Mars Direct; Mars Semidirect; Design Reference Mission. Need Updating?

Oldfart1939 · 2017-02-22 17:54:44

After perusal of the history on various Mars Mission topics, I thought that perhaps a new thread would be in order, wherein we can concentrate comments/criticisms of previous mission architectures from a 2017 viewpoint. Back in 1990 when Zubrin and Baker first postulated MARS DIRECT, especially in light of the disastrous "90 Day Plan," it was a tremendous breath of fresh air and thus was very exciting. A lot has happened in the intervening years and some of the tools now available simply weren't.

To briefly summarize what's now possible that was not possible, or at least, had not yet been demonstrated: Reusability, and possibility of Orbital assembly of subunits. Include in this: weight reduction and availability of portable and powerful Nuclear reactors. These all are capable of dramatically reducing the cost and some of the complexity of a Mars Mission.

We now have, thanks to Elon Musk and SpaceX, reusable orbital heavy lift capacity. We also have the completed ISS capable of berthing both supply and crew capsules. We also have some COMPETITION for Boeing and Lockheed-Martin. We also know that water in some form or other, is actually ON MARS! All of these components should allow an updated and more refined approach to doing the first mission to the Red Planet. We even have the Dragon 2 capsule which is designed not only to fly up to the ISS, but also return astronauts back to Earth along with substantial mass of experiments and wastes to be disposed of.

As Robert Zubrin has constantly stated, for an actual manned mission to take place, it must happen within an 8 year window or less, due to the way political winds blow. So...my challenge to this group: build a mission from existing or relatively easy to modify/upgrade components! In another thread, (one of the Apollo xx Redux threads), I suggested using an orbitally assembled system based on various components within the SpaceX product line--a "mix and match, or 'Tinkertoy' " system. And--everything based on the fairing diameter of 5 meters from the Falcon 9/Falcon Heavy system.

This is a blank sheet of paper. Let's all go for it!

Last edited by Oldfart1939 (2017-02-22 17:55:36)

RobertDyck · 2017-02-22 19:03:23

I tried to do this initially for the Energia launch vehicle. Back when Boris Yeltsin was president of Russia, and early days of Putin, before he became a problem. I proposed an ITV based on TransHab, and Mars habitat that is all inflatable. Micrometeoroids burn up in Mars atmosphere 30km above the surface. They burn up 100km above Earth's surface, but that still means no micrometeoroids on Mars surface. So a soft surface hab requires scuff protection from astronauts bumping into it, or Martian sand storms. Those sand storms can have extremely high wind speed, but low pressure, but that means sand-blasting the hab for months. The outer most layer of TransHab is Ortofabric, the same fabric as EMU spacesuits. That fabric is a double layer with 400 denier PTFE fibre outside, with a backing of 400 denier Nomex, and the backing replaces 2 threads with Kevlar ever 3/8 inch. This provides micrometeoroid protection, Nomex is the same material as fire fighters jacket and pants so it's fireproof, and Orthofabric retains its strength in the extreme temperature swings of Low Earth Orbit. However, Mars doesn't have micrometeoroids, has a CO2 atmosphere so does not need fire protection, and although temperature at the equator can get down to -80°C, Viking 2 lander recorded a low of -111°, that still means the -150°C of LEO just dosn't happen. So I propose using Tenara fabric. While Orthofabric is a double layer plain weave, Tenara is a single layer twill weave. That's the same weave as jeans. And the thread or "yarn" that Tenara is made from is the same 400 denier PTFE as the outermost layer of Orthofabric. In fact Tenara is made by the same company as Orthofabric, in the same factory, on the same machines.

One reason for proposing all soft habitats was to fit within the 5.7 metre fairing of the upper stage of Energia. That applies to the 5.2 metre fairing of Falcon 9 / Falcon Heavy. But some people didn't like the idea of all soft. They wanted at least one hard wall habitat.

Energia could lift 88 metric tonnes to 200km orbit @ 51.6° inclination. I spoke to one engineer in Canada, an immigrant from east Ukraine who received his engineering training at university in Moscow. He told me Energia with its upper stage could lift 120 metric tonnes to LEO. I found a NASA website that listed Energia performance with the upper stage to C3=0 and C3=15. Again, not as great as Saturn V or SLS block 2, but a lot more than Falcon Heavy. If we use Falcon Heavy, we require orbital assembly. I got the idea of Energia from Robert Zubrin's book, written when Boris Yeltsin was president. I hadn't heard of Energia until Robert Zubrin's book. Great book.

