Landing on Mars

GW Johnson · 2012-12-02 14:23:57

In my crude model, if you enter at 1.6 degrees depressed below horizontal, you "terminate" at M3 and near 1.6 degrees below horizontal. That might not be "right" but it's close to true. Point is, with a shallow entry, you come out of hypersonics moving a lot closer to horizontal than vertical. You do not have to slow from 0.7 km/s to zero in only 5 km altitude-as-path-length. You have a whopping lot more path length available to slow down. I was using under 1 gee for my calculations.

GW

GW Johnson · 2012-12-02 17:11:33

Louis:

In my modeling, with entry from low Mars orbit, the delta-vee for a transfer ellipse from 200 km to the surface was only 50 m/s. That's attitude thruster system work. You don't need a big stage to do this.

If you did aero-braking of some kind from some more energetic orbit, I think what you want to end up with is "about circular" at "about 200 km". Justus & Braun claim the entry interface altitude for Mars is 135 km. So, you cannot orbit that low. I see some others using 300 km, but I do not see anything to prevent mission success when using "a few months" at 200 km circular. Pick an inclination. Cover a lot of Mars's surface. Do multiple landings in a single mission.

Elliptical orbits lead to steeper entry angles, which means lower altitudes at local Mach 3, for any given ballistic coefficient. That's bad, as near as I can tell. You do need sufficient time and path length to rocket-brake to something survivable for landing. There is a human limitation to how fast this can happen.

The model I have is for entry from low altitudes at low speeds, and it "models" as constant path angle in 2-D Cartesian coordinates. This sort of thing worked pretty well for warheads in the 1950's and early 1960's. It is not extrapolatable to other entry conditions. So, don't ask.

As for post-local-Mach-3 braking, I analyzed it as vacuum-trajectory stuff, which is a small over-estimate of required rocket impulse, since the drag numbers are so low compared to practical landing masses for manned (or big cargo) missions on Mars. Plus drag coefficient gets cut roughly by factor 2 when you use retro thrust during the supersonics.

All this stuff is posted over at "exrocketman": http://exrocketman.blogspot.com for a nominal 60 metric ton manned lander with a crew of 3. Or an unmanned cargo lander with a very large payload fraction-to-ignition weight. Same basic hardware works on the moon or Mercury. Nothing more than MMH-NTO storable chemical propellants. Not reusable, though. Real reusability comes from a variant of NERVA for the lander engine.

GW

Russel · 2012-12-03 02:30:09

Hmm..

'The analysis here is similar to that of reference 1. Entry requiring heat protection is assumed “done” at Mach 3, as before. This velocity is at the same 1.63 degrees entry from low Mars orbit as in reference 1, leading to a vertical velocity component of 19 m/s, and a horizontal velocity component of 675 m/s. '

I find that a little hard to intuit. I'm sure the model is fine, but I need some intuitive explanation of how you can come through a large gravity field pointing down, with little lift to drag, and still end up flying basically horizontally.. or am I reading this wrong?

Last edited by Russel (2012-12-03 02:49:09)

Russel · 2012-12-03 07:27:09

On another wild side track...

This is spurred by the thought that one of the basic numbers that affects the entire scale and mass of the "getting back off Mars" problem, is what's the least mass you can get away with for a simple capsule with half a dozen people in it? I've seen lots of figures out there but none that break it into irreducibles. Like, how much volume, the minimum thickness of metal etc..

And take that proposition even further, what about personal descent craft? How about ultra-minimal cocoons with a hypersonic, high temp drag tail.. and then a chute.. and finally the whole thing pops off and you're left with an astronaut with a small rocket pack on his back left to slow up the last few hundred m/s.

Completely mad?

My gut feeling is its workable. If you're going to throw stuff away it might as well be the barest minimum entry hardware. Of course, landing accuracy might be a problem so you'd need to have a reasonably fast recovery. Maybe you could add a teensy "segway".. hmm..

While I'm at it, one thing that keeps going through my head is a big glider. A few hundred Kgs of hardware with heat resistant fabric for wings. In some ways pretty conventional. Thing is you'd be going for the maximum lift/drag so serious heating (>300C) wouldn't be an issue.

At the end of the journey when you're down to ~200m/s you engage small rockets and land vertically. Too crazy? My point is it wouldn't be *that* hard to assemble these things en-route and for a 6 person crew you'd be talking in the vicinity of 2 tonnes of hardware.

Or even in Earth orbit assembly. And strap them onto the sides of your space hab. Now that would look truly interesting

Oh and you'd have enormous range

GW Johnson · 2012-12-03 10:42:06

Hi Russell:

I'm using a variant of the old early-1950's scale-height approximation warhead entry model, which is a 2-D Cartesian thing. I corrected the convective heat transfer part, the dynamics were OK as they were. Gees and velocity vs range-or-altitude look good. I do it as slant range plots. There is no radiation heat transfer model, so a rule of thumb might be to triple the calculated convective loads.

Real warheads follow a curved path during entry, it is true, but the Earth curves away beneath them, too. So, as it turns out, the flight path angle at entry and the flight path at M3 are pretty close to equal. That's why the stupid oversimplified model works so well anyway.

It's only good for what it was intended, entry from low orbit or fractional-orbit trajectories. But it sure is easy to do on a spreadsheet. One needs a better model for aerobraking and other sophisticated stuff like that.

GW

Russel · 2012-12-03 19:55:09

GW,

Having thought it through, I can roughly intuit how it works. It does break down as you lose speed, thus lift.

Which is what is making me think more carefully about weird and wonderful lander designs.

Here's another curly one.

Supposed you decoupled the need for blunt-object type drag from the actual crew space?

Concrete example. Supposed the crew traveled in something that's actually designed to have as little as possible drag, and thus heating - in other words more like the nose of a Concorde. You could even imagine a thermally detached leading nose that induces a shock cone.

Attached to this craft, but at a suitable distance are 2 or more other major structures which do contribute blunt-object drag. Think of them as retro rockets sitting behind a regular circular heat shield but you could imagine other forms. So the bulk of the heat goes nowhere near the crew.

You could refine this by adding elements that are steerable so you can control overall lift/drag.

Anyhow, before anyone's got an optimal mission they've got to stop cringing and start designing novel landers. That's my thinking.

Russel · 2012-12-03 19:56:41

Oh and btw, I hasten to add, I'm talking purely about crew descent/ascent. Nothing wrong in landing the bulk of the non-crew items in just a big ugly blunt body.

Russel · 2012-12-04 05:25:33

Hey. I think I've got a solution to all of this. And it involves thinking upside down.

Start with a conventional heat shield in the order of 4-7m diameter with the heat bearing surface facing downward.
On top of that heat shield place tankage for O2 and CO. Roughly 25 tonnes of propellant.
On top of the tankage place a cluster (2 to 3) of rocket engines, with the nozzles pointed up.
Extend a support leg, again pointed upwards.

So in normal atmospheric entry the heat shield protects the tanks and engines above it.

Take four of these, and place them into a square configuration with some distance between them.

Build a truss structure that binds the four into one overall vehicle.

Now, take a conventional crew capsule (some variant of the Dragon is fine). Place it with its heat shield pointed downwards (the conventional orientation).

Attach its docking port at the top, to the center of the X shaped truss formed above (just the center for now, we can refine this).

Now you've got a formation of four rocket engines and one crew capsule bound by a truss. Now you've got he combined drag of all 5 units.

Ok, for the fun part. After the majority of heating has passed, you perform a 180 degree flip using a small amount of asymmetric thrust.

Now your heat shield are pointed up (and your capsule is upside down relative to where you entered).

The next step is fairly obvious - powered descent, landing on the four landing legs that were originally oriented up.

This now solves a number of problems. Least of which is you can use very well understood technology and there's a minimum of in-space assembly.

There are a couple of minor problems to be solved, but are not hard problems.

