Debt Ceiling Negotiations? BS For Sure!

 This is an excerpt from the 14^th amendment to the US Constitution.  It deals with the public debt. 

 

Section 4. The validity of the public debt of the United States, authorized by law, including debts incurred for payment of pensions and bounties for services in suppressing insurrection or rebellion, shall not be questioned. But neither the United States nor any State shall assume or pay any debt or obligation incurred in aid of insurrection or rebellion against the United States, or any claim for the loss or emancipation of any slave; but all such debts, obligations and claims shall be held illegal and void.

 

There is a lot of legal discussion and court opinions about the intended meaning of this section of the amendment.  But look very closely at the yellow highlighted portion.  Note the phrase “for payment of pensions”.  That includes your Social Security and your military retirement pay.  That’s what those are:  pensions.  It says these will be paid,  yet the negotiations are about cutting them,  among other things. 

So you,  the public out there,  had better be paying attention to this.

Note also how the sentence begins,  and ends:  “The validity of the public debt of the United States …..shall not be questioned”.  The clear implication of that is “we will pay our debts”.  Period.  No question about that. 

Note also the phrase “authorized by law”,  which means so very clearly that these public debts were previously agreed-to,  and appropriated by,  Congress,  the legislative body that makes law. 

I’m no lawyer,  just a layman with a logically-trained mind.   To me,  what the wording of that amendment says is that the time to argue about spending cuts is before you vote to spend it,  not after.  The amendment says we will pay our legally-incurred bills,  and according to the Constitution,  it is Congress that does the paying.

To me,  this also says the Congressional act instituting a “debt ceiling” is unconstitutional.  Arguing over whether to pay an authorized debt,  is not in accord with the amendment.  Period.

The debt ceiling is merely a dodge for the minority party to get its way with regard to spending cuts,  when it did not have the votes to pass those cuts at the time the spending was being budgeted.  This is playing politics instead of doing the people’s business properly,  which is what they swore to do in their oaths of office.

My final opinions:

Oath-breaking should be a felony crime.  It harms the entire country.

There should be no debt ceiling to argue about.  It’s unconstitutional.

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Update 5-24-2023:  A version of this was published today as a column in the Waco Tribune-Herald.

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Update 5-30-2023:  So,  Biden and McCarthy reached a deal.  Doesn't say much for the negotiator teams they appointed,  does it?  But now there is a deal with which both the far right-wing extremists of the GOP and the progressive wing of the Democrats are unhappy.  That says what's in the deal is probably just about right.  

Now the trick is to get this voted up in Congress,  and to Biden for signature,  before it is too late.  Arguing over bills to pay from things already previously approved for spending,  what a ridiculous charade!  

This debt ceiling BS is a process by which the party out of power can extort what it wants when it didn't have to votes to get it in the first place.  I would also point out that only the GOP has used this against Democrat presidents.  Democrats have never done this to GOP presidents. That should tell you something about miss-prioritizing party advantage over the good of the country. 

Playing power politics with people's livelihoods and lives,  and the economies of the nation and ultimately the world,  is just egregious oath-breaking.  They swore to do the people's business properly.  This is very most definitely NOT the proper way to do it!

We are about to see another charade take place in the House,  since there are two,  not one,  Republican parties:  the extremists,  and the rest.  We are about to see whether the extremists outnumber the rest,  and just how badly they have cowed the rest.  

The clear and present danger here is that the extremists really don't care about people or the economy,  they just want to tear everything down.  And they have brainwashed their voters into thinking that is a good thing,  instead of the evil that leads to dictatorship. 

We will probably also see McCarthy lose the speakership in the next several weeks.  The extremists in the GOP really don't like him,  he's not one of them.  He just panders to whomever is loudest at the moment (which is how he got the speakership in the first place,  remember?)  Which behavior is definitely NOT doing the people's business!  And for that moral failing,  good riddance.

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Update 6-2-2023:  So,  between the House and the Senate,  they passed the debt ceiling bill just in time,  which despite all the public drama,  was probably the real plan all along.  The country avoided default,  but greatly alarmed all the financial and credit people,  which may drive up interest charged for loans to the government anyway,  making the budget deficit problem even worse.  

And the Republicans got some spending cuts that they didn't have the votes to get,  when the spending was originally appropriated.  This brinkmanship extortion is NOT the way to do the people's business!  I say do away with the debt ceiling.  Its abuse has been mostly one-sided by the GOP anyway. 

Neither party has ever had the right approach to financial policy all by itself.  And after the extremists took control of the GOP,  they most certainly have not had the right approach.  The extremists have driven out most of "real" conservatives,  replacing them with "culture war" demagogues instead of policy proposals.   (This descent into idiocy began decades ago.)

What you really do is balance the budget,  and sock-away the surpluses,  during the good times.  You will need those surpluses during the bad times,  when government must deficit-spend like crazy,  because no one else will spend at all.  

If no one spends,  the economy grinds to a complete halt,  just like at the start of the Great Depression in 1929.  The surpluses to spend in bad times helps you keep the accumulating deficit-induced debt under better control.  

And,  whether times are good or bad are almost entirely NOT within the power of government to control!  Most of that is corporate misbehavior,  which unfortunately is largely inappropriately,  or in some areas entirely-not,  regulated (a failing of Congress).  The lag time for policies to show effects is actually a bit longer than a Presidential term in office.  The US economy is a gigantic thing.  It takes the dead dinosaur about 5-6 years,  just to fall over.

All I can say is stop re-electing these clowns!  See beyond the propaganda and lies.  Vote for the "other guys" instead.  You cannot do worse than what we have.  You might do better.  Simple as that.

