Yet another Mars architecture

JoshNH4H · 2013-07-02 14:08:34

I would imagine that crew and cargo would land separately, yes. Doing so enables the aforementioned free return in the case of two-in-a-row landing mishaps, as I said. I have no strong feelings regarding the minimum cargo of individual entry vehicles, so long as there aren't too many of them and you're not splitting things up that shouldn't be split. I would say that if you're using more than three or four cargo entry vehicles something is amiss. This is compatible with cargo landers with payloads in the range of 15-20 tonnes.

In addition to civilian space spending equivalent to the rest of the world combined (and military space spending that I would bet is 75% of world military space spending) the US has something else that nobody else does: A truly private space industry. We have Bigelow, Virgin Galactic, SpaceX, Planetary Resources, and any number of other enterprises. The industry may be small at present, but it's growing rapidly and seems to gather a lot of interest in the investment community. I'm not saying that there is no other country that could do it, but what other country would?

So, you want to develop reliable entry technology for the first mission.

I have asked you this question, in various forms, four or five times already in this thread:

What advantages does propulsive landing offer relative to aerobraking?

You have yet to directly address it, and this is the crux of my opposition to its use. If you're using aeroentry for cargo, why would you use a rocket for crew when this represents a relative safety hazard and the technology will already exist*?

*This is based on the assumption that the technology already exists and will be in use for cargo

Russel · 2013-07-03 04:58:30

JoshNH4H wrote:

I have asked you this question, in various forms, four or five times already in this thread:
What advantages does propulsive landing offer relative to aerobraking?

And I've already addressed this four or five times. The primary reason is its safer. And let me explain why, again.

An 8000Kg lander with a 6m diameter heat shield will slow to Mach 3 at around 6Km altitude with about 30 seconds before impact. Give or take.

By comparison MSL with an entry mass around 3000Kgs and a heat shield 4.5m in diameter needed to slow to Mach 2.2 before chute deployment which was nominally about 9 to 10Km altitude.

http://ntrs.nasa.gov/archive/nasa/casi. … 006430.pdf

To give a crewed lander the same sort of margin. That is to get the craft down to Mach 2.2 at 10Km means inflating the heat shield diameter to more like 9m.

Yes, quite doable, and yes still possibly less mass than a fully propulsive lander.

But, now the single riskiest item is the parachute. Its massive, or you've got more than one. All of which leads to more failure modes. I can live with that if there's no one on board. You want to know the statistics of how many people are killed by parachutes here on Earth?

GW is quite right in that beyond a certain point (in terms of increasing lander mass) you're better off ditching parachutes and going directly to retro propulsion as soon as possible. I don't agree with him on the exact numbers for two reasons. One is that when you take gravity into account your "projectile" doesn't keep the same angle to the horizontal. As it slows down below about 30Km, gravity kicks in. The other is that real margins reflect the variability of the Mars atmosphere.

This is why in my mind, landing large masses on Mars does mean having a reliable aerobraking system. But parachutes are not part of the solution. Higher velocity/temperature decelerators deployed further up might be useful. At the very least no one can avoid SRP.

With a mostly propulsive landing you've got complete authority over what speed your'e doing and when. As I said before, you can land on Olympus if you want to. Certainly a few Km above MOLA and it makes no difference - you've simply set your propulsion to the settings needed. But with aerobraking those last few Km count. Or put another way, if you land propulsively, higher landing sites simply mean less gravity loss.

You see, when all sources of error are accounted for - the main one being atmospheric variability - a landing that depends on aerobraking gives you only a couple of Km and a matter of seconds to play with. A mostly propulsive landing ensures you are where you want to be with more precision. Less variability. More margin.

Mars has big diurnal variations in its atmosphere and combining that with direct entry means over sized heat shields (just as on MSL). Think about that one. Diurnal variations trouble a mostly propulsive lander less, but even so, this is part of the reason why I prefer using low Mars orbit as a starting point. It gives you more choice as to when you land.

As well as parachutes there are other things that can go horribly wrong in a conventional landing. Separation of the back shell, and more importantly separation of the fore shell (main heat shield). Now you can throw time at this, but that's less time you have to throttle up your terminal descent.

Thrusting through a heat shield has twin problems. First you're carrying that much more mass to land. Meaning more landing fuel. The other is that if you're going to run into show stopping turbulence its going to happen because you're riding on the plume from the engine. This is why I've opted for the concept of putting the landing engines outboard and above the center of mass of the vehicle. That on the face of it gives a lot more stability.

Now, with a very large vehicle you may find you have no choice but to fire through the heat shield, or else let the heat shield disassemble under you. But it still makes sense to keep the landing engines on the periphery. Now in that scenario you're still experiencing some of the backwash from the landing thrust. Net result is that you're sitting inside a tighter envelope of speed and thrust setting. In essence, putting the engines outboard and above the center of mass means you can use higher throttle, sooner. But you can only do this if you're not exposing the engines and their supports to the extremes of hypersonic heating.

Any way you look at that it means safer. Any way you look at that it means a larger vehicle has less margin. And a vehicle that keeps the engine behind the heat shield has less margin.

There are other reasons why a mostly propulsive lander is safer. They have to do with simplicity and fewer systems that have to be kept in order, and monitored by a crew that may only have seconds to correct errors in automated systems. With a mostly propulsive lander its a lot easier to "ride the stick" if you have to.

And another issue. And I mentioned this before too. With a conventional lander your main source of error as far as exactly where you land is the atmosphere. And because you're likely dropping heavy items onto the surface you're forced to land some distance from base. With a mostly propulsive lander you don't have the raining trash issue so all things consider your landing is likely to be a lot closer to base. Things can go wrong between the lander and the safety of base and if your trip is 1Km rather than 5Km that's also a safety issue.

And if I really wanted to get into boring detail I'd get into how all the accumulated junk of heat shields and parachutes on the surface of Mars will eventually require a clean up. Anything you have to do out there comes with risk.

Moving away from the actual landing itself, lets take a look at what else I'm doing here. Now, my lander is somewhat more massive in terms of mass at low Mars orbit. Its 15 tonnes and the alternative might be 8-10 tonnes. That's a few tonnes saving relative to the many tens of tonnes you're otherwise transporting to Mars. I'll live with that.

But, where my architecture gets more decadent is in using a lander (one that has been used in a previous mission) itself as a transport for fuel to the surface. Now, admittedly I could have just dispensed with that and landed the fuel in a conventional pod. But the reason I'm doing this is that if you want spares, you've got spares. You send one ascent vehicle to Mars and you're ready to return home and the ascent vehicle loses an engine soon after take off and you'd better get used to survival rations. If you take off on my ascent vehicle and it does the same thing you've got yourself a week or two in which to swap engines. Safer? Yes.

Now, look, you could provide some spares separately and at less cost. And I remain open about that particular feature. It is nice however to have a vehicle that's already landed on Mars that can in a pinch be converted to a fully functional ascent vehicle. Take away this element of the architecture and a reusable lander and ascent vehicle actually requires less mass.

How about multiple redundancies in space brought on by the fact that you've got propulsion and fuel storage in the lander as well as in the transit vehicle? Safer? Yes.

How about approach to Earth. You've taken damage and lost some fuel. Transfer the fuel to the lander. As a last resort abandon the transit vehicles and use the lander to achieve Earth orbit. There's a host of scenarios like this where if you've got your lander with you, you're safe. If you've relied upon an expendable ascent vehicle, or worse, launched directly from Mars, then there's always the satisfaction of knowing your place in history, right? Yep, the first person in history to go all the way to Mars, and, um.. almost returned home.

Last edited by Russel (2013-07-03 10:45:59)

GW Johnson · 2013-07-03 09:39:13

OK, I don’t have any sort of trajectory code for landings. I have to do this as bounding or feasibility calculations. This just me sitting on the front porch with pencil-and-paper, and maybe a calculator.

I know that my approximate entry analysis is just that: approximate. Of course the tail end of the trajectory bends downward. With a capsule flown off-angle in pitch, you can generate a bit of aero lift, and offset that. We did exactly that with Gemini and Apollo. L/D = 0.1 is about all you can get, but it’s enough.

So, my way-oversimplified, straight-line 2-D Cartesian, entry model is not all that bad. It gets you into the ballpark, which is all it has to do. We’re looking at broad trades here, not fine details.

What I found, running systematic variations for aerobrake entry at Mars, is that you want the min credible de-orbit burn, from the min credible low orbit altitude, in order to minimize trajectory angle at entry. If you don’t, you hit before the hypersonics are over.

I also found that entry vehicles around or under 100 kg/sq.m ballistic coefficient tend to slow to M3 at 15-25 km altitudes, admittedly a bit variable due to the inherent variability of the Martian atmosphere. That’s high enough for chutes to have time to work, and it is typical of all the probes sent to Mars until very recently. I’ve been using the average atmosphere data in the Justus and Braun EDL paper for my calculations.

I also found that bigger vehicles with higher ballistic coefficients penetrate deeper before slowing. No surprises there. At 400 kg/sq.m I had about 5 km altitude at local M3. That’s about 2 minutes from impact at the very most. Not time for a chute to deploy, much less work. That’s the “classic EDL dilemma” everybody yammers about, but the way around is clear: no chutes, just thrust-to-landing.

Whether you wait to M3 to fire up the thrust is an unexplored issue. But you will have to use supersonic/hypersonic retro thrust. No way around that.
Actually, the plume stability issue is easily overcome, it’s just that NASA hasn’t looked at this since the development of Mercury, ca. 1960. But they did look at it then, and anyone easily could again. I’ve already written about how that can be resolved, and won’t repeat that discussion here.

That same Justus and Braun EDL paper describes the very old density scale height-based approximation for non-lifting entry estimates. The dynamics they showed in their paper looked pretty realistic, but the heating estimates were quite inconsistent. I chased this analysis back to its originator: H. Julian Allen at NACA in the 1950’s, and it is something I remember seeing some 4 decades ago in graduate engineering school, too.

I corrected the errors in the heating estimates in my spreadsheet version, and went to timeline-integrated heat totals, instead of the old closed-form estimates. My oversimplified-estimate data for Apollo returning from the moon look just about like the “real McCoy”, and not just for the heating, but also for the end-of-hypersonics location where drogue chute deployment could begin.

