Preliminary Evaluation of Artemis-2 Heat Shield

Not much credible information is yet available,  as only a handful of official NASA photos have been released.  There is one unofficial CBS news photo,  whose digitally-enhanced zoom-in,  fills-in some of the information gaps.  But I do my best with what there is!

Figure 1 depicts the expected re-entry flow and heating pattern for a lunar return by Orion,  which “flies” at angle-of-attack,  to generate a small lift force that is used to fine-tune the trajectory shape.  It does this by rolling about the wind vector,  to point the lift vector in the desired direction.  This was also done on Apollo,  and on the earlier Gemini missions.

Figure 1 – Expected Flow and Heating Pattern for an Orion Lunar Return

The Orion capsule has now flown 3 times as of this writing,  once before the Artemis program even existed!  All 3 were re-entries at near escape speed,  which is what a lunar return actually is!  That first test was called EFT-1,  launched using a Delta-IV rocket (now retired).  It had the heatshield fabricated from Avcoat,  hand-gunned into the cells of a fiberglass hex bonded to the capsule outer shell,  just like Apollo!  The other two Orion flights (so far to date) were launched with the SLS rocket,  and had cast and machined Avcoat tiles bonded to their outer shells,  unreinforced by any fiberglass hex!

Figure 2 below is a photo of what I believe to be the Orion flown on EFT-1.  I cannot vouch for the pedigree of this photo,  or that some of the char has not been deliberately removed for investigating the material beneath.  But the presence of the hex is clearly visible,  so this cannot be either the Artemis-1 or Artemis-2 Orion capsules!  If an Orion at all,  and I believe it is,  this has to be EFT-1! 

Note that a fair amount of ablation and erosion damage seems to be evident on the lateral side.  I cannot say for sure,  but this would appear to be the lateral side nearest the offset stagnation point,  where attached flow would feature heating comparable to that at the rim of the base heat shield.   The base heatshield itself is not visible in this view.

Figure 2 – Photo,  Pedigree Unknown,  of What is Thought to Be EFT-1

NASA changed the manufacture of the Orion capsule heat shield for Artemis,  retaining the Avcoat material,  but deleting the reinforcing hex,  to save time,  effort,  and money.  They built both the Artemis-1 and Artemis-2 heat shields this new way,  without waiting for the results of the uncrewed test flight that was Artemis-1 (and THAT was their fundamental mistake).  They were surprised by the unpredicted nature of the damage that Artemis-1 exhibited!  It shed both small and large chunks of char during re-entry,  leaving alarming craters in the heat shield!  Some of those are shown in Figure 3 below. 

This Artemis-1 flight was a skip re-entry with two heating episodes separated by a slight cooldown,  conducted at a very slightly shallower-than-normal angle,  compared to a straight-in re-entry.  Shallower reduces total heating a little,  but it does incur at least some risk of bouncing off the atmosphere like a skipped rock,  into an extended elliptical orbit,  whose period (of 5-10 days) exceeds the remaining crew life support duration! 

Flying the capsule at angle of attack is a way to control that skip effect.  Early in the re-entry,  you point the lift vector down,  to stop any skipping-off.  Late in the re-entry,  you point the lift vector up,  to keep the trajectory from “drooping” downward too soon.  

Figure 3 – The Official NASA Photo of the Artemis-1 Heat Shield,  After Recovery

After Artemis-1,  NASA spent nearly 2 years doing tests and calculations to “officially” convince itself that if they deleted the skip and came in slightly steeper (at higher heating),  that the alarming chunk-shedding would not occur!  This was based on the hypothesis that the pyrolysis gases produced during the second heating pulse could not percolate through the char layer easily enough,  and ended-up cracking it,  and blowing-off chunks of it,  with the obstructed gas pressure from beneath.

That hypothesis ignores the effects of fluid-scrubbing shear action in attached flow,  which would want to peel chunks off from between any cracks.  It also ignores any embrittlement and weakening of the charred material,  after cooling down some,  between the heating pulses.  Anyone who has ever dealt with the fragile mantles of a Coleman gasoline lantern,  would know exactly what embrittlement effect I am talking about,  and that it is quite real!

So,  NASA flew Artemis-2 crewed,  with the very same heat shield as Artemis-1,  just deleting the skip in favor of a slightly-steeper straight-in re-entry.  Their analyses said that would eliminate the chunk-shedding.  The real question is,  did that really work?

The few official photos released so far say that it did work.  However,  there is an unofficial photo that says “maybe not near as well as assumed”.  You judge for yourself. 

Figure 4 is an official photo taken during crew extraction from the capsule,  floating in the sea.  Bear in mind that you cannot see the capsule base heat shield at all in this view,  and you can only see the lateral side where the windows were located.  Those have to be away from the worst lateral heating,  to protect them from being destroyed by that heating.  

Figure 4 – Official NASA Photo of Artemis-2 During Crew Extraction

Figure 5 is an official NASA photo of mission commander Reid Wiseman pointing at some kind of an eroded crater in the lateral wall ablative insulation.  You can see the window behind his head,  and you can see there is no heavy charring (black) in this view.  This view is of the windows-side of the capsule,  where heating and its effects are greatly reduced by being in separated wake flow.  You can see nothing of the base heat shield at all,  or the more highly-heated opposite lateral side,  where flow was attached. 

Figure 5 – AI-Doctored Version of an Official NASA Photo of Mission Commander Reid Wiseman Pointing to Damage Spot

Figure 6 is an official NASA photo taken by a Navy diver,  of the base heat shield,  while the capsule was still in the water.   It looks rather pristine.  However,  you can see by the streaks that the stagnation point was just out of view,  top center left of photo.  The odd feature lower left is the mark or trace,  left on the heat shield from the flow about one of the tie-down pads. 

We see no craters from lost chunks of char in this view,  which supports the conclusion that NASA was right to delete the skip in the re-entry trajectory!  But we cannot see the region nearest the stagnation point,  where heating is higher,  and where fluid shear is higher.  That would be because the acceleration to sonic at the rim,  takes place over a much shorter distance.  That raises the surface shear forces that the heat shield material and its char layer “feel”. 

A suspicious person might say that only photos supporting NASA’s hypothesis have been released.  That is because we have not seen any photos of the offset stagnation region on the base heat shield,  or of the lateral wall on the side adjacent to that stagnation region (opposite the windows),  where heating is almost as high as on the base heat shield.  Know that eventually,  all the photos must be released,  in the report on their heat shield investigation.  That report must be publicly released!  But it may be a year before we see it!

Figure 6 --  Official NASA Photo of Artemis-2 Heat Shield Taken by a Navy Diver

As I said above,  there is an unofficial photo now circulating,  that was taken by CBS News seconds before splashdown,  while the capsule was still hanging from its parachutes.  It was distant and somewhat blurry,  but it does show some sort of white mark near the heat shield rim. 

