Speculative - cislunar transport infrastructure, 2054

Regarding the ice lakes on planets in the inner solar system: The definition of "inner solar system" I am informed of is "between the sun and the asteoid belt". There is at least one ice lake on Mars and Mars' orbit is between the sun and the asteoid belt. Consequently there is at least one ice lake on a planet in the inner solar system and it is accessable - which means that it could be used for what you have in mind Wannabespacecadet. If I remeber right then at least one more ice lake has been detected on Mars in the recent months - closer to one of the poles than the first detected ice lake which is at the equator (or nearly there).



Dipl.-Volkswirt (bdvb) Augustin (Political Economist)

Is there an echo in here?

I should have looked at my own link. The DeltaV chart shows 9,300 m/s from Mars to the Moon. So a LH2 NTR would be needed for SSTM at any speeds above minimum energy trajectories.

Ceres may turn out to have water, but since it is main belt and not near Earth, it takes more deltaV to get water from there than it does to get it from the Moon.

(EDIT) Dr. Paul S. Hardersen was on the most recent space show talking about mining asteroids.
http://thespaceshow.com/detail.asp?q=453
He currently serves as an Assistant Professor in the Department of Space Studies at the University of North Dakota. He received his PhD in geology in May 2003 (specialization: asteroid near-IR spectroscopy) from Rensselaer Polytechnic Institute in Troy, New York. I sent him an E-mail asking about evidence of water or hydrated minerals on near Earth asteroids and this was his response:

I shouldn't think you need too much of the 6,400 m/s at the Mars end.
IIRC, Mars has an atmosphere. Aerobraking? In fact the chart indicates that with "AB".

I still think the advantages of water NTR are:
1) Simpilicty of engine (debatable until the technology is proven & matured)
2) Simplicity of storage in depot and vehicle
3) Simplicity of off-Earth production

If we have sufficient off-Earth infrastructure, then a mix of LH2 NTR & LOX/LH2 is the way to go. But until we do, water NTR with some LOX/LH2 and/or LOX/Methane on Mars (where we probably care about radio-active exhaust fumes), looks promising.

Peter, you wrote also:



So obviously there is no "echo in here".

You don't have no reason to polemicize, Peter - to do so is the next mistake, error to be listed.

Of course nobody is perfect and nobody is safe against not contradicting himself unwillingly. But this means that it should be clarified which of the contradicting issues under which conditions is wanted to be considered to be meant seriously by others. In this special case here only one of the issues can be meant in serious.

...



Dipl.-Volkswirt (bdvb) Augustin (Political Economist)

Well, I am thinking about the return trip, from Mars to the Moon. I agree that getting from the Moon to Mars is pretty easy, just like getting from the Moon to LEO is easy, using aerobraking. That is because it is "down hill".

Hey, I thought you didn't like aerobraking.

Yes, that is what I am doing, debating the maturity of NTR. That and the fact that water NTR does not perform better than, or even as well as, LH2/LOX.

Can you aerobrake in the Earth's atmosphere, and still end up in Lunar orbit? Probably wouldn't save too much delta V anyway. Refueling in Lunar orbit has to be a big advantage.



I'm 'unwillingly' contradicting myself

Well, looking at the handy dandy deltaV chart I see that it takes 6.4 km/s just to leave Mars. You can save 0.6 km/s by aerobraking from "Mars low energy transfer orbit" to "Earth C=3 orbit", but you still need another 0.7 to enter lunar orbit and another 1.6 km/s to land on the Moon. Of course you could aerobrake directly into LEO and refuel there, if there were a depot there. So I see 6.4 km/s from Mars to LEO, 7.1 km/s to lunar orbit and 8.7 km/s Mars to lunar Surface. That is a little better than the original 9.3 km/s for Mars to Moon with no aerobraking at all, but not much. Especially compared to the paltry 3.8 km/s you need to go from the Moon to Mars with aerobraking at Mars. Maybe the best solution would be a reusable booster based at Mars. For the return trip, the vehicle is stacked on the booster which acts as a first stage and takes the vehicle to Mars orbit, then returns to Mars for refueling. The vehicle can then swing by Earth for the 0.6 km/s aerobrake and return to the surface of the Moon with a deltaV of only 4.6 km/s, or 3.7 if the booster takes it all the way to Phobos transfer orbit. The booster would only need to supply a deltaV of between 4.1 and 5 km/s, but it would have to be pretty big to carry the fully fueled Moon shuttle as a payload. Making LOX and LH2 should be really easy on the surface of Mars. (EDIT) Well, making LH2 is not all that easy even on Earth, but it shouldn't be much harder on Mars. (/EDIT)

Maybe we need to change the name of this topic to Martian transport.

