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To make a Nextel 440/titanium rib parasol work, you need 3 things: (1) some sort of insulation of very low thermal conductivity between the cloth and the ribs (otherwise the ribs will overheat due to the conduction path from fabric to rib), (2) a high emissivity form of the Nextel cloth so that re-radiation thermally has a prayer of equalling peak convective heating during the entry pulse, and (3) a low enough ballistic coefficient to limit the Nextel cloth hot-side face temperature to about 2000 deg F (about 1090 C) to prevent alumino-silicate solid phase change with severe shrinkage cracks and material embrittlement. Or, if this is a one-shot throwaway parasol, you may let it reach about 3000 deg F (about 1650 C), which is close to its meltpoint.
Temperature-wise and mechanical property-wise, titanium is no better than low-alloy steel. It's just lighter in density by roughly half. It has very poor properties as a casting or a forging, and is not generally ductile enough for forming operations, quite unlike steel.
Any insulation between fabric and rib will necessarily be porous to obtain that low conductivity, and thus rather weak in compressive strength. How do you plan to keep the heat shield pressure loads from destroying your insulation?
Those engineering objections are why I tend to be very skeptical about flexible heat shield concepts.
GW
Transistors appeared commercially during my teen years. My pre-teen years were spent pedalling to the corner store with TV vacuum tubes to test in their tube tester. Does anybody remember those things? We had a black-and-white Dumont vacuum-tube TV from 1952. Refrigerator-sized cabinet with a 12-inch screen. It had an unshielded Klystron tube in its power supply. If you sat too close, you could absorb a lifetime's dose of X-ray in about an hour. Of course, nobody knew that back then.
My experiences with computers were entirely different from that described by Robert Dyck. There was nothing like that available in my high school. The vocational electronics guys were doing vacuum tube TV stuff. Radio was similar in those days. I went physical stuff, not electronics.
When I got to undergrad engineering school in college, there were card batch input mainframes with fan-fold printed output, or teletype interfaces with fan-fold printed output. There were no pocket calculators of any kind, only the same slide rules I used in high school. We avoided using the mainframe like the plague, whenever possible.
I learned how to program in college in Fortran II, but never learned any "job control language" (what today we would call an operating system), because that stuff was machine-specific, and thus simply wasn't taught. That lack still cripples me today.
In industry, I became really adept at programming in Fortran IV, and later one of the advanced BASIC's, called QuickBASIC, which actually resembles the Fortran 77 I never learned. I know absolutely nothing of the modern languages, although I learned how to do very structured programming in Fortran IV and in QuickBASIC. But, there are some problems that simply require "spaghetti code", such as balancing a ramjet engine.
You cannot write "spaghetti code" in the modern languages, I do know that about them. So I already know that my QuickBASIC ramjet cycle analyses cannot ever be translated to modern languages!
Believe it or not, I still have a way to program and operate in QuickBASIC, thanks to my son, for whom modern desktop operating systems are quite transparent. I am still doing ramjet stuff in that environment, even today, and for real clients. The hardest part was figuring out how to print out from such dead languages when there is no such thing as a parallel-port printer anymore. My son solved that problem. I can do it now.
GW
I'm an OLD guy. By the standards of most participants on these forums, I'm a REALLY OLD guy. I designed my first airplane, and my first half dozen supersonic missiles, with a slide rule, back in the 1970 time frame. If any of you know what one of THOSE is. For 300 years previous, the slide rule WAS our calculator.
It's what we designed the Saturn 1 and Saturn 5 with (and their engines), and what we designed the Apollo hardware with (and its engines), by the way. As well as the X-15 and the SR-71 and the B-70 and the F-100 (and their engines).
By the time of ASALM-PTV, I was using a calculator to do what I formerly did with a slide rule. Think Mach 6 on airbreathing propulsion (NOT SCRAMJET!!!!!). We set that record BY ACCIDENT in 1980 with an ASALM flight test, and it stood until 2004. Very little of that effort had anything to do with computers (by which I mean mainframes, there were NO desktops in those days).
Because of that history and my age, I will never be comfortable "googling" things first. My comfort zone was, is, and always will be, hardcopy paper. Sorry, at my age, that ain't gonna change. Period. You guys gotta make the difference there.
I see by the curves in the previous post that there really is a significant effect of partial pressure Pp on flammability, it's NOT JUST all % oxygen. Hard to argue with real data, ain't it? Real life is a bitch, most of the time. But the slopes of those curves are low, so it really is MOSTLY % O2. Just not ALL % O2.
Once again, the "bends" experiments were run at P's > 1 atm. NO experiments were run at P's < 1 atm. "ASSUME" is spelled the way it is because it makes an "ASS" out of "U" and "ME". Dare trust nothing, as Hercule Poirot would say. So, run the damned experiments at P < 1 atm. Find out "for sure". Quit extrapolating. That will eventually kill people.
And THAT is why the X-15 and the SR-71 and the Saturn-1 and the Saturn-5 and the F-100 and the B-70 were as successful as they were. We simply went and found out what we needed to know. We did whatever it took to find out. We did not ASSUME we knew anything at those conditions. Extrapolating outside your database is nearly always fatal. THAT's our history.
GW
Yep, I think it's a real handsome contribution to the technology. I do recall seeing a free-piston version among some of their early feasibility demo hardware.
