Author Topic: Thoughts about manned lunar with existing rockets and fuel depots  (Read 10652 times)

Offline KelvinZero

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This link is mainly based on the lunar portion of the document found here.
http://forum.nasaspaceflight.com/index.php?topic=21518.0

The dry mass components for the apollo missions were not that big (someone check my numbers though)
Ascent stage: total:4.7t, fuel2.35t     => dry:2.35t
Descent stage: total:10.3t fuel:8.2t   => dry:2.1t
Command Module: 5.8t                    => dry:5.8t
Service Module: total:24t fuel:18.4t  => dry:6.1t

Third stage total: 120, fuel:105.2     => dry:15t

If I understand the document, we don't need the third stage. Instead we refuel the second stage. if someone could check that for me it would be great.

Discounting the third stage, the dry masses are not particularly great. If I understand that document a single atlas or delta class launch, plus a lot of orbital fuel could land 4-6 ton, and assuming that does not include lander then that is enough for apollo sized missions.

I actually think apollo sized is more than enough, if it is sustainable. However, could we go even bigger, eg with two launches aside from fuel depot launches? Or with VASIMR to deliver unmanned portions to lunar orbit?

I would like to hear some discussion of a modular architecture with some ability to scale up and down the missions without super heavy lift.

Offline mmeijeri

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We could go bigger than Constellation with two launches and propellant transfer, even if it is only for the lander and even if it just uses storable propellant. If we start from that basis we can then upgrade to our heart's content (RLVs, SEP, cryogenic depots, refuelable upper stages, ISRU, aerobraking, tethers, NTR, even HLV).
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Offline mmeijeri

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By the way, what is so important about restricting ourselves to two launches, not counting propellant? Not that that is a difficult constraint once you have propellant transfer.
« Last Edit: 05/13/2010 07:06 am by mmeijeri »
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Offline KelvinZero

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Nothing important about restricting ourselves to two launches. Just an example. Having a range of options, including a minimalist one, is what makes it interesting.

One thing I thought might limit us is if in-orbit assembly is more complicated than just docking the components listed above together. They are already required to come apart and some are already designed to reconnect in orbit. for the ascent, descent, command and service stages/modules, we could make them 3-5 times bigger than Apollo while still remaining in the 20-30 ton range per launch.

However if we need something larger than the upper stage for pushing it to the moon, then we need some additional complexity, perhaps the ability to strap multiple upper stages together?

Offline mmeijeri

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One thing I thought might limit us is if in-orbit assembly is more complicated than just docking the components listed above together. They are already required to come apart and some are already designed to reconnect in orbit. for the ascent, descent, command and service stages/modules, we could make them 3-5 times bigger than Apollo while still remaining in the 20-30 ton range per launch.

Even the Constellation pieces are well within the range of EELV Mediums when you offload the propellant. If you split the lander into a crasher stage and the lander proper, then you can divide them up further. Atlas Heavy would provide far, far more capacity than the dry mass of any conceivable payload we could hope to launch in the next twenty to thirty years.
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Offline KelvinZero

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Was just trying to find stats on the constellation earth departure stage to see if it would fit, but couldnt find any masses here.
http://en.wikipedia.org/wiki/Earth_Departure_Stage

Actually, I had forgotten/not grasped that constellation itself required two launches, a few days of in-space storage of fuel using new techniques, and a time limit on the second launch because of this.

Offline Downix

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If you are doing two launches, you can do it without orbital refueling entirely by slow-boating the lander.  Instead of it being a fast trip, it takes a very slow, month-long meander to the moon.  Once it is safely in orbit, your human crew departs.  Not the easiest ride, but more than doable.
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Offline KelvinZero

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Got any numbers on that?

I guess it has one problem that the lunar module is not available for back up life support and thrust, as used by Apollo 13.

Offline Downix

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Got any numbers on that?

I guess it has one problem that the lunar module is not available for back up life support and thrust, as used by Apollo 13.
To get to the moon, the best option is to use the ITN, which uses little fuel. http://en.wikipedia.org/wiki/Interplanetary_Transport_Network

You may need a third launch for an EDS Centaur, but I don't think that would be necessary
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Offline KelvinZero

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Ah yes I have heard of ITN and how it can allow cheap slow transport to the moon. Im guessing that this would make the fuel to get the lander and ascent stage relatively negligible. I wasn't confident I would get the math right figuring how much this would reduce the EDS and service module mass, total and dry, if their sole purpose was to taxi the command module to-from the moon.


Offline Bill White

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This paper suggests a 25% - 33% increase in net delivered payload using single impulse ballistic trajectories.

http://ccar.colorado.edu/nag/papers/AAS%2006-132.pdf

It appears this technique could be used to pre-position the entire lander.
« Last Edit: 05/14/2010 02:22 am by Bill White »
EML architectures should be seen as ratchet opportunities

Offline jimgagnon

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Was just trying to find stats on the constellation earth departure stage to see if it would fit, but couldnt find any masses here.
http://en.wikipedia.org/wiki/Earth_Departure_Stage

Actually, the other night I was looking for the same information and came up blank for hard, public data on the EDS. A nice placeholder for your purposes could be the Ares 1 Upper Stage (AIUS) -- with a J-2x and more propellant than a S-IVb, it would make a wonderful EDS.

These slides will provide the hard data you need on AIUS:
  http://www.slideshare.net/astrosociety/ares-i-upper-stage-update-presentation
« Last Edit: 05/14/2010 02:38 am by jimgagnon »

Offline KelvinZero

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This paper suggests a 25% - 33% increase in net delivered payload using single impulse ballistic trajectories.

http://ccar.colorado.edu/nag/papers/AAS%2006-132.pdf

It appears this technique could be used to pre-position the entire lander.

