Developing the cis-Lunar economy and infrastructure

Terraformer · 2012-06-12 11:29:30

However, keep in mind that if your engine fails on your ERV, you're not getting back to Terra either...

From EML1, it's only a 90 day trip with a delta-V of 3.6km/s. Perhaps achieving orbit at the other end would be viable with a combination of retrorockets and aerocapture? On the plus side, half of your transfer vehicle will be hydrogen, so you don't have to worry about solar flares during your flight...

If water is present on Phobos, then the return fuel could already be present in Martian orbit when the crew arrive. In that case, I'd remain on orbit - with substantial water shielding, so much that it's the safest place to be from a radiation perspective - and make a multitude of sorties to the surface. Once the next suitable window comes up, they can execute an equally speedy return to Terra using the substantial fuel reserves available.

Nothing is thrown away, the entire mission can be redone for a few dozen million dollars, and we get proper exploration done. Now, if only we knew whether there was actually water in Phobos...

RGClark · 2012-06-12 11:37:19

louis wrote:

Excuse my ignorance...
How does the SSME match up with the Falcon series? Is there anything there which could get you to Mars that quickly?

If it were the shuttle ET, then it would require 14 launches of the Falcon Heavy to deliver the ca. 720 mT propellant load to LEO. That's in the range of $1.4 billion. The plan is to keep the launch costs in the few hundred million dollars range.

Remember though this was for delivering a 100,000 lb. payload to Mars. I believe you can get a workable system that would only take 1 to 2 Falcon Heavy launches for the propellant, which would cost perhaps $200 million to bring the propellant to LEO. But perhaps you are asking about using a kerosene fueled engine?

Keep in mind the vehicle has to be refueled at Mars. Probably it could work using methane in the engine. But for first experimental missions it would be much easier to use LH2/LOX, because of the water ice proven to be wide spread on Mars.

For a small crew size say 2 to 3 and for transit times only about 70 days with 12 day stay on Mars, we might be able to get a crew hab at only ca. 5 mT mass. I'll write about this in a following post.

For Centaur-style stages, we may suppose at least 10 to 1 propellant to dry mass ratios. So let one stage have a 30 mT propellant at 3 mT dry mass and one be at 20 mT with a 2 mT dry mass. Take the Isp of the RL-10B2 engines used on the Centaurs as 465.5 s, which has already been achieved using nozzle extensions. Then with a 6 mT payload you could reach the 8.8 m/s delta-v needed to reach Mars in 70 days:

465.5*9.81(ln(1 + 30/(3 + 22 + 6)) + ln(1 + 20/(2 + 6))) = 8,800 m/s.

Then you could launch the 33 mT propellant load plus dry mass of the larger Centaur, plus the 20 mT propellant load of the smaller Centaur on a single Falcon Heavy. Then you could launch the 6 mT hab and 2 mT smaller Centaur dry mass on a Falcon 9.

Then the launch cost would only be in the range of $150 million. Of course this doesn't solve the problem of how you get the ca. 18 m/s delta-v needed for the return trip to only have a 70 day transit time.

Bob Clark

Last edited by RGClark (2012-06-13 00:31:46)

RGClark · 2012-06-12 11:45:09

Terraformer wrote:

However, keep in mind that if your engine fails on your ERV, you're not getting back to Terra either...
From EML1, it's only a 90 day trip with a delta-V of 3.6km/s. Perhaps achieving orbit at the other end would be viable with a combination of retrorockets and aerocapture? On the plus side, half of your transfer vehicle will be hydrogen, so you don't have to worry about solar flares during your flight...

I haven't seen transit times that short even from EML1 with only a 3.6 km/s delta-v. Do you have a ref for that?

Bob Clark

Terraformer · 2012-06-12 11:51:39

No, they were lost in the Great Crash. Perhaps they'll be on Hop's website (clowder.net?). I do remember them being that short though...

Terraformer · 2012-06-13 00:35:06

Ah, I can see where the delta-V is coming from now: you're using a gravitational flyby of Terra to provide most of your velocity. With the Terran escape velocity of 11km/s and a burn of 3km/s, your total delta-V^2 = 14^2 - 11^2 = 75, so your delta-V = sqrt(75) = 5sqrt(3), which is ~8km/s.

