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.
Update 27 January 2026: It turns out I need to collect 6.75% sales tax, not 6.25%. That makes it price $100.00 + tax $6.75 = bill $106.75.
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.
--------------
One of the many things I got to work on, when I was a defense industry engineer, was the foreign technology exploitation of the
Russian surface-to-air missile (SAM), known to NATO as the SA-6 “Gainful”. In Russia,
this was known as the 2K12 “Kub”
(“Cube”) system, with the missile itself
a product of Vympel (no longer in existence).
It was the Vympel model 3M9
/ 9M9 surface to air missile.
The SA-6 "Gainful" anti-aircraft SAM first
appeared in public in the 1967 May Day parade in Red Square, after the supporting technologies for it disappeared
from the Russian unclassified literature in 1962. It reached Russia’s client states in time for
the 1973 “6-Day War” between Israel and Egypt.
That event led to my involvement,
as described below. It had
heavy, voluminous vacuum-tube type
electronics, that also just happen to be
nuclear-hard, on the battlefield.
SA-6 had been replaced by a rocket SAM with solid-state
electronics in Russia by the 1980’s, but
it remains in the possession of some former client states. It was in the hands of Serbian militias in
the Balkans during the 1990’s Balkan wars,
where it downed the only F-117A stealth fighter ever lost to enemy
fire. It is now considered quite
obsolete, and is on public display as
cutaway items (see
Figure 1 and Figure 2),
yet those few still out there remain quite the potent SAM threat, even today.
In the first figure,
the forward cutaway areas show the seeker/radar electronics. The yellow area forward of the big mounting ring
(not a part of the missile) is the location of the blast-fragmentation
warhead. Aft of the ring one can see the
four supersonic spike-type side mounted inlets and the forward control
fins. The yellow area aft of the fins is
the forward portion of the solid rocket booster that is packaged inside the
ramjet combustor. The fixed tail fins
are aft of the aft mounting ring.
In the second figure,
the yellow zone aft of the aft mounting ring is the aft end of the solid
rocket booster propellant in the aft end of the ramjet combustor. The fuel-rich solid propellant gas generator is
located aft of the forward fins, between
the inlets, and is not very apparent in
either figure shown here.
These public display mockups are not entirely correct. There was no room about the ramjet nozzle in
which to package the actuators for moving surfaces on the aft fins. Control was actually achieved by moving the
forward fins, against fixed aft fins. And,
the booster nozzle was nested inside the rather large ramjet
nozzle, and ejected at boost-sustain
transition (as were the inlet port covers).
At the time of its first appearance in the 1967 May Day
parade in Moscow’s Red Square, the CIA
misidentified SA-6 as a solid rocket system with "exaggerated
fairings". This SAM was really a
ramjet missile with 4 side-mounted supersonic inlets. How CIA analysts managed to make that mistake
still mystifies me to this very day.
Unlike other ramjets of that time (Talos and Bomarc for the
US, SA-4 “Ganef” for the Russians, all SAM’s),
SA-6 “Gainful” was NOT a liquid-fueled ramjet, it did not feature a single nose inlet, and it did not have a staged-off or
“carry-along” rocket booster. It was the
first flying example of something now termed an “integral booster” or “integral
rocket-ramjet” (IRR).
The SA-6 ramjet sustainer was unlike anything seen
before, except in the odd patent. At the time,
it was termed a “ducted rocket”,
although that terminology is vague and subject to multiple incorrect
interpretations. It is more properly
termed a “fuel-rich solid-propellant gas generator-fed ramjet”, or just “gas generator-fed ramjet” with the “fuel-rich
solid propellant” part of that terminology simply understood.
Some History
By the 1973 “6-Day War”,
this Russian system had been given to client states, notably Egypt. Because it had about 3 times the
"legs" of a solid rocket system of that size, and 1950/1960-vintage heavy and voluminous vacuum-tube
electronics, it knocked down Israeli
Phantoms like ninepins, and cost them
about 30% of their air force, in only
those 6 days.
That was a technological "Pearl Harbor" that
really upset the entire West. These
events do happen! (And that one has not
been the only one during my lifetime.)
Once it was understood that SA-6 really was a ramjet, that spawned a lot of ramjet work in the US
defense industry that persisted until about the year 2000. Which is exactly why I was the “ramjet guy”
at Rocketdyne-McGregor, with my training
in propulsion and high-speed aerodynamics.
The Israelis captured damaged SA-6 hardware on the
battlefield, took it apart, made drawings, and tried to understand what it really was. They didn't fully understand what had proved
to be a ramjet, but that was NOT
liquid-fueled! So they went to the CIA
with it.
The CIA went to DOD,
and this foreign technology exploitation project became a joint
Army-Navy program under the name "Group Work". Different pieces went to different
contractors, among them the ramjet fuel
supply and combustor to the plant where I worked in Texas. That plant was then known as Rocketdyne-McGregor, later Hercules-McGregor, and was finally closed under the name ATK. Our contract was run out of Army MICOM in
Huntsville, Alabama.
Being the only front-line working Rocketdyne-McGregor "ramjet
guy", of course I worked on the “Group
Work” project to exploit the SA-6 technology. This was 1978.
I had the Israeli drawings (they were labeled in Hebrew, by the way,
but I can read millimeters) plus examples of the inlet and gas generator
injector dome hardware, and real chunks
of Russian fuel propellant to work with.
My team colleagues were Rocketdyne-McGregor propellant
chemists L. Gale Herring and Jim Muesse,
and an Israeli chemist from Technion: Moshe Gill. Our project engineer was Sam McClendon, and our program manager was Bill Miller. For “chemist”, think propellant formulation chemist.
I was the only working/line mechanical engineer among the 3
chemists. Between the four of us, we actually figured out how this gas
generator-fed ramjet with a pressed-propellant fuel grain really worked, and how it was built. The end-burning fuel propellant charge was a
compressed stack of individual discs. We
duplicated materials, processes, conducted static generator firings, and did airbreathing ground tests of the
ramjet in subscale.
It was my suggestion to “stage” the compression of the
pressed fuel propellant grain that broke-open the path to success. The secret was not to use full compression
forming the individual discs, so that
the stack could then consolidate properly under full compression. Poor consolidation of the stack led to motor
explosions. The chemists found out how
to tailor the burn rates within each disk,
so that burn rate properties could be varied down the stack. I was the one who did the gas generator
interior ballistics and ramjet combustor performance analysis for the tests.
