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#1 Re: Human missions » Yet another Mars architecture » 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.

#2 Re: Human missions » Yet another Mars architecture » 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.

#3 Re: Human missions » Yet another Mars architecture » 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.

#4 Re: Human missions » Yet another Mars architecture » 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.

#5 Re: Human missions » Yet another Mars architecture » 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 smile

#6 Re: Exploration to Settlement Creation » Buzz calls for Mars Colony... » 2013-07-07 05:06:25

And I wouldn't be too fussed about a supple suit. I mean yes it would be nice. But given an agile teleoperated robot I'd rather sit in a shirt sleeve environment. That would cover a surprising number of tasks that don't simply require something to be pushed, shoved or kicked.

#7 Re: Exploration to Settlement Creation » Buzz calls for Mars Colony... » 2013-07-07 05:02:00

I'd be a lot happier if a manned landing on Mars were preceded with a lot of prior study of potential sites, which probably means landing a lot of robots. Hard part I guess is deep(ish) drills.

Two things I'd look for are a) ice and b) some special local landscape feature that makes it easier to build a base... like a small canyon, or a cave, or suitable rock into which to build one.

Barring really good knowledge about local resources, I can't see an initial landing being anything other than highly mobile.

Just my stray thoughts for the moment.

#8 Re: Human missions » Yet another Mars architecture » 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 smile

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.

#9 Re: Human missions » Yet another Mars architecture » 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 smile

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 smile Please, go ahead..

#10 Re: Human missions » Yet another Mars architecture » 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 smile

#12 Re: Human missions » Yet another Mars architecture » 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.

#13 Re: Human missions » Yet another Mars architecture » 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.

#14 Re: Human missions » Yet another Mars architecture » 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.

#15 Re: Human missions » Yet another Mars architecture » 2013-07-02 07:33:25

JoshNH4H wrote:

Reliable, lightweight aeroentry is a vital technology for the colonization of the inner solar system.  Why develop it later rather than sooner?  This is the question that you need to answer (I am after all defending the Mars Direct plan).  This is the question that I don't think you have an acceptable answer to.

*Yes, the US Congress and the American public.  What other country could go to Mars any time soon?

Reliable lightweight aeroentry is vital to landing large masses on Mars. And I'm not suggesting it be developed later rather than sooner. Seriously, I'm not. It just doesn't matter as far as the crew goes. It will always be safer to land the crew separately (And indeed that is what you're suggesting if you have the kind of mission where the crew arrive 2 weeks later).

I will differ on one issue and that's I think you don't need to land anything on Mars larger than about 15 tonnes. Practically anything you can think of can be built up from that scale. I see Mars drive as a response to the "big lander problem" but I believe they're going a bit too far in the other direction. In any case they don't really make the landing problem go away either.

You're defending Mars direct. Fortunately I'm not smile

What I do see though is a lot of frustration people have with NASA coming up with ridiculously large mission mass. And thus a lot of architectures that are attacking the problem from particular directions. Mars direct is largely a response to the NASA overkill. I kinda like it in some ways. Its got a certain minimalism. But, for me the solution ultimately lies in something a bit more considered, a bit more conservative, but still not wasteful either.

However we get to Mars, its going to take a decade or two. And that's time enough to for everyone to sit around and bang heads together and not get too wedded to their particular approach.

NASA isn't the only body that theoretically could do this. I think ESA has the resources to do this eventually. Japan, China could all play the part.

The thing about the US is that if you've got a great architecture, whose going to build it. If its NASA then you've got to get through its strange meld of hard core rationality, plain old fashioned bureaucracy, and US style corporate welfare. Not that the ESA doesn't have its own politics at play.

Point is, even the most spartan Mars architectures are not pocket money. Eventually the big space agencies will have to be won over. And if that's the case they're going to be won over with with architectural proposals that don't just save launch costs, but also simplify development effort and most of all don't make them look like they're taking unnecessary risk with their crew.

And that's about as far into politics as I dare venture. I'm interested in this as a problem worth solving that hasn't been solved well by anyone, yet.

We do have time to get the technology right, to scrutinize what we're doing in unprecedented detail and do it safely, and I might add, in style.

#16 Re: Human missions » Yet another Mars architecture » 2013-07-02 07:04:20

JoshNH4H wrote:

I think the point you've not fully understood is that I've always operated on the basis that anything critical to safety will be landed on Mars and verified as operational before the crew leave Earth. So the safety issue you imagine isn't there. It a very simple principle. An unmanned mission lands what you need on Mars. Crew follows. There's no such thing as crew landing with life critical equipment.

What do you do if your lander crashes?  Give up?  Tell the astronauts and the contractors that you appreciate all their hard work but this mission is cancelled?  Do you push your mission timeline back 26 months (At a cost of potentially billions of dollars, plus the more intangible risks of bad press and loss of interest by investors or Congress and the American public*)?  No, you don't.  You pull out spares and build new equipment where necessary and send it in the same launch window as your crew.  If possible send it so that it enters while the crew still has the option for a free return trajectory.  I would say that this is okay even if it means using less delta-V and extending the crew's transit time by 2 weeks (It's a trade-off: Mildly more radiation for the crew for the increased safety of knowing that you're not screwed if the aerobraking fails again-- likely or not it's an eventuality that any good mission design will account for).

Again, I reject your mass multiplier of 2 for aeroentry.  I disagree with the statement that the decelerations will be higher than powered landing.  Inflatable heatshields are, if not entirely necessary to land large payloads, what I would consider an enabling technology.  They make it possible to have a larger heat shield area given constant fairing diameter.  This means that you will shed more velocity in the higher atmosphere and lowers the peak heating, as well as peak decelerations.  I would expect rocket deceleration for the last half a km/s or so.  This means that you only have one heat shield and a bit of fuel+engines.

Reliable, lightweight aeroentry is a vital technology for the colonization of the inner solar system.  Why develop it later rather than sooner?  This is the question that you need to answer (I am after all defending the Mars Direct plan).  This is the question that I don't think you have an acceptable answer to.

*Yes, the US Congress and the American public.  What other country could go to Mars any time soon?


I'm struggling to follow you here. The first question you ask is "what do you do if your lander crashes?". Which lander in which scenario is my first question. So lets go through some possibilities here.

I think most Mars architectures call for some things to be landed robotically. Correct me if I'm wrong there. Now, for instance if you were to land an ISPP plant and then wait two years before landing humans, what do you do if the robotic lander crashes, or the equipment fails? Give up? Or do you fix the error, launch again, and tell the crew they're going to have to wait? As I see it, this particular outcome is a definite possibility in any architecture that seeks to establish mission critical infrastructure ahead of a manned landing. Can't be avoided.

Now, are you considering the type of mission where mission critical infrastructure is launched during the same launch window as the crew? If so you've got two basic options. One is you land all your eggs in one basket. That's the land everything on one very big lander approach. As we discussed above, this has some serious safety issues. Now an alternative which you've raised above is to send the mission critical stuff slightly ahead of the crew. Meaning separate landings.

Thing about this approach is now you've got two separate landings. And you're headed down the track where the manned lander (the one delayed by 2 weeks) is fairly minimal. And to the extent you're not weighing that lander down with stuff that could have been landed 2 weeks sooner, and thus you've got a lighter lander and one that you can afford to build more margin into, then I'm happy with it.

Problem though is that you've reached a point where if something goes wrong with the first lander then the crew are either committed to hanging around in Mars orbit for a long time, or a free return trajectory. And what you've done here is trade the possibility that the crew will have to wait another 26 months on Earth with the possibility that the crew will have to go through the return journey and then again wait until someone figures out what went wrong and there is an opportunity to fly again.

Now, I'm committed to safety wherever technically possible. Which means a principle of anything mission critical functioning, or people waiting. You've pulled me up on the consequences of that, but the reality is that mission scrubs and long delays are a possibility in absolutely every architecture.

I could have chosen to trade off safety in various ways. For instance, a free return trajectory is well within the capability of my system. So if you were really in a blind hurry you could indeed send everything on the one launch opportunity. Ok, so the non manned lander crashes. What then? I've sent two fully fueled transit vehicles towards Mars. They don't capture. Instead they return home. Nice thing here is that instead of having to be very fussy about the trajectory in order to guarantee a free return, I'm in a position to plan non free return trajectories that can be converted to a free return. Because I've got fuel reserves both in the lander - that was tasked to ferry the crew to low orbit, and in the two vehicles. At last resort I can limp home in on transit vehicle.

The other nice thing about not having to be too fussy about having a free return trajectory in the first case is to minimise the length of time the crew spend in space. An issue you've noted.

Now, honestly, faced with a choice between that scenario, and simply opting to send the non manned mission out 26 months earlier, I'd choose the later even before scratching my head about it.

You simply don't get a choice. There will always be situations where you have to scrub missions. And its better to do that with your crew waiting on Earth than all the other alternatives.

Quite apart from the ethical considerations of sending people into space in its own right, do you stop to think about the risks to reputation and public enthusiasm that are not just oh dear we'll have to wait but oh no we just cratered a 60 tonne lander and there were people inside? People will forgive mistakes if its just stuff. They won't forgive mistakes if needless risks were taken with people and you can stop and think about the risks taken with the shuttle program.

By far and away the safest Mars architecture is one where do everything to protect the crew, including avoiding hurling people through the thin Martian atmosphere strapped to a very large brick. And if your concern is that Mars should be an ongoing venture, the biggest risk to that is not delays - its the commission of inquiry that will follow a loss of crew and the revelation that you could have done it more safely.

If you don't get why an aerobraking landing is necessarily more physically stressful you don't get the fact that the Mars atmospheric density profile is essentially an exponential curve. The only way you avoid higher g forces is to slow down earlier. That's what a mostly propulsive landing does. To get the similar effect with drag you need bigger.. MUCH bigger. So that you're already getting significant drag higher up - around 80Km. Worse part is that anything you use to create drag either has to have a variable size, or you again suffer the fact that if you've created enough drag high up, you've created way too much drag further down.

Inflatable heat shields designed to create significant drag in the higher atmosphere will be huge - 50m for a small manned lander and well over 100m for the sorts of landers you're contemplating. And what do you do when the stress becomes too high? You have to shed it and rely upon a smaller drag device. This staging effect adds mass, complexity and risk.

How about a rigid heat shield? Well for a smaller lander (under 10 tonnes) you might consider 6m diameter providing you want to hit 4g lower down. But if you want to make the same heat shield big enough to create enough drag higher up to lower the g forces, then you're going for an in Earth orbit assembled 30m affair. Obviously the mass goes over the top. Try and do the same thing with an inflatable hybrid and you need to lose the inflated section before it shreds itself. More risk.

Essentially the laws of physics are against you. There is no known way to build something that provides enough drag higher up in the Mars atmosphere that would give you the benefit of lower g forces. And do so without defeating the purpose which is trying to avoid weight.

Now, there is lift. You'll notice that MSL used lift. But it did so at the expense of throw away weights.

You could really light weight wings. But you'd either need those to fold or more likely to jettison them. Again, more mass.

Even though a minimal lander might cost you 5 tonnes in low Mars orbit, there is the extra structure and the reason for that is that it's reacting against the shell of the heat shield. That shell has to be stiff in order to avoid buckling and cause aerodynamic instability. Then you have to have a back shell under aerobraking conditions. And every time you decide to make the heat shield bigger, you also scale the back shell. You're lucky to get out of it for 7 tonnes. Now, add your parachute or other drag devices if you deem them necessary (they may just make things worse) and then the terminal descent fuel which at 500m/s using lower Isp fuel gives you a further mass multiplier of 1.22. So a 7 tonne craft is now 8.5 tonnes. And that's not the end of the story.

See, I'm creeping up on 2 to 1 here? Now, we may do better than that. But not significantly better. Not when the scheme of comparison include the system wide consequences. A mostly propulsive landing will always require more mass in LMO. How much more doesn't particularly matter. What matters, and I've said this before, is that when you don't think about landing in isolation, and think about the whole architecture, it makes a whole lot of sense.

Now, some failure modes, please?

