Landing on Mars

Russel · 2013-02-04 04:29:29

When randomly browsing I discovered a new term I'd not heard before.. a "Molly bolt" design.

http://www.isset.org/nasa/tss/aerospace … d_mars.htm

It's the fourth image down the page.

Seems its hard to come up with a novel idea Although what I have in mind the "petals" don't join up to form an unbroken surface.

Russel · 2013-02-04 04:45:05

GW,

You're correct. The "air brake" panels don't have to be round, but I suspect they work out better structurally if they've got a curve in 2 dimensions. It did cross my mind to make it into a ramp but I haven't quite figured out the geometry of the hinge/linkage that would do that.

Would they be useful earlier on in descent? My code simply allows me to introduce the odd "hack" so I can increment the drag etc depending on certain conditions.

I think when it boils down to it, slowing down the descent higher up ends up being counter productive because, in simple terms, you just end up with more time for gravity to do its thing.

That of course is simplistic. It also depends on lift. But getting lift higher up is also harder.

One thing I like about having "air brake" panels is that if you control them individually you can also force lift. Problem is that given the overall design, I suspect you have to fly directly into the stream or else turbulence will mess up the undisturbed layer of steam under your craft. And that problem too goes away as peak heating passes.

So for the moment it looks like this thing just flies straight right on down 30Km and through the period of peak heating - no lift.

Below that you gradually spread the panels asymmetrically generating gradually more lift.. then onto a phase below 20Km when they're fully spread and you're at maximum g.

They would also serve some purpose remaining spread down to close to Mach 1.

GW Johnson · 2013-02-11 12:23:42

Hi Russel:

I think steam cooling was one of the 3 things they were going to experiment with on the old X-20 Dyna Soar that got cancelled long ago. It should work. Might be heavier than ablatives, but who knows?

The cavity heat shield idea should work, except you will have to cool the ever-loving thuinder out of the rim leading edges. Otherwise, they burn away very rapidly.

I'm like you, I don't see much benefit to altering drag or generating lift/side force at entry speeds, ahead of the deceleration pulse. (Other than the tiny adjustments to correct trajectory.) The benefit is probably coming out of the deceleration pulse, where gravity is starting to "take over".

Do you or anyone else have a way to estimate stagnation point material surface temperatures?

GW

Russel · 2013-02-12 04:31:24

Since this web site took a short break I posted some of this in nasa spaceflight dot com forum. I'll include a link if the moderators don't object. So I got to work on the idea a bit more.

When I went over the data on MSL here http://www.ssdl.gatech.edu/papers/confe … 3-0908.pdf and here http://ntrs.nasa.gov/archive/nasa/casi. … 006430.pdf I came to the conclusion that the total heat load on the MSL heat shield was closer to half a GigaJoule than a full one.

Given that MSL has a much higher entry velocity but much less mass, it turns out that the kinetic energy involved is roughly the same as a 10 tonne lander entering at 3.5Km/s.

Turning water into steam and then raising that steam to about 800C involves about 3GJ/tonne . So in theory at least you're talking about 200Kg of water. Perhaps. Ablative shields also tend to be their own radiative shield thanks to all that carbon being liberated. Anyhow I'm working on a tonne of water, but it could be a fair bit less.

As to keeping the leading edge cool. If its got a reasonable radius of curvature you can probably establish a laminar flow of steam that follows the surface from the inside around the bottom. Between that and the water inside you should have some chance with regular high temp alloys.

I don't know how to establish actual temperatures. All I can do is read between the lines of actual studies, like that one on the MSL heat shield.

What specific situation are you thinking?

With the air brake I figure it would be useful below 1700m/s. The usual rule of thumb puts the temperature of the gas at about 1700K or 1400C. How hot that makes the metal I don't know. Presumably lower than that figure.

Russel · 2013-02-12 04:36:01

Oh btw, I did read up on the Dyna-Soar but couldn't find much on the cooling methods.