We did so well that we got a second contract (1979) to use
other contractors' results along with our own,
to characterize the actual engagement envelope of the SA-6 missile weapon
system. We used our ramjet results with
other groups’ results for the booster,
the vehicle aerodynamics and weight statement, the supersonic inlet aerodynamics, and a ramjet cycle analysis from Ken Watson
at WPAFB that reflected significant injected axial momentum for the fuel.
I was the working engineer for that second effort, too,
as well as its project engineer. I
did the trajectory predictions and target engagement analyses that
characterized the SA-6 SAM system correctly.
(The Mach 1.5 takeover speed in “Jane’s” is incorrect; it is actually just about Mach 1.8 takeover, for an inlet design with an inlet
shock-on-lip speed of Mach 1.9.)
This intensive effort was just another part of what I did
for ramjet work in general at Rocketdyne / Hercules in McGregor, Texas.
(Plus I also did a lot of solid rocket work.) At the time,
the highest classification of “Group Work” information associated with its
propulsion was “confidential”. It had a
12-year interval to downgrade to “unclassified” (by about 1990 or 1991). Between that,
and the cutaways on public display,
it may be presumed that everything about this propulsion system is
now both unclassified, and entirely in
the public domain. Which means
it is OK for me to publish this.
Bill Miller and Sam McClendon are still with us as of this
writing, although long retired and quite
elderly. Gale Herring and Jim Muesse
passed away long ago. I will be 70 this
coming summer, if I live that long. If chemist Moshe Gill is still alive in
Israel, then he and I are the only two
still alive in the West who know how to make that fuel propellant. It’s actually very good stuff. And I also know how to make a castable
equivalent. At the very least, this high-magnesium stuff makes an excellent
ramjet combustor igniter. And I used it
as such for many years.
I am thus the only one left alive in the West who knows how
that SA-6 ramjet combustor works in detail,
and why flameholding (as distinct from mixing) was NOT an issue with the
magnesium-rich fuel effluent from its gas generator unit. I also know how every piece of that engine
and fuel supply was constructed, and why
they were done the way they were done.
Evaluating all of those things was what our “Group Work” contracts were
all about.
Technical Items
Figure
3 shows a diagram of the inboard profile of an SA-6
missile.
Figure 3 –
3M9M Gainful Cutaway (via Vestnik-PVO)
The items
numbered in the figure are listed below.
The items in red text are things I have added to the original source, based on what I personally know from “Group
Work”.
1 1SB4M CW monopulse
semi-active homing
seeker with Doppler closure rate capability
2 3E27 CW radio two
channel proximity fuse
(30 m nominal radius)
3 3N12 57 kg
blast-fragmentation warhead
4 1SB6M autopilot
5 ramjet intake ducts
6 9D16K sustainer gas
generator charge (67 kg
LK-6TM reducing propellant) 65% Mg 20%
Na-nitrate
7 frangible seals round phenolic caps with O-ring,
held not-shouldered
by shear wires, shouldered
home by boost
pressure
8 boost stage solid propellant charge
172 kg VIK-2
propellant
9 exhaust nozzle ramjet
nozzle very large, booster
nozzle nested
inside on O-ring in ramjet throat,
boattail part of
booster nozzle, eject ass’y also
includes aft
support for cartridge-loaded booster
grain
10 cruciform tail fins
11 cruciform wing
Ramjet Inlets
The ramjet intake ducts were something determined by another
contractor to be what is called a “shock-follower inlet” during “Group Work”. This is a mixed-compression inlet whose minimum
(throat) area is oversized-enough to “start” (swallow the shock system) without
resort to starting bleed cutbacks in the cowl sidewall, or inlet throat bleed slots. Basically,
the inlet throat must swallow all the potential subcritical airflow, in order to “start” easily. (This easy starting is inherent in all of the
all-external compression inlet designs.)
For those who do not know,
a mixed-compression inlet has both external compression and
internal compression features. External
compression is that accomplished by the shock wave system upon the exposed
spike or ramp features ahead of the cowl lip.
Internal compression is that accomplished by the contraction of
flow channel area (and the associated shock wave pattern) aft of the cowl
lip.
There are all-external
compression supersonic inlets with the min (throat) area right at the
capture station, and there are mixed-compression
supersonic inlets with the throat downstream of the capture station, from a practical technological standpoint.
That “shock follower” design approach substantially reduces
supercritical pressure recovery performance,
which is objectionable for the excessive pressure losses at very high
flight speeds. However, max speed in an SA-6 is only about Mach
2.8, thus eliminating that objection! Given that situation, this approach is then a good compromise
design for easy construction, easy
starting, and low cowl lip drag, simultaneously!
The supersonic inlets were made from standard-size round aluminum
tubing available in Russia, fitted with
a protruding spike centerbody positioned upon 3 fins, in turn connected to the duct wall. The axial spike position was set by a tooling
jig, and the centerbody fins then spot-welded
to the outer tube. The lip of the cowl
capture station was beveled to a sharp edge around the external circumference.
These ducts were positioned on the airframe such that best
capture performance was obtained, with all-subsonic
transfer of the air aftward to the combustor forward dome. There,
flow was turned 30 degrees inward,
off axial, for the symmetric
4-port entry into the forward end of the combustor, through the periphery of its forward dome.
This 30-degree turn was accomplished by hot-bending the
sand-filled inlet tubing over a mandrel, after the dome bolt-up fittings had been
spin-welded to them. This process left
small compression wrinkles on the inner surfaces of the turns. But, because the ducts are round, the pressure losses caused by these
imperfections were undetectable, in
terms of post-turn pitot pressure (relative to pre-turn pitot pressure). That is precisely because round ducts have no
corners to hold vortices and separations.
Transition Gear: Inlet
Port Covers
Ramjet combustor pressures are very similar to inlet pitot
pressures: a few, to a few dozen, psia.
Booster pressures are a few hundred to a thousand-or-so psia. That requires inlet obturation during
boost, lest hot booster gas jet out of
the inlets in reverse! The SA-6 design
did this function with totally passively-ejected inlet port covers. No pyrotechnics or controls were needed at
all.
