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I notice that most ISRU designs call for a vehicle which produces ISRU propellents to be the vehicle that ultimately uses that propellent in an accent from Mars with potentially a TEI as well.
Generally we assume the ISRU vehicle is 'fully fueled' before people land/launch or in some way make an irrevocable commitment. That necessitates that the vehicle have all the tanks necessary to hold the propellent load produced so it is then not so big a leap to just use the vehicles in accent as well. The whole 'vehicle already fueled' thing sounds very safe, but we would still be looking at a LOM if the ISRU fails, having the crew abort just means we avoid a LOC on top of that. Perhaps their are advantages to moving propellent on the Martian surface.
Say that we have two landed vehicles both able to SSTO from Martian surface, one carries ISRU equipment in a removable module along with surface science equipment, surface supplies, rover and tow able tanker, and a supply of methane fuel. Second lander contains Hab and two empty tanks. Durring the surface stay the methane is transferred a bit at a time in the tanker, then the tanker is purged and the ISRU produced O2 is transferred the same way to the Hab vehicle. Crew can then take off in it's hab equipped vehicle for direct Earth Return or meeting with an orbital return stage. Given the great length of the surface stay a very small tanker can easily move the necessary propellent.
All ISRU equipment is left behind on the surface along with the empty lander and a vehicle capable of moving propellent around. Critical assets for eventually establishing a fully reusable delivery system as we would need to eventually be able to refuel vehicles on Mars with you guessed it tanker trucks. So developing both the expertise and the servicing infrastructure to do that with right from the start would be wise. A future mission could refuel the empty lander and eventually return it as well. This is a kind of in between expendable/reusable architecture, the vehicles could all be designed to be reusable but are initially landed expediently in the sense that they sit on the surface for years but with the potential to retrieve them later. This is the SpaceX way, build it for future reuse but be willing to use it expendable initially.
ISRU is not monolithic though, we have a atmospheric only high TRL oxygen generation tech, and a low TRL excavation based water technology. The former is sufficiently simple and robust that might actually be acceptable as a mission critical asset aka we would allow the crews survival to rest upon it. The latter tech though seems to me to be something so site specific and so prone to failure (mainly of the excavator vehicle) that it shouldn't even be depended on for Mission success. Rather we will take fuel with us to Mars until well after an ice harvesting system has become well established.
The naive strategy many have is to "land all assets at the same place" so we can call it a 'permanent base' the landing vehicles are no more a base then a pile of broken down pickup-trucks in your yard is a base. Bases are things that get supplied continually, and continual resupply necessitates cheap reusable vehicles and durable permanent structures, living in your travel vehicle is never permanent. Bases need to be built only from what a vehicle can offload, that's the only way to get reusable cadence going on the vehicle. It also creates far more political momentum for permanence, because the public can clearly see the distinction between transport vehicle (which they are used to seeing thrown away in space) and 'base' assets which can look and feel more permanent, just like ISS.
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Most missions send a second Mars Ascent Vehicle for return on the next launch window once the crew safety margin is determined to have the lowest risk for mission failure, due to insitu resources possibility if having not worked out as planned.
I am all in favor of recycling the cargo landers as the means to creating a growth path for Mars Residents as well as to aid in making a permanent base. There are some that favor first mission to be of scientific with no plan to follow up in the same location or to have a mobile habit design to rove the planet.
What is TRL?
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I can't imagine TRL is low, we already DO all this stuff on the ground today when fueling a rocket, but we do it in an environment in which spilling some liquid methane OR liquid oxygen creates a huge fireball. On Mars neither Methane nor LOX should be able to generate combustion with the atmosphere or the regolith, only if we spill both on top of each other do we have a problem.
I am sure the first missions will be scientific, even if our intent was colonization we would explore WHERE to put our colony first. Settling in the first place you see is ALWAYS a mistake. We need vastly more data on Mars environments to characterize local resource, geological and hydrological before we could make an informed choice, and their is lots of reason to believe these resource vary considerably from location to location. To the greatest extent possible vehicles sent in thouse scouting/exploration missions would need to be reusable and part of a fleet of vehicles that perform a regular cadence of flights back and forth, each time exploring a new location which is occupied briefly.
Then once we have an idea what spots are good we down-select and start returning to the most desirable locations and start adding to what ever surface assets were left behind when scouting. I think most people suffer from this 'forever upward' mentality, that says that as soon as a place as a human presence of X people any decline in the number X is 'failure' and 'Apollo' and 'not permanent' particularly if X goes back down to zero. The idea that a place can be occupied, then unoccupied, then occupied again is just inconceivable to some folks.
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I think the arguments about not settling in the first place you go, needing to see more than one place, are cogent. I also think it unwise and unethical to bet lives on technologies you are not virtually-100% sure of. Too many things are going to be very site-specific, such as ice availability. I also do not believe there will ever be more than one government-funded manned exploration mission to Mars. Not many in this community would agree with me, but most of government has become too stingy even to do its constitutional duties these days. That speaks volumes about manned missions to Mars.
Taken together, all of that says you send your one and only exploration expedition to Mars and you visit multiple sites with it. Anything less is just Apollo-style flags-and-footprints nonsense. You send enough stuff to do the basic mission even if all the ISRU stuff fails completely, simply because "it is the first time", and nothing else is ethical. You pick the "best" site of those you visit, set up your mini-base and all your ISRU gear, and leave your core-of-a-base there when you pick up and go home.
If ISRU propellant production really works, then you can fuel your landers as suborbital vehicles to visit even more sites. That's fantastic "gravy" beyond the basic mission. That's the smart way to approach this.
I think this sort of thing ought to be done by the "buddy system" for crew safety: divide up your crew so that one part goes exploring, while the other part watches over them with a reserve vehicle that could be a rescue craft. On a first (and likely the only) exploration mission, the best place to base initially is from is low Mars orbit. You need enough landers to have the reserve rescue vehicle, even if one of them has already failed. Later in the stay, when you set up your mini-base at the "best" site, you can base from that site on the surface.