One is creating lift to drag. The simplest and most robust is to mount the crew capsule off center. A more complex mechanism might allow you to gimball the crew capsule, giving you more control. Even further you can use the rockets themselves for fine control. One option (though not essential) is to allow the center portion of the truss to rotate so that the capsule can be 180 rotated. Of course if you don't do that its no big deal. There aren't normally huge G forces and if you're really keen you can just have rotating seats.

As you probably guessed this keeps your options open. Even though the platform has a lot of redundancy, you're still in a position to abandon it and simply use the capabilities of the crew capsule. Of course I could refine this further and take that out for saved weight, but I don't feel like that because this platform serves another purpose.

Here's the bigger picture. The platform itself (four engine pods and truss) has excess tankage. Without the crew capsule it has the ability to deliver in the order of 20 tonnes of oxygen.

With the crew capsule it has the ability to arrive in low Mars orbit with excess O2/CO fuel on board.

This excess fuel is there for a return to surface, but if that option isn't taken then the fuel provides an extra boost for the whole system (including space hab) into a higher orbit. I did some rough figures and that put it about 300-400m/s boost.

The platform itself has the option to do a empty (no crew) landing at this stage. Whereas the space hab fires up its own propulsion and goes home.

So in conclusion, I've taken my "fuel ferry" concept and refined it.

It only needs to make 2 trips into Mars orbit (one for fuel, one for crew) per mission.

Its now powered on CO. The space hab instead carries its own methane, but the bulk of the mass of return fuel (the oxygen) is sourced from Mars itself.

I returned the space hab into low Mars orbit because its neither here nor there when you add it all up. High mars orbit requires less oxygen refueling but also requires more CO burned to get it there. In the end I plumped for simplicity and crew safety which is provided by the low orbit.

You can refine this further if you do find that its less energy intensive to produce a CO/methane mix on Mars. Personally I'd go for simplicity and worry about methane on a later mission when we find a good source of hydrogen on Mars itself.

The nice thing about my mission design is that basically everything can be re-used. The lander/ascent platform (if not the capsule as well) remains at/around Mars. The space hab can be re-used. Etc.

Seriously would like some comment/refinement, or suggestions as to where to take this?

Last edited by Russel (2012-12-04 05:28:46)

GW Johnson · 2012-12-09 10:03:50

I like the idea of heat shield on one end, engines and landing legs on the other. This does presume that the vehicle (and its heat shield) are intended to fly more than one time. It is still considered (by most) too conceptually difficult to fire an engine through a hole in a heat shield, although I disagree. The problem with a hole in a heat shield is throughflow. Stop that throughflow, there is no difficulty.

The idea of heat shield on one end and engines on the other presumes you can successfully flip end-for-end in supersonic flight without risking a breakup. Here on earth, that's not really possible in a flightweight aircraft. But the air on Mars is thinner, and the dynamic pressures are lower, so maybe it's possible there. This will demand really strong attitude control thrust/mass levels, and it will demand a compact and rather dense structure, in order not to be too fragile. That's inherently higher ballistic coefficient.

Point masses connected by a truss likely will not meet that aerostructural requirement, due to the length-squared factoring effect that dimension has on bending moments and stresses. Besides, any time you connect things with a truss instead of direct contact, you just added the weight of the truss to the inert weight fraction. This is why reentry shells have tended to assume conical to short, squat conical shapes. The ones with the lowest ballistic coefficients actually approach the form factor of a frisbee flown broadside.

I tend to think that a vehicle form which might successfully swap ends in supersonic flight would be roughly spherical to a short cylinder of L/D near 1. The heat shield and pressure shell structures get to do double duty conferring the structural strength necessary to resist the broadside air loads. Plus, the lower moments of inertia make thruster control and a fast flip easier. There is also the preference for a low cg position in the vehicle as it lands, for stability on rough terrain.

All in all, you might get a lower inert weight fraction by solving the problem of rocket plume through a hole in the heat shield. The same compact shapes can be used that we already know work. All you have to do is stop the throughflow by sealing the compartment containing the engine. A static gas column is a better insulator than any ablative ever made.

But, I dunno for sure. Neither idea has been seriously pursued.

GW

Russel · 2012-12-16 07:13:15

I'm quite deliberately taking advantage of the lower dynamic pressure in this situation. And I'm not particularly worried about stability. Trust in the computer and all that. After all, if the spacex people think they can take a regular cigar tube rocket and flip it over, control it while it goes through reentry, then flip it over again and land on its engine, I'd say that pretty soon the problem will be well understood

One could imagine other form factors that would tick the same boxes but part of the original motivation is to fly something with low(er) mass to drag than would be possible given the commonly assumed maximum dimensions. Remember also that the interconnecting truss isn't entirely dead weight since it contributes to drag, along with its own light weight heat shielding.

The point of having 3 or 4 "pods" and a common truss is that you've got commonality of design. Each "pod" can be lofted separately and its a relatively simple assembly task.

Remember also that in its Mars-landing configuration you're flying with relatively little fuel - that is relative to the thrust available. That fact also contributes to the lower mass to drag. Its different of course if you have just taken off from Mars with a full fuel load. Depend on the abort situation you might find yourself having to deliberately burn off fuel at some point. Again, the capsule itself is detachable. And for the moment I feel more comfortable with the capsule as a separate, flyable entity. Maybe as the concept develops we could reduce mass and make the capsule part of the overall structure. If so you then don't actually need a conventional "capsule" shape for the crew compartment.

Indeed, I'm still working on that idea. I just need to be sure the whole thing has "abort from anywhere" capability.

There are of course alternatives to the upside-down configuration, and those include thrust-through-hole ideas. Some of those revolve (literally) around a gimballed engine/nozzle arrangement. So your "pods" rotate into the air stream. This has the advantage of being able to start off thrusting at an angle. And I'm sure you don't need 180 degrees to do this.

Remember also I'm trying to control the drag on the ascent too so that makes this more complex. Also I'm still giving some thought to being able to move the center of mass around in flight.

The one thing I'm most comfortable with though is the inherent modularity involved.

Oh and btw, even a four "pod" craft probably requires at least 3 engines per pod for the sake of engine-out capability. arranging those into a heat shield and forming a complete seal is a challenge in itself. You've also got the problem of which bits of your engine you want radiatively cooled. And that might leave you in a situation where the nozzle has to project. That and having more than one nozzle is the original motivation for the upside-down configuration.

And yes, speaking of thrust-through-shield I have considered the option of actually losing a little propellant (some liquid CO will do the trick) to both protect the engine and also to make life a little easier for the heat shield. But again you have to worry about cooling the engine in use.

Now going back to the overall mission design.

What I'd like to say is that unless someone can prove me terribly wrong, it seems to me that the whole problem of capturing propellant from Mars itself ignores that if you're using something like methane/LOX for the return journey, about three quarters of the mass you're trying to save in doing ISPP is actually oxygen.

Seems to me far easier to generate oxygen and ferry it than to go down the route of methane production on the surface. I'm using CO as a fuel for the ferry/landing/ascent vehicle which is "wasteful" but that doesn't matter that much because there's plenty of it to go around.

Here's the basic math. To make a tonne of propellant for the return journey (assuming methane/LOX) you need to make 777Kg of oxygen, and 222Kg of methane.

To make that methane you need 55.5Kg of hydrogen. (plus overhead for storage or boil off)

On a arguable typical return journey you need 40 tonnes of propellant and thus 2.2 tonnes of seed hydrogen. But, you must first land that hydrogen. However you do that you must pay for it in energy/weight. And allow for boil off. Conservatively you're getting closer to 4 tonnes of hydrogen/gear in Mars orbit. Perhaps more by the time you've landed the extra weight involved in the synthesis gear.

But instead, you could just send the methane along. The methane would have weighed 8.88 tonnes. So now you're talking about a modest saving, for a large increase in risk/complexity.

Now using CO as a propellant has its advantages/disadvantages too. But you'll probably find that the extra mass of propellant you need to manufacture at Mars (to get the ferry/ascent stage into orbit) is mostly balanced out by the extra weight/complexity/energy costs of managing hydrogen and methane synthesis.

