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On the compression thing: I cannot claim to be familiar with everything that has been proposed. I am familiar with recip and turbine. And basic thermodynamics. Yes, we have compressors on submarines that charge air banks from 1 atm to 500 atm. They are gigantic, heavy things, carried by a sub weighing 4-7000 tons. That kind of thing does not miniaturize well. And, we're talking about a compression ratio final/input of 500.
Most of the aircraft compression machinery we have is under-30 for compression ratio, most of them under 15 or so. What goes on in a shop air compressor and what goes on inside a piston engine are comparable at compression ratios in the 6-11 range. All of these devices are light enough to fly, and all of them miniaturize well. It is the lower compression ratio that allows them to be miniaturizable and be flightweight.
Mars's atmosphere is 7 mbar, except up on the highlands and mountains where it's nearer 2 mbar. But let's go with 7 mbar. Picking a challenging miniaturizable/flightweight compression ratio like 30, that's a bottled gas pressure of 210 mbar, or about 20% of 1 atm. Most of the chemistry processes I know of take place at 1-30 atm. Well, it seems like that might be a serious problem. Although, exploring low-pressure chemistry is something we can do, right here. But dollars to doughnuts, I'll bet you it ain't ready yet to go to Mars.
On the other hand, lets look at the submarine high-pressure air bank compression ratio of 500, and use that at 7 mbar. 3500 mbar, about 3.5 atm, that's usable with chemistry processes we already use industrially. The only problem is the huge tonnage of machinery for a throughput that scales down directly with inlet density (about 0.6% of that here). Heavy, inefficient. I see not much promise down that path.
Take some dry ice frost, pack it tightly into a steel can with a valve and an outlet pipe, and seal it up. Heat it with low-grade energy (solar thermal would work, albeit slowly). The CO2 vaporizes, but cannot expand, so the pressure rises by the specific volume ratio at a constant process temperature. That's a factor in the hundreds to around a thousand (don't have a thermodynamic properties table for CO2 handy).
There's a compression ratio comparable to the sub's high-pressure air bank, done with steel cans, high-pressure tubing, ordinary gas bottle plumbing, and a solar-thermal panel. Admittedly, it's a batch process, not so amenable to robotic operation. But it could be done. All you need is a source of dry ice.
That's the numbers for the kind of problem we are talking about here. If you want to make methane and oxygen out of CO2 and H20, on Mars, the poles are the place to do it. Unfortunately, that's not where the first landing(s) will take place.
That's also a part of my point: every site on Mars is different. There is no one "average Mars" we can put in a simulation chamber here on Earth.
GW
Please don't misunderstand ... I'm not against ISRU on the first mission. I'm against betting lives on unproven equipment needlessly.
Nothing never-before-done-in-situ can be considered proven. We can't really do "real ISRU" till we land on Mars. Simulations can be quite good sometimes, but it just ain't the real thing. Because our estimates of site conditions are only estimates.
It is essential to thoroughly try out everything we can dream up for ISRU from mission-1 on. I just don't think it'll work as good as folks wish. More than 50% of our early rocket shots in the 50's were failures. This is no different: it takes real trials to find and fix all the "gotchas", and believe me, there will be "gotchas".
I do have some qualms about using a nearly-pure CO2 atmosphere to make stuff like fuel. The density is so low. And how do you compress in any practical equipment from 7 mbar to 10,000 mbar or more? We have never built compressors like that before, recip or turbine, other than near-zero throughput lab devices. Not very many of them, either.
One "out" for compression might be to refrigerate a large volume of atmospheric CO2 to solidify it, pack it into a much smaller volume sealed container with no free volume, then re-heat it. Re-gasification confined like that is automatic compression of the product gas. Energetically, that's not a very efficient process because of the refrigeration, but it might be more practical to do. I just don't know.
We might be better off just mining solid dry ice near the poles, and doing confined-heating compression that way to get CO2 gas bottled at pressures we can really use. But that's not viable ISRU unless you land near one of the poles. See what I mean?
GW
Hi Rune!! How are you?
This asteroid mining stuff not only might be fun, it might even be practical. To see some billionaires pony-up to try it out is really encouraging.
I hope they visit sites like our forums. There's a lot of good ideas being bandied about here.
GW
Well, you try to rig the rocket so that if one engine explodes, the rest don't. That's the kind of suspenders-and-belt (and armored codpiece!!) design that is needed to take on the challenge of a Mars mission.
GW
I think that one could build an ice-covered pond, covered in turn by around 6-15 inches of regolith, and have stable fresh water underneath. The ice would be stable under the regolith. We've already seen that on Mars.
If the ice were a few meters thick, the pressure in the covered water would be high enough that a human diver on pure oxygen would not need a pressure suit. Just a wetsuit to stay warm at 0 C, and a pure-oxygen SCUBA, would work.
The way to keep the water liquid under the ice is the same as that required to grow aquatic Earth plants: use a sunlight-simulating electric lamp. Hence, it is possible right now to grow aquatic plants anywhere on Mars. No terraforming required. This is the best way I can imagine to turn acres and acres of surface to productive agriculture without building any sort of pressure domes.
