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Terraformer:
I had 3.455 km/sec for 200 km circular orbit, and 3.405 km/sec for the apogee speed on a transfer ellipse that just hit the surface. That's 50 m/sec to de-orbit, something an attitude thruster can do. Using the Justus & Braun report, atmospheric interface occurs at 135 km on Mars. That was at 1.63 degrees below horizontal, and 3.469 km/sec entry speed (it increased as we descended).
From there I used the 2-D Cartesian model described in Justus & Braun of the old 1956-vintage density scale height analysis. It took me from 3.6 km/sec to 0.6-0.7 km/sec at varying altitudes, depending upon ballistic coefficient. The range of altitudes was 29 km at beta=100 kg.sq.m, down to 3 km at beta = 2000 kg/sq.m. Nonlinear curve. Not too steeply bent. I never saw peak gees above about 0.7.
The landers (including MSL/Curiosity) cluster around beta 100, being 63 up to about 115 for MSL. A manned lander is probably beta 1200 kg./sq.m, in the 60 ton entry mass class. Plus or minus maybe 500 kg/sq.m.
I haven't run data for a stronger de-orbit burn. That will certainly lower the entry interface velocity, but it will also steepen the trajectory, for sure. Steeper path tends to makes it dive deeper before the Mach 3 point. Interesting question, I need to look at that.
GW
OK guys, I've run one Mars entry study using a range of ballistic coefficients with the old 1956-vintage scale-height method. Those are posted over at "exrocketman", with a gob of spreadsheet images and plots. The net outcome shows end of entry nearer 20-30 km altitudes in the 100 kg/sq.m class like the Mars probes, but under 10 km when in the 1000-1500 kg/sq.m class, more like a manned lander. This is for entry from LMO circular orbit, not direct entry. Direct penetrates deeper even at the same angle, and that's likely to be steeper as well.
I'm going to refine the density scale height correlation for the lower altitudes, and try this again, centered on about 1200 kg/sq.m. That's just about what I'd expect a manned lander to be, at around 60 tons entry mass. I think it'll be pretty close to end-of-hypersonics (at Mach 3), pretty close to 7 km altitude. That's too low to deploy a parachute, there isn't time to get it open, much less slow down significantly with it, in that thin "air".
It's looking like rocket braking-to-touchdown to me. These numbers are for no thrust during the hypersonics, BTW. Just the supersonics.
As for brainstorming configurations, here are two notions I'd like to pitch into the punchbowl and see who screams: (1) who says a heat shield has to be round? (2) who says we have to drop the heat shield to land, if it actually serves a purpose landing?
Those are in addition to the idea I threw in the other day, that you can fire a rocket through a hole in a heat shield, if the compartment containing the rocket is sealed gas-tight to prevent throughflow through the hole.
Anybody else come up with some hare-brained configuration ideas, with those three crazy items in the mix?
GW
"exrocketman" is http://exrocketman.blogspot.com
I haven't posted my stuff yet, but I did estimate a range of ballistic coefficients for a 6 ton (metric) dragon, as 380 to 453 kg/sq.m, depending upon whether you believe the hypersonic drag coefficient is closer to Apollo's 1.3 vs closer to Mercury/Gemini's 1.55.
I ran the old 1956-vintage "back-of-the-envelope" entry model reported by Justus and Braun, in their "typical" Mars atmosphere, at a nominal 500 kg/sq.m, and got a terminal (Mach 3) altitude for the entry of 15 km, from a circular 200 km orbit entry. A capsule entering at interplanetary speeds would penetrate even deeper.
My estimates for ballistic coefficient of Mercury, Gemini, and Apollo range from 246 to 374 kg/sq.m at entry, with Gemini the lowest and Apollo the highest. That's for hypersonic drag coefficients of 1.3 for Apollo, and 1.55 for Mercury and Gemini, based on the the hypersonic drag data reported in the old Hoerner "drag bible". The recent Mars landers all cluster around ballistic coefficient 100 kg/sq.m, with Curiosity near 115, and most of the rest nearer 63 to 90-something.
The problem with Mars entry is the parachute phase. Once you come out of hypersonics at about Mach 3, the parachute terminal velocity is still supersonic in the Mach 1 to 2 range. The 200 mph stuff I see on the internet news is just BS. It takes finite time to deploy a supersonic chute, and then slow down with it. If your terminal altitude is too low, the chute does you no good, you might as well simply rocket brake, only.