That white mark attracted a lot of attention,  and led to the digital enhancement of that photo,  and a digital zoom-in to examine that white mark more closely.  But that enhanced photo does indeed also show the near-stagnation region of the base heat shield,  and the lateral wall away from the windows,  where flow was attached,  and the heating much higher.  That original blurry CBS News photo is Figure 7 below,  and the zoomed-in enhancement is Figure 8 below. 

You cannot see very much in Figure 7,  but you can in Figure 8!  The white mark was left by the melting and destruction of one of the tie-down pads.  This is something NASA says was expected,  although it did not occur on Artemis-1,  as you can see in Figure 3 above.  It does seem to have left an alarmingly-deep cavity eroded into the rim of the base heat shield!  Expected or not,  that cavity would appear to be of a depth comparable the thickness of the heat shield.  And I do find that alarming!

Figure 7 – Unofficial,  Blurry Photo Taken by CBS News Seconds Before Splashdown

What nobody has been talking about are the other things I see in Figure 8.  I have circled four places that I believe show where chunks of char were shed from the base heat shield.  These are smaller chunk-shed craters,  to be sure,  and scaling up from the limited view,  not anywhere near as numerous as those seen on Artemis-1!  Yet they are there,  and the revised non-skip re-entry so very clearly did not entirely stop them from occurring!  Which simply says that something else was going on with that char chunk shedding,  besides pyrolysis gas percolation through the char! 

There’s one other circled spot,  and two arrows pointing to large locations,  on the base heat shield in Figure 8,   where I cannot tell what happened,  but I can see that some sort of damage is clearly there.  It will take better photos than this to evaluate those damages. 

What I see on the lateral side adjacent to the stagnation point are one unidentifiable small dark spot,  and two whitish bright spots of considerable size.  The two bright spots are clearly places where all the char was lost,  exposing bare metal to full heating!  And that metal looks distorted by that heating!  Very alarming indeed!

Those bright spots are not windows,  those are the bare metal of the outer capsule shell,  to which the Avcoat tiles were bonded!  Simple thermal insulation separates it from the inner metal shell,  which is the capsule crew cabin pressure vessel. 

Figure 8 – Digital Enhancement and Zoom-In of Unofficial CBS News Photo

There would seem to be four things going on here that affect the damages seen,  not just the one thing that NASA hypothesized.  They are:

#1. Pyrolysis gas percolating out against permeability resistance,  wanting to blow chunks off. (That is the NASA hypothesis.)

#2. Fluid surface shear forces wanting to peel chunks out from between cracks in the surface. 

#3. Char layer shrinking,  cracking,  and embrittlement upon cooldown,  between the heating pulses of a skip-type re-entry.  (Like the fragility of a gasoline lantern mantle.)

#4. The presence of the reinforcing hex actually ties the char layer tighter to the pyrolysis and virgin layers beneath,  and it also acts to limit the spread of cracks in the surface.

My conclusions about the rational things to do,   depend upon what response NASA takes to this Artemis-1 and -2 outcome:

#1. If NASA ever wants to resume flying skip-type re-entries,  then put the reinforcing hex back into the Avcoat!  Period!  There actually is a way to do that,  without hand-gunning Avcoat into every hex cell like Apollo!

#2. Non-skip re-entries have higher peak heating.  The Avcoat tile thickness is insufficient at the attached-flow locations:  near the tie-down pads,  and on the lateral wall opposite the windows.  NASA must thicken it at those locations!

Avcoat tiles with reinforcing hex,  but without hand-gunning:

Avcoat is an epoxy-novolac polymer loaded with solids.  Those solids include some small amount of carbon fibers,  and a lot of tiny micro-balloons made of phenolic resin.  Higher micro-balloon content lowers density,  raises ablation rate,  and increases the porosity and permeability of the char layer (also decreasing its strength).  It also greatly increases the apparent uncured mix “viscosity”,  which is already extremely thixotropic (almost a solid). 

The Avcoat mixture is so thick,  that it is almost “crumbly” coming out of an air-powered caulking gun,  whose nozzle matches the size and shape of the hex cells.  The hex is a fiberglass cloth with a phenolic resin matrix.  The glass softens at a higher temperature (~ 900 F) than that at which the polymer starts to pyrolyze (~ 300 F),  which is why it is an effective char layer stiffener and retention aid.  The carbon fibers add a bit to those reinforcing effects.  Apollo and Orion EFT-1 used hex panels bonded to the outer shell,  into which cells the Avcoat was hand-gunned.  This was extremely labor-intensive!

For the Artemis-1 and Artemis-2 Orion capsules,  the Avcoat mixture was cast into blocks,  from which tiles were machined.  There was no reinforcing hex!  The bonds and gap fillers of those tiles are not in question here,  those performed just fine!  The lack of reinforcing hex is the question!

You can make tiles with hex in them,  but only if you can stop being bound by “either/or thinking”.  That would be either doing it the Apollo/Orion EFT-1 way,  or doing it the Artemis-1/-2 way.  However,  there is a third path!  Read on:

Put a chunk of the hex the size of the cast block you want to make,  into some tooling on the outlet of a plastics extrusion press.  Load your Avcoat mixture into the press,  and use that press to force it through all the cells in the hex,  all at once!  Remove the loaded-hex tooling from the press,  put the bottom and top on that tooling,  and cure that block of Avcoat that now contains the reinforcing hex!  Then,  machine your tiles from those blocks,  and bond them to the Orion outer shell,  the same way as in Artemis-1 and -2!  Tiles that are hex-reinforced,  but with NO hand-gunning!

I gave this idea to the NASA heat protection group in Houston well over a year ago,  as of this writing,  and again directly to the new NASA Administrator Jared Isaacman only a few weeks ago.  I was able to confirm (1) that the heat protection group got it,  and (2) that they thought I was right.  Nothing confirms that Isaacman ever saw my letter to him. 

But I never heard another word out of NASA about this alternative,  and NASA has not yet done anything like it.  So,  I must conclude:

Money and schedule clearly still outweigh crew’s lives for the decision-making NASA management levels.  And “not invented here” is still quite strong at NASA,  as well as its contractors.  Those two things are my real ongoing reservations about NASA!  And they have been,  ever since the first of two lost Space Shuttle crews!  Crews lost precisely because of those very same two management culture flaws!

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Closely-Related Postings:

“About the Artemis-2 Mission” posted 31 March 2026

The updates to that article have these same photos in them,  just not as explicitly annotated as here.  And I have since added how the flow and heating patterns vary.  In near-escape Earth entry with a blunt heat shield,  plasma radiation heating is larger than convective heating,  and it decreases with distance from stagnation less rapidly.

Search code DDMMYYYY format      31032026

Search keywords         aerothermo,  launch,  radiation,  space program 

“Ramjet Data Re:  Heat Shields” posted 1 March 2026

Shows how my old experiences with ablatives in solid rocket motors,  and especially ramjet combustors,  have strong overlap with the re-entry phenomena involved here.  Char retention by the layers below,  is a crucial key,  in both venues! 