Impaler · 2012-06-13 04:41:40

I've got a SpaceWorks study that says a 1,770 m/s Delta-V Mars departure from EML1 can get you to Mars by use of a Lunar & Earth sling-shot with 3 burns, one at L1 and 1 at each perigee 100 km and 400 km altitude respectively. The C3 is 28.2 km2/s2 and the outbound duration while not stated explicitly should be in the 200-250 day range if they were making apples-2-apples comparisons with the LEO departures. So adding an additional 1,830 m/s might considerable improve the transit time while potentially leaving enough propellent for an insertion burn at Mars. The main draw-back is that your departure window becomes constrained by the cycle of the Moon which

Another potential launch point could be the Earth-Sun L2 point, it's not much further out then the orbit of the Moon and it's even higher up the Earths gravity well. Scientific probes have already been placed their and I suspect the Mars departure Delta-V is extremely low. Most importantly of all it's close enough that the Transit vehicle can be replaced and then the crew sent by small taxi craft in a reasonable time frame, I'd estimate a week tops.

Pre-positioning the Transit vehicle by SEP is the best way to make a fast human transit feasible, with it you could probably do a fast <1 month transit with the same IMLEO as a slower 200+ day Hollman direct from LEO. Ultimately I think we will get enough of a handle on Radiation that the Hollman is the desired choices for higher mass delivery.

Terraformer · 2012-06-13 05:00:38

I presume you mean Hohmann?

Radiation isn't usually that much of an issue for cargo. It's reasonable to suppose that we'll be using fast transfer for people and animals and Hohmann transfers for cargo. As I've pointed out before though, your insertion fuel is going to be considerable, so you've got a lot of mass for radiation protection anyway.

When you reach Terra you're going to be travelling at or near escape anyway, so there's nit going to me much advantage from starting further up in the gravity well.

What I find very interesting personally is that a delta-V of 2km/s can send a payload to Ceres...

It should be very evident to everyone, even Zubrin, that Lunar infrastructure is a very useful thing to have.

When it comes to returning from Mars, though, we're looking at either SML2 (Sun-Mars Lagrange 2) or Deimos to launch from, to take advantage of a Mars flyby.

RobS · 2012-06-13 08:06:29

Zubrin is often more of a debater than a scientist/engineer, and I think that's true where the moon is concerned. Moon plans are a threat to Mars plans because we usually can't do both, so he minimizes the value of the moon.

Lagrange points are useful if you are spiralling a lot of stuff slowly out of the Earth's gravitational well. They provide a region to park things. I wouldn't use solar ion propulsion if the latter uses xenon; currently xenon is worth more per ounce than gold. The world only produces a dozen tonnes of it per year. On the other hand, if one switches to argon, which is cheap, that problem is solved. Currently xenon is a practical propellant only when the launch cost to LEO is many thousands of dollars per pound. I'd favor development of solar thermal anyway; the technology has been tested on the ground, its thrust (up to 100 pounds currently) is much more than ion, the thrust can be "stored" up and used in a series of perigee kicks, and hydrogen boiloff isn't an issue because you're using your hydrogen supply for a few minutes every few hours.

The main advantage of a Lagrange parking point, though, is slow transport of stuff out of most of the Earth's gravity well. Developing that system would also help lunar transport because the same system could move cargo to low lunar orbit, reducing the cost of transport to the moon's surface. It takes something like 300 meters per second to "break out" of EL2, the point between the earth and moon. I haven't seen similar figures for other Lagrange points. Trans-Mars injection still occurs deep in a gravity well, though. Earth escape is 3.1 km/sec from low Earth orbit. From that same location, a Hohmann trajectory to Mars is 3.7 km/sec and a 6-month trajectory is 4.3 km/sec. Neither is substantially more than escape velocity because at the higher speeds you escape with more residual velocity. Possibly one could use a smaller initial delta-v to leave EL2 and head for the moon; maybe 100 m/sec or so. There, one could fire one's engines in its gravity well and gain some advantage, then fly past earth and fire engines again. But falling toward the Earth, you lose some of the gain because you fall into the gravity well faster and thus gain less speed from the pull of the Earth. The gain may not be worth it. You may want to spend a few extra days to go from EL2 to the moon, use its gravity to fling you to the Earth (no engine firing), and do your trans-Mars injection close to the Earth, just as you would if you started in low Earth orbit. In that case, the main advantage of using EL2 would be (1) spiraling stuff to near-escape and/or (2) loading up with lunar fuel at EL2. If we are using Falcon Heavies ($1000/pound) and reusable Falcons (maybe $250/pound) or gas guns, I'm not sure the use of lagrange points is really economic or worthwhile. Just launch everything to LEO and leave from there.

Impaler · 2012-06-13 21:21:40

Utilizing the EML1 point has advantages in Moon exploration primarily because it greatly widens the launch windows and landing latitudes, you can reach EML1 at any time and from any initial LEO-inclination at nearly the same cost. And then from EML1 the cost to land on any part of the moon (like the pole) is again the same you simple time your departure from EML1 correctly. Essentially EML1 breaks the orbital inclination linkage between the LEO parking or launch inclination and the accessible landing latitude on the Moon. When the Apollo mission left Earth from a 28 degree inclination that meant the Service/LEM stack arrived at the Moon at a similar inclination and the LEM fuel supply allowed it basically no inclination change during assent and descent. In the proposed the Altair/Orion moon mission the stack was supposed to have extra Delta-V making it capable of a higher Delta-V LOI that could put it into a low Polar orbit and allow the Altair to then land and return from the polar areas.

