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I quite agree with the notion of a public-private consortium of some kind. I quite agree that there are multiple launchers that could mount an expedition using orbital assembly in LEO. I quite agree that this could be done for well under $100 B, probably under $50 B, perhaps as little as $30B, but the right kind of team has to do it, and they have to have the right objective, and use the right approach.
And Apollo as we did it back then, is none of those things (hindsight is very 20-20 vision, I know).
If the consortium is more private than public, the politics plays less of a role. But as long as one of the government agencies leads and funds this, we will be saddled with politics. I suspect we're stuck with that for the exploration mission(s) and the first ISRU bases. After that, government may phase out of the game. The trick is to bootstrap through the politics, and the current economic bad times may be both a problem (no money to do anything right now) and an opportunity (a way to bootstrap out of economic depression).
We in the US experienced boom times in some sectors of the economy during the mobilization required to go to the moon, back during the 60's. That sort of thing would happen again, trying to go to Mars. It doesn't lift all boats, but it definitely lifts some of them. Even the Europeans might jump at the chance to get involved. It's a way to jump start your engineering and some of your manufacturing back into good times. They need that as much as we do. Japan, too.
GW
Sorry, I don't really think ISRU "has to wait". Didn't mean to imply that at all. There is some work that could be done starting with the very first landings, or even robotically, before we go.
I'm just trying to point out that such efforts are more of a wild guess than most folks want to admit (because it doesn't sell the project, I understand), precisely because the subsurface has yet to be sampled effectively. What ISRU equipment do you take, and exactly how do you plan to use it? That depends upon what is really there to utilize at your landing site. Kinda hard to choose when you don't really know what is there. And every site will be different, too. Don't forget that!
The first ISRU efforts are not going to be the ones that blossom into sustainability support for bases or colonies. History shows we humans do things more by trial and error, with a lot more error than success. The ISRU approaches that do work will be found later, by people already on the surface of Mars, trying to use the surprise resource bounties they found subsurface.
That's our history talking. That's the way we've always done it. Now we'll just do it in a stranger, more hostile environment. So what? We'll still do it.
GW
Looking at the LCROSS list, I see potential there. 5.5% water is intriguing, to say the least. Carbon compounds seem to be in short supply. There's a bunch of minerals and metals that could be useful, perhaps in minable percentages.
It would be really slick to set up some sort of fuel processing station using the water to create hydrogen and oxygen, and carbon shipped in from somewhere else (not Earth) to create methane. This thing could be mostly-automated, as might some sort of mineral/metal mining rig. Vanadium in particular is necessary for high-alloy steel.
Don't count nuclear stuff out "politically". That crap will have to be resolved, because nuclear power for electricity and propulsion will be necessary to do anything significant beyond Mars. Those destinations are just too faraway not to use it, and too interesting not to go. Reducing Isp potentials and frontal thrust densities is not the way to get there, excepting extreme-slowboat trajectories entirely unsuitable for human travel.
I think the moon is a really convenient, safe place to the the nuclear propulsion work that would take us way beyond where we were with NERVA 4 decades ago. And it did everything but actually get flown. We know it works, better than chemical rockets. I think it would power a dandy reusable single-stage "landing boat" for Mars. Something tough as an old boot, capable of landing in rough (interesting) terrains, and flyable hundreds if not thousands of times.
All these ideas, plus doing real astronomical science, are pretty good arguments for going back to the moon. Can't pitch it as "exploration" because we've already been there, but we could pitch it as a piece of the larger effort support exploration of Mars and those other faraway destinations. That's how the Dutch East India company justified planting its initial settlements in the Indonesia region. Good argument then, and now.
GW
Martian dirt with human garbage and sewage would be fertile enough to growth Earth-type vegetation, given water and sufficient atmosphere. I doubt 0.38 gee is that much a problem for vegetation, might still be insufficient for human health, we just don't know yet. The UV environment there is more than just a mite harsh for the plants. It would take some kind of transparent dome over your vegetable patch, one that filters out some but not all the UV. Tough structural design and materials selection problem; the "dome" might not end up being dome-shaped. I dunno what atmosphere might serve; it might not be breathable by us. Fair fraction of an atmosphere total, but less than one, would likely do. Lots of experiments could be down down here to see what might work. Harsh UV can be simulated with lamps. Even radiation can be used. Just not the lower gee, and it ain't zero, so experiments on ISS won't be very informative.