One of the things they noticed with their earlier pumps was that there was plenty of increased crank rpm capability to handle much larger flow rates. That is exactly where they went the higher-thrust applications, like their LOX-RP1 Lynx.
The "biggie" with the ULA engine project is two-fold: (1) handling LH2, and (2) scaling to much larger sizes (25,000 to 50,000 lb thrust). So far, they've demo'd the hydrogen.
GW
I don't exactly know what "145 to 220 fsw" means, "fsw" being jargon for something. But, I'm guessing it means "feet of sea water" since the research had to do with deep sea diving. Call it a good educated guess.
Assuming it's feet of sea water, 145 ft submergence at sp gr gravity = 1.025 leads to a gage pressure of 64.2 psig, and a total or absolute pressure of 78.9 psia. Ratioed (absolute not gage pressures!) to surface conditions gives a minimum decompression total pressure ratio of 5.368:1. Lumping all the inert gases together at 79.1%, the r to surface total pressure is 4.25, way above the recommended pre-breathe r limit of 1.2 for nitrogen. OF COURSE the rats got the bends. How could they not?
If we assume a 1 atm habitat pressure, that same overall pressure ratio would be talking about a 2.74 psi oxygen suit. I don't think we're really talking about going that low in suit pressure, but some of us would like to use 3 psia O2. But use it (2.74 psia) anyway. Pp N2 in the hab is 11.62 psia. The r to the total pressure in the suit would be the same 4.25. OF COURSE the astronaut would get the bends if he didn't do some nitrogen blow-off time!
That's part of why "they" want an 8 or 9 psi suit at NASA. For that, r = 11.62 Pp N2/9 psi suit O2 = 1.29, which is still a bit on the high side. You still need at least several minutes of blow-off time, you can't just don, seal, and go straight out the airlock at low O2 suit pressure.
The REAL question here is how fast can we safely blow off the inert mixture vs straight nitrogen?
If the hab atm is 20.9% O2 and 79.1% N2, the effective dissolved Pp of N2 in the blood is only a little less than that in the hab atm at 11.62 psi. If on the other hand you used a 3-inert-gas mix at equal parts, each individual gas's effective dissolved Pp in the blood is now 33% of the straight N2 value, or 3.87 psi each, although their total is still 11.62 psi.
Diffusion physics says the driving force for diffusion rates is the difference in Pp's, which in the suit would be 0's for the inerts. The risk of smooth outgassing vs fizzing (bubble formation) I believe would be driven by higher Pp differences, less at lower values. Those differences for the individual gases are lower with the mix than it is with the single inert gas. Now, other interactions may intervene, to make the total Pp difference of the mix more important to bubble formation than the Pp difference of each gas individually. That would be the point of re-running the experiments.
It's just that we do it on the low side of 1 atm, when they were done on the high side of 1 atm before, and it's not necessarily the same. Assuming it is the same is a bad idea. That's why "assume" is spelled the way it is, because of what it makes out of "u" and "me".
The cited article was unclear about that. For the high-side-of-1 atm experiments, they saw some differences in the gases, but not enough to really do anything but lump them together. NOBODY KNOWS the interactions on the low side of 1 atm. That's what we're arguing about here.
And THAT is why I recommended re-running the experiments at pressures from 1 atm down. Find out for sure what really works and what does not. It would be cheap and easy to do so.
GW
I don't know the details of the pump drive process, they keep that very proprietary. But in broad outlines, it's a piston heat engine driving piston pumps with a common crankshaft. The heat engine derives its heat from the rocket's waste heat. It's tied into the regenerative cooling somehow, I believe. But the details are unknown to me. They say they avoid the gas tap-off losses that some turbo-pumped designs suffer.
The piston pump assembly and heat engine are actually quite lightweight. More or less comparable to the turbo-pump assemblies on more conventional rocket engines. A lot of the casing is aluminum, kind of like an air-cooled VW motor. If there's an inert weight penalty, it's not very much at all. There's some photos on their website.
Besides the physical robustness of piston/cylinder hardware vs spinning turbomachinery in thin pressure vessel shells, there is another very distinct advantage: extremely-rapid "spool-up" and "spool-down" times. Almost all liquid engines have to be ignited at low thrust/low propellant flowrate conditions to avoid explosion upon ignition. Then they have to "spool-up" to full thrust for launch, which takes a time usually measured in seconds. The piston-pumped engines also have to light at low flow rate, but the spool-up only requires 1's-to-10's of microseconds. Throttle-up/throttle-down response is similarly fast.
That's about all I know, really. I've seen their hardware up close and personal. It's very impressive, what they do and how they do it. Their smaller engines, like the ones that went in the airplanes, have containment built into the thrust chamber assembly, so that a burst chamber won't down the airplane. That does cost some inert weight, but for utter safety, it's well worth it, in my opinion.
GW
This usable-atmosphere thing is mostly an arguing-in-a-vacuum thing here on the forums, because the proper experiments don’t seem to have been run. Navy deep-sea diving experience is relevant, yes, but focuses on high total pressures, not low. So it’s not a perfect fit.
