_Augustus_

NASA SLS/Orion/Payloads

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6 hours ago, T-10a said:

Did the math:

RS-68 powered SLS:  9856.47 m.s.-1 dV

RS-25 powered SLS: 10813.41 m.s.-1 dV

That extra 1000 m.s.-1 can be really important, and allow greater payload to orbit.

(DO NOTE: This was done with assuming only the core of the SLS's mass, no strapon SRBs, and TWR is irrelevant.)

A couple of problems. First of all, the RS-25 only has 1 second higher SL isp than the RS-68A. If you're using the vacuum isp of both engines, then you have to factor in far higher pressure losses for the RS-25 than for the RS-68A. Second, you only need 2 RS-68As on an SLS-sized core to deliver almost the same amount of thrust as four RS-25s, so your dry mass estimate for the RS-68 calculation was probably too high. If you add a third RS-68A, then your launch TWR jumps significantly and you have lower gravity drag than with four RS-25s.

Not saying the RS-68A is a better engine than the RS-25; just saying those numbers might be a bit misleading.

12 hours ago, tater said:

Course there are only so many RS-25 lying around, so they need another engine for SLS, anyway.

I meant the Shuttle-C concept with the cargo on the side like the orbiter, not a SDLV, like Ares, BTW. That way all the infrastructure would have stayed identical.

mac-rebisz-20150210-shuttle-c-007.jpg?14

This would have given capability with less delay, assuming a SHLV was needed.

Shuttle-C would have been a great idea, had they continued flying the Shuttle and wanted to keep all the facilities identical. But it is a good thing they ended the Shuttle program.

If NASA had been REALLY serious about reusing off-the-shelf Shuttle hardware, there were two very good options:

  • Build a Shuttle-Derived Atlas, using three SSMEs at launch with two on a jettisonable skirt, and a smaller version of the ET. The sustainer SSME completes orbital insertion minus a few hundred m/s, and the Orion's own service module provides circularization. Orion could have been used for flying crew to the ISS. For cargo flights, remove the skirt and add SRBs and an ICPS-esque upper stage. For BLEO flights of Orion, send up Orion first, and then send up a cargo mission but without payload, and dock the ICPS to Orion for the LEO exit burn. Catch the jettisonable skirt with a chute and a helicopter and/or chute the SRBs down, if it's cost-effective.
  • Build a Shuttle-derived ROMBUS, using one or two SSMEs on the core and 4-6 jettisonable tanks, again using the Orion's SM for circularization. Chute the tanks down for recovery if it's cost-effective. For cargo missions, replace 2 or more of the tanks with SRBs, using the exact same core attachment points, and fly as above.

Either option would have gotten a man-rated vehicle flying in under 3 years, using the exact same SRBs and SSMEs.

11 hours ago, Kerbal7 said:

If the SLS and Orion are a bridge to nowhere then why is NASA planning missions around it? I know y'all think our friends at NASA are too stupid to tie their own shoes and your muskssiah is going to build you a golden rocket but...

If SLS and Orion are a bridge to nowhere, then what are these people talking about?? 

They are talking about capabilities that SLS does not have.

EM-1 could be done, with crewed Orion, using what -- one RTLS Falcon 9 and two partially-expended Falcon Heavies?

EM-2's PPE delivery mission could be done with a single reusable F9 launch plus a single recoverable Falcon Heavy. The crewed component of EM-2 is meaningless if replaced with a crewed EM-1, since there are no plans to dock Orion to the PPE during EM-2.

Edited by sevenperforce

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NASA is designing missions around SLS because it's what Congress has given them. It's not more complicated than that.

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10 hours ago, tater said:

Mars Base Camp requires 10+ SLS launches. About half of those need to happen within days of each other.

That would require extremely fast turnaround time - faster than even SpaceX has ever done - and/or multiple pads, which they don't have or plan on having.

1 hour ago, sevenperforce said:

EM-2's PPE delivery mission 

It was announced at the National Space Council meeting that they'd be launching PPE before EM-2 on a commercial LV, which actually makes sense seeing as the thing is powered by SEP......

