kbd512 Light Presurized Rover and Habitat

kbd512 · 2016-03-14 18:52:24

SpaceNut wrote:

Since we will want this to land wheels able to engage the backshell diameter will be in the 8 M or so to get the overall length as needed to fit.

The "backshell" is going to be a piece of Nextel fabric or something similar. The hard aeroshells are part of what add so much mass and volume to our current EDL solution. This payload is delivered using HIAD + forced inflation parachute + minimal retro-propulsion. Four of these vehicles have to fit in one Falcon Heavy payload shroud. That's never going to happen with hard aeroshells, either.

SpaceNut wrote:

The typical ISS module is larger than what you are proposing

Clearly. That's why I didn't propose landing ISS modules on Mars.

SpaceNut wrote:

The retro rocket system must fit within the remaining shape that the truck does not take up. I think that the engines could attach to the frame with the tanks to fuel them over that and to even about the trucks roof or shell that the rover would have to give the needed space to fire the engines for the required time. Some sort of landing strut would need to keep the engines and truck from being damage from the final impact of landing being probably spring or gas filled as a shock obsorber.

Dump HIAD and use pancake type solid motors attached to the base unit of the vehicle.

SpaceNut wrote:

Once down drop a ramp to the ground, detach the engines and tanks to drive away from the landing craft. Or we could do a monstorous Skycrane.....

Forget the ramp and the sky crane. Let's make this simpler and easier to do by not trying to soft land any unnecessary mass. The rover will land on its wheels.

SpaceNut wrote:

Not sure how well the life support will scale down to for just 1 person but I think that the following numbers might help as we would require the systems simular to the ISS example.

The habitat is not an ISS aluminum can, so no, none of what ISS uses currently works for this vehicle. However, even NASA thinks that what's currently flown aboard ISS doesn't work all that well and I think that's why they're working on so many different projects to replace the current life support equipment.

SpaceNut wrote:

While we will only be capturing the sweat and urine for processing is that mass really worth the space required to recycle it as its got to be collected before we can process it even if all we do is filter and electrolysis it for oxygen and hydrogen to process the co2 out of the cabin back into oygen before wasting the methane.

This is a power consumption question. If it's easier to obtain fresh oxygen and water than it is to recycle, then that's what we'd do. Astropoop will only see vacuum desiccation before disposal.

The normal range for an adult urinary output is between 400 to 2,000 mL of urine daily -- with a normal fluid intake of about 2 liters per day. Values for normal urinary output may vary slightly between laboratories. A urine output of 500 mL per day is generally considered adequate for normal function

Our grey water tank will be considerably larger than the average soda bottle.

Then we will need to alicate space for exercise equipment

With two EVA's per day?

SpaceNut · 2016-03-14 19:41:22

The Nextel Fabric comes in a veriety of temperatures with this page giving some of the uses....it is a blanket ...so how would it attach to the heatshield? http://www.3m.com/market/industrial/cer … abric.html

Usually the protective aeroshell (payload shroud) is jetisoned long before getting to Mars orbit so what is to protect the rovers for aerocapture or aerobraking....

The inflateable is a heatshield expanding diamater to allow for more atmospheric drag.. Development of Inflatable Entry Systems Technologies

Rigid vs. Inflatable MSL-Class Mission

Mars EDL High Mass Mars Entry, Descent, and Landing Architecture Assessment

Aerocapture Summary and Risk Discussion

Also 4 stacked inside a payload shroud means that the crew is not in them for landing since a backshell fabric has no hatch entrance to get crew into them.

kbd512 · 2016-03-14 23:11:43

SpaceNut wrote:

The Nextel Fabric comes in a veriety of temperatures with this page giving some of the uses....it is a blanket ...so how would it attach to the heatshield?

