Mars Crew Landers - Two Stage or Single Stage?

[Paul451]

Is there an existing nomenclature that differentiates between an SSTO ascent vehicle that is landed on a descent stage, vs an SSTO that is self-landing, vs a genuine TSTO ascent vehicle?

The last one differentiates itself from the other two by definition, 2 stages != 1 stage.

But it's actually three stages. Lander, ascent booster, ascent upper-stage.

What people have been calling "TSTO" (a la Apollo-LM) is actually an SSTO launcher that is landed on a separate descent stage.

The middle case seems to lack catchy acronym. SSTAFO? (single stage to and from orbit)

There are multiple variants, none of which have clear nomenclature, AFAICT:

Apollo-LEM style single-stage lander that carries a separate single-stage ascent vehicle.

ITS-style single-stage ascent vehicle that is fully self-landing.

Crasher/Uncrasher-type two-stage landers. Where the main deorbital velocity comes from one stage, which is discarded, leaving the smaller lander to handle just the last hundred m/s or so and touchdown. That secondary lander could then be a self-landing single-stage ascent vehicle. Or a self-landing first-stage of a two-stage ascent vehicle (ie, the "middle" stage is both the second-stage of the lander and the first-stage booster for ascent.) Or purely a descent stage that carries a separate ascent vehicle (which could itself be SSTO or TSTO.)

And you could have a single-stage lander that carries a separate TSTO ascent vehicle.

And finally (?), a self-landing TSTO ascent vehicle. Where the single-stage lander is the launch-booster for the TSTO ascent.

So you could have anything from a single all-in-one-vehicle (like ITS at Mars) up to a four-stage system (crasher plus lander plus booster plus orbiter.)

Obviously you can just throw letters at these until you cover all variants, but is there a simple logical naming system.

(DL/AO, DLAO, DB/LAO, DB/LAB/O, DB/L/AO, DB/L/AB/O, DL/AB/O, DLB/O.)

[colbourne]

I would have thought it best to design the lander/ascent craft  to use cheap expendable solid fuel boosters. These give the best of both SSTO and multi stage designs.

Solid boosters have the worst Isp, which really hurts in a stage that you have to take with you all the way through launch, escape burn, capture and landing. Better is to have a mass efficient stage. Even better is to produce the fuel locally.

As perchlorates are common on Mars is there not potential to produce these boosters locally ?

[TrevorMonty]

I like fully propulsive landing system, it does require in orbit refuelling to keep DV down to 4km/s.  Even with LH/LOX round trip of >8km/s is marginal when you add weight crew cab. Fuel could be delivered to orbit by ITS shape reusable tanker, airbraking bring DV down to manageable level.

[clongton]

Has anybody asked themselves why the only organization that is actually going to go to Mars has selected CH4 in lieu of LH for its propulsive fuel? I continue to see people discussing LH as if that is going to be the fuel used on and around Mars. It's not. SpaceX has no plans to use LH and they are the only ones that are actually going to go there. Mars crew landers will be reusable and will be powered by in situ CH4, not LH.

[Hanelyp]

Is this lander single use disposable? partly reusable? fully reusable?  If reusable, is it refueled on the Martian surface?  In orbit from the interplanetary transport?  Both?

[redliox]

Is this lander single use disposable? partly reusable? fully reusable?  If reusable, is it refueled on the Martian surface?  In orbit from the interplanetary transport?  Both?

I left that open-ended for members to interpret on their own.  I would presume single-use Martian landers would serve as modules for a base to maximize their use.  The two-stage question applies more toward crew vehicles that need to ascend back to orbit.  Assuming two-stage (whereas a single-stage perhaps reusable), most likely its would be disposable although there's no reason the upper stage could be used if you can handle hauling it either to Earth or a station in High Mars Orbit.