#17 Re: Human missions » Yet another Mars architecture » 2013-07-01 07:57:13

GW Johnson wrote:

When I ran my Mars lander studies,  I was able to achieve 5 km altitude at local Mach 3,  simply by shallow-angle entry from low Mars orbit.  The required delta vee for ascent (complete with gravity and drag loss estimates) is about 3.7 km/s to 200 km.  That's the absolute minimum for powered landing flown as ascent-in-reverse,  but you'll actually need around another km/s more,  to cover all the real-world maneuvering and hovering. 

To aerobrake from LMO,  with my rather large vehicles (at 400 kg/sq.m ballistic coefficient,  some 60 ton size at entry),  I needed about 50 m/s de-orbit burn,  and hit the "air" at about 1.6 degree down angle.  Local Mach 3 at 5 km altitudes is 0.7 km/s velocity.  So,  I braked by 3 km/s without using but 50 m/s worth of propellant.  From that point,  theoretically I need only 0.7 km/s to land,  but the actual real-world maneuver delta vee was closer to 1.4 km/s. 

Still,  mass ratio for 1.4 or 1.5 km/s delta vee,  or mass ratio for at least 3.7 and very likely 4.7 km/s delta vee,  just to land?  You decide which one is the bigger,  more expensive thing to launch and bring from Earth.  Bear in mind,  the ascent will require about 3.7 km/s,  no matter how you choose to do it.  That propellant all has to come from somewhere. 

The aerobrake entry velocity is 3.7 km/s,  which is a lot of local Mach numbers,  but also a lot less energetic than from LEO (11 km/s).  Peak heating rates are proportional to entry velocity cubed (not squared).  Mars is a lot less demanding.  PICA-X ablatives are a nice choice at 0.27 sp. gr. and about 1.6 inches thick,  for one-shot  vehicles.  My ceramic heat shield study showed that black-surfaced alumino-silicates would be fine,  even for direct entry at 5.6 km/s,  as long as the angle was shallow.  Skin temperatures at the stagnation point fell well under the 1290 C phase-change limit for shrinkage-cracking in alumino-silicates. 

My "stuff" would be reusable,  damage-tolerant,  and easily repairable for fully-reusable vehicles.  Think under an inch thick at densities near 0.03.  Like shuttle tile,  but a whole lot tougher,  and buildable in large bolt-on panels.  Two-component ceramic composite,  laid up a lot like fiberglass is. 

There's no need to try inflatable heat shields at Mars (plenty of use here from LEO though).  You just go from hypersonics to rocket landing.  There's a lot of tradeoffs yet to be done to select such a flight path,  but the feasibility is there right now,  and with very large capsule-like vehicles massing many tons. 

GW

As you probably know from the landing thread, I coded a simulator for landing. I got generally similar results. Where we differed was because of the initial hypersonic trajectory curving under gravity. Nevertheless the results that matter to me were to do with a small manned lander (5 to 10 tonnes). In order to give that lander enough margin (better than 5Km at Mach 3) I invariably had a lander with a base diameter over 6 metres. That suited me because a 10 tonne lander and ascent vehicle needed tankage of about this scale anyhow. But it was still a beast.

Its not quite as simple as an ascent in reverse. On ascent, atmospheric drag works against you. On landing atmospheric drag works for you. If you refer to my above posts you'll see the justification for the figure of 3.8Km/s. That takes into account drag, plus the fact that the planet and atmosphere is rotating relative to the inertial frame of reference in which the initial 3.5Km/s velocity applies.

So I'll stand by the 3.8Km/s which does include allowance for final maneuvering. And with it a mass ratio of 3.

The propellant has to come from somewhere, but so does every other mass.

Let's take another look at the problem of landing people on Mars and getting them back to Mars orbit. This time I'll come up with a fair representation of a conventional landing and attempt to compare it. It uses an expendable ascent vehicle which has been landed previously inside a conventional lander. It also uses a conventional and expendable lander.

The expendable lander has a mass of 7 tonnes. Taking a requirement of 500m/s for maneuvering and an Isp of 250 for the lander and the lander needs 1.5 tonnes of fuel. So 8.5 tonnes in LMO.

The ascent vehicle has a mass of 7 tonnes. Taking a requirement of 4.5Km/s for ascent to orbit and an Isp of 350 and it needs 18 tonnes of fuel. So 23 tonnes before takeoff.

The ascent vehicle needs to be landed in the first case. Given a landed mass to entry mass of 2 to 1 that means the ascent vehicle started life as 14 tonnes in LMO

18 tonnes of fuel for the ascent vehicle comprises 4 tonnes of methane and 14 tonnes of oxygen.

Now, for the sake of the exercise I'm going to suggest that for every tonne of oxygen produced on Mars you need 50Kg of ISPP and power supply. And for every tonne of methane produced you need a further 100Kg of ISPP and power supply. This is rough but it reflects the relative complexity and energy intensiveness of simply splitting CO2 into O2 versus the steps taken in producing methane from hydrogen.

So now we need 700Kg of plant for oxygen, and 400Kg of plant for methane production. Now, in fairness, the plant is reusable, or will be when its well tested and well understood. So I'm going to allow it to provide the fuel for 2 ascents. Meaning the mass per mission is 550Kg.

Of course we had to land that so we now have 1100Kg in LMO being for the propellant production.

We of course need to land hydrogen. And this time I'm going to go beyond my better judgement and only allow for a 25% loss factor. So 1 tonne of hydrogen needed or 1.25 tonnes landed.

That would be 2.5 tonnes in LMO but for the fact that a hydrogen lander is going to involve large tanks than if you used methane. 4 tonnes of methane might require 1 tonne of hydrogen, but 1 tonne of hydrogen requires 1.5 times the volume of the methane. Given a tankage factor of about 50Kg per m3 (that's a very light weight, composite tank) it turns out that the hydrogen needed requires an 870Kg tank. So we're actually dealing with 4 tonnes of hydrogen, tank and requisite landing vehicle.

Adding this all up we get

8.5 tonnes in LMO for the lander
14 tonnes in LMO for the ascent vehicle
4 tonnes in LMO for the hydrogen
1.1 tonnes in LMO for the propellant plant

Now, lets do the same sums involving a integrated lander and ascent vehicle doing a mostly propulsive landing.

We start with a vehicle with a mass of 5 tonnes on entry. Its using methane/LOX at an ISP of 350 and requires a delta V of 3.8Kms so a mass ratio of 3.
Thus the fully fueled lander has a mass of 15 tonnes in LMO. Ok, so its heavier.

Now the ascent vehicle is one and the same. And since its 5 tonnes it needs less fuel for ascent. With the same mass ratio as above, it needs 13 tonnes of fuel.

13 tonnes of fuel breaks down to 2.888 tonnes of methane and 10.111 tonnes of oxygen.

By the above rules we need 500Kg of ISPP plant and power source. According to the same rules as above (used over 2 missions) that's 250Kg of plant per mission.

And I'm going to land the ISPP plant via a conventional lander so with a 2 to 1 mass multiplier thats 1 tonne in LMO.

Of course the expensive step is delivering the fuel. For that I need a 3 tonne lander (minus crew compartment and life support) and 3 tonnes of methane. That's 6 tonnes. Again with a mass ratio of 3 I need a further 12 tonnes of fuel in LMO. So all up in LMO I'm up for 18 tonnes. Ok, I could have landed the methane more economically, but I'll get to that.

Adding this all up we get

15 tonnes in LMO for the lander
0 tonnes for the ascent vehicle
18 tonnes in LMO in order to land the methane
0.5 tonnes in LMO for the propellant plant

Now, at this level we're comparing 27.6 tonnes for the conventional approach versus 33.5 tonnes for my approach.
A little commentary is in order.

Despite the complex sums involved in methane versus hydrogen, that choice is a second order factor in terms of overall mass.

The expensive factor in a conventional approach is landing an ascent vehicle, which because its being landed in a higher stress environment, has to be heavier, all else being equal.

The expensive factor in my approach is landing the methane fuel itself in a lander that uses mostly propulsive braking. I could have instead adopted the approach of landing the methane conventionally. That would have saved me the better part of 10 tonnes, and put me ahead of the conventional approach overall. But I didn't do so ultimately because this leaves a host of contingencies. Having an identical vehicle parked on the surface of Mars means at the very least you'll have spare parts, if not a fully functional vehicle. Its also serving another purpose and that's storage tanks.

Now for the bigger picture. The most commonly accepted approach in getting back from Mars is having a return vehicle, fully fueled and provisioned brought to LMO. That's a mass of tens of tonnes. To get to a near escape orbit around Mars that vehicle needs a delta-V of 1.3Km/s. Another mass multiplier of 1.45. So relative to the same vehicle parked in a near escape orbit you've lost tens of tonnes of fuel.

Now, what I've done is park the return vehicle in a high Mars (near escape) orbit so I've incurred a very large saving reflected back to low Earth orbit. That's the primary insight here. In order to make this happen though I need a ferry that will get me from LMO to HMO. That ferry comes in the form of the ascent vehicle. Now a 5 tonne vehicle requires only 2.25 tonnes to act as a ferry between orbits. That's an enormous saving.

Now of course you could take a conventional ascent vehicle and do the same thing. Only it would be (somewhat) heavier and would have to be required to do multiple burns (mind you, ANY such vehicle would, given the possible contingencies involved).

I'm also requiring a separate habitat be sent to low Mars orbit in advance. That's the staging point. But such a habitat whilst based on the same design as the main transit vehicle does not have to use the same degree of provisions or shielding. In other words about half the mass of a fully provisioned vehicle of its type. That extra mass I can justify within the overall budget because of the very large mass savings on the main transit vehicle being parked in high orbit. And because its sent to a low Mars orbit in advance, it and the fuel that gets used by the landers can be sent on a lower energy trajectory ahead of time.

One other thing. The primary propulsion for the transit vehicles is the lander. Now, since the lander must travel with the crew and must return to high orbit, there is minimal further cost in using it as the primary propulsion for the return journey. Instead I could provide the return vehicle with a scaled up propulsion unit and ditch the lander. That would actually save some mass. But at the expense of not being able to deal with some contingencies particularly on approach to Earth space. The lander remains at all times a last-ditch lifeboat. And I like it that way.

Ok, enough. I hope that clears things up some.

Now, regarding the thermal protection. My interest there is about the transit vehicle. One option is aerocapture. Using the thermal mass of the water on board and a modest thermal barrier I believe it would be possible to do aerocapture. I lean against aerocapture for two reasons. One is that the solar arrays would have to be stowed. The other is that the entire vehicle might have to be strengthened to cope with the stress of aerocapture. (Or its just possible it might be inherently strong enough thanks to all the layers of hull and radiation shielding anyhow).

Now assuming that aerocapture can be justified, what I need is a material that is light weight, modular and stands up well to long periods in space. That's why I'm curious about your ceramic composite.

Now the interest bit is you say it can be repaired. How? And can that be done in space? Or is it better to keep spare panels?

As I said in an earlier thread, the energy involved in Mars aerocapture is about the same as the energy involved in a full landing from low orbit. But it gets worse doing aerocapture into Earth orbit. That's where the material will be tested. Any guesses at the temperatures?

#18 Re: Human missions » Yet another Mars architecture » 2013-06-29 03:42:49

JoshNH4H wrote:

A couple of introductory comments, then a more detailed response:

First, regarding Ad Astra: I'm sure you see the relevance of the first part of that section of my reply, if not the importance.  I went rather off topic for the second part, but while I had my methods laid out I figured I would address a claim by Ad Astra that I find infinitely frustrating due to its blatant falseness.  In any case, thought it may not live up to the extremely hyperbolic claims made by its developers, VASIMR is more or less flight-ready at a small scale-- assuming the claims made by Ad Astra are true.

As an aside, based on the numbers given in this report NEXT's performance is actually worse than I assumed in the above model.  To improve upon the numbers I gave would require better engine performance, either in the Isp range or specific power, as well as better specific power for the power generation system.  Again, please note that the numbers I used may be unrealistically forgiving on the power system.  That report used 6 kg/kW, which is 167 W/kg.  This is 1/3 of my assumed value, and changes the full-system mass multiplier for your proposal to about 3.3, which is actually worse than chemical thrusters out of LEO.  This is pretty relevant to the discussion because it eliminates the reasoning behind assembly in L2.  I know that you don't consider this to be central but where numbers can lead to consensus I see every reason to pursue them.