Actually, it might even be made to work on the moon. Or, the asteroids.
There is the energy cost of running the lights. That's what solar PV and nuclear power are for.
GW
There's already a light gas gun launching small (around 5 pound) payloads at M17 for USAF for hire. It's a private venture. I saw their paper at the 14th annual Mars Society convention in Dallas last summer, same advanced technology session as my Mars mission design paper.
Bigger diameter, higher launch angle (more than anything else), a bit more fuel-oxidizer combustion power, and you reach orbital altitudes at almost-orbital speed. From there it's a very small rocket burn to circularize in LEO. Almost nothing but an old standard Mark 25 solid JATO motor, for a pretty substantial payload (hundreds to thousands of pounds). I don't remember the specific numbers, but it looked pretty good to me. Perfect for refueling an already-existing vehicle, especially if what you shoot up there is water, as tougher-than-an-old-boot ice. The real problem is launch gees. Thousands of them. Just like an artillery shell. Something we already know know to handle, just not with fragile stuff. Gun launch cost looked like about $100-200/pound, compared to Falcon-Heavy at $800-1000/pound.
As for ISRU on the first mission or two to Mars, by all means send such gear for trials, but absolutely don't count on it working right! Chances are, it will not work right, perhaps not at all, especially on the first mission. Probably not even the second.
Accordingly, it would be entirely stupid to count on ISRU for crew survival and return on the first mission. Nothing is more expensive than a dead crew. Ask NASA. They've seen it 3 times now (Apollo-1, Challenger, and Columbia). Nearly saw it on Gemini-8, Gemini-7, and Apollo-13.
It takes on the average 1.5 to 2 full scale, all-up trials of new equipment before it comes close to working "right", and that's with some very talented, artful people working the problem. That's nearly 20 years' aerospace engineering experience talking. Rocket science ain't science, it's about 50% art never written down. It's about 40% science actually written down somewhere. And, it's about 10% blind dumb luck, and you have to plan for that.
It's no different in any of the other disciplines, either. That's why the non-flight engineering disciplines use such whopping huge safety factors. Those of us designing things that fly could not afford that luxury. Fundamentally, that's why flying things are more expensive.
GW
Why would we want to resurrect an ancient 1960-vintage kerosene-LOX technology when we already have a better one? The F-1's were in the neighborhood of 265 sec Isp at sea level. Newer kerolox engines are approaching 290 sec Isp for sea level performance. Thrust depends only upon size, once you hit the chamber pressure regime needed for better Isp.
Admittedly, for a first stage, thrust is way more important than Isp. But that barrier's already been breached with the newer, higher chamber-pressure, engines that yield the same thrust per unit size, for better Isp.
It would be more fruitful to scale those newer engines up, or just stack up more of them.
GW
Louis:
No, the real problem is one of making heat transfer occur as fast as the other propulsion processes, when it truly and fundamentally does not want to be that fast. Skylon's engine is basically a liquid air cycle engine. No one else has ever made liquid air that fast, ever. But, Reaction Engines just might. I'm rootin' for 'em.
GW
edit to add content
http://newmars.com/forums/viewtopic.php … 18#p111418
RGClark wrote:GW Johnson wrote:For Bob Clark: airbreather thrust, particularly ramjet, is very strongly (dominantly) dependent upon flight speed and altitude air density. The nozzle thrust is calculated same way as a rocket (chamber total pressure, gas properties, pressure ratio across the nozzle, and nozzle geometry), the pressure is just lower and the expansion ratio a lot less. You do need to worry about the difference between static and total chamber pressure, unlike most rockets.
The ram drag is the drag of decelerating the ingested stream of air into the vehicle. Its massflow multiplied by its freestream velocity (in appropriate units of measure) is the way that is done. But, nozzle force minus ram drag is only "net jet" thrust. There are several more propulsion-related drag items to account.
There is spillage drag for subcritical inlet operation (which also means reduced inlet massflow!), additive or pre-entry drag for ingested stream tubes in contact with the vehicle forebody, and the drag of boundary layer diverters or bleed slots, quite common with supersonic inlets. None of those are simple to calculate "from scratch" (we use wind tunnel test data to correlate empirically a coefficient for each as a function of Mach and vehicle attitude angles), and taken together they are often quite a significant force.
If you subtract that sum of drags from net jet thrust, you have the "local" or "installed" thrust, corresponding with just plain airframe drag. Most airframers work in that definition. If you don't, then you have to add that sum of propulsive drags to the airframe drag to get the corresponding proper drag for "net jet" thrust-drag accounting (not very popular outside the propulsion community).Thanks for the detailed response. That's actually a little too much detail for what I need. I read your post on ramjet boosters:
Sunday, August 22, 2010
Two Ramjet Aircraft Booster Studies
http://exrocketman.blogspot.com/2010/08 … e-boe.htmlI noted that you were able to get better payload with more shallow launch angle but it created a problem for retrieving the first stage booster, since it went so far downrange. If I'm reading it correctly you were able to double the payload mass with the shallow angle, presumably using aerodynamic lift.