Probably 312, but I'm not an engineer. If actual thermal environment modeling makes it possible to use lighter fabrics, that's what we're using. The fabric back shell would attached to HIAD the same way that folding tents have little pockets for the flexible rods to go through. It would look like a gear with the "teeth" being the loops woven into the backshell. HIAD would have its own "teeth" woven into the backside of the fabric. A flexible carbon fiber rope would be threaded through the teeth and tungsten weights attached to the "rear" (as defined by reentry attitude) of the reentry vehicle. Pyro or a suitable mechanical device would cut the rope to separate the backshell and HIAD for parachute deployment. After sufficient deceleration, the rover's articulating suspension would deploy into landing configuration.

SpaceNut wrote:

Usually the protective aeroshell (payload shroud) is jetisoned long before getting to Mars orbit so what is to protect the rovers for aerocapture or aerobraking....

That statement is not congruent with MSL's cruise and EDL sequences. However, Nextel is far better than aluminum at absorbing high velocity impacts. The Falcon Heavy payload shroud will be discarded after the Falcon Heavy rocket ascends to orbit.

I'm not at all opposed to aerocapture for this mission if peak heating is within the limits of an inflatable decelerator. That cuts another expensive SEP propulsion package from this mission architecture. For cargo delivery missions, the simpler, the better.

SpaceNut wrote:

The inflateable is a heatshield expanding diamater to allow for more atmospheric drag..

The general idea behind HIAD is to decelerate high enough in the upper atmosphere to permit a parachute to be effective. CO2 cartridges will inflate a fabric torus attached to the edge of the canopy to forcibly expand the canopy. The ringsail will look a lot like a supersonic parachute, apart from having a larger diameter and an inflated torus running around the edge of the canopy.

Obviously this has to be tested, but testing parachutes should prove far less expensive and complicated than supersonic retro-propulsion. If it's possible to make the canopy large enough with a mass lower than retro-rocket motors, then we use passive methods for soft landing. I'm thinking about 50m canopies.

The MAV's require retro-propulsion. It's hard to imagine how big the parachute would have to be to soft land 10t+ payloads. In my opinion, it's not worth the trouble and probably trades unfavorably with retro-propulsion. If you have to land something really heavy, you need retro-propulsion.

NASA could give my X wing concept a test, but not for the first mission. The reconfigurable aerodynamic decelerator tool requires cutting edge aerospace materials research and development. Subsonic retro-propulsion is still required, but the propellant requirement is greatly diminished. It could be a real standout amongst competing technologies with the right materials and design, but this is twenty years down the road tech.

SpaceNut wrote:

Also 4 stacked inside a payload shroud means that the crew is not in them for landing since a backshell fabric has no hatch entrance to get crew into them.

Like the MAV's, the rovers are launched 2 years ahead of the crew. There's no reason for the crew to be inside these things during EDL. There's also no reason to send crews in DSH's if the surface hardware isn't ready. Each DSH has a MDV attached to it for crew EDL. The rovers run down the MDV capsules after the astronauts land. Your landing accuracy can be relatively poor and the rovers can still chase your capsule down.

After you land:

1. sit tight (already taken care of)

2. chill out (shouldn't be a problem on Mars)

3. wait for the cavalry to arrive (doesn't take that long if you land in the right ZIP code and if you don't it won't be a problem for very long)

SpaceNut · 2016-03-16 16:44:35

A total of 113 PICA Tiles of 27 unique shapes are part of MSL’s Heat Shield. The tiles are 3.2 centimeters thick and adhesive materials were used as gap fillers in between tiles. At peak heating, the heat shield has to withstand a thermal load of 197W/cm².

MSL uses a 19.7-meter diameter Disk-Gap-Band parachute decelerator. The system is mortar deployed when a speed of Mach 2.05 is detected and the atmospheric pressure reaches 570Pa.

