You are not logged in.
Standing vs lying down in Earth's essentially-constant surface gravity field (relative to human dimensions) induces a slight difference in blood pressure head-to-toe standing that does not happen lying down. We did evolve to cope successfully with this. That gradient is linear, and essentially models as dP = h*g*dens, the buildup of fluid pressure with depth, based on g*dens the weight density, with a constant value of g. The gradient da/dheight = 0 for this field.
If "g" also varies with height because of a strong gradient da/dr = 2*r*w (because a = r*w^2) from using short r and high w, the variation in blood pressure head-to toe is both nonlinear and more pronounced. We did not evolve to cope with that.
Because it is a change from what we evolved in, it will have some effect on health for long exposure times (on the order of 2-3 years), just at an unknown level. That may or may not be a significant effect. My point is, no one knows, as regards 2-3 year exposure times. Short centrifuge experiments simply cannot address long-term effects.
To reduce the risk of artificial gravity gradient effects upon long-term health, the obvious choice is just reduce the gradient. If you hold a constant and choose a longer r, w will be much smaller, because of the w^2 effect. Even though r is larger in the gradient, the much smaller w effect dominates, reducing the gradient.
This is independent of the choice of a. Only the final design dimensions differ. I see some on the forums think 0.3 gee is enough, in contrast to the 1 gee I have looked at. You may be right, I do not know.
GW
Hi Spacenut:
From a physics standpoint, there's no effective difference between "real gravity" and centripetal/centrifugal acceleration, excepting the gee gradient along the radius. Grossly speaking, acceleration is acceleration, it quite literally doesn't matter how it is produced.
The gradient effect leads to people blacking out, if you spin too fast at too low a radius. It's essentially a tidal effect upon blood pressure when standing or sitting-up. The gradient of which I speak is the derivative of centripetal acceleration with radius of a = r w^2, which is 2 r w, where r = radius, and w = spin rate (consistent units, of course).
Basically, you want very little difference in gee between your head and your feet, when standing. I have yet to see a trustable criterion for that, but I'm sure there already is one, somewhere out there in the biomedical field.
From a practical standpoint, there are real limits on the spin rate that can be tolerated without deranging the balance organs in the ears. People can be be "trained" (acclimatized, really) to high spin rates on the order of 10-20 rpm, yes. No one knows for how long this can be maintained, though.
Untrained "ordinary folk" can tolerate 3-4 rpm pretty readily, and for very extended periods of time, essentially indefinitely. That's just experiences since WW2 talking. That long-term experience I trust. The higher figure, I do not trust.
I would suggest from evolutionary history that we design for 1 full gee in the daily work stations, which is 56 m radius at 4 rpm. If you believe (and it may well be true) that 0.33 gees is "good enough" to be therapeutic in terms of microgravity diseases, then you can get away with 19 m radius at that same 4 rpm. Neither radius is a "killer" in terms of practical design, unless you are way-too-wedded to traditional spacecraft design concepts.
Sleeping quarters do not have to be located at that full design gee level. The bed rest studies suggest, among other things, that when people are prone, their health derives no benefit from any gee, 1 gee or lower. What that suggests is that spacecraft sleeping quarters and supply storage areas be located closer to the rotation center, while daily work stations be located farther out, at the full design gee. Simple enough.
GW
Robert Dyck:
I do understand the aspiration to be an aerospace engineer. I shared it, it's just that I did so decades ago. I am now too old for any "established" outfit to consider hiring me (the fatal age discrimination sets in at age 40-45, and I am now 64). That's why I am now a teacher. And an occasional consultant.
Further, like your experience, down here all across the US nobody wants anything but a young kid they can under-pay. Young kids like that have "book learning", but zero practical experience. That's been true since the mid-1970's. Quality and ability have no value anymore in the US aerospace industry, by-and-large.
It's the practical experience, not the book learning, that enables engineers to do "the impossible". Doing "the impossible" is what has to happen in order to send men to Mars, or anywhere else beyond cis-lunar space.
Exceptions to the general age discrimination are very rare, indeed. Even Spacex employs no one over age 45 (basically Musk's age). None of the things I did has proven to have any value in today's job market. No matter how sophisticated, no matter how rare the specialty. (And I had many very-deep specialties, not the least of which was ramjet propulsion.)