Just some odd thoughts.
GW
What's still missing is knowledge of what's more than 10 cm under the surface, in most places around Mars (or the moon, for that matter). That's where the real resources are that get used. The surface itself is rather barren and hostile. Decades of robots have yet to dig meters down, where those necessary answers are.
We could build robots like that, and it would make a huge difference. But in the end, whether we do or don't, there is a human capstone exploration that complements and completes what the robots have done, but still precedes any experimental base missions doing actual ISRU.
"Attractive landing sites" depends upon what's important to you. Flat plains are a good engineering proving ground for still-experimental vehicles, but usually are far from interesting geology and subsurface resource potentials that one might use for ISRU. Tough choice. To support the kind of exploration I advocate, the vehicle engineering proof testing needs to be already done (more likely on the moon than Mars, it's much closer).
That's part of the Apollo mistake we made in the rush to beat the Russians at the flag-and-footprints game. The first three landings on the moon were 99% engineering checkout and about 1% doing science of any sort (no rover car). We shouldn't go to Mars using that model. It's too dangerous to do vehicle experimentation that far away, and the knowledge return from the mission is too low (costliness critics become correct).
GW
I was using weight (mass) divided by wing planform (or flow blockage for broadside entry) area, which is very similar to ballistic coefficient. Load factor in aircraft design would be lift divided by weight. Load factor times wing loading would be lift divided by wing are, same as lift coefficient times dynamic pressure. Ballistic coefficient is what I was really trying to get at. If the craft has wings, wing loading is what they usually compare. If no wings, they usually do ballistic coefficient.
Aerogels might be a tad fragile, they are typically very insubstantial structures. Would need a surface layer to stop the flow right through it. Decades ago, I built a ramjet combustor liner out of a low-density silicate hobby potting compound from Cotronics in NY, reinforced by Nextel 312 fire curtain silicate cloth. I sealed the surface with a coat of Cotronics ceramic (silicate) adhesive cement. Because of the thermal gradient through the material, there was always uncracked cooler stuff supporting the stuff that cracked that was over the solid phase change temperature. Because of that, I could use it right to its meltpoint at 3200 F surface temperature. This was tough enough to withstand very violent pressure oscillations driving the engine into rich blowout instability. The temperature and the force levels are way beyond anything the fragile unreinforced shuttle tile ever endured.
GW
Too bad we never really explored the moon during Apollo. One landing per trip, unable to dig deeper than half a meter with a little manual core sampler? That didn't answer the two real questions (1) what all is there? and (2) where exactly is it? No two sites are alike, not here, not there, not anywhere.
It's time to go back with some men and robots together, and visit maybe a dozen sites on the moon in a single trip. That using landers, based from orbit. Stay for a while at each site, and dig / drill deep. Find out where the water really is, how pure it is, and how concentrated. Are there any unexpected organics? Are there any metals readily available?
I think until you know the answers to those two deceptively-simple exploration questions, you can't do any effective experiments toward in-situ resource utilization, because you don't know what resources are there, or where they are at your site. And until you do some experimental bases doing ISRU, colonies make no sense at all.
I do think we need to return to the moon. But to just repeat Apollo flag-and-footprints is just plain stupid. We need to stop thinking like Apollo. The worst mistake of all was one landing, one mission. Hindsight, yes, but it's past time to learn from those mistakes.
GW
PS - I'd be careful of polymer materials exposed to sunlight on the moon. UV destroys them, so does radiation. Both are quite worse there. Solar PV is often proposed as polymerics replacing crystalline. I don't even think that's a good idea here.
I see a lot of cart-before-the-horse problems in a lot of these discussions. That concerns me more than anything else. How can you experiment with in-situ resource utilization, until you know what resources are really there? How can you know what is really there, until you have looked beneath the surface? Meters, maybe kilometers, beneath? No two sites are alike. Not here, not anywhere.