Light weight nuclear reactor, updating Mars Direct

Yet another Mars architecture

GW Johnson · 2017-02-22 19:42:28

I think Energia will never again be seen. It's gone. Let it go.

GW

Oldfart1939 · 2017-02-22 20:16:07

Oldfart1939 wrote:

Take a look at my post #56 on the SpaceX, NASA, and the New Administration thread; the utilization and/or modification of existing or already proposed hardware should be combinable into a Moon vehicle system.
Proposed architecture: Dragon 2 capsule initially modified for an Apollo 8 redux with increased food and consumables simply by lengthening the capsule. Convert the trunk segment into a booster stage with enough fuel to undertake Moon orbit insertion and orbital departure on a return trajectory. Have an in-orbit assembly of an Earth orbit to lunar trajectory throwaway booster. Could in principle, become a technology demonstrator for the tether system of artificial gravity production? This system could utilize the surplus crew capacity of 7, scaled back to 2-3-4 astronauts in order to conserve oxygen and consumables.
For the Apollo 11 redux, lengthen the trunk stage(s) to include a lunar landing motor, fuel supply for landing in a separable stage that could be left behind with landing legs, etc. The other trunk stage would contain a motor and fuel for lunar departure. The Dragon capsule would have the 8 small Raptor engines and MMH/NTO fuel for an Earth landing. This whole system is again, built at the ISS, thereby making the boys at NASA's eyes cross with ecstasy. This only tasks SpaceX with building the intermediate stages modified from the currently unpressurized cargo trunk.This whole system isn't some multi-billion dollar wet dream. Just a rationalization of existing capabilities.

Abstract out the word "moon," and insert "Mars" in the above. Modification of several of these trunk stages, assembled in orbit could--in principle--build a Mars landing vehicle.

Oldfart1939 wrote:

When brainstorming, we should regard NOTHING as "set in stone," and be looking at alternative pathways to the same result. If, with the existing hardware (Falcon Heavy) we are limited to the payload fairing Diameter of 5 meters, do a redesign based on those parameters. The other one being 53 metric tons to LEO. Moving forward from these limitations shouldn't be insurmountable. As size increases, the tether concept increases in scope, too.
I've been conceptualizing something based on the existing components and several which could be constructed in a reasonable time frame. My design encompasses having several functional modules constructed specifically for in orbit mutual docking/assembly.
Component (1) would be an expanded and uprated Dragon 2--maybe called a Dragon 2+, with a base diameter of 5 meters. It would be capable of docking at the ISS through the docking adapter already designed. It would have a removable heat shield which will not be needed any longer , but was in place to protect the vehicle and crew in event of a launch failure. In place of the unpressurized cargo trunk, a separately launched crew quarters and supplies module, module (2) would be joined and bolted to the crew capsule. To this, a 3rd module--Mars landing module (3) would be attached, containing the engines and fuel/oxidizer components (I'm suggesting MMH and NTO for long term storability). It would be equipped with landing legs in a manner similar to the Dragon 9 v.1.2+ currently made. Finally, an Earth departure stage powered by cryogenics--a methylox combination using the new raptor engines. This would be flown to orbit immediately prior to ED. I'd err on the plus side by building 2 of these vehicles and having them both depart at the same time, then accomplish an in-flight maneuver to join together with tethers for production of artificial gravity at 0.5 Earth, just in order to maintain some extra strength at Martian 0.38 g. At Mars arrival, they would de-spin, cut the tethers and aerobrake into mars orbit. They would land sequentially several hours apart. Each vessel could in principle carry a crew of 7, but would use the capacity to carry 4 or 5. At Mars, that would give us 2-3 Triads and a mission commander at a total of 8 or 9--or 10 if a backup member is thought warranted. All of this would be preceded by sending the supply, nuclear reactor, and habitat structures, in addition to an ERV.
This is just my first "draft" of a proposed mission. The ERV would definitely be ISPP powered, and the large combined crew size would allow for some to stay on Mars for a longer time frame--an additional 18 months for the next Hohmann transfer window to open.
Yes, this mission architecture is not the minimalist model, but could be accomplished with 4 Falcon Heavy and a couple Falcon 9 v.1.2+ flights. Since all are to LEO, re-landing the booster stages at Canaveral should be possible. It also allows for this to be done over several months for all the in-orbit assembly to take place. Some of the cost could be offset by carrying some supplies to the ISS, as well.
Just a note: The affordability of this concept is made possible through reuse of recovered booster stages, and my estimate of overall cost is under $500,000,000. Or a Half Billion. Call it the cost of a single SLS?