So, my model ain’t too bad. I think you can trust my numbers for aerobrake-to-rocket-final-descent. I see no problems up to at least 60 ton sizes, which means smaller vehicles are even easier. 0.05 km/s to deorbit plus about 1.4 for terminal descent. Total delta-vee to land: about 1.4 to 1.5 km/s.

What happens as you make the de-orbit burn bigger and bigger is two-fold: velocity at atmospheric interface is lower and lower, but trajectory angle is very much higher and higher. The angle increases faster than the velocity decreases.

For even very low ballistic coefficients, you whack the surface of Mars hypersonically, if that entry angle is steeper than about 2-3 degrees or so. That’s physics. The only way around this is to keep on retro thrusting all through entry. That’s what we’re discussing in this thread.

The ultimate limiting case is completely killing all the orbital velocity in your deorbit burn. Then you fall, from rest at orbit altitude, vertically downward. Vertical velocity builds due to gravity. Ignoring air drag, and using constant gravity, from 200 km at 0.38 gee, velocity at impact is near 1.22 km/s, which you have to kill with a last-ditch, last-second burn of that same magnitude. Total: 4.87 km/s delta-vee required for the vertical free-fall descent.

That’s a minimum “credible” figure for the delta-vee required beyond the 3.65 km/s for de-orbit at 200 km. There’s no terminal maneuver kitty in that. By the way, that’s also about M5 just as you begin your last-ditch burn. Aeroheating is getting very severe at Mach numbers like that, by the way.

The other extreme “option” is to let gravity accelerate you downward to a tolerable speed, say around 500 m/s, then burn at thrust equals weight, to descend at constant speed. That’s about M2 deep in the atmosphere, and it corresponds to a 400 sec (6.7 minute) descent from 200 km. Aeroheating is significant but not severe. Exposed metal works, even aluminum. But not any faster than that.
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But, you have to burn for 400 sec to do that! Assume an average vehicle mass of 8 metric tons during that burn. 8000 kg. On Mars at 0.38 gee, that’s about 29.8 KN of thrust required to balance the pull of gravity. That’s almost 12,000 KN-s of total impulse, which at 300 s Isp (typical of kerolox, MMH-NTO, and methane-lox), is nearly 40 tons of propellant, far more than the spacecraft masses!

So, I have to conclude that this kind of all-low-speed controlled descent is simply not feasible. Even LH2-lox would make no difference to this outcome, Isp is around 450-or-so, for 26 tons of propellants. You’re not going to do it with any imaginable chemical propulsion. Not even NERVA at 900-1000 s Isp could do this.

Therefore, something close to free-fall from a maximum de-orbit burn is what you have to propose for a non-aerobraking descent. But, here’s the thing: if you have to protect from Mach 5 aeroheating anyway, you might as well protect from entry-speed aeroheating, and just do aerobraking. It’s exactly the same kind of protection.
Comparing the two extremes, aerobraking has less delta-vee required, less total propellant, and a smaller, lighter vehicle at entry interface. Meaning less to ship from Earth, which is lower cost.

Now, that’s not to say there might not be an optimum somewhere between these two extremes. As I already said, that trade is unexplored.

The safety architecture doesn’t enter into that decision. That comes later, and applies to both aerobrake entries and non-aerobrake entries, or anything in-between. I really like the idea of having multiple vehicles that could be refueled, adapted, and used as redundant ascent vehicles. How they got there makes no difference.

Finally, even at only Mach 5, beware of shock-impingement heating on adjacent nacelles and on the struts connecting them. Structures like that don’t survive well, even with ablative protection, above about Mach 3.5-4 here on Earth. There is a reason why all entry spacecraft to date have been single-body capsules or very “clean” winged designs. I really don’t think engines on struts around the periphery of a heat shield is going to work.

GW

Russel · 2013-07-03 12:04:19

GW,

I'm not taking any great issue with your conclusions regarding large vehicles. Especially at lower Betas. My concern has always been about smaller vehicles and particularly crewed ones. Where I ideally want a few more Km margin. Which means larger heat shields and thus a heavier vehicle. And if its to be reused a surprisingly heavier ascent vehicle. Now at one stage I resort to the tactic of making the back shell unfold and that was actually quite effective at increasing drag and achieved what I wanted. But also quite complex and a bit marginal structurally.

As for how much delta-V it takes to land propulsively let me spell this out again.

Starting from a 200Km orbit you do a initial burn that brings you into a slightly elliptical orbit which in the absence of an atmosphere would graze the surface. Now in this elliptical orbit the velocity at 100Km altitude is slightly higher than the velocity you started with. Its a tad over 3.6Km/s.

The benchmark case would be a free fall from zero velocity. However you don't need to free fall from 200Km. And you don't need a vacuum either. So when I calculate this I take the initial braking burn to be an impulse at 100Km. Now from 100Km a free fall assuming a vacuum is going to result in impact at 871m/s. So at a first take, the delta-V required for a theoretical two impulse burn is 4.5Km/s.

Now when I wrote that simulator I just wanted to get a feel for how much extra drag I needed in order to slow a vehicle to a certain speed. So I cared about relativity rather than absolutes. What happened then was I stumbled across a paper written as a Masters thesis which was basically a Mars landing simulator. (Sorry I can't find the link but I'll go hunting.) And from there I got some insights into where my simulator was incomplete. I still haven't recoded it to properly translate between the two frames, but I have learnt how to tweak it to get more realistic results.

The main problem is that the roughly 3.6Km/s orbital velocity is relative to an inertial frame. The planet itself rotates within that frame. And so too does the atmosphere. Its not an insignificant amount. At the Martian equator the planet is spinning at 240m/s.

The other factor at play is that if you enter the atmosphere at a relative (to the air) speed of (say) 800m/s you're actually losing more energy due to drag than gravity would add. Your velocity climbs initially, then comes down as the atmosphere thickens. At 15Km/s you're down to more like 600m/s.

What this means is an initial burn of 2.6Km/s delivering you into the atmosphere at a relative velocity of 800m/s. When you're ready to throttle up at say 20Km, you're doing 700m/s relative to the planet. If that were an impulsive burn you'd still have a theoretical free fall impact (in vacuum) of 390m/s. However its not a free fall and drag still accounts for roughly half of that.

So that's 2.6Km/s + 0.7Km/s + 0.2Km/s = 3.5Km/s And add to that maneuvering and that's where I get 3.8Km/s from. I admit that this is a working position.

Bear this also in mind. You can get more out of drag. As it stands you encounter peak heating of about 0.4W/cm2. You could shave another 0.2 to 0.3Km/s off that figure by simply coming in faster. At 1.2Km/s though the peak temperatures are a problem.

And that final 300m/s of maneuvering fuel is surplus to the need to come to a rest. Its the extra for getting the landing site right.

Bear in mind also that you're sitting in a vehicle that can do 4.5Km/s on a full tank. 10 tonnes of fuel gets you 3.8Km/s. 11 tonnes gets you 4.1Km/s.

Edit: Its based on methane/LOX at Isp of 350.

Edit 2: Having reread your post over again, its clear where you're coming from and you're making it more difficult than it needs to be.

The point here is the most efficient use of thrust is to wait as long as you can at orbital velocity and then apply as close to an impulse as you can. Which means an initial braking burn as deep into the atmosphere as you can. Now, I figure on 100Km altitude. That's still very tenuous gas up there. I suspect its possible to go further than that. Remember the discussion before about there probably being a limit to SRP that is also dependent on density? Well, that's factored in.

then you try to ride the engines. Er, no. And I thought I'd made that clear. You're down to some reasonable relative velocity. You let drag do the work. Hence my painstaking typing of the phrase "mostly propulsive". Instead you throttle down to idle. Initially you climb closer to Mach 4. Then you start going back down through Mach 3. That's very roughly at around 20Km. The key is to balance out forces. You see you're angle of declination climbs and with that more of the drag vector is oriented upwards and against gravity. Its a completely different game to when you've arrived at 70Km still doing around 3Km/s where you need to get as much energy out before you hit the lower atmosphere.

I know it seems like black magic, but there is key range of velocities where if you're already well into atmospheric interface (not at 125Km but more like 90Km) to start with, the atmosphere will do most of the work for you. Too fast and you've still got too much energy to burn below 20Km. Too slow and you're using too much fuel.

The point about the second braking burn is that you're trying to approximate an impulse (as high thrust as possible) whilst also bringing yourself past Mach 1 at a comfortable altitude. Which for me is about 5Km. A constant deceleration from that point takes just on 40 seconds. Now the thrust profile is a complex thing. You've actually got an excess of thrust to weight at this point. And you probably need to raise the throttle gradually for stability reasons. But with a notional maximum deceleration of 18m/s (A bit under 2 Earth gs) the drop from 700m/s to 300m/s can happen in just over 20 seconds. Its a decent approximation of impulsive. So at around 20Km altitude you start to pull up on the throttle and by 15Km you're braking at the maximum rate defined by the structure.

And I hasten to add, these are indicative numbers. Near enough to make the point, but obviously subject to optimisation.

So, in the end, the process is simply put as two near-impulsive burns separated by conventional drag. And its that drag that accounts for over 1Km/s in your reckoning.

I might also add here that the actual protocol is relative to the elevation of the surface you're landing on. Higher elevations give you less assistance from drag. But the difference here is that high elevations are straightforwardly possible - you just add more fuel. When you rely entirely upon aerobraking, higher elevations just get harder and harder.

Last edited by Russel (2013-07-03 13:37:43)

Russel · 2013-07-03 13:13:22

Btw, although I'm proposing to use a lander (on its last mission) to ferry fuel to the surface, I'm still not convinced myself that's worth the cost in fuel. It may make more sense to make that a more conventional landing. But when I did try this approach before I kept running into issues that had to do with where you store surface fuel.. in a rover like vehicle.. in a stationary tank? Etc. Unresolved.

If I don't land a lander as a fuel ferry then obviously there's got to be spares supplied somehow else.

But as far as an ascent vehicle goes, the least mass option is to land it with the crew. A conventional lander, plus an ascent vehicle carried down by its own lander will always end up costing more in mass than a lander that doubles as an ascent vehicle. And I'll repeat, you can't beat the very large mass savings that come from putting your transit vehicle and return fuel into a high Mars orbit. Having a decent ascent vehicle that acts as an inter-orbit ferry is part and parcel of that.

