SpaceX Dragon spacecraft for low cost trips to the Moon.

ARD · 2012-02-06 19:15:50

JoshNH4H wrote:

Sounds basically feasible, though real development work would have to be done before it would be possible to comment on whether the engineering would actually work out. The one thing that I'm not comfortable with is where you're going to get the oxygen from. It's probably not frozen in the craters, and electrolyzing your way to it would be difficult. I suppose you could use a TiO2 catalyst to photodissociate water in open sunlight, but that will only be about 6% efficient and possibly mass-intensive.

Heat up the regolith until it starts breaking up into oxygen gas and a metal slag? The unit to do that might be heavy, but one can theoretically use direct solar heating to get it done, and one of the 'waste' products is iron.

JoshNH4H · 2012-02-06 19:23:57

Well, 'heavy' is a very important factor here. I think photodissociation with a TiO2 catalyst would probably be more mass effective. It might be hard to get a lot of throughput with this method, but the same is true with all forms of solar power.

GW Johnson · 2012-02-07 22:36:29

Back to using Falcons and Dragons to go to the moon. You will need a lander of some tonnage, probably not unlike the old Apollo lander. You might even build it out of a Dragon with extra tanks, or maybe a from-scratch design. Whatever.

But, I don't see why you need a Centaur or any other departure stage. You don't need any more engines, you just need delta-vee. The Dragon will have the new Super-Draco thrusters on it with 120,000 lb of axial thrust, according to their website. There's eight of them, plenty of redundancy there. Just add a big dumb propellant tank, and plumb it up to the Dragon's system. Use the Dracos for all the delta-vee from LEO to lunar orbit, and back, sucking from the big dumb tank.

Launch the lander on one Falcon-Heavy, launch the Dragon and the big dumb tank on the other. Rendezvous in LEO, and dock the big dumb tank to the rear end of the Dragon, and the lander to its nose. Make up your plumbing connections. Then go to the moon.

Lots of minor details to work out, but I don't see any show stoppers here.

GW

Last edited by GW Johnson (2012-02-07 22:38:07)

Rune · 2012-02-07 23:11:25

I think they mean to eventually refuel these H2/LOX stages at depots at either LLO, EML1, or LEO, supplied from the moon. Or all of the above. But since what we are discussing here is how to build the infrastructure to actually mine and process those fuels, I coincide that that is looking at the problem the wrong way. Since the fuel is going to be lifted form Earth until it is produced elsewhere, it only makes sense to use storable propellants, at first. Dense hypergolics while you are at it to improve simplicity and reliability are nice. The isp advantage is weighted against the duplicity in propulsion systems, the heavier cryo tanking, and general boiloff issues to call a real, solid, "worth-the-extra-complexity" delta-v advantage at this stage, methinks. Note that no commercial spacecraft (or otherwise, I think) intended to operate in space for prolonged periods of time and numerous burns uses H2/LOX propulsion.

Hell, skip the whole upper stage completely and launch stuff directly to TLI with the Falcon Heavy. It is basically a three-stage vehicle anyway, so the throw weight to TLI should be around half than the one to LEO, give or take a few tons. And if Falcon 9 is safe enough to put a crewed Dragon on top, Falcon Heavy should be, too. Or a lander. Or a return propellant tank to put the Dragon on its way back home, if you can't pack it in the crewed launch. A crewed Dragon would have the capability to dock with all this stuff at LLO (it'll have active docking systems to go to the ISS), so lunar orbit rendezvous is certainly feasible. Orbits don't decay as much at the moon, so the components can wait there for the crewed launch, and a preliminary launch of commsats can give you the kind of sensor environment to be very sure of your dockings. What is that, three Falcon Heavies per mission at most, and the only development required a ~20mT lander? Practically loose change for governments. Very hard for private investors, if all they get out of it is a few boots on the moon (certainly less than seven pairs, in a 20mT lander) and a few rocks back.

Rune. And that would be the absolutely minimum hardware to develop to return to the moon. Decent MCPs suit would be cool to have, though, for a few million extra bucks and the extra publicity of superhero-like astronauts.

Last edited by Rune (2012-02-07 23:20:27)

GW Johnson · 2012-02-08 10:54:20

Now you're thinking outside the box. Good show, Rune.

GW

Rune · 2012-02-08 11:18:42

SpaceX would be stupid if they didn't certify Dragon on top of a Heavy, at the very least for free return cislunar flybys. Especially when Space Adventures has already sold one seat and is very close to selling another for their cislunar flyby mission using the old soyuz plan. For the price of a single of those seats ($150 million), you can pay for the launch costs of a heavy, plus some spare change. The other 6 should be enough to pay for the Dragon and lower the ticket price a bit, and still leave a healthier benefit margin than the one Space Adventures is convinced it'll make. To say something.

To refresh some minds, I recall the Space Adventures architecture gets the crew form the ISS, mount them in a dedicated soyuz, attached to a Fregat derivative that puts the whole stack into a free return lunar flyby trajectory. 3 Soyuz launches required, plus orbital docking and use of the ISS, for just a crew of three. Hence 150 million for each seat (there's a pilot involved to operate the craft). It kind of explains why I believe the manufacturing cost of a Soyuz is around 50 million, but when it gets to the market with a payload, that puts the price closer to 100 million. NASA, of course, gets asked as much money as they will pay for astronaut seats. They got that part of capitalism right.

