Author Topic: NASA HLS (Human Landing System) Lunar Landers  (Read 1198711 times)

Offline jongoff

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Re: NASA HLS (Human Landing System) Lunar Landers
« Reply #80 on: 12/10/2018 09:41 pm »
With a fuel depot infrastructure the single stage lander would not be too large for SLS or even a smaller vehicle like New Glenn or Falcon Heavy as it could act as it's own departure stage.

Yeah, it seems odd that they're talking about refueling and reuse of stages in lunar orbit but act as though you couldn't do the same thing in LEO--even if the rocket isn't a lander with legs on it.

~Jon

Offline A_M_Swallow

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Re: NASA HLS (Human Landing System) Lunar Landers
« Reply #81 on: 12/11/2018 12:31 am »
Since the LOP-G is also in vacuum putting sun shields on the fuel tanks of the LOP-G depot, lander and tanker vehicles would not be difficult. To survive the thrust and exhaust gasses the sun shields would have to be stiff rather than a curtain.

Online Steven Pietrobon

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Re: NASA HLS (Human Landing System) Lunar Landers
« Reply #82 on: 12/11/2018 05:27 am »
The deck of charts are now publicly available here: https://www.nasa.gov/directorates/heo/nac-heoc

But these appear to be the most relevant to the current discussion (these are from Crusan’s slides)



Good to some some delta-V numbers. For the crew+cargo missions, I don't think that adding all the delta-V's is correct. Each stage does not perform TLI twice! What matters is the delta-V for each separate part:

LLO (crew): 3.2 + 0.9 + 2.0 + 2.0 + 0.9 = 9 km/s
LLO (cargo, storable): 3.2 + 0.9 + 2 = 6.1 km/s
LLO (cargo, electric): 3.2 + 0.64 + 2 = 5.85 km/s

NRHO (crew): 3.2 + 0.45 + 2.75 + 2.75 + 0.45 = 9.6 km/s
NRHO (cargo, storable): 3.2 + 0.45 + 2.75 = 6.4 km/s
NRHO (cargo, electric): 3.2 + 0.03 + 2.75 = 5.98 km/s

If we're sending cargo, then going through the Gateway just adds delta-V for no advantage, since we're not reusing the hardware. For crew, going through the gateway adds 0.6 km/s. They point out that it is only 6.7% more of the total, but its 6.7% more on top of 9 km/s where the rocket equation hurts the most.

Also, adding up all the delta-V's in this way is misleading since we are staging the Orion capsule either in LLO or NHRO. For example, Direct Ascent Apollo has the same delta-V as Rendezvous Apollo, but the latter gives a significant performance advantage. Thus, a proper analysis would work out all the dry and propellant masses, going back to determine IMLEO and then make a comparison.
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Offline GWH

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Re: NASA HLS (Human Landing System) Lunar Landers
« Reply #83 on: 12/11/2018 06:04 am »
What are the rendezvous opportunities like with NRHO? If a launch slips a day or more will that mean a long wait for the orbits to line up again (compared to LLO)?

Offline Proponent

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Re: NASA HLS (Human Landing System) Lunar Landers
« Reply #84 on: 12/11/2018 10:28 am »
Hypergolics do have one advantage they can be easily stored long term in space so a SEP tug could be used to take propellant from LEO to where it's needed.
But it doesn't lend itself to ISRU like hydrogen or methane and LOX which does make it kind of a dead end for Mars and beyond.

Long-term storage of liquid hydrogen, a deep cryogen, would be problematic, but moderate cryogens like lox and methane are sometimes referred to as "space storable."  And you wouln't have to worry about them freezing.

Offline Proponent

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Re: NASA HLS (Human Landing System) Lunar Landers
« Reply #85 on: 12/11/2018 10:35 am »
Anybody here with inside knowledge on accurate numbers for the ESM?

Quote
Its dry mass is 3.5 metric tons and it can carry 8.6 tons of propellant.
https://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/20170009574.pdf

Aerospace Research Central shows a newer version of that paper.  Maybe somebody with ready access to it could check for up-to-date numbers?