GW
The yellow "fireball" at the base of the rocket was too big. It wasn't just image "bloom" on the camera. Something apparently cracked open on the combustion chamber or nozzle entrance at ignition. The hot gas leak apparently destroyed the rear of the vehicle within seconds.
Suspicion: that re-engineered Russian engine is still not a good design. The test track record is spotty enough to support that contention.
GW
Now you know why they'd prefer to capture an itty-bitty one with a robot, and relocate it near the moon. That way they can use an Apollo-on-steroids moon rocket design to reach it. They need not do anything they haven't ever done before.
The only problem with that approach is two-fold: (1) what they did before will never, ever, ever take men to Mars (or anywhere else outside cis-lunar space), and (2) after so many years, there's real doubt they can actually recreate the ability to reach cis-lunar space with men.
Too many of the mission objective and design decisions are being made by the buffoons in congress, not the engineers (who no longer have the applicable experience, since all the ones that did have died or retired).
Actually, visionary private outfits like Spacex have a better chance of actually sending crews to Mars, slim as that chance currently appears to be. Government agencies like NASA know this, and will do everything in its power bureaucratically to prevent being upstaged like that. That's just the "human nature" of bureaucrats.
GW
I know of no tunnels on the rail lines from Salt Lake City east toward Florida. The rocket whose dimensions were set by rail tunnel passage was the German WW2 V-2, not the shuttle SRB's. That's misinformation off the internet, which is notorious for that. The V-2's were shipped by rail from the underground Mittelwerk in the mountains of Bavaria. There's several tunnels on those rail lines.
The shuttle SRB segment size was set by propellant mix size, pure and simple.
There is (or at least was) another solid motor plant on the Mississippi River, operated by what was then UTC-CSD. Its products were barged to the Cape. Mix size set similar segment size limits for its products, too. The largest-diameter solids I ever heard of were 140 inches, not shuttle SRB's 120 inches. All such motors are segmented due to mix size.
All of these manufacturing and mix size limits would apply to hybrids as well. It's just a different world from the liquid rocket business.
As for segment joint design, solid (and hybrid) flames are far "dirtier" with soots and solid/molten oxides than anything the liquid boys ever even thought of, much less had any experience with. Liquid design experience simply does not apply to the solids arena. You use one and only one O-ring seal at such joints. If it passes a low-pressure (3-5 psi) whole-motor leak check test, it will hold at 1000's of psi. There's 70+ years' solid motor manufacturing experience supporting that idea.
Everything about the shuttle SRB joints was wrong from the very beginning, due to NASA's liquids-only experience being misapplied to the design decisions on a solid, and NASA's response to the Challenger disaster only made it worse. The only reason they never had another segment joint failure was they never flew soaked-out that cold again. I could go on and on about NASA's incompetence with solids, but I won't.
Solids are what we currently know how to build in very large sizes. A useful NASA program would be to scale up the hydrids to similar large sizes. That way, the abortable hybrid could replace the solid as an even safer booster (and solids can be quite safe in and of themselves). But agencies and corporations resist change. It's a profit-paranoia thing for the companies. It's a fearful-amount-of-incompetence thing for the agency.
GW
My entire point is why use a truss if you have propellant tankage anyway? Let cylindrical tank units be your "truss", and do it reusably, exactly as Tom described. That concept Tom described assumes your hab is your lander is your base. Some of the other concepts do not make that assumption, but they all work basically exactly the same way.
I know the "perfect" shape for a pressurized propellant tank is a sphere, but a very close second is cylinders with hemispherical end domes, which are more convenient for many designs, especially as the end skirts can serve many functions where these are docked together.
The experience of the last 60 years makes it quite clear there is nothing wrong with long slender cylindrical tank designs. These can be docked end-to-end to make a really long vehicle resembling the "Discovery" as depicted in the old movie "2001". You use parallel docking as well, to be able to shed empties without losing your length. Spun as a baton for artificial gravity eliminates the need for the heavy centrifuge depicted for the movie "Discovery".
The inert fraction of tankage like that is a little higher than the "perfect" sphere, yes. But the additional vehicle inert fraction added by a truss is just higher still. The more important factor here is vehicle inerts, more so than individual component inerts.
As I have said before, you want to reuse as much of your orbital transport as possible, over multiple missions. That way you amortize costs, and no one mission has to absorb it all. That's a very important facet to selling missions to stingy congresses and agencies.