So I'm actually quite happy with this approach. Now if someone comes up with some cheap way of getting hydrogen/ice off Mars, that changes everything.

There is actually one more wrinkle to this. And that's the option of using much less hydrogen and synthesising something that is mostly CO but with some methane - arguably worth the extra Isp but with the density benefits of a fuel that is mostly CO. I'll leave that to the guys who get paid for this

Point is, it wouldn't be a bad idea to design engines with the intent of at least running a mixture. Heck if you could design an engine that runs on any mixture from CO up to pure methane you're making things even simpler, but that isn't necessary or I suspect pragmatic.

Now, going a little further into, shall we say, half baked territory.

Lets suppose I develop the concept further with a crew capsule that integrates into the structure that bridges the "pods". Now you've saved more mass but you've also created something that lives its life out entirely between Mars surface and Mars orbit. It would also by its nature be fairly flat and have a low center of gravity (remember the size of the heat shields would be fairly large relative to the size/height of tankage needed). What I'm saying here is now the design leaves out a crew capsule as an element in the trans Mars and trans Earth journeys.

Now what do we have left? Well its pretty simple. Ignoring for a moment the Mars habitat and related gear, you've basically got a space hab that lives in space and has aerobraking capability.

So your mission looks like this. From Earth you ascend in a dragon capsule (no need to design anything new here). Dock with the space hab which is already in the correct orbit and head towards Mars. At Mars you perform aerobraking into a low orbit. From there you dock with the Mars ascent/descent vehicle, transfer crew and undock, then land. And of course the same in reverse. And back in Earth orbit you transfer to the dragon and return home. Its pretty simple and your space hab has the minimum of mass.

Bear this in mind though. This means that you've got no option but to aerobrake with the space hab. You can't climb into a capsule and take an alternate descent path. This may be a good or bad thing depending on your view of the risks of aerobraking versus direct. All I'm saying is if you refine the design of the ascent/descent vehicle to the point where the crew capsule is integral, that's what you get in return. And oh the other thing that motivated me towards having a crew capsule that is separate and detachable is that when this vehicle is acting as a fuel ferry you need it as light as possible.

One other idea I'll toss in here. The space hab does not need a large engine. But to make the most of the fuel one trick you can use is that you start with a relatively high orbit and then only thrust as you approach periapsis. This of course means spending a few days acquiring the velocity needed. Same thing for Mars.

Another related trick is that while the vehicle is unmanned, and you have spare solar power, you can also employ a small ion engine and gradually raise the orbit in advance of the crew.

So, lots to mull over there

JoshNH4H · 2012-12-16 13:19:22

I have always favored the use of inflatable heat shields-- Not just on Mars, but on Earth as well. In my opinion it is simply the logical way to go about increasing your surface area without using too much mass or requiring too large a payload shroud when you launch from Earth.

That said, the consensus here seems to be that even an Apollo-like heat shield will lead to pretty acceptable velocity decreases from LMO.

I would suggest that a delta wing-lifting body entry would give both more control (thus a smaller landing ellipse) and better entry characteristics. I don't know what temperatures are experienced during entry, especially on Mars. However, it would seem to me that a coating of either TiO2 or a reflective metal on the heat shield would significantly reduce thermal stress by reducing radiative heat transfer. Also, simply as a matter of curiosity I wonder about "inflating" the gas surrounding the heat shield.

Blunt heat shields work because they push the hottest gas away from being in contact with the material. By spraying small amounts of water into the cooler gas in contact with the heat shield you both cool that down and push the region of highest temperature farther from you. For Mars entry, I would bet that basalt/carbon fiber cloth, perhaps composited with silica aerogel (the basalt/carbon fiber provides compressive strength while the aerogel reduces heat conduction, of course with a TiO2 coating, would be easily sufficient for a heat shield. Just to be clear, I'm thinking either basalt OR carbon fiber, though if there are reasons to mix the two then of course it should be done.

These ideas are not necessarily related out compatible but strike me as being useful techniques to improve the ease of entry.

Given M3/700 m/s at 15 km... What about dropping an anchor?

Russel · 2012-12-16 17:13:10

Josh,

Clearly some things (like a Mars hab) are one way trip items. But others like the Mars ascent vehicle should be highly re-usable. So where I'm being led is the thought that if you're going to get satisfaction out of an inflatable heat shield its going to be in landing (and aerobraking) very large and otherwise one-way-trip things like the Mars habitat.

So the picture I keep seeing is a trailed inflatable simply there to buy you more altitude by the time you can go into powered descent.

As far as smaller, reusable craft is concerned. I think it would be a huge advantage to have a versatile, reusable craft that can do things like ascend into orbit, or travel large distances.

I keep thinking of going back to old fashioned glider shaped wings. Which could lead to a quite laid back, if lengthy descent - with temperatures down to a few hundred degrees. The trick of course is just enough thrust in the right places to get you vertical take off and landing.

Having a delta wing, especially for something potentially large enough to return to orbit around Mars, basically walks you back into the territory of having to do something special to get it off Earth. Mind you if the shuttle can do this...

JoshNH4H · 2012-12-18 17:32:57

Russel-

With regards to a Mars ascent vehicle (MAV), many mission architectures don't call for one at all. Mars Direct is the most obvious of these; It calls for an Earth Return vehicle (ERV) that launches from the surface of Mars.

Further, just because a heatshield is inflatable does not mean that it is not reusable. In fact, there is really no reason why an inflatable heatshield that was intended to be reusable would be any different from a solid one, so long as it was deflatable, which by nature anything inflatable is.

Having said that, it is considered a major plus to the safety of any mission that there are two vehicles to get back to Earth if it is found that the first one has failed. Mars Direct plans to have the second mission's ERV fully fueled and ready by the time the first mission would choose to leave the planet. Reusing a MAV eliminates this safety factor and also makes each mission less safe than the last. There is no orbital rocket on Earth that can be reused in the way you're suggesting; Though a Martian orbital launch vehicle will have an easier time of it than a Terran one, I don't think that reuse of an MAV is a safe choice. Please remember that the use of an MAV as opposed to an ERV mostly eliminates the potential benefits to be gleaned from producing the propellant on Mars.

If we're talking about standard, solid heatshields and launch vehicles with standard fairings, delta-wings would presumably be a problem. However, an inflatable delta wing heatshield is still very possible.

Russel · 2012-12-19 09:00:12

JoshNH4H wrote:

Russel-
With regards to a Mars ascent vehicle (MAV), many mission architectures don't call for one at all. Mars Direct is the most obvious of these; It calls for an Earth Return vehicle (ERV) that launches from the surface of Mars.
Further, just because a heatshield is inflatable does not mean that it is not reusable. In fact, there is really no reason why an inflatable heatshield that was intended to be reusable would be any different from a solid one, so long as it was deflatable, which by nature anything inflatable is.
Having said that, it is considered a major plus to the safety of any mission that there are two vehicles to get back to Earth if it is found that the first one has failed. Mars Direct plans to have the second mission's ERV fully fueled and ready by the time the first mission would choose to leave the planet. Reusing a MAV eliminates this safety factor and also makes each mission less safe than the last. There is no orbital rocket on Earth that can be reused in the way you're suggesting; Though a Martian orbital launch vehicle will have an easier time of it than a Terran one, I don't think that reuse of an MAV is a safe choice. Please remember that the use of an MAV as opposed to an ERV mostly eliminates the potential benefits to be gleaned from producing the propellant on Mars.
If we're talking about standard, solid heatshields and launch vehicles with standard fairings, delta-wings would presumably be a problem. However, an inflatable delta wing heatshield is still very possible.

I've never liked the idea of a direct return purely because the mass of whatever it is that keeps your crew alive - the "space hab" function is a mass you now have to land on Mars, and then get off Mars again. Doesn't make sense to me. From what I've seen of mars direct, this also led to a design where they had to cut corners on the return vehicle. A bit ugly really.