What I'm trying to determine if I can increase my payload just going to the range turbojets can get to, ca. Mach 3+. I intend to use the jets to get to medium altitude for a turbojet, ca. 15,000 m. But I need to get to a good angle as well as reaching its max speed. Another problem is that I don't know if it can get to max speed while climbing.
I looked at the case of the SR-71 and the XB-70 Valkyrie. These had more thrust than I wanted but that added weight because jet engines are so heavy. In any case I noted the climbing rate. From that it seemed doable, considering the high effective Isp, that you could reduce propellant mass that way. The problem is this is for a SSTO application and I can't afford the weight. What I wanted was the engines to put out in the range of 1/7th the vehicle weight to reduce the jet engine mass. What I don't know is how will that effect the climb rate, and will it even be able to reach supersonic now.
Note that an advantage of the SSTO is that you can get the better payload by flying a shallow angle and not have to worry about recovering the booster stage.Bob Clark
I discussed previously on NewMars the reasons why I think it should be possible to do a partially airbreathing SSTO with current jet engines in this post copied below from before the server crash.====================================================
...Looking at the numbers though I'm convinced now you can even make a single stage to orbit vehicle with a combined ramjet/rocket engine, and without having to use scramjets.
The idea is to combine the turbo-ramjet/rocket into a single engine. This is what Skylon wants to do with their Sabre engine. But the Sabre will use hypersonic airbreathing propulsion up to Mach 6.5 before the rockets take over. This will require complicated air-cooling methods using heat exchangers with flowing liquid hydrogen for the Skylon.However, just being able to get to say the Mach 3.2 reached by the SR-71 would take a significant amount off the delta-V required for orbit. Of course if the ramjet could get to Mach 5 that would be even better but key this would be doable with the existing engines of the SR-71. Note too the engines of the XB-70 Valkyrie bomber could operate at Mach 3 and as far as I know they didn't have ramjet operation mode. So it might not even be necessary for the engines to have a ramjet mode, turbojet might be sufficient.
The problem with using jets for the early part of the flight of an SSTO has been they are so heavy for the thrust they produce, generally in the T/W range of around 5 to 10. While rocket engines might have a T/W ratio in the range of 50 to 100. But a key point is the jet engine will be operating during the aerodynamic lift portion of the flight where the L/D ratio of perhaps 7. The XB-70 for instance had a L/D of about 7 during cruise at Mach 3. So if we take the T/W of the jet engine to be say 7 and the L/D to be 7, then the thrust to lift-off weight ratio might be about 50 to 1 comparable to that of rockets.
BTW, it is surprising there has been so little research on this type of combination with the jet and rocket combined into one. You hear alot about turbine-based-combined-cycle (TBCC) where it combines turbo- and scram-jets and rocket-based-combined-cycle (RBCC) , where the exhaust from a rocket is used to provide the compression for a ramjet. But not this type of combined turbojet/rocket engine. It doesn't seem to have an accepted name for example. It would not seem to be too complicated. You just use the same combustion chamber for rocket as for the jet. Probably also you would want to close off the inlets when you switch to rocket mode.
For the calculation the delta-V and propellant load would be feasible, note that for a dense propellant SSTO might require as much as 300 m/s lower delta-V than a hydrogen fueled SSTO, in the range of about 8,900 m/s, so I'll use kerosene as the fuel. Hydrogen might have an advantage though in being light-weight if what you wanted was horizontal launch. Say you were able to get to Mach 3+ with the jets, 1,000 m/s. The delta-V to supplied by the rocket-mode is then 7,900 m/s. But note also you can get to high altitude say to 25,000 m. This might subtract another 300 m/s from the required rocket-mode delta-V, so now to 7,600 m/s.
A bigger advantage than this of the altitude is the fact that you get the full vacuum Isp during rocket-mode, call it an exhaust velocity of 3,600 m/s for kerosene rockets. Note this results in a mass-ratio for the rocket mode portion of e^(7,600/3,600) = 8.3, less than half that usually cited for a kerosene-fueled all rocket SSTO. Note the fuel required for the jet-powered portion would only be a fraction of the dry mass rather than multiples of it based on the fact the 1,000 m/s jet-powered speed is only a fraction of the 10,000 m/s or so effective exhaust speed of jet engines.
Note this brings the kerosene fuel load to be about that of hydrogen fueled SSTO's, except you still have the high density of kerosene. With modern lightweight materials this should be well doable.
Bob Clark
=======================================================
Void:
Midoshi pointed out to me some time ago that atmospheric pressure is not the same as vapor pressure, in terms of the phase diagram for water. And he was right, too.
What keeps ice from sublimating at 0 deg C is an applied (partial) pressure of water vapor of 6 millibars. Dry mostly-CO2 will not serve that purpose: the exposed ice just sublimates, and also any liquid water phase just boils away. That's what Mars has right now.
