totm dec 2023 Artemis Discussion Thread

[mikegarrison]

1 hour ago, kerbiloid said:

1. How could they know that the produced is actually methane, and it will survive in cryotanks for two years until their return.

Nothing is absolutely sure. Maybe the Second Law, but who knows? Maybe not even that.

However, you don't hit the launch button the instant the methane tank hits "full". You wait and see whether that's sustainable for, say, a Martian year. Then you launch on the next available window. Or something like that, anyway.

Knowing whether it is actually methane is trivial. For one thing, your methane-producing ISRU device probably can't produce anything *but* methane.... Chemistry doesn't work differently on Mars than here.

Edited November 5, 2019 by mikegarrison

[kerbiloid]

1 hour ago, mikegarrison said:

Knowing whether it is actually methane is trivial. For one thing, your methane-producing ISRU device probably can't produce anything *but* methane.... Chemistry doesn't work differently on Mars than here.

They would probably hope that it does, because a gas plant here on Earth is a large collection of pipes and columns containing a lot of chemicals just to purify the pre-made methane pumped from the Earth from admixtures like carbon and sulfur oxides which will present in the Martian product, too.
Coolers to freeze away high-boiling substances, diethanolamine which in one column binds the carbon oxide, in another regenerates, venting it away, etc.
And a small amount of water or something else can easily freeze at the methalox temperature blocking the fuel pump.

As they need all this in cramped conditions of a small ISRU, this is of course a Zubrin-rated plan, but unlikely anybody from any administration can sign the flight plan with "I guarantee it works" _{famous last} words.

Until a crewed Martian/Lunar surface station starts ISRU-producing the fuel under the human control and with an abort option like at least a ground habitat no ISRU plans will go.

***

(I mention this here just due to the "droptanks - the good or the evil?" sense)

Edited November 5, 2019 by kerbiloid

[sevenperforce]

16 hours ago, jinnantonix said:

The drop tank layout used here just fits inside the standard Falcon Heavy fairing shell, but it can definitely be improved to reduce the number of drop tanks.  Double sized tanks, two tanks each for the TLI-NRHO and NRHO-LLO and LLO-Surface tank drops would be feasible.

I'm thinking two tanks for the transfer burn and two tanks for the descent.

16 hours ago, jinnantonix said:

The Falcon upper stage does not have sufficient fuel to reach TLI, let alone do the fly-by and LOP-G insertion.  I would imagine hydrogen boil off would play a part in that calculation.  Doing all the FH burns within a few hours of launch eliminates any issues and risks around that.

If the dV requirements on the sortie vehicle are lowered, total throw is reduced, and so the Falcon upper stage has more propellant reserves. There's no hydrogen, of course, and while there's LOX boil-off it's not too bad. The higher specific impulse of kerolox makes it better to use kerolox until the dry mass of the MVac becomes problematic. You also reduce total throw by sending the reusable capsule ahead separately.

16 hours ago, jinnantonix said:

Note that the FH upper stage for the lunar lander launch has some fuel remaining after reaching LEO, which could potentially be used, giving the craft a little extra boost - still nowhere near enough to get the lunar lander to TLI, the OMS still needs to do 3 burns and provide the 44 ton craft approx dV =1500 m/s  to get to NRHO.

You might be able to do this to help with an elliptical starting orbit before launching the TLI stage, but then you run into node phase issues, so it's better to just reserve performance and recover the side boosters.

11 hours ago, mikegarrison said:

I'm not sure I understand how "drop tanks" and "reuseable lander" go together. I guess the next flight brings more drop tanks and then they have to be attached to the lander in lunar orbit? This seems very challenging. It's also an absolutely critical failure point, because if anything goes wrong and the lander does not have enough useable fuel to get back to the LOP-G, the crew is probably dead.

Reuse is a noble goal, but there's little reason to reuse tankage in cislunar space unless you have a way of refilling locally.

This particular concept only reuses the crew capsule and RCS thrusters; a new descent/ascent vehicle is launched with each sortie. It mates to the reusable crew capsule. Part of that mating process includes a small tank that presses the crew capsule's RCS and carries enough propellant to get from LLO to LOP-G, and that's something that can be verified to work before leaving LOP-G.

