totm dec 2023 Artemis Discussion Thread

[tater]

When the SLS program is over, sensible people will look at the total dollars spent, and divide by the number of launches. I'd argue that only crew launches (or real missions for cargo variants) should count (so the first flight we just had doesn't count). I'd also say that some of the spending during Constellation also gets added in.

This is the only way to ascertain the true cost per mission of the program.

[sevenperforce]

On 10/17/2023 at 11:19 AM, Exoscientist said:

Instead, for a commercial replacement keep the hydrolox core stage but replace the expensive SSME’s with RS-68’s, to save on costs. Five RS-68’s would be required to enable lift-off.

 It could also be done with 7 SSME’s but the so-called “improved” versions now have bloated prices of ~$150 million per engine, compared to the original price of $50 million each. Seven of them would be a billion dollars just in engines alone.

 The RS-68 in contrast costs in the range of $10 to $20 million each, So 5 would be just $50 to $100 million. Note the RS-68 was considered for the SLS but its ablatively cooled nozzle would not handle well the extreme heating from the side boosters. But without the side boosters, that’s no longer an issue.

 A regeneratively cooled nozzle version of the RS-68 was investigated. It would raise the vacuum Isp to  ~421.5s compared to the 412s of the current version, and increase thrust about 20 tons. You would want to do this at some point to improve reusability.

 For the upper stage, we could use the Delta IV core, or the Boing EUS, or the Ariane 5 core. I favor the Ariane 5 core because it has both a high mass ratio and high vacuum ISP.

The Ariane 5 core is twice the max height of the VAB.

On 10/17/2023 at 11:19 AM, Exoscientist said:

I estimate in the range of 100 tons to LEO.

Well, let's not get ahead of ourselves.

To start with, let's do a one-to-one comparison with the current SLS to see if we have improved anything. If it's not improved, then swapping out upper stages is something that could as readily be done on the current SLS.

Taking the specifications of SLS Block 1 as follows...

	Booster	1st Stage	2nd Stage
Dry Mass, kg	102058.5	85275	3490
Propellant, kg	623689.5	952543.9	30710
Thrust, kN	17595	9116	110.1
Isp, seconds	268	452.3	462

...the Silverbird Astronautics calculator gives an estimated 97.2 tonnes to a 185x185 orbit from Cape Canaveral, with the 95% confidence interval running from 79.0 tonnes to 118.6 tonnes. That's just 2% over the actual stated LEO performance of SLS Block 1 according to NASA, so we can be pretty confident that Silverbird is giving us good results in this configuration.

So now let's redo it, but get rid of the boosters altogether and replace the four RS-25s with five RS-68s. First stage dry mass goes up by 19 tonnes (each of the RS-68s is almost twice as heavy as each RS-25), first stage vacuum thrust goes up to 17800 kN, and first stage vacuum specific impulse drops to 414 seconds:

	1st Stage	2nd Stage
Dry Mass, kg	104275	3490
Propellant, kg	952543.9	30710
Thrust, kN	17800	110.1
Isp, seconds	414	462

Running the numbers to the same orbit, we get a truly abysmal 35.7 tonnes, with a 95% confidence interval running from 28.7 tonnes to 44.6 tonnes.

On 10/17/2023 at 11:19 AM, Exoscientist said:

I have argued large solid rocket boosters are not cost effective.

I don't like large solid boosters any more than the next person, but you can't discount the heavy lifting they are (literally) doing in the SLS design.  Removing the boosters would mean removing 40% of the total propellant in the stack and you can't make up for that much of a loss with a higher-thrust core.

If we are speculating wildly about a better SLS design that would still use the same core tankage, one tempting option would be to ape the old Saturn S-1D Mega Atlas design:

One RS-25 in the center, four RS-68s on the outer thrust ring. Use Atlas Vs or Falcon 9s as side boosters and mount Delta IV common booster core oxygen tanks on top of them, so that you feed the two of the RS-68s their LOX from the side boosters. All four of the RS-68s get all of their liquid hydrogen from the core. Two of the RS-68s lose LOX and shut down at side booster separation, and then the other two shut down a minute or so later and the entire skirt is dropped once the thrust to weight ratio of the stack is optimal. No crossfeed necessary (although crossfeed would definitely improve performance). The core common bulkhead would need to be shifted slightly but the tankage mass and vehicle height would remain the same.

[tater]

1 hour ago, sevenperforce said:

The Ariane 5 core is twice the max height of the VAB.

*with the A5 core added as a new stage to SLS (know that's what you mean, just not clear)

 

1 hour ago, sevenperforce said:

ape the old Saturn S-1D Mega Atlas design

I frickin love that thing.