Offline clongton

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Re: NASA HLS (Human Landing System) Lunar Landers
« Reply #86 on: 12/11/2018 02:25 pm »
Long-term storage of liquid hydrogen, a deep cryogen, would be problematic, but moderate cryogens like lox and methane are sometimes referred to as "space storable."  And you wouldn't have to worry about them freezing.

One of the main goals that will be worked on when we return to the moon is to establish ISRU propellant plants so that spacecraft can be powered by fuels that do not need to be brought up from earth's deep gravity well. Carbon is necessary to create methane but it is not found on the moon in anything but rarefied amounts. Water however is apparently abundant. From water, hydrogen and oxygen can be extracted to create LOX/LH2 propellants. That's too good an opportunity to pass up. So long as the water actually is as abundant (and reasonably extractable) as it appears to be then lunar-sourced LOX/LH2 will be the in-space propellant of choice for the foreseeable future. And now that we are finding more and more water in diverse places (for example 1916 Boreas [a NEO] was just announced today) I don't see that changing for a very long time.
« Last Edit: 12/11/2018 02:26 pm by clongton »
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Offline TrevorMonty

Re: NASA HLS (Human Landing System) Lunar Landers
« Reply #87 on: 12/11/2018 02:42 pm »
NASA seems think storing LH is big problem while LM ( Lunar lander and Mars orbiting base station) and ULA (ACES Upper Stage) don't see any problems storing it for weeks to months without cryocoolers. Its about time NASA funded a demo mission by one of these companies to prove it one way or another.

Besides higher ISP performance of LH LOX it is also fuel we have lot of flight experience with. There is no flight proven methane boosters or USs.

Offline A_M_Swallow

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Re: NASA HLS (Human Landing System) Lunar Landers
« Reply #88 on: 12/11/2018 04:51 pm »
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NASA seems think storing LH is big problem while LM ( Lunar lander and Mars orbiting base station) and ULA (ACES Upper Stage) don't see any problems storing it for weeks to months without cryocoolers. Its about time NASA funded a demo mission by one of these companies to prove it one way or another.

Besides higher ISP performance of LH LOX it is also fuel we have lot of flight experience with. There is no flight proven methane boosters or USs.

A 6 month lunar surface mission aiming to have at least 1 litre of hydrogen at the end of the test. An additional design aim to lose less than half of the say 2 litres that were delivered to the Moon.

Probe likely to include hydrogen, flask, sun shield, instrumentation, (insulated) legs, own radio link, controller, solar panel, (heated) batteries and optional cryocooler. Optionally the transmitter and electronics may be switched off during the lunar night to save power.

CLPS can deliver the payload to the Moon. Electronics that works at cryogenic temperatures was developed for the James Webb Telescope. For example https://www.nasa.gov/centers/goddard/news/topstory/2008/jwst_shakerattle.html

Offline speedevil

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Re: NASA HLS (Human Landing System) Lunar Landers
« Reply #89 on: 12/11/2018 05:25 pm »
A 6 month lunar surface mission aiming to have at least 1 litre of hydrogen at the end of the test. An additional design aim to lose less than half of the say 2 litres that were delivered to the Moon.

Because loss is a surface phenomena, not volume, a litre of hydrogen has with the same tank technology, a tenth of the storage duration of 1000l.

So, you want to scale down the duration unless you actually want six month long hydrogen missions with a kilo of fuel.

It is not immediately clear to me that taking methane down, and just using it with ISRU oxygen is not a better plan than all of the infrastructure that goes into finding, extracting water and storing it as hydrogen, then using it in comparably bulky tanks.



Offline TrevorMonty

Re: NASA HLS (Human Landing System) Lunar Landers
« Reply #90 on: 12/11/2018 05:56 pm »
1
NASA seems think storing LH is big problem while LM ( Lunar lander and Mars orbiting base station) and ULA (ACES Upper Stage) don't see any problems storing it for weeks to months without cryocoolers. Its about time NASA funded a demo mission by one of these companies to prove it one way or another.

Besides higher ISP performance of LH LOX it is also fuel we have lot of flight experience with. There is no flight proven methane boosters or USs.