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Startling Images

This showed up in my wife's Facebook.  I could not resist.  The image of the hog is startling enough by itself.  But when you add the image of Miss Piggy to it,  that puts it right over top!  Enjoy.  

Heat Shields

Updated same day (5-1-2023):  replaced 3 tables embedded in text that did not indent correctly,  with images that are neat and easier to read.

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Update 5-8-2023:  Corrected the text adjacent to,  and in Figure 2 itself,  to indicate peak heating occurring before peak gees,  not later.  This was based on models run of Apollo-like objects at LEO entry and escape-speed entry.  

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Entry heating is a serious problem for any kind of space activities that require returning something to Earth,  or entering the atmospheres of other bodies that have them.  The “breakthrough” in mitigation schemes came in the early 1950’s when H. Julian Allen and A. J. Eggers realized that blunt shapes endured less heating load.  This enabled the development of ICBM warheads that could survive entry by the mid 1950’s,  and immediately thereafter the development of film payloads returnable from spy satellites.  Shortly after that,  it was used for human passengers returning from orbit in the early 1960’s.   

There are two kinds of heat loads applied to a piece of the surface material of a body during entry:  convective and radiative.  The convective heating is scrubbing by hot gases,  and is more-or-less proportional to speed cubed,  proportional to the square root of the ambient atmospheric density,  and inversely proportional to the square root of the blunt nose radius of the body.  Radiative heating is the heat shining upon the body from something else (the glowing plasma sheath) that is very hot,  and is more-or-less proportional to speed raised to the 6^th power.  However,  this is rather insignificant until it suddenly starts to dominate the total heat load at about 10 km entry speeds.  See Figure 1. 

Convective heating is maximum at the stagnation point,  something like factor 3 lower away from stagnation,  but still scrubbed by attached slipstream flow,  and around a factor of 10 lower still,  on surfaces immersed in separated-wake regions.  Radiation heating is strongest at the stagnation point,  but does not decrease as rapidly as convective heating does, around other places on the body.

There are two or three ways that incoming heat may be lost from that same piece of surface material on that entering body.  Heat absorbed within the material may be conducted further inward into interior cooler structures,  and heat may be re-radiated from its hot surface as infrared (IR) radiation,  back to the external environment.  The third involves either of a couple of mass transfer effects.  See again Figure 1. 

For non-ablative materials,  a sacrificial coolant may be percolated through a porous surface material to cool it.  The coolant absorbs some of the heat load,  then boils away,  and this vaporizing mass flow is carried away,  with that heat it absorbed,  in the slipstream.  That is called “transpirational cooling”. 

For ablative materials,  there is a layer within the material that undergoes pyrolysis as its temperature gets hot enough.  Pyrolysis products are the carbonaceous char left behind,  and copious quantities of gaseous species that percolate out into the slipstream.  It takes a “latent heat of pyrolysis” to do this physical transformation,  so the departing gaseous pyrolysis products carry away significant heat.  That is how “cooling by ablation” works.  Depending upon the density of the material,  it may also interrupt (or carry) conduction heat flow inward to the substrate.  Lower density is lower thermal conductivity.

One should be aware of the effects of the “plasma sheath” of the slipstream that is close to the surface of the body,  behind the detached bow shock wave.  This is very hot gas,  hot enough to be ionized to one level or another (ionization being the definition of “plasma”).  Visually,  it glows with incandescence.  Once brightly incandescent at about 10 km/s entry speeds or higher,  it radiates considerable heat that strikes the immediately-adjacent body surface.  This is the mechanism by which the radiation heating term arises,  which is more-or-less proportional to speed raised to the 6^th power.  See again Figure 1.

That same effect of ionization increasing with speed affects the transmission of radio waves through the plasma sheath,  starting at speeds well below 10 km/s (nearer 6).   That is the cause of the entry radio blackout intervals.  It affects both radio communications and ground-based radar (which sees the plasma sheath,  but not anymore the solid body inside that sheath),  nor can an on-board radar see the surroundings through that sheath.   

At speeds high enough,  a similar thing happens with respect to infrared (IR) radiation:  the sheath goes opaque to it.  That stops any cooling by re-radiation of IR to the surroundings.  This happens with visible light,  too.  It gets very hard to see through the plasma sheath,  if speeds are high enough.

All of these phenomena are summarized briefly in one place,  in Figure 1.  Be aware that in steady state,  the various heat flows must add to zero,  and that the small differences between very large numbers can have catastrophic effects.  Peak heating numbers during entry are quite extreme. 

Figure 1 – The Phenomena Involved With Entry Heat Shields

Entry is not a steady-state process.  One hits something called the entry interface altitude (where heating first becomes perceptible),  but due to the extremely low densities that high up,  one does not decelerate much at all,  initially.  Neither does the heat load build up,  initially.  Low density affects both. Then the larger atmospheric densities further down,  acting in concert with the still-very-high speeds,  suddenly cause very large deceleration forces (and heat loads) on the body.  They “peak”.

After this,  the body is moving very much slower,  and the deceleration and heat loads drop off,  despite the rapidly-increasing atmospheric densities.  This is shown illustratively in Figure 2.  (Although,  if you come in too steep,  you may hit the surface before the max deceleration and heating can occur.)