Rune. Oh, and thanks, I'm actually blushing a bit.

Last edited by Rune (2012-02-08 12:41:18)

JoshNH4H · 2012-02-08 20:55:13

My only potential qualm with launching straight off a Falcon heavy is the lower amount of mass on TLI as compared to adding another stage on top. I'm not sure if the extra payload will be needed, but it's always nice to have more mass if you can.

Also, just wondering, how did you get to the figure that payload would be half as much? TLI is between 3 and 3.25 km/s, and adding this onto a rocket with a delta-V of 9.5 km/s or less would lead me to believe that the payload would be significantly less than half. I know that there isn't really readily available information to plug into the rocket equation for the Dragon rockets, but did you base this on, say, the Saturn V or is it more of a guess? For comparison, the Saturn V's TLI payload was 38% of its LEO payload (45 tonnes vs 119 tonnes).

Rune · 2012-02-08 21:24:13

JoshNH4H wrote:

My only potential qualm with launching straight off a Falcon heavy is the lower amount of mass on TLI as compared to adding another stage on top. I'm not sure if the extra payload will be needed, but it's always nice to have more mass if you can.
Also, just wondering, how did you get to the figure that payload would be half as much? TLI is between 3 and 3.25 km/s, and adding this onto a rocket with a delta-V of 9.5 km/s or less would lead me to believe that the payload would be significantly less than half. I know that there isn't really readily available information to plug into the rocket equation for the Dragon rockets, but did you base this on, say, the Saturn V or is it more of a guess? For comparison, the Saturn V's TLI payload was 38% of its LEO payload (45 tonnes vs 119 tonnes).

I kind of half-remembered the old Saturn's numbers, and guessed I had remembered them optimistically. Still, 20mT out of 53mT (37.73%), I'd say my guess was pretty close to the mark, right? (Though the Saturn has a more efficient upper stage and this changes things somewhat, in Saturn V's favor). I also half-remembered an Apollo LM was less than that (14mT, turns out), so a similar lander would be easier to build.

And yeah, an upper stage of course would increase thrown weight. As well as development costs, and complexity, and eventually the cost of each launch. You just have to run the budget numbers (the very detailed, real budget numbers, with salaries and taxes and such included) to pick between them. Rough guess again, in the beginning you want to avoid large development costs. Until you have the depots in place and can refuel from of-earth sources, for example. That would be a very good time for refuelable, advanced upper (more like, in-space) stages to come available. Of course, for that, testing has to come first, as the first missions are launched to establish the refueling infrastructure.

Rune. Plus, I was going for cheapest, fastest and dirtiest. I think I won at that. ;P

Last edited by Rune (2012-02-08 21:25:20)

JoshNH4H · 2012-02-08 22:24:47

Your design was definitely a winner in terms of cheap and dirty, that's for sure. I just have serious questions about capability. I think an EDS would be necessary. It doesn't necessarily have to be super-complicated. For example, strapping a Falcon 1 upper stage onto the payload would be a perfectly serviceable EDS. The total mass of the F1 rocket at liftoff is 35 tonnes, so I would imagine that a Falcon 1 upper stage would be pretty small (5-10 tonnes?). Minimal development, still pretty quick and dirty, higher payload. Capable of Moon landing, too, in all likelihood.

Rune · 2012-02-09 11:18:24

What do you know, the wiki had the answer all along: wiki's page on the Flacon Heavy.

16mT to TLI, or in other words, just enough to fit a LEM in there with a couple of tons to spare (let's call that shroud/adapter). Or a fully loaded Dragon with an extra 6mT of propellant in the trunk. Or, you know, variations. With an Isp of 275, it would take the Super Dracos a mass ratio of 1.61 to give the required 1.31km/sec of TEI. Almost a perfect match, that, if the dragon's own orbital maneuvering propellant is enough to insert into LLO and maneuver there (it should), and you could always take out some cargo (I'm assuming a fully loaded crewed Dragon takes all of a Falcon 9's payload weight).

I call it doable in just two Falcon Heavies with no upper stage. The only development required to "carry on with Apollo where we left it" is a simple fuel tank in the Dragon's trunk plumbed to the main fuel system and a "LEM 2.0". And there were some cool missions planned. Let's not forget that there were versions of the LEM designed to carry cargo to the lunar surface (just the descent stage, about 5mT to the surface), and those missions would require only a single launch.

Rune. Well, I was damn close, what can I say.

Last edited by Rune (2012-02-09 11:44:48)

JoshNH4H · 2012-02-09 13:37:23

Haha, well fair enough. From TLI, though, it's still a further ~3 km/s to the lunar surface. Two falcon heavies is 32 tonnes to TLI, as compared with 45 tonnes for the Saturn V. It's not quite clear to me where all of the delta-V is coming from.

