Kudos to SpaceX for Flight 13

After a last-second automatic abort on the 16^th,  and a bad weather delay Thursday,  SpaceX was able to launch its Starship/Superheavy vehicle successfully on Friday (24 July 2026).  This was Flight 13,  called by some “Lucky 13”,  the second flight for the new version 3 vehicle,  which is bigger,  and uses a new,  upgraded version of the rocket engines. 

Unlike the previous test flight,  nearly everything about this flight appeared to work perfectly.  Kudos to SpaceX for a job very well done!  Very well indeed!

The upper stage Starship vehicle flew perfectly,  with all its engines working,  and re-lighting just fine.  It deployed some 20 examples of the latest version Starlink satellites,  performed the in-space re-light of an engine (which proved not possible on the previous Flight 12),  re-entered the lower atmosphere without any problems over the Indian Ocean near Australia,  descended to low altitude,  re-ignited its engines,  made its flip to nose-up,  and a very soft splashdown landing in the sea.  

It even survived toppling over into the water without exploding!  So,  it might even be towed to Australia for direct inspection,  a real bonus for evaluating the heat shield performance!

The Superheavy booster flew most of its intended flight plan quite well,  excepting the final touchdown burn into the sea,  just off the SpaceX launch site near Brownsville,  Texas.  The “hot staging” seemed to work fine,  with 3 of the 33 booster engines keeping the propellants settled in free fall,  inside the tanks during its flip-around,  so that 13 engines could be ignited to perform the boost-back burn. 

This failed on Flight 12,  due to multiple lost engines,  but it worked just fine this time on “Lucky 13”.  There was something wrong about that Flight 12 flip-around maneuver,  which may have in some way induced those failures-to-ignite.  It occurred in the wrong plane,  for one thing,  according to reports.

The coasting booster then arced over,  back toward the launch site,  and fell into the lower atmosphere,  stabilized tail-first by its grid fins,  toward a low altitude touch-down burn event.  This last burn is where the trouble happened!  Instead of igniting the intended 13 engines to slow down its very supersonic speed,  a handful of engines failed to ignite!  The booster descended stably,  but did not decelerate adequately on too-few engines,  and so it struck the sea at something still near the speed of sound!  That intended soft booster splashdown was the only failed objective in this entire test flight,  as near as I can tell.

We did not see quite so many engine ignition failures in the version 2 Starship/Superheavy test flights that happened before Flight 12 (which was the very first version-3 flight).  Something is going on with version 3 that will have to be found and fixed,  before SpaceX can risk attempting to catch the booster with the arms on the launch tower.  They did that catch  successfully more than once,  with version 2 in the earlier flights.  A truly spectacular thing to watch,  too!

But a failure to decelerate to a slow hover would end up destroying both the booster and the launch-and-catch tower!  And right now,  there is only one of those facilities in existence that is compatible with the version 3 booster!  They have to solve this engine ignition problem with version 3,  before they risk attempting another booster catch.

The upper stage Starship is another story entirely,  now after its complete success during “Lucky 13”,  and near-complete success during Flight 12!  It would be a reasonable bet that they might attempt a ship catch,  on the next flight.  Musk himself even hints at that,  on social media. If they do attempt this feat,  that would be the very first recovery of a Starship that has been into space at near-orbital speeds.  Such would be invaluable,  both for heat shield evaluations and improvements,  and for eventually demonstrating the readiness of this vehicle to actually do real orbital missions.

There will be tension within SpaceX over this.  Some will want to see another Starship upper stage water landing success,  before risking the catch facility.  Others will want to forge ahead as fast as possible.  That is normal.  We will soon hear what they decide to do next. 

I wish them success,  especially since doing spaceflight successfully,  really is a very,  very hard thing for any outfit to do.  And once again,  kudos to SpaceX for a flight test that was almost entirely successful! 

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Note:  this same article was submitted to the Waco “Tribune-Herald” 7-25-2026 as a possible column for the op-ed page.  I am on their board of contributors. 

Meanwhile,  here are the search items for this article:

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Update 8-1-2026:  The Waco "Tribune-Herald" ran the article as a column on the op-ed page in the Saturday paper today (8-1-26).  They ran it pretty much as submitted,  and posted here.  They did not use the illustration that I drew of the booster flip event.  