There is a simple approximate estimating technique,  first done by H. Julian Allen about 1953,  and declassified circa 1958.  This was originally for fairly-steeply-entering warheads,  and so was formulated as a 2-D Cartesian planar analysis.  However,  if you “wrap” the ranges around the curved Earth,  it still gets you “into the ballpark” for the shallower entries we associate with space vehicles today.  It presumes a simple exponential function representing density versus altitude,  which is adequate for the variation of density at the altitudes where these entry phenomena actually occur.  The results one gets with it clearly show that the max deceleration and max heating pulses are not simultaneous:  peak heating occurs slightly earlier than peak deceleration.  See again Figure 2.

Figure 2 – Transient Nature of Entry  (Corrected 5-8-23)

Not shown is the effective average pressure across the body cross section.  This maximizes at peak deceleration.  Think of it as the force to decelerate the body at the peak deceleration gees (basically gees multiplied by body weight),  spread over the blockage cross section area of the body (P = F/A).  That gets you into the ballpark for the surface pressures seen by the body heat shield.  Peak pressures at the stagnation point might be around factor 2 higher.  The heat shield must be capable of structurally withstanding surface wind pressures of that magnitude. 

So not only peak heating rate per unit area capability,  but also max survivable pressure capability,  are critically important to selecting the right heat shield material.  The pressure effect was mostly unrecognized early on,  but became quite important for escape-speed returns to Earth,  and for entries at other planets directly from interplanetary trajectories.

Old NASA data

I found online a slide presentation overview of what NASA knows about heat shields.  This was the Paolo Santini Memorial Lecture,  given by Ethiraj Venkatapathy,  as indicated by the notations in Figure 3.  It mentions some knowledge that precedes the formation of NASA in 1958,  and does not go into the very extensive military warhead heat shield efforts of the early 1950’s.  It does mention an expedient tried on the V-2 rocket to keep the warhead from “cooking off” prior to target arrival.  It also mentions the metallic and coated-metallic surfaces tested hypersonic on the X-15 in the 1950’s and 1960’s.  There are similar heat protection issues with metal-skinned missiles flying at high-supersonic to hypersonic speeds.

The military warhead efforts in the Figure 3 data are only summarized as “wrapped in silica phenolic” circa 1958.  That material is a very good ablator,  if rather heavy and expensive,  that is still often used in solid rocket (and modern ramjet) nozzle construction for missiles.  It is very closely-related to the material finally used for the Mercury capsule heat shield:  fiberglass cloth-reinforced phenolic resin shingles,  bonded together and to a substrate. 

A different scheme was attempted successfully for the Gemini capsule:  a silicone-RTV elastomer loaded into the hexagonal cells of a fiberglass honeycomb.  This basic elastomer-in-honeycomb notion is clearly an ancestor of the Avcoat used on Apollo.  While still heavy for Gemini,  the micro-balloons in the Avcoat used on Apollo reduced its density (to 0.51 g/cc) and weight substantially.  An even lower-density form (about 0.25 g/cc) designated SLA-561V was used on the Mars Viking landers.  See again Figure 3. 

A different scheme was used for the Pioneer-Venus and Galileo probes,  because of the vastly-higher peak heating loads and surface pressures.  This was a tape-wrapped carbon-phenolic composite material.  It was very capable,  but also heavy and expensive.  It did show the advantage of carbon materials as ablators.  So did the Genesis probe,  which used a carbon-carbon composite as ablator,  over a low density carbon insulator (basically a carbon fiber felt). See again Figure 3. 

Mars Pathfinder used a Viking heat shield,  while the Stardust probe used an initial form of the then-new Phenolic-Impregnated Carbon Ablator (PICA) heat shield.  See again Figure 3.  

Figure 3 – A Overview of NASA Heat Shield Knowledge

What got left out of this list were the Space Shuttle heat shield materials:  the two different types of low-density refractory ceramic tiles,  the ceramic cloth insulation blankets,  and the carbon-carbon composite nose cap and aerosurface leading edges,  which were slow ablators at entry conditions from low Earth orbit (about 7.9 km/s at entry interface).  The low-density ceramic tiles served the cooling-by-reradiation function,  and provided the low thermal conductivity effect,  due to their high void space fraction reflected in their low densities.  That last mostly cuts off conductive heat flow into interior structures,  allowing aluminum substructure,  but the material is also inherently weak and fragile. 

Discussion of all the ablative materials

A closer reading of the history of Project Mercury,  reveals that the initial choice was a beryllium heat-sink heat shield,  which actually flew on some of the early unmanned suborbital tests.  By the time of the first manned suborbital flight in 1962,  the glass-phenolic ablative shield was “standard”,  and it proved adequate for entry from low orbit.  It’s still quite heavy,  but was made a bit less expensive by substituting glass fiber cloth for silica fiber cloth. 

The silica version is the better ablator,  but the glass version was adequate for this design.  A sample cut from an ablated Mercury heat shield is shown in Figure 4.  One can see the glass cloth layers in the material,  especially near the pyrolysis zone.  It is the phenolic resin pyrolyzing that creates the carbon char layer.  This is definitely a fiber-reinforced composite material,  with the resin reinforced by the layers of woven glass cloth.  These need to be oriented so that surface wind shear forces do not pry apart the layers of the reinforcing cloth (the same restriction is true with silica phenolic in missile nozzles).  

Figure 4 – The Glass-Phenolic Ablative Used for Mercury Capsules

The Gemini heat shield material was a silicone elastomer injected into the cells of a fiberglass honeycomb,  and cured there.  This is quite the different material from the Mercury heat shield,  but is clearly related in its fundamental concept to the Avcoat-in-honeycomb used on Apollo.  It is also a reinforced composite material,  with the cured silicone elastomer being reinforced by the walls of the small cells of the honeycomb.  The silicone was an ambient-curing elastomer from Dow Corning:  DC 235. 