Just eyeballing the numbers, though, it looks like it might be possible to get a Dragon to the lunar surface and back if the second launch is just fuel. To get a more capable mission, though, I would think that we would want to do a third launch- which, incidentally, would give us 48 tonnes, right around the Saturn V's TLI payload. With modern engineering, we're looking at a mission significantly more capable than the Apollo missions (The longest of which was on the Moon for just a tiny bit over 3 days. Let's be honest here, for the Apollo missions science and laying foundations for colonization took a tremendous backseat to walking around and sticking our tongues out at the Soviets.). I bet we could actually get a very nice missions going with 3 Falcon Heavies, and a more Apollo-Style one going with two.

GW Johnson · 2012-02-09 17:01:21

I had the weight statements and delta-vee figures worked out for Dragon with a "modest dumb tank" inside the unpressurized cargo space. I had 2.3 km/sec total delta-vee that way. You could do a bigger tank, it just needs to dock somehow to the rear. These would be minor mods to the existing design.

The figures I caloculated are in figure 11 of the posting-version of my convention paper. It's over at http://exrocketman.blogspot.com, dated 7-25-11. The by-date navigation tool on the left can take you right to it. This stuff was reverse-engineered from the data posted on Spacex's website.

I haven't done anything for Falcon-Heavy yet, but Falcon-9 I reverse-engineered into a weight statement and some delta-vee estimates in the posting on "exrocketman" dated 12-14-11. It's the first figure in the article. That's the posting on re-usability in launch rockets. Falcon-Heavy is a Falcon-9 with two extra first stages strapped on. The key to their design is propellant cross-feeding among the three 9-engine units. The center one is still nearly full when they drop the outer two off, yet all 27 engines have been burning.

The data I calculated for Dragon should be pretty realistic, unless what I found does not take into account the cosine correction for the canted Draco thrusters. I'm not sure on that one.

What I did for Falcon-9 should be really close.

GW

Rune · 2012-02-10 05:47:52

JoshNH4H wrote:

Haha, well fair enough. From TLI, though, it's still a further ~3 km/s to the lunar surface. Two falcon heavies is 32 tonnes to TLI, as compared with 45 tonnes for the Saturn V. It's not quite clear to me where all of the delta-V is coming from.

Well, the CM+SM combo clocked in at a bit over 30 tonnes and our lunar Dragon is restricted to 16 by the launcher, so there's the source of discrepancy. Part of that comes from the SPS (the hypergolic main engine) being actually sized to lift a much bigger CM straight of the lunar surface (that was the original plan, direct return), part of that is the older, heavier tech, part of that was the fuel requirement (2,800m/s delta-v on it's own) to brake the whole stack, LM included, into LLO.

Of course, I think it's a bit optimistic to think you could get a 7 crew, 10mT lunar Dragon that can actually return to earth and keep them alive for two weeks under 16mT. But, with a lander a bit heavier than the old LM that brakes into orbit itself, and a much reduced crew, you could get 2-3 crew members to the lunar surface easy, I think.

GW Johnson wrote:

The data I calculated for Dragon should be pretty realistic, unless what I found does not take into account the cosine correction for the canted Draco thrusters. I'm not sure on that one.

I don't think anyone is sure on that one. Not enough data to do anything but rough guesses. Plus, orbital firings could be done with the smaller, not canted attitude Dracos, for a penalty on Oberth savings. And even their isp is, at this point, anyone's guess. But there is enough margin form an empty Dragon to a full one, that it seems possible with a reduced crew.

Rune. It's a mouthful of acronyms, speaking about Apollo. Look at that first paragraph!

RGClark · 2012-02-12 02:17:23

Just saw this discussed on Nasaspaceflight.com

Elon Musk on SpaceX’s Reusable Rocket Plans.
February 7, 2012 6:00 PM

The key, at least for the first stage, is the difference in speed. "It really comes down to what the staging Mach number would be," Musk says, referencing the speed the rocket would be traveling at separation. "For an expendable Falcon 9 rocket, that is around Mach 10. For a reusable Falcon 9, it is around Mach 6, depending on the mission." For the reusable version, the rocket must be traveling at a slower speed at separation because the burn must end early, preserving enough propellant to let the rocket fly back and land vertically. This also makes recovery easier because entry velocities are slower.
However, the slower speed also means that the upper stage of the Falcon rocket must supply more of the velocity needed to get to orbit, and that significantly reduces how much payload the rocket can lift into orbit. "The payload penalty for full and fast reusability versus an expendable version is roughly 40 percent," Musk says. "[But] propellant cost is less than 0.4 percent of the total flight cost. Even taking into account the payload reduction for reusability, the improvement is therefore theoretically over a hundred times."

http://www.popularmechanics.com/science … ns-6653023

Then for the Falcon 9, the payload would be reduced from 10 mT to 6 mT. If the reduction in payload really is this high, then maybe it would be better to recover the first stage at sea. The loss in payload is coming from the reduction in the speed of staging as well as the need to retain a portion of the fuel for the return to base. Recovering at sea would not have these disadvantages because you could let the first stage make its usual trajectory at returning to the sea but use just small amount of propellant for the final slowdown before the sea impact.
In this article Musk does mention that returning back to the launch point allows the turnaround time at least for the first stage to be just hours. But will we really need that short a turnaround time at this stage of the game? A turnaround time of a few days would seem to be sufficient.
Perhaps the idea that retrieval at sea would be so expensive comes from the experience of the shuttle with the SRB's. But these were quite large and heavy at ca. 90 mT dry compared to that of the Falcon 9 first stage at less than 15 mT. Also, it is well known the labor costs for the shuttle were greatly inflated compared to a privately funded program.
The only additional requirement is that you would need a cover that could be extended to cover the engine section and would be watertight.

