Recent Heat Protection Items: 3 of 3

I was recently an invited speaker at a recent American Carbon Society meeting held at North Carolina State University in March 2024.  The meeting topic was thermal management,  and it mostly (but not entirely) dealt with carbon-based materials.  Dr. Cheryl Xu of the school’s Department of Mechanical and Aerospace Engineering ran across me and invited me to attend.  The meeting was managed by Dr. Xu and by Dr. Weiming Lu of the American Carbon Society,  a senior employee of Collins Aerospace. 

I submitted 3 presentations in both slide set and submittable paper format,  which were all accepted.  Only one could be presented live,  the other two would be poster presentations.  In point of fact,  all three were made into posters.   Being about the oldest person present,  I felt quite obsolete among specialized experts used to dealing with complex software.  But,  apparently because of my having turned an awful lot of heat protection materials into char during my defense career,  they seemed glad to have me there. 

The “live” one was “Early Ballpark Analysis:  Entry”,  which took entry as the specific example for a generalizable concept screening process that makes true brainstorming feasible in terms of time and budget.  Plus,  being able to get ballpark estimates without software makes possible the recognition of garbage-in,  garbage-out problems when using software.  Those were my messages.  They seemed well-received.

My other two presentations were “Old Ceramic Composite”,  an update on a paper I gave previously at a Mars Society convention a few years ago,  and “Ramjet Ablative Liners”,  about some experimental combustor insulations I tested years ago as part of a very productive ramjet direct-connect test series.   

I have already posted a sort of trip report on the conference,  the entry screening paper,  and the ramjet ablatives paper,  on this site.  They are just below as “Presenter at Workshop” 23 April 2024,  “Entry Concept Screening” 3 May 2024,  and “Ramjet Ablative Liners” 1 June 2024. 

What follows is the text of the submittable paper for “Old Ceramic Composite”.  A related article here on this site is “The ‘Warm Brick’ Ramjet Device”,  posted 2 November 2021.

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Old Ceramic Composite           GW Johnson        12-31-2023

Abstract

At one job many decades ago,  the author had need of an insulating liner for an airbreathing combustor,  which could be fired many times without needing refurbishment.  Inspired by NASA’s Shuttle tile,  he tried low-density ceramics,  and was successful on his second attempt with a fabric-reinforced ceramic composite. 

That effort long ago was done on company R&D funding,  with a low budget.  There was no money for real materials development and characterization activities,  it had to be made from commercially-available products,  and not the most expensive ones at that.  What worked was a commercial ceramic pipe insulation paste,  reinforced with a commercially-available ceramic fire curtain cloth used in aircraft engine nacelles. 

Many years after that effort,  the author realized that this material had potential as a refractory entry heat shield material.  He identified some ways and means to use it for that purpose.

Background

The author was doing feasibility testing for a towed aircraft infrared decoy concept back in 1984.  This thing used a fuel-fired, ram-fed airbreathing combustor as a hot gas source for creating the infrared signature.  It was intended mostly for high subsonic to transonic flight,  at low altitudes.  That corresponds to the ground attack and ground support scenarios.

The author needed a cheap combustor insulator that could support multiple burns without any refurbishment,  and the then-available ablatives were not “it”.  Inspired by NASA’s low-density and consequently low thermal conductivity shuttle tile,  he decided to try low-density ceramics himself.  This was low-budget company R&D,  though.  He could only use ordinary commercial materials,  and certainly not the most expensive ones. 

Experimental Material

The author selected a moldable pipe insulation paste from Cotronics Corporation called 360M,  which was a water-based slurry of alumino-silicate flakes and fibers,  cured by drying out.  Handbook data indicated far lower density and hot thermal conductivity than most dense ceramics. 

It was quite porous and he did not want hot combustion gas entering those pores,  so he needed a surface sealant.  He used as a “paint” a ceramic adhesive from Cotronics,  designated 901,  for that sealing function.  It,  too,  was water-based,  cured by drying.  Both were recommended as max 100 C oven cures,  although the author cured them at 103 C “just to make sure”.  Both of these materials are still available from Cotronics today,  although you generally have to ask them about the 360M paste,  it is considered obsolete.

The author’s liners were retained inside the combustor shell,  but not bonded to it.  See Figure 1.  The first example worked fine,  until the fuels testing reached the point of determining rich and lean combustor blow-out limits.  Rich blowout is marked by very strong instabilities,  showing up as harsh pressure oscillations.  Those cracked the fragile insulator all the way through,  and the combustor promptly spit the whole unit out in pieces,  starting a grass fire that had to be stomped out. 

