My wife got this for me off her Facebook. It is one of the better things I have seen. Enjoy.
Update 12-16-2021: This was just too good. We made our own. It's hanging in the entryway.
I have real difficulty with the fact that, even after all these years, it is still necessary to explain to people what really destroyed Space Shuttle Challenger and killed her crew, back in January 1986. It was really two very seriously-bad upper management decisions at NASA, one long before the launch:
(1) to insist on poorly-designing the O-ring seal joints with 3 interacting serious errors, and
(2) to fly soaked-out colder than had ever been tested, when everybody’s engineers did not want to.
First, you have to understand what really happens in federal government contracting. There is only one customer, and he thinks he is always right about every decision that he makes. If you do not do it exactly the way he wants, no matter how wrong he might be, then you lose the contract and you don’t get paid. And, the government is quite often wrong about how best to do things! That’s not to say the contractors are always right, but they are wrong a lot less often than the government.
You also have to understand that NASA never did know, and still does not know, the art of building reliable solid propellant rockets. Essentially, no one at NASA ever did that kind of work. They buy these things from contractors who (by definition) know much more of the science, and especially the art, than anyone at NASA knows. The “science” is that knowledge which was written down. The “art” is the knowledge that was not written down, usually because no one wanted to pay for the writing.
I can tell you from experience as an insider within the business, that “rocket science” isn’t really “science”! It is only about 40% science, about 50% art, and about 10% blind dumb luck. And that’s in production work! In new product development work, the art and luck percentages are even higher.
Further, this same sentiment applies to pretty much any type of engineering effort, not just rocket work. That explains a lot, about a lot of things, doesn’t it?
Poorly-Designed O-Ring Seal Joints
What I show in Figure 1 is how such joints should be designed and built. This is the design that most solid rocket motors use, very successfully, whether large or small. In most rocket motors, you need only join the aft and forward closures to the case cylinder. Only in some of the really large motors, the case cylinder itself is divided into segments that must be joined, usually to limit the size of the case-bonded propellant masses that must be cast and cured within them.
The sketch in the figure is what mechanical engineers call a “radial static seal”. It is “radial” because the O-ring lies between an inner and an outer surface, that must include a gap of tightly-controlled size between the two parts, for assembly. One part stabs into/inside the other, in order to join them, in this case by a row of pins. It is “static”, because the parts, once joined, do not move anymore. There are strict but well-published guidelines and procedures for sizing the O-ring groove dimensions, the gap for assembly, and the size of the O-ring, as well as its material composition and its hardness. These guidelines and procedures are used precisely because they work so very well. Examples: Refs. 1 and 2.
Something also shown in the figure is peculiar to solid rocket motors, especially those that are segmented-case designs. There is a joint in the insulation (and thus also the propellant) that leads to the sealing surface gap, that in turn leads directly to the O-ring in its groove. You DO NOT obstruct this path with sealants, putties, greases, or anything else! But there does need to be a right-angle bend, to stop radiant heat transfer from the flame in the motor from heating the O-ring directly.
The air in this path is what gets suddenly compressed upon motor pressurization, and which in turn forces the O-ring to the far side of its groove, where it gets squeezed against that surface to seal. This is called “seating the O-ring”, and until it is properly seated, it CANNOT seal, and so it briefly leaks!
Figure 1 – A Properly Designed O-Ring Seal Joint
It is the air in the path that gets compressed against the O-ring, with hot booster gases and hot solids filling most of the path volume that the air formerly occupied. But the air cools by convection to the steel much more effectively and faster than to the O-ring itself. THAT is how the O-ring is not damaged by the hot air, or the hot gases! The hot solids are stopped by the right-angle bend. This is a rapid transient on a time scale equal to, or shorter than, the motor pressurization event.
What you DO NOT want is contact of the hot gases (and especially the hot solids) upon the O-ring! The “hot sandblast” effect of that outcome would cut through the O-ring almost instantaneously.
Note that these two design requirements of (1) one O-ring and (2) an unobstructed pressurization path, will interact very strongly with how one verifies proper assembly of the motor! You must do a pressure leak check of the motor to verify sealing, but you must do it by pressurizing the entire motor! However, you NEED NOT pressurize the motor to its full operating pressure to do this verification!
You only need an atmosphere or so of pressure difference to seat any O-ring and then verify its sealing. If it holds at that low pressure, and you followed the design guidelines correctly, it will hold at full motor operating pressure! THAT is what you verify when you do motor case hydroburst testing, long before you ever cast propellant to make a live motor! That’s the way the real solid rocket motor manufacturers prefer to do it. And it works to very high reliability levels, as indicated in the figure.
However, that is NOT what NASA insisted upon doing! In the mistaken belief that a second back-up O-ring increases sealing reliability, they insisted upon the two O-ring design indicated in Figure 2. Thiokol complied, lest they lose the contract. In the mistaken belief that they had to pressure leak check at full motor operating pressure, NASA did not want to risk fully pressurizing a live loaded motor (and rightly so). And so NASA insisted on a way to apply air pressure at full motor operating pressure, between each pair of O-rings at every joint, instead of any motor pressurization. This is shown in the figure.
What this does is drive the downstream (backup) O-ring to the correct side of the groove, thus seating it for motor operation. But, it also drives the upstream (primary) O-ring to the wrong side of its groove, from which motor pressurization upon ignition must unseat it, drive it across its groove, and re-seat it on the correct side! Until and unless it re-seats on that correct side, the upstream (primary) seal ALWAYS leaks! Period! There is NO WAY AROUND that outcome! And THAT lets hot gases and solids reach the primary O-ring, simply because the re-seating process takes a longer time than pressurization!
Figure 2 – The Improperly-Designed 2 O-ring Joint That Flew, Up Through Challenger
NASA made a third mistake: in the mistaken belief that it would prevent hot gases and hot solids from reaching the O-ring, they insisted on obstructing the pressurization path by filling the insulation joint with “heat protective” putty (zinc chromate putty actually). This is also shown in the figure.
This last mistake makes a bad risk even far worse, because high pressure gases always (ALWAYS!!!) “wormhole-through” a not-solid material (like putty or grease) at a single point! THIS effect is also shown in the figure. That re-distributes the “push” of the gas from a broad front all around the O-ring, to a single point upon the O-ring, as indicated in the figure. The delay unseating the ring, pushing it to the other side of the groove, and reseating it, almost guarantees that the compressed air leaks past it, so that booster hot gases and solids can reach the O-ring. And those will cut a hole right through it.
“Half-moon slices” right through the primary upstream O-ring were seen, upon SRB motor disassembly, in a rather significant percentage of the SRB’s recovered and refurbished. That verifies what I just said about the upstream O-ring being cut! There is no surprise there, once you understand the process!
The difference between this point load problem, and what NASA analyzed in its structural calculations for the O-ring seal is quite stark! The structural analysts were assuming pressurization on a broad front. They did not model the point load effect of the hot gases and solids wormholing-through the putty obstructing the pressurization path. Quite simply, what was built was NOT what was analyzed!
Unnecessary Risk to Fly Too Cold
If the motor is sufficiently cold-soaked, the primary upstream O-ring loses its flexibility and resilience (as do all of them). Pushing the entire embrittled O-ring across its groove all at once is risky enough, but if you concentrate the “push” at one single location by the wormhole effect, you essentially guarantee snapping the O-ring apart at that point! This cold brittleness effect was amply demonstrated by Dr. Feynman at the Rogers Commission hearings (assisted by Gen. Kutyna), when he stirred his sample of the O-ring material in his glass of ice water, and then demonstrated its non-resilience.