Search code DDMMYYYY format      01032026

Search keywords         aerothermo,  ramjet,  space program

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Trump the Incompetent Liar

 I originally wrote this as a column and submitted it to the Waco "Trib" newspaper almost a week ago.  They have so far not chosen to use it.  So I have posted it here.  

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Trump lied about the economy.  He actually inherited a recovering economy with low inflation,  and then damaged it severely with his tariffs and trade wars.  Further,  you are now facing simultaneous global recession and out-of-control inflation,  because of the closure of the Strait of Hormuz. 

Trump lied about tax cuts.  You got a small one (if any at all),  while his billionaire and corporate cronies got billions!  So,  the national debt is going completely out of control,  especially with high inflation driving up interest rates.  The interest on the debt now dominates the annual “budget”.  Which proper budgets Trump’s enablers in Congress repeatedly refuse to pass.

Trump lied about immigrants!  Most of them,  including most that are here illegally,  are more law-abiding and honestly tax-paying,  than the bulk of you readers out there!   So say the crime statistics!   And those same people are integral to our economy.  They harvest your food,  build your homes,  mow your lawns,  and serve as housekeepers.  And more!

Yet Trump’s lies about how evil all immigrants are,  have been the “justification” to weaponize ICE and CBP into something to abuse and “disappear” people that he does not like,  while trying to intimidate the rest of us into submission!  That is the very definition of a dictatorship’s secret police force!  And while you may have voted for Trump,  you did NOT vote for that!

Trump has lied about everything since the start of his Iran war,  including especially the “why”,  and the “why now”.  The experts dispute quite strongly that the Iranians were anywhere near ready to build a nuclear bomb.  Trump was frustrated because he couldn’t get a deal about the enriched uranium (after he himself scrapped the one we had before).  Then Netanyahu walked up wanting to start a war,  because of Iran’s proxy Hezbollah. 

Trump lied about how thoroughly we destroyed Iranian military capabilities.  Those lies are painfully evident.  While reduced in numbers,  Iran can still strike targets all over the region,  and ships in the Strait of Hormuz!  Which was open until Trump started his Iran war!  And now,  we have mostly run out of interceptors,  so those fewer Iranian weapons get through!

Trump has been lying about the progress of the peace negotiations with Iran!  Evidently,  he did not even know who he was really dealing with!  Or why the things he thinks he obtains during these talks,  repeatedly get quickly taken away. 

Those Iranian officials at the table in Islamabad are NOT the de-facto rulers of Iran now,  the Iranian Revolutionary Guard Corps (IRGC) is!  The IRGC is a huge,  well-armed,  terrorist army of violent extremists,  recruited specifically because they are violent extremists!  They would die to the last man,  before stopping the use of the other proxy terrorist armies (Hamas,  Hezbollah,  the Houthis,  and many more),  before giving up their enriched uranium for their terrorist bomb,  or before opening that Strait!

Trump simply does not understand non-transactional people like that,  not at all!  And he (and Netanyahu) actually made the IRGC the de-facto rulers of Iran,  by killing off most of the clerics and many of the regular government officials,  in the first few days of this war! 

Further,  by bombing civilian areas,  Trump has essentially united the people of Iran against us,  instead of helping them to overthrow their regime.  Regime change is actually the only way to get what we’d like to have from Iran!  It is now out of reach!

It appears to me that Trump started this Iran war because he could not get the then-ruling clerics in Iran to deal transactionally with him.  Just like Maduro in Venezuela.  And soon Cuba,  if and when Trump gets clear of his Iran debacle! 

In order to get clear,  Trump must now commit war crimes by deliberately bombing civilians and their infrastructure,  all across the country!  All the aboveground military targets actually have been destroyed,  although clearly many remain,  deep underground! 

You who voted for him,  you did NOT vote for American war crimes!  America’s name is now mud around the world,  because of the stupid things Trump has done,  including to our allies!  They no longer trust us at all,  and justifiably so,  nor do they even want to trade with us,  anymore!  And one of them,  Ukraine,  has developed a cheap,  mass-produced anti-drone drone,  that we do not have!  In 20-20 hindsight,  Trump’s mistreatment of Zelenskyy,  and starving Ukraine of assistance against Putin’s invasion,  looks very stupid indeed!

There is absolutely no excuse for that level of lying,  that much chaos,  and that egregious level of incompetence,  which we have seen out of Trump and nearly every appointed figure in his entire administration!  And all of you readers bloody well know that what I said is true!

Now,  readers,  please go out and do your job as proper citizens in November!  Vote all of Trump’s enablers out,  so that we can get rid of all this chaos and incompetence before it destroys us,  and get back to being a responsible world power.  With a decent economy.  But without almost-a-king!

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below is a photo of a Ukrainian-developed drone interceptor,  plus some more comments

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A Ukrainian Shahed interceptor drone. Photo credits: Bohdan Myroshnychenko

This drone is actually cheaper to build than the Iranian "Shahed" drones that it can destroy.  The Ukrainians developed this on their own,  faced with chronic mass "Shahed" drone attacks from Russia.   The Iranians have been supplying Russia with those "Shahed" drones for some time now.

Those who know me,  know that I consider what Trump has done to our alliances and our allies,  to be treason of the aid and comfort type.  The most egregious case has been trying to ram Russia's terms down Ukraine's throat,  in his so-called "peace process" negotiations.  But every alienated ally is another count of that treason.  

Those who know me also know that I consider Trump's egregious and infamous executive over-reach to be really the attempted establishment of a Trump dictatorship over us.  He is doing that so that he won't have to leave office until he dies.  That would leave us with a whole series of Trump-wannabees as dictators over us.  I estimate that imposition to be about 70% complete now.

The weaponization of ICE and CBP into a secret police force that "disappears" some of us,  and intimidates the rest of us,  is a piece of that dictatorship.  As are the "detention centers" which are really concentration camps for those waiting to be "disappeared".  As is the identification and vilification of some group,  as the enemy to be "disappeared" (in Trump's case,  immigrants).  All such are features of all dictatorships.  

Do not listen to the words,  from anybody!  Look only at the events and actions!  They verify everything that I have claimed.  

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Elliptic Departure and Arrival

Update 4-22-2026:  This post is fundamentally about using extended elliptical orbits and "space tugs" to reduce the departure (and arrival) velocity requirements on lunar and interplanetary craft of any type.  That technique provides significant benefit even if you throw away the tug stage,  as Artemis-2 just demonstrated. 

However,   if you combine it with a reusable "space tug" stage,  plus a space station in low circular where that space tug stage can be based,  the benefits greatly magnify.   Such a station needs to do both lunar or interplanetary craft assembly,  and the filling or refilling of such craft and the tug with propellants.  It does not require continuous manning,  unless other tasks needing that,  are also done there.  That space station base is also covered in the article.