Found the reason for the backshell,
http://www.lockheedmartin.com/us/produc … shell.html

Backshell
The backshell is half of the large and sophisticated two-part aeroshell capsule. In addition to protecting the rover during cruise and descent, the backshell provides structural support for the parachute and unique sky crane, a system that lowers the rover to a soft landing on the surface of Mars. The biconic-shaped backshell is made of an aluminum honeycomb structure sandwiched between graphite-epoxy face sheets. It is covered with a thermal protection system composed of the cork/silicone super light ablator (SLA) 561V that originated with the Viking landers. SLA-561V has been used on the heatshields of all Mars landers mission of past, but this is the first time it will be used on the backshell of a Mars mission. Lockheed Martin used the proprietary ablator on the backshells of the successful Genesis and Stardust missions.
Heatshield
The heatshield is the forebody or ‘wind’ facing and the backshell is the lee side during entry of the aeroshell system. Because of the unique entry trajectory profile that has environments that are more extreme than previous Mars missions and will create external temperatures up to 3,500 degrees Fahrenheit, the heatshield uses a tiled Phenolic Impregnated Carbon Ablator (PICA) thermal protection system instead of the Mars heritage SLA-561V. This will be the first time PICA has flown on a Mars mission. Invented by NASA Ames Research Center, PICA was first flown as the monolithic thermal protection system on the heatshield of the Stardust Sample Return Capsule that is now in the Smithsonian Air and Space Museum.
Because of its large size, the aeroshell experiences tremendous entry loads primarily as a result of the dynamic pressures from the atmosphere. More specifically, the heatshield is subjected to approximately 105,000 pounds of compressive force distributed across its surface. This will cause the heatshield to deflect, creating thousands of pounds of bending and shear loads that must be reacted by the backshell.

SpaceNut · 2016-03-16 17:12:59

Slide 7 shows Adept in action after atmospheric entry before parachute use...
https://spaceflightsystems.grc.nasa.gov … nsUAVs.pdf

For Atlas 5's 500-series, the fairings are about 17 feet in diameter. A "short" version, which will be used Thursday, is 68 feet long and weighs 7,770 pounds. A "medium" version that stands 77 feet tall and weighs 8,900 pounds also is available. A layer of cork is applied to the outer surface of the fairing to shield against the heating of ascent. An electrically conductive white paint is then applied over the cork to avoid electrical charges. The inside has an acoustic protection system to lessen the intense sound during launch for the payload.

kbd512 · 2016-03-17 10:16:45

SpaceNut wrote:

MSL uses a 19.7-meter diameter Disk-Gap-Band parachute decelerator. The system is mortar deployed when a speed of Mach 2.05 is detected and the atmospheric pressure reaches 570Pa.

I want to use a disk gap torus parachute decelerator. The torus will be designed to rapidly inflate after a mortar deploys the parachute. The effect is to literally force the parachute to unfurl, even with a canopy loading insufficient for deployment. That's the only feasible way I can think of to completely do away with heavy, expensive, and complicated retropropulsion. The atmosphere is too thin for any other solution. A secondary design goal of this parachute deployment technology is to control the vehicle by tugging on a particular section of the torus using specially designed risers.

If anyone has a better idea for achieving soft landing using less vehicle mass, please post it.

My mass budget for the reentry vehicle is 1t.

Current technology removed about 95% of MSL's kinetic energy upon reentry using existing EDL technology before retropropulsion was employed. We need to figure out how to kill all the kinetic energy without employing retropropulsion through the use of better materials and design. We're increasing the payload mass from .9t to 2.5t and we're reducing EDL vehicle mass from 1.5t to 1t. The only feasible way I can think of to do that is with inflatable decelerators and forced inflation parachutes. MSL was 2.4t upon reentry, so this vehicle and payload is 1.1t heavier.

SpaceNut wrote:

Found the reason for the backshell
Because of its large size, the aeroshell experiences tremendous entry loads primarily as a result of the dynamic pressures from the atmosphere. More specifically, the heatshield is subjected to approximately 105,000 pounds of compressive force distributed across its surface. This will cause the heatshield to deflect, creating thousands of pounds of bending and shear loads that must be reacted by the backshell.

The primary reason for using HIAD and ADEPT was to reduce ballistic coefficient of the reentry vehicle, or so I thought. My understanding is that the low resistance to shearing and torsion loads that lightweight solid ablator materials have as mechanical propeties, which necessitate rigid and heavy backshells, are part of what drive the development of lighter and more flexible materials to better contend with the pressure loads generated by reentry.

