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Polywell must be having troubles, too. That topic has gotten awfully quiet.
Cold fusion still suffers from a toxic reputation, probably undeserved. There was something unexplained going on there, just not the same things in all the attempts to duplicate it, which is what tarred its reputation. But, it surely would be nice if we knew what those unexplained things were, wouldn't it?
I think I'd go for some kind of gas-core hybridized fission-fusion rocket engine, with the energy and particle flux of fission "sparking" at least a little bit of continuous fusion. Feed the exhaust stream directly into a gigantic MHD generator. Or, use it for thrust. Same device, either way. Nobody is working on devices like that. Maybe they should be. (The first article is a simple gas-core fission rocket engine, something we're reasonably sure could be made to work.)
Even if only the straight fission gas core device is ever made to work, the MHD energy conversion into electricity is much higher efficiency than any heat engine power plant process known. Plus, the fission rocket can always serve as a thrust device.
It's two birds with one stone (thrusters and electricity generators), and with technologies that are much more "sure" to work. Plus, it could lead to fusion devices. How could we go wrong?
GW
Terraformer:
Exactly!!! The trouble will be getting heard to provide input. Government programs have a long history of "not invented here" attitude, and the myopia that only established giant contractors have any credibility. I'm hoping Bob has a contact on the inside at DARPA. Even if he does, it's long odds against us.
I've been looking at the TSTO spaceplane concept for about 3 years now. It's definitely feasible, even without an engine breakthrough like Skylon. The baseline concept is a rocket airplane for both stages, with the first stage incredibly big, which will be the limiting factor for deliverable payload size. You can substantially reduce the size of that first stage by adding airbreathing assist, in particular ramjet, because of its speed range from about Mach 1.6 to about Mach 5-or-6, if you design it right. Gas turbine is limited to about Mach 3.5-ish, although it works from takeoff. Scramjet is just not ready for prime time.
I'm not at all sure vertical launch is the right thing to do with such a lifting craft, because of the required takeoff thrust. Horizontal takeoff lets you use much smaller rocket engines, saving weight, and reducing the massive rate of weight change during takeoff. Doing supersonic climb on ramjet to a pullover/acceleration run in ramjet, saves an enormous weight in rocket propellants you would otherwise have to have on board.
I'm beginning to believe the best next step is to relight the rockets and climb out of the sensible air to a 3 km/s release of the second stage in effective vacuum. That not only reduces the delta-vee required of the second stage, it also solves a whole host of supersonic/hypersonic store-release aerodynamical difficulties.
The second stage could be a non-lifting rocket pod, or a winged/lifting body spaceplane of some kind. Any winged stage should do a runway landing. A rocket pod second stage fits within a cargo bay easier. Having the second stage enclosed in a cargo bay solves a whole host of hypersonic aerodynamic heating and cluster-vehicle safety issues. The tip of such a rocket pod could be a modernized version of the old Mercury or Gemini capsule, with a modern ablative heat shield (like PICA-X) such that the capsule could be reflown several times.
If the second stage is a winged spaceplane, I'd recommend strong consideration be given to sidestepping-entirely the hypersonic glider problem (extremely-intense aeroheating and extremely-poor landing characteristics). You do that by resurrecting an early 1950's concept: fold the wings into the wake behind a fuselage that enters dead-broadside, belly-first. Then they can be straight subsonic-airfoil wings, and you can land at speeds more like 70 knots than 200+ knots, one whale of a lot safer and more reliable.
Sidestep all of the supersonic/hypersonic airloads problem, too. You don't unfold them until you are subsonic, hanging from a chute. Chutes are cheap. Potentially, reusable ceramics could even be used at the stagnation point if the ballistic coefficient is low enough and the belly is blunt enough (no more carbon-carbon nose cap and leading edges).
--GW
Hi Bob:
If you are thinking of contacting DARPA regarding their spaceplane, maybe we do have some things to offer.