No point going all that way, and having to abort your explorations, just because you were too stingy with the expense of sending lots of landers. You have to stay anyway, until the orbits are "right" to come home. So, go prepared. That's just basic common sense. Besides, engines that push landers can push things to Mars, too. A modularized approach to vehicle design makes great sense, if you multi-purpose your assets like that.
That whole concept leads you directly to a modularized orbit-to-orbit transport design, and re-usable one-stage landers. It leads you to on-orbit assembly by docking in LEO. Life support requirements for a 2.5 year mission lead you to the rest.
GW
GW Johnson
McGregor, Texas
"There is nothing as expensive as a dead crew, especially one dead from a bad management decision"
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How is a government mission that visits multiple sites in one 'mission' (using a dozen launches to assemble in LEO) and then never returns any less 'flags-and-footprints' then the whole arc of the Apollo program that visited multiple sites but just expended a Saturn V for each site? As far as I can see both the objectives of you're program (multiple site visits) and resources (multiple HLV) are comparable.
Multiple landers and a crew doing multiple landings over the course of a 500 day say certainly sounds appealing for the multiplication of sites, but it has to come at some cost in IMLEO both in a upfront redesign of the lander and then in some marginal cost per site visited, I'm skeptical we can do that on the first mission.
To do what your describing you need to be able to land with cargo, refuel from ISRU and take off with all my original cargo again to get back to LMO. I don't see how we can leave behind all the surface equipment needed for a thorough manned exploration at each site and ISRU equipment for refueling each time (equipment that needs to work much faster now then in Mars Direct scenarios) we use a reusable vehicle and still come out ahead. The surface equipment would mass more then the dry lander were getting back so we would be needing more IMLEO while simultaneously risking the crew multiple times on each set of surface equipment (habs rovers etc, not supper risky but a little risky). And it seems to lead to LOC if ISRU doesn't work at each site, but I know your don't consider that ethical, so now you need to bring sufficient assent propellent.
Now to do this I'll assume that propellent for decent will then be supplied from a cache in LMO brought from Earth and it is fairly small so your really coming out ahead on that propellent vs the dry lander. But that still means a SSTO vehicle capable of 4.4 kms, which nearly a 75% propellent mass. That is doable if my vehicle disgorged all it's cargo on the surface and was refueled from an existing base and ascends cargo-less (or nearly so, a few passengers should be doable). But I don't see how it works with the original cargo in tow. Can you give me a mass breakdown of how you see this working?
Last edited by Impaler (2014-11-16 01:53:18)
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The difference between my idea and Apollo is that I do what Apollo did (visit multiple sites, except I do it in one trip) PLUS I pick the best of the explored sites, and base the rest of the 400-500 day stay there. That part looks more like the proposals I see debated here, and is something Apollo never did.
I've run some design sizing studies on landers, and posted some of what I found over at my "exrocketman" site. The "consensus" of these studies is that one stage reusable chemical landers are rather feasible for trips between low Mars orbit and the surface, but not for high-orbit basing. The basing orbit needs to be low: around 200-300 miles. The initial ones visiting multiple sites would be fueled by propellant brought from Earth. Propellant from Earth also fuels the second-part basing-on-the-surface trips.
If ISRU propellant proves successful (during the second-part stay at the selected base site), then you accumulate propellant on the surface there. It can support suborbital trips to yet more sites. If on the other hand ISRU propellant doesn't pan out, you still did your basic mission.
The trip home from low Mars orbit uses propellant brought from Earth and parked there. That way a return is assured, no matter whether ISRU propellant works or not.
The velocity requirements for a trip to Phobos and back from low Mars orbit is pretty close to what the lander can do. If you decide based on ground truth that suborbital trips to new sites aren't so very attractive after all, but you are accumulating ISRU propellants, then ship it up a few tons at a time to low Mars orbit with your landers. It could support a Phobos trip, or be left there in orbit for the next expedition to use.
The lander is just a big conical capsule shape. Low density ceramics would work rather well as a reusable heat shield. EDL is just hypersonics going straight to retro thrust landing. The ballistic coefficients are just too high to bother with chutes. You inherently come out of hypersonics at low altitudes (5-10 km).
GW
Last edited by GW Johnson (2014-11-17 10:19:00)
GW Johnson
McGregor, Texas
"There is nothing as expensive as a dead crew, especially one dead from a bad management decision"
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The method and the means to make multiple site visits could be done by sending a refueling station to each proported site that was selected for a visit so as to allow the single crew once there to do a suborbital hop to the next site and so on until all refueling station has been visited but these are most likely a on time visit for each site.
Then again the sites which have refueling units could have several pairs of crew in small landers land at each site only to chose that final site from them to call home and make just one hop to the final site.
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I agree that SSTO to Low Mars Orbit looks a lot more feasible then going higher such as to Phobos, my math shows 75% propellent mass to reach LMO from the surface assuming MethoLox. I don't know how much of initial LMO descending mass is consumed in decent though, some propellent for retro, some for terminal landing hover, a good deal of the vehicle to be heat-shield, landing legs. If a vehicle is bringing it's assent propellent down with it that makes the vehicle very heavy which stresses the EDL and it would seem to leave hardly any usable payload especially if payload it both down and up.
I looked up your blog and found this post which seems to be what your going off of
http://exrocketman.blogspot.com/2013/08 … boats.html
I'm having a bit of trouble understanding the mass fractions, they look to be just for the assent stage. Can you give me your mass breakdown for the whole round trip, how much propellent is used in decent and how much propellent needs to be picked up in LMO per surface sortie? I see propellent of nearly 50mt for all but the HydoLoX. The Sortie payload of ~3mt seems very small, I know your looking at a 3-man 1 month stay but the Assent stage of the LEM massed ~2mt DRY for an incredibly spartan 3 day stay. I think you would be hard pressed to have 2 people stay for 1 week with that kind of mass budget let alone 3 for a month, and I know you would want more science equipment too like the deep drill rigs Apollo never got to do.