If that same mission configuration was changed from a LOR (Lunar Orbital Rendezvous) to LPOR (Lagrange point Orbital Rendezvous) the total Delta-V goes up slightly from the absolute minimum Apollo type equatorial mission profile but not compared to the LOR Polar landing. And the window advantages are considerable, descent and assent from the Lunar surface at ANY TIME. LPOR is a good choice even for a one-off all disposable mission without any refueling at all. As soon as you add refueling it's essentially mandatory.

EML1 staging for a Mars mission though only makes sense if your using SEP to pre-position fuel and vehicles. If it's all chemical then going to and stopping at any orbit is just adding vehicle assembling, multiple launch campaign and Cryo-boil-off issues with no real change in Launch mass. Only the maximum size of the Launch vehicle is reduced and it's widely believed that a launch vehicle will get cheaper per kg the larger it is. Once you have access to a high ISP tug system that can move a chemical departure stage and payload to EML1 your Initial Mass in LEO can be cut in half which is a huge advantage regardless of the size of the Launch vehicle.

Last edited by Impaler (2012-06-13 22:33:40)

RGClark · 2012-06-13 23:29:47

Impaler wrote:

I've got a SpaceWorks study that says a 1,770 m/s Delta-V Mars departure from EML1 can get you to Mars by use of a Lunar & Earth sling-shot with 3 burns, one at L1 and 1 at each perigee 100 km and 400 km altitude respectively. The C3 is 28.2 km2/s2 and the outbound duration while not stated explicitly should be in the 200-250 day range if they were making apples-2-apples comparisons with the LEO departures. So adding an additional 1,830 m/s might considerable improve the transit time while potentially leaving enough propellent for an insertion burn at Mars. The main draw-back is that your departure window becomes constrained by the cycle of the Moon which
.

Do you have a link for the SpaceWorks study?

Bob Clark

Terraformer · 2012-06-13 23:56:19

Impaler, you're forgetting the impact of Lunar fuel in orbit and EML1. For a Hohmann, yes, it wouldn't make sense to use fuel to travel to EML11, refuel, and then launch, but if you want a fast transfer using chemical, that's whsat you have to do.

If you have a mature Lunar infrastructure, you severely weaken the link between fuel cost and launch cost. If buying fuel costs, say, $100/kg in LEO, that's only $100 million for 1000 tonnes of fuel.You can afford to splurge on fuel to get a much quicker trip.

Impaler · 2012-06-14 02:40:32

Just thought of a good way to deliver and assemble large Habitat cylinders to the Lunar surface. Start with a cylindrical pressure vessel habitat (ISS Destiny/Columbus derived), and attach by standard docking mechanisms to each end two identical cylindrical boosters that I'll call end-cap boosters. These boosters are different from any yet conceived in that they fire perpendicular to the axis of their cylindrical shape with their engines fully inside the cylindrical form. During launch a semi-circular shroud covers the engine nozzles and the boosters is outwardly indistinguishable from a conventional cylinder section. With a nosecone attached to one booster the whole assemblage can go on the end of a HLV without any airo-shell.

The end-caps are launch wet and during launch the thy act as the 3rd stage after separation of the second stage at high altitude the payload pitches over on its side and the twin boosters fire simultaneously and perpendicularly to the original axis at launch. Using HydroLox and with a 50% total mass fraction the boosters can impart 3 km/s Delta-V performing a significant amount of the LEO launch and allowing a very large final payload. The end-caps remain attached after reaching LEO and the whole vehicle is picked up by a tug craft presumably using (you guessed it) SEP to reach EML1. Their the end-cap boosters are refueled from a propellant depot, each being fueled separately from ports on their exposed end faces. Descent to the Lunar surface is now conducted with a generous 3 km/s Delta V capability, each end-cap boosters deploys 4 curved landing legs that are stowed in wrap-around fashion under the shrouds so as not to impinge the internal tank volume.

Upon landing the payload is now suspended above the lunar surface and supported on each end by the end-cap boosters legs, these legs have wheels that are aligned with the long axis of the payload so astronauts can tow the complete package the final distance to the habitat assembly area. Their the astronauts would position the hab adjacent to existing modules and employ simple ratchet releases in the legs to lower the payload to the surface. The end-caps would be disconnected and then ratcheted back up on their legs and towed away. Now the Hab cylinder would be rolled into contact with other hab elements by use of pulley and cable.