Figure 1 – Basic Data About the Two Experimental Liner Materials

Taking a cue from his experiences building fiberglass canoes and kayaks with his father,  the author decided to try fabric reinforcement for his low density ceramic liner as the means to withstand the nasty oscillating pressure loads.  As it turned out,  3M Corporation offered as a commercial product a ceramic fire curtain cloth for aircraft engine nacelles.  This cloth was going to be used anyway,  as the heated source of infrared signature from the decoy. 

A that time,  3M offered Nextel 312 alumino-silicate fibers as yarns woven into cloth.  If memory serves,  the cloth designation,  specifically chosen for its high air permeability,  was AF-14.  Since then,  more ceramic materials have become available from 3M for a variety of purposes.  But Nextel 312 is still available in the AF-14 cloth from 3M today. 

The author molded the ceramic composite liner as alternating layers of 360M paste and wraps of the cloth,  again cured at 103 C,  just like the plain liner.  This one was very successful,  resisting many rich blow-out instability events,  several dozen test burns of varying lengths,  and accumulating several hours of burn time,  all without any refurbishment at all.   When it proved successful,  he just used it;  he never characterized it for any of its properties.  There was no money for that.

About The Old Decoy Testing

See Figure 2.  It shows the test combustor rig hardware,  Figure 3 below shows the test facility used for this work.  The combustor shell has the dimensions of 2-inch iron pipe,  about 1.5 inch diameter as insulated,  and about 3 inches long,  inside.  Ahead of it is a duct piece where fuel-air mixing takes place,  and ahead of that is an inlet piece where air capture and diffusion,  and fuel injection,  take place. 

The author and his team used an automotive spark plug and an aftermarket automotive high-energy ignition set to achieve fuel-air ignition,  mounted in the recirculation zone just downstream of the step area change that was our flameholder.  They successfully tested with hydrogen,  propane,  aviation gasoline,  commercial jet fuel,  and technical grade ethanol as the fuels. 

Figure 2 – The Decoy Test Article

Figure 3 – Improvised Free Jet Test Facility

The test facility comprised a PVC pipe stilling chamber with a wooden convergent-only nozzle block that created a jet of high-speed air,  up to about Mach 0.9.  It was powered by what in 1984 was the largest-capacity rental air compressor trailer,  anywhere in Texas.  The temperature of the air stream was measured as 190 F in the exiting jet. 

The test article was mounted on a heavy steel stand made of plate and pipe.  This was positioned about a stilling chamber diameter downstream of the air nozzle block’s exit,  immersed in the exiting jet.  That’s basically a “poor man’s free jet test facility” for ramjets.  Fuel control was by the sight,  sound,  and smell,  as sensed by the author,  standing a few feet away in the hot jet blast behind the test article. 

As a minimum,  it took the author and a technician to operate the facility and run tests.  Sometimes the team included a second technician,  and up to two junior engineers as assistants.

Simple Molding Tools,  Near-Pristine Used Liner

See Figure 4 for an image of the molding tool set that the author used constructing the combustor liner and nozzle pieces as ceramic composites.  The molding plug mandrels are wooden items he made using his drill press to spin them,  in lieu of a lathe. (As previously mentioned,   this was low-budget company R&D stuff.)  The finish on the mandrel plugs is just a coat of varnish.  The combustor shell was its own liner outer mold shell.  He had a separate outer shell for molding the nozzle blocks. 

Figure 4 – Very Simple Molding Tools

In Figure 5 is a view of the combustor liner after all the testing was done.  Note the near-pristine appearance of the ceramic composite.  The nozzle was similarly near-pristine.  The surface temperatures approached the 3250 F melt point of the material,  as evidenced by some very localized surface-flowing downstream,  most of that in the nozzle block throat,  where heat transfer film coefficients are highest. 

This material normally has an ultimate max service temperature of about 2350 F,  limited by shrinkage cracks upon cooldown.  There are some such cracks visible,  but they do not go all the way through thanks to the fabric layers,  and there was very little in the way of fluid shear forces to “pick” at these cracks,  since the internal flow velocities were so low.    

Figure 5 – Liner Condition After Testing Concluded

Attempted Material Characterizations,  Decades Later

Many years later,  the author analyzed the combustor performance for typical test conditions,  and then used these to help set up a cylindrical steady-state heat balance model of the insulated combustor.  This hardware was immersed in 190 F air as its external coolant,  turbulently flowing by at high subsonic speeds.  On the inside,  full-rich flame temperatures dropped across the internal boundary layer to very near the meltpoint of the ceramic,  at its internal surface. 

