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Re post 43 above: I have since realized that what everyone on these forums really needs is a 1st-look spreadsheet that roughs out the bounds on what can be done with both SSTO and TSTO, reusable or not. You do that first, with assumed Isp and inert fraction values, just to bound what might be feasible or not. The better your assumptions, the better your results. The more ideas you screen this way, the fewer you must put more effort into, trying to better estimate inerts, and trying to size engines and see if they fit.
I have come up with such a spreadsheet, and will get Tom to put it in the drop box for download, and post links to it here. I want it added to the orbits+ course materials, as part of lesson 8's materials.
GW
I got distracted, and did not notice the time. When I looked up, it was already past 8:00 central. Sorry.
GW
Most 2-stagers today, stage down near 60-some km altitude and a speed near Mach 3 to 6 (which is 1 to at most 2 km/s). That's also near 60-some km downrange, with the trajectory bent over to around only 10 degrees above local horizontal. The bulk of the dV to orbit (the other 6 km/s) is best supplied by the second stage at its higher Isp with a real vacuum nozzle design on its engine(s). The first stage has to have sea level or at best compromise-ascent nozzles on its engines.
If there are SRB's, these are usually staged of a little before the first stage core, so they can give a higher thrust over a shorter time interval, at a fixed packaged total impulse (propellant mass times its ascent-averaged Isp = to the time integral of thrust). It would not make sense trying stage all that stuff away all ay once. Way too risky!
The regular rocket launcher can be made reusable by recovering its first stage for a rather slow re-entry at around Mach 3 to 6-ish. But, with a decelerating entry burn to hold entry to Mach 3 at most, that relieves you of needing a heat shield for it, even if it is made of aluminum alloys. It does a powered vertical landing, as both SpaceX and Blue Origin have demonstrated. That first stage can have a stage inert fraction of around 5%, without that heat shield, and restricted to hard pad landings with minimal landing legs.
The second stage must be a fully-qualified re-entry vehicle for 8 km/s entries at the least, which means you must cover half or more of its surface with heat shielding. That right there pushes you to an inert fraction in the 7-10% range. Then it must have some means to land: (1) vertical powered landing, (2) horizontal dry lake bed landing if a lifting body shape, or (3) horizontal runway landing with wings.
Choice (1) means you must add some sort of minimal landing legs, restricted to hard pad landings. That will push you to prey near 10% inert fraction, which is just about where SpaceX has been with its Starship test vehicles.
Choice (2) has more surface area and thus more surface to protect with heat shielding. It will also need suitable high-speed landing gear to land on that lake bed. That will drive to to inert fractions somewhere in the 12-15% range. And depending upon weight during the return (it may have payload aboard), touchdown speed for lifting bodies was demonstrated long ago to fall in the 250+ mph range, perhaps close to 300 mph if still loaded with payload. That's why you need a huge dry lakebed for your landing field! All such high-sped landings will be inherently hazardous! It is easy to lose control while rolling that fast, and tumble the vehicle.
Choice (3) has yet more surface area and the mass of the wings, which will also need heat shielding. It will also need landing gear, but with wings, your touchdown sped can fall in the 150-200 mph range, which is much less hazardous, and can be feasible at airports with runways longer than about 10,000 feet. That sort of buildout is going to put you in the 15-20% range of inert fraction.
Your second stage will not have to fight much in the way of losses, so that 6 km/s second stage is its required dV. You have to carry enough payload to be worthwhile, and you will have to deliver that dV at an Isp high enough for inert fractions in the 10-20% range. You'll do a lot better with the higher Isp of LOX-LH2, although SpaceX is sort-of making Starship work at about 10% inert and LOX-LCH4.
If instead you decide to do your stages as some kind of vertical-launched cluster of spaceplanes, there are 2 things you MUST deal with: (1) your first stage inert fraction is going to be in the 10-15% range, because you need very little heat shielding for it, but you must be a lifting body or have wings. It only shoulders 1-2 km/s of delivered dV, but it shoulders almost all of the 1-2+ km/s worth of losses (there will be more drag loss! So that's something like 2-4 km/s worth of ideal dV capability, where its payload is the second stage (much smaller) spaceplane. And its inert fraction is 10-15%, depending upon the details. The second stage spaceplane faces the same design values and choices as listed above for (2) lifting body or (3) winged spaceplane.
