Ramjet Data Re: Heat Shields

Update 3-13-2026:  Latest word is that NASA will attempt to launch Artemis-2 with a crew somewhere around 1 April.  That Orion capsule has the same flawed heat shield design as the un-crewed Artemis-1 test.  They will eliminate the entry skip on this one,  believing that skip-reheating to be the cause of the spalled-out chunks of char that left craters behind in the Artemis-1 heat shield.

The first Orion capsule to fly in space was EFT-1,  which had a heat shield built of the same basic material,  but manufactured by hand-gunning into the cells of hex,  the same way Apollo was.  It performed fine,  but like Apollo,  was labor-intensive and therefore expensive.  Artemis-1 was built of the same material,  but without the reinforcing hex,  as cast blocks of the material machined into bonded tiles.  The tile bonds and gap fillers performed fine,  but the basic material spalled out chunks unexpectedly (and unpredicted).  

It was unreinforced,  lacking the hex,  and in my judgement,  THAT is the real "smoking gun" here,  not the skip reentry!  My contention is that deletion of the reinforcing hex weakened the char enough to let it spall out chunks,  whether by internal pyrolysis gas release pressure,  or by fluid shear across the surface,  or both.  Likely both.  If Artemis-2 shows similar cratering,  then we know it was the lack of hex "for sure",  and not the second heating pulse of a skip reentry.  

Although,  we are risking a crew to see that result!

The real mistake here was the decision to build Artemis-2 the same way as Artemis-1,  BEFORE Artemis-1 ever flew!  And now,  as I hear it,  they are already building Artemis-3 to the same flawed design!

The following is a handout describing how to put the reinforcing hex back in,  without incurring the labor-intensive hand-gunning.  I sent this information to the heat protection group in Houston,  and they thought I was correct,  but it went nowhere.  Now I have sent it to the new Administrator,  Jared Isaacman.  I have not heard anything back so far.

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original article:

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The topic here is how experimental experiences, with ramjet ablative combustor liners,  relate to the unexpected and unpredicted entry heat shield problems seen on the Artemis-1 test flight.  There actually is some correspondence,  the overlap suggesting a “fix”.  What might actually be done to correct the heat shield problem is presented here.

The author is qualified to speak on this topic,  having spent 20 years in aerospace defense new product development work,  most of it in solid rockets and ramjet propulsion.  He is no narrow specialist,  as the next 20 years in teaching (at all levels),  in civil engineering,  and in aviation work demonstrates.   He is now long-retired,  but occasionally consults. 

The Problem                               

The Artemis-1 heat shield partially failed by unexpected and unpredicted cratering,  but did not have a burn-through.  It was constructed of Avcoat tiles bonded to the capsule structure.  Chunks of the charring heat shield spalled out,  creating alarmingly-deep craters in it.  The same Avcoat material was used as the Orion EFT-1 flight test heat shield,  and on the Apollo capsule years ago.  The difference with Artemis-1 was that the Avcoat of those earlier heat shields was hand-gunned into the numerous cells of a fiberglass-type hex attached directly to the capsule structure.  The change to bonded tiles was intended to greatly reduce the labor cost of the hand-gunning,  and also reduce its inherent variability.

The bonding of the tiles to the Artemis-1 capsule,  and the gaps between them,  are not in question here.  Those performed quite well.  Nor is the cost reduction in question,  it is quite real,  obtained by casting blocks of Avcoat that were final-machined to tile shapes. 

The only real difference is the question explored here:  those blocks of Avcoat had no reinforcing hex in them!

A Place to Look is Outside the Organization!                     

One needs a “fresh pair (or pairs) of eyes” to see past the thinking “ruts”!  The problem with that is that one may not really know the people who are from outside.  Therefore,  one tends to trust such outsiders less!  Another aspect of that is the “not invented here” prejudice!  One must strongly resist that!

Ramjet Ablatives Is a Related Experience

Unlike solid rockets,  ramjets have very long burn times.  And unlike rocket motor free volumes out near their case diameters,  ramjets have high fluid shear forces all along their combustor walls!  The ramjet has long proven to have the more challenging environment for ablative protection.  The reinforced-rubber insulations used in rocket cases are quite inadequate for ramjet combustors!

The reentry heat shield sees hotter local effective temperatures,  experiences lower surface pressures (because the atmosphere is so thin at entry peak heating altitudes),  and quite likely sees fluid shear forces along the heat shield surface that are on the same order as those of the ramjet combustor.  That last seems likely because both are subsonic (behind the bow shock for the capsule,  with higher sound speeds but lower densities). 

This author ran some experimental ablative candidates through short-burn,  full-scale tests in a combustor of the size for a ramjet replacement of the AIM-120 AMRAAM motor.  In the process of understanding the results,  he identified some 5 key issues to consider,  as are indicated by notes in red in Figure 1 below.  (All figures are at the end of this article.)  These issues were porosity for gas release,  the polymer pyrolysis temperature,  the “glue” effect of viscous melt materials,  the “reinforcing aggregate” properties of particulate solids,  and the reinforcing effects of fibers reaching into (or within) the char layer.

The baseline ramjet insulator,  Dow Corning’s DC 93-104,  did the best,  as expected.  The shortfalls of the experimental alternatives were understood,  once those issues were all identified.  These issues were all addressed in the formulation of the DC 93-104 (as indicated by the blue notes),  but were not all addressed in the other materials.  Those included a Japanese “equivalent” to DC 93-104,  and two pre-preg cloth layups.