Any failure of the primary upstream O-ring, whether by hot sandblast cutting, or by cold brittle fracture from the point jet force load, then puts a single-point hot sandblast jet impacting onto the downstream O-ring, simply because it is nearby! Thus, a sort of “cascade failure” is a very high risk indeed!
The post-Challenger “fix” was a third O-ring in every joint. This just set up the cascade failure as a longer chain, as indicated in Figure 3. The only reason the Challenger disaster did not repeat is that they never flew that cold again. But the 1/51 failure rate demonstrated by loss of Challenger speaks for itself!
Figure 3 – The Cascade Failure Risk Was Compounded By the Redesigned Joint
Fatal Consequences We All Saw
The photography obtained during the launch and loss of Challenger confirms everything claimed here. The seal failed upon motor ignition and pressurization, as shown quite clearly in Figure 4. The dark grey plume is carbon soot-bearing hot gases spewing through the two failed O-rings at the aft segment joint.
Figure 4 – Seal Leak Upon Ignition Seen In Photography
This leak miraculously “cured” itself by plugging-up with aluminum oxide-carbon slag from the metallized propellant. This slag-plugging just happened to hold pressure like that, until the Challenger encountered a wind shear while at “max-Q”, where it was also most highly stressed by aerodynamic forces. The slag plug failed, letting the hot motor gases and solids rush through the hole again. This is shown quite clearly as the anomalous bright-but-small extra plume in Figure 5 below.
This jet of leaking hot gases and solids finally got so big that it cut through one of the aft struts holding the SRB to the center tank. There is always hydrogen leaking from the center tank’s hydrogen tank, and in this case the leaked plume probably burned a hole in that hydrogen tank. With the strut cut, the bottom of the SRB moved outboard. That pushed the nose of the SRB inboard, such that the nose of the SRB poked a hole in the side of the center tank’s oxygen tank.
Suddenly dumping oxygen into a base-burning hydrogen-air fire caused an explosion in the wake behind the center tank that both overheated and structurally overloaded it. The tank collapsed, letting both SRB’s and the orbiter fly free. The released propellants burned explosively as this happened. All this happened in an instant, so it looks like just the one sudden explosion.
Figure 5 – Leakage Resumed After Being Shaken By Wind Shear at Max Q
The released SRB’s continued to “fly” out-of-control under their own thrusts, as we all saw. This is shown in Figure 6. The orbiter’s engines were pointed through a center of gravity that suddenly no longer existed, so they forced the orbiter to pitch-up violently, before starving for lack of propellant from the suddenly-missing center tank. The pitched-up orbiter went broadside to the supersonic wind, which tore it to pieces. This is how those pieces, that we all saw fall into the sea, came to be.
Figure 6 – The SRB Did Not Explode, But It Punched a Hole In the Center Tank
The two O-ring joint was a NASA-mandated design mistake, compounded by mandating putty obstructing the O-ring pressurization paths. The “customer is always right” in government contracting, except that he was lethally and fatally wrong about this one! See also Ref. 3.
The decision to fly cold-soaked colder than the SRB’s had ever been tested, was also a NASA management decision. Both NASA and Thiokol engineers objected, but were over-ruled. Thiokol upper management also over-ruled their own engineers, and told NASA to go ahead and launch. Thus emboldened by Thiokol management, NASA launched the thing, thus killing its crew.
The stand-down to “correct” this problem was nearly 2 years long and horribly expensive. Which just goes to prove what I like to say to anyone who will listen: “there is nothing as expensive as a dead crew, especially one dead from a bad management decision”.
The only problem with that return-to-flight effort is that they did not correct the real problems upon return-to-flight, they actually made them worse with a 3-O-ring joint, and by keeping the putty obstructions. The ONLY thing they did “right” was never to fly that cold again! Which is very likely the ONLY reason that the Challenger disaster did not repeat itself before the Shuttle got retired, since there were more than 51 more flights after the Challenger disaster!
By the way, the crew did not die in the tank explosion and subsequent ripping-apart of the orbiter by air loads. The telemetry showed no high-gee accelerations at all! The crew was still alive in the orbiter cabin until it finally hit the sea, which is about a 200-gee stop, since it hit dead broadside. See Figure 7.
Figure 7 – The Crew Was Still Alive In This Cabin Section (Arrow) That Is Falling Back
I say what I said about the crew because the flight deck back-seaters leaned forward and flipped on the breathing-air packs for the front-seater pilots. They would not have done that unless they knew the cabin had depressurized, and that would have been significantly AFTER the explosion and ripping-apart of the orbiter. They were tumbling clear of the explosion cloud by that time, as illustrated in the figure.
Those two flight deck pilots had breathed-up all the oxygen in their breathing packs by the time they hit the sea, something confirmed by the empty breathing packs that were recovered. Which means they were alive when they hit the sea! By extension, so were the back-seaters, plus the three down on the mid-deck.
They did not have pressure suits, parachutes, breathing bottles, and a hatch they could blow open (basic bail-out gear). More importantly, there was no way to take the spin off the tumbling cabin. Spinning like that, there was no way to reach and exit the hatch, even if they had the other basic bailout gear! But a small drogue parachute from the nose of the cabin section would have taken off the spin! That plus the basic bail-out gear just listed could have saved that crew! It took almost 5 minutes to hit the sea. They had the time to bail out.
I submitted that means for bail-out to NASA, but I was ignored. Coming from an outsider, my idea was “not invented here”, as far as NASA was concerned. Yet, something rather like it might even have worked for Columbia some years later: the 3 mid-deck occupants were still alive inside a tumbling cabin section as it approached impact near Tyler, Texas, well after the breakup during re-entry. Time was short for a bail-out, but without the de-spin drogue, they could not reach the hatch at all.
#1. Parker O-ring Handbook ORD 5700, copyright 2021, original release 1957, Parker O-Ring and Engineered Seals Division, Lexington, KY, available from parkerorings.com
#2. Seal Design Guide, Apple Rubber Products, Lancaster NY, available from AppleRubber.com
#3. Wikipedia article “Rogers Commission Report”, in this case accessed 11-26-2021
There are different design rules for static radial and static face seals, and different rules yet for dynamic radial seals (as on a piston moving inside a cylinder, like a syringe or a hydraulic cylinder). The Shuttle SRB joints fall into the static radial classification.
The appropriate set of rules specifies O-ring sizes and hardness, groove dimensions, and when to use back-up rings. You just follow the design rules, and make sure that only compressed air reaches the O-ring (and on a broad front), upon solid propellant motor ignition.
You accomplish that broad-front pressurization with the 90-degree bend geometry to stop the hot solids and radiant heat transfer, and by NEVER obstructing the O-ring pressurization path with anything! Even too close a fit between the hard parts, can cause problems with the transient pressurizing flow.
You verify your seal design, your case structural design, and your leak check procedure, during case hydroburst testing, long before you ever cast a live motor! You NEVER delete the hydroburst testing step in your development effort. Never! Not for any reason at all!
Then you test live motors at every environmental extreme condition in which you think you might possibly operate. If any redesigns (of anything) are needed, you go back and verify them in all the tests, from hydroburst all the way forward. No motor goes to production, until its exact design configuration has been verified in every test at every test condition!
Once your design has passed all those tests, you stick with your verified leak check procedure as if it were a religious mandate! You add rigorous quality control (of the “total quality management” type), for production. That includes X-raying every single item, to verify that there are no casting voids in the propellant, no unbonds between propellant and case liner, and no other propellant grain cracks or other problems. And then you NEVER operate a motor outside the conditions for which it was tested!