A sample cut from an ablated Gemini heat shield is shown in Figure 5.  There is a whitish surface atop the black carbonaceous char.  This is mostly molten silica product made from oxidized silicon coming from decomposition of the silicone elastomer,  which elastomer also forms the carbonaceous char.  Molten glass from the fiberglass honeycomb is a small part of this whitish surface material,  which effect is also seen as some whitish flecks on the surface of the Mercury heat shield sample. 

This material is not only heavy,  it is also rather expensive,  because of the hand labor involved.  Each cell must be “hand-gunned” full of the elastomer,  and there are hundreds of thousands of them on a heat shield of any significant size. 

Figure 5 – The Silicone-RTV in Honeycomb Ablative Used for Gemini Capsules

The next step forward with the filled-honeycomb-cells composite concept is the Avcoat-in-honeycomb used on Apollo,  and in one or another form subsequently.  The form used on Apollo was Avcoat 5026-39G,  which was an epoxy-novalac resin filled with both quartz fibers and phenolic micro-balloons,  hand-gunned into the hundreds of thousands of cells in a phenolic honeycomb,  bonded to the substrate surfaces of the vehicle. 

The epoxy-novalac resin (and the phenolic of the phenolic honeycomb) provide the source of the carbonaceous char.  The composite reinforcement is from the walls of the small honeycomb cells.  The quartz fiber filler in the resin provides a source of molten silica for densifying the surface of the char,  as well as a fiber-strengthening function for the virgin material and its char.  The phenolic micro-balloons provide the void space to lower the density rather significantly,  and thus the weight of the finished heat shield.  They also provide a compressive strengthening function,  similar to the aggregate in concrete. 

This material proved adequate in terms of heat load capacity and erosion resistance for Apollo returning from the moon at just about 10.9 km/s speeds at entry interface.  It is rated for 600 Watts/sq.cm,  at a significant fraction of an atmosphere of surface pressure.  A sample cut from an ablated heat shield is shown in Figure 6.  There is enough silica whitening to render the black char’s surface a light gray color.  One can very easily see the small honeycomb cells.

Figure 6 – The Avcoat-in-Honeycomb Ablator Used for Apollo Capsules

The same basic material,  rendered even lower in density (presumably with a higher micro-balloon content) is the SLA-516V material used for the Mars Viking lander heat shield. 

A very close variant of the Apollo material was initially chosen for the new Orion capsule,  designated Avcoat 5026-39HC/G,  which is the same epoxy-novalac resin filled with the same quartz fiber and phenolic micro-balloons,  hand-gunned into the cells of the same phenolic honeycomb.  There are more than 300,000 such cells in the heat shield of an Orion,  so the labor to hand-gun this stuff is very large and expensive,  and the quality of the results varied among the various individual “gunners”.  This heat shield flew on the first Orion flight test EFT-1,  and was very successful. 

To address the labor expense and variability,  a variation was flown on the second Orion flight test,  which was the first Artemis program flight EM-1.  For this heat shield,  the Avcoat was made in tiled blocks of cured filled resin,  without the honeycomb.  300 of these tiled blocks were bonded to the capsule substrate for that flight.  Without the reinforcing effect of the honeycomb,  this was less successful than hoped.  The erosion rate was higher and more variable than expected,  with charred material coming off erratically in larger discrete chunks,  instead of steady loss of fine char granulate eroding away.  Apparently,  deleting the honeycomb reinforcement was a design mistake!  This issue will have to be addressed before flying the first manned Artemis mission,  EM-2.

The success of the Galileo and Pioneer carbon-based heat shields,  plus the success of the carbon-carbon composites of the Space Shuttle nose cap and aerosurface leading edges,  led to serious development efforts toward carbon ablators.  Those culminated in NASA’s PICA material,  subsequently improved and used by SpaceX as PICA-X,  on its Dragon capsules. 

The basic notion ended up as a carbon fiber preform of very high void fraction,  impregnated with a phenolic resin that had lots of bubbles in it,  once cured.  The carbon fibers were the reinforcement to a composite material,  in which the phenolic was the matrix.  The phenolic would form a char under pyrolysis,  and together the carbon fibers and the carbon char would erode very slowly.  The bubbles in the phenolic,  plus the fact that it incompletely filled the void spaces in the preform,  led to low densities in the 0.25 to 0.28 g/cc range. 

This material would handle rather substantial heat loads at modest ablation rates,  survive at significant pressures,  and was very lightweight as heat shield materials go.  The variations involve exactly how you make your carbon fiber preform,  and exactly how you go about creating the bubble voids in the phenolic.  Those variations do significantly impact the heat load capacity,  density,  and strength.

NASA’s original version used a carbon fiber preform from a company known as FiberForm.  It was a felt of high porosity fully carbonized (and thus rigid),  using certain Rayon fibers as the carbon fiber source.  It would handle over 1000 Watts/sq.cm of heat load,  at around half an atmosphere pressure. 

Similar carbon fiber felts are available from other manufacturers,  and some of those are not fully carbonized,  leaving them flexible enough to be conformable.  These variations affect performance significantly,  usually leading to unacceptably-reduced char erosion rate performance as a heat shield.  The biggest problem with the NASA version of PICA has become the unavailability of suitable Rayon fiber due to environmental concerns in recent years.