I see no value to attempting 3 stages as all spaceplanes.
The best choice might actually be a real spaceplane atop a Superheavy-like booster.
GW
I did see one posting on LinkedIn by Steve Yoon of NASA that the scheduled first flight of block 3 Starship/Superheavy is May 12.
GW
By way of history, the original Atlas configurations were not technically 1-stage vehicles. It took off using 3 engines, and dropped 2 of them in the engine skirt, once the trajectory bent over and a lot of propellant had been burned off. From there it flew on the one remaining engine all the way to orbit insertion speed (or warhead release speed), as one long burn. Free-fall restarts were not available then.
Atlas was a full ICBM-capable rocket. The other single-stagers at the time were only IRBM-capable, such as Thor. Redstone/Jupiter/Juno was even shorter range, about like a Scud today. Thor eventually became the first stage core of the Delta launcher series. Titan was always a 2-stager, even at its beginning. It was the other liquid propellant ICBM that we had.
My estimates indicate that SSTO is feasible, at comparable payload fraction, fairly easily with LOX-LH2, but only just barely with LOX-LCH4. Feasibility requires the higher c* and Isp available with hydrogen. But that is true only with 1-shot stage inert fractions in the 5% range. You cannot do that low an inert fraction in a reusable stage, because it must also be fully qualified as a re-entry vehicle, and it must have some way to land. Both of those add greatly to inert fraction. It would fall somewhere above 10%, maybe as much as 20%, depending upon the exact design approach.
At 10+% inert fraction, SSTO is no longer feasible at all, even with hydrogen! The MR and Vex say so, because the required dV is just about 9 km/s to cover all the losses. At ~400 s Isp (as an average ascent value for an unseparated nozzle at sea level), that's Vex near 3.92 km/s, and that's a MR near 9.9, and that's a propellant fraction of just about 90%. That leaves you with 5% payload fraction if your inert fraction is 5%, which is a 1-shot non-reusable inert fraction! 10% inert fraction leaves 0% payload fraction, and even then may not qualify for re-entry, much less some means of landing! Anything over 10% inert fraction, and payload fraction is negative! Completely infeasible for that dV!
Talking about achieving 5% inert fractions for an ascent stage that is also a fully qualified re-entry vehicle, plus has some landing means, is just utter nonsense, with any materials and engine technologies that exist today!
And I have already shown that aerospike nozzle technology is NOT the magic-savior solution to this problem, despite all the marketing hype, because all free-expansion nozzle designs perform lousy out in vacuum, and they are lousy from somewhere down in the atmosphere above their design altitude, all the way out to that full vacuum. And that's exactly where the rocket must supply most of its dV, not down in the atmosphere!
GW
This was the headline summary in today's AIAA "Daily Launch" email newsletter:
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The Next Web
SpaceX has spent more than $15 billion on Starship
SpaceX has spent more than $15 billion developing its Starship megarocket and is pushing for a launch cadence that would make space access resemble an airline schedule rather than a government programme, Reuters reported on Friday, drawing on the company’s confidential pre-IPO prospectus.
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The link was to a Space.com site. I looked at the full article, which was a Reuters article. While Starship has been a drain at a few $billion very year, SpaceX was showing huge profits (several $hundred-billions) until it started acquiring the XAI thing and linking it to Starlink. It recently went negative on profits, but not by all that far. The real budgetary drain has been the acquisitions and AI thing. The article gives cost to orbit numbers for Falcons in the $2700-3000 per kilogram range. The target for Starship-based launch operations is $30-300 per kilogram. It is quite clear that the main role for Starship is to be a to-orbit freighter. But if it really is that cheap to operate, AND refill-on-orbit proves feasible, then sending few to the moon or beyond starts looking quite feasible, too. That has been the vision all along.
GW
I have been expecting a finding like this for several years. Ever since the Viking results were "explained away", and the Allan Hills meteorite ALH-8001 claimed microbe fossils were "explained away".
We now think we know more about the gross climate history on Mars than we did then. The thinking now is that Mars pretty well lost a thick atmosphere to solar wind and coronal mass ejection erosion around 3 billion years ago, which would be about 1.6 billion years since formation. That would be due to the lack of a shielding magnetic field. Before that loss it was far warmer and wetter and more Earthlike, although without oxygen in its atmosphere.