THAT is the way to achieve no-more-than-1-in-a-million failure rates, with solid propellant rocket motors!
The “bean counters” and “management professionals” will absolutely hate that prescription as “too expensive”, but killing a crew with a bad design just costs a whole lot more, than the cost of following that prescription. We’ve already seen that with Apollo-1, Challenger, and Columbia.
Simple as that.
And just as hard to sell to the “bean counters” and “management professionals”, as you might fear.
One of the most unusual ramjet projects I ever worked on was a non-propulsive device. This was a very miniature ram-fed airbreathing combustor, that was to be the hot gas generator for an infrared (IR) decoy. This decoy was to be towed behind an aircraft in lieu of a whole series of dispensed flares. It was intended to work by having enough IR output to cause the aircraft to drop out of the missile field-of-view first. See Figure 1 for that concept. I was working for my friend Byron Hinderer doing this.
Figure 1 – Towed IR Decoy Concept, called “Warm Brick” at Tracor in 1984
I did this at what was then Tracor Aerospace, in Austin, Texas, during 1984. We called this decoy “Warm Brick”, and my job was to determine if this concept was even feasible (it was). Our idea was to heat a porous refractory material until it glowed brightly in the IR. We preferred fuel-air combustion to minimize decoy mass, and ram combustion is the simplest of the airbreathers. Plus, I had lots of experience with ramjet combustion at what was then Rocketdyne/Hercules in McGregor, Texas.
To the very best of my knowledge, no patent was ever taken out on this concept, and Tracor never did anything at all with it. Even if there had been a patent, and it had been renewed, any such patent would have run out by now. So, what I reveal here should offend no one, and infringe no patents.
As implemented for the feasibility tests, this concept took the form of a “gasoline lantern mantle” made out of commercial ceramic fire curtain cloth, as the IR emitter. This was to be mounted behind a wake-producing spoiler, mounted at the aft end of the burner and inlet assembly. The decoy might carry its own fuel tank, or it receive fuel down its tow line, if a heavier tow line could be tolerated.
To test the scientific and engineering feasibility, I designed a very generalized inlet and burner hardware set that was flexible enough to allow evaluation with a variety of gaseous and liquid fuels. See Figures 2 and 3. The intended flight conditions were relatively low altitude from mild subsonic to barely-supersonic speeds, typical of an attack aircraft threatened by surface-to-air missiles.
Figure 2 – Assembly Sketch for the Initial Version “Warm Brick” Ram Combustor Test Device
The assembly sketch clearly depicts the long fuel injection-and-mixing duct allowed between the inlet diffuser and the sudden dump into the combustor. There was an inlet piece and a fuel injector piece, both made of aluminum for ease of rework, and an inlet tube and a combustor shell, both made of steel. The combustor shell was sized for fabrication from 2-inch schedule-40 pipe, but ended up being made of 300-series stainless to those same dimensions. We tried automotive-style spark ignition.
One can easily see how the molded low-density ceramic liner insert was to be trapped in place by the nozzle block. The arrangement shown in the assembly sketch of Figure 2 (directly-pinned nozzle block) was quickly replaced by a pinned steel nozzle shell ring, as shown in the hardware photo (Figure 3). This revision happened about the same time that the first (unreinforced) liner was replaced with the second liner (reinforced ceramic composite).
Figure 3 – Photo of the “Warm Brick” Ram Combustor Test Hardware as Revised
The design concept called for a small combustor fed by a simple pitot inlet, with a convergent-only nozzle that would likely function unchoked at most conditions. I chose a center-duct coaxial air entry with sudden-dump flame stabilization, similar to the successfully-flown ASALM-PTV liquid-fueled ramjet test vehicle. Geometric ratios were initially set equal to those used in ASALM.
Based on Reference 1, I chose a minimum ¼-inch (6 mm) step height around the dump. The combustor length was sized “empirically” (rules of thumb based on ASALM-PTV geometry) so that the annular separation bubbles would close, and the axial core would be “burned out”, before any of these flows entered the nozzle. That was basically an assumed 11-degree spreading angle, on both sides of the mixing layer between the entering mixture and the recirculated flame. That’s too crude, in hindsight.
We wanted sufficient porosity in the emitter so that the burner operation would be unaffected by the presence or absence of the emitter. The fire curtain cloth gave us that, in the sizes tested, because the surface area of the ellipsoidal shape was so large relative to the final burner throat area. Its effective porosity-driven “free” open area was very much larger than any of the burner throat areas we tested.
There were two crucial unanswered questions: (1) emitter/hot gas coupling (could we really get the emitter hot enough to radiate effectively?), and (2) obtaining stable combustion at all in a burner that small, with any fuel whatsoever! There was an extensive paper trade study done, to determine the desired fuels. In test, these fuels, and some other fuels that were easier to use, were investigated.
This combustor was nominally 1.5 inch (38 mm) inside diameter, as insulated, and 3 inches (76 mm) long inside. The smallest size ramjet combustor in my experience up to that point had been some heavyweight solid-propellant ducted rocket ramjet work (in a completely-different geometry) at 4.6 inch (117 mm) inside diameter, and length/diameter 6-to-8. The largest was ASALM-PTV at a 20 inch (51 cm) combustor case diameter. “Warm Brick” was smaller than anything of which I had any knowledge!
I didn’t want to periodically replace an ablative liner in the test burner, and I didn’t want to attempt an air-cooled liner shell for full-rich combustion in something that small. So I opted for an unknown, inspired by the Space Shuttle’s heat shield tiles. Could I put a low-density ceramic insulator in this combustor, and not melt it? The answer turned out to be “yes”, but it took some adaptive effort.
The project operated in three logical parts: (1) obtain stable combustion with a variety of fuels in the burner alone, (2) add the emitter and determine how best to shape, fabricate, and attach it, and (3) document infrared radiometric output. The real prerequisite for part (1) was the combustor insulator, since we started with gaseous fuels, thereby avoiding the fuel vaporization issue.
I selected free-jet test mode as the best way to accomplish all three parts of this project with the same hardware and test setup (see Figure 4). All that I had personally done while at Rocketdyne/Hercules was direct-connect testing, but I knew about free-jet testing, both from my research, and some experimental association with Marquardt, while I was with Rocketdyne/Hercules.
We used a commercially-rented air compressor trailer as our air source, to be run real-time. In 1984, this 750 SCFM unit was the largest of its kind in Texas. It fed a PVC pipe stilling chamber, terminating in a simple convergent-only nozzle block made (conveniently) of wood.
Figure 4 – Test Setup: Stilling Chamber Exhausting To Left, Fed From Right
The test article was bolted to a heavy pipe stand-and-sting, with its inlet immersed in the free jet of air. That free jet typically measured 190 F (88 C) stagnation temperature, at full-power compression conditions.
The first part of the investigation began with bottled hydrogen gas fuel (series 1). This and all the other trials are summarized in Table 1 below. Series 1 wasn’t very successful for two reasons: (1) the nozzle was too wide open for a stable flame, and (2) free jet air speeds higher than about 0.25 Mach blew the spark column out from the electrodes of the spark plug, even though it was located flush within the annular recirculation zone.
The device didn’t ignite at all until I obstructed the nozzle with a scrap of wood, and it still went out after ignition, if I removed the obstruction. So, I built a smaller-throat nozzle block. We still had to ignite at low airspeed and gradually work up to higher speeds, limited at that time to about half a Mach number by the stilling chamber nozzle. I also tried liquid ethanol unsuccessfully at this time (series 2).
