Author Topic: Storable Propellant Earth Departure Stages  (Read 71424 times)

Offline QuantumG

  • Senior Member
  • *****
  • Posts: 9238
  • Australia
  • Liked: 4477
  • Likes Given: 1108
Storable Propellant Earth Departure Stages
« on: 08/29/2013 01:49 am »
According to this article, which has it's own thread, NASA wants a 43 ton lunar lander.. for some reason.. and they say the only way to get a payload that big into lunar orbit is with the SLS.

I don't get it. What's wrong with just using a storable propellant Earth departure stage? Let's be conservative and say it only has an isp of about 312s, and a propellant mass fraction of 90%, how big would it be?

According to my math, please check me, I figure it would be about 10,458 kg dry and 104,579 kg when full. This would provide the 3107 m/s of delta-v to get through TLI, with the lunar insertion to be done by the lander (as in the NASA architecture).

I know 105 tons sounds like a lot, but it's only two Falcon Heavy launches, and because we're using storable propellant there's no time pressure. If you really wanted to you could do it with Falcon 9 v1.1.


   

« Last Edit: 08/29/2013 01:52 am by QuantumG »
Human spaceflight is basically just LARPing now.

Offline Warren Platts

Re: Storable Propellant Earth Departure Stages
« Reply #1 on: 08/29/2013 02:39 am »
It only has an isp of about 312 s?
"When once you have tasted flight, you will forever walk the earth with your eyes turned skyward, for there you have been, and there you will always long to return."--Leonardo Da Vinci

Offline QuantumG

  • Senior Member
  • *****
  • Posts: 9238
  • Australia
  • Liked: 4477
  • Likes Given: 1108
Re: Storable Propellant Earth Departure Stages
« Reply #2 on: 08/29/2013 02:54 am »
It only has an isp of about 312 s?

For the sake of argument.

Human spaceflight is basically just LARPing now.

Offline RocketmanUS

  • Senior Member
  • *****
  • Posts: 2226
  • USA
  • Liked: 71
  • Likes Given: 31
Re: Storable Propellant Earth Departure Stages
« Reply #3 on: 08/29/2013 03:16 am »
That could work.
First FH launches EDS part filled.
Second FH launches tanker to complete filling of EDS.
Need to figure out the mass of the tanker too.
Also the altitude would be higher than 185 km x 185 km so less mass to that higher orbit.

So if you want to get a 43 mt lander through TLI you would need another tanker launch ( three FH's ). Or use a smaller lander with less payload.

What is the dry mass of the lander?
What is the max payload mass?

I assume it would use a version of the Super Draco.

So a FH would be used to launch the 43 mt lander into LEO.
Would the FH 2nd stage have enough performance to get the lander to the needed altitude and dock with the EDS? ( FH top expected performance 53 mt and lander is 43 mt assuming it would fit a stander fairing and not need a larger more mass sized fairing. )

For crew, Orion would be launched on FH to meet up with the EDS.

For the lander four FH's are needed. $135M x 4 = $540M
For the Orion three FH's are needed. $135M x 3 = $405M
Possible savings.

Could work for commercial if they had a commercial Lunar capsule.

Offline QuantumG

  • Senior Member
  • *****
  • Posts: 9238
  • Australia
  • Liked: 4477
  • Likes Given: 1108
Re: Storable Propellant Earth Departure Stages
« Reply #4 on: 08/29/2013 03:22 am »
Heck, throw ten Falcon Heavy launches at it if you want, you might start to approach the cost of an SLS launch :)
Human spaceflight is basically just LARPing now.

Offline RocketmanUS

  • Senior Member
  • *****
  • Posts: 2226
  • USA
  • Liked: 71
  • Likes Given: 31
Re: Storable Propellant Earth Departure Stages
« Reply #5 on: 08/29/2013 03:34 am »
And if the lander used storable propellant too, mass was 43 mt and did the LLO insertion burn it would get something around 8 mt?

Forgot to add in the cost of the tankers for the EDS.

Offline Robotbeat

  • Senior Member
  • *****
  • Posts: 39270
  • Minnesota
  • Liked: 25222
  • Likes Given: 12114
Re: Storable Propellant Earth Departure Stages
« Reply #6 on: 08/29/2013 04:04 am »
Put a 40kW electric tug on it (could be built using essentially a modified off-the-shelf commsat bus with a bigger or a couple extra solar arrays and some extra Xenon tanks... big commsats now do about 20kW), and you could put a Apollo-or-larger lander (up to maybe 20mT) in LLO with just an RS-68A Delta IV Heavy. A much bigger one if you used the cheaper Falcon Heavy. It'd take a few years to get all the way there (less if you used a Falcon Heavy and could tolerate a lower Isp, or if you could use a 100kW modified commsat or something it'd take proportionally less time), but so what? A lander in LLO with a single launch of an EELV Heavy. Or a really big lander in LLO with a Falcon Heavy. Have to use storable propellants with the lander, but that's not a huge deal.

Might even be cheaper than developing a hypergolic Earth Departure Stage that is refuelable.

Or heck, if you had a 18mT reusable single-stage hypergolic lander (probably needed to equal the performance of a 14mT two-stager), you launch a big tank of hypergolic propellant on a single Falcon Heavy along with a 100kW modified-commsat-tug, and you could refuel like 3 times per Falcon Heavy launch (alternately, between 6 and 9 refuelings per SLS launch, depending on which version of SLS... but you'd need a bigger tug or you'd need to launch multiple tankers per flight...).

A yet more efficient architecture would use a hydrolox un-crasher stage and a near-zero-boiloff depot fed by fully reusable tugs and a future fully reusable launch vehicle like the envisioned F9R with reusable upper stage.
« Last Edit: 08/29/2013 04:19 am by Robotbeat »
Chris  Whoever loves correction loves knowledge, but he who hates reproof is stupid.

To the maximum extent practicable, the Federal Government shall plan missions to accommodate the space transportation services capabilities of United States commercial providers. US law http://goo.gl/YZYNt0

Offline kkattula

  • Member
  • Senior Member
  • *****
  • Posts: 3008
  • Melbourne, Australia
  • Liked: 656
  • Likes Given: 116
Re: Storable Propellant Earth Departure Stages
« Reply #7 on: 08/29/2013 04:18 am »
Why bother with refueling and a tanker?  Just use 2 x FH-payload-sized storable EDS.

One starts the TLI then stages, and the other completes it. With a smaller lander, the second EDS could also do LOI.

Offline Robotbeat

  • Senior Member
  • *****
  • Posts: 39270
  • Minnesota
  • Liked: 25222
  • Likes Given: 12114
Re: Storable Propellant Earth Departure Stages
« Reply #8 on: 08/29/2013 04:28 am »
Continuing my idea of an architecture based off of SEP tug delivery of hypergolic propellant and reusable F9R, reusable single-stage lander in LLO:

If you had a version of Dragon which was refuelable and had at least 800m/s delta-v, you could launch to EML1/2, refuel with a tanker there, then fly to LLO, and take a lander to the surface and back, refuel both lander and Dragon, fly Dragon back to EML1/2 to refuel and then finally back to Earth... could be totally reusable, assuming you could give a capsule enough propellant to do 800m/s... The fueled lander with its huge delta-v capability could be a sort of dispatchable lifeboat if the Dragon got stranded.

...Um, that sounds too complex. Makes more sense to just leave the Dragon or whatever at EML1/2 and take the lander to get between EML1/2 and LLO and back, with refueling at LLO (and EML1/2, though that's not strictly required, it'd be more efficient).

Anyway, MY POINT IS: If you're just sending a lander (without crew) to lunar-orbit rendezvous and aren't in a huge hurry, makes more sense to send it via slow-boat SEP, now that the technology is mature. A single launch of DIVH or especially a Falcon Heavy should be enough to put at very least an Apollo-sized lander in LLO if you leverage electric propulsion as a sort of uber-EDS. Xenon (and Argon, for that matter) is, after all, storable.
« Last Edit: 08/29/2013 04:29 am by Robotbeat »
Chris  Whoever loves correction loves knowledge, but he who hates reproof is stupid.

To the maximum extent practicable, the Federal Government shall plan missions to accommodate the space transportation services capabilities of United States commercial providers. US law http://goo.gl/YZYNt0

Offline RocketmanUS

  • Senior Member
  • *****
  • Posts: 2226
  • USA
  • Liked: 71
  • Likes Given: 31
Re: Storable Propellant Earth Departure Stages
« Reply #9 on: 08/29/2013 04:32 am »
Put a 40kW electric tug on it (could be built using essentially a modified off-the-shelf commsat bus with a bigger or a couple extra solar arrays and some extra Xenon tanks... big commsats now do about 20kW), and you could put a Apollo-or-larger lander (up to maybe 20mT) in LLO with just an RS-68A Delta IV Heavy. A much bigger one if you used the cheaper Falcon Heavy. It'd take a few years to get all the way there (less if you used a Falcon Heavy and could tolerate a lower Isp, or if you could use a 100kW modified commsat or something it'd take proportionally less time), but so what? A lander in LLO with a single launch of an EELV Heavy. Or a really big lander in LLO with a Falcon Heavy. Have to use storable propellants with the lander, but that's not a huge deal.

Might even be cheaper than developing a hypergolic Earth Departure Stage that is refuelable.

Or heck, if you had a 18mT reusable single-stage hypergolic lander (probably needed to equal the performance of a 14mT two-stager), you launch a big tank of hypergolic propellant on a single Falcon Heavy along with a 100kW modified-commsat-tug, and you could refuel like 3 times per Falcon Heavy launch (alternately, between 6 and 9 refuelings per SLS launch, depending on which version of SLS... but you'd need a bigger tug or you'd need to launch multiple tankers per flight...).

A yet more efficient architecture would use a hydrolox un-crasher stage and a near-zero-boiloff depot fed by fully reusable tugs and a future fully reusable launch vehicle like the envisioned F9R with reusable upper stage.
Or check this thread out. Crew or cargo and could start with the cargo version.
http://forum.nasaspaceflight.com/index.php?topic=32669.msg1088701#msg1088701

Offline QuantumG

  • Senior Member
  • *****
  • Posts: 9238
  • Australia
  • Liked: 4477
  • Likes Given: 1108
Re: Storable Propellant Earth Departure Stages
« Reply #10 on: 08/29/2013 04:35 am »
...Um, that sounds too complex.

Exactly. Better is the enemy of good enough. The point is that trivial solutions like a storable propellant earth departure stage is cheaper than SLS.
Human spaceflight is basically just LARPing now.

Offline Robotbeat

  • Senior Member
  • *****
  • Posts: 39270
  • Minnesota
  • Liked: 25222
  • Likes Given: 12114
Re: Storable Propellant Earth Departure Stages
« Reply #11 on: 08/29/2013 05:13 am »
...Um, that sounds too complex.

Exactly. Better is the enemy of good enough. The point is that trivial solutions like a storable propellant earth departure stage is cheaper than SLS.

If you do EML1/2 instead, you could do dual-launch Falcon Heavy for the entire mission:
1) FH Dragon to EML1/2
2) FH with a 23mT lunar lander (equivalent in performance to the 15mT Apollo lander, but capable of staging to and from EML1/2), launched using a modified commsat acting as an electric tug (probably would want at least 100kW of power if you launch it a year in advance, though 50kW is fine if you can wait two years).

Just two FH launches, just two dockings (like Apollo), no refueling required, just a modification of an existing platform (the comm-sat) acting as your electric EDS for the lander.
Chris  Whoever loves correction loves knowledge, but he who hates reproof is stupid.

To the maximum extent practicable, the Federal Government shall plan missions to accommodate the space transportation services capabilities of United States commercial providers. US law http://goo.gl/YZYNt0

Offline QuantumG

  • Senior Member
  • *****
  • Posts: 9238
  • Australia
  • Liked: 4477
  • Likes Given: 1108
Re: Storable Propellant Earth Departure Stages
« Reply #12 on: 08/29/2013 05:18 am »
Now you're just cheating.. for whatever reason, the assignment was to put a 43 ton lander into lunar orbit. The SLS advocates claim they have the only system that can do that. It's trivial to show they don't. QED.
Human spaceflight is basically just LARPing now.

Offline Downix

  • Senior Member
  • *****
  • Posts: 7082
  • Liked: 22
  • Likes Given: 1
Re: Storable Propellant Earth Departure Stages
« Reply #13 on: 08/29/2013 05:30 am »
From my calculations you will need 158 metric tons of storable propellant to make this work from LEO.

With the Falcon Heavy, with its on-record payload capability of 53 metric tons, you would need 3 launches *just to get the propellant into position* This is not counting the stage nor lander itself, which brings the total launches to 5 (assuming you launch the stage partially-filled). 5 launches will cost *more* than an SLS launch for the same performance, costing you $540 million vs the SLS at ~$500 million, just for the launch vehicle. The Falcon Heavy would also require the cost of the DUUS-alternative, while the SLS cost would include it.

You are throwing money away with this route.
chuck - Toilet paper has no real value? Try living with 5 other adults for 6 months in a can with no toilet paper. Man oh man. Toilet paper would be worth it's weight in gold!

Offline QuantumG

  • Senior Member
  • *****
  • Posts: 9238
  • Australia
  • Liked: 4477
  • Likes Given: 1108
Re: Storable Propellant Earth Departure Stages
« Reply #14 on: 08/29/2013 05:39 am »
From my calculations you will need 158 metric tons of storable propellant to make this work from LEO.

Please show your work.. I did.

Quote
5 launches will cost *more* than an SLS launch for the same performance, costing you $540 million vs the SLS at ~$500 million, just for the launch vehicle.

You are throwing money away with this route.

Only in pixie land is that the cost of an SLS launch.
Human spaceflight is basically just LARPing now.

Offline Downix

  • Senior Member
  • *****
  • Posts: 7082
  • Liked: 22
  • Likes Given: 1
Re: Storable Propellant Earth Departure Stages
« Reply #15 on: 08/29/2013 05:51 am »
From my calculations you will need 158 metric tons of storable propellant to make this work from LEO.

Please show your work.. I did.
I checked yours. You neglected the weight of the lander itself. To get your figures, you launched an empty stage (10.4 metric tons) but to calculate out the delta-v needed you need to add the lander weight to that, giving you 53.4 metric tons. Using two different delta-v calculators, 104 gets only ~2,050 delta-v.

To propel the stage dry-mass along with the 43 metric ton lander, you need significantly more propellant.
Quote
Quote
5 launches will cost *more* than an SLS launch for the same performance, costing you $540 million vs the SLS at ~$500 million, just for the launch vehicle.

You are throwing money away with this route.

Only in pixie land is that the cost of an SLS launch.

That's the cost on-record. Sorry to hurt your bash-SLS fest.
chuck - Toilet paper has no real value? Try living with 5 other adults for 6 months in a can with no toilet paper. Man oh man. Toilet paper would be worth it's weight in gold!

Offline QuantumG

  • Senior Member
  • *****
  • Posts: 9238
  • Australia
  • Liked: 4477
  • Likes Given: 1108
Re: Storable Propellant Earth Departure Stages
« Reply #16 on: 08/29/2013 06:01 am »
I checked yours. You neglected the weight of the lander itself.

No I didn't.

Dry mass: 10458
Payload mass: 43000
Total mass at burnout: 53458
Isp: 312
Delta-v: 3107

Apply the rocket equation, you get: 147579 in LEO.
Subtract the payload, you get: 104579
Subtract the dry mass, you get: 94121

Quote
Quote from: QuantumG
Only in pixie land is that the cost of an SLS launch.


That's the cost on-record.

Perhaps you mean that's the marginal cost that they made up, including none of the development costs, all of which could be avoided by just not building the pointless monster rocket.

Quote
Sorry to hurt your bash-SLS fest.

You didn't, if anything you helped.
« Last Edit: 08/29/2013 06:02 am by QuantumG »
Human spaceflight is basically just LARPing now.

