totm dec 2023 Artemis Discussion Thread

[RCgothic]

The capability options were interesting but additional complexity for greater performance isn't going to help with the cost, cadence, comanifest volume or schedule issues.

[tater]

The problem from the start with SLS remains. It is a "rocket to nowhere."

Making a jack of all trades (master of none!) SHLV (or BLEO vehicle) would be an entirely reasonable path to take. But to be a "jack of all trades" you must first at the very least define the trades it needs to be able to work.

It was always fancifully pitched as a Mars vehicle (as was Orion), but in order for that to be the case, it would have had to be designed from the start to have an extremely high cadence. If you need to launch (making a number up) 6 times within a period where LH₂ is at risk of boiling off, then we're talking about a flight every few weeks. Minus that cadence, it's not a master of all trades, it's useless.

For cislunar, what's the mission? If the mission is to visit a useless, distant lunar orbit every year or two? Yipee. If we all agree the only reason to be near the Moon is to be ON the Moon, then it must be capable of either the rapid cadence to fly 2 simultaneous flights, or it must be capable of flying the desired mission single stack. Saddled with Orion, this means ~70t to TLI as a minimum requirement.

The cadence was never plausible, the throw was always too low, even ignoring cost. Add in that it's about an order of magnitude more expensive than it should be per flight... and it's worse than useless.

[sevenperforce]

6 hours ago, RCgothic said:

The capability options were interesting but additional complexity for greater performance isn't going to help with the cost, cadence, comanifest volume or schedule issues.

Oh, certainly. Obviously none of these things can be done, not at this point. I was just musing about what omelets could have been made if we had been willing to break a few eggs.

The comanifested volume is one of the big issues. The Altair lander would have been 10 meters wide and 15 meters high. Of course the descent stage would have been hydrolox which drove up volume, but even with a denser propellant choice I don't see how you can fit a lander underneath Orion on SLS.

[tater]

1 hour ago, sevenperforce said:

Oh, certainly. Obviously none of these things can be done, not at this point. I was just musing about what omelets could have been made if we had been willing to break a few eggs.

The comanifested volume is one of the big issues. The Altair lander would have been 10 meters wide and 15 meters high. Of course the descent stage would have been hydrolox which drove up volume, but even with a denser propellant choice I don't see how you can fit a lander underneath Orion on SLS.

Yeah, the fundamental fail was rebranding Constellation without the same capabilities. That program explicitly treated Ares V (~SLS) as cargo-only.

[Exoscientist]

 Will the Artemis lunar landing missions cost $7 billion per mission?

Amortized SLS per flight: ~$2 billion
Amortized Orion capsule per flight: ~$2 billion
Starship HLS, $2.9 billion for two missions: > $1 billion per mission.
Boeing EUS: space writer Eric Berger estimates it as close to $1 billion per flight:

Eric Berger @SciGuySpace
The story also uses a simple cost estimator model to put a per-unit price on Boeing's
Exploration Upper Stage. The result? $880 million. That's just the upper stage, and doesn't
include the core stage, SRBs, integration, ground systems, etc. etc.
9:41 AM · Nov 5, 2019
https://twitter.com/SciGuySpace/status/1191727259834601472?s=20&t=_8rb3a-Fx5m15SVm6dMsqA

 According to this graphic the total cost of the Beoing EUS is ~$10 billion, so ~$1 billion per flight is certainly likely 

Lunar Gateway: This article suggests $4 billion just for the initial components, so ca. $1 billion per mission is likely:

Is America’s next big leap off earth just a ‘moondoggle’?
Space experts and the industry are divided over the Lunar Gateway, NASA’s planned space
station for the moon.
By JACQUELINE FELDSCHER 06/13/2019 04:57 AM EDT
The Gateway is expected to cost almost $4 billion through fiscal 2024 – which is basically just a down payment on the first two modules. That does not include the Space Launch System rocket and Orion vehicle, which are expected to cost $10 billion and $5.8 billion respectively over five years.
https://www.politico.com/agenda/story/2019/06/13/nasa-lunar-gateway-000898/

 Therefore a conservative estimate for the total cost is $7 billion per lander mission.