Somewhere in all of this, I first drove the combustor into what proved to be a very violent rich blowout instability, and completely shattered my first (unreinforced) liner! The combustor visibly shook on its sting, and it spit the pieces of its liner out the nozzle, igniting a local grass fire! Later, we estimated a pressure amplitude near 0.8 atm, at audio frequencies (a few hundred Hertz), for this instability.
A photo of the liner molding tools that I used is given in Figure 5, which includes the basic combustor shell as the outer forming tool for the combustor liner. Both it and the nozzle block were laid up as (commercial) low-density molding compound troweled onto the wooden plug, and inserted into the corresponding shell for cure. I used Cotronics Corp. 360M low-density molding compound for this.
Figure 5 – Tools Used for Molding Ceramic Combustor Insulation Liner Inserts
These parts were cured at 215 F (102 C) in an oven to drive off the water. The circuitous paths for exiting steam led to a low density ceramic matrix. The resulting parts were coated with a paint-like ceramic cement (Cotronics Corp. 901), and cured again, in the same oven. The unreinforced liner showed evidence of hot gas flow behind the insulation, and into the cracks, shown in Figures 6 and 7.
Figure 6 – Recovered Pieces of Shattered Unreinforced Liner, Bonded Together
Figure 7 – View of Fracture Surface, Showing Hot Gas Flow Damage with Sooting
I built a second ceramic composite liner reinforced by layers of the fire curtain cloth (woven from 3M Nextel 312), which survived all instabilities and any other test abuses thereafter. It survived many hours of accumulated burn time in near-pristine condition, as seen in Figures 8 and 9. The shrinkage cracks did not preclude functionality. There was some melting evident in the throat of the nozzle.
Figure 8 – View Into Near-Pristine Reinforced Liner, After Hours of Burn Time
Figure 9 – View Into Reinforced Nozzle Block, After Hours of Burn Time
Once we had the burner working at all, we tried some test sample pyrometers in its exhaust plume, with both propane and acetone as fuel (series 3, and acetone proved worthless as a fuel). These pyrometers would be old nails, or else planar samples of potential emitter materials. We even tried gasoline as fuel (series 4), but results were poor, and it became very obvious that poor vaporization was the cause! I tried propane again (series 5) as the most successful fuel, and got enough radiometer output to be encouraging, from a sample of the fire curtain cloth immersed in the jet exhaust.
So, I created a fuel-line hot-soak bucket to correct the poor fuel vaporization problem for test purposes. This took the form of an electrically-heated bucket of old motor oil, in which a coil of the fuel supply line was immersed. That rig is shown in Figure 10. It may resemble a moonshine still, but it is not!
Figure 10 – Fuel Vaporization Preheat Bucket Rig
At this point, I had a crudely-successful burner, but an unproven fuel supply method. I checked out the combined burner and fuel vaporization bucket, first on propane (series 6), then on aviation gasoline (series 7), and finally on a “home-made version of JP-4” that was actually half Jet-A and half aviation gasoline (series 8). Plus, I added instrumentation to the burner unit (enough manometer pressures and thermocouples to attempt an actual “engine” cycle analysis).
Results, including the exhaust pyrometer samples, were favorable enough to warrant continuing the project further. It still required a lower-airspeed ignition. I stood in the jet blast for all these tests, looking directly into the flame zone, and sniffing for unburned fuel, to set mixture. That “settled” the fuel injection and ignition issues well enough to test emitter coupling issues for the very first time.
The first actual emitter was made of Nextel 312 fire curtain cloth, coated with the Cotronics 901 adhesive as a “paint”. It was sewn together, with alumino-silicate thread, from bias-cut gores much like a balloon, to form an elongated semi-ellipse approximation. The seams were left on the outside of this first emitter, as shown in Figure 11. It was the first of several series 9 tests with pre-heated propane, at air speeds up to about Mach 0.47. Those test conditions are depicted in Figure 12.
Figure 11 – Test Setup for First-Article Emitter
For all subsequent tests, the seams in the sewn emitters were placed to the inside, as is depicted in Figure 13. That photo shows post-burn appearance of two series 12 emitters tested with ethanol fuel, but all the internal-seam emitters appeared similar, regardless of series and fuel.
These articles were brittle and fragile post-test, as expected for alumino-silicate materials soaked to temperatures exceeding the solid phase-change temperature of about 2350 F (1290 C). That fragility alone confirmed a high surface temperature for radiation purposes! This was also verified by radiometric measurement, which also indicated very “non-gray” behavior, in that the effective color temperature (radiation peak wavelength) was substantially cooler than the actual temperature.
The spoiler just ahead of the emitter clamp mounting provided protection from direct wind blast forces. Plus, it also provided effective hot gas recirculation effects external to the emitter surface. Both acted to raise emitter material soak temperature, and therefore IR output, quite successfully.
Figure 12 – Test Conditions Explored with Series 9 Propane
Two tests were made as series 10 in this same configuration with the “home-made JP-4” fuel. Results were similar to the series 9 propane runs, except for a small liquid-wet “cold spot” at the very end of the emitter bulb. This was due to still-unvaporized kerosene hitting the emitter on-axis.
Figure 13 – Post-Test Emitter Appearance from Series 12 Ethanol Tests
Sometime during this checkout process before the series 9 propane runs, I successfully modified the inlet to a larger lip radius, in order to decrease its “buzz” instability tendencies at higher backpressures. That also greatly improved ignition characteristics, and it further pushed the rich blow-out instability limits to richer mixtures! The test set-up for cold-flow inlet calibration is shown in Figure 14.
Both the original and modified (larger lip radius) inlets were cold-flow tested with this rig. Data were cross-plotted in a variety of ways. The data plot format for “typical” supersonic ramjets was rather undiscriminating at these subsonic speeds: stream tube area ratio versus Mach and stagnation pressure recovery ratio versus Mach. Plots in the more primitive-variable format were actually more useful for this mostly-subsonic system. These included the diffused Mach to freestream Mach ratio, and the static “pressure gain” ratio.
These results guided the 1984-vintage data reductions of the series 9 propane runs with emitters. From those, installed hot-burn test inlet performance data matched the cold-flow tests. The streamtube area recovery ratio shows a very strong influence of the so-called “highlight” area versus the true minimum area, when used as the reference area for the calculation.
After the fact, this was entirely expected, based on Reference 2, which (of course) recommends the highlight definition. At the time I did these tests, I had used something pretty close to the minimum area for the reference. It shows explicitly in the data, as a recovery ratio substantially greater than unity, which is completely out-of-line with the usual expectations for ramjet inlets.
See Figures 15 and 16.
Figure 14 – Cold-Flow Inlet Calibration Test Rig
After the series 9 and 10 tests, the air nozzle in the stilling chamber was replaced with a second wooden unit of slightly smaller throat diameter, as depicted in Figure 17. This enabled free jets of nearly Mach 1 speed at the maximum compressor output. Two more test series were conducted with this change, specifically to obtain data at those higher simulated air speeds. These were series 11, using both propane and hydrogen fuels, and series 12, which used the finally-selected ethanol fuel.
The series 12 tests employed both radiometer measurements, and imaging with a thermal imager camera. The fuel vaporizer rig was less successful with a high latent heat pure-substance fuel (ethanol), than it had been with distillate fuels, or with the easily-vaporized propane. With ethanol, it was essentially long-period unstable, with an oscillating fuel flow output. The cycling time was a few seconds.
