Wednesday, August 18, 2021

Propellant Ullage Problem and Solutions

For liquid rockets employing free-surface tanks sitting on the launch pad or in thrusted flight,  the propellants in the tanks are pushed by gravity or the vehicle acceleration into the same position covering the drains into the engine pumps.  But between burns,  the vehicle is essentially in free-fall.  The propellants,  because they no longer fill the entire volume,  get pulled by surface tension into multiple spherical globules floating around inside the tanks. 

This situation essentially removes the propellants from the inlets to the engine pumps.  The engines cannot be restarted without correcting this situation.  Nor could a ship-to-ship refueling operation be conducted,  as the drain pipe inlets are dry.  No pump of any kind can pull propellant out of tanks if the pump inlet has only vapor in it.

Initially,  rocket stages needing to ignite in free fall were the upper stages of multi-stage launch vehicles.  The first solution to this problem was adding small solid propellant rocket cartridges called “ullage motors” to the stage.  This is depicted in Figure 1.  The solid propellant cartridges were entirely unaffected by either gravity or its total absence. 

When the ullage motors fired,  their total thrust accelerated the vehicle by an amount equal to total thrust divided by vehicle mass.  After a period of time,  the globules settled into pools of liquid in the tank bottoms,  as depicted in the figure.  This has a long history of success with both storable liquids and with cryogenics,  over fairly short time intervals.  Attitude thrusters can also be used for this.

Figure 1 – The Ullage Problem and the Ullage Motor Solution

A “time constant” for this process would be the time for a globule to fall from one end of the tank to the other,  under the small acceleration induced by the ullage thrust. A figure-of-merit for the settling time into a pool with no voids in it,  would be roughly three times the time constant.  That plus the time it takes to get the engine fully ignited and stabilized,  is the min burn time required of the ullage motors.

The pressure at the main engine pump inlet (for each line) is the pressure inside the tank,  plus an increment that is the depth shown  in the figure times the liquid density times the vehicle acceleration.  That pressure divided by Earth gravity-times-the-density must equal or exceed the “net positive suction head” specified for the engine (or refill) propellant pump.  The pressure in the tank is that of the vapor,  plus that of any injected pressurant gas.

The first ullage motors were solid propellant devices,  but that is not the only way to provide ullage thrust.  Liquid propellant attitude thrusters can be used for this purpose,  if designed to operate in free-fall.  These are usually bladdered-tank systems such as those in Figures 2 and 3 (discussed just below),  and which are almost universally pressure-fed instead of pumped. 

Bladdered Tank Approaches

Bladdered tank designs contain the liquid inside a bladder,  in turn inside the tank.  Pressurant gas injected between the tank and bladder squeezes it to the liquid,  preventing the formation of void space in which free-floating globules can form.  The difference in pressure between the pressurant gas and engine pump inlet drives the expulsion of liquid from the bladdered tank.  This has a long history of success with near-room-temperature storable liquid propellants,  but none with cryogenics. 

Figure 2 – The Bladdered Tank Solution,  Done As Axial Eversion

This can be done in pretty much any geometry,  but such a bladder as it crushes under pressure will crumple and wrinkle,  which significantly lowers the liquid expulsion efficiency of the design.  Expulsion efficiency is defined here as volume expelled / volume loaded.  A way to achieve high expulsion efficiency is to “evert” the bladder,  so that one half of it collapses inside the other half.  If the symmetry of this eversion can be preserved,  the expulsion efficiency can theoretically be very nearly perfect.  There are two eversion geometries for cylindrical tanks:  axial (Figure 2),  and lateral (Figure 3). 

Figure 3 – The Bladdered Tank Solution,  Done As Lateral Eversion

The axial eversion path offers propellant expulsion from a tank end with a centered connection,  but offers a long eversion path,  increasing the probability of asymmetric eversion with wrinkles.  That leads to less expulsion from the tank than desired.  The shorter lateral eversion path offers higher symmetric eversion probability,  for higher expulsions nearer those desired.  However,  it requires side feed and expulsion connections.

Both geometries require the bladder be bonded to the tank on one side or one end.  Both feature a very sharp bend with a very short radius of curvature indeed,  at the eversion point.  The trick for reusability is to ensure the strain at the eversion point is elastic,  otherwise,  the bladder will become the wrong shape,  and will no longer fit the tank correctly.  That guarantees wrinkles and lower expulsion than design.  It also increases the likelihood of bladder failure.  Note also that the eversion point moves!

Key here is very large elastic strain values for the bladder material.  Generally speaking,  with most elastomers,  these strain capabilities are quite large at ordinary temperatures,  but quite low at cryogenic temperatures.  That is why this bladdered expulsion propellant system has historically been used with more-or-less room temperature storable propellants,  but not with cryogens!  If the right material with the right properties can be found,  it would then work with cryogens. 

Corrosiveness of the propellant is also an issue to consider.  This is especially important with the nitric acid systems,  and to some extent with the hydrazines.  Reliability of the design,  especially if it is to be reusable,  is extremely important.  This is especially true for very toxic propellants such as NTO and the hydrazines.  Leaks simply cannot be tolerated.

In any event,  the pressure of the expulsion gas that is fed in must exceed whatever propellant pressure exists within the bladder.  If not otherwise controlled,  the rate of propellant expulsion is proportional to the square root of the difference between the feed gas pressure and the tank pressure inside the bladder.  It is also proportional to the propellant outlet area. 

The feed gas volume flow rate at pressure,  must be as large as the volume rate of flow of expelled propellant,  while still providing that expulsion pressure difference.  This is generally a rather significant pressure difference,  in order to achieve useful expulsion rates.

The bladder material must be able deform easily,  so that it does not resist this pressure difference,  instead just resting against the liquid while the gas moves them both. It therefore cannot be stiff,  despite whatever thickness it must have,  to survive.   We are talking about materials with high tensile strength,  very low Young’s modulus,  and truly enormous elastic strain capability,  at all the cryogenic temperatures it will see in service. 

The radius of curvature at the eversion point,  right on the inside of that eversion bend,  is essentially zero.  That puts the outer side of the bend into considerable tensile stress and strain.  If the material is going to split,  that is where it will happen.  The thicker the bladder has to be,  the worse this eversion point bend-splitting risk is.

Piston-Driven Displacement

About the only remaining alternative approach for positive expulsion would be a gas-driven piston,  essentially a syringe.  This is depicted in Figure 4.  As far as I know,  this approach has never been used in a flying system,  other than as an engine start primer. 

There are multiple constraints on this type of design for a rocket vehicle.  One tank dome has to be inverted,  in order for there to be a place to locate the piston skirt,  when the tank is full,  such that max liquid volume is obtained.  The piston must have such a skirt,  in order to remain properly aligned,  and not jam. 

At the other end of the piston travel,  the piston face must match the contour of that dome,  so that maximum liquid volume may be expelled. 

The piston skirt and side must be of nontrivial thickness in order to house the groove (or grooves) for the O-ring seals.  Remembering that the square edges of the grooves are stress concentrators,  there must be enough “meat” to sustain the loads on the piston,  repeatedly for reusability. 