There's my ceramic composite, which even on a high-ballistic coefficient hypersonic glider could be used everywhere but the nose cap and leading edges. We've been discussing conformal tankage, perhaps Al-Li or maybe even SS.
For interior airframe structure, you'll want composites, and I know 3 different ways to make joints in that stuff that are lightweight and reliable. The best of the 3 is also the least labor-intensive, actually.
Plus there is all the vehicle sizeout stuff you and I have looked at. And there's my experiences with ramjet if they want to add airbreather assist to it. (There are almost no American full-capability ramjetters left.)
Let me know what you think.
GW
The energy is theoretically there for fast travel with fusion. The technology is not. Fusion reactors have been "20 or 30 years away" for the last 60 years that I have personally experienced. The closest right now is laser ignition fusion, but with boron and protons, not deuterium-tritium or tritium-tritium. It'll take something different (like the boron maybe) to make controlled fusion work. We've done magnetic bottles for 60 years now without success.
I'm not saying it can't be done, but the problem turned out to be a lot harder than anybody ever suspected going in. The technology to do it is still missing because some of the actual science to support that technology is still missing. That's why the major efforts ongoing are still science projects, not engineering development or real prototype programs. (There are some scientists who might call their projects an engineering prototype, but that's because they weren't trained in development work and don't really understand what it is.)
That being the case, crowdfunding may be premature. It's more appropriate for development-flavored projects than true basic science efforts. There's quite a difference between the two.
There is one fusion application that could work very soon, and be could engineering-prototyped essentially "today": nuclear explosion propulsion, just done with modern thermonuclear devices instead of 1950's fission devices. (They do have to be shaped charges with directional radiation "blast", just like in 1959, though, and that's the development work that is needed.) I don't think you would want to bring one of those vehicles inside the van Allen belts, though. The EMP side effect would be disastrous. Best place to test / base it might be the moon.
GW
Hi Decimator!
Very interesting news.
It was fears about us doing exactly that 4 decades ago, that induced the Russians to fit their Almaz spy space station with a 23 mm cannon.
GW
Very good news. Keep us posted.
GW
Lessee, 3 km/s to 7.8 km/s, that's 4.8 km/s required, factored up about 5% to 5.3 km/s for gravity losses. LOX-LH2 at around 1000 psia chamber pressure might get you an exhaust velocity of 4.48 km/s. I get a mass ratio near 3.3 to arrive on LEO with pretty-much dry tanks.
That's a propellant fraction of about 70%. For reusability, I'd hesitate to claim an inert fraction any lower than 25%. So, you're looking at a winged device or rocket pod device that carries about 5% dead-head payload.
If it weighed 10 tons at ignition, it can carry about half a ton to orbit. That's the weight of a man in a suit with some water, food, and oxygen for a day or two, and some hand tools. And we'd still need to squeeze a deorbit burn out of this thing.
Pretty minimal, but it is do-able.
GW
No reason why not. It means more of the cargo mass is life support supplies.
Plus, it would make sense to build a radiation shelter. Sooner or later there will be a lethal solar flare. These years are solar max after all.
GW
Midoshi:
I see that the idiotic government shutdown may have stalled your MAVEN launch. The news reports indicate that your launch window closes in Dec. After that it's 2 years' delay waiting for a launch window again. That's terrible! How does it look "from the inside"?
GW
Void:
Go take a look at my Budget Moon Mission posting over at "exrocketman". Bob Clark and I both think we can put significant crew and cargo tonnage on the moon for an order of magnitude-or-two less money than NASA's old Constellation plan. But, you have to shed a lot of preconceptions to do it, including some most of us don't even know we have.
GW
It would be pretty neat to put a small 1 or 2 man craft up there, using this staged airplane concept. Then refuel it on-orbit with propellants shot up there with a light gas gun. Then you could go service satellites in-situ with it.
