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.
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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