[redliox]

Going to do a thought experiment using a few assumptions.  In this case it's going to be about how fuel choice affects ascent from Mars.  I'll set a few baselines.  A Mars lander could be conjured up in many ways, but for the moment I'm assuming 2 chief things:
-Will be at least twice as heavy as the Apollo LEM
-Partly using Mars Direct as a guideline

So I'm imagining a 30 metric ton lander sitting on Mars.  Of that 30 tons 9 would be dry mass; if its two-stage the ascent vehicle would probably be 6, largely because it would include the crew cabin and its needs.  We'd have 21 mt of propellant most likely manufactured on Mars.  Thanks to modern probes, we can be confident to find water abundantly.  There's 2 fuel choices
-LOX/Hydrogen (ISP of 400)
-LOX/Methane (ISP of 350)
(I'm assuming slightly lower specific impulses; hydrogren can be as high as 450 and methane 360 or better)

For a single stage lander with a dry weight of 9 mt, 21 mt (of fuel) yields the following delta-v:
-LOX/H2:  4.7 km/s
-LOX/CH4: 4.1 km/s

This would be more than sufficient for both low orbit and Phobos', with hydrogen giving some reserve margin.  Not enough to reach escape or a high orbit, but still useful.

With a two-stage lander with an ascent stage of 6 mt, 21mt (of fuel) would yield:
-LOX/H2: 6.3 km/s
-LOX/CH4: 5.5 km/s

Both would be able to escape Mars, although methane only just able.  However, the lander will more likely target high orbit to rendezvous with a transit vehicle (which is kept high to minimize its own fuel needs for the Earth return).

My calculations are on the optimistic side, but I think it would be safe to conclude several things:
-orbits as high as Phobos can be reached within reason
-orbit of the transit vehicle (ERV, whatever) would affect whether a single or two-stage Mars lander required/possible
-low/medium/Phobos orbit should utilize methane, high orbit hydrogen
-singe stage/reusable Martian vehicles too heavy to assume return to Earth

[Robotbeat]

Problem with your analysis is you assume same dry mass for both methane and hydrogen. Very poor assumption. Methane/LOx is twice as dense as hydrolox, which means the lander can be like half the volume (not quite half, of course). Also, it takes less energy AND less water to produce a given mass of methane/LOx than hydrolox.

Plus, of course, methane is much easier to store & cryocool than hydrogen.

This is why methane is looked on so favorably for a lander.

[Robotbeat]

A Vertical landing rocket stage (first or second... Second includes heat shield) is essentially an Earth lander. This means you can make an incredibly mass efficient lander if you really want. So you can do direct to Earth if you want.

[redliox]

Problem with your analysis is you assume same dry mass for both methane and hydrogen. Very poor assumption.

Not really.  The same design for hydrogen could store either oxygen or methane easily (but the inverse not so much) and insulate it well; the engine to burn the fuel is the real issue.  Also it is now known water is fairly common on Mars; both hydrogen and methane could be manufactured from it.  Since both propellants can be obtained using water, it is reasonably possible either could be utilized.

[Toast]

The same design for hydrogen could store either oxygen or methane easily (but the inverse not so much) and insulate it well

I think you missed his point--we aren't going to design a hydrogen-fueled rocket, then decide to change it over to methane later. Sure, a tank that could hold hydrogen would be more than capable of holding methane, but it's not an optimal solution. The point he was making is that a rocket designed bottom-up to use methane could be more efficient (in terms of size and density, not isp) than a rocket designed bottom-up to use hydrogen. Because methane is denser, an ascent stage based on it could be much smaller (and lighter) for a given mass of fuel than an equivalent hydrogen stage.

[redliox]

The same design for hydrogen could store either oxygen or methane easily (but the inverse not so much) and insulate it well

I think you missed his point--we aren't going to design a hydrogen-fueled rocket, then decide to change it over to methane later. Sure, a tank that could hold hydrogen would be more than capable of holding methane, but it's not an optimal solution. The point he was making is that a rocket designed bottom-up to use methane could be more efficient (in terms of size and density, not isp) than a rocket designed bottom-up to use hydrogen. Because methane is denser, an ascent stage based on it could be much smaller (and lighter) for a given mass of fuel than an equivalent hydrogen stage.