If were in agreement that ISRU propellent can not be relied upon for crew safety then I don't think their is much point in bringing ALL landers back to LMO, for the amount of propellent delivered to LMO you could probably go with one use uber-LEM vehicles 2 stage expendable and do really brief 'snatch rocks and go' surface stays. Your objective here seems to be to validate Resource AVAILABILITY (aka ICE) not the actually production of propellent as you clearly don't have the mass budget to do that, at best were looking at 'lab' level validation, cooking some regolith in a beaker and running through a spectrograph. Thus you should be maximizing your location visits and chances to find that nice ice deposit and if that means leaving some hardware of the surface so be it.
That's why I propose a 2 vehicles down, 1 vehicle up scenario, that allows you to design and use a full 'ferry' type vessel but employ it in an expendable fashion to get closer to the kind of down mass fraction that a one way lander can do. Then if/when ISRU happens you have a direct segway into reusing the temporarily 'expended' vehicles both in Mars orbit and on the Martian surface.
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The weight statements and velocity requirements were in the posting as tables and figures, for the 4 vehicle concepts investigated. These vehicle go directly from hypersonics to retro thrust on descent. They come out of entry hypersonics too low to bother with chutes. If you click on a figure, I think most browsers will show you a greatly enlarged figure, and you can look at them all that way.
The descent delta-vee is quite modest, even for retro-thrust landing. Most of the delta vee is for ascent. I assumed the same payload mass both ways, just so ascents could bring a little ISRU propellant up from the surface. The mass ratio covers all the two-way delta vee in a single stage vehicle, intended to be refueled and flown again.
These were for a low Mars orbit-based mission. Being only a concept study, these didn't have all the margins and safety factors we might really need. The odd thing was, they all came out about the same physical diameter and height, in spite oif different gross weights. The common factor was payload mass to be carried: a few tons.
This kind of vehicle could support explorations of multiple sites from orbit, or suborbital hops from surface-to-surface if based and fueled there. Construction of a semi-permanent based would require bigger payloads and bigger vehicles, of this same basic layout. Colonization missions would require something bigger still.
If easy-to-mine ice deposits can be found to make LOX-LH2 on the surface with solar electrolysis, then the engines of choice would be fueled that way. The "smart thing to do" would be to use these or similar design concepts as a startpoint for a common airframe that could be re-equipped at any time with the tankage and engines for any of the propellant combinations that proves best suited to surface conditions at a base.
GW
GW Johnson
McGregor, Texas
"There is nothing as expensive as a dead crew, especially one dead from a bad management decision"
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I was looking for something more specific then 'very little' DeltaV for decent, could I get a number? Cause as far as I can see their is still zero allocated for that (the propellent fractions exactly match that needed for assent) and this needs to be accounted for to know our Initial Mass in LMO.
Also after some additional research on NASA's SEV concept it looks my skepticism of 3 tons for a short excursion hab is unwarranted. The SEV is only designed for 2 people but it's intended to have a 2 week endurance. Per my earlier thread that landing the crew IN an exploration vehicle provides the best return as they take all their equipment with them and the rover habit is the only one they need I recommend using a surface SEV-like vehicle on this lander concept. The surface configuration of the SEV is expected to have 150 mile range which should be plenty of time and range to conduct a good surface exploration. The surface mobility is provided by a wheeled chassis weighing an additional ton but this could be detached before returning to orbit. An additional 1 tone of modular science equipment would also be wise. So that puts landed-ed payload at 5 tons, ascended still at 3.
The configuration I'm imagining is a bit like MSL, the rover is slung underneath the lander proper. Durring decent a heat shield (preferably inflatable) which is jettisoned before landing. the lander rather then using a sky crane simply has long legs and lands tall with the rover off the ground, then squats down to bring the rover in contact with the surface and the rover roof is unbolted from the lander. At surface activity completion the rover positions itself back under and in contact with the lander and it is bolted back on by it's roof, the wheeled chassis is dropped, the landers engines ignite and assent begins, finally the landers legs are dropped and orbit is made and the crew transfers back to the LMO base habitat.
Neither of us is assuming any plane-change here which could be a significant issue if we want to visit polar locations and the LMO 'base' is equatorial, it may be desirable to enter a Polar Mars orbit so all landings can be 'dropped' over any desired portion of Mars and then assent can be into a polar orbit to rendezvous with the base. This likely requires near instant launch windows from the surface though. Separating the SEV and having it do any acceleration for rendezvous would make sense.
Lander dry mass is around 10% for engines and landing gear giving dry mass to 2, then with 3 tons of cargo up we need 15 tons propellent for Metho-Lox and a total take off mass of 20 tons. Adding in the additional 1 ton chassis and 1 ton equipment means 22 landed tons, lets assume 2 tons of additional decent and terminal hover propellent brings it to 24 tons at subsonic. From their your mass depends heavily on the efficiency of the heat-shielding, rigid heat-shields have averaged 30% of entry mass, inflatables might bring that down to 10% if they are as successful as some folks hope they will be, I'll be conservative and estimate 20% which is 6 tons even bringing the whole thing to 30 tons. That's considerably more conservative then your own numbers which had a 20% TOTAL structure mass for the whole vehicle heat-shield included, if your more or less optimistic about heat-shields or the mass generally needed to devote to decent then adjust accordingly.
