New Mars Forums

The density vs altitude trend at Mars is similar in shape but not values.  Max heating and max deceleration simply occur at lower altitudes on Mars,  closer to the surface.  As it turns out,  the plasma radiation heating gets quite significant at "typical" Mars entry conditions of around 7 km/s off the interplanetary trajectory,  when it is rather insignificant at similar-in-speed 8 km/s Earth orbital entry conditions.  It does become quite significant at 11 km/s Earth entry conditions,  worse even faster.

Blunter heat shields of the spherical-segment type minimize convective heating but maximize radiation heating.  That is why the historical capsules have a spherical segment heat shield shape.  Those typically have "nose radius" (radius of curvature) equal to or greater than the heat shield diameter.  Mercury,  Gemini,  Apollo,  and now Orion and Dragon,  all have that heat shield shape.   Blockage area hypersonic drag coefficients are all near 1.4 for those shapes.  Bear in mind that larger nose radii reduce convective heating but enhance radiation heating!  That’s just the correlation equations for stagnation convective heating and stagnation radiation heating.  The reverse is also true:  smaller nose radii enhance convective heating,  but reduce radiation heating.  Those effects are fairly dramatic,  but not as dramatic as speed-at-entry.

With the higher plasma radiation contributions,  entry at Mars was different.  All the heat shields used there so far had blunted-cone shapes.  These were pretty near 20 degrees off flat plates (70 degree half-angle cones),  and the nose radii were somewhere in the vicinity of ~10% of the diameter.  The backshells were heat protected,  too,  with an entry vehicle shape quite squat compared to the historical Earth entry capsules.  These shapes have higher hypersonic blockage drag coefficients,  nearer 1.7,  which leads to lower ballistic coefficients,  also necessary at Mars with its thin atmosphere.

As Earth entry speeds exceed 11 km/s,  it may prove necessary to modify the heat shield shape to something resembling more a blunted cone than a spherical segment,  because of the quite-significant radiation heating effects.  Such entries would be free returns coming back from Mars,  direct off the interplanetary trajectory.  Such would be a bit over 12.2 km/s off a Hohmann transfer ellipse,  and nearer 13 km/s for a faster transfer ellipse with a 1-way travel time in the 4-7 month range.

Bear in mind that higher ballistic coefficient correlates with higher peak heating and higher peak gees,  experienced deeper down in the atmosphere,  so that the end-of-hypersonics (at Mach 3) occurs at a lower altitude.  On Earth,  that altitude is still way up high,  above 30 km altitude,  even for ballistic coefficients as high as 500 kg./sq.m.  On Mars,  it is quite low,  ~10 km even at only 200 kg/sq.m.  All the probes so far have been at or under 100 kg/sq.m.

I think that the LOFTID engineers will tell you the same things I did.  For only 8 km/s,  all they needed was a spherical-segment blunt shape.  Even at the lower drag coefficient,  with its higher ballistic coefficient,  that was enough to get them an order-of-magnitude lower ballistic coefficient with the inflatable,  than that of historical capsules.  Plasma radiation heating was not a significant problem at that speed.  But wake zone heating was,  so they had to put insulating blankets over their reinforced-polymer inflatables. 

Thos engineers get to run codes for that and get more precise answers.  I just run the ancient by-hand models in a spreadsheet.  My answers are not as precise,  but they are well within the ballpark,  and usually fairly close to actual test numbers.  Just as they were in the 1950's and 1960’s. 

My "spread" of the stagnation convective and radiation heating rates,  around the vehicle away from stagnation,  is rather crude,  but it is still within the ballpark somewhere.  Even at very low heating with very low ballistic coefficients,  cooling by thermal re-radiation requires surface emission temperatures well above what any reinforced polymer inflatable structures could possibly withstand.  All those things fail utterly and start charring between only 250 and 350 F.  Of course,  the inflatable would burst and deflate if that happened!

The re-radiation temperatures are very sensitive to thermal emissivity estimates,  even at those low heating rates,  because of the temperature-to-fourth-power dependence of the thermal radiation equation.  Making those wake zone structures "black" helps,  but probably not enough.  Certainly not enough for higher-than-8 km/s entry speeds at Earth,  and also certainly not for Mars,  where radiation is a bigger contribution,  because the wake zone radiation heating is a bigger fraction of stagnation radiation heating,  than that of convection.  For wake zones,  I use stagnation/3 for radiation,  and stagnation/10 for convection.  The “stagnation values” are different for convection and radiation.  They are not the totals!  You spread them as separate items.  But at any one location,  you sum the two up to a total,  to do the thermal re-radiation calculation.