  Robert Clark

 

Edited November 19, 2023 by Exoscientist

[tater]

Why the black text? Can't read it dark theme

[sevenperforce]

On 10/20/2023 at 8:01 PM, sevenperforce said:

On 10/20/2023 at 12:52 PM, RCgothic said:

The capability options were interesting but additional complexity for greater performance isn't going to help with the cost, cadence, comanifest volume or schedule issues.

Oh, certainly. Obviously none of these things can be done, not at this point. I was just musing about what omelets could have been made if we had been willing to break a few eggs.

The comanifested volume is one of the big issues. The Altair lander would have been 10 meters wide and 15 meters high. Of course the descent stage would have been hydrolox which drove up volume, but even with a denser propellant choice I don't see how you can fit a lander underneath Orion on SLS.

Revisiting this quickly just for general edification.

The Altair lander was a two-stage design with a hypergolic ascent stage using an AJ-10, a cryogenic stage using a pressure-fed deep-throttling RL-10 for lunar orbit insertion and lunar surface descent, and a separate airlock and cargo capability on the descent stage:

Total mass launched to TLI would have been 45.9 tonnes, plus the 26.5 tonnes of Orion. The pressure-fed version of the RL10 was built and test-fired as the CECE demonstrator and was capable of throttling down to 8% with a max specific impulse of 445 seconds. To brake the entire stack into low lunar orbit, Altair would have needed to burn 13.5 tonnes of hydrolox. Orion would have detached, leaving the weight of the sortie vehicle at 32.4 tonnes; it would need to burn another 11.3 tonnes of hydrolox to get down to the lunar surface. So in total, Altair would have had a propellant capacity of roughly 24.8 tonnes of hydrolox.

Schematically, the Altair lander had a central CECE RL10 with a height of 1.53 meters and a diameter of roughly 1 meter (allowing for gimbal); it also shrouded the ascent vehicle's AJ-10 engine. Thus the hydrolox all would have been in the octagonal outer envelope surrounding these two engines: 

The RL10 has a mixture ratio of 5.88:1, meaning that those 24.8 tonnes of hydrolox include 3.6 tonnes of liquid hydrogen, occupying 50.71 cubic meters, and 21.2 tonnes of LOX, occupying 18.58 cubic meters. The portion of the vehicle containing the tanks can be approximated as a right toroidal cylinder with an external diameter of 10 meters, an internal diameter of 3.33 meters, and a height on the order of 5.2 meters (based on some rough pixel counting). In total that's a "tank carrying volume" of 363 cubic meters, giving a volumetric efficiency of ~19.1% (not unexpected given how fluffy hydrogen is and how suboptimally the tanks need to be arranged here).

So if we somehow had a beefed-up SLS capable of launching an arbitrary payload to TLI, what could we fit in the 10-meter-high, 8.4-meter-wide space under Orion on top of the EUS?

Let's start by making this a slimmed-down sortie lander, assuming pre-emplaced surface assets. First let's get a better grip on Altair.

NASA's assumptions about an Artemis ascent stage assume 9-12 tonnes, including extra props to get to NRHO, and we know Altair was bloated, so let's take Altair's ascent vehicle (with props) as 9.5 tonnes. This means the dry mass of the descent stage was 11.6 tonnes. Subtract the airlock and we save around 1.5 tonnes, dropping the descent stage to 10.1 tonnes dry mass. Remove the 350-kg engine and the expected 500 kg of unpressurized payload to the surface, and we get ~9.2 tonnes for bare structure, landing legs, and tankage weight. So the structural-and-tankage ratio here is a disappointing 2.7:1. These tanks are roughly the same diameter as the tanks on Centaur (which boasts a propellant fraction of 10:1) and held 20% more props, but these would also have been pressure-fed tanks with five times as many hemispheric bulkheads, so I'd ballpark the tankage weight on Altair at 2.5x that of Centaur or 5.2 tonnes. This suggests that the load-bearing structure, RCS propellant, RCS thrusters, radiators, landing legs, and so forth all come in at 4 tonnes.