The depth of the groove plus the wall clearance has to be such that the O-ring seal(s) is compressed radially enough to seal.  That compression distance is small if the O-ring is hard,  but the compression force is high.  The compression force is low if the O-ring is soft,  but the O-ring may not be compressed enough to effectively seal,  if this is taken too far. 

High compression force is high friction force to move the piston.  In particular,  the static friction is typically much higher than sliding friction,  leading to slip-and-jerk behavior characteristics,  which are quite undesirable. 

The width of the O-ring groove has to be wide enough not to compress the O-ring,  so that pressurant gas can fill one side of the groove evenly,  and force the O-ring against the other side,  thus effecting the seal.  This works correctly if there is only one O-ring.

If there are multiple O-ring seals,  pressurant gas cannot reach those seals further away from the pressurant gas side.  If such seals are installed,  their grooves must be narrower,  so that radial compression forces deform the O-ring enough to seal against both sides of the groove.  This requires high radial pressures,  obtaining large friction values,  and causing severe slip-and-jerk behavior.  Net result:  multiple-seal redundancy is not always a good idea!

Figure 4 – Positive Displacement With A Piston

Venting Boiloff Vapors With Cryogens

All the issues and observations made so far apply to near-room-temperature storable liquid propellants.  The boiloff behavior of cryogenic propellant materials introduces yet another very complicated issue to deal with:  adequate venting of boiloff vapors.

In the free-surface tank where ullage thrust is used,  the tank venting system is located on the forward dome.  The tank may be filled fully at launch,  or very nearly so,  but when in free-fall,   there will be considerably more physical tank volume than there is liquid volume inside it.  Vapors add to the atmosphere not occupied by liquid,  which must be vented periodically,  if tank pressure is not to rise rapidly. The forward dome location is the logical place to install such venting equipment. 

When venting,  ullage thrust must be applied to resettle the globules into a pool of liquid,  with a separate vapor atmosphere.  Otherwise,  liquid as well as vapor will be vented. 

We may conclude that free-surface tanks with ullage thrust are inherently compatible with cryogenic propellants.  History bears that conclusion out.  Periodic venting will also require ullage thrust,  in addition to engine ignitions and propellant transfer operations.

Using cryogenics in the bladdered tank approach is going to require a flexible vent line between the bladder and a location on the tank shell.  The equipment to control tank venting can be mounted at the shell location.  But,  there are two very serious problems:  (1) the vent line is quite long,  essentially full tank length,  if axial eversion is used,  and (2) how does one ensure that only vapor enters the vent line,  when vapor can form essentially anywhere within the bladder?

Problem 1 can be reduced in severity by using lateral eversion.  The vent line is much shorter,  but must still be flexible,  and it will affect the everting bladder geometry,  requiring an inconvenient pocket in the tank shell to hold it,  when the bladder is filled.  There is no practical way to put the vent line back into the pocket during tank refill,  without opening the tank at the pocket location,  and physically flaking the flexible vent line in place. 

Problem 2 has no known solution,  yet,  other than the application of ullage thrust.  But if you add ullage thrust,  you might as well just build a simple free-surface tank! 

The flexible vent line is as severe a cryogenic elastic strain capability problem,  as is the bladder itself.  These are technologies requiring development and demonstration.  They are not ready to apply!

We may therefore conclude that neither bladdered tank approach is compatible in any practical way with a boiloff vapor vent,  which is absolutely required if cryogenic propellants are to be used.  There are materials technologies that must be developed and demonstrated to enable this design approach.

The piston displacement approach will require a venting installation on the piston itself,  as there is no other feasible place to put it.  This may prevent the piston from recessing fully into the forward dome when the tank is full,  thus lowering the volume of liquid propellant that could otherwise be loaded into a tank of a given volume. 

It also requires a long flexible vent line from the piston to the forward dome,  where the venting controls can be mounted.  This flexible line will also act to hold the piston off the forward dome.  There is no way to flake this vent line into position between the piston and forward dome during refill,  without opening an access port in that forward dome.

This idea also suffers the cryogenic elastic strain capability problem for the vent line,  which the piston was supposed to eliminate by eliminating the bladder.  And,  it still suffers the same problem with how to ensure only venting vapor,  when liquid is inherently adjacent to the piston,  and the vapor can form throughout the liquid volume.  As with the bladder tank,  you could apply ullage thrust,  but then it would be easier just to build a free-surface tank. 

The flexible line technology at cryogenic temperatures requires development and demonstration.  It is not ready to apply!

We may therefore conclude that the displacement piston approach is also incompatible with a boiloff vapor vent,  which is required if cryogenic propellants are to be used.  There are materials technologies that must be developed and demonstrated to enable this design approach.

Bottom Line:

The most practical solution to the ullage problem when using cryogenic propellants,  is the free-surface tank with ullage thrust provided.  This is also the historically-proven solution,  and all the technologies to enable it are ready to apply.  It is compatible with boiloff vapor venting.  Ullage thrust must be supplied for every free-fall engine ignition,  every tank refill in free fall,  and every boiloff vapor venting event.

The bladdered tank and piston displacement approaches are not compatible with boiloff vapor venting installations,  which are required,  and they run severe cryogenic elastic strain capability risks for the various structures that are required to be flexible at cryogenic temperatures.  Such materials technologies are not yet ready to apply.

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Update 2 October 2021

There is another option for providing an ullage solution in a free-surface tank with cryogenic propellants.  That is to spin the tank to provide “artificial gravity”,  so that a free surface forms again.  The dynamics of spinning objects are stable only about those axes with maximum and minimum mass moments of inertia.  For objects that are cylindrical,  that would be spinning end-over-end (like a baton) at maximum moment of inertia,  or spinning about the long axis like a rifle bullet,  for minimum inertia. 

The two geometries are quite different in their effects. If you spin end-over-end,  the result is as illustrated in Figure 5.  The moving tank walls “intercept” the floating globules,  enforcing their acceleration into the spinning motion.  The result is the formation of a new free surface inside the tank,  reflecting the direction of the centrifugal force of the spinning motion.   It is likely there will be one tank on each side of the center of gravity,  so that propellants are slung to opposite ends of the tanks,  as illustrated.  The drain and vent roles will be reversed for at least one set of tank plumbing connections.

Because the mass moment of inertia is maximum for this spin direction,  a maximum torque-time product is required to spin-up the vehicle to any given rotation speed.  However far the spin-up thruster is from the center of gravity is the moment arm for the torque that thruster provides.  The torque-time product divided by the moment arm length is the thrust time product (total impulse) required of the thruster to spin-up the vehicle for ullage.  A similar total impulse is required to de-spin.  Note that engine restarts cannot be done successfully while the vehicle is spinning,  although propellant transfers might be. 

