) for rendezvous and docking operations.  Our planned CubeSat demonstration will use exactly this 80 gram combination for primary control as it demonstrates our “Low Mass” “Formation Flying” technologies.  These would be perfect for a small “Planetary Ascent Vehicle” in a sample return mission.  With masses this small, multiple probes could be used to return samples from several areas on a planetary surface, as well as providing redundancy.  This capability will also allow us to add Lunar Sample Return to our Google System as discussed previously.

Related to this, progress with our EMTB = “ElectroMagnetic Tractor Beam” is also good.  This is our “Zero Propellant”, extreme precision technology for formation flying plus rendezvous and docking.  We have a number of interesting demonstrations which we are in the process of videotaping.  We will soon post these videos – pushing viewers well ahead of DOD SBIR “Experts”.  Those who view these videos will know that this technology works – unlike the DOD “Experts” who evaluated our detailed discussion of the technology and our accomplishments with it – and concluded that “It Is Impossible”.

Good progress with the 6 DOF sensing portion of this technology was achieved today.  This is also related to  our very low interference, “Wireless Data Link” for use in small rocket systems.  We achieved good passive sensing (for the “Wake Up” and Wireless Battery Charge functions at isolated nodes) and excellent signals at the maximum  range needed for our small “Clustered” launch vehicles.   The smallest “Cluster” is also one our more interesting applications.  By using our wireless link between the satellite or spacecraft and the launch vehicle, the stabilization, communications and navigation resources necessary for the launch can be retained in the spacecraft with no duplication or mass penalties.  Very high performance satellites can then be flown with very light propulsion systems. 

Today's tests also verified the extreme precision of attitude measurement provided by this wireless system at moderate range (better than 1/1000 of a degree).  At maximum range, noise does not allow this precision, but based on today's tests, the maximum range is space with 10 cm coils will exceed 100 meters. CubeSats separated by the length of a football field will be able to  accurately determine their relative 6DOF positions using only this sensing system.

An unrelated technology, with related uses, has also made progress this week.  We have achieved good performance from a compact version of one of the interferometers and associated sensors in our PCALI upgraded prototype.  Our PCALI (PolyChromatic Absolute Laser Interferometer) produces an absolute distance measurement with sub nanometer resolution Without Moving the Target!  Traditional Laser Interferometers require that the target be moved to a reference position very close to the Laser Source before measurements begin.  They actually only quantify the change in position from this initial point, not absolute distance.  Since the use of White = Polychromatic light is well known in classic optics as a way to detect an “Absolute Zero” path reference, we use a related technique to determine the absolute distance to any reflector.  In a few seconds, this distance is determined and any changes from that position can be quantified in microseconds.  This is valuable in satellite formation flying, where arranging the  close initial position is very difficult.  Note that without our Absolute measurement,  any disruption of the light path or electronic function requires that this initialization be repeated.  Combined with the gentle, infinite resolution forces produced by our EMTB system, satellite formations can be stabilized to nanometer accuracy. 

We are actively seeking funding for our CubeSat demonstration of these combined technologies in space, and have small enough, low power lasers and optical systems for that purpose. 

Since many Terrestrial measuring situations could also use this “Absolute” measurement capability, we are also looking for “Funded Demonstration” opportunities to showcase this technology.  Laser Interferometers are the “Gold Standard” for precision Machine Tool linear measurements.  But our variant will allow the measuring beam to be “Time Shared”.  One beam can be redirected to measure the position of many reflector targets in a workspace – possibly a group of large robots when inspection or machining precision exceeds the accuracy that can be reliably sustained by these units. We also have arranged to work with “Spectron Engineering” (Denver, Colorado) to add precision “Robotic Theodolites” to optical measuring setups.  These Theodolites have been used by DOD and commercial aerospace organizations for years to quantify performance of HUD and HMD displays.  These units nicely compliment the capabilities of our PCALI interferometers, and add the beam steering for multiplexed use.  Long range applications of the theodolite often produce better than 25 micron position resolution.  When combined with the interferometer – even under poor atmospheric conditions – long range resolutions can be better 0.01 micron, and well below one nanometer under better conditions.   