What would an Altair built for SLS look like, then? Well, the minimal-mass ascent vehicle would come in at 9 tonnes...except that we only need to go to LLO, not all the way to NRHO, so we only need 1,870 m/s instead of 2,600 m/s. At 316 seconds of specific impulse, a 9-tonne ascent vehicle would burn 5.12 tonnes of propellant to develop 2,600 m/s of Δv, whereas we will only need ~3.23 tonnes of props, meaning our ascent vehicle need only have a wet mass of 7.11 tonnes. Let's borrow @Exoscientist's idea and use the old discontinued Standard-sized Cygnus as our pressure vessel OML, making it 5.75 meters high including the engine (3.65 meters without). Let's ditch any extensible solar panels and have it run on batteries and fixed panels after detaching from the descent module. Since we are using hypergols, we can fit our 2.2 tonnes of dinitrogen tetroxide in a pair of 1.2-meter external spherical tanks placed as far forward as possible, and fit our 1.3 tonnes of hydrazine (let's take extra for monoprop RCS) in four 0.9-meter spherical tanks clustered underneath:

 

This gives us a good estimate for the absolute maximum amount of volume (purple) we can allocate to a lander and descent stage. The cylindrical region below the engine bell has a volume of approximately 188 cubic meters, the toroidal region surrounding the engine bell has a volume of approximately 95 cubic meters, and the available volume on either side of the ascent vehicle (leaving the front and back open for both thrust balance and egress) is roughly 21 cubic meters, for a total available structural volume of 304 cubic meters. 

Of course the descent vehicle needs engines. To conserve space (and perhaps utilize a little of the wasted volume around the top of the EUS tank), let's use four BE-7 engines (two would do it, but we need redundancy). With gimbal allowance, each of those takes up a height of 80" and a diameter of 47.4", but since we can let them hang down slightly around the top of the EUS tank we only end up losing about one cubic meter of volume for each engine, bringing us to ~300 cubic meters of structural volume.

Setting volumetric efficiency at 19.1% as before, this gives us 57.3 cubic meters of usable tank volume. Bulk density of hydrolox is on the order of 0.365 g/cc, so that's just under 21 tonnes of propellant. Tankage weight will be a little better than Altair since we aren't using pressure-fed engines -- let's say 3.7 tonnes. If the load-bearing structure of Altair was 4 tonnes for 363 cubic meters of structural space, then we can ballpark the load-structure here at 3.3 tonnes. All four engines together probably come to something like 300 kg. So we have a descent stage with 21 tonnes of propellant and a dry mass of 7.3 tonnes, and an ascent stage with a total wet mass of 7.1 tonnes.

So let's throw this to TLI on a magically-beefed-up SLS and see what happens. With this approach, Orion doesn't need quite as much propellant since it will be braked into LLO by the lower stage. Orion ordinarily carries 8.6 tonnes of props, but we can drain about 2.5 tonnes out and still have enough for the return voyage (there really won't be a meaningful decrease in dry mass with this approach) to make it lighter. Total injected mass to the moon is 59.4 tonnes. Ballparking the BE-7 at 449 seconds specific impulse, the lower stage burns through 11.1 tonnes of hydrolox to brake itself and Orion into low lunar orbit. Orion detaches, leaving the lander at a wet mass of 24.3 tonnes with just 9.9 tonnes of hydrolox remaining. Fortunately, that gives it 2.3 km/s of Δv, which is sufficient margin for reaching the lunar surface. We can probably shave a total of 2.5 tonnes of hydrolox off and still make it work, reducing our lunar injection mass to 56.4 tonnes.