[sh1pman]

NASA had a chance to lower the SLS rocket’s cost—but it stuck with Boeing

Quote

As a result, NASA has gone with a contractor that significantly under-performed on the SLS core stage, which is years behind schedule, billions of dollars over budget, and yet to prove itself in flight. Now it has bet the future of its deep space exploration program for at least the next decade on the same company. NASA fans can only hope that Boeing builds rockets as well as they do lobbying coalitions.

https://arstechnica.com/science/2019/11/nasa-rejects-blue-origins-offer-of-a-cheaper-upper-stage-for-the-sls-rocket/

[tater]

https://boeing.mediaroom.com/2019-11-05-Boeing-Proposes-Fewest-Steps-to-the-Moon-for-NASAs-Human-Lander#assets_20295_130548-117:20659

 

Boeing's SLS launched lander. Designed to maximally rip off the taxpayer.

 

[jinnantonix]

13 hours ago, sevenperforce said:

I'm thinking two tanks for the transfer burn and two tanks for the descent.

If the dV requirements on the sortie vehicle are lowered, total throw is reduced, and so the Falcon upper stage has more propellant reserves. There's no hydrogen, of course, and while there's LOX boil-off it's not too bad. The higher specific impulse of kerolox makes it better to use kerolox until the dry mass of the MVac becomes problematic. You also reduce total throw by sending the reusable capsule ahead separately.

Ah, of course Merlin's use RP-1 and LOX, my bad.  I'll run a test assuming upper stage has a Merlin 1D Vacuum engine with Isp = 348 .  The estimated payload is 34t from LEO to TLI and another ~500 m/s for LOP-G rendezvous.

If that works, then the two large tanks concept is a go.  I did some quick modelling and it seems to still fit in the standard FH fairing, but it is snug.

Agree that with this design, recouping cost by recovering stage one would be preferable.

 

Edited November 6, 2019 by jinnantonix

[jadebenn]

57 minutes ago, tater said:

Uh, guys?

That EUS doesn't have RL-10s.

[sh1pman]

8 minutes ago, jadebenn said:

That EUS doesn't have RL-10s.

The artist was underpaid.

[Barzon]

51 minutes ago, tater said:

Designed to maximally rip off the taxpayer.

Can you just like, stop? I, and other do not care that you do not like SLS, or Gateway, or any related hardware. So can you please stop moaning about it, and let us enjoy having a BEO crew rated spacecraft and LV for the first time in 45 years.

[tater]

19 minutes ago, Barzon said:

Can you just like, stop? I, and other do not care that you do not like SLS, or Gateway, or any related hardware. So can you please stop moaning about it, and let us enjoy having a BEO crew rated spacecraft and LV for the first time in 45 years.

We don't have either of those yet. We will in---what year is Artemis II flying? My post was about their lander being flown on SLS.

The ripoff here is Boeing---a company that has been mismanaging SLS up to this point (according to the NASA OIG, don't take my word for it)---is proposing using their overpriced core stage (they're getting a development style cost-plus contract for what should be fixed price production cores), and their upper stage (which should have been given to "not Boeing" due to the OIG report above, but which is being given without bidding to... Boeing) to launch their lander.

Sorry, that's a rip off, period.

The call for landers should ALLOW an SLS launch all-up if they want, but should also REQUIRE that the vehicle also be able to have assembly in space from 2 or more components. Any accounting that doesn't make the launch cost of the Boeing proposed lander be multiple billions is very simply lying via unethical accounting. The cost of launching that lander is at bare minimum ~2.25 billion dollars (2 launches in 1 year with nominal program costs of 2.5B$, and marginal launch costs of 1 B$ for each of 2 flights that year (which is pretty optimistic). That doesn't even count the billions in dev cost, so I'm gifting them many billion there.

[tater]

Note that any 2-3 stage lander that comes in at 37t or under could in fact be sent with SLS all-up if doing so is more cost effective, or if it was to be determined that for some reason that assembly at Gateway was too long a pole. Definitionally the Boeing lander is in that mass class (because it can be sent to TLI by SLS). The others are aiming at components that are in the 15t or under range, so they tend to be 2 stage, plus a transfer stage.

If they dumped EUS in favor of a Block 1b with ACES (full diameter), then they might be able to do Artemis without Gateway at all, in 2 SLS launches. Send the lander to LLO directly, and send Orion directly on Block 1b since the ACES would have ~11 ton residual props after TLI, which is just shy of 1200 m/s pushing the stage plus Orion (which is about what Apollo LOI burn was). So ACES puts Orion in LLO, leaving the Orion SM with enough dv to get home.