How about this one?
http://www.universetoday.com/2010/03/06/astronomy-without-a-telescope-%E2%80%93-the-hitchhikers-guide-to-the-solar-system/

"Edward Belbruno proposed a low energy lunar transfer to get the Japanese probe Hiten into lunar orbit in 1991 despite it only having 10% of the fuel required for a traditional trans-lunar insertion trajectory. The manoeuvre was successful, although travel time to the Moon was five months instead of the traditional three days."


Offline Patchouli

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Another trick might be to split the EDS into two parts one launch brings up the main H2 tank and engines and another the main O2 tank.
The H2 side would have a small LOX tank as well.
The masses just don't split well since LOX is most of the mass in a lH2 rocket.

Use a vehicle like F9-H or Atlas phase two and it should be possible to assemble a 60T EDS in two launches.

Fuel transfer at least LOX might be easier then joining two halves of a rocket stage in orbit.

One simple refuelling solution might be to have an Ares I US equipped with a sun shade launched on a Delta IV-H with just enough lox to get it's self in orbit and then have several modified F9 upper stages refuel the LOX tank.

The CxP EDS was designed to spend up to four weeks in orbit.
« Last Edit: 05/14/2010 04:00 am by Patchouli »

Offline Bill White

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This paper suggests a 25% - 33% increase in net delivered payload using single impulse ballistic trajectories.

http://ccar.colorado.edu/nag/papers/AAS%2006-132.pdf

It appears this technique could be used to pre-position the entire lander.

How about this one?
http://www.universetoday.com/2010/03/06/astronomy-without-a-telescope-%E2%80%93-the-hitchhikers-guide-to-the-solar-system/

"Edward Belbruno proposed a low energy lunar transfer to get the Japanese probe Hiten into lunar orbit in 1991 despite it only having 10% of the fuel required for a traditional trans-lunar insertion trajectory. The manoeuvre was successful, although travel time to the Moon was five months instead of the traditional three days."

I believe these are different examples of the same basic concept of using multi-body gravitational paths rather than traditional 2 body Hohmann trajectories.

But yes, I believe that using these techniques can significantly reduce the initial mass needed in LEO for a successful lunar mission, thereby significantly reducing overall mission costs.
EML architectures should be seen as ratchet opportunities

Offline mmeijeri

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Ah yes I have heard of ITN and how it can allow cheap slow transport to the moon. Im guessing that this would make the fuel to get the lander and ascent stage relatively negligible. I wasn't confident I would get the math right figuring how much this would reduce the EDS and service module mass, total and dry, if their sole purpose was to taxi the command module to-from the moon.

It reduces the delta-v to 3.2km/s, essentially by eliminating the insertion burn. This is a sizeable saving, but it's still far from cheap. It would compensate for most of the inefficiency of storable propellant however.

We have a (dormant) thread on the mathematics behind it:

Online study group for dynamical systems theory and innovative trajectories
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Offline KelvinZero

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Another trick might be to split the EDS into two parts one launch brings up the main H2 tank and engines and another the main O2 tank.
The H2 side would have a small LOX tank as well.
The masses just don't split well since LOX is most of the mass in a lH2 rocket.

Use a vehicle like F9-H or Atlas phase two and it should be possible to assemble a 60T EDS in two launches.

Fuel transfer at least LOX might be easier then joining two halves of a rocket stage in orbit.

One simple refuelling solution might be to have an Ares I US equipped with a sun shade launched on a Delta IV-H with just enough lox to get it's self in orbit and then have several modified F9 upper stages refuel the LOX tank.

The CxP EDS was designed to spend up to four weeks in orbit.

Oh ok, I just saw 4 days on some wiki site. Four weeks is a fairly long time. Would that have needed new research also?

As for assembling engines and tanks in orbit, what about just strapping two standard size EDS together? more parts means more risk but less vehicles to design has many advantages, and this approach could extend to arbitrary numbers of rockets strapped together?

If you knew you were going to be strapping multiple rockets together all the time it would surely be better to just design a bigger rocket, but it seems nice to keep it small, with the option to go big if you really want to without any ten year gap to design a new rocket. You would probably only do this for unmanned cargo. If something went wrong there could still be a chance to recover it on some later orbit.

( I really have no idea of the actual complexity of these tasks.. 'strapping two EDS together' is easy to say.. :) )

Offline Archibald

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minimalistic approach
« Reply #17 on: 05/14/2010 05:48 am »
Nothing important about restricting ourselves to two launches. Just an example. Having a range of options, including a minimalist one, is what makes it interesting.

Minimalistic ? I use to like this approach http://www.nss.org/settlement/moon/ELA.html

I would love to see Early Lunar Access upgraded with EELVs, either the Atlas V heavy or the phase 1 EELVs (35 tons to LEO).
Han shot first and Gwynne Shotwell !

Offline KelvinZero

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We have a (dormant) thread on the mathematics behind it:

Online study group for dynamical systems theory and innovative trajectories

Thank you mmeijeri, that will be very helpful.

Now could you pick up my brain and put it back in my head? :)

Offline sdsds

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Was just trying to find stats on the constellation earth departure stage to see if it would fit, but couldnt find any masses here.
http://en.wikipedia.org/wiki/Earth_Departure_Stage

It's easiest to just trust Ed Kyle:

http://www.spacelaunchreport.com/ares5.html

The June 2008 "LV 51.00.48" iteration of the Ares V design included an EDS with a dry mass of 24 t.  Filled with 253 t of propellant it was tasked with performing as the second stage to orbit and then putting a 65 t + reserves payload through a 3175 m/s TLI.

Putting that into LEO would require the equivalent of a dozen EELV launches, assuming you dry-launch the EDS and use it as propellant depot that magically has zero boil-off.
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