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1 hour ago, sevenperforce said:

Shuttle-C would have been a great idea, had they continued flying the Shuttle and wanted to keep all the facilities identical. But it is a good thing they ended the Shuttle program.

Shuttle-C would also give Congress the ability to cancel manned spaceflight a lot earlier than they did while keeping all the satellite launch capability of the Shuttle.  NASA was almost certainly scared silly of it and had no intention of taking it anywhere.  If you started the program *after* the shuttle got canned, I can't see it progressing any differently than SLS.  Building a new rocket based on shuttle parts but with enough changes (and political meddling) that you have the costs of building from scratch without the freedom to really work around the problems.

 

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7 hours ago, T-10a said:

Did the math:

RS-68 powered SLS:  9856.47 m.s.-1 dV

RS-25 powered SLS: 10813.41 m.s.-1 dV

That extra 1000 m.s.-1 can be really important, and allow greater payload to orbit.

(DO NOTE: This was done with assuming only the core of the SLS's mass, no strapon SRBs, and TWR is irrelevant.)

1. If the rocket has two side mounted boosters they operate close to Sea Level and the difference is trivial. Again assuming Vac ISP all the way up is deceptive.
2. In general you would not design any rocket to burn past 8500 dV since you can basically leave that for the second stage. (RL10b-2 or RL10-C which have ISPs of 466 or 452 and are much lighter.
(for example you could use the remaining fuel to re-land your first stage).
3.

Thrust RS68A = 3,140  kN  (At Sea level)
Thrust RS25 =  2,279  kN (At sea level)

4.

Nozzle diameter RS68A  = 2.43 m
Nozzle diameter RS25  = 2.40 m

Lets think really big, assuming an infinite hexogonal layout and minimum distance of 1/10th diameter between engines what is the Maximum unit force per area (pressure) when number of engines exceeds 2,  which either engine can produce.

P  = T / (1.21d2*31/2) =
RS68A = 266 kPa
RS25 =  188 kPa

This means that you can (with the most compact layout of engines) launch 42%  more mass on a RS68A stack on average than on an RS25 Stack (of course you could use that stack to carry any dV you lost switching from RS25, but that would be dumb). By putting more thrust onto the core you lower the need for boosters and now you can be more flexible about the chosen boosters, for example F9 booster, which is relandable, has a higher ISP. Again your core should be relandable, which means you only want to burn to about 3500 m/s before MECO, and this means that the mass you launched with did not go into providing fuel for itself, but in providing weight for the second stage fuel mass and RL10C? or some variant of RL10 which would carry what. . . .300 t to Orbit.

NASA can use RS25 if they want, my space agency would be using RS68 because the future of space exploration is about getting bulk into space.
 

 

 

 

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What exactly is so complicated about the rs25 in comparison to any other engine? 

Also LOL at an rs25 engine test being a milestone for SLS yet SX almost saving fairings when launching a satellite is just something we’ve been doing since the 60’s - cracking logic 

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Shuttle-C made sense when STS was flyng, but after shutting down STS it would make no sense at all.

Just now, Jaff said:

What exactly is so complicated about the rs25 in comparison to any other engine? 

RS25 is the SSME. It's the single most complex rocket engine ever designed, since it was designed to be reusable, high-thrust, man-rated, (somewhat) throttlable, and with a sustained burn from sea-level to vacuum.

Just now, Jaff said:

Also LOL at an rs25 engine test being a milestone for SLS yet SX almost saving fairings when launching a satellite is just something we’ve been doing since the 60’s - cracking logic 

 

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5 minutes ago, Nibb31 said:

Shuttle-C made sense when STS was flyng, but after shutting down STS it would make no sense at all.

I meant started before Shuttle stopped flying, and continued as needed after the Orbiter was stopped. Crew would then transition to commercial crew ideally in such a counterfactual. NASA would still have no need for the capability, as they have no payloads for it, but it would have kept the Shuttle contractors busy (which is the entire point of SLS, it's not about making a rocket).