So far as I can tell, there's only really two classes that fits every possibility. Either a) you take off directly from Mars so you go "direct", or some or all of you return vehicle is in Mars orbit. Correct me if I've missed something.

I'm focused on getting the crew down onto Mars in a manner that is as safe as possible and has as much margin/redundancy as possible. That goal before all else.

Linked to that is the concept that if you're designing a mars ascent vehicle, why not also design technology that has a dual purpose. One of those is to get your crew into some kind of orbit. The other is to get fuel to your return vehicle. Amongst other factors, what you find is that the level of thrust you need for Mars ascent is far greater than the level of thrust you need for your return vehicle if its going to start from orbit. Therefore a lot of the mass involved in the ascent vehicle would be dead weight if it returned to Earth.

This leads me to a strong feeling that this element of the overall mission, if not the space-hab itself should be fully reusable.

It also happens that if you develop the engines needed for a Mars ascent vehicle for crew use, you've already done the work you need to build a reliable fuel ferry too.

I'd dearly love for someone to show me an inflatable heat shield that is re-inflatable - even multiple times. And one that can deal with the obvious practicalities. Either de-inflating it during descent, or else having to deal with it landing on the martian dirt and all the practical headaches that come with that.

Besides, I've also got the constraint that it must be fully automatic, because if I want to use inflatable structures they have to work on a fully robotic craft. Which is why I've toyed with fixed, but lightweight structures, including mesh.. even thinking about stuff that folds/unfolds.

Again, if someone can show me an inflatable heat shield that's reusable a number of times and avoids the obvious pitfalls I'd really love to see it. The furthest I've gotten down that path is a somewhat rigid drag-inducing structure above the craft, that also can then use inflatable technology. Just for the sake of a concrete example, picture a ring frame that attaches to the top of the capsule, the diameter of which is much larger than the slipstream of the capsule, and from that ring you inflate separate "balloons". The purpose of this whole exercise is to take a ring that has a diameter of about 8m and expand on it so that it has an effective diameter of say 14m - with a hole in the middle through which the superheated slipstream passes. Point is that if you can get it to deinflate properly at least it won't touch the ground. Anyhow that's the idea, but making it stow neatly and making it work again and again, is a toughy..

Stepping back to the issue of safety.

What I'm thinking is actually more redundant than you'd first guess. Firstly I'm parking a fully refueled space-hab return vehicle in Mars orbit. So when the crew arrives, they're actually parking a spare. Secondly, the ascent vehicle isn't one of a kind. There are at least two. Third the crew have the opportunity to both inspect and test the ascent vehicle. Fourth, the way I'd have it there would be full abort capability all the way to orbit.

As always when you consider the sum of all risks there are inevitable trade offs. I personally don't like direct entry to Mars. I don't like missions where the safety of the crew is reduced byt eh need to land the crew along with a large mass - I think they should always come first and have their own descent vehicle. And I don't like missions that suffer from redundant mass in places which means ultimately less food, less oxygen and less shielding.

I'm not sure what you mean by "the use of an MAV as opposed to an ERV mostly eliminates the potential benefits to be gleaned from producing the propellant on Mars."

When you do the math, both a direct ERV and an indirect MAV benefit from mars produced propellant. The major difference is that whatever you park in orbit that will fly home with you is mass you don't have to land and mass you don't have to get back into orbit. That's a substantial fuel saving all else being equal.

As for an inflatable delta wing - or inflatable aerfoils of some description. Not an entirely silly idea. If you can draw me a picture of the basic engineering that would help. As it stands I'm more comfortable with inflatable elements that trail - for stability reasons. You know, one suggestion put to me is simply a swing wing but the mechanism is simply driven by a simple gas cylinder. That could be fast enough. Even crazier ideas involve being able to use thrust into the airstream from engines hidden by a heatshield that are able to be retracted by similar means.

GW Johnson · 2012-12-19 14:03:07

Myself, I'm leery of relying on inflatables ahead or behind a vehicle during entry. That is even more true here than Mars (the speeds are higher). Such things have flown before, but only very subsonically. Back in the late 50's, Goodyear even built something for the military that it called Inflate-a-Plane. It really flew, but only under 80 mph.

Ballutes have been tried supersonically (Mach 2-ish), with mixed results. They work better with power inflation than ram air inflation. But to my knowledge no inflatable has ever been tried hypersonically, where aeroheating gets somewhere between very serious and utterly extreme.

You have to worry about both the ballute and its towline. Especially the towline is subject to extreme heating where the vehicle's "bottle shock" closes upon the towline between it and the ballute. Shock impingement heating is so extreme a regime that no one in the business considers it a survivable regime under any circumstances. That's what nearly cut the tail off the X-15 on the scramjet pod flight that hit Mach 6.7.

Further, the trailing ballute is immersed in the vehicle's wake, where it sees less dynamic pressure BY FAR than the vehicle, and it sees incandescent wake gases already heated by passage of the vehicle. There are serious questions about how much decelerating drag you can get in such a trailing ballute approach in this hypersonic regime. Those experiments have never been done.

As for hypersonic aeroheating in general, it's less about the specific materials you choose, and more about controlling the rate of heat absorbed, and the size of the sink it can go into. Example: exposed Inconel-X skins on the X-15 worked just fine at Mach 6-ish, but only because the vehicle interior was a really large cool heat sink, even after cryo propellant exhaustion. Slow the conduction rate with a low conductivity material layer, and you can use a small heat sink. That's what they did on the shuttle.

Entry heat protection is always (and has always been) a truly transient phenomena. There are no steady-state solutions, no "magic" materials, and very likely never will be, not in the lifetime of anybody alive today, for sure.

With capsule-like manned landers and heavy cargo landers at Mars (no inflatables), you will always have a high ballistic coefficient: somewhere in the several hundred to a few thousand kg/sq.m range. If you restrict the entry interface speed to that from LMO (not direct interplanetary transfer), AND you restrict the entry angle to values under about 2 degrees (again, not generally possible from direct transfer, but inherent from LMO), then there is enough drag deceleration to slow you to local Mach 3 (about 6-700 m/s) at altitudes between 5 and 10 km from the surface in the "average" Mars atmosphere.

That's way too low to deploy any chute or ballute and get it open before you hit, much less any time for it to actually slow you down any at all. But it's high enough for direct rocket braking to set you down and never exceed decelerations larger than about one-something gees, under two for sure. That's quite practical, and all it requires is supersonic retro thrust. I've posted before about that technology, and how there is very good reason to believe it will work just fine and quite easily the very first time we attempt it.

Vehicles like this with storable or mild cryogenic propellants of "decent" density are well with technological reach without developing any "new" technologies. All the tinkertoys already exist. You could probably even do it with hydrogen, although voluminous tanks would be a serious packing problem for aerodynamic flight, just as they always have been. Build these vehicles tougher for multiple flights, and refuel them on the surface, and you have your reusable Mars lander shuttle.

Although, in a reusable vehicle, inert weights are simply going to be higher, because it literally takes more material to be stronger and tougher. Example: the most reusable rocket vehicle in all of history was the X-15, each of which flew more times with less maintenance and repair than any of the space shuttles. It's inert weight fraction was right at 40%, the shuttle's was lower. It had 1 fatal crash in 199 EXPERIMENTAL flights. The shuttle killed 2 crews in about 130 flights that were supposedly "routine" after the first 4, but turned out not to be.

I think the best of these reusable capsule-shaped Mars lander shuttles refuelable on the surface of Mars might be powered by a resurrected NERVA. Isp near 900 sec at engine T/W 3.6, just as we built them last in the early 70's. Put 4 small engines canted about 10 degrees firing through simple ports in the heat shield. Seal off that engine room space to stop the throughflow, and you can leave those ports wide open all through entry. Simple as that.

What I like about the reusable Mars shuttle capsule powered by NERVA is that it has the performance to work acceptably well as a reusable shuttle, even if you do not refuel it on the surface. That kind of capability held "in reserve" on the first mission is the "suspenders-and-belt" that can get them home alive without terminating the basic mission, even if the surface refueling fails to work. We've never actually produced any in-situ propellants, you know?