Bear in mind that before around 2.5 billion years ago, some 2.1 billion years after formation, the Earth did not have any oxygen in its atmosphere, either. Oxygen was slowly put there by pre-existing life in the ocean that had "learned" how to do photosynthesis. Earth's atmosphere stayed thick, providing warmth, because Earth had an adequately strong magnetic field to greatly reduce the atmospheric stripping rate of the sun's emissions. Thicker atmosphere early-on probably balanced the dimmer young sun (according to astrophysics as we know it). Thinner atmosphere in more recent times let the climate remain hospitable as the sun brightened with age.
Estimates of when life began on Earth are still but guesses, and you must allow for the possibility that it began and was made extinct multiple times, by volcanic/tectonic phenomena far more violent early in its history. But the best guess is that life started about half a billion years after formation, or some 4.1 billion years ago. That would be single cell stuff, in the ocean, most likely.
By around 2.5 billion years ago on Earth (some 2.1 billion years after formation), there may have been multi-cellular plant life, and it appears that either single cell or multicellular plants (or both) had started oxygenating the atmosphere with photosynthesis. Whether there were any single-cell things we might call animals is unknown. But oxygenation of the atmosphere (and parts of the ocean) made multi-cellular animals, and life on land outside the ocean, possible. It took that long apparently, on Earth.
Assuming (big assumption!) that the progress of life on Mars was similar, by the time the atmosphere thinned and the ocean froze up and evaporated away at about 1.6 billion years since formation, life should still have been single-cell (and perhaps multi-cellular forms, of things that we might call "plants"). Maybe single cell "animals", maybe not. But with the exposed surface gone lethal, and the ocean drying up, any vestiges of that life would persist only underground, away from the harsh radiation, and where there was still water to support the chemistry. That may indeed still be the case today, actually! We simply do not know yet! And we will never know, until we actually go there, and dig deep looking for it.
But such a hypothesis as mine, would explain the forms seen in the Allan Hills meteorite as the microbe fossils, that they were claimed to be back then. And they might still explain the Viking results, although reactions with perchlorates would also tend to explain that finding. And that hypothesis is most certainly consistent with what Curiosity just found with its chemistry experiment.
What all that says is that it would seem unlikely that any astronauts would run across Martian life (and all the risks that might derive from it), while just fooling around on the surface. But if they drill or mine, or explore caves, that outcome might well be different!
GW
SpaceX uses the boil-off vapors as the tank pressurants, and manages it to desired tank pressure levels by deliberate controlled venting. In addition, they use it for cold-gas attitude thrusters as well.
What they did in the tank-tank test inside the one vehicle, was use the cold gas thrusters to create ullage thrust. And that ullage thrust provides the slight acceleration needed to settle the propellants into the aft tank bottoms, where the pumps can draw a suction on liquid without vapor spaces in it. That ullage thrust also slightly alters the orbital path, so you CANNOT do it that way at a space station.
But it could be done with two vehicles firmly docked together. Which is how they plan to do it, refueling one Starship docked to the other. The acceleration is slight, so you must provide ullage thrust for some period of time (minutes?) before the propellants settle, and continue to provide it all during the pumping operation. It will alter the orbit of the docked pair of vehicles.
GW
Bob:
The images are posted in an "exrocketman" article of mine. It is titled "About the Artemis-2 Mission", dated 31 March. Search code is 31032026. The photos I found are mostly down in the updates added toward the end of the article. The most telling figure is the CBS photo taken while hanging on the chutes about to splashdown, and enhanced digitally and enlarged. The other most telling photo is the photo posted by NASA's Steve Yoon, taken of the heat shield underwater by the divers.
Most of the base heat shield looks really good in the diver photo, but it does not show the side where the stagnation point was. You can tell where it is by the streaks that point to it. The lateral wall shield on the windows and hatch side look really good too, except for one little crater, pointed-at by Reid Wiseman. The enhanced close up made from the CBS photo shows one of the four tie-down pads was melted, creating a bit of cavity. NASA says that is not unexpected. But you can see at least 4 small chunk-shed craters, and some other unidentifiable damage, all near the stagnation zone. You can also see part of the lateral wall shield on the side opposite the windows, near that stagnation zone. The damage is extensive, with patches of complete char loss, exposing heat-distorted metal. That's the outer shell, not the cabin pressure shell. There is insulation between the two layers of metal.