Last edited by Russel (2013-02-12 04:46:41)

Russel · 2013-02-12 04:54:27

You might find this an interesting read.. its a Masters thesis with a whole lot of Mars EDL simulation going on.

http://dspace.mit.edu/bitstream/handle/ … 5.pdf?...1

Russel · 2013-02-12 09:35:09

Here's another article I've come across. Another proposal for a reusable Mars "ferry" (ascent/descent vehicle). Lots of interesting data too.
And also some interesting figures in the section dealing with supersonic retro propulsion.

http://www.nss.org/settlement/mars/AccessToMars.pdf

Rxke · 2013-02-12 12:45:01

interesting stuff, keep 'em coming! 8)

Russel · 2013-03-03 07:27:12

Any matetials experts around?

I got a bit further into the detail of my ascent/descent vehicle.

There are essentially four major structures. One is a toroidal tan with a major diameter of 5.7m and a minor diameter of 1.8m That forms the base of the vehicle. After a little research Ive figured it can be built with a 2-3mm shell of titanium alloy. Its built in 16 segments with a cross plate between each segment forming the structural connection to the rest of the vehicleat each cross plate. As well Ive allowed for internal longitudinal stiffening ribs to take the largely vertical forces. As a rough estimate this tank comes in at around 2 tonnes. Could be a bit less with carefull design.

The oxygen tank is 3.4m in diameter.

Remember that this tank will hold at most 2 bar of pressure.

Next major element is the spherical LOX tank positioned at the center of the vehicle with it center point raised about 2.1m above the center of the toroidal tank.

Again Im not aiming for more than a few bar of pressure and using AlLi alloy this tank comes in at roughly 600Kg. I was originally figuringbon using thetank itself as a structural element but for the momentvthats too complex for me to estimate - butvIm sure its possible.

Instead Im building a truss structure aroundvthe LOX tank. There are two 16 segmented rings arrangrd in the horizontal plane each 45 degrees off the equator.In other words thes rings are at 1.2m above and 1.2m below the equator ofvthe tank.

From these ring the oxygen tank is supported via insulating standoffs.

A truss structur connects first these two rings and then the primary structural points for the main toroidal tank which are at the top dead center of each of the 16 cross plates that the main tank is built around. Iwont go into even further details with the truss save to say it provides 16 main strutural loading points that are set just off the equator of the oxygen tank and those points are 1.2m above the top of yhe toroidal tank.

Each of these 16 main loading points has 6 tubes meeting.

From these main loading points extend 8 electric actuator screws (4 pairs) each pairvpushing an air brake panel outwards.

From these main loading points extend 8 vertical landing gear legs. Each with electric screw jack, latch and shock absorber.

And of course the engine mounts refer to these main loading points.

Now the truss structure, landing gear actuators for the air brakes I dont have more than a very rough idea of mass. Using titanium alloy and some sanity checks on thickness of the tubes Im thinking of an extra 800-1300Kg.

Now for the fun bit. If I use a nickel based alloy forvthe air brakes Im confinedvto a working temperaturevof 1000C and hopefully more like 900C. As a simple sanity test a 1mm sheet of nickelcalloy adds about 170Kg per air brake panel Id like to keep their mass to 400Kg and anything less is more payload upstairs. I think its possible. The original intention here was to build an airbrake with some degree of porosity. That helps.

Now Im sure you can cut mass by just building a very porous structure with say hastelloy and then adding thin ceramic cloth. Not only does that add a bit of insulation but it should be half the mass of the alloy per unit area. I think you might be ablr to get away with cm sized holes in the metal.

A couple of other things come to my rescue here. First if we start at 1700m/s and peak at 4 gees. then were down to 1300m/s in 10 seconds and the raw gas temp is now within the limits of the material. In practice it will take about 10 seconds to fully crank out the panel so you have to read between the lines here. Next the nickel alloy itself has its own heat capacity and it turns out yhat each panel can soak up another 60MJ before getting near 1000C and in practice I think it wont get this far. Also by the time the air brake panels get to 850C theyre already radiating over 50Kw per sqm off the upper surface.

Anyhow that leaves one major structural element and thats the crew cabin. Now Im figuring on a 2.2m diaoblate sphere and even with a safety margin that comes in at 800Kg.

Put the four main structural components together... Yhe main tank.. The LOX tank.. The framing.. The air brakesx4 and we get to about 6.3 tonnes.