SpaceX has selected a version they call PICA-X (from three possible variants) for use on their Dragon space capsules.  I have been unable to determine exactly what they did,  but they have the rigid carbon fiber preforms that they need,  seemingly made in-house at SpaceX.  They make their own PICA-X tiles from them,  and bond these to the capsule substrate.  I was unable to locate details,  but the talk is that the SpaceX PICA-X material is easier to manufacture,  and far less expensive (by a factor of 10),  than NASA’s original PICA,  while equaling (or slightly exceeding) the ablation performance of the original.

NASA’s original PICA was used in tile form on the Mars MSL lander (Curiosity rover).  Neither that application,  or any of the SpaceX Dragon capsules,  have shown a problem with tiles coming off.  All in all,  PICA-X seems to be a very reliable material,  apparently without the Rayon availability problems of the original PICA.  That is not to say that future variants could not be even better. 

Figure 7 shows a sample of PICA being tested in the arc-jet tunnel,  at entry-like conditions.  The top of the char layer is soaked-out to white-hot incandescence. 

Figure 7 – A PICA Low-Density Carbon Ablator Being Arc-Jet Tested

 

The best heat shield ablatives currently available

Discussion of refractory ceramics

NASA’s experience with refractory ceramics (which they term “insulative”,  although not all ceramic tiles are) derives largely from 30 years of experience with the Space Shuttle.  Their low density tiles (around 90% void space) were made of silica,  alumino-boro-silicate,  and alumina fibers.  The alumino-boro-silicate fibers were from Nextel,  makers of fire curtain cloth.  All this stuff is limited to returns from low Earth orbit.  High energy orbit entries,  and entries at or above escape,  simply require ablatives,  period. See Figure 15 in the addendum below,  obtained from that NASA presentation found on-line. 

These tiles ended up densified somewhat at the bond side,  and over time with two different dense coatings on the exposed side,  in two colors.  Windward side tiles were tinted black to raise emissivity above 0.8 for efficient re-radiation.  Leeward side tiles were tinted white,  because high emissivity was not needed,  but on-orbit passive vehicle thermal control was.

Initially,  the Shuttle leeside was all white tiles.  Later on,  this was replaced by  flexible thermal blankets of two kinds,  one more heat resistant,  except on the higher-risk OMS pods.  These blankets proved no more durable than the tiles,  but required less labor to install correctly.  They did present their own problems with edges protruding up,  and stitching and seams wearing out quickly. 

Because these tiles (and blankets) were very insulative (very low thermal conductivity),  heat conduction into the substrate structures was minimized enough to permit the use of an aluminum airframe construction.  That was required to make the Shuttle design feasible at all.

The nose cap,  and wing and tail leading edges,  endured temperatures too high for even the black-surfaced ceramic tiles,  and were made instead of carbon-carbon composite-based ablative structures.  The material was a carbon cloth impregnated with phenolic resin,  and furnace-pyrolyzed to an all-carbon composite structure.  This was repeatedly soaked in furfuryl alcohol and hot-dried,  to densify the composite and even-out its properties.  Clearly,  this stuff is not cheap or easy. 

The composite has a high density and thermal conductivity,  but the stagnation zone is a thin line or small patch,  on a much larger part.  It gets very hot near stagnation (around 3000 F during a 7.9 km/s entry).  Heat conducts and re-radiates internally to cooler regions of the part,  where steels can be used to secure it.  Being black,  the exterior surface,  once hot,  re-radiates efficiently to the environment.  Meanwhile,  a little of the composite ablates away with each flight.  After a few flights,  you must replace it,  or else risk loss of craft and crew,  when the thinned part collapses under entry airloads.

Steel and titanium structures supporting and attaching the carbon-carbon composite material,  is how the nose cap and leading edge structures were made.  These required internal insulation to interrupt radiant heating of the titanium portions,  and (more importantly) radiant heating of the aluminum airframe structures to which these parts were bolted.  Clearly,  proper design is not simple!

The exposed-side surface densification (and colorant) coatings were initially a glassy material (termed “reaction-cured glass” or RCG) applied to the surface,  but which did not penetrate into the porous insulative low-density tile structure.  These proved rather vulnerable to impact damage.  They were replaced by a different densifying (and colorant) surface coating (termed toughened uni-piece fibrous insulation or TUFI),  that did penetrate into the low-density tile structure.  These proved much more resistant to impact damage.

Another ceramic heat shield NASA has been working on is known as TUFROC,  for Toughened Uni-piece Fibrous Oxidation Resistant Ceramic,  intended for use on the unmanned X-37B currently operated by the US Air Force.   This craft is a small derivative of the Space Shuttle.  Whether the TUFROC tile system has ever actually flown in it,  is not very clear,  but it may have flown as protection for the leading edges.  What USAF does on-orbit with that craft is not disclosed to the public.  Little is known.

According to what NASA has disclosed,  the TUFROC design calls for two pieces mechanically tied together.  There is some sort of dense,  hard ceramic cap,  overlaying a lower-density porous fibrous ceramic interior.  They did not disclose how these tiles are mounted to the airframe,  but it is most likely similar to what was used on the Shuttle:  expansion pads and spacing bars,  all glued down with RTV silicone to the airframe,  and the tile glued down with RTV silicone to that substructure.

Whatever the TUFROC capping material is,  it can take much higher temperatures (reportedly 1922 K = 3000 F) than the aluminosilicates used on the Shuttle (2000 F rating).  Whatever the fibrous substrate is,  it can also take higher temperatures than the aluminosilicate Shuttle tile material.  NASA does not say what it is or what it can take,  but consider that the backside temperature of the dense cap material will not be that much reduced from the exposed face temperature,  since the dense material would have a high thermal conductivity. 

