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There's a common error I keep seeing regarding stability of water in near-vacuums like Mars. That 6.1 mbar triple point pressure is a water vapor partial pressure, not a total atmospheric pressure, and it refers to stable equilibrium conditions relative to a liquid pool at ) C. Look in any standard steam table, or ask user Midoshi, he will confirm what I am saying. That pressure in the steam tables is specifically labeled as "water vapor partial pressure".
The sum of all the partial pressures (water vapor and all the dry gases) is the total atmospheric pressure, which in turn is what you measure with a transducer,, gauge, or manometer. At higher liquid pool temperatures, the equilibrium water vapor partial pressure that is required is even higher. That's exactly the numbers in the steam table.
That being said, liquid water can indeed exist temporarily on Mars at 6-7 mbar total atmospheric pressure. It's just not in equilibrium, and it will eventually evaporate/boil away. The more there is, the longer it can last, and the longer a stream channel it can make. But it does finally "go away".
It's the same for the exposed ice (which we have already seen with the holes dug by the Phoenix lander): ice requires 6.1 mbar water vapor pressure in the atmosphere for 0 C ice to be in equilibrium. Colder ice temperatures require less water vapor partial pressure above it to be in equilibrium. Anything less than the equilibrium partial pressure of water vapor, and the ice sublimes. We've already seen it in action at Mars (Phoenix).
There is something about sufficient regolith cover that isolates buried ice from the lack of water vapor partial pressure in the (exceedingly-dry) Martian atmosphere above that regolith. We've already seen that, too, in the Phoenix holes. Apparently, it only takes a couple of inches (around 5 cm) to have that effect on Mars at 0.38 gee gravity. The regolith density ought to look just about like that of sand or dirt here: around 105 lb/ft3 (about 1683 kg/m3) effective mass density (there's considerable void volume due to the spaces between the grains).
BTW, a recorded humidity (relative) of 0.03% is pretty darn close to being equivalent to a 0.03% volume fraction (0.0003 without the % sign) of water vapor, which is 0.03% of a Martian atmosphere's worth of partial pressure. The calculated Martian partial pressure of water vapor at 6 mbar is then 0.0018 mbar. At 7 mbar atmospheric pressure, the water vapor partial pressure is about 0.0021 mbar.
Yep, that atmosphere is very, very, very, very dry.
GW
Hmmm. Well, what about launching up to a 53 ton vehicle to LEO with Falcon-Heavy, and then refueling it there with a Falcon-9 or maybe something else, depending on the mass to be transferred. Atlas-V-552 can put 25 tons to LEO, and I think Delta-IV puts about 18-20 in its most-strapons configuration. F-9 is either 10.1 or 13 tons, depending upon which version of its engines. That's a lot of refuel mass.
I'd work backward from the payload you want to return, and see how "bad the damage is", and what combination of mission options and rockets might be made to work. One mission / one launch is not the only viable model here.
GW
Bob:
Why not a scaled-down design similar to my chemical manned lander concept? Conical aeroshell, blunt heat shield, ascent rocket on the core axis sticking out a bit. A fleet of small rovers tucked away inside amongst the rocket-braking propellant.
Do it as a direct descent from transfer orbit, launched by a Falcon Heavy. My guess is it'll come out of hypersonics too low for a chute or ballute, so just go for direct rocket braking from M3 on down.
Rovers scoot around gathering samples, and bring them back to the ascent rocket. It'll need a contraption to load the samples into the return capsule. Direct return to Earth, and a free entry from transfer orbit. Or, stop in LEO and go get it with the X-37.
GW
Bob:
Gotta be lifted in pieces that will fit on top of things like Delta-IV, Atlas-V, Falcon-9, and (soon) Falcon-Heavy. This is on-orbit docked module assembly, maybe a bit more. It would really help if we had a supple MCP suit to wear for this work. And we could have one soon, because we had one working as a feasibility demo back in 1969.