My latest blog posting posits a habitat atmosphere near 1 atm total pressure and 21% oxygen (which answers the fire hazard thing, so that issue just went away). But, I do it with a mix of diluent gases instead of just nitrogen, to cut down on individual diluent dissolved partial pressures in the blood. It is those partial pressures that drive smooth vs bubbly outgassing as atmosphere composition and total pressure change, that’s just the physical chemistry of diffusion. And thus that’s a falsifiable scientific hypothesis to test.
I’m recommending lower suit oxygen pressures to make more, and more varied, suit designs feasible, which flies in the face of current practices, yes. But, all such suit design approaches have a place and an application. The notion of one-design-fits-all is a self-limiting dead end. It always has been, in all other walks of life! To expect differently here is the insanity that it really is.
Bureaucracies have historically proven to be very fond of self-imposed insanity, though. No surprises there. Only an opportunity to do something different, it might well be better. Much of human history proves that.
One could start in an ordinary chemistry lab with simple bell jar experiments. Use brine of blood salinity as a surrogate for blood. Simple open-top beakers. Expose the brine to 1 atm and a proposed gas mix under a bell jar long enough to equilibriate. Then quickly place it in a bell jar of oxygen only, and reduce that bell jar pressure quickly to proposed suit pressure levels. The brine sample either fizzes or it doesn’t, and it’s visible through the glass bell jar for all to see (simple go/no-go outcome). You can even be compulsive about it and quantify the decompression rates to prevent fizzing.
If an atmosphere mix, or set of atmosphere mixes, doesn’t fizz at quick decompression times, then that’s what you are looking for. Try it (them) with mice instead of brine samples. Either the mice become ill with the bends, or they do not (another simple go/no-go outcome). Check the stats on this over a very large sample, for confidence.
If it bears up, try it experimentally with human volunteers in ground test vacuum tank chambers.
If it works on the ground, try it in space on the ISS. One could temporarily modify the atmosphere in the one module where the EVA airlock is, to see if pre-breathe time can be dropped by using a multi-gas mix in an all-up demonstration. Existing suits can be operated at the lower experimental pressures just as easily as the normal higher pressures. This need not wait for a new suit design.
There I went and wrote you the outline of a program plan to find out what we can and cannot do, and whether we really need high-pressure suits with complicated and difficult glove designs as the “only feasible thing to do”. No new suit designs, no new facilities. No expensive gravy-train R&D programs. Just get the job done, quickly and efficiently, with what we already have.
Does anybody have bell jars and bottled gases? You could run the initial feasibility tests yourself!
GW
This was a press release recently from XCOR. The XR-5H25 engine is a subscale demonstrator for a full scale engine that could power the Centaur upper stage on Atlas-5. This subscale demonstrator has 2500 lb thrust. Previous releases indicated the full scale design is 25,000 lb thrust, with growth potential to twice that value.
This press release is about the very significant milestone of operating a LH2-LOX engine live-fire on a piston pump, not a turbopump assembly. (Orbital Sciences has “fingered” a turbopump failure as the prime suspect in their recent Antares launch loss.) This XCOR piston-pump approach is the same manned-aircraft-safe pumping design approach that was in the Rocket Racer rocket-powered airplane some years ago.
That was a very reusable rocket engine, and that's the title of this thread, essentially. I don't know of anybody else working on something truly innovative like this. More power to them. The sooner "everybody" does things like this, the sooner fully-reusable rockets safe enough to fly men will be a reality. -- GW
Mojave, CA, November 20, 2014 – XCOR Aerospace today announced it has completed the latest test series for the liquid hydrogen engine it is developing for United Launch Alliance (ULA). This is an important milestone in the long-running LH2 (liquid oxygen and liquid hydrogen) program. It is also a step toward running the engine in a fully closed cycle mode.
In its most recent milestone, XCOR successfully performed hot fire testing of the XR-5H25 engine’s regeneratively cooled thrust chamber, with both liquid oxygen and liquid hydrogen propellants supplied in pump-fed mode, using XCOR's proprietary piston pump technology.
“This test marks the first time liquid hydrogen and liquid oxygen have been supplied to a rocket engine with a piston pump,” says XCOR Chief Executive Officer Jeff Greason. “It is also the first time an American LH2 engine of this size has successfully fired liquid hydrogen and liquid oxygen together in pump-fed mode. We are happy to be making solid progress on the engines. This will also bring us to a new phase in our plans for orbital flight.”
"ULA has an ongoing effort to develop rocket engines for our next generation upper stage, and we are thrilled to see that progress continuing with XCOR," added ULA Vice President George Sowers.
Upcoming test series will fully integrate the nozzle with the engine and piston pumps. Fully closed cycle testing will follow soon afterwards and will complete the sub-scale demonstration engine program.
The XR-5H25 engines are being developed under contract to ULA as potential successors to the Delta and Atlas series upper stage engines currently used. These engines will also help power orbital launches.
In post #22 above, I was not suggesting solutions to anything current, in fact I said so, regarding fast, efficient, high-throughput electrolysis. I merely suggested that the very desirable storage properties of water suggest that we work on the technologies needed to successfully employ it, as something for future application (not immediate).
There are two technology paths that "jump off the page" at you for doing this: electrolysis for LOX-LH2 chemical engines, and direct use of water in something nuclear. Neither are ready now. The "nail in the coffin" for pursuing these kinds of things as long-term technology developments is that we are finding there is water "everywhere" out there, making "fuel-as-you-go" feasible, if the extraction and purification challenges can be met. And they can.