Regarding the mass multipliers for Mars Entry, I reject your figure of 2 for aerobraking.  Zubrin estimates that the mass of a heatshield will be 15-20% of the mass that it needs to land, and inflatable ones are by their nature quite light.  Fuel mass should be pretty low for the mild deceleration delta-V required, whether the fuel be methane or Hydrogen.  I'd say Methane may be preferable in this case, though I don't have much evidence in support.  I'd estimate a mass multiplier in the region of 1.5. 

The logical interpretation of "50% Hydrogen loss" is that you lose half of your hydrogen.  You're describing a situation with 33% Hydrogen loss.

Now, delving more into prime content:

I just don't see any benefit from making the mission components reusable.  It's a safety hazard, it increases the seed mass and the per-mission mass (if you want to reuse the rocket it's gonna be more massive, and you want to send fuel for it), and it requires additional development seeing as you're already developing the reentry technology.

This is the fundamental part of my argument.  Stated succinctly: You've proven beyond a reasonable doubt that your mission is different from the reference mission, and you've even argued that the increase and costs and the decrease in safety might not be too large (At least not the first mission; I do still intend to address powered landing).  However, you have by no means demonstrated that it is better

Before I launch (no pun intended) into a discussion of powered landing, I was wondering what your take on the L2 stuff was.  I think I demonstrated fairly conclusively that using electric propulsion to travel to L2 does not represent any mass savings over chemical rockets from LEO.  Do you agree with my assessment or are you of the opinion that it is invalid for whatever reason?

Regarding reusable rockets for powered descent:

You say that you're considering the landing of crew separately from the landing of cargo.  The problem is that these things are not totally separate.  There is no reason why one needs to be safer than another.  Without a landed powersource, the crew is dead, period.  Now, you claim that it is not necessary to develop "man-rated" landing equipment because you can always send vital equipment at the next launch window.  Let's say we need each piece of equipment to have a 1% total chance of not making it.  This means that, given two chances, the aerobraking would have to have a reliability of 90%.  Honestly, 90% isn't that much different in engineering terms from 99%, or 99.9%.  The difference is that you put in more time and money to really verify your method in all situations, vs. skimping on proper testing and simulating.  If you have a 90% reliable method, you have a method that works, and it's just a matter of making it work every time. 

This is a perfect time to develop this vital technology; Why skimp?  We're going to need high-reliability aerobraking eventually.  It's an investment that will pay for itself many times over, and we might as well start as soon as possible.

Now, you claim that the primary reason why EtO (Earth to Orbit) rockets have not been made reusable is that the delta-V to Earth orbit (~9.3 km/s) is much higher than the delta-V to Mars orbit.  While this is true, it's not the only or even primary reason.  Even a two-stage rocket has about 4.7 km/s per stage, which isn't that big of a difference from what you're proposing.  The reason why it's hard to make a reusable rocket is that multiple firings can damage the engine; the thermal cycling of launch and re-entry, followed by spending time on the cool surface of Mars (not to mention the cryogenics in the tanks) weakens the tanks and any weld connections that may be there.  Most rocket engines wear out after the equivalent of more than a couple minutes of continuous firing.  None of these issues are unresolvable, of course-- but the primary way to resolve them is to trade payload for rocket structure.  This makes your mass multipliers significantly worse. 

And again--Rockets are complicated machines.  You can't use them over and over and over again and just trust them not to break.  Especially not without a large ground team on standby to fix anything that may have gone wrong during a flight.  That's how you get someone killed.  That's how you get colonization initiatives ended.

So, I pose this question to you: Given the obvious and hard-to-fix safety issues, why shouldn't we just aerobrake?

I think the point you've not fully understood is that I've always operated on the basis that anything critical to safety will be landed on Mars and verified as operational before the crew leave Earth. So the safety issue you imagine isn't there. It a very simple principle. An unmanned mission lands what you need on Mars. Crew follows. There's no such thing as crew landing with life critical equipment.

I'd put it to you that the shoe is on the other foot here. You land crew with cargo and if the cargo doesn't function, you've got a safety issue. Even if you have landed successfully. Safety alone dictates you take the alternate route and land anything critical to safety before the crew leave Earth.

Once you've understood this, its possible to decouple the quite separate issues of landing large masses, and landing crew. In doing so you can provide more safety for the crew than you can possibly do were you to land the crew on a large vehicle.

I stand by a mass multiplier of 2 to 1. That's a 50% landed mass to entry mass. No landing on Mars so far has come close to this. NASA thinks that to get 30+ tonnes to Mars you need 100 tonnes at entry. Zubrin is a lot more optimistic. I'm adopting the middle ground. I think that a mass multiplier of 2 to 1 is possible with large payloads.

However with smaller vehicles you'd be quite lucky to get to that 2 to 1 (landed versus entry mass) figure. I was being fair. Zubrin btw is primarily interested in large vehicles.

I tend to agree that an evolved heat shield may amount to 15-20% of the overall entry mass. However one factor I keep repeating that doesn't seem to be fully understood is that apart from the heat shield the entire vehicle needs to be heavier because its being exposed to much higher forces during aerobraking than would be the case for a mostly propulsive landing.

So for instance, if you wanted to land an expendable ascent vehicle using aerobraking (inside an expendable lander), the payload itself - that is the ascent vehicle - must be heavier than if it were landed mostly propulsively. Hence one of the advantages of the lander and ascent vehicle being one and the same vehicle is that you can make the ascent vehicle lighter than could possibly be the case if the ascent vehicle were landed conventionally. Lets be more concrete about this. A mostly propulsive landing can be engineered with a peak deceleration of 18m/s/s. (About 2 Earth g). A landing involving aerobraking involves upwards of 40m/s/s. The reason for this in simple terms is that you've burned off most of the velocity in the initial burn and the final burn is of comparable scale to a terminal descent from aerobraking - so around the 500m/s mark, give or take. From there to zero vertical velocity in 40 seconds without needing to exceed the 18m/s/s mark. Now with aerobraking you're a prisoner to the density profile of the atmosphere and you're having to burn through all that velocity much closer to the ground.

Now, adding more drag. A larger heat shield. Parachutes. Ballutes. Anything you can imagine. They can only slow you down faster in the denser layers of atmosphere. Which means the more margin you want to build in, the higher the stress on the vehicle structure. And that affects everything. Framing, tankage, even down to the little bolts that secure the plumbing.

There are of course ways out of this. Generating more drag higher up will give you lower stresses. But this means something huge, and light, and thus fragile. And probably something you will have to jettison in favor of a smaller drag device (going from ballute to parachute) at a relatively high altitude and as well as creating more things that can go wrong, will be pushing the limits of the second tier drag device. There is also lift that buys you more time but also buys you into more complexity and more pushing of the limits of materials. Even a trailing ballute device engineered to provide lift still has the problem that as the air becomes more dense it will generate too much drag and either self destruct or put too much load onto the vehicle.

Worse, all of this gets harder and harder to do, the larger your vehicle is. Buying margin for your crew is a lot easier and less expensive on a smaller vehicle. Taking a vehicle that masses 60 tonnes at entry and trying to give it the same margin as is possible with a smaller vehicle is, well, frankly bordering on show stopping. And there will always be the temptation to cut corners. You either end up with outrageous situations where you've reached final maneuvering (most of your vertical velocity is gone) with a Km to go, or you end up discovering that all your margin is gone and the only way to make up for it is to land people on lower terrain.

The initial burn of a mostly propulsive landing takes away most of the velocity and most of the difficult and thus safety critical materials issues. I'll say that again. Aerobraking is a difficult and dangerous process because you're stressing materials to their limits. And no I'm not particularly worried about ablatives there. I'm talking about the materials you're going to use for ballutes and the like. I don't mind those things if they're part of landing a non manned cargo. (After all, your crew isn't going to leave Earth until critical cargo has been landed safely and is functional, right?). That way the risk of very large drag devices isn't an issue of crew safety. And lets not forget also that landing large masses on Mars absolutely begs for these kinds of devices.

Another point about safety. When you land propulsively you're given back the ability to set the parameters. You've restored authority over the situation. You can land anywhere on Mars. You can choose the margin. It does cost fuel sure. But you're in a position where any other drag device you deploy merely adds to the overall level of comfort but you can still land safely even if it fails. Crew survival doesn't then absolutely depend on the correct functioning of such devices. With a small vehicle and an aerobraking landing that depends on drag devices you're taking a risk on those devices. With a much larger craft that has to land cargo and at the same time maintain margins for crew you're further pushing the materials. You're further increasing risk.

Final point on this issue. Much of uncertainty in an aerobraking landing is due to uncertainties in the atmosphere itself. You can to a certain extent compensate for that. That's what MSL was programmed to do by using a small amount of thrust to essentially fly its way out of these uncertainties. Nevertheless this problem exists because you're relying on air to slow you down. A mostly propulsive landing is affected less by the uncertainty in the atmospheric profile and also has more ability to correct for it. If you care about crew safety its little details like that that matter. And landing crew with a large vehicle makes this problem also much more difficult. And more difficult still if the large vehicle is relying even more heavily on cutting edge drag devices, and has more inertia and is thus harder to "fly".

Lets go to the issue of engine reliability. Reliable engines are not black magic. It simply requires competent engineering. The problem with most engines are they are deliberately designed down in mass knowing they only have to work once. There are wel l known examples of engines that can survive multiple restarts and long burns. The price to be paid of course is extra mass. But that's extra mass on top of a component that overall is not a large fraction of your overall mass budget. Put it this way, given the choice between an engine with a 100:1 thrust to weight ratio (relative to Earth gravity) that is designed for single use and an alternative engine with a 70:1 thrust to weight ratio (its heavier) that is ultra reliable. Which one are they going to specify, even if the mission only (nominally) requires one firing?

I'll also repeat something I said before. There are many Mars architectures that rely upon methane/LOX engines that are capable of multiple firings and long burns. We cannot avoid either designing such an engine or testing it.

Putting an Earth return vehicle on low Mars orbit, or even resorting to landing it on Mars does not obviate the need for reliable engines. Instead a sensible mission involves the capability of multiple firings if not for course corrections then for unforeseen maneuvers or contingencies. Engines deliberately designed down (and I do mean deliberately.. its a choice) to the point of being single use, should never be sent to Mars.

Having a minimal mass ascent vehicle means less propellant manufactured on Mars. Less propellant (even if its just oxygen) means less mas both for the ISPP plant and for the energy source. Some architectures seem to take ISPP too much for granted. I'm merely making the point that if your sole concern is how much mass you need to get into low Earth orbit, sooner or later you need to deal with the mass of the ISPP system and thus the mass of the ascent vehicle itself. And the lightest possible ascent vehicle is the one you've landed via mostly propulsive means, since its a gentler landing.


Even in a single stage to orbit Mars ascent vehicle you still have to incorporate features that are specific to landing. Otherwise you've no abort to land. And designing a lander and ascent vehicle together guarantees that the ascent vehicle is guaranteed to have abort to land capability. How many Mars architectures have been thought through this far?

I'm not asking my engines to be reused over and over again for multiple missions. They're being used once for a crewed landing. Once for a crewed ascent. During those two critical times, there is redundancy. At all other times there is even more redundancy. Although I'm using the lander as the primary propulsion source for trans-Earth that's in the context where both the transit vehicles have their own secondary propulsion using 2 of the same engines. Beyond return to Earth the lander is recycled as a vehicle that will and fuel on Mars. That's a non manned mission. And a non safety critical process for exactly the same reason discussed above. And beyond this the lander becomes spare parts.

What I'd ask you to do is to specify exactly the failure mode and scenario and then I'll address a response to that.

And for the sake of completeness I'll point out again that one thing I like about a mostly propulsive lander and ascent vehicle is that its inherently simplicity. Not just less things to go wrong. Its also about servicability. Its all out in the open. Its easy to visually inspect. Its easy to repair. And the lander that ferried fuel to the surface is a ready made source of spares, including whole engines, which at this scale can be man handled. The crewed landing is effectively being done on pristine engines. You land, you inspect, you swap components if necessary. You've got an ascent vehicle that has a firing sequence that if there are problems in the first seconds you do a soft landing and go back and fix the problem. And having transit vehicles in both low and high Mars orbits gives you a very long effective launch window so these sorts of problems can be fixed.