NASA has been researching the new ultra-high-temperature ceramics (UHTC).  These are high-density materials of high thermal conductivity,  so that a similar scheme to the Shuttle carbon-carbon-composite parts must be used in order hang onto such hot parts.  This technology must be viewed as very immature. 

NASA has also been looking at future ablators.  These include advanced PICA-like ablators,  something called “graded ablators” (which likely has to do with layering different materials together),  conformable PICA (which has recently been found to ablate faster than rigid PICA),  fully-flexible forms of PICA and something called SIRCA,  3-D woven carbon materials,  and a replacement for carbon phenolic.  None of these are ready-to-apply technologies.  (But they could become ready!)

SpaceX is using some sort of tiled heat shield on its new “Starship” vehicle.  It is not yet clear what those tiles are.  They are black for efficient re-radiation,  and they are located on Starship’s windward surfaces during entry.  I have seen them called both “ablative” and “ceramic” in the talk online.  I have even seen one suggestion that they are TUFROC,  although NASA has given that technology to Boeing,  a serious competitor to SpaceX.  If the PICA vs PICA-X history is any guide,  it seems likely the SpaceX tiles are something developed and manufactured in-house at SpaceX.  They could well be an adapted form of the PICA-X ablator.  We just do not yet know.

The best refractory ceramics available

Discussion of re-radiatively-cooled metals

The notion of using metal surfaces exposed during entry goes back to the 1930’s and 1940’s,  long before the heating issues during entry were understood.  The basic notion is to let the skin get hot,  then let it re-radiate thermally to the environment.  There would be some (or perhaps lots) of conduction into cooler structures inside.  The design must establish an equilibrium where the re-radiative (and conduction) heat flows balance the entry heating encountered. 

There is some merit to that notion,  but as it turns out,  even from low Earth orbit,  only on leeside surfaces where the entry heating loads are far lower.  It also has merit for high-supersonic and low-hypersonic flight down in the atmosphere (usually higher in the stratosphere where the densities and heating loads are lower).

There are two things of critical importance for this notion:  how hot can the surface get,  and how efficiently can it thermally re-radiate?  The effective temperature of the surroundings is a part of that efficiency,  but it is primarily controlled by the surface’s spectral emissivity,  a number between 0 and 1.  Surfaces with high emissivities in the IR band re-radiate thermally very efficiently.  Those with low emissivities do not.  The emission is far better at higher temperatures,  that being controlled by a temperature to the 4^th power term in the radiated energy equation.  See Figure 8.

Figure 8 – Thermal Emission from Hot Surfaces

As the figure indicates,  there’s not much re-radiation to be had down near 1000 F temperatures.  Accordingly,  the emissivity makes only a small difference there.  At around 1500 F,  the amount of re-radiation available is becoming quite significant,  and the emissivity makes a clear and compelling difference.  The closer to 2000 F we can operate,  the more effect we can get out of this kind of cooling,  but we really have to have a high emissivity to obtain it. 

You don’t get that high emissivity with an ordinary paint.  At surface temperatures nearer 1500-2000 F,  such a paint would be burnt away.  It takes some sort of metallurgical surface treatment or coating to achieve this,  especially since many metal alloys are quite shiny-silvery in color,  bespeaking quite the low emissivity.  The more reflective the surface,  the lower the visual emissivity,  and the lower the thermal (IR) emissivity is likely to be.  The visual band (0.3 to 0.7 microns wavelength) is just not that far from the infrared bands (0.8-14 microns).

What we are looking for are metal alloys with high max service temperature limits at or above 1000 F,  that still have significant strength when soaked out that hot,  and that can be shaped and welded without too much trouble.  The 1000 F value rules out aluminums (max 350-400 F),  titaniums (max 750-800 F,  and mild carbon steels (max 750-800 F).  That leaves as places to look:  the low-alloy and intermediate-alloy steels,  the stainless steels,  and the high-temperature alloys (iron and other bases). 

Of the low-alloy steels available,  Figure 9 would suggest only D6AC and AISI grades 4140,  4340,  and 8740 as candidates.  These have very nice high strength hot,  but are limited to temperatures in the low end of the attractive IR emissions range:  1000-1100 F.  All would need metallurgical surface treatment. 

Figure 9 – Data For the Low-Alloy Steels

The intermediate-alloy steels of Figure 10 are not very attractive,  being limited to service temperatures only in the 800-900 F range.  “Chrome-moly aircraft steel” (5Cr-Mo-V) is one of these.  

Figure 10 – Data For the Intermediate-Alloy Steels

The stainless steels offer more promise.  These are formable and machineable,  and weldable if one takes care to chose the weldable grades,  usually bearing an L suffix.  They are work-hardenable,  but anneal back to the soft state upon heating.  Strengths are not all that high,  especially hot.  For long-term loading,  creep-rupture effects dominate over short-term strength.  See Figure 11.  

Figure 11 – The Stainless Steels (Both Austenitic 300’s and the Martensitic/Ferritic 400’s)

If service temperatures to about 1600 F are acceptable,  both 316 and 347 stainless are easily available.  The 316L grade is weldable.  Its hot strength is only about 25 ksi tensile ultimate,  which may be too low for a structural skin application on a windward surface.  Strengths are higher, in the 75-80 ksi range,  if you can keep them cooled near only 400 F,  with a refractory or ablative covering.  Grade 304 would be comparable in strength,  and equal in service temperature rating (1600 F),  driven by oxidation (forming surface scale).  The 304L form is weldable.  304/304L also serve well at cryogenic temperatures.

If you need to go hotter,  then grade 310 can take you to 1800 F at 12 ksi strength,  with an oxidation limit of 1900 F.  It is not as available as 304,  316,  and 347.  I am unsure if it is a weldable grade.

