Update
23 March 2024:
For the readers of this and other similar articles about ramjet
propulsion, be aware that GW’s ramjet
book is finally available as a self-published item. Its title is “A Practical Guide to Ramjet
Propulsion”. Right now, contact GW at gwj5886@gmail.com to buy your copy.
He will, upon receipt of payment by surface mail or Western
Union (or similar), manually email the
book to you as pdf files. This will take
place as 9 emails, each with 3 files
attached, for a total of 27 files (1 for
the up-front stuff, 1 each for 22
chapters, and 1 each for 4
appendices). The base price is
$100, to which $6.25 of Texas sales tax
must be added, for an invoice total of
$106.25.
This
procedure will get replaced with a secure automated web site, that can take credit cards, and automatically send the book as
files. However, that option is not yet available. Watch this space for the announcement when it
is.
GW is working
on a second edition. No projections yet
for when that will become available.
--------------
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. 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:
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.
------------------------
Update: From AIAA’s “Daily Launch” email newsletter
for 1-20-2022:
Spanish Typhoons Now
Equipped With MBDA Meteor
Aviation
Week (1/19) reports Spain is now the third Typhoon partner
nation “to induct MBDA’s Meteor beyond-visual-range air-to-air missile into
front-line use.”
-----------------------
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