The first manned mission is one of "exploration". It should finish answering the two deceptively-simple questions: (1) what all is there? and (2) where exactly is it? You don't do that with a landing at a single site, the way we did it on the moon. And I do mean those two questions exactly as I worded them. That is not Texas slang, although it does sound like it.
Men should be working with robots. It is not men vs robots, that is a false zero-sum budget game. Robots see only what they are programmed to see. Men can see what is actually there, if you don't train it out of them. Robots can go where men cannot. So you start with robot probes, and you add men to the mix in the final exploration mission.
I think it is stupid to go to all the trouble to send men to Mars, and just make one landing. Let's not do any more flag-and-footprints nonsense. It was a waste on the moon, it would be a waste on Mars.
Base instead from orbit, and visit dozens of sites, all in the one trip. Send down a lander, rover, drill rig, men and robots, and stay for a week or two at each site. It would really help if the lander is one stage, reusable. That's nuclear, by the way. Check it with the rocket equation for yourself. We all but flew the engine 4+ decades ago, and then quit, like fools.
A mission like that, "capstoning" all these decades of robot probes, could actually answer the two questions. It could be done with the rockets we have, using the orbital assembly techniques we have. Hotter nuke propulsion would help, but is not absolutely required to do this.
Then, the second mission plants a base or two at the most promising sites. That's when you find out how to live off the land, and what you might produce for trade back home. That might take more than a single mission to do. But once self-sustainability and a profit commodity have actually been identified, then a colony makes sense. Not until then.
GW
Oddball ideas for reusable / recoverable upper stages, or pieces of them
None of this is developed: we are merely at the idea generation stage. All concepts needs evaluation before any selection takes place. Classic “brainstorm” process.
Fact: lower wing loading / ballistic coefficient reduces skin temperatures during re-entry to about 2000 F (1094 C) if refractory protection is used. This makes wider choices of materials survivable.
Idea: inflatable extended aerosurface. Possibilities: reusable or sacrificial. Might not take the form of wings per se, but just a shield between the stage and the oncoming re-entry air stream, most likely deployed from the front end, so that bags plus stage protects the engines at the rear. Would require active attitude control throughout re-entry. Any such system is an inert weight penalty against payload.
Sacrificial inflatable: multilayered polymer bags, with outer layers the sacrificial ablative that is eroded away during re-entry. Not at all sure how to thermally isolate the layers from each other. Typical polymeric material ablation decomposition (pyrolysis) is around 600 F (316 C) material internal temperature. All polymeric materials are thermally destroyed as structural materials at 200-300 F (94-149 C) internal material temperatures. It is rather likely the typical Mach 1 / 20,000 foot airloads peak would rip these structures away if not jettisoned by then. In any event, ocean splashdown forces probably would rip them away. But, it would be nice to retain the inflated bags as floats on the sea surface, or as impact attenuators for landing on the land.
Reusable inflatable: multilayer bag, inner layer polymer inflatable, outer layer refractory ceramic fiber cloth. Not at all sure how to thermally isolate the layers, since the refractory is so much hotter than the polymer can withstand, and the refractory is not a gas-tight structure so it cannot be inflated. How to deflate and stow for re-use is also a big unknown and a huge technical risk, as compared to the sacrificial inflatable. This re-stow would most likely have to be done during descent before the (somewhat ill-defined) max airload point at Mach 1 / 20,000 feet.
Idea: refractory or ablative material applied to tankage lateral surface, stage re-enters broadside to oncoming stream, to reduce ballistic coefficient to inert weight divided by lateral flow blockage area. Might require active attitude control throughout re-entry, and would definitely require engines to be protected by retaining an interstage skirt ring as a shield (a weight penalty against payload). Simple spin stabilization is a possibility for attitude control.