It might be possible with minimal terraforming to achieve an atmospheric pressure in the Hellas Basin of 23 mbar, but, how much of that pressure is water vapor partial pressure? I dunno, but, until you have 6 mbar of water vapor partial pressure in Mars's atmosphere at 0C liquid temperature, liquid water and ice are simply not stable. Period.
That being said, I really am a fan of trying to terraform Mars. But, I am sure it will take a lot more than 23 mbar in Hellas Basin to accomplish that end.
How about 2 psia (0.14 atm, 140+ mbar) total atmospheric pressure? If nearly-pure oxygen, that would be good enough for humans to breathe without a pressure suit, but would still require an "oxygen mask" (really a pressure-breathing rig). Assuming it was not destroyed during terraforming, 6 mbar of that (about 4%) would be the original CO2 atmosphere. But, at 1% (just a wild guess) absolute (not relative!!!!!!) humidity, that would still be a water vapor pressure too low at 1.4 mbar to stop exposed ice sublimation, and exposed water boiling away.
Any terraformed Mars will require a still-denser atmosphere to ensure stable ponds and lakes, even if ice-covered. Maybe 4 times the 2 psia oxygen I proposed just above.
That's about the limit of what I know. Maybe some others could shed better light on this.
GW
Bob Clark:
There's a related thread on PRI under human missions. Closely related to what y'all are discussing here. Which is resource recovery/mining.
GW
Louis:
They're a long way from flying yet, but I really do hope it works. I know how to do the same job with a parallel-burn combination of rockets and ramjets in a two-stage airplane, but this would be loads better. Single stage at realistic mass fractions (and that is one big hell of an if!!!!!) would always be better.
Skylon's biggest problem has always been lack of funding. Fix that, and one only must contend with technical results. Technical results is always a faster path than funding, for everything I can remember. And, I'm an old guy.
GW
edit adding in content:
http://newmars.com/forums/viewtopic.php … 02#p117002
Beyond that, we're looking for Star Trek-style impulse engines and warp drive. The physics ain't there yet for them, much less the technology.
Scramjet (supersonic-combustion ramjet) on the other hand features (at worst) a constant area chamber, and usually gently-expanding at around 5 degrees half angle. There is no throat contraction at all, but locating the final expansion bell axially can make-or-break obtaining a burn at all. This technology is still very far from being ready-for-prime-time, the recent X-51 flights notwithstanding. There’s no good way to reconcile these conflicting geometries except by one-shot/throwaway ejected components, and even that is very most certainly not a trivial exercise, or a “sure thing”.
And, we have not addressed inlet geometry incompatibilities between ramjet and scramjet at all. They are huge, more especially in the internal ducting lines, surprisingly enough. The external compression features are actually quite similar, which is terribly misleading. Failure to get this right causes as many violent explosions in scramjet test articles as does too-low a scramjet takeover Mach. It is quite catastrophic, and (so far) quite incompatible.
Integrating ramjet-or-scramjet with rocket is even worse. Most rockets have a very large area ratio contraction from chamber to throat: on the order of 10+. And the exit bell area expansion ratios exceed 10, often by a very, very large margin. I know of no variable-geometry techniques to accomplish this kind of geometry change, except one-shot/throwaway ejectable insert items.
You can go from rocket to ramjet that way, but you absolutely cannot go back to rocket. How are you going to change back to rocket in a combined-cycle engine, if your trade studies say that you want to do that? I don’t think anybody on Earth has a practical answer to that, excepting maybe the Skylon folks with their SABRE engine, and I am very definitely not even sure of that! (I hope they do, but I am most definitely not going to bet the farm on it.)
And that’s why I think parallel-burn options for the differing engine types are way-to-hell-and-gone far superior to any combined-cycle proposals I have ever heard of. I have seen many of those for the last 4.5 decades. None has ever led anywhere, before. Not a good track record.
But parallel-burn works, both ways. Try that. We can do it right now. All-existing technologies. Not trivial, but very definitely do-able.
GW
Would it be worthwhile to point out that the first mission or two to Mars need not be "sustainable" in the sense that y'all are discussing here? The ones following, yes, because those will be more-or-less permanent bases of some sort. And we will know a whole lot more about ISRU after that first mission or two. If (AND ONLY IF) we are very smart about what we do on those first one or two missions!
No more Apollo-style "flag-and-footprints" nonsense, please!
Don't forget about light gas gun technology for launching water (as ice) into orbit for processing into propellants. It's just about ready for that job on Earth right now. It's more than powerful enough for the moon and Mars. Right now! And processing water into hydrogen and oxygen can be done (slowly, yes, at low power levels) by solar PV. Right now!!!
Just an odd thought or two to consider, from an old guy.
GW
Mark Friedenbach:
Thanks, Mark. I didn't know myself that anybody had seriously been looking at this NEO resource production cost issue. I'm an old guy (very old now), and I've not kept up very well with what's been going on in much detail. Old guys are like that, we're getting tired.