Nevertheless, using ethanol fuel produced an output spectral power distribution closer to what is needed from the non-gray decoy. The radiometer data clearly showed this. We attributed this difference (with a high degree of confidence) to the lack of yellow carbon glare in the ethanol flame. This yellow carbon glare was quite noticeable in the propane tests, and even more so when using gasoline or jet fuel. The series 12 ethanol runs looked to the eye “positively white” in comparison.
The ethanol fuel injector was stopwatch-and-bucket calibrated for those series 12 tests. Those calibration data are shown in Figure 18.
Figure 15 – Calibrated Inlet Performance Derived from Series 9 Data, Part 1
Figure 16 -- Calibrated Inlet Performance Derived from Series 9 Data, Part 2
Figure 17 – Air Nozzle Re-Work for Higher-Airspeed Test Capability
Figure 18 – Flow Calibration Data for the Series 12 Ethanol Fuel Runs
After these tests, the fuel vaporization problem was conceptually addressed as a hot-gas tap from the forward end of the combustor to the lower-pressure zone at the minimum area of the inlet. Fuel would be injected into this very hot recirculated gas stream to effect rapid vaporization. While the design analysis looked good, that concept never received any testing due to budgetary constraints that essentially stopped all experimental work on the project after late 1984. Some prototype flyable hardware was designed, and a few of those parts manufactured, before all work on the project was completely stopped. It never resumed. So NOTHING is confirmed about any of this!
The ceramic liner material was never characterized, it “just worked”. Density, strength, and thermal conductivity were never measured in any way! However, it handled as if it were about as dense as commercial Styrofoam products. The strength was considerable, considering the rich blow-out instability abuse it endured. Immersed in a 190 F (88 C) air stream, the combustor shell would “barely boil spit” after an hour-long burn test at full rich mixture (theoretically around 3800 F or 2100 C), with but 0.2 inch (5 mm) thickness of the insulation! That indicated very low thermal conductivity indeed!
Table 1 – Summary of “Warm Brick” Burner Tests
These recent compressible-flow cycle analysis results defined the bulk flow conditions inside the combustor well enough to attempt a heat transfer model with a reasonable expectation of success. That model was cylindrical convective-conductive, and based on standard compressible flow models inside and outside the combustor shell. Radiative loss was near zero, as there was no effective path by which thermal radiation could leave the interior. The shell radiation cooling potential was very low.
While the steel shell has a well-known thermal conductivity, the ceramic composite liner did not, so I ran this model parametrically versus conductivity values from “very low” to “very high”. The “best” value of thermal conductivity was that which matched both my recollections of perceived shell temperature, and my observation that the liner surface was often close to melting (3250 F, 1790 C).
Those thermal conductivity results are given in Figures 21 and 22. The highlighted value of 0.02 BTU/hr-ft-F equates to 0.035 W/m-C. Density and strength still lack actual characterization! I have often wondered whether this material might serve as a re-entry heat shield material, the way that the somewhat-similar low-density ceramic Shuttle tile did. But that is another topic for another venue.
#1. Curran, Edward T., “An Investigation Of Flame Stability In A Coaxial Dump Combustor” (dissertation, AFAPL/RJ WPAFB, Dayton, OH), AFIT/AE/DS 79-1, Jan. 1979.
#2. Seddon, J., and Goldsmith, E. L., “Intake Aerodynamics”, AIAA Education Series, 1985, ISBN 0-930403-03-7.
Figure 19 – Spreadsheet Setup for “Warm Brick” Cycle Analysis at Series 9 Propane Conditions
Figure 20 – Spreadsheet Cycle Analysis Results for “Warm Brick” at Series 9 Propane Conditions
Figure 21 – Heat Transfer Model Results for “Warm Brick” Liner Thermal Conductivity
Figure 22 – Heat Transfer Model Results Plotted vs Radius
Epilogue: Some Practical Combustion Device Lessons Learned
Cycle analysis with one-dimensional flow models turned out to be less important than the actual scale-dependent physical chemistry of flame stability, for this “Warm Brick” device. Residence time is proportional to dimension, all else equal, while chemical reaction rates are scale-independent. This alone suggests there is a minimum size below which a thing “just won’t work” with a particular fuel.
Mixing is another very strong determinant of flame stability. Mixing is not proportional to scale, nor is it scale-independent, but it is something in-between. Again, this also suggests that there is a size below which a thing “just won’t work” with any particular fuel. That is precisely one issue (of many) in flameholding.
Those considerations explain why the required nozzle contraction ratio (and internal flow velocities) were so low in the “Warm Brick” device for stable ignition and burning, relative to everything I knew about, before I attempted this project. However, these experiences with the Warm Brick subminiature combustor predate the in-depth understanding of flameholding and flame stability that I was later able to achieve, after returning to Rocketdyne/Hercules. That knowledge is summarized in the “exrocketman” article titled “Ramjet Flameholding” (on this site) and dated 3 March 2020.
The vaporization of fuels of different latent heats and boiling behavior revealed a surge instability in the hot-bucket fuel rig (referring again to the crude hardware in Fig. 10 above). The basic layout was a source of fuel at pressure, led through a copper line coiled in the hot bucket, and from there to the metering orifices inside the test article. See the cartoon in Fig. 23.
The source of fuel-at-pressure was a standard 5-gallon propane bottle (usually around 200-250 psig), or a welding gas bottle (initially 2200 psig), or a pressure tank of liquid fuel pressurized with compressed dry nitrogen (usually pressurized in the 100-300 psig range). All of these pressurization schemes are regulator-controlled. That regulator was physically located about 5-to-10 feet downstream of the test article, and within arm’s reach of the exhaust plume. This allowed me to manually adjust the fuel flow during the test by varying the regulated pressures, while standing immersed in the exhaust where I could smell for unburned fuel. For the open-nozzle tests, I could literally see the flame up the tailpipe.
Fig. 23 – Conceptual Layout and Operation of Fuel Supply
When using hydrogen directly from the welding gas bottle, there was no vaporization problem, as this was simply compressed hydrogen gas. We did not use a pre-heater bucket with this fuel, but the rest of the component layout in Fig. 22 is correct.
With propane in the 0.47 Mach air tests, we found the line just downstream of the regulator, and the sides of the propane bottle, to be cold. This is because the vaporizing pool of liquid propane in the bottle must draw about 150 BTU/lbm of latent heat from itself and from its surroundings, mostly from itself (gets cold). If it cannot draw sufficient heat to vaporize, then it won’t vaporize, pressure drop notwithstanding! The energy to change phase (latent heat) simply must come from somewhere!
There was a cold-line risk of re-condensation on the way to the test article, which we “cured” with the hot oil bucket preheater. We kept the line length from bottle to preheater as short as practical. We also found bottle “freeze-up” occurred at the higher flow rates with the Mach 0.9 airstream tests. We “cured” that by the camper’s expedient of putting the propane bottle in a tub of hot water.
With gasoline and jet fuel, the driving pressures helped us pre-heat the liquid fuel without getting any boiling in the fuel line. Without preheat, there was insufficient air stream heat in the test article to get the fuel to vaporize and burn. With about 300 F preheat, we got all but the “tag-end” of the distillation curve to vaporize upon being injected, due to combined atomization and pressure-drop boiling.
With the gasoline and 300 F preheat, our nominal 100-300 psig driving pressure was apparently barely enough to prevent any significant boiling in the line, so we did not encounter any noticeable problems with vapor lock-induced fuel flow rate surges. With the jet fuel and its lower volatility, we had no real risk of vapor lock surging, but we did see a little more “tag end” unvaporized fuel, indicating a higher preheat temperature was really needed. Both of these are about 150 BTU/lbm latent heat materials.