The same basic orbital transport design could serve any destination in the Solar System inside the asteroid belt, with just the technology we have ready to employ today. It does require that we quit thinking rocket-and-payload, one-launch-per-mission. But that change also eliminates the need for costly gigantic rockets that serve only one purpose, which frees up funds for the missions. Chicken-and-egg, but only because we are still refusing to jettison traditions that no longer serve.
It is also a flexible-enough design approach to switch out engines for better technologies as they come along, and even to switch out tanks (not all cryogenics are the same). Let it change and modify as time goes by, just as with naval vessels.
GW
If I had to guess, I'd guess that the Los Alamos rocket was a hybrid. They used a new "energetic" fuel and a liquid oxidizer (likely LOX). That's how you achieve safety by separating fuel and oxidizer. Pressure-fed LOX eliminates turbopumps, too, at the expense of a heavy tank. The fuel probably resembled an AP composite solid propellant with the AP left out. Might or might not have had some HMX or RDX in it. Military composite solid propellants have used those as minor additives for decades now.
As for getting into space, there's nothing wrong with using solid or hybrid boosters off the pad into the first part of the trajectory. The issue early on isn't Isp, it's raw thrust, and solids are really good for that. You add them to a two-stage liquid core, and stage them off before the first stage liquid core stages off. It works really good. This is what Atlas-5 and Delta-4 really are. It's also what SLS is supposed to be. There's a reason for that: because it works pretty good.
There's a square-cube scaling law at work here because real materials are only so strong. The shuttle SRB's have given solids a bad name because of a bad joint design made worse by its putative "fix", and because they were so big as to get dinged up too badly upon ocean impact way too often. Smaller SRB's survive better, and if the joints are done right, are less susceptible to failure and less susceptible to ocean impact damage. ULA just throws them away, but they really could be reused.
Reusing the liquid core first stage is what Spacex is pioneering. That requires a 10,000 fps reentry survival. As long as the stage tumbles, airloads will crush it. Under control, they can be landed.
Liquid core second stages are an entirely different matter: a 26,000 fps reentry. Entry heating is an even bigger issue than airloads. That one may or may not be solved anytime soon.
GW
The "trick" with getting radius is to use stuff you have to have with you anyway. Don't add structure for this, whether truss or cable, that's just dead inert weight. Use the propellant and stores you already have to have anyway. Just dock those modules together into a "long slender baton" shape. Then spin it end over end.
ISS is structurally unsound to be moved anywhere of any significance. Certainly not to Mars. Plus, it's the wrong branched configuration of the docked modules to be spun up for any purpose whatsoever. See also paragraph 1, again.
If 1/3 gee works, then use 19 m at 4 rpm. That's even far easier to do than what I have been looking at for the last few years, which is 1 full gee: 56 m at 4 rpm. BTW, we may find that 1 full gee is needed for successful pregnancy. Nobody yet knows. It's a criterion for colonization ships, not exploration ships.
GW
Centripetal acceleration is rw^2 where r is the radius and w is the angular rate (spin rate). Its derivative is the acceleration gradient per unit radius, 2rw. You DO NOT want a large gradient (gees per meter of radius), because that leads to blood pooling in the feet, and blackouts. I don't really have a good criterion for that, other than the 1 gee at extreme radius in which we evolved.
I think you really want something closer to 1 gee at 56 m radius, for which the radial gradient is closer to something humans have proven themselves capable of withstanding (crudely 0.02 gee per meter). 2% is a relatively "small" number. I'd go with that.
That says you can do this artificial gravity "thing" quite easily at R=56 m for 1 gee at 4 rpm (or 0.33 gee at 4 rpm and 19 m radius), for which the gradient is da/dR = 2rw. Again, gee is proportional to rw^2. So, if you think you can get away with 0.33 gee, you can use a lower radius. But only a little lower, by the gee ratio you think might be applicable.
0.33 gee says R= 56m /3 = 19m. For the same gradient. For the same human blackout susceptibility due to blood pooling in the legs.
You don't get 19 m radii from a 4 m shroud diameter for a one-rocket / one launch scenario. Period. End of issue. You need to give up on that concept. Build your vehicle in LEO from smaller components docked together, using the rockets we already have.
And forget cable-connected designs. There are simply too many failure modes.