Be careful, though: I would say this about guys who "built their careers" on some technological item. We started "seriously looking" at scramjet propulsion ca. 1960, and many famous folks that I knew personally "built their careers" looking at it, without any significant success (and I do mean zero) until the USAF X-51 test very recently. That is over 50 years. (The NASA X-43A tests in 2004 do not count, being only 1-second burns after rocket boost to full speed, with no thermal energy balance, and no acceleration under airbreathing power.)
Point is, guys who "build careers" on a difficult problem actually have a vested interest in not solving said problem. I can say that with a lot of believability, because I'm just about the US's last living all-around expert in (subsonic-combustion) ramjet propulsion, and scramjet is something I kept up with long ago as a possible second specialty. This vested interest problem in scramjet is not any different in any of the other technology niches. Never has been. Never will be. Human nature.
I'm not so sanguine as you are about rare-earth ore concentrations in stony or metallic asteroids. This is all based on remote observation and (most significantly) inferred numbers. The actual history is that "ground truth" has always been vastly different than any of the remote observations ever led us to believe, for well over a century now.
There is also subsurface versus surface to consider. Few coal mines show high carbon concentrations on the surface for any kind of remote sensing to pick up. Same for oil. It really does take a deep drill rig or some equivalent thereof. And it will out in space, too. That’s also been the history of things. So, be very skeptical of anybody’s claims to “know” what’s really out there.
But, yes, it's well worth going to find out. It always has been, for well over half a millennium now. That’s another lesson from history. Knowing that, and actually believing in it, is a source of faith, I suppose.
Keep the faith....
GW
As I said earlier, for some decades yet, the whole thing about space resources centers around LEO. What is the price of a gallon of clean water in LEO? (Or the price of any other commodity in LEO.) I gave some data in an earlier post for the delivery cost of water to LEO.
That brings up total cost: there are two components to it. One is the delivery or transportation cost, for which we have pretty good figures Earth to LEO. Figures for transport from the moon or other places are more speculative, but are at least estimatable in ways we can at least debate. The other major cost item is processing/recovery costs on-site at the resource location. For water on Earth, in the industrialized countries, this cost is almost trivial at pennies per 1000 gallons, plus or minus. (Not so in the third world, lamentably, I might add.)
Here is where we have no data at all: what will it cost per unit of product to produce that product on the moon, Mars, an NEO, or anywhere else? Not only do we have no data, but also we don't yet even know how it can be done at all. Nor do we have a clue what the set-up costs will be, that have to be amortized over the (also still-unknown) life of the system. And, it's probably a completely different system at each different site, even on a single planet like Mars. Every site is different: you don't mine coal the same way in Wyoming that you do in West Virginia, for example.
That being said, it is still fruitful to dream up ways this task could be done. The ones that are simplest, with the smallest energy and effort inputs, and the smallest waste factors, are likely to be the least expensive ones. At least, that's been the history down here at home, so it's a good place to start.
Accordingly, I am concerned by proposals that might require enclosures of very large size that also have to input very large amounts of energy, as in capturing and "cooking" or "grinding" small asteroids. Enclosures like that are expensive to build, to launch, and to send to remote bodies. The power supplies for the proposed operations are even worse. Example: shuttle's 25 KW solar array was 65 feet long, several feet wide, heavy, and expensive. We might get 50 KW for the same costs today, but that's a far cry from the 10's to 100's of MW for melting rocks.
That's why I recommend looking first for really high-grade "ores", such as nearly-pure water ice, that would require very little quantity of nothing but low-grade heat at low temperatures to process the resource, at worst a bulldozer or backhoe to recover it (as on the moon or Mars), and a lightweight/low-pressure or otherwise "negligible" container for processing and transport. For example, ice can be shipped almost but not quite naked through space, and supplies its own structural strength to do so. This is highly to be desired. You just need a few millibars worth of containment pressure of water vapor to stop sublimation.
I looked first at ice because it is a source of water, oxygen, and hydrogen, things any human presence in space must have to survive. The two gases are propulsion energy even for a robotic presence. But, minerals and metals you have to look at in the very same way. Picking up iron meteorites on Mars might be a good way to accumulate that commodity. So would deflecting a nickel-iron asteroid into a suitable orbit.
The asteroids and comet cores will have some variable amount of volatiles (ice, ammonia, and CO2, largely), and a lot of stony solids. I doubt if most of these would be very rich in real metals of interest, like iron. As limited as we are for the next few decades, it's the very small ones we might process for products, limited primarily by launch of equipment from Earth (as we are). Very-thinly spread metals imbedded as part of the stony materials are not (yet) of much interest, as the yield is just too low for the processing effort, just like low-grade ores here on Earth.
I would look in some way inside these bodies for high ice content in an otherwise undifferentiated structure, in sizes under 50 meters. You enclose the thing gas-tight in a thin shell, apply solar heat, and melt the ice. The initial spin of the body becomes your enclosure's spin (how to practically achieve that I'm not yet sure). The atmosphere inside your enclosure becomes some fraction of an atmosphere's worth of ammonia and CO2 as the volatiles sublime. Inside the spinning enclosure, the ice liquifies to water and separates centrifugally from the stony particulate content. Your enclosure does need to be strong enough to take the "whacks" as the asteroid breaks up and its pieces fall radially outward under centrifugal force. No plastic films here!