These were short-burn tests,  all under 1 minute long.  It was already known that many ramjet applications needed longer burn times than that,  but that thicker layers of the ablative were simply not a design option,  because of the enclosed volume reductions.  The solution to the burn time problem came from “in-house”,  not Dow Corning,  and is illustrated in Figure 2.  As used in the 20-inch diameter ASALM-PTV combustor at the same liner thickness,  the retention ribbons held the char in place as an insulator,  for up to 15 minute burns!  While not as “good” an insulator as the virgin material,  the retained char was “good enough” to essentially achieve acceptable steady-state results in ground tests and in flight tests!

That kinked-ribbon retention feature was not included in any of the experimental insulation tests discussed here,  primarily because all the burn times were under 1 minute!  Not being present in these tests,  it cannot obscure or impact the fundamental ablation phenomena behind the results obtained!  Furthermore,  the two pre-preg cloth alternatives had to be wrapped onto an inflatable-bladder mandrel for installation,  more like rocket motor insulations.  They could not use the kinked-ribbon retention feature anyway!  The hope was that the cloth fabric would provide enough retention.  But it did not.

Applying This Knowledge to Ramjet Combustors

The designer must deal with all 5 of the issues in the ramjet environment,  but only 1 of the 5 issues actually dominated the picture,  as indicated by the gold stars in Figure 3.  This dominant issue proved to be randomly-oriented high-temperature-capable fibers tying the char to the virgin beneath.  The need for enough porosity to release the pyrolysis gases through the char was not much of an issue in the thin layer designs we were using.  In heat shields that are thicker,  it might be more important!  There is some impact of the presence (or not) of sticky,  viscous melt on the surface,  to prevent the more rapid erosion of small particles by the fluid shear.

If ramjet combustor design was all that we were considering,  then the loss of some char chunks as seen with the Japanese material,  might be countered either by the long-burn kinked-ribbon retention feature,  or by adjusting the formulation to include more carbon fiber,  or both.  Even so,  it provided useful 1 minute burn time,  even with some chunk loss.

Applying This Knowledge to Heat Shields

There is definitely overlap of the physics identified between the ramjet combustor and the heat shield applications!   While heat shields do not deal with all the same issues,  the porosity and char loss do indeed overlap a little.  But the need for radial fiber reinforcement would seem to dominate both applications,  preventing as it does the loss of chunks of char!  This is illustrated for ramjets in Figure 4.

Quite apparently the real mistake made with the Artemis-2 bonded-tile heat shield design,  was deleting the reinforcing hex from those bonded tiles!  And that says the most important thing NASA could do is to look for ways and means to put that reinforcing hex back into those very bonded tiles!  Preferably doing it without doing the labor-intensive and expensive hand-gunning of the Avcoat into each individual hex cell! 

So,  the main question here is:  can that really be done?

There Really Is a Practical Fix         

The answer is “yes”! 

But to even consider it,  one must avoid the trap of “either/or” thinking!  That trap considers only tiles without hex versus doing it hand-gunned completely like Apollo and Orion EFT-1.  That thinking trap is illustrated in Figure 5,  along with this author’s way out of that trap:  use an extrusion press to load all the cells in the hex with Avcoat,  all at once!

That way,  one still uses bonded tiles,  and avoids all the hand-gunning labor!  And the tiles with hex in them should cease shedding chunks of char,  no matter how the entry is flown,  skip or not!  Plus,  the high micro-balloon content that leads to low densities,  but also to high mix viscosity difficulties,  is something the extrusion press can handle.  All that is needed is the right tile mold tooling to hold the hex in place on the press.  That is just tooling design.  Plus,  the inherent human labor variability is eliminated.

Such a solution is unavailable to people hampered by the bad habit of “either/or” thinking.  The author gave this solution to NASA more than a year ago!  But they so very clearly did not use it,  or even contact him about it!

Final Comments

These ramjet results and their application to heat shields are as much “engineering art” as they are anything!  One will not find this in published reports or academic texts.  A lot of this is qualitative,  not quantitative,  and inherently so!  The difference between “engineering art” and “engineering science” is explained in Figure 6.  Note the large fraction of the necessary knowledge that is “art”,  especially in development work!

Note also that the engineering art is passed-on,  one-on-one,  on-the-job,  from the old hands to the “newbies”.  That is,  it is passed-on if,  and only if,  there are any old hands on the staff to do that teaching job!  Too many organizations prefer to hire only “newbies” that they can under-pay,  instead of retaining the “old hands”,  excusing this with “because they are too expensive”.  But it loses you half (or more) of your essential knowledge!

The current status is this:

It is too late to fix the Artemis-2 heat shield,  that rocket is on the pad to launch!

               Artemis-2 is forced to risk worse or deeper cratering with a crew

               Probability of a fatal burn-through is not zero !

               Had almost 2 years,  so why not?  Schedule and money,  same as Challenger!

Must fix the heat shields of Artemis-3-on,  to avoid a fatal failure,  sooner or later!

               Must not repeat Challenger mistake:  must value lives above schedule & money

               Trust results (even qualitative),  not the same sources that failed to predict cratering

What one has to learn and remember is this very harsh lesson:  There is nothing as expensive as a dead crew!  Especially one dead from a bad management decision!  

In different ways,  this bad decision thing is really is the root cause of both the Challenger and the Columbia disasters!  “Avoid that mistake like plague” is the best advice this author can give you,  even if it means listening to outsiders,  or overcoming “not invented here” prejudice!  Please feel free contact him for more details.  He has test data and lots of photos from those old ramjet tests. 