 

[Exoscientist]

On 10/17/2023 at 4:35 PM, RCgothic said:

The key to a cost-effective replacement for SLS is for the love of Jeb never award any more business to Boeing, Aerojet Rocketdyne or Northrop Grumman.

 

  I’m of the opinion we’re already at the stage where we no longer need the expensive, governmentally financed, “Old Space” approach to either LEO flights or beyond LEO flights.  

 I think everyone in the U.S. space community recognizes now only the New Space approach is worthwhile for launchers to LEO, as its development costs can be done for 1/10th that of the Old Space approach. Thus we have Rocket Lab and Relativity Space and others following in SpaceX footsteps to produce their own privately financed  launchers.

The European space community is just barely coming into that realization with still using the Old Space governmentally financed approach to developing the Ariane 6, Vega-C, and future launchers - and at greatly inflated and noncompetitive prices. However, fortunately they do have some small launch start-ups that will soon be coming into operation.

 So New Space certainly can do flights to LEO at reduced development costs, in the range of 1/10th those of the Old Space approach. But what’s really surprising is that now beyond LEO flights can be done for 1/100th the price of Old Space approach.

 I was led to this conclusion after a calculation that Mars Sample Return can be done at a price of 1/100th that of the NASA estimate of $10 billion, i.e., for costs in the range of only $100 million. For a cost this low the fully privately financed costs can be covered by advertising alone.

 Then following the same calculational approach, the development cost for a privately financed, commercial manned lunar program can be done instead of the tens of billions of dollars of Old Space but for only a few hundreds of million dollars. And instead of each mission costing multi billions of dollars, they can be done for less than $100 million.

   So how is it possible that beyond LEO missions, manned and robotic, can be done by commercial space at 1/100th the cost of usual governmentally financed approach?

 A little thought shows why this is possible. First, quite key is development of new components can be done at 1/10th the cost of usual approach. BUT developments of spacecraft over several decades now means we have rocket stages and spacecraft existing at a wide range of various capabilities. 

 Now remember the development costs of a spacecraft or rocket are much higher than the production cost of the individual rocket or spacecraft. But if the existing rocket or spacecraft has already existed for several years, you are no longer paying off the development cost. You are close to just paying on the production cost, plus profit, on that rocket or spacecraft. In short, by using existing rockets or stages well into their operational history, you have a price well less than having to develop them from anew.

 

  Robert Clark

 

 

Edited October 19, 2023 by Exoscientist

[Exoscientist]

9 hours ago, sevenperforce said:

Well, let's not get ahead of ourselves.

To start with, let's do a one-to-one comparison with the current SLS to see if we have improved anything. If it's not improved, then swapping out upper stages is something that could as readily be done on the current SLS.

Taking the specifications of SLS Block 1 as follows...

	Booster	1st Stage	2nd Stage
Dry Mass, kg	102058.5	85275	3490
Propellant, kg	623689.5	952543.9	30710
Thrust, kN	17595	9116	110.1
Isp, seconds	268	452.3	462

...the Silverbird Astronautics calculator gives an estimated 97.2 tonnes to a 185x185 orbit from Cape Canaveral, with the 95% confidence interval running from 79.0 tonnes to 118.6 tonnes. That's just 2% over the actual stated LEO performance of SLS Block 1 according to NASA, so we can be pretty confident that Silverbird is giving us good results in this configuration.

So now let's redo it, but get rid of the boosters altogether and replace the four RS-25s with five RS-68s. First stage dry mass goes up by 19 tonnes (each of the RS-68s is almost twice as heavy as each RS-25), first stage vacuum thrust goes up to 17800 kN, and first stage vacuum specific impulse drops to 414 seconds:

	1st Stage	2nd Stage
Dry Mass, kg	104275	3490
Propellant, kg	952543.9	30710
Thrust, kN	17800	110.1
Isp, seconds	414	462

Running the numbers to the same orbit, we get a truly abysmal 35.7 tonnes, with a 95% confidence interval running from 28.7 tonnes to 44.6 tonnes.

 Do the calculation with the Ariane 5 core as the upper stage. A commercial space approach wouldn’t use such a tiny upper stage such as the ICPS for a rocket this size when larger, appropriately-sized upper stages are available, and at about the same price.

  Robert Clark

 

[tater]

35 minutes ago, Exoscientist said:

 Do the calculation with the Ariane 5 core as the upper stage. A commercial space approach wouldn’t use such a tiny upper stage such as the ICPS for a rocket this size when larger, appropriately-sized upper stages are available, and at about the same price.

Would not fit in the VAB as he said, so why bother doing the math?