Hence ripoff, again. ACES would be far superior to EUS, yet they're throwing EUS to Boeing.

[jadebenn]

1 hour ago, tater said:

The call for landers should ALLOW an SLS launch all-up if they want, but should also REQUIRE that the vehicle also be able to have assembly in space from 2 or more components.

What!? Why?

The whole point of this RFP was to allow companies the flexibility to propose solution as they see fit. I though this was near universally-agreed to be a good thing. Now you want to artificially constrain them?

[tater]

25 minutes ago, jadebenn said:

What!? Why?

The whole point of this RFP was to allow companies the flexibility to propose solution as they see fit. I though this was near universally-agreed to be a good thing. Now you want to artificially constrain them?

Because I have no confidence that the costing of SLS launches will be properly considered.

It's quite simple: If the Boeing suggestion were to win any contract using SLS as the LV, then the process is demonstrably broken by definition.

There is no possible reality where SLS is more cost effective than literally any other solution, including building the lander from parts using EVAs at ISS, frankly. (I'm only half-joking here, sadly)

Any use of SLS to compete with commercial vehicles needs to do the same sort of accounting. If BO and SpaceX and even ULA need to amortize development costs at any level by having a margin above their cost that they sell launches for, SLS should be priced the same way for the purpose of price competition with commercial providers. If SpaceX makes 20% on a FH launch, then figure out the total actual cost of SLS (2.25B$ or more assuming 2 in a year and a lowball marginal price), and add 20%.

If their pitch is that it's more reliable, so worth an extra 2 billion per lander launch (that's a lowball, too), then I want to see the work where they conclusively demonstrate the cost-benefit of doing that.

Edited November 6, 2019 by tater

[tater]

So to be clear, if you can show that NASA and Boeing can be completely transparent in costing the lander SLS launch in a way that no one can argue---or if they will then subsidize the competing LVs to the same degree that SLS has and will continued to be subsidized by NASA labor then fine.

Basically that entire "program cost" divided by the number of launches including 2021+ (not counting the first one as a gimme), so if they have 1 launch in 2022-2023, and 2 in 2024 (the lander mission in question), then they need to take the 3 years of program costs, and divide by the 3 launches (Artemis II, Artemis III, and IIIb (lander)), we'd get the marginal cost of the launch, plus 2.5 billion for each.

That makes each launch cost more like 3.something billion.

If going forward they launch 2 every year, that drops to (2.5+1.X+1.X)/2 ~= 2.25 B$ each. If they can drop the marginal cost under a billion, it can get closer to 2 B$ each.

[Ultimate Steve]

12 hours ago, sh1pman said:

NASA had a chance to lower the SLS rocket’s cost—but it stuck with Boeing

https://arstechnica.com/science/2019/11/nasa-rejects-blue-origins-offer-of-a-cheaper-upper-stage-for-the-sls-rocket/

There are good reasons, though.

NASA sets out three reasons for not opening the competition to Blue Origin. In the document, signed by various agency officials including the acting director for human spaceflight, Ken Bowersox, NASA says Blue Origin's "alternate" stage cannot fly 10 tons of cargo along with the Orion spacecraft.

Moreover, NASA says, the total height of the SLS rocket's core stage with Blue Origin's upper stage exceeds the height of the Vertical Assembly Building's door, resulting in "modifications to the VAB building height and substantial cost and schedule delays." Finally, the agency says the BE-3U engine's higher stage thrust would result in an increase to the end-of-life acceleration of the Orion spacecraft and a significant impact to the Orion solar array design.

 

[tater]

The reasons they give are reasonable, true. But they don't address the cost to deal with them vs the cost of using EUS, so we honestly don't know which makes more sense.

Ie: it might be cheaper to alter the VAB than develop and build EUS. The schedule delays might be real, but what are the Boeing Schedule delays going to be? Also, fixed cost vs cost plus? It might not be expensive to strengthen the solar panels, or the Be-3U could be throttled down, etc.

If the BO stage is just the extant 7m stage, that's one issue, but what about an 8m stage? It's not like EUS already exists. A paper 8.4m BO stage is just as paper as a Boeing one (well, unfair there, Boeing has actually built real rockets, unlike BO, so I'll take that back).

What about ULA? Has ULA bid on an upper stage? Sort of hard for ULA to bid against Boeing on anything, since they are tied together. So no giant Centaur with ACES (which would be better than EUS in every way, and might make SLS approaching useful, even if grossly overpriced).