GW
I was talking turbine inlet temperature with respect to the Eurofighter 1800 K issue. It might be possible to get to 1800 K with a ceramic turbine, but only in small diameters, or only for very short engine lifetimes. There are no metal alloys capable of sustained, long-life operation at turbine inlet temperatures above 2200 F, even with active cooling by air, or by sacrificial liquids.
When you compare 2200 F turbine inlet temperature (exit from the combustion chambers) with the incoming unburnt air temperatures (inlet temperature), you see where the Mach 3.6 (to at most about Mach 3.8) limitation on turbine comes from. As you get to those speeds, there is no "room" left to heat the air in the combustors by burning fuel. No heat addition, no frontal thrust, period. Any altitude.
Yep, I was the one quoting speed limits for the SR-71. And with very good reason. I know the J-79's that pushed the SR-71 (and the A-12) were advertised as "air turborockets", but they really were not. What they had was unique at the time (late 1950's): an air bypass direct from about stage 3 or 4 of the compressor to the afterburner duct, and even that was limited to only 0-25% of the inlet captured massflow. It was simply impossible to isolate the turbine core from the super-high inlet air temperatures if flight speeds became excessive.
A "real" air turborocket can do that isolation function. There are none. There never have been any. Although, possibly, there could be.
BTW, the blackbird could fly steady-state higher than 80 kft, but only at speeds around Mach 3 to 3.5. Nearer 100 kft, actually.
I roughed-out a design for a parallel-burn turbine / ramjet aircraft way back in 1985 that was intended to fly over hostile territory at Mach 5 / 100-150 kft. The FBI confiscated all my design notes, even though I did this from 100% open sources. Turns out, somebody else inside the government was looking at exactly the same thing. "Surprise, surprise", -- Gomer Pyle.
To do the same mission today, I would use parallel-burn rocket and ramjet. Lighter, simpler, cheaper. Like the old Rocket-Racer aircraft, but with a big ramjet in the fuselage, and an airframe actually intended for hypersonic flight. (And it ain't what you think.) The problem isn't propulsion, it's heat protection, when you exceed Mach 3.5 to 4. Plain old ramjet can take you to Mach 6. We've already done it, nearly 40 years ago.
I see little point to flying within the atmosphere at speeds above about Mach 5 or 6. The heat transfer (and the drag) are just too high. Which is why I think scramjet is useless as a launch technology, only useful as a one-shot tactical missile technology for Mach 8 to 12 stuff. Better to go exoatmospheric by the time you want to exceed those Mach 5 or 6 speeds. Simple as that.
As for launch costs and re-usability, take a look at the expendable-booster unit costs I have posted over at "exrocketman". ULA's Atlas-5 and Delta-series all have around $2500/lb in the 15-20 ton to LEO category. Falcon-9 at 13 tons is $2400/lb. The Russian and French launchers fit the same curve of unit cost to LEO as these. That curve projects $1000/lb to LEO at 53 tons, and $1000/lb is Spacex's current projection for Falcon-Heavy! The same curve projects under $500/lb in the 100-ton class, while NASA's best estimates project $2500/lb for SLS. (Of course, NASA has long been infamous for way-underestimated costs.)
You hang the propellant margin on that for reusability, and you reduce payload. You raise your unit cost to LEO by exactly the same factor, all else being equal. But, it ain't. Pure and simple.
If you can reduce your logistics tail, then you can reduce total launch cost for a given first stage thrust size, which also impacts unit payload cost to LEO. Reduce that more than you drive it it up for re-usability propellant margins, and you can get a net reduction in unit cost to LEO with reusable systems. It's all about simplifying and reducing support manpower costs. THAT is where Spacex has pioneered so successfully, unique in the industry.
I think for smaller-mass payloads (like crews) to LEO, there is a niche for spaceplanes about the size of Dreamchaser at under $2500/lb, given the right booster rocket. I think for larger-mass payloads (like 15-50 ton modules to be assembled on-orbit), it'll be very hard to beat the simple expendable booster rocket, at about $2500/lb down to $500/lb, depending upon size. I think that will likely be true for most of the 21st century, unless we achieve some massive physics breakthrough like "warp drive". (I am not holding my breath for that breakthrough, though.)
GW
1800 K is too high to be an inlet temperature for an aircraft under Mach 2.5 max speed.
There are no alloys on the planet capable of withstanding the airloads and centrifugal stresses of turbine operation at more than about 2200 F (2660 R, 1478 K, 1205 C). Whatever that temperature was, it either wasn't K, or it wasn't turbine inlet.