HIAD and ADEPT are flexible to a degree that PICA is not. The inflated tori of HIAD can't burst and the deployment arms for ADEPT can't shear off their connection points to the flexible TPS panels, but I think NASA has already figured this out. If the basic mechanical properties of the materials used or the basic vehicle geometries used were not feasible for the deployable decelerators NASA is actively developing, I think development would be terminated.

GW Johnson · 2016-03-17 11:59:23

570 Pa (0.57 KPa) atmospheric pressure on Mars is 5.7 millibars! If the surface is around 6-7 mbar (as it was at the Viking lander sites), you aren't but hundreds to a couple of thousand meters above the surface! That's a second or two or three from impact at local Mach 2 (around 600 m/s, angled steeply downward)! Chute might (!) deploy, but it will have no time to decelerate anything.

Are you sure that ringsail chute on MSL was set to deploy at a pressure that high? The only places I can imagine where such a thing could possibly work would be only the very-low lowlands, where surface pressures might be around 10-12 mbar. Then you'd be around 5+ km up off the surface. Still an awfully short time-to-impact to expect a chute to do very much.

You can figure the forces acting on your item from its entry mass and the peak gee during hypersonic entry, which occurs right toward the end of the hypersonic entry stuff with the heavier ballistic coefficients, only a little earlier if lighter. You can reduce that peak gee, and come out of hypersonics higher up, if you come in at a shallower angle to local horizontal (slower entry speed does the same thing).

Shallower angle: it only requires flying during entry at an L/D of about 0.1, as has been done since Gemini. But this shallower angle may be inconvenient to unachievable on typical direct-entry from interplanetary trajectories.

It's really easy to do from low orbit. Use a surface-grazing transfer ellipse, and on Mars your entry speed is but 3.7 km/s, and your entry angle at interface is 1.-something degrees. Your de-orbit burn is just a dozen-or-so m/s, too.

So I have to ask the question: if EDL is so damned difficult with direct entry mission architecture, then why not consider making it easy with staging your mission out of LMO? That's where you need to park your deep space hab for the return trip, anyway.

GW

Last edited by GW Johnson (2016-03-17 12:05:27)

kbd512 · 2016-03-17 18:05:04

GW Johnson wrote:

570 Pa (0.57 KPa) atmospheric pressure on Mars is 5.7 millibars! If the surface is around 6-7 mbar (as it was at the Viking lander sites), you aren't but hundreds to a couple of thousand meters above the surface! That's a second or two or three from impact at local Mach 2 (around 600 m/s, angled steeply downward)! Chute might (!) deploy, but it will have no time to decelerate anything.

GW,

I don't understand what you're getting at. The DGB parachute attached to MSL slowed the descent from 470m/s from a deployment altitude of 10km to 100m/s at an altitude of 1.8km. Then retro propulsion began. I have to kill 100m/s worth of velocity a bit higher up in the atmosphere. The entire point behind using HIAD is to kill more velocity further up in the atmosphere. MSL had a 16m ringsail attached to it. I want to use a 24m ringsail. My vehicle mass at parachute deployment approximately 2.7t compared to MSL's 1.8t.

GW Johnson wrote:

Are you sure that ringsail chute on MSL was set to deploy at a pressure that high? The only places I can imagine where such a thing could possibly work would be only the very-low lowlands, where surface pressures might be around 10-12 mbar. Then you'd be around 5+ km up off the surface. Still an awfully short time-to-impact to expect a chute to do very much.

Since MSL's parachute was deployed at 10km in altitude, if I can slow my lower BC vehicle to an equivalent velocity about 2km to 4km higher, is there any reason to believe a soft landing without retropropulsion is not achievable?

You can figure the forces acting on your item from its entry mass and the peak gee during hypersonic entry, which occurs right toward the end of the hypersonic entry stuff with the heavier ballistic coefficients, only a little earlier if lighter. You can reduce that peak gee, and come out of hypersonics higher up, if you come in at a shallower angle to local horizontal (slower entry speed does the same thing).

GW Johnson wrote:

Shallower angle: it only requires flying during entry at an L/D of about 0.1, as has been done since Gemini. But this shallower angle may be inconvenient to unachievable on typical direct-entry from interplanetary trajectories.