Figure 1 – Related Experience with Ramjet Ablatives

Figure 2 – Raising Ramjet Burn Times with Thin Liners

Figure 3 – Only One of the Issues Dominated the Short-Burn Ramjet Tests

Figure 4 – There is Physics Overlap from Ramjet Liners to Heat Shields

Figure 5 – There Really Is a Practical “Fix”

Figure 6 – “Rocket Science” Really Is Not Just Science

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Get Acquainted Info: High Speed Vehicles

This article is for people who know little about high speed flight vehicles.  It gets across some key concepts about:

#1. frontal thrust density and top speed capabilities, 

#2. how the same inlet components are used quite differently in ramjet versus turbojet installations, 

#3. why achieving combined cycle engine designs can be so difficult,  and

#4.  how heat protection is the true driving issue for high-supersonic and hypersonic flight.

There are other articles posted here and available elsewhere,  that go into considerably more detail about these topics.  But this one tries to illustrate the basics,  to get started.

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Get Acquainted Info: High Speed Vehicles  

There are many concepts to understand about high-speed flight.  Frontal thrust density is a very important issue.  And,  there is no “magic” to waveriders.  See these 2 illustrations:

 

The number of propulsion nozzles at the back of a vehicle also seriously affects frontal thrust density.  This applies to both rockets and airbreathers (of any type).  See:

The over-simplified behavior of inlets on a supersonic ramjet vehicle is shown: 

 Bear in mind that pitot-normal shock inlets,  which have no shock-on-lip behavior,  actually have 6 behaviors to understand,  and external-compression feature-fitted inlets have 9 different behaviors to understand.   You do not initially need to understand all that detail!

But,  it is the basic as-illustrated inlet behavior above,  that drives supersonic ramjet performance.  Ramjet takeover from the booster needs to occur no lower than shock-on-lip speed.  The lower the shock-on-lip speed is,  the smaller the booster can be,  leaving more room for ramjet fuel and the nonpropulsive items.  Considerably higher speed is still efficient:

For supersonic flight,  gas turbine engine installations use the same supersonic inlet components,  but they use them quite differently!  These are usually low-bypass “turbojets”,  and they are usually fitted with afterburners. 

Unlike the ramjet,  which when operating properly,  accepts a fixed scooped air massflow from the inlet,  the turbojet demands a variable air massflow corresponding to its rotor speed(s),  determined in turn by the throttle control setting.  The turbojet inlet has to vary the captured air massflow to match engine demand,  which inherently requires subcritical inlet operation,  with variable-but-significant amounts of spillage around the cowl lip. 

The dominant pressure-rise feature in a turbojet installation is the compressor,  not the inlet!  (The only pressure rise feature in a ramjet is the inlet.)  See:

High speed flight involves lots of aero-heating.  Adjacent and captured air temperatures are high.  As you go hypersonic,  shock impingements multiply heating rates substantially.  See:

Shown just below are the heating rates to,  from,  and within,  any given piece of exposed material.  There is steady-state equilibrium (applicable to hypersonic cruise),  and there is transient behavior (applicable to atmospheric entry),  to worry about. 

Radiation occurs only when there is a view of something hot or cold from the affected surface.  The emissivity “e” can make radiative transfer either inefficient if low,  or efficient if high.  It varies between 0 and 1.  (The sigma represents Boltzmann’s constant.)

For convective transfer,  heating rates can be to,  or from,  the surface.  The “film coefficient” h is larger near stagnation zones,  and smaller on lateral skins.  The values of h all decrease as the air thins drastically at very high altitudes. 

Thermal conduction can be to,  from,  or within the piece.  The conduction within acts to set the temperature distribution of the piece from one end to the other.  The other two determine how much heat enters or leaves the piece.  See:

It should now be obvious that the main enabling factor for high supersonic,  or especially hypersonic,  flight is really thermal management,  more so even than propulsion.

And “scramjet propulsion”,  whether combined-cycle or not,   does not make your job any easier,  because it is geometrically incompatible with ramjet and gas turbine,  including even most of the inlet.  In fact,  combining any of these propulsive cycles,  including rocket,  is difficult at best,  because of the severe geometric incompatibilities,  not to mention the speed-of-application differences.  See:

The two that do combine well are rocket and ramjet,  for the “integral rocket ramjet” (IRR):

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On Pressure Vessels

One can build an unpressurized liquid tank to any desired shape,  unless it is so large that the depth of liquid inside exerts a significant pressure all by itself.  The same is not true of pressurized vessels,  even small ones,  beyond a “single handful of psi” gage pressure inside it.  There are very good reasons these tanks are only made in spherical shapes,  or as right circular cylinders with hemispherical or spherical-segment ends. 

This document presumes the reader knows what “gage pressure” means,  and what mechanical stresses and strains are.

Figure 1 shows a cylindrical pressure vessel holding gage pressure P inside.  This could be a tank,  or even a pipe or a tube.  The “hoop stress” is given by the formula shown top right,  which is sometimes referred to as Barlow’s pipe stress formula,  which is a measure of the stress trying to  split the cylinder open longitudinally.  Technically,  you use the cylinder inside diameter for this formula (a requirement working with pipe and tubing),  but as long as the material thickness is small compared to the diameter,  something true of rocket cases and propellant tanks,  there is little-or-no perceptible difference between inside,  outside,  and average diameters. 