As far as shock impingement goes, I do take that message to heart. That's why I'm thinking in terms of Mach 4 at most. At least in densities of any consequence. Now the configuration puts the engines above the center of mass and the engine bell shadows the guts of the engine. Its the boom that sits in the path of shock waves. And the answer to that, at least at these sorts of speeds is a selective thermal protection layer on a small portion of the boom. If you really wanted to design it to the limit you could use fuel as coolant for the boom - since you're expecting a small amount of fuel flow for the whole trip.

Russel · 2013-07-03 13:26:17

I think the paper is here..

http://dspace.mit.edu/bitstream/handle/ … sequence=1

JoshNH4H · 2013-07-03 21:35:02

Russel-- I don't understand the issue. I've been saying all along that retros are fine for the last 700 m/s, and obviously you will use a larger heatshield for crew to increase margin. You're being very uncharitable towards aerobraking because you have a preference for propulsive entry, but it's really not merited.

Russel · 2013-07-04 02:32:05

JoshNH4H wrote:

Russel-- I don't understand the issue. I've been saying all along that retros are fine for the last 700 m/s, and obviously you will use a larger heatshield for crew to increase margin. You're being very uncharitable towards aerobraking because you have a preference for propulsive entry, but it's really not merited.

Josh,

I started this thread to discuss an overall architecture. And more than that to encourage wider discussion about comparative architectures. There are many more possible architectures than those currently being championed. And too often what you see are people responding to a particular problem, often at the expense of causing other problems. What I'm trying to do is to get a more holistic approach. I'm happy to discuss partly versus mostly propulsive landings, but that's a small part of this.

As far as the overall architecture goes, I think there is a lack of discussion about the detail. Simple things like how you get from low Mars orbit to high Mars orbit. What you do exactly on return to Earth space. What you contingencies are. And so on. Those considerations flow back into the design of your vehicles and the overall development effort.

Early in this thread I pointed out that my architecture as much as anything is about simplifying the development effort - by having only two vehicles involved in the "getting there and back" problem. I've not spoken much about the development of large, aerobraked vehicles that land stuff on Mars on a one-way trip because in essence, that's something common to every architecture. And I'm perfectly happy to talk about that. And I'm even willing to say that I agree that such big landers can be optimised to 50% landed mass, and possibly a bit beyond that.

As far as the nature of the manned lander goes. Is it partly or mostly propulsive. We've reached a point where there is little to be argued about regarding mass. A single use lander, accompanied by an ascent vehicle that itself sits on top of its own lander, will always be more mass than a mostly propulsive lander/ascent vehicle. So lets exclude that option. That leaves us with a lander and ascent vehicle that is partly propulsive - it uses a heat shield. Where do we get to with that? Well, a vehicle that is a few tonnes less mass than fully fueled mostly propulsive lander. Does those few tonnes count? In isolation yes. In context no.

One of the things drilled into me in engineering is the phrase "the devil is in the detail". You might start out with something that is conceptually simple but end up with a monster. A jet engine is conceptually very simple. But look at the decades of effort and thousands of parts that make a modern engine. And all that detail has its price. The engine that failed in a Qantas A380, failed because of a slightly misaligned drill hole - a tiny mistake in a huge file of numbers.

All else being equal - and this is pretty much the case here when comparing the partly and mostly propulsive landers - I'd settle anytime for the simpler, more serviceable and to my mind, safer machine.

I'd like someone to tell me the deal-making advantage of having to rely upon a heat shield - without simply deferring to "conceptual simplicity". A heat shield is a necessary evil for larger landers. But with a smaller lander you've got a choice. And I really don't want to go into here all the operational and safety advantages that come with a mostly propulsive concept. Been there, done that.

But I will repeat one issue, since I'm talking about the devil in the detail. Supersonic retro propulsion is a problem that will be solved. But, it may be solved with certain caveats. Certain no-go zones in terms of the placement of the retros, the density and velocity. I've done two things that I believe make this problem easier. One is to move the engines away from the body of the vehicle. So the turbulence has less to react against. That's simple physics. The other thing is I'm able to maintain constant ignition through the whole procedure. That may or may not matter. But what people may discover is that you probably can operate a retro into a higher velocity stream than you can start one up at. If you've ever watched videos of people blowing out oil well fires with a jet engine then you'll understand what I mean.

The inevitable result is that SRP parked behind a shield will almost certainly operate within a tighter envelope than those don't and remain lit for the whole procedure. Which means a vehicle that uses a heat shield and then goes directly to SRP is going to have a shorter, harder terminal brake, and with less margin.

NO way I would like to ride that myself.

I'll stand by the conclusion that with comparable margins, the mostly propulsive lander is only somewhat more massive than one that uses a heat shield. And the resulting ascent vehicle is lighter. Yeah ok, I'm repeating myself again.

Now you can call me uncharitable but I'm just not enraptured by the tendency to ignore the details. For instance, I've noted several times here that trash is an issue. But gotten no response. Ok, you start with a Mars hab. That's got to be landed with a heat shield and once that comes to rest you probably don't want to land a similar vehicle anywhere within a few tens of Km. Now what happens when you land a smaller, crewed lander? Well you have to target a landing ellipse for the heat shield that's well away from base. Then what happens with the next lander? Things get complicated, and people start needing longer distance rovers. Anyhow I'll stop at that for now. I'd just urge people to consider that the detailed design.. the boring engineering stuff.. really does matter.

And in the end, simplicity of actual detailed design (as opposed to simplicity of concept) does matter.

Now, please, can we talk about overall architectures? Are we going to park a return vehicle in low Mars orbit, or high Mars orbit? Are we going to return that vehicle to Earth space in a reusable condition, or trash it? Where are the designs for small "taxi" vehicles designed to transfer people between orbits? Should we integrate visiting Phobos in the process? Should aerocapture be a requirement before we go to Mars, or simply a later refinement? What level of redundancy do we need? Are we going to trust a singular life support system, or (as I have done) split it over two self contained units? What happens when we develop leaks? How about leaking fuel? should we build in the capability to transfer fuel?

That's what I'd like to see a discussion about. Thanks

RobertDyck · 2013-07-04 02:59:26

Russel wrote:

I started this thread to discuss an overall architecture. And more than that to encourage wider discussion about comparative architectures. There are many more possible architectures than those currently being championed.
...
Now, please, can we talk about overall architectures? Are we going to park a return vehicle in low Mars orbit, or high Mars orbit? Are we going to return that vehicle to Earth space in a reusable condition, or trash it? Where are the designs for small "taxi" vehicles designed to transfer people between orbits? Should we integrate visiting Phobos in the process? Should aerocapture be a requirement before we go to Mars, or simply a later refinement? What level of redundancy do we need? Are we going to trust a singular life support system, or (as I have done) split it over two self contained units? What happens when we develop leaks? How about leaking fuel? should we build in the capability to transfer fuel?
That's what I'd like to see a discussion about. Thanks

Ok, I'll bite. I've proposed several different architectures over the years. I joined the Mars Society in 1999, one year after it was founded. Some people didn't like what I proposed because it wasn't Dr. Zubrin's. Others appreciated how much I had worked out. And as I said, over the last (how many years?) I've proposed a few.

I've been thinking that my architectures address many of your key points, but avoid the problems others are pointing out here. If I repeated my ideas, would you consider them, or just reject them? Not Invented By Me?

Russel · 2013-07-04 07:19:10

RobertDyck wrote:

Russel wrote:
I started this thread to discuss an overall architecture. And more than that to encourage wider discussion about comparative architectures. There are many more possible architectures than those currently being championed.
...
Now, please, can we talk about overall architectures? Are we going to park a return vehicle in low Mars orbit, or high Mars orbit? Are we going to return that vehicle to Earth space in a reusable condition, or trash it? Where are the designs for small "taxi" vehicles designed to transfer people between orbits? Should we integrate visiting Phobos in the process? Should aerocapture be a requirement before we go to Mars, or simply a later refinement? What level of redundancy do we need? Are we going to trust a singular life support system, or (as I have done) split it over two self contained units? What happens when we develop leaks? How about leaking fuel? should we build in the capability to transfer fuel?
That's what I'd like to see a discussion about. Thanks
Ok, I'll bite. I've proposed several different architectures over the years. I joined the Mars Society in 1999, one year after it was founded. Some people didn't like what I proposed because it wasn't Dr. Zubrin's. Others appreciated how much I had worked out. And as I said, over the last (how many years?) I've proposed a few.
I've been thinking that my architectures address many of your key points, but avoid the problems others are pointing out here. If I repeated my ideas, would you consider them, or just reject them? Not Invented By Me?

I'll promise to cannibalize any good idea Please, go ahead..

GW Johnson · 2013-07-04 11:57:55

Russel:

You know what? What you describe and what I describe are not so very far apart after all. You are deorbiting into the same surface-grazing orbit that I use. You just seem to be burning hard for retro propulsion near 100 km. It also sounds like you are getting benefits from aerobraking, just in the lower Mach range. That kind of thing lies between the two extremes that I was looking at.

My baseline mission concept was an orbit-to-orbit transport LEO to LMO and back, with at least the habitat and engine modules reusable. The landers (whatever they were) could push their own propellant supply to Mars one-way as separate vehicles, everybody to park in LMO. I was looking at re-using as much as I could, and staging multiple landings at different sites from LMO in the one trip to Mars. It's a very old concept, dating to the 50's and 60's in one form or another.

I hate to discard so much as a chute. Haven't yet figured out how to make a chemical lander one-stage/reusable except with propellant ISRU on Mars, and I'm not sure I trust that enough to bet lives on it. Not yet. But it really needs to be done. I see no point at all to the one trip/one-landing concept. If one is going to go to all the trouble to send men to Mars, then one really ought to explore a whole lot while there. That means a bunch of landings all over the planet. That kind of information return (a real planetary survey of ground truth) is what must be in-hand to plan viable bases and colonies.

You seem to know something about staging out of high Mars orbit vs LMO. Some sort of delt-vee reduction. I think you had in mind shuttling the landers back and forth. There's a cost there for shuttling, but the savings with transit vehicle might overcome that.