What I have just outlined (to me) suggests that the very first mission ought to stage from LMO, and it ought to carry enough lander propellant to carry out some of the mission, even if surface refueling fails. They ought to use NERVA-powered reusable landers to make that happen. We really ought to use NERVA for the Earth to Mars transfer, too. That's what it was developed for. If you do, the same NERVA engines that power the landers can push lander propellant to Mars. It also has the "oomph" to go visit Phobos and maybe even Deimos from LMO.

GW

JoshNH4H · 2012-12-20 01:47:32

GW-

Myself, I'm leery of relying on inflatables ahead or behind a vehicle during entry. That is even more true here than Mars (the speeds are higher). Such things have flown before, but only very subsonically. Back in the late 50's, Goodyear even built something for the military that it called Inflate-a-Plane. It really flew, but only under 80 mph.
Ballutes have been tried supersonically (Mach 2-ish), with mixed results. They work better with power inflation than ram air inflation. But to my knowledge no inflatable has ever been tried hypersonically, where aeroheating gets somewhere between very serious and utterly extreme.

Because NASA and other space agencies are very much aware of the shroud diameter issue, there is actually a lot of interest in inflatable heat shields. NASA recently experimented with an inflatable heat shield. It expanded from a shroud diameter of .56 m to 10 m. It entered the atmosphere at Mach 10 and successfully landed in the ocean after a test that seems to have went as well as the scientists involved could have hoped. More information here.

I would like to point out that Mach 10 is ever so slightly less than Martian orbital velocity. Based on this page describing the lifting abilities of the sounding rocket on which it was launched, the payload was probably somewhere around 350 kg.

Also sometime in the early 60s there were designs for an "escape pod" inflatable reentry capsule, called the FIRST Glider. None were ever built or tested, but they seem to have believed that it was practical. Transonic velocities would have been reached at 43 km, and maximum deceleration would have been 2 Gs.

Entry heat protection is always (and has always been) a truly transient phenomena. There are no steady-state solutions, no "magic" materials, and very likely never will be, not in the lifetime of anybody alive today, for sure.

Given family history, I probably have around 70 years in me if I eat right, exercise, etc. That's a bet I'd be willing to take, given the rate at which materials science, especially with Carbon, is advancing. In combination with heat radiators elsewhere on the craft some craft capable of withstanding pretty high aeroentry velocities should be possible, though obviously that is technology that is far from what we have now.

Speaking more of present and near future technology, I've heard that larger shield areas for the same mass reduces heating because more velocity is shed higher in the atmosphere. This is something that inflatables can be helpful with.

Given traditional solid state heatshields and high ballistic coefficients, a powered landing seems pretty necessary. However, in the circumstance of a non-NERVA transfer, I'm not sure how much sense it makes to use NERVA for simply leaving the Martian surface, or for the powered phases of EDL (Entry, Descent, Landing).

While propellant has never been manufactured on other planets, Zubrin makes a good point when he says that it is a very testable technology: We know the composition, pressure, and temperature of the Martian atmosphere, and that can be replicated well on Earth, where the technology can be tested and redesigned until the engineers are satisfied that it will work.

While nuclear rockets are a boon for a Mars mission, they are not a requirement. I love the NERVA technology, but the use of chemical rockets just represents additional launch mass. This isn't a mission killer. Depending on who is launching the mission, I think it's important to allow for the possibility of both nuclear and non-nuclear propulsion options.

Russel-

Responding to your post now. You raise valid points, but I am not entirely in agreement.

Russel · 2012-12-20 04:32:25

Josh,

I've been following IRVE for a while and I'm glad they're getting somewhere. And I'd really, really love to combine inflatable with fully reusable and robotic.

GW is quite right about being able to land large payloads from low mars orbit and go directly to a powered landing. There's still some second order problems - getting the lift to drag right, and the packaging of the rocket nozzles and heat shield(s).

The open question for me though is whether some inflatable elements are worth the extra weight - they might be.

To me there are 2 places inflatables might be a good idea.

One is when you want to land something on Mars that is going to just stay there - one way trip items.

The other is supplemental drag for aerobraking. In this case the inflatable part would be consumable and you'd be in a position to replace it when you return to Earth orbit. Nice thing about inflatables for aerobraking is that again, you capture more drag higher up so a Km or so error isn't going to make as big a difference as it would further down. Also with inflatables you get to "trim" your aerobraking.

Going back to solid heat shields, I don't see why you can't capture the benefit of lower ballistic coefficient and thus lower heating through having a lighter and thinner heatshield. As I understand it, provided you keep below certain temperatures there is effectively no ablation at all. This is one of the underlying assumptions in building something big (surface area wise), even if it has to be assembled. Its to do what GW is suggesting and that's come in as slow and shallow as possible and avoid extremes of temperature even if the extra mass of superstructure seems at first glance non-optimal.

GW,

I'm aware of the impinging shock behind the craft and that's why I've moved away from a center tow line or tow structure towards an arrangement where the superstructure goes through the stream sideways at several points, before the stream gets to converge. Yes, I'm aware that that that's possibly the weakest point. However the more drag you can create the cooler (in relative terms) is the stream of gas you're talking about.

As far as having a trailing ballute, or a trailing anything for that matter, because of what you speak of, what you need basically is a donut.

Leaving aside nuclear solutions, if you simply built something that was a classic capsule shape with thrusters firing through 10 degree ports, how many would you have to position in order to provide enough redundancy? My feeling is somewhere between 8 and 12 in order to retain navigability with 2 or possibly 3 engine failures.

Here's another thing to keep in mind. Something that is designed to carry upwards of 20 tonnes (or more) of propellant cargo into orbit has a lot of margin left over when its only transporting 5 tonnes or so of crew and capsule.

As far as fuel goes, I'm quite comfortable with just plain O2 and CO. Both are reasonably dense.

As far as building something robust. Its not the frame I'm worried about, its the reusability of the rocket engines themselves given long periods of inactivity that bothers me. Again, this is what engineering is for. Besides you don't cut corners when you don't need to. You just burn more fuel. And since CO is reasonably "cheap" to procure that's not an issue.

In fact, would I be game enough to send a nuclear power source to Mars, I wouldn't use it as a primary power plant for a rocket. Rather I'd rely upon having lots of available energy to manufacture lots of spare propellant, making lots of problems easier - including having a more robust vehicle. And now we've got so much spare fuel we can afford to use it for longer distance trips around the planet.

Remember also that when you're thinking about inert mass for a mars landing/ascent/ferry vehicle one thing that counts in your favour is that if you have one, you then have a much lighter space hab.

Oh and btw. One thing that I have up my sleeve is this. If you build a Mars ascent vehicle with the tankage to act as a fuel ferry and its less than half loaded with crew on board, you could if you really wanted to use it as a boost vehicle, getting the space hab out of low mars orbit and up into a higher orbit. All of that and still being able to re-land itself after.

GW Johnson · 2012-12-20 13:13:44

I'd heard about IRVE, too. That will lead somewhere, but it's a long way from a usable technology. Give it time.

Reducing heating at low ballistic coefficient by decelerating higher up is a known phenomenon. To do it with a spaceplane requires a wing loading not unlike a Piper Cub. That's about 8 pounds per sq. ft, or about 40 kg/sq.m. A capsule shape would have to have a ballistic coefficient in that same range. That's an awful long way away from where we are right now in construction techniques. However, that's a place the inflatables could take us, once developed and "wrung-out".

Interestingly enough, it just might be possible to build a spaceplane with a wing loading that low, by building it the same way a Piper Cub is built: steel tube truss frame, covered by ceramic fiber (instead of linen) fabric. There are a host of problems to solve, such as low-conductivity stand-offs between the fabric and the tubing, and a gaseous atmosphere inside to be at least part of the heat sink. Yet the fabric can easily survive reusably at 2300 F ( 1260 C), and doesn't melt in one-shot use (it does embrittle, though) at 3200 F (1760 C). This is the same alumino-silicate fire curtain cloth they have long used in modern aircraft engine nacelles. I used it myself as the reinforcement in a very low-density ceramic composite liner for a ramjet, about 3 decades ago.