Broadside with ablative: could be layer of cork, could be layer of a hard char-forming rubber, such as DC-93-104. Might even be “intumescent paint”. The rubber would require retention ribbons to retain the char layer once charred through. Probably not suitable for aluminum tanks, well-demonstrated heat protection scheme for steel ramjet combustor cases. Not sure about the paint, the cork would have to be thick enough not to char through. For any of these, a serious inert weight penalty against payload is incurred, because this is installing a re-entry heat shield all over the lateral surface of the stage. Plus, an interstage skirt ring must also be retained as a shield to protect the engines. Rolling the stage rapidly for stabilization might also help distribute the heat loads onto all the lateral surface, lowering the effecting char rates.
Broadside with refractory: could be tiles or blankets of low-density ceramic. If tiles, suggest ceramic fiber-reinforced composite to avoid the fragility experienced with shuttle tiles. Either way, a serious inert weight penalty against payload is incurred, because this is installing a re-entry heat shield all over the lateral surface of the stage. Plus, an interstage skirt ring must also be retained as a shield to protect the engines. The thickness and weight of the required shield might be somewhat reduced by spinning the stage for stabilization during re-entry, thus spreading the heat load evenly all around the stage. These kinds of materials will be very susceptible to impact force damage at sea or on land, and to internal porosity contamination by sea water or dirt.
Other ideas??? Such as recovering just the engines, not the tankage? Or recovering engines (dense and heavy) separately from the tankage (low density, lightweight, usually rather fragile)? Fragile tankage can be made stronger structurally by internal pressurization (such as early Atlas)!
The trade matrix for evaluation is going to be rather large and complicated, if the “brainstorm” process model is followed properly. We need a lot more ideas to evaluate, for one thing. It would probably take to real engineering analysis or test to fill some of the cells out in the trade matrix. Once filled out, some weighting factors get assigned to the various evaluation categories, and the trades can be evaluated. Assigning weight factors is not simple, either. But properly done, this process usually gives a good answer that can really be built.
Somebody take this on and run with it. I have to go back to work, and I have a book to write on the art and science of ramjets, before shuffling-off this mortal coil.
GW
Josh:
The idea of using what aircraft designers call low wing loading, and ballistics guys call low ballistic coefficient, is a good one for reducing peak skin temperatures during re-entry. It actually increases the total amount of heat that must be absorbed (a heat sink design issue), but skin temperatures are really the more critical issue, because that sets what materials can be used.
I’m not sure of the actual numbers, yours are probably better, but shuttle looked like a combat jet fighter, as I recall. That would be in the neighborhood of 100-200 pounds per square foot of wing planform (488-976 kg/sq.meter). (That’s converted at 4.88157 kg/sq.m per lb/sq.ft.; I’m used to working in US customary, because all those around me worked in it, although I speak metric, too.) The old one-man Mercury space capsule was near the 200 lb/sq.ft (976 kg/sq.m) figure, too.
What you want in order to achieve lower skin temperatures is a wing loading closer to that of a Piper Cub or Cessna 150: around 6-12 lb/sq.ft (29.3-58.6 kg/sq.m). That would be very difficult to achieve indeed in a practical design, but if you could, skin temperatures under 2000 F (1367 K, 1094 C) are possible . In principle, that makes a fabric-covered steel truss structure just like the old Piper Cub possible as a re-entry vehicle, as long as you use ceramic fiber fire curtain cloth for the skin covering.
Nextel 312, and whatever 400-series Nextel product has replaced it in recent decades, is alumino-silicate fiber, good to 2300 F (1533 K, 1260 C) without solid phase change cracking. Meltpoint is actually 3200 F (2033K, 1760 C), but you’ve lost structural integrity once the solid phase change occurs.
For a stage tank coming back, consider turning the thing sideways (broadside) to the oncoming stream. That gets you max stream blockage area for the weight. Especially for voluminous hydrogen tanks, your mass / blockage area might get down into that range without any aerosurfaces at all.
One would have to seriously question whether it could take the pressure loads, though. Especially as it decelerates through Mach 1 just about 20,000 feet altitude, while still hot, and weakened from that heat. That’s where Skylab and Shuttle Columbia’s cabin both broke up. Tough design problem.
