An Ex Rocket Man's Take On It

Launch Cost Data

2012-01-09T18:40:00.000-08:00

I took a look around on the internet, visiting various company and industry watchdog sites, looking for payloads deliverable to low earth orbit (LEO), and the launch costs associated with them. I found pretty good payload information. Launch costs are little more speculative, especially as some of the data was not quite current.

I looked up some pretty reliable data for the Spacex family of Falcon rockets. Falcon 1 and Falcon-9 are already flying, and the cost data were available directly on the Spacex site. Falcon-Heavy is supposed to fly for the first time this year.

I found some pretty reliable payload data for a few members of ULA's Atlas-5 family. That's now a joint venture of Boeing and Lockheed Martin. The Atlas family was formerly operated by Lockheed Martin. The cost data came from a watchdog site, but I think it is fairly current. I have data for the -401, -551, and -HLV configurations.

I also found some payload data for a few Delta 4 configurations on the same ULA site. Delta is a Boeing operation, built in ULA facilities. I found cost data for only Delta-4 Heavy on a watchdog site, but it's about 5 years old. So, if anything, today's launch costs for Delta-4 Heavy would be higher still.

I processed these data as payload mass delivered from Canaveral to LEO, and as US dollars per unit mass of payload (payload mass divided by launch cost). The results are plotted here in metric units ($/kg vs metric tons) and in US customary units ($/lb vs US tons). I expected to see the "scale effect" operating (lower cost per mass at higher masses), and I did.

But, there are differences among the vendors. Remember, when assembling something large in LEO, it is the cost per unit mass that determines your overall project launch costs. Once the Spacex Falcon-Heavy is flying, it will be very hard to beat, in terms of unit cost and absolute payload tonnage. Judge for yourself ----

In metric:

In US customary:

On Somali Pirates: "I Told You So"

2012-01-02T12:43:00.000-08:00

From an article by Neal Ungerleider, on MSNBC.msn.com, 2 January 2012

“Commercial shipping is one of the most cutthroat industries in the world, and shipowners don't have the budgets to spend on space-age defenses. The recent drop-off in piracy has been due to the increased use of armed guards, not technological innovation, which in turn have been made economically feasible by skyrocketing ransoms and lengthening captivity periods. One insurance company issued a stat a few months back that 80 percent of pirate attacks were being repelled by armed guards, and no vessel employing them has been hijacked.”

http://technolog.msnbc.msn.com/_news/2012/01/01/9834437-fighting-somali-pirates-with-science

Now for the “I told you so” part:

This was the first, best recommendation that I made, of three, nearly 3 years ago on 4-11-09, on this site, in an article titled “3 Solutions to the Somali Pirate Problem”.

The Old Train Still Runs!

2011-12-30T10:47:00.000-08:00

For the first time in a few years, I set up my old electric train in the shop. I got this train as a very young boy, Christmas of 1952. My dad and his next door neighbor built the train board layout in 1954. Some of the items, including two more freight cars, were added between then and about 1960.

The engine and tender is a 4-6-4 coal-fired steam locomotive of the type that was used to pull passenger trains on the old New York Central railway. Lionel called this particular engine 2046, and used it in more than one of their train sets. My original set included a silver tank car, silver box car, black gondola car, and a caboose. The yellow barrel car and the red explosives box car got added later.

The first image shows a good close-up of the engine and tender:

The second image shows the whole train board layout. The inner (third) loop of track is something I added about a decade or so ago. The original setup had two concentric loops connected by 4 switches, all 1954-ish vintage track and switches.

Along the way, the two extra freight cars, the crossing equipment, the water tower, and the beacon got added, as Christmas presents, if I remember correctly. My paternal grandmother gave me the gantry crane, which still works. It rotates, moves up and down, and the electromagnet still picks up iron things.

I put some miscellaneous small toy cars into this set-up. I also built the loading ramps out of scrap wood from an old VW bus wooden headliner. My painted-paper landscape simulation from over 20 years ago has deteriorated past repair. I need to replace it, and add some more hand-made buildings. I'll do it, once I retire.

The third image is my finger pointing at the entry in an original 1953 Lionel catalog for the exact set that is my original train: 1505WS, which was $49.95 in 1952-dollars, and still the same price in 1953. My good friend Harry Petersen in Minnesota found that catalog and sent it to me. He, too, is a model railroad enthusiast.

Below I have embedded a video clip taken with my wife's camera of this train running on that train board. It no longer smokes, but the whistle still blows. This thing is 59 years old this year, and it still runs! In spite of all the mistreatment I gave it as a child. Lionel certainly made a good product.

This posting is just for fun. Hope you enjoyed looking at it.

GW

Latest Production Version of the Kactus Kicker

2011-12-28T10:00:00.000-08:00

For those of you wanting to know about my cactus tools, here are some pictures of the latest production model with the tougher snout and bigger barge front. These are from my wife's computer, file number 2010-04-25. I believe they were taken after construction of serial numbers 047 and 048.

This is a "machine" with no moving parts, towed on a simple chain bridle behind any tractor with a drawbar. It kills prickly pear cactus "in situ", without pick-up and disposal of the debris, and without chemicals. It's just driving-a-tractor work. You do it several times, for a full eradication. See http://www.txideafarm.com and go to the cactus eradication sub-page for a good description of how it really works.

photo 040 How to Hitch-Up

It really is just that simple. Flip the loop in the tow bridle over a trailer ball on your towbar. If you do this with a 3-point rig, be sure it is braced for sideways loads over 1000 pounds per tool (you can tow more than one at a time). You will incur forces like that when you turn.

photo 041 What the Bridle Looks Like All Hitched-Up and Ready to Tow

Be sure the bridle is not in the lift configuration, pinned up with a bolt over the tool's center of gravity. It needs to make a big Vee, you tow from the corners of the deck. The snout just stabilizes it like a gigantic, super-tough sled runner out front. The chain through the snout braces just limits up/down and side-side travel on really rough ground. It should be slack, otherwise.

photo 042 How-To Pry-Up the Tool to Get at What's Underneath, or Store It

Back up the tractor and slack the chain, then un-hitch it. You will need about a 6-foot prybar and a 2-foot piece of small angle iron. Use the prybar as the photo shows to get leverage to lift the tool up onto its rear edge, then prop it in place with the angle iron under one of the skids. The "tongue load" on the snout is just too high to do this without a good prybar.

photo 045 Proper Stowage Without Killing Grass

Once propped up, you can remove any debris accumulated under the tool that makes it ride off the ground. Old barbed wire and certain kinds of vine-like weeds are prone to do this. Just kick or hoe them out from underneath, and you can lower the tool with the prybar, re-hitch, and resume work. This is also a very good way to store the tool in the pasture between treatments, since it cannot kill a whole big patch of grass while tipped up on edge like this. This is how I store mine.

photo 037 How to Pick-Up the Tool with Its Own Bridle

If you pull the tow bridle aft, you can pin it together with the extra 2"L 3/8 UNC bolt, nuts, and washers that I provide with every tool. If you have serial number 047 or 048, you might have to re-rig the snout travel-limiter chain slightly to do this, but I generally already have it rigged for lifting easily, right from the shop (from serial number 049-on). The center of gravity is just between the rear of the snout tube and the front edge of the big ballast bar flat. Pin the bridle together there, and pick it up at the pin point as shown.

The snout travel-limiter picks up the forward load of 3-places, the chain towers being the other two. Be careful, this thing weighs 600-700 pounds. But most tractors now have hydraulic buckets. Just use a tow chain with hooks, and pick the tool up with the bucket, and put it right where you want it (pick-up bed or flat trailer).

GW

FTL Neutrinos Update

2011-12-21T09:41:00.000-08:00

I posted an update to the faster-than-light neutrinos article. Scroll down to it dated 10-9-11 and titled “Faster-Than-Light Neutrinos? Maybe! Their Meaning? Arguable!”

GW

Reusability in Launch Rockets

2011-12-14T18:47:00.000-08:00

A group of folks I correspond with (at the forums on NewMars.com) has been discussing reusable launch rocket possibilities. One of the names they use is “big dumb booster”, or BDB. My own opinion is that reusability is incompatible with the low inert mass fractions used in the stages of typical launch rockets today: too light is simply too fragile. I do know from their website that Spacex is interested in reusing the first stage of their Falcon-9 booster, but that their results so far are unsuccessful. So, my analysis results here should be of interest, both to my correspondees, and to Spacex.

Spacex’s Falcon-9 is a two-stage rocket with kerosene-oxygen engines in both stages. It features an interstage ring and a payload shroud (on the satellite version) that I assume both get jettisoned at staging. The same engines are used in both stages, except that the one in the second stage has a longer bell than the nine in the first stage, and the first stage engines see atmospheric backpressure.

Baseline Falcon-9 Performance Estimate

I looked up most of the basic engine and vehicle data from Spacex’s website, for Falcon-9 as a baseline case, and reverse-engineered the rest. Here it is, summarized, in Figure 1:

Figure 1 – Baseline Falcon-9 Data

These performance data were computed with the simple rocket equation, and some experiential “jigger factors” that knock down ideal velocity increments to more realistic values. The other choice for analysis is a real trajectory computer code, either two-dimensional or three-dimensional, which is a complicated thing to set up and to use. I used the simple analysis approach to set up actual computer trajectory analyses, for the Scout launch vehicle at LTV Aerospace, about 4 decades ago.

Here, I used a “jigger factor” of 1.10 to knock down the first stage ideal velocity increment, because that stage sees air drag, and flies mostly vertically, so that gravity drag is significant. For the second stage, I used 1.05, reflecting flight in vacuum, mostly but not entirely horizontal. The final summed velocity increment I estimate for Falcon-9 is about 26,900 feet/second, or 8.19 km/second, which is remarkably close to the orbital velocity at low altitudes (about 7.9 km/second). It’s close enough that any simplified design trades made under these assumptions are realistic enough to be useful.

I looked at two potential solutions to the trade-off between extra structural weight for reusability, and reduced payload fraction that increases the price per unit payload delivered to orbit. One was to retain the basic two-stage design, and increase the size of the first stage to compensate for added inert fraction, at constant mass ratio. The other approach was to replace the two-stage design with an equivalent three stage design, keep the top two stages as throwaways, and increase the first stage size to compensate for increased first stage inert weight fractions. Both were done at constant delivered payload weight.

Two Stage Analysis with Heavier Structural Inert Fractions in the 1st Stage

The payload is exactly the same as baseline. I assumed the payload shroud weight to be proportional to the maximum payload weight it contains at 15.18%. There are no changes to the second stage weight statement or performance values. The interstate ring weight I assumed proportional at 0.815% to the weight it carries, in this case the second stage ignition weight. It is the first stage weight statement that varies, but at constant mass ratio, so the propellant weight fraction is the same as baseline in all cases. The equation relating mass ratio MR and propellant weight fraction fprop is:

fprop = (MR – 1)/MR

Now, 1 – fprop is the total of the inert mass fraction and the stage payload mass fraction, where the first stage payload comprises the ready-to-ignite second stage, the interstage ring, and the payload shroud. I looked at the baseline, twice, and three times the first stage inert weight fraction, scaling up the first stage ignition weight to match. The resulting weight statements are given in Figure 2. Bear in mind that the delivered stage performance data are identical to baseline, since the mass fractions are identical to baseline.

Three Stage Analysis with Heavier Structural Inert Fractions in the 1st Stage (Only)

I had to allocate velocity increments among the three stages in some logical fashion. I chose to make the second and third stage mass ratios 5 like the Falcon-9 second stage, and my first stage mass ratio 4, like the Falcon-9 first stage. I used “jigger factors” of 1.10 and 1.05 on my first and third stages, similar to the Falcon-9 first and second stages. I used an intermediate factor of 1.07 for my second stage. My first stage Isp was 289.5 sec, like the Falcon-9 first stage. My second and third stages used Isp = 304 sec, like the Falcon-9 second stage. The corresponding exhaust velocities are 9314.4 and 9780.9 ft/sec.

I computed the sum of the estimated actual velocity increments to be factor 1.5406 too high, so I knocked down each stage’s velocity increment by this factor, and recomputed mass ratios as 2.45935 for my first stage, and 2.84251 in my second and third stages. I ran the design study to the same payload as Falcon-9, with the same shroud weight, and two interstage rings at 0.815% of the stage weights above each ring. I assumed that interstage ring 1-2 and the payload shroud drop off with stage 1, and that interstage ring 2-3 drops off with stage 2.

The payload is exactly the same as baseline at 23,050 lb. I assumed the payload shroud weight to be proportional to the maximum payload weight it contains at 15.18%, for 3500 lb. There are no changes to the second or third stage weight statements or performance values as I changed first stages. The interstate ring 2-3 weight I assumed proportional at 0.815% to the weight it carries (in this case the second stage ignition weight) for 606 lb. Interstage ring 1-2 is 0.815% of stage 2 ignition weight, for 1999 lb. It is the first stage weight statement that varies, but at constant mass ratio, so the propellant weight fraction is the same as baseline in all cases, and so is the performance.

For the “baseline” three-stage inert fractions, I assumed 5% for my first stage, very similar to the multi-engine first stage of Falcon-9. I used the same 4.2% for my third stage as for the single-engine second stage of Falcon-9. My second stage has an intermediate inert fraction of 4.6%, chosen to reflect only a few engines in the second stage. The weight statements for the trade study are given in Figure 3. Bear in mind that all three versions of the three-stage vehicle have exactly the same estimated velocity performance, also shown in the figure.

Figure 2 – Weight Statements for the Two-Stage Reusability Trade Study

Figure 3 – Weight Statements for the Three-Stage Reusability Trade Study

Note that in both Figure 2 and Figure 3, I have included the overall payload weight fraction, computed as payload weight delivered to orbit Wpay, divided by the stage 1 ignition weight, which is the launch weight WL. (In the context of this analysis, the term “weight” really refers to mass.) In both trade studies, payload fraction decreases as stage 1 inert weight increases, exactly as expected. I was surprised and pleased to see that the baseline throwaway 3-stage option had a slightly higher payload fraction than the corresponding baseline throwaway 2-stage option. This and the slopes of the trends did seriously impact the final conclusions.

Trajectory Comparison

The final trajectories are compared in Figure 4. Both the 2-stage and 3-stage vehicles follow similar paths to the same orbital insertion conditions, at the same altitude (in the vicinity of 200-300 miles, or 300-500 km, up). Only potential re-use of the first stage was considered, for either configuration. A first stage fallback is indicated for each. Reentry velocity is simply assumed the same as the first stage burnout velocity. They would be comparable, in any event. Noting that reentry gets really challenging much above 10,000 feet/second (near Mach 10), I see little point to trying to make the second stage of the 3-stage vehicle reusable. It simply comes back too fast to be readily survivable.

Figure 4 – Comparison of Trajectories for 2-Stage and 3-Stage Vehicles

Figure 5 – Comparison of 2-Stage and 3-Stage Results

Payload Fraction Results Comparison

The payload fraction vs first stage inert fraction data are plotted in Figure 5 for both the 2-stage and 3-stage vehicles. The trends are reasonably linear-looking over the ranges computed, but at different slopes. As expected, the 3-stage vehicle design is less sensitive to first stage inert fraction than the 2-stage design (3-stage having the shallower slope). I did not really expect to see the baseline 3-stage vehicle to have a slightly-higher payload fraction at the baseline throwaway inert value, but it did.

Between the higher baseline throwaway payload fraction, and the shallower slope with first stage inerts, it appears that at 10% inerts in the first stage, the 3 stage vehicle has a payload fraction near 2.8%, while the 2-stage vehicle is down near 2.2% at the same 10% inerts. 10% inerts in the first stage is of enormous interest, because that is close to the inert fraction of the Space Shuttle solid booster motors, which actually were reusable most (but not all) of the time. That’s about the level where your tankage becomes strong enough to be pressure vessel-capable, as well as survivable for ocean impact on parachutes. Tankage that is pressure vessel-capable might as well be used as a pressure-feed system, eliminating the weight, cost, and reliability risks of turbopump machinery.

Conclusions

It is clear the 3-stage option is more tolerant of higher inert weights in the first stage. Combine this with a lower first stage fall-back speed, and reusability seems more certain at 10% inerts, and with a higher payload fraction (nearly 3% 3-stage vs only a bit over 2% 2-stage).

Accordingly, 3 stages is a better option than 2 stages, if the first stage is to be reused. The drop from 2-stage non-reusable payload fraction is actually quite small (3.1% to about 2.8%). This is because the all-throwaway 3-stage vehicle actually has a better baseline throwaway payload fraction than the 2-stage (3.3% at 5% inerts, vs 3.1% at 4% inerts).

This does raise the question of whether 4 stages might allow first stage reusability at even better payload fraction, or else allow the same payload fraction with both first and second-stage reusability. I leave that for others to investigate.

The main lesson here is that you really do have to do something different in order to get a different result. Reusability will require a greater inert weight fraction to cover recovery gear, and to confer the strength to survive better. Practical reusability simply cannot happen in the 4-8% inert range.

This study points toward 10% inerts in the first stage, at the very least. The more first stage inerts you have to “cover”, the more stages you need to use, to be tolerant of lowered mass ratio in each stage.

But at least we know the job really can be done, and here is one well-proven way to do it (more stages).

Gas Fracking: Good or Bad? Depends!

2011-11-13T09:55:00.000-08:00

Recent news reports published by the internet news services tell how the EPA is seriously investigating complaints related to natural gas fracking (hydro-fracturing) near Pavillion, Wyoming. Those complaints include contamination of water supplies by methane and by toxic fracking chemicals.

I looked up the geography and geology of Pavillion: it lies in the western half of Wyoming, a region dominated by the Rocky Mountains. There will be sediments in the basins, but the fundamental underlying geology is contorted and fractured mountain zone rock.

As a result, I am entirely unsurprised that both natural gas and fracking chemicals are finding their ways into the groundwater. I am surprised that how this can be, is a still a matter of legal debate.

I am no geologist, but even I can understand what is happening, and how, and I published it as a guest column in the Waco Tribune Herald last May. Here is the original submitted text for that column, with some emphasis added now:

Coming Even Cleaner on Fracking” (submitted 5-26-11/published 5-28-11)

The “Trib’s” editors recently ran a very nice editorial on the controversy surround the process of “fracking” (short for “hydraulic fracturing”) for natural gas in shale. This article neatly laid out the two sides of the public debate, which is centering mainly on whether or not there are undesirable side effects.

I find it very interesting that the studies are "still inconclusive", seeing as how the field data is very indicative of what actually happens. It's not a simple either-or situation, it’s geology-dependent, and this is completely left out of the current public debates.

Here in Texas and nearby states, the rock layers are old seabed sediments, more or less level, and are relatively intact. Few paths exist across these layers for oil and gas to migrate upward. That is why fracking has few side effects in this part of the country. The most notable exception has been very minor earthquake tremors induced from the disposal of used fracking fluids by deep well injection.

In Pennsylvania and the other states in the Appalachian mountain zone, there have been widespread complaints about natural gas getting into groundwater, leading to fire and explosion incidents when turning on the water tap. These are real incidents, and are easy to understand if one simply looks at the geology below the surface.

In a mountain zone, the rock layers are highly contorted, fractured, and thoroughly broken-up. There are many paths for oil, and especially the far-more-mobile gas, to migrate to the surface. It is entirely unsurprising, and in fact quite predictable, that this very mobile gas, once released from a deep shale, should migrate upward and contaminate near-surface water supplies. It does so by dissolving into the water under earth pressures, similar to a carbonated beverage.

The solution to the exploding kitchen faucet problem is simple: fracking for gas is OK in continuous-layered sea bottom sediment zones, but not OK in highly-fractured mountainous zones. So, we don't frack there, period. Those gas deposits await a still-undiscovered recovery technology with fewer side effects, more suited to that kind of geology.

This does mean that the agencies regulating gas leases actually do have to regulate, and sometimes to deny permits, unaccustomed as they apparently are to such activities.

The processes of fracking and fracking-fluid disposal were specifically exempted from EPA regulation under the Clean Water Act. This happened in that secretive energy company meeting at the White House during the last administration. It is known as the Halliburton exemption.

However, the injection of diesel fuel into the earth is actually still regulated. While fracking fluid is mostly water plus a little sand or glass beads, the most common liquid trace additive in all these "secret" recipes is diesel fuel. If those recipes were widely revealed, the use and disposal of these fluids would come under direct EPA regulation again, meaning only that they take a little better care doing what they already do.

In that event, fracking for gas would still be quite profitable, just not quite as much as it is without any regulation at all. But fewer folks suffer the side effects, and that’s a good thing.

Student Pulsejet Project

2011-11-12T12:52:00.000-08:00

TSTC welding student Justin Friend made the Waco paper Friday 11-11-11, with a pulsejet thruster he built from plans he found on the internet. He had already tested this device himself, but brought it out to the TSTC airport apron for a demonstration test Thursday morning. Most of the attending crowd were aviation maintenance and welding program students and faculty. This was a personal project for Justin, not a class project. I called Bill Whitaker, the editor at the Waco “Trib”, and he sent a reporter.

Justin is a college algebra student this semester with my colleague Otto Wilke, who attended the demo test, as did I. (Another math department colleague, Doyle Ware, also went with me to see the test.) Justin sought sheet metal geometry help from Otto, and operating and safety advice from me, once he found out that Otto and I are engineers. His welds were obviously very good, as the pulsejet tube runs very hot in places. No cracks or flaws of any kind have turned up to date.

This pulsejet tube is valveless, so there are no moving parts at all. It is a “folded pulsejet”, so that the back-spit from the short inlet contributes to its thrust. This one is a nominal 50 pound thrust device, big enough to push a go-kart around. It runs on propane. Justin cut the parts from flat stainless steel sheet, rolled them up, and welded them together, excepting the return-bend tubing. This return bend is on the exhaust side, and is the hottest part of the structure. It glows in broad daylight when running throttled up. I got copies of photos Justin made during his first tests. Two are here.

P 1030654

This device has a spark plug on the side of its combustion chamber, and a propane injection manifold tube across the inlet right at the dump into the chamber. One starts the spark, some starting air from a leaf blower, and the propane, to light it off. Once running, starting air and spark are no longer necessary. It throttles up and down a wide range of thrust by simply raising and lowering the propane feed pressure.

P1030658

This thing is dangerously noisy: I estimate around 130-135 decibels, so ear protection is a necessity. You can feel the sound waves beating on your stomach. At the TSTC demo runs, you could feel the concrete airport apron shake beneath your feet. Unlike all other forms of jet engine, pulsejets “sing” at a definite frequency, the rate of the pulsed fuel-air explosions inside the tube. This size tube “sang” at about 80 Hertz, like an earth shakingly-loud operatic bass.

I haven’t heard noises that loud in decades. Being a part of this young man’s project was a huge amount of fun. I actually knew something about this engine and could help Justin with it, because decades ago I researched the military work done on them in the 1950’s and 1960’s. I always wanted to build one myself, but never actually did it. (That may change, this was just too much fun.)

Added 11-13-11:

Here is the first of two QuickTime Movie (.MOV) files I got from Justin. You can see the tube slowly warm up and quit spewing unburned fuel from the inlet (a thin whitish spray). You can also get a sense of the 80 Hertz "singing" tonal quality of the sound, but no real hint of how loud it was. Upon shutdown, the flames from the inlet are propane residuals from the fuel line venting into a very hot environment.

Here is the second movie file from Justin. In this one, the tube is quite warm and operating very well at near-full thrust. You can see him disconnect the spark while it runs, with absolutely no effect upon the operation of the tube.

Added 12-7-11:

Since this article originally posted, Justin has mounted his 50-pound thruster to an old golf cart, and driven it at the Hearne, TX airport (an uncontrolled field). The video is of a pass he made after the tube was fully warmed up. The speed is close to the control limit for that cart.

Justin has since begun procuring parts and materials for a much larger thruster.

October 29, 2011, Update

2011-10-29T15:00:00.000-07:00

Here follows various updates regarding various projects I have been pursuing. The list is not fully comprehensive.

Mars Mission / Paper Presentation at Mars Society Meeting August 2011

That project is completed. I stand by my earlier statements that we could put men safely on Mars and return them safely, for under something like $50 billion. It would take a space agency that we do not currently have, and a contractor base that we do not currently have, to accomplish this. The real take-home lesson is that the agency and contractor base we have built and maintained all these decades is the wrong setup.

I did revamp the program back-up to solid-core nuclear “slowboat” from the original VASIMR-based fast-trip. The baseline is still gas-core nuclear fast-trip. As it turns out, VASIMR is just another electric ion drive, no better than the others in terms of performance potential. None of those is suitable for fast manned trips to Mars. I did add some better orbit trajectory estimates. See the 9-6-11 posting “Mars Mission Second Thoughts Illustrated” for those details. The original posting of paper content was 7-25-11 “Going to Mars (or Anywhere Else Nearby) the posting version”.

Ethanol Vehicle and Engine Work

That is now completed. I have decided that it is easier and more effective for most people just to use stiff gasohol blends E-20 to E-35, than it is to come up with shade-tree conversions for still-higher blend ratios. For some vehicles and engines, conversions are easy, for others, not so much.

This work is documented well enough in postings 5-5-11 “Ethanol Does Not Hurt Engines” and 2-12-11 “”How-To” for Ethanol and Blend Vehicles” for others to use the information. Anyone can learn to make E-30-something blends in any vehicle. For me, this has become routine operation of an F-150, a Nissan Sentra, two lawnmowers, a wood chipper, and a garden tiller on E-30-something blend, all completely factory stock. I still run my slightly-modified Farmall tractor on straight E-85, but have re-mothballed the modified 1973 “ethanol VW” and the unmodified 1960 “blend VW” against future needs.

My two-stroke chain saw and weed eater seem to run just fine on the E-10 blend they now sell as regular unleaded gasoline, with one exception. The weed eater is having age-related fuel line replacement troubles, but also seems to suffer from some poor design choices, as well, regarding how these lines connect to the fuel tank and the other components.

I have never run stiffer blends in either the weed eater or the chain saw. Is there is a materials incompatibility problem with the weed eater? I don’t know yet. Is there a “bad design” issue with loosening connections? Yes, of that I am sure. Is there an overheat problem that stiffens fuel line hose? I think so, and it’s another “bad design” issue unrelated to fuel composition.

Ramjet Engineering

In June of 2010 I paid a visit to Jeff Greason and his crew at XCOR Aerospace, Mojave, California. This was in regard to a future space launch project of theirs that involves ramjet propulsion. They had been unable to locate an all-around ramjet expert, until they ran across me by accident. I had not done such engineering since the old rocket plant in McGregor closed, at which time I was laid off: November 1994. It was not very long after that layoff, that most serious military ramjet work simply dried up in this country (although, not overseas).

As it turns out, of the few of us that I considered to be all-around experts with significant real design and test experience, most (or maybe all) the others are now dead. I seem to have outlived them all, thus becoming pretty much this country’s last living expert in that kind of propulsion.

Over the last year or so, I dug out some of my old ramjet stuff and got back into the “swing of it”, in order to be “back up to speed” when XCOR needs my help, perhaps next year. I began by looking at high-speed systems for orbital launch, which aligns with their project. This took the form of pencil-and-paper stuff in the odd evenings, and eventually evolved into a two-stage horizontal takeoff/landing aircraft, the lower stage being parallel-burn rocket and ramjet. We paralleled each other in this.

As best I can figure, the ramjet strap-on assist idea for vertical-launch rocket vehicles is more of a low-speed design system. That idea is less worked-out than the high-speed two-stage airplane, which does seem both feasible and attractive. But I do believe that my top-level conclusions are correct. I do not yet have the software tools necessary to do this kind of work for XCOR or anybody else, seeing as how no computer folks still support the old DOS-based programming languages that I learned long ago.

I have put a lot of “typical” ramjet performance estimates up on this “exrocketman” site, in articles too numerous to catalog here. More will be forthcoming, as I develop stuff into results one can truly trust. I have put some efforts into converting my old “smarts” and programs into modern Excel spreadsheet format. I now have a sizing spreadsheet that works for the old lower-speed “stovepipe” designs, plus a performance mapping option verified to work as the nozzle unchokes. It has become clear that the nested iteration-loop character of these calculations demands a real computer code, not a manually-converged spreadsheet, which is simply too slow and labor-intensive to be practical.

I already had a sizing code in an advanced BASIC language that works for high-speed designs. The corresponding performance-mapping code does not yet work. The corresponding codes for low-speed designs are nowhere-near in working order at all. There is available to me only a single obsolete computer with an old Windows-98 operating system, that will even run the programming language. This is a real practical problem yet to be solved.

More recently, I made contact with Aerojet, who is the current inheritor of the gas generator-fed ramjet work I did long ago at the old McGregor missile propulsion plant. It seems there might be a need for my skills once again. We’ll see. I have heard nothing positive back since that initial contact, though.

Meanwhile, I have been documenting my old procedures and methods in a series of reports that I keep in a “ramjet how-to” notebook. It’s not complete, but I do have a document regarding the high-speed engine cycle analysis, and one giving estimated spike inlet performance, plus another one documenting low-speed engine cycle analysis, with flameholding and heat protection “how-to” for both speed regimes. I think eventually this notebook will become a book on the “how-to” of ramjets.

One thing I do know: it is very easy to document science, but it is very, very hard to document art. Most of ramjet engineering is art, not science! And the science is hard enough, being way more complicated than rocketry.

Cactus Tool Stuff

I get more email inquiries now, but still very few sales. Many folks seem to be finding the cactus page on the “txideafarm” site. These seem to be generally younger folks. But, as it was with the old ranchers looking at magazine ads, few seem willing to believe it really works.

But, it does work! My place, and everything my friend Dave Gross has done, proves it.

I build and sell about 1 or 2 tools a year, these days. I sell about 2-4 plans sets per year. The rigors of fabricating pieceparts is taking an increasing toll on my aging body. Steel is now about 3 times as expensive as when I started doing this, 6 years ago. This all makes me wonder if I should not give up fabrication in favor of just selling plans. Maybe it is time to license construction to a bigger company.

Dear readers, please weigh in on this: should I continue selling tools, or just plans? There’s comment buttons available; use them.

Reno Air Race Crash

I posted an article 9-23-11 about the fatal crash of the Galloping Ghost at the recent Reno air race. That article says the elevator trim tab failed, most likely due to aeroelastic or other structural divergence effects at high speed.

Tab failure in the P-51 at race speeds, modified or not, leads to a sudden pitch-up condition. That leads quickly to pilot gee-induced unconsciousness, at best. The cure for spectator fatalities is not to position any spectators under expected flight paths.

I still see no reason to revise this posting until the NTSB has a chance to report its findings, perhaps sometime in 2012. And I doubt there will be any need to revise it then.

Energy Resources

I had published an article here 3-14-10 “Drill Here, Drill Now, Pay Less?” that dealt with a purported vast US oil resource named “the Bakken”. That resource is really shale tar, and is not (and will never be) “drillable oil”. Since then, I have become aware of the fact that not all of that rock unit is shale. There is a substantially-more porous dolomite layer in the Williston basin that actually does contain a light crude recoverable with hydro-fracturing technology.

In an article dated 9-5-11 “Surprise, Surprise: Oil Boom in the Williston Basin (“the Bakken”)”, I took on the size and recoverability of that resource. It is significant, but no “game-changer”, as some would have you believe. I concluded that yes, we should go get this oil. But, no, it will not save us from foreign oil dependence. I still see no reason to change those conclusions.

The fundamental economic problem is that our western economy was designed to run on cheap energy, primary of which is transportation fuel. Energy today, particularly transportation fuel, is no longer cheap. Therefore, we have economic recession/depression (choose your word). Government policies (from either side) have nothing to do with boom or bust conditions. Only energy prices really matter. The proof of this thesis is in the data posted 2-4-11 “Oil Prices, Recessions, and the War”.

It is hard to argue with data, is it not?

My most important point is that the best way to win this “war on terror” is to not need middle eastern oil any more. It’s not so much about the economics, it’s about victory. What is so damned hard to understand about that concept?

Concluding Remarks

Enough rambling. Please weigh-in by means of the various comment buttons. It is the only way I know that anyone even sees this stuff at all.