The question on the BO upper stage is that could it improve margin on something more like Block 1? Say it's far cheaper than EUS, but only comanifests 5 tons. 10 tons isn't enough to be useful anyway for a start. Also, will the BO upper stage have their version of ZBO, or other life extension? Could a BO stage 2 have enough prop reserves after TLI to do an LOI burn for Orion, for example? Then we get a cheaper SLS that can place Orion directly in LLO (the Orion CSM can return from LLO, it is limited only if it needs to do the LOI burn AND the TEI burn (which limits it to a higher orbit)).

Alternately, let's concentrate on making SLS useful for the only BLEO use of Orion that requires a LV with SLS level C3---the Moon. Orion as a return vehicle for a crew Mars mission needs no SM to speak of, and certainly doesn't need SLS, any Mars craft will be far too large to build at Gateway (because the LV to deliver the propellants for it would have to make SLS look like a toy, it's the lander problem up thread, but the lander is a a Mars lander. And Orion. And a spacecraft fro the crew to live in for a year or more. And props for MOI. And the props for TEI from mars. Hence Mars vehicle built with SLS is built in LEO. So for SLS to be useful at the Moon, Orion needs a decent SM so it can do LLO. That means a version of Block 1 that can loft a slightly heavier Orion SM. Just a few tons more propellants, and the Orion CSM would be capable of LOI and TEI from LLO with margin. A better Block 1 would in fact be quite useful.

Comanifesting cargo? Meh. If they can't fit a lander, it doesn't matter. Anything small enough to fly on Block 1b might as well go with FH.

 

Edited November 6, 2019 by tater

[jinnantonix]

19 hours ago, sevenperforce said:

I'm thinking two tanks for the transfer burn and two tanks for the descent.

@sevenperforce I have modeled this in KSP.  Theoretically the FH (expendable) upper stage with Merlin 1D Vac has enough propellant to get a 32t payload to TLI, and then another approx 1000m/s for the LOP-G rendezvous.  However, the FH (expendable) completes its second stage burn about 500 m/s shy of LEO.  Add to that the need to burn some fuel on establishing the rendezvous.  The margins are very tight.  IMHO the margins are too tight.  

The previous design with the 3 drop tank pairs had about a 10-20% margin on remaining propellant.  I am not sure even that is viable from a safety perspective.  The design assumes the AJ10-190 OMS burns 7 times for a total of about 2 hours, over a mission time of just under 3 weeks.  On paper the OMS can do it, but if the engine fails during the lunar descent or ascent, the crew dies.  I still think redundancy is needed.

Regarding Boeing's submission

Solar panels, side mounted - sure they are landing at the south pole.  But there is no sun in Shackleton Crater.

I think Boeing are having a lend if they think NASA will accept a design the requires SLS Cargo as a launch vehicle.  Too expensive.

Edited November 6, 2019 by jinnantonix

[tater]

4 hours ago, jinnantonix said:

Solar panels, side mounted - sure they are landing at the south pole.  But there is no sun in Shackleton Crater.

No one is landing inside a crater that is permanently shadowed. That is not, and has never been part of Artemis. The idea is to land in the areas that have nearly constant sunlight that are next to the dark polar craters. The sunlight (coming from the side) is the point.

 

The 4 lettered areas are possible sites where the area is illuminated 80% of the time.

 

 

Edited November 6, 2019 by tater

[wumpus]

25 minutes ago, tater said:

No one is landing inside a crater that is permanently shadowed. That is not, and has never been part of Artemis. The idea is to land in the areas that have nearly constant sunlight that are next to the dark polar craters. The sunlight (coming from the side) is the point.

While the 80% solar areas are great, I wouldn't ignore the ability to not require any shielding from solar radiation (especially flares and similar).  Even if Artemis can't do it, I would still expect any longer-term habitats to go "where the Sun don't shine".

[tater]

6 minutes ago, wumpus said:

While the 80% solar areas are great, I wouldn't ignore the ability to not require any shielding from solar radiation (especially flares and similar).  Even if Artemis can't do it, I would still expect any longer-term habitats to go "where the Sun don't shine".

Any long term lunar hab will be buried. Being where the sun isn't shining is not the problem, dark craters still see GCRs, though with the horizon higher, so instead of a 50% reduction vs deep space, they maybe see a 60%+ reduction.

Artemis will not land people in the dark. That's not a thing, and never has been a thing.