[Patchouli]

If those tiny landers work they could be useful later on as secondary payloads for an Orion or Dragon to release in mass while parked in LLO for scoping out places with potential resources.
Also would give the poor sap who has to stay with the mother ship on early missions something to do.

[rpspeck]

RCS and Maneuvering CubeSat Details

Even tiny commercial pneumatic valves provide TOO MUCH gas flow to work well for attitude control in a CubeSat, or our similar sized “Very Low Mass” lunar landers.  A typical, 40 gram air valve from Clippard or similar supplier, is rated to produce a gas flow of 0.6 to 1.0 SCFM (Standard Cubic Foot per Minute) with a pressure drop of 100 to 200 psi.  A “Standard” cubic foot of gas is one expanded to standard atmospheric pressure and temperature.  One Standard cubic foot of air will mass 42 grams and the 1 SCFM gas flow is a 0.71 gram per second mass flow. (Please excuse the mixed units – the valve ratings are often in the “English” units given).  Incidentally,  this valve would have a 0.01 CV rating (Coefficient of Velocity (or flow)) and a 0.6mm (0.024 inch) orifice.   

The 0.71 grams per second mass flow, with the typical 60 seconds ISP obtainable in a “cold gas”  reaction motor using Nitrogen, would give a 42 gram weight thrust, or 0.415 Newtons force. For linear propulsion, this would accelerate a 1 Kg spacecraft at 42 “milli-g”, or 415 cm/sec^2.  In 2.8 seconds it would have achieved a “Delta V” of 1.16 meters per second  and consumed  1.99 grams of Nitrogen.

This is enough Delta V to produce about 0.01 degree Plane Change, generating a 1 Km lateral position adjustment for rendezvous (+/- 1 Km cyclic), or  a 4 Km Altitude, eccentric cycle.  A second similar pulse could stabilize a CubeSat with +/- 4km Altitude change.  The later, using 4 or 5 total grams of CO2 from the 12 grams in a small cartridge, could accomplish orbital rendezvous starting with an old launch vehicle like the Dnepr (+/- 4 km altitude, 0.04 degree inclination error spec). Phasing, following the first pulse, would adjust orbital spacing by 6 km per orbit or 100 km per day, with the second pulse stopping this motion prior to docking.  Since these are not small distances, a CO2 cartridge which will fit in a CubeSat can produce many Km of orbital offset, followed by by docking demonstrations. 

Back to the attitude question, with jets positioned on the sides of a cube sat, the 0.415 Nt force will produce 0.021 Nt*m Torque.  But the Moment of Inertia of a CubeSat will be near 0.001 Kg*m^2, and this torque will produce an excessive rotational acceleration of  21 radians/sec^2.  After one second, the satellite will be spinning at 3.3 revolutions/sec = 200 rpm!  It will have already completed 1.6 rotations in that one second.  Admittedly, a 10 millisecond (1/100 sec) pulse will produce only 2 rpm rotation, but it is unlikely that the breaking pulses can be timed accurately enough to get the rotation below 0.1 rotation (36 degrees) per minute, and frequent pulses (with a lot of propellant use) would be necessary to keep the pointing error less than 10 degrees.  The practical use of gas jets of this type will be to “Desaturate” momentum wheels, with the later being used to make precise attitude adjustments.  With momentum wheels not much smaller than the 10 cm satellite cube, mass 1% of the satellite total, and 10,000 rpm max speed, a set of three wheels could “absorb” up to 100 rpm of satellite rotation.  The jets would be used to unload accumulated angular momentum, if that should occur, and 0.36 grams of gas would be used to offset that maximum in a ½ second burst.