Cramped, but doable...if it was possible to make an SLS Block 2+ with 10-15 tonnes more capability than SLS Block 2, which is already never going to happen.

[tater]

56.4t. Nice.

Interesting as in this thread, and the SLS/Orion thread, and the previous (closed) SLS thread the math was done less definitively and came up with similar ballpark figures (60-70t as the min TLI capability for any single-stack lunar mission including Orion.

If the dry mass of an expended Starship could be brought down to ~49t, SS/SH could send this Orion/lander stack to TLI. This is very possible I think as the steel/engine mass of the tank section is ~35t (3.6mm 304L). Assuming the upper part is staged off as a fairing (would be oddball since Orion needs to stick out on top with LES. Bottom line is that SS/SH can almost certainly send an Orion based system to the Moon in 1 stack, no refilling required.

Alternately, send the lander in SS, Orion on NG, dock, then send it.

[Exoscientist]

2 hours ago, sevenperforce said:

Cramped, but doable...if it was possible to make an SLS Block 2+ with 10-15 tonnes more capability than SLS Block 2, which is already never going to happen.

The SLS Block 2 is supposed to be able to make 130 tons to LEO. Surely that should be enough for any reasonably sized lunar lander.

  Bob Clark 

[RCgothic]

1 minute ago, Exoscientist said:

The SLS Block 2 is supposed to be able to make 130 tons to LEO. Surely that should be enough for any reasonably sized lunar lander.

  Bob Clark 

Saturn V could manage upwards of 140t to LEO and did not have a reasonably sized lander.

[Exoscientist]

 For the Saturn V the metric tons to LEO was 118 tons. For a manned lunar architecture, just getting there is the important thing since you can send cargo and surface habitats by cheaper commercial launchers.

  Bob Clark 

[tater]

15 minutes ago, Exoscientist said:

 For the Saturn V the metric tons to LEO was 118 tons. For a manned lunar architecture, just getting there is the important thing since you can send cargo and surface habitats by cheaper commercial launchers.

From Apollo 11 transcript:

PAO: "This is Apollo Control. The Canary Island station has acquisition of Apollo 11 now. We'll continue to stand by live for any air-to-ground communication. We're showing an orbital weight of the combined vehicles of 297,914 pounds." That's 135,131 kg.

Subsequent flights got better numbers.

24 minutes ago, Exoscientist said:

The SLS Block 2 is supposed to be able to make 130 tons to LEO. Surely that should be enough for any reasonably sized lunar lander.

Nope.

The throw to TLI is not enough with Orion as part of the stack. Not even close. The math @sevenperforce just did above shows what a sortie lander could maybe be slashed down to saddled with Orion, and that's >10t more than B2 can do. On top of that, that is for a sortie lander—so you still need to send a habitat, rover, etc ahead. All that for the low-low price of a gajillion dollars and maybe another decade or two!

[Exoscientist]

4 hours ago, tater said:

From Apollo 11 transcript:

PAO: "This is Apollo Control. The Canary Island station has acquisition of Apollo 11 now. We'll continue to stand by live for any air-to-ground communication. We're showing an orbital weight of the combined vehicles of 297,914 pounds." That's 135,131 kg.

Subsequent flights got better numbers.

Nope.

The throw to TLI is not enough with Orion as part of the stack. Not even close. The math @sevenperforce just did above shows what a sortie lander could maybe be slashed down to saddled with Orion, and that's >10t more than B2 can do. On top of that, that is for a sortie lander—so you still need to send a habitat, rover, etc ahead. All that for the low-low price of a gajillion dollars and maybe another decade or two!