Bob Clark

louis · 2012-02-12 05:15:05

RGClark wrote:

Just saw this discussed on Nasaspaceflight.com
Elon Musk on SpaceX’s Reusable Rocket Plans.
February 7, 2012 6:00 PM
The key, at least for the first stage, is the difference in speed. "It really comes down to what the staging Mach number would be," Musk says, referencing the speed the rocket would be traveling at separation. "For an expendable Falcon 9 rocket, that is around Mach 10. For a reusable Falcon 9, it is around Mach 6, depending on the mission." For the reusable version, the rocket must be traveling at a slower speed at separation because the burn must end early, preserving enough propellant to let the rocket fly back and land vertically. This also makes recovery easier because entry velocities are slower.
However, the slower speed also means that the upper stage of the Falcon rocket must supply more of the velocity needed to get to orbit, and that significantly reduces how much payload the rocket can lift into orbit. "The payload penalty for full and fast reusability versus an expendable version is roughly 40 percent," Musk says. "[But] propellant cost is less than 0.4 percent of the total flight cost. Even taking into account the payload reduction for reusability, the improvement is therefore theoretically over a hundred times."
http://www.popularmechanics.com/science … ns-6653023
Then for the Falcon 9, the payload would be reduced from 10 mT to 6 mT. If the reduction in payload really is this high, then maybe it would be better to recover the first stage at sea. The loss in payload is coming from the reduction in the speed of staging as well as the need to retain a portion of the fuel for the return to base. Recovering at sea would not have these disadvantages because you could let the first stage make its usual trajectory at returning to the sea but use just small amount of propellant for the final slowdown before the sea impact.
In this article Musk does mention that returning back to the launch point allows the turnaround time at least for the first stage to be just hours. But will we really need that short a turnaround time at this stage of the game? A turnaround time of a few days would seem to be sufficient.
Perhaps the idea that retrieval at sea would be so expensive comes from the experience of the shuttle with the SRB's. But these were quite large and heavy at ca. 90 mT dry compared to that of the Falcon 9 first stage at less than 15 mT. Also, it is well known the labor costs for the shuttle were greatly inflated compared to a privately funded program.
The only additional requirement is that you would need a cover that could be extended to cover the engine section and would be watertight.

Bob Clark

Well I can't enough of Musk personally...he seems to hit the nail bang centre on the head. He knows exactly what needs to be done to revolutionise space travel and he's doing it.

Bob, "time is money" as they say. I don't think he's worried about scheduling - it's the labour input that drives up the cost of recovery from sea. A return to a spaceport is what is required. Rest assured he will have done all the calculations. I think he's right.

6MT is still a huge amount to get up there.

I am presuming that for the Mars transit he must be thinking in terms of orbital assembly.

RGClark · 2012-02-12 17:59:57

louis wrote:

RGClark wrote:
Just saw this discussed on Nasaspaceflight.com
Elon Musk on SpaceX’s Reusable Rocket Plans.
February 7, 2012 6:00 PM
The key, at least for the first stage, is the difference in speed. "It really comes down to what the staging Mach number would be," Musk says, referencing the speed the rocket would be traveling at separation. "For an expendable Falcon 9 rocket, that is around Mach 10. For a reusable Falcon 9, it is around Mach 6, depending on the mission." For the reusable version, the rocket must be traveling at a slower speed at separation because the burn must end early, preserving enough propellant to let the rocket fly back and land vertically. This also makes recovery easier because entry velocities are slower.
However, the slower speed also means that the upper stage of the Falcon rocket must supply more of the velocity needed to get to orbit, and that significantly reduces how much payload the rocket can lift into orbit. "The payload penalty for full and fast reusability versus an expendable version is roughly 40 percent," Musk says. "[But] propellant cost is less than 0.4 percent of the total flight cost. Even taking into account the payload reduction for reusability, the improvement is therefore theoretically over a hundred times."
http://www.popularmechanics.com/science … ns-6653023
Then for the Falcon 9, the payload would be reduced from 10 mT to 6 mT. If the reduction in payload really is this high, then maybe it would be better to recover the first stage at sea. The loss in payload is coming from the reduction in the speed of staging as well as the need to retain a portion of the fuel for the return to base. Recovering at sea would not have these disadvantages because you could let the first stage make its usual trajectory at returning to the sea but use just small amount of propellant for the final slowdown before the sea impact.
In this article Musk does mention that returning back to the launch point allows the turnaround time at least for the first stage to be just hours. But will we really need that short a turnaround time at this stage of the game? A turnaround time of a few days would seem to be sufficient.
Perhaps the idea that retrieval at sea would be so expensive comes from the experience of the shuttle with the SRB's. But these were quite large and heavy at ca. 90 mT dry compared to that of the Falcon 9 first stage at less than 15 mT. Also, it is well known the labor costs for the shuttle were greatly inflated compared to a privately funded program.
The only additional requirement is that you would need a cover that could be extended to cover the engine section and would be watertight.