Or maybe, use the staged airplane to send up the men. Then refuel via light gas gun an orbital repair tug vehicle, previously launched and left up there. Man the refueled tug, and go service satellites with it. Leave the tug and come home in the rocket pod/spaceplane second stage. Again, and again, and again.
That's what the shuttle did, except we had to launch and recover the entire orbital repair tug (cargo bay and arm and operator's cabin) every single flight. That turned out to be not a very smart way of doing an otherwise very beneficial job.
GW
Spacex's site has dramatically changed recently. I'm still trying to figure out how to navigate it. This stuff is not intuitive to an old guy like me.
But, Bob is right: Merlin 1D has a slightly-higher thrust and specific impulse than Merlin 1C, and a significant weight reduction. Falcon 9 v1.1 is bigger (longer tanks), thrustier, and heavier-at-ignition than what they flew before. The most remarkable thing is the re-arrangement of the first stage engine cluster. This changes the geometry of the tail of the bird considerably, and might even reduce its drag slightly. It looks a tad "cleaner", anyway.
From the published news reports, it appears that Spacex made some progress toward its goal of recovering and reusing Falcon-9 first stages. They got it back into the air without breaking up, and then lost it to an uncontrolled spin. That's a control problem, not something fundamental. And that's definitely progress.
They seemed to have kicked up a storm among observers over pieces of debris off the second stage, after payload deployment, when they vented the tanks. Apparently, a bunch of pieces of foil-backed insulation look like tankage pieces on radar.
GW
I quite agree with Bob. There's no reason a return to the moon should cost $100M's and require all-new government vehicles and a giant launch rocket no one else needs. Bob knows more than I do about the specs and characteristics of items like the Centaur upper stage. You can trust his numbers.
But, I did a generic / clean-sheet-of-paper bounding calculation to show that we could land up to 7 men and many tons of stuff for about a 2 week stay, all with commercially-available rockets and capsules that should or could be flying in a year or two. The price would be closer to $1-2B than $50B. It would be robust enough to credibly start the construction of something permanent, not a flags-and-footprints reprise.
As Bob has said on his site, you have to free your mind of a lot of preconceptions, in order to succeed in designing something that would work and yet not break the bank. My study is posted over at "exrocketman", for those interested.
GW
I find it disturbing that NASA still thinks it can send people into a 2.5+ year mission without at least some artificial gravity. Besides the zero-gee transits, each of which is near the demonstrated max exposure time, there is no data at all (of any reliability) to suggest that 0.38 gee for that 1+ year on the surface of Mars is "enough". If bed rest studies were so good, then our bones would soften and decalcify in bed down here the way they do in space. They don't.
You do not have to build giant Battle Star Galactica's no one could afford, or resort to Rube Goldberg cable-connected contraptions of questionable safety at best, to provide artificial gravity by spin. It is easily integrated into a practical vehicle design, even at one full gee, which requires 56 m radius at the tolerable 4 rpm spin rate.
You shape your vehicle of docked modules as a baton that spins end-over-end, with the habitat at one end. This is an inherently stable configuration, as anyone who has ever seen the baton twirlers at a football game can testify. It integrates well with staged-off empty tanks: your baton just gets slimmer. There's no need for a truss, just connect the tanks axially and laterally.
I really have to question why otherwise-intriguing mission designs (of all types and from many sources) so often continue to ignore this medically-devastating issue, yet an issue so easily resolved.
GW
Hi Bob:
I'm beginning to think the NASA scientists stirred, or are being used to stir, this radiation-danger thing up in the press. It seems fundamentally to be a political / budgetary thing, cloaked in "technical difficulties" that just do not bear up to close examination.
We could have sent men to Mars successfully with a high probability of getting them home alive, starting about 1995. We'd learned enough about microgravity and radiation by then to know how to cope. As I said above, that crew shouldn't fly again outside LEO. I'd say here that they could expect increased problems with cancers in old age, relative to the rest of us. Somewhat increased, but not a lot.