Offline Archibald

  • Senior Member
  • *****
  • Posts: 2611
  • Liked: 499
  • Likes Given: 1096
Re: Storable Propellant Earth Departure Stages
« Reply #17 on: 08/29/2013 06:47 am »
According to this article, which has it's own thread, NASA wants a 43 ton lunar lander.. for some reason.. and they say the only way to get a payload that big into lunar orbit is with the SLS.

I don't get it. What's wrong with just using a storable propellant Earth departure stage? Let's be conservative and say it only has an isp of about 312s, and a propellant mass fraction of 90%, how big would it be?

According to my math, please check me, I figure it would be about 10,458 kg dry and 104,579 kg when full. This would provide the 3107 m/s of delta-v to get through TLI, with the lunar insertion to be done by the lander (as in the NASA architecture).

I know 105 tons sounds like a lot, but it's only two Falcon Heavy launches, and because we're using storable propellant there's no time pressure. If you really wanted to you could do it with Falcon 9 v1.1.


A very good idea, kind of "cheap and dirty" lunar architecture. I see your point - Falcon Heavy is cheap, storable propellants are straightforward technology, they don't boiloff with time. I have no doubts this would work pretty well.
So what's wrong, do you ask ?
Simple...
NASA obsession with LOX/LH2. Never, ever, would they consider something else for the TLI. That paradigm also apply to SSTO, btw. It is a little annoying, because issues with liquid hydrogen have long been obvious...  ::)
« Last Edit: 08/29/2013 06:48 am by Archibald »
Han shot first and Gwynne Shotwell !

Offline RocketmanUS

  • Senior Member
  • *****
  • Posts: 2226
  • USA
  • Liked: 71
  • Likes Given: 31
Re: Storable Propellant Earth Departure Stages
« Reply #18 on: 08/29/2013 06:55 am »
I used this calculator, it must be broken  ;D.
http://www.quantumg.net/rocketeq.html

With storable propellants I got 15,848 kg on the Lunar surface.
That is with the 43 mt lander doing the LLO insertion burn and landing burn.

There was little margin for unused propellants and pressurant gas. A little increase would take care of this. 

According to this article, which has it's own thread, NASA wants a 43 ton lunar lander.. for some reason.. and they say the only way to get a payload that big into lunar orbit is with the SLS.

I don't get it. What's wrong with just using a storable propellant Earth departure stage? Let's be conservative and say it only has an isp of about 312s, and a propellant mass fraction of 90%, how big would it be?

According to my math, please check me, I figure it would be about 10,458 kg dry and 104,579 kg when full. This would provide the 3107 m/s of delta-v to get through TLI, with the lunar insertion to be done by the lander (as in the NASA architecture).

I know 105 tons sounds like a lot, but it's only two Falcon Heavy launches, and because we're using storable propellant there's no time pressure. If you really wanted to you could do it with Falcon 9 v1.1.


A very good idea, kind of "cheap and dirty" lunar architecture. I see your point - Falcon Heavy is cheap, storable propellants are straightforward technology, they don't boiloff with time. I have no doubts this would work pretty well.
So what's wrong, do you ask ?
Simple...
NASA obsession with LOX/LH2. Never, ever, would they consider something else for the TLI. That paradigm also apply to SSTO, btw. It is a little annoying, because issues with liquid hydrogen have long been obvious...  ::)

Reason they would not used such a system is they don't want to go to the moon or any crewed BLEO program. That has been the real reason we have been strung along. They had the option with Atlas V and DIV for Lunar before investing in the HLV.

Offline mmeijeri

  • Senior Member
  • *****
  • Posts: 7772
  • Martijn Meijering
  • NL
  • Liked: 397
  • Likes Given: 822
Re: Storable Propellant Earth Departure Stages
« Reply #19 on: 08/29/2013 09:26 am »
Exactly. Better is the enemy of good enough. The point is that trivial solutions like a storable propellant earth departure stage is cheaper than SLS.

You don't even need a storable departure stage if you launch the crew first and have them await a second launch that brings them their EDS inside the ISS. The lander is going to be the heaviest component, but only if it's fueled. If the lander is launched mostly dry, you can move it all the way to L1/L2 before finally loading it with storable propellant.

I think the simplest near term architecture is as follows:

- Separately launched Centaur / DCSS based EDS edit: or F9US
- Crew staging point at ISS
- Dragon
- Hypergolic horizontal lander that isn't loaded with propellant until it gets to L1/L2

The lander could be reusable, but doesn't have to be. Just a reusable ascent stage with an (obviously expendable) crasher stage would be nice already.

And why bother with refueling? Because given the required amounts of propellant it would create a large and fiercely competitive market for launch services, which could drive down launch prices by an order of magnitude or more over the course of a decade. If that were to happen, fully commercial spaceflight would likely be a reality.
« Last Edit: 08/29/2013 01:28 pm by mmeijeri »
Pro-tip: you don't have to be a jerk if someone doesn't agree with your theories

Offline Archibald

  • Senior Member
  • *****
  • Posts: 2611
  • Liked: 499
  • Likes Given: 1096
Re: Storable Propellant Earth Departure Stages
« Reply #20 on: 08/29/2013 09:34 am »
Quote
Reason they would not used such a system is they don't want to go to the moon or any crewed BLEO program. That has been the real reason we have been strung along. They had the option with Atlas V and DIV for Lunar before investing in the HLV.

they, they. who is they, btw ? the illuminatis ?
Han shot first and Gwynne Shotwell !

Offline Downix

  • Senior Member
  • *****
  • Posts: 7082
  • Liked: 22
  • Likes Given: 1
Re: Storable Propellant Earth Departure Stages
« Reply #21 on: 08/29/2013 03:20 pm »
I checked yours. You neglected the weight of the lander itself.

No I didn't.

Dry mass: 10458
Payload mass: 43000
Total mass at burnout: 53458
Isp: 312
Delta-v: 3107

Apply the rocket equation, you get: 147579 in LEO.
Subtract the payload, you get: 104579
Subtract the dry mass, you get: 94121
You are right, I messed up my math. Does not matter the app used if you use the wrong numbers in the wrong spots.

This is why I normally don't get into math debates after 8pm, my brain turns into a pumpkin.
Quote
Quote
Quote from: QuantumG
Only in pixie land is that the cost of an SLS launch.


That's the cost on-record.

Perhaps you mean that's the marginal cost that they made up, including none of the development costs, all of which could be avoided by just not building the pointless monster rocket.
All of the Falcon Heavy's development costs (which need to be recouped) can also be forgone if they never develop it either, so your argument is moot. Unlike the FH, the SLS is not driven by a desire to recoup the investment, but instead for marginal cost to deliverable.

Basic civics. You cannot compare a government vs a corporate development, their metrics are radically different.
chuck - Toilet paper has no real value? Try living with 5 other adults for 6 months in a can with no toilet paper. Man oh man. Toilet paper would be worth it's weight in gold!

Offline Lobo

  • Senior Member
  • *****
  • Posts: 6915
  • Spokane, WA
  • Liked: 672
  • Likes Given: 436
Re: Storable Propellant Earth Departure Stages
« Reply #22 on: 08/29/2013 03:42 pm »
It only has an isp of about 312 s?

For the sake of argument.



By Storabe, you mean hyperogolics like MMH/N2O4?

Well, having large quantities of toxic storables at the pad for fueling could be less than desirable, especially for a crewed launch.  Hypergolics are expensive and difficult to handle, which is one of the several reasons I've read that Titan IV was retired and replaced with EELV's using cryogenics. 

I think methalox can be stored in space though, almost indefinately, so I'd probably be more keen to ponder that route for an EDS with long loiter capability.

Offline mmeijeri

  • Senior Member
  • *****
  • Posts: 7772
  • Martijn Meijering
  • NL
  • Liked: 397
  • Likes Given: 822
Re: Storable Propellant Earth Departure Stages
« Reply #23 on: 08/29/2013 03:59 pm »
Well, having large quantities of toxic storables at the pad for fueling could be less than desirable, especially for a crewed launch.  Hypergolics are expensive and difficult to handle, which is one of the several reasons I've read that Titan IV was retired and replaced with EELV's using cryogenics. 

Everybody who launches stuff launches hypergolics as well, and knows how to handle them. There are (much) less toxic alternatives, but it's not enough of a problem to switch. As for crewed launches, Orion and Dragon both use hypergolics, as did the Shuttle. And the propellant for an EDS or lander need not be launched together with the crew. In fact, the ability to offload propellant without new technology development is a large part of the attractiveness of hypergolics for this application.
« Last Edit: 08/29/2013 04:01 pm by mmeijeri »
Pro-tip: you don't have to be a jerk if someone doesn't agree with your theories

Offline RocketmanUS

  • Senior Member
  • *****
  • Posts: 2226
  • USA
  • Liked: 71
  • Likes Given: 31
Re: Storable Propellant Earth Departure Stages
« Reply #24 on: 08/29/2013 04:47 pm »
Quote
Reason they would not used such a system is they don't want to go to the moon or any crewed BLEO program. That has been the real reason we have been strung along. They had the option with Atlas V and DIV for Lunar before investing in the HLV.

they, they. who is they, btw ? the illuminatis ?
Good one.

However I mean Congress and Bush and then Congress and Obama.
If either of these had wanted a return Lunar landing we would have seen it by now. We have had the tech and funds to do it if we did not waste it on Ares I/V. If we had designed the Orion for Lunar it would have been less mass and smaller diameter. Atlas V human rated. Atlas V 552 and Delta IVH for cargo, lander, tanker. Could have had in space fulling with storables.

Offline Lobo

  • Senior Member
  • *****
  • Posts: 6915
  • Spokane, WA
  • Liked: 672
  • Likes Given: 436
Re: Storable Propellant Earth Departure Stages
« Reply #25 on: 08/29/2013 04:54 pm »
Everybody who launches stuff launches hypergolics as well, and knows how to handle them. There are (much) less toxic alternatives, but it's not enough of a problem to switch. As for crewed launches, Orion and Dragon both use hypergolics, as did the Shuttle. And the propellant for an EDS or lander need not be launched together with the crew. In fact, the ability to offload propellant without new technology development is a large part of the attractiveness of hypergolics for this application.

That's true, but in smaller quantities.  for RCS/OMS systems.  The last US LV to use the kind of hypergolic quantities an EDS would need was Titan IV, and the cost and difficulty of handling the hypergolics was a reason given for it's rise in costs and eventual retirement.  Titan III was pretty cheap, so either hypergolics were cheaper in the 60's and 70's, or they didn't have the same safety concerns with handling it, or whatever.  Something went up later on in the Titan program.

Proton and Long March both use hypergolics in large amounts, but I don't know they have the same safety concerns we do here.

Not saying those issues are showstoppers by any means.  Just citing a reason against a hypergolic EDS other than it's lower ISP. 




Offline Robotbeat

  • Senior Member
  • *****
  • Posts: 39270
  • Minnesota
  • Liked: 25222
  • Likes Given: 12114
Re: Storable Propellant Earth Departure Stages
« Reply #26 on: 08/29/2013 05:08 pm »
Large commercial satellites are launched with tons of hypergolic propellants, especially large military birds. I don't see this as a serious concern.
Chris  Whoever loves correction loves knowledge, but he who hates reproof is stupid.

To the maximum extent practicable, the Federal Government shall plan missions to accommodate the space transportation services capabilities of United States commercial providers. US law http://goo.gl/YZYNt0

Offline Ben the Space Brit

  • Senior Member
  • *****
  • Posts: 7206
  • A spaceflight fan
  • London, UK
  • Liked: 806
  • Likes Given: 900
Re: Storable Propellant Earth Departure Stages
« Reply #27 on: 08/29/2013 09:40 pm »
I suppose that there must be a 'tipping point' for all types of propellent - where the physical mass of the stage + payload is so great that no matter how long you run the engines, you will never reach escape velocity.  What is that mass for hypergols and heavy cryogens like liquid methane?
"Oops! I left the silly thing in reverse!" - Duck Dodgers

~*~*~*~

The Space Shuttle Program - 1981-2011

The time for words has passed; The time has come to put up or shut up!
DON'T PROPAGANDISE, FLY!!!

Offline mmeijeri

  • Senior Member
  • *****
  • Posts: 7772
  • Martijn Meijering
  • NL
  • Liked: 397
  • Likes Given: 822
Re: Storable Propellant Earth Departure Stages
« Reply #28 on: 08/29/2013 09:43 pm »
Unless I misunderstand what you mean there is no such tipping point.
Pro-tip: you don't have to be a jerk if someone doesn't agree with your theories

Offline Ben the Space Brit

  • Senior Member
  • *****
  • Posts: 7206
  • A spaceflight fan
  • London, UK
  • Liked: 806
  • Likes Given: 900
Re: Storable Propellant Earth Departure Stages
« Reply #29 on: 08/29/2013 09:48 pm »
Unless I misunderstand what you mean there is no such tipping point.

You do misunderstand because every rocket performance list I've ever seen has given a maximum mass it can move to a certain orbit.  So, just how big does propulsion stage have to be when its' own mass exceeds its ability to push that mass through TLI?
"Oops! I left the silly thing in reverse!" - Duck Dodgers

~*~*~*~

The Space Shuttle Program - 1981-2011

The time for words has passed; The time has come to put up or shut up!
DON'T PROPAGANDISE, FLY!!!

Offline Robotbeat

  • Senior Member
  • *****
  • Posts: 39270
  • Minnesota
  • Liked: 25222
  • Likes Given: 12114
Re: Storable Propellant Earth Departure Stages
« Reply #30 on: 08/29/2013 09:48 pm »
Indeed! You may get greater gravity losses if your engines aren't big enough, but that's it. And even that can be mitigated with multiple passes.
Chris  Whoever loves correction loves knowledge, but he who hates reproof is stupid.

To the maximum extent practicable, the Federal Government shall plan missions to accommodate the space transportation services capabilities of United States commercial providers. US law http://goo.gl/YZYNt0

Offline mmeijeri

  • Senior Member
  • *****
  • Posts: 7772
  • Martijn Meijering
  • NL
  • Liked: 397
  • Likes Given: 822
Re: Storable Propellant Earth Departure Stages
« Reply #31 on: 08/29/2013 09:50 pm »
You do misunderstand because every rocket performance list I've ever seen has given a maximum mass it can move to a certain orbit.  So, just how big does propulsion stage have to be when its' own mass exceeds its ability to push that mass through TLI?

That limit is for a given size of the stage, but a larger stage will be able to move a larger payload, it's not related to the type of propellant per se.
Pro-tip: you don't have to be a jerk if someone doesn't agree with your theories

Offline Ben the Space Brit

  • Senior Member
  • *****
  • Posts: 7206
  • A spaceflight fan
  • London, UK
  • Liked: 806
  • Likes Given: 900
Re: Storable Propellant Earth Departure Stages
« Reply #32 on: 08/29/2013 09:52 pm »
You do misunderstand because every rocket performance list I've ever seen has given a maximum mass it can move to a certain orbit.  So, just how big does propulsion stage have to be when its' own mass exceeds its ability to push that mass through TLI?

That limit is for a given size of the stage, but a larger stage will be able to move a larger payload, it's not related to the type of propellant per se.

Different propellent have different energies (Isp).  There must be a mass where attempting to move it with the number of engines that a tank of that size can reasonably support will never add sufficient energy to the system to reach escape velocity before the propellent is exhausted.  It could be that this figure is so large that it isn't a real issue but the limit must exist.
« Last Edit: 08/29/2013 09:53 pm by Ben the Space Brit »
"Oops! I left the silly thing in reverse!" - Duck Dodgers

~*~*~*~

The Space Shuttle Program - 1981-2011

The time for words has passed; The time has come to put up or shut up!
DON'T PROPAGANDISE, FLY!!!