My guesses for a Dragon equipped with Super Dracos and landing legs says it comes out of the hypersonics somewhere near 15 km altitude, with no retro thrust during entry. I have yet to run numbers, but I'd think it could rocket-brake from there to touchdown pretty effectively, since the Super Dracos are very "thrusty". You just need enough on-board propellant to carry it out. That propellant mass will displace deliverable cargo.
I'm getting entry decel gees near 0.7 from circular. And braking gees near 3-ish. Nothing final yet. But rocket retro thrust can simply be on the capsule. You just need T/W enough to fly "smartly" upward on Earth, maybe a tad more. That's what the Super Dracos were designed to do, for launch escape.
A "skycrane" is not the only answer. Although it might have been the lightest answer for Curiosity. And the rocket-braking gees are not all that bad.
It'll be worse still for a big lander in the 60 ton class. But not that bad. We could easily ride it.
GW
Hi Rune:
Ah yes. School, summers, and girls. I do remember, even after all these decades.
My calcs are for entry from LMO, not direct entry. Higher betas and higher entry speeds very quickly push you toward hypersonic impact with the surface. So does steepening the depression angle, which may actually be the most dangerous actor. It's not much problem from LMO, but could easily be a problem with direct entry. In those cases you must retro thrust during entry. There is no other option.
I've got the orbit mechanics all worked out, and the back-of-the-envelope entry model seems to produce reasonable results. I ran a exploratory spread of betas vs M3 altitudes last night, bracketing all the way from the small probes to 100 tons. As soon as I get it saucered-and-blowed, I'll post it.
What do you think of a 7.1 meter diameter cone about 7 me tall, for 60 tons? That's right at beta 1174 if you use an Apollo-like drag coefficient of 1.3. I get an outer envelope effective packing density of 654 kg/cu.m with numbers like that. Seems tight, but still sort of reasonable for at least a one-way lander. This does not take into account the possibility of an inflatable heat shield like the IRVE tests. That thing isn't "ready for prime time" yet.
HA > H transliterates as "the number of horses' asses always exceeds the number of horses".
GW
This is very, very preliminary, but I've come to believe a conical lander of a shape similar to Apollo, but around 60 metric tons, will have a ballistic coefficient (beta) near 1200 kg/sq.m. From 200 km LMO, entry will start about 3.47 km/s at about 135 km, and about 1.6 degrees angle of depression. That requires a mere 50 m/s de-orbit burn. My back-of-the-envelope entry model suggests that this thing comes out of the hypersonics at M3 and around 7 km altitude. That would be about 720 m/s at 7000 m, angled about 2 degrees down.
That's only mere seconds from impact, so parachutes will definitely do no good. So, I'm thinking heavy rocket braking right from the end of the hypersonics, aimed at reducing the horizontal velocity quickly to zero, before one runs out of altitude. I will have to build a good model and develop it, but some very crude pencil and paper stuff says the thrust levels will be around 1000-2000 KN on that 60 ton vehicle, and thrust accelerations will be in the 2-3 gee range. The time from end of hypersonics to touchdown will be a minute or less.
That seems challenging but doable. It seems that my supersonics-to-touchdown heavy rocket braking model may be a worthwhile objective. This is with no rocket braking during the hypersonics, and no parachutes or ballutes. I will pursue this as time permits in the late evenings. Have patience, don't hold your breath.
As I get stuff done, I will post it over at "exrocketman". There's an update to the ballistic coefficient correlation over there right now.
GW
http://exrocketman.blogspot.com
"HA > H explains an awful lot of what's going on in this world"
Lessee if I am understanding things correctly.
Is not "soil" for growing plants not a mixture of rock dust and organic material, with some of that organic material supplying texture and water absorption properties, and some of it the nutrients?
If I am correct about "soil", then we mix rock dust from Mars with wet sewage from the astronauts, plus a little fibrous cellulosic waste (shredded paper works, but shredded stems and leaves work better, later on). We put this under an appropriate atmosphere at an appropriate pressure, shield it from too much UV and nuclear radiation, and plant crops in it.
As for Martian microbes, the odds are they are either quite closely related to us by the panspermia hypothesis, or else they are utterly alien. In either case, we're likely OK, and so are they. It wouldn't take very long to find out on one of the probes, or from a sample return, but we need to get on with that task.
Some of us have pretty much figured out how to get people there within 5-10 years. So has Spacex.