Now this doesnt include the thevdedcent fuel tanks and engines. Lets allow a tonne.

Then there is the hydrazine tank, the life suppot tankage snd lots of other ancilliaries.

Can it come in at 10 tonnes with some RCS fuel but minuscthe crew and propellants. Still possible

Now I need time to learn CAD.. May be a while

GW Johnson · 2013-03-03 10:16:18

Russel:

Watch out specifying titanium for anything but physically very small parts or else big heavy castings or forgings. There are no formable titanium alloys in most editions of Mil Handbook 5, and for a reason. All the 6-4V alloys are castable/machinable only. The two beta-phase titaniums (RM-21 is one of the names) are formable, but have very serious aging problems at room temperature. Grain growth, embrittlement, loss of tensile strength. That was the trouble with the skins on the SR-71, and with a stretch-formed ramjet motor case I once worked on.

In my researches I found a couple of things that might help your calculations. One is a rough rule of thumb for entry hypersonics: shock layer gas temperature in deg K numerically equals velocity in meters/sec. Another is that up to about 10 or 11 km/sec entry velocities, convective heating dominates, which means the heating estimates in my BOE model are actually pretty good just as they are. The sources that said these things (some quite old, like 1953) said they were for Earth, but should work at Mars in spite of the different gas composition. At least crudely accurate.

I haven't doped out the old pencil-and-paper ablation model, but I have doped out the old re-radiative model, which applies to any sort of re-radiating hot skin surface, metal or ceramic. It's just the emissivity-modified Stefan-Boltzmann law (abs temp to 4th power). Re-radiated stagnation heat flux must balance convective stagnation heat flux, and the skin temperature must rise to that balance level, no matter what the shock layer gas temperature is. The trick is guessing an appropriate spectrally-averaged emissivity: no usable material is really "gray". The closest to "gray" behavior is carbon black. Your skin temperature limitation is acceptable temperatures for whatever material you are using. All are different.

I put the LEO folding-wing spaceplane study up on "exrocketman" last night. I re-shaped its belly radius slightly so that I could use alumino-silicate low-density ceramic as the heat shield, even at stagnation conditions. This stuff should work good for Mars entry, and may not even need to be black-surfaced, I haven't looked into that yet. There really is a way to build a super-lightweight heat shield that does not ablate, need not be installed as myriads of separate tiles, and is a lot tougher against impact than anything on the old Shuttle. (That material is still quite experimental, I haven't made it since 1984.) Shuttle tile would work, of course.

I'm not sure Al or Al-Li is what you want for LOX tanks that serve more than once. Down here, we don't do that. LOX tanks with long service lives are almost invariably made of 300-series stainless steel. I'd have to look to see which one gets used and why, but the problem you have to face is brittle behavior soaked out cold, repeatedly. There's one or two of the 300 series SS's that stay non-brittle soaked cold at LOX temperatures, repeatedly.

I dunno, but I suspect the preferred alloys for the hydrazine and NTO tanks are also stainless. I seem to remember that's what they were in the old capsules, and I think on the old Shuttle, too.

GW

GW Johnson · 2013-03-03 13:24:06

An update:

I checked my hand-made low-density ceramic material from decades ago for entry on Mars. If black-surfaced (like windward-side Shuttle tile), the stagnation-point skin temperature should be about 1110 F for entry from 200 km LMO, and about 1780 F for direct entry transfer from Earth at shallow entry corridor. Black-surfaced Shuttle tile should be exactly the same.

If left white (like lee-side Shuttle tile), the LMO entry stagnation skin temp is about 1760 F, while the direct entry is unacceptable at 2700 F (this stuff has a 2300 F max limit to forestall structural degradation from the solid phase change; so also does Shuttle tile).

The vehicle has a ballistic coefficient of 400 kg/sq.m and a nose radius of 12.4 m. LMO entry conditions from 200 km are 3.469 km/s at entry interface, depression angle 1.63 degrees. For direct entry I used a "typical" 5.6 km/s at entry interface, and the same shallow 1.63 degrees.