The high temperature metals include iron-based,  nickel-based,  and cobalt-based alloys.  The data on these are given in Figures 12 and 13.  At 1400 F,  Hastelloy B has good strength and is at its oxidation limit,  although there are strength data to 1600 F.  A stronger candidate at 1400 F is Waspalloy,  well within its oxidation limit. 

At 1600 F,  Inconel 625,  Inconel 718,  M-252,  and Udimet-500 all have significant strength,  and are at or under their oxidation limits.  There is also Rene 41 (proposed for the X-20 Dyna-Soar).  Rene 41 has process problems reported (long exposures weaken it),  but it has strength reported at 1800 F,  despite being above its oxidation limit there. 

At 1800 F,  there is L-605,  which is within its oxidation limit. 

At 2000 F,  there is N-155,  which is not within its oxidation limit,  and Alloy 188,  which is.  Alloy 188 has the highest oxidation limit at 2100 F of any of the high-temperature metals.  It will go hotter than any of the stainlesses,  and has higher strength that hot,  than the stainlesses do at their limits. 

Figure 12 – Strength-Temperature Data for the High-Temperature Metals

Figure 13 – Machineability and Weldability Data For the High-Temperature Metals

The best re-radiatively-cooled metals

The oddball case of the X-15A-2

The famous X-15 rocket plane was first flown in 1959,  and completed 199 missions by the time it was taken out of service in 1968.  There were 3 vehicles:  X-15-1 bureau number 56-6670,  X-15-2 bureau number 56-6671,  and X-15-3,  bureau number 56-6672.  X-15-3 was destroyed in a fatal crash.  X-15-2 was badly damaged in a crash landing,  and subsequently rebuilt as the X-15A-2,  with external propellant drop tanks.  Both the X-15A-2 and X-15-1 are now on public display.

This craft had Inconel X-750 skins over titanium internal structure,  and had a very black,  highly-emissive metallurgical surface coating.  These skins were convectively heated by low hypersonic flight,  and were cooled by re-radiation of IR thermal energy.  This sufficed to about Mach 5.5-to-6 speeds.  

To go beyond Mach 6,  the X-15A-2 was coated with a catalyst-cured silicone rubber ablative,  designated MA-25S,  which also saw use on the Space Shuttle,  and is a protective coating in common aircraft use.  It is rated to about 700 F,  and will char slowly while surviving 2000+ F fire exposures for several minutes.

There are two forms:  a Type I sprayable,  for area coverage,  and a Type II that is solids-loaded and trowelable,  for small areas,  or making repairs to the Type I.  This stuff is rather pink in color,  and resembles the pink rubber of a pencil eraser.  As it turns out,  there is a fire and explosion danger,  if liquid oxygen is spilled upon this material. 

The X-15A-2 was coated all-over with sprayable Type I MA-25S,  except for the wing,  tail,  and fin leading edges,  which were coated in the moldable Type II.  Because of the liquid oxygen risk (and the test pilot refusing to fly a pink airplane),  this ablative coating had a white sealer coat of paint applied to it.  I am unsure what paint was used,  but I suspect it was some sort of ceramic high-temperature paint. 

Multiple flights were made,  with and without the external tanks,  culminating in the speed record-setting flight to Mach 6.7 at 19.3 miles (about 100,000 feet).  On that flight,  the craft carried a scramjet test article on its ventral fin stub.  See Figure 14.  The scramjet article is mounted to the forward end of the ventral fin stub.  It simulated a cone-spike inlet geometry,  but was not an actual engine.  

Figure 14 – Launch of X-15A-2 with Scramjet Article From B-52 Carrier Plane

There was considerable shock-impingement and shock-interference heating problems due to the presence of the scramjet under the tail of the airplane.  NASA TM-X 1669 indicates that heating was locally increased by a factor of 9 in the impingement zones,  and by a factor of 7 in the interference zones.  On the fin stub and under the tail,  the silicone ablative was completely stripped away,  and numerous holes burned through the Inconel skins,  some of them quite large.  Had the exposure continued even a little longer,  the aircraft might have been fatally damaged.

The silicone ablative was seriously charred in other areas,  most notably the wing,  tail,  and fin leading edges.  Anywhere that there was a local hot spot for any reason,  the white coating was lost,  and the underlying silicone ablative was damaged.  Canopy framing and instrumentation probes were other locations showing local hot spot damage.

An informed speculation says that the white paint color may have impeded re-radiation by low emissivity,  leading to higher surface temperatures than the experiences with the black metal would suggest.  However,  there was nothing about the metal re-radiation or the silicone ablative that could have resisted the shock-impingement and shock-interference damage!  The key for future designs is to eliminate those effects.  That requires very careful aerodynamic design for locating tail surfaces,  and the presence of no parallel-mounted nacelles (like the scramjet article) at all!

Addendum

Figure 15 was obtained from the NASA presentation found on-line regarding the Shuttle thermal protection system (TPS).  It pretty much makes the case (very visually!) about where reusable refractories may be used,  and where ablatives must be used,  for entry heat protection!  This is pretty much based on the 2000 F max surface temperature limitation for the low-density ceramic Shuttle tiles. 

One point:  the earlier space capsules (Mercury and Gemini) entered at conditions similar to the Space Shuttle,  in terms of the altitude-velocity “space” depicted in the figure.  Also bear in mind that locations on the Shuttle that endured stagnation-zone heating were not tiled,  they were protected by the carbon-carbon composite slow ablative.   Ballistic coefficients would have been crudely comparable for the Shuttle and those capsules.  