GW

If you look at about 700 F as quoted for the inflatable membrane interface,  plus about 590 F drop through its insulation layer,  also as quoted (1000 C),  you get a backside surface temperature for the inflatable heat shield of about 1290 F.  Which is NOT VERY FAR AT ALL from what I estimated with my crude little by-hand entry modeling technique (about 1300 F)!

Which proves my other point that even wake zone backside surfaces need heat protection,  at entry speeds as low as 8 km/s,  if they are not 2000-2500 F-capable superalloys. 

The front surface quotation was 1400-1600 C (2550 F to 2910 F) is above my little estimate,  as was the 9 gees.  Silicon Carbide is a refractory re-radiator,  not an ablative,  so my surmise that their entry averaged a little steeper than the 2 degrees I analyzed,  is therefore confirmed indeed!  I could probably very closely duplicate their numbers at something in the 3 to 5 degree average angle range,  but I am NOT going to waste my time doing it!   

My main point is that going to very low ballistic coefficients,  which LOFTID successfully did,  does NOT relieve you of heat protection requirements during entry.  Their 25 kg/sq.m corresponded to around 40 W/sq.cm,  not far from what I estimated,  and far below the ~200-ish W/sq.cm you get at ballistic coefficients nearer 300 kg/sq.cm. 

But the re-radiation temperatures are still quite high!  That's the temperature to 4th power variation of the Boltzmann equation for thermal radiation, inverted to yield a temperature that produces a certain re-radiation rate.  It means for a factor-12 reduction in heating rate to be re-radiated,  your surface re-radiation absolute temperature is crudely only 12^0.25 = 1.86 times smaller!

There is simply no way with any materials technologies that we have,  including these new inflatable heat shields,  that we are going to be able to delete the need for heat protection,  by going to a very low ballistic coefficient.  Thermal re-radiation physics and the well-established correlations for entry heating,  simply will NOT allow that!

GW

The consensus of opinions and Musk postings "out there" is that they may fly Flight 14 late this month (August).  It would appear to be a full orbital mission,  with an attempted tower ship-catch back at Starbase in south Texas,  if the FAA approves the flight plan. 

As I thought,  they are not going to risk the tower with a booster-catch,  this time around.  They need a soft landing (or preferably two) of the booster before they risk that,  and that means they must solve the "booster Raptor relight problem",  whatever that really is. 

I only suspect it may have something to do with the roughly 2000 F air caught in the engine bay and heat-soaking the engines,  during booster entry at approximately Mach 5.  It would potentially counter the cryogenic cooldown they seem to need,  in order to light them.  But I may be wrong about that.  Just note that nobody has said anything in public about that problem.  They may,  or may not,  understand it yet.

The ship will be flying over populated areas in Mexico and south Texas during its entry,  descent,  and landing after an orbital flight,  which will be of great concern to the FAA in the flight plan approval process.  Not to mention the interruption of a lot of airliner flights to and from Mexico.  The post-hypersonics "swoosh" maneuver during descent is how they will attempt to fly around Brownsville,  instead of over it,  as they approach Starbase.

If Shotwell is successful at getting Musk to keep his loud mouth zipped about this flight,  they may actually get this plan approved.  We will soon see.  He does have a demonstrated track record of pissing off the FAA,  going all the way back to first very-sub-orbital ship-only flights.  She knows that.

GW

Harold:

No,  I never forget.  That was a bad joint design to begin with.  I think we have beat that to death before.  You DO NOT DO 2 O-RINGS in a solid joint!  You do only 1!  If it holds at 20-30 psia,  it will hold at 2000-plus psi. 

Ask me sometime about the Tarter-Terrier gas generator canister.  Same stupid problem.  Just embodied differently. 

GW

Manned Orbiting Laboratory (or MOL) was a USAF program aimed at a manned spy satellite capability.  It was to be launched by a Titan-3 upgrade to the then-existing Titan-2 launcher,  and it consisted of a modest house trailer-sized laboratory and a Gemini-B capsule. 

Gemini-B was a revised design with shorter life support and electric power duration,  and with the seats displaced angularly more at the heads,  so that a hatchway through the heat shield would be accessible by the two astronauts.  This hatchway led directly into the laboratory,  so that the crew could go into it without depressurizing and going outside.  The capsule had a month of life powered-down,  but only a few days powered-up. 