Bob Clark
Well I can't enough of Musk personally...he seems to hit the nail bang centre on the head. He knows exactly what needs to be done to revolutionise space travel and he's doing it.
Bob, "time is money" as they say. I don't think he's worried about scheduling - it's the labour input that drives up the cost of recovery from sea. A return to a spaceport is what is required. Rest assured he will have done all the calculations. I think he's right.
6MT is still a huge amount to get up there.
I am presuming that for the Mars transit he must be thinking in terms of orbital assembly.

Still I'd like to see the trade study.

Bob Clark

Rune · 2012-02-12 20:58:50

RGClark wrote:

Just saw this discussed on Nasaspaceflight.com

I actually linked to it on the reusable launcher thread, thought it was a more suitable place. Nice story with lots of info!

RGClark wrote:

Then for the Falcon 9, the payload would be reduced from 10 mT to 6 mT. If the reduction in payload really is this high, then maybe it would be better to recover the first stage at sea. The loss in payload is coming from the reduction in the speed of staging as well as the need to retain a portion of the fuel for the return to base. Recovering at sea would not have these disadvantages because you could let the first stage make its usual trajectory at returning to the sea but use just small amount of propellant for the final slowdown before the sea impact.
In this article Musk does mention that returning back to the launch point allows the turnaround time at least for the first stage to be just hours. But will we really need that short a turnaround time at this stage of the game? A turnaround time of a few days would seem to be sufficient.
Perhaps the idea that retrieval at sea would be so expensive comes from the experience of the shuttle with the SRB's. But these were quite large and heavy at ca. 90 mT dry compared to that of the Falcon 9 first stage at less than 15 mT. Also, it is well known the labor costs for the shuttle were greatly inflated compared to a privately funded program.
The only additional requirement is that you would need a cover that could be extended to cover the engine section and would be watertight.

The boosters don't stage slower so they have the range to get back, they stage slower because if they didn't, they'd keep on blowing on reentry long before they hit the water like they do now, unless you built them as stout as SRB's (actually, stouter, the SRB's stage slower). Plus, their lower speed lends itself to a 'pop up first stage' launch profile, so no flyback required at all, which always looked kind of silly to me in the first place. And no matter how you cover the engines, humidity at least would find a way in, it always does, so you still have to protect for corrosion the engine. Probably have to do that anyway to some extent, if you are going to use it form the cape a lot of times.

My main point of contention is actually the second stage performance. That has to reenter at high mach, so beefy structure and heavy ablative heatshield. And they want it to provide ~8kms/sec? It is starting to look like a SSTO, air-launched. Which actually, lends itself to the idea of treating the first stage boosters like airplanes. Short turnaround times, little to no work required to get it back in the air. Just do a visual inspection, get horizontal, stack it together and transport to the launch tower, fill 'em up and back to flight. A single booster (the heaviest part, so the most expensive) could service several second stages that way.

Rune. Damn, Falcon R is starting to look suspiciously like the DH-1.

RGClark · 2012-02-14 01:09:20

GW Johnson wrote:

Back to using Falcons and Dragons to go to the moon. You will need a lander of some tonnage, probably not unlike the old Apollo lander. You might even build it out of a Dragon with extra tanks, or maybe a from-scratch design. Whatever.
But, I don't see why you need a Centaur or any other departure stage. You don't need any more engines, you just need delta-vee. The Dragon will have the new Super-Draco thrusters on it with 120,000 lb of axial thrust, according to their website. There's eight of them, plenty of redundancy there. Just add a big dumb propellant tank, and plumb it up to the Dragon's system. Use the Dracos for all the delta-vee from LEO to lunar orbit, and back, sucking from the big dumb tank.
Launch the lander on one Falcon-Heavy, launch the Dragon and the big dumb tank on the other. Rendezvous in LEO, and dock the big dumb tank to the rear end of the Dragon, and the lander to its nose. Make up your plumbing connections. Then go to the moon.
Lots of minor details to work out, but I don't see any show stoppers here.
GW

Two Falcon Heavy's would be 106 mT to LEO which is getting into Saturn V class, and SpaceX already said it would take two Falcon Heavy's to send a manned mission to the Moon.
Could you get your plan to work with a single Falcon Heavy?