For LMO entry I got peak stagnation q/A = 2.6 W/sq.cm, peak decel 0.72 gees, and a M3 end-of-hypersonics at 17.2 km altitude. For the direct entry I got peak stagnation q/A 10.82 W/sq.cm, peak decel 1.88 gees, and M3 end-of-hypersonics near 15 km altitude.

Don't need ablatives, we can do this job with low-density ceramics (or superalloy metals but that's a lot heavier), either mine or Shuttle tile. If direct entry is involved, use the black high-emissivity surface (gross spectrally-averaged emissivity 0.80). Otherwise, you can use the plain white natural surface (gross spectrally-averaged emissivity 0.20).

In point of fact, plain carbon steel will stand up to this, if you don't load it at all structurally. It's just heavy. The superalloys were made for service in this temperature range. The low density ceramics are really super light, though. Mine is likely 30% heavier than NASA's more fragile version, but certainly no more than factor 2 heavier. Twice nothing is still nothing.

There's no reason at all why we cannot build fully-reusable, very-long-service-life "Mars ferries" or "Mars landing boats". This can be done single stage one-way chemical and still make round trips if refueled on the surface or from an orbiting fuel dump. It can be done single stage round trip without refueling with solid core nuclear. Take your pick. It's just rocket equation stuff and realistic weight allowances from here.

GW

Russel · 2013-03-03 19:53:09

GW,

I think the largest individual part is a tube 1.6m long. And Im aware that if built from titanium a lot of these parts would require a lot of machining. There are benefits to that approach though including being able to be fussier about shapes. For instance the panels that form the torus can have their inner stiffening machined in place.

Nevertheless I have to wonder what ogher materials would suit.

For instance the toroidal tank initially I considered allowing portions to heat to 800C. But looking again I realised that wherever water (or cryogenic fuel) contacts the inside surface the outside surface would be far below that temp. I moved towards titanium with a target maximum temp of 400C. This would involve an in ternal spray system that keeps the relevant inner surfaces

Now if I were totally confident about the heat transfer or had a reliable coating then that might lead us to something easier to build with but what I dont know.

As for the rule of thumb about gas temp thats why Im working on 1700m/s which should translate to 1700K or indeed 1400C. Given the heat capacity of the metal and radiative cooling Im pretty happy working with a peak material temp of 1000C but it could easily work out lower..

Speaking of materials there are now commercial carbon fibre overwrapped aluminium tanks so theres still opportunites for mass reduction here and there. Certainly the cre compartment could become compsite.

Russel · 2013-03-03 19:59:50

If ceramics or ceramic coatings advance to the point of being trustworthy enough that we can land and ascend on Mars a dozen times I might rethink things.

Likewise there is a traeoff between deployable air brakes and simply increasing the base area to compensate. Last time I tried a 7.5m basr with airbrakes could be equalled by a vehicle with an 11m base.

Russel · 2013-03-03 20:04:31

As for the light weight plane type lander the question is how do you get it back into orbit?

One idea that kept popping up was a reusable core but the wings (or most of yhem) are expendable. How do you feel about this?

Russel · 2013-03-03 20:05:57

Have you considered wings that are flexible enough to shef load?

GW Johnson · 2013-03-04 10:47:59

Hi Russel:

The airplane was for entry from LEO at Earth. It is boosted to LEO by a commercial launch rocket. I fold the wings and tails to enter dead broadside like a space capsule, with extreme bluntness, which lowers peak convective heating. The spaceplane design has a wings-folded ballistic coefficient of 196 kg/sq.m, and a "nose" radius of 14.2 meter. The unfolded wingspan is close to 20 m. I got the peak entry heating down to 25 W/sq.cm that way. The re-radiating skin at 0.80 emissivity was 2300 F, right on the hairy edge for ceramic tiles or my composite ceramic.

My composite was used (differently, as a low-conduction insulation) in a combustor that accumulated 5 or so hours of burn time in dozens and dozens of separate burn tests. It looked the same the day I pulled it out as when it went in. It survived all kinds of violent rich blowout instabilities in those tests. I made it as a monolithic sleeve, not individual tiles. I made the combustor's nozzle of the same stuff. Monolithic insert, pinned in place.