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Call to Action

I received this image from HOTNIR,  which is the Heart of Texas Network for Immigrant Rights,  located in and around Waco,  Texas.  The image is self-explanatory:  we so very clearly have government officers behaving like a secret police force intended to mistreat and “disappear” the demonized few (immigrants),  and intimidate the rest of us into submission.  

I am sick and tired of seeing government agencies weaponized and abused into the tools of dictatorship!  On this,  our nation’s birthday,  I call upon all Americans to come together,  and rise up to depose our wannabee-king and all his enabling minions!  Your last chance to do this peacefully at the ballot box is this November! 

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Quick First-Cut at Orion Entries

I cannot model skip trajectories or trajectories that change angle during entry.  I can only model straight-in entries at constant angle below local horizontal.  I am using the old H. Julian Allen scale height entry model (ref. 1),  updated to include plasma radiation heating (ref. 3). 

Sources on-line differ in the exact numbers given for most vehicles,  Apollo and Orion included.   I did the best I could,  with what I could find.  Figure 1 was a convenient comparison.  Because the shapes are similar,  I presumed the same Rn/D ratio,  and the same blockage-area-basis hypersonic drag coefficient C_D (ref. 4).  Assuming the load of samples could be heavier,  I increased the mass a bit to increase the ballistic coefficient β slightly. 

Figure 1 – Source Data for Entry Models of Apollo and Orion

I simply presumed speed at entry interface of 10.95 km/s returning from the moon.  Entry angles were 2.0 degrees (Figure 2),  1.5 degrees (Figure 3),  1.0 degrees (Figure 4),  and 0.7 degrees (Figure 5).  Entry interface altitude was 140 km,  per the Justus & Braun scale height model of Earth’s “typical” atmosphere (ref. 2).  (Apollo modeled at 2 degrees.)

Figure 2 – Estimate for Orion at 2.0 Degrees Below Local Horizontal

Figure 3 – Estimate for Orion at 1.5 Degrees Below Local Horizontal

Figure 4 – Estimate for Orion at 1.0 Degrees Below Local Horizontal

Figure 5 – Estimate for Orion at 0.7 Degrees Below Local Horizontal

I accumulated these annotated data from the figures,  figured some heat shield pressures with them,  and cross-plotted the results as a function of the constant average entry angle.  Peak deceleration gees vs entry angle is plotted in Figure 6. 

Peak decelerating force is simply entry mass times peak gees times the standard acceleration of gravity at Earth.   Average pressure on the heat shield is that decelerating force divided by the capsule blockage area.  The peak pressure at the stagnation point is about 4/3 of the average pressure.  This peak pressure estimate is plotted in Figure 7,  along with a reported limiting pressure for the kind of Avcoat used on Apollo. 

The entry spreadsheet model figures both convective and plasma radiation heating rates per unit area at stagnation,  and it totals them,  for the peak stagnation total heating.  These total heating values are plotted in Figure 8,  along with one reported max heating limit value,  for the kind of Avcoat ablative heat shield that was used on Apollo. 

That Apollo material was hand-gunned into a reinforcing fiberglass hex already bonded to the capsule exterior.  So,  too,  was the heat shield used on the original Orion flight test (EFT-1).  The version of Avcoat used on the Orion for both Artemis-1 and Artemis-2,  was bonded tiles machined from blocks of cast Avcoat,  without any reinforcing hex.

Figure 6 – Peak Deceleration Gees Vs Constant Entry Angle (Apollo 10-11 near 2)

Figure 7 – Peak Pressure on Heat Shield Vs. Constant Entry Angle (Apollo ~ 0.56 near 2)

Figure 8 – Peak Stagnation Heating Vs. Constant Entry Angle (Apollo ~ 380 near 2)

Orion flies entry at some modest angle of attack,  to generate  lift force perpendicular to the oncoming wind.  This is for fine trajectory shaping and control,  accomplished by rolling the capsule to point the lift vector in the desired direction.  This has been done since Gemini in the mid-1960’s.  It was done with Apollo.