You have to be very careful of images found on the internet these days. Many are not what they say they are, many others have been doctored with AI or other means. The last photo I posted is one of those, as an example of what to be wary of. If it is an Orion at all, it is EFT-1. You can see the hex even in the lateral wall heat shield that was hand-gunned. Neither Orion for Artemis-1 or -2 had that, they had bonded hexless tiles instead.
GW
The vibratory effect was essentially the basis for the "worm" drill that failed so miserably on a recent lander. It works in fine-grained stuff that is ALL fine-grained. But Mars is not. The worm drill bumped into a rock just subsurface underground, and it could not navigate around that rock. The rock completely stopped it.
My point: the subsurface material on Mars is NOT a uniform material in terms of sizes. This is a mix of stuff from wind-blown dust all the way to house-sized boulders (or larger), because Mars is heavily cratered and covered with ejecta (like the moon. Why? Because of the very long time it has had a very thin atmosphere that cannot stop meteors very well at all.
The bedrock projects up through this in many places all over Mars, usually as craggy-looking stuff. It is both weathered, and cratered. In the northern lowlands, this surface is also overlain by sea bed deposits resembling sands, and maybe some clays. Or maybe not any clays. But that, too, has some cratering disturbance. Just not as much. Which suggests the ocean persisted a bit longer than the atmosphere did.
The "best estimates" I have seen published say that Mars essentially lost its atmosphere down to the thin thing we see today, dried up, and froze, something like 3 billion years ago (1.6 By since formation). Before then, it was supposedly far more Earthlike. The leading contender for "why" it lost its atmosphere is no significant magnetic field to ward off solar winds and mass ejections from stripping it away. But nobody really knows "for sure" what happened.
If life here (as best we understand its history) is any guide, there was microbial life in the Martian ocean and lakes. There may still be vestiges of it surviving underground here and there. But the planet likely went uninhabitable before multi-cellular life could evolve. Life on Earth was single-cell microbial as best we know, until about 0.6 billion years ago (4 By since formation). Our atmosphere was pretty much devoid of oxygen until photosynthesis changed that, somewhere around 2.5 billion years ago (2.1 By since formation).
Assuming similar life would follow a similar gross history there, life on Mars never even had a chance to put much oxygen into its atmosphere, which is thought to be a prerequisite for multi-cellular forms of life, at least multi-cellular animal life.
But if things happened the way I suggest on Mars, that would explain the long-derided supposed microbe fossils found inside the Allan Hills meteorite ALH-84001, which came from Mars.
Whoops, I wandered off topic! Sorry! But what you must deal with, operating upon that surface to build things, depends integrally on the past history that created it.
Screws do not work in soil containing big rocks (6-inch and up), especially close-packed big rocks. I have literally seen that, here in Texas, on the farm as well as out on construction sites.
The microwave idea would require an enormous amount of electrical energy. You literally have to melt ton-class masses of rock and rock dust around your pile.
If you use piles at all, the gas-driven hammer would be best. You must use harvesting dry ice and thawing it confined in a tight space, to gasify it at pressure, in order to compress it. Normal compressors like those here (about 6 to 10:1 pressure ratio) will NOT work in a 6-7 mbar atmosphere! They do not work at 110,000 feet here, either. Compression ratio 10 on 6 mbar is 60 mbar (still next to nothing), and the mass you compressed for the power expended is quite tiny, because the inlet density is so low!
Rather than piles, myself, I would go for some sort of slab construction laid on a well-graded surface. Heavily reinforced, of course. Because of the internal pressures it has to support. Which means we need the equivalents to concrete and rebar. We do not have them!
So that gets down to the piling of rocks for dry-fit masonry, upon a well-graded surface. If the thing is big, it will require a proper "foundation" of close-bedded stones in layers, decreasing in size to gravel at the surface. Road beds, when properly done, use that technique. It will have to be quite massive to resist the blow-out loads of internal pressurization on Mars. And the lower gravity reduces the weights of the stones that produce the friction that resists those blow-out loads!
Let's just say there must be some serious in-situ construction experimentation before real recommendations for building construction techniques can be made, and let it go at that.
But THAT is a big part of why I say that there is an "experimental base phase" between initial exploratory landing, and the building of anything permanent like a city or a colony. Or your colonization effort is doomed to failure.
GW
Turns out there was a glitch. The second stage put the satellite in the wrong orbit. Not just "wrong", completely unusable.