The Mars lander I was looking at had a big blunt rounded heat shield and a conical afterbody, much like a Mercury capsule in overall outlines. I forget now the dimensions, but this was a 4 man job, and still carried a few tons of cargo. Its ballistic coefficient was 400 kg/sq.m, higher even than Apollo.

GW

Russel · 2013-03-14 14:41:50

An interesting article

http://www.islandone.org/Launch/boron-sharp-article.htm

Russel · 2013-03-14 15:38:14

As you're probably aware, I've been working on a lander/ascent vehicle and I set myself the challenge of using CO as a fuel. Bottom line for now is that I think it can work albeit you have to be carefull with the structural mass and there doesn't appear to be much prospect for additional cargo/samples - unless you want to bump the fueled mass closer to 100 tonnes.

Since part of the motivation is to reduce the effort/energy in producing fuel there probably is a limit.

In deference to the interest in using methane as a fuel I considered what changes this would make to the overall design. And the answer is not much in general form but somewhat lighter - perhaps a tonne.

Another realization is that methane makes a pretty good coolant. Not so much in boiling but for the fact that the gas has a rising heat capacity as it gets hotter.

Ok firstly you need to take on liquid hydrogen from orbit. But this can be kept under 2 tonnes so overall the vehicle is only slightly heavier to land. As you know Im reluctant to land people with any uneccessary mass.

It turns out that the outer toroidal tank can carry the hydrogen down. This time the tank needs to be half the volume it had to be for CO.

Oxygen for descent is produced on the surface.

Landing fuel is hydrogen. So that means a seperate set of LH2/LOX engines for descent. But even allowing for this it equates to less hydrogen use overall. Reason is that with a purely methane powered craft hydrogen would have to be used to make mthane that is then consumed to transport the landing fuel.

Doing this means being able to put cant into the landing engines separate from the ascent engines and with hydrogen fuel you can afford a bit more margin for landing.

On landing the hydrogen is unloadef since the outer tank is only modestly insulated.

Instead of seperate tanks for landing fuel I now have a seperate smaller tank (about 1m3 ) or liquid methane coolant.

Instead of exposing the outer tank to reentry its now insulated and the surface exposed to reentry is wrapped in a high temperature material - probably metal. And fairly thin at that. Enough to protect the insulation.

Instead of relying on steam I pump liquid methane from the coolant tank into the protected space below.

So the boundary layer under the tank is kept below 950C. So no huge constraints on material.

On ascent the outer toroidal tank is about two thirds full of liquid methane. In orbit the remainder is used as a top up for the coolant tank.

One noteable shift in design is that the outer tank now only experiences low temperatures so it can be made from aluminium alloy.

One other feature is that this design is tolerant to methane containing some fraction of CO. Exactly how much I haven't gone into detail about.

SpaceNut · 2013-03-14 19:55:24

Russel wrote:

Oh btw, I did read up on the Dyna-Soar but couldn't find much on the cooling methods.

Transpiring heatshield....

Russel · 2013-03-15 19:20:41

I keep getting nowhere searching for actual engineering detail on transpiring heat shields.

Apparently they require a lot of small holes to keep the layer even and they run into issues of clogging. Even found a patent whose core idea was to run cleaning fluid through the system during launch.

I don't think I need anything as elaborate.

Imagine the bottom of the vehicle has a generally concave shape (like the inverted bottom of some pressurised tanks). In the space within that concave depression you find the engine nozzles.

Now originally I was thinking about directing steam with laminar flow nozzles.

Having thought more about this I realise its as simple as spraying liquid into the protected space under the vehicle. The heat of rentry - entirely radiative at first boils the liquid. As it does it creates pressure. The pressure matches the dynamic pressure of reentry keeping convective heat transfer well away from much of the underside. The coolant gas continues to expand and mixes to some extent and will naturally flow outward. Particularly in the case of methane as its temperature rises so does its heat capacity. So as it passes 500C under the skirt it can absorb a lot more energy.

Turns out the hottest zone should be the rising portion of the outer skirt.

So you control temperature with a simple valve and something not much fancier than a garden sprinkler.

Hence it should be much more reliable.