Faster-Than-Light Neutrinos? Maybe! Their Meaning? Arguable!

2011-10-09T11:05:00.000-07:00

Recent news reports from the world of science indicate one team working in a particle accelerator has clocked neutrinos traveling slightly faster than lightspeed. They are begging other teams to confirm or disprove this result independently, as is proper and normal in the business of science.

Commentators and experts have been weighing in on what such a result might mean, if confirmed to be true. “Everybody” points at Einstein (specifically his 1905 Special Theory of Relativity) to say that there is a speed limit these neutrinos seem to be violating. It’s either/or, not both.

Speed limit? That is an interpretation of Einstein’s theory, not a result. It is a very old interpretation, and I personally disagree with it. Here’s why:

In Einstein’s original 1905 paper, he sets up and solves the equations that describe the appearance of object A to an observer in reference frame B, moving at some relative but constant velocity V, when seen by light photons traveling at vacuum lightspeed c.

He did this for speeding subatomic particles, just like those neutrinos. Others since have extended the theory to large objects.

The theory postulates that c is a value which all observers measure as the same, no matter their motion, which really is something actually experimentally demonstrated with certainty. The theory’s mathematical results (sometimes called the Lorentz-Fitzgerald contraction equations) describe the object’s mass, dimension, and rate of time passage:

M = Mo/√(1-V^2/c^2) where M is what is seen and Mo is the resting value at V = 0

L = Lo*√(1-V^2/c^2) where L is what is seen and Lo is the resting value at V = 0

T = To*√(1-V^2/c^2) where T is what is seen and To is the resting value at V = 0

For these, M is the object’s mass, L its dimension in the direction of travel, and T is its local rate of time flow. Lateral dimensions are unaffected by V.

If one plugs a V greater than vacuum lightspeed c into these equations, the results are not real numbers, when a real-number result is what one seeks, being the only result that has meaning in the context of this problem. For almost a century now, the common interpretation has been that the not-real result means it is not possible to travel at speeds V exceeding vacuum lightspeed c. This is the origin of the common statement about “Einstein’s speed limit”.

When solving formulas in any branch of science and engineering, there are always fundamental assumptions about the problem, even if they are unspoken. Getting a not-real result with a modeling equation can nearly always be traced back to violating a fundamental assumption, even if it is not obvious.

Why should not-real results from Einstein’s theory be any different? What was his fundamental assumption that we violate when we plug in V greater than c in those equations?

Remember, the equations describe the appearance of object A to an observer in reference frame B. That presupposes that observer B can actually see object A (there it is!)

At V greater than c, we get a non-real result, which most likely simply means observer B cannot see object A, since we assumed he could. This interpretation is based on all those other experiences with formulas and problem-solving, and getting real versus not-real results.

When you think about it, how could observer B see object A traveling so fast? Object A is traveling faster than the photons with which observer B sees. The same sort of observational thing happens with supersonic aircraft: you cannot hear them coming because the sound waves by which you hear don’t arrive until much later.

Now, is the moving object A really heavier, shorter, and moving slower in time, or does he just look that way? How do you tell? You have to bring object A back to rest in observer B’s frame of reference.

Once the relative velocity V is back to zero, the equations say mass, length, and rates of time flow look completely normal and quite equal to both object A and observer B.

And furthermore, it does not matter who was really moving: V is relative only (that’s in part where the name of the theory came from).

Yet, somebody’s time flow rate really was slower. Their clocks (which are totaling devices, not rate of flow devices) will disagree. This is an experimentally validated and very certain result. Special Relativity does not resolve that problem, which is often called the Twin Paradox.

It is Einstein’s 1915 work on General Relativity that provides the answer: whoever did the accelerating to V and then back to zero is the one who experienced less total accumulated passage of time. Yet, his sense of time flow was entirely normal, to him, throughout the journey! This is the direct consequence of speed of light, not rate of time flow, being constant for all observers.

Could object A be observed if it were flying faster than light? To me, the equations say “no” with the not-real result; remember that they were derived on the assumption the object can be observed.

Could object A actually travel faster than light with respect to reference frame B? That’s a very good question! If the not-real result actually just means he cannot be seen, then that same not-real result says nothing about whether he can actually fly that fast!

So, that’s my maverick interpretation: Einstein says nothing about an actual speed limit. I seem to stand alone in this. But I always did like shouting from the corner that the Emperor has no clothes.

But, if I am right, we actually can travel faster than light, given sufficiently powerful technology. But, navigation will be hell if we see by photons, because the entire universe becomes unobservable!

We’re going to need some additional theory!

Update 12-21-11

The December 2, 2011 issue of “Science” (volume 334 issue 6060) has an interesting “News Focus” article on pages 1200-1201. This magazine is the peer-reviewed journal published by AAAS. The article title is “Where Does the Time Go?” and its topic line is “Superluminal Neutrinos”. There is now a lot of effort at a lot of places to either replicate the experiment or de-bug the procedures OPERA used to produce its faster-than-light results. Depending upon those outcomes, this result may never be explained. But this kind of activity is exactly what should be going on. The process of doing science really works.

The article describes the experimental concept as a simple timing across a fixed distance, although the elements of accomplishing that are not that simple. These are pulses of neutrino creation at one location correlated as pulses of neutrinos received at another location. Timing is by speed-of-light corrected GPS measurements, and by electrical transmission speed corrections in all the equipment. They are looking at how much the graphs of these pulses overlap.

One item questioned in the article is the GPS calibration, which really wasn’t done often enough. Countering the notion of miscalibrated GPS is the systematic time shift required to correct the calculated neutrino speed downward, when a more randomized error would be expected.

In the sixth paragraph (on page 1200) is an assumption the OPERA scientists made about the location where the neutrinos get created, of about a km delay. That assumption should be investigated. While the article says the error associated with it is “small”, it is a small effect we are arguing about (around 60 nanoseconds of time over a distance of about 700 km).

If it were a surveying error, the article says it would have to be on the order of 18 meters, which is not really credible. To me, this points right back to the creation-delay assumption.

It has been my engineering experience that 90+% of assumptions made are faulty or inappropriate. That why I suggest investigating assumptions first, or at least very high up on the priority list.

The other assumption that should be investigated is the interpretation of special relativity implying a speed limit, as discussed above. While not a popular topic in the physics community, it should still be done.

I would be delighted if the neutrinos were actually faster than light, and we needed some new theory. That would be a real breakthrough, and no telling where it might lead.

GW

Comic Opera Buffoons and Puppet Theater

2011-10-08T12:09:00.000-07:00

Take a few minutes to watch this cartoon, it is well worth it.

1948 Cartoon

http://nationaljuggernaut.blogspot.com/2009/09/this-cartoon-seemed-far-fetched-in-1948.html

I last saw this cartoon as the “movie day” movie in a public school classroom sometime in the late 1950’s or early 1960’s, but I don't remember which grade or whose class.

I will say this: the "ism" this cartoon warns against is exactly what is being preached today by BOTH our political parties, ON BOTH SIDES OF THE AISLE. The "issues" that divide them are nothing but puppet theater to distract us from the real and lethal problems that we face. Those are problems that require us to work together, but that cooperation does not contribute to re-election campaigning.

Holding public office has ceased to be a public service, and has instead become a high-dollar for-profit business, especially at the national level. It is also prevalent at the state level, and below in some places. As a result, we typically choose from among various comic-opera buffoon candidates, who are motivated only by selfish interest, instead of real statesmen who would do the people's business in preference to their own.

From the letters I see published in the newspaper, the ridiculous forwarded email "hit pieces", and most of the mainstream media opinion columns and broadcasts, it appears that a majority of the public has fallen for these lies. Furthermore, it appears most of the voting public actually believes the nonsense shouted by these buffoons. That scares the daylights out of me.

If I am right about this, and no one else out there wakes up and sees the real truth, then we are truly doomed.

GW

UARS: Why Another Uncontrolled Satellite Crash?

2011-09-26T12:07:00.000-07:00

Saturday, the 6-ton UARS satellite crashed. No one is sure exactly where: the odds say it crashed in the Pacific Ocean, but the uncertainty says some debris might have come down over the Pacific Northwest.

Official estimates were that about half a ton of this satellite would crash to Earth in 26 pieces, one as large as 300 pounds or so. My own estimate of surviving debris mass is larger.

I’ve been watching satellites and satellite re-entries all my life. In all those years, no person has been hurt, because most of these hit the sea (the only victims being fish).

But the risk is still there. So, why do we continue launching satellites that can crash uncontrolled? That's a good question!

Our 85-ton Skylab space station crashed onto Australia in 1979, leaving lighter debris on rooftops in a coastal town, with the heavier pieces carrying further inland. About 75 tons were eventually recovered!

Before that crash, the “official wisdom” was that it would mostly “burn up”. It clearly didn’t do that. Had its debris field been more centered on that Australian town, it is likely someone would have been hurt or killed.

That same year, the 10-ton Pegasus 2 satellite also came crashing back to Earth. This one hit the ocean, as do most.

In 1978, the Russian satellite Cosmos 954 crashed to Earth in north-central Canada. This one was the special case of a nuclear reactor-powered satellite, for which the preferred disposal method (a really high, decay-proof orbit) failed. There was serious radioactive contamination over an area of back-country Canada hundreds of miles wide.

And, most folks remember vividly the crash of Space Shuttle Columbia in Texas in 2003. Again, there were a lot of pieces, some quite large, on the ground from Dallas to Tyler. No one on the ground was hurt, but they very easily could have been.

In contrast, the Russians deliberately crashed their Mir space station into the Pacific in 2001. They used a rocket motor to de-orbit the station and put it down exactly where they wanted: away from land and people.

Excepting disasters and reactor disposals, most satellites could (and should) be equipped with a small rocket motor for a controlled crash in a safe place. Small solid rocket motors are cheap, light, compact, widely available, and they last for decades without any maintenance, waiting to be used.

So why don’t we do this, especially considering the nail-biting experience with Skylab?

Simply because no rule says we have to. That’s something very easy to fix, and without a new law.

For civilian / commercial satellites, the Federal Aviation Administration (FAA) should simply require controlled de-orbit provisions. It’s their jurisdiction and they already have rule-making authority. A word from the President to do it, is all it would take.

For military satellites, a simple order from the Commander-in-Chief is all that is needed.

Mr. President, fix this. Give the word to the FAA and the Joint Chiefs. I bet most of the other satellite-launching nations would soon follow suit.

Air Races, Air Shows, and Risks

2011-09-23T11:22:00.000-07:00

The recent fatal crash of the modified World War 2 P-51 “Galloping Ghost” at the Reno air races is a horrible incident. Lots of things have been said on the news and on the internet about it, but all of this is speculation based on incomplete information.

The National Transportation Safety Board (NTSB) will investigate this and determine its cause. They will use all the available information, including stuff not yet reported or on the internet. Until they publish their findings, perhaps a year from now, all else is mere speculation.

That being said, some speculators are more informed than others. I would place more trust in the speculations of an actual aircraft engineer over the speculations of most other members of the public. Being such an engineer, here are my speculations:

Facts

“Galloping Ghost” suddenly pitched up and climbed, rolled over inverted, and dove to the ground, all in a matter of scant seconds. Impact was not directly upon, but was immediately adjacent to, spectators, many of whom were killed by pieces thrown from the wreck. There was no post crash fire.

The left trim tab was photographed departing from the airplane’s horizontal tail before impact. The pilot’s head was not visible in the canopy before impact. The retractable tail wheel was seen extended before impact.

This aircraft was modified in several ways from its World War 2 configuration to compete in the races. Most notable were “clipped wings” reducing span (and aileron size) by 5 feet each side, and removal of the belly air scoop and radiator in favor of a sacrificial coolant system. These enable higher top speed, at the cost of higher landing speed and perhaps a reduced maximum roll rate, not a loss of basic stability.

Less obvious were changes to canopy size, wing fillet size, and smoothing of protuberances, for drag reduction. The race speeds significantly exceed 500 mph, when the original level-flight top speed for the P-51 during World War 2 was 435 mph.

I do not know what the original “never exceed” speed was for the P-51, but these race speeds would be approaching or exceeding that limit. Flying too fast risks structural failures of wing and tail components by a phenomenon called flutter.

Similar Previous Incident

About a decade ago, a similarly-modified P-51 named “Voodoo-5” experienced a very similar incident: sudden high speed pitch-up (at high acceleration) into a climb, with the pilot losing consciousness briefly due to excessive gee forces. He woke up, with no memory of events, at 9000 feet altitude, and regained control, landing successfully.

“Voodo-5” was found to have lost the same left trim tab as “Galloping Ghost”. Loss of the tab at high speed caused failure of part of the elevator control linkage, leaving only the right elevator for pitch control. Aerodynamic flutter was blamed for loss of the trim tab.

At 400+ mph, P-51's exhibit a relatively unusual nose-up tendency that you fight with down trim and down stick. Other aircraft exhibit high-speed nose-down "tuck", or no trim-change tendencies at all.

In the P-51 with that nose-up tendency, sudden loss of half your elevator effectiveness at very high speed causes the aircraft to suddenly and violently pitch up, at something near 10-15 gees. The pilot passes out, or can even be killed with a broken neck, depending on helmet weight and head restraints, or the lack thereof.

Speculations Regarding “Galloping Ghost”

“Galloping Ghost” lost the same left trim tab, and pitched up similarly at high gee. Some on the internet say telemetry from the aircraft indicated 11.5 gees. The differences between “Voodoo-5” and “Galloping Ghost” are (1) “Galloping Ghost” also experienced a roll, and (2) it appears her pilot never woke up, or was perhaps already dying of a broken neck.

It also appears that the high pitch-up gee level forced deployment of her tail wheel. The roll motion on the way up caused her to peak in inverted flight, as photographed. I think I see a light-colored helmet on the dark dashboard in that internet inverted-flight photo, but I could be wrong. She then continued her pitch-roll motion into a dive-to-impact.

There has been speculation that the pilot’s seat failed in “Galloping Ghost”, which might explain why his head was not visible in the canopy. I would be surprised at seat failure in a fighter plane at only 10-15 gees, but I guess it could happen. It did not in “Voodoo-5”, though.

Waiting for the Truth

The NTSB will opine officially maybe a year from now, but I'd almost bet they say that P-51 trim tabs are vulnerable to flutter-induced departure at race speeds beyond the original design's never-exceed speed. Few designers provide aerodynamic or mass balancing, or any other anti-flutter structural treatment, to a trim tab. Maybe they should. If so, the NTSB will say so.

Inappropriate Fear-Mongering

The public safety issue raised by some reporters has less to do with any given aircraft being "pushed too far", or being modified "too radically", and more to do with simple spectator crowd placement. The wording in those reports seems deliberately chosen to inflame fears, and is a disservice to the public, much like yelling “fire” in a crowded theater when there isn’t one.

At air shows, spectators may not legally be located beneath expected aircraft flight paths. At the Reno air races, they can be (and are) located under flight paths. Perhaps they should not be, similar to the air show restrictions. While these are the first spectator deaths at Reno since the 1950's, that risk has always been there.

There was a fatal crash at an air show the day following the “Galloping Ghost” incident. No one but the pilot was killed, because no one was underneath the falling plane.

Mars Mission Second Thoughts Illustrated

2011-09-06T21:51:00.000-07:00

As I said in a previous posting (8-9-11), I had some second thoughts about the back-up propulsion for my fast trip Mars mission paper, presented at the Mars Society convention in Dallas, Texas, August 4-7, 2011. My backup had been the VASIMR electric propulsion scheme, thinking it a breakthrough in thrust for the power consumption. Based on what I saw at the meeting, it is no breakthrough, and is really mostly unsuitable for fast trips to Mars.

My second thoughts centered around an alternative slow-traveling vehicle requiring artificial gravity, because the manned mission duration would exceed the 1 year known to be tolerable. This vehicle would be powered by the same solid-core nuclear thermal technology I assumed in my landers, derived from the NERVA tested successfully 4 decades ago. I planned this alternative around simple minimum-energy Hohmann transfer orbits, because it is easy.

That still leaves the gas-core nuclear thermal-powered fast trip vehicle, which is still my baseline. I took a closer look at the orbits and the near-straight line “shots” across the solar system at the higher travel speeds. This verified my earlier crude ballpark estimate of the fast trip velocity requirements. All of this is illustrated here, at a level of analysis no deeper than is required to confirm the concepts and their feasibility. For example, I used circles to approximate the actually slightly-elliptical orbits of the planets. To first order for a feasibility check, this is “good enough”.

Baseline Trajectories

The baseline “fast trip mission” sends a fleet of three unmanned ships to Mars parking orbit ahead of the manned ship. This fleet is propelled by the landers themselves, and comprises all the propellant required to support the landing operations, plus enough to send these assets one-way to Mars by Hohmann transfer. Figure 1 shows the initial Hohmann transfer for these unmanned assets. Note that there is an opposition during the unmanned flight to Mars.

I looked for ways to center my manned fast trip about that first opposition, without adding too much extra time in orbit to the manned mission. This did not prove feasible, so that the manned fast trip is centered about a second opposition some 779 days after the first one. The mission calls for 16 weeks at Mars making landings, which would be 56 days to either side of the opposition. A little time spent making rough calculations gave me an “optimal” one-way flight time pretty near 83 days for the “average” mission these approximations represent. This is shown in Figure 2.

Baseline Vehicles

The total mission time (for the men) is under 9 months, so no artificial gravity is required. Note that the total time the propellant tanks sent unmanned must maintain the liquid hydrogen is well over two years – rather challenging! The vehicle designs are as shown in Figure 3, and are essentially unchanged from my paper. The direct launch costs are pretty much as I estimated in the original paper.

Guessing that total program costs are about 6 times the direct launch costs gives something like $50 billion to mount this mission, given the right team. Those figures are similar to the ones in the original paper. That “right team” issue is also discussed in more detail in that paper. See the 7-25-11 posting for an on-line version of that original paper.

Backup “Slowboat” Trajectories

If the manned vehicle is comprised of the same basic modules, but with solid core nuclear engines instead of the gas core engines, then single stage two-way flight, even on a Hohmann transfer, is not possible. But, a single stage transfer to Mars can be flown, and the empty tanks left there at Mars. In this way, a single stage return to Earth can be made, without relying on propellant already sent unmanned to Mars. This is a safety issue: what if rendezvous should fail in Mars orbit? The crew needs a way to return anyway.

The Hohmann transfer to Mars is identical to that in Figure 1. All four ships travel together as a single fleet: 3 unmanned and the one manned vessel. It is not possible to return by Hohmann transfer until the second opposition approaches, as illustrated in Figure 4. These oppositions are separated by 779 days, which leads to the timelines shown for the return in Figure 4. Thus, total manned mission duration is about 2.66 years, requiring the use of artificial gravity to protect the health of the crew, and considerably more packed supplies for the longer mission. Time at Mars about doubles, allowing for 16 2-week landings instead of 16 1-week landings, as in the baseline.

Backup “Slowboat” Vehicles

The unmanned vehicles are unchanged from my paper. The manned vehicle is necessarily bigger than the baseline design, driven by the substantially lower performance of solid core nuclear thermal rockets (SC-NTR) vs. gas core nuclear thermal rockets (GC-NTR). The solid core vehicle is substantially longer and about twice as heavy as the baseline gas core vehicle. The “payload” is larger, too, driven by the need to pack about 3 times as much supplies, with some of that bulky, heavy frozen food. These are depicted in Figure 5.

The return vehicles, command module (also the radiation shelter), habitat module, and supply storage modules are the same, I just needed 3 storage modules instead of one. I did take a closer look at the habitat module, since more space is needed for the longer mission to maintain psychological health. The easiest way to do that was to make the habitat an inflatable, along the lines of the Bigelow Aerospace modules already in experimental flight test now. Equipment and floor structure would be stowed along the axis for launch, and folded out into position once the module is inflated, as illustrated in Figure 6.

The same module could be used on the baseline fast trip vehicle, there is no need to build two different designs. It is imperative not to mount equipment on module walls, as they need to be accessible for very rapid meteoroid puncture repairs. (The same is true of non-inflatable modules.)

I wrestled with several ideas on how to provide adequate radius at acceptable spin rates for artificial gravity, at the one gee level which we already know would be adequate. The breakthrough was to spin the long ship end-over-end, using the long module stack as its own spin diameter. For the trip to Mars, the propellant stack is 34 modules long, each figured as 5.2 m diameter and 13.9 m long, based on the payload shroud dimensions for the SpaceX Falcon-heavy launch vehicle. Spinning end-over-end at only 1.2 rpm provides right about 1 gee at the forward end of the inflatable habitat (at its lower deck as illustrated). The stack is shorter returning to Earth, but should be long enough to provide close to 1 gee at no more than the acceptable limit of 4 rpm.

About Radiation

The original paper covers solar flare radiation shielding in the command module. This is done by surrounding the flight deck with water and wastewater tanks, plus perhaps a little steel plate. One provides space in there for all 6 crew, and a day or two of supplies to outlast the typical solar storm. This enables critical maneuvers to be flown, no matter what the solar weather, a major flight safety issue.

A little research since then provided credible dose estimates for the cosmic ray background radiation, composed of particles so energetic that ordinary shielding is more-or-less impractical. The dose varies between 22 and 60 REM per year in a steady “drizzle”, depending upon the strength of the solar wind, which tends to deflect some of it. The original radiation dose limits for astronauts was set to 25 REM/year, which was the World War 2-vintage max dose for civilian adults. It has since been revised to 50 REM/year, based on what I can find on the internet. The actual dosage rate only sometimes exceeds the newer limits, and then only by a small amount. Trips to Mars thus appear quite feasible without incurring any immediate health risks from cosmic rays, or even any significant prospect of long-term effects.

The Program As Revised

Changing to SC-NTR backup propulsion puts the artificial gravity and frozen food storage issues into the design mix. This has to be made to work, and they are things we have never before done. Using the baseline GC-NTR propulsion puts that very propulsion into the mix as something we never did do before, excepting some feasibility experiments. Those are the two development items to be worked in parallel, so that one or the other is ready in time to fly. (This is the same basic parallel path development idea that was in the original paper, where the baseline was GC-NTR, and the backup was VASIMR and its power plant.) All the other items are simple design / build / checkout efforts based on known technologies, and that includes the SC-NTR. The high-level program plan is just a bunch of parallel paths, as illustrated in Figure 7.

So, we are looking at somewhere in the vicinity of $53 B to $70 B to send 6 men to Mars to make 16 widely-separated landings all over the planet, in the one trip, with maximum safety and self-rescue capability designed-in at every step, and with all-reusable assets left in space to be refueled and reused by subsequent missions. The whole thing could be done for prices like that, in only 5-10 years, given the right kind of contractor teams, and the right kind of an agency to lead them. That is one incredible amount of “bang for the buck”!

As I said in my original paper, right now we do not have that agency, and only a couple of the right kind of contractors, at best. But, if we fix those lacks, we could really do this. The numbers show it is definitely feasible.

The last time we as a nation embarked on a mission to explore another world (the moon), we had nearly two decades of sustained economic boom, from all the jobs created just to get the mission done. That may not be causal, but it is definitely correlated. Why not do it again?

Figure 1 – “Slowboat” Transfer to Mars, Baseline and Backup

Figure 2 – “Fast Trip” Transfer To and From Mars, Baseline

Figure 3 – Baseline Manned and Unmanned Vehicles

Figure 4 – “Slowboat” Transfers to Earth, Backup

Figure 5 – Backup Manned and Unmanned Vehicles

Figure 6 – Inflatable Habitat Module, Baseline and Backup Vehicles

Figure 7 – Program Outline Plan

Surprise, Surprise: Oil Boom in the Williston Basin (“the Bakken”)

2011-09-05T14:03:00.000-07:00

Resources on the internet about this formation have been revised recently. There appears to be an oil drilling boom going on in eastern Montana and western North Dakota. They are horizontal-drilling and hydro-fracturing for light crude (meaning low viscosity liquid). One of the descriptions says the crude they can recover seems to be just about the same gross physical properties as diesel (density, viscosity).

That's a surprise to me. Two years ago I researched this formation as a "shale unit, very low porosity and microscopic permeability", and everything I read about the hydrocarbons in it said a consistency more like tar. Hydro-fracturing simply would not work on a near-solid resource like that. It would have to be mined, like coal.

What I read now says the Bakken comprises a dolomite layer around 100-140 feet thick, bounded above and below by shale layers. Typically, the shale is the “original” source for the hydrocarbons. The dolomite is listed as 5% porosity and microscopic permeability (1-10 microdarcy's, just almost impermeable). It is in the dolomite layer (not the shale) that they are horizontal-drilling and hydro-fracturing. Estimates vary about how much of the total resource they might possibly recover this way, by over an order of magnitude, depending upon who made the estimate and what agenda they have.

For the Burgess Shale natural gas hydro-fracturing here in Texas, the estimate is that about 3% of the gas down there is actually recoverable. For the liquid in the Bakken dolomite layer, I'd simply guess that factor as 3% or less, which is nearer the 1% end of the estimate range of 1% to 50% that I saw on-line yesterday. Almost-nil permeability just has that effect, hydro-fracturing notwithstanding.

I suspect that there are residual tars left behind in both of the shale units in the Bakken formation, and that the source for the light fractions in the sandwiched dolomite layer is the lower shale member. Somehow, I don't see light fractions migrating downward from the upper shale member, so its lighter fractions are most likely now lost to us.

So, how much recoverable light oil might there be, and how much good might it do, if we can recover around 2% of it?

Oil in the Dolomite Layer:

If you guess that there's something like 500 x 500 statute miles of this formation, averaging 100 ft thick, at 5% porosity, then there might be as many as 6 trillion barrels of light oil down there.

500 mile dimension x 5280 ft per mile = 2.64E6 ft. 500 mi x 500 mi is then 6.97E12 sq ft. Multiply by 100 ft thick to obtain 6.97E14 cu.ft of dolomite rock. The hydrocarbon volume equals the pore space volume at 5% of rock volume, assuming the pores are 100% full. That's 3.48E13 cu.ft of hydrocarbons. Cu.ft volume of hydrocarbons x 7.48 gal per cu.ft is 2.61E14 gal hydrocarbons; divide that by 42 gal/barrel. That's 6.2E12 (about 6 trillion) barrels of hydrocarbon volume down there in the pores of the dolomite layer, supposedly all hydro-fracturable, very light crude.

Assume we can recover 2% of it. That's about 1.24E11 barrels of light oil that could be recovered, or about 124 billion barrels in ordinary terms. That's quite significant. I could be off by a factor of 2-3 in rock volume assumptions, more likely toward the smaller than the larger, so these figures are rather optimistic.

At our 7-8 billion barrels / year consumption in the USA, then potentially, this could power us for about 16-17 years. That really is significant, even if it is optimistic by a factor of 2-3. If it is all light oil. If we really can recover 2% of it. If the rock pores are really full. Lots of "ifs".

Let's say this oil boom lasts 20-30 years (typical for a very large field). The average production rate from the mature field (which takes several years to achieve) might be as much as around 4-6 billion barrels a year, again possibly optimistic by factor of 2-3. That's still a lot, optimistic or otherwise.

Replacing Foreign Imports:

About 1/3 of our consumption is domestic production, about 1/3 comes from Mexico and Canada, and about 1/3 comes from OPEC (which includes Venezuela, along with that idiot running it; and our “friend” Iran, with that insane group of religious fanatics running it). That's about 2.5 billion barrels per year from each source. We might very well be able to replace much of the OPEC oil with domestic from the Bakken dolomite layer, even as the other sources decline. For a little while.

But, no matter how politically expedient, it is still clearly not at all wise to count on it “ending” our dependence on foreign oil. Although, you can bet more than one GOP/Tea Party candidate will run on "why not save ourselves from oil dependence with the Bakken, if the environmentalists and Democrats would just get out of the way?" They did exactly that in '08: remember “drill, baby, drill?”

Even with the new oil boom that I did not expect to see, it’s still a comic-opera puppet-theater issue intended to distract the public from the real truths that threaten us. It’s still just a fake electioneering issue for a bunch of comic-opera buffoon candidates. Beware! I warned you!

About the Tar Shale Layers:

I saw no thickness figures on the two shale units, in the new data that I found this year. I bet they're quite thick, though. You'd have to deep strip mine it, and what I saw said it averages 2 miles down. Figure shale at 0.5% or less porosity, for maybe another handful of trillions of barrels of potentially-recoverable hydrocarbon. This tar shale stuff would be very hard to extract and process, though, and so it would be a supremely expensive product.

And, we would get it for the environmental cost of a permanent crater some 500x500x2 miles in size, which is bigger by far than the volume of Lake Superior. That shale tar is what I was thinking about when I posted what I did about "the Bakken" last year (the 3-14-10 article). That’s still true, oil boom notwithstanding.

Conclusions:

Yep, we need to go get the hydro-fracturable light oil.

Yep, it’ll surely help with imports.

Nope, it will not “save” us.

There is no permanent answer among depletable (fossil) fuels, and never will be.

Balanced News

2011-09-02T16:10:00.000-07:00

TRY CHECKING YOUR PERSONAL IDEOLOGICAL BELIEFS AGAINST THE DATA

Left-wing, right-wing, doesn’t matter. Here are the actual data:

Spending

So who increased spending the most?

(No doubt it’s too much.)

(War spending does not explain all of this.)

Deficits

So which is the better trend?

(And this is before any of the recently-negotiated cuts “kick in”.)

The Stimulus and Jobs

So did the stimulus help job growth or not? (And remember, each man spent about half of the total stimulus)

Did we stop too soon?

Could we have better-targeted the money to steepen the job growth trend?

A Very Good Question:

If your beliefs about what has been happening do not square with the actual data, then what are you going to do?

Change your thinking and use the data? Or keep your beliefs?

Post-Meeting Results (Mars Mission)

2011-08-09T09:18:00.000-07:00

The Mars Society convention was a lot of fun and very informative. My paper was presented in the advanced technology track on Saturday afternoon. It was well received: I had several people come up to me afterwards and tell me so.

Most of the other mission designs included some sort of Mars base assets with the first manned mission. This is quite different from what I assumed regarding the nature of exploration versus subsequent activities. So, I pitched my paper as less of a “real” mission plan, and more of a “mine” for different and potentially useful ideas. In part, that’s why it was so well received.

Changes to My Paper (see the 7-25-11 posting not far below)

I did learn some very interesting things, two in particular. One has to do with my alternate for the manned ship’s “hot rod propulsion”. It seems VASIMR is not really an improvement on electric propulsion, just another way of doing it. Its weakness is indeed what I thought: the mass of the nuclear electric power plant required. Its thrust per unit power supplied is just a lot worse than I thought it was when I did the calculations. It’s just not suitable for really fast missions.

The other interesting thing is the notion of a light gas gun for launching hardened payloads into orbit very inexpensively. It should be possible to launch large quantities of propellants and tough hardware for something on the order of $300/pound, if they can be hardened to withstand 3200 gees. This is based on a smaller gun already launching small experimental scramjet payloads for the Air Force at Mach 9. Refueling of my reusable manned ship looks really good in such a situation. Once there is a water mine and propellant station on Mars, the same thing is true for refueling the lander assets left in Mars orbit.

In any event, about the only change I might make in my paper is to replace the VASIMR alternate with a solid core nuclear thermal version, and include artificial gravity and frozen food in the habitat configuration. Its one-way trip time would be 6 to 8 months, and the stay at Mars a little longer than the baseline 16 weeks. The technology development, to be run in parallel with the baseline gas core nuclear thermal rocket effort, would be the artificial gravity habitat. I think a pair of rigid arms out to inflatable living spaces, and spinning the entire T-shaped ship, might work well enough.

Another Interesting Idea

A third very interesting idea is to store and ship hydrogen as frozen water. In this form, it is very strong and so is proof against accidents or mishap. You thaw and electrolyze what you need as you go, which does require power, although solar thermal thawing offers a big help. The oxygen liberated by electrolysis can be used for a lot of things. It takes very little pressure to prevent sublimation of the ice, and a simple sunshade keeps it very cold.