Offline QuantumG

  • Senior Member
  • *****
  • Posts: 9238
  • Australia
  • Liked: 4477
  • Likes Given: 1108
Re: Storable Propellant Earth Departure Stages
« Reply #33 on: 08/29/2013 10:01 pm »
Another way to say what Ben is saying: you'll need an engine for this big EDS stage that can produce sufficient thrust, what is it?

I don't know. I'm just speculating that there's an available engine, or something could be modified for the job.
Human spaceflight is basically just LARPing now.

Offline Robotbeat

  • Senior Member
  • *****
  • Posts: 39270
  • Minnesota
  • Liked: 25222
  • Likes Given: 12114
Re: Storable Propellant Earth Departure Stages
« Reply #34 on: 08/29/2013 10:02 pm »
You do misunderstand because every rocket performance list I've ever seen has given a maximum mass it can move to a certain orbit.  So, just how big does propulsion stage have to be when its' own mass exceeds its ability to push that mass through TLI?

That limit is for a given size of the stage, but a larger stage will be able to move a larger payload, it's not related to the type of propellant per se.

Different propellent have different energies (Isp).  There must be a mass where attempting to move it with the number of engines that a tank of that size can reasonably support will never add sufficient energy to the system to reach escape velocity before the propellent is exhausted.  It could be that this figure is so large that it isn't a real issue but the limit must exist.
No limit, if you allow staging.

And even with staging, you're limited by the tank payload fraction, not by the engine thrust. You can have arbitrarily low thrust and still get to orbit. If you have a 90% mass fraction, you can do about 7km/s delta-v, way more than enough to get to escape velocity even at the low-thrust limit (though that requires a few more passes and obviously longer time).

Think about electric propulsion, which operates often at the approximately zero-thrust limit. Technically the low-thrust limit is about 7-7.5km/s to escape from LEO, but that assumes no optimization of the trajectory. If you shape the trajectory so that you have a whole bunch of orbits and burn just at perigee, you can approach the infinite-thrust delta-v needed for escape (which is about 3.2km/s) even with arbitrarily low thrust (though it will take an incredibly long time as you approach the limit).

https://en.wikipedia.org/wiki/Tsiolkovsky_rocket_equation
And this now has a section on arbitrarily-low-thrust delta-vs:
https://en.wikipedia.org/wiki/Delta-v_budget
« Last Edit: 08/29/2013 10:09 pm by Robotbeat »
Chris  Whoever loves correction loves knowledge, but he who hates reproof is stupid.

To the maximum extent practicable, the Federal Government shall plan missions to accommodate the space transportation services capabilities of United States commercial providers. US law http://goo.gl/YZYNt0

Offline mmeijeri

  • Senior Member
  • *****
  • Posts: 7772
  • Martijn Meijering
  • NL
  • Liked: 397
  • Likes Given: 822
Re: Storable Propellant Earth Departure Stages
« Reply #35 on: 08/29/2013 10:04 pm »
Different propellent have different energies (Isp).  There must be a mass where attempting to move it with the number of engines that a tank of that size can reasonably support will never add sufficient energy to the system to reach escape velocity before the propellent is exhausted.  It could be that this figure is so large that it isn't a real issue but the limit must exist.

Isp isn't that important, density impulse is more important, but there hypergolics score better than LOX/LH2. In the end it boils down to the size of payload fairings. That might argue in favour of dense propellants and against LOX/LH2, but if you use staging at a Lagrange point as you should, then LOX/LH2 is absolutely fine too.
Pro-tip: you don't have to be a jerk if someone doesn't agree with your theories

Offline mmeijeri

  • Senior Member
  • *****
  • Posts: 7772
  • Martijn Meijering
  • NL
  • Liked: 397
  • Likes Given: 822
Re: Storable Propellant Earth Departure Stages
« Reply #36 on: 08/29/2013 10:09 pm »
Another way to say what Ben is saying: you'll need an engine for this big EDS stage that can produce sufficient thrust, what is it?

You would need to develop a pump-fed engine. That's part of the reason I'm not keen on a hypergolic stage for use from LEO (rather than L1/L2), the other being a needless performance penalty. I think new pump-fed hypergolic engines for manned deep-space missions are an excellent idea, but I'd want them off the critical path. Pressure-fed engines are fine for use from L1/L2, even for limited manned missions, including moon landings.

You could also use existing Russian engines of course, say a cluster of first stage engines modified for vacuum use. They have 340s engines. Thrust won't be a problem at all.
« Last Edit: 08/29/2013 10:16 pm by mmeijeri »
Pro-tip: you don't have to be a jerk if someone doesn't agree with your theories

Offline Robotbeat

  • Senior Member
  • *****
  • Posts: 39270
  • Minnesota
  • Liked: 25222
  • Likes Given: 12114
Re: Storable Propellant Earth Departure Stages
« Reply #37 on: 08/29/2013 10:11 pm »
LR87 with a big bell on it should get plenty of thrust plus about the Isp that QuantumG is looking for.
Chris  Whoever loves correction loves knowledge, but he who hates reproof is stupid.

To the maximum extent practicable, the Federal Government shall plan missions to accommodate the space transportation services capabilities of United States commercial providers. US law http://goo.gl/YZYNt0

Offline mmeijeri

  • Senior Member
  • *****
  • Posts: 7772
  • Martijn Meijering
  • NL
  • Liked: 397
  • Likes Given: 822
Re: Storable Propellant Earth Departure Stages
« Reply #38 on: 08/29/2013 10:13 pm »
The advantage of Russian engines is that they use staged combustion, or at least some of them do, and that they exist already. The disadvantage is that they're, well, Russian. Then again, that's also true of RD-180, so it's not decisive.
« Last Edit: 08/29/2013 10:36 pm by mmeijeri »
Pro-tip: you don't have to be a jerk if someone doesn't agree with your theories

Offline joek

  • Senior Member
  • *****
  • Posts: 4860
  • Liked: 2780
  • Likes Given: 1095
Re: Storable Propellant Earth Departure Stages
« Reply #39 on: 08/29/2013 10:34 pm »
According to my math, please check me, I figure it would be about 10,458 kg dry and 104,579 kg when full. This would provide the 3107 m/s of delta-v to get through TLI, with the lunar insertion to be done by the lander (as in the NASA architecture).

I know 105 tons sounds like a lot, but it's only two Falcon Heavy launches, and because we're using storable propellant there's no time pressure. If you really wanted to you could do it with Falcon 9 v1.1.

Are there safety and pad infrastructure considerations associated with launching that much storable propellant (presumably nasty hypergolics) which would present significant challenges?  Granted, the Russians deal with mass quantities of the stuff regularly, and the US did previously with Titan, but it seems to me those considerations may be a significant factor?

Offline mmeijeri

  • Senior Member
  • *****
  • Posts: 7772
  • Martijn Meijering
  • NL
  • Liked: 397
  • Likes Given: 822
Re: Storable Propellant Earth Departure Stages
« Reply #40 on: 08/29/2013 10:36 pm »
If you think toxicity is an issue, you could use kerosene / peroxide too. Or even peroxide as a monopropellant.
Pro-tip: you don't have to be a jerk if someone doesn't agree with your theories

Offline Robotbeat

  • Senior Member
  • *****
  • Posts: 39270
  • Minnesota
  • Liked: 25222
  • Likes Given: 12114
Re: Storable Propellant Earth Departure Stages
« Reply #41 on: 08/29/2013 10:49 pm »
According to my math, please check me, I figure it would be about 10,458 kg dry and 104,579 kg when full. This would provide the 3107 m/s of delta-v to get through TLI, with the lunar insertion to be done by the lander (as in the NASA architecture).

I know 105 tons sounds like a lot, but it's only two Falcon Heavy launches, and because we're using storable propellant there's no time pressure. If you really wanted to you could do it with Falcon 9 v1.1.

Are there safety and pad infrastructure considerations associated with launching that much storable propellant (presumably nasty hypergolics) which would present significant challenges?  Granted, the Russians deal with mass quantities of the stuff regularly, and the US did previously with Titan, but it seems to me those considerations may be a significant factor?
No, not a significant factor. If you launched the propellant on a F9R or something, each launch would have about as much hypergolic propellant as a single Delta II launch (launching a bird into LEO) or a D4H launch with a big spy sat. Different from an entire first stage with hundreds of tons of hypergols like Proton (600 tons of hypergolic propellants).

I don't think it's significant.
Chris  Whoever loves correction loves knowledge, but he who hates reproof is stupid.

To the maximum extent practicable, the Federal Government shall plan missions to accommodate the space transportation services capabilities of United States commercial providers. US law http://goo.gl/YZYNt0

Offline QuantumG

  • Senior Member
  • *****
  • Posts: 9238
  • Australia
  • Liked: 4477
  • Likes Given: 1108
Re: Storable Propellant Earth Departure Stages
« Reply #42 on: 08/29/2013 10:52 pm »
Clearly, any risk is significant.. we can't ever take any risk in order to perform spaceflight. Better that NASA just spend billions and not fly anyone or anything.

Human spaceflight is basically just LARPing now.

Offline Lobo

  • Senior Member
  • *****
  • Posts: 6915
  • Spokane, WA
  • Liked: 672
  • Likes Given: 436
Re: Storable Propellant Earth Departure Stages
« Reply #43 on: 08/29/2013 10:55 pm »
Large commercial satellites are launched with tons of hypergolic propellants, especially large military birds. I don't see this as a serious concern.

Because I SAID it was, darnit!!

Stop asking so many questions...

;-)


Offline Lobo

  • Senior Member
  • *****
  • Posts: 6915
  • Spokane, WA
  • Liked: 672
  • Likes Given: 436
Re: Storable Propellant Earth Departure Stages
« Reply #44 on: 08/29/2013 11:02 pm »
Unless I misunderstand what you mean there is no such tipping point.

You do misunderstand because every rocket performance list I've ever seen has given a maximum mass it can move to a certain orbit.  So, just how big does propulsion stage have to be when its' own mass exceeds its ability to push that mass through TLI?

Not sure if there's such a point for hypergolics.  Titan had a hypergolic core and 2nd stage, as does Proton.  Proton uses a hypergolic 3rd stage too.  Although Titan used Centaur when it needed the 3rd stage.

Not sure if there's any physical reason you can't make a Saturn V class hypergolic rocket. 
But that is a boat load of nasty chemicals to handle and get to the launch pad.
My point earlier is a large upper stage is an order of magnitude larger volume that some RCS thruster tanks.  And here in the US, we moved intentionally away form hypergolic boosters to cryogenic ones.
I'm assuming there are reasons for that?  Was there a hypergolic bidder in the EELV competition?  Or only cryogenic?  Was the USAF's RFQ specifying only cryogenics?
I don't know. 


Offline Robotbeat

  • Senior Member
  • *****
  • Posts: 39270
  • Minnesota
  • Liked: 25222
  • Likes Given: 12114
Re: Storable Propellant Earth Departure Stages
« Reply #45 on: 08/29/2013 11:07 pm »
Unless I misunderstand what you mean there is no such tipping point.

You do misunderstand because every rocket performance list I've ever seen has given a maximum mass it can move to a certain orbit.  So, just how big does propulsion stage have to be when its' own mass exceeds its ability to push that mass through TLI?
Can you rephrase this?
Chris  Whoever loves correction loves knowledge, but he who hates reproof is stupid.

To the maximum extent practicable, the Federal Government shall plan missions to accommodate the space transportation services capabilities of United States commercial providers. US law http://goo.gl/YZYNt0

Offline joek

  • Senior Member
  • *****
  • Posts: 4860
  • Liked: 2780
  • Likes Given: 1095
Re: Storable Propellant Earth Departure Stages
« Reply #46 on: 08/29/2013 11:15 pm »
If you think toxicity is an issue, you could use kerosene / peroxide too. Or even peroxide as a monopropellant.

True, but the issue is a bit broader: Current infrastructure and payload processing is oriented around spacecraft, not bulk propellants.  What, if any, significance does that have in the decision?

Clearly, any risk is significant.. we can't ever take any risk in order to perform spaceflight. Better that NASA just spend billions and not fly anyone or anything.

Reductio ad absurdum.  Again: How do current infrastructure and payload processing capabilities weigh in the balance?  It's not as if we can simply drag a couple hoses with your propellants of choice to the spacecraft-tanker-EDS-whatever.  Or can we?  And if we can, is that a solution we would want to bank on for the future, or simply to satisfy short term objectives?

Offline mmeijeri

  • Senior Member
  • *****
  • Posts: 7772
  • Martijn Meijering
  • NL
  • Liked: 397
  • Likes Given: 822
Re: Storable Propellant Earth Departure Stages
« Reply #47 on: 08/29/2013 11:19 pm »
True, but the issue is a bit broader: Current infrastructure and payload processing is oriented around spacecraft, not bulk propellants.  What, if any, significance does that have in the decision?

With EELVs we're not talking about radically larger amounts of propellant compared to spacecraft, especially if you launch propellant to L1/L2 directly.
Pro-tip: you don't have to be a jerk if someone doesn't agree with your theories

Offline joek

  • Senior Member
  • *****
  • Posts: 4860
  • Liked: 2780
  • Likes Given: 1095
Re: Storable Propellant Earth Departure Stages
« Reply #48 on: 08/29/2013 11:27 pm »
True, but the issue is a bit broader: Current infrastructure and payload processing is oriented around spacecraft, not bulk propellants.  What, if any, significance does that have in the decision?
With EELVs we're not talking about radically larger amounts of propellant compared to spacecraft, especially if you launch propellant to L1/L2 directly.
Per QuantumG's OP, we would be talking about a radically larger amount of propellant, with the vast majority of the mass being propellant instead of spacecraft (~50t on an F9H).

Offline Robotbeat

  • Senior Member
  • *****
  • Posts: 39270
  • Minnesota
  • Liked: 25222
  • Likes Given: 12114
Re: Storable Propellant Earth Departure Stages
« Reply #49 on: 08/29/2013 11:31 pm »
True, but the issue is a bit broader: Current infrastructure and payload processing is oriented around spacecraft, not bulk propellants.  What, if any, significance does that have in the decision?
With EELVs we're not talking about radically larger amounts of propellant compared to spacecraft, especially if you launch propellant to L1/L2 directly.
Per QuantumG's OP, we would be talking about a radically larger amount of propellant, with the vast majority of the mass being propellant instead of spacecraft (~50t on an F9H).
Mmeijeri said EELVs, you just said FH which is twice as big as the next biggest EELV. Atlas V is probably what you'd use (if you used an EELV) for a tanker, so 18 tons or so, which is only about two or three times a Delta II launch or a launch of a big spacecraft. Not radically more, less than half an order of magnitude.
Chris  Whoever loves correction loves knowledge, but he who hates reproof is stupid.

To the maximum extent practicable, the Federal Government shall plan missions to accommodate the space transportation services capabilities of United States commercial providers. US law http://goo.gl/YZYNt0

Offline mmeijeri

  • Senior Member
  • *****
  • Posts: 7772
  • Martijn Meijering
  • NL
  • Liked: 397
  • Likes Given: 822
Re: Storable Propellant Earth Departure Stages
« Reply #50 on: 08/29/2013 11:37 pm »
Per QuantumG's OP, we would be talking about a radically larger amount of propellant, with the vast majority of the mass being propellant instead of spacecraft (~50t on an F9H).

OK, that's one order of magnitude larger. I think using a kerolox or hydrolox stage to L1/L2 and hypergolics from there onward is better, but not necessarily because of this. I'd be surprised if 50mT was a problem, but I don't know that for a fact.
Pro-tip: you don't have to be a jerk if someone doesn't agree with your theories

Offline joek

  • Senior Member
  • *****
  • Posts: 4860
  • Liked: 2780
  • Likes Given: 1095
Re: Storable Propellant Earth Departure Stages
« Reply #51 on: 08/29/2013 11:45 pm »
Mmeijeri said EELVs, you just said FH which is twice as big as the next biggest EELV. Atlas V is probably what you'd use (if you used an EELV) for a tanker, so 18 tons or so, which is only about two or three times a Delta II launch or a launch of a big spacecraft. Not radically more, less than half an order of magnitude.