You pump the now-clean layer of water to lightweight "tanks" and let it refreeze there. Compress, separate (not sure how), and store the ammonia and CO2 in gas bottles. Then bag up the stony particulates for shipment, if they represent something you can use. If there were iron, I doubt you could separate it magnetically, because nickel-iron isn't magnetic.
Done automated or by remote control from a practical distance without an hour's time delay 2-way, a rig like this could move from NEO to NEO and process them into usable products for several years. That's the kind of thing that just might be able to supply water to LEO for under $7500 per gallon. If we're lucky and we're good at designing this, the price might be well under $7500 per gallon in LEO.
The smarts for doing this does not lie within a government lab anywhere in the world, not NASA, nor any of the others, it lies within a visionary private group. The smarts for everything we ever did never lay within government labs, it was in the contractors and vendors they hired to do the job. Many of those have grown too large and stultified to qualify as visionary anymore. Truly visionary private groups are rare indeed. Spacex is one. The new asteroid mining venture (PRI) is another. The tourist space companies are some others. That's the fertile field you cultivate to get anything revolutionary done.
Ammonia and CO2 might not have much value in LEO. But for a colony or base on Mars, ammonia is a good fertilizer for crops, however they are grown. And compressed CO2 from an NEO might be cheaper and easier than trying to compress atmospheric CO2 from 7 mbar to 2000 psi. It's a different place; you have to think differently. As in a multi-way trade network among the several colonies we would like to plant eventually.
GW
My idea about Saturn was to take advantage of mostly-pure water (as compared to very thinly-spread water on the moon and most NEO's), not much processing required if the "ore" is very rich. The capture and processing, as I described the very-slightly elliptical or inclined orbit, could be done automated. Even the shipment from Saturn could be done automated or by remote control.
Without men aboard, what difference does it make if the one way transit time is 5 years? Just ship a load out every few months, and a while later, start seeing shipments arrive every few months.
I know Saturn has a gravity well, so does Mars, the only other known source of high-purity ice off Earth.
What we're addressing here is the price of a gallon of clean water in LEO. With the shuttle, at $1.5B per launch to a max of 25 metric tons, and 8.34 pounds in a gallon, that was about $227,000/gallon from Earth. With Atlas-V 551/552 or Falcon-9 at about $2400-2500/pound, that's about $20,000/gallon from Earth. Projected with Falcon-Heavy at $800-1000/pound, it's about $7500/gallon from Earth.
Water from the moon, Mars, NEO's, or Saturn should be under about $7500/gallon "turnkey" delivered to LEO to be competitive, once Falcon-Heavy enters service the next year or so. I'd think that could be done in an automated fashion from any of those destinations, but I don't yet really know.
GW
I was just thinking of very little objects, a few meters or smaller. Things to enclose objects like that we can build, right now.
The Saturn rings idea: last I heard these particles were very small objects, like snowballs and snowflakes. It occurred to me that a very slightly elongated or inclined orbit might allow my "craft" to move into one of the rings and then out for a while, with a very slow relative velocity while in the ring: maybe 5-10 mph.
While in the ring, you scoop up material, and while outside the ring, you close the opening and process the captured material for the products, especially water. After you accumulate a bunch of water, it might be worth shipping inward toward, maybe, Mars. I was looking at the rings for a high yield of water, and very little of the solids.
Just a screwy idea.
GW
Metals I understand. Some stony minerals might also prove useful, who knows yet? Water (and other volatiles like ammonia and CO2) I think vary very greatly from object to object, and are likely the "matrix" that sticks the sand, gravel, cobbles, and boulders together, sort of a natural "icecrete".
What we have been calling "asteroids" have lesser volatiles, what we have been calling "comets" have more, but I'd bet real money these are really just a spectrum of volatile content, not two distinct classes of objects. The drier ones are the really loose rubble piles.
I would think enclosing a small asteroid/comet object inside a pressure "shell" of some sort, rated in a dozen or so millibar pressure capability, and heating the body to the ice-melting point 0 deg C, would separate the metals and minerals as solids, and the volatiles as liquid water and gases. Spin the vessel a little to separate these materials centrifugally, and then pump the gases and liquids where you want them.
I'm wondering if the rings of Saturn might not be a happier hunting ground for volatiles, especially water. Does anybody know if we have a composition, and particle density, for any of those rings yet?
GW
Peroxide can be handled OK, but not at full strength. 50% peroxide in water is the max accepted strength for safe long-term storage. You distill it up to 90+%, use it, then dilute what you didn't use back down under 50%. At 90%, you've only got 3-5 days before it spontaneously decomposes, and very violently, too. That's why the scheme never really caught on for the submarines. But, ignition of peroxide with hydrocarbon is hypergolic.