Figure 5 – What Happens With End-Over-End Spin

The other option is rifle-bullet spin,  as illustrated in Figure 6.  This one slings the propellants radially outward against the periphery,  as shown,  to form a cylindrical free surface inside each tank.  The total impulse required to spin-up and de-spin is less,  because this is the minimum mass moment of inertia.  The pre-existing vent connections can still serve that role in both tanks,  as long as they are near the center of the tank dome.  The drain connections will require a second set of drains along the tank peripheries,  where the liquid pools are located,  in addition to the aft dome center connections,  used when under thrust.  It would best preserve spin stability to install the periphery drain connections in a symmetrical manner.  Otherwise cross products of inertia become large instead of zero.

There is one other problem to solve,  associated specifically with this spin direction.  There are no surfaces that “intercept” the globules when you spin-up the vehicle!  In effect,  you spin-up the hardware,  but not the propellant.  Only the random motions of the globules bring them into contact with the moving tank wall.  Friction forces collect some of the spatter from those collisions onto the wall.  In this way,  eventually the propellant finally gets “spun up” and affixed to the tank wall.  But this random and inefficient process takes a very long time indeed!

The solution is a set of radially-oriented perforated baffles,  similar to anti-slosh baffles.  When you spin-up the hardware,  these baffles “intercept” the floating globules,  forcing their immediate spin-up,   and thus their immediately getting slung outward against the tank wall.  This is shown in Figure 7

This solution costs some extra inert mass in the form of the extra peripheral drain connections,  and the radial perforated baffles.  But it does reduce the mass of spin and de-spin thruster propellant that needs to be budgeted.  Given a fast switch from peripheral to aft dome drains,  this rifle-bullet spin geometry might possibly serve for engine reignitions as well as for propellant transfers.  For transfers,  the very convenient nose-to-nose or tail-to-tail docking geometry could serve.   See Figure 8

Figure 6 – What Happens With Rifle-Bullet Spin

Figure 7 – Spinning-Up the Propellant As Well As the Hardware

Figure 8 – Using Rifle-Bullet Spin For Propellant Transfers Between Two Docked Vehicles

Updated Bottom Line:

The rifle-bullet spin technique could easily supplant the application of ullage thrust for propellant transfers between docked vehicles.  It might not serve as well for engine reignitions,  where simple ullage thrust is already well-proven and easily had.   While this spin technique needs demonstration,  there is nothing here to suggest that it wouldn’t be a successful and short effort.

The obvious application here is the tanker problem for refilling SpaceX “Starships” on-orbit.  Such is not required for low orbit operations,  but it is required for high-orbit operations,  and for outside-of-orbit operations.  That last includes trips to the moon and to Mars (or anywhere else outside of Earth orbit).  This spin technique may help make possible the otherwise-very attractive tanker scenarios I have already explored in multiple other articles on this site.

A not-so-obvious (but still important) application would be for an on-orbit propellant depot facility.  Such a depot would have to serve a variety of vehicles,  and so would have to store a variety of propellant combinations.  Some of those are going to be cryogenics.  The bladdered tank solutions in Figures 2 and 3 apply to the storable materials,  but the cryogenic materials s could be stored in tanks that spin like rifle bullets.  That is an idea worth exploring further,  perhaps in a future article.

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Monday, August 2, 2021

The Ramjet I Worked On The Most

Update 10-1-21:  The choked variable-area throttle valve technology used for the ramjet AMRAAM is documented in “Use of the Choked Pintle Valve for a Solid Propellant Gas Generator Throttle”,  dated 10-1-21,  and published on this same site.

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The ramjet I worked on the most was a solid-propellant gas generator-fed ramjet intended to upgrade the AIM-120 AMRAAM.  AMRAAM is a long-range radar-guided air-to-air missile propelled by a solid rocket motor.  It is 7 inches outside diameter,  12 feet long,  and a bit over 300 pounds at launch.  Typical co-altitude head-on engagements (at middle altitudes) have AMRAAM launch at about 20 nmi range,  and go autonomous at 10 or 12 nmi range-to-target.  The ramjet upgrade allowed that launch range to increase past 60 nmi.  The notation “nmi” means “nautical mile”,  where 1 nautical mile is 6076.1 feet,  same as 1852.0 meters.  See Figure 1.  

Figure 1 – The Ramjet Upgrade Concept for AMRAAM

The ramjet propulsion upgrade for AMRAAM was run out of what was then known as the Aero Propulsion Laboratory at WPAFB,  in Dayton Ohio,  as a series of “6.2” applied R&D programs to determine what was feasible and what was not. These eventually led to a “6.3” program to demonstrate readiness for flight test evaluation.  You can think of “6.2” as being applied research and development (R&D),  and “6.3” as a more sharply-focused sort of engineering development. 

Several contractors variously competed and teamed for these programs:  Rocketdyne/Hercules (the one I worked at,  now closed),  CSD (Chemical Systems Division) at UTC (United Technologies Corporation),  ARC (Atlantic Research Corporation,  now part of Northrup Grumman),  the Marquardt Company (TMC, now closed),  LTV Aerospace (LTV),  and Hughes Aircraft Corporation (HAC).  Of these,  Marquardt had a long history of developing and producing ramjet engines,  all of them liquid-fueled.  CSD was also a liquid ramjet source.  Rocketdyne/Hercules,  ARC,  and CSD were all well-known solid propellant contractors.  HAC and LTV were airframe “primes”.  

Genesis of Ramjet AMRAAM 

This ramjet AMRAAM effort (and some others) were sparked by the appearance of the Soviet solid gas generator-fed ramjet surface-to-air missile known in the west as the SA-6 “Gainful”.  This missile used (1) gas generator-fed ramjet propulsion,  (2) a solid propellant rocket integral booster housed within the ramjet combustion chamber,  (3) a means to obturate the air inlets during boost,  and (4) an ejectable booster nozzle to get best performance out of both the booster rocket and the ramjet sustainer,  which otherwise have vastly-incompatible nozzle geometries. 

The SA-6 first appeared in public,  in the 1967 May Day parade in Red Square.  At the time,  the CIA did not recognize it as an airbreather,  classing it as a rocket vehicle with some exaggerated fairings.  Those fairings turned out to be supersonic air inlets for the ramjet sustainer engine.  This was not understood until the 1973 Mideast war,  when it knocked down Israeli Phantoms at 2 to 3 times the range expected for a rocket missile that size.  This was a bit of a technological “Pearl Harbor” for the West.  I worked as lead mechanical engineer in two contracts that exploited this foreign technology under the project name “Group Work”.  This is described in Ref. 1.

The understanding of the SA-6 as a ramjet sparked USAF interest in a ramjet propulsion upgrade for the AMRAAM,  USAF interest in a high-altitude/high-speed ramjet cruise missile denoted as ASALM (“Advanced Strategic Air-Launched Missile”),  USN interest in a ramjet strike missile denoted as ALVRJ (“Air-Launched Low-Volume Ramjet”),  and some others that came later.  There were also many requests for information from several missile primes about possible ramjet propulsion applications. 