 

 The phrasing there leads me to believe that includes the dry mass(empty mass) of the S-IVB third stage as well as fairings and adapters. This page gives the LEO payload as 118 metric tons:

10 Things: Rockets We Love-Saturn V.
3—...and Busloads of Thrust
Stand back, Ms. Frizzle. The Saturn V generated 7.6 million pounds (34.5 million newtons) of thrust at launch, creating more power than 85 Hoover Dams. It could launch about 130 tons (118,000 kilograms) into Earth orbit. That's about as much weight as 10 school buses. The Saturn V could launch about 50 tons (43,500 kilograms) to the Moon. That's about the same as four school buses.
https://solarsystem.nasa.gov/news/382/10-things-rockets-we-love-saturn-v

 The dry mass, i.e., empty mass, of the S-IVB third stage was 13.5 tons:

https://en.m.wikipedia.org/wiki/S-IVB

 As shown in this diagram that mass reaching orbit would also include the fairings and adapters.

 

 About the mass of the capsule of our proposed commercial Moon rocket, we definitely would not use the Orion capsule. It is far too expensive, as ca. ~$2 billion per flight alone. 

 As I recall the total payload mass sent to TLI for Apollo including the command module and fully fueled service module and lunar lander was in the range of 43 metric tons.

  Bob Clark

[tater]

35 minutes ago, Exoscientist said:

 The phrasing there leads me to believe that includes the dry mass(empty mass) of the S-IVB third stage as well as fairings and adapters. This page gives the LEO payload as 118 metric tons:

Doesn't matter, that's what's in orbit. 297,914 pounds. And before the TLI burn. The dry mass matters, as the vehicle accelerates all of that to TLI, then they sep the CSM, dump the petal fairings, extract the LM, etc. It's all sent to TLI. If the bulk of SLS "payload" to LEO is in fact residual props in the upper stage, why would you neglect the stage required to hold those props, and the engines needed to sent it to TLI?

[Exoscientist]

 Payload does not mean the dry mass of the upper stage or fairings. For our proposed commercial Moon rocket we would also get higher mass to orbit if we included the dry mass of the upper stage and fairings and adapters.

  Bob Clark

[tater]

3 hours ago, Exoscientist said:

 Payload does not mean the dry mass of the upper stage or fairings. For our proposed commercial Moon rocket we would also get higher mass to orbit if we included the dry mass of the upper stage and fairings and adapters.

Payload in that case is released.

For a Moon rocket, where S2 (or S3) is doing the TLI burn, the LEO payload is meaningless as the CSM/Lander are not staged off in LEO. The total mass in LEO, and the residual props in the upper stage are what matters, as that determines what it can send to TLI. So unless SLS dumps all XXX tons of payload in LEO, minus the ICPS/EUS/whatever, LEO payload doesn't do anything, TLI payload.

Unless TLI payload is north of 60-70t (or 56.4t for that bare bones sortie lander concept), nothing else about SLS matters. It can do that, or it's a waste of time and money.

Edited October 24, 2023 by tater

[Exoscientist]

 The “Angry Astronaut” on the Artemis program’s financial troubles. Once more he slays all the sacred cows:

 

 

  Bob Clark

[RCgothic]

Agreed with most of what he said up until 17mins or so.

I like ALPACA, but not more than either of the other two HLS solutions and I'm willing to wait. In my opinion getting back to the moon fast is not as important as getting back in a cost effective and scalable fashion, so I'm not especially upset ALPACA wasn't picked looking at the potential of the ones that were.

Secondly, Starship doesn't need to be fully crew-rated for Earth entry descent and landing to perform creed lunar missions, and that wouldn't hold things up particularly in the absence of SLS. The crew can meet Starship in LEO via Dragon/Starliner, or even putting Orion atop an expendable Superheavy is something that looks comparatively quick and cheap. With Orion's oversized LAS and Superheavy's potentially high flight rate (even expendably), getting Superheavy crew-rated for launch really wouldn't that big a deal.

Thirdly, even ALPACA would have needed orbital refuelling. This is just something we need to master, and he doesn't even doubt we will. So once again my disagreement is on schedule not feasibility, and so each to their own.