I used FH as an example as that is what was posed by QuantumG in the OP--as I stated.  If you wish to change the goalposts and posit an EELV, feel free, but please take it elsewhere.  That, however, does not invalidate the question of whether using FH (or equivalent)--where propellants are the primary payload--and associated infrastructure requirements would be a significant factor in determining the choice of propellants.
« Last Edit: 08/29/2013 11:48 pm by joek »

Offline Robotbeat

  • Senior Member
  • *****
  • Posts: 39270
  • Minnesota
  • Liked: 25222
  • Likes Given: 12114
Re: Storable Propellant Earth Departure Stages
« Reply #52 on: 08/29/2013 11:48 pm »
Goalposts? You were responding to Mmeij who said EELV.

And besides, the propellants and hypergols are not added at the pad but during spacecraft processing. If anything, this would support the use of hypergols, since the infrastructure is already in place to fuel up payloads with hypergols.
« Last Edit: 08/29/2013 11:50 pm by Robotbeat »
Chris  Whoever loves correction loves knowledge, but he who hates reproof is stupid.

To the maximum extent practicable, the Federal Government shall plan missions to accommodate the space transportation services capabilities of United States commercial providers. US law http://goo.gl/YZYNt0

Offline Robotbeat

  • Senior Member
  • *****
  • Posts: 39270
  • Minnesota
  • Liked: 25222
  • Likes Given: 12114
Re: Storable Propellant Earth Departure Stages
« Reply #53 on: 08/29/2013 11:54 pm »
Also, if you're going to launch a 43mT lander that's likely going to be hypergol, then a 50mT launch of hypergolic propellant is in the same league.
Chris  Whoever loves correction loves knowledge, but he who hates reproof is stupid.

To the maximum extent practicable, the Federal Government shall plan missions to accommodate the space transportation services capabilities of United States commercial providers. US law http://goo.gl/YZYNt0

Offline QuantumG

  • Senior Member
  • *****
  • Posts: 9238
  • Australia
  • Liked: 4477
  • Likes Given: 1108
Re: Storable Propellant Earth Departure Stages
« Reply #54 on: 08/30/2013 12:11 am »
Also, if you're going to launch a 43mT lander that's likely going to be hypergol, then a 50mT launch of hypergolic propellant is in the same league.

Exactly. Ya know what else is an order of magnitude bigger joek? The whole pointless SLS program! Remember the damn context.
« Last Edit: 08/30/2013 12:11 am by QuantumG »
Human spaceflight is basically just LARPing now.

Offline mmeijeri

  • Senior Member
  • *****
  • Posts: 7772
  • Martijn Meijering
  • NL
  • Liked: 397
  • Likes Given: 822
Re: Storable Propellant Earth Departure Stages
« Reply #55 on: 08/30/2013 12:18 am »
I don't think FH is crucial to Trent's point, it's about whether SLS is necessary to get a 43mT lander to lunar orbit as NASA claims. That simply isn't the case as Trent's example shows. It's been a while since I did the calculations, but by my calculations you can get a 100mT lander to L1/L2 with just EELVs, hypergolic refueling at L1/L2, and EELV-derived EDSs. And then you wouldn't need radically larger amounts of hypergolics per launch, though you would need a lot of launches.
« Last Edit: 08/30/2013 12:18 am by mmeijeri »
Pro-tip: you don't have to be a jerk if someone doesn't agree with your theories

Offline joek

  • Senior Member
  • *****
  • Posts: 4860
  • Liked: 2780
  • Likes Given: 1095
Re: Storable Propellant Earth Departure Stages
« Reply #56 on: 08/30/2013 12:27 am »
Also, if you're going to launch a 43mT lander that's likely going to be hypergol, then a 50mT launch of hypergolic propellant is in the same league.
Exactly. Ya know what else is an order of magnitude bigger joek? The whole pointless SLS program! Remember the damn context.

Yeah, I get the SLS bit and context.  So back to the point... If you are considering which propellants to use, would not gound handling and infrastructure considerations be a significant consideration in your decision?

Offline Robotbeat

  • Senior Member
  • *****
  • Posts: 39270
  • Minnesota
  • Liked: 25222
  • Likes Given: 12114
Re: Storable Propellant Earth Departure Stages
« Reply #57 on: 08/30/2013 12:28 am »
I don't think FH is crucial to Trent's point, it's about whether SLS is necessary to get a 43mT lander to lunar orbit as NASA claims. That simply isn't the case as Trent's example shows. It's been a while since I did the calculations, but by my calculations you can get a 100mT lander to L1/L2 with just EELVs, hypergolic refueling at L1/L2, and EELV-derived EDSs. And then you wouldn't need radically larger amounts of hypergolics per launch, though you would need a lot of launches.
right, with fully expendable EELVs and no FH and no electric propulsion, deep-space hypergolic refueling makes a lot of sense. You could halve or third the number of required launches for the lander if you used SEP, though...
Chris  Whoever loves correction loves knowledge, but he who hates reproof is stupid.

To the maximum extent practicable, the Federal Government shall plan missions to accommodate the space transportation services capabilities of United States commercial providers. US law http://goo.gl/YZYNt0

Offline QuantumG

  • Senior Member
  • *****
  • Posts: 9238
  • Australia
  • Liked: 4477
  • Likes Given: 1108
Re: Storable Propellant Earth Departure Stages
« Reply #58 on: 08/30/2013 12:32 am »
Yeah, I get the SLS bit and context.  So back to the point... If you are considering which propellants to use, would not ground handling and infrastructure considerations be a significant consideration in your decision?

Sure, but my question to you is: what's easier? Handling toxic propellants or getting SpaceX to launch on time?
Human spaceflight is basically just LARPing now.

Offline Robotbeat

  • Senior Member
  • *****
  • Posts: 39270
  • Minnesota
  • Liked: 25222
  • Likes Given: 12114
Re: Storable Propellant Earth Departure Stages
« Reply #59 on: 08/30/2013 12:37 am »
LOL.
Chris  Whoever loves correction loves knowledge, but he who hates reproof is stupid.

To the maximum extent practicable, the Federal Government shall plan missions to accommodate the space transportation services capabilities of United States commercial providers. US law http://goo.gl/YZYNt0

Offline RocketmanUS

  • Senior Member
  • *****
  • Posts: 2226
  • USA
  • Liked: 71
  • Likes Given: 31
Re: Storable Propellant Earth Departure Stages
« Reply #60 on: 08/30/2013 12:52 am »
Version of the Super Draco for the EDS engine(s).

Seven at 15,000lb thrust each would give 105,00lb thrust.

That should be good enough t/w for TLI.

For the lander it could take the six day verses the three day voyage.
That should give the EDS good enough margin.


Offline mmeijeri

  • Senior Member
  • *****
  • Posts: 7772
  • Martijn Meijering
  • NL
  • Liked: 397
  • Likes Given: 822
Re: Storable Propellant Earth Departure Stages
« Reply #61 on: 08/30/2013 12:58 am »
right, with fully expendable EELVs and no FH and no electric propulsion, deep-space hypergolic refueling makes a lot of sense. You could halve or third the number of required launches for the lander if you used SEP, though...

Yeah, and if you procured propellant at L1/L2 competitively, then that might close the business case for development of a commercial SEP tug, especially since most of the technology is available off the shelf on comsats.
Pro-tip: you don't have to be a jerk if someone doesn't agree with your theories

Offline mmeijeri

  • Senior Member
  • *****
  • Posts: 7772
  • Martijn Meijering
  • NL
  • Liked: 397
  • Likes Given: 822
Re: Storable Propellant Earth Departure Stages
« Reply #62 on: 08/30/2013 01:00 am »
SuperDraco probably has very low Isp.
Pro-tip: you don't have to be a jerk if someone doesn't agree with your theories

Offline Robotbeat

  • Senior Member
  • *****
  • Posts: 39270
  • Minnesota
  • Liked: 25222
  • Likes Given: 12114
Re: Storable Propellant Earth Departure Stages
« Reply #63 on: 08/30/2013 01:03 am »
SuperDraco probably has very low Isp.
...and would have heavy tanks.
Chris  Whoever loves correction loves knowledge, but he who hates reproof is stupid.

To the maximum extent practicable, the Federal Government shall plan missions to accommodate the space transportation services capabilities of United States commercial providers. US law http://goo.gl/YZYNt0

Offline RocketmanUS

  • Senior Member
  • *****
  • Posts: 2226
  • USA
  • Liked: 71
  • Likes Given: 31
Re: Storable Propellant Earth Departure Stages
« Reply #64 on: 08/30/2013 01:07 am »
SuperDraco probably has very low Isp.
...and would have heavy tanks.
Why would the tank mass change do to using a version of the Super Draco?

Offline A_M_Swallow

  • Elite Veteran
  • Senior Member
  • *****
  • Posts: 8906
  • South coast of England
  • Liked: 500
  • Likes Given: 223
Re: Storable Propellant Earth Departure Stages
« Reply #65 on: 08/30/2013 01:07 am »
The Morpheus engine has an Isp 321, 5000 lbf, pressure fed and burns methane/LOX.

It was designed for a lander but you can use it in a tug if you want.

Offline QuantumG

  • Senior Member
  • *****
  • Posts: 9238
  • Australia
  • Liked: 4477
  • Likes Given: 1108
Re: Storable Propellant Earth Departure Stages
« Reply #66 on: 08/30/2013 01:08 am »
Because SuperDraco is pressure fed.. if it were pump fed, it wouldn't be a SuperDraco.

You're not getting a high mass fraction with a pressure fed engine.
Human spaceflight is basically just LARPing now.

Offline RocketmanUS

  • Senior Member
  • *****
  • Posts: 2226
  • USA
  • Liked: 71
  • Likes Given: 31
Re: Storable Propellant Earth Departure Stages
« Reply #67 on: 08/30/2013 01:09 am »
Because SuperDraco is pressure fed.. if it were pump fed, it wouldn't be a SuperDraco.

You're not getting a high mass fraction with a pressure fed engine.
So tank pressure is higher?

Offline Robotbeat

  • Senior Member
  • *****
  • Posts: 39270
  • Minnesota
  • Liked: 25222
  • Likes Given: 12114
Re: Storable Propellant Earth Departure Stages
« Reply #68 on: 08/30/2013 01:14 am »
Because SuperDraco is pressure fed.. if it were pump fed, it wouldn't be a SuperDraco.

You're not getting a high mass fraction with a pressure fed engine.
So tank pressure is higher?
No, SuperDraco is pressure-fed, and the pressure in the tanks must be at least the pressure in the combustion chamber for pressure-fed engines. So you need high pressure tanks, which are heavy, or you need a much lower pressure (and larger and much less efficient) combustion chamber. A PUMP-fed engine can have a quite low tank pressure while still having a high chamber pressure (and thus smaller and lighter chamber and much more efficient), and thus much lighter tanks, ala Falcon 9 and Falcon Heavy.

(Falcon 1 upper stage Kestrel engine was pressure-fed, by the way!)
« Last Edit: 08/30/2013 01:15 am by Robotbeat »
Chris  Whoever loves correction loves knowledge, but he who hates reproof is stupid.

To the maximum extent practicable, the Federal Government shall plan missions to accommodate the space transportation services capabilities of United States commercial providers. US law http://goo.gl/YZYNt0

Online sdsds

  • Senior Member
  • *****
  • Posts: 7194
  • “With peace and hope for all mankind.”
  • Seattle
  • Liked: 2039
  • Likes Given: 1962
Re: Storable Propellant Earth Departure Stages
« Reply #69 on: 08/30/2013 01:16 am »
Use a small cluster of Aestus II (RS 72) engines? Hot-fire tested; Vac. Isp of 340 sec.

http://cs.astrium.eads.net/sp/launcher-propulsion/rocket-engines/aestus-rs72-rocket-engine.html
— 𝐬𝐝𝐒𝐝𝐬 —

Offline QuantumG

  • Senior Member
  • *****
  • Posts: 9238
  • Australia
  • Liked: 4477
  • Likes Given: 1108
Re: Storable Propellant Earth Departure Stages
« Reply #70 on: 08/30/2013 01:18 am »
Use a small cluster of Aestus II (RS 72) engines? Hot-fire tested; Vac. Isp of 340 sec.

http://cs.astrium.eads.net/sp/launcher-propulsion/rocket-engines/aestus-rs72-rocket-engine.html

mmm.. that's nice.
Human spaceflight is basically just LARPing now.

Offline mmeijeri

  • Senior Member
  • *****
  • Posts: 7772
  • Martijn Meijering
  • NL
  • Liked: 397
  • Likes Given: 822
Re: Storable Propellant Earth Departure Stages
« Reply #71 on: 08/30/2013 01:03 pm »
I said the Russians had 340s hypergolic engines, but I can't find that information anymore. I may have been confusing it with the Aestus 2 number. But Aestus 2 uses the gas generator cycle, so a staged combustion engine should do better still.
« Last Edit: 08/31/2013 11:21 am by mmeijeri »
Pro-tip: you don't have to be a jerk if someone doesn't agree with your theories

Offline muomega0

  • Full Member
  • ****
  • Posts: 862
  • Liked: 70
  • Likes Given: 1
Re: Storable Propellant Earth Departure Stages
« Reply #72 on: 08/30/2013 02:18 pm »
340..well that is a start, but so far away from 460, no?

According to this article, which has it's own thread, NASA wants a 43 ton lunar lander.. for some reason.. and they say the only way to get a payload that big into lunar orbit is with the SLS.

I don't get it. What's wrong with just using a storable propellant Earth departure stage? Let's be conservative and say it only has an isp of about 312s, and a propellant mass fraction of 90%, how big would it be?

According to my math, please check me, I figure it would be about 10,458 kg dry and 104,579 kg when full. This would provide the 3107 m/s of delta-v to get through TLI, with the lunar insertion to be done by the lander (as in the NASA architecture).
I know 105 tons sounds like a lot, but it's only two Falcon Heavy launches, and because we're using storable propellant there's no time pressure. If you really wanted to you could do it with Falcon 9 v1.1.
A very good idea, kind of "cheap and dirty" lunar architecture. I see your point - Falcon Heavy is cheap, storable propellants are straightforward technology, they don't boiloff with time. I have no doubts this would work pretty well.
So what's wrong, do you ask ?
Simple...
NASA obsession with LOX/LH2. Never, ever, would they consider something else for the TLI. That paradigm also apply to SSTO, btw. It is a little annoying, because issues with liquid hydrogen have long been obvious...  ::)

Yes, obvious, as outlined in the HLV Evolution, Mars DRM 5 with Ares V required five, no six flights due to boiloff.
The six month delay in the dual launch hurts component reliability and requires 210 mT to perform a 120 to 140 mT mission.  Ouch!  Even with so many more assumptions to check, the math is not too far off.  Two upper stages to rendezvous at the moon...why not just stretch the tanks and take advantage of amplification factor?

The good news is that there is an easy, very economical solution:  add some power, cryocoolers, and of course a conical sunshield to the components of an upper stage to create a ZBO LH2 LEO depot that will survive for decades, thereby reducing the LV size required.
IOW:  the best cis-lunar architecture is a L2 Gateway+ZBO depot centric architecture.

Please take some time and review Boiloff:  Active vs Passive Depots to see how easy and economical it is to address LH2 boiloff.  Comments welcome  :)

Saturn V first stage was RP-1, but the rest of the stages were LH2/LOX.  So why is NASA "obsessed" with LOX/LH2? 
Different propellent have different energies (Isp).  There must be a mass where attempting to move it with the number of engines that a tank of that size can reasonably support will never add sufficient energy to the system to reach escape velocity before the propellent is exhausted.  It could be that this figure is so large that it isn't a real issue but the limit must exist.
Isp isn't that important, density impulse is more important, but there hypergolics score better than LOX/LH2. In the end it boils down to the size of payload fairings. That might argue in favour of dense propellants and against LOX/LH2, but if you use staging at a Lagrange point as you should, then LOX/LH2 is absolutely fine too.
ISP dominates for the transfer stage.  Higher ISP reduces the IMLEO for the mission, assuming of course that boiloff is included, if any.