Bottom left,  the cross-section of the cylinder,  or any section through a sphere,  is shown.  For a cylinder,  the “axial stress” is that which resists the pressure trying to part one end of the vessel from the other.  For a sphere,  it’s just the “membrane stress” that resists splitting the sphere apart,  no matter the section orientation.  That stress works out to be just half the hoop stress,  as shown.

Figure 1 – Pressure Vessel Stresses In the Vessel Material

For a cylindrical pressure vessel fitted with hemispherical ends,  Figure 2 takes this notion a bit further,  showing as it does the stress distributions on a tiny patch of material,  located on either side of the joint.  The axial stresses match,  while the hoop direction stresses differ by a factor of 2.  There are (at least theoretically) no stresses perpendicular to the material itself (the radial direction from the cylinder axis,  or from the hemisphere’s center). 

This hoop stress mismatch at the joint corresponds to a strain mismatch in the circumferential direction,  in turn corresponding to a radial displacement mismatch between the cylinder material,  and the end (or “head”) material.  The cylinder swells radially under pressure twice as much as the end or “head” swells radially,  as measured at the joint.  This distorts both the cylinder and the head locally at the joint,  as the materials bend locally,  in order to try to stay joined.  This induces large bending stresses locally,  which add in certain ways to the hoop and axial (or membrane) stresses already described.  That makes the joint quite vulnerable to local overstress failure. 

The ”fix” for this is to locally thicken the cylinder and head materials at the joint.  In effect,  the extra material “sops up” the extra imposed stresses.  For boilers,  there are very specific guidelines for how much local thickening is needed,  and how far “local” extends away from the joint.  Those rules are the ASME boiler code,  which is legally mandatory everywhere in the country,  for designing and building boilers.  Every provision represents a life lost learning that lesson.  This is serious business!

Figure 2 – There Are Distortions With Extra Stresses At the Joints

There are choices allowed for how to implement those localized thickenings at the joints.  Those are depicted in Figure 3,  and also apply to solid rocket motor case designs.  You can increase the thickness toward the inside,  or toward the outside,  or even some of both,  just as long as enough  extra material is supplied.  Finite-element stress-strain analysis can refine this further.

The same figure also shows a variation on the spheroidal end,  where only a sphere segment is used as the end membrane.  This requires a connection ring that resists radial swelling at about the same rate as the end resists radial swelling at its attachment joint.  That way,  no thickening of the membrane is required,  the ring supplies that for the membrane.  You still need a local thickening of the cylinder material at its attachment to the ring,  because of the radial swelling mismatch.

This spherical segment and ring approach lets one enclose more volume within a given length,  without making the assembly any heavier than a full hemispherical end.  This is how most solid rocket motor case closures are designed.  It is a well-proven solution.  The flatter the membrane,  the heavier the ring gets,  though.  It’s a trade-off.

Figure 3 – Full Hemispherical Ends Vs. Spherical Segment Ends With Rings

There are many possible reasons for wanting to use other shapes for one or both cylindrical pressure vessel end heads.  As long as membrane stress is insignificant (meaning very low pressure indeed!),  you can do that!  But as soon as the pressure (and the stresses it induces) become significant,  those other shapes rapidly become infeasible.  This is shown in Figure 4,  where the two spheroidal options are depicted,  along with elliptical and conical ends.  A flat end fails even worse than the elliptical,  from similar stresses that are just higher,  plus a sharp corner effect that locally greatly magnifies the local stresses even further,  right at the corner.

There is one positive benefit to a conical end,  if an axially-directed load must be carried from the cone tip into the cylinder.  This is an efficient load path for such a load.  But it is still a lousy pressure vessel choice!  It will require a lot of internal stiffeners to keep it from trying to “go round”.  Those are going to add significant weight,  there is no way around that problem!  You must trade off the axial load path advantage against the big weight gain incurred to make a conical shape a pressure vessel.

Figure 4 – Which End Shapes Work and Which Do Not,  and Why

It is a common belief that an elliptical shape is as good as the ring and spherical segment.  This is not true from a pressure vessel design standpoint,  as the figure shows.  The elliptical head will try to “go round”,  inducing severe bending stresses.  It is volumetrically efficient,  which is why many unpressurized railroad tank cars use elliptical heads on cylindrical bodies.  But these tank cars are not pressurized!  Their evident abundance is deceptive regarding the pressure vessel issues.

That same effect is why you want to use circular cylinders as your basic tank body for a pressure vessel,  not some elliptical (or other) cross section shape.  Those other shapes will try to “go round” upon pressurization,  leading to enormous bending stresses and very rapid failure.  If you must put pressurized storage within some oddly-shaped volume,  you must fill it with multiple small circular cylinders!  The non-circular cross section is not,  and will never be,  a successful pressure vessel design!  This is why air mattress floats are made the way that they are,  for example:  multiple cylinders,  connected at the ends so as to fill together all at once.

Addendum:  Exact Analysis

The exact analysis for right circular cylinders and spheres is given in Figure 5,  along with the geometries that allow these formulations to be made from very simple measurements.  

Figure 5 – Exact Formulations

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Ramjet Ablative Liners

In March 2024 I was an invited speaker at an American Carbon Society meeting held at North Carolina State University in Raleigh,  North Carolina.  I brought 3 presentations to that meeting,  only one of which could presented live.  The other two were converted on site to poster presentations,  so that all three were gen in one form or another.

My report on the meeting is given in the posting on this site dated 23 April 2024 and titled “Presenter at Workshop”.  The live presentation and associated text document was about using old-time by-hand analyses for initial concept screening,  to enable a real brainstorming process,  without the expense of creating an actual design and multiple computer models,  for each and every concept.  The example for this was by-hand calculations of re-entry dynamics and conditions.  That presentation is also documented on this site as the posting dated 3 May 2024 and titled “Entry Concept Screening”. 