[Exoscientist]

 Actually it would. The limits for the length of an  upper stage for the SLS are based on specs for a supposed second mobile launcher for the SLS which has to be provided to the manufacturer of the mobile launcher. But that second mobile launcher has not been constructed yet, and the manufacturer has been greatly criticized for costs overruns and delays. A commercial approach would probably just choose a different manufacturer for the second mobile launcher such as SpaceX.

 The doors of the vehicle assembly building are quite huge and can accommodate the size of the SLS core and Ariane 5 as a second stage:

Vehicle Assembly Building.
There are four entries to the bays located inside the building, which are the four largest doors in the world.[14] Each door is 456 feet (139.0 m) high, has seven vertical panels and four horizontal panels, and takes 45 minutes to completely open or close. https://en.m.wikipedia.org/wiki/Vehicle_Assembly_Building#Capabilities

  Robert Clark 

Edited October 19, 2023 by Exoscientist

[tater]

42 minutes ago, Exoscientist said:

 The doors of the vehicle assembly building are quite huge and can accommodate the size of the SLS core and Ariane 5 as a second stage.

OK, assume that is true—still, Ariane 5 no longer exists. So restart production, plus build an interstage, plus new ML (current Block 1B is years behind and ~$1B)?

Aside from the fact that it would never happen (because it does not put jobs into districts), how does it help, exactly? It's 170.5t, and has a 0.6:1 TWR—without a payload on top.

The only USEFUL payload mass for SLS is ~70t to TLI. So the payload needs to be 70t, and we will assume your stage is doing the TLI burn since it's now stage 2. So the stack on top of SLS core is now 240.5t (ignoring the LES). ICPS is 32.7t.

Using the same numbers for SLS @sevenperforce did for Block 1, but with stage 2 replaced with Ariane 5 core, it gives:

Estimated Payload:  75316 kg

95% Confidence Interval:  63040 - 89794 kg

75t. Block 1 does 97.2t per Silverbird (above actual by a couple % as @sevenperforce said).

So your stage idea is less capable than current Block 1. The second idea of RS-68s is lower than B1, so it's not gonna do better with a heavier stage 2, particularly one with a low TWR having to start burning earlier because of gravity losses.

And once again, making "better" stages for SLS is meaningless unless it can accomplish a useful mission when done. The target is not "better than B1B," or even "better than Block 2." The question is, "Can an alternate SLS throw 70t to TLI?"

Nothing else matters. It can either do an all up in 1 stack lunar mission, or it can't.

 

<EDIT> The 70t to TLI assumes Orion CSM on top. You can posit "not Orion CSM" on top—but add another decade and 10s of billions of $. So yeah, that's not happening.

Edited October 19, 2023 by tater

[SunlitZelkova]

48 minutes ago, tater said:

<EDIT> The 70t to TLI assumes Orion CSM on top. You can posit "not Orion CSM" on top—but add another decade and 10s of billions of $. So yeah, that's not happening.

As long as we’re doing a total franken rocket, couldn’t we put Crew Dragon on top?

It would probably need a proper service module though, and I’m not sure how that would impact mass. And of course, a beefed up heat shield.

Considering it was originally intended to take 7 people to the ISS, IIRC, it should be far more roomier and comfortable than Orion.

[tater]

1 minute ago, SunlitZelkova said:

As long as we’re doing a total franken rocket, couldn’t we put Crew Dragon on top?

It would probably need a proper service module though, and I’m not sure how that would impact mass. And of course, a beefed up heat shield.

Considering it was originally intended to take 7 people to the ISS, IIRC, it should be far more roomier and comfortable than Orion.

A Dragon (or someday Starliner?) refit for lunar return entry velocities is fine, but at that point, why bother with SLS? Might not be possible anyway, due to the LES capability, and having to pull away from a LV with SRBs—not to mention it currently uses the trunk as part of the LES, and with trunk as SM, that likely doesn't work, either.

Keeping "Artemis," but using existing, or soon existing vehicles, sans SLS/Orion, I would go straight to EOR missions, and I'd tend to dump the idea of taking the capsule to the Moon and back. Make a more comfortable transfer vehicle, and propulsively return to LEO, where the capsule waits.

[SunlitZelkova]

4 minutes ago, tater said:

A Dragon (or someday Starliner?) refit for lunar return entry velocities is fine, but at that point, why bother with SLS? Might not be possible anyway, due to the LES capability, and having to pull away from a LV with SRBs—not to mention it currently uses the trunk as part of the LES, and with trunk as SM, that likely doesn't work, either.