Since this is too much thrust for direct attitude control of a CubeSat, how big a craft could it handle? 

Our 20 kg initial mass, “Above LEO” launch vehicle has an initial roll moment of inertia of 0.012 km*m^2.  The assumed jets would produce a snappy, 1.7 radians/sec^2 rotational acceleration on this vehicle, with the minimum pulse producing a practical 0.017 radians/sec = 0.97 degrees/sec roll.  This is certainly attractive! 

But in deep space, even that rotational acceleration is excessive.  The Aviation Standard “Two Minute Rate Turn” is only 3 degrees per second.  This may seem very slow, but it requires a  considerable 27 degree bank angle to accomplish in a coordinated aircraft turn at 219 mph, and this goes up to a 45 degree bank “Steep Turn” at 438 mph!  At this rate the image in a “Standard” 35 mm camera “Pans” across the film, from edge to center in 7.5 second and back out of view after 15 seconds – and about the same for a “Good View” through an aircraft window.  This is about the largest “Pan” rate which will let an observer really inspect the view.  It is also a turn rate which avoids disorientation and  vertigo in pilots and passengers.

Using these leisurely rates saves a lot of fuel, but still completes a “Turn Over” in 60 seconds. A 12 second thruster burst at 0.004 radian/sec^2, followed by a 54 second delay and a 12 second breaking deceleration would accomplish the flip at reasonable speed, with low mass jets and little fuel use.  With the attitude jets mounted 1 meter from the craft cg, the 0.415 Newton thrust would produce 0.415 Nt*m torque, and accomplish this acceleration with a spacecraft moment of inertia just over 100 Kg*m^2 . 

This value (100 Kg*m^2) equals the calculated moment of inertia for our Human Transport System, with full fuel, prepared to land a human astronaut on the Moon! The 80 gram projected mass for our 6 DOF Reaction Control System could thus serve for fully redundant attitude control of our Human Lunar Lander.  (The “Flip” described would use no more than 17 grams of propellant gas.)

This example is typical of our operational developments, with small spacecraft systems proving adequate for optimized human spaceflight, and the “Virtuous Cycle” of subsystem mass  reductions leading to other subsystem mass reductions making our projections more and more conservative!

Now we only need to find customers who need these capabilities, sponsors who want their names associated with advanced spaceflight technology, or Adventurers who want to GO THEMSELVES!

[rpspeck]

Temporal Orthogonality

It is really nice when you can count on a flow of ideas which make your systems better, lighter and lower cost! I admit that I find that flow most reliably connected with prayer, but I am thankful for the results.  I will skip an update on my shrinking GLXP lander mass for today (but the 1.2 gram gearmotor  actuators I am evaluating look very good, and my last post noted that it is hard to make a cold gas attitude jet “too small” for attitude control of a spacecraft in this mass range.)

I am presently working on the “Reference” signal for my ElectroMagnetic Tractor Beam (EMTB).  Since this also serves for 6 DOF relative position determination when a pair of spacecraft are maneuvered to create or sustain a desired formation, or for docking, it is necessary to create three “Orthogonal” signals which can be nicely separated from each other to generate the desired data.  A set of three receiving coils (1, 2 and 3) are placed on one spacecraft and a set of three driven coils (A, B and C) on the other spacecraft.  Comparing the A, B and C  signals received by each of the receiving  coils provides almost all the information necessary to determine the 6 DOF Relative Positions.   One additional piece of information is needed to break a “Front Back” indeterminacy, but this is not difficult to obtain. 

The isolation of the three signals received by each coil is easiest if the signal voltages have “Temporal Orthogonality” (as do all the harmonics considered in the “Fourier Transform”).  This can be accomplished if the signals are widely different frequencies, even if the “orthogonality” is not mathematically perfect.  But since embedded computers (and some form of digital signal processing) will certainly be involved, the isolation is easier if the “Orthogonality” is directly imposed on the signal generators and reproduced in the digital filters.