LH2 versus methane versus hypergolic transfer stages
Methane vs LH2 Transfer Stages (not IMLEO)
Depending on the deltaV required, the payload mass fractions for Lox/methane versus LOX/LH2 range from 60 to 80% for the injection stage.   ISP=342, T/W=0.4, dV from 2000 to 4000, mixture ratio of 3.5, which seems quite generous.

The 60 to 80% is for a large number of cases.  For example, for one case, the Methane payload mass fraction is 0.435, and LH2=0.565, so Methane only achieves 77% in the comparison.  So if one required 100,000 kg IMLEO for LH2/LOX, the lower payload mass fraction of methane requires 125,000kg, and for lower ISP hypergolics, its 150,000kg.

But it is not about performance, rather it is about cost, at least in most of NASA, when it comes to obsession.  At $5,000/kg, each mission of methane would cost $125M more and hypergolics $150M more than LH2 based systems, with vary generous assumptions given to methane and hypergolics, including $/kg.

So since engine costs dominate, suppose your methane/hypergolic engines were $1M/ea, one could pay quite the premium for the LH2/LOX engines and still be cost effective. Of course, a major savings is the elimination of unneeded product lines.

LH2 is storable with ZBO equipment without killing mass fraction--place it on a depot and not a refuel-able transfer stage acting like a depot.  Substitute sustainable for Exploring sooner.

Like Apollo, a kero based IMLEO LV may indeed provide the lowest $/kg, and like Apollo, a LH2/LOX based approach appears to be the most cost effective.
« Last Edit: 08/30/2013 04:19 pm by muomega0 »

Offline QuantumG

  • Senior Member
  • *****
  • Posts: 9238
  • Australia
  • Liked: 4477
  • Likes Given: 1108
Re: Storable Propellant Earth Departure Stages
« Reply #73 on: 08/30/2013 10:24 pm »
In this case, we're comparing using storable propellant to building a whole new heavy lift launch vehicle. We know all the drawbacks of storable propellant - including lower isp, less availability of engines, and more launches required - and the analysis still tells us that it's cheaper than SLS. Improvements over storable propellants are nice, but they're not necessary to make the point.
Human spaceflight is basically just LARPing now.

Offline RocketmanUS

  • Senior Member
  • *****
  • Posts: 2226
  • USA
  • Liked: 71
  • Likes Given: 31
Re: Storable Propellant Earth Departure Stages
« Reply #74 on: 08/30/2013 11:11 pm »
In this case, we're comparing using storable propellant to building a whole new heavy lift launch vehicle. We know all the drawbacks of storable propellant - including lower isp, less availability of engines, and more launches required - and the analysis still tells us that it's cheaper than SLS. Improvements over storable propellants are nice, but they're not necessary to make the point.

We could have been back to the moon by now.
With Atlas and Delta launch when we could the lander, EDS, propellants for them. Lots of loiter time. For crew do the same for it's EDS. Then when the launch window is open for Lunar send crew up, have tanker waiting to fill it's CM so it could be launched on the Atlas or Delta ( what ever they human rated ). More cargo flights would have gone anyway if going for a base for longer stays over short exploration stays.

1 ) LEO crew taxi and cargo version for LEO/ISS.
2 ) Lunar lander, EDS, tankers, rovers, habs and other related hardware
     ( Send rovers first testing out system before ever sending crew ).
3 ) Crew Lunar capsule and Lunar ascent/crew cab for Lunar lander.

Develop in that order over time as funding was available.
This leaves time and funding for the HLV development for Mars later on.

Hypergolics could have gotten us back to the moon and affordable on a yearly budget. HLV's might with cryo propellants make it better and or less cost later on, however we didn't need to bet that we would get a HLV. Let alone the funding for a Lunar program. VSE might have seen a Lunar landing of at least cargo by now and the crew and cargo LEO taxi's if we had gone this route. One thing we don't know is if we can life on the moon or how long we can stay. This could have answered that question without to great an investment and a long term commitment.

Offline mmeijeri

  • Senior Member
  • *****
  • Posts: 7772
  • Martijn Meijering
  • NL
  • Liked: 397
  • Likes Given: 822
Re: Storable Propellant Earth Departure Stages
« Reply #75 on: 08/30/2013 11:38 pm »
Hypergolics could have gotten us back to the moon and affordable on a yearly budget.

Absolutely. And if you had used a hybrid of prefueled cryogenics and hypergolic refueling at L1/L2, you could have created a large and fiercely competitive market for launch services right away, with only a small (~10%) mass penalty. And since this could have led to commercial development of SEP tugs soon, you might even end up with lower IMLEO sooner, not to mention lower launch costs per kg.

Quote
HLV's might with cryo propellants make it better and or less cost later on, however we didn't need to bet that we would get a HLV.

Cryos and depots, not cryos and HLV. And the biggest mistake was not betting there would be an HLV, but closing the door on that propellant market, which could have funded commercial development of RLVs instead of government development of an HLV.
« Last Edit: 08/30/2013 11:44 pm by mmeijeri »
Pro-tip: you don't have to be a jerk if someone doesn't agree with your theories

Offline RocketmanUS

  • Senior Member
  • *****
  • Posts: 2226
  • USA
  • Liked: 71
  • Likes Given: 31
Re: Storable Propellant Earth Departure Stages
« Reply #76 on: 08/31/2013 01:21 am »
Hypergolics could have gotten us back to the moon and affordable on a yearly budget.

Absolutely. And if you had used a hybrid of prefueled cryogenics and hypergolic refueling at L1/L2, you could have created a large and fiercely competitive market for launch services right away, with only a small (~10%) mass penalty. And since this could have led to commercial development of SEP tugs soon, you might even end up with lower IMLEO sooner, not to mention lower launch costs per kg.

Quote
HLV's might with cryo propellants make it better and or less cost later on, however we didn't need to bet that we would get a HLV.

Cryos and depots, not cryos and HLV. And the biggest mistake was not betting there would be an HLV, but closing the door on that propellant market, which could have funded commercial development of RLVs instead of government development of an HLV.
No, just straight to LLO, no EML1/2. That could have been developed later for Mars , just like the HLV.

The HLV with cryo propellants are for larger mass and sized payloads once there could have been a need for them ( commercial development of the moon ). That would then require Lunar made propellants.

We are not building a multi lane super highway. We just need a way to repeatedly get to the moon at first. First see if there is a way to capitalize on the moon's resources. Then build a business plan for the long term. Now they can built that super highway to support such an enterprise.

Use of hypergolic EDS for Lunar and in space fueling would have increased the flight rate for Atlas and Delta. That would lower the per launch cost. Also if the program continued ( including going to Mars ) that would invite others to develop lower cost means to LEO. Mainly for the propellants, then later cargo and crew when their new launchers have improved and proven them selves.

Offline Warren Platts

Re: Storable Propellant Earth Departure Stages
« Reply #77 on: 08/31/2013 01:35 am »
Hypergolics are a step in the wrong direction. As has been pointed out repeatedly, it causes increases in IMLEO that are measured in the hundreds of millions of USD per mission. While boiloff is not a nonissue, it can be managed. Show me the money. Let's see some spreadsheets that demonstrate that hypergolics save money. I don't believe it. The only question is whether we should go with H2 or CH3 as a rocket fuel. And even that's not really debatable IMHO.
"When once you have tasted flight, you will forever walk the earth with your eyes turned skyward, for there you have been, and there you will always long to return."--Leonardo Da Vinci

Offline RocketmanUS

  • Senior Member
  • *****
  • Posts: 2226
  • USA
  • Liked: 71
  • Likes Given: 31
Re: Storable Propellant Earth Departure Stages
« Reply #78 on: 08/31/2013 01:50 am »
Hypergolics are a step in the wrong direction. As has been pointed out repeatedly, it causes increases in IMLEO that are measured in the hundreds of millions of USD per mission. While boiloff is not a nonissue, it can be managed. Show me the money. Let's see some spreadsheets that demonstrate that hypergolics save money. I don't believe it. The only question is whether we should go with H2 or CH3 as a rocket fuel. And even that's not really debatable IMHO.
It is about being able to get to the moon sooner. If we do get an HLV then we could use LH2/LOX for the EDS with the lander or crew vehicle on top for a single launch to LLO. However we haven't had an HLV since the Skylab launch. And no one knows yet what a new one will cost or when we will get one. So we might as well design the lander to work with either EELV's sized or HLV. We put our funding's in the wrong order.  Why waste money on a HLV we might never use or was not to the specs we would have needed?

We could have been back, we are still here. ::)

Offline Patchouli

  • Senior Member
  • *****
  • Posts: 4490
  • Liked: 253
  • Likes Given: 457
Re: Storable Propellant Earth Departure Stages
« Reply #79 on: 08/31/2013 03:03 am »
It only has an isp of about 312 s?

For the sake of argument.



By Storabe, you mean hyperogolics like MMH/N2O4?

Well, having large quantities of toxic storables at the pad for fueling could be less than desirable, especially for a crewed launch.  Hypergolics are expensive and difficult to handle, which is one of the several reasons I've read that Titan IV was retired and replaced with EELV's using cryogenics. 

I think methalox can be stored in space though, almost indefinately, so I'd probably be more keen to ponder that route for an EDS with long loiter capability.

Use Propane plus N2O as your storable propellant  as the combination gives an ISP of 312 which compares favorably with NTO/UDMH but is a lot easier to handle.
« Last Edit: 08/31/2013 03:05 am by Patchouli »

Offline ChrisWilson68

  • Senior Member
  • *****
  • Posts: 5266
  • Sunnyvale, CA
  • Liked: 4992
  • Likes Given: 6459
Re: Storable Propellant Earth Departure Stages
« Reply #80 on: 08/31/2013 04:19 am »
Hypergolics are a step in the wrong direction. As has been pointed out repeatedly, it causes increases in IMLEO that are measured in the hundreds of millions of USD per mission. While boiloff is not a nonissue, it can be managed. Show me the money. Let's see some spreadsheets that demonstrate that hypergolics save money. I don't believe it. The only question is whether we should go with H2 or CH3 as a rocket fuel. And even that's not really debatable IMHO.

QuantumG's point is that a very simple architecture using hypergolics and a minimal number of launches and rendezvous will get the same outcome as the dual-SLS architecture from NASA, but much more cheaply.

Whether there are other architectures that would be better in various ways is irrelevant to this point.  It's an existence proof that NASA's dual-SLS lunar architecture isn't the best way to achieve such a lunar mission.

I haven't seen any comment here that offers any argument against that.

So, yet another potential reason for developing SLS is blown out of the water.

Offline mmeijeri

  • Senior Member
  • *****
  • Posts: 7772
  • Martijn Meijering
  • NL
  • Liked: 397
  • Likes Given: 822
Re: Storable Propellant Earth Departure Stages
« Reply #81 on: 08/31/2013 10:19 am »
No, just straight to LLO, no EML1/2. That could have been developed later for Mars , just like the HLV.

L1/L2 is actually easier than LLO, for lunar applications as well as for Mars. It's the most natural staging point beyond LEO. You don't need a space station there or anything, it's just a staging point. And of course, building a station there would still be a simpler step than building a moon base.

Quote
The HLV with cryo propellants are for larger mass and sized payloads once there could have been a need for them ( commercial development of the moon ). That would then require Lunar made propellants.

Depots do not require lunar propellants, though they do work very well with them. And you don't need an HLV for lunar development either, you can land larger payloads than Constellation was aiming for even with just EELVs. And smallish HLVs like EELV Phase 1 and Falcon Heavy would be natural consequences of a large-scale propellant launch program, so that would be an additional argument against a government-built HLV.

Quote
We are not building a multi lane super highway. We just need a way to repeatedly get to the moon at first. First see if there is a way to capitalize on the moon's resources. Then build a business plan for the long term. Now they can built that super highway to support such an enterprise.

Absolutely. That's what I'm aiming for too.

Quote
Use of hypergolic EDS for Lunar and in space fueling would have increased the flight rate for Atlas and Delta. That would lower the per launch cost. Also if the program continued ( including going to Mars ) that would invite others to develop lower cost means to LEO. Mainly for the propellants, then later cargo and crew when their new launchers have improved and proven them selves.

Yep, and the same is true if you used the hypergolics only for the spacecraft and only used deep-space refueling at L1/L2. It would mitigate some of the downsides of hypergolics (low Isp, need for new or foreign engines). While these downsides are not decisive, it's still nice to mitigate them.
Pro-tip: you don't have to be a jerk if someone doesn't agree with your theories

Offline spectre9

  • Senior Member
  • *****
  • Posts: 2403
  • Australia
  • Liked: 42
  • Likes Given: 68
Re: Storable Propellant Earth Departure Stages
« Reply #82 on: 08/31/2013 10:53 am »
Missions with more launches and more IMLEO might be cheaper but are they better?

Such missions are discussed on these forums ad infinitum yet no missions spread over a large number of launches have ever been given serious consideration by NASA.

Offline QuantumG

  • Senior Member
  • *****
  • Posts: 9238
  • Australia
  • Liked: 4477
  • Likes Given: 1108
Re: Storable Propellant Earth Departure Stages
« Reply #83 on: 08/31/2013 11:15 am »
Missions with more launches and more IMLEO might be cheaper but are they better?

Yes.

Quote
Such missions are discussed on these forums ad infinitum yet no missions spread over a large number of launches have ever been given serious consideration by NASA.

Ya know, other than the whole ISS program.
Human spaceflight is basically just LARPing now.

Offline spectre9

  • Senior Member
  • *****
  • Posts: 2403
  • Australia
  • Liked: 42
  • Likes Given: 68
Re: Storable Propellant Earth Departure Stages
« Reply #84 on: 08/31/2013 11:43 am »
Got me QG  :)

It's a mission to nowhere but it's still a mission.

Multi launch missions are better only when multi launching is shown a feasible. Between all the commsats and spysats I just don't see where the launch rate ramp up comes from.

Perhaps SpaceX will crack that particular puzzle.

Offline mmeijeri

  • Senior Member
  • *****
  • Posts: 7772
  • Martijn Meijering
  • NL
  • Liked: 397
  • Likes Given: 822
Re: Storable Propellant Earth Departure Stages
« Reply #85 on: 08/31/2013 12:01 pm »
I think it's true that storable EDSs based in LEO are a step backward (at least for the US), though that alone is not a decisive point. But it isn't true for L1/L2-based deep space EDSs and for spacecraft, where hypergolics will remain the natural choice for the foreseeable future. In the bigger picture I think storable EDSs would be fine, though a hybrid approach would be better.
« Last Edit: 08/31/2013 12:35 pm by mmeijeri »
Pro-tip: you don't have to be a jerk if someone doesn't agree with your theories

Offline Robotbeat

  • Senior Member
  • *****
  • Posts: 39270
  • Minnesota
  • Liked: 25222
  • Likes Given: 12114
Re: Storable Propellant Earth Departure Stages
« Reply #86 on: 08/31/2013 11:51 pm »
Over a hundred launches have been done over the years in support of ISS. There's supposed to be about 5 US launches a year just for commercial cargo (SpaceX and Orbital have some catching up to do), plus another two for commercial crew. Another two Soyuzes a year plus 4-5 Progresses (doing hypergolic propellant transfer to the ISS!) and the odd HTV or ATV, and you have quite a flight rate (14 a year or so). All to the same place.

Missions which require several launches are not just feasible, they're mind-bogglingly common. We're supposed to have about 7 launches per expedition (though they're staggered), once everything's up and running smoothly (there's significant margin in there, though). And two expeditions per year.