My work on ASALM is described in Ref. 2.  ALVRJ rolled out at LTV in the summer of 1974,  when I was a summer hire there,  working on the “Scout” satellite launcher,  while still in graduate engineering school. ALVRJ was a CSD liquid ramjet with an integral booster pushing an LTV airframe and front end.  At that time,  I already had an M.S. degree in Aerospace Engineering,  specializing in high-speed aerodynamics (and aerothermodynamics),  and was starting work toward a Ph.D. degree.  I had passed the written qualifiers with flying colors in all topic areas,  but ran into a roadblock on my oral qualifiers in late 1975.  I ran out of patience and money,  and decided to go to work in industry.  (I got my Ph.D. in General Engineering much later in life.)

How I Got Started In Ramjet Work

I was originally hired at Rocketdyne/Hercules to be an understudy structural engineer,  based on my high performance on the written structural qualifying exam.  I had studied under Ron Stearman for that particular exam;  he was the nephew of the man who designed the famous Stearman biplane,  and the head structures guy in my academic department.  I got started at Rocketdyne/Hercules as a structural engineer on the ASALM-related work that we had to do,  as related in Ref. 2,  among other things. 

It was not long before the program managers at Rocketdyne/Hercules became aware of my background in aerodynamics,  aerothermodynamics,  and general propulsion.  At that point I got “co-opted” to work on a project they had,  toward something termed “ducted rocket”,  which had air inlets.  The “ducted rocket” is really a solid-propellant gas generator-fed ramjet.  Initially,  these were IR&D (“Independent Research and Development”) projects undertaken for later reimbursement by the government. 

Rocketdyne/Hercules had a big IR&D effort aimed at the USAF 6.2 programs for the AMRAAM propulsion upgrade. That is how I met W. H. “Bill” Miller,  who became not only my boss on various IR&D and contract efforts,  but also my good friend.  Same for Sam McClendon,  who was Bill’s preferred project engineer.  Both were University of Texas at Austin graduates,  as I was.

Initially,  there were only a few USAF requirements for a ramjet upgrade to AMRAAM.  It had to stay 7 inch OD (outside diameter) and 12 feet long,  and it could not exceed about 355 pounds at launch.  Otherwise,  the sorts of technologies that could be applied were “wide open”.  That changed later:  toward reduced smoke technologies,  and rocket-ramjet transition technologies that eliminated all ejecta.  This was peculiar to USAF;  USN had no such qualms about smoke or ejecta. 

We at Rocketdyne/Hercules had gotten started (just before I came aboard) with a “cooperative IR&D” effort in concert with Marquardt,  supplying them gas generators to test in their ramjet direct-connect facility,  while we built one of our own.  If you are not worried about characterizing inlets,  that direct-connect mode of testing is the very best,  most cost-effective,  way to test ramjets on the ground.  You can test for the effects of both fuel species and “geometry” upon ramjet performance,  with great fidelity,  in direct-connect mode.  The term “geometry” includes flameholding geometry,  fuel injection geometry,  and overall engine geometry.  That covers a great deal of ground,  as Ref. 3 indicates. 

I was involved in this initial effort in two ways:  (1) running what are called “cycle codes” to predict ramjet performance,  and (2) participating integrally in the shakedown of our direct-connect facility at Rocketdyne/Hercules.  Bill Miller made the initial decisions about what we built,  and he made the right ones,  in my best estimation.   He chose to use a blowdown air supply,  and simple pebble-bed air heat. 

These choices were to reduce costs by eliminating the need for computer-controlled anything,  but they also turned out to offer a very significant advantage from a technical standpoint,  particularly when testing highly-metallized fuels:  we fed real air to the engine,  when the vitiated systems do not.  If the fuel is metallized,  those metals can see the vitiation combustion products (water and carbon dioxide) as additional oxygen content in the “air”,  which leads to erroneous and misleading performance data.

Initially,  we came at this AMRAAM ramjet design with high-magnesium fuel propellants,  same as was in SA-6,  except that ours were castable (the propellant in the SA-6 was pressed).  We were trying to team with LTV as prime and CSD as the ramjet engine maker,  with ourselves in the role of gas generator supplier.  We had some very good magnesium propellants,  which include LPM-212 as an HTPB-binder/AP-oxidizer blend,  and LPM-269,   which used a unique silicone rubber binder,  plus some AP oxidizer,  and about 60% magnesium powder. 

In subsequent years,  I used that same silicone-magnesium propellant as a very reliable and safe-to-handle combustor igniter material.  It also deposited a magnesium-silicate slag on the test hardware’s ablative liner,  that greatly extended its useful life to dozens of tests.

These propellants were roughly 20% AP and 60% magnesium,  with around 3% of carbon black and yellow iron oxide.  These were “smoky” because of the magnesium oxide particulates,  plus some other particulates,  but not nearly as smoky as a “standard” aluminized solid rocket propellant,  because of the air dilution effect of the airflow through the engine!  This not-so-smoky effect had already been seen in the videos taken of the SA-6 in flight during the ’73 war.

Regardless,  the USAF decided they wanted reduced smoke,  and awarded the fixed-flow DR-PTV program to “the other guys” (ARC),  so we began to look further at HTPB-bound,  AP-oxidized fuel-rich solid propellants.  That moved us toward HAC as the prime,  and Marquardt as the ramjet engine contractor,  with Rocketdyne/Hercules as the gas generator supplier.  The LTV AMRAAM upgrade design featured two inlets about 180 degrees apart,  while the HAC design featured inlets only 90 degrees apart.  The inlet performance characteristics are similar,  but definitely not the same.

I don’t know from whom LTV got their inlet recovery data;  HAC got theirs from Marquardt,  as the “AM 149-A-3” inlet design.  The importance of inlet performance and how it dominates ramjet performance is described in Refs. 4 and 5.  Our high-magnesium formulations were designated as LPM-“formulation number”,  while our low-to-zero-magnesium formulations were designated by LPH-“formulation number”.  LPM stood for “Lab Propellant Magnesium”,  while LPH stood for “Lab Propellant Hydrocarbon”.  There was often a suffix number representing the mix number of the same formulation,  initially.  Formulation numbers were 3-digit,  starting at 101.

As it turns out,  the inlet entry symmetry vs asymmetry has a very big effect on what is feasible,  and what is not,  as detailed in Ref. 3although we did not really know this at the time that decision by USAF to go reduced-smoke was made.  We learned it in testing later.  Almost anything in the way of engine geometry works with high-magnesium propellant effluent,  while very little works well,  with low-to-zero magnesium in the formulation.  This is quite unlike the case with liquid fuels

That flameholding issue got complicated by the issue of ramjet combustor ignition,  which often occurred from gas generator igniter debris,  shed still-burning into the combustor in some designs,  but not others!  And it was further complicated by the presence or absence of dedicated combustor ignition devices,  whether pyrophoric liquid injection systems (at Marquardt) or pyrotechnic devices (at Rocketdyne/Hercules).  All of that took a while to sort out,  in experimental tests. 