I think what RobS is proposing with exhaust gas dilution is fuel+oxygen, diluted by combustion gases. It can be made to work, but you have to cool the gases before you can feed them back. Since heat transfer is intrinsically slow compared to chemistry, that'll be the limiting factor on any design.
As for making hydrogen peroxide, I dunno. Never looked into it before. But I know it's done in very large quantities industrially. I've seen tank car trains of under-50% going by at RR crossings.
GW
I ran a crude ideal chemistry balance that assumed a "hydrocarbon" at an empirical H to C ratio of 2:1, and ideal conversion to nothing but CO2 and H2O. It balanced out as: CH2 + 3H2O2 = CO2 + 4H2O. By mass, the required peroxide to fuel ratio is about 7.3. The steam to fuel mass ratio is about 5.1, and the CO2/fuel mass ratio is 3.1. It might balance a bit different with methane fuel, but crudely in the same ballpark, I think.
That's a lot of steam to dilute the fuel-oxygen 6000F reaction down to something cooled steel can contain. Burning fuel and air, the nitrogen in the air fulfills the same dilution role, so that fuel air reactions at stoichiometry are in the 4000 F neighborhood, not 6000 F. Air/fuel by mass is typically near 14-15 with hydrocarbons.
The diesel-peroxide submarine propulsion scheme was not as efficient as sucking air on the surface or through a snorkel, but it did allow short bursts at high power while very deeply submerged. They did it in the same diesel engine as ran diesel and air normally. But I know little of the details. Normally, diesels run very lean of stoichiometry, and control power by mixture ratio, there being no air throttle at all.
Exhaust gas feedback in cars is a very small percentage, and is done mainly to finish burning fuel-like pollutants that couldn't get burned completely before the exhaust valve opened. It's a hot enough feed that your EGR valve will fail from corrosion, sooner or later. Mine certainly did, just not long ago. It upsets air/fuel mixture just a tad when that happens, knocking about a mpg off of about 30 mpg.
GW
The longer form of what I posted just above is now posted for all to see at http://exrocketman.blogspot.com, in the article dated 4-23-12. The details and all the rationale are there. I included no illustrations.
GW
I didn't say it in my long post just above, but yeah, by all means take some experimental ISRU gear on the first mission and try it out. It's just not smart to count on it for crew survival on that first mission, because the probability is, it won't quite work right. Maybe not at all. That's just the nature of engineering development.
If the first mission really does get done right (and I think that is a low probability, given NASA's track record since 1972), then the second mission really could be based on the surface at one, at most two, sites, instead of LMO. Some better ISRU machinery prototypes could really get "wrung out" on a mission like that, but it's still just plain stupid to count on them for crew survival. My experience with engineering development (19 straight years) is that "second time up" still does not work well enough to serve. Doesn't matter what you are attempting, that's just pretty much a "given" in the real world.
That's why I'd like to see two properly-sequenced government missions before a corporate visionary takes over, like Musk. (Boeing and Lockheed-Martin = ULA sure as hell won't.) By that time, he will have both the lander and the ISRU technology, to really succeed at planting a proper base, one that might actually become a nascent colony.
Do it wrong or out of sequence, that colony just plain will not happen in the next century, at least. It'll take that base/colony a significant while (measured in years) to become self-supporting. That's been the history of things. That's also why ULA won't plant it: no short/near-term profit in doing it, unless the government pays them to do it. And I can pretty much guarantee you that it won't.
GW
I rather think that if we're lucky, there will be two government-funded "exploration" missions to Mars in the next 20 years. If we're really lucky, it'll be a coordinated and sequenced pair of mission objectives from the same government or consortium of governments.
If we're not so lucky, it'll be a duplicated first landing "exploration" by different governments. If we're not very lucky at all, there'll only be one government mission. Most likely, there won't be any. Not the way things have been going since 1972.
A government "exploration" mission is the enabling prerequisite for a privately-funded trip, generally speaking from 500 years' worth of history. The exception today is Musk/Spacex. But, he lacks a manned landing vehicle. His Falcon-9, Falcon-heavy, and manned Dragon, plus an inflatable spacehab module (perhaps from Bigelow), is pretty much all we need to get men to Mars orbit. But to the surface? Hmmmm.....
I consider it possible but rather unlikely that Musk might beat a government mission to put men on Mars, precisely because of the lander problem. One-way suicide missions are not the answer, although a Dragon might just be able to pull one of those off.
With chemical propulsion, and the rockets I named, it is fairly easy to stack up enough propellant modules, some engines, a spacehab module(s), some crew return Dragons, and maybe a lander or two in LEO. You end up throwing away all the propellant modules and perhaps all but the crew return capsules. But you don't need a gigantic rocket to do this. That gigantic rocket development development is why NASA is starved for money everywhere else, and it's just not necessary. Once you're at or above 25-ton shuttle payload sizes, only cost per unit payload to LEO matters anymore. We're there already, with Falcon-heavy and the Atlas-5 551 configuration.