Upper stages will likely require some sort of fixed or deployable aerosurface to hit ballistic coefficients (wing loadings) that low. That is substantial added inert weight, which is why I keep saying upper stage inert fractions under 10-15% are total nonsense in reusable designs. No one is yet really listening.
First stages are far easier, as Mach 10-or-less entry speeds are far less challenging, both heat- and force-wise. Under about Mach 3, you can even heat-sink your way through it with plastics, if the deceleration is about 4 gees or more. Steady state would require steel or titanium, but re-entry is not steady state.
On another note, I’m not at all sure the achievable Isp with liquid methane is very much higher than kerosenes. Both are hydrocarbons at crudely 2:1 hydrogen:carbon. It’s just that methane is inherently far cleaner of contaminants than kerosene, for far fewer practical injector-plugging troubles, and it doesn’t coke-up flow passages so easily upon overheating. The cryogenic nature of it can be used to absorb and dispose of (re-use, actually) waste heat, too.
There is very little difference among RP-1, Jet-A or A1 (same as JP-5), JP-10, JP-8, and K-1 camp fuel / lantern / heater kerosenes, except the filtering for cleanliness. Of those, the two commercial aircraft jet fuels are the cleanest, and they’re still noticeably dirty. Because of that, I am fast becoming a liquid methane fan. But be careful, the mods to a rocket engine to burn liquid methane instead of kerosene are not trivial; just ask XCOR, they’re the experts in LM-LOX.
BTW, Jet B (same as JP-4) is just kerosene cut with essentially a gasoline, to thin it down and lower its freezepoint.
GW
Merry Christmas and Happy New Year, guys.
Hoerner also had a lift book, in addition to his drag book. Same self-published thing. Probably obtainable from Amazon, although I have never looked for them there. I have both in my library.
There's way more in them than most of y'all would ever need. It's arranged a little backwards, as Hoerner was a German. Sort of reverse, like the grammar. Fun to look at, though.
Look in the supersonic section for projectile drag. pretty close to just about any launch rocket drag. His plots go up as far as M6. Most rockets leave the sensible air at about M2 to 3. Good enough.
The real trick is getting the "transonic drag rise" modeled - that's part of the max Q worry we've heard about for decades. Biggest CD at near-M1 just before the density starts tailing off fast as you rise up past 30-40,000 feet. Biggest forces on the structures.
GW
I have two big points to make about the costs of spaceflight, manned or unmanned.
One is the impact of a small supporting logistical tail, vs the traditional gigantic one. ULA supporting the shuttle at a billion dollars (or maybe more) for 25 metric tons max per flight is the wrong approach. Spacex at $2500/lb on Falcon-9 is more like it. Atlas-5 is similar, but watch that cost rise if something happens to Spacex! That gigantic entity of Boeing plus Lockmart requires a lot of cash to feed it. Too big is just plain bloated.
The 53 metric ton Falcon-Heavy is supposed to fly next year for the first time, priced at something around $800-1000/lb. With that coming to the market, why do we need a NASA SLS at billion-dollar shuttle prices and only 100-150 tons? We know how to dock and assemble in orbit now.
The other big point is the type of lead agency and contractors that can support such endeavors without exploding overruns. We don’t have that, and it kinda shows. What we have done for half a century since Sputnik is the wrong way to do it. It’s not about flag-and-footprints, it’s about real exploration.
I gave a paper on that very topic at last August’s Mars Society convention in Dallas, Texas. You can find that paper online at the Mars Society’s site, in its electronic archive. Or you can read a version of it on my blog site http://exrocketman.blogspot.com. Scroll down to the paper at date 7-25-11 titled “Going to Mars (or anywhere else nearby)”, and see also my second thoughts about the backup scheme, in the article dated 9-6-11 titled “Mars Mission Second Thoughts Illustrated”.
The gist of the exploration definition is getting the answer to two deceptively-simple questions: (1) what all is there? and (2) where exactly is it?
That wording is not Texas slang, I meant it exactly as written, word for word.