I'm not going to rely on reentry angle alone to slow my vehicle further up in the atmosphere. The BC must be a little lower. The question is how much lower. What diameter must HIAD have to decelerate to the same velocity as MSL was traveling at, at parachute deployment, approximately 2km to 4km higher up? That's what I need to know.

GW Johnson wrote:

It's really easy to do from low orbit. Use a surface-grazing transfer ellipse, and on Mars your entry speed is but 3.7 km/s, and your entry angle at interface is 1.-something degrees. Your de-orbit burn is just a dozen-or-so m/s, too.

I'm not a great fan of aerobraking and aerocapture because it places extreme performance requirements on the vehicle, but I think those performance requirements are actually achievable in this case and even though the parachute is really big, it's not insanely big.

GW Johnson wrote:

So I have to ask the question: if EDL is so damned difficult with direct entry mission architecture, then why not consider making it easy with staging your mission out of LMO? That's where you need to park your deep space hab for the return trip, anyway.
GW

1. Minimize cost - this payload could only be orbited using SEP. That's expensive and while the mass required is pretty reasonable using higher Isp operation, I want the delivery hardware for the rovers to be stupid simple. I have a 15t payload cap for chemical rocket delivery because I'm relying on commodity rockets and throw capability maxes out between 17t and 20t. The 15t mass cap gives us margin for cycle variances. Some opportunities require more dV than others.

The MAV's already require SEP-CTV's for delivery. A SEP-CTV is expensive. The propellants and solar panels are both prohibitively costly. If Eagle Works can determine how microwave thrusters work, then no expensive propellant is required. However, our thrust level is basically the same as SEP. Power conversion is a little better with Q thrusters, but not much.

2. Minimize launches - If launch costs were not a major issue, which is certainly not the case, then additional launches to deliver additional payloads is still something I still want to pare down to the bare minimum required to get the job done so that NASA can afford to have backups. All launches can fail and some inevitably will. If we minimize the required number of launches to deliver the surface exploration hardware, and thus minimize cost, we can afford to purchase some backups.

I have to deliver 3 MAV's (just like the MDV and rovers, each MAV only holds 2 crew) and 4 rovers. That requires 4 Falcon Heavy flights at Opportunity #1. Not including the SEP-CTV's, that's $500M in launch costs just to send the surface hardware to Mars. That's equivalent to a STS flight. I guesstimate the cost of the SEP-CTV between $125M and $175M. Now we're up to the cost of two STS flights and nothing's been delivered to Mars yet. Add in the cost of the MAV's and rovers and our total outlay for Opportunity #1 is roughly equivalent to three or four STS flights.

Edit: Forgot about the TEI kick stages. Those have to be delivered during Opportunity #1, too.

This is the minimum cost architecture at work here and it's quite obvious that it's still not cheap. We're not going to Mars on a shoestring budget, but we can easily do it with our existing budget if we focus like a NIF laser on affordability.

NASA's going to want to fly some astronauts aboard ISS for training and whatnot, so we need a two STS flight funding margin so NASA can continue to fly humans in space while we're fabricating hardware for the next launch opportunity.

3. Minimize complexity - Maximizing use of existing chemical propulsion technology and minimizing in-space propulsion requirements is all about minimizing complexity. Every little additional capability built into any hardware set above and beyond what's actually required to complete a task increases our outlays for research and development that much more.

Lots of things are nice to have, but the question that must be answered is what is actually required to do a real surface exploration mission.

Last edited by kbd512 (2016-03-17 19:37:48)

SpaceNut · 2016-03-17 19:48:09

I think that you are wishing to deploy HIAD in orbit just before starting to deorbit to head towards the plan.....

Pump Up the Volume is sort of showing how it would be before activating it.

The current inflatable design comes from a partnership with a private company, HDT Global. It's a series of inflated rings made of braided Kevlar - the same material as bulletproof vests - that are stacked together. Each is lined with silicon on the inside. The scientific name for the donut-shaped ring is a torus. The Kevlar gives each torus its strength, while the silicon liner keeps the compressed gas inside - sort of like a bicycle tube and tire. Kevlar straps keep the rings attached to each other and to the payload hardware.