The other problems are long-voyage food (good old heavy frozen food solves that one), and space-enough-to-stay-sane (a Skylab volume for no more than 6 solves that one). Then there is the lander. If you let it be big, it can be one-stage and reusable: a real ferry or landing boat.
That's worth doing, because you get to land men more than once at each of more-than-one site. It's still enormous trouble to send men two-way to Mars. Why go to all that trouble for just one landing? That just flag-and-footprints nonsense.
GW
Hi Void:
I quite agree that the moon is a good place to learn things. In particular, the in-situ resource technologies and the "how-can-we-make-construction-materials-out-of-regolith?" problem could benefit from being done on the moon. For one thing it's not hard to reach. For another, it's close enough men can ride there in a cramped capsule. And thirdly, rescue from Earth can actually be feasible on the moon.
In the long run, I think one of the really major attractions of a base on the moon will be a safe place to test nuclear propulsion articles without having to do plume capture, or risking fallout-on-the-neighbors. Not much in the way of air or water to pollute, either.
As for artificial gravity items on the moon, why not just build a Frisbee-shaped partly-or-wholly-buried stationary pressure vessel of a building, and set spinning concentric rings within it, angled for a proper sense of down perpendicular to the floor. We already know how to build non-spinning buildings, and the rings are rather similar to escalators or moving sidewalks. Why spin the whole building if you don't have to? That's really hard to do if you are not in free-fall.
None of that is "exploration", and it shouldn't be pitched as such to get funding, whether governmental or private. It supports exploration-later-with-better-technologies. And that is actually a very valid reason to go and do it.
And, as Bob Clark has so eloquently pointed out elsewhere in the forums and on his blog site, you don't need a giant rocket program to do it. We could do this right now with the rockets and space capsules we have right now. Although, having Falcon-Heavy flying next year does make it easier. Just give up on the one-launch/one-mission preconception, and start taking advantage of what we have learned the last half-century about rendezvous and docking-as-assembly.
When you do that, the only excuse for a giant rocket is cheaper cost per unit delivered payload. The sensitivity to size is pretty low, once you get past about 5 tons to LEO, being only about a factor of 2 from there to 100 tons. And government-only designs seem to be running about factor-4 more expensive than commercial launchers, throughout the range from 5 to 100 tons. My conclusion: SLS at 100+ tons to LEO will not be cost effective. Docking assembly with the rockets we have will be cheaper.
GW
Not sure about working overseas vs ITAR. I'm too old to move anymore, anyway.
But, the payload bay idea is a good idea. There are two limitations to consider, both with fundamental impact on the basic design concept. (1) the bay contains rocket pods a lot easier than winged or lifting-body shapes, and (2) store separations from a weapons bay are only known to work up to about Mach 2, which means that a launcher moving high supersonic or hypersonic must exit the sensible air to open its payload bay and launch the second stage it contains.
Exiting the sensible air requires either rocket thrust or coasting with a lot of deceleration, trading speed for altitude, in the face of drag. The most important staging variables are ranked as (a) velocity, (b) steeply-climbing path angle, and (c) altitude at staging. That ranking opposes the notion of stage release at an apogee point, because you've lost all your speed, and that's the most important item.
It's a tough problem. I'm beginning to believe that staging should not occur at ramjet top speed, but a tad later, after a big rocket-propelled pull-up. Not just a pull-up, but a pull-up maintained past the edge of the sensible air. But I have no numbers to back that notion up. That scenario would match the payload bay concept, as long as the second stage fits the bay (item 1 above).
The downside is that more propulsion from rocket raises first stage size and weight. There's a limit to how big an airplane can be, since mass scales as dimension cubed, while strength scales only as dimension squared. Yield and ultimate material stresses do not scale.
GW
If you mine the ice, I'd bet the perchlorate isn't in it. Perchlorate is a salt. Here on Earth, when sea water freezes, the ice is fresh water. The salt stays behind in the liquid.