Keeping "Artemis," but using existing, or soon existing vehicles, sans SLS/Orion, I would go straight to EOR missions, and I'd tend to dump the idea of taking the capsule to the Moon and back. Make a more comfortable transfer vehicle, and propulsively return to LEO, where the capsule waits.

Another plus of using a lander as a shuttle is you can make it big like Starship. I really hope SpaceX is taking into account the lesson of Apollo 13 to have a section of the spacecraft “lifeboat capable”.

That said, I don’t think returning to a dual spacecraft architecture is feasible just for safety reasons. We got lucky on Apollo 13, but there seems to be a consensus that either a single lander or “all up” direct ascent lander is best for lunar base operations. LOR was chosen because it was the fastest, not sustainable.

Basically every Soviet lunar landing concept outside of the L3 (the LOR one for the N1), which was also designed around speed and not sustainably, utilized a single lander and direct ascent, and the original Apollo was a direct ascent lander too. Post-Apollo American concepts have mostly centered around a LEO shuttle, a space shipyard, and a lunar shuttle which docks at the shipyard. Sometimes there are big monolithic landers, but these require big expendable heavy lift launch vehicles and aren’t good for a long term lunar base. And of course, then there’s Constellation.

I kind of wonder if eliminating Gateway and using Starship HLS in a direct ascent role would allow a greater payload to the surface. Travel to it in Dragon and then go directly to the surface.

I wonder if SpaceX is considering the possibility of conducting maintenance on Starship HLS’ either on the surface or in orbit.

I wonder if they have a division like Bellcom, which was the think tank NASA charged with coming up with crazy applications for Apollo hardware in the mid 60s.

Getting off this tangent… somehow Starship feels like a safer lunar transit vehicle than Orion.

[AckSed]

As long as we're throwing out ideas for a liquid-fuelled booster, how about we go really Kerbal and slap two Super Heavy boosters to the side of SLS? I think I read somewhere that its performance was approaching SRB. As a bonus, you now have partial reusability.

[Exoscientist]

8 hours ago, SunlitZelkova said:

As long as we’re doing a total franken rocket, couldn’t we put Crew Dragon on top?

It would probably need a proper service module though, and I’m not sure how that would impact mass. And of course, a beefed up heat shield.

Considering it was originally intended to take 7 people to the ISS, IIRC, it should be far more roomier and comfortable than Orion.

 Yes. That is a key point. The Orion capsule is also bloated Old Space, with amortized cost estimated at ca. $2 billion per launch(!) Yes, that is just the capsule, not even the SLS rocket. 
 
 As I recall, the crew Dragon for a lunar mission would need only minor modifications such as a more powerful communications system for communication from the Moon. It was already given sufficient heat shield capability in its design for return from escape velocity, which is higher than just return from LEO.

  Robert Clark

[tater]

3 minutes ago, Exoscientist said:

 Yes. That is a key point. The Orion capsule is also bloated Old Space, with amortized cost estimated at ca. $2 billion per launch(!) Yes, that is just the capsule, not even the SLS rocket. 
 
 As I recall, the crew Dragon for a lunar mission would need only minor modifications such as a more powerful communications system for communication from the Moon. It was already given sufficient heat shield capability in its design for return from escape velocity, which is higher than just return from LEO.

The problem is that it needs a service module for this use case. LES for Dragon pulls the trunk off (the fins are for LES contingency) with the capsule. Even with a trunk to SM alteration, it's not like it will have much capability.

The best alternate cislunar CONOPS would not be making new 1-stick to the Moon systems, but to assemble vehicles in LEO using existing infrastructure.

[AtomicTech]

38 minutes ago, Exoscientist said:

 As I recall, the crew Dragon for a lunar mission would need only minor modifications such as a more powerful communications system for communication from the Moon. It was already given sufficient heat shield capability in its design for return from escape velocity, which is higher than just return from LEO.

Definitely! You could build a transfer vehicle with just a few launches of FH and have a temporary (or maybe even permanent) station wherever you decided to set-up shop!

[magnemoe]

10 hours ago, SunlitZelkova said:

Another plus of using a lander as a shuttle is you can make it big like Starship. I really hope SpaceX is taking into account the lesson of Apollo 13 to have a section of the spacecraft “lifeboat capable”.

That said, I don’t think returning to a dual spacecraft architecture is feasible just for safety reasons. We got lucky on Apollo 13, but there seems to be a consensus that either a single lander or “all up” direct ascent lander is best for lunar base operations. LOR was chosen because it was the fastest, not sustainable.