The other two presentations,  converted to poster presentations,  existed as potentially-live presentations and text documents when I went to NCSU.  One had to do with ablative ramjet insulations that I tested in actual ramjet direct connect tests a few decades ago.  The other was an update to a previous presentation and posting regarding a low density ceramic composite burner liner that I tested a few decades ago.  That earlier posting was dated 18 March 2013 and titled “Low-Density Non-Ablative Ceramic Heat Shields”.

To quickly find any posting here on this site,  all you need is its posting date and its title.  Use the blog archive tool on the left side of this page.  Click first on the year,  then on the month,  then on the title if need be (if more than one posting was made that month).

To see an enlarged figure in any given posting,  click on the figure.  There is an “X-out” button top right,  that takes you right back to the article.

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Presented here is the text document of the ramjet liner presentation,  based on those tests long ago. 

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Ramjet Ablative Liners               GW Johnson        12-24-2023

Abstract

For several years,  the author worked on the Hercules-McGregor plant’s “Airbreathing IR & D” project,  where “IR&D” means “Independent Research and Development”,  and is reimbursed by the government.  The focus of that project was a fuel-rich solid-propellant gas generator-fed ramjet,  which included an integral booster housed within its combustor.  

During 1989-1992,  a series of ramjet tests were performed in the McGregor direct-connect test facility,  that addressed 3 objectives.  First,  the combustion performance of multiple experimental fuel propellants was evaluated.  Second,  a fuel rate control approach with no moving parts was evaluated.  Third,  multiple possible alternate combustor ablative insulation materials were evaluated,  as possible alternates for Dow Corning’s DC 93-104 silicone ablative,  which is the subject here.

Presented here are the heat protection results from that testing series.  Included are a unique way of extending the time that heat protection can be obtained,  and the means by which chemically incompatible materials can be bonded together. 

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. 

Background:

Figure 1 illustrates how drastically-different the conditions are in a ramjet,  versus those in a solid rocket.  It was already known that the types of insulation that work in rockets simply could not cope with the heavy fluid shear “scrubbing” and very long burn times in the ramjet. 

The materials commonly used in rocket cases are randomly-oriented fibers reinforcing some type of rubber.  In the old days,  asbestos fibers were used,  replaced in recent years with things like Kevlar fibers.  One common rubber was EPDM (ethylene propylene diene monomer),  easily compatible with cast composite propellant binder systems.  These are shown in Figure 2,  along with the candidates being considered for ramjet testing. 

These materials were usually B-staged as partially-cured sheets,  so that a layered wrap upon a mandrel could be inserted into a primed case,  that mandrel inflated to exert pressure against the case wall,  and then heat-cured to vulcanize the rubber.  This was very cost-effective processing,  but the materials were inadequate for ramjet application. 

Figure 1 – Rocket Vs. Ramjet Conditions

Figure 2 – Candidate Materials

Replacing the randomly-oriented fibers with layers of fully-woven cloth was the approach that was hoped to be adequate,  while preserving the preferred cost-effective processing.  We already knew about Dow Corning’s silicone material with the higher pyrolysis temperatures,  so we switched to their poly dimethyl silicone (PDMS) polymer,  and attempted carbon and silica fabric reinforcements. 

Dow Corning’s DC 93-104 ablative is PDMS polymer loaded with 3 solids.  Two are silica and silicon carbide granulates,  the third is carbon fibers in random lengths. This is a very “thick” (viscous) material that can be troweled,  or pressure-cast around installation tooling. 

We had significant experience with it in the ASALM-PTV program,  where we insulated with DC 93-104 and then cast integral boosters,  into 20-inch OD combustors,  for ramjet propulsion designed by the Marquardt Company.  We applied this same experience to our 7-inch OD ramjet engine.

We also identified a Japanese more-or-less equivalent to DC 93-104:  Type 0 Shin Etsu. It processes identically to DC 93-104.  We also tested it.

About the Ramjet Testing:

Over several years,  we also obtained a lot of experience and hardware for a 7-inch OD ramjet engine being considered for an advanced propulsion replacement for the AIM-120 AMRAAM missile.  This used the same DC 93-104 liner that ASALM used.  Figure 3 shows the geometry and dimensions of the ramjet engine,  and typical test conditions.  There is even an approximate thermal analysis plot of the temperature distributions expected from the flame.  There are drops across the thermal boundary layer,  across the char,  across the virgin,  and finally a tiny one at the case.  

Figure 3 – Typical 7” Test Conditions

Note that the thermal conductivity of the char is about factor 3 times that of the virgin material,  leading to the strong temperature gradient slope change at the pyrolysis zone.  Despite this,  the char is still more of an insulator than any sort of thermal conductor.  That is an important result!

Such tests in full flight-weight 7” hardware usually burn 30-60 seconds,  but sometimes can run longer,  which is challenging even for DC 93-104.  That problem was solved on ASALM,  and that same solution was used for the AMRAAM engine.  The IR&D tests being shorter-burn tests,  we did not need to use that solution in these IR&D tests.  There is more information about that solution below. 

Figure 4 shows a color-highlighted table of some 8 tests conducted for 3 different reasons:  (1) to test experimental fuels,  (2) to test an experimental fuel control technique (unchoked-throat self-throttling),  and (3) to test and compare experimental and “stock baseline” combustor insulations.  

