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.
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. 3, although 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:
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:
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