[tater]

1 hour ago, Exoscientist said:

 The “Angry Astronaut” on the Artemis program’s financial troubles. Once more he slays all the sacred cows:

I've never managed to watch one of his videos, not gonna start now. The cost issues of SLS/Orion have been known since forever, though, and the OIG reports are clear. It's been beaten to death in this thread, and the SLS thread that even just the marginal launch costs of SLS are absurd, much less the all-in costs that amortize development. The real cow that needs slaying—or dead horse that needs pounding into a pulp—is that it's a rocket to nowhere. It lacks the capability to do any useful mission in 1 launch by itself. Block 2 is not going to happen, and even if it did, it adds no useful capability. Block 1B adds the useful capability of improved operations, but at an additional cost per launch of a few hundred million bucks for EUS (not counting dev cost and ML-2 cost). The operational improvements are huge compared to the Artemis I flight, since the vehicle can circularize the paring orbit, and phase the TLI burn, vs the narrow windows requires with ICPS. Still, you get Orion in a distant lunar orbit, yipee. Additional throw to TLI is pretty useless, maybe give Orion a better SM... except that likely takes a long, long time to develop, changes the ML, etc.

11 minutes ago, RCgothic said:

I like ALPACA, but not more than either of the other two HLS solutions and I'm willing to wait. In my opinion getting back to the moon fast is not as important as getting back in a cost effective and scalable fashion, so I'm not especially upset ALPACA wasn't picked looking at the potential of the ones that were.

The point is sustainable missions, and SLS won't ever do that. I agree refilling is critical to anything interesting in cislunar. Once refilling is demonstrated and operational, SLS is obviated anyway, and a EOR architecture can replace it.

So I agree, any complaints about refilling stages have as the only possible solution single stack alternatives (or a single stack for the CSM, and another for the lander). In the latter case, Starship can certainly send some other lander to TLI in 1 launch, expended.

[Exoscientist]

On 10/24/2023 at 11:04 AM, RCgothic said:

Agreed with most of what he said up until 17mins or so.

I like ALPACA, but not more than either of the other two HLS solutions and I'm willing to wait. In my opinion getting back to the moon fast is not as important as getting back in a cost effective and scalable fashion, so I'm not especially upset ALPACA wasn't picked looking at the potential of the ones that were.

Secondly, Starship doesn't need to be fully crew-rated for Earth entry descent and landing to perform creed lunar missions, and that wouldn't hold things up particularly in the absence of SLS. The crew can meet Starship in LEO via Dragon/Starliner, or even putting Orion atop an expendable Superheavy is something that looks comparatively quick and cheap. With Orion's oversized LAS and Superheavy's potentially high flight rate (even expendably), getting Superheavy crew-rated for launch really wouldn't that big a deal.

Thirdly, even ALPACA would have needed orbital refuelling. This is just something we need to master, and he doesn't even doubt we will. So once again my disagreement is on schedule not feasibility, and so each to their own.

 

 OK. I’m not a fan of ALPACA either. I don’t like the need for orbital refueling for a single mission. For the same reason I don’t like the Starship HLS for its even larger number of orbital refuelings. 

 Think of the Starship HLS this way: assuming fully fueled at 1,200 ton propellant load plus 120 ton dry mass, that’s 1,300 tons for just the propulsion section. In contrast the propulsion section(s) of an Apollo-sized lander would only be ca. 13 tons. The Starship HLS would weigh 100 times more than needed for a manned lander to the Moon. 

 Granted, it could carry a larger passenger section and more payload, but we don’t need it for that purpose. The plan is for it to only carry four crew to surface anyway. And existing commercial launchers could carry any needed cargo or habitats one-way to the Moon.

  Bob Clark

Edited November 2, 2023 by Exoscientist

[RCgothic]

4 crew every other year isn't one tenth as ambitious as either HLS option could reasonably support.

And once again the problem is SLS/Orion.