GW
IMHO the ISS has taught us two very important things: (1) how to assemble very large structures from docked modules, and (2) some fuzzy but useful practical limits for exposure to microgravity.
Item (1) is essential to going to destinations beyond the moon. The only practical way to do things like that is assembly of the vehicle in orbit. The most mass conservative architectures result when you do an orbit-to-orbit transit vehicle built for at least partial re-use, with appropriate landers for use at destination. These are the lessons of Apollo, even though we didn't do Apollo itself that way.
Item (2) is just observational. I notice they rotate crews every 6-7 months. It is not that hard to fully recovery from microgravity disease. There is a cosmonaut who recovered from 440 days, but it took him a long time for a difficult recovery. That says we need artificial gravity, and not knowing any better, it ought to be 1 full gee, or pretty near.
What the ISS could do for us next is answer the question "how much gee is enough?". That will require the medical centrifuge that they never launched. I hear it still exists, and if an Atlas-V-552 can't fling it up there, then a Falcon-Heavy soon will be able to.
GW
Back to 3-D printers: I saw recently that they were 3-D printing in powdered metal now. That's pretty slick. And yesterday I saw that NASA is trying out some sort of 3-D printing rig in the zero-gee airplane. The idea has definite potential.
GW
My two-way delta-vee of 8 km/sec is probably a little high, except for realistic plane changes. Ballpark realistic under real 3-D conditions, I think.
I used heavier-than-min-calculated consumables mass to cover using frozen food. That freeze-dried non-refrigerated stuff they use now doesn't last more than about 1-1.5 years. Frozen does, it'll go decades if need be. My numbers are just ballpark guesses.
GW
Sounds like somebody at Spacex is coming to some of the same conclusions as I did. I haven't looked at solar electric yet, but they are. Those are good things. If Spacex stays on that path, they will beat NASA to Mars and everywhere else. Bravo for them.
Their only trouble with nuclear thermal is that all things nuclear in the US are a government monopoly. Red tape in abundance at best.
GW
Today's internet news: the IRVE-3 test was a success. Mach 10 speeds, temperatures to 1000 F. No other hard data could be gleaned from the story. Looks very promising, though.
GW
Hi Bob:
Chemical SSTO does indeed seem possible once one can achieve stage inert mass fractions in the 5% range. I do have some very serious doubts about such stages ever being reusable in any way. That's the sort of inert weight of something like the Gossamer Condor, fragile beyond belief. No wood and fabric, metal and fabric, or all-metal airplane has ever been strong enough to fly routinely with a good safety record below an inert fraction of about 40%. The ones with the best survival-of-battle-damage records are closer to 50% inerts.
I know space launch vehicles are not airplanes, but you see my point. There is some sort of a lower limit for inert weight fractions that really are going to be reusable. I think they're way above 10% (shuttle SRB's were near that 10% figure, and were not actually fully reusable - there was a significant loss rate of motor case segments).
Last time I thought about space-faring airframes, I guessed about half and half composites and metals, and pivoted off that 40% all-metal aircraft figure with composites twice the strength-to-weight. You cannot use all-composite, they have a very low tolerance for heating.
I was using about 27% inert for an airframe that might conceivably fly into space reusably thousands of times (or more). At something in the neighborhood of 27%, no chemical rocket is capable of SSTO, but some nuclear rockets might be. The only other option was airbreathing propulsion, but I found scramjet a dead end, even assuming it could be made technologically ready (they've been trying since about 1960).
That was a non-intuitive dead end, too. Few people even today know it's a dead end. To get the thrust out of the engine to climb, you have to stay deeper in the atmosphere as you accelerate. But that's where drag is higher. Beyond about Mach 10 or 12 or so, you incur drag losses faster than you can maintain climb rate, and those drag losses eat up all your gains from the higher-than-rocket Isp. Dead end. I found that answer for almost nothing (pencil, paper, some hours of my time). The X-30 required billions of dollars to find the same answer (but they never actually publicly admitted what they found).
So, I think there's vertical launch rockets and there's TSTO spaceplanes (unless you allow nuke rocket spaceplanes). Any airbreathing assist occurs during the lower altitude phases early in the trajectory for both.
On a vertical launch rocket, you've pretty well left the usable atmosphere around 70,000 feet, at speeds around M 2 to at most 3. I could put ramjet strap-ons on the first stage, but they'd be simple pitot-inlet designs, running from about M0.5-ish to at most about M2.5-ish. That's 1945 technology, except that I would use a 1970-vintage dump combustor instead of a baffle stabilizer. Much better suited to variable-speed operation.