Again, specific failure modes please?

As for your last question. Its aerobraking that has obvious and hard to fix safety issues. Its time to think laterally.

As for L2. Using L2 does not depend on solar tugs. But having a base at L2 might benefit from solar tugs - provided that we're doing this on a very large scale and over many missions. Otherwise the development cost of a solar tug doesn't save enough to be worth it.

My architecture does not depend on the means by which you get to high Earth orbit, or for that matter which high Earth orbit it is. It could be L2, or it could be simply a high, elliptical orbit. The latter being slightly less energy.

I have noted that the only real advantage (to me) in using L2 is that what you have at L2 isn't traversing the radiation belts on each orbit. Arguably a problem for systems. Definitely a problem for people. But again, the latter has workarounds. And the former probably means running systems in safe mode at times.

The big advantage for others in having a base at L2 is proximity to the moon. I'm not fussed. Its also having a steady platform at which you can do manned assembly. Again I'm not fussed because what I've proposed is bump together assembly.

As for large unmanned cargo deliveries I have no problem with simply very large heat shields. As much fun as inflatable heat shield extensions are, there is doubt in my mind that they cannot be matched by light weight but still rigid heat shields of the order of 30m diameter. Which of course requires manual assembly. And that to me means using low Earth orbit. Doing so (with a rigid heat shield) means taking the most advantage of emerging light weight ceramics (GW would be happy to hear me say this). Anyhow I'm not wedded to any particular solution and your mileage may vary.

#19 Re: Human missions » Yet another Mars architecture » 2013-06-25 12:06:35

GW Johnson wrote:

Hi Russell:

I think we're dancing around the extremes of a trade study here.  It is quite possible to maximize aerodeceleration hypersonically at Mars,  and then cope with a low altitude when you come out of hypersonics.  In my studies,  I just went to a retro-thrust powered landing,  since for 0.7 km/s (local Mach 3) at 5 km altitude,  even at low path angle you are a mere two minutes from impact. 

In your study,  retro thrust is used throughout the entry,  which is going to reduce peak heating and total heat to absorb,  however it is resisted.  This will also increase the altitude at which one comes out of the hypersonics.  That makes chutes and ballutes feasible,  and therefore a thing you could consider.  Although,  I don't recommend a chute or ballute simultaneous with retro thrust,  except for last-second touchdown of extremely heavy loads.  It's too easy for flight control to be upset by wind gusts laterally. 

The trade study would be when to start retro thrust during hypersonic entry.  You say at the beginning of the hypersonics,  I say at end of hypersonics,  maybe there is an optimum in between!  The idea here is to land the greatest tonnage of payload for the least tonnage of propellant.  That would be true whether or not you plan to make return propellant while on the surface of Mars. 

To my knowledge,  the "when-during-hypersonics-do-we-start-retro-thrust?" trade study has never been run by anyone.  People have shied away from concepts like that because of (what I consider to be relatively groundless) fears over firing through holes in heat shields and stability of retro plumes versus vehicle attitude control. 

There is something very serious to consider about hypersonic flight conditions and vehicle configurations:  extreme shock-impingement heating above about local Mach 5 or 6 speeds.  This is a demonstrated risk from the old X-15 program.  You must fly a very clean shape,  not an assembly of nacelles,  because each nacelle sheds a shock system that impacts the other nacelles,  and the structures connecting them. 

The X-15 flight with the scramjet test article replacing the ventral fin was the flight that reached Mach 6.67.  The scram nacelle bow wave impinging upon the adjacent undersurface of the fuselage nearly cut the tail section off the bird in a matter of seconds,  once about Mach 6 got exceeded.  And don't forget that the skin was exotic refractory Inconel-X. 

Josh:

There's rocket engines,  and there's complicated rocket engines.  Hydrogen-oxygen things tend to be quite complicated and therefore expendable one-shot items.  Its only advantage is 450+ s Isp.  Methane-oxygen might be complicated,  or maybe not.  It is 300-class Isp. 

XCOR has the habit of making very reliable,  very long-life reusable engines without much of an Isp penalty.  They've certainly done it with kerosene-oxygen (300-class Isp),  and I know they're working on methane-oxygen.  Their engines don't look much like those NASA has always procured heretofore.  They look like something you could safely put in an airplane.  Some of them use piston pumps,  not turbopumps. 

The easiest,  simplest,  most reliable engines of moderate thrust that I know of are the hypergolic-ignition units that use hydrazine-nitrogen tetroxide (also 300-class Isp).  Could be any of the hydrazines,  but MMH seems favored.  These can range from 1-pound thrusters all the way up to the huge OMS units on the shuttle.  Pretty reliable technology,  we've been using it since the late 50's.  Something like it powered the old Titan-II’s that launched Gemini. 

Unfortunately,  MMH-NTO doesn't yet sound like something we might make in-situ on Mars.  But it is quite dense and easily shipped,  and extremely easily-storable for years at a time.  I dunno,  but MMH-NTO sounds like an ideal Mars lander/ascent vehicle fuel.  Until we figure out how to make it there,  we'll have to bite the bullet and ship it from Earth. 

For the transit vehicle,  cryo boiloff problems can be solved with only a modest loss for using LH2-LOX (450-class Isp) or LH2-nuclear (750-1100+ class Isp as solid core,  depending upon details).  But I doubt cryo tankage like that could also be successfully adapted to resist the heat and air loads of aerocapture at Mars.  Too many conflicting requirements for the designs. 

MMH-NTO (300-class Isp) is a serious mass ratio penalty for transit,  but could easily be protected for aerocapture at Mars.  There's a lot off tradeoffs there.  Including LH2-LOX to Mars,  aerocapture,  and MMH-NTO for the return.  Haven't seen anybody look at that yet. 

I'm not yet sure at all how the ISRU propellant restrictions feed back into the lander design,  much less aerocapture vs transit vehicle propellant restrictions.  My point is that there are very serious restrictions on your mission design imposed by these choices.  Your choices are not free. 

If I was doing it without aerocapture or ISRU,  I'd used LH2-nuclear for the transit vehicle,  and MMH-NTO for the landers.  I'd stage the landers out of LMO and refuel them there for reuse with propellant I brought along.  I build this from 20-50 ton modules launched to LEO,  with rockets we already have by next year,  and docked there.  I’d spend my money on the mission that way,  instead of developing a giant rocket that can launch 100-ton modules,  but that nobody else has any use for,  yet. 

There’s a brute-force baseline,  with the best stuff we have.  Now,  try lowering the launched mass by aerocapture and ISRU,  while staying within the propellant-choice restrictions they impose on your designs.  That’s the way to reduce mission costs. 

GW

Sorry for the long quote but there are points here and there I want to pick up on.

Getting to mach 3 at a reasonable altitude and relying only on aerodynamic braking means having enough effective drag area. Which makes for an awkward trade between the diameter of the heat shield and overall mass. Using propulsion to some extent side steps that issue. In any event a vehicle that is designed without a heat shield and with SRP in mind probably will place the engines well outside the main body of the vehicle - effectively on booms. And that arrangement tends to add to drag. Another reason for placing the engines above the center of mass is that then adds to stability.

As far as drag devices go, what I have in mind that might be useful is a trailing drag device deployed not long after the initial braking burn. So starting at around 80Km into a relative air speed of at most 600m/s. And then letting it go at around 20Km whilst throttling up. During the time the drag device is deployed there is either no thrust, or low thrust. Prudence may call for low thrust purely to lessen the risk of restarting into a denser air stream.

The trades as to the exact thrust profile are beyond my resources. I'd like to see it done but I think before that's done we need to gain a better understanding of how real thrusters behave. That's why I hope they use Dragon to deliberately test this. I suspect that what we will find is an envelope. A graph plotting on one axis velocity and on the other density. And within that there is a boundary inside which SRP can be used. Or even a set of boundaries defined by the level of thrust. And it would not surprise me at all if allowable velocity rises as density decreases, and vice versa. So for me the trickiest trade is probably about knowing how soon and how fast to throttle up. Even with aerodynamic braking the same problem will apply. Do you attempt SRP at local Mach 3 or do you attempt it at a higher Mach number but with less density air. Or more likely do you ramp up your thrust according to some criteria.

Fortunately I don't anticipate going anywhere near Mach 5 once the initial braking burn has been done.

Regarding engines, its my understanding that a well designed methane/LOX engine will develop an Isp of 350 in a vacuum. Which is another point. The density of air on Mars is near enough of a vacuum that this won't affect the Isp of a rocket engine designed for vacuum. But the same engine approaching at a relative velocity of Mach 3 will probably experience some reduction in Isp.

As far as reliability goes, I don't see an issue in designing reliable methane/LOX engines. Ones that can spend long periods idle. Ones that can be restarted multiple times. Ones that endure long burns. Its a matter of good design and a lot of testing. As I pointed out earlier, the reliability I need for methane/LOX engines in my lander is the reliability needed for the methane/LOX engines in most other architectures. And I suspect that the hardest requirement is actually long periods of idle. Again, I'd very much want to be in the position of being able to visually inspect the engine before I flew on it. And having a lander that also becomes the propulsion unit for the flight home is all part of that assurance.

I am not sure what you mean by transit vehicle propellant restrictions.

As for a nuclear LH2 engine, where I'd like to see one tested within the context of my architecture is this. I propose to transfer to low Mars orbit a transit vehicle which is also a fuel depot and a lander (the one used in the previous mission but is now the lander that lands the next cargo of methane) All of this mass - around 45 tonnes if I recall correctly - can be sent on a low energy trajectory well ahead of the crew. If the booster responsible were nuclear then the modest power requirement would mean a compact core. Eventually if you could prove the technology and demonstrate its safety then a scaled core version could become the booster for the manned transfer. I would still rely upon stored methane as the return fuel. In this case the stack would look like this. A lander docked on top. Then a transit vehicle docked to a further transit vehicle and finally at the bottom connected by the same universal docking port the nuclear booster. The crew would initially travel in the upper transit vehicle which would put a lot of mass between them and the core. I think I recall that in NASA's reference design they wanted 10 tonnes of shielding for the manned version. Eep.

#20 Re: Human missions » Yet another Mars architecture » 2013-06-25 11:10:12

GW Johnson wrote:

Oops,  I forgot.  My oddball ceramic heat shield is retained (two ways) upon backplates that are part of the vehicle shell assembly.  The idea is to bolt-on directly to the interior structure,  big extended panels of shell plating that have integral heat protection.  They'd come off the same way for interior access. 

GW

Ok well in response to that one what I'm thinking here is that for my purposes most of the surfaces that need some form of heat protection are actually cryogenic tanks. The surface is actually kept cold by the contents so the purpose of the thermal protection is really to reduce the rate of boil off. Hence small shingles even with gaps is quite sufficient.

#21 Re: Human missions » Yet another Mars architecture » 2013-06-25 02:57:48

GW,

When I was considering a lander capable of a mostly propulsive landing, what I considered was the overall architecture. Hence the thread title.

The thought process goes roughly like this. A lander capable of landing mostly propulsively has a dry mass of around 3 tonnes, before you add the crew and basic life support. That's my starting point. Now, that vehicle does have tankage but the actual tankage can be less than a tonne. The rest of the dry mass is structure, attitude control, helium tanks and so on.

Now take the same lander and instead convert it to a lander that relies upon aerobraking and four things happen.

Firstly the dynamic pressure of entry increases and along with that structural loads and mass. So g forces go from a peak of 1.5g under the mostly propulsive case to closer to 4g.

Secondly relying mostly upon aerobraking means adding more drag area and thus despite less tankage, the vehicle actually needs a larger base.

Thirdly is the mass of the heat shielding itself. Now I believe (and I tend to share your optimism) that this is a second order problem. I think that light weight ablative materials capable of withstanding a few entries are going to cost a few hundred Kg. If your ceramic material is robust then halve that.

Fourth is that with hypersonic atmospheric reentry you need to add a back shell which adds further mass.

All of this adds up to a lander that probably tops 7 tonnes with crew. That includes life support, RCS fuel, but not terminal descent fuel. Terminal descent fuel is probably in the order of a further 400m/s so that adds a further tonne. More if you're not using methane and instead opt for lesser but less complex (in a small engine) fuels.