Basically every Soviet lunar landing concept outside of the L3 (the LOR one for the N1), which was also designed around speed and not sustainably, utilized a single lander and direct ascent, and the original Apollo was a direct ascent lander too. Post-Apollo American concepts have mostly centered around a LEO shuttle, a space shipyard, and a lunar shuttle which docks at the shipyard. Sometimes there are big monolithic landers, but these require big expendable heavy lift launch vehicles and aren’t good for a long term lunar base. And of course, then there’s Constellation.

I kind of wonder if eliminating Gateway and using Starship HLS in a direct ascent role would allow a greater payload to the surface. Travel to it in Dragon and then go directly to the surface.

I wonder if SpaceX is considering the possibility of conducting maintenance on Starship HLS’ either on the surface or in orbit.

I wonder if they have a division like Bellcom, which was the think tank NASA charged with coming up with crazy applications for Apollo hardware in the mid 60s.

Getting off this tangent… somehow Starship feels like a safer lunar transit vehicle than Orion.

Apollo was 4 stages, the capsule, the service module and decent and accent module for landing on the moon. 
Focus for Apollo was speed and this also required low weight. 
Now new landers will be sent to the moon on their own and the crew launches separately. This saves on launch cost and the lander launcher does not be to be human rated. 

Most of the new landers will be single stage as this enables reuse after refueling. Another option is an replaceable crasher stage, an drop tank is also possible as fuel flow on lander is more like aircraft term not hydro plants. 

[sevenperforce]

10 hours ago, Exoscientist said:

11 hours ago, tater said:

11 hours ago, Exoscientist said:

Do the calculation with the Ariane 5 core as the upper stage.

Would not fit in the VAB as he said, so why bother doing the math?

Actually it would. The limits for the length of an  upper stage for the SLS are based on specs for a supposed second mobile launcher for the SLS which has to be provided to the manufacturer of the mobile launcher. But that second mobile launcher has not been constructed yet, and the manufacturer has been greatly criticized for costs overruns and delays. A commercial approach would probably just choose a different manufacturer for the second mobile launcher such as SpaceX.

 The doors of the vehicle assembly building are quite huge and can accommodate the size of the SLS core and Ariane 5 as a second stage:

Vehicle Assembly Building.
There are four entries to the bays located inside the building, which are the four largest doors in the world.[14] Each door is 456 feet (139.0 m) high...

With the existing mobile launcher design, there is an 18-meter height available for the upper stage.  The current mobile launcher base comes 7.6 meters above the ground and has a launch umbilical tower that is 108 meters higher than that, reaching 5.7 meters higher than the max height of Block 1B. 

The Ariane 5 core is 23.8 meters high. Putting it into SLS as the second stage would lift the height of this new "Block 1F" (F for Frankenrocket) above the height of the current launch umbilical tower, which is already what limits the height relative to the VAB doors.

The only option is to build a crawler-transporter that carries SLS lower to the ground. I'm not sure how we are going to manage that. SLS is heavy. 

12 hours ago, Exoscientist said:

22 hours ago, sevenperforce said:

To start with, let's do a one-to-one comparison with the current SLS to see if we have improved anything. If it's not improved, then swapping out upper stages is something that could as readily be done on the current SLS.

Do the calculation with the Ariane 5 core as the upper stage. A commercial space approach wouldn’t use such a tiny upper stage such as the ICPS for a rocket this size when larger, appropriately-sized upper stages are available

As I said, you need to start with a one-to-one comparison to see if ditching boosters and using RS-68s improves anything on the lower stage, before proposing a different upper stage. Otherwise you're changing multiple variables at once, which muddies the water. If changing the lower stage makes things worse, then you should drop that plan entirely, before talking about combining a bad lower stage with a new upper stage.

Also, the Ariane 5 core is not "available" at all; it's not even being made.

Also also, the Ariane 5 core had a Vulcain 2 on it, which is not a vacuum engine and gets really poor vacuum specific impulse compared to a proper upper stage.

But if you're calling the ICPS "tiny" then you should also be calling the Ariane 5 core "tiny" in comparison to the EUS for Block 1B. The EUS carries 76% more propellant than the Ariane 5 core. The Ariane 5 core is extremely narrow compared to EUS. A reasonable space approach wouldn't use such a skinny, narrow stage as the A5 core for an upper stage when larger, appropriate-width upper stages are available.

1 hour ago, Exoscientist said:

10 hours ago, SunlitZelkova said:

As long as we’re doing a total franken rocket, couldn’t we put Crew Dragon on top?

It would probably need a proper service module though, and I’m not sure how that would impact mass. And of course, a beefed up heat shield.

As I recall, the crew Dragon for a lunar mission would need only minor modifications such as a more powerful communications system for communication from the Moon. It was already given sufficient heat shield capability in its design for return from escape velocity, which is higher than just return from LEO.