The critical enabling factor will be propellant transfer from tank to tank in zero-gee!  At a space station,  you do NOT want to do such transfers by spinning large structures,  or with orbit-altering ullage thrust!  We can already do that task for storable propellants with bladder-expulsion tanks.  But there has not been such a solution for cryogenic tanks.  

Until now.  

I came up with a scheme that would work quite well.  The article covers it,  as well. 

This article thus gives you the path to an affordable,  but very capable,  space program!  There is now no excuse not to do this. 

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original posted article follows

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Artemis-2 used an elliptic departure orbit to greatly reduce the trans-lunar injection departure burn required of its Orion service module.  It used the SLS core stage and SRB’s to enter a surface-grazing ellipse with a very-low apogee,  and then used its ICPS second stage to raise the perigee of that ellipse out of the atmosphere with a small ICPS burn at apogee.  Upon reaching the new perigee,  it made another substantial ICPS burn to raise the apogee into a very extended departure ellipse.  It demonstrated what I propose here!

The ICPS had just enough propellant to be disposed-of safely,  after the Orion capsule and service module separated,  after getting onto the extended ellipse.  The Orion and its service module made one circuit about the extended ellipse,  preparing for the lunar trip.  Upon reaching perigee,  a modest service module burn put it onto the lunar trajectory,  with propellant reserves for course corrections both ways.  This lunar trajectory was a free-return loop around the moon,  with a direct re-entry back at Earth.  See Figure 1. 

Figure 1 – How Artemis-2 Used Elliptic Departure to Reduce the Lunar Departure Burn

Changes for a Permanent Capability

The perigee Artemis used was still too low at about 185 km altitude,  too close to the entry interface altitude of 140 km,  to be a permanent orbit.  A more permanent low circular orbit would about 300 km or more,  as shown in Figure 2.  An extended elliptical departure orbit was selected by this author that had more than 100:1 stronger pull of Earth gravity at apogee than lunar gravity,  plus an integer ratio of its period to the period of the low circular orbit!  Such would be stable for multiple circuits,  and the integer period ratio ensures that anything left behind in low circular will be there just as you arrive back at that perigee!  That makes rendezvous and docking much easier and faster. 

Figure 2 – Proposed Elliptic Departure/Arrival Orbit With Basing in Low Circular

It would be unattractive to base directly in the extended orbit,  when its perigee speed is very nearly Earth escape speed.  The most demanding portion of the ascent is surface-to-orbit,  for which low circular is demanding enough,  just reaching circular orbit speed with any significant payload.   It takes a much bigger launch vehicle to reach near escape speed with that same payload!  That’s more expensive,  and may require dedicated designs!

Instead,  one bases in the easier-to-reach low circular,  and uses a convenient stage or vehicle as a tug,  to take its payload craft from there to the departure ellipse.  From there,  the departure burn demanded of the payload craft is quite small,  especially for lunar missions.  The “tug” for Artemis-2 was the ICPS second stage of the SLS launch vehicle.  Using a tug to get onto a departure ellipse is still a substantial reduction of the departure burn,  even for faster-than-Hohmann Mars missions,  as the figure shows.

As we already now know,  reusability dramatically lowers mission costs!  To accomplish that as a tug-assisted elliptic departure,  we need a reusable stage or vehicle to use as our tug,  and we need to base it in low circular orbit.  Such is more easily reached for re-supply from the surface.  The tug can stay on the departure ellipse,  after releasing the payload craft to make its modest departure burn.  This tug then returns around that ellipse,  and burns unladen (for low propellant expenditure!) to return to low circular, where it is based.

Basing at a Space Station

That base ought to be an appropriate space station located in the low circular orbit at low inclination (to reduce plane change requirements for lunar or interplanetary missions).  What we need of that station is twofold:  (1) the means to assemble mission craft from docked-together components,  and (2) the means by which to fill (and refill) such craft,  and the reusable tug,  with appropriate propellants. 

We already know that we need manipulator arms and a framework to support them,  from the space shuttle and ISS experiences.  We will need the means by which to transfer both room temperature storable propellants,  and cryogenic propellants,  from tank to tank in zero-gee,  without spinning big structures or using orbit-changing ullage thrust!

Given that we can accomplish those things,  the advantages are enormous,  as detailed in Figure 3.  The notation V_near is also known as c₃,  the speed with respect to Earth needed at end of burn,  close to Earth,  to accomplish the lunar or interplanetary mission.

Figure 3 – Reusable Tug From Low Circular Greatly Reduces Final Departure Burns

That means that our space station is a frame to which multiple manipulator arms are affixed,  with arm operator cabins,  mounting and holding fixtures,  plus the support equipment for such crewed activities.  It must also have a multiplicity of appropriate propellant storage tanks,  the plumbing for propellant transport point-to-point on the station,  and docking facilities for the “supply tanker” transports bringing propellant supplies to the station.

The solution to the transfer of storable propellants in zero-gee has long been known:  bladder expulsion using gas pressure,  to squeeze the bladder within the tank walls.  There are no well-known cryogenic propellant solutions,  other than large structure spin or ullage thrust,  since there are no polymers with the enormous strain capability required (over 100% elongation),  for bladder service at cryogenic temperatures!  Neither structure-spin nor ullage thrust would be useful at a space station,  for any number of reasons.

There is,  however,  a not-well-known solution involving spin,  but only spinning the propellant inside stationary tanks!  One does this with vanes,  driven by electric motors.  If one does this by spinning half the propellant one way,  and half the other way,  then all the spin reactions and gyroscopic forces sum to zero at the tank mountings!  See Figure 4.  

Figure 4 – Proposed Cryogenic Delivery Tank Using Spin of Only the Propellant

Note also that the most practical tank design,  for (only) the payload of cryogenic propellants delivered to the station,  would be that same vane tank approach!  Storables could use bladder expulsion,  same as the main propulsion tanks.  Or they could use vane tanks,  but the bladder expulsion approach is both lighter weight and long-proven.

Now the form of the space station becomes clearer:  a long truss space frame,  along a portion of which are disposed a number of storable bladder tanks and cryogenic vane tanks,  and along another portion of which there are assembly arms,  holding fixtures,  and arm operator cabins.  Plumbing and power lines get routed within the frame.  You put the crew quarters and re-boost propulsion at one end,  and leave the other open,  for tanker vehicle docking,  and for future propellant capacity growth.  See Figure 5.  

Figure 5 – Proposed Refill and Assembly Space Station Concept

Benefits of This Approach Going to the Moon

For 1-way lunar landings,  your craft only needs 0.1 km/s to depart from the ellipse where the tug took it.  It needs around 0.9 to 1.1 km/s dV to enter low lunar orbit at some convenient inclination.  From there,  about 1.7-1.9 km/s dV will land it.  That’s a total of likely-under-3 km/s dV required of your 1-way lunar landing mission craft! 