Now in order to use the same lander for ascent you need to deal with the issue that 4.5Km/s of delta V gives you a mass ratio of 3.6 which takes you to over 25 tonnes on take off. But now you've added extra tankage and structure. Which illustrates why its easy to go beyond 30 tonnes. Most of that mass is fuel and this has to be accounted for in the design of the ISPP system. And it has a knock on effect in terms of the mass of fuel that needs to be imported. This is true with hydrogen or methane.

You can of course eject the heat shield and back shell during landing. But what you end up with is a vehicle which as an ascent vehicle is somewhat fragile. And your abort to land options are more limited.

In essence you have to trade the fuel you'd use to land a lander mostly propulsively against the extra mass of fuel (and corresponding systems) that need to be placed on the surface. All of which have their own mass multipliers.

Now, viewed in isolation from the rest of the architecture the lander problem is not a huge deal. You build an ascent vehicle and wrap it in a conventional lander and land it on the surface prior to your crewed mission. I might add even in this case, even without the landing wrapper, the ascent vehicle is necessarily heavier than one landed propulsively. Then you refuel the ascent vehicle on the surface. Then you send the crew in a conventional expendable lander.

Its such an obvious process that it tends to be take for granted in many Mars architectures.

The problem starts getting interesting though when you consider what happens to your crew when they get into low Mars orbit at the end of a mission. Obviously there has to be some form of habitat there waiting for them. And that's where you start running into the philosophical differences between direct and semi direct camps. What I came to realise about this problem is that there is a lot of mass to be saved from keeping as much mass as possible in a high (near escape) Mars orbit. So what I've proposed is an architecture where most of the mass you bring from Earth is parked in a high Mars orbit. There is also a functional habitat that arrives in low Mars orbit. Its the same vehicle but its simply not as heavily provisioned. Its mostly being used as a tanker. But it also serves the purpose of a safe haven in orbit.

Once you think in those terms it becomes apparent you need a taxi of sorts. And the lander I've proposed is that taxi. And minimalism again counts in terms of the amount of fuel needed to transfer between orbits. Or for that matter transfer to and land on Phobos or Deimos.

I would consider a lander that uses a hypersonic heat shield and then ejects that and the back shell during landing and then goes on to become an ascent vehicle (even if by doing so it is necessarily a heavier vehicle during ascent. As I said, this still raises the concern that as an ascent vehicle is has less margin in abort to land scenarios.

What ultimately steers my hand though is how safe the landing phase is. A conventional hypersonic aerobraking landing is by its nature one where you have to make trades between the size/mass of the vehicle and the drag forces you experience in the lower atmosphere (below 15Km) and thus the final altitude at which you go into propulsive terminal descent. You're also committing yourself to a much more complex landing sequence where in essence, more things can go wrong. With a mostly propulsive lander there is much more freedom to control where you want to be and when. Its the difference between controlled flight and being a brick with a parachute attached. And with a lander that maintains propulsion its a lot easier to land safely at high altitudes. Even Olympus if you wish.

Try doing that with a parachute smile

Now back to thermal protection systems. As you've correctly pointed out the heating is a cube law. Which means that the first 2Km/s of braking overcomes much of the problem. That's achievable with an initial braking burn - perhaps down to 80Km. What I've tried to do from there on is to limit the heating to about 0.5W/cm2. That's basically 3 suns of heating. I did that because setting that limit in the simulator and then controlling the amount of propulsion to stay within that target was easily doable. I can do less heating but it does fractionally add more fuel.

What I'd like to do is to rely more heavily upon aerobraking, but again without adding too much in the way of structural stresses. The key challenge for me is that this vehicle is all out in the open. Its not hiding behind a heat shield. So things that stick out leg landing feet are going to get hot. Again, given the rule of thumb that the temperature in Kelvins is roughly equivalent to velocity I feel fairly confident in limiting the peak relative air speed to about 800m/s.

However, I could go further. And that's where light weight thermal protection comes in. So far I've left it at that because what is interesting me more now is the overall architecture.

Oh and one other issue regarding thermal protection. In my days before being an engineer I spent a lot of time being an apprentice mechanic. Which meant I got to appreciate stuff that was simple and designed to be serviceable. And I got to hate Austin 1800s smile In terms of a lander, what I found is that wrapping the whole thing in a heat shield and and back shell tended to make a lot of stuff harder to service, or even discover if it was damaged. I like systems that can be quickly visually inspected. And I'm definitely aiming for a lander design where whole engines can be manually replaced - even in a space suit. So for the moment what appeals to me is thermal protection in blanket form that can be manually wrapped, unwrapped and repaired.

Now, when it comes to ceramic materials the real question is, can you do it in a modular clip on, clip off fashion. That would appeal.

#22 Re: Human missions » Yet another Mars architecture » 2013-06-22 01:41:14

JoshNH4H wrote:

Russel-

The important point that I think you forget is that there is more to be landed than just the crew.  There will be a lot of other equipment that would need to be landed, and the mass multipliers for choosing not to use aerobraking would be truly horrific.  Most of the equipment sent to the surface will be important mission technology; it wouldn't be sent otherwise.  Loss of a hab could prove equally fatal as hitting the ground at 5 km/s. 

I reject your estimate of 50% loss of Hydrogen.  Given refrigerators and well-designed equipment, I would expect it to be closer to 10-20%, tops.  source.  Note that additional insulation, as well as the use of a reflective sunshield, could reduce boiloff to levels below those estimated by Dunnspace.  By cooling the Hydrogen below its freezing point one can improve the density further, and potentially eliminate boiloff entirely.  Do note that one would need the same size tank as for liquid Hydrogen because I would expect most if not all of it to melt.

Regarding loss through machinery: It all depends if you design for Hydrogen being present or not.  Once on Mars, boiloff doesn't matter because a properly insulated tank will result in boiloff at a rate slower than the Hydrogen will be used for fuel production.  Leaks and such should account for less than 1% of product mass.  Once synthesized, meth-lox can be refrigerated fairly easily.  In fact, if you design it right you can use the same cooling system for the Methlox as for the Hydrogen. 

Regarding the size of the ISPP plant, Zubrin actually built one that would be capable of producing 20 times its mass in methlox fuel.  This was done at a small scale and on a low budget ($50,000).  I would bet a huge mass reduction would be possible with some development, perhaps fivefold.  That would be leveraging of 100:1.  Power requirements are relatively modest, too.  It's also worth noting that this machinery would be fairly easily repairable by the crew were anything to break. 

I would like to reiterate:  Safe Martian aerocapture and aeroentry is not optional in the long term.  It's something that needs to be developed, and as GW has said it's not that difficult to do.  A one-use descent vehicle and a one-use ascent vehicle (Land it the orbital window before so that you have a chance to send another with the crew if necessary, in case of any failures).  One-use because rockets are complex and prone to failure.  I'd like to point out that there has never been a truly reusable rocket on Earth; The reason is that it's really difficult, requires a lot of development, and is expensive to develop.  If you've ever been to Cape Kennedy, they have an RS-24 (Space Shuttle Main Engine) on display.  This engine is "refurbishable" in that one doesn't dispose it after use.  It looks like it has a million pieces.  They spent a whole lot of money developing and operating it, and they still couldn't be sure that it would work for more than one launch.  What about a Mars mission do you think will change that difficulty?

wrt L2: It is a benefit, if you have electric engines with high enough thrust.  But I don't know just how much, and it adds another operation for unknown benefit.  It does sure reduce the delta-V a lot.  Why not do L1, out of curiosity?  According to wikipedia it has the same delta-V as L2 to Mars.  I'd like to point out that the delta V from LEO to L1 with low thrust is 7 km/s, instead of 3.77 km/s with high thrust.  This will reduce the usefulness of electric propulsion methods, though not eliminate them entirely.  For reference, the mass multiplier for H2/LOX from LEO to Mars transfer is going to be about 2.6-2.7 including rocket structure.  From LEO to L2 using electric propulsion (Isp=3000 s), assuming a power source with a specific power of 500 W/kg, all inclusive.  Ad Astra is very closed-mouth about the specific power or specific thrust of their device.  I will estimate that it will consume 250 W/kg.  Ad Astra says that their engine will produce 5.7 N/200 kW supplied to the RF generator.  Because they are generally evasive, duplicitous, and nontrustworthy in their published figures, I will assume that means 250 kW input to the engines as a whole, giving a thrust-to-power ratio of .000023 N/w. 

Assuming that you want your trip from LEO to L2 to take 6 months of continuous engine firing, you need .00045 N/kg of ship mass.  This means that for every kg of wet mass, you need 20 W of power.  This is 120 g of panel and engine per kilogram of wet mass.  The mass ratio will be 1.3, so that's another 230 g/kg of fuel.  Allow 50 g/kg of other structure (e.g., 5% of wet mass) and you end up with a mass multiplier of 2.  It's a savings, but it's not huge.  Engine mass is inversely proportional to travel time; The minimum travel time, when you have nothing but engine, fuel, and structure, is a bit over a month.  Keep in mind that one must then launch to Mars, requiring an additional .74 km/s, for a second stage mass-multiplier of 2.5 if you use H2/LOX to boost.  However, given that you are skeptical about the ability of Hydrogen to keep for six months (I'll be honest, I'm enjoying this part), using Methlox instead results in a mass multiplier of 2.6.  The conclusion here is that, based on my calculations, Pure chemical is equivalent to stopping at L2.  And yes, I did inflict a structural mass penalty for chemical as well--Not 5% of the wet mass but 12%. 

As an aside, please note that this calculation suggests that the 39 days to Mars from LEO proposal by Ad Astra is simply ridiculous, having a total delta-V of 8 km/s and assuming Aerocapture at Mars.  The mass multiplier for a 39 day trip would be 7, assuming additional structural masses to be zero.  Assuming 5% of the total mass is other structure, the mass multiplier is 10.  Most of this will be heavy engines and panels, which are a significant cost inandof themselves.

In this post I have assumed that firing time is the same as transfer time.  While this is a good enough assumption in Earth-Moon space, for interplanetary distances it's not.  Covering the 80 million km between Earth and Mars in 39 days would require a constant acceleration of .028 m/s^2, which would require 7.32 kg of engines and panels per kilo of wet mass.  The total delta-V would be 94.3 km/s, for a mass ratio of 23.2.  That is to say, for every kilogram of rocket, fuel, and structure, you can carry about -8 kg (Yes, that is negative eight kilograms) of payload.  Given that we can't build our crew cabin out of exotic matter, I think this really demonstrates the duplicitiveness of Ad Astra's advertising.  Even if the engines have no mass at all this mission is impossible.

I've not forgotten that there is more to be landed than just crew. Indeed I've acknowledged that. But I've also treated it as out of scope. Why have I done that?

Because everything else that must be landed on Mars is on a one way trip. And conventional, heat shielded entry vehicles are almost certainly the way to land those masses. And I've deliberately left that part of the enterprise out of scope because essentially it costs what it costs. In other words, landing the crew in a mostly propulsive lander does not change the cost of landing other materials. It simply changes the cost of landing the crew. Hence I can logically separate the two.

It would be unfair to describe the mass multipliers for a mostly propulsive landing as horrific. In simple terms, the mass multiplier is 3, versus 2 for a conventional landing. The question is not the mass multiplier, its the mass being multiplied. As I've made clear, in rough terms, to get a crew down to the surface of Mars and back again, it costs around 35 tonnes per crew in Mars orbit. To do that conventionally the cost is around two thirds that. That 10 to 15 tonnes extra, delivered to low Mars orbit is a real cost, and I acknowledge that. But mass isn't the only thing that matters. Other metrics such as flexibility and safety matter more. Most missions that rely upon conventional landers have other points of failure. Its those points of failure I'm trying to address. So this isn't just about the landing method per se, its about having the right vehicles. And its about those vehicles performing multiple roles, and thus allowing for simpler, but more focused development.

As far as loss of hab goes. Yes, it doesn't do any good to your mission timeline when things get cratered. But its not loss of crew if all that mass is landed, and verified before a crew leave Earth. Most people agree on this point anyway so I didn't bother to mention it. I'll repeat the point though that in landing crew along with a large object, such as a hab, you actually increase the risk to the crew or the cost, because in the one case you're dealing with higher velocities and lower margins (and quite probably lower landing sites as a consequence) and in the other case you're giving a huge entry vehicle more margin because there is crew on board, but obviously that costs.