SpaceNut · 2013-03-16 19:59:29

http://en.wikipedia.org/wiki/Waverider X51-A's Scramjet

The temperature problem can be solved with some combination of a transpiring surface, exotic materials, and possibly heat-pipes. In a transpiring surface, small amounts of a coolant such as water are pumped through small holes in the aircraft's skin (see transpiration and perspiration). This design works for Mach-25 spacecraft re-entry shields, and therefore should work for any aircraft that can carry the weight of the coolant.

http://www.freshpatents.com/Materials-f … 252045.php

http://www.patentstorm.us/patents/53301 … ption.html

Void · 2013-03-17 22:28:42

I don't normally consider that I should post to threads like this for lack of ability. However I was trying to grasp what you are trying for and came upon this. Part is not useful, the rotons. However apparently they also were working with cooling methods which might parallel to some degree what you are after. So here is this:

An abandoned spacecraft attempt:
http://en.wikipedia.org/wiki/Rotary_Rocket

In addition, the rotating exhaust acted as an effective wall at the outer edge of the engine base, and the entire base area effectively is pumped down below ambient due to ejector pump effect, creating an effective suction cup at the bottom in atmosphere. This could be alleviated using makeup gas to develop base pressure, requiring effectively an additional rocket engine to fill up the base of the main rocket engine.
At the rim, 96 miniature jets would exhaust the burning propellants (LOX and kerosene) around the rim of the base of the vehicle, which gained the vehicle extra thrust at high altitude –effectively acting as a zero-length truncated aerospike nozzle.[2] A similar system with non-rotating engines was studied for the N1 rocket. That application had a much smaller base area, and did not create the suction effect a larger peripheral engine induces. The Roton engine had a projected vacuum ISP (specific impulse) of ~355 seconds (3.5 kN·s/kg), which is very high for a LOX/kerosene engine –and a thrust to weight ratio of 150, which is extremely light.[3]
During reentry, the base also served as a water-cooled heatshield. This was theoretically a good way to survive reentry, particularly for a lightweight reusable vehicle. However, using water as a coolant would require converting it into superheated steam, at high temperatures and pressures, and there were concerns about micrometeorite damage on orbit puncturing the pressure vessel, causing the reentry shield to fail. These concerns were resolved using a failure resistant massively redundant flow system, created using thin metal sheets etched via chem etch with a pattern of micropores, with a channel system such that it was robust against failure and damage.
In addition, cooling was achieved two different ways; one way was the vaporization of the water, but the second was even more significant, and was due to the creation of a layer of "cool" steam surrounding the base surface, reducing the ability to heat. Further, the water metering system would have to be extremely reliable, giving one drop per second per square inch, and was achieved via a trial/error design approach on real hardware. By the end of the ROTON program, some hardware had been built and tested. The reentry trajectory was to be trimmed, similar to the Soyuz, to minimize the G loads on the passengers. And the ballistic coefficient was better for the Roton and could be better tailored. When the Soyuz trim system failed and it went full ballistic, the G levels did rise significantly but without incident to the passengers.

If this is just noise to you please just pass it by, I am not well equiped to exactly understand what you need.

GW Johnson · 2013-03-17 23:10:20

Hi guys. We're just home from a week in Japan visiting friends and relatives. Jet lag is a killer.

Russel: what I have been able to find indicates that transpiring heat shields will indeed work, but system-wise, they are simply heavier than ablatives, due to the sacrificial coolant mass required. Maybe you want to look at something else. The only thing heavier was a simple metal heat sink.

X-20 was to have been an ablative nose cap and leading edges, with a high-temperature metallic re-radiating skins. The ablative could have been augmented by transpiration. It was a carbon tile thing, with embedded radial metal reinforcing rods, essentially the predecessor to the reinforced carbon-carbon composite used on the Shuttle for its nose cap and leading edges. They were looking at Rene-41 for the non-ablative re-radiating hot skins on the X-20. That would be the predecessor to the metallic SHARPS material proposed for X-33.

So, ablatives were the lightest solution, even at silica-phenolic densities. That's why we went that way with the space capsules, and even today with Dragon and Orion, although today's PICA and PICA-X are far lower density than the old phenolics. Then they invented non-ablative low-density ceramics, and used them on the Shuttle. These worked just like the Rene-41 would have worked on X-20, but were a lot lighter. Their only problems were low strength (damage susceptibility), and limited allowable skin temperatures for re-radiation (2300 F).