We did have a real fuel surge problem running neat ethanol as fuel. This material has a far higher latent heat at about 378 BTU/lbm, and it has a single normal boiling point, instead of distillation behavior. At our delivery conditions, the pressure was insufficient to prevent boiling in the line, leading directly to vapor lock-induced flow rate surging! Fuel delivery rates oscillated through about a factor of two, on a long period of several seconds. It would vapor lock, unlock, and relock to cause this surging.
We could not reduce preheat temperatures and still expect to get any flash vaporization upon injection, in hindsight due to that higher latent heat. We could not increase the feed pressures to preclude the boiling without re-working the test article for much smaller injection orifices. That latter is the real design solution to this problem, but we did not use it for these tests! We were able to get our infrared radiometer data from the high points of the oscillating-intensity burn.
While high pressure preheat to get flash vaporization from an atomizing injector is an approach that really works, the equipment to do it is usually large and heavy, too much so for a miniature decoy. The alternative would be to mix the fuel with hot combustion gas to get vaporization, downstream of the metering point. The design difficulty is then to get good mixing of the fuel-rich gas stream with the inlet airstream, without suffering large pressure losses. That seemed the better approach for the flight decoy design. We were never able to test this, though! It is still just a concept!
For an aero-engine application, high-pressure fuel pre-heat with atomizing flash vaporization is likely the better design approach. The sizing of required preheat depends upon raising the liquid to a temperature such that the enthalpy drop across the injection orifice exceeds the latent heat of vaporization. The size of the orifice and the feed line pressure determine flow rate. But, the feed line pressures must always exceed fuel vapor pressure at that high preheat temperature! If this is not done, then vapor lock-induced surging will occur, and at very significant magnitudes. Fuel control then becomes impossible.
As indicated, we never got to test the concept of vaporization by injection into a hot combustion gas stream, followed by injecting that hot mixed stream into inlet air. There is a lot of promise in that notion, but it is fraught with practical difficulties, as well.
Final Comments: IR Emission Characteristics and Towed Decoy Physics
The IR emission characteristics topic has been mostly ignored here, except to say these ceramics were decidedly “non-gray” in their spectrally-dependent emissivity properties. They were non-gray enough to reduce expected radiation in the 1-2 micron band very markedly, to near what they emitted in the 3-5 micron band, despite operating at a temperature somewhere near 3000 F (1650 C). The effective “color temperature” (really the wavelength at peak spectral distribution of radiation) was much closer to typical tailpipe temperatures at full power (but less than those with full afterburning).
Suffice it to say that a great deal of infrared power was radiated by a very small object, whose color temperature and radiated-power in-band looked like a very large jet engine tailpipe at full power. This little emitter would blister my face with radiated heat from some 20 feet away. The large radiated power would be the temperature-to-fourth-power effect, while the color temperature would be the non-gray emissivity effect. Both are critical effects.
Exploring this IR emission topic in more detail would be the subject of some future article, or perhaps even a book relating these experiences. This is an application of otherwise well-established physics.
Another unaddressed topic is aeromechanical in nature: how to tow hard-body decoys stably on towlines, at speeds from very subsonic to low supersonic. The answers are not what one would expect, based on the towed gunnery targets that have been flown for some decades now. Straight tow is the easiest to achieve at all speeds, meaning the tow line extends mostly straight back from the aircraft, although you DO NOT tow the body by its nose! Low or high tow are far, far more difficult to achieve, especially as speeds become high subsonic and the aero forces exceed the weight force. Stable side tow is nearly impossible, even at low subsonic. This applies to radar decoys as well as IR decoys.
Exploring how to tow hard bodies behind aircraft might be the topic of a future article or articles, or even part of a book. The basic rules were invented by my friend Byron Hinderer. I researched the details, and documented what did not work, as well as what did, in the wind tunnel while at Tracor.
The final unaddressed topic deals with what is called “engagement analysis”, where the geometry of the aircraft, the tow, the approach geometry of the attacking aircraft or missile, and the characteristics of the decoy (IR or radar) and the seeker, all interplay. The desired result is an estimate of the kill probability for the attack. The decoy designer wishes to reduce that kill probability to near zero.
Exploring engagement analysis with IR decoys and IR threats might be some future article. Or it might also be part of a book on these experiences. This topic I learned and practiced while at Tracor.
I was once an all-around ramjet design, development, and test engineer, among many other things, including rocket work. This was mostly at a plant in McGregor, Texas, once known as Rocketdyne or Hercules. Part of that reservation is where SpaceX tests rockets now.
I did just about everything there was to do, for this ramjet work. There are very few indeed with knowledge and experience this comprehensive, I was definitely not a narrow specialist! But my knowledge and abilities, in each of all these different specialty disciplines, was actually quite substantial and deep!
My design analyses usually took the form of custom hand-calculations, not just sitting there blindly running other people’s computer codes. (Although, I did use computer codes, and even wrote some myself.) I have informally published several articles on my blog site that describe how some of this ramjet work was done.
Ramjet & Closely-Related Articles (there are others, but these are the best):
11-2-21 The “Warm Brick” Ramjet Device (nonpropulsive application to an infrared decoy) [also the 11-2-21 update to this catalogue list]
10-1-21 Use of the Choked Pintle Valve for a Solid Propellant Gas Generator Throttle
8-2-21 The Ramjet I Worked On the Most
7-1-21 Another Ramjet I Worked On
11-9-20 Fundamentals of Inlets
3-3-20 Ramjet Flameholding
2-16-20 Solid Rocket Analysis (applies to ramjet for boosters)
2-4-20 One of Several Ramjets That I Worked On
1-2-20 On High-Speed Aerodynamics and Heat Transfer
11-12-18 How Propulsion Nozzles Work
7-4-17 Heat Protection Is the Key to Hypersonic Flight
6-12-17 Shock Impingement Heating Is Very Dangerous
12-10-16 Primer on Ramjets
12-21-12 Ramjet Cycle Analyses
These are located on http://exrocketman.blogspot.com, along with many others on a wide variety of subjects.
There is a navigation tool on the left of that page. For the article you want, you only need its publication date and its title. Use the navigation tool: click on the year, then the month. Then click on the title if you need to. The data you need are in these lists.
If you click on one of the figures, you can see all of them enlarged. You see nothing but the figures, though. There is an “X-out” from this view, upper right of screen.
At the end of any given article, there is also a list of search keywords assigned to it. If you click on “ramjet”, you will only see the articles bearing that keyword. The same is true of the other keywords.
Here follows a photo of one of the ramjets I worked on: ASALM-PTV. It is hanging under the wing of an A-7 Corsair-II, an aircraft my father designed. I always considered this photo a sort of “family portrait”.
ASALM-PTV Ramjet Vehicle Underwing of A-7 Corsair-II
Some of those ramjet articles overlap with the next list. That next list is of aerothermodynamics and heat transfer-related articles. Some of these relate to high-speed atmospheric flight, and others to atmospheric entry from space. Those two scenarios are quite different, in that atmospheric flight is a steady-state equilibrium problem, while atmospheric entry is mostly a transient heat-sinking problem. The search keyword for these is “aerothermo”. Clearly, I was adept at multiple specialties.