Like it or not, the sanest method of getting people to Mars is on purpose built vehicles. Ones with low mass and where you can then afford to suffer other mass multipliers because ultimately the last few Km of descent does matter, a lot.

In factoring a 50% loss of hydrogen, I'm not saying you're losing half your hydrogen, I'm saying you're losing a third for what you started with. Just to make that clear. Secondly its not just boil off losses. Hydrogen has a habit of permeating most materials. Even in a liquid state. And because you've got complex shapes inside a ISPP including manifolds, heat exchangers and so on its nearly impossible to avoid some loss. The only way to really reduce loss with hydrogen is to quickly form it into water, and that requires previously generating a commensurate volume of oxygen. Now, that's feasible. But, then there is always a cost.

Now I won't quibble with you about the precise percentage of hydrogen loss. In the scenario above you need theoretically 722Kg of hydrogen. I accounted for 50% over and above that for both boil off and other losses. That made it 1083Kg. If the loss accounted for is only 20% then you need to import 866.4Kg. The net saving is 216.6Kg. Its called diminishing returns. As I said, given the risks you'd normally adopt a more conservative margin for hydrogen than you would for methane. Nevertheless I can adopt a margin of 20% and it changes nothing to the argument. The cost of converting hydrogen to methane ultimately boils down to more mass having to be landed. At the risk of repeating myself, the ISPP plant needed to be built will be more massive (over one that purely produces oxygen) and the power source mass will also scale. Now, I'd like to think that such an ISPP plant would service 2 or 3 missions. But its a risk to assume it would last longer. And the main reason I think hydrogen isn't as good a mass-save as you'd assume is that the vehicle that lands the hydrogen has to be larger and heavier. And that's a mass multiplier that gets worse for a conventional lander because a conventional lander has higher forces and thus that extra tankage and support structure for the hydrogen scales.

And again, the wider picture is that even with some optimistic assumptions, landing 866Kg of Hydrogen versus 3000Kg of methane translates to a mass saving in the order of 10 tonnes back in low Earth orbit. Now, it must be said that any mass saving is, on the face of it, a good one. But compared to the 200+ tonnes of mass launched to low Earth orbit that's a difficult judgement call as to whether you want to spend the extra $20M or so on launch costs, or you want to face the development costs, and risks of going directly to importing hydrogen on your first mission.

Again, I'll state that hydrogen does make good sense if its obtained from the planet itself. That's a goal I'm fairly confident will happen. But as I said also, I think that the process of scraping Martian dirt and processing it, at scale, is probably going to require some human supervision. Making 100% propellant production a goal for the 3rd mission. Now, as and when that happens, my architecture evolves accordingly. We then don't need to land any fuel, and as a consequence we only need one lander landing. And as confidence in the technology builds it might be possible to reuse the lander for more missions. Again, that's another reason why I feel its a good idea to take the lander back into Earth space with you because it gives you an opportunity to check out and refurbish.

Let me make it clear. I don't think the benefit of importing hydrogen (over the 4 times heavier methane) vanishes entirely because of the problems I've noted (particularly as a consequence of scaling the lander that lands the hydrogen). But practicalities do reduce the benefit obtained. And in the end, you have to maintain a sense of perspective and make judgement calls where a few tens of millions in launch costs are not worth certain added complexities and risks.

By the way, I'm actually in favor of testing hydrogen to methane synthesis even at an earlier robotic stage. I hope that clears things up.

As far as safe Martian aerocapture goes. That would be wonderful. The hardest decision for me is choosing not to use aerocapture on either end. Were I to be convinced of the long term stability of light weight ceramic materials as GW is talking about, I'd actually be tempted to do aerocapture into a high Earth orbit first. And the reason is that on return to Earth you can put the crew into the lander and it has the capability to both capture and put the crew into low Earth orbit, and presumably dock with the space station. If the aerocapture for the main vehicle doesn't work, well, you're going to have to build a new one. But if aerocapture does work then you can extend the same concept to Mars. This time you give the lander more fuel and its capable of doing both propulsive capture and docking in low Mars orbit (see the beauty of having some form of habitat there?). Those two procedures knock about 40% off the original mass. So its not to be sneezed at. But when you look at the numbers, the whole mission is perfectly doable even without aerocapture. Its just going to be cheaper with.

However you ran aerocapture and aeroentry together and that's a bit unfortunate. Aerocapture has the potential to save nearly a hundred tonnes IMLEO. Aeroentry - at least for the crew - is a second order saving.

I've done a lot of thinking about every single technique that can be used to trim mass. Before coming back to realise that mass simply isn't everything.

Now the reason why there has never been a truly reusable rocket on Earth is that Earth has a much larger gravity well. Its simply a lot more difficult to do. 9Km/s versus 4Km/s. That's the long and short of it. The reality is that a mostly propulsive lander for Mars requires a bit less fuel to land than it does to ascend. You can't do that on Earth of course. The mass ratios would kill you. But on Mars the mass ratio is about 3.6 going up and about 3 going down. Well within the realms of well understood engineering.

I'm not sure why we're talking about Ad-Astra. I think their engine is way cool. I also think it would require a legendary power source to make practical as trans-Mars and trans-Earth vehicle. More to the point though when it comes to dealing with Earth's gravity well you don't need a fancy new engine. A NEXT ion drive (or a bank of them) would do the job satisfactorily. I'm not advocating doing that. Indeed I'm happy with the 200+ tonnes IMLEO that would be required to do this with conventional boosters and a fair chunk of that obviously would be ditched before the whole project left the Earth domain. I'd be even happier with that of course, if aerocapture could be relied upon. I've just not specified aerocapture as essential because when its all said and done, if we spend a few hundred million launching rocket fuel (per mission) then I can live with that until its clear how to safely improve on that.

There is a case for L2 and its with solar tugs between there and LEO. But the development cost for a MW class solar tug I believe is such that it might be ruled out as a possibility for the first mission. I could be wrong. These guys who want to go to the moon again, and establish L2 might win their argument. And in which case you'd piggy back off that infrastructure.

Let me also bring in something that is more speculative but reinforces my feeling that the design has some truth to it. When you build a space hab or transit vehicle if you wish, there's a minimum mass, partly because of consumables and water. And also because mass is a desirable thing when you're avoiding radiation. Now, in my architecture you're travelling home with a reserve of fuel that's intended for Earth orbital capture. So up until relatively late in the trip, the fuel you've kept in reserve creates extra shielding, at least for a portion of your combined vehicle. If you do without that you've got to deal with the trade off somewhere. That's one thing. The other is that because you're more than likely to carry several tonnes of water you've got yourself a good heat sink. It turns out that in aerocapture you can simply point the blunt end of your vehicle into the air stream and youu've got yourself a big kettle. Now you probably need a modest amount of thermal insulation (light weigh ceramic) and the rest is handled by the water. You might end up with several tonnes of water heated to 60C (anyone for a hot tub?) but essentially its a fairly non stressful way to do things. Again, provided you get the thermal protection right. Can such a thermal protection layer also compensate for and act as part of the radiation shield. I really don't know.

#23 Re: Human missions » Yet another Mars architecture » 2013-06-21 21:34:17

GW Johnson wrote:

Hmmmm. 

Entry heating protection at Mars is nowhere near as difficult as at Earth.  Peak entry heating varies as entry interface velocity cubed,  and it's a whole lot lower at Mars,  by factors of 2 to 3. 

There's a really good semi-reusable solution with Spacex's PICA-X ablative,  which could be flown several times,  maybe even a few dozen times,  before replacement.  And,  it's fairly low density at 0.27 g/cc,  you need around 2-3 inches of it for "long" life,  and to cut off the conductive heat load. 

Except for the demonstrated fragility and high maintenance costs,  low density ceramics of the shuttle-tile type could be a longer-life solution.  It takes about half an inch or so of that stuff to cut off conduction,  and be processible for bonding the tiles.  Except for damage repairs,  the life is theoretically infinite. 

I have an oddball low-density ceramic-composite material that I made and tested for a different use over 30 years ago.  It has the density of Styrofoam (roughly 0.03 g/cc) and very low conductivity (around 0.035 W/sq.cm).  It's a two-component composite laid up sort-of like ordinary fiberglass,  and features redundant retention,  even in large panels.  It proved very tough,  very damage-resistant,  when I tested it so very long ago.  I'll present a paper on it at the convention in Boulder CO this August.  Much less than an inch of it would be needed. 

That gets you through the hypersonics.  That's where aero deceleration is most effective on Mars,  completely unlike Earth.

If your entry angle was shallow (say near 1.6 degrees at interface),  you come out of hypersonics at local Mach 3.  On Mars,  that's around 0.7 km/s at around 5 km altitudes,  typically (in anything big enough to carry people or tons of cargo).  That's too low to deploy a chute at all,  much less have it decelerate you any noticeable amount,  so why bother?  Just go to direct rocket braking-to-touchdown from there. 

You will use less fuel in a lower mass vehicle than in any other imaginable scenario,  done that way.  I already looked at a multitude of approaches.  That was the best.

Firing through openings in heat shields is no problem,  just seal the engine compartment behind the heat shield so that there is no through-flow through the hole.  You might need a tad of coolant gas injection into the sealed compartment to make up for the volume-filling transient as you descend,  but it's a very minor massflow. 

Plume stability for supersonic retro thrust is an issue often raised,  but easily addressed by using multiple nozzles canted off centerline just a few degrees.  Spacex knows of this stabilizing effect already,  why else would the super Draco thrusters be arranged to fire canted at 45 degrees around the edge of the Dragon heat shield?  You don't have to fire at 45 degrees around the edge,  you can fire through ports in the heat shield canted at only 5-10 degrees and still get stability. 

Landing big stuff on Mars is not really such a difficult thing,  but you do have to do something different than we have been doing all these decades.  Fear-of-the-new more than actual technical issues is holding us back. 

GW

GW,

Your approach - an entry from low Mars orbit with a shallow angle - is basically my starting point.

As you've seen in my earlier posts, I had focused on how to design a reusable ascent and descent vehicle based around a conventional hypersonic heat shield. Its indeed possible, but the critical problem is the takeoff mass. Which then scales the amount of fuel you need to get it into orbit. And as you recall, the ascent fuel was headed into the several tens of tonnes region.

Taken in isolation that's not a problem - just a bigger ISPP plant. But when I went away and thought about how a reusable lander and ascent vehicle integrates into an overall architecture (which is the main subject of this thread) I kept coming back to the fact that the takeoff mass of the ascent vehicle (reusable or not) is an issue that has to be addressed.

Another point here is that no matter how I come at it, the problem of ascent fuel (the fuel, not the oxygen) doesn't have an easy solution. Despite the enthusiasm for hydrogen, the perceived benefits of importing hydrogen are reduced or negated by the engineering realities. And if you're importing methane, again its the mass of the ascent vehicle that is the root cause of most of your problems. Either way you can't afford to have an overly heavy lander and ascent vehicle.

Now, in designing a lander and ascent vehicle that uses a conventional heat shield I kept coming up with a dry mass in the order of 10 tonnes. And even I would concede that figure was, if not optimistic, certainly going to involve some careful design and a lot of testing. Even with ceramics or a thinner ablative, there is still the support structure. And in brief, the kinds of stresses involved translate to more mass. More mass translates to more fuel on take off. Which again scales the tankage and so on.

In brief, the problem seems to be how to build something as minimalist as possible as an ascent vehicle and then secondary to that is how to land that same vehicle without adding substantial mass. And the mostly propulsive approach does that. It doesn't mean you don't need *some* thermal protection, but as you're probably aware the temperatures involved in Mach 3 or 4 are a whole different ballpark when it comes to materials.

Remember that at the temperatures involved there's lots of things that don't really need protection. Much of the bulk of the vehicle is tankage and the fuel itself keeps that cold. You only really need a thermal blanket to lower the rate of boil off. There's really only a few external systems that need special protection (electronics etc). And the whole thing is likely to be more amenable to hands on maintenance - which I think is an aspect that's often neglected. Designing stuff to be field serviceable. One thing that keeps coming up when I visualise the lander is that the conventional heat shield tends to make stuff less accessible to maintenance and repair.