GW
At the triple point of water, the condensed phase temperature is 0C (ice or liquid), and the vapor pressure (not total pressure), is 6.1 mbar. Look it up in any standard steam table, metric or US units. What is listed there is equilibrium vapor pressures. Says so right on the column headings.
Actual situations are usually not equilibrium, but then, they do not persist "very long" (a relative notion at best). I doubt the ice surfaces on the outer moons are in equilibrium. They are in fact slowly subliming. There's just lots of ice. 4 billion years ago, those same moons were very probably a lot larger.
GW
We seem to be getting spammers again, at least in this thread. Anyway, I looked at reconfiguring my chemical Mars designs to work at Mercury. Turns out not to be all that hard. Cargo payloads are reduced a bit in favor of extra descent propellant on an airless world, but I can delete the heavy heat shield. Posted today at "exrocketman". Have fun.
Bob - this is the same one-shot chemical design. MMH-NTO storables, very lightweight tanks in terms of inerts.
I will be looking at our moon next. But I have the expectation that configuration variations of exactly the same hardware set will work at destinations further than we can yet send men: all the way out to Titan.
I looked at the film and the images. I think the moving round things in the film are rock dust particles moving in the wind. There were no "microscopic" imagers of any kind on any of these probes, only "magnifying imagers". You are looking at things on the order of 0.01 to 0.1 mm in size, not bacteria sizes (nanometers down to angstroms).
That being said, I do indeed favor the view that the Allan Hills meteorite contained bacterial fossils. I do indeed favor the view that the unusual chemistry on Mars today has skewed the search for compounds associated with life from Viking onward. I think the NASA scientists who detected those fossils in the Allan Hills meteorite were very badly treated.
What I expect to come out as the truth, probably long after I am dead, is that early Mars abounded with microbial life, but never had the time (around 3 billion years) for complex forms to evolve, before it acidified, dried out, froze, and lost its atmosphere. It is quite possible, indeed quite likely, that deep underground, there is still microbial life on Mars, resembling quite strongly the deep underground microbes found here in recent years.
I rather doubt there's anything alive on or near the surface, due to the ionizing radiation, and due to the extreme dessication that a near-vacuum of an atmosphere enforces. But I think if you look with a microscope, you will find fossils of what once was alive on the surface.
If we terraform Mars, that underground life could come blossoming back to the surface. Actually, I think it will be found to be quite similar to life here, although the base pair sequences in the DNA or RNA will be quite different. This would be true even if the panspermia hypothesis is true, because the isolation has been 4.5 billion years. I doubt it could infect or hurt us, or that our bacteria could hurt it.
But I am an engineer, not a biologist. I just read things in the journals. Why believe me?
GW
Re: Fuzzy pictures / disappointment with NASA --
Read the book "The Death of Common Sense" to find out why government bureaucracies petrify into uselessness and complete inability to function, over time. Yes, this has happened in the US, and yes, it has happened to NASA (that's why they cannot do the amazing wonders rapidly, like they did around 1960-1970).
Of course the pictures could and should be better for the price tag. But it would take the NASA of 1960 to do that. It's a wonder that NASA-JPL can still build probes that can land on Mars at all. Sometimes I think that's the only part of NASA that still works at all.
I yield the soapbox,
GW
Re: Vincent post 88 on partial pressure --
What you said is true, but there's a tad more to the picture. A liquid water pool at 0 C in a closed system will evaporate slowly toward equilibrium, and humidify the space above to a partial pressure of water vapor of 6.1 mbar, regardless of any dry gas inside the container. The total pressure in the container will then be the sum of that water vapor partial pressure, and the original dry gas pressure, which is now its partial pressure. The same is true of a 0 C ice cube, it's just slower.