Ice mining off-world doesn't scare me. We have mined coal and other minerals for centuries. That's where you start. But it takes several years to adapt the technologies so they will be available when we need them! I don’t see that effort going on, and that’s what bothers me. When we go to Mars, we’re going to need it, sooner or later. Probably sooner than we ever thought.
Picking through post #24, Impaler also has a long-term technology development suggestion, although that was not his purpose when he said what he said: we really need much better electrical power supplies than anything imaginable today. That would also make electric propulsion methods look much more attractive for faster travel possibilities.
Of course there are more technology needs for successful interplanetary travel: such as habitat modules, closed-cycle or at least mostly-recycling life support methods, experimentation with amounts and with ways-and-means of artificial gravity, and better ways-and-means of radiation shielding.
A part of the very purpose of a government agency like NASA is to work on things exactly like these, so that they are ready to employ when the time comes. But are they? No. Not to any significant degree. Follow the money! See for yourself.
NASA says they want to put men on Mars in the 2030's. But they also actively promote the public fiction that a crowded-as-a-phone-booth and relatively-unshielded capsule is going to carry crews there in zero-gee, for 8 months one-way trip time, on a 2.5 year mission, when even the best "astronaut food" lasts 18 months at most. Right. And none of the things listed above is being seriously funded now, in order to be ready after 2030, when we will need them.
So pardon me if I disbelieve NASA projections and PR, and pardon me when I hold up my hand saying "the emperor has no clothes" when NASA is re-building moon trip hardware, not building Mars trip or asteroid-trip hardware, and yet calling it hardware for going to Mars. The fact that they have to re-create the capability to send men to the moon tells you how far we have sunk in the last 40 years.
And the disparity between NASA's version of moon hardware and what is really needed (see Bob Clark's "budget moon mission" ideas, etc.) to put men on the moon tells you an awful lot about the non-transferrence of the experience and engineering art of the Apollo era folks to the current generation.
Outfits like Spacex and Bigelow are coming much closer to working on the right stuff to go to Mars, but even they are focused on the initial steps near LEO right now. Musk's "Mars Colonial Transport" is an idea, not a real design, right now. They don't have time to work seriously on it until (1) Falcon-Heavy is flying reliably and often, and (2) they've solved their production backlog problem.
My whole point is (and often has been, in the various threads) to point "outside the box" at the things we should be working on right now to enable real interplanetary voyaging several years from now. The sooner we start, the sooner we really get to go.
That being said, this thread title has to do with propellant transfers. No one offered a single comment (!!!!) about my post #15 above, where I suggested what we need is an adaptation of existing hydraulic quick-disconnect fitting technology, combined with receptacle (housing-in-the-vehicle-shell) design to control and contain (actually manage) the inevitable, inherent spillages at connect/disconnect operations. And these aren’t what you fear! Not with the technology I suggest.
What I and every other farmer do on the farm with hydraulic fluids, hoses, and fittings every day could be done as easily with dangerous gasoline, using the exact same fittings, although there's no need for that here on Earth. So RP-1 or any other kerosene (and likely any storable liquid) could certainly be handled that exact same way.
I lose about 1-3 cc at every disconnect, but it doesn't spray anywhere, it just drips off the fittings. The adaptation to cryogenics should not be that hard, excepting maybe peculiar LH2 itself. That last requires quite a bit of thought and experiment. But no magic, it can be done. And should have been done by now, but hasn't.
GW
I don't know why others want 4+ km/s. Different assumptions breed different results.
If I thought there was a way to electrolyze water rapidly and efficiently (to my knowledge there is not), I'd go LOX-LH2 and ship a lot of my surface exploration and return-to-earth propellant as water. Really easy to store in space as ice covered in an opaque plastic bag, and strong enough to be its own structure, for a real inert weight savings. Could be supplemented with mined ice on Mars's surface at some sites, not others, rather easily and rapidly. Just electrolyze in advance what you need for the next burn.
If there were such a thing as a water-NERVA you could skip all the electrolysis. Just melt the ice, filter out the dirt, ready-to-use. Especially in a gas-core engine, but there ought to be a configuration for a solid core that would work.
Fast efficient electrolysis and a water-NERVA sound like good development projects for outfits like NASA. Easier to do the Mars trip with them than without them.
GW
Mars surface escape velocity is listed as 5.03 km/s. That makes the surface circular orbit speed 3.56 km/s, a factor of square root of 2 smaller. That's pretty close to the circular orbit speed at low altitudes like 200 km or 200 miles. Orbit speed (without any losses or plane-change deltas is a lot closer to 3.56 km/s than it is 4+ km/s. Gravity and especially drag losses on Mars are just a lot lower than they are here.
Descent requirements depend very strongly on whether you can make effective use of supersonic parachutes or not. Below a ballistic coefficient of about 100 kg/sq.m you can, above it, you cannot (there isn't enough time to deploy, much less aero-decelerate with a chute). Somewhere close to 1.5 km/s for ballistic coefficients in the 300 kg/sq.m class is "reasonable, but by no means precise.
For descent/ascent to/from very low orbit with no plane changes, your total delta vee is in the 3.6+1.5 + loss "kitty" = around 5.1-5.4 km/s or so. If you are fueled/refueled from an on-orbit supply, you will inherently be heavy on descent. Unpleasant fact of life, that simply must be dealt with.