Figure 4 – 7” OD Test Results

 

Of these,  the first test was an unintended no-burn using the PDMS-carbon cloth insulation,  leaving it intact.  It was re-used “as-is” in the second test (blue),  a short-burn run with a “clean” fuel. 

The third test was another “clean” fuel short-burn test) with the PDMS-silica fabric insulator (green.  It was in such good condition,  we re-used “as-is” in the fourth test (also green),  another short-burn test of a highly-metallized fuel. 

The 5^th test was a long-burn test of a low-boron fuel,  using DC 93-104 as the case insulation (orange). 

The 6^th test was another low-boron fuel test,  this one a short burn,  using the Shin Etsu “clone” of the Dow Corning material (yellow).  The 7^th test was a longer-burn test of an older baseline fuel with the unchoked “throttle”,  and a re-used Shin Etsu insulator (again yellow) from test 6. 

The 8^th test (no color) did not give us any useful data,  it being almost-a-no-burn,  with a very disappointing highly-experimental fuel,  on another DC 93-104 insulator,  not a re-used one.

Specific Test Results:  PDMS/Carbon Fabric

Figure 5 shows our hybrid flight-weight combustor/heavyweight lab motor hardware,  as mounted on the thrust stand in the Hercules-McGregor direct-connect test facility.  This is a hybrid,  using a flight-weight combustor with a heavyweight lab motor as the fuel-rich solid propellant gas generator,  in a gas generator-fed ramjet.  (ASALM was a liquid fuel ramjet).  IR&D stuff is never “pretty”,  and this one certainly is ugly-looking.  Figure 6 shows the post-burn appearance of the insulator.  It was mostly used-up after only an 11 second burn.   That was very disappointing,  as we usually got over a minute out of DC 93-104,  even without the very-long-burn solution.

Figure 5 – Hybrid Test Hardware:  Flight-weight Combustor and Heavyweight Lab Motor Gas Generator

Figure 6 – Results From One PDMS/C Fabric Test

 

Specific Test Results:  PDMS/Silica Fabric

Figure 7 shows the white textured appearance of the as-built PDMS/Silica fabric insulator.  Figure 8 shows the  post-burn appearance after two short-burn tests,  one after the other,  on the same combustor insulation.  The first was a 10-second short burn repeat “clean” fuel test.  The second was a 15 second short burn of a very-highly metallized boron-titanium fuel.  Total accumulated burn for the two tests was only 25 seconds.  There was not enough liner left to risk a third re-use.  Again,  that was disappointing. 

Figure 7 – Pre-Test Appearance of the PDMS/Silica Fabric Insulator

Figure 8 – Post Test Appearance of the PDMS/Silica Fabric

 

 

Most of the 7-inch IR&D tests looked like what is shown in Figure 9 for the boron-titanium test.  The fuel is 28% metallized boron-titanium.  The gas generator throat is choked.  The liner is the PDMS/Silica cloth being reused “as is” from the previous test.  The sparklers are not from the fuel,  they are coming from the liner as it erodes.  Otherwise,  the clean plume is quite amazing for such a highly-metallized fuel.  Incidentally,  this was the very first time anyone ever burned high-percentage boron efficiently,  in a ramjet engine!

Figure 9 – Typical 7” Test Appearance:  BTi/PDMS-SiO2

 

 

Specific Test Results:  DC 93-104

Figure 10 and Figure 11 are post-test,  just two different views.  This was a long-burn test of a low-boron fuel,  rate-controlled by the unchoked throat.  Note the typical “mud crack” cracking pattern,  which is what we usually see with DC 93-104 liners.  With that kind of cracking,  it is abundantly clear that the char layer is very strongly held by the virgin material beneath.  That would be the effect of the randomly-oriented carbon fibers connecting the char layer to the virgin.   There was enough left after this 40 second test to possibly have risked a short-burn reuse,  but we did not do that. 

Figure 10 – DC 93-104 Post-Test Appearance, one view from rear

Figure 11 – DC 93-104 Post-Test Appearance, another view from the front

 

 

Specific Test Results:  Type 0 Shin Etsu

Figure 12 and Figure 13 show two post-burn views of this insulator.  The first one is after a 15-second short burn with a low-boron fuel,  unchoked.  The second one is after a 42-second long-burn unchoked-throttle test with an old baseline fuel,  in which the same Shin Etsu liner was reused “as-is”.  Total accumulated burn was 57 seconds,  about what we usually get from DC 93-104.  

Figure 12 – Type 0 Shin Etsu Post-Test Appearance,  after low-boron short burn

Figure 13 – Type 0 Shin Etsu Test Appearance,  after re-use in a long-burn unchoked throttle test

 

There was some instability in the 42-second burn,  resulting from a flow-control aero-grid failure in one of the 2 inlets.  Despite this flow asymmetry,  the burn was “good” and the liner survived the pressure pulsations. 

What we saw with Type 0 Shin Etsu was poorer char-retention strength atop the virgin material,  but with a “slicker-looking” and apparently-harder char layer.  It would generally shed most of the chunks of char layer after the actual burn,  during air rundown,  not during the burn itself!  This was seen as a “puff” of black in the air-only plume after the flame went out. 

The Extended-Burn and Chemical-Compatibility Solutions:

These liners (and the one in ASALM) were only about 0.20 inches thick.  In the 7-inch hardware,  we would get about a minute’s burn with DC 93-104 before the charring reached the case wall,  releasing the char layer to break up into chunks and fly downstream.  Only the Type 0 Shin Etsu matched that duration performance.  Both are thick pressure-castables. 