These port covers were hemispherical caps with cylindrical
skirts, convex to the booster chamber, and
sealed with an O-ring into the inlet tubes.
They were held in place with small radial shear wires in the skirts, not quite shouldered-home into the inlet
tubes. Boost pressure would shoulder them
fully home, breaking the shear
wires. Then upon booster pressure tailoff
at burnout (a very rapid bleed-down of pressure to slightly below local
atmospheric), the low-supersonic inlet
pitot pressures at takeover Mach (about 1.8) would pop these port covers out of
their inlet tubes. All four port covers
would easily fit through the ramjet throat simultaneously.
Solid Propellant IRR Booster
The solid propellant booster for the SA-6 was a double-base
propellant in a hollow cylindrical form,
cartridge-loaded, and burning on
both the outer and inner surfaces. It
was mechanically supported at its forward end by the gas generator cover, and at its aft end by the ejectable booster
nozzle structure.
This is a fairly-neutral thrust-time profile created by this
booster grain design approach, as
evaluated by another contractor. At web
burnout, the fast chamber pressure drop
activated a switch, that in turn
simultaneously activated the gas generator igniter and the booster nozzle
ejection pyrotechnics. Both of these
were two-item-redundant initiator systems.
The gas generator igniter was a ring-shaped device located
in the gas generator chamber immediately-adjacent to the grain ignition
surface. It was not quite the same as
the “BKNO3” igniters often used for rocket motors in the West. It was required to generate sustained hot gas
flow washing the grain ignition surface,
for a significant period of time.
Beyond that, I no longer remember
exactly what it was.
The ejectable booster nozzle was a small-throat/large exit
diameter device nested within the ramjet nozzle upon an O-ring seal. It was held in place by 3 “fingers” between
the nozzle assembly and 3 slots cut into the ramjet nozzle exit cone. These “fingers” were held in place by a
confining clamp about the booster nozzle,
but within the external aerodynamic shell of that booster nozzle
assembly.
There were two pyrotechnic charges on that clamp, either of which could open it. Once opened,
the “fingers” and the booster nozzle were pushed aft out of the ramjet
nozzle, by the decaying booster
pressure. From that point, everything was entirely passive.
Ramjet Combustor
The SA-6 ramjet nozzle throat area itself was about 90-95%
of the insulated combustor flow area.
That proportion is utterly infeasible with fuels that require
flameholding with air (such as the liquid hydrocarbons), but is entirely feasible with fuels that are
hypergolic-ignition with air. This
proportion allows very much larger inlet airflow for more thrust, but inherently means that the ramjet nozzle
supersonic expansion is limited indeed. Being
hypergolic, no combustor igniter is
needed to start the burn.
The so-called “hydrocarbon” gas generator fuels, and the liquid hydrocarbon fuels, require flameholding and a combustor igniter
to start the burn. They are anything
but “hypergolic” ignition! They
require throat/combustor area ratios in the 60-65% range for reasons of flame
stability and practical inlet duct size.
The SA-6 combustor is about 12.5 inches diameter, and about 6 diameters long. It is fed by 4 inlet airstreams, and some 12 entering fuel stream ports, as indicated in Figure 4. The more the entering air and fuel streams
are divided, the quicker mixing can
occur. The gentle air entry angle of 30
degrees reduces mixing turbulence intensity,
but this is compensated by the long combustor proportion.
High mixing rates and thus high combustion efficiencies are
thus achieved by the long combustor with multi-subdivided fuel and air
streams. Yet geometries like this are
incompatible with flameholding, for
non-hypergolic fuels. That is not a
problem for the SA-6, with its gas
generator effluent dominated by hypergolic magnesium vapors. But one cannot just substitute a conventional
flameholding fuel into this combustor:
it will not flamehold, which
means it will not burn!
The ramjet combustor is a steel tube lined with a thin layer
of something resembling silica phenolic ablative. Ordinarily,
this thin layer would be inadequate as a combustor insulator, because the magnesium-air flame is as hot as
any other fuel burning with air, and
because of the condensed particles of magnesium oxide product. That cloud of hot particles drives very high
internal radiative heating of the wall. Because the external airflow does not exceed
flight speed Mach 2.8, external
aero-heating is not the driving issue.
Mach 2.8 flight is just not that bad.
Internal heating is the real driving issue for design.
However, the same
condensed magnesium oxide product coats the ablative liner very quickly with a
layer of solid white magnesium oxide “slag”.
Because the stoichiometric air/fuel ratio is about 3, there is a lot of this slag coating very
quickly deposited.
As we found in the subscale ramjet tests, this is a very effective protectant for the
steel ramjet case that is the missile aft airframe. We were able to test and retest on the same
subscale insulation sleeves many times,
for a much longer accumulated burn than would ever be seen in flight.
The two adjacent fuel entry ports for each inlet entering
stream provide some impetus for re-turning the air back to axial direction, reducing the total pressure loss, despite the lack of physical walls. This relies upon collision of the airstream
with the associated pair of fuel streams.
It is in this way that “synergistic” behavior is obtained: more downstream total pressure recovery than
one might otherwise expect. The 4 ports
dedicated to dome center injection really do also increase this effect.
All of these physics are far outside of what applies to more
conventional fuel-air flameholding. This
SA-6 combustor and its fuel and air entry design is ONLY applicable to truly hypergolic
fuels!
There is a fiber-reinforced phenolic cover over the 12-port
dome that is the gas generator outlet.
It keeps the fuel grain from being hot-gas ignited, during the booster burn. That cover is also the forward mechanical
support for the IRR booster propellant grain.
It is held in place with an aluminum screw, and sealed by an O-ring.
After the booster burns out,
when the gas generator ignites,
the pressure rise in the gas generator breaks the aluminum screw at
about 200 psia. The gas generator dome cover
then departs downstream right after the inlet port covers. It all fits through the ramjet nozzle throat, even simultaneously.
Fuel-Rich Solid Propellant Fuel Supply
This is a pressed propellant system rich in magnesium, resembling (but not at all the same as)
a magnesium flare composition. Average propellant
composition values are 65% magnesium by mass,
20% sodium nitrate by mass, and
5% each (by mass) carbon black,
naphthalene, and mineral
oil. The naphthalene is the hard white solid
often described as a moth repellant in the West.