Early Hydrocarbon Test Details

The first tests with hydrocarbon fuels were done in the Marquardt hardware,  which featured two side inlets 90 degrees apart,  entering at 45 degrees off axial.  Marquardt had a nozzle contraction ratio A5/A4 of 0.67 initially,  and 0.57 later in their tests.  Their inlet/combustor area ratio A2/A4 was 0.56,  similar only in magnitude to the forward dome stepback ratio x/d4 of 0.57.   Combustor length/diameter ratio L/d4 was 6.7.  I no longer remember their combustor inside diameter d4,  but it might have been in the 5 to 6 inch range.  See Figure 2.  

Figure 2 – The Test Geometries for the Early “Hydrocarbon” Fuel Database

Being a liquid fuel ramjet house,  they started with an inlet injection port in each of the two inlets,  where liquids are almost invariably injected.  They also tested a vertical twin direct dome injection geometry,  a horizontal centered twin,  and then the same dual adjacent and dual opposite injection geometries that Rocketdyne/Hercules pioneered (based on flow visualization experiments).

Rocketdyne/Hercules started with a 4 side inlet rig,  entering at 30 degrees off axial,  same as the SA-6.  Inlet/combustor area ratio was similar to that at Marquardt,  at 0.58,  and the forward dome stepback ratio was either 0.12 or 0.55,  set by the presence or absence of a spacer ring between the gas generator and combustor hardware.  The 0.55 value was similar to that used at Marquardt.

The nozzle contraction ratio was smaller,  at 0.37 to 0.44,  depending mostly upon the ablated inside diameter of the test combustor,  which was used for several tests before being replaced.  The as-made inside diameter d4 was 4.6 inches,  with 0.7 inch thick silica phenolic as the ablative insulation. 

Combustor length to diameter could be varied quite strongly in the Rocketdyne/Hercules hardware,  but was almost invariably near 7.6 during the cooperative IR&D tests,  and 6.6 later.  The injection was a single center port with the 4-inlet rig.  It was quite successful with high-magnesium propellants,  but a bit less so with hydrocarbon propellants unless the nozzle were stopped-down,  and the gas generator effluent made rather hot. 

This rig was replaced with a two-side-inlet rig after the cooperative IR&D effort,  made from generalized 3-inlet hardware,  entering at 45 degrees off axial.  There were actually 3 inlet arms,  of which only two were hooked up,  the other being blanked off.  Thus,  either two inlets 180 degrees apart,  or two inlets 90 degrees apart,  could be tested. 

As a two-inlet rig,  inlet/combustor area ratio was similar to that at Marquardt,  at 0.56,  and the most common length/diameter ratio was 6.6.  Stepback ratio x/d4 was either 0.52 or 0.12,  again with or without the spacer ring.  Most tests were conducted with 2 inlets 90 degrees apart.

This is the rig in which single center port injection,  the Marquardt vertical twin,  the dual opposite,  dual centered,  and the dual adjacent injection geometries could be tested,  with 2 inlets 90 degrees apart,  entering at 45 degrees.  Tests with 2 inlets 180 degrees apart entering at 45 degrees,  with a single center injector,  did not fare well with hydrocarbon fuels (those tests are not shown here).  The raw dataset for these tests is given in Table 1.  Conclusions reached are given in Table 2.


Table 1 – Early “Hydrocarbon” Fuel Database


Table 2 – Conclusions Reached from Early “Hydrocarbon” Testing


This generalized rig was then replaced by a closer subscale replication of the inlet divergent passages actually to be used for AMRAAM,  in which the best twin injection (dual adjacent) proved to be about equal to the 5-port injector used on DR-PTV.  The 5-port was really easy to modify (generating a patent for me) for integration with a throttle valve,  and so that combination became baseline for the original VFDR program proposal. 

What we at Rocketdyne/Hercules learned from all these tests,  plus subsequent full-scale tests in our expanded facility,  eventually became the genesis for the side entry flameholding knowledge given in Ref. 3. To this I brought some numerically-substantiated flow visualization results,  plus some perfectly-stirred reactor modeling efforts.  All of that is discussed in that reference.

The propellant formulations tested in the early subscale tests (cooperative IR&D with Marquardt plus the Rocketdyne/Hercules IR&D leading up to the VFDR proposal) were all variations within the same basic formulation family.   These were all AP-oxidized,  with HTPB binders.  Hydrocarbon resin particulates replaced some of the binder in most (but not all) of these formulations.  Any metal-bearing additives replaced some of the hydrocarbon resin content.

The gas generators tested during the early cooperative IR&D effort also featured added hot-gas propellant grains to enhance generator and combustor ignition characteristics.  These were sometimes tube grains inside the nozzle housing,  and sometimes overcast materials added to a trimmed and restricted fuel propellant grain.  They acted to increase effluent temperature during a short ignition transient.  Getting ignition in the combustor depended on this transiently-high effluent temperature,  a high inlet air temperature,  and a near-stoichiometric equivalence ratio,  in that order of importance.

The Rocketdyne/Hercules Test Facility in McGregor,  Texas Grew Over Time

The term “cycle analysis” is a reference to the standard thermodynamics cycle models in those textbooks:  things such as Brayton Cycle,  Otto Cycle,  Carnot Cycle,  and others.  For ramjets,  the math model that gets the “right” answers is one composed of a series of empirical and theoretical component models strung together,  and analyzed with standard compressible flow analysis (which presumes ideal gas behavior).  This is discussed extensively in Ref. 5.

The “typical pressure ratio” models in some textbooks provide good answers for gas turbine machines,  but will generate unrealistic answers for ramjet!  That is because (1) gas turbine performance is dominated by the compressor pressure rise and turbine pressure drop values that are entirely missing in ramjet,  and (2) the only pressure-rise item in a ramjet is inlet recovery,  which equals the sum of all the pressure loss factors,  and all of these are very strongly dependent upon the flow state entering each of them.  “Typical averages” is just the wrong concept for ramjet work! 

Given appropriate inlet performance data and a properly-sized engine geometry,  one can predict for any given flight condition,  the ingested airflow and provided fuel flow values,  along with the total temperature of that ingested air.  Those three are enough to run a very realistic direct connect test,  simply by providing those values of air temperature and flow rate,  and that fuel flow rate.  As long as the inlet is well-known,  all the other variables can be optimized:  the combustor and fuel injection geometries,  the choice of fuel,  and other real-world engineering “details” like heat protection.  Direct-connect testing done this way is far more cost-effective than semi-freejet testing or full freejet testing.

I got started doing the “cycle analyses” to set up tests (and predict system performance) with a series of computer codes supplied by my friends Ken Watson and John Leingang,  at the Aero Propulsion Lab at WPAFB.  Ken wrote these.  They were:

Code                     purpose

AB                         point performance of high speed ramjets

ABTRAJ                trajectory with AB as a propulsion subroutine

RJ                          point performance (improved) with sizing included (high speed)

RJTRAJ                 trajectory (improved) with RJ as a propulsion subroutine

ZTRAJ                   a variant of RJTRAJ set up for running on desktop PC’s

I have since (in recent years) written my own codes for sizing and point performance,  tailored for running on desktop PC’s.  These were written in an antiquated language that I was familiar and conversant with,  that being QuickBASIC 4.5.  I covered both the high speed range that Watson covered (flight speeds never under about Mach 1.6,  up to about Mach 6 max),  and the low speed range (subsonic to about Mach 2 max at most).  These are:

Code                     purpose

RJLOSZ                 low speed range sizing

RJLOPF                 low speed range point performance

RJHISZ                  high speed range sizing

RJHIPF                  high speed range point performance

These codes (whether mine or Watson’s) all have to balance the ingestable air flow into the engine versus what combusted flow will fit through the nozzle,  at the combustor pressure the inlet can deliver.  The adjustment is either by spilling air massflow at the inlet entrance,  or by a deeper shock position,  and stronger shock loss,  in the divergent inlet diffuser passage.  But not both,  and you cannot change that path from one case to the other,  once started. 

