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).

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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.