A 6 month lunar surface mission aiming to have at least 1 litre of hydrogen at the end of the test. An additional design aim to lose less than half of the say 2 litres that were delivered to the Moon.

Probe likely to include hydrogen, flask, sun shield, instrumentation, (insulated) legs, own radio link, controller, solar panel, (heated) batteries and optional cryocooler. Optionally the transmitter and electronics may be switched off during the lunar night to save power.

CLPS can deliver the payload to the Moon. Electronics that works at cryogenic temperatures was developed for the James Webb Telescope. For example https://www.nasa.gov/centers/goddard/news/topstory/2008/jwst_shakerattle.html
I was thinking more like few 100kgs at EM1 , well away from reflected heat of earth.


Offline TrevorMonty

Re: NASA HLS (Human Landing System) Lunar Landers
« Reply #91 on: 12/11/2018 06:09 pm »



It is not immediately clear to me that taking methane down, and just using it with ISRU oxygen is not a better plan than all of the infrastructure that goes into finding, extracting water and storing it as hydrogen, then using it in comparably bulky tanks.

The expensive part and most energy intense is splitting hydrogen from oxygen, its waste throw H away once hard work has been done. Cooling should be easy with access to deep cold of water bearing craters.

Offline A_M_Swallow

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Re: NASA HLS (Human Landing System) Lunar Landers
« Reply #92 on: 12/11/2018 08:29 pm »
A 6 month lunar surface mission aiming to have at least 1 litre of hydrogen at the end of the test. An additional design aim to lose less than half of the say 2 litres that were delivered to the Moon.

Because loss is a surface phenomena, not volume, a litre of hydrogen has with the same tank technology, a tenth of the storage duration of 1000l.

So, you want to scale down the duration unless you actually want six month long hydrogen missions with a kilo of fuel.

This is a proof of concept test rather than a test of final hardware. They 2 kg flask size is to keep the total mass under 35 kg, the upper payload limit of the initial CLPS landers. There are obvious variations on this test depending on need and money.

Storing the fuel for six months corresponds to two manned trips to the Moon a year. The fuel could be in the lander's tanks or come from ISRU or be delivered by a cargo lander for use by the manned ascent stage. IMHO A manned lander and tanker lander for hydrogen will only be designed if NASA knows hydrogen can be kept on the Moon and possibly in an orbiting propellant depot.

The experiment could send reports back every two weeks allowing monitoring of any boil off. This would provide information for shorter missions.

Quote
It is not immediately clear to me that taking methane down, and just using it with ISRU oxygen is not a better plan than all of the infrastructure that goes into finding, extracting water and storing it as hydrogen, then using it in comparably bulky tanks.


A storage system that can handle hydrogen will only need minor modifications to handle methane and Lox.

Oxygen can also be extracted from solids like silicon dioxide (sand) and aluminium oxide (sapphire) but splitting water is simpler.

Online Markstark

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Re: NASA HLS (Human Landing System) Lunar Landers
« Reply #93 on: 12/12/2018 01:24 am »
The deck of charts are now publicly available here: https://www.nasa.gov/directorates/heo/nac-heoc

But these appear to be the most relevant to the current discussion (these are from Crusan’s slides)



I don't think that adding all the delta-V's is correct. Each stage does not perform TLI twice!

Hey Steve - which case are you referring to in the chart? The places where I see TLIs listed twice are in instances where two rockets involved (1 crew, 1 cargo). So the multiple TLI are for the multiple rockets rather than # stages. At least that's how I read it. But I could be misreading it. Thanks!

Offline speedevil

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Re: NASA HLS (Human Landing System) Lunar Landers
« Reply #94 on: 12/12/2018 02:44 am »



It is not immediately clear to me that taking methane down, and just using it with ISRU oxygen is not a better plan than all of the infrastructure that goes into finding, extracting water and storing it as hydrogen, then using it in comparably bulky tanks.

The expensive part and most energy intense is splitting hydrogen from oxygen, its waste throw H away once hard work has been done. Cooling should be easy with access to deep cold of water bearing craters.

Apologies for being unclear, I was referring to sourcing O2 from regolith, not water.