Last edited by Oldfart1939 (2017-02-22 20:28:32)

SpaceNut · 2017-02-22 20:54:37

As noted by all the current capsuleare only good for going to the ISS andnothing else that said we do need a new design for going deeper into space.

That also requires artifical gravity for an appriopriately sized habitat. All of which requires a larger launcher than 53 mT, new EDS stages, Engine to fuel matches for use, new return from Mars stage capable of being used many years after a long stay on mars.

A large enough mars landing craft that scales upward using propulsive retro propulsion which means a huge ship to hold lots of fuel. The ability to return back home once we are on that surface after making the needed fuel to do so.

Then targeted insitu manufacturing useage and science to go with such a program to go and eventually stay....

Oldfart1939 · 2017-02-22 22:06:10

In my post #4 above, and especially in he second quote box, I have addressed most of these points. Artificial gravity? Yes. Larger size vehicles? Yes. Prepositioned ERV? Yes.

kbd512 · 2017-02-22 22:23:37

Cygnus modules are miniature tuna cans. The standard module has sufficient pressurized volume to store the food and water required for a year, presuming some recycling of air and water. A Falcon Heavy upper stage or Vulcan Heavy upper stage can perform TMI since Cygnus is within mass constraints. Cygnus can tether off to the expended upper stage for artificial gravity. It's a one-way ticket to Mars for two astronauts. Cygnus is not fancy and it's rather small, but it'll get you there. If anyone here wants to do this mission in the next decade, then we're using what we already have.

RobertDyck · 2017-02-22 23:59:57

The architecture I proposed a long time ago was modified from Mars Direct. The reason I mention Energia is one criteria was to fit a launch vehicle that had lower lift, and smaller fairing. I waffled on hard vs soft habitats. GW argued for at least one hard hab. But hard hab means large diameter payload; the word "fairing" does not apply when launched "naked". So all the same issues apply to Falcon Heavy.

Updating Mars Direct:
ERV: Dragon Rider capsule, not Dragon v2 because launch abort is not required when launching from Mars. If you don't achieve orbit, you're dead anyway. And Cygnus pressurized cargo module docked to the nose of Dragon to house recycling life support equipment. Life support is the same as ISS. The capsule will use non-recycling life support for Earth atmospheric Entry/Descent/Landing. Use SAFE-400 to power ISPP, and mobility components from Curiosity rover to transport the reactor. Use ADEPT as the heat shield for aerocapture and atmospheric entry, for both the ERV and surface habitat.

Mars Orbit Rendezvous aka Hybrid-Direct:
MAV: capsule is just a fairing, astronauts ride in spacesuits. Deep Space Habitat aka ITV parked in Mars orbit. A single Dragon Rider capsule attached to the ITV as emergency escape pod. ADEPT for aerocapture into both Mars orbit and Earth orbit. At Mars leave parked in highly eliptical high orbit, but at Earth aerobrake down to ISS. "Escape pod" is used only if aerocapture into Earth orbit fails. Dock at ISS until next mission. TMI is achieved with expended LH2/LOX stage. That stage can be an SLS Interim-Cryogenic Upper Stage, aka Delta IV upper stage, lauched as payload on Falcon Heavy.

All this has been posted many times. Is that the kind of mix-and-match or Tinkertoy system you're looking for?

Oldfart1939 · 2017-02-23 11:47:33

Robert-

You're on the right track. But what we probably should do is make a set of lists--3 of them; first, what is available NOW; second, what is already under construction and due for completion, and finally, what do we still need to develop. Regarding what we need to develop, I don't mean entirely new heavy lift vehicles, or entirely new from scratch components. Use existing hardware capable of some major redesign/alteration to suit our purposes.

Available now:
(1) 100 KWe Nuclear powerplant.
(2) Falcon 9 v.1.2+.
(3) Dragon 1 cargo version.
(4) Dragon unpressurized cargo trunk.

Available near-term:
(1) Falcon Heavy, upgraded to Full Thrust.
(2) Dragon 2, Human Rated spacecraft.
(3) Red Dragon Mars landing rated.