Your best source of usable water on Mars isn't wet, salty dirt, it's buried freshwater glaciers. Big glaciers, not the little buried ice lenses that the polar lander uncovered. Big glaciers aren't everywhere, so locating them is the critical prospecting that has to be done before you land the stuff to create a base.
None of the rovers or orbiters can do that yet. Takes deep drilling to get reliable ground truth, of the sort you can bet lives on. Suspenders-and-belt, because nothing is as expensive as a dead crew.
It's likely there will be one manned mission to Mars (whoever does it). The real trick is configuring that mission to land and explore and find those buried glaciers, then land the base there, all in the one mission. We're talking about multiple landings based from orbit, until the base site is determined, then land everybody at the base, and stay there until it is time for the crew to come home.
GW
Wouldn't it be easier and most likely cheaper to build that wheel in LEO, instead of transporting all that stuff to the moon?
The old wheel-shaped space station design concepts of the 1930's and 1940's make really good sense, if you want to investigate "how much gee is enough?" for humans, plants, etc.
Same structure can investigate multiple gee levels at once if you shape it like more like a Frisbee than a bicycle rim, with many decks, all at various radii.
GW
That's amazing that such a thin atmosphere could be such an effective shield against GCR as it is. I would not have expected that.
My statement about "non-shieldable" was referring to spacecraft design. Wasn't thinking about atmospheric shielding down on the planet.
GW
Hi Terraformer:
I think you're asking about an experimental hypersonic ramjet airplane as a demonstrator vehicle. A thing like that could "easily" be done, and would approximate the old X-15 in terms of vehicle size and program costs. I don't have a dollar figure for that, but I'd guess nearer 10's of $M than a $B. It'd have to be done by a skunk works-like outfit, the "big boys" would turn it into a do-nothing "gravy train" just to suck on the government's tit. That's always at least 10 x more expensive, too.
Ramjet combustors with blockage-element flameholders were built and used in missiles up to about 1 m diameter long ago (Bomarc & similar). What's needed for an accelerator is a dump-combustor ramjet to render the flameholding insensitive to varying flight speeds and conditions. Those were flown up to about 0.5 m diameter in ASALM-PTV ca 1980.
Those dump combustors would need to be scaled up to around 1 meter dia for the experimental airplane, and maybe 2-3 m dia for a real TSTO orbital system of significant payload. That's cut-and-try experimental work, there is no reliable science for it, but it can be done. Engineering art, pure and simple, proponents of computational fluid mechanics notwithstanding. (Real flameholding is simply not in their models.)
There are now very few of us left in the US who know that art. I'm one. If Joe Bendot is still with us out in LA, he's another, but he's significantly older than me, and I'm a senior citizen now. If anybody's going to do this, they'd better get on the stick while we're still around. It's engineering art, and we taught no apprentices, we just all got laid off long ago.
The hardest part about such a project isn't the ramjet, it's the configuration layout to avoid disastrous aeroheat damage from shock impingement heating, and still demonstrate a form that could work for the orbital system. Depends on your airframe material, but with the Inconel skins on the X-15, it was Mach 6-ish for the speed above which the airframe gets cut apart by the heat from the shock waves. Mach 6 is also just about the upper limit for using ramjet, too.
GW
Bob:
I found a NASA document on-line that describes the astronaut rules for radiation exposure, which allow higher levels than for civilians here on Earth. There's a monthly limit of 25 REM, an annual limit of 50 REM, and a career total limit that varies by gender and age.
Using GCR that varies sinusoidally from 24 REM/year (2 REM/month) at solar max to 60 REM/year (5 REM/month) at solar min, and 2.5 years for a total mission time, the total absorbed GCR can range 60 REM to 150 REM on the mission.
This ignores the half-sky shielding effect for the year spent on Mars or in close orbit, waiting to return.