It's kind of ridiculous that we're spending so much energy on fitting everything in just one or two launches for a mission to the Moon when you look at the dozens of launches we've been doing for ISS and continue doing...
Chris  Whoever loves correction loves knowledge, but he who hates reproof is stupid.

To the maximum extent practicable, the Federal Government shall plan missions to accommodate the space transportation services capabilities of United States commercial providers. US law http://goo.gl/YZYNt0

Offline A_M_Swallow

  • Elite Veteran
  • Senior Member
  • *****
  • Posts: 8906
  • South coast of England
  • Liked: 500
  • Likes Given: 223
Re: Storable Propellant Earth Departure Stages
« Reply #87 on: 09/01/2013 02:06 am »
{snip}
Missions which require several launches are not just feasible, they're mind-bogglingly common. We're supposed to have about 7 launches per expedition (though they're staggered), once everything's up and running smoothly (there's significant margin in there, though). And two expeditions per year.

It's kind of ridiculous that we're spending so much energy on fitting everything in just one or two launches for a mission to the Moon when you look at the dozens of launches we've been doing for ISS and continue doing...

A few problems.
The payload capabilities of Dragon and Cygnus are surprisingly small.
The modules will need docking or berthing interfaces.
ISS modules needed space walks to join wires and pipes together.  A robot may be able to do this now.

Two or three extra launches may permit construction of a mini spaceship yard with habitat, arms, tools and possibly a tug.  Designed properly the yard could assemble more than one vehicle.

Offline QuantumG

  • Senior Member
  • *****
  • Posts: 9238
  • Australia
  • Liked: 4477
  • Likes Given: 1108
Re: Storable Propellant Earth Departure Stages
« Reply #88 on: 09/01/2013 02:19 am »

Yes, you're going to need docking. This is no different to the proposed Constellation architecture, which required docking on the EDS for the crew capsule.

I don't see how Dragon and Cygnus are relevant. Progress does propellant transfer without any spacewalks.

« Last Edit: 09/01/2013 02:29 am by QuantumG »
Human spaceflight is basically just LARPing now.

Offline Robotbeat

  • Senior Member
  • *****
  • Posts: 39270
  • Minnesota
  • Liked: 25222
  • Likes Given: 12114
Re: Storable Propellant Earth Departure Stages
« Reply #89 on: 09/01/2013 02:27 am »
{snip}
Missions which require several launches are not just feasible, they're mind-bogglingly common. We're supposed to have about 7 launches per expedition (though they're staggered), once everything's up and running smoothly (there's significant margin in there, though). And two expeditions per year.

It's kind of ridiculous that we're spending so much energy on fitting everything in just one or two launches for a mission to the Moon when you look at the dozens of launches we've been doing for ISS and continue doing...

A few problems.
The payload capabilities of Dragon and Cygnus are surprisingly small.
The modules will need docking or berthing interfaces.
ISS modules needed space walks to join wires and pipes together.  A robot may be able to do this now.

Two or three extra launches may permit construction of a mini spaceship yard with habitat, arms, tools and possibly a tug.  Designed properly the yard could assemble more than one vehicle.
I was talking merely about the objection that more than 1 or 2 launches is too much. We do a dozen launches a year to ISS already. We can do maybe 5 or 6 larger EELV Heavy-sized launches, spread out over a year or 6 months or whatever to do a lunar mission if we have to, though there are ways to do it with just 2 or 3 EELV-Heavy or Falcon Heavy launches.
Chris  Whoever loves correction loves knowledge, but he who hates reproof is stupid.

To the maximum extent practicable, the Federal Government shall plan missions to accommodate the space transportation services capabilities of United States commercial providers. US law http://goo.gl/YZYNt0

Offline spectre9

  • Senior Member
  • *****
  • Posts: 2403
  • Australia
  • Liked: 42
  • Likes Given: 68
Re: Storable Propellant Earth Departure Stages
« Reply #90 on: 09/01/2013 03:27 am »
Two Delta IV heavies haven't launched in a single calender year.

Falcon Heavy is still on the drawing board although it does share components with the currently produced F9 v1.1.

This is my main concern with architectures that require many medium launch vehicles.

ISS has used the following.

Shuttle
Proton
Soyuz-FG
Soyuz-U
Ariane 5
H-IIB
Falcon 9
Antares soon to join the list.

If you can figure out a way to use that fleet for BEO missions in a way that is acceptable to NASA it would be a spaceflight miracle.

Offline ChrisWilson68

  • Senior Member
  • *****
  • Posts: 5266
  • Sunnyvale, CA
  • Liked: 4992
  • Likes Given: 6459
Re: Storable Propellant Earth Departure Stages
« Reply #91 on: 09/01/2013 03:37 am »
Falcon Heavy is still on the drawing board although it does share components with the currently produced F9 v1.1.

I think that's a bit of an underestimation of the status of Falcon Heavy.

F9v1.1 doesn't just share some components with Falcon Heavy -- nearly every component of Falcon Heavy is on F9v1.1.  Early flights of Falcon Heavy won't even use propellant cross-feed, which is really the biggest hardware change.  Aside from cross-feed, it's really just the physical attachment of the existing cores to each other, and that was designed into the F9v1.1 cores.  It's a different configuration, and it needs to be tested to in the Heavy configuration before we'll verify it works as expected, but that's very different from "on the drawing boards".

Offline spectre9

  • Senior Member
  • *****
  • Posts: 2403
  • Australia
  • Liked: 42
  • Likes Given: 68
Re: Storable Propellant Earth Departure Stages
« Reply #92 on: 09/01/2013 03:40 am »
The side boosters have extended tanks.

The middle core needs to hande the load of a heavier payload.

Saying it's just 3 Falcons strapped together is an underestimation.

Offline Robotbeat

  • Senior Member
  • *****
  • Posts: 39270
  • Minnesota
  • Liked: 25222
  • Likes Given: 12114
Re: Storable Propellant Earth Departure Stages
« Reply #93 on: 09/01/2013 03:40 am »
So what? Delta IV was originally designed to launch a lot more frequently. Falcon Heavy is far more real than SLS. Lots of Atlas Vs have launched every year (5 this year alone, 6 last year), and if we're talking propellant launches (and don't do SpaceX), the cheaper Atlas V would be used for the vast majority of launches. It doesn't have to be the same calendar year, either. Atlas V can do 18mT to LEO.

ISS proves that lots of vehicles can be launched to support HSF and that mixing a bunch of vehicles gives you a HSF program robust and continuous in the face of launch vehicle failures or cancellations.

And to a large extent, the low launch rate of Delta IV Heavy is due to the low demand.
« Last Edit: 09/01/2013 03:47 am by Robotbeat »
Chris  Whoever loves correction loves knowledge, but he who hates reproof is stupid.

To the maximum extent practicable, the Federal Government shall plan missions to accommodate the space transportation services capabilities of United States commercial providers. US law http://goo.gl/YZYNt0

Offline Patchouli

  • Senior Member
  • *****
  • Posts: 4490
  • Liked: 253
  • Likes Given: 457
Re: Storable Propellant Earth Departure Stages
« Reply #94 on: 09/01/2013 03:41 am »
Two Delta IV heavies haven't launched in a single calender year.

Falcon Heavy is still on the drawing board although it does share components with the currently produced F9 v1.1.

This is my main concern with architectures that require many medium launch vehicles.

ISS has used the following.

Shuttle
Proton
Soyuz-FG
Soyuz-U
Ariane 5
H-IIB
Falcon 9
Antares soon to join the list.

If you can figure out a way to use that fleet for BEO missions in a way that is acceptable to NASA it would be a spaceflight miracle.

I favor using an upgraded Delta IV-H and Falcon Heavy for cargo and Atlas V and Falcon 9 for crew.
One issue would be getting a large enough fairing for large BEO payloads but an 8 meter hammerhead fairing on the Delta IV Heavy might do the job.
« Last Edit: 09/01/2013 03:41 am by Patchouli »

Offline QuantumG

  • Senior Member
  • *****
  • Posts: 9238
  • Australia
  • Liked: 4477
  • Likes Given: 1108
Re: Storable Propellant Earth Departure Stages
« Reply #95 on: 09/01/2013 03:42 am »
The SLS plan called for a launch every other year, two launches per lunar campaign. That's a mission every 4 years. This is not a challenging timeline for a multi-launch campaign.
 
Human spaceflight is basically just LARPing now.

Offline Patchouli

  • Senior Member
  • *****
  • Posts: 4490
  • Liked: 253
  • Likes Given: 457
Re: Storable Propellant Earth Departure Stages
« Reply #96 on: 09/01/2013 04:02 am »
The SLS plan called for a launch every other year, two launches per lunar campaign. That's a mission every 4 years. This is not a challenging timeline for a multi-launch campaign.
 

The existing fleet of MLVs could easily support a flight rate or three or four lunar missions per year.

Just the Delta V production line if ramped up there could be 40 CBCs produced per year.

Assume the DOD is going to take at least five to ten of these that leaves 30 CBCs which is enough for ten Delta Heavy flights a year to be dedicated for BEO exploration.

That's just one MLV though the Delta and Atlas share a lot of production facilities now.
« Last Edit: 09/01/2013 04:09 am by Patchouli »

Offline MP99

Re: Storable Propellant Earth Departure Stages
« Reply #97 on: 09/01/2013 11:01 am »
Over a hundred launches have been done over the years in support of ISS. There's supposed to be about 5 US launches a year just for commercial cargo (SpaceX and Orbital have some catching up to do), plus another two for commercial crew. Another two Soyuzes a year plus 4-5 Progresses (doing hypergolic propellant transfer to the ISS!) and the odd HTV or ATV, and you have quite a flight rate (14 a year or so). All to the same place.

Missions which require several launches are not just feasible, they're mind-bogglingly common. We're supposed to have about 7 launches per expedition (though they're staggered), once everything's up and running smoothly (there's significant margin in there, though). And two expeditions per year.

It's kind of ridiculous that we're spending so much energy on fitting everything in just one or two launches for a mission to the Moon when you look at the dozens of launches we've been doing for ISS and continue doing...

I was talking merely about the objection that more than 1 or 2 launches is too much. We do a dozen launches a year to ISS already.

So, just to put a number on it, how much of ISS's annual budget is spent on planning and executing those launches & berthings/dockings?

cheers, Martin

Offline mmeijeri

  • Senior Member
  • *****
  • Posts: 7772
  • Martijn Meijering
  • NL
  • Liked: 397
  • Likes Given: 822
Re: Storable Propellant Earth Departure Stages
« Reply #98 on: 09/01/2013 11:18 am »
And to a large extent, the low launch rate of Delta IV Heavy is due to the low demand.

Exactly, Delta and Atlas were each designed to be launched 20-40 times a year, and the fact that there isn't enough demand for that is a large part of the reason they are currently so expensive. Due to consolidation the production capacity is now lower, but much higher than is being used.

And then there's also Falcon 9 and soon Antares. In addition, given sufficient demand you could expect to see EELV Phase 1 developed, leading to a large increase in yearly launch capacity.
Pro-tip: you don't have to be a jerk if someone doesn't agree with your theories

Offline mmeijeri

  • Senior Member
  • *****
  • Posts: 7772
  • Martijn Meijering
  • NL
  • Liked: 397
  • Likes Given: 822
Re: Storable Propellant Earth Departure Stages
« Reply #99 on: 09/01/2013 11:21 am »
One issue would be getting a large enough fairing for large BEO payloads but an 8 meter hammerhead fairing on the Delta IV Heavy might do the job.

You don't need 8m fairings for exploration, even though EELVs could perhaps be modified to support them.
Pro-tip: you don't have to be a jerk if someone doesn't agree with your theories

Offline ChrisWilson68

  • Senior Member
  • *****
  • Posts: 5266
  • Sunnyvale, CA
  • Liked: 4992
  • Likes Given: 6459
Re: Storable Propellant Earth Departure Stages
« Reply #100 on: 09/01/2013 11:38 am »
One issue would be getting a large enough fairing for large BEO payloads but an 8 meter hammerhead fairing on the Delta IV Heavy might do the job.

You don't need 8m fairings for exploration, even though EELVs could perhaps be modified to support them.

The question is what hardware needs 8m fairings and what is the cost of redesigning that hardware to fit in smaller fairings (perhaps with more on-orbit assembly) versus the cost of switching from low-cost launchers to SLS.  Certainly not any hardware for the lunar missionthis thread is about.

Offline mmeijeri

  • Senior Member
  • *****
  • Posts: 7772
  • Martijn Meijering
  • NL
  • Liked: 397
  • Likes Given: 822
Re: Storable Propellant Earth Departure Stages
« Reply #101 on: 09/01/2013 01:17 pm »
The question is what hardware needs 8m fairings and what is the cost of redesigning that hardware to fit in smaller fairings (perhaps with more on-orbit assembly) versus the cost of switching from low-cost launchers to SLS.  Certainly not any hardware for the lunar missionthis thread is about.

There are several ways to land Constellation-sized payloads or larger without a lander that needs 8m fairings and these ways can even be combined: fully hypergolic lander, horizontal lander, crasher stage.

And even for Mars there are alternatives: fully propulsive landings, transpiration cooling, strongly lifting entry.
Pro-tip: you don't have to be a jerk if someone doesn't agree with your theories

Offline A_M_Swallow

  • Elite Veteran
  • Senior Member
  • *****
  • Posts: 8906
  • South coast of England
  • Liked: 500
  • Likes Given: 223
Re: Storable Propellant Earth Departure Stages
« Reply #102 on: 09/01/2013 08:10 pm »

Yes, you're going to need docking. This is no different to the proposed Constellation architecture, which required docking on the EDS for the crew capsule.

I don't see how Dragon and Cygnus are relevant. Progress does propellant transfer without any spacewalks.


The USA is short of cargo spacecraft with the sophisticated guidance control systems allowing them to dock.  The Constellation crew capsule like the Dragon and Cygnus was designed to have one.  A 20 tonne module full of food would probably not have the RCS needed.  The tug's job would be to accurately bring the module the last 10 miles.

Offline DGH

  • Full Member
  • *
  • Posts: 168
  • Liked: 7
  • Likes Given: 4
Re: Storable Propellant Earth Departure Stages
« Reply #103 on: 09/02/2013 09:18 pm »
According to this article, which has it's own thread, NASA wants a 43 ton lunar lander.. for some reason.. and they say the only way to get a payload that big into lunar orbit is with the SLS.

I don't get it. What's wrong with just using a storable propellant Earth departure stage? Let's be conservative and say it only has an isp of about 312s, and a propellant mass fraction of 90%, how big would it be?

According to my math, please check me, I figure it would be about 10,458 kg dry and 104,579 kg when full. This would provide the 3107 m/s of delta-v to get through TLI, with the lunar insertion to be done by the lander (as in the NASA architecture).

I know 105 tons sounds like a lot, but it's only two Falcon Heavy launches, and because we're using storable propellant there's no time pressure. If you really wanted to you could do it with Falcon 9 v1.1.


   



Why not use the fuel to load up the Lander and Orion once in LEO.
This would massively improve SLS numbers to TLI.

Offline gbaikie

  • Full Member
  • ****
  • Posts: 1592
  • Liked: 49
  • Likes Given: 5
Re: Storable Propellant Earth Departure Stages
« Reply #104 on: 09/04/2013 07:03 am »
According to this article, which has it's own thread, NASA wants a 43 ton lunar lander.. for some reason.. and they say the only way to get a payload that big into lunar orbit is with the SLS.

I don't get it. What's wrong with just using a storable propellant Earth departure stage? Let's be conservative and say it only has an isp of about 312s, and a propellant mass fraction of 90%, how big would it be?

According to my math, please check me, I figure it would be about 10,458 kg dry and 104,579 kg when full. This would provide the 3107 m/s of delta-v to get through TLI, with the lunar insertion to be done by the lander (as in the NASA architecture).