It means you land and you dig deep and you drill very deep. Drilling kilometers down, perhaps. You have to do this in a lot of sites, too. A real planetary survey. We never even did that on the moon, so we still don’t know what is really there, even today. And none of 4 decades’ worth of robot landers has actually answered those questions for Mars.
The gist of the “right team to do it” question is that the NASA we need is not the NASA we have, and the contractor base we need is not the contractor base we have. If we had the right team, we could go to Mars at any time for under $50B, and make dozens of landings in one trip. The right contractors would look more like a Spacex, an XCOR, or a Scaled Composites. I still don’t see any credible agency or entity to lead it, not in the US, nor in Europe or Japan. Japan may come the closest, but still misses the boat by a wide margin.
It would take too long to justify all these assertions here. I suggest you look at my convention paper, or at the two cited blog site articles.
There is a third idea in the conversation thread here: reusability. Implementing reusability in one form or another is a lot less effective than reducing logistical tails, toward reducing spaceflight costs. It’s also a very tough technological nut to crack, but it can be done, at least for lower stages.
There’s a third article on my blog site, dated 12-14-11 and titled “Reusability in Launch Rockets” that addresses what might be most fruitful things to attempt.
GW
Then again, you have to consider where the exposure standards came from: best guesses based on folks exposed. Linear dose rate models do not work. The low-exposure limits are very, very crude best-guesses based on aging Japanese A-bomb survivors, and troops exposed in tests in Nevada in the 50's.
Most old guys like me were exposed to a lot more radiation than 1 rem a year just from watching early-model TV's with unshielded Klystron tubes in them. Momma always said don't sit too close, but never knew why. Now we know: a lifetime's X-ray in a week to a month, sitting within 6 feet. I'm still here. Most of us this age still are (for a little while yet).
So, I don't really see much problem with 25-50 rem /year exposure "for a while". The career limits are more doubtful. I honestly don't know. The original WW2 standards were 25 rem a year, no career limit. A lot of those guys had problems, a lot didn't. Hard to know why and how.
I guess my point is that the low-dose standards are really guesses, not such hard science after all. The harder science is the high-dose standards. Things like 25-50 rem in a week, that's going to kill some percentage of those exposed, and within days. It's a certain thing. We've seen it and measured it, directly.
GW
Josh:
Use the simple jiggered rocket equation analysis technique to help you pick the problems to run with your trajectory code. dVo = Vex ln(MR), where Vex = Isp*gc is a useful approximation, and fprop = (MR-1)/MR, for which 1 = frop + fpay + finert. Actual dV = dVo/factor, to model gravity and drag losses. For lower stages flying in air and nearly vertically, I use factor = 1.10. For upper stages flying in vacuum and more horizontally, factor = 1.05. This kind of thing will get you started by landing you in the right ballpark.
Trajectory analysis takes more effort, but is more reliable. Usually, weight statements are no problem. Real thrust vs flowrate and real backpressure effects require some real knowledge of rocket engines and nozzles. The toughest nut to crack is realistic air drag. The best source of actual drag data that I know is Sighard F. Hoerner's "Fluid Dynamic Drag", which his widow published from her home until she died. I doubt it's available anymore, except in a library. Hoerner was one of the aerodynamics guys on the ME-109 before WW2. He had all kinds of stuff in his book, including hypersonics.
GW
For annual cosmic radiation exposure limits used for NASA astronauts, see Table 1 from http://srag.jsc.nasa.gov/Publications/T … chmemo.htm for the 50 rem that I have been using. You can glean from other sites that cosmic radiation fluxes in near-Earth space are modulated by the strength of the solar wind to between 60 rem and around 30 rem.
Most of the 22-year solar cycle, in-space exposures are within the 50 rem limit. On the surface, remember that exposure is cut in half, because the planet beneath your feet is a shield against half the sky. Shielding might not be true for small bodies like asteroids.
The problem is Table 2, the career exposure limits to cosmic radiation, which limit you to around 2 or 3 years in space. Remember, there really isn't a practical shielding technique for radiation this energetic, because of the secondary showers it creates.