Overall that's a considerable improvement vs the reusable lander, we can bring 5:3 ratio as many of these disposables vs just the propellent for the Metho-Lox equivalent reusable, and 2:1 when looking at the whole vehicles. I've only done the math for one fuel though and others might not show the same ratios. The overall mission could consist of 6 landers and 4 men, each pairing of men taking a lander to a different location while their compatriots in orbit tele-operate other rovers or even the manned rover when it's own crew is sleeping. The six landers total 180mt which is about on par with the amount of surface destination assets many conventional Mars Direct style missions employ. Total surface exploration is 12 weeks or 24 man-weeks, far short of 70 weeks and 280 man-weeks if we stayed the whole time on Mars with the whole crew. But the briefness of the stays means the crew can push very hard during that time and probably get twice their long-term sustainable exploration pace done, they can then recoup in orbit for 9 weeks minimum before doing another. That means the nominal 24 man-weeks should still achieve ~20% of the exploration activity level of the one site landing. Given the location multiplier of 6 and the huge value that gives I feel your concept is quite justified if done with this modified expendable lander rather then the reusable configuration.
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One question to ask as well is the fuel used for descent might not be the same one used for ascent when some engines are capable of dual fuel use.
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I think everyone here is assuming the same propellent TYPE and the same actual engine firing for both assent and decent with single stage in both directions. Engines generally get designed around one chemical propellent type (please do not bring up NTR) and can't operate with another, or if they do it comes at a performance hit. I don't see any advantage to even trying to do a dual fuel lander, if you had a ISRU based propellent available to refill with you would just bring that fuel from Earth for decent so you have a single fuel and a simpler engine. If you don't have ISRU then again you just use one propellent that you brought with you.
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A number of years ago I spoke with one engineer for Boeing. One of the great advantages of the Mars Society is you get to meet these people. She complained how difficult it was to design a coupling for propellant transfer. The purpose was in-space transfer of propellant from an unmanned re-supply cargo ship to the International Space Station. She explained there can't be any residual what so ever, because ISS uses UDMH fuel with N2O4 oxidizer. That's the same fuel that Apollo used for its service module, and the same fuel that Soyuz uses today. Shuttle used MMH fuel with N2O4; MMH is a modification of UDMH, designed to be more stable so therefore safer for ground crews to handle. The issues, according to her, were that UDMH and MMH are both stabilized forms of hydrazine, which is hypergolic. And N2O4 is toxic. So if any residual got on a spacesuit, tracked in through the air lock, then it wouldn't be safe inside the station. But I pointed out that Russia already does this; their Progress cargo ships transfer propellant to ISS every time. In fact they've been doing this since Mir, and before that with Salyut. If the Russians can do this, why can't America? She didn't have an answer, just complained how hard it was.
But notice what happens with ISRU propellants. Every design for Mars uses liquid oxygen and liquid methane. They're soft cryogenics, so require cold temperatures for storage, but that temperature is available in space with nothing but a sun shade. No active refrigeration required. The surface of Mars would require refrigeration, but not interplanetary space. And if you track some in to an airlock on the outside of a spacesuit, then those propellants are neither hypergolic nor toxic.
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Contaminating the Astronauts spacesuits during a refueling operation would be a concern and I can see why that is a barrier to going out and plugging in hoses via EVA. Generally most propellent depot ideas involved a fully autonomous robotic coupling which should be both faster and safer then EVA in zero G.
But on a Planetary surface I'd be very doubtful of doing it fully roboticly particularly when the transfer is so mission critical you would want people doing it hands on (or maybe tele-operationally is the sweet spot). If we have a cryogenic propellent in a cryo-tank on a truck then perhaps a well insulated hose that is kept off the ground (eliminating conduction with the ground) will simply be sufficient to get the propellent from truck-tank to the launch vehicle. To minimize contamination of the crew when the hose is being attached and detached I would put an additional disposable 'bunny suit' over the regular space-suit, hell you might want to do that all the time for dust mitigation. Suits are probably being attached to suit-ports too rather then going into airlocks so contamination of the habitat might not happen directly, rather it would be by 'soaking through' the suits soft parts, in any case cryogenic propellent is nasty stuff you do not want contacting the suit. This bunny suit could be the kind of big zipper-across-the-front thing and made of a material like tyvec and just nominally puffed up with some ambient martian atmosphere, the tricky part would be making it able to withstand having liquid oxygen splashed on it and not instantly combust/explode, tyvec might not be up to that. That may simply not be possible to do safely and we would be limited to doing only Methane transfers which should be a lot safer if spilled on a polymer or metallic suit.
Alternatively some kind of purging of the tubing might be possible before it detaches so spills of LOX are further minimized, normal Martian air should be able to purge a tube of LOX by evaporating it while the Astronauts stand at a safe distance and only go in to detach it when it's done cycling.
Last edited by Impaler (2014-11-25 17:12:38)
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Sorry guys, I've been very ill the last several days. Hope that's finally coming to an end.
Impaler:
About my Mars lander studies, if memory serves, I was using just over 3 km/s for fast ascent to very low Mars orbit (200-300 km), and at min about 1.4 km/s for descent, sometimes as much as 2 for extended hover modes. I would start retro burn upon coming out of hypersonics at local Mach 3, typically near 10 km altitude. This is based on "average" Mars conditions. I'd have to go look, but I think my vehicles were typically sized to deliver about 4.7 km/s worth of ideal delta-vee. It was enough for the same vehicle to go from low Mars orbit out to Phobos and back, too. And you're probably right: a 3 ton payload is probably not enough. But I had to start somewhere.
General discussion just above:
Propellant transfer safety is all in the design of the plumbing. The various propellant combinations must have hardware sets that are all geometrically incompatible with each other to prevent mixing the wrong stuff. The actual fittings will likely resemble the ball-valve-protected quick-disconnect fittings used in hydraulic equipment. I have stuff like that out here on my farm. It does a good job, but it is not perfect: you lose a small amount at every disconnection.
That means the receptacles in which these fittings are mounted must be shaped to capture "the drip" (which means there must be gravity) at least long enough to allow its safe evaporation (requiring venting, too). I'm not quite sure how to handle that in zero-gee; where automated is probably better. The receptacles have to be large enough to accommodate both gloved hands at once (makes MCP look more attractive, doesn't it?). One hand pushes on the line, the other slides the coupling ring to make or break the connection. Same as out here on the farm. Go to Tractor Supply Co and look at the Pioneer 4000 series hydraulic fittings, or the very similar John Deere fittings they sell. Then imagine much larger.