Up to about Mach 6, where the ideal gas assumptions are starting to break down significantly, you can figure subsonic-duct inlet stagnation temperature rather easily from Mach number and OAT, using an air specific heat ratio of 1.4 (decelerated like that, inlet total is freestream total, and inlet static is only a little bit less). It goes up very fast and nonlinear with Mach number.
I think the results are quite telling: at 65,000 ft altitude/standard day conditions, Mach 6 air total temperature is 2906 R (2446 F, 1614 K, 1341 C). You can only heat the air in a combustor to about 4500-5000 R (4040-4540 F, 2500-2778 K, 2227-2505 C) before too much ionization sets in to derive any thrust out of the nozzle at all. Without max heating, frontal thrust density falls, even as frontal drag density is going up as the square of Mach number. (ASALM-PTV reached Mach 6 on the one flight test, barely, at lower altitude.)
1800 K for the Eurofighter sounds more like an afterburner temperature to me.
Scramjet is a way to sidestep this problem by never decelerating the air subsonic inside the engine. Unfortunately, you incur a whole host of problems doing this that have been very intractable for over 60 years. You can count the number of successful test flights of scramjet engines on the fingers of 2 hands, worldwide. It is not a technology in any way ready to apply.
The old Project Pluto engine was a subsonic "combustion" ramjet heated by a nuclear reactor instead of combustion. Its core mountings were operating about 10-20 degrees F (R) below their meltpoints at flight power. Hardware life was never very long in test, and test it they did, on the ground in Nevada, not far at all from the Jackass Flats nuke rocket facility. This thing was a low-altitude strategic cruise missile at Mach 3, under 5000 ft altitude. Its radiation trail, and the trailing shock wave, would have killed more people than the 5 megaton warhead it was to carry. No, nobody is going to fly a technology like that.
The problem with high altitude airbreather thrust density isn't the heat, it's the low ambient air density. You inlet is scooping up a large volume that just has no appreciable mass. Thrust is massflow. Low massflow, low thrust. Simple as that.
GW
If you are setting up a real colony, you are sending "mass quantities" of both people and supplies/equipment/materials to Mars. You just don't do that with anything we have been launching these last 50+ years. Those methods are too expensive, and they always will be, even if their price drops by a factor of 10.
Robots vs humans will not get you that factor-10 price break; plus, a lot of us (myself included) don't trust machines. I dislike flying on Airbus aircraft, precisely because the pilot cannot override the computerized control system. That already led to at least one fatal crash, several years ago.
A lot of the supplies and materials (especially liquid propellants) are capable of extreme acceleration. These can be shot into orbit with light gas gun technology, and recovered there by an orbital space tug. It tows them to the orbit-to-orbit transport. Other stuff, like the people, have to ride rockets or spaceplanes at low gee. That'll be under $1000/lb soon, maybe under $500/lb. But not with government fiascoes like SLS. With commercial things like Falcon-Heavy.
The orbit-to-orbit transport is the key: we can build landers for anything. Conventional rocketry works for small exploration missions, even setting up small bases. But to plant a real colony? That's a very large item (100's to 1000's of tons). It gets too ridiculous too fast to ever be practical with the stuff we have been using.
But there is something that works best/most efficiently in very large sizes (5000 tons and up), and we have known exactly how to do it since 1959 (and I'm quite sure that 1955-vintage technology can be updated to make it even better). It's called nuclear explosion propulsion. Until there is "warp drive", nuke explosion propulsion is your ONLY PRACTICAL method of planting LARGE colonies. Anywhere. Build a couple of these ships, and use them for decades to centuries as orbit-to-orbit transports of immense size, for setting up large colonies in a whole lot of places.
GW
You can forget scramjets for launch purposes. No airbreather has any frontal thrust density in the thin air above about 70,000 feet. Scramjets do not reach speeds high enough to be useful until you are simply too high along your trajectory: the air is too thin to produce the frontal thrust necessary to climb (THAT is why the X-30 "Orient Express" of about 1990 led nowhere). You're better off with plain ramjet at lower altitudes. The only real application I see for scramjet is lower-altitude / relatively short-range missile work.
GW
The "official" elevation of Boulder, CO, is 5430 feet above mean sea level according to a quick internet search. That corresponds to an altitude pressure of 24.50 inches mercury in a standard atmosphere. Sea level standard is 29.92 inches mercury or 14.696 psia. Also 101.325 KPa.