At 330 s Isp for the lunar lander using storables,  that is about a single-stage mass ratio of 2.53,  or a propellant mass fraction of 60%.  If the lander has rough field capability,  its inert mass fraction might be near 15% at most.  Which leaves a payload fraction near at least 25%,  with storable propellants!  This is how you send cargo 1-way to build a base there!

Astonishing!

And it’s only about 6 km/s to return to the extended elliptic Earth orbit,  all the way from the surface of the moon,  unrefilled!   How to return such a vehicle unladen of payload,  but as a single stage,  is a topic for another time.  But it can be done,  even with storables!

What About Mars?

For a 2-way Mars orbit-to-orbit transport using low Mars orbit,  you only need around 1.5 km/s to depart the ellipse,  and about 2 km/s to enter low Mars orbit, 1-way.  That’s only 3.5 km/s required 1-way,  and so only 7 km/s required for the complete 2-way trip,  all unrefilled in a single stage!  Which would amount to a reusable chemically-powered orbit-to-orbit transport,  even with zero infrastructure at Mars!  That’s a single-stage mass ratio of about 8.7 at only 330 s Isp for storables,  or a propellant fraction of 89%.  If the inert fraction for a vacuum-only ship is 5%,  the payload fraction could be 6%!  And with only storable propellants in a single stage!

Astonishing!

To deliver payloads 1-way to Mars for direct entry and landing,  you are looking at about 1.5 km/s to depart,  plus some modest course corrections,  and the final landing burn of 1 to 1.5 km/s.  That’s at most 3 km/s dV demanded of the craft.  At 330 s Isp with storables,  that’s a mass ratio of only 2.53,  for a propellant fraction of about 60%.  If the lander,  which must survive entry as well as be configured for rough field landing,  has an inert fraction of 20%,  that is still near 20% payload fraction,  even with only storable propellants!  That could well be how to send cargo 1-way to build a base there!

Even more astonishing! 

See also Table 1.

What About Practical Tug Designs?

Now,  what do we need of the tug?  Assuming it might have to travel the extended ellipse for 2,  maybe 3,  circuits,  that’s about 10-12 days’ time in space.  If it uses cryogenic propellants for their high performance,  it needs a useful “stage life” without serious evaporative loss, of only some 10 or 12 days!  That rules out common bulkheads between LOX and LH2 tanks,  and it rules out bare single-wall tank shells exposed to sunlight in space!  But it does NOT rule out using LOX-LH2 for its highest performance!  You just need separate LOX and LH2 tanks,  which must be well-insulated externally,  plus with a shiny foil outer covering to shade them from sunlight heating.  That’s probably closer to a stage inert fraction of 10% than the usual vacuum stage inert of 5%.

We do not need months or years of “stage life”,  only a couple of weeks,  or so!  Which means we really do not need the weight and power-required penalties of cryocooler equipment!  We just need a minimal tank redesign from what is otherwise basically Centaur stage technology.  And later on,  we might need to scale it up for larger-mass mission craft!

See Figure 6.  

Figure 6 – Probable Tug Tank Construction for LOX-LH2 Propellants and 2-Week “Stage Life”

What About Arrival Versus Departure?

The dV requirements for arrival are almost exactly the same as departure,  but the timing requirements are different!  It only takes several seconds for the tug to undock from its payload craft and move several meters away.  That means for departures,  the tug can feasibly fire for reaching ellipse perigee speed,  undock,  and let the craft fire for its departure,  all in the one pass!  The tug then makes 1 circuit about the ellipse before burning at next perigee unladen, to get back into low circular,  going back to the station. 

Or,  if mission preparation time is needed by the payload craft,  both tug and craft can make one circuit about the departure ellipse,  with the craft departing,  and the tug burning to return to circular,  at the next perigee.

Arrival is different:  the tug cannot “be there” just as the craft arrives and burns into the ellipse perigee!  The craft needs to make a circuit about the ellipse,  before the tug can (1) rendezvous with it at its next ellipse perigee,  and (2) the tug must then burn to get onto the ellipse with the craft.  It then takes significant time to actually get docked together.  So the docked pair must make a second payload craft circuit about the ellipse,  before the tug can burn,  to put them into low circular,  and take them right to the station. 

And the laden vs unladen weight statements are more beneficial for departure,  and not as beneficial for arrivals.  You’d like the bigger propellant burn to be the first,  but that cannot happen when using a tug to assist arrival.  Arrival retrieved craft sizes must then inherently be smaller,  for a given tug design.  See Figure 7 for the estimated departure and arrival data for a tug sized to put a 50 metric ton craft onto the departure ellipse defined above,  as done with simple linked rocket equation calculations in a spreadsheet. 

Figure 7 --  Typical Tug Rough-Sizing and Performance by Spreadsheet

About Resupply Tankers

The resupply tanker vehicles sent to the station from the surface could be the upper stages or payload items of almost any existing or planned launch vehicle!  If sending up a storable propellant,  the payload can be a simple bladder-expulsion tank,  quite separate from the vehicle’s main propulsion propellant.  If sending up a cryogenic propellant,  the payload needs to be a vane tank,  also quite separate from the main propulsion propellant. 

Done that way,  all the tanker vehicle need do with the space station is dock,  and then hook up to the station plumbing,  to deliver its payload.  It goes without saying that reusable tanker vehicles would be much preferred for the long term.

The main caveat for tanker vehicles is that you do not want to use the same payload propellant tank (of either type) for shipping different propellants!  Propellants do get into the inherent slight porosity,  of even metal tank walls!  You need to be able to swap out the payload propellant tank,  if the same vehicle upper stage gets re-used to deliver other propellant species!  Period!  That nevertheless could be something quite convenient,  whether the payload tank is a vane tank,  or a bladder-expulsion tank!  See Figure 8. 

Figure 8 – Plausible Tanker Supply Vehicle Configurations

Prior Related Postings

As you can see from the list below,  I have been thinking about the various aspects of,  and problems feeding requirements into,  this tug-assisted elliptic departure space station scenario,  for some time.  The dates are shown in MM-DD-YYYY format.  There is an archive search tool on the left side of this page.  All that you need in order to use it are the year,  month,  and title.  I suggest that you jot down the ones you would like to see.  Click on the year,  then the month,  then the title if there were multiple postings that month.

This site also has a keyword search option,  in the sense that if you select a keyword,  then you see only those postings that are labeled with that keyword.  The current list of keywords is:  aerothermo,  airplanes,  asteroid defense,  bad computers,  bad government,  bad manners,  cactus-killing,  climate change,  current events,  education,  ethanol,  forensics,  fossil fuel,  fun stuff,  Gulf oil disaster,  guns,  health care reform,  idiocy in politics,  IR,  launch,  Mars,  Mideast threats,  North Korean rocket test,  nuclear crisis,  old cars,  pulsejet,  radiation,  ramjet,  space program,  spacesuit,  towed decoys,  trains,  treason;  the three highlighted are the ones most applicable to this article’s topic.