Aerothermodynamics & Heat Transfer Articles:
4-1-20 Entry Heating Estimates
1-2-20 On High-Speed Aerodynamics and Heat Transfer
1-9-19 Subsonic Inlet Duct Investigation
1-6-19 A Look At Nosetips (Or Leading Edges)
1-2-19 Thermal Protection Trends For High-Speed Atmospheric Flight
11-12-18 How Propulsion Nozzles Work
7-4-17 Heat Protection Is the Key to Hypersonic Flight
6-12-17 Shock Impingement Heating Is Very Dangerous
11-17-15 Why Air Is Hot When You Fly Fast
8-4-13 Entry Issues
3-18-13 Low-Density Non-Ablative Ceramic Heat Shields
1-21-13 BOE Entry Analysis of Apollo Returning From the Moon
1-21-13 BOE Entry Model User’s Guide
8-19-12 Ballute Drag Data
8-19-12 Blunt Capsule Drag Data
7-14-12 “Back Of the Envelope” Entry Model
I was also a rocket propulsion engineer, mostly in solid composite propellants. However, from the chamber outlet through the nozzle, the ballistics of all rockets are the same, including liquid propellant rockets. If you can allow for any gas bled off and dumped overboard for turbopump operation, then the very same ballistics apply, right down to the chamber pressure vs flow rate calculation.
Further, the estimation of vehicle performance from the simple rocket equation can be made quite accurate, if you know how to apply “jigger factors” in the appropriate places for gravity and drag losses, and if you know what values of these “jigger factors” to apply. I have been very successful at doing this kind of work. The following list shows that, and mostly shares the “launch” and “space program” keywords.
Rocket Ballistics and Rocket Vehicle Performance articles:
5-1-22 Investigation: "Big Ship" Propellant From the Moon vs From Earth (added to list as part of Update 5-1-22)
4-2-22 Earth-Mars Orbit-to-Orbit Transport Propulsion Studies (added to list as part of Update 5-1-22)
3-15-21 Reverse Engineering Estimates: Starship Lunar Landings
3-9-21 Reverse-Engineering Starship/Superheavy 2021
3-5-21 Fundamentals of Elliptic Orbits (delta-vee requirements)
2-9-21 Rocket Vehicle Performance Spreadsheet (rocket vehicle performance)
7-13-20 Non-Direct to the Moon with 2020 Starship
7-5-20 How the Spreadsheet Works (Starship to Mars)
7-5-20 2020 Starship/Superheavy Estimates for the Moon
7-3-20 Cis-Lunar Orbits and Requirements
6-21-20 2020 Starship/Superheavy Estimates for Mars
5-25-20 2020 Reverse Engineering Estimates for Starship/Superheavy
2-16-20 Solid Rocket Analysis (solid ballistics & more)
11-21-19 Interplanetary Trajectories and Requirements
10-22-19 Reverse-Engineering the 2019 Version of The Spacex “Starship” / “Super Heavy” Design
9-26-19 Reverse-Engineered “Raptor” Engine Performance (liquid ballistics)
9-16-19 Spacex “Starship” as a Ferry for Colonization Ships
9-9-19 Colonization Ship Study
11-12-18 How Propulsion Nozzles Work (rocket, ramjet, & turbine; plain & free-expansion)
9-11-18 Velocity Requirements for Mars
8-23-18 Back-of-the-Envelope Rocket Propulsion Analysis (rocket vehicle performance)
4-17-18 Reverse Engineering the 2017 Version of the Spacex BFR
10-23-17 Reverse-Engineering the ITS/Second Stage Of the Spacex BFR/ITS System
3-18-17 Bounding Analysis for Lunar Lander Designs (rocket vehicle performance)
3-6-17 Reverse-Engineered “Dragon” Data (rocket vehicle performance)
8-31-13 Reusable Chemical Mars Landing Boats Are Feasible (rocket vehicle performance)
In 2009, I attended an asteroid defense conference in Granada, Spain, as a poster paper presenter. I have since written some articles about asteroid defense. Unfortunately, the asteroid defense capability picture hasn’t changed much since my 2009 attendance at that conference. Again, the latest are the best and most up-to-date. Be aware that “NEO” (Near Earth Object) includes comets as well as asteroids as threats. Comets may be the more difficult to defend against, because of the surprise nature of the detection and orbits. These articles all share the “asteroid defense” keyword.
Asteroid Defense Articles:
8-30-20 Asteroid Threats (current status assessment: not good)
6-3-20 On the Manned Spacex Launch
7-14-19 Just Mooning Around
12-13-13 Mars Mission Study 2013
4-21-09 On Asteroid Defense and a Good Reason for Having National Space Programs
I have also applied my wide-ranging knowledge to the problems of atmospheres to breathe while in space, and the kinds of spacesuits that might best serve our needs. Again, the latest is the best and most up-to-date. But I have been looking into these issues for some time, as indicated by the dates on these articles. These all share the “spacesuit” keyword.
Space Suits and Atmospheres Articles:
1-2-22 Refining Proposed Suit and Habitat Atmospheres (update 1-2-22) best case and easiest-to-remember cases, plus an independent estimate of the utter min suit pressures feasible
1-1-22 Habitat Atmospheres and Long-Term Health (update 1-1-22) adds a long term hypoxia criterion for the habitat in addition to short term criteria for the min-P suit
3-16-18 Suit and Habitat Atmospheres 2018
11-23-17 A Better Version of the MCP Spacesuit?
2-15-16 Suits and Atmospheres for Space
1-15-16 Astronaut Facing Drowning Points Out Need for Better Space Suit
11-17-14 Space Suit and Habitat Atmospheres
2-11-14 On-Orbit Repair and Assembly Facility
12-13-13 Mars Mission Study 2013
1-21-11 Fundamental Design Criteria for Alternative Space Suit Approaches
One of my favorites is the MCP (mechanical counter pressure) version of the spacesuit. This was pioneered by Dr. Webb in the 1960’s as a possible suit for the Apollo missions to the moon. It is not a full pressure suit at all, but essentially a tight garment that simply squeezes the body. It is porous, so that you sweat right through it to cool, just like ordinary street clothing. But this design was tested quite successfully in 1968 for 30 minutes in a vacuum chamber, at way above the equivalent “vacuum deathpoint” altitude. Photo follows:
Webb’s MCP Space Suit: Helmet, Backpack, and Supple Garment Total 85 Lbs
Besides vacuum death and microgravity disease, there is also a radiation hazard to worry about in space. But, it is not quite what you think: there are two completely different hazards to worry about. On Earth, we have two kinds of protection: the atmosphere, and the magnetic field. In low Earth orbit, we have only the magnetic field. Outside the magnetic field, going to the moon or anywhere else, there is no protection. Yet these things can be quantified, and some of it shielded fairly effectively. What got me started on this topic were the dangers posed by the nuclear disaster in Fukushima, Japan. Keyword “radiation”.
Radiation Hazard Articles:
10-5-18 Space Radiation Risks: GCR vs SFE
4-11-15 Radiation Risks for Mars Trip
5-2-12 Space Travel Radiation Risks
3-24-11 Radiation and Humans
3-17-11 Follow-Up On the Japan Nuclear Crisis
3-15-11 On the Nuclear Crisis In Japan
On a lighter note, I have long been interested in pulsejet engines, especially valveless pulsejets. While teaching math at TSTC, Waco, I became involved with mentoring a student who was also interested in pulsejets. I and a colleague assisted this student in making his own valveless pulsejet engines, which attention and involvement also turned this student into an “A” student in math! Keyword “pulsejet”.