This would apply to chemical as well as nuclear propulsion. Nuclear uses only hydrogen. Chemical uses both hydrogen and oxygen at a mass ratio of 1:6. The ratio in the water is 1:8, so that leaves excess oxygen left over for other uses, even with chemical systems. There is just more available in a nuclear scenario.

A Very Serious Near-Term Problem

Consider: 100% of the humans who ever walked on the moon were Americans, sent there by NASA in its human spaceflight program. 100% of the so-far successful landers on Mars were (and are) American, sent there by NASA in its robotic exploration program. Almost 100% of the probes sent to other celestial bodies are American, sent there by NASA in its robotic exploration program.

The human and robotic programs began together in the late 1950’s; they are synergistic. You cannot successfully do one without the other. Regardless of your opinion of NASA and its effectiveness today, it is the premier entity for the exploration of space, and therefore it is irreplaceable.

Here is the problem: there has been no human exploration target since Apollo ended in 1973. Manned operations in Earth orbit, while essential and even inspiring, are not exploration. We have had men and machines in Earth orbit, beginning with Sputnik in 1957. Going back to the moon is not exploration in the public’s eyes, because “we’ve already been there”. This perception is quite real, even though we didn’t really explore the moon (in the sense of my paper) during Apollo.

The public supports exploration: that is why the probes, the Mars landers, and the Hubble pictures are so popular. Of all the probes and landers, it is the Mars probes that hold the public’s fascination best. This is because Mars has fascinated people for centuries. It is not just the best target for human exploration, it is the only one. Those other near-Earth targets are at best but steps along the way to Mars. (The next destination after Mars is the stars, with the outer solar system destinations but steps along the way.) Reality has nothing to do with perception, and experience says you cannot fight perception.

We have a budgetary and political tsunami about to sweep America, with a great likelihood of doing massive damage to all aspects of all of our lives. One’s politics and outlook on this do not matter, discretionary spending is about to be drastically cut or eliminated, no matter how useful or necessary, for the sake of election politics. That means NASA, among many other things. And NASA has had no viable target or plan for manned exploration since Apollo. A vague “give us X-billion dollars for the next 20-40 years and we might reach Mars” is not a manned exploration program. If the manned spaceflight program is cancelled, the robotic program will eventually fall, as well.

Folks, this cannot be allowed to happen.

It is feasible to send men to Mars right now, with the technologies and hardware we have right now or within the next very few years. We don’t even have to have giant launch rockets. We can do this for under $50 billion, not the trillions everybody out there seems to think it will take. But we cannot do this with the “business as usual” techniques of the last 4 decades, and that includes the way NASA works. Massive management change is required, and that is the hardest part, not the actual flying to Mars.

If you are a space exploration enthusiast, then help get the word out. Technologically we are ready to send men to Mars. And we can do it for a few billions, not multiple trillions, of dollars. The real change required is managerial (and political, not surprisingly).

What Should the Government’s Manned Space Exploration Strategy Be?

2011-08-02T11:30:00.000-07:00

Going beyond the moon requires fundamental changes in the way we approach spacecraft and mission design. Mars makes an excellent target for starting this new process.

Inner Solar System

The key is going back to fundamentals to figure what we really want out of exploration, then looking at required technologies, given that crew survival, safety, and self rescue are THE paramount design requirement for every single phase. You do that for the most challenging mission (Mars) FIRST, and force every single piece of hardware to be totally reusable.

Exploration and the Greater Scheme of Things

This produces one set of "tinkertoys" that takes you anywhere within range: Mars, Venus, NEO's, and Mercury (the entire inner solar system). Plus, you don't have to keep launching components, just propellants and supplies.

This is my Mars mission paper to be presented to the Mars Society convention in Dallas, August 4-7. A shorter version is posted at http://exrocketman.blogspot.com, dated 25 July, 2011.

Inner Solar System Hardware Designed Around a Mars Mission

Main Asteroid Belt

To go further soonest, we simply upgrade the inner solar system “tinkertoys” with knowledge obtained between now and then. Add a provision for artificial gravity, and either solve the food preservation problem, or add frozen food (which is bigger and heavier). That should make two-or three year trips feasible, most likely limited by the accumulation of cosmic ray exposure. It puts the main belt asteroids and some comets within reach.

The Artificial Gravity Problem As We Know It Now

Giant Planets and Outer Solar System

If a better way to shield against radiation can be devised, and even faster “hot rod propulsion” developed, the giant planets and outer solar system become reachable. It would help greatly to know how to build closed-ecology life support by this time.

At this point, we have to repeat the design process from scratch, because upgrades to the inner solar system “tinkertoys” are no longer feasible. We will know more about what to do when that time comes. The target for design should be the Kuiper Belt, where Pluto is.

I suggest some version of the old Project Orion nuclear pulse propulsion. One must build much larger ships for this: it has the odd characteristic of working more efficiently, the larger the mass to be moved.

Nuclear Pulse Propulsion from 1959-1965 USAF "Project Orion"

A Place To Do the Supporting Work

The kinds of “hot rod propulsion” that we will need to explore these places are very dangerous to develop and test on Earth, because the energy sources for the drives are all nuclear. We need a safe place to test, but it has to be dynamically stable (you cannot test a rocket engine in zero gravity where every test is a flight test).

I suggest the airless, waterless, uninhabited moon. Pick a smaller crater with high ring walls, plant a base adjacent to it, and put the test stands down inside the crater. It’s reachable from Earth without any exotic propulsion at all, and the mildest of the “hot rod propulsion” techniques just makes it less expensive. Perfect!

End of an Era Need Not Be End of a Capability

2011-08-02T11:06:00.000-07:00

This one appeared as a guest column in the Waco Tribune-Herald newspaper of Waco, Texas. It is not about exploration, but about useful and necessary infrastructure in Earth orbit.

The last shuttle flight is complete, the orbiters are headed for museums, and thousands are being laid off at Cape Canaveral and Mission Control. There is going to be a hiatus in America’s ability to launch its own astronauts that may continue for a few years.

Actually, I do believe the private companies like SpaceX will fill that void sooner rather than later. I also believe there will be a second, privately-owned, space station up there, very soon.

But, these private ventures, which build upon 50 years’ expertise, will require smaller teams on the ground at the Cape and in Mission Control. Except for local employment prospects, that’s a good thing, because it means spaceflight will become less expensive.

I predict that more than one commercial spacecraft will be flying soon, and that some of our familiar launch rockets will be updated and man-rated to launch them. That’s what NASA’s commercial spaceflight initiative was supposed to achieve, and it looks to me as if it is succeeding.

Here is what we will have: space capsules as taxis to orbit, space stations conducting research and business, and NASA looking outward beyond orbit with men for the first time in 40 years. But, something is missing from that picture!

Oh, yes, the “space repair truck” function of the space shuttle will be missing. Remember it?

It was the self-maneuvering shuttle with the adaptable human crew, that enormous cargo bay as a work area, and that manipulator arm, which repaired so many important satellites, most notably the Hubble Space Telescope.

In hindsight, having to launch that capability in every mission makes less sense than having two or three vehicles like that up there all the time. When you need one, send the crew and some thruster fuel up with one of the new space capsules: same results, far less expense.

All we need is a crew cabin, a space frame about the size of the shuttle cargo bay, that manipulator arm, some thrusters, and thruster fuel tanks.

These could be assembled in place by docking-together modules small enough to be launched by the rockets we already have. This is not a gigantic project, there is no new technology here, just a planned series of launches to regain a capability that we lost with Atlantis’s final landing.

It makes sense to have one in the highly-inclined orbit near or docked to the ISS, one in the more standard orbit eastward from the cape (the kind of orbit Hubble is in), and one in polar orbit to service those satellites. This may not be exploration, but it would certainly be helpful to what we are already doing in space.

There will always be satellites needing repair, and we already have one space station to repair, maintain, and upgrade. We might even use this capability to help build the new exploration vehicles.

I recommend this idea to NASA as something worthwhile and necessary. Maybe some of those laid-off workers could be re-hired to carry it out.

ABOUT ARTIFICIAL GRAVITY FOR LONG SPACE MISSIONS

2011-07-31T12:19:00.000-07:00

This article discusses artificial gravity by centrifugal force for spaceflight. It supports the manned Mars mission article posted 7-25-11. This article demonstrates why “just spin the vehicle” is a naïve statement.

The concept of using a spinning design to provide a substitute for gravity (“artificial gravity”) is a very old concept, going back many decades in science fiction, and to the 1930’s and 1940’s in real scientific proposals. The late 1930’s concept of a spinning wheel space station (from the von Braun group in Germany) is an example of an early, but very real, scientific proposal.

To provide artificial gravity by spin requires a radius for the spin, a spin rate, and a design value for how much gravity (“gee”) is to be produced. These variables are related by the familiar equation from circular motion:

a = R ω^2

or:

n = R ω^2 / gc

where a is the acceleration level, R is the spin radius, ω is the angular velocity (rate of spin), n is the acceleration expressed non-dimensionally in gees, and gc is the standard acceleration of gravity. Consistent units of measurement must be used, of course. If metric, R is meters, ω is radians per second, a and gc are both meters/second2.

To find R, one must know the desired a (or n), and must have a value of ω. In other words, we need to know the answers to two simple questions: “how much gee is enough?” and “how much spin rate is too much?”

So far, we only have real human experience at one gee (here on Earth), and zero gees (out in space). There are surrogate studies, based on bed rest, but these are only surrogates, not exact models. These are no more trustworthy than dosages of drugs based on doses in mice, without human trials.

So, until real fractional gee studies are done, the design value must be one full gee. There is no other ethical, moral choice.

Tolerance of spin rate is something dependent on exposure time. I am unaware of any real long-term experimental results for this, but anecdotes based on real experiences would suggest that 4 rpm is the upper limit, perhaps plus or minus 1 rpm. People do get seriously motion sick when exposed to faster rates for long periods of time.

4 rpm = 4 rev/min = 4 rev * (1min/60 sec) * (2 π rad/rev) = 0.4189 rad/sec

There is one other issue to worry about, especially for faster spin rates. That is the gradient of gees head to toe for someone standing in an artificial gravity device. If that gradient is too large, there could be blood pooling in the legs, leading to possible fainting. We have no real criteria for this, the real human experiments never having been done. That gradient is the first derivative of the circular motion equation with respect to R:

da/dR = ω^2

or:

dn/dR = ω^2 / gc

Using the equation from circular motion in its second form above, plus the gradient equation, and using 4 rpm for the spin rate (converted to radians per second) leads directly to figure 1. Reading these curves at one gee required leads to R = 56 meters, and a very small gradient (0.0179 gee/m).

Figure 1 – Artificial Gravity Data for Spin at 4 RPM

What we learn from this is that, under these requirements, the gradient issue is of no real concern, while the size of the artificial gravity device is quite large. Size, and the weight, complexity, and expense that go with it, are extremely serious design issues for spaceflight.

There are essentially three ways to handle this dilemma: (1) cable-connected modules spinning about their common center of gravity, (2) truss-connected modules spinning about the center of gravity of the structure, and (3) simply making the entire ship (or space station) very large, and spinning it about its principal axis. These choices are shown in figure 2.

Figure 2 – Ways and Means to Provide 1-Gee Artificial Gravity at 4 RPM

The cable-connected module option is perhaps the simplest and lightest practical solution, except that maneuvers are impossible in the spinning configuration. This rules out mid-course corrections and collision-avoidance maneuvers. One must stop the spin and re-dock the modules to maneuver.

The truss-connected module option allows limited maneuvering of the spinning structure, as indicated, at the center of gravity. Precessional effects must be taken into account. Even so, how then does one make a major “burn” without disassembling the structure? That question has no answer at this time.

Building a ship (or space station) with a 56 meter radius is a major construction undertaking. This cannot be, and can never be, easy or inexpensive under any imaginable circumstances. It does allow maneuver easily, by simply arresting the spin. This would be appropriate for very large objects, such as colonization ships or space stations, but it is difficult to imagine how to do it with small docked modules, for something smaller overall.

We do know from space stations in Earth orbit, that exposure to zero-gee does damage to many body systems, the most well-known being bone decalcification. Unfortunately, the heart and circulatory systems also suffer. There are serious concerns about effects on the immune system as well. However, our experiences also indicate that these effects can be held to acceptable, recoverable levels by appropriate vigorous exercise, for at most about a year’s exposure.

What that says is that for journeys under a year duration, we may fly without artificial gravity, given the space and equipment to do the appropriate vigorous exercise daily. Beyond that time, we must provide artificial gravity.

We cannot count on the surface gravity of places like Mars (0.38 gee) to counter the effects of weightlessness enough to be therapeutic, simply because we have never done the experiments to find out “how much gee is enough?” That means the 1 year limit applies to the round trip, not just the one-way voyage there.

The classic min-energy Hohmann trajectory to Mars is 8.5 months one-way. We can do somewhat better than that without expending a great deal of extra energy: 6 months. But, even with a 6-month one-way trajectory to Mars, the total round trip is over a year, even with a trivial stay at the planet. Most such mission designs require a considerable stay at Mars, from months to nearly 2 years, because of orbital mechanics constraints.

CONCLUSIONS:

So, until the appropriate experiments have been done, our ethics-bound criteria must be:

Provide one full gee at no more than 4 rpm, for any round trip over 1 year

Fly without gravity for up to a one year round trip, given appropriate exercise

This poses a very severe design dilemma for manned Mars missions: either fly the men very much faster than 6 months one way, or else provide one full gee artificial gravity at enormous construction difficulty and expense.

For exploration, which otherwise requires much smaller vehicles than colonization, I favor flying faster.

So, as always, propulsion is the key. In my opinion, the solutions to faster propulsion will likely be nuclear. It’s past time to face up to that fact, and simply get on with it.

Recovery - final

2011-07-26T23:20:00.000-07:00

I am doing very well now, and should be discharged from therapy soon. I went back to work (albeit limited duties and limited time) Monday (yesterday), day 42 since surgery. Still going to therapy 3 times a week. I can sit feet-down about 6 hours now without causing too much pain, and I can walk a few thousand yards. Driving is no longer much of a problem, to at least 60 miles. There's one more week to push it further and be ready to present my paper at that meeting. I think it will be fine.

We have made great progress with the little kitten, Kit T. Kat. He now lives with the four grown cats pretty much 24/7. I moved the last of his stuff into their room Monday (yesterday). He still irritates them by grabbing and chewing too much, but is finally starting to learn when to back off. It took a lot of being hit with paws (not much claws) to achieve this, the little cuss is quite hard-headed. The grown cats have actually been quite patient with him.

GOING TO MARS (or anywhere else nearby) the posting version

2011-07-25T14:14:00.000-07:00

This is not the text or the presentation form of my paper to be presented August 6 at the Mars Society convention in Dallas. But, it is a shortened version of the major points. I did include all of the figures here, which pretty much speak for themselves.

The title of the paper is GOING TO MARS (or anywhere else nearby), for which the document date is July 17, 2011. It is a refinement of pencil-and-paper bounding/feasibility work I did in fall and winter 2010, sitting in the recliner with just a calculator and a spreadsheet, in the evenings.

SUMMARY OF RESULTS:

The study was aimed at going to Mars, but the same vehicles could be used anywhere in the inner solar system inside the orbit of Mars: Mars, Venus, Mercury, and the near Earth objects (NEO’s) which include some asteroids.

I made entirely-different assumptions about what was important and what were appropriate constraints on the mission and vehicle designs:

This is a clean-sheet design. The differences with other studies are: (1) legacy hardware and contractors are absolutely not required, (2) new developments are minimal, (3) maximum self-rescue is designed into every phase, and (4) all known requirements and risks for life support absolutely take precedence over any other factors in the design.

I got startlingly different results for what is possible and how hard and expensive it might be, results far different than anything I’ve seen published in decades:

One 9-month round trip, 6 men, 16 one-week landings, nothing thrown away, minimal technology developments (but fairly certain of success), no new launch rockets, under $8 billion in direct launch costs, possibly as soon as 5 years.

A wild guess for the overall program might be 3-4 times that amount (around $50 billion in 2010 dollars), if done by a team of lean, efficient, well-focused contractors, with a lean, efficient, well-focused government agency leading it. Compare that to Apollo 4 decades ago: 6 brief landings of 7 attempted, for around $200 billion in 2010 dollars. (We know more now.)

A further startling result:

All mission assets are reusable, and left in place around Earth and Mars for the next missions to refuel and reuse.

The assumptions made for this study really are quite different from those of any other Mars mission study this author has ever seen, or heard about. It should therefore be no surprise that the results obtained here are at considerable variance with the expectations of most folks knowledgeable of this field of endeavor. That difference does not invalidate these results!

STARTING POINT:

I started with a proper definition of the true mission objective. That brings up the question of what is exploration, what do we want from it, and where does it fit within the larger scheme of things? I used the best model from 500 years ago for the voyages to the New World from Europe, and its subsequent exploitation and settlement. See figure 1, which points out that an exploration mission must answer two crucial but deceptively simple questions. Experimental and economic bases and colonies come later, as the figure shows. Note the differing roles of government and industry in those different phases. They are very important!

Figure 1 – The Most Successful Model in History for Exploration and Colonization

Carrying this out means these landings are not just Apollo-style “flag-and-footprints” missions with “a towsack full of surface rocks” as the “science return”. It also means we make many landings in one trip, to make the effort worthwhile. That is quite a different mission objective. I have not seen that idea proposed since the 1959 Disney movie about going to Mars.

MISSION ARCHITECTURE:

Apollo did identify two different ways to reduce required launch rocket size, but only used one of them. See figure 2. Since both are effective, and we currently have no gigantic launch rockets, why not use both architecture options in a single mission design? That’s what I did in this study: see figure 3.

Figure 2 – Choices of Mission Architecture for Apollo to the Moon

Figure 3 – Use Both LEO Assembly and LMO Landers to Reduce Launcher Size

IMPACT OF LIFE SUPPORT ISSUES:

Figure 4 lists the known life support issues, and indicates their characteristics and time dependence effects. One should note that we do not yet know how to do a closed-cycle ecology. That would be a major technology development, especially since once we can do it here, we have to adapt it for spaceflight.

Another thing to consider is that we do not yet know how to preserve packed food acceptably beyond about a year and a half; that is another major technology development.

A third thing to consider is that we have no direct experience with the potential therapeutic effects of fractional gee for the artificial gravity issue. Surrogate studies (bed rest) simply cannot be trusted enough to risk lives. Surrogates are not exact models!

The only real, trustable experience we have is at 1 gee and zero gee. But, artificial gravity solutions at 1 gee and 4 rpm will be big (56 m radius), heavy, difficult to build, difficult or impossible to maneuver while spinning, and horribly expensive. (See the 7-31-11 posting for more details on this issue.)

The net message is that everything in figure 4, combined with what we know and can do today, points toward planning zero-gee manned missions with flight times under one year. Since minimum energy mission plans call for 2-3 years round trip, we will have to find a way to fly the men very much faster: we need some “hot rod propulsion”.

Figure 4 – Life Support Issues Are Time-Dependent

EFFECT OF AVAILABLE LAUNCHERS:

Some typical launchers are shown in Figure 5. Most of these are flying. There are a couple no longer available, and a couple not yet flying, shown in the figure. SpaceX’s Falcon family has by far the best cost per unit payload to LEO in the industry. Its Falcon-9-heavy should make its first flight next year, meaning it should be available in a timely fashion for this design.

I was using 2010 data right off SpaceX’s website for this, to set my dockable module size at 32 metric tons maximum. In February 2011, they revised their payload figure to 53 metric tons, which just makes things better. However I did not go back and redo my study for that change. The figure shows that update in red. The price range for 2011 also updated, but is still centered near the 2010 value, so I did not update that in the figure. Oops, there's a typo in fig 5: the payload for Falcon-1 is 1.01 metric tons, not the value shown. I used 1.01 tons in my analyses, no errors there.

Figure 5 – Selected Launcher Data

LANDER AS A CRITICAL DESIGN ITEM:

The total weight of landers and landing equipment thrown to Mars is a critical issue for vehicle size, the number of launched modules to create those vehicles, the total number of launch rockets required, and thus total effort and expense.

Landing coverage on Mars requires that we consider plane changes in the lander “burns” for descent and ascent. I used 30 degrees (arbitrary), so that from an ecliptic-derived 25 degree inclined parking orbit, I could cover plus or minus 55 degrees of latitude. With a little plane change capability in the transfer vehicle to achieve a higher parking orbit inclination of 60 degrees, polar coverage is possible.

The landers will be too large to benefit much from aerobraking on descent, but will need a bit of heat protection. Min mass thrown to Mars requires these landers to be reusable to the greatest extent possible. Rocket braking descent and rocket ascent with 30 degrees plane change sets the two way velocity change requirement at 12 km/s. This is out of reach of chemical propulsion single stage with a rugged inert weight fraction of 20% and a decent payload fraction of 10%.

Figure 6 shows the best choice to be a solid core nuclear thermal rocket (NTR) for a single stage reusable lander. This is a technology extensively tested 4 decades ago that came within a year or two of flying, as the NERVA engine. It is not a major technology development, although it needs to be recreated along with the supporting expertise (engineering art).

Figure 6 – Why a Nuclear Rocket Single Stage Reusable Lander Makes Good Sense

How we use that lander ties directly into the objective of exploration (answer the two questions “what all is there?” and “where exactly is it?”) and also the selected mission architecture. Crew safety also gets into this as top priority.

We send that lander down with an armored inflatable Quonset hut along the lines of the Bigelow Aerospace space station modules, to be erected at a safe distance from the radioactive reactor core of the lander engine. We use a crane for loading access, which is part of the unusual equipment to be made part of this lander’s inert structural weight, along with some sort of refueling probe and the landing legs.

The crew comprise a pilot/engineer, a geologist/geochemist, and a biologist/biochemist. They are cross-trained well enough to assist, or even fill in for, one another. We send lab equipment they can use to do most of the science on-site, during their nominal week-long stay on the surface at each site.

We send a rover car big enough to use as an equipment transporter as well as an exploration vehicle. It has a real drill rig on it, capable enough to sample at least a km down. We also send small robot mini-rovers with this crew to assist in exploring and sampling.

We split the Mars mission “vehicle” into a fast manned vehicle capable of returning just as fast on its own fuel if rendezvous at Mars fails, and a number of “slowboat” unmanned vehicles that bring the landers (with equipment) and all of the landing propellant supply. The one-way “slowboat” velocity-change requirement, with some plane-change capability, is about 8 km/s.

Since the lander engine is solid core NTR, and 8 km/s is well within its single-stage capability at reasonable inert mass fractions, we also use the nuclear landers as the unmanned transfer propulsion. This avoids the introduction of a third propulsion item for those vehicles.

Figure 7 – Baseline Site Camp Scenario with the Single Stage Nuclear Lander

This lander needs landing legs and a squat form factor for stability on potentially rough ground. A cartoon of it with all the data and characteristics is given in figure 8, and its estimated performance in figure 9, relative to min (no plane change) and max (30 degree plane change) velocity-change requirements.

CREW SAFETY ISSUES REGARDING LANDER USE AND NUMBER:

Crew safety concerns demand that 3 persons stay in orbit (doing science from there), while monitoring 3 persons on the surface. We must have one lander “in reserve”, ready to use as a rescue vehicle, whenever a lander is on the surface. We must have 3 landers, so that if one becomes inoperative, we need not terminate the entire mission. The rescue crew is 1 man.

3 landers sets the number of “slowboat” unmanned vehicles, each with 1/3 of the landing propellant supply, that must be assembled from 32 ton modules by docking in Earth orbit. This is much like the modular assembly of the International Space Station (ISS).

I would have liked multi-engine landers, but the nuclear engine design does not scale or throttle well enough to do that. It will just have to be tested more and proven fully reliable.

Figure 8 – Configuration Layout and Data for the Reusable Single-Stage Lander

Figure 9 – Estimated Performance and Requirements for Lander

IMPACT OF LIFE SUPPORT ISSUES DRIVES TRANSFER HABITAT DESIGN:

The living space requires something around half the available volume per person of the old Skylab space station of the 1970’s. This station was “super roomy” for 3 men up to half a year. It weighed about 85 tons. A design of 3 32-ton modules docked together matches the weight of Skylab pretty closely, although this form factor is a bit skinnier.

Safety demands that all 6 crew be able to shelter for a day or two from solar flare event radiation. (The slow drizzle of cosmic rays is mitigated to acceptable levels simply by round trip travel times under one year.) Solar flare radiation can be shielded by water or wastewater, and maybe a little steel plate placed strategically.

Safety also demands that critical mission flight maneuvers be conducted, solar flare event or not. Therefore the vehicle flight deck is the radiation shelter, and it needs to be surrounded by the habitat water and wastewater tanks. One transfer habitat module is rigged for these functions, as shown in figure 10. The second module must be wide open to provide the open living space, and a place for eating, exercise, and recreational activities. The third is a combination crew dorm and packed storage module, which also provides the privacy function.

Figure 10 – Three-Module Crew Habitat Section Layout

CREW RETURN DRIVEN BY SAFETY AND EXISTING TECHNOLOGY:

I “reversed-engineered” the data on SpaceX’s website to estimate performance of its Dragon capsule, and how much more velocity-change I could get with extra thruster propellant tanks in the unpressurized section, see figure 11. It may need a bit thicker heat shield, as the emergency “bail-out” free return velocity is over 16 km/s. This craft looks like an excellent crew return vehicle as well as emergency return. One modified Dragon could carry all 6 crew; I used 2 for the safety of redundancy.

Figure 11 – Modified Spacex Dragon Capsules as Crew / Emergency Return Vehicles

MANNED VEHICLE “PAYLOAD”:

The figure 10 crew transfer habitat plus the two figure 11 modified Dragon return vehicles are the payload of the manned vehicle to be assembled (and recovered) in Earth orbit, all “single stage” in the sense of jettisoning nothing. The remainder of this vehicle is its propellant and its “hot rod propulsion”.

A LOW-RISK PATH TO “HOT ROD PROPULSION”:

I picked a nominal 75 days to cover a nominal 100 million km, essentially straight-line, flight path, to and from Mars, on both sides of a single opposition. These add up, with 16 weeks at Mars, to a nominal 9 month mission. That’s under the year limit for life support without artificial gravity, with a time margin in case something untoward forces us to slow down a little.

Figured as a simple square-wave velocity-distance trace, the average coast cruise is almost 16 km/s. The departure and arrival velocity changes are roughly equal this magnitude at each end of the trip. For safety’s sake, we must carry the propellant for the return trip in this vehicle, just in case rendezvous should fail at Mars. That’s double the one-way requirement.

So, the total velocity requirement is then about 62 km/s, way beyond the capability of solid-core NTR single-stage at believable inert weight fractions. We have to have “thrusty” propulsion, around 6000 s of minimum specific impulse, to meet this requirement, because we also need around 0.05 gee vehicle acceleration to make these burns effectively “impulsive”.

This “hot rod propulsion” is the only development item. To minimize the development risk, we must pick “not from scratch” items, and we need two worked in parallel, to ensure one will be ready when we need it. It would help if they both use a cryogenic liquefied gas as propellant, so that we could use a common propellant tank module for all vehicles, as in figure 12.

Figure 12 – Rough-Out Layout of the Common Propellant Module

There are two good candidates meeting all those requirements: the hydrogen-propellant open-cycle gas-core NTR of 4 decades ago (see figure 13 for the estimated characteristics), and the argon-propellant VASIMR engine being developed by former astronaut Franklin Chang-Diaz in Houston. Both would be multi-engine designs technically, as well as for safety. I personally know more about the gas core NTR, so I show it in the illustrations. (See "changes" in the 8-9-11 post above.)

Figure 13 – The Open-Cycle Gas Core Nuclear Thermal Rocket

The gas-core NTR was never tested propulsively 4 decades ago, but did come within about 2-3 years of it, so it is not “from scratch”. Controllable gas phase fission, and the necessary fireball containment flow pattern were successfully bench tested back then. Putting them together and testing a series of engine designs is the technology development path here.

VASIMR is being tested at the 200 KW power level, and may soon flight test on the ISS, so it is not a “from-scratch” development. Its strong point is high efficiency. Its weakest point is the electrical power source. The tradeoff between scale-up and number of thrusters is as yet undetermined for this mission. It will need a space-worthy, flight-weight nuclear-electric power plant, likely in the multi-MW range. We already know how to build those down here. The technology development item is making those power plants space-worthy and flight-weight.

HOW ALL THIS FITS INTO A PLAN AND VEHICLE DESIGNS:

There are four vehicles to assemble in LEO from docked modules of a maximum 32 metric tons. Three are unmanned “slowboats” that bring the landers and all the propellant for the landings to Mars on one-way trips. The other is the fast-trip manned vehicle with the “hot rod propulsion” that is recovered intact and complete in LEO after the mission is done. All assets are left in place in orbit around Mars and Earth for future missions to refuel and re-use, which is of tremendous technical and economic benefit. See figure 14 for the final mission plan. Figures 15-18 describe the vehicles.

Figure 14 – The Basic Mission Plan

Figure 15 – Unmanned Vehicle Configuration for the Mars Mission

Figure 16 – Estimated Performance and Requirements, Unmanned Vehicle

Figure 17 – Manned Vehicle Configuration for the Mars Mission

Figure 18 – Estimated Performance and Requirements, Manned Vehicle

FINAL COMMENTS:

There is another technology needed to carry this out: a dexterous, rugged, lightweight space suit. This is presumed to exist by the time these vehicles are being assembled in LEO. That suit will be needed for lander refueling operations, and for use on the surface of Mars. I would suggest a mechanical counterpressure design, but I also strongly suggest that the required compression level requirement be honestly investigated, not simply presumed equal to the traditional 1/3 atmosphere of the gas balloon suits we have been using. See the article on this same blog site titled Fundamental Design Criteria for Alternative Space Suit Approaches and dated 21 January 2011 for more details about that.

This is not an optimized design, it is just a ballpark exploration of what is feasible under these very realistic assumptions and constraints. The crew size, number of landings made at once, details of rendezvous in LMO, and a host of other details, all need exploration and optimization with far more rigorous calculations than this.

However, all of these other possibilities will be found to operate within the same basic ballpark that this study identified: that is, multiple vehicles to Mars, manned items flying very fast, combined use of LEO and LMO rendezvous to reduce launcher size to something reasonable, and multiple landings from a single mission.

The real critical factor for success is not any of the technical things, but in having lean, efficient private and public entities to do this work.

The same basic vehicles and components can be used to visit anything in the inner solar system within the orbit of Mars; i.e., anywhere the round trip manned vehicle flight time is under 1 year. One simply adjusts the number of common propellant tank models in each stack to meet or exceed velocity change requirements.

For Mercury, the same four vehicles are used. For Venus and the NEO’s, men cannot land, so only the manned vehicle need be used. The moon is too close to require this kind of vehicle, although it could travel there.

The Good News

This particular study design adds up to a direct launch cost for Mars of about $8 billion (2010 dollars), based on cost figures right off the SpaceX website, and the total number of launches to orbit all of the required modules. A set of “lean” contractors led by a “lean” government agency could actually get this done rather quickly, in perhaps as little as 5 years. In this context, “lean” means dedicated, focused, efficient organizations, unhampered by outside bureaucratic or political interference.

Done by “the usual crowd”, the cost would be many, many times higher, and the timeline at least 15-20 years, if it could be done at all. We’ve seen this before. Ever since Apollo, really.

The Bad News

The NASA that could lead and manage this project as presented herein, is very most definitely not the NASA that we have. We have an enormous government space agency trying to do an entire plethora of things, in which every little project is a budgetary line item and “political football” in Congress. This is a Congress that now arrogantly “designs” heavy lift rockets (that we may or may not even need) by pork barrel politics instead of engineering realities!

It is this author’s opinion that we have not done the government space agency function correctly for some 40 years now, which in large part explains why men have not flown anywhere new in all that time. (I have little hope we can really change that.)

But, that’s another topic.

Budgetary Malfeasance in DC

2011-07-12T19:11:00.000-07:00

This budget impasse in DC is the most egregious example yet of election politics being more important to our elected representation than their sworn duty to do the people's business. Both side's positions on this issue are based on lies to us.