For the case of methane (or O2), a storage vessel that ends up with 50% of the total initial mass after 6 months is utterly trivial and off the shelf, for containers over some 30l.

An off-the-shelf dewar will get you that performance, for around $1000 shipped. You will need a simple sunshade.

No cryo refrigeration, no development.
« Last Edit: 12/12/2018 02:51 am by speedevil »

Online Steven Pietrobon

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Re: NASA HLS (Human Landing System) Lunar Landers
« Reply #95 on: 12/12/2018 07:01 am »
Hey Steve - which case are you referring to in the chart? The places where I see TLIs listed twice are in instances where two rockets involved (1 crew, 1 cargo). So the multiple TLI are for the multiple rockets rather than # stages. At least that's how I read it. But I could be misreading it. Thanks!

Yes, that's my understanding as well.
Akin's Laws of Spacecraft Design #1:  Engineering is done with numbers.  Analysis without numbers is only an opinion.

Offline ncb1397

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Re: NASA HLS (Human Landing System) Lunar Landers
« Reply #96 on: 12/12/2018 08:14 am »
Considering how changing an in-space stage and service module into a lunar lander would already warrant many changes to the design anyway, like new avionics and landing legs, slightly stretching the tanks to meet additional delta V requirements would be a minor change.

I think that propellant tank changes would hurt component commonality with the service module as being a large part of the current design, it likely touches a lot of the other internal components. Landing legs and some additional sensors would largely be external to the current design and thus doesn't interfere with anything that is already there.

One possible way to avoid that is just doing a clean sheet ascent stage and using ESM derived for the lander and transfer stage. This deals with the dV shortfall for the descent stage and optimizes the mass of the ascent stage where weight savings are most felt. You would need the following specs:

tug - 3 t version of the 3.5 t ESM (no consumable tanks, 2 solar wings vs 4)
lander - 3.5 t version of the 3.5 t ESM (water/O2/N2 tanks are reduced to half count, 4 solar array wings with 1 panel per wing vs 3, additional bottom mounted landing sensors and legs)
ascent stage - 2.5 t dry, 4.3 t propellant load
payload - .35 t, mainly two 80 kg astronauts + 2 80 kg EVA suits. With lighter escape suits for a lunar base, 3-4 astronauts.

Total stack would be about 31 t that can be launched complete or in 7-12 t components (SLS co-manifest/Falcon Heavy/New Glenn/Vulcan/etc). Self ferry would require about 100 m/s to reach the gateway via the long-duration trajectory and so that needs to be looked at whether you can account for that hit somehow or you need to top off tanks via refueling. That being said, self ferry as a complete stack wouldn't likely be a problem as the transit stage has a little wiggle room (1000 m/s vs 850 m/s requirement per NAC presentation).

This leads to the following delta-vs assuming 316 isp engines.

1st stage: ~1000 m/s
2nd stage: ~1830 m/s
3rd srage: ~2850 m/s

The 1st stage might need to help with the descent a bit like the how the Apollo CSM did for later Apollo missions. Additionally, any delta-v shortfall could be accommodated by adjusting the size of the NRHO temporarily.  As far as landing thrust to weight on the shuttle OMS, you get thrust to weigh of about 2 at landing near descent tank depletion. The Aux engines are about .2. Not sure about throttleability on that engine, but you would need that for hover.

The other remaining big question is the feasibility of a 2.5 t dry, 4.3 t propellant load ascent stage. The Apollo LEM ascent stage was 2.15 t dry with a RCS/APS load of 2.65 t. 1.65 t of extra propellant at a cost of .35 t of dry mass or a ratio of 4.7 to 1 shouldn't be a difficult lift.

As far as the cargo variant, this would of course leave the ascent stage off. The transfer stage would do the LLO injection from a TLI and the lander's payload would essentially be the mass of the ascent stage + .35 t (about 7 t). Given that the Destiny module was ~105 cubic meters of volume at a mass of about 14 t, you would be able to fit decently sized habs, rovers etc (enough for NASA's exploration needs anyways). The payload on Falcon Heavy though would be a bit smaller as it couldn't push the 31 t stack to TLI. Injection would need to be sub-TLI and the transfer/descent stage would have to make up the difference with a lighter payload (not sure what that is, but it is likely substantially greater than 0).