If on the other hand, you put some 0 C liquid water in a closed container and suddenly (magically) remove all the atmosphere above it, you have a very nonequilibrium situation. That water will boil violently and suddenly freeze in the act of boiling (latent heat comes from internal energy). The resulting ice will sublime, getting colder as it does (again, latent heat coming from internal energy), putting a few mbar worth of water vapor as atmosphere in the closed container. Eventually, with transfer of a tad of heat into the container to bring the ice back to 0 C, there is a 6.1 mb water vapor atmosphere in the container at final equilibrium.
That thought experiment does not correspond exactly to Mars, but it's not that far away, either. My point is that liquid water exposed on the surface of Mars in any significant volume will boil violently away, as it cannot significantly humidify the atmosphere above it. That's because the atmospheric currents carry the vapor cloud away as fast as the boiling can produce it. It does take time for this to occur, so you can have a temporary river or lake on Mars today, but it will boil away, sooner or later. Water droplets may not actually boil, but they will evaporate fairly quickly. And thin layers of ice sublime away in hours or less, we've already seen that.
For that not to occur at 7-8 mbar local total pressure, the local atmosphere must be at least 6.1 mbar water vapor, leaving only 1-2 mbar CO2 pressure. I do not see how that could occur, even in a closed container, since any gas can only hold a limited amount of water vapor without condensation. Plus there's an enormous mass of 7-8 mbar CO2 on Mars, not 1-2. You might see a local snow flurry above a liquid pool boiling away on Mars, but it would all eventually be carried away on the wind, even the ice (because it would sublime).
GW
@ Bob Clark:
Honestly Bob, I dunno how it would size out on LH2-LOX. I'd probably start at the same gross entry mass, and adjust the weight statements for less propellants and more cargo mass, due to the higher Isp. The volumes would be quite different, especially for the LH2, so there'd be a lot less available volume within the descent shape, and the ascent vehicle would be a lot longer at the same diameter. Whether that would seriously change the bi-conic descent shape is anyone's guess.
GW
Excluding the B-52 it dropped from, the inert mass fraction of the basic X-15 was right at 40%, somewhat like many other very fast aircraft. Even in the X-15-A-2 configuration with the big external tanks, it was still near this inert figure.
Kinda takes structural weight to confer the strength to be tough as an old boot at Mach 6, don't it?
Supersonic USAF bombers were about 48-53% inert, and supersonic USN bombers like the A-5 were about 58% inert. Even the subsonic carrier-based A-7 was 50% inert. 17 US tons dry, with 17 more US tons of fuel and ordnance. Tough old birds.
GW
I second the motion, Bob. God speed, Neil. And all the rest we have lost.
GW
OK, guys, I have revisited the chemical Mars lander problem under different assumptions regarding ballistic coefficient scaling versus lander mass, and further, I took it to rough dimensions and volumes, not just rough weight statements. I got 3 men and 5 tons of supplies and equipment to the surface in a lander massing 60 metric tons at entry, along with an ascent vehicle (as dead-head payload) whose crew cabin is also an abort capsule, either in descent or ascent. The same 60 ton (at entry) lander without men or an ascent vehicle delivers over 28 metric tons of cargo to the surface, even very low-density cargo. There's over 170 cu.m usable cargo volume available.
These are one-shot vehicles, not reusable. Storable propellants MMH and NTO. I included a cosine factor for 10 degree engine cant, to get flowfield stability firing retro thrust into the oncoming supersonic stream. No retro thrust during hypersonic entry from LMO, but no aero-decelerator, either; there's just not enough time to deploy one, much less have it do any good. End of hypersonics, just drop the heat shield, and fire up the rockets. Direct rocket braking to touchdown. One neat little nuance: I used the engines of the ascent vehicle as my descent engines, just sucking from descent propellant tanks.
I posted this stuff as a technical article over at "exrocketman", dated today (8-28-12), and titled "Manned Chemical Lander Revisit". Search/screen keywords are "Mars" and "space program". Have fun. There's a bunch of stuff I've posted over there. Feel free to use any of it.