That's the kind of numbers I was using.
GW
Hi Terraformer:
The cost of membrane desalination includes the cost of periodically replacing the membranes when the become too contaminated. The membranes are expensive already. When you add in the cost of transporting replacement membranes to Mars (or elsewhere), boiling might actually be the better choice for off-world desalination.
Although, on Mars, extreme cold is readily available. The first ice that forms from freezing seawater is always quite fresh. Why not freeze local brine for freshwater ice, even if you have to reject over half your throughput? On Mars, you pay zero for the freezer. Your energy costs are (1) pumping liquid and (2) some heat (80 cal /gram) to melt the freshwater ice plus (3) perhaps some heat to get the brine in the first place.
GW
I've looked at the reference Impaler supplied above. The various values of acceptable R seem to be related to shorter pre-breathe and higher risk vs longer pre-breathe and lower risk, with the fundamental assumption being "it's a two-gas system". I looked for, but didn't see anything about relevant R values for a 3-gas system. So, I think that the issue of a multi-gas system is still open. That's a very good report Impaler provided, and I downloaded it for more "digestion".
Heliox for deep diving has been known for a long time. The trouble with helium is "Donald Duck voice". Argon and nitrogen get toxic if their pressures get too high, which should not be a problem for space habitats. Not sure about the others, but I'd guess neon is OK if not to much is used.
The suggestion I have offered is use nitrogen first, then as much argon, helium, and neon as you dare with each. No one dissolved partial pressure is any higher than their partial pressure in the habitat atmosphere (each a minor fraction of the total pressure), and those are just lower numbers with a multi-gas system. That has to restrict bubble formation relative to a single diluent situation, where the diluent partial pressure is instead a major fraction of the total pressure.
I don’t think the NASA guys who wrote that report considered that possibility, so their report does not rule the multi-gas approach out. The “other gases” for them were all trace constituents. As for “alveolar oxygen”, that’s what I am calculating when I allow saturated water vapor at body temperature to displace part of the “dry” habitat atmosphere (or suit atmosphere) inspired into the lungs.
GW
Sorry guys, I've been very ill the last several days. Hope that's finally coming to an end.
Impaler:
About my Mars lander studies, if memory serves, I was using just over 3 km/s for fast ascent to very low Mars orbit (200-300 km), and at min about 1.4 km/s for descent, sometimes as much as 2 for extended hover modes. I would start retro burn upon coming out of hypersonics at local Mach 3, typically near 10 km altitude. This is based on "average" Mars conditions. I'd have to go look, but I think my vehicles were typically sized to deliver about 4.7 km/s worth of ideal delta-vee. It was enough for the same vehicle to go from low Mars orbit out to Phobos and back, too. And you're probably right: a 3 ton payload is probably not enough. But I had to start somewhere.
General discussion just above:
Propellant transfer safety is all in the design of the plumbing. The various propellant combinations must have hardware sets that are all geometrically incompatible with each other to prevent mixing the wrong stuff. The actual fittings will likely resemble the ball-valve-protected quick-disconnect fittings used in hydraulic equipment. I have stuff like that out here on my farm. It does a good job, but it is not perfect: you lose a small amount at every disconnection.
That means the receptacles in which these fittings are mounted must be shaped to capture "the drip" (which means there must be gravity) at least long enough to allow its safe evaporation (requiring venting, too). I'm not quite sure how to handle that in zero-gee; where automated is probably better. The receptacles have to be large enough to accommodate both gloved hands at once (makes MCP look more attractive, doesn't it?). One hand pushes on the line, the other slides the coupling ring to make or break the connection. Same as out here on the farm. Go to Tractor Supply Co and look at the Pioneer 4000 series hydraulic fittings, or the very similar John Deere fittings they sell. Then imagine much larger.
Propellant transfer on planetary surfaces (or in orbit) is something we can do. But we haven't actually done it yet. The Russians have started. Nobody is doing cryo that way yet. But we could. You just have to think your way through the hazards and design arround them, with the flexibility in your basic design to make incompatible hardware sets for the various propellant combinations. By that I mean every fuel and every oxidizer get its own set.
For the storables, I think a rubber over-glove might help keep suits uncontaminated. For the cryogenics, it is thermal injury that is the real hazard, as evaporation will be inherent.
GW
Sorry to have been absent several days. I have been very ill. Hopefully, that is coming to an end.
Impaler:
You and I agree about SLS being a boondoggle, yes. NASA is particularly susceptible to them, as we-the-people have allowed its extreme politicization. I don't see an end to that anytime soon. I don't see NASA doing anything productive about sending men to Mars or anywhere else outside cis-lunar space until that politicization does end. So no one should be holding their breath.
On the other hand, I do know about Spacex's intent to build its own giant rocket. I just don't think they will do anything in that area for some time to come. Their hands are busy coping with an accelerating launch manifest for Falcon-9, and getting Falcon-Heavy flying, plus building another spaceport in south Texas. The spaceport won't start until the test site on Texas can handle Falcon-Heavy (an effort coming to an end, I believe), and Falcon-Heavy won't fly until they can wring-it-out on the ground at the test site. On top of all that, they cannot divert engineers to the point that Falcon-9 starts failing. No, it'll be a long time yet before they make any significant start on their giant rocket. That's why I usually don't say much about it.