For the TSTO spaceplane, I would put both rockets and a high-speed ramjet engine in the first stage. Rocket off the runway to about M1.5-M1.6 takeover with an external-compression inlet. Ramjet climb at low Mach (1.6 to 2-ish), pullover and accelerate to about M5-6-ish at 60,000 feet, light the rockets, too, and pull up at 45 degrees to release the second stage. Cut rockets and chop ramjet throttle to decelerate in a 180 degree turn, and cruise back to base at M1.5-ish for a glide landing. Rocket power for emergency go-around or divert. The second stage is just a plain old rocket plane.
The ramjet would have the supersonic inlet, a C-D nozzle of throat area about 65% of engine flow area, and a dump combustor design. All fixed geometry, except possibly a translating inlet spike to maintain shock-on-lip operation. It likely would completely fill (be) the fuselage. The wings and fillets would contain the rockets and all the propellants.
GW
Russel:
What you describe as trailed bodies sounds sort of like a series of badminton birdies connected together as a linear string. These are flying in each other's wakes, in regions of ever-increasing dynamic pressure deficit as you look downstream.
Try it with actual badminton birdies connected with string, and film it with a high-quality camera. I think you will find on close analysis of the photos, there isn't much tension in the string. It'll likely go completely slack. In the limit of hypersonics, I think these things will want to rear-end each other, as only the lead "birdie" sees most of the drag.
You can offset the drag deficits looking downstream by going to extremely low mass. But, very lightweight structures usually cannot survive the heating. There's exceptions: that IRVE series of inflatable heat shields is one. They seem to have some sort of sealed Kevlar "inner tubes" as the inflation members, some sort of a low-density insulating layer, and what they describe as an elastomer-impregnated fabric surface layer as the actual ablative surface. I have no idea what the low density insulating layer is, and very little idea what the flexible ablative elastomer/fabric combination is.
IRVE is essentially all one big badminton "birdie" as an extended heat shield surface for a fixed mass, to lower the ballistic coefficient. To do your trailed thing this way, there must be some sort of size and mass distributions that put the higher-drag lower mass items further out to the rear. The only way I really see to do that is to make the ones further downstream bigger and bigger, yet lighter and lighter. That way, at least a part of the object pokes out into undisturbed slipstream at full dynamic pressure. That's the only way I can think of to keep tensions in all the towline segments.
So, your train of drag shapes is starting to look like a train of carefully sized inflatable cones, perhaps porous, with the larger ones further behind. Might work, but how one might actually build such a thing, I dunno.
GW
Lessee, NeoSM post 246 has seen some actual studies that suggest up to 7.5-10 rpm may be tolerable in stages. The old 3 rpm figure was definitely too conservative. My 4 rpm figure was an impression, an educated guess, from the old centrifuge studies for pilot gee tolerance, plus experiences playing on a merry-go-round as a child. Nice to see it was roughly correct.
I'd also suggest my orbit-to-orbit transport designs are more appropriately included under the interplanetary transportation topic, or here under human missions as a separate thread.
That last one I posted over at "exrocketman" was more of a demonstrator for "easy" artificial gravity than a proper mission design, although the vehicles so sized resemble those I got for the paper at last year's convention. The key to artificial gravity is modular spacecraft design. That allows one to adjust the form factor at any given mass to get the radius you need to use 3-4 rpm for 1 full gee. With 1 gee, you arrive fit at both ends of the trip.
Russel post 247 has an interesting idea: the trailed drag item, like a long windsock or streamer. I have done a thing like that before, and supersonically, too. It worked, but it's not a lot of drag. There's more drag the longer it is, but not as much as you think, because it's immersed in the vehicle's wake, where there is a dynamic pressure deficit (that persists to very far distances behind the vehicle). What we did was grocery store grape bag netting, in tubular form, as the stabilizing drogue on a towed ribbon decoy. It worked in the wind tunnel to Mach 1.4. We switched to a ribbon chute drogue for packaging reasons.
For entry hypersonics, different forms and materials would be required. That wake zone is not only dynamic pressure-deficient, it is also incandescently hot. A stream of intensely-ionized plasma, as a matter of fact. I don't know what form or shape to use, but a long ribbon of something could be streamed out for kilometers behind. It would have to be made of some kind of ablating sacrificial material that was also very flexible.