These ingredients are “dry-mulled” together as powders, then compacted in a press. The naphthalene and the mineral oil make this
compactibility happen. The carbon black
and the mineral oil are the mold lubricants.
This has nothing to do with any conventional cast propellant processing
in the West, and a whole lot more to do
with magnesium flare processing, except
that Teflon is not involved, and two separate
press operations are required, unlike
with flares.
What we found at Rocketdyne-McGregor was that burn rate was
controlled more by the magnesium particle size,
not so much the sodium nitrate oxidizer particle size. This is understandable, since magnesium dominates the
composition, and by far. It proved easiest experimentally to tailor propellant
burn rates with magnesium particle size than anything else. The SA-6 gas generator proved to have higher
burn rate initially, and lower burn rate later,
for a net slightly-regressive burn profile.
As it turned out, the
choice of sodium nitrate as the fuel propellant oxidizer was in part driven by
the radar return of the missile exhaust plume in ramjet operation. This made tracking where the missile really
was (versus where it needed to be) rather easy.
The sodium ions stood out like radar beacons in the exhaust plume.
The fuel propellant grain was mostly a stack of
otherwise-identical fuel propellant discs,
each with its own burn rate.
These were a sort of flat disc shape modified with circular
“whorls”. The running joke at
Rocketdyne-McGregor was that the fuel propellant grain was a stack of “camel
turds”. See Figure 5.
During the first contract,
these discs were initially pressed to full compaction at 17,000
psi, then stacked into a fuel
grain, and re-pressed at the same 17,000
psi to “consolidate” the stack. That
process proved to be completely ineffective!
Flame would flash into the disc interface areas, over-pressuring the test lab motors to the
explosion point, within milliseconds of
ignition. Bonding agents did not
alleviate this.
The solution to this dilemma proved to be a notion based on
a child’s mud pie: barely “hang” the
discs together at a reduced compaction pressure (about 5000 psi), then stack these up in desired burn rate
order, and finally compact the stack (at
17,000 psi) to full density. 5000 psi
got density variations through the stack matching the samples of Russian
propellant. Compacting instead to 7000
psi eliminated all the observed variations seen in the Russian samples. But the main lesson was that the stack would
not consolidate unless there was about 30% compaction potential still left in
the discs.
Either way, the final
compaction of the stack to 17,000 psi eliminated the flashing of flame onto the
interfaces between discs. There is a
“mold bloom” on stack diameter that occurs between the two compactions. It necessitates either two slightly-different
tooling diameters, or else filing-off
the grain stack OD to fit the case diameter for the second compaction
operation.
The SA-6 gas generator case was another standard-size tube
of an annealed titanium alloy. There was
no liner or insulator; instead, the grain stack was wrapped in what amounted
to Kraft paper, then inserted into the
case. With tooling installed, this stack was compacted to 17,000 psi inside
that case, squeezing it out against the
Kraft paper and stretching the case slightly.
This crushing action on the paper against the case completely inhibited
burning down the sides of the grain stack.
The two-stage pressing operation eliminated burning on the interfaces
between discs.
All the discs were dimensionally identical except two: (1) the forwardmost disc had a flat face
matching the flat forward closure plate of the gas generator, and (2) the aftmost disc had a configured
ignition surface featuring a center peak. In the SA-6, burn rates were higher at the ignition
end, and lower in each disc to the
lowest rate at the forwardmost disc. Burning
surface is also slightly regressive.
Burn rates were about 20-30% regressive ignition-to-burnout. This crudely compensated for nozzle throat
area reduction due to magnesium oxide slagging,
for a crudely-neutral pressure trace with time, at about 300 psi (20 bar). The flow rate trace
with time was slightly regressive, like
the burn surface.
The forward case closure was installed such that two screw
holes in it were open. Something
resembling Elmer’s Glue-All was injected by one screw hole to fill the void gap
between grain and closure, until it
extruded out the second screw hole.
Screws were inserted into these holes,
and the glue allowed to set.
The gas generator igniter ring was installed adjacent to the
ignition surface of the fuel grain,
followed by the aft generator closure,
which had the 12-hole injection dome into the combustor. This injection dome was sealed by installation
of the closure cap, that doubled as the
forward booster grain support.
Subscale testing revealed that the fuel injection ports
really did “optimize” as sonic-only ports.
Ramjet direct connect tests confirmed a ramjet performance
decrement, if the port exit Mach number
was supersonic. Each of these ports was
a graphite insert in the basic steel structure.
From a ballistics standpoint, there are three processes operating during
the SA-6 gas generator burn. One is the
regressive distribution of burn rates down the stack of discs. Another is the slag buildup in the fuel ports
with time. The third is the slight regressive
variation in effective burning surface of an essentially end-burning grain, starting from a surface that is not flat.
These result in a grossly-neutral pressure-time
history, and a slightly-regressive
effluent massflow-versus-time history. This
is a fixed delivery history, there is no
control of fuel flow rate. Such
is a good compromise for the effects of increased scooped airflow versus flight
speed, over-balanced by decreased air
density with altitude, as the SAM climbs
out on its way to its target.
The SA-6 is a SAM intended to hit aircraft targets from under-5000
to about-60,000 feet altitudes. Its
fixed fuel delivery trace is a sort of middle-of-the-road compromise to cover
those extremes. It will be near-perfect
fuel/air ratio as the SAM accelerates toward a middle-altitude target, too rich against very high altitude
targets, and too lean against targets
near the surface. But, because the fuel effluent is hypergolic with
air, there are no risks of lean or rich
blowouts!
The gross characteristics of the gas generator effluent as a
fuel are stoichiometric air/fuel ratio about 3.0 (by mass), and lower heating value about 8300
BTU/lbm. This effluent is heavily-dominated
by magnesium vapor, which is hypergolic
with air, even at room
temperatures.
Ramjet Ground Tests
In the early years when “Group Work” was done, the McGregor plant had a small direct-connect
ramjet test facility. Its airflow
capacity was up to 5 lbm/sec at up to 750 F total temperature, for a usable time of about 1 minute or so. This was a very simple blowdown-type facility
using bottled air and a two-stage regulator control feeding a calibrated choked
venturi meter upstream of the air heater.
The heater was a simple pebble-bed heater, electrically heated to the desired
temperature before the test.