There is an analogy between what I call the “leaky bucket”, and any budget , from an individual to the US government. This picture makes it clear:

As drawn, the bucket will sooner or later run dry, any time the outflow exceeds the inflow. The same is true for any bank account. This is running a deficit.

The bucket will also sooner or later fill to overflowing, any time inflow exceeds outflow. Again, the same is true of any bank account. This is running a surplus.

The water level in the bucket will stay steady, any time the outflow exactly equals the inflow. Same is true for any bank account.

If the account is full, then one can run a deficit temporarily in time of need, but only until the account is emptied. A recession is one such need: government spends when no one else will, so that the economy does not grind to a complete halt (a true disaster) for lack of any spending at all.

The upper control valve (inflow) is analogous to the various forms of revenue. For a government, that is taxes. For the US government, there are tax rates, and there are tax breaks and loopholes that lower revenue. Those are the available tools.

The lower control valve (outflow) is analogous to rates of spending. Again, for the US government, these are the various funded programs. Spending is controlled by deciding how much and which items to fund.

The water bucket fills fastest if the outflow is turned off, and the inflow is turned up to maximum. But, in the real world, there are definite limits to how much spending can be cut, and how high taxes can be. Most of the time, you have to do both. Here’s why…

When taxes get too high, people get hurt. Those hurt worst and first are the poorer folks (overtaxing businesses does cut jobs). Carried to extremes, everyone suffers.

When spending cuts get too deep, people get hurt. Again, those hurt worst and first are the poorer folks. Carried to extremes, everyone suffers.

The sacred cows for the Democrats are the entitlement programs, especially Social Security, Medicare, and Medicaid. They would prefer to balance the budget by raising taxes. It’s an ideology thing.

For the Republicans, the sacred cow is “no tax increases”. They would prefer to balance the budget by spending cuts. Ideology again.

We cannot cut spending enough to balance the budget by that means alone: too many people would get hurt. That’s simply a fact of life.

Similarly, we cannot balance the budget by raising taxes alone: too many people would get hurt. That is also a fact of life.

Clearly we have to do both. It’s way past time that our elected representatives put aside their ideologies and get on with the job of governance.

We need some statesmen who will actually do the peoples’ business, not more politicians putting party advantage ahead of the good of the people.

So, where are they?

Recovery, part 4

2011-07-07T18:21:00.000-07:00

Recovery, Part 4 Thursday 7-7-11

On Wednesday, 7-6-11 (day 23 since surgery), I drove myself to therapy with a backup plan (Ellen in the other seat) in case I couldn’t get that far (15 miles one-way). I made it, but I didn’t push it by driving home, too. The truck is still daunting: I still need to “heel-and-toe” the pedals, but that opportunity exists in most passenger cars. Getting in on the driver’s side is actually easier than the passenger side: I get to “lead” with the “bad” leg.

Therapy was brutal but effective: I got “over-the-top” for the first time on the bicycle exercise machine, and ended the session with 3-100 degrees motion in the knee (deep in the pain zone). 0 is straight out, the hardware limit is 140 degrees flexed. The weakest muscle group is the quads that kick out the lower leg. They’re still partly asleep from the anesthesia. The therapists are astonished at how fast I have come this far. The Chinese cuss words help when therapy gets really brutal, we are all laughing about that. I am pushing this as hard as I can.

At home, I haven’t used the walker for over a week, and I have been bathing without the shower chair for almost a week. No pain meds during the day, and for the last few days, only a half dose at nights to enable sleep. I still use the ice pack machine occasionally to control swelling in the evenings. But the swelling is now just in the knee itself: can’t yet see my kneecap.

Next therapy session is Friday 7-8-11 (day 25 since surgery). There will be no one available to drive me, so this one is “for real”: I have to drive myself two-way. After Wednesday’s experiment, I believe I can do it. Watch this space for updates. Not going to try climbing onto the Farmall until those sleepy quads wake up.

Walking range is well over 1000 yards cumulative in a day, maybe up to 200 yards in one trip. My stride is still a bit limited. Standing time is still kind of short: maybe an hour. Sitting up in a chair with feet on floor is still limited, too. In either scenario, the pain builds with time, until I have to relieve it by laying down or going feet-up in the recliner. I can sit up straight for only about an hour, yet. Not ready for classroom duties, not just yet.

If you haven’t yet heard, last Friday (7-1-11), we got “adopted” by a kitten about 2 months old. I think he was a “dumper” out here, a friend in McGregor got “adopted” by what looks like a sibling the very same morning. We named our little male “Kit T. Kat” (say it fast and see how that comes out), with the T. standing for “Thermonuclear”, since he is definitely atomic-powered in his playfulness. There’s a second layer to the nuclear joke: if you look at his given initials K. T., that’s the standard abbreviation for “kiloton”, another standard measure in atomic stuff. Getting the four grown cats to accept him as one of them is going to be a long, tricky, careful process. But he already has one of the two dogs scared of him.

Recovery, Part 3

2011-07-03T11:20:00.000-07:00

I am progressing fast, but still a very long way from fully recovered. Day 19 (Saturday 7-2-11) I actually drove a car about 3 miles. But that's about the limit before leg posture-induced pain gets to be unbearable. Plus, the gas and brake pedals need to be the same height, and close to the floorboard, so I can "heel-and-toe" them. Lifting and bending that leg are still severe problems to overcome.

I haven't needed the walker for several days now, and I was also able to bathe without the shower chair for the last couple of days. The swelling in my leg has gone down to mostly just the knee itself. I still have to wear the compression hose to ward off blood clots for another week or so. I also still need the ice pack machine at least once a day, in the evenings.

Therapy is extremely brutal, but effective. They have me doing strengthening exercises as well as flexibility. I really push things very hard in therapy, knowing as I do now what is necessary. That's probably why I am recovering a little faster with this knee than the other one 2.5 years ago. It is excruciating, but well worth it. There is no other way.

I can get down to nearly straight at roughly 5 degrees with some pain, and Friday I went beyond 90 degrees flex with a whopping lot of pain. When I hit 0-120, therapy discharges me. The hardware limits are 0-140, which I achieved with the other knee a couple of months after discharge from therapy. I still have a very long way to go with this right knee. But I will get there!

Recovery, Part 2

2011-06-29T11:20:00.000-07:00

Wednesday 6-29 (day 16 since surgery on Monday 6-13): I successfully drove son’s Ford Focus across the yard to the shop. I also successfully climbed into my F-150 and manipulated the pedals forcefully enough to drive safely. I cannot do this activity for very long, perhaps for a 5 mile trip at the most. But, it is very satisfying that I now have such a capability if there is an emergency. With the other knee, I did not reach this capability until day 29.

I have been walking without a walker on smooth paved surfaces in daylight since about day 11. My gait is hobbled because the knee joint motion is still limited to about 5-75 degrees (0 is full extension, 90 is flexed to a right angle, and the hardware limits are 0-140). We’re working that issue much harder now in therapy, since I got the staples out on Monday (day 14).

I still have to use pain medications to sleep at night, and even so, it is a hard thing to do. I wake up a lot during the night. But, I haven’t used the pain medications at all during the days, since day 11. This recovery is definitely proceeding faster than the other one 2.5 years ago.

In Recovery

2011-06-22T16:34:00.000-07:00

The knee replacement surgery was very successful. I am home and undergoing outpatient therapy. I seem to be about a week ahead of the recovery timeline experienced with the first knee, 2.5 years ago. This is based on some diary-like notes I made back then.

The photo is me scurrying around the house on the walker. I probably won’t need the walker much longer. As of this writing, the knee is stable enough to bear full weight walking, for a majority of the time. 3 days ago, that was entirely not true. Status changes very quickly.

Photo 006-2

There’s a lot of reading and paperwork to catch up, plus getting my paper ready for the Mars Society convention in Dallas August 4-7. Until recently, I didn’t feel much like doing any of this. I still need a good long nap after therapy sessions. But, things are good enough now to start taking on those chores.

Another Red-Letter Event

2011-06-12T10:06:00.000-07:00

My wife Ellen was an adoptee. Her parents made no secret of her Japanese birth mother’s identity (Toshie Uchida), and did encourage Ellen to find her. They did not know where she was, though. After both of Ellen’s parents passed on, Ellen felt like an only child, and determined to find Toshie, or at least find out what happened to her.

But, with the help of dear friends in Japan, and a private investigator in Chicago, Ellen did finally locate her near Peoria, Illinois. Ellen made contact, and found out she has a half-sister and a nephew that we never knew about. Emails and phone calls led to a trip to Peoria June 4 through 6. All three of us went. This turned out quite well, as you can tell by the looks on everyone’s faces. Here are four photos from the trip.

The first shows Ellen, her sister Joyce, and her mom. We all marveled at them gabbing away, while around that table, and others. It was clear they were related.

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The second is of most of the men associated with these three. Left to right are me, Ted Meyer (Toshie’s husband for many years), our son James, and Joyce’s husband Tim Ricci. Ted was Navy and Marine, then a railroad hand for decades, now retired. Tim is a policeman.

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The third shows my son James and his cousin David (Joyce and Tim’s son). Both of these two are into computers and gaming.

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The last is a close-up of the two sisters, Joyce and Ellen. The picture says it all. As it turns out, both are involved with correctional work professionally. Amazing.

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This trip was a wonderful visit. I think there will be many more.

Another Red-Letter Day

2011-06-12T08:19:00.000-07:00

The two VW beetles have served their purposes for me, and are going back into mothballs. I have completed the preservation activities for both engines and both fuel systems. There really is nothing to preserve about the brakes: both have seen recent service, and both are already DOT-5 silicone converted systems.

The blue 1973 “ethanol VW” served very well for 5 years as an E-85-only demonstrator, and last year in flex-fuel configuration for experimental confirmation of my earlier conclusions from the F-150 about the limits of “stiff” gasohol blends in unmodified vehicles. It is still working quite well, but has a very “high-time” engine and a very “high-time” transmission. It is time to put it back to rest, before I break something. I remain convinced that the ethanol extended its useful life considerably: it should have burned a valve or lost compression entirely, long before now. It has a little over 250,000 original miles on chassis and transmission, and well over 100,000 miles on the last engine rebuild.

The old white 1960 “gasohol VW” had been de-mothballed to replace the blue VW in ethanol experimentation as a daily driver. Turns out I didn’t need it after all, my wife’s old Nissan Sentra is currently filling that role. As for the VW, I “woke it up” completely unmodified on E-33 gasohol blend, and it has performed excellently on that fuel. This vehicle is so old as to have plain carbon steel valves and seats, so it requires a lead substitute with today’s unleaded fuels. I have successfully used 9 cc per gallon Marvel Mystery Lube in that role since the mid 1970’s. The car has a little over 227,000 original miles on it, but the engine and transmission are both “low-time”.

Ethanol fuel research continues with the Nissan, my F-150, my Farmall tractor, and all my lawn and garden equipment. All are completely-unmodified equipment running on “stiff” gasohol blends (E-25 to E-35) except the Farmall, which is a straight E-85 machine. I still maintain that blends up to E-35 are fine as "drop-in" fuels for every car in the fleet. The vehicle will get the same mileage it got on gasoline, it will run cleaner inside the engine and exhaust, and I am convinced it will last longer precisely because of reduced flame sooting.

If you want a good summary of how I converted the blue 1973 "ethanol VW", or how I routinely make my own gasohol blends at the pump, or what I know about ethanol fuel compatibility in engines and fuel systems, then look below for two articles posted on this site. Together, they make a pretty good “handbook” for you. See “How-To” For Ethanol and Blend Vehicles, dated 2-12-11, and see Ethanol Does Not Hurt Engines, dated 5-5-11. The keywords for both are “fun stuff, old cars”.

End of May Update

2011-05-31T21:46:00.000-07:00

My Mars mission study has been accepted for presentation at the Mars Society convention in Grapevine, Texas, August 4-7 this year. I will barely be back on my feet after knee surgery at that time, but I wouldn’t miss this meeting on a bet. The basic study was posted here 12-20-10, with an update posted 1-8-11, correcting an error I made in a Falcon-9 payload number.

All of the launcher and capsule data I used were obtained from the Spacex website for 2010. Sometime after January 2011, Spacex updated these data, showing a significant increase in payload capacity estimated for the Falcon-9-heavy, which is now supposed to fly for the first time out of Vandenburg late this year. That change just makes my design even more cost-effective, but I am not going to change my presentation because of it.

I did a little nontraditional thinking, dominated by crew safety, and came up with a tremendous science return for a lot less investment than anyone will likely believe. This is way more than “flag-and-footprints”. You never know what you can do until the blinders come off. It is going to be a lot of fun.

Latest news releases in AAAS’s peer-reviewed “Science” journal confirm everything I posted here about the nuclear plant disaster in Fukushima, Japan. You have to go inside the reactor building ruins before you find radiation too dangerous for adult humans, and then it’s just barely enough to cause prompt radiation sickness. The whole argument over risk to the Japanese public seems to center on the putative risks of very low-level exposures. With no evidence to support or deny a very slight risk of slightly-above-background-level exposures, it seems rather silly to worry much about it.

Now that the Mars mission study is done, and I will be off for a while with the knee, I can return to the question of ramjet-assisted launch. I’m pretty sure that horizontal takeoff with a two-stage vehicle is feasible with simple parallel-burn rocket and ramjet in a winged first stage. I’m pretty sure the staging velocity can be pushed to right at Mach 6 with a simple subsonic-combustion hydrocarbon-fueled ramjet. The upper stage rocket can be either winged or plain gravity-turn ballistic, as desired. To hit Mach 6 staging, a spike inlet will be required, which pushes practical takeover velocities to around Mach 1.5. Parallel-burn can be used right at staging to achieve path angle. I think the staging altitude is probably closer to 60,000 feet than 80,000 or 100,000 feet altitudes. That study needs a rerun with a resized engine and revised staging altitude. The earlier versions were posted here last year.

I don’t know yet whether ramjet assist is beneficial for a vertical-launch gravity-turn staged vehicle. But, if it is, it will be leaving the sensible, usable atmosphere closer to Mach 2 or 3 speeds than Mach 6. That’s a lower-speed ramjet design, most likely a simple normal-shock (pitot) inlet and a simple convergent-only nozzle. Takeover will likely be at slightly-subsonic to merely-transonic speeds. I have most of the guts of a sizing and performance code programmed, but it still does not quite work right in performance mode. Maybe I’ll have a chance to work on that while recuperating. Properly sized with an appropriate integral booster rocket, this kind of ramjet would be a strap-on staged off substantially earlier than the typical first stage burnout, in a two-stage vehicle. But it just might help.

Watch this space for future postings.

Presentations Available for Lunch and Dinner Groups (v.8) GWJ 5-7-11

2011-05-07T13:47:00.000-07:00

This is what I have available as of the current date. I would be happy to present any of these to any interested group. Just contact me.

Name: Ethanol Experiments (set 12 – final, 1-24-11) (15 slides)

Status: Ready

Summary: After characterization on gasoline during 2005, an old VW beetle was converted in October 2006, at no cost, to use E-85 ethanol fuel, and later tested as a “flex-fuel” on blends from E-15 to E-57 in 2010. An old farm tractor was also converted to E-85 with flex fuel potential. In 2007, tests began on blends from E-15 to E-45 in an unmodified pickup truck. Not long after, multiple pieces of unmodified 4-stroke lawn and garden equipment were operated routinely on E-22, then E-34 blends. In 2010, an unmodified 1998 Nissan Sentra was operated routinely on blends from E-20 to E-35, and once accidentally near E-50. Technical data of a very favorable nature were obtained on the truck and the VW. How-to information for do-it-yourself blend operations with unmodified vehicles is also included.

Name: Fuel Savings Advice (12 slides)

Status: Ready

Summary: This short presentation offers both technical truth and practical advice for anyone interested in saving fuel, independent of any switch to alternative fuels. The big items are matching vehicle to mission, and carpooling if possible. A smaller item is a national speed limit, but, surprisingly, 65 mph will do as much good as 55 mph. This is because of the tradeoff between road load reductions and efficiency losses as speeds decrease, an effect not considered when the old 55 mph limit was imposed in 1974.

Name: “One Book” Presentation (18 slides)

Status: Ready

Summary: This was prepared as a part of the Waco ‘One Book” project, spring 2009, related to the Homer Hickam book “Rocket Boys”. The presentation was to explain how and why I became interested in engineering, and what I did during my career, which included rocket and missile work. Many old aircraft, rocket, and other technical-item photographs were found and included, some quite rare. These were things that were simply “part of my life”, starting in the 50’s.

Name: Asteroid Results from Spain (26 slides)

Status: Ready

Summary: I presented a poster paper at the 1st IAA international conference on planetary defense, held April 27-30, 2009, in Granada, Spain. This conference dealt with detection of, and mitigation of, asteroids and comets that may impact the earth. Detection efforts are in progress, but incomplete, largely for lack of funding. Concepts for mitigation schemes were discussed, but with as-yet little real development work, again largely due to lack of funding. Civil defense coordination and warning mechanisms for this international problem do not yet exist. Not discussed at the meeting: propulsion to support such schemes does not yet exist, although I identified some promising ideas abandoned long ago. Some slides of pictures taken while touring Granada are included at the end.

Name: Global Warming (vers.B 14 slides, obs. Vers.A 13 slides)

Status: Ready

Summary: The earth’s climate is warming, as evidenced by polar sea ice thinning and glacier retreat on land. Icecaps on land could melt, causing massive sea level rise. Rainfall patterns (and agriculturally-fertile zones) could shift. Independent of “fault”, the options are to act, or not to act, in two areas. These are attempted mitigation, and coping with massive change. Concepts for mitigation are currently under public discussion and debate, but concepts for coping are not. The decision to act, or not, should be made on facts and logic, not politics and belief-preferences. Some experimental facts are given, and some recommendations made.

Name: Peak Oil and Coal (15 slides)

Status: Ready

Summary: The empirical “Hubbert curve” model for resource depletion is presented and explained. This includes its proper application and its limitations. Hubbert’s successful 1956 prediction with this model, of peak US oil production in 1970, is verified by historical US production data. The relative impact of Alaskan oil is included in the data. An opinion is presented regarding planetary “peak oil”, and its implications. Coal production in the UK followed a similarly-shaped curve, so the Hubbert curve model can probably be used for planetary “peak coal” predictions. An opinion is presented regarding that scenario, although the dates are less certain for coal.

Name: How Things Burn (27 slides, initial version)

Status: Ready

Summary: This presentation covers the “kitchen physics” and “kitchen chemistry” of how various gaseous, liquid, and solid fuels burn with air. The intended audience could be laypersons, or engineering students just getting started in this discipline. To the classic “fire triangle” is added (1) mixture limits and autoignition behavior, (2) how fuel phase changes interact with combustion, (3) three basic “burn rate” regimes, (4) the influence of theoretical thermochemistry predictions, (5) an explanation of how piston engines really work and their problems, (6) the “combustion aerodynamics” of several through-flow combustors, (7) how pool and pile fires really work, and the “firestorm”, (8) solid coal and carbon particles as the odd slow-burning case, (9) the radiation physics of fires, (10) simple models for real through-flow combustors, and (11) practical effects with gas-solid slurry fuels. (Emissions phenomena and some practical case studies will be added to the next version. Future versions will likely include rockets, and various kinds of explosives.)

Name: Ramjet-Assisted Staged-Rocket Launch Vehicle (20 slides v.11-18-09)

Status: Ready

Summary: This documents the results of a back-of-the-envelope three-stage vehicle sizing study that trades higher blended performance for the higher structural fractions that allow true re-usability. Subsonic combustion ramjets with integral boosters are added to the second stage, and burned during both first and second stage operation to improve performance substantially. The third stage is a plain rocket. This is the kind of design that could result from a “clean sheet of paper” start, looking at technologies not currently “traditional” for space launch.

Name: “Lob-Up” Nuke Launch (18 slides 11-30-09)

Status: Ready

Summary: This documents the results of a back-of-the-envelope two-stage vehicle sizing study that explores how to safely use nuclear thermal rocketry for surface launch. The design takes advantage of the ramjet-assist results to trade high performance for reusable structure in the lower stage, which lofts the upper stage and payload vertically out of the atmosphere. The nuclear upper stage then fires horizontally from its inertial apogee to reach orbital velocity with the payload. It never returns to earth, and its exhaust plume never enters the atmosphere.

Name: Carrier Plane Launch (19 slides 12-18-09)

Status: Ready

Summary: This documents the results of a back-of-the-envelope vehicle sizing study that uses a horizontal takeoff and landing hypersonic airplane as the first stage of a three-stage launch vehicle to orbit. The carrier aircraft is a scale-up and extension of the technologies in the SR-71, and the design staging point is M6 at 100,000 ft altitude. The upper two stages are simplified LH2-LOX rockets, arranged in a triple cluster for easy carriage under the airplane. These stages feature heat shields, pivoting wings, and small turbojets for flyback as 3 separate items. Final circularization and de-orbit of the 3rd stage is with storable hypergolic propellants.

Name: Oil Prices, Recessions, and the War (18 slides 2-5-11)

Status: Ready

Summary: Using gasoline price history data, this document explores the connection between fuel prices and middle east foreign policy issues, and between fuel prices and economic events. Both monopoly cartel pricing effects and supply-demand market effects are explored, including the possible effects of planetary “peak oil” just beginning to crop up. The feasibility and effects are explored of displacing (with domestic alternatives) about one third of the petroleum used for all liquid transportation fuels. (This presentation replaces “The Dark Side of Oil” as an expanded update.)

Name: The Mars Mission Design (30 slides 5-7-11)

Status: Ready

Summary: This mission and vehicle concept design study explores what could be done if one presumes a “clean sheet of paper” approach. Existing or very near-term technologies and hardware are “fair game”, but no restrictions to “legacy” concepts or hardware, or to existing contractor infrastructure, are imposed. The most startling result is that a super-heavy-lift launch rocket is not necessary to send men to Mars, at substantially less risk than was accepted for the Apollo moon missions. The second most startling result is that the direct launch costs can be quite “low”, compared to Apollo or the Shuttle. This is with no technology or hardware “breakthroughs” assumed for launch. It does assume carrying through with two of three nuclear propulsion technologies from 1959-1973 times.

Contact: Gary W. Johnson, PE, PhD
gwj5886@gmail.com
http://www.txideafarm.com
http://exrocketman.blogspot.com

AAAS Confirms 3 Topics

2011-05-07T12:08:00.000-07:00

The American Association for the Advancement of Science (AAAS) publishes a refereed journal called “Science”, in which new findings are publicized, duplicated, debated, and verified. Issue 6024 of volume 331, dated 25 March 2011, has some very interesting items. The 1 April 2011 issue (volume 332, issue 6025) has more confirmation of my articles, and more details than anything I have seen (so far) in the public media about the nuclear disaster in Japan.

Relative to “Oil Prices, Recessions, and the War” dated 4 February 2011:

In issue 6024 on page 1510 is a “News Focus” article “Peak Oil Production May Already Be Here”. The “Science” article has a plot of actual production history data which verifies that non-OPEC oil production has already peaked. Closer inspection of the same plot reveals that OPEC oil production is peaking, or has just about peaked.

The article also makes the point that “unconventional oil” (such as the Alberta tar sands), cannot make up the difference between energy demand and a peaking conventional oil supply. These are all points I made in the “exrocketman” article.

It feels very nice to be verified by items published in a refereed scientific journal.

Relative to Pre-Clovis Hunters in Texas on “exrocketman” dated 9 April 2011:

In issue 6024 on page 1512 is a “News Focus” article about the Buttermilk Creek archeological site, in Texas. The actual technical report on those findings is on page 1599 of the same issue. The scientific debate and verification now begins, but it looks like they have made a pretty good case for people in Texas long before the Clovis hunters.

The institutions making up this team include Texas A&M, Baylor University, and the University of Minnesota. We can be very proud of these institutions (especially our two local ones).

It’s mighty nice to see “local folk” make good.

Relative to 3 articles on “exrocketman” about the Japan nuclear plant disaster:

In issue 6024 on pages 1504-1507 are three “News Focus” articles about the Japanese nuclear plant disaster and associated cleanup difficulties.

Issue 6025 on pages 24-25 has a News Analysis article that has more confirmation of my articles, and more details than anything I have seen (so far) in the public media.

These are also closely related to three articles of mine on “exrocketman”: “On the Nuclear Crisis in Japan 3-15-11, “Follow-Up on the Japan Nuclear Crisis” 3-17-11, and “Radiation and Humans” 3-24-11.

These “Science” articles confirm what I have been saying on “exrocketman”: that comparisons to the Chernobyl disaster are overblown. There is no credible risk in the US no matter what happens in Japan, or what tiny amounts of radiation might be detected over here. The risks in Japan are only serious within a few miles of the plant.

The Fukushima Daiichi plant has 6 reactor units, as has been widely reported. It has 7 spent fuel rod pools, one for each reactor unit, and 1 more for the entire plant, which has not been widely reported. None of these spent fuel rod pools had any sort of reactor-like containment, in spite of containing quantities of spent fuel rod assemblies resembling reactor cores. These pools were housed in buildings of ordinary construction that were destroyed by the tsunami, as was all the electrical and cooling water infrastructure for the entire plant.

According to this latest article, it is the spent fuel rod pool for unit 4 that seems to be the true problem, radiologically. The core meltdowns reported earlier seem to have little to do with the spread of radiation beyond the plant. Apparently, this pool went dry and its load of zirconium alloy fuel rod casings overheated and caught fire in the air. The article even mentions “aerosolization” of fuel materials and fission products from burning fuel rod assemblies, exactly as I wrote.

According to the articles, the long half-life and rather dangerous cesium-137 contamination found within a few miles of the plant has to be coming from the unit 4 spent fuel rod pool, because it has to come from older (spent) fuel. The very short (8-day) half-life iodine-131 contamination found as far away as Tokyo’s water supply is far less a threat, and cannot be coming from fuel rod pool 4. The article suggests reactor unit 3 as a possible source.

Finally, the article discusses upgraded design criteria for earthquake and tsunami risks, exactly as I advocated in my third article. This is the first time I have seen this issue raised in any public news forum. It is long overdue, of course. But, at least someone is beginning to look.

Once again, it is nice to be verified by data published in a refereed journal.

This may be a blog, but I always try to tell the truth. To the very best of my ability, I do!

Ethanol Does Not Hurt Engines

2011-05-05T17:39:00.000-07:00

A lot of folks seem to think ethanol will hurt their engine. This is not true. Here is how I know that:

The following is an excerpt from the “ethanol properties” downloadable document located on the ethanol projects sub-page on my website http://www.txideafarm.com. It addresses what materials can be used with ethanol and with “stiff” ethanol blends.

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“Materials compatibility is the next most important item. Ethanol is known to dissolve or damage certain materials, and can cause enhanced corrosion of metals, especially bare aluminums. You have to know what these pitfalls are, in order to avoid them.

Materials that cannot be used with ethanol but can be used with gasoline are old-time lacquered cork carburetor floats, the antique zinc-based “pot metal” castings for fuel pumps and carburetors, and Lexan or Plexiglas if there is warm vapor contact.

Materials that cannot be used with either gasoline or ethanol are natural rubber, butyl rubber, and polystyrene plastics. Both fuels dissolve these materials.

Materials that can be used with both fuels include neoprene rubber (any color), steel, aluminum, most polypropylene-type plastics, and both Lexan and Plexiglass if limited to liquid contact. Aluminum should be protected by a surface coating. Update 6-9-11: Teflon is also good with both fuels. Thus the typical aircraft fuel hose (stainless steel braid over teflon) is fine with ethanol.

The corrosivity of ethanol to aluminum and steel is not serious as long as the ethanol is dry, and it is far less than that of methanol, whose corrosivity gave both alcohols a bad reputation. The presence of 2+% water content greatly enhances ethanol’s corrosivity, however. The presence of water in the fuel can also cause phase separation problems (see below). Thus, if ethanol is used neat or in blend, it should be kept very dry.

As a rule of thumb, “anything good for gasoline is good for ethanol” is a pretty good guide, as long as Lexan or Plexiglas are not involved, and no truly antique parts are used (those are the zinc-based pot metal castings and the lacquered cork floats).”

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To this I would add the following about the “scare stories” being circulated by the small engine and boat motor lobbies. They quite often claim that gasohol fuels cause damage requiring engine overhaul. This is simply NOT TRUE. But there are two troubles to anticipate and avoid:

First: dirty fuel systems, which can be traced to really poor housekeeping on the part of the owner. The presence of ethanol in the fuel, with its solvent properties, will mobilize any pre-existing dirt in the fuel system, something gasoline will not do. This is quite the common problem with lawnmowers. That mobilized dirt will clog either a fuel filter or a carburetor jet, or both. You simply drain out and replace the dirty fuel (including that in the carburetor bowl), and clean or replace the clogged item. You DO NOT need to overhaul the engine.

Second: pre-existing water in the fuel system, something quite common in boats, and in lawnmowers kept outside in the weather. Ethanol in the fuel will scavenge water bottoms up to an unpredictable point (5-25% water by volume). Beyond that “limit”, the fuel will “phase separate” into a wet ethanol layer underneath a dry hydrocarbon layer. The engine is not carbureted to run on straight ethanol, and so will not run on the wet ethanol layer. Neither will it run successfully on the dry hydrocarbon fraction, which is devoid of any octane boosting ethanol or methanol added at the refinery. Simply drain and replace the fuel, including that in the carburetor bowl. You DO NOT need to overhaul the engine.

------

If your fuel system is clean and dry, you may add any gasohol blend to an unmodified four-stroke engine as a “drop-in” fuel, up to about 35% ethanol, no modifications required at all. It will get the same fuel consumption it got on gasoline, it will run cleaner internally, its oil will stay clean longer, and so the engine will actually last longer. It will soot less out the exhaust, too, which means it is causing less pollution (US EPA’s oil lobby-induced fears notwithstanding).

I have been using E-22 to E-35 blends in unmodified cars and four-stroke lawn equipment for 5 years now (Brazil has been using E-22 since 1980). These vehicle and equipment items range from vintage 1960 to vintage 1998. I have been using straight E-85 in an old car (vintage 1973) and a really old farm tractor (vintage 1944) just as long. If there was a problem with ethanol, I would know about it by now.

There is no problem!

GW

The Good, the Bad, and the Ugly

2011-05-02T08:31:00.000-07:00

Seems like good news and bad news always come together. Thus any conclusions must inherently be ugly. Oh, well, here it is.

The good news:

That was the most amazing and unexpected good news last night: Osama bin Laden is dead.

My congratulations to the US Navy Seals for a job well done, and to the intelligence community that made this possible. Keep up the good work! It's not over yet, there are at least two more top figures to go: Al Zwahari, and that other creep in Yemen. Plus all their followers, protectors, and imitators.

The spontaneous celebrations I saw on television in NYC and DC proved one thing: we are all still Americans. That's something that has been almost totally obscured by the vicious politics in recent years. (Congress, please take note, and start doing the people's business again, instead of playing those destructive politics-as-usual games.)

Putting an end to bin Laden is the result of a sustained effort spanning two presidential administrations. My thanks to both gentlemen, and all those who serve or served under them.

The bad news:

According to the released information, bin Laden was living in a gigantic compound just down the street from the Pakistani military academy. We didn't tell the Pakistanis we were doing this. There were no security leaks in this operation. I don't find those 3 items to be coincidence.

As near as I can tell, the Afghani and Pakistani Taliban are the same web of organizations. They protected Al Qaeda in Afghanistan, and I think they do in Pakistan, too. I believe the Pakistani Taliban not only controls large portions of Pakistan, they control a big portion of its government.

I'm hoping this upset will induce the non-Taliban Pakistanis to clean their house of Taliban and Al Qaeda, but I think that's a low probability outcome. The Taliban/Al Qaeda nuclear weapons will come from the Pakistani arsenal, I predict. The ties between the ISI (the Pakistani secret police and spy agency) and the Pakistani military on the one hand, and the Taliban/Al Qaeda groups on the other hand, are just too strong.

My conclusions:

The authorities are right to warn us to be vigilant, that the risk of a terrorist revenge attack is high.