GW
Impaler:
Looking at published launch prices and max payload capabilities (to LEO) for all the launchers (every country) produces a very interesting curve. The commercial launchers (subject to price competition) are all in the vicinity of $2500/lb at 10-20 tons. Spacex's published data for Falcon-Heavy falls in the vicinity of $1000/lb at 53 tons. If you extend/extrapolate that trend, the unit price should be at or under $500/lb at around 100 tons. It's probably nonlinear, but still, way under $1000/lb.
Yet NASA's best estimates for SLS fall in the vicinity of $2500/lb, which you have to take with a very big chunk of rock salt. Their history of estimating costs is abysmal. That's the difference between a government-run design and commercially-competitive designs.
From all that I conclude SLS will never be cheap: what's the cost difference between 100 tons launched to LEO at $2500+/lb on SLS, and 100 tons launched in five 20 ton chunks at $2500-2400/lb on ULA and F-9 rockets, and docked together? Not much. Not much at all.
The noncompetitive historical government missile-derived launchers like Titan-IV fall 4+ times higher than the commercial curve. The one-and-only routine spaceplane (shuttle) delivered about 15 tons for a $1B launch price: around $30,000/lb.
Skylon might (maybe someday) do better, but its inherently small payload fraction will preclude “cheap” commercially-competitive use as a tanker, which needs to be lots of tonnage. Spaceplanes will have a niche for small payloads only. Like crews.
So Spacex, ULA, their foreign counterparts like Ariane, and maybe(!!) NASA's SLS are what you have to work with for missions to Mars or anywhere else outside cis-lunar space, now, and for the foreseeable future. Period.
And, you have to consider achievable flight rates. Falcon-Heavy will be cheaper, yes, but not very many of them will be available for a long time yet. Spacex is having trouble keeping up with the Falcon-9 business it has already booked. This shows in slipped launch dates.
There might (or might not ever) be an SLS flight rate of 1 per year.
ULA actually has the highest flight rate, primarily because it has two rockets built in two different factories. But that won't last. ULA will soon have to replace the Russian engines on Atlas-5, due to the bad international relations we will be having for the foreseeable future.
The space station was a $10B+ item precisely because most of it was assembled from 15-ton items launched at near $30,000/lb. The same job could be done commercially today with 15-20 ton items at $2500/lb, which would slash the launch price by more than a factor of 10.
Looking at that overall picture of what's available and flight rates for the foreseeable future, I don't see any future in waiting for SLS which will most likely be several times more expensive than we have right now, or for the inexpensive Falcon-Heavy, whose flight rates cannot sustain a large effort. To me, that picture says use what you have now, or you'll never fly at all.
So, I disagree with your statement that current commercial rockets are "staggeringly inefficient". They are quite demonstrably the most cost-efficient means that we have. Further, smaller launchers will never be cheaper per pound. If you add Falcon-1 to the curve, it's above $4000/lb at 1 ton size. So bigger really is better, but you must use what you have. And the commercial unit prices are simply the very best available, by at least a factor of 4.
GW
Hi Spacenut:
What you posted about landers, split missions, and SLS capabilities is intriguing. I cannot say I have digested and understand it all yet.
To answer your question, the 3-ship fleet in my posting is an orbit-to-orbit manned transport and two unmanned vehicles. The manned vehicle has the habitat, engines, and core structure which are are reused for other missions. It never lands anywhere. You just add loaded propellant tanks and go. The other two unmanned vehicles are propellant tank farms pushed by a Mars lander. That's unmanned one-way stuff, and can be sent first. This fleet has to rendezvous in low Mars orbit. This is actually the old 1950's picture of a manned Mars mission.
Landing operations with single-stage reusable landers draws on the in-orbit propellant tank farm. The landers are more-or-less conical capsule shapes, and they are the surface habitat for crews on the surface. One crew stays in orbit doing science while the other spends about a week at a site. Rotating surface crews provides the safety of rescue capability with the other lander. You do this during about the first half of the stay at Mars, visiting as many sites as can be supported by the tank farm you have bought. More money is more real exploration, pure and simple.
In the second half of the stay at Mars, all the crew goes to the surface with both landers at the "best" site for surface ISRU propellants, etc. If ISRU propellant manufacture works well enough, it can support suborbital trips to yet more sites to explore. If not, at least the baseline mission gets accomplished. Between the two landers and maybe a couple of inflatables, there should be plenty of living space for a few months on the surface.
The crew ascends in the landers to low orbit, for rendezvous with the tank farm and orbit-to-orbit vehicle. Landers and empty tanks are left in orbit for use later by the "next mission" whenever and whoever that might be. The orbit-to-orbit transport returns to LEO for recovery and reuse. I include small free-return-capable entry capsules in that vehicle for an engine failure accident returning to LEO. But they should never have to be used.
The orbit-to-orbit transport as I envision it is a baton shape with the hab at one end and the engines at the other. It spins end-over-end for artificial gravity in the habitat section. I always use chemical rockets for departure and arrival burns. One could add electric thrust during transit to shorten trip times, though. Locate it at the spin center, and direct its thrust perpendicular to the spin plane.