As for starting engines in SRP, a lot depends upon propellant selection. If they're hypergolic, it'd be hard to see how they wouldn't ignite, no matter the slipstream. Isp figures look just about the same for kerolox and MMH-NTO at 300-310 or so. methane-lox with reasonable non-vacuum bells would be similar, I think. But MMH-NTO is hypergolic, and storable without cryogenic considerations for very long times. I've seen no ISRU proposals to make MMH-NTO on Mars, though.

GW

Last edited by GW Johnson (2013-07-04 12:00:37)

RobertDyck · 2013-07-04 13:15:46

First, I'm concerned any architecture that uses all expendable equipment is far too easy for a politician to cancel. Richard Nixon cancelled Apollo as soon as he was elected, but the Shuttle continued to fly long past it's original intended life. Well, it was supposed to fly 100 missions per orbiter, so with 4 orbiters it should have flown 400 missions, but that was expected in 10 years: 1981-1991. It wasn't cancelled until 2011. Three times its expected life. For this reason, I want any Mars architecture to use reusable equipment. I would start with partially reusable, replacing expendable equipment with reusable as we establish a permanent presence on Mars.

I also argue that dropping the Earth Return Vehicle on Mars surface, only to lift it off again, is a waste of fuel. It's better to park the interplanetary transit vehicle in Mars orbit. Rendezvous and docking in orbit is a mature technology.

I read about the Russian Energia rocket in Robert Zubrin's book "The Case for Mars". I thought it was a great idea. A few other Mars Society members talked about it, and made digital images of a Mars Direct mission launched on one. But after a while I had to ask if anyone had actually talked to the Russians. I didn't get an answer, so I assumed not. So I took it upon myself to talk to the Russians. In December 2000 I contacted them, in February 2001 the head of the International Division of Russian Space Corporation Energia replied. He said no one from the Mars Society had contacted him before me. Yes their big rocket was available, but it would have to include restoration of infrastructure as well as per launch cost. He didn't confirm the cost, but when I talked to an employee of their American office in December 2000, he said NASA had contacted them "a few years ago" about using that rocket for a human mission to the Moon. They conducted a study but didn't get a contract. At that time the cost was between $60 million and $100 million US dollars to restore infrastructure. I also found a NASA website that listed Energia: $120 million per launch, including the upper stage, in 1994 dollars. Ok, that tells me when "a few years ago" was. In 2002 I also contacted the manufacture of the RD-0120 engines, he said he still had all the plans and jigs, would have to replace a CNC milling machine, but was willing to swallow retooling cost on the condition of getting a solid order for new engines. Great! That keeps infrastructure cost down. Unfortunately there was an accident. As part of the break-up of the Soviet Union, ownership of the Baikonur Cosmodrome was handed over to Kazakhstan on January 1, 2000. In 2002 it was time to re-roof of the vehicle assembly building. They had theft from the site, so workmen stored 10 metric tonnes of roofing material on the flat roof. There was a rain storm. On April 25, 2002, the roof collapsed. The Buran space shuttle and all Energia stages were in that building at the time. They're gone now. The Ptichka space shuttle orbiter is still intact, but all other Russian shuttles are gone. I've talked to American contractors who did not want to accept the Energia could be restored, but until the accident it could. At 2/3 the launch mass of a Saturn V, it would have required 3 Energias for Mars Direct. At the prices I just cited, even after inflation, that was the cheapest launch vehicle. But now it's 11 years without a roof, 11 springs with wet melting snow against the inside walls of a steel building. And it's bridge crane. And more infrastructure may have degraded. Russians build stuff extremely durable, designed to just last. But still, 11 more years without maintenance. Even I have to face the fact it may be too late for Energia.

I mention all this because I started with existing launch vehicles: Altas V heavy, Delta IV heavy, Ariane V, Angara, Proton, and the lovely Energia. Now there's Falcon 9! But Energia solved a lot of problems. It could throw a significant mass vehicle directly into TMI. It will be hard without.

More lamenting Energia: it could lift 88 tonnes to 200km orbit, without its upper stage. With EUS it could throw 29.3 metric tonnes directly into TMI.

Ok, architecture. I start with a reusable Interplanetary Transit Vehicle (ITV). Launch a Mars Assent Vehicle directly to the surface of Mars, unmanned. (I was counting on Energia for that.) Use ISPP to fill its fuel tanks. Assemble in LEO, using ISS as a construction shack. The reusable ITV would be launched in one throw by Delta IV heavy, or Atlas V heavy, or Proton, or Ariane V, or Angara. Now we would use Falcon 9, and it can lift a little more weight. Using Russia's automated docking system, it would dock with ISS. The expendable propulsion stage would be lifted to ISS by Energia, but Falcon Heavy may be able to do it. The ITV would aerocapture into Mars orbit: highly elliptical, high orbit. Using a parasol of Nextel 440 cloth. That's the ceramic fibre cloth that Ames Research Center used for DurAFRSI, the advanced thermal blanket that was never installed on the Shuttle. Use titanium alloy ribs for the parasol; both springy and high temperature. Descend to Mars with a capsule, just a single seat for each astronaut. But the Mars lander would include an inflatable habitat. And an open rover, with a couple pressure "pup tents". The tents would use spacesuit PLSS for life support, powered by the rover for extended use.

The hab would use 400 denier Tenara fabric for it's outer surface. White EMU spacesuit use Orotofabric, a double layer fabric with 400 denier PTFE cloth outer layer, backed by 400 denier Nomex with 2 threads of Nomex replaced by Kevlar every 3/8". The backing is combination micrometeoroid protection, and fireproof. You don’t need fire protection in a 7 mbar, 95% CO2 atmosphere. And you don’t need micrometeoroid protection on the surface of Mars. You do need scuff protection, and protection from dust storms. Tenara is architectural fabric, the same 400 denier PTFE yarn made in the same factory by the same machines as the facing of Orothofabric. It's a single layer but twill weave instead of plane weave. That's ideal for Mars. Lower mass, and since it's in commercial production, a lot cheaper.

The MAV would have an unpressurized cabin, basically a fairing with seats. And extra large propellant tanks. The MAV would act as the TEI stage, to push the ITV out of Mars orbit toward Earth. That allows all return propellant to be made by ISPP. It would aerocapture into Earth orbit, then aerobrake down to ISS. Use a tiny bit if fuel to circularize orbit, and rendezvous/dock with ISS. In case aerocapture fails, the ITV would carry escape pods. I was thinking of 2 Soyuz descent modules, with tiny service modules and no orbital module at all. However, a single Dragon could replace them. As long as the ITV docks with ISS safely, the capsule would remain attached for the next mission.

A separate cargo lander would send a pressurized laboratory with pressurized rover. The lab wouldn't have life support, counting on the hab to attach. As a backup, life support in the pressurized rover could supply the lab. All lab equipment would have to be thrown outside to reconfigure it as a hab, but this means a backup hab would be waiting on the surface before astronauts leave Earth. For this purpose, the lab would include a full set of food for the surface stay, although compact food that may not be the most tasty. I was counting on Energia for the cargo lander as well; direct throw.

Mars atmospheric entry would be conventional: aeroshell with heat shield, parachute, propulsive landing for the last few metres, and legs with shock absorbers. Don't try to "bounce and roll" with humans on board. And forget the "sky crane", just use a lander with legs like Viking or Mars Phoenix.

One possible optimization: the MAV could use bladders for LOX and Liquid Methane. Fluoropolymer bags, with springy carbon fibre support members that would be folded and wrapped around the lander for atmospheric entry. Once on the surface the supports would "spring up" to form a basket to hold the bladders. Aluminized PCTFE would work great for the bladder. The purpose is landing would have almost empty fuel tanks, so collapsing them would keep them out of the way during hypersonic entry. You could further optimize by placing one bladder inside the other. LOX is colder than L-CH4, and both are colder than Mars ambient, so place the LOX bladder inside the L-CH4, with a refrigeration pump for the LOX only.

The ITV would be optimized for zero-G, complete with exercise equipment. And would carry food for the transit from Earth to Mars, and the return home. The Mars lander would carry food for the surface stay. Since the lander would be carried by the ITV, if a free-return is necessary, you already have food for the entire extended loop home. And yes, that means the lander food store would have to be accessible while it's docked to the ITV.

Upon return, a reusable spacecraft would ferry astronauts and their Mars samples back to Earth. At the time the Shuttle was still flying, so I envisioned the Shuttle. You could use a Dragon, but that would defeat the "reusable" architecture. Better yet a Dream Chaser.

First mission:
- 1 Energia for MAV (or SLS)
- 1 Energia for lab & pressurized rover (or SLS)
- 1 Falcon 9 for ITV
- 1 Energia for TMI stage (or Falcon Heavy)
- 1 Falcon 9 for lander & unpressurized rover
- 1 Falcon 9 for Dragon
- 1 Atlas V 402 for Dream Chaser

Second mission:
- 1 Energia for MAV (or SLS)
- 1 Energia for lab & pressurized rover (or SLS)
- 1 Energia for TMI stage (or Falcon Heavy)
- 1 Falcon 9 for lander & unpressurized rover
- 1 Atlas V 402 for Dream Chaser

Dream Chaser is design for 7 crew, not just 4. And I'm arguing for 4 crew per mission. That leaves some room for cargo. Food and supplies to replenish the ITV would have to be launched; that would be in the extra space in Dream Chaser. It could be left docked to ISS until the Mars ITV returns, or used for ISS operations.

Note this reduces total launches by only 2 Falcon 9s. But a spacecraft costs more than the launch, so reusing the ITV saves a lot of money. And Robert Zubrin argued for human exploration rather than robotic, but that was in 1990. I argue the robotic exploration is finished. We actually have enough information to start the permanent Mars base with the first human mission. So the first two missions would each deliver a hab and pressurized rover. The base would be left with 2 labs, 2 pressurized rovers, and 2 unpressurized.

That means a third mission would require:
- 1 Energia for MAV (or SLS)
- 1 Energia for TMI stage (or Falcon Heavy)
- 1 Falcon 9 for lander & unpressurized rover
- 1 Atlas V 402 for Dream Chaser

This would still provide a 3rd habitat and 3rd unpressurized rover. That really starts to build a robust Mars base.