I think I know the answers to damage susceptibility: no side-mount configurations, and fabric-reinforced ceramic composite. My particular material also eliminates bonding troubles, and thermal expansion incompatibilities. But, it is still experimental. The last time I made this stuff was 1984.

NASA solved re-radiation with black surfaces for the windward side of its Shuttle, but only away from stagnation conditions on the nose cap and leading edges. I think I know a way to hard-surface my material with black instead of the white I used 29 years ago. But, it needs to be tried experimentally.

For entry from LEO, one can reduce the stagnation heating to around 20-30 W/sq.cm by careful attention to shallow angle, low ballistic coefficient, and very blunt (flat) heat shield surfaces. It's normally an order of magnitude higher with the old capsules and the Shuttle. For at-the-very-most-25 W/sq.cm stagnation peak heating, equilibrium re-radiation skin temperatures at emissivity 0.80 or higher can be a snit under the 2300 F limit for alumino-silicate low-density ceramics.

At Mars, things are far less demanding. I was calculating around 11 W/sq.cm for a 5.6 km/sec shallow entry from direct interplanetary transfer, which works quite easily for black-surfaced alumino-silicate at stagnation, but not the white surface (emissivity around 0.20). This was at 400 kg/sq.m ballistic coefficient, and a radius ratio about like a Mercury capsule.

From LMO, things are even easier at about 2-3 W/sq.cm, for which either black or white surface alumino-silicates are quite feasible. Same shape and ballistic coefficient.

For any surface that does not bend, I think I know a way to coat it with ceramic composite. Think density in the styrofoam range, like shuttle tile, but with sensible strength and toughness, too. It has to be integrated into the design, though. My stuff on a thin metal backplate, which panel in turn bolts to the spacecraft framing.

GW

GW Johnson · 2013-08-31 17:26:07

This one has been inactive for a while now. But, I wanted to let everyone know I have looked at a reusable single-stage chemical Mars lander, and found it to be feasible. I posted this in a short form over at http://exrocketman.blogspot.com in a posted dated 8-31-13, and titled "Reusable Chemical Mars Landing Boats Are Feasible".

As for the ceramic heat shield material in my last post above in this thread, I determined it really will work, and presented that outcome in a favorably-received paper at the 16th convention in Boulder. It was part of my reusable lander feasibility study, with PICA-X ablatives as a backup.

GW

louis · 2013-08-31 20:25:13

GW Johnson wrote:

This one has been inactive for a while now. But, I wanted to let everyone know I have looked at a reusable single-stage chemical Mars lander, and found it to be feasible. I posted this in a short form over at http://exrocketman.blogspot.com in a posted dated 8-31-13, and titled "Reusable Chemical Mars Landing Boats Are Feasible".
As for the ceramic heat shield material in my last post above in this thread, I determined it really will work, and presented that outcome in a favorably-received paper at the 16th convention in Boulder. It was part of my reusable lander feasibility study, with PICA-X ablatives as a backup.
GW

Interesting! I am not a rocket scientist but got a feel that an Armadillo style rocket could get people off Mars. I think it would be feasible to build the bulk of the parts for such rockets on Mars with 3D printing and scaled down furnaces.

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#390 2013-03-03 20:05:57

Re: Landing on Mars

#391 2013-03-04 10:47:59

Re: Landing on Mars

#392 2013-03-14 14:41:50

Re: Landing on Mars

#393 2013-03-14 15:38:14

Re: Landing on Mars

#394 2013-03-14 19:55:24

Re: Landing on Mars

#395 2013-03-15 19:20:41

Re: Landing on Mars

#396 2013-03-16 19:59:29

Re: Landing on Mars

#397 2013-03-17 22:28:42

Re: Landing on Mars

#398 2013-03-17 23:10:20

Re: Landing on Mars

#399 2013-08-31 17:26:07

Re: Landing on Mars

#400 2013-08-31 20:25:13

Re: Landing on Mars

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