Well there is some good reason to nominally use a design optimized for hydrogen.  Many missions, including both Lockheed's Mars Camp and Zubrin's Mars Direct, assume long-term storage of cryogenic propellant to be a significant part of the mission.  This means storing LOX and whichever preferred propellant for months at a time with as minimal boil-off as possible.  A hydrogen tank's insulation needs make for a good starting point if you wish to do this without over-the-top future tech.

Indeed as Toast points out, a key difference is density.  Methane will be easier to store compared to hydrogen; it should be favored.  However you can't rule out hydrogen either, as it can be manufactured directly with electrolysis on Mars.  If you want a fully independent supply of methane you need some hydrogen to seed the process.  So, frankly, we need both to some degree.  May as well ensure the tankage can handle the more difficult of the 2 fuels.

Bear in mind, the thought experiment I did was to show how fuel choice can enhance a mission.  Either hydrogen or methane could tackle most Martian orbits, but hydrogen is the better choice via delta-v for high orbit...which is where most schemes currently are positioning the orbiting return vehicle.  The point was to show what options are available hypothetically.

Personally I favor methane, but it might not be an absolute choice.

[Robotbeat]

You don't need any seed liquid hydrogen for making methane.

 Storing liquid hydrogen on Mars is hard. It is super bulky and wants to boil off on you.

There's been a big change from years' past in that we now know Mars has water everywhere. This means we don't have to bother with hauling hydrogen all the way from Earth with all the huge headaches that involves. That makes the original Mars Direct obsolete.

It also means you don't have to have just disposable landers.

It also means your lander, for the same mass sent from Earth (and the same dry mass) can actually get more mass in orbit with methane than with hydrogen.

This is an important point that often goes overlooked! You're only sending dry mass from Earth, so, in fact, methane outperforms hydrogen (for the delta-v of a Mars ascent vehicle) when measuring by dry mass. We think too much in terms of wet mass and not enough in terms of dry mass. Isp is talking about wet mass. But the propellant we get on Mars! We don't bring it from Earth. So in that case, we should judge on dry mass instead.

Another thing: it takes less energy to produce methane than to produce hydrogen, on a /mass/ basis.

[clongton]

Why is anyone even talking about using hydrogen engines on this vehicle? Hydrogen is not going to be used. It will be methane. That decision has already been made.

[MickQ]

Snip

It also means your lander, for the same mass sent from Earth (and the same dry mass) can actually get more mass in orbit with methane than with hydrogen.

And therefore you can put the same mass into a higher orbit ?

[redliox]

Snip

It also means your lander, for the same mass sent from Earth (and the same dry mass) can actually get more mass in orbit with methane than with hydrogen.

And therefore you can put the same mass into a higher orbit ?

That's basically what it comes down to.  Some NASA studies assumed targeting low Mars orbit, but more recent ones have the orbiting vehicle in an elliptical high orbit.  A high orbit allows for the Earth Return Vehicle, in whatever form it may take, to arrive and depart with less fuel; putting fuel in Mars orbit is a major reason why some plans needed numerous launches.

[Oli]

This is an important point that often goes overlooked! You're only sending dry mass from Earth, so, in fact, methane outperforms hydrogen (for the delta-v of a Mars ascent vehicle) when measuring by dry mass. We think too much in terms of wet mass and not enough in terms of dry mass. Isp is talking about wet mass. But the propellant we get on Mars! We don't bring it from Earth. So in that case, we should judge on dry mass instead.

Another thing: it takes less energy to produce methane than to produce hydrogen, on a /mass/ basis.

Funnily enough for something like ITS (~7km/s) the benefits of hydrolox start to show. I kind of expected it to be hydrolox up to the announcement.

I'm not sold on subcooled methalox for in-space propulsion. Sounds much more of a pain to me than hydrogen.

[Robotbeat]

This is an important point that often goes overlooked! You're only sending dry mass from Earth, so, in fact, methane outperforms hydrogen (for the delta-v of a Mars ascent vehicle) when measuring by dry mass. We think too much in terms of wet mass and not enough in terms of dry mass. Isp is talking about wet mass. But the propellant we get on Mars! We don't bring it from Earth. So in that case, we should judge on dry mass instead.

Another thing: it takes less energy to produce methane than to produce hydrogen, on a /mass/ basis.