I know 105 tons sounds like a lot, but it's only two Falcon Heavy launches, and because we're using storable propellant there's no time pressure. If you really wanted to you could do it with Falcon 9 v1.1.


   



Why not use the fuel to load up the Lander and Orion once in LEO.
This would massively improve SLS numbers to TLI.


One also wouldn't need to design the Orion to withstand to gee loads of fully fueled Orion during a SLS launch to orbit.

What will be the maximum gee load caused by SLS launch, anyone know what this is supposed to be, btw?

Offline newpylong

  • Full Member
  • ****
  • Posts: 1499
  • Liked: 200
  • Likes Given: 343
Re: Storable Propellant Earth Departure Stages
« Reply #105 on: 09/04/2013 04:22 pm »


One also wouldn't need to design the Orion to withstand to gee loads of fully fueled Orion during a SLS launch to orbit.

What will be the maximum gee load caused by SLS launch, anyone know what this is supposed to be, btw?

Orion itself does not have to do, the Service Module shrouds handle half the weight of Orion and the LAS.

Offline Robotbeat

  • Senior Member
  • *****
  • Posts: 39270
  • Minnesota
  • Liked: 25222
  • Likes Given: 12114
Re: Storable Propellant Earth Departure Stages
« Reply #106 on: 09/04/2013 04:31 pm »
One issue would be getting a large enough fairing for large BEO payloads but an 8 meter hammerhead fairing on the Delta IV Heavy might do the job.

You don't need 8m fairings for exploration, even though EELVs could perhaps be modified to support them.


You might be able to swing Apollo again without 8m but forget BLEO on EELV fairings.
You're wrong. Atlas V can support 7.2m diameter fairings by customer request.
see: http://www.ulalaunch.com/site/docs/product_cards/guides/AtlasVUsersGuide2010.pdf

And Apollo used a much smaller fairing than that (used a cone-shaped fairing that was--at its very widest--6.6m and at its narrowest 3.9m in diameter, basically comparable to the standard 5.2m diameter fairings often used for EELV launches). There aren't even standard facilities that can test things bigger than ~7m diameter, plus transport is almost impossible except via boat. BLEO exploration can be done perfectly well with 7.2m diameter fairings.

The next manned Moon lander is almost certainly not going to be hydrolox-based, at least if NASA builds it. Everyone I talked to (at GRC) said it would be hypergolic (if it happens), and that fits with what I've seen in the most recent documents here (though methane would work, too). Anything other than hydrolox would be significantly smaller in diameter than the old Altair was (and that was a significant transport problem that--to my knowledge--was never solved). The Deep Space Hab talked about in documents here are ISS-sized components for which 7.2m would be enormously overkill.

Skylab itself was 6.6m in diameter, and we barely could test and transport it.
« Last Edit: 09/04/2013 04:37 pm by Robotbeat »
Chris  Whoever loves correction loves knowledge, but he who hates reproof is stupid.

To the maximum extent practicable, the Federal Government shall plan missions to accommodate the space transportation services capabilities of United States commercial providers. US law http://goo.gl/YZYNt0

Offline mmeijeri

  • Senior Member
  • *****
  • Posts: 7772
  • Martijn Meijering
  • NL
  • Liked: 397
  • Likes Given: 822
Re: Storable Propellant Earth Departure Stages
« Reply #107 on: 09/04/2013 04:39 pm »
Going beyond LEO doesn't require anything larger than existing EELV fairings, which are huge already. There are ways to design the necessary transfer stages and spacecraft that would require larger fairings, but there are also ways to design them so that they don't.
« Last Edit: 09/04/2013 04:40 pm by mmeijeri »
Pro-tip: you don't have to be a jerk if someone doesn't agree with your theories

Offline clongton

  • Expert
  • Senior Member
  • *****
  • Posts: 12048
  • Connecticut
    • Direct Launcher
  • Liked: 7331
  • Likes Given: 3744
Re: Storable Propellant Earth Departure Stages
« Reply #108 on: 09/06/2013 01:10 am »
Going beyond LEO doesn't require anything larger than existing EELV fairings, which are huge already. There are ways to design the necessary transfer stages and spacecraft that would require larger fairings, but there are also ways to design them so that they don't.

Agreed, but there is also a point at which the expense of doing that outweighs the expense of a larger fairing with simplified payload construction. It's a balancing act.
Chuck - DIRECT co-founder
I started my career on the Saturn-V F-1A engine

Offline Robotbeat

  • Senior Member
  • *****
  • Posts: 39270
  • Minnesota
  • Liked: 25222
  • Likes Given: 12114
Re: Storable Propellant Earth Departure Stages
« Reply #109 on: 09/06/2013 01:18 am »
Going beyond LEO doesn't require anything larger than existing EELV fairings, which are huge already. There are ways to design the necessary transfer stages and spacecraft that would require larger fairings, but there are also ways to design them so that they don't.

Agreed, but there is also a point at which the expense of doing that outweighs the expense of a larger fairing with simplified payload construction. It's a balancing act.
that may be true but simply transporting the payload on the ground becomes a big pain even if you're inside the EELV fairing size limit. And beyond 7m, the only option is boat.
Chris  Whoever loves correction loves knowledge, but he who hates reproof is stupid.

To the maximum extent practicable, the Federal Government shall plan missions to accommodate the space transportation services capabilities of United States commercial providers. US law http://goo.gl/YZYNt0

Offline clongton

  • Expert
  • Senior Member
  • *****
  • Posts: 12048
  • Connecticut
    • Direct Launcher
  • Liked: 7331
  • Likes Given: 3744
Re: Storable Propellant Earth Departure Stages
« Reply #110 on: 09/06/2013 01:28 am »
Going beyond LEO doesn't require anything larger than existing EELV fairings, which are huge already. There are ways to design the necessary transfer stages and spacecraft that would require larger fairings, but there are also ways to design them so that they don't.

Agreed, but there is also a point at which the expense of doing that outweighs the expense of a larger fairing with simplified payload construction. It's a balancing act.
that may be true but simply transporting the payload on the ground becomes a big pain even if you're inside the EELV fairing size limit. And beyond 7m, the only option is boat.

There is also the option of final assembly at the Cape, mitigating the transportation issues.
Chuck - DIRECT co-founder
I started my career on the Saturn-V F-1A engine

Offline RocketmanUS

  • Senior Member
  • *****
  • Posts: 2226
  • USA
  • Liked: 71
  • Likes Given: 31
Re: Storable Propellant Earth Departure Stages
« Reply #111 on: 09/06/2013 01:41 am »
Going beyond LEO doesn't require anything larger than existing EELV fairings, which are huge already. There are ways to design the necessary transfer stages and spacecraft that would require larger fairings, but there are also ways to design them so that they don't.

Agreed, but there is also a point at which the expense of doing that outweighs the expense of a larger fairing with simplified payload construction. It's a balancing act.
that may be true but simply transporting the payload on the ground becomes a big pain even if you're inside the EELV fairing size limit. And beyond 7m, the only option is boat.

There is also the option of final assembly at the Cape, mitigating the transportation issues.
We start out with what we have.
Then when the high launch rate is justified for wide body HLV then the needed infrastructure can be put in.

There is also the possible new airships that might see serves. If they do come available for commercial use they would be able to transport large volume and mass from the factory to the launch vehicle integration site.

Offline mmeijeri

  • Senior Member
  • *****
  • Posts: 7772
  • Martijn Meijering
  • NL
  • Liked: 397
  • Likes Given: 822
Re: Storable Propellant Earth Departure Stages
« Reply #112 on: 09/06/2013 12:55 pm »
Agreed, but there is also a point at which the expense of doing that outweighs the expense of a larger fairing with simplified payload construction. It's a balancing act.

I think that's true of 3m fairings, but not of 5m ones.
Pro-tip: you don't have to be a jerk if someone doesn't agree with your theories

Offline muomega0

  • Full Member
  • ****
  • Posts: 862
  • Liked: 70
  • Likes Given: 1
Re: Storable Propellant Earth Departure Stages
« Reply #113 on: 09/06/2013 01:22 pm »
One issue would be getting a large enough fairing for large BEO payloads but an 8 meter hammerhead fairing on the Delta IV Heavy might do the job.

You don't need 8m fairings for exploration, even though EELVs could perhaps be modified to support them.


You might be able to swing Apollo again without 8m but forget BLEO on EELV fairings.
You're wrong. Atlas V can support 7.2m diameter fairings by customer request.
see: http://www.ulalaunch.com/site/docs/product_cards/guides/AtlasVUsersGuide2010.pdf

And Apollo used a much smaller fairing than that (used a cone-shaped fairing that was--at its very widest--6.6m and at its narrowest 3.9m in diameter, basically comparable to the standard 5.2m diameter fairings often used for EELV launches). There aren't even standard facilities that can test things bigger than ~7m diameter, plus transport is almost impossible except via boat. BLEO exploration can be done perfectly well with 7.2m diameter fairings.

The next manned Moon lander is almost certainly not going to be hydrolox-based, at least if NASA builds it. Everyone I talked to (at GRC) said it would be hypergolic (if it happens), and that fits with what I've seen in the most recent documents here (though methane would work, too).
Using lower ISP landers and Earth Departure stages will reduce the number of missions by 30% or more in the same time frame due to the increased IMLEO Costs.  It the explore sooner  vs sustainable more missions with infrastructure first debate.  LH2 is storable...just needs the $$ to demonstrate and deploy.  It would take a significant increase in LOC to justify lower ISP, and the numbers do not reflect a significant difference.

One lander option is the uncrasher stage described by jon goff that features partial reusability that would reduce IMLEO costs further by reducing "crasher" lander mass.

Offline Robotbeat

  • Senior Member
  • *****
  • Posts: 39270
  • Minnesota
  • Liked: 25222
  • Likes Given: 12114
Re: Storable Propellant Earth Departure Stages
« Reply #114 on: 09/06/2013 03:17 pm »
One issue would be getting a large enough fairing for large BEO payloads but an 8 meter hammerhead fairing on the Delta IV Heavy might do the job.

You don't need 8m fairings for exploration, even though EELVs could perhaps be modified to support them.


You might be able to swing Apollo again without 8m but forget BLEO on EELV fairings.
You're wrong. Atlas V can support 7.2m diameter fairings by customer request.
see: http://www.ulalaunch.com/site/docs/product_cards/guides/AtlasVUsersGuide2010.pdf

And Apollo used a much smaller fairing than that (used a cone-shaped fairing that was--at its very widest--6.6m and at its narrowest 3.9m in diameter, basically comparable to the standard 5.2m diameter fairings often used for EELV launches). There aren't even standard facilities that can test things bigger than ~7m diameter, plus transport is almost impossible except via boat. BLEO exploration can be done perfectly well with 7.2m diameter fairings.

The next manned Moon lander is almost certainly not going to be hydrolox-based, at least if NASA builds it. Everyone I talked to (at GRC) said it would be hypergolic (if it happens), and that fits with what I've seen in the most recent documents here (though methane would work, too).
Using lower ISP landers and Earth Departure stages will reduce the number of missions by 30% or more in the same time frame due to the increased IMLEO Costs.  It the explore sooner  vs sustainable more missions with infrastructure first debate.  LH2 is storable...just needs the $$ to demonstrate and deploy.  It would take a significant increase in LOC to justify lower ISP, and the numbers do not reflect a significant difference.

One lander option is the uncrasher stage described by jon goff that features partial reusability that would reduce IMLEO costs further by reducing "crasher" lander mass.

...you can reduce the IMLEO required for hypergolics to below that of hydrolox if you just use an electric stage (hypergolics allow you to take your time), so your very first point is untrue. Electric stage (launched on the same flight as the lander) can allow a comparable-to-Apollo-capability lander to be put at EML2 with a single launch of an EELV. If you could use Falcon Heavy, the capability would be at least twice that of Apollo on a single launch. Hydrolox's 40-50% improvement in Isp doesn't overcome this advantage (electric propulsion has an Isp 1000% greater).

The state of electric propulsion being what it is, there are high-TRL components right now that you could put on a comm-sat bus to function as a 100kW electric stage (though 40-50kW may be sufficient). Modern arrays are available over 25kW right now (>12.5kW per wing), so putting an extra set of arrays on a modified comm-sat bus would be sufficient to allow you to put a hypergolic lander at EML2.

I'm not against hydrolox, but right now, the boil-off of hydrogen would be too much for the trip times needed for an electric stage. Electric propulsion can operate with >70% total efficiency at over 4km/s delta-v with a total system specific power better than 50W/kg.



You'd still use hydrolox for an EDS to put the crew at EML2, but the lander can be put there (and kept there for station-keeping) using electric propulsion.
« Last Edit: 09/06/2013 03:27 pm by Robotbeat »
Chris  Whoever loves correction loves knowledge, but he who hates reproof is stupid.

To the maximum extent practicable, the Federal Government shall plan missions to accommodate the space transportation services capabilities of United States commercial providers. US law http://goo.gl/YZYNt0

Offline mmeijeri

  • Senior Member
  • *****
  • Posts: 7772
  • Martijn Meijering
  • NL
  • Liked: 397
  • Likes Given: 822
Re: Storable Propellant Earth Departure Stages
« Reply #115 on: 09/06/2013 03:45 pm »
Even without the SEP tug, hypergolics with quasi-ballistic trajectories can have roughly equal IMLEO compared to LOX/LH2 without them. And while SEP tugs are basically off-the-shelf technology, you would still have to develop (integrate) them. I'd rather leave that to the market. If you buy transportation of crew, empty spacecraft and storable propellant to L1/L2, you leave the mode of transport up to the market.

My guess is that cryogenic stages to get from LEO to L1/L2 are what the market would choose initially for heavy spacecraft, later augmented with SEP, but that's just a guess. Propellant and smaller spacecraft would likely start with a single launch to L1/L2. In any event, I think it is better to use deep-space refueling for the spacecraft and both launch and move them through TLI dry.
« Last Edit: 09/06/2013 03:49 pm by mmeijeri »
Pro-tip: you don't have to be a jerk if someone doesn't agree with your theories

Offline muomega0

  • Full Member
  • ****
  • Posts: 862
  • Liked: 70
  • Likes Given: 1
Re: Storable Propellant Earth Departure Stages
« Reply #116 on: 09/06/2013 05:06 pm »
One issue would be getting a large enough fairing for large BEO payloads but an 8 meter hammerhead fairing on the Delta IV Heavy might do the job.

You don't need 8m fairings for exploration, even though EELVs could perhaps be modified to support them.


You might be able to swing Apollo again without 8m but forget BLEO on EELV fairings.
You're wrong. Atlas V can support 7.2m diameter fairings by customer request.
see: http://www.ulalaunch.com/site/docs/product_cards/guides/AtlasVUsersGuide2010.pdf

And Apollo used a much smaller fairing than that (used a cone-shaped fairing that was--at its very widest--6.6m and at its narrowest 3.9m in diameter, basically comparable to the standard 5.2m diameter fairings often used for EELV launches). There aren't even standard facilities that can test things bigger than ~7m diameter, plus transport is almost impossible except via boat. BLEO exploration can be done perfectly well with 7.2m diameter fairings.

The next manned Moon lander is almost certainly not going to be hydrolox-based, at least if NASA builds it. Everyone I talked to (at GRC) said it would be hypergolic (if it happens), and that fits with what I've seen in the most recent documents here (though methane would work, too).
Using lower ISP landers and Earth Departure stages will reduce the number of missions by 30% or more in the same time frame due to the increased IMLEO Costs.  It the explore sooner  vs sustainable more missions with infrastructure first debate.  LH2 is storable...just needs the $$ to demonstrate and deploy.  It would take a significant increase in LOC to justify lower ISP, and the numbers do not reflect a significant difference.

One lander option is the uncrasher stage described by jon goff that features partial reusability that would reduce IMLEO costs further by reducing "crasher" lander mass.