Last night I saw a very nice way to modify the circuits in my demonstration system to accomplish this, and make the coil preamps and analysis circuits simpler as well!  Since the circuits are getting simpler, the mass and power of the assembly is shrinking as well.

[tnphysics]

Will the analysis be done in hardware or software, and what kind of electronics will be used- rad-hard or standard plus shielding (the latter allows parts to be used that are more powerful by factors).

And what is the new fuel, and why not a hydrazine RCS?

[rpspeck]

Most of the analysis will be done in software.  All commercial parts will be used, with shielding as required for a short stay in space. (The Radiation doses acquired by the Apollo Astronauts were all quite modest.)

Hydrazine of  all sorts is being avoided for our projects because of its hazards.  I can't envision any being allowed in a CubeSat or similar “Secondary Payload”. 

The recent computations for our spacecraft attitude control were all done  for cold gas thrusters.  We prefer “warm” CO2 for its higher storage density, and lower storage pressure, in spite of modestly lower performance.  We hope that small DOT certified containers for CO2 and N2O will be allowed in a CubeSat, given their excellent safety history and very limited  hazard potential even in the worst case.

My last posting relates to generating 6 DOF positioning forces between spacecraft using no propellant at all!  Since this is an operating system, these forces and the effects they can produce are limited by the appropriate conservation laws. Solar power is of course consumed, and at least one of the spacecraft needs to have active propulsion if modification of the interacting system's cumulative momentum is desired.

[rpspeck]

Wireless Recharge

We have the “Wireless Recharge” system operating for our “Non Radiated Field” Data Links.  This is not a particularly radical technology, but this demonstrated ability to use the Transmit/Receive Transducer Coils of this Data Link system to charge the batteries at each node, as well as verify all the calibrations and operating parameters at each isolated node, is a significant milestone.

We intend to cluster propulsion modules to form a variety of customized “Deep Space” experimental systems.  Our Google Lunar Lander system will probably use use a cluster of seven modules (each having a fuel tank, motor and relevant control systems).  Customized systems for interplanetary experiments may have as many as 50 modules.  The wireless links will allow such assemblies with no cable harness or connectors.  (This can also allow the modules to be assembled on orbit with simple techniques.)  Modules can also be selected, replaced and even reprogrammed – with opening - at any time during vehicle integration.   

The ability to conduct recharge, test and calibration without opening any of the modules makes preparation for launch much easier – even  if it occurs many months  after module production and at a far distant launch site. 

Our prototype system accomplishes recharge at moderate distance even with our low power Transmitter prototype.  In addition to battery recharge – to accommodate battery “self discharge” - this mode will also be used with appropriate Data Codes to switch the modules from “Zero Power” mode to standby. 

In the standby mode, microamps of current will  keep a more sensitive detector in operation.  For a preset number of days after standby activation, the Link Nodes will be ready to respond to a much weaker signal – generated from the other Link Nodes in the vehicle – to switch them into full operation, with microcontroller data handling at each node and the potential to operate motors, release mechanisms, valves and sensors.

[rpspeck]

Radio Link Progress

We are well on our way to finalizing the “Command, Control and Telemetry” Data Link for our Google Lunar Lander system.  We are approaching the < 100 Kelvin Radio Temperature we want in our 10 gram Command Receiver (20 grams with heavier, prototype circuit boards). The allows reliable, moderate data rate control with 5 nanovolts (0.005 microvolts) of received (50 Ohm) RF.  Using a fairly simple “2 meter” antenna on the lander, 2 Watts ERBP (Effective Radiated Beam Power) is needed at the control transmitter on the Earth, and a small fraction of that will suffice with a good transmitting antenna. 