Need to be developed:
(1) Living quarters-expanded cargo trunk for Dragon 2+.
(2) Earth Departure stage, modified from Dragon trunk sections.
(3) Mars landing stage with extendable legs.
(4) Mars habitat stage, prepositioned on Mars with transponder system. Includes prefabricated "hard wall" habitat & airlocks.
(5) Earth return strategy and hardware.
(6) Sabatier and Moxie reaction units and storage technology for products.

Now we get to the interesting part of the discussion. Do we use the modified Mars Direct plan, with an orbiting ERV, or do we do a straight Mars to Earth ERV? How long do we stay? How big a mission? How big a rover do we need (versus "want")? How many prepositioning cargo missions do we need? What strategy ISPP do we adapt? Until we have an assessment of what's available now and in the pipeline, this is all an exercise in futility.

Last edited by Oldfart1939 (2017-02-23 19:00:13)

SpaceNut · 2017-02-23 18:11:18

I would also add in subassembly of ship at the ISS and possible re-fuel reuse to parts of it that return from Mars to help manage cost.

I agree that the choice of mars return is either via direct or by linking back up with orbiting ship. I thik that a direct return erv due to mass required to do direct would need to be able to supply a crew for the 6-8 month journey home is quite a tall order to not only land first on mars but also to get back off the surface.

The GPS and beacons could be of the cube sat variety as this does not to be all that powerful only capable of a set of solar panels to unfold from a small packaged size to be able to transmit enough signal with.

Oldfart1939 · 2017-02-23 19:01:50

SpaceNut--

In an edit, I added in the development of both a Sabatier reactor and a Moxie system, along with the technology for storing their products.

SpaceNut · 2017-02-23 19:48:24

Those would be under the category of all three for seperate reasons for each as we do not know the long term duration of success for either under mars conditions. All the testing that we can do will help before putting them through the paces of doing the job for real. Based on the ISS we will need to build in key elements that typically fail with electronically controlled options to access the extra parts to try make it fool proof.

Oldfart1939 · 2017-02-24 10:31:49

I believe that Robert Zubrin, in his latest iteration of Mars Direct, alludes to bringing return fuel and only producing LOX on the Martian surface? Not sure whether he advocates RP-1 or LCH4? Either way, the LOX represents 80% of the propellants needed for the ERV.

RobertDyck · 2017-02-24 12:27:12

No. After all the excitement of Mars Direct, some in NASA had a case of "Not Invented By Me". So they came up with NASA Design Reference Mission (DRM). NASA hired Dr. Zubrin as a consultant. NASA insisted on bringing return propellant, not NASA. But the included a Mars Ascent Vehicle (MAV) to go from the surface to Mars orbit, with an Interplanetary Transit Vehicle (ITV) parked in Mars orbit. The ITV would use propellant from Earth, but the MAV would use ISPP (LCH4/LOX). So astronauts are committed to ISPP anyway. A tank of propellant from Earth dramatically increases launch weight, and that requires more TMI propellant, which also has to be launched from Earth. So that dramatically increases the number of launches. The manufacturer of launch vehicles likes that, but those paying the bills don't. Mars Direct had an estimated cost of $20 billion in 1990 dollars, plus $2 billion per launch thereafter. Or if they commit upfront to 7 launches, $30 billion. According to Robert Zubrin's book. NASA Design Reference Mission had a cost estimate of $55 billion. Congress at the time looked at the price and said, "You haven't even built anything yet, and the costs has already doubled!? No!"

Dr. Zubrin gave a nic-name to NASA's DRM: Semi-Direct. As far as I know, that's the only mission architecture Zubrin was involved with that would bring return propellant from Earth.

Oldfart1939 · 2017-02-24 14:05:51

Robert-

There is definitely an update by Zubrin IN PRINT, and available through Amazon; it's entitled "Mars Direct;" it is a small paperback update to his thinking. Some have criticized it as being something of an anti NASA political rant, but it does contain some thoughts about use of Falcon Heavy and SpaceX involvement. I personally recommend it.

Oldfart1939 · 2017-02-24 16:23:14

Oldfart1939 wrote:

Need to be developed:
(1) Living quarters-expanded cargo trunk for Dragon 2+.
(2) Earth Departure stage, modified from Dragon trunk sections.
(3) Mars landing stage with extendable legs.
(4) Mars habitat stage, prepositioned on Mars with transponder system. Includes prefabricated "hard wall" habitat & airlocks.
(5) Earth return strategy and hardware.
(6) Sabatier and Moxie reaction units and storage technology for products.