To get a longer protection duration,  one has to mechanically retain the char,  after the virgin material bonded to the case wall has been fully pyrolyzed away.  The means to do this is illustrated in Figure 14,  and was developed at Hercules-McGregor on the 20-inch ASALM program,  working together with Marquardt.   If you take thin strips of stainless steel about 0.1 inch wide,  kink them on a 1 inch spacing to extend up about halfway through the liner,  and spot-weld them to the case spaced about an inch or so apart,  this will retain the char layer in-place,  long after the virgin has charred through.  The kink tips are buried deep enough inside the char so as not to overheat.  

Figure 14 – Details Matter:  Long-Burn and Propellant Compatibility Solutions

 

This worked for up-to-15-minute burns in ASALM!  It was adopted for the 7” AMRAAM engine as well,  on the contract programs.  We think it might help address char chunk loss with the Shin Etsu,  as well.

The other issue with PDMS silicone is chemical incompatibility with composite propellant binders,  very important if integral boosters must be packaged within the combustor!  This problem was also solved on ASALM.  One has to chemically isolate the materials with an inert separator sheet,  yet still retain good adhesion to both the ramjet liner and to the propellant.  This was achieved with thin Teflon film (essentially a large size “Saran Wrap”),  acid etched on both sides to provide cleaned surfaces with rough texture. 

This separator is bladdered onto a liner surface primed with DC 1200 from Dow Corning,  and cured in place.  Then the exposed inside surface can be primed appropriately for the propellant binder system,  before propellant is cast and cured.  The bonds are strong,  yet the chemical isolation is absolute.

Conclusions:

DC 93-104 is the best by far,  compared to bladderable fabric-reinforced rubbers

               For burns > ~ 1 minute,  must use the kinked strip retention system

               Kinked strip retention system tested in 20” diameter for burns up to 15 minutes

 

Type 0 Shin Etsu is almost as good as DC 93-104

               Char-virgin strength is weaker

               Sheds char chunks during air rundown

               Retention strips might ease these troubles

As the conclusions indicate,  the greater pressure-cast processing and separator-sheet preparation efforts are worth the better protection afforded by DC 93-104,  especially with the post-char-through protection afforded by use of the kinked retention strips.  The kinked retention strip detail is not something Dow Corning thought of,  but it makes their product perform very much longer! 

The strips would likely enhance the performance of the Japanese material,  as well.  It is not an exact clone of the Dow Corning material,  with the apparent weaker char layer retention,  but we believe the strips might help correct that,  based on these IR&D test results.

References:

#1. Dow Corning Product Information Sheet “Dow Corning 93-104 Ablative Material” ,  available as a pdf file from their website.

#2.  MSDS (Material Safety Data Sheet) for DC 93-104 kits,  MSDS number 000001189166,  issued 2 April 2015,  last revised 6 April 2015,  available from Dow Corning as a pdf file.

#3. US Patent 3,623,904 “Elastomeric Composition Containing Silicon Carbide for Use as an Ablative Coating”,  issued to James A. Ramseyer and assigned to Dow Corning Corporation,  30 November 1971.

#4. Cheryl L. Resch,  “Ablation Models of Thermal Protection Materials”,  published in the Johns Hopkins APL Technical Digest,  Volume 13,  Number 3,  dated 1992.

 

Presenter at Workshop

I attended the American Carbon Society Symposium and Workshop on thermal management,  held at North Carolina State University,  March 18 and 19,  2024.  I was an invited speaker at that conference,  and I took 3 presentations,  ready to present.  Only one was “live”,  we turned the other two into poster presentations on-site. 

The live presentation (photo) had to do with old-style by-hand methods of design analysis being used up-front for concept screening,  to enable efficient use of the concept brainstorming process to increase chances of project success.  That enables concentrating the real design efforts and heavy-duty design analyses with software packages,  to be reserved for only the one or two best concepts,  thus using resources and schedule time efficiently.   Analysts who can do this sort of by-hand analysis can also more readily-recognize “garbage-in,  garbage-out” problems with software packages!

The specific example used for this presentation was the old H. Julian Allen by-hand simplified re-entry analysis,  used for warhead design about 1953,  and declassified in the late 1950’s.  I have taken that analysis and re-implemented it in the form of an Excel spreadsheet file,  with worksheets representing the atmospheres of Earth,  Mars,  and Titan.  Those models came from the Justus and Braun paper regarding entry,  descent,  and landing,  presented several years ago. 

There are only 4 easily-estimated pieces of data required to represent the entering object:  its speed at entry interface,  its trajectory angle below horizontal at entry interface,  its hypersonic ballistic coefficient,  and its effective nose radius which determines how bad the stagnation heating will be.  The altitude at entry interface is part of the atmosphere model.  My spreadsheet creates plots,  the most useful of which tell you the peak heating rate,  followed closely in time by the peak deceleration gees. 

Those lead easily and immediately to the peak pressure on the heat shield material,  and (by way of a thermal balance) the surface temperature that must be withstood.  Those in turn constrain your material selection.

The other two presentations were about a unique ceramic composite heat shield material I created out of essentially hardware-store materials,  decades ago,  and about the ramjet combustor ablative materials I tested decades ago,  in a particularly-productive direct-connect test series.

The spreadsheet entry analysis,  and others about orbital mechanics,  compressible flow,  high speed heat transfer,  and rocket engine performance,  are all things I can make available.  Contact me.

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Up to this point,  I was able to post the same remarks on LinkedIn and stay within a 400 word limit.  Here on “exrocketman” I can say more and provide more informative detail. 