There are still good uses for light weight ceramics and I'll be keen to see your paper. As it happens, I keep toying with the idea of allowing the transit vehicle (space hab) to have the capability to aerocapture, and besides having its own native heat sink (tonnes of water) the thing I'd need is a light weight ceramic. I've already allowed for aerobraking (after capture) but that's something that can be done gradually. Aeorcapture back to Earth orbit would be a real advantage, but for the moment I've gone for propulsive capture but with the understanding that aerocapture is an option and therefore planning for propulsive capture leaves you with more margin.

Now, as far as SRP goes, my understanding of the problem is you wish to avoid having the turbulent gas stream reacting against the vehicle. Now the Spacex solution appears to be to mount the engines above the plane of the heat shield and at a reasonable degree of cant. Meaning their solution (if it works) solves the problem but only where you aren't demanding a lot of delta-V from the propulsion system.

I can't rely exclusively upon the same approach because I'm demanding a lot more performance from the system. Instead what I would like to do takes a little description. Imagine a vehicle that's essentially 4 tanks arranged symmetrically around the vehicles vertical center line. These are your methane and LOX tanks. There is a center volume between these tanks and that's a good place to put the crew compartment. The actual frame of the vehicle is (if you're looking down at it) a four cornered square truss. That also supports the landing legs, which remain deployed at all times. Along each truss there are 2 engines. One of those as at the far end and its canted at a modest angle (say 15 degrees). The other engine, which is inside the landing leg, but still outside the projection of the body of the vehicle is pointed vertically. Both engines are mounted above the center of mass of the vehicle.

Now the trick here is to maintain efficiency, but use cant where necessary. For a start the main braking burn can use all 8 engines (or even just the inner 4 if that turns out to be optimal). It doesn't matter because the air is too thin to matter. During the mid part of the descent you'd either run with no engines, or use a low thrust on the outboard 4 engines. As you approach a key point (lets say 25Km) you throttle up on the outboard 4 engines. You'd rely upon that method to reduce the Mach number before then throttling up on the inner engines. All of this technique to minimise cosine losses. Its just possible though that so long as you mount the engines high enough and far enough from the center of the vehicle, that there's no problem and you can do what you please.

Now for the wider picture. I wanted to keep as much mass as possible in high Mars orbit. But doing so meant for various reasons having a staging point in low Mars orbit. The reason I originally started with the idea of 2 transit vehicles is that one such vehicle could then make the journey to low Mars orbit and back. So then you'd offload mass to the other vehicle and leave it in high orbit and then aerobrake gradually into low orbit. Still, the final ascent back to high orbit did consume some fuel and presented other difficulties. What would be good is a Mars orbital "taxi". A vehicle capable of transferring crew between orbits but otherwise fairly minimal. I also kept running into the issue of how to provide fuel in low Mars orbit for the lander. And remember that's still an issue even for a conventional lander. And of course in any scenario there's still the question of ascent fuel. So I needed a "tanker" in low Mars orbit.

The key to the solution was to make that "tanker" one and the same as a transit vehicle, and to send it on a prior unmanned, low energy trajectory. Which made the lander even more interesting. If I could have a lander that would act as the taxi then things start coming together. But again, that meant having a low mass lander. And this is where the bigger picture comes into play. A conventional lander might use less fuel going down, but its going to be heavier, and thus use more fuel as a inter-orbit taxi.

More than this, there is also the issue that the transit vehicles need a propulsion system, and as it happens the lander happens to be the right size. Now, one option is to never return the lander to Earth space. In that case it would go on to land the ascent fuel and that's its final purpose (apart from a backup and a source of spares). But that would mean more engines for the transit vehicles. More importantly I felt that having the lander there attached to the transit vehicles on the return journey adds more resilience. The lander would not in itself have a long term life support system, but it would be very useful if for instance you suffered a fuel leak on the return journey. Then you could transfer fuel back to the lander's tanks, abandon the transit vehicles and use the lander to do the propulsive capture back into Earth orbit.

And then of course, its a useful craft for visiting the Mars moons.

A very handy vehicle indeed, and only possible if its as light weight as possible. Now, I can't make that happen if it has a heat shield and is designed to cope with the stresses of a standard re-entry.

#24 Re: Human missions » Yet another Mars architecture » 2013-06-20 00:18:37

JoshNH4H wrote:

Assuming that these are the fundamentals of what you're proposing (nothing else major pops out to me) I would simply like to raise objections that you probably expected, but that I don't think were sufficiently addressed in your post.

The mass penalty for a propulsive landing is very significant.  Launch costs may have been decreasing of late but I don't think you can justify that kind of mass increase in the mission.  Especially not given that you don't intend to import Hydrogen.  By the way, the sabatier reaction is not complicated.  Zubrin demonstrated via experimentation that it's actually quite simple.  Hydrogen storage itself is not an issue, as long as you're willing to accept some level of boiloff.  Given proper insulation and a reflective coating to minimize sunlight, plus perhaps some level of active refrigeration I would expect it to be quite small.  This gives you a fourfold reduction in imported fuel mass over importing methane.  By the way, Oxygen may be present but the Reverse Water Gas Shift reaction is probably the best way to get at it.  While it has been shown that electrolysis of Carbon Dioxide is possible, it's a very inefficient procedure relative to water electrolysis.  It's also less well-developed, meaning that the equipment will have worse specific power and specific daily output.

Further, enabling technologies for landing on Mars are not far away.  The big one is inflatable heatshields, and work is being done on these at NASA as we speak.  Beyond that, we need to get a better understanding of the dynamics of the Martian atmosphere, which could be accomplished by sending balloon-type probes to different levels and observing their drifts, and allowing them to measure temperatures (Thermometers and GPS units are very light, easily less than 50 grams together.  I'd imagine a barometer could be built to be extremely lightweight using piezoelectrics.)   We might also want to fit out a couple with mass spectrometers to analyze local composition, including humidity.) or perhaps even through the use of a more typical weather satellite, of the kind we have orbiting Earth.  Beyond that, improved non-propulsive deceleration (both in the form of aerocapture and aerobraking) is an important enabling technology for the colonization of the inner solar system, and it's better that we develop it sooner rather than later, so as to squeeze out maximal benefit.

With regards to L2 vs. LEO or another Earth Orbit, it doesn't really matter where you do the assembly.  If you do it in LEO you have the benefit of shorter round-trip communication times with the hardware for more direct control/backups.  While the outbound delta-V is higher, it takes fuel to launch to L2 and this will cost every bit as much as launching straight out of LEO.

I would raise significant concerns regarding the use of a reuseable vehicle on the very first mission, when there are no people out there who can fix the rocket or even check it out properly.  The second use represents a huge safety issue, easily leading to loss of mission and loss of crew if anything goes wrong.

Yes, you're correct. These were expected objections smile

The mass penalty for a mostly propulsive landing is what it is. Two thirds of your mass on entry is going to be fuel. So question is the difference in mass between a mostly propulsive landing and a conventional landing. And the more important question is whether that can be justified in terms of:

a) overall dollar budget
b) development cost and complexity
c) simplicity of design
d) robustness, resilience and safety

Considering only the mass involved in transporting the crew from low Mars orbit to Mars surface and back again, my proposal looks like this.

The crewed lander is posited as being 5 tonnes of mass, without fuel, but including crew and life support consumables.
The lander which delivers methane to the Martian surface is posited as having a dry mass of 3 tonnes.
The landed methane. That's 3 tonnes.
The fuel needed to land the methane. 12 tonnes.
The fuel needed to land the crew. 12 tonnes.

So accounted for in this way, the actual process of landing a crew on the surface of Mars, and bringing them safely back into a low orbit, involves 35 tonnes delivered to low Mars orbit. I hasten to add that with the architecture I propose, all the mass (fuel, hab and lander) delivered to low Mars orbit has first been delivered to high Mars orbit and then aerobraked to low Mars orbit. (And strictly accounted for, the mass of the crew does not belong in the above accounting, but I'll leave it this way to be conservative.)

Now before going to look at alternative approaches to the problem, I'm going to make the observation that the 35 tonnes delivered to low Mars orbit in order to perform this function, compares to the 70 tonnes of mass delivered to high Mars orbit, which is the transit vehicles and their fuel supply. In other words, if a third of your mass budget is spent on safely landing the crew and returning them as far as low Mars orbit, then I'm comfortable with the result.

I should also add that there is additional mass in the habitat vehicle that is transferred to low Mars orbit and is used in part as a fuel depot. Its difficult to know where to account for that because in this context, much of the mass of that vehicle is actually life support and serves the purpose of providing a home in orbit for the crew. But I do acknowledge that these numbers are neither perfect nor final.

A conventional single use lander suffers its own mass multipliers. It needs to withstand the forces of a conventional landing. Meaning several tonnes of basic structure. Add to that the heat shield. Add to that unavoidable propulsion system which still needs to provide more or less the same thrust even if for less duration. And now its easy to see why Dragon heads north of 5 tonnes. Factoring in the uncertainties and the need for larger drag devices for Mars, I'd suggest that 7 tonnes is a conservative figure for a fully fueled lander of this type. Some have suggested much higher than this.

In addition you need an ascent vehicle landed separately. Now, with a minimalist vehicle (such as my own) then you have a take off mass of 5 tonnes (plus fuel). Some of that mass is crew. And some of it is life support consumables that can be produced on the surface. So lets be fair and call it 4 tonnes of ascent vehicle.

Then there is the imported methane fuel (I'll consider hydrogen shortly). A 5 tonne ascent vehicle needs 13 tonnes of fuel. That's 2.88 tonnes of methane. If its imported you probably need 3 tonnes of methane to start with. Now to import the methane you need a vehicle to land it in. The most economical way of course is to land the methane with the ascent vehicle, bringing the total ascent vehicle landed mass to 7 tonnes.

A conventional landing craft has a mass penalty of about 2 to 1, so 7 tonnes of payload means 14 tonnes at entry.

Which takes us down to a comparison between 35 tonnes delivered to high Mars orbit, then aerobraked. Versus 21 tonnes at Mars atmospheric entry (we assume direct entry). The former amounts to 111 tonnes in LEO. The latter amounts to 57 tonnes in LEO.

Total saving: 54 tonnes in LEO.

Now, that's nothing to be sneezed at. At $2000/tonne that's a saving of a bit over $100M in launch costs per mission.
But that should be seen against the backdrop of billions in development costs. And development costs are driven by complexity and time.
I do agree that in the long term, if we decide we want to return to Mars many times over, then an indigenous source of hydrogen is the key.

----

Now, to do a fair comparison regarding the importation of hydrogen I'm going to use as a baseline the conventional mission described above. A conventional 7 tonne single use lander and a previous landing of an ascent vehicle with methane fuel on board which added up to 14 tonnes.

From this baseline we derive a version where hydrogen is imported instead of methane.

I'm first of all going to introduce hydrogen losses into the equation. Its not just boil off, its also the losses incurred by passing hydrogen through the plumbing and heat exchangers of a ISPP plant. In short, hydrogen is innately leaky. I'm going to assign a factor of 50% for overall loss. That's even with very good insulation and some measure of active refrigeration, but allowing for leakage and for a safety margin (we're assuming the hydrogen is our first and only source of ascent fuel). With that we need 1125Kg of hydrogen.

Landed in a conventional lander, 1125Kg of hydrogen versus 3000Kg of methane translates to a difference of 3.75 tonnes at Mars entry. With the direct entry delta-V of 4.6 assumed above, that translates to a saving of 12.8 tonnes initial mass in low Earth orbit.

Now that's the scale of the maximum possible theoretical saving. About 13 tonnes of mass, and about $26M of launch costs but in return all the issues and risks inherent in hydrogen.

Of course if you decided you wanted to use mostly propulsive landers for other reasons, and still wanted to deliver hydrogen rather than methane. Then the saving in IMLEO terms in delivering hydrogen would be roughly 6 tonnes. Of course that's the theoretical maximum saving.

Beyond that theoretical best case the real world detail doesn't look pretty.