Only some of the more recent postings have received a search code for direct access.  That search code is something I have to give you.  It is the posting date in DDMMYYYY format.

Date............Keywords...............................Title

11-11-2025..space program.......................Where Should the New Space Stations Be Located?

10-16-2025..space program.......................Going Back to the Moon

7-26-2025....space program.......................Tank Design for Easy Cryogenic Transfers In Weightlessness

5-01-2025....space program.......................Vehicle Assembly and Refueling Facility in LEO

1-25-2025....space program.......................Initial Study for Tug Missions LEO to LLO

1-2-2025......launch, space program...........SpaceX’s ‘Starship’ As a Space Tug

12-1-2024....space program........................Tug-Assisted Arrivals and Departures

10-1-2024....space program........................Elliptic Capture

12-9-23........launch, space program............Overall Study Results: Propellant From Moon

5-1-22..........Mars, space program..............Investigation: “Big Ship” Propellant From Moon vs From Earth

4-2-22..........Mars, space program..............Earth-Mars Orbit-to-Orbit Transport Propulsion Studies

2-1-22..........space program........................A Concept for an On-Orbit Propellant Depot

8-18-21........launch, space program...........Propellant Ullage Problem and Solutions

3-23-21........space program........................Third Spacex Tanker Study

3-21-21........space program........................Second Spacex Tanker Study

3-17-21........space program........................Spacex Tanker Investigation

7-3-20..........launch, space program............Cis-Lunar Orbits and Requirements

11-21-19......Mars, space program...............Interplanetary Trajectories and Requirements

2-11-14........Mars, space prgm,  spacesuit..On-Orbit Repair and Assembly Facility

10-2-13........space program.........................Budget Moon Missions

8-2-12..........Mars, space program...............Velocity Requirements for Mars Orbit-Orbit Missions

8-2-11..........space program..........................End of an Era Need Not Be End of a Capability

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Search code (DDMMYYYY)                 14042026

Search keywords                                       space program

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On the Iran War

The following is an open letter to my federal representation:

I (and more than half the country) find the strategy for the Iran war flawed,  and that its start was illegal and unconstitutional.  The flawed strategy shows up,  in that it took a month of war,  for Trump’s goals to finally be stated coherently.  The illegal start was that this is not part of the war on terror that the war powers delegation act authorized,  this was an attack on another country (terrorist regime or not),  something only Congress can authorize!

 I do not know who our people are negotiating with,  but they are not the rulers of Iran anymore,  the Iranian Revolutionary Guard Corps (IRGC) is!  Being the terrorist army of violent extremists that they are,  they do not want to stop this war,  they would rather die than stop!  Trump and Netanyahu’s “regime change” killing the clerics and officials,  has not been successful at all,  they made it worse with only the IRGC now in charge.  And they have turned the Iranian people against us,  more so than against their own evil regime.

I rather doubt the ceasefire is going to hold,  especially with both Netanyahu and Iran already violating it,  on its first day.  After almost a month and a half of bombing,  there’s mostly just civilians and their infrastructure left to bomb,  and that’s a war crime under US and international law!  Hegseth’s ongoing purge of high officers who might oppose Trump,  has left in charge those more willing to commit that war crime for him!  And THAT is very dangerous indeed!

It is way past time for Congress to step in and set this right,  and finish this war correctly!  Which means I want you to get up and start taking actions to correct Trump’s egregious errors! 

First responders inspect a residential building hit in an earlier U.S.-Israeli strike in Tehran, Friday, March 27, 2026. (AP Photo/Vahid Salemi)

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Search code DDMMYYYY format:     09042026

Search keywords:  bad government, idiocy in politics

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About the Artemis-2 Mission

See the 4-13-2026 update at the end,  for some early photos of Artemis-2 post landing,  plus my preliminary assessment of heat shield damage from them.

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The following picture is a NASA illustration of the mission.  It does not give reliable distances,  nor does it show when the vehicle crosses,  or is outside,  the Van Allen radiation belts.  Outside the Van Allen belts,  the vehicle is also outside the Earth’s magnetic field,  and therefore at risk for radiation exposure,  should a solar flare’s coronal mass ejection (CME) event occur,  and happen to hit them.  

The next two figures together show the sequence of actual orbits being used,  after I looked around to various sources on line,  for real distance information.  I could not fit all the information into one figure.  That is why there are two figures that I hand-drew.

The NASA space launch system (SLS) rocket is a two-stage launch vehicle whose first stage core is assisted by two large solid boosters (SRB’s),  as was the space shuttle.  This first stage booster and its SRB’s essentially put the interim cryogenic propellant second stage (ICPS),  plus Orion capsule and its service module (SM),  onto a transfer ellipse with an apogee at the desired apogee altitude,  and a perigee that is pretty close to being a surface-grazing orbit.  There is no second stage ICPS burn to get onto that transfer ellipse. 

The ICPS makes a small burn at that desired apogee,  which raises the perigee up to the desired perigee altitude just outside the atmosphere,   but still quite low.  That apogee is actually within the inner of the two Van Allen radiation belts!    Without that perigee-raising burn at the initial apogee,  the Orion would inevitably re-enter the atmosphere on its first-pass return to perigee! 

The ICPS second stage then makes another,  larger burn at the newly-raised perigee point,  which raises the apogee to a very high elliptical orbit apogee,  actually outside both of the Van Allen belts.  They extend from about 1000 km,  to over 50,000 km,  altitudes.

The ICPS stage then makes one last “small” burn to put it into a “graveyard orbit”.  That graveyard orbit was not defined anywhere that I could access,  so I could not determine the magnitude of that burn.  All of this that I could find,  is in my first hand-drawn figure,  here: 

My second hand-drawn figure just below shows the transition from this high elliptic orbit to the actual transfer trajectory to the moon.  I used an ellipse from the low 185 km perigee at Earth to an apogee at the average distance of the moon to approximate the actual figure-8 trajectory,  for the purpose of estimating the departure delta-vee (dV) supplied by the Orion Service Module (SM) propulsion.  That dV is shown in that second hand-drawn figure. 

The actual transfer trajectory starts with this ellipse,  but gets distorted into the figure-8 shape by the gravity of the moon passing by,  making those details a 3-body problem one can only solve by finite-difference methods on the computer.  It loops around the moon in almost a polar orientation,  somewhere near 6500 km altitude,  behind the moon.  Then it free-returns to Earth.  A minor course correction from the SM is needed,  just before it gets jettisoned,  to ensure hitting Earth’s atmosphere at just the right angle,  for the free-return direct entry.

That direct re-entry returning from the moon is more demanding than one from low Earth orbit like the space shuttle endured,  because the speed at entry interface (about 140 km altitude) is higher,  at essentially the perigee speed of the transfer ellipse model shown above,  or right at 10.94 km/s,  maybe even 10.98 km/s.  That speed is very nearly Earth escape speed,  which is 11.18 km/s at the surface,  and 11.07 km/s at the entry interface altitude of 140 km.  Only about 90 to 130 m/s different!