After looking around a bit, it occurred to me we don't really have to do that much re: EDS and a Mars LS. The Russian Proton M uses Asymmetrical Dimethyl Hydrazine and NTO as propellants. We don't need the first stage due to having Falcon Heavy, but the second and third stages could be suitably modified by SpaceX into an EDS and MLS, or simply by purchasing the motors and adapting to the Dragon 2 Cargo trunks.

This kinda' accomplishes "getting there," so all we need now is a Earth departure stage fuelled by ISPP.

RobertDyck · 2017-02-24 17:54:34

Spacesuits designed for LEO have to endure temperatures from +250°F to -250°F (~ +121.11°C to -156.67°C). LOX boils at -182.96°C at 1 atmosphere pressure. Boiling temperature goes up with pressure. At 10.14 bar pressure and 119.860°K (-153.29°C) O2 has 974.773 kg/m³ liquid density, and 38.9914 kg/m³ gas density.
Spreadsheet with O2 density data: AirLiquide O2 Physical Properties

LCH4 boils at -161.48°C at 1 atmosphere pressure. At 1.873 bar pressure and 119.690°K (-153.46°C) CH4 has 410.377 kg/m³ liquid density, and 3.1962 kg/m³ gas density.
Spreadsheet with CH4 density data: AirLiquide CH4 Physical Properties

This means by simply ensuring propellant tanks remain in shade, we can keep them liquid with reasonable pressure.

Last edited by RobertDyck (2017-02-24 17:57:03)

SpaceNut · 2017-02-24 20:36:28

Gw ran the numbers for the Red Dragon landing uncrewed as it has no life support for man on Mars to which it can but there is no possible way for it to get back to orbit, so that is a big problem. That said we need to design a Mars lander to fit the bill for manned use.

RobertDyck · 2017-02-25 00:52:11

However, again, do not use propulsive orbital insertion at Mars. Instead use aerocapture with ADEPT. This is part of the original Mars Direct. Again, aerocapture has only been tried once. It failed not because of any technology problem, but because of a metric to US measure conversion error. The engineer gave altitude to aerocapture in US measure, software engineers converted that to metric. The result was Mars Climate Orbiter dipped too deeply into Mars atmosphere, did not recover, crashed into Mars. There is a new crater on Mars somewhere that is Mars Climate Orbiter. NASA has ordered all work in metric, so this problem is solved.

Oldfart1939 · 2017-02-25 09:04:42

Last night I tried to find some hard data on the total fueled mass of the Falcon 9 FT second stage, but was unable to do so. I was wanting to make a comparison to the Proton M third propulsive stage in order to do some modeling of a Mars departure stage up to Low Mars Orbit. My question was whether the Falcon Heavy could send a fully fueled Falcon 9 second stage into LEO?

Some idle speculation here: should Musk wish to test his new raptor Vacuum engine in true vacuum, maybe he's planning a possible performance upgrade to the Falcon Heavy system through incorporation of a Raptor powered 2nd stage? There's a significant performance boost available through the difference in Isp and engine thrust developed. That would achieve 2 goals: improved mass to orbit, and "proof of concept" for the Methylox system.

RobertDyck · 2017-02-25 10:18:18

SpaceFlight101: Falcon 9 FT (Falcon 9 v1.2)
Scroll down for the grey table with second stage specifications, or search the page for "Falcon 9 FT Stage 2".
Inert mass 4,000 kg
Specific Impulse 348s
Propellant mass 107,500 kg
which consists of: LOX 75,200 kg + RP1 32,300 kg

Information on the Raptor is available through Wikipedia: Raptor (rocket engine family) - Comparison to other engine designs
Raptor Vacuum Isp=382s
Thrust is dramatically higher: Merlin 1D vacuum = 934 kilonewtons (210,000 lbf), Raptor vacuum = 3,500 kilonewtons (790,000 lbf).

This means a standard Falcon 9 FT upper stage is too heavy for Falcon Heavy. You could launch one partially fuelled. Or launch with no payload, and the upper stage launches itself into orbit. The upper stage can be commanded to shut down its engine, and is designed for in-space engine restart. It even has cold-gas nitrogen RCS thrusters. But no rendezvous radar. Thrusters are intended for orientation control, for extended mission to deploy a satellite into high orbit. Not to rendezvous with ISS or a spacecraft assembly.