The main advantage here with the pebble bed heater is that
the air supplied to the test article is really air, not “vitiated air” from a combustion heater
with oxygen replenished. That requires
very sophisticated controls, the pebble
bed does not. With reactive metal
fuels, especially magnesium, the excess water vapor and carbon dioxide in
vitiated air are a reactive oxidizer,
not an inert.
If you test metallic fuels in a vitiated air system, you will get bad answers in the form of
over-optimistic performance numbers. Our
facility was almost unique around the country at that time, in using real air, so as to get realistic performance with
metallic fuels, as well as the hydrocarbon
fuels.
The test hardware we used was essentially based on 6-inch
commercial pipe hardware (tubes and flanges).
Motor explosion safety was by use of neckdown bolts on items containing
propellant. This hardware was
specifically designed not to throw shrapnel in the event of a motor explosion
event.
The thrust stand in those days was not calibratable for
accurate tare forces, so we relied on
performance computed from static pressure measurements, knowing the total/static ratio at the
combustor exit from its nozzle contraction ratio, and that its nozzle throat was choked. The stand was a small table atop 3 flexures
underneath it, and restrained by a
thrust transducer. We took thrust
data, but did not rely on it for performance,
during those early days. That changed later (see below).
This facility had both photography and altitude simulation
capability. This was before cheap
video, so the photography was real film
sequence cameras and real movie film cameras.
It usually took about half a week to a week, after the test, to get the photographic results.
The altitude simulation used a second-throat supersonic
diffuser sealed to the nozzle housing with a rolling-diaphragm seal, and belling-out to a steam ejector that
pumped the decelerated stream back up to barometric. Open-air nozzle testing
was preferred, if at all possible in
terms of nozzle choke.
Instrumentation in those early times was two moving-paper
oscillograph recorders, with a total of
about 24 channels. We used some 8
thermocouples and some 16 pressure and force channels, to record the conditions throughout the air
system, test article, and altitude rig (if any). These traces were reduced manually with
scales and pencils, and hand-drawn
plots. This took about 40 man-hours per
test to accomplish. (That also changed
several years later to digital data acquisition and computer-processing, but still using my same test item cycle
analysis, which included transient air
system effects).
In the 3 years leading up to our “Group Work” testing, I had developed the procedures, test order document format, and (most importantly) the test item cycle
analysis for pressure (and thrust) based performance. At this stage of our testing
development, I witnessed every test from
outside the blockhouse, and about 75
feet away from the stand, which was definitely
a 5-sense experience!
Over the course of some 8 years testing (of which the “Group
Work” tests were a only small part), I
witnessed about 120 such tests. We had a
very good experimental test record: only
4 of these 120 exploded. Quite
exciting. You cannot help but turn and
run; the fight-or-flight impulse just
takes over.
The test hardware we used for the “Group Work” tests was an
air-entry section that had 4 inlets equally spaced, entering at 30 degrees. The inlet/combustor area proportion was
comparable to SA-6 at about 40% area ratio.
This was 6-inch pipe size steel hardware, insulated with silica phenolic sleeves to a
4.6 inch combustor ID. We used enough
sections downstream to get about 6-7:1 L/D ratio for the chamber, and tested with a graphite nozzle insert that
was sonic-only for test data reduction simplification.
We had other inlet entry sections that entered at 45
degrees, and which could be two inlets
either 90 or 180 degrees apart, at
inlet/combustor area ratios near 50%.
These (and the 4-inlet rig) could have the lab motor gas generators
“stepped back” so that the fuel injection plane was at, or ahead,
of the air entry station. 4 and 6
inch lab motors could be used. We did
not use these other inlet entry sections for “Group Work”. Instead,
we fired a variety of magnesium and hydrocarbon fuel propellants in all
of them for other projects, both
contract and IR&D.
The gas generators we used for “Group Work” were standard
heavyweight 4-inch diameter laboratory motors,
spaced with adapters such that the injection orifice plane was right at
the inlet dump plane, to be as close to
the SA-6 geometry as we could possibly get.
In testing, it did not matter
whether we had one injection port or four,
so we usually just tested with one.
These gas generators used graphite discs with drilled holes to make the
sonic ports.
It was well after the “Group Work” testing that we began
expanding our test facility.
Eventually, it had two
independent air lines, each 10 lbm/sec
flow rate, with the 750 F pebble bed
heater on one, and a 1200 F pebble bed heater
on the other. The test stand air entry
got revised for larger flight-like hardware,
and well-calibrated for tare forces,
so that both pressure and thrust-based performance were obtained. This used the same altitude ejector, but with a new, larger second-throat supersonic
diffuser. I played a major role in all
of this expansion. It was all scrapped
when the plant got closed.
A sketch of our test facility arrangement, as it was in those early years, is given in Figure 6.
Engine / Inlet Proportions
From the very beginning,
we did our ramjet nomenclature and performance definitions in accord
with the standards set forth in CPIA 276.
The relevant station numbering and combustion efficiency definitions are
shown in Figure 7.
The SA-6 has four side-mounted inlets, similar to the side-mounted entry shown in
the figure. These are equally-spaced 90
degrees apart around the circumference of the vehicle. These inlets are round, unlike the 2-D inlets illustrated in the
figure. The swept out capture area AC
is the sum of four circles, each the
inside diameter of the inlet tubing.
This is sort of like the round nose inlet in the figure, but it is just that there are four of them.
For the SA-6, station
2 is after the 30-degree bend, but
before the sudden dump into the chamber.
The round tubing from which these inlets are made is the same inside
diameter at station 2 as it is at station C.
Thus A2 = AC.
The SA-6 combustor is a constant diameter tube, so that A3 = A4. Dimensions are such that A2/A4
~ 0.40, and so also AC/A4
~ 0.40. A5/A4 is
just about 0.90 to 0.95, and A6/A4
~ 1.0. I can no longer find my old set
of drawings to pin these ratios down to their exact values.
Missile Engagement Characteristics
Our second contract was also run out of MICOM in
Alabama. The Army gave us the results
from the other contractors for the vehicle weight statement, vehicle aerodynamics, inlet aerodynamics, and booster performance. They confirmed to us that this SAM flew as
semi-active radar homing using proportional navigation with modest gravity bias.