GW
http://exrocketman.blogspot.com
PS -- hey Rune! This latest one is the one you really want to examine critically. Turns out landing big things on Mars is not so very hard after all. I think you'll really like what I did.
quote:
Note that the pressure never fell below the 6.1 mbars pressure required for pure water to remain liquid, i.e., not boil off.
Also, eyeballing the temperature graph it looks like surface temperatures remain above 0C for perhaps 3 hours per day.
Bob Clark
Don't get too excited about 7+ mbar pressure observations vs the possibility of liquid water. What keeps ice from sublimating, and liquid water from boiling away, at 0 C is not 6.1 mbar total pressure, but 6.1 mbar water vapor partial pressure.
The atmospheric pressure might reach 7.8 mbar total, but just how much of that is water vapor partial pressure? If it's under 6.1 mbar (and it is, because the dry surface pressure of the CO2 is pretty close to 6-7 mbar all over the plains of Mars, leaving at most 1.8 mbar water vapor partial pressure), then there is no equilibrium liquid water on the Curiosity site, except as the briefest of transients (by definition not equilibrium).
A question: why would/should the atmosphere pressure be cycling up and down (post #76 above), with almost exactly the same trace shape, at exactly the same frequency? Which frequency is 1 cycle per sol? Is that not an instrumentation artifact of some kind? Perhaps a reading sensitivity to instrument soak-out temperature?
This is nearly-raw data from an experimental instrument platform. I'd be very careful about saying it was an accurate reading of true conditions on Mars.
GW
You're going to have a lot of carbonate slag clogging up any imaginable heat engine, if you use "molten carbonate fuel". Or a silane burning with CO2. Same thing may well be true of a fuel cell using that kind of fuel. I dunno for sure, but I had some experience with a silicone / magnesium ramjet fuel decades ago, burning with air. Fortunately, the ramjet had no moving parts. If it had had some, they would have been slagged solidly immovable. Good for internal heat protection, bad for part functionality.
GW
A stationary compressor of some sort might be feasible, if it can spend a lot of time charging up the bottles. Most propulsion applications require high mass flow throughput rates to create thrust or power, certainly any kind of imaginable heat engine. Cycle pressures in such things are a few std atm. Compressing from 7 mbar to a few thousand mbar is a compression ratio in the thousands, not something that can be made mobile, or that can be made to provide high mass flow rate. We have things like that in submarines, but they are huge, massive, and produce a slow trickle, in proportion.
I would suggest compression by other means than pistons or turbine blading. 7 mbar is just too thin for that. CR 30 applied to near vacuum is still near vacuum, far at variance with what any imaginable heat engine would need.
There are things that burn with CO2, yes. Magnesium is one. I guess H2 might be another, although I am unfamiliar with doing that. Why not store really energetic reactants, like LOX and liquid methane, and using compressed atmospheric CO2 to dilute the reaction down to temperatures your engine could stand. Both ordinary piston and ordinary turbine engines become possible on Mars, done that way.
If you are doing ISRU creation of LOX and methane from local water and CO2, might as well compress some extra CO2 for the dilution gas.
There are some absorption-based methods of compression for CO2. Not enough mass flow for mobile use, but it's entirely feasible as the first stage in a practical system: something that gets you from 7 mbar to the 300-500 mbar that our ordinary machinery can handle. You have to put these tinkertoys together like that, in order to make this thing work.
GW
I am unconvinced that the financial benefits of launch rocket reusability will ever be dramatically large. The real cost-reducer has been the transition to smaller logistical tails. But I could be wrong.
That being said, I don't see the desirability of going for reusability and SSTO in the same design, since both of these depend critically upon inert mass fractions, but with opposite signs on their trends. I rather suspect that practical SSTO designs will always be throwaway items, driven there by 4-5% inert fractions to get 2-3% payload fractions. KeroLox, LH2-LOX, no difference. It's chemical, all you can get is 310-470 sec Isp.