After an hour-long burn at the indicated conditions,  one can presume that heat transfer is steady-state.  Back in 1984,  the author licked his thumb and touched it to the shell,  to find that it would just barely boil spit,  meaning it was near 215-220 F in thermal equilibrium.  The shell itself was 300-series stainless steel,  with a high thermal conductivity,  and thus very nearly isothermal,  as the numbers indicated.  

Not knowing what the actual hot thermal conductivity of the ceramic composite was,  the author ran his thermal analysis repeatedly in a spreadsheet,  across a large range of possible thermal conductivity values.  The one that “fit” the observations of “near-melt” and “boil-spit” the best (and this is crude, of course),  was 0.02 BTU/hr-ft-F,  which is in metric 0.035 W/m-C.  See Figure 6 below for all the data,  and Figure 7 below,  for the temperature distribution plot obtained at the “best” thermal conductivity value.  

That hot thermal conductivity is actually lower than the hot thermal conductivity reported in the Cotronics handbook for the 360M paste alone!  Because low thermal conductivity correlates with low density (or high void volume fraction),  this implies the density of the author’s composite was lower than the handbook density,  as well.  He honestly does not know what that finished density really was,  but it subjectively felt like it was somewhere near that of industrial-grade styrofoam.

Figure 6 – Simple Cylindrical Steady State Heat Balance,  Parametric On Liner Thermal Conductivity

Figure 7 – Temperature Distribution for “Best-Fit” Thermal Conductivity

While the heat balance analysis was actually done in US customary units,  I converted the temperature distribution to metric for Figure 7.  The thermal gradient indicated for the evidently low thermal conductivity liner is quite remarkable!  It’s in the vicinity of 270-275 C per mm of thickness,  or in US units 12,000-13,000 F per inch!  And that effect is a very real result!  The hot gas temperature and outer shell temperature are known well enough to support that assertion.

Possible Explanation for Low Thermal Conductivity and Density

The apparent low thermal conductivity and density,  estimated years afterward for this ceramic composite,  are well below the Cotronics handbook values for their cured 360M material.  See Figure 8 below.  There are only two differences here between the normal use of that paste as trowel-on pipe insulation,  and what the author did with it:  (1) he cured his liner at a higher-than-recommended temperature by about 3 C,  and (2) there is a cloth of some inherent porosity embedded within it. 

Cotronics recommends at most 100 C for the cure temperature that drives the water out by evaporation.  The paste is aluminosilicate solids slurried in water,  with something in the water that enables the solids to adhere together,  once the water is gone. 

The elevation where the author did this work is about 500 feet above mean sea level,  so the normal boiling point of water there is very slightly below the sea level value of 100 C.  He cured at 103 C,  so the slurry water in his layup actually flashed into steam,  at least partly.  He suspects that the steam “wormholing” its way out of the layup,  acted to increase its void fraction,  thereby lowering both finished density and the resulting thermal conductivity. 

Only the water in the slurry could penetrate into the pores of the fabric,  and into the pores between the fibers in its yarns.  The solid flakes and fibers could not do that,  and yet being similar materials,  the cure chemistry did allow them to bond with that fabric.  But the large inherent porosity in that fabric would necessarily add to the void fraction in the finished article.  That would also act to lower density and thermal conductivity of the finished product. 

At least,  that is the author’s hypothesis.  There is no proof,  other than “it worked”.  

Figure 8 – Rough Estimates of Material Properties Made Decades Afterward

 

Adaptations for Use as an Entry Heat Shield

To use this stuff as an entry heat shield material,  three things must be addressed:  (1) achieving high thermal emissivity to make re-radiation cooling more effective,  (2) the risk of shrinkage cracks in a high fluid shear environment,  and (3) mounting on a exterior surface.

The two plots in Figures 9 and 10 show calculated results for a black surface of high emissivity in Figure 9,  and in Figure 10 for the stock white surface of low emissivity.  In making the plots,  “high” thermal emissivity was assumed to be 0.80,  and “low” as 0.20.  The plot ordinates are equilibrium surface temperature such that re-radiation to Earth temperatures exactly balances stagnation heating,  versus how much stagnation heat flux as the abscissa.  Also shown are the max surface temperature limits imposed by the onset of cracking and melting.  