Hmmmmm. I wonder if the inflatable heat shield cover fabric isn't canvas impregnated with something like DC-93-104, the silicone hard char-forming rubber? That kind of thing might work for a trailing ribbon, too.
Russel is probably right in these early days , in his preference for landing crew and cargo separately. That eases the entry vehicle design problem. It does incur the precision landing problem, which in Mars's super-thin atmosphere will require active guidance and steering during entry hypersonics. There is no time or cross-range capability left once the hypersonics are over, you are simply too low.
With enough propulsion, you can do entry retro thrust to terminate hypersonics at much higher altitudes. The low thrust level simply "offsets" some of the mass, making the ballistic coefficient "appear" lower. You're not going to do much of that with NTO-MMH at 280 sec Isp. But later on, with a nuke rocket in the 900-1000 sec Isp class, well, big single-stage fully reusable "landing boat" designs start to look really good. 60 ton vehicle, 6 ton dead-head payload, down and back up on one fill-up. That sort of thing.
I just wish we had a working nuke rocket that used water instead of LH2. Sure would make in-situ refueling easier.
GW
OK, as promised above, I refined that impromptu Mars mission study, and published it over at http://exrocketman.blogspot.com, as the article dated 7-19-12. The vehicle designs are "reasonable", and I accomplished my purpose of showing how easy it can be to incorporate 1 full gee of artificial spin gravity.
Scroll on down and I have posted articles on a simplified entry model, and good atmosphere data for Mars and Titan. There's some gravity data, too. Enjoy.
Now I can go back to my real project: roughing-out the design of some big but practical Mars landers.
GW
Take a look at what Paul Webb did ca. 1969 with the pantyhose-material MCP suit. It was a multi-piece garment, multiple layers. It was porous, non-gas-tight at all. The test subject sweated right through it, both in atmosphere tests and in the vacuum tank. They tested half an hour at simulated 87,000 ft, no ill effects, sweating through the garment with no ice formation.
The two mid-60's vacuum exposure accidents in NASA's vacuum facilities produced evaporative frost and frostbite injuries to wet tissues only. The mouth, nose, and eyes are at risk. Ordinary skin does not do that. Nor does it swell or burst (nor do the eyes), except in response to internal tissue edema, which takes 10 to 30 minute's exposure to set in.
Same results for the 3 cosmonauts who died in the Soyuz depressurization accident during reentry. 10 minutes vacuum exposure. No noticeable freezing injuries, no swelling of anything, just blue color from total anoxia. Non-rescusitatable. They tried. Just too long gone. About 1.5-2 minutes is it. Brain death too far gone.
GW
If you are referring to retro thrust during supersonic or hypersonic flight, then, no I don't think an aerospike would help provide directional stability to the plume. But cant angle would. Aerospike (and the other semi-free expansion designs) merely match expansion to a varying backpressure. Stability has to do with which way the plume turns when it reverses. That process induces moments and side loads on the vehicle. If the plume is flip-flopping around unsteadily, you could tumble the vehicle.
GW
If you stabilize the object so it doesn't tumble, and build it stout enough so that it's tough enough to take the air loads, then your only real worry is the dings and dents from whacking the sea, even on a chute. Most of the time, shuttle SRB segments could be re-used, but not always. These were steel pressure vessels designed to operate routinely at about 900 psi chamber pressure. It was ocean impact that ruined the ones we couldn't re-use. Lightweight unpressurized liquid tankage gets smashed to rubble by ocean impact, even if the air loads don't crush it. That's why no Falcon stages have been re-used yet.
GW
Think MCP-as-vacuum-protective underwear. Then add whatever coveralls and hats and gloves and boots you need for physical protection from heat, cold, and other hazards. It's exactly the same sort of outer clothing we use down here, as a matter of fact. The MCP rig does not need a cooling system: the tight cloth is porous, you sweat right through it into vacuum.
GW
Shipping livestock is no problem in a big nuclear pulse propulsion (old Project Orion) atomic ship. That is the kind of ship you need to plant colonies anyway. I'd go for about 20,000 tons, built of steel armor plate, in something resembling a marine shipyard. Launch it just once. Leave it in space and re-use it for a century or two as an orbit-to-orbit transport. Build single-stage re-usable "landing boats" for it that are powered by solid core nuke rockets. Ship huge herds of any critters you want, anywhere you want to go. These ships are big enough to spin for 1 full gee about the long axis, not head-over-heels.