A reusable launch rocket will be 2 or maybe even 3 stages, so that inert fractions in the 10-20% range can be tolerated and still get 2-3% payload fractions. I'm not at all sure that the second and thirds stages should be reusable, but I think a reusable first stage is feasible, if it's not too large.
We have enough trouble landing million pound airplanes without driving the landing gears up through the airframe. You can't whack the air broadside hypersonically, or the sea at around 50 mph, without damage to lightweight structures. So they cannot be lightweight. Metals and composites are only so strong.
Mass scales as dimension cubed, while part cross-section area scales as dimension squared. Better to stay smaller with reusable items.
I agree with Antius that a nuclear thermal rocket provides the Isp for a practical surface-launch SSTO, probably as a winged spaceplane of medium size. I'm not sure the public could be sold on reactor safety in the event of an abort or crash, not with solid core. A gas core lightbulb engine is a better sell for that application. The safest version for crash or abort is open-cycle gas core, but the somewhat radioactive plume will be a hard sell. Those things being true, I'm not so very sure the best application for nukes is surface launch. It would be a lot easier to sell them as orbit-to-orbit propulsion to and from Earth. Not trivial, but an easier sell.
GW
I would suggest resurrecting the LH2 NERVA of 1973 first, which should take a small, talented team, no more than about 5 years. This is Kelly Johnson-style Skunk Works stuff. As the design finally was, the plume was pretty clean. Didn't/doesn't need plume capture, just shoot it in the open. Get that thing operational, and use it ASAP for orbit-to-orbit transport purposes, not launch vehicle upper stage work. That takes care of the reactor-falling-from-orbit fear, pretty much.
Working in parallel, we need variants on the NERVA idea that upgrade its T/W and Isp, things like DUMBO and TIMBERWIND. These may need plume capture, at least at first. These would still be LH2 engines, and would be upgrades to the NERVA's being used, about 5-10 years down the line. Again, you have to have a small talented team doing this, not a government bureaucracy more interested in process than results. Skunk Works-type organization.
Once these things are operational, we need the water-propellant variant, which is a third item to be worked in parallel, starting right now. A water NTR could be refueled anywhere, because ice seems to be everywhere in the solar system. Plus, water is so easy to handle and store, even in space. THAT is where we need to go with solid-core NTR. It's a practicality and ISRU thing. Lowered Isp is quite acceptable if you can mine and melt ice and thus be free of refueling restrictions.
Within 5 years of cranking this solid-core NTR effort up, the gas core NTR needs the practical work it never got. I'd suggest doing that one on the moon, as we ought to be easily capable of planting a test base there with the rockets we have, especially if augmented by solid core NTR's. Open-cycle gas-core NTR always has a somewhat radioactive plume, and there's a risk of it even in the light-bulb concepts, should containment fail. But we need gas core engines. That's how you fly faster than the low-energy transfer orbits we use today. High Isp and high T/W at the same time is the payoff, something electric propulsion does not appear to be capable of offering.
And then there's the nuclear explosion drive. That works better in ships the size of naval battleships of WW2-vintage mass (20,000 tons and up). I'd reserve that for colony-planting missions down the road a bit. But we're going to need it. So we're going to need to test it out and get it working. No better, safer place to do that, than the moon.
GW
I found some ballute stuff and posted it over on "exrocketman", too.
GW
RobS:
Aerobraking will always be tricky because the "window" requirements on trajectory angle will always be very tight.
What I have been doing with a 2-D Cartesian model for simple entry will never work for estimating aerobraking, because that requires a circular or spherical coordinate derivation. Mine is constant path angle 2-D.
I remember from the references where I got the simple model that aerobraking on Mars is descent to altitudes nearer 70 km, while aerocapture requires descent to altitudes nearer 25 km. While it's all thin-as-all-get-out densities, aerocapture densities are far higher. That's the Justus & Braun report on model atmospheres for Earth, Venus, Mars, Titan, etc, that Rune turned me onto.