It is hard to tell from reports filed by reporters not technically competent, but it sounds like the second orbit-adjusting burn demanded of the second stage about an hour after it shut down, did not happen.
GW
Congrats to Blue Origin. It has joined ranks with Spacex in the heavy launch to orbit business, complete with reusable first stage boosters. They still need to establish the same kind of good track record as SpaceX has with its Falcons, but as good as their technology looks, I think that will happen, and fairly soon.
Meanwhile ULA's Vulcan is grounded for a fault in its SRB nozzles, and has yet to demonstrate any reusable anything. So, my conclusion is that "new space" is pulling ahead of "old space" in the heavy launch business. Not to mention several small "new space" launchers out there.
GW
I think the closeup photo is an enhanced version with clearer focus, but a reduced view dimension, of the blurry photo that shows the whole capsule.
I an not at all sure this has anything to do with being "hoisted up". I suspect without proof the blurry photo was taken just before splashdown, still hanging from the main chutes.
Take a good look at the clearer restricted-view photo. This was intended to show the damage near one attachment pad that did not survive. That is what the whitish "stain" is, staining from the melting metal. There is extra erosion there, too.
But, look up at the lateral side near that same place. Do you, or do you not, see exposed and distorted metal, and maybe a burn-through, where the lateral-side heat tiles were thinner? Maybe too thin? Those whitish areas are not windows, they are exposed locations of the metal outer shell to which the heat shield tiles were attached.
GW
Another follow-up to post 89. I finally found enough photos to make a preliminary assessment of the Artemis-2 heat shield.
Both NASA and I were right, in the sense that NASA said that a single heating pulse no-skip re-entry would reduce cratering damage (it did). I said that cratering would still occur, albeit reduced and smaller, without the reinforcing hex (and it did).
I did see something else nobody expected: complete localized heat shield loss and metal distortion or burn-through, on the more windward lateral side of the capsule, while flying at angle of attack to generate a side force for fine trajectory control (something done since Gemini).
It would appear that flow along that side, supposed to be a separated wake zone, was instead at least intermittently attached, with resulting far-higher convective heating than the thinner Avcoat tiles there, could resist. This was the side opposite the windows. That windows side looked to be in good shape.
GW
I put an Artemis-2 mission article up over at my "exrocketman" site, titled "About the Artemis-2 Mission", posted 31 March 2026, search code 31032026. Today (13 April) I added an update with photos, that lead to a preliminary assessment about that heat shield.
Both NASA and I were right.
NASA was right that eliminating the two-heating-pulse skip entry would reduce heat shield damage. It did. I was right in saying that at least some chunk-shedding crater damage would still occur, and it apparently did. The photos I found so far prove that.
Unexpectedly, I also saw heat shield damage to the lateral wall of the capsule, on the side closer to the stagnation point, away from the windows. That is where the Avcoat bonded tiles were thinner. Hex in those tiles would greatly reduce that, but maybe a bit thicker is also needed, too. The damaged metal was the outer skin to which the tiles are bonded. That is NOT the inner cabin pressure shell!
What that means is that if NASA would put the hex into the bonded tiles, less damage would occur, even with a two-heating-pulse skip entry. I have already told them how to do that, without resorting to manual hand-gunning of the Avcoat in the hex cells. They need to do that before flying again to the moon, with a 10.9 km/s entry. Artemis-3 will come back at only 7.9 km/s from LEO. Even what they have now will be adequate for that.
GW
Both SpaceX and Blue Origin are companies that first and foremost have to make a profit, not lose money. SpaceX was profitable with Falcon-9/-Heavy, but has bet its future "farm" on rideshares with Starship/Superheavy. They still have their hands full just trying to make Starship/Superheavy work at all as a transport to LEO. Blue Origin has its hands full trying to make New Glenn into a profit center. Coming up with a contracted NASA lunar lander is a smaller piece of that overall larger puzzle, for both of them. I cannot fault their priorities.
Of the two, I suspect Blue Origin might be a little closer to satisfying the NASA lander contract. That is because SpaceX bit off a much-larger piece of "iffy" technology advancement, trying to do the all-reusable Starship/Superheavy. Plus, my reading of the events suggests the ratio of Musk time to real time (3 to 4) is a bit bigger than Bezos's ratio (2 to 3).