Sooner or later, somewhere in the US or Europe, they will strike back as hard as they can.

Sooner or later, the “terrorist nuke” really is a credible threat.

Please, no more "birther" nonsense!

2011-04-10T09:59:00.000-07:00

The real answer to the “birthers”, from MSNBC 4-10-11:

By Michael Isikoff National investigative correspondent
NBC News NBC News
updated 2 hours 17 minutes ago 2011-04-10T16:37:33

The Hawaiian state health official who personally reviewed Barack Obama's original birth certificate has affirmed again that the document is "real" and denounced "conspiracy theorists" in the so-called "birther" movement for continuing to spread bogus claims about the issue.

"It’s kind of ludicrous at this point," Dr. Chiyome Fukino, the former director of Hawaii's Department of Health, said in a rare telephone interview with NBC.

Fukino, sounding both exasperated and amused, spoke to a reporter in the aftermath of Donald Trump's statements on the NBC Today show last week questioning whether Obama has a legitimate birth certificate.

Trump, who says he is considering a run for president, repeated his claims on CNN's "State of the Union" Sunday, saying that "nobody has any information" about the president's birth and that "if he wasn't born in this country, he shouldn't be president of the United States."

No matter what state officials release on the issue, the "birthers" are going to question it, said Fukino. "They’re going to question the ink on which it was written or say it was fabricated," said Fukino. "The whole thing is silly."

Story: The Isikoff Files

As the top Hawaiian official in charge of state health records in 2008, when the issue of Obama's birth first arose, Fukino said she thought she had put the matter to rest. Contacted by NBC, Fukino expanded on previous public statements and made two key points when asked about Trump's recent comments.

The first is that the original so-called "long form" birth certificate — described by Hawaiian officials as a "record of live birth" — absolutely exists, located in a bound volume in a file cabinet on the first floor of the state Department of Health. Fukimo said she has personally inspected it — twice. The first time was in late October 2008, during the closing days of the presidential campaign, when the communications director for the state's then Republican governor, Linda Lingle (who appointed Fukino) asked if she could make a public statement in response to claims then circulating on the Internet that Obama was actually born in Kenya.

Before she would do so, Fukino said, she wanted to inspect the files — and did so, taking with her the state official in charge of vital records. She found the original birth record, properly numbered, half typed and half handwritten, and signed by the doctor who delivered Obama, located in the files. She then put out a public statement asserting to the document's validity. She later put out another public statement in July 2009 — after reviewing the original birth record a second time.

Story: Trump: I have ‘real doubts’ Obama was born in U.S.

"It is real, and no amount of saying it is not, is going to change that," Fukino said. Moreover, she added, her boss at the time, Lingle — who was backing John McCain for president — would presumably have to be in on any cover up since Fukino made her public comment at the governor's office's request. "Why would a Republican governor — who was stumping for the other guy — hold out on a big secret?" she asked.

Her second point — one she made repeatedly in the interview — is that the shorter, computer generated "certification of live birth" that was obtained by the Obama campaign in 2007 and has since been publicly released is the standard document that anybody requesting their birth certificate from the state of Hawaii would receive from the health department.

The document was distributed to the Obama campaign in 2007 after Obama, at the request of a campaign official, personally signed a Hawaii birth certificate request form downloaded on the Internet, according to a former campaign official who asked for anonymity. (Obama was "testy" when asked to sign the form but did so anyway to put the issue to rest, the former campaign official said. The White House has dismissed all questions about the president's birth as "fictional nonsense.")

The certification that the campaign received back —which shows that Obama was born in Honolulu at 7:24 p.m. on Aug. 4, 1961 — was based on the content of the original document in state files, Fukino said.

"What he got, everybody got," said Fukino. "He put out exactly what everybody gets when they ask for a birth certificate."

Hawaiian officials say that the certification is, in fact, only one piece of abundant evidence of Obama's birth in Hawaii. Joshua Wisch, a spokesman for the Hawaii attorney general's office, noted that a public index of vital records, available for inspection in a bound volume at the Health Department's Office of Health Status Monitoring, lists a male child named "Obama II, Barack Hussein" as having been born in the state.

In addition, as Factcheck.org and other media organizations have repeatedly pointed out, both of Honolulu's newspapers, the Honolulu Advertiser on Aug. 13, 1961, and the Honolulu Star Bulletin, on Aug. 14, 1961, both ran birth announcements listing Obama's birth on Aug. 4 of that year.

Even Fukino accepts that her comments are not likely to end the matter for the die-hard birthers. Trump and other skeptics have questioned why the original birth certificate has not been released.

But Wisch, the spokesman for the attorney general's office, said state law does not in fact permit the release of "vital records," including an original "record of live birth" — even to the individual whose birth it records.

"It's a Department of Health record and it can't be released to anybody," he said. Nor do state laws have any provision that authorizes such records to be photocopied, Wisch said. If Obama wanted to personally visit the state health department, he would be permitted to inspect his birth record, Wisch said.

But if he or anybody else wanted a copy of their birth records, they would be told to fill out the appropriate state form and receive back the same computer generated "certification of live birth" form that everybody else gets — which is exactly what Obama did four years ago.

Pre-Clovis Hunters in Texas

2011-04-09T09:37:00.000-07:00

So, humans were camped in Texas at Buttermilk Creek, making stone tools long before they knew how to make Clovis points. The “Clovis first” theory of North American archeology is now shown to be wrong, after 80 years’ dominance. Kudos to the Baylor and A&M teams, among others, who found and “pedigreed” this very important site.

The Buttermilk Creek thing is no real surprise to me and a few others. There have been reports on this topic published and debated in AAAS's peer-reviewed journal "Science" for some years now. Some of these mention other sites in way-far-south South America, some deemed as old as 26,000 years. Not everyone agrees with those dating claims, although the youngest reported date for that South American site still predates Clovis. The persistence of these not-yet-accepted reports tells me that humans settled the Americas long before the Clovis folks, probably during the height of the last ice age, if not before that.

My sort-of-educated guess is that they traveled along the coastlines by boat, living off the sea by fishing, sealing, and whaling. The remains of their camps are probably buried in the sediments deep beneath the sea, since sea levels during the ice age were as much as 480 feet lower than they are today. These sites would be widely scattered and almost impossible to find, excepting the wildest strokes of luck. (I've seen geological reports in that same journal about fossil beaches that far below sea level, and as high as 350 feet above sea level, which explains why the current evident polar ice melting trends alarm me so.)

Seeing as how the concrete and steel we use today won't reliably last 5000 years as archeological evidence, then what does constitute believable evidence of an advanced civilization (human or not) tens or hundreds or even thousands of millennia ago? As a result, we simply do not know whether our ice age ancestors were sophisticated folks with an advanced civilization of some sort.

Nor do we know that they were not!! If they had built houses and sailing ships of wood and plant fiber, how would we know? Those kinds of artifacts simply do not survive very long. Many folks who should know better have never actually faced up to that question.

And a very interesting question it is.

Radiation and Humans

2011-03-24T17:25:00.000-07:00

Even after a week, the “Chicken Little / Sky Is Falling” reporting continues on the nuclear power plant accident in Japan. This has (1) completely overshadowed the true disaster (the tsunami which devastated a far larger area at least as thoroughly as the two atomic bombings in World War 2), and (2) has induced people to panic and do stupid things (like buying potassium iodide over here in the US).

There is too much ignorance about nuclear technology and radiation characteristics, out among the population in general, and the media reporters in particular. There is also too much ratings/money chasing sensationalism among the reporters. It dis-incentivizes telling the truth, and that’s a bad thing.

As the two previous articles indicate, not all nuclear radiation is the same. There are different intensity levels and different decay times. There is also a huge dilution effect that renders a potent local release very dilute and relatively inoffensive when spread across an ocean and around the world.

Long-Term Exposure Limits

Furthermore, not all human medical responses to radiation are the same. The steady slow accumulation of general “background” radiation exposure has one set of limits, for long-term late-in-life effects. Further, these limits have evolved substantially since nuclear age began. These are depicted in figure 1. The current numbers are 5000 milli-REMS (mR) for adults in any one year, with a lifetime accumulation limit of 1000 mREM x age in years. For minors (under age 18), the limit is 500 mR per year. For fetuses (and thus pregnant women), it is also 500mR per year, with the added restriction of no more than 50 mR in any one month. (See also units of measure note below.)

This 5000 mR annual limit for adults has evolved from earlier, higher limits set when we knew a lot less about the medical effects of radiation. During World War 2 and up to 1950, it was 25,000 mR per year. Between 1950 and 1957, this was reduced to 15,000 mR accumulated in any one year. The current level was set in 1957. The minor and fetus limits are newer. Today, astronauts get to absorb the old World War 2 limit in any one mission, which really means they shouldn’t do that more than once a year.

Steady Background Radiation

These exposure limits apply to general background (steady) exposure. That radiation has a natural component and a man-made component. The natural component includes natural radioactive materials in the Earth’s crust, plus cosmic rays from space. It is there all the time, and varies widely from location to location. The average for the entire Earth is in the neighborhood of 240 mR per year, and 300 in the US. Very locally at very few locations (such as uranium mines), it can be 10 or even 100 times more.

The man-made component to background radiation includes the well-known atomic weapons testing fallout, the leakage of radiation from nuclear plant accidents, and leaks from nuclear industry processes in general. When nuclear weapons were being tested above ground 1945-1963, there was a significant radioactive fallout contribution, the nuclear total peaking in 1963 at about 15 mR per year. Since atmospheric testing ceased in 1963, the nuclear total has dropped back to about 0.5 mR per year, practically insignificant compared to the natural background, and that includes accidents like Chernobyl.

It also includes the effects of accidents like Three Mile Island (which leaked almost nothing) and the current crisis in Japan (which has leaked something, but nothing like Chernobyl). This figure alone explains why it is silly to fear radiation exposure in the US, no matter what happens in Japan.

The man-made component of background radiation also includes what amounts to fallout from the burning of coal for electric power. Radioactive elements are included in coal deposits, and these are in both the gas plume and especially the fly ash. About 1/3 (around 120 mR per year) of the total background comes from coal plants. Most folks do not know or appreciate the fact that coal plants are a greater radiation hazard than all the nuclear bomb explosions or reactor accidents that ever were, combined. But you can clearly see it in figure 2.

The other exposure isn’t really background, because not all of us are exposed to it: medical imaging that uses radiation (X-ray or nuclear radiation). This can be up to around another 100 mR per year. This wasn’t as widespread decades ago, and a lot of these technologies didn’t even exist then, so in figure 2, I only show this on the modern exposure.

Figure 1 – Allowable Exposure Limits for Background Radiation

Figure 2 – Typical Background “Then and Now”

Short-Term High-Level Exposures

Nuclear bombs and reactor accidents release localized high doses of radiation that decay over time, as depicted in figure 3. These present risks of relatively prompt casualties, as well as enhanced risks of long-term injury. The human medical response to this type of incident is simply different. The arbitrarily-chosen standard is the dose accumulated in one week (168 hours). Table 1 is an abridged description of these effects. These take the form of a list of mR per hour rates, with the total dose absorbed (not shown) being the average hourly rate multiplied by 168 hours. (These weeklong accumulations are far higher than any yearly background exposure limits, even from World War 2.)

The peak hourly rate is crudely about twice the average hourly rate. Peak hourly rates are directly comparable to hourly dose rates reported or predicted for the fallout from nuclear weapon explosions. For bombs ranging from 200 kilotons to 1 megaton yield, “very close in” (a few miles) the typical peak dose rate is around 500,000 mR per hour within 1 or 2 hours of the blast. “Far away” (dozens to hundreds of miles) the peak dose rate is closer to 500 mR/hour, something like 9 to 14 hours after the blast. Decay to far more survivable levels happens in a week or three. Local leakage exposures from the worst nuclear reactor accidents are hundreds of times less intense than that from bombs.

One Numerical Example

Say the average hourly dose rate was 179 mR/ per hour, and this was received over one week. The total absorbed dose would be some 30,000 mR. Compare that to the World War 2 dose of maximum 25,000 mR in one year. It is 5 times the modern limit of 5000 mR per year for an adult, and 50 times the modern limit for a child. Close to a major event, you can accumulate more than a year’s dose very quickly. This stuff is very dangerous, make no mistake about that.

Figure 3 – The Nature of Decaying High-Level Short-Term Exposures

Table 1 – Medical Consequences of High-Level Short-Term Exposures (Abridged) (click on table to see a larger one, browser "back" returns you to article)

Conclusions Regarding the Dangers Posed by the Japanese Reactor Accident

The workers in the plant in Japan trying to rectify the situation are at very serious risk. Adults a few miles away are at only a little risk, but infants and fetuses are at some risk. Dozens to a hundred miles away from the leaking plant, there is almost no risk at all from airborne radiation. The only credible risks would be contaminated food or water, and they’re fairly small, excepting maybe for infants.

Quite frankly, the human-made background radiation from coal plants is the more serious risk.

A Note On Units of Measure

There are specific definitions for units of radiation known as the “Roentgen”, the “rad”, and the “REM”, but quite frankly, they are all equivalent. In the literature, they are all reported as the abbreviation “R”, reflecting that interchangeability. The milli-REM ( mR) used here is 1/1000 of a REM (or rad or Roentgen). Two other common units are the Sievert (Sv) and the Gray (Gy). Effectively, 1 Sievert is 100 R, and so also is 1 Gray. The Sievert is actually 100 REM while the Gray is 100 Rad, but effectively the REM and the rad are the same (just “R”).

A Roentgen is the energy of gamma radiation contained within 1 cc of air. A rad (radiation absorbed dose) is the radiation energy transferred to some mass of material, typically for us, a human. REM (Roentgen equivalent man) is the rad dose multiplied by a quality factor that models the biological impact. For gamma radiation and beta particle radiation, the quality factor is 1 (1 rad = 1 REM).

Follow-Up on the Japan Nuclear Crisis

2011-03-17T16:47:00.000-07:00

The egregious “Chicken Little / The Sky Is Falling” reporting continues unabated over Japan’s nuclear plant crisis. I am getting very disgusted with all the fear-mongering-for-profit.

News media: please learn something about what you are talking about, before opening mouth and inserting foot. The run you have precipitated on anti-radiation iodine pills in the US is completely ridiculous, even insane.

It’s not that there aren’t serious risks at this plant, because there are. But, they have little to do with what gets splattered all over the evening news. Please refer to my previous posting on this subject (15 March 2011), and to figure 1 below.

Even if the reactor cores completely melt down, there is no risk of a Chernobyl-style intense-radiation event. Most, although maybe not all, the intensely-radioactive debris will remain within the outer (#3) containment, even if it is breached.

Breaching the #3 containment into the atmosphere to leak copious amounts of dangerous stuff is extremely unlikely in the extreme. This is for precisely the same reason that most of the fibrous “tire snot” sealing materials stay inside the tire, even after a blowout.

In any event, breaching that #3 outer containment is unlikely, because it is so very tough. This type of concrete and steel construction is pretty much the same as those concrete and steel structures still standing at ground zero of the Hiroshima and Nagasaki atom bomb explosions in 1945.

There, that should “calibrate” your gut feel for just how tough these structures really are.

Breaching the floor of the #3 containment, and releasing dangerous stuff into the earth and groundwater is more likely, but only in the event of a full core meltdown, which is still pretty unlikely, in spite of the severity of the disaster in Japan.

However, this is not atmospheric dispersion, and it stays more or less confined in a definite and geographically-small area. Radioactive fallout from the air is just not a risk more than a few miles from the plant, no matter what happens.

The real risk in this particular disaster is not the reactor and its containment structures, it is the spent fuel rod disposal ponds. These have little or no containment. If they go dry, and the spent fuel rods overheat and catch fire, that really is atmospheric dispersal of some very dangerous stuff.

That is the only thing anybody should be seriously worried about.

There are some lessons to be learned here. They are applicable to retrofits of existing nuclear plants, and to new plant designs. They are also nothing but plain old “horse sense”.

The Japanese plant was designed to withstand a Richter magnitude 8.2 earthquake, and quite successfully withstood magnitude 8.9. However, the geologic record (not the historical record) suggests that the Pacific Ring of Fire experiences quakes substantially above magnitude 9.

Not much can be done to “harden” existing plants. But, pick any number you want for design criteria for future plants. How about 9.5?

The problem in Japan wasn’t so much earthquake damage, it was the tsunami. Nobody planned for a 30-foot wave to sweep away the electrical grid and to disable all the back-up generators and water supplies. They haven’t seen a tsunami like that in Japan since 1700 AD.

Yet, the geologic record suggests that there actually have been far higher tsunami waves all around the Pacific, and in other oceans as well, including the Gulf of Mexico. Pick any number you want, but numbers on the order of 100 to 200 feet tall seem to have happened multiple times, all around the world in the geologic record. Just not within recorded history.

The other problem is the lack of multiple containment, around the spent fuel rod storage ponds. This is retro-fittable at every plant in the world today. It should be done as soon as is practical. It’s a safety thing, the “bottom line” is (and should always be) secondary to that consideration.

Figure 1 – Reactor Containments vs Spent Fuel Ponds

On the Nuclear Crisis in Japan

2011-03-15T14:17:00.000-07:00

I am very displeased at the sensationalized, inaccurate reporting of this crisis in Japan, especially from the for-profit radio and television media. This is a disservice for the citizenry, and needs to stop. Now.

We all know that only bad news sells, and the more it is exaggerated, the more likely it is folks will sit through the commercials to hear the rest of the story. Besides that, the technical inaccuracies are quite egregious at times.

When there is a problem with a nuclear reactor, the appropriate expert to consult is a real nuclear engineer, not a nuclear physicist or a nuclear plant security expert. They did not study the actual engineering design of reactor systems and equipment, only the nuclear engineer did.

These exaggerations and inaccuracies feed into public fears, which in turn lets politicians take advantage of these fears, just to further their own careers. They do this instead of doing actual good for the people (which they swore to do, but so often fail).

This sensationalized, inaccurate reporting could put a damper on efforts toward increased US use of nuclear energy, at a time when we so desperately need more of it. Some truth about what is really happening in Japan could allay public fears and let our country get on with what it must do.

I personally am no nuclear engineer, but I am an experienced mechanical engineer. Here is some of what I do know about nuclear plants.

About Radiation

First, not all “nuclear radiation” is the same. There is more than one type of radiation, and more than one intensity level.

Some reactor fuel materials and their daughter products are quite intensely radioactive, and stay that way for long periods of time. These are generally solid materials, and in modern designs, not very flammable.

Other materials, such as core structures, containment vessels, and cooling water, can be “activated” into being radioactive materials by exposure to the nuclear radiation coming from the reactor core. These are also generally solid materials, and their radioactivity is initially far less intense, and decays far more quickly into insignificance.

The most innocuous of these are steam and air, harmless within minutes, and gaseous in form, so they cannot drop as “fallout” from the sky. See figure 1.

Figure 1 - Not All Radiation Is The Same

About Heat Production

Both the fission reaction and radioactive decay create heat within the active materials. Fission creates by far the most heat, but this ends immediately when the control rods are inserted to “kill” the fission reaction.

Radioactive decay produces less heat, but it is persistent for a time , until the radioactivity decays. Depending upon the material, this can be a very long time.

Reactor fuel and daughter products typically require many, many years to decay. This is in part why spent fuel rods are placed in pools of water: to keep them cool in spite of the heat produced by radioactive decay. The other part is that the water is a shield to absorb the radiation and protect the environment.

When a reactor is shut down by inserting the control rods, its core requires considerable cooling for a time measured in days to weeks, to offset the heat of radioactive decay. There is no fission heat being produced at all after shutdown.

Modern Reactor Designs

These typically have three very stout layers of containment: the fuel rod assembly tube, the reactor vessel itself, and a containment vessel; surrounding the reactor vessel and its closely-associated equipment. Some plants place this inside an ordinary building, others do not. See figure 2.

Figure 2 - Modern Reactor Designs Have 3 Layers of Containment

In these designs, none of the reactor fuel or core materials are flammable. The fuel is metal oxide, and is contained within a tube of exotic metal alloy. The fuel rod tube material actually melts before the fuel pellets themselves melt.

Fuel assemblies plus control rods, immersed in water, constitute the “reactor core” that is contained inside the reactor vessel. Water goes in and comes out as steam, because of heat produced in the core. This steam can be used (indirectly for radiation safety purposes) to generate electricity.

The first layer of containment is the fuel rod tube itself. Only if this gets hot enough to melt, are any of the still solid (or at worst molten) fuel and daughter products able to escape the tube.

The second layer of containment is the reactor vessel, which is a very stout steel item. It would take extreme temperatures and pressures to broach this vessel.

Pressure control in the reactor vessel is by venting, which releases the slightly radioactive steam and air, and little- to-none of the fuel or daughter product and fuel assembly tube materials. For post-shutdown cooling, water is pumped through the core in this vessel, whether that core is intact or not.

The third layer of containment is an extremely strong concrete and steel shell, built around the reactor and its associated equipment. These things are built to withstand impacting aircraft, explosive attack, tornadoes, and just about anything else but a direct hit with a large nuclear bomb.

If the reactor vessel does fail, the extremely dangerous mess is still contained within this shell. Again, pressure control is by venting the relatively innocuous radioactive steam and air. The truly dangerous stuff is nongaseous, and gets almost entirely contained within.

Many of the fuel assembly tubes, and maybe one of the reactor vessels, have broken in Japan because the plant’s core cooling ability was destroyed by the tsunami. None of the containment structures have failed, nor is it reasonable to think they ever will.

There cannot, and will not, be release of a disastrous quantity of the intensely-radioactive fuel and daughter products from the crisis in Japan. For pressure control, a very tiny amount of this material will be released, aerosolized along with the steam.

There will be a much larger release of the far-less-dangerous “activated” structural materials. Taken together, the danger of the released radiation is actually fairly low, and easily decontaminated by ordinary-but-prompt showers, and simple wash-downs of hardware.

Comparison to Chernobyl ?

This comparison is unreasonable fear-mongering. The reactor designs are completely different. The simple uncontained pile reactor at Chernobyl was a 1950’s Cold War legacy that should have been dismantled decades before it exploded . See Figure 3.

Today, no responsible or ethical engineer would design a reactor like that. It has no safety features or containment, and the operating characteristics are far less stable.

At Chernobyl, they lost control of the reactor, let it get too hot, and literally caught the core materials on fire. Both the graphite block structure and metallic reactor fuel were chemically flammable in air, and that is exactly what happened.

Without any containment at all, this fire produced enormous quantities of intensely radioactive smoke directly in the air. This smoke was composed of particles of graphite, reactor fuel, and daughter products, and exposure to it caused death within hours to weeks.

The proper comparison is to Three Mile Island, which did suffer a core meltdown inside the reactor vessel. However, at Three Mile Island, neither the reactor vessel nor the containment structure were breached. The only radiation released was the relatively innocuous and short-lived radioactive steam, and that at about the level of an ordinary chest X-ray to those exposed.

Figure 3 - The Antique and Unsafe Chernobyl Design

What Really Went Wrong In Japan

What went wrong in Japan had nothing to do with reactor core or containment design. It was an unanticipated wave size for the tsunami, which destroyed the post-shutdown cooling capabilities, specifically their electric power supplies. These were necessarily located outside the containment.

If we just design for taller tsunami waves, this problem with post-shutdown cooling capabilities never happens again. Simple as that.

"How-To" for Ethanol and Blend Vehicles

2011-02-12T09:54:00.000-08:00

Some folks have contacted me for help do-it-yourself converting cars to E-85 or for using "stiff" gasohol blends in unmodified cars. Here's what I know works ----

"Ethanol VW"

My "ethanol VW" was a 1973 beetle with an essentially-stock 1600 cc case, jugs, and crank. It was a dual-port head engine, stock valves, stock rockers, no modifications at all. I had long ago replaced the worn-out combination distributor with an aftermarket Bosch 009 all-mechanical unit. (I had also long ago undone the idiot 15-degrees-retarded timing setting, that the factory used in 1973 to try to meet EPA emission standards. This took me from 5 degrees late (mark on pulley) to 10 degrees before top dead center, static timing.)

After I got the Bosch-009, what worked best for ignition timing was a very simple 30 degrees all-in (about 2500+ rpm) by strobe, on gasoline. I used this setting successfully for decades. I decreased the valve lash setting and oil change intervals to 2000 miles from the factory-recommended 3000 miles, and changed from the recommended lash of 0.006 inches intake and exhaust to 0.006 inch intake and 0.008 inch exhaust. These changes enabled me to avoid valve-burning and excess bearing wear problems in the Texas heat. By switching to aviation-grade oil in the 1980’s, I was able to increase the oil change interval to 4000 miles. After the advent of SF-or-better grade auto oils in the mid 1990’s, I was able to return to using auto oils at the longer 4000 mile interval.

The original carburetor was a Solex 34 PICT-3. I went through a couple of them; they wear around the shaft of the throttle plate, and leak air. It upsets the off-idle transition very badly. I had finally replaced it with an aftermarket Solex 30/31 with the adapter plate for the 34 mm manifold. That worked fine for many years, but finally wore out the same way as the 34 PICT-3. I also had available, but had never used, a not-worn-out Solex 30-PICT-2, off a single port head 1600 cc Bus engine.

I went through several combinations of jet sizes with both the 34 PICT-3 and the 30 PICT-2 carburetors, before I settled on the 30 PICT-2, because it did not leak air around the throttle butterfly shaft. I had to use the adapter for the 30/31 to make it fit, and the accelerator pump cover off the 30/31, to find an accelerator pump link bar that would fit. I never even tried the 30/31 with ethanol, because it has idle circuitry that I never really understood: it uses two idle jets of different sizes.

The Converted 30 PICT-2

On the 30 PICT-2 the stock main jet was a "116.0", which is 1.160 mm dia (.045"). On E-85, I settled on a "137.5" from another aftermarket Solex-Brazil carburetor, which is 1.375 mm dia (.054") for good driveability at speed. The stock idle jet was a g55, which is 0.055 mm dia (.022"). I drilled that out to .762 mm (.030") before I was satisfied with idle settings. If the idle jet is too small, you will suck the idle well dry with too much idle circuit air flow, because the idle screw is open too wide. It's a transient effect, with a time constant somewhere around 15-30 seconds. The stock accelerator discharge nozzle is right at .50 mm dia (.020"). I drilled that out to .712 mm dia (.028") before I was satisfied with the off-idle transition.

Ignition Timing Changes

None of this works at all, if you don't first revise the timing. I found that out the hard way. On gasoline with my aftermarket distributor, timing was +30 BTDC all-in at about 2500+ rpm. I set that with a strobe as the most repeatable way. I had to add 15 degrees to that setting, before it showed the same coast-down vacuum curve on E-85. On E-85, the revised timing spec was thus +45 degrees BTDC, all-in at 2500+ rpm. In other words, you need to add right about 15 crankshaft degrees to whatever timing setting you are using on gasoline. Use the minimum that recreates your old gasoline vacuum coastdown curve.

The Converted 34 PICT-3

You may still have a 34 PICT-3 carburetor. If so, here are the best jet combinations I found, before I gave up on it due to the shaft air leak. Stock main is a "127.5", which is 1.275 mm dia (.050"). On E-85, I used a drilled-out 1.57 mm dia (.062"). Stock idle was a "g55", which is 0.55 mm dia (.022"). On E-85, I drilled that out to .965 mm dia (.038"). Stock accelerator discharge was 0.50 mm dia (.020"). On E-85, I drilled that out to .965 mm dia (.038"). I give both metric and US sizes, because it's a metric car, but all I had to work with was a set of the tiny US-sized bits one uses to clean out oxyacetylene torch tips.

Heated Intake Air

The only other thing I had to do (which you might not if it never gets cold where you are) was to fool the intake air into thinking it was always summer. Any time the outside air temperature was under 70 F, I sucked my intake combustion air from a partial sheet metal glove around the muffler, made from scrap metal roofing trim. Above 70F, ambient air works fine. If you don't do this, both driveability and mileage suffer whenever it is cold. This rig worked all the way down to 15 F for me. I did it with a tee made of scrap plastic bottles on the air cleaner intake. I just plugged-up the cold inlet in cold weather, and let it draw from both inlets in warm weather. My hot source was connected to the side inlet of the tee, which has just a tad more flow resistance. Thus it favored cold air with both inlets open.

How It Should Perform

Have fun running your late-model 1600 VW beetle (or bus) on E-85. If you do it right, you should get about 80% of your former gasoline mileage, not the 70% that the fuel energy per gallon says you should get. Ethanol simply burns more efficiently than gasoline in a piston engine. This partially offsets the lower energy per gallon of the ethanol. Tailpipe soot should gradually disappear. Your spark plugs will start looking pristine-clean, too. So also will the carburetor casting look much cleaner, inside and out. It's really amazing how much cleaner E-85 is than gasoline, in so many different ways.

Minor-to-Moderate Compression Troubles, and How to Cope

If you smell ethanol in your motor oil (it'll smell different, anyway, so I am talking about detecting really serious odor here), your rings are leaking. This will show up as uneven (by around 15 psi) or generally somewhat-low (by about 15 psi) dry compression test readings. If your readings are worse than that, you really need to do the overhaul work first. Add about 10 or 20% Lucas Oil Stabilizer to your crankcase oil, and that modest compression defect will correct itself, and the fuel smell in the oil will go away.

Use the "finger test" to reset your oil change interval, it'll get substantially longer with synthetic in the mix like that (mine pretty much doubled from 4000 to at least 8000 miles, on modern SM-rated oils in an 80-20 blend). The Lucas additive really does a good job arresting cold start wear. Before the advent of SF-grade oils, I could not get even 3000 miles without seriously failing the finger test, so I used aviation-grade oils instead. Nowadays, the SM-grade(same as ILSAC-4, by the way) is way better than the aviation grade oil.

If you don’t understand how to run the “finger test”, you better ask me, or a professional mechanic. It’ll tell you everything a lab test can tell, except for a numerical particle identity and count. But, if you see visible metal wear particles, that’s all you need to know anyway (time to overhaul completely).

Carbureted “Flex-Fuel”?

My 1973 VW beetle is going back into mothballs. Before I was done with it, I reset the carburetor back to gasoline settings by installing a screw on the enlarged main jet, reset the timing back to gasoline-suitable, and undid the heated intake air. I didn’t change the enlarged idle or accelerator discharge. I just reset the idle speed and mixture screw settings as needed, to make it run just fine on gasoline. Then I ran progressively-stiffer gasohol blends until I saw the late timing problem kick-in about E-45-ish on blend strength. The car ran just fine testing blends all the way to E-57 like that. I got the same story (late timing above about E-45) from fuel mileage figures in my fuel-injected unmodified 1995 F-150, and subjectively from my fuel-injected, unmodified 1998 Nissan Sentra.

I did have to reset the VW carburetor screws a little for the blends above about E-40. The fuel-injected Ford and Nissan needed nothing at all, all the way to E-50-something (because closed-loop injection compensates mixture strength automatically, within system flow rate limits).

The requirement for the extra 15 degrees of timing advance (and presumably the warmed intake air) seems to kick-in like a light switch, right about at E-45-ish. If you don’t make these changes, then above E-45-ish, you run weak, smooth, and fuel-consumptive, with a little less intake vacuum on coastdown. That's symptomatic of late timing.

Blend Limits for Unmodified Engines

Cold-weather start "irritations" limited me to E-30 to E-35 max in the unmodified fuel-injected cars. These take the form of starting but dying quickly. A second start then usually works just fine. The problems kick in about freezing. I have tested down to about 10 F here in Texas. This is not serious, just irritating. I have a very old carbureted VW beetle (1960 model) that is running totally unmodified on E-34, and it seems to be doing OK, too. All my completely-unmodifiable lawn and garden equipment runs just fine on E-34, and has for 4 years now.