With what I have, I just cannot quantify how much velocity you could dissipate, vs altitude & approach velocity.
GW
PS for the rest: I just posted some blunt capsule drag data over at "exrocketman" for Mercury and Apollo. Anybody have a good, traceable weight statement for Dragon?
RobS:
No, the calculation is too crude and primitive for that. Getting some L/D just makes things a tad better. My real point was to show that landing big things is feasible, even if you cannot take advantage of all the best tricks. I just go straight from plain vanilla ballistic hypersonics to rocket braking. There's no time to deploy a chute, much less have it do any good, coming out of entry that low.
The main angle-related troubles I can forsee are hitting the entry interface at too steep an angle. That's got to be very shallow (about a degree to a degree-and-a-half) or you'll whack the ground before entry is over. L/D 0.2 won't really change that very much. That sort of shallow entry is "guaranteed" from LMO with a min deorbit burn (you can screw it up with too aggressive a deorbit burn). Doing it from direct entry, it is very tricky to hit that corridor, just like it was coming back from the moon.
The models I am using are 2-D planar elliptic orbit to the interface altitude, followed by 2-D Cartesian approximation of entry down to Mach 3. That last is nothing but a spreadsheet "automation" of the old 1956 way of estimating velocity vs altitude from an exponential approximation to the density profile-with-altitude. I didn't like the closed-form heating estimates, so I re-did that part as an actual finite-difference "integration" in the spreadsheet. Doing a finite difference "derivative" in the spreadsheet matched the old closed-form peak gees estimate pretty well. Getting the dynamics about right but the heating wrong was pretty much the experience with this method back in 1956, with warheads and film capsules. I fixed the heating, so it works fairly well now. Convective heating only, though.
GW
OK, guys, I did a pencil-and-paper analysis of post-hypersonic trajectory dynamics, and rough-sized a family of chemical landers for it. The manned lander has an ascent vehicle for return to LMO with crew of 3. There's no plane changes allowed. Descent and ascent use the same engines. The unmanned lander is a one-way device, where more cargo mass replaces the ascent vehicle, in an otherwise overall-the-same vehicle.
I'm showing end of entry at local Mach 3 at 8 km altitude, no retro thrust during entry. The 2-ton heat shield gets blown off explosively in pieces. Retro thrust kills horizontal speed by about 3.4 km altitude, then kills vertical sink, then does a tail-sitter "shtick" from about 200 m. No chutes or ballutes, there isn't time to deploy them or for them to be effective. I used 10-degree cant for retro plume stability.
There's extendable landing legs, and the aeroshell segments get used as unload ramps. Depending on the mass of the crew cabin capsule, the manned version carries 2 to 6 tons cargo to the surface, but nothing significant back up. The unmanned version carries about 25 tons one-way to the surface. Both are 60 tons at entry. About 7 m diameter, short and squat, tall slender single-stage ascent vehicle, hypersonic CD 1.30, beta = 1200 kg/sq.m. MMH-NTO engines. Built in Earth orbit from pieces, including a segmented heat shield.
I posted the whole thing over at "exrocketman" minutes ago. Let me know what you think.
If I can come up with this, based on no more sophisticated tools and analyses than I have used, then NASA bloody well had better have already come up with something similar. They have the people and the analysis tools to put this on a firmer footing than I can do.
Landing big stuff on Mars turns out not to be all that hard, as long as you are unafraid to look at things besides those things you have already done before.
EDLA: Manned vehicles will be larger, with larger ballistic coefficients. From LMO, if you keep angle shallow, entry is around 0.7 peak gee, terminating (Mach 3) around 8 km altitude. Direct entry at higher speeds, or at steeper angles, quickly takes you to hypersonic surface impact. At 8 km M3, I see little opportunity for a chute/ballute to do any good. Looks like rocket braking to touchdown to me. If you do direct entry or steeper entry, you may need retro thrust during the hypersonics to make the ballistic coefficient effectively lower.