Edited November 6, 2019 by tater

[sevenperforce]

Does anyone have an actual copy of the Boeing proposal or has it not been leaked? I can FOIA it.

[sevenperforce]

Very little can be apprised of the Boeing proposal from the images and presser alone, so I am hoping the full proposal leaks some time soon. A few things can be gleaned, though. It's possible that the apparent single engine on the EUS is just a trick of perspective and that the other three nozzles are simply obscured (and that the artist forgot to apply their firing layer). This is definitely the EUS, however.

I assume these are the methalox engines already in dev by Intuitive Machines. Thus we see eight on the upper stage and likely twelve on the lower stage. Also, there is no central engine on either stage. The square print on the underside of the ascent stage in the landed image and the apparent descent ladder on the lower stage shows that there is an airlock in the lower compartment, reducing mass on the ascent vehicle and greatly simplifying egress. I had proposed this several times in the past. Additionally, the absence of engines underneath means cargo can be delivered in that way as well.

If we go by very rough estimates and say that the EUS can throw 37 tonnes to TLI, and we assume methalox isp on the order of 370 seconds, and you need a touch over 3 km/s to get from TLI to Gateway to LLO to the lunar surface, then that landing stage will burn 21 tonnes of props getting to that point. Guesstimating lower stage dry mass is completely speculative, especially given that we have the lower hab module and any potential amount of cargo, but if we wildly guess that total landed mass on the descent stage is 30% of propellant mass (The Apollo descent stage dry mass was 25% the mass of the propellant it carried), then that puts it at 6.3 tonnes down, leaving an ascent stage at 9.7 tonnes. With the same 370-second isp (and that number could be off-base) and 2.6 km/s needed to get back to LOP-G, that gives us an ascent stage dry mass of 4.7 tonnes, which is quite respectable.

 

Edited November 6, 2019 by sevenperforce

[tater]

Boeing has pushed a lander design with the airlock and crew stuff in the descent stage for a while now, this is just the latest version of that.

Given that the 2 stages must therefore have a hatch between them, it seems like making it a docking mechanism (so you could later attach a replacement Descent stage, if you could refill props on the Ascent stage) would mean that there is no reason the lander could not in fact be sent in 2 parts.

[sevenperforce]

7 hours ago, jinnantonix said:

I have modeled this in KSP.  Theoretically the FH (expendable) upper stage with Merlin 1D Vac has enough propellant to get a 32t payload to TLI, and then another approx 1000m/s for the LOP-G rendezvous.  However, the FH (expendable) completes its second stage burn about 500 m/s shy of LEO.  Add to that the need to burn some fuel on establishing the rendezvous.  The margins are very tight.  IMHO the margins are too tight.  

The previous design with the 3 drop tank pairs had about a 10-20% margin on remaining propellant.  I am not sure even that is viable from a safety perspective.  The design assumes the AJ10-190 OMS burns 7 times for a total of about 2 hours, over a mission time of just under 3 weeks.  On paper the OMS can do it, but if the engine fails during the lunar descent or ascent, the crew dies.  I still think redundancy is needed.

Assuming the stack is dropped in proper LEO, it needs 3.63 km/s to get all the way to LOP-G. My calculations give a naked FHeUS 44.8 tonnes of residuals in LEO, but FHe is quoted as being able to deliver up to 63.8 tonnes of payload to LEO, so my number may be low. Even with my numbers, however, a naked FHeUS can take 18.9 tonnes all the way to LOP-G. If a naked FHeUS has closer to 55 tonnes of residuals, on the other hand, it would be able to deliver about 25 tonnes to LOP-G.

10-20% margin on remaining propellant is way, way too high. You only need 3-5%, really. Tops.

1 minute ago, tater said:

Boeing has pushed a lander design with the airlock and crew stuff in the descent stage for a while now, this is just the latest version of that.

Given that the 2 stages must therefore have a hatch between them, it seems like making it a docking mechanism (so you could later attach a replacement Descent stage, if you could refill props on the Ascent stage) would mean that there is no reason the lander could not in fact be sent in 2 parts.

Oh but if you do that then there's no reason to use SLS in the first place, which means Boeing doesn't get as much money.

[tater]

19 minutes ago, sevenperforce said:

Oh but if you do that then there's no reason to use SLS in the first place, which means Boeing doesn't get as much money.

They are part of ULA... what's the throw of Vulcan Heavy to TLI?