Crew Dragon has some re-entry stability issues due to the greater height-to-width ratio of its OML compared to more traditional capsules. It's fine for LEO returns but I don't believe it can handle the higher angle of attack needed for cislunar returns.

More importantly, though, it doesn't only need a more powerful communications system; it needs a service module. Its onboard propellant reserves are not enough for a lunar return, and certainly not enough for a lunar orbital insertion (in fact it can't even fire its main propulsive vacuum thrusters while attached to another module).

And again, at this point, why are we bothering with SLS at all?

21 hours ago, tater said:

23 hours ago, sevenperforce said:

ape the old Saturn S-1D Mega Atlas design

I frickin love that thing.

Now you've got me wondering what the maths would look like.

The SLS core carries 29.9% more propellant than the Shuttle did. The intertank is essentially identical (not unlike the Ariane 5, the thrust of the solid boosters is transmitted in large part through the intertank in both the SLS and STS designs, leaving the hydrogen tank to hang in tension until booster separation). The SLS core masses 85.3 tonnes without engines. Crunching the numbers, you get roughly 34.4 tonnes for the tankage, 16.1 tonnes for the forward skirt section, and 34.7 tonnes for the engine boattail section sans engines. Note that SLS Block 1 also carries the 4.5-tonne LVSA adapter for ICPS.

Let's imagine a few alternate possibilities inspired by the S-1D design. With all of these, I'm going to compare solely to SLS Block 1 to get an idea of what is actually possible, with the understanding that adding EUS to the design will create further improvements. To that end, everything north of the LOX tank remains identical: the same forward skirt, the same stage adapter, the same ICPS, and so forth, and total height must remain the same as well.

For reference, here's the aft end of the current SLS:

We plug in the following values for SLS Block 1 to LEO with Orion:

	Booster	1st Stage	2nd Stage
Dry Mass, kg	102058.3	103872.1	3490
Propellant, kg	623689.5	952544	30710
Thrust, kN	17595	9116	110.1
Isp, seconds	268	452.3	462

We use default prop residuals, a restartable upper stage, and set a "payload fairing" of 6,926 kg (that's Orion's LAS) to jettison at 120 seconds. Launching from Cape Canaveral to a reference orbit of 185x185 km at 28.5 degrees using a two-burn ascent gives 85,671 kg to LEO. Note that this is under the earlier estimate both because of the addition of Orion's LAS and because I was using a number for the dry mass of the first stage that didn't include the mass of the engines or the LVSA.

Time to get creative! Let's start small. Suppose that we keep the existing RS-25s but split that 34.7-tonne boattail engine section into two parts, one of which is attached to the core stage and one of which can be jettisoned: 

When do we jettison the outboard engines? Well, the boosters burn out and separate at T+131 seconds, after the core has burned 269.1 tonnes of hydrolox. At separation, then, SLS Block 1 has a T/W ratio of 1.13, not including payload. If we want this same T/W ratio to be maintained at outboard engine jettison, where our thrust will be cut in half, we need to wait to T+331 seconds, at which point there will be just 272.7 tonnes of hydrolox remaining in the tanks. For the purposes of plugging everything into the launch performance calculator, this creates a simulated three-stage architecture, where the "dry mass" of the "1st Stage" is merely the mass of the discarded outboard engines and engine section and the "propellant" of the "1st Stage" is the total prop burned at jettison, with all four engines as the "thrust" of the "1st Stage" and only two for the "2nd Stage":

	Booster	1st Stage	2nd Stage	3rd Stage
Dry Mass, kg	102058.3	24404	79468.1	3490
Propellant, kg	623689.5	679858	272700	30710
Thrust, kN	17595	9116	4558	110.1
Isp, seconds	268	452.3	452.3	462

Plugging this into the calculator, with the same other parameters listed above, gives 94,523 kg to LEO. Now, before you go and say "wait, that's only what SLS Block 1 can already do!", remember that this is comparative to the 85.7 tonnes that the calculator gave us for SLS Block 1 before, which was notably low (in part because of the Orion LAS jettison). So in reality, we're talking about a ~10% improvement in payload just by dropping these engines at the correct time.

Now, let's go one step further and replace those two outboard engines with RS-68As and only put a single RS-25 on the core:

The RS-68A is only ever so slightly larger than the RS-25. The major issue with putting them on the aft end of the planned Ares V was that the heat flux from the solid boosters would have melted them, but hopefully with greater distance from the solid motors they will be in better shape here. I'm sure we can afford to slap some extra heat shielding on here somewhere if necessary.