For a start, the ISPP plant needed to produce methane from hydrogen is likely to be more massive than the alternative which simply produces oxygen. Possibly by a factor of 2. The reason is that it involves more parts and more plumbing and heat exchangers (or else you sacrifice efficiency). Now its fair to say that such a plant is likely to see service for years so it can't be accounted for on a per mission basis, but the issue is there. The same applies to the power source. And again the task of producing methane from hydrogen is much more energy intensive. For two reasons. One is the inefficiencies in the process, especially if done on a small scale. The other is the refrigeration costs. Often we're dealing with very hot gases but the ultimate product is a hard cryogen. I won't go further into this but its a real issue.

More importantly in terms of its effect on mass per mission, is that 3 tonnes of methane occupies a bit over 7m3 of tankage. 1.125 tonnes of hydrogen occupies closer to 16m3 of tankage. That presents problems both on the ground and more importantly in terms of how you land it. Any lander, mostly propulsive, or conventional suffers the issue that larger tankage translates to larger structural elements. Worse, the conventional lander is hit hardest because with the higher stresses, a large tank of hydrogen just multiplies the mass of structure and heat shield. This factor alone probably removes half of the real advantage of hydrogen.

There are also other system wide mass costs from importing hydrogen that go all the way back to supply of further power and refrigeration equipment right back to low Earth orbit. In the end the advantage of importing hydrogen over methane (even assuming conventional landers) may evaporate.

This is not to say that hydrogen isn't the future. It is. Its indigenous hydrogen we need. And for that we need machines that dig and process dirt. And that's quite probably an argument for having people there, to manage the prototypes. But I'm getting off track here.

------

Back to the big picture. A singular lander and ascent vehicle design is an exercise in reducing complexity and saving development costs. It also affords you a platform that serves many needs. For one thing, a lander becomes the taxi that transports crew between high and low Mars orbits, thus enabling the practice of parking the transit vehicles in high orbit. That factor alone saves the 2.6Km/s of delta V otherwise involved in bringing the transit vehicles down to low orbit and back. That would double the overall mass in and of itself.

Next I'm using the lander as the primary propulsion source for the return journey. Now the transit vehicles have their own engines but these are for backup purposes and its because we're relying upon the lander as the main engine we can then mass reduce the transit vehicle.

Next, the lander is capable of running surface to surface hopping missions. Meaning that its capable of long distance hops from base to base.

Next, the lander is capable of simple missions to the Phobos and Deimos. Indeed for a reasonable amount of fuel one could stop to claim samples from those moons on the way back to high Mars orbit.

Next, a mostly propulsive landing isn't as given to the vagaries of the atmosphere and has more capability to land on higher landing sites.

These synergies actually reduce the total mass of doing all of these things in other ways.

....

Inflatable heat shields are a great idea - if you plan on an expendable vehicle. I've no doubt that they, or something similar will come in very handy when it comes to landing large objects on Mars. But the problem with an inflatable heat shield is that it is still added, mass complexity and risk. And the bigger it is the more stability issues it has. Again, a necessary evil for large payloads, but lets make these large payloads unmanned.

The more you rely upon stuff that you eject during landing, the more you end up with the cost (and risk) of garbage.

Now, I would be willing to consider that in the case of a mostly propulsive, reusable crewed lander, a supersonic decelerator (probably a towed one) could be useful. But only to provide further margin. In other words you can land without it but it means you have more time to consider your landing site. Here you have two advantages. Firstly you're not dealing with extremely high dynamic loads and temperatures and you're not contemplating operating above Mach 3. Which means a lighter decelerator. Think in terms of being man handled and replaced easily. The other advantage is that because of the lower velocities involved you can afford to deploy the decelerator virtually from interface, or as soon as there is enough density to make it worthwhile.

I agree that by the time we get to land on Mars we will probably have real time, detailed observations of the atmosphere. Even a mostly propulsive landing would (especially if the landing site is higher) require some consideration as to the atmospheric conditions.

Which brings me to another point. One bug bear I have with direct entry is that you don't get much choice about the atmospherics. But if you based your architecture around having a safe (life support etc) haven in low Mars orbit, you're then free to choose when you land. Which means you can time it for the best atmospherics. And again, having an outpost in low Mars orbit does come at a cost, but it does have its advantages in terms of safety issues like this, plus as I said, you can keep most of your transit mass in a high orbit.

----

In some ways I'm not particularly fussed about L2 versus LEO. Certainly if you're in the business of large scale fabrication (over sized heat shields, trusses etc) then I'd definitely go for LEO. Remember that what I propose does not involve anything more than bump together assembly. No EVA required.

However the bigger issue is that you get the choice between either assembling (or rather docking) everything in LEO and then using conventional propulsion to Mars, or if you're going to adopt a solar tug then L2 is possibly the better option. Yes, you could just use a solar tug to lift the whole thing into a high elliptical Earth orbit but the whole thing would be traversing the radiation belts on every orbit. You'd then have to time the crew transfer well and get them on board quickly. Possible, yes.

L2 does give you more time to do check out systems (again that's in the context of using solar tugs).

I guess if you pushed me, I'd go for a conventional launch from LEO. At least at first. But remember that if you're seriously concerned about mass delivered to low Earth orbit, you wouldn't do that. You'd develop the solar tug first. Again, its useful to have a sense of perspective.

....

Now regarding the final point about reliability. For me there really are two enabling technologies that we need to go to Mars under any architecture that we don't have.

One is supersonic retro propulsion. A fact of life even for more conventional landers. And I need to emphasise this. Large payloads are never going to go subsonic in the Mars atmosphere. We need to understand SRP. Simple as that.

The other is enabling technology is developing a methane/lox engine that's very reliable. And that's the case even without a propulsive lander. As soon as you rely upon methane/lox engines to get you home, then you need an engine that can survive long burns, multiple restarts and long idle periods in space conditions.

It also has to be added that there's no such thing as a non propulsive lander. Sooner or later you're going to have to use retro propulsion. And you'd hope your engines are reliable.

What I've done though in having an architecture that relies upon a mostly propulsive landing for the crew is that I can achieve the goal of having one universal engine design for all purposes. And only having two basic vehicles. Which means a lot of the supporting hardware is of common design. Which means the capability to swap components.

(The above sentence ignores the big dumb H2/LOX booster which I think everyone takes for granted).

Then I've built in multiple layers of redundancy. Typically the lander itself would be flying with 8 engines. Four directly down. Four pointed about 15 degrees to the side. There's a certain amount of redundancy there that I won't go into right here. Where redundancy also comes in is that instead of landing the ascent fuel in a purpose built, expendable vehicle, I am instead landing the fuel in another lander. Which means that there's a source of spare parts and spare engines waiting on the surface. Indeed it might be possible in an emergency to retrofit the fuel lander as a crewed ascent vehicle.

The redundancy continues in space. Instead of discarding the lander, I'm bringing it back to orbit and then on to high orbit. There it provides the primary source of propulsion for the return journey. But if it fails there are still 2 engines on each of the transit vehicles. Again, its designed so that you can swap over engines if needed.

Your final fall back is to abandon one of the space habs, transfer fuel across and limp home (Ok, its cramped but you're safe). I don't see any other architecture providing that final layer of safety.

And technically there is one final refuge, and that's the lander itself.

----

I hope that answers things?

Btw, I was hoping to be asked about turbulence smile

Edit: Ok, now I've managed to get this post into some readable shape.

Edit 2: Since I brought up turbulence I might was well answer it.

This is somewhat more speculative. I'm simply making an argument that a good solution may be possible, at least for a vehicle without a major heat shield.

What little evidence I have about SRP is what can be found on the web. Some of it discusses the reduction in effective drag coefficient caused by an engine pushing the air away from the base of the vehicle. Some of it is the video showing clearly the turbulence that is caused by firing an engine into a supersonic air stream. Its fair to say that a lot of discussion about SRP is framed by the assumption that the engine or engines will be project through a conventional heat shield. Obviously I'm stepping outside that convention. And what I can't find is information that relates the nature of the turbulence to the density and relative velocity of the stream. Nor is there information about the nature of the turbulence. What the physical and time scales are and again, how those are governed by density and so on.

There are basically 3 regimes in which a mostly propulsive lander would have to work. The first is essentially outside the atmosphere. Here you have no constraint since there is no atmosphere to provide turbulence. In doing my rough math I've figured that this regime extends down at least to about 100Km. Now the reason for getting this point as low as possible is that the lower you do your initial large braking burn, the more efficient you are because of lower gravity losses.

The key question is, when does the atmosphere become dense enough for significant turbulence to apply. It might exist, but the forces might be too small to be of consequence. If I recall correctly, at around 100Km the air is still measured in micro grams per cubic metre. Its also quite possible that thermal considerations might kick in before turbulence from propulsion becomes an issue. In other words, it becomes necessary to apply (some of) the initial braking burn to keep heating below an acceptable level. The acceptable level I'm thinking of is around 0.3W/cm2. In other words something akin to being exposed to 3KW per m2. That's about 2 suns in Earth orbit. 

The third regime is below the point at which full retro propulsion is required in order to effect a safe landing. Which is probably somewhere below 15Km. The key here is that there is a trade off involved. Wait longer and normal atmospheric drag will continue to slow your descent and thus lower the effect of turbulence when you apply retro propulsion. Wait longer and you need to apply more thrust.

The second regime is obviously between those two extremes.

Now, one thing that can help is to move your engines away from the body of the vehicle. This has two potential benefits. One is that the region of lower pressure created by the engine can extend away from the body of the vehicle (especially with engines directed away from the vertical) and this should have the effect of reducing the effect of propulsion on normal drag. A lot of numbers I have seen give effective drag coefficients based on engine thrust, but do so on the assumption that the engine is directly below the vehicle and thus will create lower pressure under it. The other reason for moving the engines away from the vehicle is to remove as much of the turbulent wake of the engine as possible away from the body of the vehicle. Now the theory here is that the effects of turbulence are only important to the extent that changing air pressure can react against something.

So in essence you end up with a design that has engines mounted on a four star truss that projects beyond the body of the vehicle and also keeps the engine as high as possible. Quite probably integrated with the design of the landing legs. On each point you have two engines. One directed downwards. The other projected about 15 degrees off vertical and outwards.

Now, in regime 1, the key is to brake as late and as hard as possible. So you use all 8 engines at first, and then switch to the 4 engines that are pointed outwards and continue to complete the initial brake. My hope is that you can do this well below 100Km.

In regime 2, its yet to be determined what the optimum is. One suspects the optimal strategy is either a continued low thrust, or none at all (for the sake of engine efficiency at lower powers). An advantage of continued low thrust perhaps is to avoid re-ignition problems further down.

In regime 3 I suspect that you'd use the outer 4 engines at low thrust initially and then gradually build thrust as velocity falls. As velocity and thus turbulence becomes less of an issue you then bring in the 4 vertically projected engines and quickly go to full power. And of course as you approach final landing you'd throttle back and use the outer 4 engines.

Now all of that is based on the educated guess that turbulence can and will develop but so long as it has minimal surfaces to react against, its manageable.

And the nice thing about a mostly propulsive landing is that its easier to get away with doing that. Because your engines don't have to be hidden behind a heat shield they can be positioned where you like them and away from the body of the vehicle smile

#25 Re: Human missions » Yet another Mars architecture » 2013-06-18 22:04:49

JoshNH4H wrote:

So, in summary:  You propose that, rather than using aerobraking for atmospheric entry, we should instead do a breaking burn followed by powered descent; and that further, rather than importing Hydrogen for use in ISRU that we should import methane.  You also propose a reusable launch/descent vehicle for crew and (am I reading this correctly?) for cargo.  Rather than starting in LEO you instead suggest that the mission should start in L2.

Your justifications for this (which I intend to argue, once I can be sure that I'm arguing with what you actually said smile ) are of course contained in your introductory post.

Yes, I'm proposing a mostly propulsive landing. One aspect of which is a braking burn at high altitude. The key characteristic of this descent is making as much use as possible of aerobraking but within the capabilities of a modest TPS.

Yes, I'm proposing importing methane. Not going into all the issues with importing hydrogen, that decision was made out of sheer simplicity.

Yes, I'm proposing a reusable vehicle for both landing and ascent. With the caveat that in this instance its only used 3 times in an atmosphere. First is a crewed descent. Second is a crewed ascent. Third as an uncrewed descent fuel delivery.

Yes, I've proposed L2 as a staging point.

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