The end-cap boosters can be collected and attached to each other to create a tank farm or being wheeled they could be towed by any hypothetical lunar surface vehicle as a tanker for extending range or for collecting volatiles from mobile ISRU (assuming 'combine' style scoop-bake-dump). Alternatively they might be attached again in pairs to a crew cabin and become reusable landers.

So far all the lander concepts I've seen put payloads on the top of a large vertical LEM style lander (Altair) which leaves no way to get this safely off the lander and transport it too a construction sight. The one horizontal lander concept I've seen is another disposable LEM variant which again can't deposit any significant habitat structure that would allow for a descent base building effort comparable to ISS. The detachable end-cap configuration is the best way to deposit a cylindrical payload on its side while moving and handling it in a way that will be safe and logistically feasible, while the use of the end-caps as 3rd stages and the EML1 refuel trick is a bit more speculative it is a very effective means to boost payload delivery.

Assuming a 20 mt payload and end-cap boosters that dry mass 5 mt each and hold 15 mt each propellent (a very conservative 25% dry mass fraction), that's a total of 60 mt wet which should easily fly on a F9H despite being a few tons over the 53 ton max because it can contribute that 3rd stage Delta-V. Another F9H for fuel delivery to EML1 and the cost to put the Hab payload on the Moon is just couple hundred million.

Terraformer · 2012-06-14 04:27:40

The ULA study had horizontal landers. Have you seen it? The link should be somewhere in this thread.

Impaler · 2012-06-14 15:20:43

Terraformer wrote:

Impaler, you're forgetting the impact of Lunar fuel in orbit and EML1. For a Hohmann, yes, it wouldn't make sense to use fuel to travel to EML11, refuel, and then launch, but if you want a fast transfer using chemical, that's whsat you have to do.
If you have a mature Lunar infrastructure, you severely weaken the link between fuel cost and launch cost. If buying fuel costs, say, $100/kg in LEO, that's only $100 million for 1000 tonnes of fuel.You can afford to splurge on fuel to get a much quicker trip.

Lunar fuel if it ever even happens would only ever be used for lunar assent of crew and sample returns. With SEP tugs and the maturation of LEO launch systems that are going to be necessary for any kind of robust return to the moon the Earth-sourced propellant is going to be cheaper then lunar-sourced propellant at EML1 and probably down to LLO too. Propellant doesn't care about slow transit times and we will have solved the boil-off problem or else the whole idea is moot.

The correct economic math to determine if Lunar produced propellant is worth utilizing at any point between the Earth and Moon is to compare the kg of mass that must be sent past that point to ultimately produce the propellant on the moon vs how many kg of propellant ultimately return to said point. In addition the cost of making the ISRU equipment should also be considered as its cost multiplier over propellant at the reference point acts as another multiplier that lunar production must overcome. Propellant on the Earths surface is incredibly cheap ware as robotic equipment is already expensive, of course as you blast each one into space the cost for each increases but the increase is proportional to mass so equipment retains its original absolute cost premium while falling in relative cost. Example: Surface cost per kg Propellant $1, Equipment $100 LEO cost per kg Propellant $101, Equipment $200.

Now look at the mass losses that must occur due to HydroLox ISP limits, to get the ISRU equipment down to the Moon from EML1 you lose half the mass (3 km/s Delta V to allow for precision landings, RCS docking etc). Then to get the Propellant back to EML1 you must use a reusable tanker vehicle for their to be any viable system and even a very optimistic 10% dry mass vehicle will burn 2 units of propellent for every 1 delivered to EML1 because it needs to retain some fuel for descent to the moon. Together the 2:1 and 3:1 ratios mean the ISRU equipment needs to produce 6 times its mass in Propellant to just break even on mass and that's assuming unlimited re-use of the tanker so that it amortizes away completely.

Now multiply 6 by the expected cost multiplier of equipment at EML1 vs Propellant, this is a much fizzier number but the cheaper the launch cost to LEO the worse it gets, the earlier example of equipment that costs only $100 kg is exceedingly generous. In all likely-hood the cost differential will grow from. I'd estimate that right now the multiplier is between 2 and 10 and as launch costs improve it would rise to between 10 and 50. So add it all up and you end up with the challenge of making ISRU equipment that will produce over it's lifespan at minimum 60 times it's mass in Hydrolox propellant in-order for it to be a viable supplier to EML1. That's a big engineering challenge and I doubt it will ever be viable given that the lunar volatiles are are probably thinly dispersed and a witches-brew of many different elements that would require a considerable amount of processing.

Of course moving the point of usage to the lunar surface say for an assent or for just plain habitat consumption that drops the factor of 6 and the task becomes that much easier to accomplish. That's why I think we will only ever see lunar produced propellant or any volatile derived their used anywhere but the lunar surface.