Bob Clark

RGClark · 2012-02-16 03:01:43

...
On this listing of space vehicles you can find that the later versions of the Centaur upper stage have a mass ratio of about 10 to 1:
http://www.friends-partners.org/partner … pndexc.htm
The Isp's given for the RL-10A engines used on these stages are around 450 s, but an updated version with a longer, extensible nozzle has an Isp of 465.5 s:
RL10B-2.
http://www.pw.utc.com/products/pwr/asse … l10b-2.pdf
This page gives the delta-V's needed for trips within the Earth-Moon system:
Delta-V budget.
Earth–Moon space.

http://en.wikipedia.org/wiki/Delta-v_bu … Moon_space
...
The RL-10 engine was proven to be reusable for multiple uses with quick turnaround time on the DC-X. The total propellant load of 40,000 kg could be lofted by two 20,000+ kg payload capacity launchers, such as the Atlas V, Delta IV Heavy, Ariane 5, and Proton.
The price for these launchers is in the range of $100-140 million according to the specifications on this page:
Expendable Launch Vehicles.
http://www.spaceandtech.com/spacedata/elvs/elvs.shtml

The original architecture was to use two of the 20 mT to LEO launchers currently available with two Centaur upper stages to get a 4 mT Dragon to the Moon and back.
What can we do with a single one of these launchers currently available? Using a single one of these launchers to carry a single Centaur upper stage we could carry about 1 mT to the Moon and back:
From the delta-V table, you need 4.04 km/s to go from LEO to low lunar orbit, 1.87 km/s to go from low lunar orbit to the lunar surface, and 2.74 km/s with aerobraking to go from the lunar surface back to LEO for a total of 8.65 km/s delta-V for a single stage making the round-trip.
Then with a 465.5 s Isp, 20 mT total mass including payload, 2 mT dry mass, and 1 mT payload we get: 465.5*9.8ln(20,000/(2000 + 1000)) = 8,650 m/s, sufficient for the round-trip.

This would suffice to carry a lunar rover to operate in the permanently shadowed regions of the lunar poles or for an NEO asteroid:

Lunar Prospecting Robot To Be Field Tested On Hawaii's Mauna Kea
ScienceDaily (Oct. 14, 2008)
http://www.sciencedaily.com/releases/20 … 134111.htm

This university developed robot probably cost no more than a few million dollars. The single Centaur upper stage costs in the range of $30 million. And the 20 mT to LEO launchers cost in the range of $100-140 million, according to the Spaceandtech.com site estimates, for a total in the range of $200 million. This is a fraction of the amount spent by mining interests on exploration:

Explore Mining.

World non-ferrous expenditures for all exploration in 2007 are estimated to be about $10.4 Billion dollars.

http://www.holden.house.gov/comm/explor … ploration/

This same site also indicates that mining exploration is by nature high risk:

So just what is exploration?
It’s the collection of processes that gather information about the presence or absence of mineral deposits
The over-riding goal of exploration is to find deposits that can be worked as profitable mining operations.
It is a time-consuming, multi-stage investment in information different gathering processes.
It’s also an expensive, high-risk investment, unlike ordinary businesses investments.
Depending on the literature source, the success rate for finding profitable mining operations (when weighed against the total number of mineral properties examined by a company) have ranges from a high of 4 in 100 (that’s a 4% success rate!), to less than 1 in 100 and as low as 1 in 1000 (that’s a .1% success rate!).

For any investment venture a cost/risk/benefit analysis has to be made. Compared to the cost already spent by mining interests yearly the cost is relatively low especially for a consortium of mining interests funding the mission together.
The risk is composed of the risk of the mission failing and of it not finding the high amounts of precious minerals. At least for the asteroid missions the risk of it not finding the high value minerals is low as there are several independent lines of evidence that precious metals are located uniformly on asteroids. So that leaves the risk of the mission failing. Considering the amount of U.S. experience with planetary missions, this risk is considerably better than the 1 in 1,000 chance of success some estimates put on Earth bound mining exploration.
However, quite important when measuring cost and risk, are the benefits to justify them. The possible benefits are more mineral wealth in a single asteroid than all that mined in all of human history.
Indeed the likelihood of the high amounts of precious minerals is so good, and the benefits of success are so extraordinarily high, that it would pay to do several missions if there are failures.
That is for the asteroid missions. However, if such asteroid mining missions are to be profitable then it would be much cheaper if the large amount of propellant needed to carry out the transport could be obtained from the Moon rather than by lofting it from Earth's deep gravity well. Then to insure that propellant could be obtained from the Moon's polar regions sample return missions to the lunar poles would have to be mounted as well. The nice thing about these missions is that the same rovers and spacecraft could be used for the asteroid sample return missions. Then these lunar sample return missions could be regarded as test missions to give further assurance of the technology for returning the samples from asteroids. And if the lunar polar samples show the high precious metal amounts tentatively detected by LCROSS then so much the better.
As I said to keep costs low these missions should be privately financed. NASA is planning to launch an asteroid sample return mission in 2016. This would not return the samples though until 2023 and is budgeted at $800 million without even launch costs:

NASA to Launch Asteroid-Sampling Spacecraft in 2016.
Mike Wall, SPACE.com Senior WriterDate: 25 May 2011 Time: 07:10 PM ET
http://www.space.com/11788-nasa-asteroi … -rq36.html

When you add on launch costs and considering the usual NASA cost overruns this will probably wind up being a billion dollar mission. Also, since some proposed human missions to asteroids would have a duration of 5 to 6 months, these sample return missions could return their samples in months rather than the seven years planned for the NASA mission.