The weather barometer as reported is corrected to sea level, so 29.88 inches of mercury is what a barometer would read at the bottom of a well 5430 feet deep (down to sea level). Standard 29.92 minus 29.88 says the atmosphere was 0.04 inches mercury lower than at-elevation standard that day. Subtracting .04 inches from the at-elevation standard value of 24.50 inches says that the local at-surface (uncorrected) barometer in Boulder was about 24.46 inches mercury that day. That's 12.014 psia.
Using 20.94% by volume (partial pressure) oxygen, the partial pressure of O2 that day in Boulder should have been 2.516 psi. But that figure would be for absolutely zero humidity. The partial pressure of the water vapor at any given humidity displaces the dry air pressure by that same amount. On a very humid day, that might be 3% of an atmosphere, leaving not 12.014 psi of dry air, but only 11.654 psi of dry air. 20.94% of that reduced value is O2, for a partial oxygen pressure of 2.440 psi.
Ordinary people function just fine at lower O2 partial pressures than that, all the time. Dry air partial pressure O2 at 10,000 feet (standard) is 2.116 psi, less if humid, less if a weather low on any given day. There are people who live at 15,000, even 20,000 feet.
It's even worse inside the wet lungs, where the water vapor pressure is the equilibrium value (straight out of the standard steam tables) for body temperature, independent of total air pressure. That's about 6% of a std atmosphere displaced dry air.
BTW, I like the cold service temperature for the PCTFE plastic. That's better than anything I ever heard of before. Should work fine. Does it outgas volatiles in any way, and is that outgassing a function of external atmospheric pressure? If so, it will age (and embrittle) quite fast in the laboratory vacuum that is the Martian atmosphere. Just something else to be worried about. Myself, I dunno.
And I really like that 1.2 ratio for preventing bends. That's a really good, useful criterion.
GW
At least some vigorous outgassing like a geyser.
GW
Midoshi:
Any word on how it's going with MAVEN, or with your colleagues from India?
GW
Very interesting. Apollo/Skylab atmospheres were 3 psi O2 + 2 psi N2 for 5 psi total. That's 60% O2 at low pressure, and it was enough for adequate flash fire prevention? And the zero-decompression limit for N2 is 3.6 psi going into pure O2 in the suit? Those are handy criteria to have.
For longer-term habitations, are there any criteria governing exposure to more than 21% O2 and less than about 10 psi total pressure for successful pregnancies? I've heard there are problems with this, I'm guessing based on lab animal studies.
For the heat transfer calculations, what one has to remember here is that the clear-ish plastic domes or tents y'all are talking about are not the surfaces radiating IR away into the environment. It is the warm solid surfaces inside that are the source of the IR radiation emitted toward the environment, not the dome/tent itself. The inside warm-surface IR just has to transmit at less than 100% efficiency through your layers of plastic and interwall gas spaces.
It won't take the outer layer or two very long at all to chill very cold, and so be thermally-embrittled (vulnerable to shattering), even if everything starts out warm (which it won't). You'll have to heat this stuff as you lay it out on the cold, cold ground setting things up. Otherwise it'll break as you erect it.
GW
Hi Void:
Believe it or not, something sort-of similar to your capsule-suit idea was proposed in the mid 1950's, before there was a NASA. Take a look sometime at the old Disney "Tomorrowland" series about going into space. The EVA "suits" were powered capsules, with external manipulators operated from inside the capsule, which was a shirtsleeve environment. That's actually quite similar to your idea, and the manipulator arms presage the shuttle's manipulator arm.
Actually, I think something like that is a very good idea, particularly for construction operations like operating heavy machinery. A capsule suit becomes the operator's cabin on a bulldozer, etc. I do think there are other necessary activities that require a human protected against vacuum, but employing his fine motor skills. Such tasks would include things like fine wiring, small plumbing, and small nuts-and-bolts work. That'll never happen with a gas balloon suit.
GW
These are a few things among many that I have posted about previously. This stuff is located at http://exrocketman.blogspot.com. On the left side of that page is a navigation tool that works by year, then month, then title.
My qualifications for knowing about these things are two degrees in aerospace engineering obtained long ago before I entered industry for two decades in aerospace weapons development work, plus a much more recent PhD in general engineering, plus about 2 more decades in civilian work and in teaching. My papers presented at Mars Society conventions have always startled the audiences that heard them, and have always received positive feedback.