The test history of the Orion capsule and its heat shield points to a disturbing possibility of heat shield damage possibly happening on this Artemis-2 mission!  The first Orion flew atop a Delta-IV launch vehicle uncrewed,  for a re-entry test,  among other things.  This was before Artemis,  and was named experimental flight test 1,  or EFT-1.  It had a heat shield manufactured of the same materials,  and built exactly the same way,  as the Apollo heat shields.  It did fine,  but that is an expensive,  labor-intensive manufacturing process. 

NASA used the same basic ablative material,  but manufactured in a completely different way,  for building two Orion capsules at once:  designated for the Artemis-1 and -2 missions.  They built the Artemis-2 heat shield before flight-testing the revised manufacturing process for it on the uncrewed Artemis-1,  launched by an SLS rocket.    They expected it to do fine,  but it did not,  unexpectedly shedding chunks of char,  leaving craters in the heat shield,  as the photo just below shows.  This is an official NASA photo of the Artemis-1 heat shield,  as recovered after the uncrewed Artemis-1 flight. 

The streaks on the heat shield point to a “source”,  to photo right on the heat shield,  which would be the stagnation point deliberately located off-center,  for generating a slight lift force during entry.  That force is small,  but it can “fine tune” the re-entry trajectory shape,  by rolling the vehicle to point that force where you want it.  NASA has done this for decades,  dating back to the Gemini flights of the mid 1960’s.  That is normal.

But,  if you look close,  you can see “craters”,  some large,  some small,  all over that heat shield,  where it shed chunks of the charred material from its surface!  That outcome was entirely unexpected,  and led to serious investigations at NASA,  for what to do about it. 

It must also be said that these damage craters,  as they were experienced on Artemis-1,  were not a risk for a fatal burn through!  The interior temperatures in the cabin did not vary from normal and expected,  despite the alarming damage!

NASA decided from its investigations that the two-heating-pulse “skip” entry,  that they used experimentally for Artemis-1,  was the culprit behind the chunk-shedding,  thinking that charring-material gas-evolution during the second heating pulse is what “blew” these chunks out,  leaving the craters behind.  They eliminated the skip during re-entry for Artemis-2,  getting it down to 1 heating pulse,  and decided to fly the same heat shield design,  already installed on the Artemis-2 capsule,  with a crew.

Others are not so sure about that damage mechanism.  We shall soon see. 

If Artemis-2 shows similar damage to Artemis-1,  then we (and NASA) will know that they were wrong about this damage mechanism!  Only a flight test can tell!

The problem here is shedding a second chunk from the bottom of one of the larger craters.  Should that happen,  the probability of a fatal burn through becomes very significant indeed!   Such would be a very low-probability event,  but that probability is not zero!

Last year,  I sent my concerns about this problem,  plus a low-cost means to stop the chunk-shedding,  even for a two-pulse skip re-entry,  to the entry heat protection group at NASA-Houston,  and again this year directly to the new Administrator,  at his DC office.  I do know that the Houston heat protection group thought I was right about my concern,  and about my proposed “fix”. 

So far,  NASA has not officially chosen to explore my alternative,  and,  as near as I can tell,  has already started construction of the Artemis-3 capsule with the same Artemis-1/Artemis-2 heat shield design.

Personally,  I would not ask a crew to fly with an unresolved risk that I already thought that I knew how to mitigate!  I think that is unethical!  Apparently,  there are high-level managers at NASA who disagree with that assessment. 

All I can say to them is this:  “there is nothing as expensive as a dead crew,  especially one dead from a bad management decision”.

What I sent the new Administrator is what was posted here on “exrocketman” 1 March 2026,  under the title “Ramjet Data Re:  Heat Shields”.  I knew a lot about ablative heat protection in ramjets and in solid rockets.  Some of that overlaps re-entry heat shields!

I did the 2-body ellipse orbital calculations illustrated here,  with a simple Excel spreadsheet “orbit basics spreadsheet.xlsx”,  which is part of the course materials posted online for free download,  via the Mars Society’s New Mars forums.  It simply automates the classic 2-body textbook equations.  While that New Mars forums site is down as of this writing,  you can still get that spreadsheet from me.  Just email me for it.  Anybody can do what I did here.

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search code DDMMYYYY format:   31032026

search keywords:  aerothermo, launch, radiation, space program

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Update 4-11-2026:  They made it back safe.  Of that I am glad.  I am eagerly awaiting photos of the condition of their heat shield.  Once I see such,  I will update here again.

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Update 4-13-2026:  Enough photos have surfaced of the Artemis-2 heat shield to make a preliminary assessment.  Starting with Figure A,  which is too blurry to make out very much,  other than the side facing the camera is not the side with the hatch and windows.  There is one tie-down pad zone that seems to have suffered damage (the whitish thing near the rim.  The other one near it did not,  being the slightly-dark spot below it and to its right,  also near the rim.  The other two tie-down pads are not in this view.   It is unclear whether the capsule is being hoisted,  or is still descending on its main chutes.

Figure A – View of Artemis-2 Orion Capsule

Figure B is a clearer version but of limited view dimension,  probably enhanced for clarity,  and apparently made from the blurry photo in Figure A,  intended to examine the whitish hold down pad damage.  It looks like a cavity in the heat shield,  burned out around where that hold-down pad got destroyed. 

To its left and lower left,  I see 4 dark,  small cavities that appear to be craters left behind from char chunks lost,  similar to those seen on the Artemis-1 heat shield.  They are just smaller and fewer than the hundred or so seen on Artemis-1.  Bear in mind that these few are on a small portion of the entire heat shield!  There could well be several,  to many, more!

Up on the lateral side of the capsule,  where the Avcoat protection was thinner,  I do see exposed metal in at least two places,  where the charred heat shielding was lost,  and that exposed metal looks distorted,  as if it did indeed see overheating during entry.

Figure B – Detail Near Hold-Down Pad Damage,  Enhanced for Clarity

Figure C is a photo taken during the post-splashdown extraction of the crew.   This is on the other side of the capsule,  where the windows and the hatch are located,  opposite side from that examined in Figures A and B.   The heat shield itself is hidden in this view.  I wish to point out that the lateral side of the capsule suffered very much less apparent damage than that seen in Figure B.  It would appear that the Figure B side saw more attached flow due to maneuvering angle-of-attack,  than the Figure C side,  thus charring the thinner lateral tiles through,  and so losing some of them.  

Figure C – Photo Taken During Crew Extraction,  Other Side of Capsule

The last figure,  Figure D below,  is NOT of an Orion capsule,  but an Apollo capsule,  specifically the one from Apollo-11.  In it,  one can very clearly see the reinforcing hex in the heat shield charred surface,  plus a variety of plugs that close openings in that heat shield. 