Funnily enough for something like ITS (~7km/s) the benefits of hydrolox start to show. I kind of expected it to be hydrolox up to the announcement.
...

Not if you're still talking about a /single/ stage and you're limited by /dry mass/ and not wet mass. Since a single-stage lander is limited by dry mass (not take-off mass), methane/LOx's doubled density vs hydrolox makes it the superior choice for all delta-v.

Let's take the limiting case of a single stage, of dry mass 10 tons and a volume of 100 m^3 (tank dry mass scales linearly with volume... and engine thrust per weight also scales roughly linearly with bulk density). Methane/LOx has a bulk density of 828kg/m^3, hydrolox has just 358kg/m^3 with vacuum Isp of 469s and 386s, respectively, for exhaust velocity of 4.60km/s and 3.78km/s.

The wet mass will thus be 45.8t and 92.8t, respectively. And Mass ratios of 4.58 and 9.28 respectively.

So what's the ultimate delta-v of both?

hydrogen: 4.60km/s*ln(4.58) =7.00km/s
methane: 3.78km/s*ln(9.28) = 8.42km/s

So, in this limiting case, you'll notice that methane has higher total delta-v! Dry mass is the same. That means for 7km/s delta-v, hydrogen gets zero payload!

So this idea that hydrogen performs better at higher delta-v must be taken with a grain of salt! It's true if you're talking WET mass (which is quite relevant for multi-stage rockets, as the first stage mostly just cares about the WET mass of the stage above it), but actually usually false if you're concerned with stage DRY mass (which we are here, since we're producing the propellant on Mars and are here talking a single stage).

[Oli]

Let's take the limiting case of a single stage, of dry mass 10 tons and a volume of 100 m^3 (tank dry mass scales linearly with volume... and engine thrust per weight also scales roughly linearly with bulk density).

I'm interested in real-world dry mass fractions. Dry mass fractions scaling linearly with propellant density is just not something I have found out there. Methalox stages having 2/3 of the dry mass fraction of hydrolox stages seems to be a good estimate. Admitted, that does not include TPS, landing legs etc. but TPS actually scales with area.

By the way, in these slides (http://images.spaceref.com/news/2010/SpaceX_Propulsion.pdf) Tom Markusic (SpaceX) assumes an inert mass fraction of 0.06 for RP-1 and 0.08 for LH2 (including oxidizer I guess). I actually find that very optimistic.

So this idea that hydrogen performs better at higher delta-v must be taken with a grain of salt! It's true if you're talking WET mass (which is quite relevant for multi-stage rockets, as the first stage mostly just cares about the WET mass of the stage above it), but actually usually false if you're concerned with stage DRY mass (which we are here, since we're producing the propellant on Mars and are here talking a single stage).

When we're talking ~7km/s the propellant mass for the same payload is substantially less with hydrolox. At least with my specs. At that point it's questionable whether methalox is cheaper to produce on Mars (depends on cost of water mining relative to energy cost). Moreover, the booster on Earth can be smaller and the IMLEO mass is substantially lower (after all that is basically a third stage).

It's obviously still possible for methalox to win cost-wise, but at least to me it doesn't seem like a clear-cut case.

[Russel]

I'd like to point out the elephant in the room. Everyone seems to think an ascent vehicle has to be massive (I had to chuckle at the MAV used on "The Martian"). Everyone is arguing about fuel and Isp. But the single biggest variable is the dry mass of the ascent vehicle itself.

We have come a hell of a long way since the  LEM was dedigned. Modern, light materials. 3D printed engines. Advances in electronics and batteries. And the list goes on. Why can we not build an ascent vehicle that is the space equivalent of a beach buggy - light and tough?  A tonne or so basic mass (sans crew and consumables). Including tankage, engines and a pressurised (0.5atm) capsule.

This basic mass is multiplied many times - fuel/ISRU/power etc etc. Yet we don't have good estimates of what can actually be done with today's materials.

I know its diificult and complex. I guess that's why it gets skipped over. But the issue of the basic structural mass of the ascent vehicle is for me far more important than for instance the choice of fuel.