Impaler · 2012-06-14 15:28:33

Here is the SpaceWorks study, the EML1 departure table is down on page 50, theirs also an mp3 of a presentation of the same study but it mostly discusses issues like propellant boil-off, mass fraction not the departure trajectories.

http://spirit.as.utexas.edu/~fiso/telec … r_5-16-12/

Last edited by Impaler (2012-06-14 15:40:20)

Terraformer · 2012-06-18 07:17:17

Eh? I wouldn't call 2m thick sheets of ice "thinly dispersed"... And processing water isn't that hard; just melt it, and the more volatile components will outgas and the dust will settle to the bottom of the cylinder.

I don't know exactly how much the equipment will mass, but given the simplicity of the equipment (electrolysis is not that complex...) and the masses that are involved in Mars Direct, producing a few dozen times the mass of the equipment in fuel is no problem.

A lot of your objections have already been covered by Hop and others in this thread...

Can we get back to talking about the Lunar infrastructure now?

RGClark · 2012-06-18 19:07:39

RobS wrote:

...
Lagrange points are useful if you are spiralling a lot of stuff slowly out of the Earth's gravitational well. They provide a region to park things. I wouldn't use solar ion propulsion if the latter uses xenon; currently xenon is worth more per ounce than gold. The world only produces a dozen tonnes of it per year. On the other hand, if one switches to argon, which is cheap, that problem is solved. Currently xenon is a practical propellant only when the launch cost to LEO is many thousands of dollars per pound. I'd favor development of solar thermal anyway; the technology has been tested on the ground, its thrust (up to 100 pounds currently) is much more than ion, the thrust can be "stored" up and used in a series of perigee kicks, and hydrogen boiloff isn't an issue because you're using your hydrogen supply for a few minutes every few hours.
...

I read that report cited by the supporters of asteroid mining about moving a 500 mT asteroid to lunar orbit:

Asteroid Retrieval Feasibility Study.
2 April 2012
http://kiss.caltech.edu/study/asteroid/ … report.pdf

The news report were all about the solar electric ion propulsion used and that perhaps it would cost $2.6 billion to develop. But I was surprised in the report that it also discussed doing it with LH2/LOX chemical propulsion and how little propellant it would use. First it notes that an asteroid such as 2008HU4 at closest approach would require only a 170 m/s (!) delta-v to bring it to lunar orbit. Then in figure 19 on p. 43 is given a comparison between the propellant required for LH2/LO2, N204/MMH, and SEP propulsion for this asteroid at an assumed 1,000 mT mass. Surprisingly, for the hydrogen case it is less then 40 mT for a 1,000 mT payload! This is because of course it is only a 170 m/s delta-v. But this means for the 500 mT case that was cited in the news reports it would be less than 20 mT propellant load, and a LH2/LOX propulsion stage this size is already available in the Centaur.

Since the chemical propulsion would have greater thrust, the mission return time would also be significantly less than the 10 years for the SEP propulsion. The problem though is that you would need to get the return propellant from the asteroid. The report proposes using a carbonaceous asteroid which would have abundant H2O likely in chemically bound form, perhaps as much as 20%. So the question is are there simple methods known from ISRU studies to release H2O from say the carbonates and clays likely to be in carbonaceous asteroids?

You could enclose the ca. 7 meter diameter in a shroud and using a solar furnace either with a mirror or fresnel lens to release the volatiles, such as water, CO2 and methane. I'm thinking you could do this at low enough temperature that only the volatiles are released, and the solid material would not be vaporized. Then you would use the highest temperatures possible with a solar furnace, ca. 3,500 degress C:

Solar Furnace.
http://en.wikipedia.org/wiki/Solar_furnace

to do electrolysis to separate the hydrogen and oxygen. Another complication is that you would need cryogenic chilling equipment to liquefy the hydrogen and oxygen for the RL-10 engines on the Centaur. Probably you would also need to separate out the other volatiles such as the CO2 and methane, another complication.

Anyone know of research on this question of heating of water bearing minerals to remove the water all the way to the step of turning it separately into hydrogen and oxygen?

Another possibility if you are going to use a solar furnace is to use solar thermal propulsion. One particularly simple implementation of this would be to simply vaporize the asteroidal material without separating it out to say hydrogen as the reaction mass. At maximal temperatures of 3,500 C you would still get pretty good Isp even using high molecular weight materials. A disadvantage of this is that you would also lose some proportion of the valuable minerals.

Any of these possibilities I would estimate would cost far less than a billion dollars however.