USAF supplied us a ramjet-capable trajectory code called
ABTRAJ, from Ken Watson at WPAFB in
Ohio. I did the rest. My targets were thing like F-4’s and
F-15’s.
Against a mid-level maneuvering target, SA-6 is still powered at intercept, and doing about Mach 2.8. At mid altitudes, it is capable of maneuvering at about 44 gees
turn acceleration. This occurs about 30
km out from the launch site. A rocket
that size would be coasting, able to
pull fewer gees, and only about 10-15 km
out. That explains neatly the losses seen by the Israelis in the 6-Day War.
At the very highest altitudes, SAM speed is nearer only Mach 2, and the thin air reduces gee capability for
the SAM, and for its targets. Near the deck, speed is about Mach 2.5, and in the thick air, the SAM has full maneuver gees, as do its targets. High and low altitude ranges to intercept are
somewhat shorter, nearer 15-20 km.
This SAM is late-1950’s vacuum tube technology. Its radar is lower frequency that those used
today. It flies command-guided from the
launch site until its on-board seeker can find and lock onto the target. At that point it goes autonomous, ignoring commands from the launch site. That usually happens around 2-3 seconds from
impact, which is around a nautical mile
(or maybe 2 km) from the target.
The rocket-powered SAM with solid-state electronics that
replaced SA-6 is the SA-11. It covers
about the same engagement zone with only a solid rocket, because its payload is both smaller and
lighter. One can fire SA-6 missiles
using SA-11 guidance and control equipment at the launch site, but one cannot fire SA-11 missiles using SA-6
guidance and control equipment at the launch site.
There is one “sort-of” unexpected advantage to the older
lower-frequency on-board radar seeker of the SA-6. As it turns out, stealth characteristics of aircraft are very
frequency-dependent. These designs are
for more modern, higher-frequency
radars. It turns out that stealth
aircraft are just not very stealthy at the older, lower SA-6 frequencies. This neatly explains the shootdown of an
F-117A in the Balkans by an SA-6 missile.
Signature Issues
I have already indicated that sodium ions in the ramjet
exhaust plume make it a huge radar target.
This starts with the transition to ramjet propulsion, about 3-4 seconds after launch. That allows the ground operator to easily
track the missile, so that he may
effectively command-guide it close enough to the target that its on-board radar
seeker can go autonomous.
The infrared situation is different. The hot slag particles in the exhaust plume
are still close to chamber temperature,
up close to the missile, but these
are spread thinly in space. This is
about 1 to 1.5 micron radiation in terms of color temperature, but the total signature emitted is not very
large at all, because of how thinly those
particles are spread. It’s as if the
effective emitting area is far smaller than the perceived plume area.
As for aeroheated hot skin temperatures, these are at most the external air stream
recovery temperature, very little
different from the total temperature.
Even at 2.8 Mach at sea level,
this total temperature is only about 870 F, and colder still at slower speeds and higher
altitudes. There just isn’t very much
signature emitted from surfaces that cold,
even at 100% emissivity.
Visually, this is a
“smoky” system, although its smokiness
is far less than a rocket of equal thrust.
Such rockets leave an opaquely-dense white smoke trail. This ramjet leaves a very translucent thin
white trail. That is the air dilution
effect at work.
There is some video footage from the 6-Day War showing what
is reported to be an SA-6 traveling against a clear blue sky on a bright
day. The plume is white in color, but very thin to the eye, indeed.
On a gray, misty, overcast day,
it would be almost invisible to the naked eye.
Final Comments About All the Other Ramjet Work I Did
When I first hired on at Rocketdyne-McGregor straight out of
graduate engineering school, it was
originally to be an understudy structural engineer. That was December 1975. But my training in propulsion and especially
aerodynamics came to the attention of local managers within a couple of months, as they struggled to deal with a solid
propellant gas generator-fed ramjet application known as “ducted rocket”. Nobody else knew what the inlets were for, or how they worked.
This change for me included helping with the construction
and checkout, already underway, of a ramjet direct-connect test
facility. Initially that facility had a
clean-air airflow capability up to 5 lbm/sec, at up to 750 F total temperature, using a pebble bed heater. It had about a 1 minute blowdown
capacity, and it used a second-throat
supersonic diffuser and steam ejector for its altitude simulation option.
Plus, these efforts
included a cooperative IR&D project between McGregor and the Marquardt
Company in Los Angeles, to develop fuel
rich solid propellant gas generators, and test them with air. These efforts are how I got started in ramjet
work, and how I first got to know some
experts at Marquardt, and at WPAFB.
These were notably Joe Bendot and Bob Ozawa at
Marquardt, and E. Tom Curran at
WPAFB. Ozawa was the V-gutter
flameholder expert, and Curran the
coaxial dump flameholder expert. I
eventually became the side dump flameholder expert.
These efforts (plus a continuing McGregor IR&D effort in
airbreathing propulsion) initially focused on various cast and pressed
high-magnesium fuel propellants. We
tested at McGregor in a 4-inlets-at-30-degrees configuration, and Marquardt tested in a
two-inlets-at-45-degrees-and-90-degrees-apart configuration. Ours was driven by tales we had heard about
the SA-6. This preceded “Group Work”.
We added hydrocarbon fuels during these efforts, and immediately ran into failed combustor
ignition troubles with them, while
Marquardt did not. The sizes were not
that different, but the inlet entry
geometries were!
Later on, we tested
in the same two-inlets-at-45-degrees-and-90-degrees-apart configuration, and most (but not all) of our combustor ignition
troubles went away. But the troubles
returned, with our attempted testing in
a two-inlets-at-45-degrees-and-180-degrees-apart configuration. Symmetry versus asymmetry was the common
thread here. About this time, Hercules bought the McGregor plant from
Rocketdyne.
Ultimately, I found
that there were two problems: (1)
symmetrical inlet entry geometries incompatible with the physical
characteristics of the hydrocarbon fuel propellant effluents, and (2) providing an adequate combustor
ignition stimulus. Neither was an issue
with 50%-or-higher magnesium, which was
just hypergolic with air at any air temperature.
These compatible-geometry results are at strong variance
with the many flameholding geometries that work with liquid fuels, but then,
so are the solid gas generator fuel effluent physical properties. The main difference was the solid soot
content that reacted something like 10 to 100 times slower than the carbon
monoxide gas content.