That balancing act does not obtain in a direct-connect test analysis!  The flow rates are what they are,  and the combustor pressure (and its nozzle thrust) is simply the result.  Their realism depends upon how good a job you did,  analyzing flight system performance at the flight condition your test simulates.  There is no iterative balance.  Otherwise,  pretty much the same components and compressible flow analyses get used for the combustor,  nozzle,  and divergent inlet passage(s).

Getting good,  reliable performance out of a direct-connect test on the ground,  must address and overcome 3,  maybe 4,  major pitfalls.  Plus a whole host of minor problems. 

The first is transient air system performance:  because of volume storage effects,  what is delivered at the inlets can be significantly different from what is metered upstream. 

The second is thrust stand tare forces:  these have to be experimentally calibrated.  There is no such thing as a tare pressure,  so always believe your pressure-based performance,  and then believe the thrust-based performance,  only if it agrees with your pressure-based performance!   

The third is your theoretical thermochemical values,  which are the benchmark against which you measure the performances you achieve out of your test.  The “gold standard” here is the NASA ODE (One Dimensional Equilibrium) code,  run at the fuel/air ratio and inlet total temperature (and combustor static pressure level) of your test. 

Use the properties predicted by the code;  do NOT use a so-called “process specific heat ratio” for your test analysis!  Doing so is essentially assuming the answer you wish to find!  (I did come up with an easy-to-use convenient approximation that is within about 1% of NASA ODE in terms of combusted c* velocity.  My cycle codes use that approximation.)

The fourth depends upon which kind of fuel you are using:  a liquid,  versus the effluent from a fuel-rich solid-propellant gas generator.  The flow of liquid fuel through any given test rig can be calibrated (with water for safety!) with a stopwatch and bucket.  The flow rate can be corrected from water to your fuel,  with your fuel’s specific gravity. 

The gas generator effluent case requires that a full ballistic analysis be done of the solid propellant device firing.  It must be done to a very high accuracy standard (fraction of a percent),  which requires converging not only the surface-vs-web history and expelled mass,  but also the delivered generator c* history,  and the “real” delivered burn rate curve.  This CANNOT be done real-time during the test,  totally unlike liquid fuel flow rate!  See Ref. 6 for very real-world information about how that works. 

The Rocketdyne/Hercules direct-connect test facility started out small,  and grew over time.  At the time this early hydrocarbon fuel database was created,  it was still quite small:  5 lbm/sec max airflow at 750 F max air total temperature.  This limited us to rather subscale hardware.  We started out with 40 welding gas bottles of air,  as our blowdown air supply,  but soon went to 100 bottles,  as shown in Figure 3,  to get more tests out of a set of bottles. 

The early subscale combustor hardware was based on 6-inch schedule-40 pipe,  with welded flanged connections.  It was insulated with 0.7-inch thick silica phenolic sleeves,  which put the as-built combustor inside diameter d4 = 4.60 inches.  This easily mated-up with both 4-inch and 6-inch lab motor hardware,  as both were made to the same 6-inch welded flanged pipe connections.  A spacer ring,  between the gas generator and the combustor inlet section,  allowed us to easily vary the stepback “x” of the forward dome from the inlet entry station. 

We could vary the ramjet nozzle throat sizes used in the nozzle section.  Eventually,  these became graphite inserts.  Altitude testing was required if the nozzle would unchoke at test conditions.  This was accomplished with a supersonic diffuser pipe to slow the exit plume subsonic,  then a steam ejector pump to raise that subsonic stream’s pressure back to ambient.  There was a rolling diaphragm seal to prevent inducing extra airflow around the exterior of the nozzle housing.  Open-nozzle testing was much preferred,  by far.  

Figure 3 – Initial Subscale Direct-Connect Test Facility at Rocketdyne/Hercules

For the Rocketdyne/Hercules IR&D tests that took place after the cooperative IR&D effort,  but before the original VFDR proposal,  this facility grew substantially,  although in a very cost-effective way.   That growth happened in stages,  before,  during,  and after the original VFDR program.  That growth is shown in Figure 4.  

Figure 4 – Expanded Facility at Rocketdyne/Hercules

The first change was adding a second air line as a cold-air bypass line,  with upgraded regulators and larger metering venturis available.  This took us from a single line at 5 pps max,  to two lines,  each capable of 10 pps max.  The delivered air temperature was rather limited,  as the mass-mixing average of ambient and 750 F max.  That made full-scale testing in AMRAAM-size flight-like hardware possible,  to help win the original VFDR program,  during my first tenure.

The second change was adding a 1200 F pebble bed heater to what was the cold bypass line.  This gave us 950 F capability with both lines flowing full at max heater settings.  This was also accomplished during my first tenure at Rocketdyne/Hercules.

The third change was adding a commercial air tanker truck capability to replace the 100-bottle air supply.  The capacity of the tanker truck was far beyond what could be stored in 100 bottles.  That became the new “standard” for operating this facility,  during my second tenure at Rocketdyne/Hercules.  This supported the intermediate programs,  plus the second VFDR program.

Not shown is the change to automated data recovery.  During my first tenure,  data were recovered analog on magnetic tape,  and played back through oscillographs to create a paper record.  Reduction to engineering units was entirely a manual process.  Only the performance analysis of engineering units data was done with a computer,  as card batch input to a mainframe.  During my second tenure,  this was replaced by digital data capture and processing to engineering units with a desktop-type computer.  The performance analysis was done in that same type of computer,  with a desktop-compatible version of the same analysis code. 

These changes were enough to allow full-scale testing in AMRAAM-size flight-like hardware across a significant portion of its expected flight envelope.  Such was used on the original VFDR program.  And on simultaneous and subsequent contract programs,  including the second VFDR program.  We had flight-like hardware for the gas generator,  the throttle interstage,  the ramjet combustor,  and the inlet divergent passages (complete with choke blocks). 

For the airbreathing IR&D effort during my second tenure,  I had an adapter made that coupled a 6-inch lab motor to the flight-like 7-inch ramjet combustor.  This could be configured either as a choked center injector,  or as an unchoked port on center.  The latter proved to be a very practical,  safe,  and convenient way to test experimental propellants very rapidly!  Using a 6-inch lab motor as the gas generator,  with an internal-burning grain design,  was a really good way to test at full flow rate,  just in a short-burn ramjet test.