The nose cap, windscreen, and leading edges might need a more sophisticated treatment, who knows? But it's an intriguing idea, and most of the tinkertoys are already in place to try it. It'll take someone really innovative like an XCOR or a Spacex to try it, because it's so counterintuitive, otherwise.

Speaking of XCOR, take a good close look at the engines they offer. Some of these have the same safety and lifetime characteristics as ordinary certified aircraft engines. We're talking thousands of burns and thousands of hours of burn, here. With minimal repair/refurbishment. That's precisely the kind of engine you need for any reusable spacecraft. I think you might be surprised to learn that the best of these, already used in real aircraft, do not use turbopumps as the propellant pumps. Another really good tinkertoy available, that.

GW

Russel · 2012-12-21 07:54:07

GW Johnson wrote:

I'd heard about IRVE, too. That will lead somewhere, but it's a long way from a usable technology. Give it time.
Reducing heating at low ballistic coefficient by decelerating higher up is a known phenomenon. To do it with a spaceplane requires a wing loading not unlike a Piper Cub. That's about 8 pounds per sq. ft, or about 40 kg/sq.m. A capsule shape would have to have a ballistic coefficient in that same range. That's an awful long way away from where we are right now in construction techniques. However, that's a place the inflatables could take us, once developed and "wrung-out".
Interestingly enough, it just might be possible to build a spaceplane with a wing loading that low, by building it the same way a Piper Cub is built: steel tube truss frame, covered by ceramic fiber (instead of linen) fabric. There are a host of problems to solve, such as low-conductivity stand-offs between the fabric and the tubing, and a gaseous atmosphere inside to be at least part of the heat sink. Yet the fabric can easily survive reusably at 2300 F ( 1260 C), and doesn't melt in one-shot use (it does embrittle, though) at 3200 F (1760 C). This is the same alumino-silicate fire curtain cloth they have long used in modern aircraft engine nacelles. I used it myself as the reinforcement in a very low-density ceramic composite liner for a ramjet, about 3 decades ago.
The nose cap, windscreen, and leading edges might need a more sophisticated treatment, who knows? But it's an intriguing idea, and most of the tinkertoys are already in place to try it. It'll take someone really innovative like an XCOR or a Spacex to try it, because it's so counterintuitive, otherwise.
Speaking of XCOR, take a good close look at the engines they offer. Some of these have the same safety and lifetime characteristics as ordinary certified aircraft engines. We're talking thousands of burns and thousands of hours of burn, here. With minimal repair/refurbishment. That's precisely the kind of engine you need for any reusable spacecraft. I think you might be surprised to learn that the best of these, already used in real aircraft, do not use turbopumps as the propellant pumps. Another really good tinkertoy available, that.
GW

You may have the freedom to design a low enough ballistic coefficient (and more importantly a high enough lift to drag) that the temperatures can be such that you don't need to protect the frame of your wing. Instead it quickly assumes the same temperature as the skin.

The point I'm making is that if your "wing" (which can be all kinds of shapes so long as its light and has the required lift) can be separate from the main part of the structure which then contains the things that actually need more thermal protection. Indeed the structural element that joins the wing to the main body itself provides a long enough thermal path that the rest is easy.

I've yet to take these notions and come up with something elegant. At least something elegant, and stable.

As far as strength goes we can take advantage of two things. First, modern aircraft wings are strong precisely because they can flex. The other is that the forces you're dealing with are still smaller relative to a landing on Earth.

Is there a link to those engines? As I understand it, without turbopumps you pay a penalty in terms of less Isp or more fuel. One other thought. Thinking back about cooling and nozzles that are behind a heat shield, are there any good known engine designs that also use the propellant to cool the nozzle?

Oh and I'd be more excited about taking the piper cub idea literally if I could figure out a way to do vertical take off and landing with it and manage to package all the extra tankage you need to turn it into an ascent vehicle too.

GW Johnson · 2012-12-21 10:13:50

Hi Russell:

Go visit XCOR's website and navigate down to see the products they offer. Their baseline product has been rocket engines for others while they develop their Lynx suborbital space plane. Their engines have every bit as good an Isp as anybody else's. Engine T/W is a little less, reflecting the extra material it takes to be robust and reliable over a long lifetime. They use a heat engine-driven piston pump on some models instead of a turbopump, which is driven off the waste heat in the regenerative cooling stream. Those might be a bit heavier, but I doubt it. I don't know the weight data either way.

Not all strong wings are flexible, particularly at high loadings, like high AOA or even broadside attitudes. The last airliners to have stiff wings were the Lockheed Electra-II models now serving as USN P-3's. The only real difference is ride quality: flexible is smoother in turbulence. Given that this was an early 1950's design still serving in the 2000's, those are extraordinarily strong, tough, old birds. Stiff wings may well be required at reentry conditions. The shuttle designers thought so. So did the X-37 designers.

As for a practical space plane, you just need a powered reentry vehicle small enough to shoot up there on an existing rocket. Like X-37 and a whole lot of other ideas dating back to about 1960. They even thought about shooting the X-15 into orbit before they figured out it could not survive reentry.

The difference with the "piper cub" idea is going for very low wing loading on a real winged vehicle. The structure has to be very strong to take the high-AOA to dead-broadside air loads at very high dynamic pressure (could be a few thousand pounds/sq.ft). That's a really strong wing indeed.

For all known materials, that means the structure has to stay cool. Most steels are losing it by 800-1000 F, even the high-alloy / super alloy steels are losing it well before 2000 F. All the aluminums are "butter" by 300-400 F. Titanium is similar to steel, except lighter. Unfortunately the useful alloys are not formable, and the brittleness of castings isn't really useful in main airframe structures, generally. Especially for spars and ribs - those need real toughness, and cast materials just don't have that.

The ceramic fiber cloth covering over a truss structure might really be a way to build this low wing-loading thing. Might need some carbon-carbon nose cap and leading edge pieces like shuttle, although those are pretty heavy. It does avoid the vulnerable fragile low density ceramic tiles shuttle used in favor of something with some toughness and flexibility.

I'm not at all sure how to provide the wetted-surface heat sink medium, since these fabrics are not gas tight. Maybe a limited-porosity bleed-through of sacrificial coolant (like steam) would work. An idea like this was proposed as one of the three experimental heat protection schemes they were going to try out experimentally on the old X-20 Dyna-Soar, also vintage 1960-something.

Any ceramic skin technique like that is going to need a low-conductivity stand-off between it and the supporting airframe structures inside. Otherwise the steel heats up close to the potential-2000 F skin temperatures, and it then loses way too much strength. Go see the data in Mil Handbook 5 to see the truth of what I am saying. Those standoffs cannot be ablatives, or you will be re-covering the craft after every flight.

A low density ceramic composite with reinforcing fibers in it might be the answer to the standoff problem. The stuff I made in 1984 was almost as low a density as NASA's fragile tile, accordingly it still held a measured 14,000 F/inch gradient in a steady-state test for me. It also withstood hours and hours of accumulated burn at flame temperatures up to 4000 F, corresponding to surface temperatures right at 3200 F (I did see a small bit of surface melting here and there). The real difference with NASA's tile was that it endured the pressure violence of rich blow-out instability in the combustor it protected, all the way up to that 3200 F surface temperature. All of this was far more demanding an environment than what NASA's tiles endured.

A lot of this ceramic composite and fabric-on-truss in reentry stuff is not yet "tinkertoys in-hand", ready for general use. But it could be made so fairly quickly, based on what I accomplished with these materials experimentally so long ago. Convincing skeptical minds will be harder than doing the actual design and testing.