The mission was a month long,  and the lab not reused or recovered.  This was the first use of telescopic cameras with folded optical paths,  to enable car license plates to be read from orbit.  (They can recognize human faces from orbit now.)  Most of the lab volume was the camera telescope.  It was fairly crowded in there for a crew of 2,  but not as crowded as the Gemini itself. 

This program began before 1963,  along with the X-20 Dyna-Soar,  also a USAF program.  In 1966,  the USAF manned space program was terminated in favor of the NASA manned space program.  That killed X-20,  and MOL was then in limbo,  pending unmanned spy satellite development progress.  Gemini-B flew successfully with the hatchway cut through its heat shield,  in 1969,  I think it was.  But by that time,  the unmanned spy satellite technologies were ready to deploy and proved to be superior in terms of both cost and lifetime on-orbit.  That was the end of MOL.  Most of the USAF astronauts went to NASA,  and became shuttle astronauts,  some by way of the X-15.

The Russians had been doing something similar,  called Almaz.  If memory serves,  it flew 3 times under the name Salyut as the initial Russian space stations.  But the Russians found most of the same answers we did:  unmanned spy satellites had longer on-orbit life and lower costs.  Many of theirs were nuclear-powered.  Cosmos 954 that crashed in Canada was one of those.

GW

For RG Clark re: post 2355 --

Bob,  I worked out a model from the data in your posting about the LOFTID test,  and got something fairly representative of their blunted cone shape and size and ballistic coefficient.  And yes,  the shape makes a difference:  blunted cones are what has been used at Mars for decades to minimize plasma radiation heating,  and those drag coefficients are near 1.7.  Spherical radius-type blunt heat shields have been used at Earth where plasma radiation is not as severe,  and have lower drag coefficients,  nearer 1.4.  For the same size and mass,  that is a big difference in ballistic coefficient,  and thus in peak heating,  as well as how deep into the atmosphere it penetrates before slowing.

I got numbers more or less similar to those in the article you linked and quoted,  except I got nearer only 6 gees at peak deceleration,  and that article indicates something over 9 gees.  I used an average 2 degrees below horizontal,  they may have come in a bit steeper than that,  although steeper does drive up heating,  too.

The temperature I got for a re-radiating heat shield with no ablation is close to what they seem to say in the article.  I must be somewhere in the ballpark with my little model.  However,  when I spread heating rates around the vehicle,  I got surface temperatures seemingly too high for inflatable fabrics to survive.  I think they said something about insulating the inflatables with ceramic fabrics.  The problem with equilibrium temperature estimates is that they are extraordinarily affected by the thermal emissivity that you use.

My point:  the picture a lot of people share,  that you only need protection on the windward side,  is false.  Even the wake zone sees the same super-high effective plasma temperatures that the windward side sees right behind the bow shock.  It's just that the effective heat transfer coefficients are drastically lower.  That is why Apollo's (and Orion's) lateral sides also had ablative coatings.  The higher speed drove that.  Bare superalloy skins survived on Mercury and Gemini,  but only because entry speeds were lower.

The inflatables at the lower speeds will need wake zone heat protection,  because the inflatable materials have lower max service temperatures than superalloys.  By far.

GW

Update 2026/08/07 ... GW Johnson asked that this be added to post #2358

Rob:

The short answer to your question "could the Starship heat shield survive a return from Mars?",  is "NO".

GW

The technology and materials and procedures for a heat shield that can make hundreds of LEO entries without significant repair simply DOES NOT EXIST YET! 

NO ONE has such a thing,  not NASA,  not SpaceX,  not Boeing,  not Lockheed-Martin,  not Northrup Grumman,  not Blue Origin,  not anyone!  To criticize SpaceX over that,  is rather pointless,  until and unless such a better technology is actually ready to apply. 

The basic science and perhaps some exotic material candidates do exist,  but that "state of the art" is decades away from being ready-to-apply,  if it ever gets there at all.  That little ugly fact of life is too often forgotten by far too many!

What SpaceX does have,  is a heat shield that might fly several times,  with some repairs made in between,  much like we saw with Shuttle,  and much like what happens with X-37B.  Their backup underlayer is a good thing,  something Shuttle and X-37B do/did not have.  It appears to work.  They have not flown it twice yet,  simply because they have yet to recover a Starship for potential re-flight. 

Once they start recovering Starships,  we will see their heat shield re-flown,  and we will (finally) see how much repair is required between flights.  If that is less effort than what was required for Shuttle,  then yes,  they have made a big improvement!  So what if it is not yet "perfect"?  Whatever "perfect" really is.