Bob Clark

Impaler · 2012-06-19 03:08:49

Terraformer wrote:

Eh? I wouldn't call 2m thick sheets of ice "thinly dispersed"... And processing water isn't that hard; just melt it, and the more volatile components will outgas and the dust will settle to the bottom of the cylinder.
I don't know exactly how much the equipment will mass, but given the simplicity of the equipment (electrolysis is not that complex...) and the masses that are involved in Mars Direct, producing a few dozen times the mass of the equipment in fuel is no problem.
A lot of your objections have already been covered by Hop and others in this thread...
Can we get back to talking about the Lunar infrastructure now?

Their are almost certainly no 2m thick sheets of water-ice in the lunar craters. Your quoting people who are making the maximum possible extrapolations of the Chandara satellite result which were made before other Probes made more exact measurements. Chandara used Polarized light cameras, polarized light can't distinguish between rough surfaces and ice and I suspect it doesn't distinguish water ice from the many other ices that could exist in those conditions. The LCROSS findings while confirming water ice at ~5% concentration definitively ruled out massive pure ice sheets.

http://adsabs.harvard.edu/abs/2011JGRE..11601005N

AboveLink wrote:

On 9 October 2009 the Lunar Crater Observation and Sensing Satellite (LCROSS) impacted Cabeus crater, located near the south pole of the Moon. Prior to that impact, the Mini-RF instruments on ISRO's Chandrayaan-1 and NASA's Lunar Reconnaissance Orbiter (LRO) obtained S band (12.6 cm) synthetic aperture radar images of the impact site at 150 and 30 m resolution, respectively. These observations show that the floor of Cabeus has a circular polarization ratio (CPR) comparable to or less than the average of nearby terrain in the southern lunar highlands. Furthermore, <2% of the pixels in Cabeus crater have CPR values greater than unity. This observation is not consistent with the presence of thick deposits of nearly pure water ice within a few meters of the lunar surface, but it does not rule out the presence of small (<˜10 cm), discrete pieces of ice mixed in with the regolith. In addition, Mini-RF on LRO acquired a postimpact S band image of the region surrounding the LCROSS impact site, providing important geologic context for the site. Registering the LRO image to a near-infrared (NIR) image taken by the LCROSS shepherding spacecraft, we find that the impactor landed in the ray of a fresh, radar-bright, 1 km crater. However, the difference between preimpact and postimpact images is not above the speckle noise. This implies that the size of the LCROSS impact crater is less than Mini-RF's resolution (30 m), and/or that the impact did not excavate more decimeter-size blocks than were already present at the impact site.

All logic and evidence point to some kind of dust-hoarfrost mixture containing every conceivable chemical that would volatilize and get cold-traped on the Moon, including things like Ammonia, Carbon-dioxide and Mercury (it Bakes out of the Lunar Regolith but the vapor lacks escape velocity and gets caught in the cold-trap craters). Their is really no way that water-ice gets deposited without accompanying compounds.

We don't know nearly enough yet to characterize how feasible it would be to mine and purify these deposits, at 5% water content you would need to move and bake 60 tons of regolith to deliver 1 ton to EML1, the energy requirements are going to be huge just for the hauling. Electrolysis of water is also notoriously inefficient when done at low temperature, high temperature steam systems are thought to have potential but will be much more complex. We also don't know how much the regolith will need to be heated to get the water out as the regolith minerals might simply hydrate up all the melting water requiring a very high temperature bake to dehydrate them. Before anyone goes and pins their hopes lunar ice we will need to at least characterize the exact nature and size of these deposits using robotic rovers doing direct-measurements, and possibly sample returns as well.

Last edited by Impaler (2012-06-19 03:29:18)

RGClark · 2012-06-19 10:09:46

Impaler wrote:

Here is the SpaceWorks study, the EML1 departure table is down on page 50, theirs also an mp3 of a presentation of the same study but it mostly discusses issues like propellant boil-off, mass fraction not the departure trajectories.
http://spirit.as.utexas.edu/~fiso/telec … r_5-16-12/

Thanks for that. It's discussed on pages 41,49, and 50. It is able to reduce a 4,400 m/s delta-v to 1,770 m/s this way, a reduction of 2,600 m/s.

I wonder if you can reduce higher delta-v trajectories also. For instance in this post I discussed reducing the outbound travel time to 70 days by using a ca. 8,800 m/s delta-v:

Re: Human missions » Developing the cis-Lunar economy and infrastructure » 2012-06-11 13:56:35.
http://newmars.com/forums/viewtopic.php … 55#p113455

If you could reduce the delta-v down to to 6,100 m/s from 8,800 m/s, you could increase the payload in that calculation I gave by a factor of 2.5 from 6 mT to 15 mT.

Bob Clark

Impaler · 2012-06-19 16:04:42

It would seem to me that any arbitrarily high Earth escape velocity could be achieved after an Earth sling-shoot trajectory by simply extending the final Earth Perigee burn to accelerate away from the Earth. I'm not sure that 2600 m/s can just be dropped from any arbitrary LEO escape DV but it sounds reasonable.