E-85 is nominally 85% by volume ethanol and 15% gasoline. Its nominal volume fraction ethanol is thus 0.85. These days, unleaded regular gasoline is nominally 10% by volume ethanol and 90% gasoline. Its nominal volume fraction ethanol is thus 0.10. If you know how many gallons of fuel it takes to fill your tank or fuel can (V), and what blend fraction ethanol you want (R), you can use these figures to compute how many of the fill gallons should be E-85 (X):

X = V / [1 + (0.85 – R) / (R – 0.10)]

Examples: for a desired E-35 blend, R = 0.35. Thus X = V/[1 + .50/.25] = V/3.00. Similarly, for R = 0.30, X = V/3.75. For R = 0.25, X = V/5.00. For R = 0.20, X = V/7.50.

Assuming the tank is burned down pretty low (or the can is nearly empty), the residuals will combine with your fill blend pretty close to the R you selected for the fill. In the case of a vehicle fuel tank, this presumes that you have calibrated your fuel gage for gallons-to-fill versus marks on the gage.

This stuff “splash-blends” right in the tank or fuel can. No mixing is required. Just put in your “X” gallons of E-85, and top-off “to the mark” with gasoline. Total gasoline added should come out very close to “V – X” if you did it right.

Calibrating a Fuel Gage

Keep a mileage log over at least three tankfuls of fuel. Record as a minimum the odometer reading and the gallons-to-fill at each fill-up. Fill the tank to exactly the same mark each time. The average mileage between fill-ups is the difference in odometer readings divided by the gallons-to-fill.

While driving on each tank of fuel, as the gage’s needle reaches each mark on the gage, record that odometer reading. The differences in these recorded readings give you miles-between-marks for that tank of fuel. Dividing those by the average mileage for that tank gives you gallons-between-marks. These you average over the multiple tanks of fuel. Listing the averaged gallons-between-marks in a cumulative fashion gives you gallons-to-fill (V) for each gage mark.

Checking Blend Strengths Experimentally

I do this with a simple added-water phase separation test. This requires lab-grade glassware, those being a graduated cylinder of 100 cc capacity for the fuel sample, and a graduated cylinder of about 30 to 50 cc capacity for the added water. You must “abuse” standard laboratory practice and read these to the nearest quarter-division instead of the standard-practice nearest half-division. If you do it this way, your results will come out pretty close to plus or minus 1 or 2 percentage points on blend strength (plus or minus 1 or 2 E-number points). Smaller sample sizes do not work out accurate enough to be useful. I draw my samples from the Schrader fitting located on the fuel rail in most fuel-injected vehicles.

Draw a fuel sample between 58 and 68 cc in volume into the larger cylinder and measure it precisely (bottom of meniscus, or BOM). Compute 1/3 of this volume for the water, put about that much into the smaller test cylinder, and measure what you have precisely (BOM). Record these numbers. Then add the water to the fuel, which will begin to phase-separate immediately. Let this stand 2-4 minutes until all the air bubbles quit decanting. Then measure the total liquid volume (BOM), and the volume below the interface between the separated layers (there is no meniscus, this is a flat plane).

Now, all the water and the ethanol go to the bottom layer, which may grade from cloudy white below to clear right at the interface. The hydrocarbon will all go to the top layer, which is a clear straw-colored liquid. You cannot use the water-plus-ethanol volume directly, because mixed ethanol and water volumes are not conserved, while mixed ethanol and hydrocarbon volumes are conserved.

Subtract the wet ethanol layer volume from the total separated sample volume to determine the hydrocarbon volume floating on top, and record it. Subtract this hydrocarbon volume from the original fuel sample volume, to determine the wet ethanol volume present in the original fuel sample, and record it. Dividing this wet ethanol volume by the fuel sample volume determines the wet ethanol fraction in the original fuel, which in percentage format is a really good estimate of the blend E-number.

I typically find the E-10 “gasoline” to be really closer to E-8; indeed, the placard on the pump usually says “up to 10% ethanol”, not “exactly 10% ethanol”. E-85 typically tests as E-87, which means there is most likely about 1-2% water in the mix. That’s not surprising, as moisture from the air readily absorbs into the ethanol in the fuel. 1 or 2% water is not a problem.

Oil Prices, Recessions, and the War

2011-02-04T16:49:00.000-08:00

By putting together facts from different sources, adding in some events from recent history, and a little common sense, one can draw some startling conclusions. These should make you as mad as they do me. It’s hard to argue with factual data. My conclusions are my own opinions. You draw conclusions for yourself, and form your own opinions.

I start with a graph of US regular gasoline price history from about 1970 to the present, adjusted for inflation, as January 2011-dollar equivalent. I got this from “zfacts.com”, which has quite the variety of both facts and opinions. Price history is fact, not opinion, however.

To this time history graph, I added several historical events, a line representing the current equivalent of 1958’s 25 cent/gallon gasoline, and a second line representing a conclusion I drew from all this data regarding recessions. That modified chart is complicated and takes a while to digest, but here it is:

I was able to discern several connections between this price history and contemporary events, as well as linkages between fuel prices and recessionary events. These are listed in bullet form on the next graphic. The most important one is in capital letters. It makes liars out of most US politicians running for office, from either party. The notion of an enormous monopoly-cartel pricing effect superposed on top of a basic supply-demand price level, makes liars out of those who claim the international oil market is nothing but a “free market”, for it most clearly is not. The scariest bullet is the very last one, however.

That brings up the question of oil supplies available. Here’s the US production history (actual data):

M. King Hubbert was the geologist who used an empirical curve fit to predict a US production peak in 1965 or 1970, back in 1956. His model takes advantage of a convenient mathematical curve shape, with no scientific causality built in, but was surprisingly accurate. He did this long before oil was discovered in Alaska. The area under a Hubbert bell curve is proportional to the volume of the resource it models. Here is what that production history looks like with some Hubbert curves superimposed:

My conclusions follow:

In my opinion this makes (1) liars out of the US politicians running for office who told us we could drill our way out of dependence on foreign oil, and (2) fools out of those who believed them.

The depletion of US oil reserves brings up the question of world oil depletion, and its effect on the basic supply-demand pricing level underneath the monopoly cartel pricing effects. The Saudis have not exceeded their own 2004 production levels. They sit on the 3 largest known remaining oil reserves left on the planet. It isn’t pretty:

But wait, some say we have tremendous reserves here at home. I see this claim quite a lot in email forwards about “the Bakken”, and some other names. These forwards always claim we have “cheap oil” in quantities exceeding Saudi Arabia, it’s just that the political opposition and/or environmentalists “won’t let us drill it”. These are just politically-motivated hit pieces. They mix facts with egregious lies and very slanted rhetoric. The truth is quite different. Bulletized list follows. I might add that the cleanup costs for the wastewater generated by the Alberta tar sands products we buy, are not in the product prices we pay, because they are as yet unknown. The volume of wastewater impounded under armed guard now exceeds Lake Erie. No one knows how to clean it up. That bill will come due. Soon.

So, as long as we use oil for fuel, we’re stuck with importing it. Most of those imports come from OPEC, dominated by middle eastern countries, some of whom are downright hostile. What do they do with all that money we have paid them for oil, for the past half century? You won’t like the answer:

We’ve been paying them to kill us. For decades. That does bring up good questions about treason.

There are a lot of entrenched interests long opposed to the implementation of alternative liquid transportation fuels, for a lot of “good-sounding” reasons. Yet there only three types of fuel to worry about, and three good drop-in alternatives available at one level or another, right now.

Gasoline can be stretched quite a bit further by blending-in significant ethanol, without any vehicle or infrastructure modifications (steel and neoprene are as good with ethanol as they are with gasoline). The only relevant questions are what source(s) do we use, and how much is available?

One of the specious arguments against ethanol is the effect of corn diversion from food use to fuel production. Many claim that the upsurge in food prices in 2009 was due to ethanol production. If you look at the actual facts, you find this claim is a lie.

To use ethanol successfully, we need the cellulosic technology that is just now being scaled-up and industrialized. This was originally made possible by grants from NREL, the alternative energy lab that has been part of DOE since Jimmy Carter’s time as president. Those first NREL grants made the industrial R&D possible, that in turn has recently led to pilot plant production of cellulosic ethanol at prices similar to gasoline, or cheaper. (That NREL/DOE story makes liars out of the authors of popular e-mail forwards claiming DOE has been a worthless waste, does it not?)

The situation is similar for using biodiesels in diesel fuel and jet fuel. The algae technology needs its development finished, so it can also be scaled up and industrialized. There’s more of it available.

OK, given that we finish the development, scale-up, and deployment of cellulosic ethanol and algae oils, what could we do with them? Remember, blend fuel products based on these materials are drop-in fuels, suitable even for the legacy fleets of old cars, old trucks, and old airplanes.

Answer: displace as much as possible of that imported oil we get from generally-hostile and financially-predatory OPEC.

If we go for E-33 and B-33 blend levels in the three fuels, this is what could happen:

The US could zero-out the imports it needs from OPEC! Wow!

If the US eliminates its dependence on OPEC oil, that is a major blow to the money OPEC funnels to the terrorists and proxy armies we fight, the US being their single largest customer by far. That has the added benefit of dropping oil prices via the supply-demand mechanism, compounding the denial of funds to the enemy.

If the rest of the western world followed suit, they and we together could dry up virtually all income to the Arab states of OPEC. Since those states have no other source of revenue (they have no other export the world wants), they would have to civilize themselves and join modern society as responsible members, or else go back to the stone age. And they know that (it is their greatest fear)!

In other words, we could win this war economically, with no more invasions or armies. And, start making our economies proof against any more oil price-induced recessions, to boot.

How to Drive on Icy Roads

2011-02-03T16:28:00.000-08:00

Roads tend to get beaten "clean" in ruts, especially in the southlands. Drive in those ruts, there's better traction there. Just slow down until you feel no hint of fishtail instability, and then another 5 or 10 mph slower. It's that simple.

Your enemy is bridges and big culverts. These tend to accumulate more and slicker ice, and they ice-up first. The sand (if any) helps only for a little while, then the wetted sand and slush re-freezes into a new and harder coating that is just as slick as plain ice, and a whole lot harder to pound loose by the passage of traffic. Never, ever assume a bridge is safe! Turn your flasher on to warn the folks behind you, that you are doing something they don't expect, and slow way down before you reach the bridge. Most cars have a maximum controllable speed on slick ice in the neighborhood of 20 mph. You need to be moving at least that slow as you reach the bridge, so you can see what's really on it, in time to respond. Plain and simple.

If the bridge has clean ruts or is clean and dry, speed back up and cross. Stay within the ruts. If the bridge is icy, stay under that 20 mph. Steer to cross the bridge in a straight line, get off the gas and brake (and stay off them) and make no steering changes while you coast across. You will come out just fine on the other side, if you do these things. You will not come out fine, if you accelerate, brake, or turn, in the slightest. For long bridges, modify this as absolutely-constant speed driving, no braking, no turning, at substantially less than 20 mph. I recommend about 10-15 mph for most cars.

For roads with paved shoulders, there is usually gravel on the shoulder and a few inches beyond. If you start fishtailing before you can slow, put two wheels into that grass and gravel right at the edge of the shoulder. Your fishtailing will stop. But, don't stay there, decelerate by coasting, and put it back onto the road, before you get stuck. Don't do this on a farm-to-market road, there are no shoulders, and the ground off the pavement is generally too soft (because of the precipitation). Just drive much slower on those roads, so that you never fishtail.

Pickup trucks and front-engine, rear-drive cars are extremely prone to rear-end-breakaway skids, because the weight distribution is very bad for all-wheel traction. You need to go much slower than a ordinary car, because once it breaks away, you are out of control, and you won't get it back until you come pretty much to rest (which might entail fetching up against something really solid). It'll warn you by feeling very unsteady, by wanting to fishtail. Find the fishtail speed for your vehicle (not in traffic, please!), and then drop at least 10 mph below that. Be aware that this speed changes as conditions change.

SUV's are famous for being able to "go" when other vehicles won't, especially the 4-wheel-drive ones. But, they do not stop any better than the worst of the conventional cars, and they are far more unstable due to the high center of gravity. I've seen more SUV's upside down in medians and bar ditches, than any other type of vehicle. Slow way down!

The vehicles with more even weight distributions front-to-back can suffer from the other type of skid: front end breakaway. That's when you crank the steering wheel to turn and nothing happens. Why? You are going too fast. Steer straight and miss the turn, coast down, and drive a lot slower. The problem will go away if you slow down. If this happens on a curve, the only thing you can try is a shallower turn. Do not brake (you will spin out), do not accelerate (same result). Very gentle steering inputs will sometimes work when a big input fails. Trouble is, most of the time, you don't have room for that. So, I recommend you try your vehicle out turning on the ice in an empty parking lot. Look for the speed at which it breaks away, and back off at least 10 mph below that. Use that reduced speed figure as your maneuvering speed out there on the icy road.

Some folks put chains on. They work, but rarely are they rated for driving more than 10 or 15 mph. You can pretty much do just as well without them, at those low speeds. Drive too fast, and they come apart. The flying fragments are steel shrapnel. They will damage your car, and they will hurt innocent bystanders. I haven't owned chains in decades. No need.

Don't forget to turn on your lights. This is as much to be seen by others, as it is for you to see better. In the fog, mist, and snow, all colors are "stealth", even reds and yellows.

And don't forget to wear your seat belts. If you don't know what you're doing, or have little practice, driving on ice, chances are actually very high you will have at least a minor accident. Could easily be a major accident. Belts make the difference between a bruised ego and a hospital stay. Or death.

If you get stuck, and you can almost but not quite "rock" your way out using forward and reverse, then try using the vehicle's floor mats. Put one, textured rubber side up, under each drive wheel. For marginal cases, that's often just enough extra traction to get free. Better to get tire tread marks and dirt on your mats, than to freeze while waiting for rescue.

Finally, dress for it! Dress like you have to walk miles in the snow and wind and cold. You very well might have to.

Fundamental Design Criteria for Alternative Space Suit Approaches

2011-01-21T13:00:00.000-08:00

This posting supports the manned Mars mission design posting of 7-25-11. The supple space suit discussed here is presumed to exist for that mission.

The focus here is the so-called mechanical counterpressure (MCP) space suit. This was a design approach explored (among other things) on the PBS program “Nova Science Now”, aired Wednesday, January 19, 2011, in Texas. This approach has its genesis in the “partial pressure suits” used by USAF in the late 1940’s, and during the 1950’s. It was tested as a possible space suit design with the elastic fabrics of the 1960’s, and in the last decade has come back under consideration again, as a possible way to vastly improve astronaut mobility and dexterity.

General Considerations:

Here on Earth, our atmosphere contains the oxygen we need to support life, and it also exerts its pressure upon our bodies. This pressure has two effects: (1) to concentrate the oxygen enough to diffuse effectively into our blood across the alveoli structures in our lungs, and (2) to keep the water in our blood and tissues from boiling away at body temperature. The minimum partial pressure of oxygen necessary for effective respiration is an indistinct limit, but it is substantially higher than the value of external pressure necessary to keep our blood and tissue moisture from boiling away. This boil-off level is pretty close to the equilibrium vapor pressure of water at a body temperature of 98.6 F (37.0 C), which is 0.06192 of a standard atmosphere.

In space, there is no oxygen and there is no pressure. Unprotected persons die quickly, first losing consciousness in seconds to a couple of minutes, due to anoxia, then suffering anoxic brain death within a very few more minutes (around 10), which event is generally irreversible. Some several minutes after the onset of brain death, the heart stops and blood pressure falls below the moisture boil-off level. The water in the blood and tissues begins to boil away into space, breaking open cell membranes and splitting open tissue structures. Boil-off of this moisture draws heat from the surrounding tissue structures, chilling the body rapidly toward freezing as it partially desiccates.

Thus, a protective enclosure is necessary for us to survive in space, one which provides oxygen at a suitable pressure inside the lungs and breathing passages. This breathing gas pressure must be balanced by the same fluid pressure within the body, in turn produced by an equal pressure applied to the external surface of the body. An important fact: this external pressure can be supplied in two ways: (1) by gas or fluid pressure within a sealed garment that is essentially a balloon, or (2) mechanical pressure applied directly to (and distributed over) the skin.

Historical Specifics:

By about 1930, high-flying pilots had to be protected from lack of oxygen. One way was a modified deep-sea diver’s dress, functioning as a pressurized balloon. The atmosphere inside the sealed garment could be pure oxygen instead of air, and its pressure need only be enough to concentrate the oxygen in the lungs sufficient to support respiration. Such pressures are typically well above the moisture boil-off level. This balance of breathing gas vs internal tissue fluid pressures, with a distributed external pressure creating those internal fluid pressures, is illustrated in the right hand portion of figure 1. The creation of those internal tissue fluid pressures is very much like squeezing a water balloon all over, raising its internal pressure, as in the center portion. This works because the cells of the body are essentially tiny water balloons. Thus, in the aggregate, the body responds like the water balloon.

Figure 1 – Pressure Balance Physics with Fluid-Filled Objects and Aggregates of Same

Military pilots flying “high-gee” maneuvers tend to faint, because the blood supplying oxygen to their brains is pulled by the high accelerations toward their lower extremities (pooling in their legs). Hypoxia thus induced soon leads to blackout, often within a very few seconds. The solution is to drive that blood supply upward against the acceleration by means of the “gee suit”. These garments provide a mechanical squeezing action on the lower body and legs, similar to squeezing just part of a water balloon. This uneven compression drives the fluid within toward the uncompressed portion, instead of generally increased internal pressure. This action is illustrated in the left portion of figure 1. In the water balloon, the unsqueezed portion expands with the extra water driven there. In the body, blood that would have pooled in the legs is driven upward against the acceleration back to the brain.

In the late 1940’s, this “gee-suit” action was proposed as a temporary protection garment for test and fighter pilots having to bail out or deal with loss of cabin pressure at extreme altitudes. By this time, it was already known that an ordinary oxygen mask could not supply a sufficient concentration of oxygen, because the atmospheric external pressure governing that concentration was too low. (The critical altitude for that is a little “fuzzy”, but generally we use 45,000 feet.) Pressure breathing gear was required for flying higher, and this required a counterbalancing internal tissue fluid pressure within the body (as illustrated in the right-hand portion of figure 1). The balloon-type pressure suits of that time were simply too restrictive of movement, too bulky, and too heavy, to serve in this application.

An extension of the gee suit mechanical compression approach provided the answer used in the late 1940’s and 1950’s: the so-called “partial pressure suit”. The mechanical compression was extended over most of the body, providing internal fluid pressures more or less sufficient to balance breathing gas pressures in a helmet, in turn adequate to support life long enough to bail out from around 100,000-foot altitudes. It wasn’t perfect: hands and feet remained uncompressed, and the actual compression achieved over the torso and limbs was very uneven. Blood pooling within inadequately compressed limbs would lead to the pilot blacking out, in about 10 minutes or so, and serious swelling from edema within about an hour. But, it worked well enough to serve for the few minutes of a bailout or emergency descent. Compression was achieved by tensioning the non-elastic fabric across the skin by inflatable “capstans” (tubes). These suits were far less restrictive of movement, far less bulky, and far lighter than the gas balloon-type pressure suits of that time.

By the end of the 1950’s, the gas balloon suits had been sufficiently improved to be competitive with the partial pressure suit in terms of bulk, weight, and movement restrictions. These newer gas balloon suits (“full pressure suits”) had no restriction on protection time, as an even distribution of compression was inherently achieved on all body parts. This eliminates blood pooling problems. Thus, these were chosen as the space suits of the 1960’s, and have been the standard ever since. As evolved for in-space and lunar use since then, these have become very bulky, restrictive garments. Cooling systems are required inside the hermetically-sealed suit. Due to the stiffness and bulk of the pressurized gloves, manual dexterity is very limited. Typically, the suit pressure used is 1/3 of an atmosphere of pure oxygen, near 253 mm Hg.

In the late 1960’s, the partial pressure suit problems of uneven compression distributions and limited coverage were addressed fairly successfully by the use of elastic fabrics. By mechanically compressing the hands and feet as well as the limbs and torso, time-unlimited protection could be had. Such garments need not be one piece, as they were not gas-tight balloons, merely the equivalent of tight panty hose or shrink-wrap. Tailoring the distribution and arrangement of layers of elastic fabric in the various garment sections was the hard part. No cooling system was required: the wearer could sweat right through the porous garment into vacuum. Such a space suit was demonstrated as a prototype in the vacuum tank successfully, under the direction of Dr. Paul Webb, and partly funded by NASA. Including the oxygen supply and helmet, it weighed 85 pounds, with little movement restriction, and marvelous dexterity. Compare that with the 200+ pound full pressure suits used on the moon! The breathing pressure was in the neighborhood of 170-190 mm Hg. These efforts did not lead to application then.

The mechanical compression idea has resurfaced in the last decade, including efforts for NASA. The modern ability to tailor elastic fabrics is even better than that available in the late 1960’s. Up to a certain level of mechanical compression, these techniques now work fine. The level currently achievable is not 1/3 of an atmosphere, however. But, is that level really necessary, considering that the experiments of the late 1960’s were successful with much lower compression? This is important, because if one specifies a compression level higher than can be reached with the technology, this MCP technique could be deemed infeasible, when in fact it is feasible, and offers some very significant advantages.

Design Criteria for Breathing Pressures:

A good startpoint is the concentration of oxygen available to people at sea level. The standard oxygen content of dry air is generally thought to be 20.946 volume percent (v%). One standard atmosphere’s pressure is defined as 14.696 psia, 1013.25 mbar, 29.92 inch Hg (mercury manometer column height), or 760 mm Hg. The gas laws indicate that partial pressure percentages are the same as volume percent composition. See figure 2. At sea level, the partial pressure of oxygen in dry air is then 159.2 mm Hg.

Inside the moist lungs the air is no longer dry. A fairly realistic assumption is that the water vapor is saturated. Water vapor pressure is determined at equilibrium by the temperature of the liquid phase in contact with it, in this case, body temperature of a human (98.6 F or 37.0 C). From the standard steam tables, this vapor pressure is 47.1 mm Hg. That vapor displaces some of the dry air, so that the partial pressures of the dry air and the water vapor now add to the imposed atmospheric pressure, in this case sea level (760 mm Hg). The oxygen partial pressure in the wetted air freshly inhaled into the lungs is then 20.946% of the dry air partial pressure (712.9 mm Hg), or about 149.3 in Hg.

Between inhale and exhale, some of this oxygen is diffused across the alveoli into the blood, and some carbon dioxide diffuses from the blood back into the air in the lungs. Compared to the water vapor displacement effect, these transient effects are small, and are ignored here.

What drives diffusion of oxygen into the blood is the partial pressure of oxygen in the wet air inside the lungs. This must be larger than the dissolved oxygen pressure in the blood for diffusion to occur at a useful rate. Therefore, there is a lower limit to the in-lung partial pressure of oxygen, but it is a little “fuzzy”, and must be determined empirically. The calculation of in-lung oxygen concentrations from atmospheric air at altitude pressure, and their relationship to pure oxygen breathing pressures in a suit, is illustrated in figure 2. It should be noted that the water vapor pressure is a constant set by body temperature, having a larger percentage effect at high-altitude lower atmosphere pressures, and at lower suit oxygen pressures.

Figure 2 – In-Lung Oxygen Estimates from Air at Altitude and In-Suit Oxygen

One estimate of the lower limit for in-lung oxygen concentration comes from in-flight oxygen rules for pilots. Civilian pilots are not required to use oxygen below 10,000 feet altitude under US FAA rules (the USN uses a different figure: 5000 feet). Most other agencies use something like the 10,000 foot rule. Using the procedure outlined in figure 2 at sea level and at 10,000 feet atmosphere pressures produces the results bordered in green in figure 3. For people adapted to more-or-less sea level air, in-lung oxygen partial pressures in the range 149.3 down to 99.6 mm Hg seem quite adequate. These correspond to altitudes in air from sea level up to 10,000 feet. They also correspond to oxygen breathing gas and suit compression pressure levels of 196.4 down to 146.7 mm Hg. These are 25.8 down to 19.3% of a standard atmosphere as the suit compression levels required.

Using the current full pressure suit standard of 1/3 atmosphere pure oxygen (253.3 mm Hg), one has a wet in-lung oxygen of partial pressure of 206.3 mm Hg, substantially more than in sea level air. These are the data bordered in blue in figure 3. They do not correspond to a calculated altitude, which would be far below sea level. Clearly, 33% of a standard atmosphere is an over-stringent pressurization requirement for the MCP suit. Something closer to 25% of an atmosphere is equivalent to sea level air, and serves well as an upper limit for MCP compression design.

Figure 3 – Suit Compression Levels Corresponding to In-Lung Oxygen at Various Altitudes

There are populations of humans who live at very high altitudes. Substantial numbers of people live near 15,000 feet , corresponding to in-lung oxygen near 80 mm Hg, and further corresponding to a suit compression level of 127 mm Hg (16.7% of an atmosphere). However these people are acclimatized to these conditions. It takes substantial time for the body to so acclimatize. Without that acclimatization, altitude sickness and fainting are high risks. There are a few people living near 20,000 feet in the high Andes or Himalayas. That altitude corresponds to in-lung oxygen 63.4 mm Hg, and suit compression 110 mm Hg (14.5% of an atmosphere). The 15,000 and 20,000 foot calculated data are included in the figure 3 table bordered yellow. “Flatlanders” in space suits at those pressures would survive, but would not be functional. Therefore, the 10,000 foot data make a pretty good empirical lower limit for MCP suit compression design.

Further, consider the nominal 45,000 foot requirement for something better than a simple oxygen mask. Those data are included as one of two entries in the red-bordered portion of figure 3. At that altitude with pure oxygen in a simple mask, in-lung oxygen is 64 mm Hg, and external compression is 111 mm Hg (14.6% of an atmosphere). Those levels correspond very closely with the in-lung oxygen levels of long-acclimatized mountain folks living in the open air near 20,000 feet. That is a good rough estimate of an extreme lower compression limit to stave off certain slow death by hypoxia.

The other entry in the red-bordered portion of figure 3 is the moisture boil-off point (where the blood starts to boil, and tissues outgas water vapor as they begin to desiccate and freeze). There is no oxygen at all in the lungs at this pressure; it is all water vapor at 47 mm Hg, which has to be the external compression pressure. Those figures correspond to 64,000 feet altitude and 6.2% of an atmosphere. About 60,000 feet has long been thought to be the short-exposure “deathpoint” for risk of blood boiling.

Conclusions:

The minimum design compression for an MCP-type space suit is very likely near 19 or 20% of a standard atmosphere (147 mm Hg), corresponding very closely with the in-lung wet oxygen partial pressure of 99.6 mm Hg experienced in the open air at 10,000 feet. Compressing less risks dysfunction.

The maximum necessary design compression for an MCP-type space suit is near 25 or 26% of a standard atmosphere (196 mm Hg), corresponding very closely with the in-lung wet oxygen partial pressure of 149 mm Hg experienced in the open air at sea level. It does not hurt to compress more.

The “typical” space suit design standard of 33% of an atmosphere seems to be an unnecessarily stringent design requirement, especially if fabric technology cannot quite achieve it right now.

On the Shooting Rampage in Tucson

2011-01-13T08:39:00.000-08:00

I haven’t weighed in earlier on this event, because I wanted to find out the facts. They are that a troubled young man descended into madness, got a gun, killed 6 people, and wounded several others before being subdued.

The President in his speech in Tucson was quite correct. We in this country need to reunify and work together in a way that would make those lost proud of us. It is wrong to try to make political hay out of this by assigning ideological “blame”, and that applies to both sides. Please tone down the politics.

There were several missed opportunities to stop this. This young man’s high school classmates knew he was troubled, yet he got no help. His classmates, teachers, and administrators at his junior college also knew, yet nothing was done.

The army would not take him because of one of his troubles, that being drug use. No one did anything in response to the bizarre postings he put on the internet. The signs were there! We’ve seen this before, including that demon-haunted look in those eyes!

The final failure was the background check when he bought the gun. It is only illegal to sell a gun to a crazy person when a court has judged him crazy. The store personnel knew this young man was crazy, but the background check said he had never been so judged by a court. So, they could not refuse to sell on that basis.

And that is one of the two weak points in the background check process, the other being no check with unlicensed dealers. So, only those two items are what we should address, with regard to gun laws. One (no check) is very straightforward to fix.

The other (insanity) is not. A court judgment of insanity is sometimes not timely enough. Reasonable suspicion of insanity should be enough to delay and investigate further. The lawyers will have to sort out exactly how that should be done, but that is their job, and they need to go and do it.

The existing federal law is correct in that felons, mentally ill persons, and illegal immigrants should not be allowed firearms. The states are correct using this law as the minimum standard about which their own laws are written to reflect their local cultures. I have no problem with that concept at all.

The missing piece in our culture is actually more important to address. This young man spent all the years from childhood descending into madness, accelerating greatly this last year. Many noticed, but none reported, or stepped forward to help him.

The result of that cultural failure is 6 murdered including a child, several wounded, a federal assassination attempt, and a national tragedy that could have been prevented. We are too isolated from each other in our local communities.

I have no solution for that, but I know we need one.

Update to Manned Mars Mission Concept

2011-01-08T15:30:00.000-08:00

(this is an update for the 12-20-2010 post on a Mars mission concept)

Summary:

I found three mistakes in the data used for the “Feasibility of a Manned Mars Exploration Mission Concept” study, dated 12-15-10. These were (1) an incorrect payload capacity for Falcon-9 to LEO, (2) an over-estimate of the minimum delta-vee required for Hohmann orbit transfer to Mars, and (3) an over-estimate of the weight of a Dragon capsule to be used for crew return. The effects of these turned out to be minor, but are corrected herein.

One other item was investigated: the orbital maneuvering capability of the existing Dragon capsule was greater than expected, especially if extra thruster fuel is carried in the unpressurized cargo section. This eliminated the need for an extra service module to handle the delta-vee needed for getting the correct entry trajectory from an emergency free return / loss of vehicle scenario. The effect of this, plus error (3) above, reduced the weight, development needs, and cost of the manned vehicle a little.

The effect of the Hohmann delta-vee error would have been to reduce the size and cost of the three unmanned vehicles, except that these also need significant orbital plane change capabilities at both Earth and Mars. Adding in the plane change capability put the required delta-vee very close to the conservative estimate used in the original study. Thus, these vehicles did not change.

The effect of the Falcon-9 payload capacity error was taken care of by the more in-depth look at the real weights and delta-vee capability of the Dragon capsule, which allowed the manned vehicle to go from a 14 tank configuration in the original study to a 12 tank configuration reported here. The only Falcon-9 launches were to put the two Dragons in orbit for vehicle assembly.

One other item was looked at, but without effect on weights or launch costs. This was an allocation of what functions would be contained in each of the three habitat modules. This was with due regard for a radiation protection by packaging of water and wastewater tanks around the command “deck”, plus a little steel plate. A portion of this module would have several low-diameter long tanks disposed around the inside periphery, with a steel protection shell inside that. The portion devoted to science would have none. The other two modules are much more open inside, having neither tanks nor steel shell.

Falcon-9 Payload:

I do not know how I misread the LEO payload capacity of Falcon-9 from Canaveral on the Spacex website, unless I looked at pounds and thought I was reading kilograms. Perhaps I did, but in any event, the corrected chart is figure 1 below. Having the correct payload capacity gave me a very good estimate of the weight statement for the Dragon capsule, which turned out to be much lighter than I had been assuming in the original study. This effect is what lightened the payload of the manned vehicle, and reduced the number of propellant tanks used in its assembly.

Fig. 1 – Corrected LEO Payload Data from the Spacex Web Site

It should be noted that most of the modules in these concept vehicle designs are in the 30-32 metric ton weight range, launched with Falcon-9-heavy. Thus, most of the data for the two designs are unaffected by the errors. It was the two Dragons used as crew return vehicles that are launched by Falcon-9’s, and this did not change in the count of launches made. The empty engine shells for the two gas core engines were estimated at around half a metric ton each, and launched by two Falcon-1’s, also unchanged.