For an exploration mission to Mars, a 2 year voyage is OK, but don't ask them to fly again, under these rules. With thick roofs or underground habitations on the moon and Mars, exposure should be tolerable, but will likely violate career limits after 5 years or so, due to secondary shower effects. The only real thing a meter or 10 of regolith can protect you from is solar flare radiation, not cosmic rays.
Shoot, we get hit with cosmic rays right here on Earth, some primary, some secondary shower coming down from our own atmosphere, the mass of which is the real shield. Our magnetic field turns solar flare particles, not cosmic rays.
BTW, average Earth natural background radiation (of all types) is around 0.3 rem annually, the top third of which is radioactive emissions from coal plants. This value varies widely around the planet by a factor of 10 or more. It's pretty variable.
GW
Like I said before, I hope the guy is right. We could use some clean fusion power.
This E-cat stuff would be outside the realm of "accepted science", if it is true, not a fraud. That doesn't bother me a bit. When the universe doesn't conform to our precious theories, it's time for some new theories, I always say.
But then, I'm an engineer.
GW
Cosmic radiation is not the bugaboo that everyone thinks. The max radiation exposure occurs during solar minimum, and is just a tad beyond the yearly dose we now allow astronauts to receive anyway. At solar maximum, this exposure is cut in half by the solar wind, so that for most of the 22-year sunspot cycle, cosmic ray doses are under what is already allowed.
The real radiation danger is not cosmic rays, it is solar coronal mass ejections. Those, not cosmic rays, are what our magnetic field shields us against. Mars has none. Fortunately, these are brief events, a few hours to a day or so. About a meter of water or dirt works pretty good as a shield. Nothing special there.
GW
Josh:
That's exactly what I did writing codes like that long ago. It'll work. Going to multiple stages is not very hard, once you get the basic algorithm working. You just shift weight statement and drag, plus any thrust controls, at staging. It'll take some idiotic fractional time step to exactly hit the stagepoint. That's the hardest part, and it's not that bad.
GW
I'm thinking some sort of steel-making plant is one of the first things a permanent base will need. Shipping the plant once (whatever it really is) is cheaper over the life of the base (decades+) than shipping steel stocks from Earth, almost no matter what the cost to LEO is.
The other is some sort of plastics-making plant. Not everything should be made of steel. Aluminum can come later. If you've got plastics and steel, you can pretty well cope for a while.
On Mars, concrete is going to be a bit of a problem, as limestone does not seem to be available, although water is. There has to be some equivalent with the minerals widely available there. It may take a while experimenting before we find it. But find it we must, concrete is just too useful to do without it. An ice-regolith composite reinforced by steel bars might serve in some applications, as long as material temperatures do not exceed 0 C.
GW
I assume you are writing a finite-difference solution to the equations of motion, presumably in 2-D. It won't matter much if you do it 2-D Cartesian or round-Earth. The easiest version is a 2 degree-of-freedom flying particle in 2-D Cartesian, I'd start there.
If at any given time step, you compute all the force vectors (magnitude and direction) acting on the vehicle (thrust, weight, air drag), their sum and the current particle mass gives you the acceleration vector direction and magnitude. This force sum includes the weight force, which then automatically accounts for the "gravity loss" as you "integrate" the trajectory. Acceleration times time step is the delta-vee vector increment to the next step. This delta-vee times time step is the displacement to the next step. The velocity change and position change take place in the direction of the acceleration vector, so you do components to figure the new 2-D vector position. It is easier and more useful to keep track of velocity magnitude tangent to the current path, as that is how drag is figured. This gives you the position and velocity at the next time step, where you do it all over again.
You will need to build a "model" of your vehicle with many facets. First is a good thrust model. That can be a large topic, even for a rocket, as delivered thrust and impulse are functions of backpressure and exit area, even at constant propellant flow rate. Second is a good weight statement, such as I had been calculating for Falcon-9 and the reusability trades I just did. Third is a good drag model, which requires coefficient vs at least Mach, and a suitable reference area, for each shape the vehicle takes on as it stages. The coefficients are most definitely not constants. You can pretty well zero the drag forces above around 200,000 feet, no matter how fast you are flying, as the density is becoming too low. But, you will need a model atmosphere, so you can figure speed of sound from ambient temperature, to calculate Mach for looking up the drag.