The complainers anger me with their lack of proper perspective,  and their lack of acknowledgement of the VAST difference between what might be possible,  and what is actually practical to do.

GW

I was pointing out that if either the booster or the upper stage fails to come to a controlled hover,  in the correct position relative to the tower,  then the collision will destroy both the stage and the tower,  even if the propellants do not explode. 

But the propellants will explode because the stage will break open in the collision and spill them,  and so create even more destruction.  That's true even if the stages are nearly empty upon arrival.  There will be the equivalent of multiple sticks of dynamite,  in even just a few tons of fuel and LOX,  which all by itself is more than enough to destroy the facility.  Lots of propellant,  only some propellant,  the facility is destroyed,  either way.  The only difference is how far the debris flies. 

A full load explosion like Blue Origin experienced is more or less equivalent to a small atomic bomb,  but once you are at the few sticks of dynamite level,  it no longer makes much difference at all to the destructive effect:  facility destroyed,  either way.  And that costs a lot of time and money to replace!

I know of one rocket company that destroyed a USAF test stand facility decades ago,  with a mis-designed solid rocket booster.  It blew up during a test at the USAF facility.  They ended up having to take a $billion write-off in one year.  They were never the same financially after that.  That company went belly-up decades ago.  Traceable to that mistake.

GW

Just finished watching "Lucky 13" live on the SpaceX website.  Launch good.  Hot stage good.  6 engines on Starship,  booster did its flip and boost-back OK.  Booster lit engines hanging on grid fins at entry,  but lost a couple of them during the landing burn,  and hit "hard" in the water. Upper stage Starship ascent OK.  Deploy 20 v.3 Starlinks OK.  Relight 1 SL Raptor in space with 14 s burn OK.  Entry looked good.  Descent looked good,  including maneuvers.  3 SL Raptors lit and executed flip,  then final descent as planned on 3,  then 2, then 1 engine,  to a very gentle splashdown.  Vehicle fell over onto its dorsal side,  but did not explode.  Venting methane burned,  but no explosions.  Fantastic photography of heat shield,  as it floated nice and high in the water on its back.  Except for booster splashdown,  test flight looked to be just about perfect!

Here is an image created by GW Johnson to illustrate the text above:

(th) 2026/07/25

GW

From AIAA’s “Daily Launch” email newsletter for 23 July 2026:

SPACE
After nearly 30 years, NASA realized this near-Earth asteroid is actually a comet. The discovery may help us defend the planet some day

Astronomers discovered that a near-Earth asteroid is actually a comet that could help scientists identify more hidden comets and improve planetary defense.

-----   

My take on it:

Somewhere on these forums,  some years ago,  I opined that there really was no distinction between type C asteroids and comet cores,  except that “comets” have ices mixed in,  and the “asteroids” do not. 

Nearly everything we have seen of small bodies beyond Saturn would classify as “comets”,  and most things closer in have lost their ices to evaporation over geologic time,  and so we classify them as “asteroids”. 

There is the likelihood that ice content varied with formation distance from the sun,  yes,  but anything small that has remained in the same place over geologic time would have more or less found its equilibrium ice content.  Inward of Saturn,  that is likely pretty near zero.

The asteroid described in the article is a near-Earth asteroid,  but it very most likely has not been in that orbit for eons of time.  If it had,  it would be dry of ices. 

On the "asteroid defense" aspect:

If a body can be determined to have ices,  then a beam of laser or maser energy could be used to heat those ices under a local surface spot,  creating gas jets that might change the orbit over time.  Exactly how that might be done is not known at this time. 

This does require long warning time,  something we do not have for most perturbed "comets" coming in by surprise from somewhere out near the Oort cloud or Kuiper belt.

GW

The biggest problem that I have with all this is the scarcity of CO in the Martian atmosphere.  It is but a trace. 

For LOX-LCH4 propellants,  there is abundant CO2 in the atmosphere (above 90%),  and there may or may not be abundant water at any given site.  The Sabatier reaction allows you to make LCH4 from these,  assuming that you electrolyze the water resource for its hydrogen.

For LOX-LH2 propellants,  all you have to do is electrolyze the water.  You need nothing from the very thin atmosphere.  You just need to clean up the water.  Which you must do anyway,  for LOX-LCH4 propellants,  as well. 

There is a metal catalyst for making CO and O2 out of the far more abundant CO2,  but you must supply considerable energy to make it work,  just like with the electrolysis of water for the other two choices!  Given the lower Isp of CO + O2 combustion,  I do not see the point of that choice!  It's no better than solid propellants. The other two are far-higher Isp return.

GW