That student built a small engine that eventually pushed an old golf cart around, and then a much bigger engine which we together fired up out here on my farm homestead. Photos of the two engines follow:
Smaller Student-Built Valveless Pulsejet Engine (Later Pushed a Golf Cart)
Larger Student-Built Valveless Pulsejet Engine
5-20-12 Recommended Broad Design Guidelines For Valveless Pulsejet Combustors
4-30-12 Big Student Pulsejet an Even Larger Hit at TSTC
3-6-12 Student Pulsejet a Hit at EAA Meeting
11-12-11 Student Pulsejet Project
I have been interested in ethanol fuels since my early days in college. When I went to work for what is now Minnesota State University, after my 20-year career in aerospace defense work ended, I got more serious about it. My next job was at Baylor University in Waco, Texas, and it dealt directly in alternative fuels for aircraft. The scope of that included ethanol (and an ether) as piston-engine fuels, and biodiesel-jet fuel blends as turbine fuels, plus STC work with the FAA, and also experimental engineering research work, as well as classroom teaching.
Not too long after leaving Baylor, I began my own experimental engineering research at home, using E-85 ethanol fuel, and stiff ethanol blends, in a variety of vehicles. Those would include straight E-85 ethanol fuel in an old farm tractor and in an old-time air-cooled VW beetle, plus stiff ethanol blends in a variety of completely-unmodified cars and 4-stroke lawn and garden equipment. I basically recommend up to E-35 blend strength, as a “drop-in” fuel, for just about any 4-stroke piston engine.
The keywords are “ethanol” and “old cars” for most of these articles. Once again, the latest is the best and most up-to-date.
Articles About Ethanol and Ethanol Blends in Vehicles:
9-1-21 Making Stiff Blends At the Gas Pump
11-3-13 Aviation Alternative Fuel Compatibility Issues
11-2-13 An Update on Ethanol Fuel Use
8-9-12 Biofuels in General and Ethanol in Particular
5-4-12 Energy Storage: Batteries vs Unpressurized Liquid Fuels
6-12-11 Another Red-Letter Day
5-5-11 Ethanol Does Not Hurt Engines
2-12-11 “How-To” For Ethanol and Blend Vehicles
11-17-10 Nissan Mileage Results on Blends
11-12-10 Stiff Blend Effects in Gasoline Cars
12-15-09 Red Letter Day: Ethanol VW Experiment Complete
7-1-09 Another Antique Comes Out of Storage
I have returned part-time out of retirement to help a friend with his auto repair business. I was once ASE-certified as a condition of employment while teaching at Minnesota State in its Automotive and Manufacturing Engineering Technology department. Before that, I did most of my own automotive maintenance and repair work. Accordingly, I have posted some articles about basic car care, plus one funny. These all share the “old cars” and “fun stuff” keywords.
Automotive Care Articles:
3-4-22 Understanding Your Tires (added to this list as update 5-1-22)
12-3-20 Blinker Fluid (the “funny”, and it is a sight gag)
8-20-20 Underhood Check
7-25-20 Taking Care of Car Batteries
When I returned to the rocket plant in McGregor for my second employment there, the family and I acquired an old farm outside McGregor as our home. We have been there ever since. This place was largely covered in shin- to knee-high prickly pear cactus, so thick there were few trails through it. After grubbing it out of the house’s back yard with hand tools, I decided there had to be a better way to do this cactus eradication.
I tried a variety of mechanical drags behind my old farm tractor for some 15 years without success. The results were always the same: it looked better for a while, but returned worse than ever before, within months. My neighbor was trying shredding at 1 inch off the ground. Eventually that worked, but required the neighbor to be out there shredding, every single day, the same ground over and over, for 6 (or more) years. The neighbor also tried spraying herbicides on one patch of ground, which took 3 years to show results, but then totally reinfested within another 2 years.
I then tried to build a “scooper-upper” out of scrap steel. The idea was to bust the aboveground cactus loose from its roots, and catch it on a tarp towed behind the “scoop-upper”, for disposal in a burn pit. It completely failed to work, because when the tool hit the cactus and busted it loose from its roots, it fell forward in front of the tool, instead of backward onto the deck. The tool then just ran over the top of the cactus debris. I gave up in disgust when this failure-to-scoop happened.
I went back up a few months later to salvage the steel, and saw something totally unexpected: the cactus was dead and gone wherever the tool had been towed! Grass was growing in the cow pasture where the cactus had been. It did not take very long to understand that the aboveground cactus foliage had been crushed and damaged passing underneath the heavy tool, such that the pads dried out and died, before they could put down new roots from the thorn sites in contact with soil. They had completely composted away over those months.
I “played” with this tool to get it just “right”, and started killing acres of prickly pear quite effectively, and with very little time and effort involved. In fact, I still have this very same experimental prototype, and it still works today. This prototype led to me filing a patent on the cactus tool in 2002.
I revised the design to something more producible from real steel stocks, and built two production prototypes that worked just as well as the original experimental prototype, but were easier to build. Then, with the patent in hand as of 2004, I began building and selling these tools to the public. My first customer wouldn’t wait for a real production tool, and insisted on buying one of the two production prototypes. I still have the other one. I still use it, and it now serves as an experimental test bed for new features, too.
As time went by, it quickly became apparent that other folks had rockier land, or land with tree stumps. I changed the design twice, to add a heavier stabilizing snout, plus a “barge front” wedging surface to get over small rock outcrops. This was quite successful, and is embodied in the tools still built and sold today.
A close friend wanted to do cactus-killing for hire, and bought a “one-off” design from me. I also helped him build and modify a few more tools, until the “commercial version” was defined: a really tough snout, a big “barge front”, and retractable wheels to facilitate stepping over obstacles, plus easier loading up ramps onto trailers.
When that friend retired, I revised his “commercial” design into something that used a common core tool chassis with my “homeowner grade” plain tool. This common core chassis had the big barge front, and used either a tough snout for the “plain tool”, or a longer tough snout for the “hydraulic tool”, that was also fitted with retractable wheels operated hydraulically. I sell both versions to this very day. Both are towed on a chain bridle behind a farm tractor’s drawbar.
I am working on a third version that could be an alternative implement affixed to the hydraulic boom of a skid-steer loader. It uses an already-available “universal” adapter plate to accomplish this, as a quick-change item. There is nothing to report here yet about that project, but the “plain” and “hydraulic” tools are well-described in a series of articles on “exrocketman” under the keyword “cactus-killing”.
These two versions are shown in the photo, with the plain tool in the foreground, and the hydraulic tool in the background.
Foreground: Plain Tool; Background: Hydraulic (Wheeled) Tool
Articles Related to Cactus Eradication:
2-9-17 Time Lapse Proof It Works
7-30-15 New Cactus Tool Website
1-8-15 Kactus Kicker Development
1-8-14 Kactus Kicker: Recent Progress
10-12-13 Construction of the Plain Cactus Tool
5-19-13 Loading Steel Safely (Cactus Tool)
12-19-12 Using the Cactus Tool or Tools
11-1-12 About the Kactus Kicker
12-28-11 Latest Production Version of the Kactus Kicker
12-1-21 The Seal Failure in the SRB That Doomed Challenger
12-10-20 Spacex Test Flight Results in Explosion
9-1-20 On the Beirut Explosion
5-4-18 Some Thoughts on the Anniversary of the West Explosion
11-1-14 Two Commercial Spaceflight Disasters in One Week
7-9-13 On the Asiana 214 Crash
7-9-13 On the Train Wreck in Quebec
4-18-13 Fertilizer Explosion in West, Texas
9-23-11 Air Races, Air Shows, and Risks
6-3-10 Plenty of Blame to Go Around for the Disaster in the Gulf
5-20-10 It really was the North Koreans who sank the South Korean ship