Propellant transfer on planetary surfaces (or in orbit) is something we can do. But we haven't actually done it yet. The Russians have started. Nobody is doing cryo that way yet. But we could. You just have to think your way through the hazards and design arround them, with the flexibility in your basic design to make incompatible hardware sets for the various propellant combinations. By that I mean every fuel and every oxidizer get its own set.
For the storables, I think a rubber over-glove might help keep suits uncontaminated. For the cryogenics, it is thermal injury that is the real hazard, as evaporation will be inherent.
GW
Last edited by GW Johnson (2014-11-28 13:22:50)
GW Johnson
McGregor, Texas
"There is nothing as expensive as a dead crew, especially one dead from a bad management decision"
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My discussion with the Boeing engineer regarded an unmanned cargo ship to resupply ISS. Progress has fittings beside the hatch. So when it docks with ISS, it also connects the propellant transfer lines, and electrical connections. My argument was why couldn't an American cargo ship do the same?
For Mars, my design for a ship for the first few human missions does not use propellant transfer. The Mars ascent vehicle would have extra large propellant tanks, to carry propellant for TEI. But if the Mars Ascent Vehicle (MAV) is expendable, and transfers propellant to the Interplanetary Transit Vehicle (ITV), then what's the point of propellant transfer? Why not just use the entire MAV as the TEI stage? No need to transfer propellant, instead the tanks of the MAV are the tanks for TEI.
Of course Robert Zubrin argues that everything should be expendable. At least everything not left on Mars. So based on that argument, his design is a ERV that launches directly from Mars surface to Earth surface. But part of my justification is why land on Mars the capsule for entering Earth's atmosphere? You can save propellant by leaving that parked in Mars orbit. Furthermore, a reusable spacecraft allows as much luxury during the transit back to Earth as the transit to Mars. The Achilles Heel of Mars Direct is the tiny ERV, basically an Apollo capsule with zero-G.
But propellant transfer? That's for later. In fact, my mission design allows the TMI stage to be replaced with a reusable one. It's expendable for the first mission(s), but can be replaced once better technology becomes available. In fact, a reusable TMI stage would be used for TEI as well. That would require propellant transfer. But before you can consider propellant transfer, you require a facility to produce propellant, and a reusable tanker to deliver it. I have argued that's not for the first human mission. That's after substantial infrastructure is built on Mars. You could produce propellant on the surface of Mars, or one of it's moons. Building a facility on a moon means a substantial facility other than the primary human base. A second facility with actively operated and maintained equipment means substantial construction. That's not for the first mission, that's later. But my mission architecture allows transition to that, simply by replacing the expendable TMI and TEI stages with a single reusable stage.
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Just to be clear I've only been considering surface propellent transfer in the thread simply because ALL instances of propellent transfer I'd seen in the past were in-space transfers and it looked to me that no one had considered the viability or utility of doing it on the surface and I wanted to see what came of it as an unexplored idea space. But on a general principle it got me thinking.
When we look at a Mission profile the thing that stands out are the Rendezvous, as we all know rendezvous in lunar orbit was THE key architecture innovation that made Apollo possible. In fact the EXISTENCE of Mission architecture like we are doing here exists because of the broad recognition around Apollo architecture and how much leverage it gives. When we look at a typical Mars architecture is is FULL of rendezvous, you have Earth orbit rendezvous to assemble TMI stacks, you have surface rendezvous (which was notably rejected under Apollo and rightly so considering the landing accuracy they had, but we should be able to do now), and you generally have Mars Orbital rendezvous on the way back (sometimes on the way in too), and finally (rarely) some instances of Earth Orbit rendezvous upon return generally a high orbit like L2 rather then LEO to an existing station which is already in a normal crew ferrying rotation from Earth.
So a Mars architecture is all about how much rendezvousing your going to do, and at every rendezvous you have the opportunity to move crew if one of the vehicles is manned. It's kind of natural to move crew around like that, it's what feels normal from Apollo cause nothing else was transferred that way so people don't really think about other possible transfers. But we should aggressively try to find non-crew transferables every time a rendezvous happens, even if the stuff being moved is available in the next vehicle the scavenged stuff can become the cushion for the next stage of the mission. After all a Mars mission is just full of TIME, the crew has more then enough time to do transfers both in space and on the surface, transferring stuff can be one of the most efficient uses of that time if it pads the safety margins.
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GW: Thx for the specific numbers, the decent DeltaV of 1.4-2 kms is of significant interest to me as I've found it hard to find any references for that leg of the mission. Unfortunately your assent DeltaV of only 3 is very much out of line with what I have read before of 4.1 which is common on many Delat-V maps, how can you explain such a discrepancy, is it a lower altitude then the standard LMO? If the assent DeltaV is indeed 4.1 and we take the minimum decent figure of 1.4 then total DeltaV for vehicle needs to be 5.5 which implies ~80% propellent mass at atmospheric entry. That's getting very brutal not just for reusability but even for a two-stage disposable, it might need to be 3 stages.
If I go back to my lander math and put in the 1.4 kms DeltaV it radically grows my 2 ton landing propellent to 11 tons as the vehicle needs to expend around 33% of mass (ISP 350) no wonder people love to use parachutes for decent. So vehicle is now 33 tons at sub-sonic rather then 22, a 50% growth which then spills upward to a 9 ton Heat-shield and a 42 ton total vehicle at atmospheric entry.