Bob Clark

Last edited by RGClark (2012-02-16 07:40:48)

RGClark · 2012-02-21 08:46:57

RGClark wrote:

...
As I said to keep costs low these missions should be privately financed. NASA is planning to launch an asteroid sample return mission in 2016. This would not return the samples though until 2023 and is budgeted at $800 million without even launch costs:
NASA to Launch Asteroid-Sampling Spacecraft in 2016.
Mike Wall, SPACE.com Senior WriterDate: 25 May 2011 Time: 07:10 PM ET
http://www.space.com/11788-nasa-asteroi … -rq36.html
When you add on launch costs and considering the usual NASA cost overruns this will probably wind up being a billion dollar mission. Also, since some proposed human missions to asteroids would have a duration of 5 to 6 months, these sample return missions could return their samples in months rather than the seven years planned for the NASA mission.

Note that all the components for such a mission already exist, the launcher, the spacecraft, and the rover. All that is required is to mate them together. On that basis such a mission probably could be launched within a year. Note also all of the U.S., Russia, and Europe have the required 20 mT launcher, and the upper or space stage capable of the space traverse. And China will also with the introduction of the Long March 5 in 2014. Then the question arises who will be first?

A common charge leveled at the space program is what is it good for? If the U.S. government fully financed the mining operation then based on an estimated $20 trillion value for the minerals on a single asteroid, this would have enough value to retire the entire U.S. debt(!) Preferably though the U.S. would only be a partial investor to retain the costs savings of a privately financed venture. Even then as a minority investor, the return in value to the U.S. government could be in the trillions.

However, it may indeed be possible that a fully NASA financed venture could maintain the low costs of a privately financed one - with the right management. I consider the LCROSS lunar impactor to be the perfect NASA mission because it returned such profoundly important results and at low cost, only $79 million without launch costs, which is like pocket change for planetary missions:

Inside NASA's Plan to Bomb the Moon and Find Water.
By Michael Milstein
October 1, 2009 12:00 AM

Typically, 10 to 15 percent of a spacecraft's budget goes into instruments; on LCROSS, it's roughly 3 percent, or $2 million. When Anthony Colaprete, NASA's lead scientist for the mission, went to big aerospace companies for instruments, they laughed at his budget. So he turned to small outfits instead. He bought near-infrared spectrometers from a company that makes them for breweries to test the alcohol content of beer on assembly lines. He resisted agency reviewers who wanted him to put an anodized coating on the aluminum storage boxes. "One of their arguments was, `It's not very expensive--just do it,'" he says. "I'm like, `Well, I want to save that $1000. I'm very cheap.'"

http://www.popularmechanics.com/science/space/4277592

LCROSS: A HIGH-RETURN, SMALL SATELLITE MISSION.
Daniel Andrews, LCROSS PM
NASA-Ames Research Center, MS 240-3, Moffett Field, CA 94035, USA.
http://ntrs.nasa.gov/archive/nasa/casi. … 030093.pdf

Academy of Program/Project & Engineering Leadership.
Lunar CRater Observation and Sensing Satellite (LCROSS).

The Good Enough Spacecraft.
From Andrews‘s perspective, the LCROSS spacecraft had to be ―faster, good enough, cheaper.‖ He made clear to his team from the beginning that LCROSS was not about maximum performance. ―It was about cost containment,‖ Andrews said. ―LCROSS was not about pushing the technical envelope. It was about keeping it simple – keeping it good enough.‖
The LCROSS team had 29 months and $79 million to build a Class D mission spacecraft. (See below for a brief explanation of NASA mission risk classifications.) The low-cost, high-risk tolerance nature of the project led to a design based on heritage hardware, parts from LRO, and commercial-off-the-shelf components.

http://www.nasa.gov/pdf/474589main_LCRO … _23_10.pdf

LCROSS rode piggyback on the LRO mission so did not have to pay for the Centaur space stage, but even if you include this that would only be an additional $30 million or so.

LCROSS Program Manager Daniel Andrews and lead scientist Anthony Colaprete deserve major kudos for using innovative methods to accomplish such a successful mission under cost saving constraints. If we were to have NASA financed asteroidal and lunar prospector landers then they would be my choice to manage those missions.

Note now that if NASA funded these exploratory lander missions that proved definitively that asteroids or even the Moon contained such extraordinary mineral wealth, then under the principle that the government has the authority to grant mining rights to private companies, the U.S. government could sell these rights for a total of, say, $1 trillion, while only having to have spent ca. $200 million for the lander missions.

Bob Clark

Last edited by RGClark (2012-02-25 02:42:05)

RGClark · 2012-02-22 02:08:38

More on the low cost approach to the LCROSS mission from program manager Dan Andrews:

Going Lunar for Less.
By Melissa Salpietra
Posted 04.01.09
NOVA scienceNOW

DAN ANDREWS: The whole key with LCROSS is to use what exists. You take things that are available, you glue them together, you attach them in as simple a way as you can. You’re not doing a bunch of custom designs and development. You’re leveraging everywhere you can. And that’s a really smart way to get the most out of the money that you're given.

http://www.pbs.org/wgbh/nova/space/moon-for-less.html

LCROSS won Popular Mechanics' Breakthrough Award in 2010:

NASA's LCROSS Wins 2010 Popular Mechanics Breakthrough Award.