Eventually the TMI stage would be replaced by a reusable one. That would require a tanker at Mars to deliver propellant. Either from Mars surface or one of its moons. I leave that open. And replacing the entire TMI stage allows for any future propulsion technology.

Getting to the surface would be via Mars shuttle, based on DC-XA. I argue that technology is ideal for Mars, although not so for Earth. A lifting body shuttle is optimal for Earth, but DC-XA style propulsive lander is optimal for Mars. It would still use a heat shield for hypersonic entry. But again, this is not for the first mission, or the second or third. Only when there's a reliable fuel manufacturing and storage depot on Mars.

Prerequisites:
- demonstrate aerocapture with a Mars orbiter
- demonstrate ISPP with a robotic Mars sample return (call it technology demonstrator to make sure no one removes ISPP this time)
- complete Dream Chaser
- demonstrate some additional life support equipment on ISS
- develop a Mars surface spacesuit; I prefer MCP because it's machine washable

I would also like the Centrifuge Accommodation Module for ISS launched and installed. That may require additional solar panels, the Russian Science Power Platform would do. The centrifuge module would answer questions of long term exposure to Mars gravity.

But this is again stumbling on my plans for Energia. Since the Russian shuttle is launched by Energia, I wanted the Russians to launch their shuttle Ptichka twice: once for the Centrifuge module, the other for their solar panels. I even had a plan to pay for it. Canada would have Russia launch 2 satellites, each on a Proton rocket. Pay Russia 10% the usual cost in cash, the other 90% would be applied to restore infrastructure for Energia. That avoids the Russian argument that Kazakhstan should pay. Tell Russia they pay all the cost for shuttle Ptichka, including launch costs. But if they do, Canada would provide one free CanadArm, identical to the one we gave NASA for their shuttles. Complete with all auxillary equipment and cosmonaut training. In fact, the arm from Endeavour has been returned to Canada, give Russia that one. The condition is they have to fly Ptichka twice to install the equipment. My ulterior motive is this would demonstrate Energia is safe for human travel, launching twice with all Russian crews before we trust any Canadian astronauts or equipment. But again, it may be too late for Energia.

Last edited by RobertDyck (2013-07-08 21:46:23)

Russel · 2013-07-05 03:05:20

I guess my first question is what is the anticipated mass of the ITV and which Mars orbit is it using? If I were to guess, what you're aiming for is a high Mars orbit, closer to escape than to low orbit in terms of energy?

And the basic idea is that the ITV cycles between low Earth orbit and a high Mars orbit?

What I'd probably disagree over is the use of a Mars ascent vehicle to carry the fuel needed for TEI. Obviously if your ITV is in a near escape Mars orbit and you're using aerocapture back at Earth, then the amount of fuel for TEI is kept to a minimum. So about 1.2 to 1.6Km/s depending on the original orbit and how long you wish to take. So the mass ratio on return to Earth is about 1.5.

Now, just throwing in some concrete numbers, a 20 tonne ITV would need 10 tonnes of fuel. Problem I kept running into at this point is that not only does the ascent vehicle need the extra delta-V to get to a high Mars orbit, it also needs to carry that extra fuel for TEI with it. From Mars surface to high Mars orbit is a delta-V of about 5.8Km/s and that's a mass ratio of 5.2. The notional 10 tonnes of TEI fuel becomes 50 tonnes of fuel back on Mars. And with that comes ascent vehicle bloat.

No matter how I came at that problem, I couldn't find a way to get fuel from Mars surface for the return journey that made sense. At one point I had a multiple-reuse lander/ascent vehicle that needed multiple trips to ferry enough fuel to orbit. And eventually I figured that would be asking too much of ISPP. Very large power source etc.

Its on the basis of that math that I settled on having a habitat in low Mars orbit, which doubles as a tanker. I could have just parked a tanker there, or even expendable boosters. But in the end I felt uncomfortable with the length of time people spend in suits.

Now, I'm happy with an ascent vehicle that has no pressurised compartment for crew. A faring with seats. But a store of air. Comfortable for a few hours, with a contingency for a couple of days in an emergency. The thing I didn't like when I came at this problem myself was that launching people direct from the surface to a high Mars orbit not only requires a lot more fuel on the surface and a larger vehicle, it also places constraints on navigation. Far easier to aim for a nearly circular low orbit, refuel, climb into a pressurized space and enjoy a beer, and wait for right launch window to take you to high orbit. Now, I'm pretty sure that I haven't considered all possible options here.

As for aerocapture, I'm a little easier about it. And its doable. It does require testing though. What I'd prefer to see happen is that the vehicle has the capability to propulsively capture at both ends. Which might lead to a program where initially the vehicle is propulsively captured, but as more confidence is gained then aerocapture becomes the norm.

Aerocapture less so aerobraking means at the very least a stowable solar array. I can't get around that. Keeps the crew busy I guess.

My architecture isn't entirely dissimilar in that I have a reusable transit vehicle. It goes from a high Mars orbit to a high Earth orbit. Now, what I still have to think about is what happens then. From a high Earth orbit to ISS orbit its a relatively simple task to aerobrake. Its just a case of whether the days to weeks involved in a gentle aerobrake is too long to keep the crew confined. This is one of the reasons why I swapped from having a Mars lander that remained near Mars and one that came back with the return vehicle. Its to allow the flexibility to get the crew from high Earth orbit back to the ISS in a hurry. Its by no means essential though. But having the couple of days of emergency air on the lander is important.

I agree that having a reusable transit vehicle is important politically in the sense that having it up there keeps the game going - rather like the ISS really.

I am divided on how to approach the word "reusable". In other words, how many missions do you want to fly the same vehicle. Initially I figured on a lifespan (with refurbishment) of about 13 years. That's 5 missions. Enough to amortize the cost of the vehicle. Problem is that the vehicle itself is not actually the big ticket item - its the development program surrounding it. When you add that to the fact that inevitably "upgrades" will happen and inevitably you end up with hybrids if not total replacements. The other thing I kept coming across is that the bulk of the mass of a transit vehicle is not vehicle - its fuel and consumables. All of which has to be launched from Earth. Hence I'm comfortable with a reusable transit vehicle, but not necessarily one that lasts a decade or more.

Now, I stand to be convinced about the merits of making return fuel on Mars. Fuel that gets you into Mars orbit yes. High orbit maybe. But given the decision to send fuel to Mars and leave it there in low orbit, I then had to decide on what vehicle. And ultimately that came down to a decision to use a transit vehicle as the tanker. And I'm surprised no ones yet criticised that because I'm not absolutely happy with it either. However, a transit vehicle is a good thing to have in low Mars orbit. For the reasons discussed above, and because it has the tankage needed. Now of course using it as a fuel store for the lander means the transit vehicle parked in low Mars orbit stays there. And if you send one on every mission you end up with a small fleet of them. A little wasteful but also handy because the more there are, the more places in low Mars orbit you can find a refuge with life support.

I'm also not entirely keen on landing ascent fuel using one of my landers. That's the single biggest use of fuel taken from low Mars orbit and without that I could probably make the transit vehicle parked in low Mars orbit last two missions. Nevertheless, taking what I just said as a starting point I'm left with a bit of a juggling act.

Instead of just having one transit vehicle that shuttles back and forth, I've actually got three. Although two of them effectively act as the one vehicle. The manned vehicle. The other lone transit vehicle is sent unmanned, propulsively captured into high Mars orbit, then aerobraked into low Mars orbit. Now, what I'm thinking of is rotating vehicles (like car tyres.). So the life of a transit vehicle goes like this. First, its launched from Earth containing all the spares and consumables and fuel needed for a manned mission. Now its partner might join it in LEO (not decided on this). Various masses are distributed and the vehicle checked out. The whole kit is boosted to a high Earth orbit and waits there for the crew to arrive. A separate flight brings the crew. Both vehicles transit to Mars (high orbit) and remain there for the duration of a mission. They return to high Earth orbit. The crew transfer to low orbit and then back to Earth. Now what happens here is that newest transit vehicle now waits in Earth orbit ready for a bran new vehicle to join it. It then does the Mars and Earth mission. After the second mission and on return back to Earth, it then becomes the vehicle that is sent ahead on an un manned flight to become the vehicle parked in low Mars orbit. So its already served two complete missions at that point.

So for every manned mission, one complete new transit vehicle is launched from earth. Which then becomes half of the actual vehicle sent to Mars.

Part of the reason I've done all of this is to limit the range and size of boosters needed. I figured on being able to use a booster capable of launching a bit over 50 tonnes to LEO - basically a Falcon heavy. For each manned mission there are 6 launches. Goes something like this.

First two launches are for the unmanned portion. What is delivered into LEO is a fueled transit vehicle (maybe a lander or if not some other vehicle designed to land fuel on Mars). And a smaller hydrolox booster. These are paired and sent on a lower energy trajectory that ultimately arrives in low Mars orbit ahead of the crew. One launch is in the order of 45 tonnes. The other is in the order of 30 tonnes.

Next four launches are..

A transit vehicle, fully fueled and carrying supplies. Topping out at around 50 tonnes.
A hydrolox booster, again around the 50 tonne mark.
A fuel tank providing fuel for the other transit vehicle. Plus the lander. Another 40 tonnes.
Finally, the crew. To the ISS in presumably something like a Dragon.

So I'm thinking in terms of a shade under 250 tonnes in low Earth orbit terms. That's per manned mission, ignoring the other one-way stuff that's actually landed on Mars.
Single biggest improvement on that would be using aerocapture. With aerocapture you could take 80 tonnes off that.

As far as what gets landed on Mars (and stays there) I'd personally be pleased to have at least one fixed, rigid structure. But I've no problems with rovers etc. If only Mars drive could get their power problems worked out.. One thing I'd like to see is a fixed pressurized structure that can nevertheless be disassembled into tow-able units. Certainly I think that the kinds of materials you're talking about are inevitably going to be part of this.

I'm starting to feel like a camping/caravan holiday here

Ok, probably left out a fair bit of detail here. Have to post and run. Let me know if I've missed anything you feel should have been talked about.