Dragon Capsule Evaluation:

I assumed that a fully-loaded Dragon would “saturate” the payload capacity of a Falcon-9 from Canaveral. Allowing a little off for the aeroshell on its nose, that puts the fully loaded capsule right at 10 metric tons. According to the Spacex website, this capsule is capable of carrying 6 metric tons up to LEO, and returning with 3 metric tons on board. The website also indicates there is 10 cubic meters storage volume in the pressurized capsule, and 14 cubic meters in the unpressurized module. There is 1290 kg worth of thruster fuel and oxidizer in the capsule for its Draco thrusters, identified as monomethyl hydrazine and nitrogen tetroxide. I assumed a specific impulse of about 330 seconds for these, that being about 98% of the vacuum value commonly reported for this propellant combination.

I then ratioed masses for the pressurized and unpressurized cargo spaces by the available volumes. The idea was to put extra thruster propellants in the unpressurized section, and connect it up directly to the thruster system, which would be the only modification to a “stock” Dragon. The pressurized section could carry up extra thruster propellants for the landers. I assumed 10% inerts for these extra thruster tanks. The results are given in figure 2 below.

Figure 2 – Dragon with Extra Thruster Fuel as Emergency Free Return with Crew-of-6

The results were astonishing. Using all of the “stock” on-board propellant as delta-vee fuel (not possible, attitude must be controlled), the capsule has a little under 1 km/second capability. But with 6000 kg of extra thruster fuel tanks in the unpressurized section, the capsule has near 2 km/sec delta-vee, even with 6 suited crew and minimal life support supplies on board. Intuition suggests this is probably sufficient to adjust trajectory for a safe emergency reentry if the main propulsion on the manned vehicle fails. The only remaining question is whether the Dragon’s heat shield would survive the 29 km/sec reentry for the fast trajectory in this concept.

Effects on Manned Vehicle Design:

As stated in the summary, I made no changes to the unmanned vehicle design. For the manned vehicle, the reduced weight estimates for the two Dragon crew return capsules allowed reducing the vehicle stack by two common hydrogen tank modules. The revised vehicle stack is depicted in figure 3 below, and the revised performance plots in figure 4 below.

Fig. 3 – Revised Manned Vehicle Reduces to 526 m.Tons in LEO, from 619 m.Tons

Fig. 4 – Performance Estimates for Revised Manned Vehicle Reduce to 12 Tanks From 14

Extra Details on Habitat Module Allocations:

A lot of stored supplies (water, oxygen, and food) are required for a 9 month mission. I assumed these would be stored in the module used as a crew sleeping quarters, because the sleeping berths for 6 persons are actually quite small. Two restroom/bathing facilities (rendundancy) could be located in the same module, as these are also rather small volumes. Some of the water and wastewater tankage is here, and must be connected with the rest in the command module.

Intuition suggests that the radiation-shielded command “deck” and science stations would pretty well fill another module. This is because the command deck is surrounded by a “carpet” of water and wastewater tanks around the inside periphery of the module for radiation shielding in the event of a solar flare. This shielding effect is augmented by some thin steel plate inside the tanks. Enough volume is shielded to contain the entire 6-person crew plus one restroom facility and about 3 days of supplies. The science stations are unshielded. The purpose in radiation-shielding the command “deck” is so that the ship may be maneuvered, even during a radiation storm event.

The third module is the common room living space plus galley. Recreation and exercise activities are conducted here. This is the big open space, complete with substantial outside views. I put this in the middle, with the command module forward, fitted for docking with the two crew return capsules. The crew dorm / storage module is aft, and docks directly to the stack of common propellant modules. See figure 5 for the general layout.

Fig. 5 – Habitat for 6 for Up To 1 Year, Functions Allocated Among 3 Modules

Effects Upon Overall Mission Design:

There were no effects upon the design of the three unmanned vehicles, because orbital plane changes add to minimum delta-vee capability in such a way that my original estimate was ballpark correct.

There was a payload mass reduction in the manned vehicle such that two propellant tanks could be deleted, reducing the total launch requirement by two Falcon-9-heavy launchers. The overall mission is summarized in figure 6 below.

Fig. 6 – Mission Characteristics Summary, As Revised

It should be noted that virtually everything about this mission could be launched by Spacex rockets available now or within about 5 years, and that none of these are so-called “heavy-lift” launcher developments. There are only two rocket technologies assumed for the mission that not currently available: the solid core and gas core nuclear thermal rockets. The solid core assumed here is an update to the old NERVA devices tested quite successfully as late as 1973. The gas core is assumed based on projected characteristics of an engine design that had come within about 2 years of first article test, when the entire nuclear rocket project was cancelled about 1972. These things could be rapidly recreated, and the gas core engine finished.

The only other serious technology lack would relate to the survivability of the Dragon and its crew on a very high speed return trip emergency. The average velocity is near 29 km/sec on the 75 day trip. This is well beyond a Hohmann free return from Mars, which would be near 15 km/sec. There would be concerns about both the heat shield, and deceleration gees during reentry. However, intuition suggests both problems are solvable with minimal impact to the basic capsule design.

Concluding Remarks:

Total direct launch costs reduce by two Falcon-9-heavy from the original study, down to $7.8 B.

There would be technology development effort costs for bringing solid core nuclear rockets back “on-line” with a very minor update. These should be minimal if a lean, mission-focused contractor can be retained.

There would be significant technology development effort costs for finishing the development of the gas core nuclear rocket. These might be of significant size, even with a lean, mission-focused contractor. This effort will also likely be the pacing programmatic element for overall project schedule.

There would be a vehicle development effort for the single-stage nuclear lander, aside from the solid core nuclear engine that powers it. This effort would very likely cost far less than the solid core nuclear engine effort, and would not pace the overall schedule.

There would be a very minor effort to reconfigure the Dragon for use as a Mars crew return vehicle of a fast-trajectory mission. The issues to resolve are: integration of extra thruster fuel in unpressurized storage, and upgrade of heat shield and entry trajectory for 29 km/sec entry speeds. Spacex itself could most effectively do this, being the manufacturer.

Intuition still suggests that this could be done for around $12-15 B, if lean contractors are retained, pretty much as suggested in the original study. I would hazard a wild guess at 10 years to readiness, paced by the gas core engine effort.

Feasibility of a Manned Mars Exploration Mission Concept

2010-12-20T15:53:00.000-08:00

(see also 1-11-2011 post for an update to this study)

Summary:

A manned mission to Mars was investigated for feasibility. The objective was significant exploration, not a single Apollo—style “stunt” landing. It seemed insane to go to all the trouble and expense of sending men to Mars, and not visit several different sites. It also seemed insane to launch so much equipment and not try to reuse it. This consideration eliminates very small vehicle designs.

Earth orbit rendezvous and Mars orbit rendezvous were combined into a single mission design to save weight. Unmanned assets were sent by Hohmann transfer ahead of a fast-trip manned vehicle to save weight. Self-rescue or escape capability was designed-into every mission phase as much as possible. The landers and the manned vehicle were designed as reusable single-stage items.

The unmanned assets were sent single-stage one-way to Mars orbit using the lander propulsion to save weight. Lander and propellant tank assets were left in Mars orbit for refueling and re-use by subsequent missions. The manned vehicle was recovered in Earth orbit for refueling and reuse.

The manned mission time was under 1 year, eliminating the need for voluminous and heavy artificial gravity by spin (because the only requirements currently understood are 1 gee at 4 rpm). The flight deck in the habitat module was assumed to be radiation-shielded against solar flares by water and wastewater tanks, plus a little steel plate. Crew size was 6. Every component was small enough to be launched by a Falcon-9-heavy booster, or smaller. But, the interior volume is comparable to the old Skylab station, thus promising effective alleviation of long-confinement psychological issues.

All assumed propulsion was nuclear thermal rocket (NTR). The lander engines were assumed to be slight updates to the old NERVA solid core technology last tested in 1973. The manned vehicle engine was assumed to be a radiator-cooled gas core NTR, approximating a design that came within about 2 years of first article test in 1972, when all such work was stopped. This old design’s Isp was 6000 sec.

Only launch costs were estimated, as near $8 billion in today’s dollars for 16 landings during the one mission. A trade study evaluated cost reductions available for reducing landings per mission. Programmatic costs, and technology development and hardware production costs, were not estimated. It is thought that getting the gas core NTR technology ready would be the pacing schedule item for such a project. These estimates are only rough hand-calculations done by pencil and paper. They are just good enough to demonstrate feasibility, and to serve as a startpoint for more sophisticated analyses.

Basic Mission Design Approaches, Constraints, And Assumptions:

Fig. 1 (below) shows a list of the fundamental considerations. Top-of-the-list is to design-into every aspect a “way out”, meaning escape or self rescue, if in any way possible. This is the very essence of “man-rating”.

The next two items relate to flying the manned portion fast, in order to cut travel times under a year and allow deletion of the need for artificial gravity, in accordance with experience obtained on the International Space Station (ISS). Mission times exceeding a year require artificial gravity, because we have nothing but indirect (surrogate) data such as bed-rest experiments to suggest otherwise. It is unethical to risk the lives and health of astronauts on indirect data, even if they are willing. Artificial gravity by spin has to be designed for one full gee at no more than about 4 rpm, based on what we currently know by direct experiment. This leads to large, heavy, and costly vehicle designs. Choosing instead more advanced propulsion for the fast trip is very likely the easier, surer, course.

The fourth item says to do both Earth orbit rendezvous and Mars orbit rendezvous in order to save weight, plus obtain further weight savings by sending unmanned vehicles on min energy trajectories. The lander vehicles and the propellant supply supporting their operation can be sent ahead of the manned vehicle, which would then rendezvous with these supplies at Mars. For crew safety, the propellant supply for the return voyage cannot be sent this way, even though it would save considerable weight to do so. This is because the crew would be stranded, and would die, if rendezvous at Mars failed for any reason. If the manned vehicle has sufficient propellant on board to return, this outcome is avoided.

It makes little sense to go to the trouble of sending men to Mars, and not do some serious exploration. This mission study is based on doing 16 separate landings at widely-separated sites, of up to a week each, while the manned vehicle is there. This increases enormously the information return from the mission, a sort of “shotgun-pattern” planetary survey. It might even be possible to begin planting “prospecting” bases on the next mission, instead of further initial exploration.

This multiple-landing plan is subject to some safety constraints, as the last item indicates. The concept explored in this study is that 3 of the crew visit the surface, monitored from orbit by the other 3 crew. There must be at least one ready lander in orbit, in order to perform a rescue landing, if need be. This rule would terminate the mission upon a lander becoming unserviceable, unless at least three landers are sent to Mars. To minimize crew risks, any rescue landing is piloted by a single crewperson.

To make such a plan work, the landers have to be reusable, so that only three need be sent. To keep from sending lower stages, the landers must be single stage. That means they must be nuclear. As shown in fig. 2, the idea is to separate the lander propellant supply into three parts corresponding to the three landers, and send these to Mars using the lander engines themselves. These assets would remain in orbit at Mars to be refueled and reused by subsequent missions.

Fig. 1 -- Basic Mission Criteria

Fig. 2 – Sending Landers and Lander Propellant Unmanned One-Way to Mars

Fig.3 shows the thinking behind the design of the fast-trip manned vehicle. There are serious considerations for radiation sheltering during solar flare events. In the habitat module, there must be a space surrounded by water and wastewater tanks, and a little steel plate, in which the crew of 6 could shelter during a radiation storm. Prudence dictates that this be the vehicle’s command deck as well, so that critical mission maneuvers can be conducted, storm notwithstanding. The mission should take no more than about 9 months. The vehicle should be stocked with over a year’s supplies, “just in case”.

One of the requirements often soft-pedaled or ignored is volume of space available per crewmember. The psychological impact of prolonged confinement in tight spaces is a very real danger, one that can be confirmed by any prisoner who spent time in solitary. Most crew habitat designs I have seen provide about the same space as was in the Apollo capsule, which is about like a modest bedroom closet per man. That is simply not enough. The space should be comparable to that available to a family of 4 in a small (1200 square foot) house, ideally. The old Skylab space station, at 90 tons, comes pretty close to the size of habitat that is needed. It provides the baseline for this study.

Fig. 3 also shows two main engines, needed for redundancy, and a round trip propellant supply, needed in case rendezvous fails at Mars. There are two crew return capsules, each large enough to carry all 6 crew, but twinned for redundancy. These need to carried along, in case maneuver propulsion fails on the return voyage. In that event, a free return must be attempted in the capsules as the vehicle flies by Earth. These capsules must have enough delta-vee capability to “hit” an acceptable reentry corridor. That means they probably need a small propulsive service module or supply.

Fig. 3 – Safety and Design Considerations for the Manned Vehicle

Fig. 4 shows some practical launch vehicle constraints for assembly of the Mars mission in Earth orbit. Many different items could be selected, these simply correspond to a family of very cost-effective launch rockets. Data were taken directly from the Spacex website, as it exists at the time of this writing. The biggest impact is the size of objects to be launched. Payload mass is more important than fitting within the “factory stock” payload shroud. A lot of the Mars mission components could in fact ride “naked” on top of the launch rocket.

Fig. 4 – Launch Rocket Constraints on Mars Mission Component Designs

Rough Mission Delta-Vee Estimates:

Two scenarios needed investigation: a basic min-energy Hohmann transfer ellipse for the unmanned vehicles, and an “almost straight-line shot” fast-trip high-energy trajectory for the manned vehicle. Of these, determining a realistic delta-vee requirement for the fast trip is actually easier. One simply divides a representative straight-line path length (in this case about 100 million kilometers (km) by a tolerable trip time (for this analysis about 75 days). Assuming impulsive delta-vee events at each end of the trip, one obtains a nearly square-wave velocity trace vs range, because at these speeds, the sun’s gravitational deceleration or acceleration effects are negligible, as is path curvature. See fig. 5.

Fig. 5 – Fast-Trip Scenario Approximation

The average velocity over this trace is very nearly the delta-vee value required to start it, and also to end it. Thus, twice the average velocity is pretty close to the one-way delta-vee requirement for the trip. For a two-way trip, one doubles this again, to about 4 times the average velocity. For these numbers, the two-way fast trip delta-vee requirement is a very demanding 61.72 km/sec. For practical single-stage vehicles, this corresponds to specific impulse (Isp) requirements closer to 6000-7000 sec than the 900-1000 sec of a NERVA-type solid core NTR. Hence the selection of gas-core NTR technology with a waste heat radiator to effect engine cooling.

For the unmanned vehicles making a one-way trip by Hohmann ellipse transfer, the estimate is made by classical orbital mechanics methods. Calculations were made for the average orbital velocities of Earth and Mars around the sun. An ellipse was fitted between the average distances of Earth and Mars from the sun, and its perihelion and apohelion velocities calculated. Escape velocities were calculated for Earth and Mars, and circular orbit velocities calculated for a low Earth orbit (LEO) altitude of 300 km, and for a low Mars orbit (LMO) altitude of 200 km. At Earth, the delta-vee to escape on a trajectory to Mars was estimated as the difference between escape and circular velocities, added to the difference between transfer perihelion and Earth orbital velocities. The delta-vee to capture at Mars was calculated as the difference between apohelion and Mars orbital velocities, added to the difference between Mars escape and circular orbit velocities. This is probably over-conservative. See fig. 6.

Fig. 6 – Rough Estimate for Hohmann Delta-Vee Requirements

The planes of the transfer ellipse and the straight-line “shot” are more or less in the plane of the ecliptic. Thus the plane of LMO achieved this way will be inclined relative to Mars’s equator, by around 25 degrees. The landers must have some amount of plane change capability, in addition to the delta-vee necessary to land without aerobraking (in the extremely thin “air”). The surface circular orbit velocity is larger than that at 200 km, and makes a good rough estimate of minimum delta-vee. Factoring this value up by about 1.10, accounts roughly for gravity and drag losses. The plane change requirement is figured from an isosceles triangle on the surface circular orbit velocity. These velocity increments are summed for the one-way delta-vee requirement, and doubled for the two-way trip. A maximum plane change requirement of 30 degrees was assumed arbitrarily. By judicious choice from an inclined orbit, this capability brings the majority of Mars’s surface within reach of the landers. See fig. 7 below.

This 11.52 km/s value is the maximum. Not all sites require a 30 degree plane change. The minimum is no plane change at all, for the much smaller two-way delta-vee requirement of about 7.84 km/sec. Such missions have a substantially-smaller propellant “burn”. Such propellant savings, plus the very capable nuclear engine, would very likely make an orbital mission to Phobos possible during this same exploration mission, using the same equipment.

Fig. 7 – Rough-Estimate of Lander Delta-Vee Requirement

Lander Rough-Out:

The basic layout of the lander is propellant tank-as-airframe, topped by some sort of command cabin big enough for 3, and equipped with long landing legs disposed around the nuclear rocket engine. The width of the footprint should be comparable to the overall length of the vehicle for stability, so this vehicle is rather “squat” in its proportions. There should be some sort of deployable crane arm to provide a hoist to the surface. There should also be some sort of deployable solar panels to augment fuel cell electrical power. Given a heavy solid core engine and extensive landing leg structures, an inert fraction of 20% seems reasonable to assume. Combined with a propellant fraction of 70% and a payload fraction of 10%, the mass ratio is compatible with the solid core NTR Isp of 1000 sec, and the max plane change round-trip delta-vee requirement of 11.52 km/sec.

The payload comprises the crew of 3 with suits, 2 weeks of air, food, water, and fuel cell reactants (conservative for a maximum 1 week mission in case of trouble), a 3-man rover car, an inflatable Quonset hut with camping and cooking gear, and half a metric ton of scientific equipment, to include a small drill rig. Assuming the command cabin structure itself to be part of the payload, I rough-estimated 6 metric tons for payload. Thus the whole lander fully-fueled is 60 tons, with a propellant weight of 42 tons, and a dry-tank weight of 18 tons. I assumed half a ton of waste was left behind at takeoff, in making propellant usage calculations. See fig. 8. Note that an empty lander is within the payload weight to LEO of a Spacex Falcon-9 booster, although not within payload shroud dimensions.

Fig. 8 – Rough-Out Lander Design

This lander design with a 180-200 KN thrust solid core NTR engine does not have quite enough thrust to leave the surface of Mars fully fueled, but can easily take off partly-fueled, after landing from orbit. A slightly higher thrust specification would make fully fueled surface takeoff possible, but at the cost of a slightly heavier and larger engine. This is not really a necessary requirement for explorations conducted from LMO. Performance is compared to requirements in fig. 9 below.

Roughing out this vehicle is a supremely important prerequisite for the rest of the mission and vehicles because it is a major payload item, as is the crew habitat module. It should be noted that it is specifically the choice of nuclear propulsion that makes a single-stage lander possible. The delta-vee requirements for a single stage lander are simply out of the practical range of mass ratios for chemical propulsion. Without a reusable single stage lander, the mission exploration return is very much diminished: we are more-or-less back to a very few Apollo-style “stunt” landings. With chemical propulsion, a staged lander may only be used once (although an upper stage might be reused with a new lower stage). The number of landing sites is then no more than the number of landers carried to Mars, and this has a far greater effect on vehicle weights, mission complexity, and costs.

For the design and mission selected here, the average lander mission consumes some 37.56 metric tons of liquid hydrogen (LH2) nuclear rocket propellant. The plan for 16 such landings during the course of the mission then requires the delivery of some 600.89 tons of propellant as payload to Mars to support lander operations. That delivery requires even more propellant for the Hohmann transfer, even with using the lander propulsion as the propulsion sending these vehicles to Mars in order to save weight, launch costs, and complexity. See the unmanned vehicle rough sizing below.

For safety and self-rescue purposes, lander operations are envisioned as a series of sequential single landings, each with a crew of 3, while the other 3 stay in orbit to monitor progress and provide rescue capability with another ready lander. Thus at least 2 landers are required, and unless there is a third, the mission ends if one is rendered inoperative for any reason. That is why this mission plan sends three landers to Mars. Rescue is envisioned as risking only one crew as pilot in the rescue lander.

Fig. 9 – Rough Lander Performance Estimates

Crew Habitat Rough-Out:

There are three fundamental crew survival issues that must be addressed for any mission involving months of travel beyond Earth’s Van Allen Belts. These are (1) radiation protection (solar flares and cosmic rays), (2) protection from medical deterioration due to microgravity, and (3) sufficient habitat volume to stave off the psychological effects of prolonged confinement.

This mission is nominally 9 months, and certainly under 1 year in total duration. The dose of accumulated cosmic radiation is minor. The probability of a solar flare event is quite high. Therefore, there must be a zone inside the habitat shielded by water and wastewater tanks, and a little steel plate, which can support 6 crew temporarily during the event. Safety demands that this shelter also be the ship’s command deck, so that critical mission maneuvers may be flown, radiation storm or not.

The under-1-year mission time is within the realm of experience we have with microgravity exposures on the International Space Station (ISS). Therefore, this design need not provide artificial gravity by spin. The ISS exercise regimens will be adequate. This is very important, because provision of artificial gravity greatly adds to the habitat and vehicle size, weight, and complexity. The design requirements for such artificial gravity are still poorly understood, since the direct experimental work has never been done. We have only imperfect, indirect evidence from surrogate studies, such as bed rest. It is unethical to subject a crew to serious life and health risks, based on no better evidence than that. Therefore, the best design criteria we have are to provide one full gee at no more than 4 rpm. Any slower trajectory pushing total mission time beyond 1 year must deal with this design issue.

The habitat volume per crew issue is something ignored in design studies such as “Transhab”. Most of these designs would confine a crew in a space per person not much bigger than a typical bedroom closet. This is very likely to be psychologically very unhealthy, as any prisoner who has served time in cramped solitary confinement can testify. A design volume per person more like that in a lower middle class home would be far preferable. Units this large would resemble the old Skylab station in dimensions, and would be very difficult to launch. But, such a habitat could be assembled from smaller modules. It would be a part of the payload of the manned vehicle, the crew return capsules being the other part.

This study’s design is comprised of three modules, each 32 metric tons, docked in LEO to form a 96 ton habitat, stocked with substantially more than a year’s supply of food, water, oxygen, and other supplies. Such a habitat is very close to the mass of the old Skylab, but has a longer, narrower form factor. One of these modules would contain the radiation-shielded command deck. Each of these modules is within the near-term projected payload capability to LEO of the Spacex Falcon-9-heavy launch vehicle. Again, payload shroud constraints may be violated. See fig. 10.

Fig. 10 – Three-Piece Assembled Habitat Module

Crew Return Capsules:

For safety purposes, it is critical that these be carried with the habitat on the entire round-trip mission. This is because the vehicle propulsion might fail on the return trip, leaving no way to slow for capture. In that event, a crew return capsule capable of making a free return reentry at speeds very significantly higher than Earth escape speed offers the only avenue of crew escape. Each capsule should be capable of carrying the entire crew of 6, and there should be two such craft for redundancy. Some amount of service module propulsion is required to effect a proper reentry angle for survival.

Such a capsule already exists in its initial form as the Spacex Dragon. Dragon has crew capacity up to 7, and a heat shield rated for free Mars return. It fits a Falcon-9 launcher, although fitments need to be changed to accommodate some extra propulsion. I simply guessed this add-on propulsion module at 2 metric tons each. This plus the empty Dragon should be in the vicinity of 22 tons. See fig. 11 below.

Fig. 11 – Modified “Dragon” as the Crew Return Capsule, Two Required

Common Propellant Tank Module:

This is a more sophisticated design item than it first appears. There is a need to store LH2 for months at a time in zero-gee conditions. That requires what amounts to a double-shell tank, essentially a Dewar, with protection from solar thermal radiation, and at least a little meteor protection. As stackable modules, these require substantial structural strength. There is some sort of cryo-cooler (or a suitable equivalent) equipment required, plus the solar power to run that. There is considerable interconnect piping to meld these modules into an integrated propellant supply, plus a kit of extra pipe lengths and fittings to make those interconnections. It will be a substantial design challenge to achieve this in a 10% inert weight budget. Loaded tank size is set by the projected Falcon-9-heavy deliverable LEO payload weight of 32 metric tons. See Fig. 12 below.

Unmanned Vehicle Rough Sizing:

A part of the payload for this vehicle is an empty lander (18 metric tons), whose engine is also the unmanned vehicle propulsion. Inert weights of 3.2 tons per tank module add to this payload to comprise the dry-tank “burnout” weight. Loaded weights of 32 tons per tank module add to this payload weight to comprise the departure “ignition” weight. Thus vehicle mass ratio and delta-vee capability is a function of the number of propellant modules in the vehicle stack. Enough untapped modules need to arrive at Mars to support the mission’s lander operations. One third of that requirement (rounded up to the next largest number of tanks) is carried by each of the three unmanned vehicles, each with a lander. Those untapped tank modules are the remainder of the vehicle payload. See fig. 13 for a pictorial, and fig. 14 for estimated vehicle performance on its one-way mission.

Remember, after the mission concludes, the landers and empty tank assets are left docked in LMO. Subsequent missions need only bring more propellants, and reuse the landers, up to the lander engine lifetimes. Empty tank assets could be cannibalized for other purposes in future missions.

One other note: the same basic vehicle design is suitable for a variety of inner solar system missions. One simply stacks up enough common tank modules to meet the mission velocity requirements.

Fig. 12 – Common Tank Module

Fig. 13 – Unmanned Vehicle Stack (One of Three)

Fig. 14 – Rough Estimates of Unmanned Vehicle Performance

Manned Vehicle Rough-Sizing:

The design approach for this vehicle is very similar to the unmanned vehicles, only the payload and propulsion is different. The payload comprises the three-piece habitat module, plus two crew return capsules. To the radiator assembly (30 tons) and twin gas core NTR engines (half ton each of two for redundancy), one adds 3.2 tons per tank module for the inert weight total. This plus payload is the dry-tank “burnout” weight. To this, one adds 28.8 tons of propellant per tank module, to arrive at the departure “ignition” weight. As with the unmanned vehicles, mass ratio and delta-vee is a function of the number of tank modules in the stack. See fig. 15 below.

In this particular design, all the propellant required for the two-way trip is included in the vehicle. It would save weight to send the return trip propellant as a Hohmann transfer unmanned package, but, this incurs the risk that the manned vehicle might not be able to rendezvous with the unmanned fleet. In that event, the crew would be stranded, and would die. Abort scenarios where capture is avoided at Mars for an immediate return home would also be impossible. From a safety standpoint, it is simply more prudent to fuel the vehicle to be able to return home independently of all the other mission components.

Fig. 15 – Manned Vehicle Design for the Fast Trip

The Missing Technologies: Solid and Gas Core Nuclear Rockets

The key element to the fast-trip design of the manned vehicle is, of course, its engine. This is presumed to be a gas core open-cycle version of the basic nuclear thermal rocket. Unlike the NERVA-type solid core engine design of the lander, the gas core machine was never tested as a rocket engine. However, it did undergo component bench tests verifying containment of the uranium relative to the hydrogen, at about 1000:1 hydrogen:uranium flow rate ratio. It also underwent bench tests verifying controlled gas phase nuclear fission. The gas core engine was about 2 years away from a first-article rocket test when the program was shut down in 1972. The mission plans at that time allowed about 15 years to test and perfect the design, before potentially employing it on a manned Mars mission then scheduled for 1987. Most of this history is forgotten today.

The open cycle gas core NTR was thought to be adequately cooled by regenerative cooling, up to power levels corresponding to Isp around 2000-2500 sec. Above that power level, regenerative cooling was known to be inadequate. This necessitated use of a high-temperature radiator to cool the engine, whose characteristics are still guesswork. It was also thought there was a power limit above which the engine would vaporize itself due to propellant transparency to all radiation, up around 10,000 sec Isp. The planned Mars engine for 1987 was an Isp = 6000 sec design, well under that poorly-understood upper limit. That same projected design is assumed for this mission study. See fig. 16.

Fig. 16 – The Radiator-Cooled Open-Cycle Gas Core Nuclear Thermal Rocket Engine

Assuming that the developed engine and radiator system have characteristics even close to what I used for this study, then the performance of the vehicle powered by it can be calculated with at least some confidence. The results of have 6000 sec of Isp available is astounding, as given in fig. 17. Comparing this plot to the performance plot for the unmanned vehicle (of crudely similar size), one can see the difference in the delta-vee levels achievable: several tens vs only several km/sec.

Fig. 17 – Rough-Estimated Performance of the Gas-Core Manned Vehicle

As a comparison, fig. 18 below illustrates the updated NERVA engine used in the lander. That technology was substantially mature, and this shows in the quoted data in the figure. It should be noted that the lander design as worked out uses one engine, not a redundant two or three. The problem is one of thrust against gravity, and scalability of the nuclear design. The size used herein is not all that far from the original NERVA. There is some question whether a much smaller engine could even be made to go critical and produce power. Engine-out under gravity means that the remaining engines must throttle-up thrust levels to compensate for the lost engine. It might actually be easier to simply redesign the basic engine to be more reliable. This is an issue needing investigation before any designs can be finalized.

Fig. 18 -- Solid Core Lander Engine Based on NERVA Technology

Mission Information Return vs Mission Cost:

The direct launch costs are “retail”, based on number of payloads times the cost for the appropriate launcher. The data were obtained from the Spacex website, including projected costs for the yet-untested Falcon-9-heavy vehicle, and the Falcon-9 vehicle currently in flight test. On this basis, the total launch cost to LEO for 3 unmanned and one manned vehicle is right at $8 billion. That “buys” 16 landings, each up to a week long, at 16 separate and widely-dispersed sites on Mars, up to 30 degrees worth of plane change from the ecliptic. It also buys hardware that can be used again on subsequent missions, and other missions in the inner solar system.

There are hardware development and production costs to be considered, and programmatic costs. The habitat modules, the common propellant tank module, the modified “Dragon” crew return capsules, and the lander, are all items needing a “normal” amount of development, in the aggregate perhaps totaling around a billion dollars. The updated NERVA engine for the lander would actually require very little development. On the other hand, the gas core engine would be a very serious development item. Taken together, those two engines might total around a billion or two dollars. That is a wild guess predicated upon these projects being done by lean, efficient organizations. For “business-as-usual” with large, inefficient organizations, one should probably double or triple those estimates. Thus, the lower bound “wild guess” is then about $11B for 16 landings on Mars, all in one trip. See fig. 19.

Fig. 19 – Summary of Mission Design Characteristics

Reduced Mission Scope?

It is entirely possible to lower costs by reducing the number of landings under the same mission safety rules. The minimum is three. Only the three unmanned vehicles reduce in size, the manned vehicle is unchanged. But on a dollars-per-landing basis, that would be a very inefficient thing to do. See fig. 20 below for estimated savings from reducing the number of landings, total mission cost, and the prorated per-landing cost, using $11B as the total for a 16-landing mission.

Plus, there is the political effect of “getting much of the initial exploration done” in a single trip with 16 landings, in such a way as to enable future prospecting-base missions, and eventually, a colony. Compare that to the single-landing scenario, which would require more exploration missions before anything else could be done. Each and every one of these follow-up missions could be cancelled.

As a proper exploration strategy, allow me to suggest a “shotgun-pattern” planetary survey with the full 16 landings, perhaps to be followed by a second exploration mission of fewer landings at the most promising sites uncovered by the first mission. These fewer landings in the second mission would stay substantially longer times on the surface, rather similar to those proposed in “Mars Direct”.

Fig. 20 – Cost Trades vs, Number of Landings in a Single Mission

Once the exploration planetary survey is done, we are ready for a different type of mission in which “prospecting” bases are built. This is the type of mission where the in situ resources begin to be utilized, and the first indications are obtained as to what trade commodities there might be, and what the trade economy might be. These are the prerequisites for an actual colony in the future.

Alternatives to Gas Core Nuclear Thermal Rockets:

To do the fast trip manned vehicle with solid core technology requires a throwaway staged vehicle, which is neither cost effective, nor conducive to authorizing follow-on missions. Slowing to Hohmann-transfer speeds puts the total mission well over a year, which requires artificial gravity (and the resulting huge impacts on vehicle weight, size, and complexity, as well as costs). This path is not recommended.

One could spend efforts developing a flightweight nuclear electric power plant in the multi-megawatt range. Then the same manned fast trip could be done with VASIMR, or something very much like it. Developing such a power station is likely about the same risk as producing a gas core NTR engine. There would be a larger inert weight for the manned vehicle, and a much smaller propellant weight, of a different type, with VASIMR. The unmanned vehicles would look the same. This path is a recommended possibility, although there is less commonality with the unmanned vehicles.