The power from the Moon is a more serious limitation.  With a more complex, but similar mass 70 cm transmitting antenna, 2 watts would serve for telemetry from the Moon with a simple receiving antenna on Earth.  Clustering several good “70 cm” Yagi Ham antennas, the telemetry transmitter power can be pushed below 100 milliwatts, again at modest data rate.  (The HDTV transmission will, as noted earlier, take more power, probably use 2.4 GHZ, and require an impressive Earth Station to get reasonable transmission time.  But those components are also available.)   More details later...

[mlorrey]

RCS and Maneuvering CubeSat Details

Even tiny commercial pneumatic valves provide TOO MUCH gas flow to work well for attitude control in a CubeSat, or our similar sized “Very Low Mass” lunar landers. 

I would suggest you get away from the fluid/valve/pump type design model entirely, go digital. Build an integrated chip that has a ton of tiny cavities in its face that are each filled with a high explosive chemical. Each cavity would be individually ignited via electronic ignition. This would give you zero moving parts, and very high Isp from a pulse detonation propulsion method, with very fine grained throttling of thrust of one pixel at a time up to several...

[rpspeck]

Systems Progress and Suborbital Rendezvous

Time Flies!  Micro-Space is making steady progress on several fronts.  Our Google Lunar X PRIZE update captures some of this progress.  We are making good progress on the wireless data links which will enable the modular, clustered flight systems which are at the core of our commercial strategy.  We do not envision a large market for a single type of low mass, deep space experimental system.  The cumulative market for a number of configurations will probably be  significant, but this is only relevant if specialized configurations can be assembled – and flown – with limited custom engineering.

Among the configurations we are targeting, and plan to operationally demonstrate, are the lunar access system, the Google type Lunar Lander, and a sample return adaptation of the Lunar Lander. 

Our wireless link uses modulation, decoding and synchronization hardware we are concurrently applying  for our “Electro Magnetic Tractor Beam” (which will be employed for autonomous rendezvous of our lunar sample return “PAV” and the orbiting “Earth Return Module”) and for high efficiency communication and navigation. 

This development has been somewhat more demanding than expected – and other demands have consumed considerable time – but we have the core subsystems for this cluster of applications now working.  I will provide more details in the coming days. 

Since the market for small payload transportation to orbit is still neither stable nor well supported, Micro-Space remains committed to developing independent flight systems  to demonstrate  our abilities.  These will include “Suborbital Rendezvous” demonstration systems.  Although much more demanding than Orbital Rendezvous, near free fall conditions exist for more than 40 seconds around apogee for vertical flight above 100,000 feet.   

[blazotron]

RCS and Maneuvering CubeSat Details

Even tiny commercial pneumatic valves provide TOO MUCH gas flow to work well for attitude control in a CubeSat, or our similar sized “Very Low Mass” lunar landers. 

I would suggest you get away from the fluid/valve/pump type design model entirely, go digital. Build an integrated chip that has a ton of tiny cavities in its face that are each filled with a high explosive chemical. Each cavity would be individually ignited via electronic ignition. This would give you zero moving parts, and very high Isp from a pulse detonation propulsion method, with very fine grained throttling of thrust of one pixel at a time up to several...

You may already be aware of this work, given that you used several of the buzzwords and laid out essentially the complete strategy, but this method of attitude and station-keeping control has been investigated.  Lewis, Janson, Cohen, and Antonsson called it Digital Propulsion in their work here:
http://design.caltech.edu/micropropulsion/99d.pdf

The basic idea is to bond an array of resistors on a silicon wafer with an array of combustion chambers in a glass layer and an array of burst diaphragms and nozzles.  The chambers are filled with fuel (lead styphnate in our case) and alligned and bonded so that each chamber is matched with a resistor and a diaphragm/nozzle.  When a "bit" of thrust is required, the correct resistor on the array is energized and about 10^-6 (theoretical) to 10^-4 (demonstrated) Newton-meter of thrust is produced.  On the order of about a million single-use bits can be placed on each side of what we would now call a ~1 kg nanosatellite (the paper linked above uses the term microsatellite before the distinctions were clear).