Figure 9 – Calculated Steady-State Heat Balance For High Emissivity

Figure 10 – Calculated Steady-State Heat Balance For Low Emissivity

The switch to a high emissivity black surface would seem to be easy enough to address.  The author called Cotronics and asked if significant carbon black could be added to their 901 adhesive.  Their answer was that a fair amount actually could be added,  and they could do it.  So,  you just seal the surface pores of your layup with a black “paint” instead of a white “paint”.   Nothing exotic. 

To address the risk of high fluid shear forces tearing at shrinkage crack lips,  one simply operates at conditions below the shrinkage crack limit,  in turn well below melting.  That makes stagnation zone service for Earth entries at around 25 W/sq.cm very marginal,  but it would have application to Mars entries either from low orbit or directly off the interplanetary trajectory there. This does presume very blunt shapes with very large effective “nose radii”.  Stagnation convective heating flux is modeled per H. Julian Allen’s 1953-vintage warhead entry analysis:  Qconv/A =  K(density/nose radius)^0.5(velocity)³.

NASA‘s shuttle tile was bonded directly to the structural substrate (metal airframe surface) with an RTV adhesive.  That bondline is a single point failure mode,  however!  The same could be done with my ceramic composite,  but there is another possibility arising from the fabric reinforcement that is embedded within it! 

If made in panels upon a metal substrate panel as depicted in Figure 11,  it can be bonded with RTV to that substrate,  but the free edges of the reinforcing fabric can also be folded around the edges of that substrate,  and clamped in place on its backside.  That provides the redundant retention that NASA’s tiles never had!

Figure 11 – Achieving Redundant Retention

All this requires in the way of vehicle design change is putting the airframe’s metal skin panels onto the airframe framing as already-insulated substrate panels,  instead of installed bare-metal panels to be insulated afterwards.  That does require fasteners instead of welding or rivets,  to facilitate removing individual panels for repairs.

Development Status

The author’s ceramic composite from 1984 is basically decades obsolete,  and never received any materials development effort,  rendering it technologically immature.  Others have since gone far beyond where the author ever got.  But the concept of a ceramic cloth reinforcement embedded in a ceramic matrix still has very good promise today!  The greatest of these promises is the possibility of redundant retention.  But his notion of a ceramic matrix composite has already been used. 

Figure 12 shows what NASA developed from its Shuttle tile technology,  that is currently flying on the USAF X-37B.  This serves even in its stagnation zones,  where the older shuttle tile was inadequate.  NASA calls this “Tufroc”.  It features a ceramic composite cap of higher density,  strength,  and temperature resistance,  atop a low-density ceramic substrate,  in turn applied as a tile.   

Figure 12 – NASA “Tufroc” 2-Piece Ceramic Tile as Used On USAF X-37B

The cap piece is a reinforced composite,  similar to the author’s concept:  a carbon fiber preform reinforcing a ceramic matrix.  The substrate piece is a low-density ceramic that is a strength improvement upon the older shuttle tile material.  It is not a composite,  although it could be.  The two pieces are mechanically linked,  but the substrate is still just bonded with an RTV adhesive to the metal airframe substrate.

This material will handle Earth low orbit entries in stagnation zones,  it is much more resistant to impact and erosion damage,  and it does not suffer from surface shrinkage cracking.  However,  the bond to the substrate is still a possible single point failure!  

If the substrate piece were at least partly a fabric-reinforced composite,  then it could possibly incorporate the author’s clamped-fabric idea for redundant retention.

None of these NASA materials are ordinary commercial products.  They are special-made,  and available only in smaller sizes. The author’s stuff is made of fairly-inexpensive ordinary commercial materials,  and it could be made in rather large panels.  Nobody is doing that,  just yet.

Conclusions

The author’s material is very promising,  but totally undeveloped.  It “just worked” as a combustor liner.,  and it was made of ordinary commercially-available materials.

There are other more recent materials of good promise,  that are mature and flying.

Adding the author’s clamped reinforcing fabric concept to those newer materials might well solve the redundant retention problem. 

About the author:

The author had a 20 year career in aerospace defense doing new product development design,  analysis,  test,  and evaluation, entering the workforce in the slide rule days with a master’s degree in aerospace engineering.  Transition to the then-expensive pocket calculators was underway,  but desktop computers were still years in the future.  That career was mostly (but not entirely) in rocket and ramjet missile propulsion.  It ended with a plant shutdown and layoff in 1994,  just when the industry was shrinking drastically.  The author then had a second 20 year career that was mostly in teaching (at all levels from high school to university),  plus some civil engineering and aviation work.  He earned a doctorate in general engineering late in life,   to support that second career.  He is now retired.