I would like to start this post by pointing out that rockets are hard to design when your delta-V is greater than your rocket exhaust velocity. That's when you are really hit with the exponential nature of the rocket equation. Below this point your mass ratio is 2.7 or less and the effect of further increasing Isp is small (the partial derivative of mass ratio with respect to Isp.) At that point it's quite feasable to create an RLV, and factors like crew salary and maintainence cost will come to dominate.

So my first conclusion is that technology development will proceed until delta-V's are about equal to exhaust velocities of whatever technology is in use. There are two ways this can be accomplished. The first is to develop rockets with higher Isp. The second is to break the trip into smaller chunks with a refueling depot between each chunk.

Looking at it this way, Earth to LEO really stands out as a huge chunk that you have to swallow all at once. The assumption in this thread is that there is significant infrastrucutre and traffic in space. So my second conclusion is that there will be some kind of reusable infrastructure that reduces the delta-V from earth to LEO. This may be a hypersonic skyhook, rotating tether, catapult launch from the ground, 50km tall tower, or whatever. Such a thing could be paid for if amortized over thousands of flights per day. A true space elevator might be possible, but I wonder if it would cost less than a hypersonic skyhook that reduced delta-V to 5.5km/s plus a cheap reusable rocket that can make 5.5km/s.

Then this residual chunk of velocity will come to dominate decisions about where to place other refueling depots. For example, to get from LEO all the way to the lunar surface is 5.5km/s so the rocket/skyhook combination above could do everything in cis-Lunar space if it could refuel in LEO and on the lunar surface.

This is getting long so I'll summarize my abstract conclusions, and in another post I'll present concrete details of one possible implementation.

1) Infrastructure will be built to reduce the delta-V of Earth to LEO.

2) Rocket technology will be built to have a Ve approximately equal to this delta-V.

3) Additional fuel depots will be built only where necessary to ensure trip segments not exceeding this delta-V.

Ah yes, the crux of all our problems. Overcoming that barrier is really step zero in any space activity.

This is a continuation of my previous post. One thing I didn't mention is that there will be low thrust technologies like ion engines and solar sails for cargo, but they are just too slow for people. And not just people. Today we send a certain amount of cargo by next day air even though an ocean going freighter would be cheaper.

In my previous post I left these questions to be answered:

1) What launch assist infrastructure will reduce the Earth to LEO delta-V?

2) What type of rocket will make up the residual delta-V?

3) Where will the fuel depots be, and where will the propellants originally come from?

#2 can be further broken down into two questions: what will be the energy source, and what will be the reaction mass? In 50 years I can't see any technology other than thermal rockets having the thrust to weight ratio needed. This includes chemical thermal, nuclear thermal, or beamed power thermal (of which solar thermal is a sub-type.)

If you can use liquid hydrogen propellant then nuclear and beamed power have a significant Isp advantage. Liquid hydrogen is simply much lighter than any chemical reaction product. But with any other propellant such as water, methane, or ammonia the advantage is much less. You can get nearly the same amount of energy and molecular weight with chemical reaction products. The primary limitation becomes the temperature limits of the materials in your engine.

Dealing with liquid hydrogen is a problem and even more so in space. It is much colder than liquid oxygen, has a variety of material incompatibilities, and is very good at leaking due to it's small molecular size. In space you have to worry about losing it to boil-off, and where do you get it in the first place.