The proposed vehicle isn't the battle-star-galactica quite yet but it sure seems like it could benefit from having more stages during assent to help reduce the take off mass. One possibility is to try to strip down the SEV even more aggressively, rather then just dropping the chassis some of the systems that have been exhausted over the two week stay can be dumped an the crew rely on a more Spartan living standard for the remaining assent to orbit. Could this bring the mass down to the 2 ton range perhaps, that is nearly the dry mass of the LEM which is kind of living standards were now talking about, just a bit more spacious as they have a gutted ship now. Removing this 1 ton should ripple backward to become several at atmospheric entry.
Now lets also assume that the SEV will separate from the rest of the assent stage to complete the final bit of the assent itself. To do this it needs some Delta-V of it's own, maybe just 1 kms will be enough. As their would be no purpose to having this capability on the vehicle when in ground exploration it should be attached prior to accent, presumably replacing the bottom chassis as some of the SEV concepts do in fact do for an in-space version of the vehicle. This stage needs to be landed so it's an additional landed mass lets say 1 ton, but by using it to give ~1.2 kms on assent (i'll assume a low 300 ISP for this stage as it's so small) we can shrink assent propellent by 7 tons. Landed tons reduce by 6 and a 28% reduction ripples upward to bring the total entry mass down to 30 tons again.
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RobertDyck the capsule heatshield for entering Earth's atmosphere leaving that parked in Mars orbit would save propellant manufacture on mars surface IMO and also changes the need of the ERV design.
What we also know about the ERV is that a habitat area and comsumables also should be parked in orbit which brings us back towards the ITV design built in Earth orbit.
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Mars surface escape velocity is listed as 5.03 km/s. That makes the surface circular orbit speed 3.56 km/s, a factor of square root of 2 smaller. That's pretty close to the circular orbit speed at low altitudes like 200 km or 200 miles. Orbit speed (without any losses or plane-change deltas is a lot closer to 3.56 km/s than it is 4+ km/s. Gravity and especially drag losses on Mars are just a lot lower than they are here.
Descent requirements depend very strongly on whether you can make effective use of supersonic parachutes or not. Below a ballistic coefficient of about 100 kg/sq.m you can, above it, you cannot (there isn't enough time to deploy, much less aero-decelerate with a chute). Somewhere close to 1.5 km/s for ballistic coefficients in the 300 kg/sq.m class is "reasonable, but by no means precise.
For descent/ascent to/from very low orbit with no plane changes, your total delta vee is in the 3.6+1.5 + loss "kitty" = around 5.1-5.4 km/s or so. If you are fueled/refueled from an on-orbit supply, you will inherently be heavy on descent. Unpleasant fact of life, that simply must be dealt with.
That's the kind of numbers I was using.
GW
GW Johnson
McGregor, Texas
"There is nothing as expensive as a dead crew, especially one dead from a bad management decision"
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So do you believe that the 4.1 to 4.2 number for Mars assent which are commonly reported are erroneously high because they assume excessive (Earth levels) of gravity and drag losses? Perhaps 4.1 is the number NASA has been using because it is a fully 'padded' number involving both assent, some plane-change and docking with the ERV, I could easily see that explaining the gap.
I also began thinking about a CO/O2 based system that need to only take in atmosphere, the ISP is 290 but the ability to land completely empty and not mess with Hydrogen brought all the way from Earth would greatly reduce the EDL difficulties. I think this may be sufficient to close the circle on a fully reusable exploration lander that receives small (<10mt) amounts of propellent in orbit per trip (or a new parachute). The mission is a bit of a hybrid between the Mars-Direct style of having a fully fueled assender waiting for you and the living off the land 'hopper' concept. Each lander has a ground vehicle and atmosphere processor and the vehicle can autonomously deploy a solar blanket on a reel that can be retrieved by simply reeling it in. Three such landers are pre-deployed on Mars and fuel up, then the crew lands in a 4th identical lander. The crews stays for 20 days and during that time they can consume the whole output of the their vehicles processor to fuel their life-support and exploration activities. Then the crew returns to orbit in the other vehicle leaving the one they landed in on the surface, it now begins building a full propellent load. When full enough it conducts a sub-orbital hop to a new exploration site and begins to make another propellent batch. If each full batch takes 60 days to make then the vehicle is ready to receives a crew landing every 150 days. With 4 vehicles a continual cycle is maintained with crew on surface for 20 days and in orbit for 30 for a 50 day total cycle and achieves 10 site visits over the whole 500 day Mars stay. During the 30 day stop over of each vehicle in orbit the recombined crew can service it by EVA, replacing scientific equipment, refueling, attaching new parachutes, possibly new inflatable heat-shields if the EDL needs are that great and we can afford to bring a supply of these assets we avoid taking them down into the Mars gravity well until they are actually needed. On one level it is psychologically desirable to do this just to keep the crew that aren't in the next upcoming sortie occupied.
Lastly a 5th rescue lander is kept in reserve in orbit. Should their be a failure in the assent to orbit either through a non-start of the engines or an abort to surface during assent then the rescue lander is sent down. The rescue lander carries an additional 20 days of consumables, no rover and a triple set of air processors and solar blankets 2 of which which are removable and substitute for the rover in the cargo hold. The crew deploys these and the necessary propellent load is reached in the 20 day time frame (some equipment the help move propellent from the other landers may also be prudent which would give access to up too 5 processor units for more margin). As the crew already had access to 2 rovers one of these is then brought into the lander and serves as the crew cabin for assent as in all other assents. Should the rescue lander go unused for the whole mission duration then it is landed at the most desirable location for a future base for it's equipment to be used in subsequent missions, though it will not be able to deploy any of it without human intervention.
That's a very ambitious vehicle and a very expansive destination set, with 10 sites you could visit virtually every high profile site on Mars, you can even start thinking about doing more difficult sites as the systems get proven in flight. Start with low elevation northern Hemisphere sites and them work up in elevation and southward, if propellent production slows as a result you can trade quantity of sorties for quality. The critical factor is can ISPP be fast enough to produce the assent propellent in 60 days? So long as it's in the ~100 day range the mission profile still holds, you just visit less sites, but if it is in the range of multiple hundreds of days then the proposition falls apart, the vehicles left behind by crew sorties don't have enough time to redeploy to new sites and be ready to ascend before the crew departs for Earth and their would be no point in making them reusable.