Last edited by Russel (2013-07-05 03:24:20)

RobertDyck · 2013-07-07 09:36:26

Russel wrote:

I guess my first question is what is the anticipated mass of the ITV and which Mars orbit is it using? If I were to guess, what you're aiming for is a high Mars orbit, closer to escape than to low orbit in terms of energy?
And the basic idea is that the ITV cycles between low Earth orbit and a high Mars orbit?

yes

What I'd probably disagree over is the use of a Mars ascent vehicle to carry the fuel needed for TEI.

Mars direct has an ERV that launches directly from Mars surface to Trans-Earth trajectory. Lifting the habitat for Earth return out of the gravity well takes a lot more fuel. Leaving it in Mars orbit, only lifting fuel, astronauts and samples, requires a lot less fuel. So what I describe requires LESS fuel.

Now, I'm happy with an ascent vehicle that has no pressurised compartment for crew. A faring with seats. But a store of air. Comfortable for a few hours, with a contingency for a couple of days in an emergency.

Yup, that's the idea.

The thing I didn't like...also places constraints on navigation. Far easier to aim for a nearly circular low orbit...wait for right launch window to take you to high orbit.

Gemini 6A was able to launch directly from Earth to rendezvous with Gemini 7. Just 5 hours and 4 minutes after launch, Gemini 7 was visible as a "star" from Gemini 6A. They them moved to 1 foot distance. That was 1965. On May 29 of this year, the Russians did it with Soyuz TMA-09A, rendezvous with ISS in less than 6 hours. This is how it's supposed to be done.

enjoy a beer

NASA doesn't like beer, or any alcohol. They can be prudes, quite prohibitionist.

As for aerocapture... I'd prefer to see...the vehicle has the capability to propulsively capture at both ends.

That defeats the point. The purpose of aerocapture is to reduce weight by not carrying fuel. Not fuel for propulsive capture.

My architecture isn't entirely dissimilar... It goes from a high Mars orbit to a high Earth orbit. ...what happens then.

One of my architectures did leave the ITV in high Earth orbit. Getting to the ITV is a problem. If you use a Soyuz or Dragon, it requires a larger rocket. Existing rockets are only able to lift those spacecraft to LEO. A custom launch vehicle as a space taxi, just to HEO, could be a show stopper. More practical to stage at ISS. Besides, aerocapture requires an eliptical orbit that dips into Earth's atmosphere. Might as well keep aerobraking until you circularize.

I stand to be convinced about the merits of making return fuel on Mars.

ISPP for LOX/L-CH4 results in 1 tonne of LH2 becoming 16 tonnes of fuel. That's very important leverage. And if you're going to rely on ISPP at all, in any way, then it's already a "risk". Wasting launch mass by bringing fuel for return to Earth only increases cost. Increasing cost kills the mission.

Instead of just having one transit vehicle that shuttles back and forth, I've actually got three.

Duplication means more cost. Increasing cost kills the misison.

for every manned mission, one complete new transit vehicle is launched from earth.

Defeats the point of reusable.

As far as what gets landed on Mars (and stays there) I'd personally be pleased to have at least one fixed, rigid structure.

Yea, that's ideal. But we need to keep cost down. That means keeping launch mass down. High cost kills the mission.

fixed pressurized structure that can...be disassembled into tow-able units.

Moving an entire Mars Direct habitat over significant distance, is not practical. Moving a 2 story house thousands of kilometres, with no roads, and the rough surface of Mars? Not going to happen. Breaking a rigid structure into towable units? That breaks all pressure seals.

I included inflatables for two reasons. They're lower mass, and can be moved easly on the surface so they can be connected.

Russel · 2013-07-08 05:22:07

Whilst Mars Direct has the problem of the launch costs (on Mars) of getting a return vehicle back to Mars orbit, there is also the cost of getting that vehicle back to a near escape orbit. That's my original starting point. Which is why I opt for a lower mass (yes, less fuel) means to return crew to a return vehicle that is already in a near escape orbit.

As for the economics of sourcing return fuel from Mars surface, it goes something like this. In an ideal situation, 1 tonne of hydrogen landed on Mars will create 4 tonnes of methane, and thus 18 tonnes of fuel, and of this 5 tonnes makes it back to orbit. So a 5 to 1 payback looks great. But the following problems arise. Firstly, an ascent vehicle sized to export the amount of fuel you need for a return journey from a low Mars orbit is going to be rather large. Basic figures are from high Mars orbit back to high Earth orbit, assuming aerocapture is 1Km/s (in an ideal case). Mass ratio 1.33. But from low Mars orbit you add another 1.3Km/s. The mass ratio now being 2.1 . Its this basic difference that allowed me to feel more comfortable with having fuel reserves for aerocapture, but that's an aside).

Essentially, in a very ideal case, to get back to Earth from low Mars orbit, assuming aerocapture, you need to double your mass with fuel. Now, a bare bones return vehicle might mass 25 tonnes if you scrimp a little. And there's reasons for thinking this is too low (including shielding). I use the working assumption of a return vehicle with a mass of 40 tonnes all up. Spread over 2 vehicles, but with the freedom to concentrate mass related to shielding in one end, during flight.

Now, that 25 tonne vehicle in low Mars orbit needs 25 tonnes of fuel. So if you start for a moment and think about a basic crew ascent vehicle with a mass of (say) 7 tonnes. And a fully fueled mass of 25.2 tonnes. Then add to that the cargo of 25 tonnes of fuel for the return vehicle, plus the fuel to get the fuel up there, and you're up to 115 tonnes on the Mars surface. That assumes the extra cargo costs you nothing in structure. In reality the structure has probably doubled. And of course that means more fuel. End result is closer to 140 tonnes of ascent vehicle, fully fueled.

That's a serious bit of engineering. Now, the 100+ tonnes of fuel represents a ISPP challenge. It starts scaling things to the point where the plant and equipment is heavy enough and complex enough - well its just plain beyond what you'd risk doing on a first mission. Yes, sometime later you're going to want to build big ISPP plants for other reasons. But even Martian oxygen isn't free.

Bear in mind that in a conventional architecture, you've also got to land the ascent vehicle, and the hydrogen. And either you've completely lost the perceived mass advantages or you've actually gone backwards.

As I keep pointing out, the biggest single gain you can make in this system is to keep the return vehicle in high orbit.

Now about launching direct from Mars surface to a high orbit. It can be done in one step. And its not a huge navigational challenge. The real problem is, if anything does go wrong technically, your only alternative in most cases is to attempt to get back to a low orbit. It would be kinda handy if ending up back in low Mars orbit wasn't an end-of-crew situation. Like, having a habitable spacecraft waiting there. Eventually we're going to end up with this situation one way or another. Why not plan it.

Other issue is fuel. 4.5Km/s to low Mars orbit. 5.8Km/s to high Mars orbit. As much as I hate the extra step of refueling, it makes a lot more sense to do that than to design an even bigger ascent vehicle. Just to make it clear, we're talking here about the case where the return vehicle is in high orbit. Were it in low orbit you'd not have these issues, but you'd have the ones I generically talked about above.

So, bigger, uglier ascent vehicle (one that already is way over 100 tonnes if you want to use it as a fuel tanker for the return flight) taking you all the way to high Mars orbit.. or.. a return vehicle that uses a lot more fuel to get itself out of low Mars orbit...

Or.. a light weight ascent vehicle which only requires just over 2 tonnes of fuel to get from low Mars orbit to high Mars orbit.

Yes, I was joking about the beer. Which is kinda funny give the psychoactive substances they do take into space..

As for aerocapture versus propulsive capture. I've said that in an ideal world aerocapture is the way to go. I just find it more risk than its worth (and yet more development effort) when the price to pay for propulsive capture in terms of fuel launched over one or two missions, isn't worth it. Instead I'd prefer a gradual learning approach. Even being able to come at the problem with partial aerocapture. You see, aerocapture is about 0.9Km/s gain headed to Mars and 0.6Km/s on return (yes, this is ideal, it does get worse with a faster trajectory). Compare that to the 1.3Km/s from low to high Mars orbit, or the even larger gravity well on the Earth side and you realise there are lower cherries to pick. But again, I'm not opposed to aerocapture. Its just one thing we don't absolutely have to do before being able to actually get there.

Speaking of space taxis. That's where my lander and ascent vehicle comes in. It serves a number of roles. One of those could be low to high Earth orbit if you so wished. I've never really considered getting a crew capsule (notionally something like a dragon) into a high Earth orbit to rendezvous with a transit vehicle a big issue. We already have the hardware.

Let me make something else clearer. I draw a clear distinction between aerocapture and aerobraking. The former is necessarily a one pass affair. The latter can be handled with multiple passes. Aeorobraking I'm counting on. Aerocapture is still a hard thing to do. But again, the devil is in the detail. If I found out that because of the requirements of shielding the transit vehicle is already by design strong enough to deal with the stresses, and that the thermal protection itself contributes to the overall radiation shielding then that would tip me over in terms of would I do aerocapture on a first mission.

Regarding ISPP. I think Zubrin oversold things by talking about mass leverage ratios as if the oxygen produced by an ISPP plant is leveraged off the hydrogen. In reality the ideal mass leverage from hydrogen is 4 to 1. 1 tonne of hydrogen begets 4 tonnes of methane. The important thing to realise here is that you can actually produce oxygen on Mars without importing anything. It is in that sense "free". What you do by importing hydrogen is you reduce somewhat the mass of the fuel, in this case methane.

Look at it another way. You want a massive ascent vehicle on Mars. Say, 100 tonnes of fuel. Of that 77% is oxygen. You need to import methane, or hydrogen. methane means landing 22 tonnes. hydrogen means importing 5.5 tonnes. And then that perceived advantage starts getting whittled away by practical realities. Now, what happens here is that people start treating ISPP as a panacea to everything, so suddenly huge ascent vehicles look feasible. Well, kinda. But there's a host of other development and risk related issues that come with it. Its better to drill down and focus in on how to reduce the mass of the ascent vehicle. When you do that, the mass of fuel (not oxidiser) starts becoming very manageable.

Again, the absolute most we can hope to obtain from hydrogen import, is 4 to 1 mass leverage. Not 16.