Hypersonic Inflatable Aerodynamic Decelerator (HIAD) Technology Development Overview

HIAD Earth Atmospheric Reentry Test (HEART) are being developed which will demonstrate a relevant scale vehicle in relevant environments. HEART would employ a large-scale aeroshell (10m) entering at orbital velocity (~7km/sec) with an entry mass on the order of 4MT.

Game Changing Development of Hypersonic Inflatable Aerodynamic Decelerator a flexible thermal protection system.

The F-TPS, which covers the inflatable structure and insulates it from the searing heat of atmospheric entry, can be separated into three functional layers: an exterior ceramic fiber cloth layer that can maintain integrity at surface temperatures in excess of 1600° C, protecting the underlying insulation from the aerodynamic shear forces; a middle layer of high-temperature insulators that inhibit heat transmission; and an interior impermeable gas barrier layer that prevents hot gas from reaching the inflatable structure.

http://gameon.nasa.gov/projects/hyperso … or-hiad-2/

Boeing-ASME Recognizes Research on HIAD’s TPS Aeroelasticity

The generation 1 TPS configuration illustrated here consists of one layer of Aluminized Kapton-Kevlar (AKK), four layers of Pyrogel 2250, and two layers of Nextel 440-BF20.

http://www.aeronautics.nasa.gov/te11_su … rogels.htm

Upon encountering an atmosphere, the HIAD inflates and becomes rigid, slowing a spacecraft to which it is attached, permitting a safe descent and landing. Since the HIAD would be covered by a thermal protection system insulated by aerogels, eliminating any threat from frictional heating, larger payload masses could be transported through an atmosphere without worry.
Aerogels are ideally suited for use in a vacuum, as in space, or for habitations sited on the moon or perhaps other planets. Industry has also taken notice of the material's commercial potential, with possible applications in refrigeration, building and construction, updating historical structures, and in confined spaces where smaller quantities of more effective insulation are needed.

SpaceNut · 2016-03-17 20:46:01

I am hoping that we can get back to the Rover....

Life support of the ISS racks are sized for a crew of three.....

http://www.nasa.gov/centers/marshall/pd … _eclss.pdf

The Environmental Control and Life Support System (ECLSS) for the Space Station performs several functions:
• Provides oxygen for metabolic consumption;
• Provides potable water for consumption, food preparation and hygiene uses;
• Removes carbon dioxide from the cabin air;
• Filters particulates and microorganisms from the cabin air;
• Removes volatile organic trace gases from the cabin air;
• Monitors and controls cabin air partial pressures of nitrogen, oxygen, carbon dioxide, methane, hydrogen and water vapor;
• Maintains total cabin pressure;
• Maintains cabin temperature and humidity levels;
• Distributes cabin air between connected modules.
The Oxygen Generation System is designed to generate oxygen at a selectable rate and is capable of operating both continuously and cyclically. It provides from 5 to 20 pounds (2.3 to 9 kg) of oxygen per day during continuous operation and a normal rate of 12 pounds (5.4 kg) of oxygen per day during cyclic operation..

Nice details on the life suppor systems on the station
http://wsn.spaceflight.esa.int/docs/Fac … S%20LR.pdf

SpaceNut · 2016-03-17 21:38:22

http://www.projectrho.com/public_html/r … tPart1.pdf

O2 GENERATION
?Static Feed Water Electrolysis
? Solid Polymer Electrolysis - Liquid Anode Feed
?CO2 Electrolysis
?Water Vapor Electrolysis
? Bioregeneratio

CO2 REMOVAL
?Four Bed Molecular Sieve
? Solid Solid Amine Water Desorbed
? Electrochemical Depolarized Concentrator
?Air Polarized Concentrator
?Two-Bed Molecular Sieve
? Lithium Hydroxide (LiOH)
? Bioregeneration