Radiation: doesn't appear to be that big a problem on Mars's surface. In space, solar flares are the more immediate threat. That's 20 cm water or equivalent.
Microgravity diseases: In space, spin the slender-baton-shaped ship head-over-heels for artificial gravity. Put the hab on one end. About 3-4 rpm is tolerable. 56 m at 4 rpm is one full gee. If the ship is assembled of docked modules, it is easy to reconfigure after every burn to shed modules but still be a slender baton of the correct length.
On Mars: we simply dunno how much gee is enough. Not yet. May need a conical centrifuge exercise room in any permanent settlement. Or maybe not.
We'll need the bulkier, heavier frozen food to make the trip. That freeze-dried and sealed/irradiated stuff doesn't last but about a year, maybe year-and-a-half.
The next lander ought to be a modified Dragon sample return. After that, men need to go. If no one can afford the sample return, then men need to go. But, I'd bet Musk can pull off a sample return affordably based on Dragon/Falcon-Heavy, in a direct entry scenario. The Dragon may need a lot of extra thruster fuel, or maybe not so much, I dunno. The side-mounted Super Dracos could do retro thrust in entry, if needed.
GW
OK, guys, I have just finished posting all of my results from the big Mars lander entry sensitivity study. It is the latest in a series of articles about, or related to, landing on Mars. All of these are recent postings over at http://exrocketman.blogspot.com. List follows:
Date title subject
8-10-12 Big Mars Lander Entry Sensitivity Study impacts of beta and de-orbit burn size
8-5-12 Ballistic Entry from Low Mars Orbit impacts: beta from 100-2000 kg/sq.m
7-25-12 Rough Correlation of Entry Ballistic Coefficient scaling from <1 ton to 100-ton sizes
vs. Size for “Typical” Mars Landers
7-14-12 Gravity Data on All the Interesting Worlds traceable data for surface gee, etc
7-14-12 “Back of the Envelope” Entry Model 2-D Cartesian estimator, updated
6-30-12 Atmosphere Models for Earth, Mars, and Titan traceable “typical” data
6-3-12 Deceleration by Drag Devices (and more) concepts for retro thrust during entry
on Mars
I still need to explore what happens after entry ends, because a lot of these altitudes are single-digit km, not the 20-30 km stuff NASA-JPL has been seeing in its designs. This is seconds-to-impact stuff. No time for a chute, just massive rocket braking to touchdown. I don’t have a model built, but at least it is just straight supersonic-to-subsonic aerodynamics and vehicle dynamics. I know how to do that.
So far, none of this requires rocket braking during entry. But that topic needs exploration, too. And I still have to figure out exactly how I am going to do that.
GW
Terraformer:
I did look at the bigger retro burn, just like I said I would. It lowered velocity at entry, as expected, but it also steepened entry, as I feared. The dynamics were more sensitive to the steeper entry angle than the entry velocity. 200 km circular 3.455 km/s minus 50 m/s put 135 km entry at 3.469 km/s at 1.63 degrees. At 1200 kg/sq.m, you come out of hypersonics at M3 at about 8 km altitude. Plus and minus 500 on the ballistic coefficient lowered and raised terminal altitude by about 2 km.
If from the same orbit you de-orbit by 500 m/s, then entry is at 3.029 km/s, but at 6.4 degrees. At the same 1200 ballistic coefficient, you are still doing 4.7 Mach when you impact the surface. Entry isn't over yet.
Both entries were figured with a "typical" Mars atmosphere, and that 1956-vintage 2-D Cartesian "back-of-the-envelope" entry model. The orbital mechanics part is pretty standard ellipse stuff, I just did it pencil-and-paper.
GW