Regarding duplication. Well, duplication is why cars are cheap, no? You see the majority of the program cost is actually development effort, not the actual manufacturing cost of the vehicle. Why do I want 2? Well, in reality I don't need 2. But then I'd need a bigger one. The sort of size I'm talking about (notionally 20 tonnes each vehicle) is not spacious. You could survive in one. But with two you've got essentially 2 spaces. One is a sleeping/shelter area. And the other is the main living space. Again, you could just merge the two and call it one vehicle. No biggie. But nor is there actually much additional expense in designing one vehicle, and then making two of them.. or three for that matter. Hence the architecture uses one in low Mars orbit as the essential (I think people will eventually come to understand this is essential) refuge. You see, one common design, multiple uses. So it goes for the lander/ascent vehicle. Aside from the surface gear, and the dumb boosters, you've got literally 2 vehicles to design. And the less things you have to get right, the more chance you have of getting it right.

And I repeat, you could merge those two into one vehicle. But why? Doesn't save much. Development costs are where its at. And in keeping them separate you get one thing I've never seen in a Mars architecture and that's redundancy in life support. Redundancy in power supply. Redundancy in propulsion. IN other words, major things can fail and you still survive. This is why I've invited people to talk about potential failure modes and then stack them up one architecture against another. Main reason I feel uneasy about most Mars architectures is that their transit vehicles are almost universally a single point of failure.

As for the second last paragraph. I'm aiming to keep costs down, by targeting the thing that inevitably costs the most, and that's development cost.

As for the remark about Mars habs, well I was being a bit whimsical. But it is a real challenge to design a decent Mars habitat that in theory could be packed down, or unbolted.

Or maybe we are going to end up with the classic 2 storey affair that's never going anywhere, backed up by long distance rovers which themselves carry inflatable structures. Let me make the following observation though. Beyond a certain point one thing that would be very, very handy, is a vehicle capable of long distance flight. Even suborbital flight. But certainly over distances over 1000Km. Part of what I'm trying to do is to point out that a well designed lander and ascent vehicle paves the way to precisely this. That then makes a whole lot of things possible that weren't before. Including dealing with long distance emergencies.

Btw, are you familiar with the camper trailer? Its a trailer that has everything you need down to the fridge and the whole tent unfolds from it. Its just a thought that comes to mind when I think of long Mars journeys. I guess more seriously someone is going to figure out the utility of making life support systems and other basic gear modular and transportable. Ok, enough raving for me

Last edited by Russel (2013-07-08 05:31:18)

RobertDyck · 2013-07-11 07:46:09

I'm having difficulty reading this. You keep blathering on about low Mars orbit. You admit that to keep it simple, park the return vehicle in high Mars orbit. Low Mars orbit is for satellites only. You also admit that ISPP results in each tonne of hydrogen brought from Earth becoming several in Mars orbit, but then you shrug it off. That sounds like a politician: admit you're wrong, but keep talking anyway in the hope you'll win by fatigue. At one point you called ISPP complicated. If you think it's complicated, they stay out of the way, let someone else design it.

Then you blather about manufacturing cost. Are you seriously suggesting manufacturing multi-billion dollar vehicles just to throw them away? If that's your argument, I can guarantee Congress will not authorize it.

If you are looking for excuses to bring return fuel from Earth, again that increases total mission cost. Congress will not authorize it. Either use ISPP for all fuel for return to Earth, or forget about being a mission designer.

Here is another simple rule; one the military contractors do not want to accept. Keep cost so low that NASA's total budget does not increase at all. What so ever. That means cancel Orion, redirect funds to Mars. And either cancel SLS, or incorporate it. But all funds for a humans to Mars program must come from funds NASA currently receives for resurrected Constellation program hardware. So no net increase in NASA funding. Congress will authorize that. Congress will not authorize anything that requires an increase. At all. Period.

Terraformer · 2013-07-11 10:17:35

I was under the impression that his point was, every tonne of hydrogen brought from Terra becomes four tonnes of propellent on the Martian surface, but that has to be launched into orbit, becoming again (near enough) one tonne of propellent in orbit, and as such you haven't actually gained any advantage. Not logic I necessarily agree with, since once you start mining water it helps to have the equipment already on Mars, but I can see his point for the first mission.

RobertDyck · 2013-07-11 16:07:20

Terraformer wrote:

I was under the impression that his point was, every tonne of hydrogen brought from Terra becomes four tonnes of propellent on the Martian surface, but that has to be launched into orbit, becoming again (near enough) one tonne of propellent in orbit, and as such you haven't actually gained any advantage. Not logic I necessarily agree with, since once you start mining water it helps to have the equipment already on Mars, but I can see his point for the first mission.

If that's the case, it's not true. Every tonne of hydrogen becomes 4 tonnes of methane on the Martian surface, but all the oxygen comes from Mars atmosphere. So that becomes 18 tonnes fuel total on the surface. Any spacecraft parked in Mars orbit, would be in high Mars orbit, and highly eliptical orbit. The orbit periapsis (to be technically correct, "perigee" only applies to Earth), would be barely above Mars atmosphere. That's after aerocapture. Apoapsis (again generic version of "apogee") would almost achieve escape. That way it wouldn't take much fuel to push the ITV out of Mars orbit. Most of the delta-V would push it into trans-Earth trajectory. So if a large proportion of fuel is consumed getting up there from Mars surface, that's Ok. That's most of the way anyway.

And as you said, once you start mining water, you need equipment on Mars.

For chemcial rockets there is no difference between fuel and propellant. For nuclear thermal, ion, or plasma there is. I'm sure most of the people reading this already know that, but if I don't get the terms right, someone will correct me.

TwinBeam · 2013-07-12 19:24:40

All this verbiage to describe missions is confusing.

Maybe it could be clarified by using a format something like the following? (Example is very simple and one-way, numbers are bogus. More complex examples would leave mass/equipment behind for subsequent phases.)

================================================================
Mission: TotallyDirectToMar
================================================================
LOCATION Mass Consisting of these components
-------------------------------------------------------------------------------------------------------------------
Mass@Earth 800MT 1 heavy lift launcher, fuel, Mars transit vehicle fueled, etc etc
Mass@Mars Atm Entry 0
Mass@Mars sub-sonic 0
Mass@Mars Surface 0
-------------------------------
** Launch direct from Earth's surface to Mars atmosphere entry, and <brief notes justifying/explaining the transition to next phase>
-------------------------------
Mass@Earth 0
Mass@Mars Atm Entry 40MT 30MT fueled Lander, 10MT Aeroshield
Mass@Mars sub-sonic 0
Mass@Mars Surface 0
-------------------------------
** Aerocapture to sub-sonic and <notes explaining why that should work>
-------------------------------
Mass@Earth 0
Mass@Mars Atm Entry 0
Mass@Mars sub-sonic 30MT fueled lander, etc
Mass@Mars Surface 0
-------------------------------
** Rocket landing and <why the fuel expended makes sense>
-------------------------------
Mass@Earth 0
Mass@Mars Atm Entry 0
Mass@Mars sub-sonic 0
Mass@Mars Surface 8MT Lander, 3 crew, supplies, whatever
-------------------------------

Yeah, it's a bit bulky, but it'd make it easier to compare mission profiles, especially when you're trying to explain what you think is wrong with someone else' scheme.
For crewed mission profiles, assume a pre-positioned hab with supplies, small nuke plant and small ISPP plant (rover fuel if nothing else).

Russel · 2013-08-02 02:45:30

RobertDyck wrote:

I'm having difficulty reading this. You keep blathering on about low Mars orbit. You admit that to keep it simple, park the return vehicle in high Mars orbit. Low Mars orbit is for satellites only. You also admit that ISPP results in each tonne of hydrogen brought from Earth becoming several in Mars orbit, but then you shrug it off. That sounds like a politician: admit you're wrong, but keep talking anyway in the hope you'll win by fatigue. At one point you called ISPP complicated. If you think it's complicated, they stay out of the way, let someone else design it.
Then you blather about manufacturing cost. Are you seriously suggesting manufacturing multi-billion dollar vehicles just to throw them away? If that's your argument, I can guarantee Congress will not authorize it.
If you are looking for excuses to bring return fuel from Earth, again that increases total mission cost. Congress will not authorize it. Either use ISPP for all fuel for return to Earth, or forget about being a mission designer.
Here is another simple rule; one the military contractors do not want to accept. Keep cost so low that NASA's total budget does not increase at all. What so ever. That means cancel Orion, redirect funds to Mars. And either cancel SLS, or incorporate it. But all funds for a humans to Mars program must come from funds NASA currently receives for resurrected Constellation program hardware. So no net increase in NASA funding. Congress will authorize that. Congress will not authorize anything that requires an increase. At all. Period.

I'm afraid you're going to have to a) read it in more detail and b) if you're still struggling ask some specific questions.

Let me explain to you why low Mars orbit figures in this. The cost of getting something from Mars surface to low Mars orbit is best summed up by the mass ratio of 3.6. The cost of getting something from Mars surface to a high (near escape) Mars orbit is best summed up as a mass ratio of 5.3.

Now, whilst it is indeed quite feasible to build an ascent vehicle that can take you from the surface to a near escape orbit, that difference in mass ratios adds up to quite a few tonnes. Not just in the fuel, but in the vehicle itself - which will need to be landed in the first place.

When you do the math there is a very clear advantage to having the minimal mass ascent vehicle, and refuel it in low Mars orbit. To get a 5 tonne vehicle from low Mars orbit to high Mars orbit requires about 2.3 tonnes of fuel. The same vehicle launched direct from the Mars surface to high Mars orbit would require an additional 8.5 tonnes of fuel at Mars surface but then would also be a heavier vehicle. That extra mass then scales the fuel. So you end up with something like (roughly) a 7 tonne ascent vehicle needing 30 tonnes of fuel when originally you needed 13 tonnes of fuel on the Mars surface.

Fuel produced on the Mars surface does not come for free. It inevitably adds to the mass of your mission back at Earth.

Ok, that's one reason I'm considering having a habitable vehicle positioned at low Mars orbit. The other reasons are safety and flexibility.

What you have to remember here is that the bulk of what you're transporting in the process of getting a crew safely to Mars and back is the consumables and what I'm doing is keeping the bulk of that mass in high Mars orbit. Which has a very large fuel saving.