The time from lunar rendezvous being the adopted Apollo architecture in 1964 or 1965, to the Apollo-9 checkout of the Apollo CSM with its LM in LEO in 1969, was only 4 or 5 years! THAT is how long it took Grumman to come up with a workable lander, under a crash program where cost was no object. And higher risk-taking by NASA with its astronauts was "normal".
Artemis is NOT a crash program where cost is no object, and NASA (I hope) has learned not to take such extreme risks with its astronauts! Expecting SpaceX and Blue Origin to come up with anything workable as a lunar lander in only 4 years or so, is actually quite unreasonable! SpaceX started only 2-3 years ago, and Blue Origin "in earnest" only last year.
You CANNOT count the proposal and contract-win time, as real hardware development time! That only sets the concept they will focus upon. REAL development only starts AFTER contract award. And coming up with a concept has NOTHING to do with its development into something real! That's just life. Ugly, ain't it?
NASA projecting schedules that have no reality tells me there is no one there anymore that understands the difference between company time and real time, and that the ratio varies from company to company. I would expect that, after all the former traditional contractors agglomerated into monopolies that no longer really compete (with the government making no anti-trust moves to stop it). THAT is why "new space" has had such a hard time getting established. The game was rigged.
GW
With Spacenut's photo of the fiberglass hex with the hand-gun tool inserted, I an now sure that the hex's resin was phenolic. I saw a lot of electrical and electronic board materials in the 60's and 70's made of this very same stuff. The color is the key to identifying it: that orange is commercial phenolic resin, on plain white fiberglass cloth. Cures at modest heat and only some pressure between mold platens, if you are making flat panels. I do not know what tooling was used to make hex.
The type of phenolic that went into the glass and silica phenolic materials was different! It cures under greater heat and a lot of pressure, and is tan in color. I used a lot of silica phenolic in ramjet nozzles, and in rocket nozzle assemblies. It is tough, dense, slow-ablating, and very heavy. The fiber is in woven cloth form, and you must be very careful to orient the cloth layers correctly relative to the flow direction.
GW
Avcoat is a thick, not-quite-liquid, paste that is more like a dry mortar or cement material, comprised of epoxy-novolac polymer heavily loaded with silica fiber and phenolic microballoons. The polymer contribute the carbon to the char layer, with the silica fibers contributing some silica content. But the char is largely porous amorphous carbon. I do not know the standard percentages of the components, but I do know they can vary, especially the microballoon content. Porous amorphous carbon char handles to the touch about like a piece of charcoal from the BBQ grill that is burnt-through but not yet consumed to ash.
The microballoons contribute the porosity required to get the pyrolysis gases out from the pyrolyzing layer through the char. That is a serious issue in inch-plus thicknesses, not very much in fractional-inch thicknesses. And that is because the lower the permeability letting the gas out, the higher the driving gas pressure must be, to get out. And carbonaceous char is a very structurally weak material, especially in tensional loadings.
The microballoons lower the density, to around sp.gr = 0.51, instead of slightly greater than 1. That also increases the ablation rate (which is both pyrolysis and erosion of the char from the surface as fine grit). Higher microballoon content is lower density, higher ablation rate, and higher char permeability. It's a tradeoff, and can be varied from place to place on the heat shield, if desired.
The heat shield on Apollo and on Orion EFT-1 was hand-gunned with what amounts to an air-powered caulking gun into each and every cell of a fiberglass hex bonded to the capsule structure. I think it was probably fiberglass-phenolic, but I do not know for sure that it was phenolic. I am sure of the fiberglass. These cells are on the order of at most hlf an inch in dimension. There were almost 300,000 of them on Apollo, and nearly 400,000 of them on Orion, in part because the lateral sides also needed the protection. That glass fiber hex reinforcement provides tensile strength to retain char from breaking off, and acts to limit cracks propagating from cell to cell.
The enormous time and cost of the hand-gunning is why they decided to cast Avcoat blocks and machine precision tils from them. These were bonded to the capsule structure the way that PICA or PICA-X tiles would be bonded. The bonds and gap-fillers worked fine on Artemis-1, but the retention of the char did not. Without the reinforcing hex to hold it down and limit crack spread, several large chunks and bunch of small ones spalled off during that entry. Complicating that was this was a skip entry, with 2 heating pulses separated by a modest cooldown. A lot of these materials, particularly silica, suffer a solid phasa change at about 2300 F that causes shrinkage by around 3%, and embrittlement to the point of no strength at all: they just crumble at a touch.