[tater]

1 hour ago, Exoscientist said:

OK. I’m not a fan of ALPACA either. I don’t like the need for orbital refueling for a single mission. For the same reason I don’t like the Starship HLS for its even larger number of orbital refuelings. 

 Think of the Starship HLS this way: assuming fully fueled at 1,200 ton propellant load plus 120 ton dry mass, that’s 1,300 tons for just just the propulsion section. In contrast the propulsion section(s) of an Apollo-sized lander would only be ca. 13 tons. The Starship HLS would weigh 100 times more than needed for a manned lander to the Moon. 

 Granted, it could carry a larger passenger section and more payload, but we don’t need it for that purpose. The plan is for it to only carry four crew to surface anyway. And existing commercial launchers could carry any needed cargo or habitats one-way to the Moon.

All that matters IMO is cost.

LSS is not 120t, BTW, probably closer to 80. Steel/engine mass is maybe 52t with 3.6mm steel. If they can drop it to 3mm steel, more like 45t. That leaves room for a lot of fitting out mass.

[tater]

LSS dv based on different dry mass assumptions with 1200t of props. ~6.1 km/s is required to go from LEO to the lunar surface, and 8.1 km/s from the LEO to the lunar surface and back to LLO. 8.85km/s from LEO to the surface and up to NRHO. A round trip from LEO to the lunar surface and back to LEO is 12.2 km/s (with no aerobraking).

50t, 11.9 km/s

60t, 11.2 km/s

70t, 10.7 km/s

80t, 10.28 km/s

100t, 9.5 km/s

Note that all these mission dv values allow a flight from LEO to the lunar surface and back to Gateway.

Every mm of 304L thickness we lose drops ~11.6t from the dry mass of Starship. Not sure how thin they can go. There are loads of possibilities for alternate Starship architectures, or using a Starship variant as a hab that is never meant to leave the surface, etc.

Assuming they could do 2.6mm steel vs the 3.6mm that I think they are using now, then the dry mass is ~40t. Fitting out (descent engines up high per LSS, a crew area, etc, might get the thing up towards 60t.

A 49t (dry) version of SS, assuming 250t to LEO expended (Musk has said it's now up to ~300t expended) can go to the lunar surface with no refilling. At 300t to LEO expended, we can send a 71t SS to the lunar surface with no refilling required. This could be used to preposition assets on the surface for habitation, then go for smaller landers to deliver crew, I suppose.

 

 

[magnemoe]

17 hours ago, tater said:

Payload in that case is released.

For a Moon rocket, where S2 (or S3) is doing the TLI burn, the LEO payload is meaningless as the CSM/Lander are not staged off in LEO. The total mass in LEO, and the residual props in the upper stage are what matters, as that determines what it can send to TLI. So unless SLS dumps all XXX tons of payload in LEO, minus the ICPS/EUS/whatever, LEO payload doesn't do anything, TLI payload.

Unless TLI payload is north of 60-70t (or 56.4t for that bare bones sortie lander concept), nothing else about SLS matters. It can do that, or it's a waste of time and money.

This, CMS+lander was staged after TLI. 3rd stage had two jobs, circulate and do the TLI. Later 3rd stages was set to they impacted the moon, not sure if they staged a bit early or the 3rd stage did an secondary burn to put itself on an impact trajectory. 

[tater]

2 minutes ago, magnemoe said:

This, CMS+lander was staged after TLI. 3rd stage had two jobs, circulate and do the TLI. Later 3rd stages was set to they impacted the moon, not sure if they staged a bit early or the 3rd stage did an secondary burn to put itself on an impact trajectory. 

Yeah, S-IVB did a brief burn to circularize the parking orbit, then they checked out that everything was "go" before doing the TLI burn. The circular orbit gave them contingency to do additional work, cause they could always stay a extra orbits and adjust the start position of the burn slightly. The mass in LEO that mattered for Apollo was the total mass of the S-IVB, and everything attached to it. What that total mass was didn't matter as long as that stack had ~3.2 km/s of dv sitting there in LEO.