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The entry spreadsheet uses worksheets with the atmosphere models already set up for each of three worlds:  Earth,  Mars,  and Titan.  All use the same stagnation heating model.  There are only 4 inputs needed to model an entering vehicle. It generates plots automatically,  but you need to make sure the altitude data in the worksheet do not go past something very close to the Mach 3 point.  Or else you would have to recreate the plots from scratch,  limiting what data you select for plotting to the Mach 3 point,  in order to prevent extreme scale distortion.   This is what the Mars entry worksheet looks like:

This old model is 2-D Cartesian (you have to “wrap” its results around the planet).  The trajectory has a constant angle Ɵ with respect to horizontal,  making it a simple straight line (in the real world,  it will “droop” significantly after the peak deceleration pulse).  It uses a very simple scale-height type of exponential model for density variation with altitude:  ρ = ρ₀ exp[h/h_scale],  where ρ₀ and h_scale are merely the curve fit constants for modeling density in the altitude range of interest.  It presumes a constant hypersonic ballistic coefficient β = M_entry/(C_D A_block),  which for blunt shapes means the entry analysis math assumptions are violated below local Mach 3.  Allen came up with a simple closed-form double-exponential equation modeling speed versus altitude,  under these particular assumptions: 

V = V_atm exp{-C exp[h/h_scale]}, 

where V_atm is the object’s speed at entry interface,  C = 1000*ρ₀*h_scale/(2*β*sin Ɵ),  and the analysis starts downward from h = h_atm,  the altitude for entry interface (a property of the atmosphere model along with ρ₀ and h_scale).  The factor of 1000 converts the customary km units of h_scale to m.  While the equations create results at speeds under local Mach 3,  they are in error for not being hypersonic (β is no longer constant),  and those points should not be included in any reported results or plots.  

Allen used a stagnation convective heating correlation that is surprisingly accurate,  even today.  It is q = Q/A = 1.75 x 10^-8 (ρ/R_N)^0.5 (1000*V, km/s)³,  where ρ is measured in kg/m³,  and R_N is the effective nose radius in meters.  The value of Q/A = q  is measured in Watts/cm².  Its integral with time is in the spreadsheet.  This is convective heating only,  one would have to add a model for plasma sheath radiation heating,  for speeds at entry exceeding about 9 km/s.  That is currently not in the spreadsheet,  but is considered to be negligible at entry interface speeds of 8 km/s and less.  The analysis is summarized in this figure:

Where Do-It-Yourselfers Can Obtain Such Materials

At least the entry spreadsheet,  the orbital mechanics spreadsheet,  and one version of the rocket engine performance spreadsheet,  can downloaded for free,  using links that are on the Mars Society’s “New Mars” forums site:  newmars.com/forums/ 

These are located on that forums site in the “Acheron Labs” section,  under the topic “Interplanetary transportation”.  Scroll down a page or two,  to the thread titled “orbital mechanics class traditional”.  The list of available lessons is in the first posting there.  Subsequent posts have the links to all the lessons,  which are actually located in a drop box on-line.  All three named-above spreadsheets are available from that drop box,  as part of the supplied class materials for this course.  

The course comprises multiple lessons that acquaint the student with classical 2-body orbital mechanics of elliptic orbits,  to include interplanetary transfers,  adds in empirical corrections for losses during launch and when 3 bodies are involved,  acquaints you with entry,  descent,  and landing issues,  then takes up rocket vehicle performance estimation (and the rocket engine performance estimation methods to support it).

Be aware that I have two other courses not available from this New Mars forums site,  but instead directly from me.  One is about compressible flow,  to include flow with losses and with heat addition,  as well as shock waves and expansion fans,  plus the same rocket engine performance estimations as are in one of the orbits course lessons.  The other has to do with high-speed heat transfer,  complete with recommended models for various situations.  Both of these courses are associated with spreadsheets as part of the class materials. 

All these class materials include pdf documents that are essentially texts from which to teach yourself how to do these things.  They include demonstration problems with solutions,  and assigned problems to be worked,  plus solutions to those assigned problems,  for comparison afterwards.   For the already-adept,  there are also slide shows from which you can teach others.

GW’s Ramjet Book

Also be aware that I have offered my ramjet book “A Practical Guide to Ramjet Propulsion” as a self-published item.  Just contact me by email,  it currently comes as a series of pdf files,  which I email to you upon receipt of payment.  I hope to soon have a fully automated site,  with a final single download file for the book.  This is not an academic work,  it is a real “how-to” guide written from my direct experiences doing that kind of work in the aerospace/defense industry long ago.  It deals with plain subsonic-combustion ramjets,  to include integral boosters,  but not ejector ramjets,  combined cycles,  or supersonic combustion.  If that interests you,  please contact me (email preferred).

Other Technical Articles Posted On “Exrocketman”

There are many technical articles on a variety of topics posted here on “exrocketman”,  along with a few things posted on youtube under the channel name “exrocketman1”.  Here on “exrocketman” the blog site,  there is a catalog article posted,  that I try to keep current,  which has these things as a list for each one of multiple topic areas.  This is the article “Lists of Some Articles by Topic Area”,  posted 21 October 2021. 

All you need are the posting date and title of the article you seek,  to find anything quickly on this site,  using the blog archive tool,  left side of page.  Click on the year,  then the month,  then the title if need be (such as if multiple articles were posted that month).  Just peruse the lists and jot down the dates and titles you want to see,  then use the archive tool.

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