GW

JoshNH4H · 2012-12-21 21:11:00

Russel-

First, I would like to address what seems to be a major conceptual error in your post with regards to logistics: I know of no mission plan that calls for landing the rocket stage used to get from LEO to Mars on planet or reusing it. In fact, as often as not mission plans call for the upper stage of a HLLV (High Lift Launch Vehicle) to be used to send cargo directly to the Red Planet, depending on the size of the rocket used. This is probably not feasible if the Falcon Heavy or an equivalent is being used as your launch vehicle because payload to Mars would be too small. The "Ares" rocket that Zubrin uses as a reference launcher in Mars Direct puts 121 tonnes to LEO and 47 towards Mars.

Having said that:

Linked to that is the concept that if you're designing a mars ascent vehicle, why not also design technology that has a dual purpose. One of those is to get your crew into some kind of orbit. The other is to get fuel to your return vehicle. Amongst other factors, what you find is that the level of thrust you need for Mars ascent is far greater than the level of thrust you need for your return vehicle if its going to start from orbit. Therefore a lot of the mass involved in the ascent vehicle would be dead weight if it returned to Earth.

So, the basic idea here is to use the MAV (Mars Ascent Vehicle) as a fuel ferry from the fuel plant on the surface to the return craft, which is left in orbit? I'm skeptical about the safety of this proposal. My reasoning here is as follows:

The MAV only needs to have the payload capacity to lift (this seems to be a pretty standard crew size) 6 people. Falcon 9 does this with a payload of 10,450 kg, but it actually has a significant payload capacity, even in passenger configuration. Assuming that, for safety reasons, a full re-entry capability exists on the crew section of the MAV I'll say that payload mass to LEO is the same 10.5 tonnes. From Astronautix, the habitat parts of the ERV (Earth Return Vehicle) would mass 20 tonnes. Zubrin has been criticized for lowballing here, but for the sake of argument I'll use his numbers. Given methane fuel, the mass ratio from LMO (Low Mars Orbit) to Earth assuming aerocapture is 2.0. Assuming another 5 tonnes for fuel and rocket structure as well as on-orbit refueling, that will require 25 tonnes of fuel to be orbited from Mars. That's another three launches of your MAV, done with no refurbishment and almost no human involvement. You have to be able to land it within a couple kilometers of your landing site (where the nuclear reactor and the fuel is going to be) each time. Presumably after that, you have to mount your capsule on top (how?) and then launch into LMO. That's four launches and four mars entries (per mission?) of a mission component that is absolutely vital. Failure in any of these stages would probably result in loss of crew.

To reiterate, an ERV would be a separate vehicle from the vehicle used to transport cargo to Mars.

While using an MAV to launch fuel from the surface to orbit does not make sense, the general idea is a mass saver and is something that I think makes a lot of sense for the mission. I would recommend launching fuel and crew on one rocket, then docking with an orbiting ERV. If zero gravity fuel transfer has not been developed, the two rockets could be tethered to each other and spun in order to generate a centrifugal force that can be used as gravity.

On the other hand, considering that the crew launch capsule will weigh half (or possibly closer to a third) of what the hab will, I'm not sure this represents an increase in safety. It may be equivalent to Mars Direct in safety with higher mass, which is not a good tradeoff.

I accept that an inflatable heat shield is probably not reusable without repacking. By the very same logic, I hold that it is very difficult to design a rocket that is reusable in the way you suggest with acceptable safety margins.

With regards to inflatables, first of all I think you need to take a look at the wikipedia article on the pendulum rocket fallacy. Basically, because gravity acts equally on all parts of a rocket (or a capsule and ballute) a trailing ballute is not more stable than the kinds of inflatables which have actually been tested. I would say, in fact, that because it is possible for the ballute to move relative to the capsule it decreases the stability of the craft.

Russel · 2012-12-22 06:01:05

Josh,

Just to clear some confusions here. The germ of this thought experiment is to see how to refuel a transit vehicle in Mars orbit, such that that vehicle can do multiple trips from Earth to Mars and back. The next step was basically to design a fuel ferry for the purpose. Next I came across the same conclusion you do, that a vehicle that is scaled to transport a crew from Mars surface to Mars orbit would have to perform multiple trips in order to transport enough fuel. At one stage I figured as many as 6 trips for fuel. Certainly at least 3.

Whether that's a reliability issue or not has yet to be established. However, what I did next was to say, ok, lets build a vehicle large enough to transport fuel in one go. Then use the same type of vehicle to also transport crew from Mars surface into Mars orbit. Essentially in order to be a fuel ferry it is over-specified for a crewed ascent vehicle. That in itself is not a bad thing, and may give you needed overhead if what you're considering is safe abort from any stage.

I should add that we're talking about the same type of vehicle and not necessarily the same vehicle. I'm coming to the conclusion that inevitably it will specialize.

The real question is how much of the design can be common. But for the sake of the exercise I proceeded with the idea that at most we have one distinct type of vehicle, either docked to a capsule, or not. The crewed variant has a "permanently" docked capsule. The non crewed variant doesn't, but is able to fill its oxygen tank fully.

I hope that clears some potential confusion. Behind this is some thought being given to potential permutations. What if, for instance, you're simply interested in commonality of engine components? Well, then you specialize. You have one vehicle that is set up purely as a fuel ferry. And another purely set up as a crew landing/ascent vehicle. That has its advantages too.

For instance, if you discovered that inflatable heat shields were a viable technology you could design that into the crewed variant. Now you're in a position where you always have a crew around to do the necessary repacking/adjustment/cleaning.

The non crewed variant would then take a slightly riskier, faster, deeper approach and rely more heavily on a powered landing. But this would mean you could in theory avoid anything but a standard capsule shape - provided GW is correct and you can thrust through openings in the heat shield and you can get the lift correct. I'll add to this the thought that you might be able to create lift simply using a small amount of asymmetrical thrust from the point of interface.

The only reason I've been talking about more advanced configurations above (multiple pods, wings etc) is simply to push the edges of what can be done to bring the peak temperature down. But there's nothing stopping you building a heat shield for the fuel ferry from existing known technology that will survive many (perhaps a dozen or more) re-entries. Certainly you'd be worried about other kinds of ageing before the heat shield is exhausted.

Now, going back to the "all in one" design. It goes something like this. You've always got at least 2 of these vehicles either in orbit or on the surface. When configured as a crewed vehicle it is docked to a capsule. When its configured as a fuel ferry, its not docked to the capsule. So in the normal course of events you simply leave the capsule docked to only one of the vehicles.

The reason I left it detachable at all was simply a nod to the idea that under adverse situations you might want a form of abort where the capsule itself is free to land on its own.

On a normal landing the capsule remains attached and stays there for later use. Now, in orbit you may want to undock the capsule, swing it around and then dock it to the transit vehicle.

Of course your logistics get harder if for instance one landing/ascent vehicle fails. But at least then you still have options. I'll be happy to go into the details if you wish.

As for launching fuel and crew in one rocket. I've no problem with that since its just an extension of the same concept. Of course, there's things to worry about then too and one big issue is that now the dangers of in orbit refueling are occurring whilst there is a crew around. My mission design includes a spare transit vehicle so I can afford to take more risks with the hardware whilst reducing risks for the crew.

The difficulties in robotic refueling are there, but they are second order issues. And its something we have to get used to, and learn I suspect for other missions nearer to home. However, what I've done to simplify things is I'm simply transporting oxygen - which carries the bulk of the weight of the fuel. I'm also in a position to restock the transit vehicle with breathing oxygen at this point.

As for ballutes, towed drag etc. That's not essential. But I like the idea of something you can adjust at the last moment. Otherwise aerobraking is still a risky business. As for stability, well, its true that so long as the center of drag is behind (from the reference frame of motion) the center of mass, it will definitely work in terms of causing drag. How stable it is, on average, is a good question. We will have to test these things well, before having crew along for the ride.