Opting instead for nuclear pulse propulsion runs into the odd efficiency scaling that kind of propulsion entails: Isp is higher at larger vehicle masses. At the masses of these exploration vehicles (619 metric tons at departure from LEO for the manned vehicle, 690 tons for each of 3 unmanned vehicles), pulse propulsion Isp resembles no more than gas core NTR Isp, and is maybe not as good. At vehicle masses around 10,000 tons and up, pulse propulsion looks like Isp = 10,000 sec or higher, and this at vehicle accelerations in the 2-4 gee range. These kinds of characteristics are well suited to large scale operations like base-building and planting actual colonies.

A Note on Crew Selection:

This mission is planned around 6 crew members, going to the surface 3 at a time. Since the goal is a science information return, to be obtained with maximum crew safety, I suggest each group of 3 be one pilot/engineer, one geology specialist, and one chemistry/biochemistry specialist. That would be two of each comprising the 6 total. Each should be cross-trained enough to function as a lander pilot for emergencies. Each should be cross-trained enough to support the other science specialties.

Concluding Remarks:

The point of this study was to show that a manned mission to Mars is feasible, and safe, with launch rockets available today or within 5 years.

All of the known crew health and survival issues can be addressed in a design that can be assembled from docked modules that fit the presumed launch rockets. Mission times are short enough not to provide artificial gravity. To go further out than Mars will require artificial gravity.

There are two missing propulsion technologies: (1) an update of the old solid-core NTR “NERVA” technology, which could probably be available in under 5 years, and (2) a gas core NTR (or VASIMR equivalent), which will likely require about 10 years to make ready.

These vehicle designs that support a well-planned exploration of Mars are reusable, and could be utilized anywhere in the inner solar system.

(see also the 1-8-2011 post for an update to this study)

Fast Transit To and From Mars

2010-11-29T19:49:00.000-08:00

Based on some earlier work, I thought I would take a more systematic look at fast trips to and from Mars. I limited this study to gas core nuclear thermal rocket propulsion.

Orbital Tradeoffs for Transit:

The “typical” close approach distance from Earth to Mars at opposition is some 97 million kilometers. For very fast one-way flight times (on the order of 30 to75 days), the trip is very nearly a straight-line “shot”, at a rather low angle to a radial line. For fairly short stays at Mars (1 to 3 months), this picture doesn’t change much. Accordingly, I chose a nice round figure 100 million km, with an angle factor of 1. I looked at transit times from 30 to 75 days, computing average velocity as path length divided by transit time. A spreadsheet program automated these calculations. See fig.1.

Figure 1 – The Fast Transit Orbital Picture

The peak velocity is related to actual vehicle delta-vee. For fairly short impulsive delta-vee events, the velocity versus path length trace is nearly a square wave. The higher the acceleration, the closer the peak velocity is to the average. I assumed for this study the fairly brisk acceleration level of 0.5 gee at departure weight, knowing this would increase by the mass ratio at burnout weight, unless thrust were throttled back by the mass ratio. This takes the form of a manual convergence in the spreadsheet. For a round trip, vehicle delta-vee capability is 4 times this peak velocity.

Assumed Vehicle Characteristics:

The vehicle assumed for this generic trade study is depicted in the fig. 2 cartoon. This vehicle is single stage, 10% payload, 20% inerts, and 70% propellant. This corresponds to a vehicle mass ratio of 3-1/3.

Payload includes the crew, their habitat and supplies, their crew return vehicles, the Mars landing craft plus their propellants and all the exploration gear. All of this landing and exploration equipment is assumed to be returned to Earth, which is not realistic, but is very conservative. That “covers” the lack of a propellant “margin” in these figures.

Inerts include the vehicle connective structures, auxiliary electric power generating equipment, tank structures and insulation, cryo-cooler equipment to keep the liquid hydrogen liquid for months at a time, the basic gas core fission engine, and the waste heat radiator equipment which allows the engine to operate at Isp greater than about 2000 sec. In particular, the radiator could be a very major portion of that 20% inerts figure. It is unknown whether this is really adequate.

The propellant figure is the liquid hydrogen working fluid for the gas core fission nuclear thermal rocket engine.

Figure 2 – Basic Vehicle Assumptions

Velocity Requirements:

From the velocity trace modeled in the spreadsheet, we calculate average velocity versus one-way travel time, and peak velocity from the average using a constant 0.5 gee acceleration level. These are plotted in fig.3. At this level of acceleration, average and peak velocities are virtually indistinguishable. Required vehicle delta-vee for a one-way trip is twice the peak velocity, four times peak for a two-way trip, as plotted in fig. 4

Figure 3 – Average and Peak Transit Velocities

Figure 4 – Estimated Vehicle Delta-Vee Requirements

Assumed Propulsion Characteristics:

The focus here is on one type of propulsion: open-cycle gas core nuclear thermal rocketry. This is something that came very close to propulsive test in the late 1960’s.

At the end of the 60’s, controlled gas-phase fission had been demonstrated, as well as a powered-plasma flow model simulating a hydrogen to uranium flow ratio of 1000:1, which was as good as perfect containment. It was known that a waste heat radiator would be required, but not how it could be built, nor how much it would really weigh.

Neither the characteristics of that radiator, nor the maximum impulse capability of such systems, were known at that time. However, a first generation design concept was targeted at specific impulse Isp = 6000 seconds, because they were pretty sure they could achieve it. That is the screening criterion used in this study.

Estimated Propulsive Requirements:

From the mass ratio and the two-way trip delta-vee figures, the exhaust velocity required of the engine can be estimated. Dividing this by the gravity constant is a rather good estimate of the engine Isp required.

From the previous study, the payload for a crew of 6,twolanders,and two crew return capsules was 240 metric tons. Using this figure and the 10% payload fraction provides a constant estimate of Earth orbit departure weight, shown in fig.5 with the Isp.

Figure 5 – Required Isp and Departure Weight

The acceleration gee level at departure was assumed to be 0.5 for this study. At burnout weight on return, the acceleration is higher by the mass ratio, assuming constant thrust level. Under the constant thrust level assumption, the departure weight and acceleration set the engine thrust level. For these figures, the departure weight is 2400 metric tons, so the thrust level is 1200 metric “tons-force”, or 1,200,000 “kg-force”, which is some 11.76 mega-Newtons of force in proper SI units.

Thrust divided by Isp is propellant flow rate. Propellant mass divided by flow rate is the total burn time available to “dry tanks”. This was converted to hours for easy plotting. Vehicle departure and burnout gees, plus total burn time to dry tanks, is given in fig. 6.

Figure 6 – Acceleration and Burn Time Data

Conclusions:

This oversimplified study is for a single-stage two-way trip: everything goes and returns, no expendables. From the standpoint of Isp’s thought feasible, it looks like such a vehicle could be powered by a gas core fission rocket, if the one-way travel time were no shorter than about 65 days. This is for an “average” opposition mission.

Possible major refinements: (1) reduce lander mass in favor of more lander propellants (extra landings) by taking some advantage of descent aerobraking, and (2) leave landers, unused lander propellants, and some small portion of the vehicle tankage behind in Mars orbit, as a depot for subsequent missions, thereby reducing payload and inerts for the return home.

Mars in 39 Days One-Way !

2010-11-26T23:14:00.000-08:00

This was a discussion topic in the forums on NewMars.com. I decided to try a bounding calculation to rough-out what might be involved. My first attempt was not very good, but I have double-checked these numbers, They are good.

Orbits, Distance Traveled, and Times:

Mars and Earth move into opposition roughly every two years. The straight-line distance from Earth to Mars at opposition varies considerably, but averages 60 million miles (97 million kilometers).

For trips faster than about 60 days one-way, and stay times at Mars under about a month, the planets don’t have time to change relative position very much. Such travel is essentially a straight-line shot over a distance of about 60 million miles (see fig. 1). Longer travel times are different, because the orbits become curved, and the path along the orbit is much longer.

Figure 1 – Straight-Line “Shot” for Fast Trips Around Opposition

Constant-Acceleration Kinematics:

Along a straight-line path, at constant acceleration, the average and peak velocities can be estimated quite easily. Average velocity is merely the path length divided by the travel time. This is just high school physics math, but nothing better is needed.

I looked at two bounding cases: (1) vehicle acceleration so low that one thrusts to accelerate to the mid-point of the one-way path, skew-flips under power, and then thrusts to decelerate the second half, and (2) one accelerates “hard” at each end such that time spent in acceleration is only 10% of travel time for each of the two “burns” (the “10% tails” case).

Both cases put lower limits on vehicle thrust. The continuous-burn skew-flip case has a lower ratio of average velocity to peak velocity at 50%. The 10% tails case has average velocity 90% of the peak velocity. See fig. 2.

In both cases, one computes the acceleration required as the peak velocity divided by the burn time. The two peak velocities are different for the two cases. The skew-flip burn time is half the travel time. The “10% tail” burn time is 10% of the travel time.

Figure 2 – Kinematics for Straight-Line “Shot” Trips: Two Cases

The data inputs here are path length 97 million km, and 39 days travel time. This figures to an average velocity of 28.79 km/s. To go there and back again, requires a total of 4 burns, each to the peak velocity.

My figures from the hand estimates and from the spreadsheet are a peak velocity of 57.57 km/s for the skew-flip option, and 31.99 km/s for the “10% tails” case. These correspond to total propulsive delta-vee requirements of 230.3 km/s for the skew-flip option, and 127.9 km/s for the “10% tails” option.

The 10% tails option spends most of its time coasting at peak velocity, reflected in the higher average/peak velocity ratio. For the same average velocity, the “10% tails” option has a lower peak velocity, and a lower delta-vee requirement, by a substantial amount. However, the thrust levels demanded are substantially lower for the skew-flip option, since it has half the trip time to accelerate, instead of 10% of it.

Vehicle Components and Configurations:

In order to do a little better than a simple Apollo-style “stunt” landing, I decided to send a fairly substantial crew of 6 with two landing vehicles and two crew return capsules. I included a habitat module on the order of the old “Skylab” module. The idea would be to rotate the crewmembers on successive surface expeditions, 3 in each lander, to cover 2 separate landing sites.

My estimates for the habitat reflect stored supplies, no artificial gravity, but a radiation-sheltered room from which the ship may be maneuvered during a solar flare event. I simply guessed this module at 90 tons, same as the old “Skylab” station. There is enough volume in a module this big for a crew of 6 to stay sane and healthy for perhaps 6 months at a stretch. The design here is 3-4 months round trip.

The crew return vehicles would each seat 3 comfortably, all 6 in a pinch, rather like the original Apollo capsules. These would have an abbreviated service module with hypergolics propulsion for deorbit maneuvers. I simply guessed these at 5 tons each, for 10 tons total. See fig. 3.

Figure 3 – Payload for the Orbit-to-Orbit Transfer Vehicle

I wanted to rough-out the landers as single-stage nuclear devices. This would provide the option to leave these vehicles in orbit at Mars, to be used again.

Mars surface-to-orbit requires a little under 6 km/s delta-vee one-way. Accordingly, I roughed out a single-stage rocket for a total 12 km/sec delta-vee, using a solid core nuclear thermal Isp = 1000 sec. This allowed me 9% payload, 71% propellant, and a 20% inert fraction (to cover landing legs, a cargo crane, and maybe a bit of shielding around the reactor.

I figured the payload as 3 men and supplies plus some sort of a rover car, similar to what we did on the moon. I guessed 6 tons for this, and calculated just under 70 tons for each lander. For two landers, that adds 140 tons to the transfer ship payload. See fig. 4.

The old NERVA solid core engine tested as 1000 sec, with an engine thrust/weight near 3. This was insufficient for surface launch from Earth, but might work on Mars. A design update to incorporate the results from projects like Dumbo and Timberwind would very likely provide a 1000 sec Isp engine with thrust/weight 10+.

Figure 4 – Single-Stage Solid Core Nuclear Thermal Lander

I looked at a variety of configurations, based around a core vehicle with a single engine or engine cluster, and some number of added propellant tanks (with no engines of their own). The idea would be to assemble this vehicle in low Earth orbit (LEO), and fly it to Mars orbit, then back to Earth orbit. Maximum recovered components provide maximum reusability for future missions.

I looked at a very conservative design for the core vehicle, to cover both the robustness necessary for reusability, and the connective structure to hold propulsion, tanks, and payload items together. I assumed 10% payload fraction (the 240 tons discussed above), and 20% structural inerts fraction. That leaves 70% for propellant load.

The tanks are also conservative, in that I assumed 10% structural inerts and 90% propellant. This should cover some sort of cryo-cooler equipment items necessary to store liquid hydrogen as propellant for the presumed nuclear engines. See fig. 5.

Figure 5 – Assumptions for Core Vehicle and Added Tanks

Not every configuration I looked at uses staging (in the sense of dropping empty tanks). For those that did, I assumed only two burn phases: (1) burn all the propellant out of all the tanks and then drop the empties all at once, and (2) burn on the propellant in the core vehicle. This is suboptimal, as one could drop tanks as they become empty (it might be necessary to do this in pairs to preserve weight and balance). Accordingly, my estimates of assembled vehicle weight and propulsive Isp may be a little bit high, but only by a very few percent. See fig. 6.

The list of configurations is as follows:

(1)baseline core vehicle without tanks, as a single stage (all recovered and reusable)

(2)core vehicle plus tanks at 50-50 weight split, single stage (recover and re-use tanks as well as core)

(3)core vehicle plus tanks at 50 - 50 weight split, two stage (expendable tanks)

(4)core vehicle plus tanks at 20 core - 80 tanks weight split, two stage (expendable tanks)

(5)core vehicle plus tanks at 5 core – 95 tanks weight split, two stage (expendable tanks)

Figure 6 – Staging Assumptions for Expendable-Tank Configurations

All of these configurations use the same core vehicle and 240 ton payload assembly. The transfer propulsion is on the core vehicle. It is assumed to make all four burns required by each round trip, and be reusable for several more trips.

I ran a weight statement for each of these configurations, based on the component weight fractions, the shipped payload (240 tons), and the core-tank weight splits.

For each of the three staged configurations, the mass ratio for the first delta-vee burn is departure weight divided by departure weight minus the total tank propellant weight. The mass ratio for the core vehicle is always the same: core vehicle weight divided by core vehicle weight minus core vehicle propellant weight. Assuming the same Isp for both burns (it is the same engine), one may sum the natural logarithms of these mass ratios and use that sum as total vehicle delta-vee divided by engine exhaust velocity.

The two single stage configurations are the core vehicle alone and the core vehicle plus tanks that are never dropped. One simply runs single mass ratios based on the weight statements for these configurations.

These mass ratio-based velocity ratios may be used with the total delta-vee requirements from the kinematics to estimate propulsive Isp. Delta-vee divided by the velocity ratio is an estimate of engine exhaust velocity required to perform the mission. This exhaust velocity divided by the gravity constant gc (9.805 m/s2 = 32.174 ft/sec2) is a rather good estimate of required propulsive Isp.

From the 240 ton payload weight, and each configuration’s weight statement, one can estimate the departure weight of the vehicle that must be assembled and fueled in LEO. From that departure weight and the average acceleration required by the assumed kinematics (skew-flip or “10% tails”), one can estimate the minimum allowable actual thrust required of the engine system.

Those data are given in fig. 7. Thrust figures are minimum allowables. Remember, for the skew-flip option, a minimum vehicle acceleration of 0.0035 gees is required; higher values allow one to coast a little between the acceleration and deceleration burns. For the “10 tails” option, a minimum vehicle acceleration of 0.0097 gees is required; higher values simply shorten the “tails”.

Figure 7 – Results for 5 Configurations, and for 2 Kinematic Options

Indications of Propulsion Required:

The single-stage core-only configuration has a relatively low assembled weight in LEO (although still far too high for a single launch), but requires propulsive Isp beyond anything we understand with nuclear thermal rocketry, whether one looks at the skew-flip or “10% tails” scenario.

A variable-thrust electric approach such as VASIMR might possibly serve, although there are very serious questions whether an 8-24 ton thrust VASIMR can be powered by something that would fit this weight statement. For all the skew-flip configurations, a higher thrust phase is required for efficient orbital escape and capture at each end of the trip, something easily handled by VASIMR.

The single stage core plus non-expendable tanks, at a 50-50 core-tank weight split, has the effect of improved mass ratio and reduced impulse requirements. The 4800 ton departure weight is something that seems “reasonable” for assembly and fueling in LEO, although it is clear that many launches would be required.

For both the skew-flip and “10% tail” scenarios, the impulse requirements seem infeasible even for gas core nuclear thermal rockets (maximum in the 6000-10,000 sec range). The higher thrust levels make the VASIMR power supply problem worse, even in the heavier vehicle.

The two-stage core plus expendable tanks, at a 50-50 core-tank weight split, has the same departure weight of 4800 tons. The impulse requirements for the skew-flip and “10% tails” scenarios are reduced a little, but only a little bit. Even the “10% tails” impulse requirement is likely still infeasible for a practical first-generation gas core nuclear thermal rocket. A waste heat radiator is required, and would be rather heavy. It might not fit well in this weight statement.

The two-stage core plus expendable tanks, at a 20-80 core-tank weight split, has a larger departure weight of 12,000 tons. This is getting very ambitious for assembly and fueling in LEO. However, for the 10% tails scenario, the thrust level and specific impulse requirement looks very achievable for gas core nuclear thermal rocketry. The skew-flip scenario’s requirements still look infeasible for this type of propulsion.

The two stage core plus expendable tanks, at a 5-95 core-tank weight split, provides only a small further reduction in impulse, at the cost of an enormous increase in departure weight. 48,000 tons assembled and fueled in LEO is a rather daunting prospect, for only a reduction from 5270 sec to 4160 sec gas core Isp in the 10% tails” scenario (skew flip is still infeasible). There is no option here for solid core nuclear thermal main propulsion, and little prospect for it by extremizing the weight statement still further.

I would hazard a guess that solid core nuclear thermal propulsion would only become feasible if we multi-staged the core vehicle, including the nuclear engines. This is not an attractive possibility.

On the other hand, nuclear pulse propulsion is a method that is feasible, just not well-matched to payload-driven weight statements like these. “Small” systems in the 2000-5000 ton departure weight regime would have Isp values in the 2000-5000 sec range, based on the old design data from about 1960. Much larger systems, in the 10,000 to 20,000 ton range would have Isp values in the 10,000 to 20,000 sec range.

A large pulse propulsion vehicle in the 10,000 to 20,000 ton departure weight range would have the impulse to achieve the 39 day mission. The 2-4 gee vehicle accelerations available would cut the “10% tails” to 1% or less. Deliverable payloads would be more in line with colonization efforts that initial explorations, though, being around 10 or more times the 240 ton figure used here.

Conclusions:

The “10% tails” scenario is preferred because of the substantially lower impulse requirements, in spite of its higher minimum thrust requirements.

The impulse levels reported here are being driven by the very high average velocities associated with a 39 day one-way timeline for the flight. These vehicle results are being driven very hard by those impulse figures.

If the electric power plant problem can be solved, then a 4800-or-larger ton VASIMR-powered vehicle operating on the “10% tails” scenario might be an attractive option. This would be a staged vehicle with expendable tanks, at a 50-50 weight split. The VASIMR would have to operate briefly at lower Isp and higher acceleration to effectively “get away”, then throttle back to higher Isp at low thrust, until the required peak velocity is attained. That same variable-performance scenario in reverse is required for effective capture into Mars orbit. A similar profile is required for the return.

If the high-temperature radiator problem can be solved, then a 12,000-or-less ton vehicle powered by a gas core nuclear thermal rocket operating on the “10% tails” scenario can do this mission. The impulse is slightly lower than the 6000 second design target of the ca.-1970 design data, and the thrust is in a feasible range.

Other Options to Consider:

Mass ratio and delta-vee can be improved slightly with any of these designs by leaving the lander craft behind in Mars orbit for future re-use. Subsequent missions then need only bring propellants for these landers, not the entire craft, up to the design lifetimes of the landers’ engines.

Leaving the rover cars and other exploration gear behind would ease lander propellant consumption slightly, and provide ready assets for a subsequent return to these sites.

It might be wise to re-run this analysis at a 60 day one-way time, in order to see just how strongly that parameter drives these vehicle designs.

It would be wise to explore just how best to design a pulse propulsion vehicle to do this mission. Such designs are not driven by payload weight (mass ratio-limited at a separately-defined Isp), but by departure weight (which actually defines the Isp).

Detailed Weight Statement Data Used:

1-stage core only
Payload......0.10.......MR = 1.00/(1.00 – 0.70) = 3.333333333
Inerts.......0.20
Propellant...0.70.......ΔV/Vex = ln (MR) = 1.20397280

Core + non-expendable tanks (50-50)
Component....core.(50%).tanks.(50%).overall
Payload......0.10–0.05..0.00–0.00...0.05.....MR = 1.00/(1.00 – 0.80) = 5
Inerts.......0.20–0.10..0.10–0.05...0.15
Propellant...0.70–0.35..0.90–0.45...0.80.....ΔV/Vex=ln(MR)=1.6094379

Core + expendable tanks (50-50)
Component....core.(50%).tanks.(50%).overall...MR1 = 1.00/(1.00 – 0.45)
Payload......0.10–0.05..0.00–0.00...0.05..........= 1.818181818
Inerts.......0.20–0.10..0.10–0.05...0.15......MR2 = 0.50/(0.50 – 0.35)
Propellant...0.70–0.35..0.90–0.45...0.80..........= 3.333333333
ΔV1/Vex = ln (MR1) = 0.597837001........ΔV2/Vex = ln (MR2) = 1.20397280

Core + expendable tanks (20-80)
Component....core.(20%).tanks.(80%).overall...MR1 = 1.00/(1.00 – 0.72)
Payload......0.10–0.02..0.00–0.00...0.02..........= 3.5714
Inerts.......0.20–0.04..0.10–0.08...0.12......MR2 = 0.20/(0.20 – 0.14)
Propellant...0.70–0.14..0.90–0.72...0.86..........= 3.333333
ΔV1/Vex = ln (MR1) = 1.2729657..........ΔV2/Vex = ln (MR2) = 1.20397280

Core + expendable tanks (5-95)
Component....core.(5%)...tanks.(95%).overall...MR1 = 1.000/(1.000 – 0.855)
Payload......0.10–0.005..0.00–0.000..0.005.........= 6.89655
Inerts.......0.20–0.010..0.10–0.095..0.105.....MR2 = 0.050/(0.050 – 0.035)
Propellant...0.70–0.035..0.90–0.855..0.890.........= 3.333333
ΔV1/Vex = ln (MR1) = 1.93102115...........ΔV2/Vex = ln (MR2) = 1.20397280

Nissan Mileage Results on Blends

2010-11-17T19:18:00.000-08:00

These data came from over a year's worth of stiff blend trials in the unmodified Nissan. This was a 1998 "Sentra". I had no way to draw fuel samples until the last two tanks of fuel, so the blend strength data are best-estimate E-numbers for what was added at each fill-up. All but one or two of these were fill-ups on top of quarter tank-or-less residuals, so it's a decent estimate, no matter what.

Although scattered at about plus-or-minus 10%, there does seem to be a very slight (5%) decline in fuel mileage between all gasoline and 25-something-% blends. Maybe. The decline is 5% when the scatter is plus-or-minus 10%. Maybe that drop is not actually real.

Theoretical lower heating value for an E-25 is 91% that of plain E-0 gasoline: a 9% drop. It would seem the car is doing better than heating values would indicate, although the scatter is too great to actually draw that conclusion. This is pretty much the same result as was seen with blend trials in the unmodified Ford F-150.

One tank (only) was accidentally too rich in ethanol at an estimated E-49. This one data point is about 80% of the E-0 mileage. Such a drop is big enough to be "real", even with plus-or-minus 10% scatter. There are points without drop at about E-40 to E-42. This behavior is consistent with stiff blend trials in the unmodified Ford, and with the ethanol VW as rigged to gasoline settings. (See the November 12, 2010 post just prior to this one). The VW data are spot-check "snapshots" from intake vacuum measurements, not tank averages of mileage.

That's three wildly-different cars providing the same quantitative answer, and from two different completely-different types of data. Stiff blends in unmodified vehicles provide essentially gasoline-only performance up to blend strengths in the E-40-to-42 range.

For the Nissan, I was able to cross-check the best-estimated E-numbers against actual test measurements on the last two tanks. Both were within two percent of being correct.

I see nothing here that contradicts the conclusions in the earlier post:

(1) Above E-42, timing seems to be late (low vacuum, low performance).

(2) Below E-42, you cannot practically tell these blends from a straight gasoline.

Considering that no sophisticated test equipment was used at all, these results are quite remarkable, are they not?

Stiff Blend Effects in Gasoline Cars

2010-11-12T18:17:00.000-08:00

I recently completed the final "stiff blend" experiments with the "ethanol VW". This was the 1973 VW Type 1 sedan that I had converted to E-85-only operation in January 2007. In fall 2010, I had re-converted this vehicle to gasoline-only settings, and tried a series of increasingly-stiff (high E-number) blends. To make such testing practical, I added an adjusting screw to the E-85 main jet to make this a "flex-fuel" carburetor. This old-technology vehicle is carbureted, with distributor ignition.

For these stiff blend experiments, I used no added intake air preheat, and no extra timing advance. For most of these runs, I used the gasoline-only jet settings on the main and idle jet screws. I used the E-85 accelerator pump discharge nozzle as it was. Later in the tests, at higher E-number blends, I reset the main and idle jet screws a little, to accomplish manually the same "compensation" that electronic fuel injection does automatically. This was necessary to maintain good drive-ability.

I tested this VW in terms of intake vacuum observations during in-gear coast-down. This was based upon the well-known fact that late ignition timing shows up as lower intake vacuum levels, all else being equal. Blend strength was determined by a simple forced added-water phase separation test. The theory behind this is that ethanol requires more advanced ignition timing, because it has a longer ignition delay. Thus, you have to "light the fire" sooner, in order to achieve peak in-cylinder pressures by about 2 degrees after top dead center crankshaft position.

What I found was absolutely no differences between gasoline-only and stiff blend performance up to about 40% ethanol in the mix (E-40 blend). For blends from there up to the maximum tested (E-56), there was a definite drop in coast-down vacuum, although it was not large. My data curves show a definitely-repeatable "shape" that is most likely a systematic gage-reading error in these on-the-road tests. Essentially, all these blends show just about the same vacuum loss, once 40% was exceeded.

The VW blend experiment was inspired by earlier results in 2008 with stiff blends in my 1995 Ford F-150 pickup truck. This vehicle has distributor ignition and electronic fuel injection. I ran these stiff blends in it completely unmodified, and I still do today. The data for the truck were obtained from fuel mileage, not intake vacuum. I kept track of two driving cycles: "TSTC", which was a virtually-all highway commute 30 miles one-way to / from work at TSTC-Waco, and "SRMcL", which was general driving over shorter trip ranges around McLennan county, Texas.

Driving cycle on blend fuels did not seem to make much difference, after normalizing blend mileage to the corresponding gasoline-only average mileages on the same driving cycles. Somewhere just above E-40 blend strength, fuel mileage dropped dramatically. The performance characteristics above 40% ethanol were very smooth operation, low perceived power, and the sense that it was "sucking fuel too fast", confirmed later by the actual mileage data.

The data shown here are exactly the same data that were in the February 2008 "Ford Report 8" (see my website http://www.txideafarm.com, and navigate to the "ethanol projects" sub-page). A few more data points added later (by May 2008, but not included here) confirmed the same basic trend, and "pinned down" the critical blend strength a little better, to right at E-42.

The most amazing thing about this is that two vehicles of such different vintages and technologies gave almost exactly the same answer from two completely-independent types of data. Timing is late on gasoline settings, in otherwise unmodified vehicles, above about E-40-to-42 blend strength.

I got pretty much the same answer, although not actually quantified, from a 1998 Nissan Sentra early this year. My best estimate was that I accidentally put a blend a little above E-50 in the tank, causing perceived low power, low mileage, but smooth operation (exactly what one would expect from late timing).

It looks to me as if above about E-40 to E-42 blend strength in an unmodified car, the timing is suddenly late, as if "turned off like a light switch". Up to that point, one cannot in any way tell stiff blends from straight gasoline (either by mileage or by intake vacuum).

I would speculate that in blends under the critical level, the gasoline component controls the effective ignition delay, while for blends over the critical level, the ethanol component controls the effective ignition delay. Thus for stiff blends, the required timing setting is not smoothly dependent upon blend strength, but instead is a step function triggered by a critical blend level near 40-42% ethanol.

This finding has serious implications for the design of factory flex-fuel vehicles.

You heard it here, from me, first!

Election 2010: TX Governor

2010-10-25T15:50:00.000-07:00

My 3 issues for choosing are:

1. Factual sources of information (most definitely not campaign ads).
2. Re-elect no incumbent that you cannot personally verify did more good than harm.
3. Look for real specifics in proposed actions(names and numbers, not slogans).

Factual data as reported by the newspapers:

Governor Perry was behind the attempted land grab for the Trans-Texas Corridor, a proposed foreign-owned toll road. He vetoed the eminent domain reform bill passed by the Legislature specifically to make this possible. At stake was almost a million acres of privately-owned land to be forcibly taken by the state, and turned over to a foreign company for their profit. The bill he vetoed specifically outlawed this taking of private lands for private profit.

He was also behind the attempted fast-tracking of permits for a big cluster of old-technology coal-fired power plants to be built close to the greater Waco area. This would have driven McLennan county into air pollution non-attainment status, forcing local residents to get their cars emissions-tested for state inspection, at increased expense. It would have also driven the Dallas-Ft. Worth and Houston non-attainment areas into direct federal EPA control, something no one here wants.

More good than harm:

The two items listed above would have done great harm to the people of Texas. I see no way that the good Mr. Perry has done, and he has done some, outweighs these two items. He fails, and Mr. White wins, on this selection issue.

Proposed specifics:

I have not yet seen much (that I trust as factual data) from Mr. Perry or Mr. White in the way of specific policy proposals. It's a "wash".

Conclusion:

I recommend voting for Bill White for governor. That's what I will be doing.

On Election 2010

2010-10-23T16:04:00.000-07:00

It’s election time, which is always a very important decision, one I encourage you to make. Allow me to share three ideas, which I heartily recommend you consider, as you make your decisions.

First: believe only sources of factual information, campaign ads and internet forwards most definitely do not qualify. Stick to public debates refereed by responsible parties, and to interviews by actual professional journalists, as reported in the newspapers.

Do not place much weight on simple speeches. Those are good for drawing your attention, not so much for honest, factual, and substantive information.

Second: there is the tendency of incumbent politicians to become corrupted as time-in-office passes, in the sense that special interests buy them with lots of financial contributions. This takes time to occur, so as a general rule, regular turnover insures better, more honest representation for the common man.

There are exceptions to this generalization, but only down at the 1-3% level. If you can verify for yourself, from factual information, that an incumbent did more good than harm, then he is worth re-electing.

Third: when comparing two candidates, look at the details of what they say, not just the “sound” of it (which is really your own ideological “filter” talking, rather than anything the candidate actually said). Look for names, numbers, and other specifics. If you do not see specific proposals with actual figures and names, then you are seeing only slogans and ideology.

Better to choose the candidate with substantive ideas than the sloganeer. Ideologies and slogans historically made very bad public policy, such as happened in the former communist world, among many other examples.

In my book, looking at candidates with those three issues in mind trumps any possible politics, be they the candidate’s or yours. What you are looking for is an honest, thoughtful person, with substantive ideas, who will do right by you.

You choose for yourself. But I do offer these two thoughts:

I recommend re-electing Chet Edwards. He has by far the more substantive ideas about what to actually do. And by voting against his own party at times, I really do know Chet does more good than harm.

I also recommend re-electing Doc Anderson, and for the same reasons. Doc voting against his own party is partly why our kids will not grow up under a pall of coal ash. That is most definitely more good than harm.

As for the rest? I haven’t made up my mind yet. But, if I don’t know, then I vote “no”. A lot of them have not made their cases to me.