GW
Silanes produce slag, which does not "mix" well with turbomachinery. or even piston/cylinder stuff.
GW
Nobody ever said we need 100 launches to go to Mars! But, we might come close, if we mount a really ambitious mission. So what?
A fleet of three ships, one manned, two unmanned, each around 500 tons, might do the mission, in a very BIG way. That's 1500 tons total to assemble in LEO. And that's a REALLY BIG mission!
Say a quarter of it (around 375 tons) is Falcon-Heavy stuff at 50 tons a whack. That's about 9 Falcon Heavies. Say another quarter of it is Falcon-9 stuff at around 12.5 tons a whack. That's 30 Falcon-9 launches. The other half (750 tons) is launchable with Atlas-5 and Delta-4 at around 20 tons a whack. That's about 38 rockets from ULA.
That's a total of 77 launches to assemble a REALLY big Mars mission. And I do mean REALLY REALLY big! Figure launch costs at an average of 2500/lb ($5.5M/ton) for the ULA and Falcon-9 stuff, and nearer $1000/lb ($2.2M/ton) for the Falcon-Heavy stuff. You must fly near-capacity to get it that cheap. Even so, that's a launch price of $825M for the Falcon-Heavy launches, and about $6188M for the rest. Total launch price is $7013M = about $7B for direct launch costs, if you fly near-capacity on every rocket.
In a well-run program, launch costs OUGHT to be around 25% of your program cost. In a poorly-run program, maybe 10%. Use that 10%. On the $7B.
You get a price in the vicinity of $70B to put a wastefully-inefficient project in place, for a HUGE expedition to Mars. You really ought to be able to do it for something closer to $30B.
And I'm talking reusable orbit-to-orbit transports, reusable landers, and conducting around a dozen or so different landings at sites all over Mars, in the one single trip, plus establishing the core of a permanent base at the "best" site. And maybe even flying a lander to visit Phobos while we're there.
NASA estimated a far less ambitious trip (one site, "flag-and-footprints only) at $450B a few years ago. Bah, humbug, and what-a-load-of-BS that is/was.
It's the Spacex's and Virgin Galactic's of this world that will do this right. Not governments.
GW
PS: ANY rocket can send something to Mars, even the ancient original Atlas or Thor of 60 years ago. They did, in point of fact.
RobertDyck and Impaler and the rest of us have been arguing about habitat atmospheres and oxygen suit pressures. I put a posting up on exrocketman a couple of days ago about this topic, and I just updated it today. In that update, I created a multi-gas habitat atmosphere at 21.3% oxygen and almost 1 full atm, with a corresponding low O2 suit pressure that would apply to full pressure suits or to mechanical counterpressure suits.
If my pre-breathe factors are anywhere near right, this can be done with zero pre-breathe time. The suit pressure is low enough to make full pressure suits more supple, more like those of about 1964. It also makes the MCP suits immediately feasible, whether you look at the old 1968 stuff, or Dava Newman's stuff at MIT.
Check it out, I think the results will startle you. As of today, the blog posting on atmospheres and suit pressures is the one at the top, first thing you see.
The problem you face with launchers for stuff to be built in LEO is that not all the components are big and heavy. Especially if assembly is by docking stuff together, very little of it would be 100 ton class, only the finished, assembled, and fueled item. That means current launchers (all of them, Spacex and ULA) are appropriate for a lot of the stuff we would want to send to orbit, if not the entirety of it. Depends on how much docking assembly you end up doing in your design.
Launch is only "cheap" if the vehicle flies at full load. You pay the same launch cost for a given vehicle, regardless how light your payload might be. The current fleet is characterized as about $2500/lb, but that figure goes up by big factors if you fail to fly full load. That per unit payload figure is simply launch price divided by max load. If you fly at half load, your unit price is twice as high ($5000/lb).
Launcher selection is thus set by two things: (1) mass of payload vs launcher capacity, and (2) width of payload versus diameter limit for the launcher. That last is true whether you ride inside a shroud or "naked" on top of the rocket. It's an aerodynamic stability limit during ascent.
If NASA ever gets its act together and gets SLS flying, then we will have a launcher fleet that can carry 10-20 ton objects (the current ones from ULA and Spacex), 53 tons (Falcon Heavy), and 70-130 tons (SLS). I'm not sure SLS will ever be remotely affordable. The others already are.
GW
The weight statements and velocity requirements were in the posting as tables and figures, for the 4 vehicle concepts investigated. These vehicle go directly from hypersonics to retro thrust on descent. They come out of entry hypersonics too low to bother with chutes. If you click on a figure, I think most browsers will show you a greatly enlarged figure, and you can look at them all that way.
The descent delta-vee is quite modest, even for retro-thrust landing. Most of the delta vee is for ascent. I assumed the same payload mass both ways, just so ascents could bring a little ISRU propellant up from the surface. The mass ratio covers all the two-way delta vee in a single stage vehicle, intended to be refueled and flown again.
These were for a low Mars orbit-based mission. Being only a concept study, these didn't have all the margins and safety factors we might really need. The odd thing was, they all came out about the same physical diameter and height, in spite oif different gross weights. The common factor was payload mass to be carried: a few tons.