So I'll outline two options. One is a nuclear or beamed power (I don't think it matters which) thermal rocket using pure LH2 with a Ve of 8km/s. The other is a LOX/methane chemical thermal rocket with a Ve of 3.5km/s. Undoubtedly both of these rockets can and will be built. The question is which one will wind up being most economical and dominate the commercial fleets. I don't think I can answer that now.

In the LH2 case there will be a need for very little if any launch assist. The round trip from LEO to lunar surface and back is 7.8km/s with aerobraking so it may be best to design for that delta-V and have a vertical catapult in a mine shaft to supply the additional 1.9km/s for Earth to LEO, basically paying gravity and drag losses.

There would be no need for an LH2 depot on the moon. The LEO depot could be a massively huge sphere to take advantage of the squared-cubed law to prevent boiloff. While the moon would be lots of small settlements and having a LH2 depot at each one would be more trouble.

Of course, this would mean when you get to the Moon you drop off your load and come back right away to minimize boiloff. If you did get stuck you could refuel with an emergency supply of storable propellant on the Moon. Perhaps liquid oxygen. You could use it as low performance reaction mass and ignore it's chemical energy. This could also be used for suborbital point to point hops on the moon.

With LOX/Methane just getting to LEO would be more trouble. You would likely have a large orbiting structure catch a suborbital rocket. Then you can refuel and make it to lunar orbit. Then you can refuel and make it to the lunar surface and back.

As far as getting the propellants to the depots that's a more open question. We really don't know if there will be enough economically extractable deposits of hydrogen that can support such a large transportation system. If necessary, hydrogen could be brought from Earth, Mars, or the outer solar system. In that case, we would probably be sparing with it. Chemically bind it into methane or another storable fuel as soon as it reaches the depot and burn it in a chemical engine. We know we can get carbon and oxygen at lots of places.

And the other way too. So the 7.8 km/s rocket starts fully fueled on the moon and returns there empty. Hydrogen has been measured at the poles on the Moon, the only question is how hard would it be to extract. So far no hydrogen has been detected on any near Earth asteroid. Mars, Ceres, comets and every other place we know of where hydrogen exists requires more deltaV to reach than the Moon. If the hydrogen is produced on the Moon, then that is the logical place to put the depot and base all the operations.

As I've posted in another topic, one solution is a big RLV booster that's good for about 2.5 to 3 km/s delta-V. This thing would be huge, but use low cost, easy to handle fuel (maybe LOX/Ethanol), and low maintenance design. A mass ratio of about 3.5 would be sufficient. Say 72% fuel, 14% vehicle, 14% payload

It could lift a vacuum optimized rocket with 6 to 7 km/s delta-V, inside an aerodynamic fairing, that opens at around 70 to 80 km. The booster would then re-enter (no big deal at those speeds) and land vertically 200 to 300 km down range. At a GLW of 2000 tonnes, each flight could cost less than $ 1 million in fuel. Flight rates would determine the ammortized cost per flight for development, maintenance, support facilities etc, but could easily be about the same or less.

2000 tonnes would put 280 tonnes of rocket into sub-orbit. An LH2/LOX rocket with mass ratio of 5 would have over 7 km/s delta-V. Say 80% fuel, 13% vehicle, 7% payload, puts about 20 tonnes in orbit.

Here is a quote from a recent space.com story about the upcoming Lunar Reconnaissance Orbiter mission.
Yes, there is a hydrogen signature just where everyone really wanted to see it, but it is weak. We need to guard against reading our hopes and dreams into the readings and just consider the actual data dispassionately. It is too early to jump to the conclusion that we could actually get water from the lunar poles.

There is some disagreement at NASA about the best place for a lunar base. Most are in favor of the poles, but some are not.
http://www.newscientistspace.com/article.ns?id=dn8683

It turns out that there is hydrogen and other volatiles in Apollo samples.
http://fti.neep.wisc.edu/neep602/9301/n ... 0000000000
Up to 60 ppm hydrogen, which is 3 kg of hydrogen per 50,000 kg of regolith. A huge strip mine would be needed to supply enough hydrogen for ongoing use as rocket propellant. That would be expensive, possibly more expensive than bringing it from Earth or Mars, but not totally impossible.