Advanced spacesuits and advanced on-orbit assembly capabilities:
“On-Orbit Repair and Assembly Facility” 2-11-14
“Fundamental Design Criteria for Alternative Spacesuit Approaches” 1-21-11
“End of an Era Need Not Be End of a Capability” 8-2-11
Radiation and Microgravity Solutions:
“Space Travel Radiation Risks” 5-2-12
“Space Recommendations” 4-17-10
Colonization-level Propulsion and ship design:
“About Old Project Orion – the Nuclear Explosion Drive” 3-10-10
Living “off the land” on Mars:
““Icecrete”, a Substitute for Concrete as a Building Material on Other (Colder) Worlds” 3-11-12
“Aquaculture Habitat Lake for Mars” 3-18-12
“Pressurizable Domed Habitat Structures” 6-9-12
“Aboveground Mars Houses” 1-26-13
GW
There was, is, and will always be, a big difference between what should be done sending men into space, and what will actually get done. What gets done will fall short of what should have been done in many, many ways. It always has. It always will. My epistle above is just something that voices what should be done. How this actually gets done, will be different. And not as effective.
As for the spacesuits, I vote straight MCP suits. Except not the way NASA sometimes dribble-funds Dava Newman at MIT. They (NASA) are caught up with the "everything garment" concept: that the suit is a miniature spaceship that protects its wearer from every conceivable risk, all in the one garment design. That's just wrong, and it's why a shuttle suit was 300+ pounds, almost immobilizing to its wearers, and its stiff gloves ripped off astronauts' fingernails quite frequently.
The "right" way to do MCP pivots off the work of Paul Webb in the 1960's, but done with Dava Newman's modern tailorable-property materials. This right idea is nothing but vacuum-protective underwear (!!!!) with an O2 helmet for minimal-pressure breathing (around 0.15 to 0.20 atm, not NASA'a arbitrary 0.33 atm overkill requirement). Over that, one wears only whatever conventional earthly garments one needs for protection from heat, cold, or mechanical injury. For cooling, you just sweat right through the porous elastic underwear garment into vacuum. No airconditioner or moisture control is needed.
Further, with MCP, a rip in the suit is not fatal. You have perhaps 30 minutes to get inside, or to tape up the rip tightly with a vacuum-qualified duct tape and continue working. Sew up the rip later, after you do go inside. With the current full pressure suit approach, if there's any leak, you must get inside within the seconds it takes your suit to depressurize, otherwise, you die. Period.
MCP's ONLY vulnerability is a cracked O2 helmet. That's no change from a full-pressure suit. In every other way, MCP is VERY far superior! Only tradition is holding this back. The tradition of not doing anything different from "what we did before". And even that tradition is false: MCP sort-of worked as the primitive partial pressure suits worn by high-performance jet pilots in the late 1940's, throughout the 1950's, up to about 1960. What Paul Webb did in the late 1960's was better, but never used.
GW
Well, at this point in our history, long distance manned spaceflight presents serious challenges to us, analogous to the challenges of transoceanic travel 500 years ago. It's difficult, but not impossible. Only flying cislunar is relatively easy for us at this time, because trip times are measured in days not years.
That being the case, if we decide that now is the time to send men to Mars, it will be very difficult to get them there and back again. We're looking at a round-trip journey around 2.5 years long. We will have to solve (at one level or another) the problems of artificial gravity and radiation protection for that crew. Period. Or they won't come home alive, most likely outcome.
So, if we go to all the trouble of doing that, why would we not visit both Phobos and the surface of Mars, and maybe even Deimos, while the crew is there? To go to all the trouble to cross that vast distance, and not actually "go ashore and explore", makes absolutely no sense at all to me! It violates the very oldest basic characteristic of our species and its ancestors: which is to explore.
To do all of that in one trip is going to require basing or staging out of orbit around Mars. No way around that, because you have to send lots of propellant to do the landings and the return home. That naturally leads you to an orbit-to-orbit manned transport, equipped separately (or integrally) with some sort of landers.
The velocity requirement to land on Mars from orbit (mostly aerobrake + terminal propulsive landing, skip the chute; then rocket back to orbit) is just about same as to visit Phobos from low Mars orbit. The same lander can do both. Visiting Deimos might need more propellant. But this could be done fairly easily once the lander and some propellant tanks are in Mars orbit.