Similar to Orion,  there is more erosion near the “compression shear pad”  (same function as the “hold down pad” on Orion.  That kind of thing is apparently not unexpected. 

What I want to point out is the oxidizer dump plug,  which does not seem to have had any hex in it.  That location shows more erosion than the heat shield around it,  which emphasizes the positive reinforcement function served by the hex.  

Figure D – Minor Damages Seen on the Apollo-11 Heat Shield

My conclusions are three-fold: 

(1) Beneficial erosion reduction was indeed obtained by deleting the two-heating-pulse skip entry,  just as thought by NASA. 

(2) While erosion was reduced by deleting the skip,  there was still some char chunk-shedding going on,  which NASA hoped would not happen.  Plus,  there was some damage to the lateral side of the capsule. 

(3) The need for the reinforcing hex has been apparent since Apollo,  which did NOT do skip entries.

My recommendation still stands:  NASA,  put the reinforcing hex into your Avcoat tiles.  Use the extrusion press the way I suggested,  to do that without manual hand-gunning.  

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Update 4-20-2026:  I have been able to confirm that Figure A just above is indeed a CBS-taken photo of Artemis-2 still hanging from its chutes,  just immediately prior to splashdown.  That one is blurry,  and shows a white splotch near the rim that sparked a lot of questions and discussions.  I am inferring that CBS did digital enhancement to clear up the blurriness,  and zoom-in on the white splotch.  That would be Figure B just above.  It is quite clearly derived from the original in Figure A.  

What I see is essentially a lot of unexpected lateral-side damage in Figure B,  plus some cracking and about 4 small chunk-loss craters,  on a very small portion of the main bottom of the heat shield.  

What I have noticed is that NASA has not jumped to release photos taken all over that heat shield,  to quiet the fears sparked by the Artemis-1 damage.  The only ones we have seen show the lateral side with the windows and hatch,  where things look quite good.  None show the other lateral side where there was attached flow with higher heating while flying at angle-of-attack,  and none show the main heat shield on the bottom.  

My suspicion remains:  I think there was still some chunk-loss cratering,  and cracking,  on the main heat shield bottom,  just fewer craters,  and of smaller size.  So it really did do better with the revised trajectory,  just not good enough!

I have a new suspicion:  the damage on the back lateral side threatens a burn-through into the pressure cabin shell inside,  and was apparently quite unexpected.  Deleting the skip increased the peak heating load,  and that may have been a bit too much for that backside lateral wall,  where flow was attached at angle of attack,  causing a lot more "scrubbing action".  

If I am right,  neither suspicion threatens Artemis-3,  which is an Earth-orbital mission that sees far less heating upon re-entry.  But is does threaten Artemis-4,  the intended first landing mission to the moon.  It must return from the moon just like Artemis-2 just did.  

They really,  really need to fix this heat shield problem "right" before Artemis-4 flies!  Otherwise,  they are playing the odds with crews' lives.  Sooner or later,  that kills crews.  We have already seen it!

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Update 4-23-2026:  Here are some more photos that have surfaced since the Artemis-2 splashdown.  We have not yet seen anything “official” about Artemis-2’s heat shield condition and performance from NASA,  not as of this writing! 

Figure I below is supposed to be a photo taken of lateral-wall damage seen,  as pointed out by the mission commander Reid Wiseman.  I do not know exactly where this photo came from!  The tan color surface is stuff that is almost unheated.  The black stuff is material that saw enough heating to actually char.   There is some sort of localized crater in the lateral wall heat shield that he is pointing to. 

Note the black thing near his head in the image.  That is either the hatch opening or one of the recessed windows,  which would have been in the separated flow region on the opposite side of the capsule,  from where the stagnation point was biased toward the rim. 

Damages might well have been more severe on the other side of the capsule,  adjacent to the stagnation point that was offset toward the rim,  where heating would have been much higher.  We do not know:  no images from there have been released by NASA.

We have seen (so far) no images of the other side of the capsule at all,  where flow was attached and heating higher,   other than in the unofficial enhanced CBS image taken just before splashdown. 

Figure II below is a NASA photo taken of the heat shield by the Navy divers,  while the capsule was still in the water,  probably before the astronauts were recovered,  and certainly before the capsule itself was recovered!  It was posted on LinkedIn by NASA’s Steve Yoon.  You can tell by the surface streaks on the heatshield,  that the stagnation point was toward the top left in this image,  but actually outside the view of the camera,  from when the image was taken. 

The basic heat shield surface in this photo looks quite intact,  so far as this image discloses!  Which means NASA was correct in saying that the change in entry trajectory from skip to non-skip reduced base heat shield damage!  There is no doubt about that!

You can see an inverted-U-shape mark on the heat shield,  which is located around where one of the hold-down pads was located.  You must look very close to see the other one visible in this view,  to image top right.  Yet,  it is there.  Neither of these is the whitish “stain” and erosion cavity that was seen in the CBS image and its zoom-in enhancement.  Those other two pad locations are not visible in this image’s view.  They are on the other side.

Figure III below is something I found posted on LinkedIn,  purported to be damage on the lateral capsule wall of Artemis-2,  but I do not believe that claim!  If this is even a photo of an Orion at all,  it has to be from the initial flight test EFT-1,  before the Artemis program began!  I say that because it is quite evident that there was a reinforcing hex in the pictured lateral heat shield!  There was such on EFT-1,  since it was built the same way as Apollo. There was no such reinforcing hex installed on either the Artemis-1 or Artemis-2 Orion capsules!  Those were insulated with cast and machined Avcoat tiles that had no reinforcing hex.

We are not going to see the real evidence,  until NASA actually releases photos of the entire base heat shield,  and photos of the lateral wall heat shield that include the attached flow side,  opposite where the windows and the hatch were located.  Such may not be forthcoming for some time.  But the laws (and ethics) require that such evidence actually be released to the public!

What I suspect,  but have zero evidence for,  is that the base heat shield still showed some chunk-shedding craters over on the stagnation-near-rim side,  just less than Artemis-1.  The enhanced CBS image does indicate that to be the case.  If that is true,  it says that NASA still needs to restore the reinforcing hex to the bonded Avcoat tiles that they want to use.

The other thing I suspect,  but still have zero evidence for,  is that increasing the heating by deleting the skip,  pushed one lateral side of the capsule into a situation of inadequate insulation thickness.  That would be the side away from the windows,  where attached flow might really happen.  This would be based upon the burn-throughs and total char losses seen in the enhanced CBS pre-splashdown image.  And perhaps that Figure III image,  or maybe not. 

We cannot know any better than my “informed speculations”,  until the full suite of “official” images might be released. 

Figure I – Photo of Reid Wiseman Pointing to Damage on Lateral Wall of Artemis-2 Capsule

Figure II – Navy Diver Photo of Artemis-2 Heat Shield in the Water,  via Steven Yoon, NASA

Figure III – LinkedIn Posting Photo,  of Lateral-Wall Heating Damage to a Space Capsule

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