The initial subscale capability used two parallel,  vertically-oriented downcomers from the off-stand air manifold pipe to the inlet spider plumbing assembly located on the thrust stand.  These downcomers were short,  straight bellows tubes.  Their tare forces were not small,  but could be calibrated versus thrust level,  pressurization level,  and air temperature. 

The final air feed rig used two horizontally-opposed bellows,  from the off-stand air manifold,  to the air spider on the stand that fed the inlets.  Tare forces were smaller,  but still significant.  They calibrated exactly the same way in terms of thrust level,  pressure,  and temperature,  just with different numbers.  With this rig,  it was routine to see the same performance calculated from calibrated thrust,  as was calculated from pressure.  That routine agreement had never before been had.

I worked at Rocketdyne/Hercules in two tenures:  December 1975-December 1983,  and April 1987-November 1994.  These were separated by a tenure working at what was then Tracor Aerospace in Austin,  Texas.   My second tenure at Rocketdyne/Hercules started in program management,  but I soon returned to engineering.  In a de-facto sense,  I managed all the plant IR&D for the plant chief engineer.  That was budgeted at $1-2 million annually,  funding some 10-20 investigators each year.

My first tenure began under Rocketdyne,  but the plant was purchased by Hercules Aerospace in 1978.  Everything after that was under Hercules ownership.  The reason I left in 1983 was because Hercules insisted on limiting raises to 2-3%,  during years when the inflation rate peaked at 18%!  That amounted to an effective 15-16% salary cut each year!  Tracor hired me for a substantial increase,  and provided substantial raises each year that I worked there.  I returned for my second tenure at Rocketdyne/Hercules at almost twice the salary I had when I left.

During my second tenure at Hercules,  I was the principal investigator for airbreathing IR&D at $300-500 thousand per year.  That effort provided better propellants to VFDR program,  plus a better unchoked generator test technique.  Plus,  I supported substantially the final nozzleless booster design and corresponding propellant development.  And I did a lot of other smaller items.

USAF Programs Oriented Toward Ramjet AMRAAM                                                     

The sequence of programs related to the ramjet upgrade for AMRAAM is illustrated in Figure 5.  I have tried to indicate how the Rocketdyne/Hercules Airbreathing IR&D efforts aided this.  The list of programs is not comprehensive,  because I was not privy to what the other guys did,  especially their IR&D efforts (which I made no attempt to show).  I’m not even sure I got all the Rocketdyne/Hercules programs.  However,  the sense of this is clear.  

Figure 5 – Programs Related to the Ramjet Upgrade for AMRAAM

The IR&D effort under the date 1976 is the cooperative effort with Marquardt.  The IR&D efforts at Rocketdyne/Hercules after we lost DR-PTV to ARC,  to prepare to propose the original VFDR,  are also shown.  These two are the source of the data in Tables 1 and 2.  We did the “Ballistic Improvement and Dual Grain contracts during this interval.  Ballistic Improvement actually led to the original VFDR,  being where we matured the magnesium-bearing versions of our VFDR fuel.  That same IR&D also matured the CA-5-bearing propellants,  and added the SAEB (strand-augmented end-burner ) technology. 

The “other guys” (ARC) I think also had an original VFDR contract to work on their wire-pulling throttle while we were working the variable throat area throttle on IR&D and our VFDR contract.  The wire-pulling throttle proved unreliable (frequently blowing up),  while our variable-area throat throttle proved to be quite reliable.  That is why our VFDR program led to future contracts,  and theirs did not.  

ARC won contract efforts to investigate the unchoked gas generator “throttle” flown by France as “Rustique”,  and (I think) a contract to investigate ways and means not-to-eject port covers.  The port cover work fed directly into the “6.3” VFDR program.  Meanwhile,  we had contracts to investigate boost-sustain grain designs in the gas generator (SFDR,  for Split Flow Ducted Rocket),  and a nozzleless booster contract based on our IR&D work that identified and matured a grain design and candidate propellants.  The nozzleless contract produced the baseline booster for the 6.3 VFDR contract.

We did not get funded by the government for the unchoked-generator “throttle”.  However,  I investigated this on airbreathing IR&D,  and found it quite useful as a very safe way to screen experimental fuel propellants very rapidly.  This work produced an unclassified paper at a classified session,  right after ARC reported the progress on their contract.  We had a real engineering ballistic design analysis based on fundamentals,  and actual test data in full-scale AMRAAM hardware,  for 10 times less money than the value of ARC’s contract.  In contrast,  they had only an approximate analysis,  and never got their subscale test hardware to achieve ramjet ignition.  Our paper created quite a stir.

So,  ARC came into the 6.3 VFDR program with some non-ejectable port cover experience,  and their boron fuel propellant (Arcadene-428) that looked really attractive on paper.  Rocketdyne/Hercules came into that 6.3 VFDR contract with a well-developed nozzleless booster,  a throttle valve and control that was well-verified,  a baseline end-burning grain design and fuel propellant (LPH-453),  plus a second SAEB grain design and higher-energy fuel propellant (LPH-563A).  Plus,  from Airbreathing IR&D,  we also brought two boron fuels and one nonmetallized “clean fuel” that met NATO min smoke requirements.

USAF demanded that we form a joint venture with ARC,  or they would not award the contract (because they wanted the ARC fuel propellant that looked so attractive on paper).  During the 6.3 VFDR program,  there was a “shoot-off” of the various fuel candidates,  observed on-site by USAF,  and scrupulously held under the same conditions in our facility.  The results clearly showed that every Rocketdyne/Hercules fuel propellant provide just about the same high level of actually-delivered performance,  with the ARC Arcadene 428 boron fuel falling significantly short of that same level of delivered performance.  This was at a rather modest altitude. 

About that same time,  Hercules corporate made the decision to close the McGregor plant.  The actual closure happened under ATK ownership,  but the effect was the same:  the other Hercules tactical plant did not want the airbreathing technology or program,  so ARC “inherited” everything via the joint venture.  Once they were in sole control,  the “selected fuel” for VFDR became their underperforming Arcadene-428,  which then promptly failed in direct-connect tests to ignite at middle and high altitudes. 

The reason for that failure was its low “combustibility index”,  a phenomenon that is well-discussed in Ref. 3. None of the Rocketdyne/Hercules fuels had a combustibility index that low,  and so they could all be expected to ignite with air at middle and high altitudes,  and also at the colder air temperatures.  The one with the greatest high-altitude/cold air risk was LPH-563A,  although we had some successful high-altitude test experience with it.  Its combustibility index was twice that of Arcadene-428.

And that proven-persistent high-altitude ignition failure ultimately killed the program to provide a ramjet upgrade for AMRAAM!  The ramjet,  using the underperforming fuel,  was simply not ready for flight test!  After two decades effort and several funded programs,  the USAF decided they had no more money to spend on it.

Where the Technology Finally Went                                                   

ARC-as-inheritor looked at putting the VFDR system into an engine of HARM size about year 2005,  and about that same time sold the VFDR without the nozzleless booster to USN for their “Coyote” gunnery target drone.  Photos of it in flight show a dark plume of unburned fuel (excess effluent soot),  much as would be expected from low combustibility index,  even at very low (sea-skimming) altitude.  The technology has gone nowhere since,  in the US. 