CO2 REDUCTION
?Sabatier Reactor
?Bosch Reactor
?Advanced Carbon Reactor
?CO2 Electrolysis
?Bioregeneration

http://www.slideshare.net/astrosociety/ … rrasquillo

Average Human Metabolic Balance (lb/person-day)
• Oxygen 1.84
• Water 7.77 Drink 3.56 In food 2.54 Food Prep 1.67
• Food Solids 1.36 Oxygen 0.44 Hydrogen 0.08 Carbon 0.60 Other 0.24
• Total In 10.97
• Carbon Dioxide 2.20
• Water 8.53 Urine 3.31 Sweat & respiration 5.02 Feces 0.20
• Solids 0.24 In urine 0.13 In sweat 0.04 In feces 0.07
• Total Out 10.97 Sustaining people in space requires managing all of their “ins and outs”

Recovery of ~50% O2 from CO2 (Sabatier process)
Recover O2 from CO2: >50% recovery – Supply O2: more reliable and simpler O2 generator, high pressure electrolysis (3600 psia)

kbd512 · 2016-03-17 23:54:08

SpaceNut,

We really need to start working on a table of component masses, volumes, and power requirements. I'll try to start that effort over the weekend.

HIAD:

HEART will either demonstrate the feasibility of what I want to do or not. If it works, then it is the "game changer" that NASA purports it to be. If not, then it's back to the drawing board. I'm optimistic.

ISS ECLSS:

The ISS ECLSS is not what I intended to use for the rover's water processor. It's far too large and far too heavy. The Paragon SDC IWP technology is what I intend to use. Compared to current ISS UPA hardware, it's much smaller and uses less power. The water isn't potable, but it's as good as the UPA. In other words, I'm still looking for the other half of my rover's water processor solution.

ICES-2014-269.pdf

Note the parameterized equation for system weights and volumes.

Note that the masses and volumes are sized for larger units that support more than two crew members.

ICES_2015_submission_124.pdf

ICES_2015_submission_202.pdf

The Microlith Bosch reactor appears to be the most promising near-term CO2 removal technology for the rover from a volume, power consumption, and complexity standpoint. However, there is an up-and-coming technology from University of Florida that may replace current CO2 removal and solid oxide electrolysis oxygen generation technologies. It has much lower power requirements, operates at much lower temperatures (400C vs 850C for SOXE), and is pretty simple from a mechanical standpoint.

EA-4.pdf

Our power requirements for water processing are quite high, as I suspected they would be, so our power requirements for atmospheric control need to be reduced.

SpaceNut · 2016-03-18 19:40:26

I agree kbd512 on the creation of "a table of component masses, volumes, and power requirements" and that I will keep feeding the information that I can find on each item of unknown values that we are in search of.

It takes me a while to read all the documents but the one for the Ionomer-membrane Water Processor by Paragon Space Development Corporation is lacking in the change out rate from it becoming less effiecient at filtering the urine as the higher the calcium and other such hard mineral content is the less it will pass through the membrane.

Updating post on Concurrent Carbon Dioxide Control and Oxygen Generation which is high temperature Co2 gaseous electrolysis and not a liquid electrolysis that has already been proved to work but at huge energy expense to compress the mars atmosphere and liquify it in order to create insitu oxygen it is favorable for a habitat reclaiming process as it would eliminate the lithium scrubber or other types requiring a heating cycle by direct input into the unit.

SpaceNut · 2016-03-18 20:34:11

Following up on the high temperature electrolysis there is also anothe thought with a simular concept in that it can also make fuel and oxygen from stream and Co2 in the same method as the other.

High-Temperature Co-Electrolysis of H2O and CO2 for Syngas Production

Hydrogen Production Using High Temperature Steam Electrolysis and Gas Reactors with Supercritical CO2 Cycles

Solid Oxide Co-Electrolysis Cells

Mathematical Analysis of High-Temperature Co-electrolysis of CO2 and O2 Production in a Closed-Loop Atmosphere Revitalization System

SpaceNut · 2024-04-03 19:55:02

bump

quiet topic that will soon come to pass as 3 companies win NASA contracts to develop Artemis moon rover designs

BB1l2HCD.img?w=768&h=432&m=6

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