GCR is fundamentally un-shieldable by any technology or science that we have. But, the other source is brief bursts of solar flare radiation, which we can shield, with about 20 cm of water or wastewater packaged around some designated shelter area. So, the leakage through the flare shield adds what the planet-shielding effect subtracts, roughly. And my doses listed above serve as good estimates of the rates and totals of all forms.
Monthly is no problem. Annual is a problem at solar min conditions, with 60 REM/year vs 50 allowable. Career limits are no problem at solar max at all, and no problem at solar min for males as young as age 25, being right at the 150 REM limit. Females need to be about age 31 or 32 to have a limit equal to the solar min exposure total of 150 REM. They have no problem as young as age 25 for solar max career exposures.
The only problem I see is for annual exposures is during a mission pretty-much "centered" at solar min, when the annual exposure rate is higher than allowed under those astronaut rules. But, lots of (nearly all) astronauts will still volunteer for this mission, in part because it's not that big a violation of the limits, at 60 vs 50 REM annual. For missions "centered" the other 68-69% of the solar cycle, the exposure will not exceed the 50 REM annual limit.
My point is, under NASA's own rules, radiation is not a show-stopper for sending men to Mars. These rules are at least 20 years old now. Why this is being ballyhooed in the press of late is a mystery to me, unless it is really being used to justify deciding not to go.
However, I would recommend that any Mars mission crewmembers not fly in space outside the Van Allen belts ever again. They will be close to career exposure limits after going to Mars, especially for a mission flown near or at solar min.
GW
I ran a little bounding-type trade study for SSTO launch. Results are posted over at "exrocketman", in the 9-24-13 article.
GW
Josh:
Don't give up yet, but my numbers say you want closer to 700-800 sec Isp than 600. I did a trade study and posted it over at "exrocketman". It's the 9-24-13 posting.
GW
Recombination does not start until you reduce internal energy, either by heat withdrawal (for some other purpose), or by the start of expansion. The recombination that happens in liquid and solid rockets occurs in the exit bell, usually upstream of the throat (effectively-frozen chemistry in the supersonic part). Most of the thermochemical codes that compute Isp or c* for you take this into account. Going at it pencil-and-paper from just theoretical energy content, you fail to account for this.
If I was going to seed the flow of an LH2 NERVA to make it more opaque, I'd inject carbon monoxide just downstream of the core. I think the hydrogen would steal the oxygen from the monoxide, leaving at worst colloidal-size soot in the flow, maybe molecular size. The effective emissivity of a solid like that is usually between 0.1 to 1 if optically dense, while the effective emissivity of straight furnace gases is usually in the 0.01 to 0.1 range, at most. Could be lower.
How all that plays into a hybrid chemical-nuclear device, I am unsure at best. But, it sounds like multiple "treatment" stages, some before the throat, some after.
Aluminum in solid rockets seems to be an additional energy source in the chamber, and a loss in the nozzle flow process. In the solids, the reduced-smoke composites had little or no metal, and flame temperatures near 4500-5000 F. With 20% aluminum, flame temperatures were closer to 5500+F, which means it added almost 1000 F. The product is liquid/solid aluminum oxide, which contributes nothing to the hot gas expansion process. You have to remember that these formulations are solids-limited (thixotropic mix "viscosity"-limited), not oxidizer-limited. They are nearly always badly-underoxidized. The aluminum replaces some of the AP.
The droplets/particles are relatively large, and there is no time for any effective heat transfer. What you get at the exit is a high speed gas jet filled with 5500+ F sand blasting "grit" that lags a bit in velocity behind the gas. It's really tough on O-ring seals, as they found out back in 1986 with Challenger's SRB joint.
If you add to that, the fact that O-ring failures are nearly always concentrated at a single point, then you can understand why old hands like me see no point to using more than one O-ring in a solid rocket motor joint. The 3-ring joint they finally came up with was even less reliable than the 2-ring joint they started with. Only reason it never failed again was that they never flew it 29 F cold again.
GW