Together, a single RS-25 and a pair of RS-68As produce 9396 kN of thrust, just slightly more than four RS-25s. Good start so far. The specific impulse of the trio in combination is now going to be 421.8 seconds, based on thrust-specific Isp. The weight of the engine section is a function of the max thrust it has to transfer to the vehicle, so the total engine section weight is going to go up from 34.7 tonnes san engines to 35.8 tonnes. However, 75.7% of that is going to be in the thrust structure for the RS-68As, meaning more mass can be jettisoned. However, we'll have to burn all three for longer in order to make sure our thrust-to-weight is in good shape when we jettison. Because we are going to be jettisoning later (as in, nearly-to-LEO), gravity drag will be much lower and we can get away with a T/W ratio on the order of ~0.85. 

In the end, we get this:

	Booster	1st Stage	2nd Stage	3rd Stage
Dry Mass, kg	102058.3	40608.4	62730.2	3490
Propellant, kg	623689.5	782100	170444	30710
Thrust, kN	17595	9396	2279	110.1
Isp, seconds	268	421.8	452.3	462

Once again plugging this into the calculator with the same parameters as before, we get a disappointing 86,717 kg to LEO. This is still technically an improvement over SLS Block 1, but it's not nearly as much of an improvement as our first concept. That's not too surprising; that extra 40 seconds of specific impulse on the RS-25 over the RS-68A makes a big difference. This is a sustainer architecture and so the extra thrust on the core stage doesn't do all that much for us.

Finally, let's try a combination of these two designs: two RS-25s in the center and two RS-68As outboard:

Now the core is really starting to get beefy. The total core thrust comes up to 11.7 MN so we need to beef up our engine section weight to 44.5 tonnes. Doing all the math as above, we get this:

	Booster	1st Stage	2nd Stage	3rd Stage
Dry Mass, kg	102058.3	40608.4	79468.1	3490
Propellant, kg	623689.5	679858	272700	30710
Thrust, kN	17595	11675.3	4558	110.1
Isp, seconds	268	427.7	452.3	462

Unfortunately, this only gets us 84,567 kg to LEO. The extra thrust isn't being utilized well, and the extra weight is prohibitive.

The RS-25 is a hard engine to beat.

Just for the fun of it, let's come up with a REALLY wild proposal. Let's get rid of the SRBs entirely and replace them with Atlas V first stage boosters, and let's put a grand total of five engines on the core (four RS-68s and one RS-25). We'll stack a pair of additional Atlas V LOX tanks on top of each one, which we will use to crossfeed the RS-68As:

Things are a little bit different now. The boattail section doesn't have to transmit nearly as much booster force and so I'll leave its total mass no greater than for the 2-and-2 design (though obviously it is getting two more RS-68A engines at an unpleasant 6,740 kg each). The empty mass of each booster is 21,054 kg, plus the weight of two LOX tanks which I will ballpark at an additional 30 tonnes per pair. Here the dry mass is added to the boosters but the propellant mass is added to the core, effectively, thanks to crossfeeding. Each pair of extra LOX tanks will hold 415442 kg of LOX (we will move the bulkhead in the core but otherwise everything will be as it was). Our numbers look like this:

	Booster	1st Stage	2nd Stage	3rd Stage
Dry Mass, kg	51054	54088.4	62730.2	3490
Propellant, kg	284089	1510742	170444	30710
Thrust, kN	4152	20072.2	2279	110.1
Isp, seconds	337.8	417.6	452.3	462

To my delight, this gets 86,952 kg to LEO -- not as high as some of the other designs above, but really very impressive considering that we've completely ditched those big solid boosters entirely!

We can also do the same design but use expendable Falcon Heavy side boosters (still sticking with the Atlas LOX crossfeed tanks so I don't have to rework the math):

The numbers:

	Booster	1st Stage	2nd Stage	3rd Stage
Dry Mass, kg	55600	54088.4	62730.2	3490
Propellant, kg	395700	1510742	170444	30710
Thrust, kN	8227	20072.2	2279	110.1
Isp, seconds	311	417.6	452.3	462

This configuration gets an absolutely whopping 121,720 kg to LEO with just ICPS on top. Replace ICPS with EUS as planned for Block 1B, and it can send 144 tonnes to LEO or 63.7 tonnes to TLI, nearly 40% more than the mythical SLS Block II and well over what Saturn V could do.

[Exoscientist]

4 hours ago, sevenperforce said:

Plugging this into the calculator, with the same other parameters listed above, gives 94,523 kg to LEO. Now, before you go and say "wait, that's only what SLS Block 1 can already do!", remember that this is comparative to the 85.7 tonnes that the calculator gave us for SLS Block 1 before, which was notably low (in part because of the Orion LAS jettison).