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I would have been happy with Orion flying side-mounted on top of Shuttle-C. At least the crew could escape.

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Did they look into that?

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54 minutes ago, wumpus said:

If you started the program *after* the shuttle got canned, I can't see it progressing any differently than SLS.  Building a new rocket based on shuttle parts but with enough changes (and political meddling) that you have the costs of building from scratch without the freedom to really work around the problems.

If the program had been started after the Shuttle got canned, with the sole requirement being "reuse mothballed STS parts" and no Congressional meddling, they could have saved a metric ton of money by doing a drop-tank or drop-engine SSTO crew vehicle, capable of being retrofit for SHL. If there was ever an engine for SSTO, other than Raptor, it would be the SSME.

19 minutes ago, Nibb31 said:

I would have been happy with Orion flying side-mounted on top of Shuttle-C. At least the crew could *probably* escape.

Fixed that for you.

32 minutes ago, Jaff said:

What exactly is so complicated about the rs25 in comparison to any other engine?  

Specifically, it had a special tapering nozzle design to prevent flow separation at sea level but maintain high expansion at vacuum; the nozzle was regeneratively cooled using liquid hydrogen, it preburned both the LOX and the LH2, it used both high-pressure and low-pressure turbopumps, it had dual-redundant MECs, it had a separate hydrogen coolant loop for the combustion chamber, it used a ten-tank helium gas system for actuating all valves and purging the engine, and it had a 10.5+/- degree gimbal range along two axes using a titanium alloy gimbal assembly which also served as the thrust transfer mechanism. And with all that, it managed a vacuum ISP within 4% of the highest-isp hydrolox engine flying, the RL10.

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13 hours ago, tater said:

Mars Base Camp is the Lockheed Martin version of Elon Musk's Mars talk.

That said, Mars Base Camp is closer to the current (not Apollo) NASA way, which has them sending 2 of everything out of an abundance of caution.

Each of those LM landers is an SLS just to get to orbit, I think, and another to be sent to Mars as I recall. The central hub is sent ahead (1 SLS), and each transit hub is a launch, possibly 2. Plus 2X Orion with EUS.

So Mars Base Camp requires 10+ SLS launches. About half of those need to happen within days of each other.

Let that sink in.

I don‘t think the number of launches is correct. The Lockheed Lander is flying to mars under its own power, to be refueled by a commercialy launched tanker. Also, since the Vehicle is reusable i guess the high number of Launches for assembly is bearable,  even of they take quite some time. Atleast better than one of those NTR Mars mission concepts which require the same high number of Launches for every excursion.

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I know to orbit it's at least an SLS. I posted their whole PDF up thread someplace.

https://lockheedmartin.com/content/dam/lockheed/data/space/photo/mbc/MBC_Updates_IAC_2017.pdf

Their graphic shows 6 SLS launches and 6 commercial launches... and the double base camp is not part of those launches at all, it's already up. at point "A" on their graphic.(page 6).

 

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13 hours ago, T-10a said:

Did the math:

RS-68 powered SLS:  9856.47 m.s.-1 dV

RS-25 powered SLS: 10813.41 m.s.-1 dV

That extra 1000 m.s.-1 can be really important, and allow greater payload to orbit.

(DO NOTE: This was done with assuming only the core of the SLS's mass, no strapon SRBs, and TWR is irrelevant.)

So what? What useful tasks can it do with those numbers? 

Land on Mars? No. 

Land on the Moon? No. 

Put mass in orbit? Falcon 9 and Delta IV Heavy can do the same faster and cheaper.

Put volume in orbit? Cool. But other launch providers will probably get there first.

Launch flagship class payloads to faraway places? Perhaps. Occasionally. When NASA can afford to.

Unless you're designing for a specific mission, your rocket needs to be good for a number of missions, which SLS patently isn't. So what do we want in a general purposee rocket?