To know if the ISPP (and it's attendant power needs) can do the job it all comes down to how fast the equipment can multiply it's own mass and how effective the propellent is in getting the rest of the vehicle back to orbit. Both the raw rates of propellent from the equipment and the ISP are relevant here and interact strongly. Their are several possible propellents that could be made from atmosphere alone they range from low ISP CO-LoX at 290, Metho-Lox at 360 and Hydro-Lox at 450. Assuming a constant payload to make for fair comparisons then we have a kind of double rocket equation with time as another variable. Let Rp be the Rate of propellent mass production per day expressed as a ratio with equipment mass (including all power sources), so a 1mt equipment package that produces 1 kg a day has a .001 Rp value. Now assume that the vehicle needs to launch a purely parasitic payload that will be set to 1 for comparison, then the vehicle will have structure, tankage and engines appropriate for the propellent type call that Sf, the fraction is of the total take off mass. We will assume that all parasitic mass from EDL are shed and the vehicle expended all propellent on decent and starts totally dry. Finally we need to propellent fraction needed to make Mars orbit with each propellent, these are CO-LoX 77%, Metho-Lox 69% and Hydro-Lox 61%, label this Pf. Lastly the variable Vd will be dry vehicle mass (the portion deriving from Sf) and Td with be time in days, Td will be time in days and Vt is the vehicles total take off mass relative to payload.
Now to combine it all you get Vt = 1 / (1 - Sf - Pf - (Pf / (Rp * Td)))
Plugging in 60 days the Propellent fractions for each type and a 10% across the board Structure fraction and reducing we get
CO-Lox Vt = 1/(.13-(.77/60 * Rp)))
Metho-Lox Vt = 1/(.21-(.69/(60 * Rp)))
Hydro-Lox Vt = 1/(.29-(.61/(60 *Rp)))
Graphing these curves at http://graphsketch.com/ shows that me the knee of each curve which is where the feasibility RAPIDLY goes from working to not working, Hydro-Lox is unsurprising the best with at knee at around 0.1 Rp, a system that can make 10% of it's mass daily in propellent works and just a little less performance blows up the vehicle size exponentially as the processor starts to become the a whole second payload in itself. Also At the 0.1 knee the total vehicle take off mass is around 5x the payload. The Metho-Lox knee is is around 0.2 and the vehicle is around 7x payload. The CO-Lox system has a knee at 0.4 and is 10x payload mass. Higher Rp values of 1.0 for all systems would not significantly reduce the final vehicle sizes, as the flat line at around 3.5, 5 and 8.5 respectively, the processor mass has become trivial.
Now that we know a ball-park range of what a ISPP has to do, are any of them up to the task? If we looked at just the chemical processor half of the system their are estimates that put the conventional Sabatier/Electrolysis and Reverse Water Gas Shift (the Zubrin solutions were familiar with) ahead of the 0.2 they would need though they depend on imported hydrogen or moisture extraction from the atmosphere. Unfortunately power requirements blows everything else away if were using conventional 12.5 kw/kg solar, essentially the ISPP chemistry problem boils down to an electric power density problem. The only solution seems to be to have cold-fusion or extend propellent production time for which we get a linear reduction in required Rp but this invalidates the original premise of rapid cycling reusable landers for the first human explorers.
Hypothetically this type of lander might be a good exploration craft to assist in a sustained campaign of planet wide exploration conducted from a permanent orbital base if ~12 craft were provided and the cadence were reduced significantly to the range of 3 per year. Dependency on Hydrogen being topped off in orbit may be less of an issue as the Hydrogen can be stored and cooled centrally allowing the lander to be more leaky or to temporarily convert all the hydrogen to water right upon landing then converting it to propellent slowly.
Last edited by Impaler (2014-12-02 03:25:18)
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I don't know why others want 4+ km/s. Different assumptions breed different results.
If I thought there was a way to electrolyze water rapidly and efficiently (to my knowledge there is not), I'd go LOX-LH2 and ship a lot of my surface exploration and return-to-earth propellant as water. Really easy to store in space as ice covered in an opaque plastic bag, and strong enough to be its own structure, for a real inert weight savings. Could be supplemented with mined ice on Mars's surface at some sites, not others, rather easily and rapidly. Just electrolyze in advance what you need for the next burn.
If there were such a thing as a water-NERVA you could skip all the electrolysis. Just melt the ice, filter out the dirt, ready-to-use. Especially in a gas-core engine, but there ought to be a configuration for a solid core that would work.
Fast efficient electrolysis and a water-NERVA sound like good development projects for outfits like NASA. Easier to do the Mars trip with them than without them.
GW
GW Johnson
McGregor, Texas
"There is nothing as expensive as a dead crew, especially one dead from a bad management decision"
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At one Mars Society convention, I spoke with one of the engineers for NERVA. I pointed out that the military engine Timberwind had a much lower mass reactor; why couldn't NASA find out what they did and do the same thing? Timberwind was designed for a ground launched ICBM, so single launch only. The 1990 redesign of NERVA increased Isp to 925 seconds in vacuum, Timberwind increased it to 1,000 just by increasing temperature. But Timberwind used a pebble bed reactor, operating so hot that the pebbles agglomerated. They developed hot spots that would melt the pebbles together. I argued to use the NERVA design, and perhaps reduce the temperature to produce the same 925 second Isp, but whatever they did to reduce reactor mass, do the same. The guy I spoke with at the time didn't know what it was, but since I've discovered it was change of nuclear material. They used Americium 242m instead of Uranium.
There has been work on solid core nuclear thermal rockets. The work is theoretical, with computer models, I don't know of any physical test objects. But it has been studied. Because the engine boils the water to steam, filtration isn't quite good enough. They require steam distilled water. You could probably get away with reverse osmosis filtered water. But it can't have any solutes, no salt or calcium, because they would clog the engine real quick.