Any orbit-to-orbit transport could (and probably should) be made reusable, so that it can be used for more than one mission. That way mission costs can be substantially reduced by spreading those costs around. The same ship can be used for Mars, asteroids, Venus, and even Mercury. (You wouldn't need landers at Venus or the asteroids, but a Mercury lander will be bigger than a Mars lander, being two-way rocket only).
If we don't save costs that way, it is likely no governments will ever sponsor such missions, not until we have new physics and new technologies decades (or more) in the future. Even so, it is likely that there will be one and only one government-funded exploration mission to Mars. Anything else will be done by visionary private entities (rare as they are). Most of the world's governments, including ours, have proven rather feckless over the last 4 decades as regards manned spaceflight. We've even backslid away from being able to reach the moon with men. (That’s what Orion and SLS are really for: the moon, not Mars or asteroids. I think most of us know that.)
Similarly, the lander ought the rigged as reusable and one-stage, again, to save costs and spread them around. That way, the same machines can be used many times to make landings all over the planet. No two sites anywhere on Earth are the same, it'll be the same outcome on Mars. We need to explore and try out our first-attempts at living off local resources at many sites, before we decide what can and cannot be done. Again, if you go to all the trouble to make the trip, why make just one landing? Seems kinda stupid to me. Landers like that are simply inherently larger than most minimalist designs I see proposed. But, they can do so much more than minimalist designs!
That kind of mission is thus simply not a minimalist mission design; in point of fact, it cannot be! This is not (and can never be) something you shoot straight to Mars with one or two giant rockets. This is something you must assemble by docking in Earth orbit, and, you recover it there after every mission it does fly. Because it's assembled from smaller modules in orbit, in point of fact, you don't really need a giant rocket. The ones we have already fling stuff big enough to do this job.
There's two kinds of radiation to worry about: galactic cosmic rays, and bursts from the sun. The former are a slow drizzle of stuff so energetic that we cannot shield very effectively against it (the real trouble being the secondary shower effect unless the shield is too thin to do much good anyway). Just being in Mars orbit cuts the dose in half, because the planet fills half the sky. And exposures down on the surface are much lower still.
But, the dose from a 2.5 year mission spent all out in space will fall pretty much within the annual and career-limit criteria we currently use for astronauts. That takes care of cosmic rays on an exploration mission. It’s just that the crew who does this cannot ever fly outside the Van Allen belts again.
The solar flare stuff is erratic in occurrence, and variable in its lethality, but we know it can be as deadly as nuclear war fallout (death by radiation poisoning within a few hours, worst case). It turns out that 20 cm of water is an effective shield for this stuff, and is practical for a big ship design (but not for most minimalist concepts).
Sending men on a two-way trip is simply-and-inherently a whole 'nother ballpark from sending one-way probes. But it can be done, and right now. I just think that expecting the vehicles and mission design to look like the minimalist designs we use to send the probes is just plain wrong. If it's worth doing at all, it is worth doing right.
GW
I thought maybe radiation would dominate over convection in air that thin. Thanks to Antius for running the numbers. Conduction is the other "biggie", of course. Unlike here, convective heat loss is the least of your worries.
GW
The question of whether there exists within Phobos any volatiles at all is completely unanswered.
Betting lives on the supposition that ice exists there, is therefore a bad idea, until better information becomes available. Period. What you have to assume for a first landing is that no volatiles at all exist within Phobos. Period.
In contrast, we know already that ice in one concentration or another exists within the dirt on Mars itself. That makes a direct surface landing on Mars a surer bet than a surface landing on Phobos first. Period.
But, even so, it's still a poor bet, even going to Mars first. What we found with the Mars Polar Lander was tiny lenses of ice buried under several inches of soil, not the massive deposits that would be so much easier to mine and utilize.
So, the first mission to the vicinity of Mars (including Phobos) can assume NO utilizable resources, until verified in situ, after the fact of a landing. That means, you must bring everything you need from Earth, period, on that first mission. No exceptions. Nothing else is ethical.
If you find something local that you can use, well, that's "gravy". But the odds are you won't! Period. Based on what we know so far.
GW
Hi Midoshi:
That's great news! My congrats to the entire MAVEN team! I had spotted a news release on the internet that said MAVEN had made it. Thought I would check here, and there was your notice.
Again, congrats!!! Well done.
GW