With solid core nuke propulsion, you can do exactly the same slowboat job, except that you can keep most or all of your ship, and use it again for the second mission. If you can do that, then why not? Why launch all that stuff twice, and throw it all away both times? That's utterly stupid. Resurrecting NERVA for the Mars mission(s) is way to hell-and-gone far more important than developing another gigantic moon rocket from shuttle technology. Mars transit propulsion is what NERVA was originally for.
Radiation: two different types. Cosmic background 22-60 REM/year, astronauts currently allowed 50 REM/year, with an accumulated career limitation. Close enough even at worst I'd say we can get them there and back without killing them. But, one 2- or 3-year round trip is pretty much a career limit. No second trips. Solar flare: just requires some water and wastewater tanks around the crew as a suitable storm shelter. I'd make it the vehicle's flight deck, myself.
Microgravity disease: just build your vehicle as a long stack, put the hab on one end and the engines on the other. Spin it end-over-end. 4 rpm at 56m "radius" is one full gee. The ship should stack up way longer than 112 m long anyway. Also solves a world of problems cooking, bathing, toilet, etc. Just de-spin for maneuvers or docking operations, then re-spin for coasting cruise.
Consider sending the landers and landing supplies, and all the landing propellants, as separate vehicles assembled out of modules in LEO. Just send it/them one-way. Have it all waiting in LMO for the crew when they get there. Send the men with enough propellant to get home, in case rendezvous fails in LMO. Suspenders-and-belt. Dead crews are more expensive than anything else at all. Just ask NASA.
Myself, I'd solve the lander problem before I designed anything else. Those are dead-head payload. The more landings you make, the more landers you send, the more you have to launch from Earth, the more it all costs, and the less likely it'll ever get done. But, exploration is not "flag-and-footprints". What we did in Apollo with one trip-one landing was not really exploration at all. Don’t think like that. Too many still do.
If there were a single-stage re-usable lander, you don't have to send very many, yet you could still make a lot of landings all over Mars on that first mission. Chemical can't do that job. NERVA could. And push all the landing stuff to Mars, too.
Just some ideas to consider.
GW
Need rails? Use dirt path?
Sure, it's done all the time, especially in Australia. The only difference between a train of truck-trailers and a train on rails is friction. Any sort of wheels (pneumatic or solid) on roads, especially dirt roads, is very high friction. Rail friction is (2-3 order-of-magnitude) 100-1000 times lower.
If manufacture of the engine propellants is tough, then that's a serious issue to consider, since fuel use is proportional to the friction.
GW
It would be very hard to build a Bessemer converter that worked in a near-vacuum CO2 atmosphere. Direct reduction would be much better. Easier to do, because it's all closed reactor vessels that don't care what's outside.
You'll have to melt the sponge iron in an electric furnace, adding just the right amount of carbon and alloying elements. The electrical demand to power that is enormous.
We do it all the time here with our electric grid, but on Mars with no grid at all, that's going to take dedicated atomic power of large capacity. I would suggest a big water electrolysis plant powered by a big reactor. The electrolysis goes offline in favor of the furnace when making steel. That way the reactor itself runs a constant power.
GW
I knew of paper designs based around the NB-36 airframe for a test nuclear turbojet, but to my knowledge it was never actually flown. They did fly a test of a reactor in the NB-36, as a sort of safety prerequisite for the nuke turbojet test. That's the flight with the 500-mph-impact containment vessel. But to my knowledge, there was no heat transfer loop flown.
The turbojet test bed for the B-52 wasn't anything to do with the B-36. The initial B-52 designs (Boeing) were for a turboprop similar to the Tupolev Bear. That turboprop was tested on one or two old B-17's, replacing the bombardier/nose gunner compartment, making it a 5-engine airplane. When they (Boeing) decided to go turbojet, the engines were based on those powering the 6-engine all-turbojet B-47 (also a Boeing airplane). Not exactly the same engines, but a later model. GE's I think it was, but I could be wrong about details like that.
Convair/General Dynamics had a competitor (the B-60) to the B-52 that was based on its B-36. They swept the wings and tails, pointed the nose a bit, and put 8 turbojets under the wings, very similar to Boeing's final B-52 jet design. Convair's entry was the YB-60, Boeing's was the YB-52. As it turns out, Boeing won. The two aircraft were quite similar in characteristics, as it turns out.
Convair's "last gasp" in the strategic bomber business was the B-58 Hustler. Impressive aircraft in many ways, but a one-way suicide trip on its design mission, and no hold-at-a-failsafe-point capability. Compared to the B-36 and B-52, its practical range was way too limited, precisely because it flew so fast. Supersonic is very expensive in terms of fuel.
As General Dynamics, they (Convair) did a pretty good job, and made a lot of money selling, the F-111. Turned out to be a very good attack bomber. Not a fighter at all, in spite of the "F" designation. And there's the F-16, one of the finest fighters of all time. General Dynamic's other big division was/still is the Electric Boat Company, which builds most of the Navy's submarines, even today.
GW
ps - sorry, we seem to have wandered very far afield from the topic of terraforming Vesta.