Now where was that? I found a web site that calculated Isp for nuclear rockets. Operating at the same temperature as the 1990 version of NERVA, but using water as propellant, it had an Isp similar to SSME. Just a hair higher I believe. Do that with an Am-242m engine. One big advantage of water is propellant density: the tank is a lot smaller. And propellant is storable, no need for cryogenic cooling.
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GW: For the last time would you stop talking NTR in every single thread, I'm sick of it being brought out constantly as your knee-jerk 'hand-wavium' to get around every technical barrier. If you want to re-debate it go bump the old NTR thread and I'll do it again with the same result, but I do not what you to EVER mention NTR to me again outside a thread has "NTR" in the title.
For electrolysis the tech is already fairly energy efficient such that the time-span gap I'm looking at could not be closed even with 100% efficiency, the electric power source simply must be above a certain density to allow a rocket to make the energy necessary to reach orbit even if the processing equipment were mass-less and 100% efficient. The power-density issue is why bringing stored hydrogen is so attractive, it's not just the access to a molecule that's hard to get in the Mars atmosphere, hydrogen IS stored energy and takes a huge burden off the electric power source.
Ice mining is something I can not have any faith in to close the 60 day autonomous ISPP scenario. It uses too many moving rovers, heavy bulk regolith handling refinery processes and it is just too chancy that the ice wouldn't be available because the scenario here is of high diversity of sites and being dependent on ice mining means I cant visit the vast majority of the surface. Even sites that I have a suspicion might lack the needed ice can't be visited because the loss of the vehicle would be so costly in the reduction of the future exploration temp. Ice-mining is only viable as a base-camp activity which is not the scenario discussed here.
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In post #22 above, I was not suggesting solutions to anything current, in fact I said so, regarding fast, efficient, high-throughput electrolysis. I merely suggested that the very desirable storage properties of water suggest that we work on the technologies needed to successfully employ it, as something for future application (not immediate).
There are two technology paths that "jump off the page" at you for doing this: electrolysis for LOX-LH2 chemical engines, and direct use of water in something nuclear. Neither are ready now. The "nail in the coffin" for pursuing these kinds of things as long-term technology developments is that we are finding there is water "everywhere" out there, making "fuel-as-you-go" feasible, if the extraction and purification challenges can be met. And they can.
Ice mining off-world doesn't scare me. We have mined coal and other minerals for centuries. That's where you start. But it takes several years to adapt the technologies so they will be available when we need them! I don’t see that effort going on, and that’s what bothers me. When we go to Mars, we’re going to need it, sooner or later. Probably sooner than we ever thought.
Picking through post #24, Impaler also has a long-term technology development suggestion, although that was not his purpose when he said what he said: we really need much better electrical power supplies than anything imaginable today. That would also make electric propulsion methods look much more attractive for faster travel possibilities.
Of course there are more technology needs for successful interplanetary travel: such as habitat modules, closed-cycle or at least mostly-recycling life support methods, experimentation with amounts and with ways-and-means of artificial gravity, and better ways-and-means of radiation shielding.
A part of the very purpose of a government agency like NASA is to work on things exactly like these, so that they are ready to employ when the time comes. But are they? No. Not to any significant degree. Follow the money! See for yourself.
NASA says they want to put men on Mars in the 2030's. But they also actively promote the public fiction that a crowded-as-a-phone-booth and relatively-unshielded capsule is going to carry crews there in zero-gee, for 8 months one-way trip time, on a 2.5 year mission, when even the best "astronaut food" lasts 18 months at most. Right. And none of the things listed above is being seriously funded now, in order to be ready after 2030, when we will need them.
So pardon me if I disbelieve NASA projections and PR, and pardon me when I hold up my hand saying "the emperor has no clothes" when NASA is re-building moon trip hardware, not building Mars trip or asteroid-trip hardware, and yet calling it hardware for going to Mars. The fact that they have to re-create the capability to send men to the moon tells you how far we have sunk in the last 40 years.
And the disparity between NASA's version of moon hardware and what is really needed (see Bob Clark's "budget moon mission" ideas, etc.) to put men on the moon tells you an awful lot about the non-transferrence of the experience and engineering art of the Apollo era folks to the current generation.
Outfits like Spacex and Bigelow are coming much closer to working on the right stuff to go to Mars, but even they are focused on the initial steps near LEO right now. Musk's "Mars Colonial Transport" is an idea, not a real design, right now. They don't have time to work seriously on it until (1) Falcon-Heavy is flying reliably and often, and (2) they've solved their production backlog problem.
My whole point is (and often has been, in the various threads) to point "outside the box" at the things we should be working on right now to enable real interplanetary voyaging several years from now. The sooner we start, the sooner we really get to go.
That being said, this thread title has to do with propellant transfers. No one offered a single comment (!!!!) about my post #15 above, where I suggested what we need is an adaptation of existing hydraulic quick-disconnect fitting technology, combined with receptacle (housing-in-the-vehicle-shell) design to control and contain (actually manage) the inevitable, inherent spillages at connect/disconnect operations. And these aren’t what you fear! Not with the technology I suggest.
What I and every other farmer do on the farm with hydraulic fluids, hoses, and fittings every day could be done as easily with dangerous gasoline, using the exact same fittings, although there's no need for that here on Earth. So RP-1 or any other kerosene (and likely any storable liquid) could certainly be handled that exact same way.
I lose about 1-3 cc at every disconnect, but it doesn't spray anywhere, it just drips off the fittings. The adaptation to cryogenics should not be that hard, excepting maybe peculiar LH2 itself. That last requires quite a bit of thought and experiment. But no magic, it can be done. And should have been done by now, but hasn't.
GW
GW Johnson
McGregor, Texas
"There is nothing as expensive as a dead crew, especially one dead from a bad management decision"
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