You have to keep the time step very small. Make no more than a 10% change in any physical condition such as velocity, weight, or thrust, in any given step. 1% would be more accurate. It is possible to use this to program an adaptive time step that is very fine near launch, but "steps out" bigger as the vehicle flies fast. You can do this in the very simple forward-stepping procedure I described above quite accurately. You don't need Runge-Kutta integration, that was for the ancient computers with kilohertz or slower processing speeds.
The hardest part is initial conditions. If you are doing an adaptive time step, it may get hung up at launch. Sometimes you have to give the vehicle a trivial upward velocity to succeed with the calculation. Maybe 30 cm/sec.
And all of that is what it takes for non-lifting gravity turn flight. Winged lifting vehicles require at least a 3-degree of freedom model that includes moment sums and pitch inertia, for just a 2-D Cartesian model. They also require pitch control as a user input, and a description of lift curve slope vs Mach, plus drag-due-to-lift vs Mach. Not for the amateur.
Hope that guidance helps.
GW
Last data I saw posted / projected for Falcon-heavy said $800-1000/lb to LEO, at 53 metric ton sizes. Of what possible attraction is an SLS at 8,000-10,000/lb? Only larger payloads? So what? We know how to rendezvous and dock.
There's very little we would want to do in the next few decades that we couldn't do with Falcon-heavy. I haven't been to Spacex's site lately, so I am not (yet) acquainted with any Falcon-super-heavy. But, Falcon-heavy is supposed to fly sometime next year out of Vandenburg AFB. I know they are building a new thrust stand here in McGregor, Texas, to accommodate an all-up Falcon-heavy test, at reduced noise.
GW
Don't worry about imperial units for me, I speak both. I just think a bit better in imperial, because that's what we used the most when I worked in the defense industry long ago. It's just practice, not a problem understanding.
Hadn't thought about methanol, might be a good idea. Not so sure about methanol-in-contact-with-N2O4 or any other oxidizer. Pre-mixed flammables and explosives are a bit scary. A lot of folks don't realize that methanol is a skin-absorption nerve poison. It is possible to absorb a lethal dose through the skin of your hands, if you keep them immersed for a few hours in a day. But, it's a very nice fuel material, solvent, etc, if you just treat it with a tad of respect. It is a bit corrosive to a lot of materials, especially polymers. Steel works OK, stainless is better. Aluminum, not so much.
GW
JoshNH4H:
I'm fast becoming a methane convert myself. Easier to make, handle, and store than hydrogen, no matter where in the solar system you are. Also usable in ramjets instead of kerosene (common fuel rocket and ramjet). I know XCOR is getting excited about methane fuel, too. Cleaner by far than any kerosene available. Far fewer injector orifice clogging problems. I kinda like all of that.
GW
Sure, but it's still deep burial for clathrates. It would be around 200 meters at 2C on Earth under a dirt overburden (that's figured at about 105 lb/cu.ft for loose dirt, vs 62.4 lb/cu.ft for fresh water). On Mars at 0.38 gee, it's almost 3 times that depth. That's a lot of digging. Building pressure tanks is probably easier.
One of the first things to import is probably a steel mill. Ha! Ha!
GW
The methane clathrate is stable on Earth in deep cold water. Most of what we know about is in the 0-2 C range, hotter comes apart. Most of what we know about is 300 meters deep or deeper. Shallower is unstable.
I rather like the plain methane myself. If we will be using chemical propulsion, it make a lot of sense in the first stage, because the volume is lower (less structure). I dunno about Mars, but on Earth, the first stage of a rocket flight is thrust-limited, not Isp-limited. Lower-Isp fuels like methane make better sense because of their higher density.
It is the second stage that is more Isp-limited, where hydrogen itself makes more sense. Methane and oxygen are easier to process and store than hydrogen, so for practicality, all-methane-LOX makes a lot of sense, even for an upper stage.
GW