That hex-reinforced Avcoat worked just fine on every Apollo and that Orion EFT-1. Something similar flew on Gemini with the hex cell thing on only the heat shield (the lateral sides were bare superalloy), but the polymer was a Dow Corning silicone, and I do not know what solids it was loaded with.
The problem was that Artemis-2's heat shield was built and shipped for assembly before Artemis-1 ever flew. The spalling of chunks caught everyone by surprise. Their thermo-structural models and arc jet data did not predict this. So was it the lack of hex, or the two-heating pulse skip? NASA spent a year convincing its management that it was the skip, so they flew Artemis-2 back with almost no skip at all.
Myself, I think it's actually both effects. They need to put the hex into the tiles, but they need to do it without hand-gunning, or they might as well go back to the Apollo and Orion EFT-1 technique. I figured out a way to load all the cells at once in a chunk of hex, using an extrusion press, in order to make hex reinforced blocks for machining the bonded tiles. And I gave that to NASA, although so far they have ignored me.
I have seen one blurry photo of Artemis-2 being hoisted out of the sea. Everybody comments on the weird-looking but expected damage near one of the four hold-down pads. I thought I saw some missing-chunk craters, fewer and smaller than Artemis-1, but there! But, the photo was blurry, so I as-yet know nothing for-sure! I did see some localized outer-layer burn-through damages on the lateral side, low down, close to the heat shield, in some of the photos of the crew standing next to it. The Avcoat is very thin there.
GW
There are 2 "credible" lunar lander candidates, the SpaceX HLS, which is still nothing but a paper design, and the Blue Origin "Blue Moon" landers, the smaller uncrewed version being under construction. I do not know if the larger crewed form has seen any construction yet.
How those undeveloped statuses blend with some sort of Artemis-3 mission in LEO in 2027, is beyond me to understand! That mission is publicly said to be about docking with, and maneuvering with, one or the other or both landers, as prep for a landing mission (or two landing missions) in 2028.
Sounds like BS to me. I must conclude that there is no one left alive at NASA today, who is competent to distinguish between Musk time and real time, or Bezos time and real time.
GW
A follow-up to post 88.
I saw a NASA news conference where they indicated that the recovery divers took underwater photos of the Artemis-2 heat shield while the astronauts were being extracted. None of these were shown at the news conference, and I cannot find them on-line.
You would think that if these showed no char chunk-shedding cratering, they would have been posted quickly, to quiet critics like me, former astronaut Camarda, and many others. But THAT has NOT happened!
I have to wonder what those photos really show!
GW
Forgive me, but why is the Artemis-2 stuff listed under "unmanned probes"? There were 4 people aboard!
GW
The Navy does not operate Sikorsky Skycranes. The Army did fly a few, but most of them serve commercially now.
GW
Well, Artemis-2 made it back with 4 astronauts safe and sound. I am very glad about that. I have seen no images of the condition of their heat shield yet. Sure would like to see one! I am expecting to see chunk-shedding cratering, similar to Artemis-1, maybe reduced, maybe not.
Meanwhile projecting an Artemis-3 flight to dock with a lunar lander in LEO in 2027 is nonsense, until there really is a lunar lander to dock with. So is projecting lunar landings in 2028. Not until there is a lunar lander!
GW
I see that SpaceX has pushed the next Starship/Superheavy launch off into May. No reason has been given that I know of. But I suspect this one is different enough from the block 2 version that flew previously, that it is taking them longer to make sure they think it is ready to fly. THAT is why Musk time and real time differ by about a factor of 3.
GW
Today's "Daily Launch" from AIAA finally brings up the heat shield issue facing Artemis-2 as it heads to re-entry. The link was to a Scientific American article of some kind. Not all the supposed facts quoted in it were correct, but the concern over char cracking leading to the shedding of chunks of char was correctly pointed out.
Myself, I think the odds are good that the crew will come home just fine. Where I differ with NASA is that I think the same damage seen on Artemis-1 will happen to Artemis-2's heat shield, despite switching to a non-skip, single heating pulse entry. It might be less than what happened to Artemis-1, or it might not. But if ANY chunks at all get shed, then NASA was wrong and I was right!
GW