This shifts the stagnation point on the heat shield away from center,  toward the rim on one side,  which reduces the angle seen between the wall and the separated flow boundary and plasma sheath coming off the rim of the heat shield.  That can lead to attached flow with higher “scrubbing action” and heating rate,  on a swatch of the capsule lateral wall,  on that side.  The crew’s windows must be on the other side,  where such extra heating cannot occur because of the larger separation angle forcing local separation. 

One cannot just rescale the total stagnation heating to different locations around the capsule,  because convection and radiation rescale differently. 

For regions away from stagnation but with attached flow,  my ballpark estimate for convection is stagnation/3,  and my ballpark estimate for the separated wake region is stagnation/10. 

For regions away from stagnation but with attached flow,  my ballpark estimate for plasma radiation is stagnation (because the plasma sheath is still quite nearby),  and my ballpark estimate for separated wakes is stagnation/3 (because the plasma sheath is more remote).  Some have claimed that radiation heating has proven higher than initially expected,  in those separated wake zones.  

I roughed out the numbers from the annotated data in the spreadsheet plots,  and plotted them vs entry angle in Figure 9.  I then depicted them around the capsule in Figure 10.  Bear in mind that these illustrations are not to scale,  and the angle of attack (AOA) is shown somewhat exaggerated. 

It should be quite clear that lateral capsule surfaces with attached flow will need more heat shield thickness than regions that always stay separated.  Heating in regions with attached flow is really not very far below that at stagnation,  since the radiation heating is still quite near stagnation values.

Figure 9 --  Estimated Heating Rates for Stagnation,  Attached, and Separated Flow

Figure 10 – Rough Ballpark Heating Distributions Around the Capsule

We cannot make conclusions about the adequacy of the Artemis-2 heat shield relative to the damages seen on the Artemis-1 heat shield,  from data like this!  That will take good photographs of both capsules side-by-side.  Such have not yet been released as of this writing.  But one place to look,  besides the main base heat shield,  is very clearly the lateral side away from the windows,  where attached flow is likely to occur while flying entry at angle of attack.  Too little thickness there risks a burn-through,  at one or another level. 

I cannot model the difference between the “skip trajectory” of Artemis-1,  and the “non-skip” or “reduced-skip” used by Artemis-2,  except to say that the “non-skip” trajectory is similar to steeper angles below horizontal,  as modeled here.  The heating numbers are higher,  the steeper the angle.  And that is as true for attached flow on a lateral side,  as it is on the main base heat shield. 

I did hear it claimed during the televised entry coverage that the crew of Artemis-2 experienced something like only 4 or 5 peak gees during entry.  That is unlike the 10-11 gees experienced by Apollo crews returning from the moon.  So,  it is likely that the “best” model among those shown here,  for the Artemis-2 entry,  would lie somewhere near my constant 1 degree model.   That corresponds to heating rates well below that of Apollo,  suggesting in turn that the max heating rate limit for the unreinforced Avcoat is below that of Apollo’s.

References as indicated above:

#1.  H. J. Allen and A. J. Eggers,  “A Study of the Motion and Aerodynamic Heating of Ballistic Missiles Entering the Earth’s Atmosphere at High Supersonic Speeds”,  NACA Technical Report 1381,  44^th Annual Report of the NACA 1958,  Washington D.C. 1959. (unclassified) – this has the scale height atmosphere model and the relationship between altitude and velocity,  plus the convective stagnation heating correlation.

#2.  C. G. Justus and R. D. Braun,  “Atmospheric Environments for Entry,  Descent,  and Landing”,  MSFC-198,  June,  2007.  – this has the same Allen and Eggers entry model,  and scale height atmosphere model as Allen and Eggers,  but goes beyond just Earth.  Atmospheres for Mars,  Titan,  and Venus were obtained from here.

#3.  SAE,  “Aerospace Applied Thermodynamics Manual”,  1969.  (hardbound) – this had a simple plasma radiation heating model that was modified and added to the spreadsheet embodying the Allen and Eggers technique.

#4.  Sighard Hoerner,  “Fluid Dynamic Drag”,  self-published by the author,  1965.  – drag data for many shapes into the low hypersonic range are in this reference.

Note that the spreadsheet used for this study was described in an earlier posting to this site: “Entry By Hand”, 1 May 2026.  A related study comparing probes at both Earth and Mars entry conditions,  plus Apollo at Earth,  was posted as “Entry Study”,  1 June 2026. 

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