The soot has to be “centrifuged-out” of the flameholding
recirculation, which requires asymmetry-of-entry
to organize the recirculation into a single strong vortex. This effect is worse at smaller diameters
than larger. Simple, but subtle!
And not widely believed by others in the business at the time.
The combustor ignition stimulus had to be much better than
just waiting for gas generator igniter debris to reach the combustor. Tiny rocket motors loaded with castable
high-magnesium fuel-rich propellant,
firing directly into the flameholding recirculation zone, proved to be utterly reliable. It means you need to know exactly where that zone
is, and just how it works. We had these answers by about 1978. They served us well for many years. I cannot say the same for our competitors.
These IR&D efforts were aimed at a series of “6.2
technology development” programs from USAF at WPAFB. These were targeted at a proposed ramjet
propulsion upgrade for the AIM-120 AMRAAM missile. A couple of these were named DRED and
DR-PTV. The “DR” was “ducted
rocket”.
If memory serves,
these programs ran through about 1979 or 1980. They did establish that the ramjet AMRAAM
engine would have reduced-smoke hydrocarbon fuel, burned in an asymmetrical two-inlets-at-45-degrees-90-degrees-apart inlet
entry configuration. The “Group Work”
effort was underway alongside all this. The
differences between magnesium and hydrocarbon could not be more stark.
The baseline design for that AMRAAM engine at that time was
a fixed flowrate delivery history that we often called “FFDR” for “fixed-flow
ducted rocket”. The USAF began to insist
on eliminating the rocket-ramjet transition ejecta, that being the ejectable booster nozzle and
inlet port covers. Because the trajectory
studies showed serious rich and lean blowout risks trying to cover a wide range
of target altitude, USAF also wanted a
means of controlling the fuel flow.
McGregor developed on IR&D a gas generator throttle
valve (a variable-area choked throat device),
and the high-exponent fuel propellants to make it work effectively. Atlantic Research developed a wire-pulling
gadget that coned the burn surface in response to the wire extraction
rate. This varied burn surface area and
effective burn rate against fixed throat area.
Atlantic Research never solved an erratic flashing of fire
down the side of the moving wire, which
caused motor explosions. We solved all
of our problems (including motor explosions on a linear control, by developing a reliable nonlinear control), and so we won the VFDR (“variable flow ducted
rocket”) contract from USAF. This was about 1981 or 1982 or so, if memory serves.
We even adapted the coning notion to cast propellant strands,
instead of wires, which divorced the strand ballistic
properties from the fuel properties of the matrix propellant. This plus our throttle valve and nonlinear
control became the VFDR baseline gas generator design. It was named SAEB, for “strand-augmented end burner”.
About this same time,
McGregor teamed with Marquardt and Martin-as-prime for a liquid ramjet
program called ASALM-PTV. Marquardt was
to supply the ramjet and inlet port cover,
we at McGregor were to supply the integral booster, complete with ejectable booster nozzle.
My part in ASALM started with support to the
thermal-structural design on the ejectable nozzle, and support to the ballistics design analysis
of the booster propellant grain design. That
was my introduction to the keyhole-slot grain design, that I now like so much for application to
ramjet boosters.
It went on to a contract we had, to investigate feasibility of a variable
geometry ramjet nozzle. This led to
live fire tests of a design that worked after a 900-second burn. This was a “lollipop” in the ramjet
throat. It was not chosen to fly on
ASALM. However, ASALM was my first encounter with the coaxial
center-dump flameholder geometry.
I did get to visit the very large (100 lbm/sec) direct-connect
test facility operated by Martin at its Orlando facility. That is where I first encountered the dense
kerosene-like synthetic liquid fuel RJ-5/Shelldyne-H. Ultimately, ASALM flew 7 times in flight tests at Eglin
AFB in 1980. 6.5 of those tests were
fully successful. The very first one suffered
a throttle runaway, and accidentally set
a speed record at about Mach 6.
This was about the same time as we at McGregor were
enlarging and updating our own facility on IR&D. We went to two air lines, at up to 10 lbm/sec each. Eventually,
we added a 1200 F pebble bed heater to the second line. We added tank farm capacity to keep the
1-minute blowdown, and a larger diffuser
for larger test articles. Eventually we
went to a big air tank trailer for our “tank farm”.
Starting on IR&D,
and transitioning to a USAF contract,
we developed a dual propellant overcast grain design for a “nozzle-less
booster” for the ramjet upgrade to AMRAAM.
This eventually became the standard booster design for the ramjet AMRAAM,
because it eliminated all the booster
nozzle ejecta. Port cover ejecta were
eliminated simply by retention, for
burn-up in place. The port covers folded
inward, together between the entering
inlets.
Key to the success of our nozzle-less booster design was
overcast of a lower-rate propellant over a higher-rate propellant. The nozzle-less performance decrement was
held to only 15-20% in this way. We tested this in the full scale ramjet AMRAAM
hardware multiple times. It was very
reliable.
For the interval from December 1983 to March 1987, I left the McGregor plant to work at what was
then known as Tracor Aerospace Austin,
in the airborne countermeasures and deception business. Among the many things I did there, one relates to the ramjet work being
discussed here. This was a small ram-fed
airbreathing combustor that was a hot gas generator for an infrared decoy.
For this combustor I used the ASALM coaxial dump
technology, but I had to stop-down the
nozzle substantially, to slow the flow
speeds enough to have flame stability in such a small size. This combustor was 1.5 inches ID and 3 inches
long. In it I successfully burned
hydrogen, propane, gasoline,
jet fuel, and alcohol. I got
started writing my own ramjet cycles codes for this effort, as well as very practical experiences burning
liquid fuels in ram combustors. And I
invented a low-density ceramic insulator that was reusable for hours of burn on
this project (see
Figure 8)
Figure 8 – Small Decoy Combustor Ceramic Insulator
While at Tracor, I
stayed in contact with friends at McGregor,
helping them with teething troubles in the new upgraded test
facility, as they prepared to bid the
second VFDR program. The USAF sponsors
insisted that Hercules-McGregor do a joint venture with Atlantic Research, before they would award it. They wanted
Atlantic Research’s boron fuel because it looked so good (theoretically on
paper).