The Europeans fielded a very close equivalent to ramjet AMRAAM that they named “Meteor”,  also about 2005 or thereabouts.  Visually,  it looks very much like the ramjet AMRAAM we were pursuing.  It also used a variable-area throat throttle on its fuel-rich solid-propellant gas generator. 

The post-Soviet Russians had flown in operational flight test a ramjet variant of their AA-12 “Adder” air-to-air missile,  but they chose not to produce it,  because we chose not to fly our ramjet AMRAAM.  Given the recent operational status of “Meteor”,  they now have a motive to produce the ramjet “Adder”,  but the factory that designed it,  Vympel,  is no more.    

Some Typical Test Photos                                                        

I do not have any photos from the early days of those particular tests,  nor do I have photos from the 6.3 VFDR “shoot-off”.  But I do have some good color photos from the airbreathing IR&D efforts that brought multiple additional fuels to that “shootoff”.  These tests were run with airflows corresponding to full speeds at modest altitudes,  in the hybridized hardware that used a 6-inch lab motor as a short-burn gas generator,  for a full scale AMRAAM combustor and inlets.  (This would be at the correct A5/A4 = 0.65 per the full scale design,  by the way.)

The first one of these,  in Figure 6,  shows the nonmetallized “clean fuel” at lean conditions.  This was the fuel rich propellant that used pelletized nitrocellulose instead of magnesium,  or any other metal or metal-bearing combustion aid.  This material unofficially meets the NATO min smoke criteria,  despite using some AP oxidizer.  It can do this because of the air dilution effect. 

It delivered the same performance as all the other fuel candidates in terms of thrust and specific impulse.  It also has a high combustibility index,  indicating reliable ramjet ignition,  even at high altitude,  or with colder air.  This particular hybridized hardware used a sonic (choked) gas generator.

The second one of these,  in Figure 7,  used a version of one of the two baseline VFDR fuels,  this one being the 37% AP formulation with 5% aluminum-bearing CA-5 combustion aid,  for the plain end-burning grain design.  This was the baseline LPH-453 fuel in the original VFDR contract,  tested here at lean conditions. 

It delivered the same performance as all the other fuels in the “shoot-off”.  It also has a high combustibility index,  and a demonstrated history of reliable ramjet ignition at high altitude and in colder air.  This one is being tested with an unchoked gas generator,  using an internal-burning lab grain.

Figure 8 is from a test of one of the two 2.5% boron fuels we developed on IR&D,  in an attempt to raise theoretical heating value without sacrificing much combustibility.  This one also meets the same basic performance levels as the rest in the “shoot-off”,  fired here at lean conditions. 

The boron is encapsulated in an ethyl cellulose binder with some fluorinated graphite,  an analog to the aluminum-fluorinated graphite CA-5 combustion aid.  This test is an unchoked-generator form with an internal-burning lab grain.  It has an acceptable combustibility index.

The very best boron formulation,  with both high heating value and high combustibility,  is the 24.5% metal formulation shown at lean conditions in Figure 9.  In this one,  the boron is added as a blend of boron and titanium powders.  During combustion in the gas generator,  these metals alloy in a very exothermic manner,  replacing substantial oxidizer content without sacrificing chamber temperature or effluent composition (combustibility index).  This test is also an unchoked gas generator with an internal-burning lab grain.

We did make and test a propellant pursuant to an Army initiative,  that used no AP at all,  lots of carbon black (25%),  and a liquid explosive glycidyl azide (GAP) polymer as the binder (75%).  It turned out to have a very low combustibility index,  and performed dismally at lean conditions,  as shown in Fig. 10,  although its theoretical density heating value was quite competitive.  It barely burned at all in the ramjet.  This shows as a very dim tailpipe flame and a lot of unburned soot in the plume.

I have no ground test photo of Arcadene-428 from the shoot-off,  nor could I obtain and test it on airbreathing IR&D.  But it also calculates as a very low combustibility index with a very high theoretical heating value.  In the shoot-off at a rather modest altitude and hot air,  it underperformed significantly,  relative to all the other fuels.  And in the later ground tests during the 6.3 VFDR contract,  it failed to ignite in the ramjet at higher altitudes.  Some pertinent comparison data are given in Table 3 for all the fuel propellants discussed here,  plus the other SAEB baseline fuel,  LPH-563A,  which has 8% aluminum,  and a lower combustibility index.  It performed well at modest altitude,  but underperformed somewhat a high altitude,  yet it did not fail to ignite.    

As fuel combustibility index falls,  first you see performance degradation at higher altitudes,  and with colder air (example LPH-563A at CI = 0.28).  Then you see performance degradation at low altitudes along with failure to ignite in the ramjet at high altitudes (example Arcadene-428 at CI = 0.14).  Low enough,  it won’t burn at all,  even at the most favorable low altitude and hot air conditions (example the GAP-carbon fuel).  This behavior was experimentally confirmed on my airbreathing IR&D effort.

The low-combustibility Arcadene-428 material is the fuel that ARC used in the VFDR system that it sold to the USN for the “Coyote” gunnery target drone. Perhaps the appearance of that drone in flight,  with a smoky black plume and dim tailpipe flame,  should be quite unsurprising!  Especially since these generator effluents do not perform as well in a symmetric inlet geometry (as discussed in Ref. 3).  That appearance is shown in Fig. 11.  Judge for yourself!

                              

 Table 3 – Fuel Propellant Comparison Data at Modest Altitude and Hot Air

References (all authored by G. W. Johnson and located on this site)

#1. 4 Feb 2020,  “One of Several Ramjets That I Worked On” [SA-6 evaluation]

#2. 1 July 2021,  “Another Ramjet That I Worked On”  [ASALM work]

#3.  3 March 2020,  “Ramjet Flameholding” [geometry and conditions,  for liquids and gas generator]

#4. 9 Nov 2020,  “Fundamentals of Inlets” [application to ramjet and to gas turbine]

#5. 21 Dec 2012,  “Ramjet Cycle Analyses” [compressible flow models]

#6. 16 February 2020,  “Solid Rocket Analysis” [internal ballistics with real-world effects]

Figure 6 – AP-HTPB-PAMS-NC “Clean Fuel”,  Lean,  Very Good Combustibility


Figure 7 – AP-HTPB-PAMS-CA-5 (2% Aluminum) Baseline Fuel,  Lean,  CI = 0.71


Figure 8 – AP-HTPB-PAMS-BCFx (2.5% Boron) Fuel,  Lean,  CI = 0.76


Figure 9 – AP-HTPB-PAMS-BTi (24.5% metal) Fuel,  Lean,  CI = 0.48


Figure 10 – GAP-C Fuel Propellant,  Very Lean,  Very Low Combustibility (~0.1?)


Figure 11 --  The “Coyote” Gunnery Target Drone At Mach 2 to 2.5, CI = 0.14