Bit of history here. Where I started from I actually took a landing capsule along for the ride, along with the transit vehicle. That way you could plan to aerocapture the capsule separately from the main vehicle. Reason being its likely to be lighter relative to surface area to start with, and have more stability and have more control elements. That was then. I'm not sure now. For now I'm just as happy to let the crew aerocapture whilst inside the transit vehicle, but then give the transit vehicle some more safety features.

Stepping back from the landing problem, the overall mission design looks like this..

First step is to land on Mars anything you need there that is going to stay there. I'm going to gloss over this because aside from low thrust trajectories, ion drives etc you've basically got a shed load of gear to ship and that's the same in any mission, almost.

Second step is you deliver into Mars orbit one transit vehicle and two landing/ascent vehicles. Lets suppose for sake of this that these vehicles are specialized.

The fuel ferry descends and takes on fuel. The crewed landing/ascent vehicle is left in orbit and verified as functional. The fuel ferry ascends and delivers fuel to the transit vehicle and the transit vehicle is checked out. The fuel ferry returns the surface and beings to reload. So all of these systems are available and functional before you commit to launching crew. We're also assuming that the technology has been flown in a previous non manned mission.

Third step is to send the actual crew. This starts in Earth orbit with a second transit vehicle with Earth supplied methane/lox. On arrival at Mars, the crew transfer to the crewed landing/ascent vehicle and land on Mars.

During their stay on Mars, the fuel ferry delivers its next load of oxygen to the transit vehicle that arrived with the crew. So now you have a fully fueled spare.

At the end of the stay, the crew use the crewed landing/ascent vehicle to return to the original transit vehicle. If for any reason there is a problem, there is a backup, and spare supplies on the other transit vehicle.

Fourth step is the return to Earth and aerobraking. We then need an Earth return capsule which has been delivered into orbit previously. So another straight forward transfer.

I think you can see how this cycle repeats. The features of this design are that anything that doesn't land on Mars and stay on mars is reusable for a number of missions.

The transit vehicle requires the minimum mass for its trans Mars and trans Earth engine. It doesn't have an attached capsule so that is also a mass saving. Instead the crewed Mars landing/ascent vehicle remains on or near Mars, and likewise the earth return capsule has a similar life cycle.

That's a significant saving, because a lot of designs attach the Mars lander to each and every trans-Mars journey. And I'm happy to go through them with you one at a time and point out their shortcomings when it comes to mass/energy/cost. Point is, if you optimise well, you then have spare mass you can devote to crew safety and comfort. Believe me, one of the things I see going wrong with some mission designs is they consider the threats from externalities and hardware and then create another threat from the crew itself - bored, stressed, tired etc.

Last edited by Russel (2012-12-22 06:43:04)

Russel · 2012-12-24 06:16:30

Mery Xmas from the land of oz and may the new year bring you a damn good lander design!

JoshNH4H · 2012-12-24 21:52:38

To address what is, I think, the most glaring error in your post:

Whether [multiple launches of a Mars Ascent Vehicle are] a reliability issue or not has yet to be established.

While you later suggested that each mission would only involve one launch and one entry, subsequent missions would reuse the same craft. So what you would have is a six month flight to Mars, followed by a long period of stasis (either in LMO or on the Martian surface). There will be an atmospheric entry at some point, followed by a 4 km/s launch. After two years of stasis, repeat. Two more years of stasis, repeat again. With an LMO orbital period of two hours that's hundreds or thousands of thermal cycles for each rocket, followed by a high thrust launch (even though the Martian gravity is lower, it's still pretty high thrust and high powered). After this launch is an aeroentry followed by rocket braking and vertical soft landing. This mission would be a significant feat even with an earth-style support staff on the ground. Instead, you have 6 men and women, presumably no more than two of which will have the skills to deal with this particular rocket, and all of who will by busy trying to fulfill the science and technology goals of the mission.

With regards to your proposed architecture, I would like to make a list of payload items to LEO for your proposed architecture:

Surface Gear- Not the current topic of discussion
- Surface Gear Transfer Vehicle
Transfer Vehicle- reusable
Fuel Ascent Vehicle-reusable *
Personnel Ascent Vehicle-reusable*
Earth Descent Vehicle

Where asterisks indicate that I have serious qualms with the designation of reusability. The lifetime of the transfer vehicle is obviously limited but not so much so as the other components and could be replaced every few missions with little safety hazard. Also, just to specify, the mission involves ISRU (In-situ Resource Production) of Oxygen but not Methane, on-orbit fuel transfer from the Fuel Ascent Vehicle to the Transfer Vehicle, on-orbit transfer of fuel to the Transfer Vehicle in LEO, and on-orbit docking of the Personnel Ascent Vehicle with the Transfer Vehicle, the Fuel Ascent Vehicle with the Transfer Vehicle, and in LEO docking of the Transfer Vehicle with crewed capsules to pick up the crew and to drop them off for reentry. Just for reference, the following is a list payload items to LEO utilized in Zubrin's Mars Direct:

Surface Gear- Not the current topic of discussion, but sent and landed over two orbits, probably in conjunction with crew
- Surface Gear Transfer Vehicle (2)
Earth Return Vehicle (ERV)

None of the hardware is intended to be reused, and there is presumably an on-orbit docking between the Surface Gear/Crew Transfer vehicle once the crew are launched on a separate vehicle and a docking to transfer them back to a similar capsule intended for re-entry.

The ERV has more payload than both your Fuel Ascent Vehicle and your Personnel Ascent Vehicle. However, even ignoring the reusability issues of your architecture Mars Direct (and derivatives) have a major benefit that yours doesn't: Fewer rockets and fewer transfers. Each engine firing is a chance for something to go wrong. When your crew and all of your payload are separate in two rockets, the failure of either rocket results in the death of the crew. This approximately doubles the chance of a loss of crew (technically, it is [1-(chance of one rocket succeeding) x (chance of another rocket succeeding)] ). The use of two rockets where one can be used is done not once but twice in your architecture, both in sending things to Mars and in launching things from Mars to LMO. While you try to compensate by suggesting that the fuel rocket could carry crew to orbit (without suggesting a way to mount the capsule atop the rocket), separating the rockets does not enhance crew safety because without fuel for return the crew is also dead.

More generally, I would just like to give an idea of the amount of rocket that has to be sent to Mars to launch extra payload. The Atlas IIIA rocket stage is fueled by kerosene-lox. Assuming a different fuel mass based on a different density (1031 kg/m^3 for kerosene vs. 828 kg/m^3) but also a higher Isp (365 vs. 335 s vacuum) this Atlas stage would be able to carry just under 45 tonnes of payload. This is for 13.3 tonnes of rocket mass with an additional 195.6 tonnes of fuel (for which about 12.5 tonnes of Hydrogen would need to be imported). You've also suggested Carbon Monoxide fuel.

This report suggests that the Isp of Carbon Monoxide-Oxygen fuel is 250 s. Given a density of 863 kg/m^3, this results in a payload of 15.6 tonnes to LMO with the same Atlas IIIA stage. Even though you don't have to import any Hydrogen for the fuel, you get a significantly lower ratio of rocket payload to landed mass (1.74 vs 1.27) indicating that Methane is probably a better fuel. While the exact figures will vary from rocket to rocket, I would expect the general result to be very much the same.

In short: Your proposed mission saves on mass delivered to LEO at a cost of mission safety and higher technology development requirements. Methane requires less mass lifted to LEO than Carbon Monoxide and is probably superior because we have more experience with similar fuels.

Russel · 2012-12-25 00:29:57

Josh,

Just for starters, here is my starting point regarding CO fuel.

http://ntrs.nasa.gov/archive/nasa/casi. … 014990.pdf

Which gives you an Isp of around 260 for a pressure fed engine, and 290 for a pump fed engine.

I've assumed the lower figure for most of my calculations. That gives a mass ratio of just over 5 for ascent to Mars orbit.

I'm not, in case this has been confused, using CO as a fuel for anything other than the crew lander and fuel ferry.

My preferred fuel for the transit vehicle is methane/LOX.

I'll get onto the other issues once I've gotten through this pile of chocolate..

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