I initially returned to McGregor in program management. But, I
was soon reassigned to project engineering,
made the principal investigator for airbreathing propulsion
IR&D, and informally managed the
entire plant IR&D program for the chief engineer. For airbreathing IR&D, I went to work on boron and clean fuels, and on an unchoked-throat gas generator that
could provide constant fuel-air ratio “control” with no moving parts or
controls at all.
The unchoked-throat effort led to a very safe and convenient
way to test experimental fuels rapidly in the direct-connect facility. This really accelerated the fuels development
effort at McGregor. We did not win the actual
contract to test this technology for USAF,
so we just continued on IR&D,
which led to a paper on it at the Naval Postgraduate School. We were able to test successfully in
full-scale AMRAAM hardware, for a tenth
what USAF paid their contractor, who
never got their subscale engine to burn at all. Our paper was right after theirs at the
meeting. It made quite a stir at that
meeting.
While all this was going on,
I got involved in some aspects of a USN program called AAAM. This was a nose-inlet ramjet with liquid fuel
and a coaxial center dump flameholder.
McGregor was to supply the ramjet case and integral booster.
This was a very speculative investigation program, in that high flight speed aeroheating was too
much for aluminum, and weight allowances
would not allow steel. We flow-formed
the case from a beta-phase ductile titanium that was not in Mil Hndbk 5. This held much promise until we found that
the alloy would age at room temperature into uselessness. USN killed the project when that happened.
For VFDR, I was able
to bring two different boron fuels and a non-metallized smokeless fuel from
IR&D to the VFDR testing program by about 1992. The contract already had two McGregor low-aluminum
fuels, one having the SAEB
technology. Atlantic Research brought
their boron fuel. In the refereed tests
witnessed by USAF, all of these
performed about the same, except for
Atlantic Research’s boron, which
seriously underperformed. Clearly, theoretical paper evaluations can be quite
misleading.
Nasty corporate politics induced Hercules to close the
McGregor plant while VFDR was still underway.
The other Hercules tactical location refused to learn how to cast the
thick propellants we used, and did not
want the airbreathing programs. So
everything we had done at McGregor for the ramjet AMRAAM over 2 decades went to
our joint venture partner Atlantic Research.
(Hercules sold it to ATK, who actually
closed it, about a year after the first
wave of layoffs that took me.)
Atlantic Research (of course) made their underperforming
boron fuel the baseline VFDR fuel, and (not
surprisingly to me) in ground tests it failed to ignite at simulated middle and
high altitudes. That “killed” the ramjet
AMRAAM application as far as USAF was concerned, so the technology never flew in that
missile. Something very much like it is
flying today in the “Meteor” missile produced by the Europeans.
The gas generator and throttle technology did make it into
the USN gunnery target drone “Coyote”,
which is powered by a gas generator-fed ramjet using the underperforming
Atlantic Research boron fuel. It leaves
a black smoky trail of what looks like unburned fuel (also unsurprising to me).
Being a new technology development guy, I was laid off in the first wave, when the plant closure announcement was made
in 1994. I never worked in the defense
industry again, since a million and a
half other aerospace engineers were also unemployed that same year. The McGregor test facility (like so much
other plant equipment) was broken-up and scrapped. No other facility anywhere in the country
ever learned our methods of casting thick propellant mixes. What a waste!
All I have been able to do since is advise or assist a few
small groups about ramjet propulsion technology, or run studies for myself. I have written a set of cycle codes to do
sizing and point performance work for these ventures. These include both high-speed and low-speed
ramjet designs.
And I have written a very practical “how-to” book on ramjet
propulsion. AIAA did not want to publish
it, so I will have to devise a way to
self-publish it. Watch this space for
updates about that.
As for my friends the other experts: Bob Ozawa passed away a long time ago. E. Tom Curran is in assisted-living with his
wife. I don’t know if Joe Bendot (the
ejector ramjet guy) is still alive, but
if he is, he would be about age 90. As near as I can tell, there are none like us anymore.
Related Posts About “Ramjet”
These are the postings on this site related to ramjet
propulsion. The list is in reverse order,
newest first.
Date title
2-4-2020 One of Several
Ramjets That I Worked On (this article)
1-2-2020 On High Speed
Aerodynamics and Heat Transfer
1-9-2019 Subsonic Inlet Duct
Investigation
1-6-2019 A Look at Nosetips
(or Leading Edges)
1-2-2019 Thermal Protection
Trends for High Speed Atmospheric Flight
11-12-2018 How Propulsion Nozzles
Work
7-4-2017 Heat Protection Is
the Key to Hypersonic Flight
6-12-2017 Shock Impingement
Heating Is Very Dangerous
12-10-2016 Primer on Ramjets
11-26-2015 Bounding Analysis: Single Stage To Orbit Spaceplane, Vertical
Launch
11-17-2015 Why Air Is Hot When You
Fly Very Fast
8-16-2014 The Realities of Air
Launch to Low Earth Orbit
11-17-2013 Payload Comparisons
11-6-2013 HTO/HL Launch With
Ramjet Assist
8-20-2013 Applying Ramjet to
Launch Accelerators
3-18-2013 Low Density Ceramic
Non-Ablative Ceramic Heat Shields
12-21-2012 Ramjet Cycle Analyses
8-16-2012 Third X-51A Scramjet
Test Not Successful
8-22-2010 Two Ramjet Aircraft
Booster Studies
7-23-2010 More Strap-On Pod
Ramjet Engine Data
7-11-2010 More Ramjet
Performance Numbers for the Strap-On Pod
2-28-2010 Preliminary
Acceleration Margins for Baseline Pod
2-20-2010 Ramjet Strap-On Pod
Point Performance Mapping
2-20-2010 Ramjet Strap-On Pod
Concept
2-20-2010 Inlet
Data for Ramjet Strap-On Pod
(edit 2-8-20) Posts About Pulsejet
Pulse jet is NOT ramjet, but I do know something about that, too. There are a few posts on this site by me about pulse jet which draw readership. These are listed here for pulse jet enthusiasts.
5-20-12 Recommended Broad Design Guidelines for Valveless Pulse Jet Combustors
4-30-12 Big Student Pulse Jet an Even Larger Hit at TSTC
3-6-12 Student Pulse Jet a Hit at EAA Meeting
11-12-11 Student Pulse Jet Project