This kind of vehicle could support explorations of multiple sites from orbit, or suborbital hops from surface-to-surface if based and fueled there. Construction of a semi-permanent based would require bigger payloads and bigger vehicles, of this same basic layout. Colonization missions would require something bigger still.
If easy-to-mine ice deposits can be found to make LOX-LH2 on the surface with solar electrolysis, then the engines of choice would be fueled that way. The "smart thing to do" would be to use these or similar design concepts as a startpoint for a common airframe that could be re-equipped at any time with the tankage and engines for any of the propellant combinations that proves best suited to surface conditions at a base.
GW
OK, if flammability depends most strongly on vol % oxygen, even at lower pressures, then add the argon. What would you think of 29% oxygen in nitrogen-argon, at about 12 psi total? If I use that, I can use a low-pressure less-restrictive spacesuit on pure oxygen, and do it without a pre-breathe time, assuming the factor 1.2 Pp N2 / suit P O2, and lower similar factor (0.86, ratioed by molecular weights) for Argon.
edit 11-17-14: I ran the numbers in a traceable and organized way, and got 32.68% O2 at 8.16 psia total. This assumes a pre-breathe factor of 1.2 for N2, and 0.86 for Ar. It's based on in-lung wet Pp O2 from a spacesuit designed to mimic 5000 ft altitude, then factored up 10% for leaks.
Ran it again for a 4-gas system (O2 N2 Ar and He) and got 27.03% O2 at 9.68 psia. This used a slightly larger argon factor of 1.0, and a wild guess factor of 0.5 for helium.
GW
The difference between my idea and Apollo is that I do what Apollo did (visit multiple sites, except I do it in one trip) PLUS I pick the best of the explored sites, and base the rest of the 400-500 day stay there. That part looks more like the proposals I see debated here, and is something Apollo never did.
I've run some design sizing studies on landers, and posted some of what I found over at my "exrocketman" site. The "consensus" of these studies is that one stage reusable chemical landers are rather feasible for trips between low Mars orbit and the surface, but not for high-orbit basing. The basing orbit needs to be low: around 200-300 miles. The initial ones visiting multiple sites would be fueled by propellant brought from Earth. Propellant from Earth also fuels the second-part basing-on-the-surface trips.
If ISRU propellant proves successful (during the second-part stay at the selected base site), then you accumulate propellant on the surface there. It can support suborbital trips to yet more sites. If on the other hand ISRU propellant doesn't pan out, you still did your basic mission.
The trip home from low Mars orbit uses propellant brought from Earth and parked there. That way a return is assured, no matter whether ISRU propellant works or not.
The velocity requirements for a trip to Phobos and back from low Mars orbit is pretty close to what the lander can do. If you decide based on ground truth that suborbital trips to new sites aren't so very attractive after all, but you are accumulating ISRU propellants, then ship it up a few tons at a time to low Mars orbit with your landers. It could support a Phobos trip, or be left there in orbit for the next expedition to use.
The lander is just a big conical capsule shape. Low density ceramics would work rather well as a reusable heat shield. EDL is just hypersonics going straight to retro thrust landing. The ballistic coefficients are just too high to bother with chutes. You inherently come out of hypersonics at low altitudes (5-10 km).
GW
From what I read on Spacex's website, the payload shroud diameter is the same for Falcon-9 and Falcon-Heavy. Only the payload mass (and presumably shroud length, but I am not sure about that) are different. Why not build landers in LEO from pre-fab components shipped up with these boosters. Plus, anything a Falcon-9 can fling, so can Atlas-5 and Delta-4. No need to suffer problems with flight rates of the companies supplying the services.
Make the lander chassis out of steel framing members (for strength to weight, plus absolute strength and toughness). Have it fold up like a trundle bed frame, so it can ride up inside a payload shroud.
Do the same thing with the heat shield panels: do them as separate segments, and fit these into a payload shroud or two. They unfold, and caulk together only once. Same for the outer aeroshell panels.
The cabin and propellant tankage items are pressure vessels. These can be shipped up in a series of payload shrouds limited more by thrown weight than by volume, as none of these need be anywhere near the heat shield diameter of the lander.
Fueling on orbit requires more "payload shrouds" as tankers. Or, these could be tanks that ride "naked" except for a streamlined nose cap, on the front of the rocket. Dock such tanks to your lander, and make the connections. The lander can push itself and a huge swarm of tanks to Mars one way, to support a bunch of landings, from low Mars orbit.
Another lander plus a swarm of tanks plus a human habitat could be the manned orbit-to-orbit transport, except that's a waste of a good lander. Just use the lander engines on a lander chassis as a propulsion module docked to your orbit-to-orbit transport.
The mass of your fleet times launch prices between $1000/lb and $2500/lb is your launch cost, which in a well-run project ought to be around 33% of your total program cost. We're talking at most a few thousand tons in LEO here. Depends upon how big an expedition you wish to mount. Closer to 0.5-to-1 $B than any of the other numbers I have seen (especially that notorious $450B NASA ran up the flagpole several years ago). But the key is on-orbit assembly using commercial rockets that already serve other purposes. That way, you don't have to spend all your money developing gigantic rockets that can only serve one purpose.
Sure would be nice if we had a supple spacesuit to do this kind of on-orbit assembly, wouldn't it?
Sure would be nice if NASA was doing these things instead of spending $T's on SLS/Orion, with no space hab and no landers.
GW