The main goal is to create a super heavy lift launcher at a much lower cost than the SLS. Then an approx. 90+ tons to LEO launcher is acceptable as it makes possible a single-launch architecture for manned lunar missions.  

 In keeping with the low cost goal, using liquid fueled side boosters is acceptable as they would be much lower cost than the SLS solid boosters and would be reusable. 
 

 Robert Clark

Edited October 20, 2023 by Exoscientist

[sevenperforce]

7 minutes ago, Exoscientist said:

The main goal is to create a super heavy lift launcher at a much lower cost than the SLS. Then an approx. 90+ tons to LEO launcher is acceptable as it makes possible a single-launch architecture for manned lunar missions.  

 In keeping with the low cost goal, using liquid fueled side boosters is acceptable as they would be much lower cost than the SLS solid boosters and would be reusable. 

90 tonnes to LEO is not enough for single-launch manned lunar missions.

To be clear, the Falcon 9 SLS booster concept I describe above is absolutely NOT a reusable architecture. SLS liquid booster reuse is a VERY long pole.

[tater]

29 minutes ago, Exoscientist said:

The main goal is to create a super heavy lift launcher at a much lower cost than the SLS. Then an approx. 90+ tons to LEO launcher is acceptable as it makes possible a single-launch architecture for manned lunar missions.  

 In keeping with the low cost goal, using liquid fueled side boosters is acceptable as they would be much lower cost than the SLS solid boosters and would be reusable.

1. If it involves literally any element of SLS it is not, and will never be "much lower cost." Nonstarter for that goal.

2. As was said, 90t to LEO won't cut it for a single stack lunar mission.

[Exoscientist]

On 10/19/2023 at 1:55 AM, SunlitZelkova said:

Basically every Soviet lunar landing concept outside of the L3 (the LOR one for the N1), which was also designed around speed and not sustainably, utilized a single lander and direct ascent, and the original Apollo was a direct ascent lander too. Post-Apollo American concepts have mostly centered around a LEO shuttle, a space shipyard, and a lunar shuttle which docks at the shipyard. Sometimes there are big monolithic landers, but these require big expendable heavy lift launch vehicles and aren’t good for a long term lunar base. And of course, then there’s Constellation.

 Do you mean the entire crew capsule descending to the lunar surface atop the lunar lander? This is doable:

SpaceX Dragon spacecraft for low cost trips to the Moon, page 2: Comparison to 'Early Lunar Access'.

http://exoscientist.blogspot.com/2012/10/spacex-dragon-spacecraft-for-low-cost.html

 

  Bob Clark

[Exoscientist]

 

[RCgothic]

There just isn't any way to modify any part of SLS/Orion without making the cost and schedule problem worse, not better.

All the low-cost/high-cadence architectures that we could reasonably achieve within 10 years are Earth Orbit Rendezvous using some combination of Falcon/Superheavy/Vulcan/New Glenn and Dragon/Starship HLS/BO HLS and probably Starliner once they finally get that flying.

The biggest modifications to those basic elements we could get completed in a reasonable timescale and budget would be an expendable upper stage for Superheavy, and maybe one new service/propulsion module.

Edited October 20, 2023 by RCgothic

[tater]

5 hours ago, Exoscientist said:

SpaceX Dragon spacecraft for low cost trips to the Moon, page 2: Comparison to 'Early Lunar Access'.

The goal of Artemis is not a bare bones "flags and footprints" mission architecture. After the first landing, it's >2 astronauts, and for ~1 week on the surface or more. Assuming Gateway, in 6.5 day increments based on rendezvous opportunities.

 

[sevenperforce]

Any architecture based on SLS is also going to have the issue of limited volume for a co-manifested lander.

8.4x10 meters of volume is definitely a good bit -- probably enough for a flags-and-footprints lander -- but it's not enough to include a dedicated LOI braking stage for a single-launch architecture, and Orion certainly can't  do the braking burn. So even if you were able to beef up SLS (SLS Block 2B?) enough to throw an arbitrary amount of payload to TLI, it's still unclear how to make that work.

Then again, maybe I'm wrong and there IS a way to fit a good lander and a braking stage into that volume.

It just makes much more sense to have all of the assets delivered to cislunar space ahead of sending crew.

14 hours ago, tater said:

If it involves literally any element of SLS it is not, and will never be "much lower cost." Nonstarter for that goal.

And yet the fiddling I did yesterday is making me very curious about other possibilities.

I wonder what would happen if SLS used 4-segment boosters and put liquid kerosene tanks on top, and then put RD-180s there on the outer jettisonable engine skirt, fed from the kerosene tanks on the boosters and from the core LOX tank.