Oldfart1939 · 2017-02-25 11:37:13

Robert-

THX. I found the spaceflight101 data just before returning here to find your contribution. You are correct about the ability of the Falcon Heavy not currently having the necessary ability to orbit a fully fueled second stage as the payload; an uprated Falcon 9 as converted to the Raptor motor sounds capable of doing so, though. My thoughts ran along these lines: use a fully fueled Falcon 9-Raptor engine (Methylox fuel system) to power the EDS/MLS, equipped with extensible landing legs. That would propel my version of an expanded Dragon2 with the trunk extension converted to living quarters and supplies to Mars, with sufficient fuel to land. Also capable of being refueled by ISPP. Second thought: all it would need do is achieve LMO and meet up with a fully-fueled Proton M 3rd stage (storable propellants); do in-orbit reconfiguration and burn back to Earth. Granted, these are just some speculative ideas. We would have to send one of these 18 months ahead of the manned version in order to have ISPP, but it would be fully provisioned for a return mission. We could use the Falcon 9 2nd stage to boost the Proton 3rd stages to LMO and stockpile several of them there.

GW Johnson · 2017-02-25 15:37:38

Somewhere, somebody has to address the wisdom of exploring one site versus many. 500 years ago it was many, for the most successful ventures of that time. There's real lesson about what "exploration" really is there.

If you go with many sites, all this minimalist throwaway stuff is nonsense. You have to start thinking recoverable and reusable designs, for as much of the architecture as is humanly possible. Use it many, many times for multiple missions and purposes, and in spite of the higher thrown weights, it becomes ultimately cheaper, but only as amortized over a life cycle.

GW

Last edited by GW Johnson (2017-02-25 15:38:52)

Oldfart1939 · 2017-02-25 15:58:32

GW-

I'd have to say that in order to have recoverable expensive assets, we initially should concentrate at the most interesting site on Mars. In that manner we can develop a complete orbital/ERV vehicle refueling center. In my above posts I'm attempting to come up with not a MINIMALIST approach, but one which can be carried out the soonest with components which may be available in our lifetimes. I firmly am convinced that we will be able to come up with something!

What's your opinion of using a Falcon 9 FT version 2nd stage as the landing, and potentially ERV power pack for a spacecraft with a mass of 30,000 kg? Especially if Musk goes to an upgrade with a single raptor engine using methylox fuels? Use a second similar module power pack for Trans Mars trajectory insertion. As it stands now, getting something that powerful into LEO seems to be out of reach unless SpaceX incorporates this same uprated module as it's second stage. Thoughts?

GW Johnson · 2017-02-25 16:58:16

Oldfart1939:

To answer your question, I honestly dunno. I haven't looked at the capabilities of Spacex stages with this-or-that engines. I did look at what their Red Dragon might possibly be able to do, but that's it so far. And it’s still based on approximate data. Although it’s in the ballpark.

Without some sort of service module stage that has around 3+ km/s capability, Mars, Titan, and the asteroids are just about it for Red Dragon on Falcon-Heavy. Plus, the extra service module stage is too heavy for Falcon-Heavy to throw in one shot. And that's just one-way probe stuff.

You need the service module stage to make some of the other destinations feasible. Jupiter's inner moons are really tough. The delta-vees are really high to land there, not because of the moons, but because of Jupiter's enormously-deep gravity well and your resulting approach speeds. The other gas giants are similar, but not as bad.

Landing on Mars is a whole lot easier. Even if it's entirely retropropulsive after aerobraking, the dV budget is only around a km/s or so. Taking off back to low Mars orbit is not that hard, either. Theoretically it is about 3.6 km/s, call it 3.7 for losses and a tiny maneuver margin. You can do that with storables easily, even as single-stage reusable, although payload fractions for single stage are low, and thus vehicle sizes are large.

Personally, I'd like to see a prototype of a Mars propellant-making plant taken to a real seacoast glacier with a bunch of bottled slightly-impure CO2, and see it actually make methane and LOX from CO2 and dirty water, and just how fast. I'd like to see that done before we plan missions that inherently risk lives with it. I'm a see-it-to-believe-it sort of engineer (retired now). Professional coward, really. That's what kept me alive working in explosives plants for as many decades as I did.

GW

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Mars Direct; Mars Semidirect; Design Reference Mission. Need Updating?

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