GW
OK, I revisited the calculations and did a bit of convergence here and there with a spreadsheet to help me out. I'll post that stuff over at "exrocketman" in the next day or two. Lots of figures and spreadsheet images. It'll take me a while just to get it done. Everything is built out of multiple examples of common modules for hab, supply storage, propellant and engine. The landers are just a mass allowance. I haven't added up the final launch costs, but it's 6 men, and two separate landings 3 men at at time.
I built up two vehicles in LEO, one manned, one not. The manned vehicle returns to LEO using propellant modules sent to LMO by the unmanned vehicle. Everything else gets left behind in LMO. The manned vehicle spins head-over-heels at 3 rpm for 1 full gee artificial gravity both ways. My Hohmann transfer delta-Vees are pretty crude overestimates, intended to cover plane changes and propellant boiloff: I made the vehicles capable of 8 km/sec (1-way) LEO to LMO. I used 8.5 months 1-way, and a 27 month round trip to estimate supplies. I was using about 15 kg food+water+makeup O2 per person per day, but that's just a wild guess.
I used the old NERVA data because that's the only design that ever got any significant testing, to the best of my knowledge. It was essentially flight-qualified when they shut it all down. There were other better ideas, but none got the massive testing that NERVA got, to the best I know. Any better nuke just makes things look better. My finished vehicles are closer to 800-something tons than the 580-ish I originally thought. The NERVA modules are smaller and lighter, too. I sized for 0.05 gee ignition acceleration or better, same engine all three configurations at the same thrust.
Can't test nuke engines in orbit, must test on a celestial body somewhere. Rocket development testing cannot be done hanging weightless in space where every test is a flight test. You must have a stable thrust stand that does not move. Fact of life. If we can't do it here at home, then the moon is close enough to reach without nuke power.
GW
Most of the aerodynamics data for forces was correlated to flat planar areas. I'm guessing you are talking about dead-broadside return of these stages. If so, length x diameter is probably the reference area you want, and the broadside drag coefficients for circular cylinders apply. That stuff is a strong function of Mach between about 0.75M to about M3.
Nice to see such low numbers. That'll decelerate more, higher up in the thinner air, for sure. Ride might be rougher, just like a Piper Cub vs a jumbo jet. I'd worry about air pressure crush as it decelerates through about M1 about 20,000 feet. Structures that light are very flimsy. Air pressure crush after entry was over is what broke up both Skylab and shuttle Columbia's cabin section.
GW
Hey, I provoked a good run of conversation with that modular Mars mission post! (#226 above)! That's what it's all about.
I do think I need to re-run those calculations. I think I messed up reconciling available volume with contained mass for the LH2 propellant tanks. The concept is still "in the ballpark", though.
The inflatable habitat idea is just the thing. Bigelow comes to mind. I think it needs to be combined with two other things: (1) meteoroid "armor" made of alternating layers of foam and foil about 15 inches thick, and (2) the design constraint that access to the pressure shell be unobstructed for fast puncture repair.
Aerocapture at Mars is fine, once we've done it a time or two. Suspenders-and-belt thing. First time up, it might not work right. So, that should not be the manned trip. I'd prefer to stage the first exploration trip from orbit, anyway, so I could visit multiple sites on the one trip.
Still thinking about practical lander designs. I'm pretty sure retro thrust during the entry hypersonics will be needed. Not so sure a supersonic chute is worthwhile.
Testing nuke rockets down here requires a great deal of public education to undo all those decades of hysterical, nonsensical misinformation. I think a plume capture facility could be built (although hardly necessary), but it might be cheaper in the long run to test them on the moon. Good reason to go back there, actually.
I sure would like to see some real work done toward inflatable heat shields. Intriguing idea. Some sort of flexible elastomeric ablative is needed for the outer layer, I guess. Nifty idea for one-shot use.
GW
Here's an odd thought, applicable to mammals, anyway. Not so sure about birds and reptiles or fish.
Ship the livestock as frozen embryos in LN2 cold storage. Thaw and raise in a lab on Mars. Once the herds are started, shipment of live animals on Mars is no more time-consuming than truck shipment here.
They already do it with humans in the in-vitro fertilization thing.
GW
So am I (confident of high-grade steel in small batches by more-or-less conventional methods). Maybe not by 3-D printing. But then, who knows what they'll come up with fairly soon?
GW