As for your second paragraph. I think you need to read more carefully. Manufacturing cost is not the killer. Its development cost. And I'm actually trying to avoid throwing things away. That's why my transit vehicles are reusable. The concession made to the fact that these things ultimately have a life span is that when they have been through two missions already they are then left in low Mars orbit. There is a variation on that theme and that is to design them with a life span that exceeds ten years and as the project progresses, simply reuse them more on actual crewed transfer missions.

As far as overall costs are concerned, the biggest cost is development. Which is why I'm proposing to limit the non-Mars-surface part of the mission to essentially two distinct vehicles and no more. Compare that to a lot of other architectures.

Now on the point about hydrogen. I'll go back into that in my next post.

Russel · 2013-08-02 03:02:28

About the economics of hydrogen as a feed stock for making methane fuel. I would have thought I'd covered that pretty well above, but it seems I've failed to get the point across.

Firstly, in the ideal situation, one tonne of hydrogen will make 4 tonnes of methane. That in itself is not contentious. My problem all along has been when people take ISPP to be a panacea for all things and end up requiring ridiculous quantities of fuel to be made on Mars surface. And I also have a problem with Zubrin coming up with mass leverage ratios that take the mass of oxygen into account.

Oxygen production on Mars is a simpler, less energy intensive process than fuel synthesis. That doesn't mean it doesn't have its problems or that it doesn't entail mass costs back at Earth.

However the true basis of comparison is ultimately this. On the one hand you ship methane to Mars and then you use local oxygen. Or on the other hand you ship hydrogen to Mars and then you use the local oxygen. What is the mass leverage ratio of one option as compared to the other.

Well, in the ideal case, you ship a tonne of hydrogen to Mars and you get 4 tonnes of methane. With that you also obtain 14 tonnes of oxygen. Zubrin would claim this as a mass leverage of 18:1

Likewise you ship 4 tonnes of methane to Mars and you also obtain 14 tonnes of oxygen. By the same logic, that's a mass leverage of 4.5:1

In actual fact the ratio that matters is the ratio between the two outcomes which is 4:1

So, in the perfect world hydrogen wins over - by a ratio of 4 to 1 - in terms of Mars entry mass. Fair enough. My problem has always been that landing hydrogen has its costs. So does storing hydrogen. So does the process itself. All of which whittles away at the theoretical advantage. And all of which adds, cost, complexity and risk.

I'm not saying that hydrogen to methane synthesis isn't a useful technology - it is. But I am saying that it makes a lot more sense in the context of discovering and exploiting native hydrogen.

And the other point I'm making is that pure ratios don't give you a feel for the primary advantage of what I'm doing which is keeping the mass of the ascent vehicle to a minimum which means the masses we are quibbling over amount to a couple of tonnes if that, per crew. That's gotta be considered in the context where a) launch costs are becoming cheaper and b) sending a crew to Mars is going to involve tens of tonnes of consumables, not including fuel.

Now, regarding fuel production on Mars surface vis a vis fuel being stored in Mars orbit. The problem again is the 3.6:1 mass ratio just getting to low Mars orbit. It really doesn't add up trying to make fuel on Mars surface to be used in orbit or beyond. Its simply cheaper to lug that from Earth (or get it from Phobos or Deimos if that's even possible).

In a future where hydrogen is found to be freely available (doesn't cost a huge amount of energy) on Mars surface, then yes, it might make sense to make large amounts of fuel, in order to haul small amounts of fuel into Mars orbit or beyond. Maybe. The devil is in the detail. And I'll bet my house it won't happen until we already have quite a few years experience on Mars.

JoshNH4H · 2013-08-02 07:29:26

How do you propose to extract oxygen from the ambient environment?

Russel · 2013-08-02 08:34:16

JoshNH4H wrote:

How do you propose to extract oxygen from the ambient environment?

There are a number of possible methods. Here's a good reference to start with.

http://www.niac.usra.edu/files/library/ … ngland.pdf

Personally I prefer to see advances in electrochemical processes such as..

http://research.jsc.nasa.gov/BiennialRe … F/EA-4.pdf

Edit: Something I didn't know before. Martian air is 0.13% molecular oxygen.

Last edited by Russel (2013-08-02 08:40:13)

JoshNH4H · 2013-08-03 23:02:05

Those alternatives are both actually much harder than running the sabatier process. Here's why:

For getting the oxygen out, you have to repeatedly pressurize and depressurize, as well as cool the ambient atmosphere. To obtain 1 kg of molecular oxygen, you need to process 1060 kg of atmosphere. To obtain 75 tonnes of LOX, you need to process almost 80 million kg of the stuff. At ambient density, that's 5.7 billion cubic meters. 5.7 cubic kilometers. A cube of atmosphere 1.8 km (a bit over a mile) to a side. It needs to be compressed from .006 atm to several atmospheres, cooled, the remainder extracted, and then the process has to be repeated. Note that this process is inherently less than 100% efficient at extracting oxygen, so you will actually need to process more than this amount of atmosphere.

Please note that your source is misleading in this respect: While the ocean may have less free oxygen than Mars on a mass/mass basis, the oceans are approximately 100,000 times denser than the Martian atmosphere and therefore have significantly more oxygen in them. Please also note that the oceans have a negligible amount of Carbon monoxide. Please note that the Nitrogen in the atmosphere is actually a big problem here, because there is about 30 times as much of it as Oxygen, and the two are very difficult to separate. On the whole, this proposal is very difficult, very energy intensive, and very failure prone because of all the required components and the difficulty of testing it on a sufficiently large scale on Earth.

Carbon Dioxide electrolysis, on the other hand, is very difficult as well. It is a poorly researched procedure that is inefficient and requires high temperatures and pressures. The machines are prone to breaking down and are nowhere near the Technology Readiness Level (TRL) that they could be sent to Mars. There are reasons to believe that the procedure will always work less well than water electrolysis because liquid CO2 occurs either at high pressures ( =failure prone equipment and more mass) or low temperatures, which require heat pumps to remove waste heat. Furthermore, in any state CO2 is a poorer solvent than liquid water, which tends to increase the inefficiency of the system (highly efficient water electrolysis units on Earth tend to work with a 70/30 mixture of water and KOH).

Zubrin also suggests that the Reverse Water Gas Shift reaction:

CO2 + H2 -> CO + H2O

could be used to produce oxygen, by electrolyzing the H2O and reusing the Hydrogen produced therefrom. But wait! If you've gone this route, you need to bring Hydrogen anyway. And on top of this, you're already using the same technology (Down to steel pipes and ruthenium catalysts) used in the Sabatier reaction. Two reactors for the price of one, and less mass brought to Mars to boot. Your fears about Hydrogen boiloff and handling are massively overstated and unnecessary. This is something that we do all the time on Earth. Most of the difficulty is presumably related to human operators and oxygen atmospheres, both of which do not exist on Mars.

The point here is that Oxygen is not, as you claim, "free". Rather, it is an expensive and unnecessary result of your choice to forego Hydrogen in favor of Methane. It's a choice that has definite costs but you still favor for reasons that you have yet to elaborate, given the conclusive evidence that it does not represent an improvement in terms of safety or mission cost.

Russel · 2013-08-04 06:39:51

Without going into chapter and verse about the process engineering involved (I'm not a chemical engineer, but I do know my way around materials) what I can say is that no presently proposed process is ready for Mars. And every process that involves the use of Martian CO2 requires development of technology for efficient compression. The sabatier process is not by any means simple. And it involves working with electrolysis to boot. The point of the paper I presented is that there is ample room for improvement in electrolysis of CO2 using lower temperatures and more reliable materials. Now we can debate this back and forth but my original claims about conversion of hydrogen to methane, versus simple oxygen extraction being likely to be more energy intensive and involve more mass of equipment still stand. There is no free lunch. And nor did I claim that oxygen was "free". I said that its wrong of Zubrin to include the mass of oxygen in his leverage ratios. He simply hasn't considered importing methane or doing the comparison on that basis.

What I fear going on is a chicken and egg situation, where people refuse to fly to Mars unless they've actually committed to a development program that is actually about settlement. And I suspect that this way of thinking is what I'm dealing with here. I don't necessarily subscribe to colonisation. I do subscribe to exploration. But the problem is this. At the very least we need to extract oxygen from the Martian atmosphere. We need to do that in order to breathe. So we need to get that technology right independently of making fuel. If we settle for the less ambitious goal of importing methane we get to land successfully, leave safely, and have the time to then experiment with sabatier technology for the sake of future missions. If we insist we do it the Zubrin way we're just going to have to trust robotic precursor missions to get the technology right, and my instinct tells me that will inevitably delay the point at which we can safely set foot on the planet.

Now about hydrogen. You've either got hard cryo cooling, or you don't. If you don't, then you can use all the insulation you want but in your process plant there will be gaseous hydrogen, and in that state there is no known material that can contain it. It will leak. That's an accepted fact of Earthly process engineering. Either you've got hard cryo cooling, or you lose some hydrogen. Its as simple as that.

And hard cryo cooling (again a power hog) is another technology yet to be proven. If you can convince me it can be done (and I'm ready to accept it is possible) then of course the picture changes and there's other questions to be asked - like why bother converting hydrogen to methane, when hydrogen is actually a better fuel? This is what I mean. Suppose in both cases you have a 5 tonne ascent vehicle. One powered by hydrolox the other methane/lox.

For a 4.5Km/s dV you need in one case 1.2 tonnes of hydrogen and 7.4 tonnes of oxygen. But in the other case you need 2.8 tonnes of methane (0.72 tonnes of seed hydrogen) and 10.1 tonnes of oxygen.

You see, the version that requires you to convert hydrogen to methane might require 60% of the hydrogen, but then you need to create 36% more oxygen. And most importantly you end up with a heavier vehicle overall. And these sorts of trade offs occur all over the place. Again, conversion of hydrogen to methane makes a lot of sense if you've got a local source of hydrogen.

I'm afraid I don't have the energy to convince you that what I'm proposing overall represents more mission safety. I've already presented the arguments needed. I'll just say again that too many architectures that people get enthusiastic have little in the way of resilience and redundancy - especially on the long haul space transit part of the mission. I've tried to look beyond that and put forth ways and means of dealing with that issue and I wait patiently for others to get the point and catch up.

Last edited by Russel (2013-08-04 06:43:29)

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