As far as what this leads to in the bigger picture. I could see this lander working to boot strap a small base with propellant production capability. You would then try to migrate to the 50 t + fully re-usable single stage lander that could go back and forth from the gateway to the surface ad nauseum with the only limitation being propellant production capacity. With the crew portion left off, the same vehicle could act as tanker for refueling interplanetary craft based at the gateway (gateway -> high Mars orbit-> gateway is a modest dV cost). You could also see it working in tandem with something like BFR, where some location is not reachable directly by the larger craft, this lander would provide about 5 km/s (ascent+descent stage) of extra manueverability at a cost of ~20 t).
« Last Edit: 12/12/2018 08:32 am by ncb1397 »

Offline TrevorMonty

Re: NASA HLS (Human Landing System) Lunar Landers
« Reply #97 on: 12/12/2018 08:20 am »



It is not immediately clear to me that taking methane down, and just using it with ISRU oxygen is not a better plan than all of the infrastructure that goes into finding, extracting water and storing it as hydrogen, then using it in comparably bulky tanks.

The expensive part and most energy intense is splitting hydrogen from oxygen, its waste throw H away once hard work has been done. Cooling should be easy with access to deep cold of water bearing craters.

Apologies for being unclear, I was referring to sourcing O2 from regolith, not water.

For the case of methane (or O2), a storage vessel that ends up with 50% of the total initial mass after 6 months is utterly trivial and off the shelf, for containers over some 30l.

An off-the-shelf dewar will get you that performance, for around $1000 shipped. You will need a simple sunshade.

No cryo refrigeration, no development.

To extract O, the regolith is heated with H gas. O bonds with H forming water, with iron being one of by products of this reaction. NB this how to produce iron on moon.
Water now needs splitting by electrolysis to extract O and H. H can be used again to extract more O from regolith by repeating process.

If you have access to water may as well produce LH and LOX for little more energy input.
« Last Edit: 12/12/2018 08:27 am by TrevorMonty »

Online Markstark

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Re: NASA HLS (Human Landing System) Lunar Landers
« Reply #98 on: 12/12/2018 10:53 pm »
I guess I'm not understanding why NASA wouldn't want to make the entire lunar lander reusable.

What's mentioned in the presentation: If you use hypergolic propellants a single stage lander, even when using a space tug for lunar orbital transfers, would be too heavy and physically large to lift on commercial launchers and can't be co-manifested on SLS.

Seeing how NASA limits itself to hypergolics, it seems more likely that NASA instead wants to minimise programme risk and development cost, foregoing the technologies I mentioned earlier that would enable full reuse. With only partial reuse, SLS only flying once a year and no potential for ISRU, there is no actual plan to actually do frequent surface sorties that would justify the added development costs for fully reusable lander. In other words there is no plan to actually turn this into a serious way to do lunar exploration and settlement.

I listened in on the call during Jason Crusan’s Gateway presentation this week. He spent a lot of time discussing the Gateway Lunar Lander. He was specifically asked about ISRU and he said that a fuel combination has not been selected for the lander at this point and that the trade space is open to various types including methalox, hydrolox and hypergolics. There’s an imminent lander study broad agency announcement where they will fund 6-9 month studies to refine the design including propellant selection. He also mentioned that there is a desire to have all three elements be reusable at some point but it will not be a requirement for the initial descent elements. Hope this helps.
« Last Edit: 12/12/2018 10:56 pm by Markstark »

Offline speedevil

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Re: NASA HLS (Human Landing System) Lunar Landers
« Reply #99 on: 12/13/2018 09:09 am »
<snip regolith oxygen extraction question>
To extract O, the regolith is heated with H gas. O bonds with H forming water, with iron being one of by products of this reaction. NB this how to produce iron on moon.
Water now needs splitting by electrolysis to extract O and H. H can be used again to extract more O from regolith by repeating process.

If you have access to water may as well produce LH and LOX for little more energy input.
Sure, if that is the method you are using.
Melt phase electrolysis is also another possibility - broadly how aluminim is made.

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