The individual LCROSS 2010 Breakthrough Award recipients are:
Daniel Andrews, LCROSS project manager at Ames
Anthony Colaprete, LCROSS project scientist and principal investigator at Ames
Stephen Carman, LCROSS spacecraft project manager at Northrop Grumman
Craig Elder, LCROSS spacecraft manager at Northrop Grumman
"We are honored to win this award," said Steve Hixson, vice president of Advanced Concepts - Space and Directed Energy Systems for Northrop Grumman Aerospace Systems in Redondo Beach, Calif. "It is a significant acknowledgement of the high caliber of our engineering skills and our close partnership with Ames, which developed the LCROSS payload and conducted mission operations. It also validates our ability to build small, inexpensive spacecraft with high science value very quickly, awakening the industry and the nation to the viability of this mission class."

http://www.nasa.gov/centers/ames/news/r … -86AR.html

Program manager Dan Andrews also was awarded NASA's Systems Engineering Excellence award for the mission:

2010 Systems Engineering Excellence Award.
01.06.10
http://www.nasa.gov/centers/ames/news/f … award.html

Bob Clark

SpaceNut · 2012-02-22 06:03:38

ARD wrote:

JoshNH4H wrote:
Sounds basically feasible, though real development work would have to be done before it would be possible to comment on whether the engineering would actually work out. The one thing that I'm not comfortable with is where you're going to get the oxygen from. It's probably not frozen in the craters, and electrolyzing your way to it would be difficult. I suppose you could use a TiO2 catalyst to photodissociate water in open sunlight, but that will only be about 6% efficient and possibly mass-intensive.
Heat up the regolith until it starts breaking up into oxygen gas and a metal slag? The unit to do that might be heavy, but one can theoretically use direct solar heating to get it done, and one of the 'waste' products is iron.

JoshNH4H wrote:

Well, 'heavy' is a very important factor here. I think photodissociation with a TiO2 catalyst would probably be more mass effective. It might be hard to get a lot of throughput with this method, but the same is true with all forms of solar power.

What you are talking about is vacuum pyrolysis. Turning Moon Dust into Oxygen Which we have talked about in the past under the Oxygen generation topics.....

RGClark · 2012-02-28 07:14:08

There will be a media demonstration of the Scarab lunar rover using a new fuel-cell technology on Wednesday, Feb. 29th at the NASA Glenn center:

Media Invited to NASA Glenn to See New Fuel Cell Demonstration on Mobile Rover.
Source: Glenn Research Center
Posted Thursday, February 23, 2012

CLEVELAND - A demonstration of a fuel cell that will allow rovers on extraterrestrial surfaces to go farther and last longer will be conducted at NASA's Glenn Research Center on Feb. 29 at 11 a.m.
The new type of fuel cell will extend the range of surface operations for rovers that will explore new worlds as part of future NASA missions. Unlike a conventional fuel cell that needs a pump to remove the water produced inside the device, this non-flow-through fuel cell uses capillary action to wick away the water. By eliminating the pump, a non-flow-through fuel cell is simpler, lighter, and more reliable.
The rover that will demonstrate the fuel cell in Glenn's Simulated Lunar Operations (SLOPE) facility is called SCARAB. It was developed by Carnegie Mellon Robotics Institute, Pittsburgh, under a grant from Glenn, and is regularly used for Human Robotic systems project mobility research in SLOPE.

http://www.spaceref.com/news/viewpr.html?pid=36206

Perhaps one of the reporters will inquire when the test vehicle can be turned into a flight ready version.

Bob Clark

JoshNH4H · 2012-03-13 22:48:41

SpaceNut wrote:

What you are talking about is vacuum pyrolysis. Turning Moon Dust into Oxygen Which we have talked about in the past under the Oxygen generation topics.....

Nah, I'm talking about a different technology. Wikipedia has more: http://en.wikipedia.org/wiki/TiO2#Photocatalyst

GW Johnson · 2012-03-15 09:26:09

Just to let y'all know, Spacex has been testing Merlin engines frequently at its McGregor facility, just a few miles from my front porch. I've been listening to the tests the last several days. They did a full Falcon-9 first stage just a day or so ago.

That stuff then goes to Canaveral. It'll be the next launch, which I see is now scheduled for April 30. This is the final COTS mission, condensed from two in the original plan. Actual docking and token (test) delivery of material to the ISS.

The rendezvous is automatic. The docking is done by the ISS crew, using its remote manipulator arm.

Looks to me like the Falcon-9/Dragon system is coming right along, and very smartly.

I know they're testing the super-Dracos out at McGregor, but I can't hear those from 8 miles away. Not loud enough. That'll be the landing thrusters for Dragon-as-planetary-probe, and the launch escape system for manned-Dragon.

I also know they're building a new test stand out there for Falcon-Heavy. Last I heard, it's supposed to be down in a hole, with water-curtain noise attenuation.

Just an update.

GW

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