Firstly, there's no longer any good reason for it to be man rated. Orion's only a couple of tons too heavy to be put up on an expendable Falcon 9 (strongly suspect 22t will be exceeded by block 5) nevermind Delta IV Heavy or Falcon Heavy. So don't. Separate Crew and Cargo and rendezvous if necessary. Ares had this right.

Secondly, it needs to put mass and volume in orbit cheaply. If you can't do that then you might as well go home because otherwise you're doing the equivalent of bringing a knife to a gunfight against the new space guys.

Thirdly, if you want to go beyond Earth's sphere of influence you need to be able to support a rapid launch cadence. Yes, you can get to the moon by launching an enormous stack Apollo style. But if you want a moon base or a mission to Mars you have to be able to go to orbit much more often than a few times a year. You don't need 140t to orbit if you can do 25t every two weeks.

IMO the best way to achieve 2 and 3 is reusability.

Unless you rock those boxes there's really no point.

3 hours ago, insert_name said:

SLS loses one of its only payloads. One step closer to cancellation.

Edited by RCgothic

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10 minutes ago, _Augustus_ said:

Another payload bites the dust.......

What was the other one?

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20 minutes ago, RCgothic said:

So what? What useful tasks can it do with those numbers? 

Well, to be fair, those numbers were meaningless.

But I agree with the rest of what you said.

1 hour ago, insert_name said:

Unless SpaceX beats them to it. And, probably, gets it there faster. Falcon Heavy expendable can put 7.88 tonnes on a Hohman transfer to Jupiter, assuming an LEO exit burn of 6.3 km/s. How heavy is Europa Clipper without the lander and add-ons?

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21 hours ago, tater said:

This is a nonsensical statement. If the cost of propellants, and the pro-rata cost of the vehicle is less than the cost to throw away a rocket for the same launch, it's profitable. Propellants are cheap. Several hundred thousand $ for BFR. A few hundred million for the 2 vehicles, combined. Call it 300 M$. If they can relaunch even just 10 times, the total cost/launch is 31 M$. For up to 150 tons. 5 times? It;s in F9 territory. just once? Sadly, that's Delta IV Heavy kinda pricing---but for 150 tons.

The problem is that it risks gobbling up the entire market in one launch and then spending the rest of the year in cold storage, because the launch market booming due to decreased costs is not as guaranteed as some like to think.

The bulk of costs being due to keeping the lights on and the personnel on the roster, gee, where have we heard that before!?

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Except that it’s still cheaper to fly with minimal payload.

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23 minutes ago, tater said:
35 minutes ago, DDE said:

The problem is that it risks gobbling up the entire market in one launch and then spending the rest of the year in cold storage, because the launch market booming due to decreased costs is not as guaranteed as some like to think.

The bulk of costs being due to keeping the lights on and the personnel on the roster, gee, where have we heard that before!?

Except that it’s still cheaper to fly with minimal payload.

This discussion should probably be moved to the SpaceX thread, but...

What would the most mass-conservative way to add crew-carrying capacity to BFS with 0-0 L/L abort and lifeboat capability?

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^^^Move that to the other thread.

On topic, anyone know the wet mass of the LockMart MADV upon Martian landing?

 

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16 minutes ago, tater said:

^^^Move that to the other thread.

On topic, anyone know the wet mass of the LockMart MADV upon Martian landing?

 

To get a rough idea, I'd say just add together the mass of an Orion and fully fueled ACES/Centaur-5 (to account for the heat shielding and whatnot).

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Looks like their MADV is about 105 tons, wet, and Mars EDL is wet. Crew BFS is supposed to be ~85 tons dry, cargo on top. So at least twice as massive as the LM design at the upper limit. For a similar type of mission architecture (small crew, brief stay), BFS doesn't actually mass much more, though it might be stuck on the surface :) . Dunno what the min prop load is to make LMO, maybe it could carry extra props instead of cargo and be able to land, then make orbit (seems like it would be possible with the same number of launches as Mars Base Camp (6 SLS, 6 heavy commercial launches) to send a few BFR in a sort of MBC mission using alternate vessels.

 

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