2600 m/s is right about the DV for achieving GTO from LEO. I guess when you 'fall' from EML1 to do the slingshots the orbit approximates a highly elliptical GTO orbit and you effectively get back the 2600 m/s worth of work that would have gone into creating a GTO orbit. But the DV for EML1 from LEO is a full 3770 m/s using high thrust propulsion so compared to a direct departure from LEO the two legs of LEO-EML1 and EML1-TMI is 1100 m/s high DV.

So again EML1 (at any desired departure speed) only makes sense if your doing it with some non-chemical High ISP propulsion system that can save propellant mass fraction. If your using all chemical propulsion the best solution would just be to stage the vehicle in GTO and then make TMI at perigee. At 450 ISP you would need 45% propellant fraction to do a GTO insertion from LEO, so any high ISP system which can achieve GTO at lower then 45% mass fraction (fuel and system combined) would be able to reduce IMLEO for any given size payload.

GTO would be a bad place to try to assemble a vehicle though, its going through the Van-Alan belts constantly and your launch windows would be instant in-order for the launch to match the orbit of the target. Perhaps a better alternative is Earth-Sun L2 (ESL2), it doesn't have the Van-Alan radiation issue (though it has to deal with CME radiation the vehicle will have that after TMI too so not anything special). From what I've read the DV for reaching ESL2 is not that much high then EML1, but I don't have any idea how much DV you need for TMI out their or if an Earth sling-shoot is possible or desirable.

RGClark · 2012-09-01 15:32:04

I was trying to get a lower roundtrip delta-V for lunar missions by flying directly to the lunar surface rather than going first into lunar orbit then descending. Here's a list of delta-V's of the Earth/Moon system:

Delta-V budget.
Earth–Moon space.

http://en.wikipedia.org/wiki/Delta-v_bu … Moon_space

If you add up the delta-V's from LEO to LLO, 4,040 m/s, then to the lunar surface, 1,870 m/s, then back to LEO, 2,740 m/s, you get 8,650 m/s, with aerobraking on the return.
I wanted to reduce the 4,040 m/s + 1,870 m/s = 5,910 m/s for the trip to the Moon. The idea was to do a trans lunar injection at 3,150 m/s towards the Moon then cancel out the speed the vehicle picks up by the Moons gravity. This would be the escape velocity for the Moon at 2,400 m/s. Then the total would be 5,550 m/s. This is a saving of 360 m/s. This brings the roundtrip delta-V down to 8,290 m/s.
I had a question though if the relative velocity of the Moon around the Earth might add to this amount. But the book The Rocket Company, a fictional account of the private development of a reusable launch vehicle written by rocket engineers, gives the same amount for the "direct descent" delta-V to the Moon 18,200 feet/sec, 5,550 m/s:

The Rocket Company.
http://books.google.com/books?id=ku3sBb … -V&f=false

Bob Clark

SpaceNut · 2018-08-21 17:38:27

I think that the water ice in the shadows of the craters have been answered and that it will lead the way once we are capable of landing on the moons surface...

SpaceNut · 2019-01-01 15:56:30

So after getting a minimal crew to the moon following 3 falcon 9 heavy launches;

what is the trade after setting up a base station toe / foot hold to stay?

How do we make the revenue to be able to keep building with the core concept of using the Falcon 9 Heavy?

What are the key insitu mining operations after providing fuel and air for the crews that earth would want and man can use even on the moon?

Are there any limits on some mined resources and uses?

Mars_B4_Moon · 2022-01-04 12:39:55

I think Mars is more important than Moon missions however I don't ignore the Moon. Seems like Ion Drives can be used for slow non essential cargo just like we use Ships on Earth to Deliver Cargo across the world because its more efficient than flying the stuff by jet.

Could some from of levitation and then thrusters become the main propulsion engine for heavy cargo and unmanned A.I piloted space vehicles, could you use a hybrid of tech or could you even one day put tourist people inside Lunar Cruise Ships to travel inside these new machines?

The Chinese also have plans for Nuclear reactors in Space and Ion Power, the Russians might have plans for a Space-Tug
http://newmars.com/forums/viewtopic.php?id=9893

So it seems that perhaps the Back-To-The-Future style Hoverboard is not so silly afterall and finally might become a real thing on the Moon?