Some of the probe designs for returning from the far solar system would have had smaller ballistic coefficients,  and a small sample-return capsule doing a free return from Mars would have similar  ballistic coefficients as well as similar extreme velocities.  That is really why the trajectory lines for Mars return and far solar system return are so close together.   A manned vehicle coming back from Mars will more likely have a higher ballistic coefficient and a less extreme return velocity.  Its curve would be to the left and somewhat below the Mars return line shown in the figure.

What makes the stagnation zone problem so difficult,  is that there is a limit to the heat rate that can be re-radiated from a “black” surface,  that is determined by just how hot you can let that surface get (in the Shuttle tile case:  2000 F).  The stagnation zone heating rate even from only low Earth orbit can be (and in most cases is) very much higher than the possible re-radiation heat rate.  With the conduction inward mostly cut off by the low density and low thermal conductivity of the ceramic tiles,  re-radiation was the only way to reject the applied convective loading.   They had to be equal!  That limits speed.

However,  away from those stagnation zones at the nose cap and leading edges of the Shuttle,  heating rates are crudely a factor of 3 lower,  which is why the black tile solution was feasible for the windward-side surfaces on the belly,  and the bottoms of the wings.  On the leeward-side surfaces,  heating rates are crudely a factor of 10 lower than stagnation,  which is why the leeward-side tiles could be the inefficiently-re-radiating “white” color that allowed passive thermal control on-orbit. 

For a small object with a very low ballistic coefficient,  stagnation heating rates are somewhat lower,  because the peak deceleration and peak heating occur higher up in the lower-density atmosphere,  if the entry angle is shallow.  If the object is also very,  very blunt (large “nose” radius),  that also lowers stagnation heating.  For such an extremized case,  “black” shuttle tile with a 2000 F limit could be used to protect even the stagnation zone.   Steep entry angle easily negates this.

The point here is that the lines on the figure representing the various entry trajectories can vary somewhat with varying ballistic coefficient,  which is a function of object size (or mass).  The red line dividing “reusable TPS” from “ablatives only” can move quite a bit with varying ballistic coefficients and bluntness,  and significantly if low-density ceramics with higher temperature limits become available.  The NASA figure indicates limits which have been “typical” up to now,  not “cast in stone” in perpetuity. 

Figure 15 – Figure From NASA Presentation Showing Entry Heat Shield Choices

References consulted but not formally cited

Agrawal and Chavez-Garcia,  “Fracture In Phenolic Impregnated Carbon Ablator”,  paper given at the 42^nd AIAA Thermophysics Conference,  Honolulu,  HI,  June 2011.

Ethiraj Venkatapathy,  “Ablators:  From Apollo to Future Missions to Moon,  Mars,  and Beyond”,  the Paolo Santini Memorial Lecture,  given at the 70^th International Astronautical Congress,  Washington,  DC,  October 2019.

Panerai,  et. al.,  “Analysis of rigid and flexible substrates for lightweight ablators based on X-ray micro-tomography”,  manuscript found on-line via Elsevier,  dated 2016.

Nowlin and Thimons,  “Surviving the Heat:  The Application of Phenolic Impregnated Carbon Ablators”,  Conference Session B9,  paper number 3131,  University of Pittsburg Swanson School of Engineering,  dated 2013.

Poloni,  et. al.,  “Carbon ablators with porosity designed for enhanced aerospace protection”,  international paper financed by the Swiss National Science Foundation pertinent to project 200021_160184.  No presentation or publication date given,  but ref. 1 dates to 2020.

Rodriguez and Snapp,  “Orbiter Thermal Protection System Lessons Learned”,  AIAA paper 2011-7308,  AIAA Space 2011 conference and exposition,  Long Beach,  CA,  2011.

Sylvia Johnson,  “Thermal Protection Materials:  Development,  Characterization,  and Evaluation”,  presentation at HiTemp2012,  Munich,  Germany,  2012.

Watts,  “Flight Experience with Shock Impingement and Interference Heating on the X-15A-2 Research Airplane”,  NASA TM X-1669,  October 1968.

MA-25S Product Data Sheet,  labeled as coming from “Thermal Protection Products”,  no date given.

Mil Hndbk 5C,  “Military Standardization Handbook:  Metallic Materials and Elements for Aerospace Vehicle Structures”,  Sept. 1976.

“High Temperature Characteristics of Stainless Steels”,  a designer’s handbook series no. 9004,  distributed by the Nickel Development Institute (NiDI),  and produced by the American Iron and Steel Institute (AISI), no date given.

Related Articles

I have also posted a number of related articles on this “exrocketman” site.  Use the navigation tool on the left side of this web page to find them quickly and easily.  All you need (I suggest jotting them down) is the title and posting date,  to use the navigation tool.  Click on the year,  then on the month,  then on the title,  if more than one article was posted that month. 

Early High-Speed Experimental Planes, 3 July 2022

About Hypersonic Vehicles, 1 June 2022

On High-Speed Aerodynamics and Heat Transfer, 2 January 2020

Heat Protection Is Key to Hypersonic Flight, 4 July 2017

Shock Impingement Heating Is Very Dangerous, 12 June 2017

Entry Heating Estimates, 1 April 2020

Thermal Protection Trends For High Speed Atmospheric Flight, 2 January 2019

Low-Density Non-Ablative Ceramic Heat Shields, 18 March 2013

BOE Entry Analysis of Apollo Returning From the Moon, 21 January 2013

“Back of the Envelope” Entry Model, 14 July 2012