The cislunar space is a useful label for "the volume between geostationary orbit and the moon's orbit" where beyond cislunar space lies translunar space your paths to Mars, to Venus, the Sun or going far out to Jupiter or on your long way on some extra solar robotic mission to Alpha Centauri. Some people think to regard the Lagrange points L4 and L5, the stable regions of the Moon's Trojan points, as CisLunar Space points. Ion-drives might consume 1 kW of power, they can be 80% efficent and have exhaust velocities around 20–50 km/s as Ion Drives improve they can get power from solar or nuclear or whatever power source might become part of every spacecraft system, however Ion thrust engines are only useable in no air or practical only in the vacuum of space. The recent SpaceX launched DoubleAsteroid Test (DART) with insutruments from NASA, the ESA and Japanese is now once again using an Ion-Drive, VASIMR is an electrothermal thruster uses radio waves to ionize and heat an inert propellant but it also has a lot of issues to solve. MPD thrusters could produce extremely high specific impulses (Isp) with a very high exhaust velocity triple the value of current xenon-based ion thrusters. Perhaps now as we see technology develop, new ion drives, reports about a saucer type spacecraft/ aircraft designed for the Moon, 3-D printers that might 'Print' materials on the surface or combining these technologies with other tech and chemistry ideas such as using Lunar metals or thermite reaction to combine with oxygen and Lunar dust but virtualy all pure metals will burn in the right conditions, a New Engine would have to be Designed to deal with this new fuel but perhaps a way could be found to improve ISP and have more efficent thrust/weight ratio maybe helped by 'levitation' or 'Ion Drives' adding another hybrid way of getting materials and people and astronauts into Orbit?
From NASA's website
'The MPD's ability to efficiently convert megawatts of electric power into thrust makes this technology a prime candidate for economical delivery of lunar and Mars cargo, outer planet rendezvous, and sample return, and for enabling other bold new ventures in deep space robotic and piloted planetary exploration'
https://www.nasa.gov/centers/glenn/about/fs22grc.html
MPD

Future Uses
Future power-rich robotic and piloted outer planet missions will require exhaust velocities approaching 100,000 meters per second (over 200,000 mph). These higher velocities can be achieved using noncondensable hydrogen plasmas, which are currently under investigation at NASA Glenn. As research continues, the efficiency of the MPD thruster will be increased, which will allow missions with reduced propellant requirements or increased range. Higher exhaust velocities and thrust levels will lead to shorter trip times and reduced mission cost, which is especially beneficial for cargo and piloted missions. As large amounts of power become available in space, MPD thrusters may become the method of propulsion that carries humans to other planets in our solar system.

I was reading news again on that Saucer idea with flying craft using the Moon's natural charge and reading up on Ion-drives and wondered if some kind of hybrid technology craft could be used to escape the Moon and save on fuel. So I though I might re-post news here about a real-life flying saucer shaped thing that is expected to be on the Lunar surface one day. The escape velocity or escape speed of another body can be much smaller, a man or astronaut might be able to jump off an asteroid, the G "universal gravitational constant" M = mass of the planetoid or object or moon (kg) R = radius of the planet or moon (m). Mars is going to be different to the Moon, sometimes more difficult and other times more easy depending on what situatuion is faced, Mars has a 24 and half hour day, it has an atmosphere, it has a surface gravity more than a third of Earth, it is 0.107 the Mass of Earth, a little over half the radius 0.53 times the radius of Earth, the Escape Velocity on Mars is 5,027 m/s. Its value far less is Moon has no atmosphere it has less mass, therefore escape velocity from the moon is almost 2400 m/s if I remember correct, or just under 2.4 km/s a high powered gun there's a slim chance the bullet could go all the way around and hit you in the back of the head. The news I read was

'Scientists are testing a new concept for a hovering rover that LEVITATES by harnessing the moon's natural charge.'

https://www.dailymail.co.uk/sciencetech … harge.html

Massachusetts Institute of Technology have designed a 'flying saucer' that can levitate on the moon, asteroids, and other airless planetary surfaces.

https://www.wionews.com/science/researc … ids-441035

Hoverboards Might Actually Work...but Only On the Moon. The levitating vehicles of our dreams might actually be possible, just not on Earth.

https://gizmodo.com/these-hoverboards-m … 1848252112

Many years back on this forum there were people who did not really like the ideas of Ion-Drives they said it lacked power, I think there were very few missions using them. NASA had just started using Dawn, Russians had used the in Satellites, the ESA began designs Smart-1 and Japan was building something but it was still a 'fringe' science idea to use them to colonize the solar system. The ideas like using a Train or Ion Drive to achive escape velocity or 'selenocentric' Orbit were dismissed as science fiction or ridiculous. I wonder now if that technology has become more feasible?

One of the reasons I am against wasting too much time on the Moon is I believe a lot of Mars designs do not apply to the Moon and I also believe a lot of tech that will be used on the Lunar surface does not apply to building the first Martian village.

Last edited by Mars_B4_Moon (2022-01-04 12:47:27)

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