...you can reduce the IMLEO required for hypergolics to below that of hydrolox if you just use an electric stage (hypergolics allow you to take your time), so your very first point is untrue.
The 30% value assumes lower ISP for both the transfer stage *and* lander, so a methane transfer stage and lander vs LH2 transfer stage and lander, not mixed like Jon assumed (LH2 transfer+uncrasher and methane crasher).

so in your case you are substituting EP for the uncrasher, a different case :)

One also has to keep more production lines open....fixed costs eat NASA's lunch.

Electric stage (launched on the same flight as the lander) can allow a comparable-to-Apollo-capability lander to be put at EML2 with a single launch of an EELV. If you could use Falcon Heavy, the capability would be at least twice that of Apollo on a single launch. Hydrolox's 40-50% improvement in Isp doesn't overcome this advantage (electric propulsion has an Isp 1000% greater).

The state of electric propulsion being what it is, there are high-TRL components right now that you could put on a comm-sat bus to function as a 100kW electric stage (though 40-50kW may be sufficient). Modern arrays are available over 25kW right now (>12.5kW per wing), so putting an extra set of arrays on a modified comm-sat bus would be sufficient to allow you to put a hypergolic lander at EML2.
so for the EP "uncrasher"..
The solution depends on a number of factors:  number and type of missions, orbit and release alititude, etc.   40 to 50 kW power bus with EP-tanks is quite a high IMLEO.  the orbit at lunar and altitude will affect the battery mass.  Its not the slam dunk being projected.   It would serve to test out EP...so why not test out LH2 as well (depot first)?

I'm not against hydrolox, but right now, the boil-off of hydrogen would be too much for the trip times needed for an electric stage. Electric propulsion can operate with >70% total efficiency at over 4km/s delta-v with a total system specific power better than 50W/kg.

You'd still use hydrolox for an EDS to put the crew at EML2, but the lander can be put there (and kept there for station-keeping) using electric propulsion.
Explore sooner...i get it.  All the (objective) studies point toward high ISP solutions.   One needs a ZBO LH2 depot to substantially cut IMLEO costs.   When considering EP from LEO, or in a low lunar orbit...do not forget about the mass and reliability impacts of eclipse or the transfer tug.
If you have power, why use EP or any fuel for station keeping? 

Offline Robotbeat

  • Senior Member
  • *****
  • Posts: 39270
  • Minnesota
  • Liked: 25222
  • Likes Given: 12114
Re: Storable Propellant Earth Departure Stages
« Reply #117 on: 09/06/2013 05:31 pm »
...40 to 50 kW power bus with EP-tanks is quite a high IMLEO....
Where do you get that idea? No, it's not. Do the math.

It's about 1-2 tons (maybe three, if you include plenty of margin and structure and heavy docking adapter) for the solar arrays, PPU, thruster and tanks.
Solar arrays can do about 150kW/kg with modern IMM and UltraFlex arrays (like are being used on Advanced Cygnus due to launch next year), and both the PPU (plus radiators) and thruster are better than that.
Depending on the size of the tanks. Xenon tanks can have a full:empty mass ratio of somewhere between 10:1 and 20:1 using the standard supercritical tanks (would allow about 5-10mT of Xenon propellant with a mass of just 500kg for the tanks... valves and such can be the sort of lightweight stuff already used on commercial comm sats). Store it mildly cryogenically and you can do better, but that's beyond current practice.

EDIT:Also, eclipse only matters on the very beginning of the flight to EML2 (because eclipse times are much less of the total length of the orbit when you're higher up). And there's a simple solution: don't thrust during those times.
« Last Edit: 09/06/2013 05:40 pm by Robotbeat »
Chris  Whoever loves correction loves knowledge, but he who hates reproof is stupid.

To the maximum extent practicable, the Federal Government shall plan missions to accommodate the space transportation services capabilities of United States commercial providers. US law http://goo.gl/YZYNt0

Offline hydra9

  • Full Member
  • ***
  • Posts: 349
  • Liked: 16
  • Likes Given: 6
Re: Storable Propellant Earth Departure Stages
« Reply #118 on: 09/06/2013 05:38 pm »
Using the DUUS (Dual Use Upper Stage) NASA would be able to place a fuel depot weighing at least 30 tonnes at a Lagrange point or into Low Lunar Orbit.

Any future lunar lander should be a-- single stage reusable vehicle-- that could eventually utilize lunar water resources. Lockheed-Martin has proposed such a LOX/LH2 single stage vehicle.

So all NASA would have to do is to launch a fuel depot into LLO or to L1 or L2. Then, perhaps, six months later (if the ground  infrastructure for an immediate two launch is not available), they could launch the MPCV + an empty or  hydrogen only fueled lunar lander to the fuel depot for the lunar mission.

But hydrogen boil-off shouldn't be a problem since NASA has already invented cryocooler technology that can continuously re-liquify hydrogen ullage gasses for years using a moderate amount of solar electricity.

The fuel depot itself could be derived from the tanks of the  reusable lunar lander in fashion similar to the way the ULA wants to use ACES fuel tanks for fuel depots.   

Of course, once fuel is available on the lunar surface then only one SLS launch would be required for a manned lunar mission since the reusable lunar lander would be fueled for its round trip with fuel manufactured at a lunar outpost.


Marcel F. Williams

Offline muomega0

  • Full Member
  • ****
  • Posts: 862
  • Liked: 70
  • Likes Given: 1
Re: Storable Propellant Earth Departure Stages
« Reply #119 on: 09/06/2013 06:43 pm »
...40 to 50 kW power bus with EP-tanks is quite a high IMLEO....
Where do you get that idea? No, it's not. Do the math.

It's about 1-2 tons (maybe three, if you include plenty of margin and structure and heavy docking adapter) for the solar arrays, PPU, thruster and tanks.
Solar arrays can do about 150kW/kg with modern IMM and UltraFlex arrays (like are being used on Advanced Cygnus due to launch next year), and both the PPU (plus radiators) and thruster are better than that.
Depending on the size of the tanks. Xenon tanks can have a full:empty mass ratio of somewhere between 10:1 and 20:1 using the standard supercritical tanks (would allow about 5-10mT of Xenon propellant with a mass of just 500kg for the tanks... valves and such can be the sort of lightweight stuff already used on commercial comm sats). Store it mildly cryogenically and you can do better, but that's beyond current practice.

EDIT:Also, eclipse only matters on the very beginning of the flight to EML2 (because eclipse times are much less of the total length of the orbit when you're higher up). And there's a simple solution: don't thrust during those times.
My dry mass is about 10 mT for this 50 kW EP tug.   It includes energy storage, heat rejection, your 150 W/kg SA, power conversion/bus at 300V--not the lower ISS voltage, tanks, support structure, gimbals, comm, thrusters, etc and power is dependent on a few more factors.  Did you compare your estimates to HEFT, where the SA is only 14% of the mass?

If you do not thrust, then thruster reliability becomes an issue during eclipse, which is why EP is great from L2 to deeper space, and perhaps as the uncrasher, but EP from LEO....much more work and $$ than a ZBO LH2 depot.   
« Last Edit: 09/06/2013 07:09 pm by muomega0 »

Offline A_M_Swallow

  • Elite Veteran
  • Senior Member
  • *****
  • Posts: 8906
  • South coast of England
  • Liked: 500
  • Likes Given: 223
Re: Storable Propellant Earth Departure Stages
« Reply #120 on: 09/06/2013 09:27 pm »
Part 2 of the ARRM asteroid mission hopes to use a SEP.  Using the same specification parts in the SEP tug will keep costs down.

Offline muomega0

  • Full Member
  • ****
  • Posts: 862
  • Liked: 70
  • Likes Given: 1
Re: Storable Propellant Earth Departure Stages
« Reply #121 on: 09/09/2013 01:21 pm »
Part 2 of the ARRM asteroid mission hopes to use a SEP.  Using the same specification parts in the SEP tug will keep costs down.
How the components scale or how common they are is the key to lowering costs. 

Missions in the sun do no have to include any significant energy storage( i.e. Mars, Asteroid mission).  Contrast these sun missions to those that depart from LEO to L2, where significant amount of mass and hardware elements must now be included for the eclipse. The structure to support this eclipse mass will vary significantly depending on the power level and eclipse time.   Yes, the thrusters can be pulsed on and off, but reliability must be shown and trip time increase on operation costs included.  Quite a bit of work for the EP folks to step and meet the challenges.

If or when NASA is ever able to truly compare all the architecture options, it will be interesting to see how the mass and cost options develop over time assuming one actually sees technology maturation programs.

One concept is that EP can be used for the chemical uncrasher stage.  Yet,   It does not appear cost effective that NASA would include the high ISP solution of EP so it can use lower ISP for the lander to 'explore sooner', however.

Offline MP99

Re: Storable Propellant Earth Departure Stages
« Reply #122 on: 09/09/2013 07:21 pm »
Solar arrays can do about 150kW/kg

Typo! (Or, I'll take a dozen.)

Edit: ignoring the tug-returns-itself-to-LEO option, I'm wondering how easy it would be to re-use the solar arrays from those tugs either at an EML station, or as part of a TMI stack. With more power, you can get higher Isp for the same thrust, especially where insolation is so much lower at Mars' distance.

Also wondering about having the tug take itself onwards to Mars, so it can be docked to for the trip home - not for propellant, just for the extra power.

cheers, Martin
« Last Edit: 09/09/2013 07:32 pm by MP99 »

Offline Robotbeat

  • Senior Member
  • *****
  • Posts: 39270
  • Minnesota
  • Liked: 25222
  • Likes Given: 12114
Re: Storable Propellant Earth Departure Stages
« Reply #123 on: 09/09/2013 10:52 pm »
Solar arrays can do about 150kW/kg

Typo! (Or, I'll take a dozen.)

Edit: ignoring the tug-returns-itself-to-LEO option, I'm wondering how easy it would be to re-use the solar arrays from those tugs either at an EML station, or as part of a TMI stack. With more power, you can get higher Isp for the same thrust, especially where insolation is so much lower at Mars' distance.

Also wondering about having the tug take itself onwards to Mars, so it can be docked to for the trip home - not for propellant, just for the extra power.

cheers, Martin
heh, you got me! ;) (though 10kW/kg MIGHT be possible... With significant effort)
Chris  Whoever loves correction loves knowledge, but he who hates reproof is stupid.

To the maximum extent practicable, the Federal Government shall plan missions to accommodate the space transportation services capabilities of United States commercial providers. US law http://goo.gl/YZYNt0

Online sdsds

  • Senior Member
  • *****
  • Posts: 7194
  • “With peace and hope for all mankind.”
  • Seattle
  • Liked: 2039
  • Likes Given: 1962
Re: Storable Propellant Earth Departure Stages
« Reply #124 on: 09/12/2013 05:41 am »
One issue would be getting a large enough fairing for large BEO payloads but an 8 meter hammerhead fairing on the Delta IV Heavy might do the job.

You don't need 8m fairings for exploration, even though EELVs could perhaps be modified to support them.

I don't understand the concerns about large fairings. Imagine a storable propellant Earth departure stage more-or-less appropriate for SLS  ginned up by stacking three Briz-Ms, each about 4 m in diameter and combined about 8 or 10 m high (depending on interstages). Gross mass ~70 tonnes. Why would one need anything wider?
— 𝐬𝐝𝐒𝐝𝐬 —

Offline mmeijeri

  • Senior Member
  • *****
  • Posts: 7772
  • Martijn Meijering
  • NL
  • Liked: 397
  • Likes Given: 822
Re: Storable Propellant Earth Departure Stages
« Reply #125 on: 09/12/2013 07:50 am »
Why would one need anything wider?

I think people are suggesting the payloads may have to be wider. Of course, I disagree with that.
Pro-tip: you don't have to be a jerk if someone doesn't agree with your theories

Offline Patchouli

  • Senior Member
  • *****
  • Posts: 4490
  • Liked: 253
  • Likes Given: 457
Re: Storable Propellant Earth Departure Stages
« Reply #126 on: 09/13/2013 04:31 am »
One issue would be getting a large enough fairing for large BEO payloads but an 8 meter hammerhead fairing on the Delta IV Heavy might do the job.

You don't need 8m fairings for exploration, even though EELVs could perhaps be modified to support them.

I don't understand the concerns about large fairings. Imagine a storable propellant Earth departure stage more-or-less appropriate for SLS  ginned up by stacking three Briz-Ms, each about 4 m in diameter and combined about 8 or 10 m high (depending on interstages). Gross mass ~70 tonnes. Why would one need anything wider?

Partly for habs as 4.5M is much too narrow and cramped but mostly for Mars EDL systems.

As for transport of large items that was solved when they built the super guppy.

If they need something even larger NASA could order a 747 LCF.
The cost would be a very small compared to the cost of a lunar program and the savings in time and reduced damage over road transport might more then makeup for it.
« Last Edit: 09/13/2013 04:35 am by Patchouli »

Offline Robotbeat

  • Senior Member
  • *****
  • Posts: 39270
  • Minnesota
  • Liked: 25222
  • Likes Given: 12114
Re: Storable Propellant Earth Departure Stages
« Reply #127 on: 09/13/2013 04:46 am »
4.5m isn't narrow and cramped (and if it is, you can consolidate equipment... 4.5m is a huge space by itself). In fact, one of the lessons learned from Skylab is if you have too big of open spaces, space sickness is worsened and the astronauts can get stuck halfway in between while also being a less efficient use of space.
Chris  Whoever loves correction loves knowledge, but he who hates reproof is stupid.

To the maximum extent practicable, the Federal Government shall plan missions to accommodate the space transportation services capabilities of United States commercial providers. US law http://goo.gl/YZYNt0

Offline gbaikie

  • Full Member
  • ****
  • Posts: 1592
  • Liked: 49
  • Likes Given: 5
Re: Storable Propellant Earth Departure Stages
« Reply #128 on: 09/13/2013 05:40 am »
From my calculations you will need 158 metric tons of storable propellant to make this work from LEO.

With the Falcon Heavy, with its on-record payload capability of 53 metric tons, you would need 3 launches *just to get the propellant into position* This is not counting the stage nor lander itself, which brings the total launches to 5 (assuming you launch the stage partially-filled). 5 launches will cost *more* than an SLS launch for the same performance, costing you $540 million vs the SLS at ~$500 million, just for the launch vehicle. The Falcon Heavy would also require the cost of the DUUS-alternative, while the SLS cost would include it.

You are throwing money away with this route.

SLS is costing about 2 and half billion per year in order to develop a rocket launch with capability of 70 ton to LEO with first planned launch in 2017. So it will have spent over 10 billion dollars before it launches.
So you going to spend over 10 billion dollars which is not included in your launch costs, for something equal in size of Shuttle launch which was partially reusable launch vehicle, and the SLS is not reusable.
So 5 launches of Falcon Heavy is 53 metric tonnes times 5 which is 265 metric tonnes.
 The Falcon Heavy could be launched in 2014, and may launched few times before the year 2017.
If want the 130 ton SLS NASA will have to pay for it's development after 2017.
There is no reason to assume the 70 ton SLS will launched at lower cost than compared to the Shuttle, and there is even less reason to assume a 130 ton SLS will launch for less cost than a Shuttle launch.
And it's likely that the 130 SLS verison of SLS will not actually developed, considering the cost of the 70 ton vehicle will need have it's operational cost paid for, while paying more to develop the 130 ton version.
It's development costs at 2 1/2 billion is not making a rocket which being launched, will undoubtedly cost more, once one using the rockets to launch payloads.

Meanwhile SpaceX is developing a partially reusable launch vehicle, and not likely that by 2017 SpaceX will have partially reusable Falcon Heavy.
Since NASA isn't developing it's 130 ton rocket, in 2017, what you would be comparing in Heavy Falcon which could be partially reusable to the SLS 80 ton expendable rocket.
A NASA will have the difficult task of trying to explain to Congress why a 130 ton rocket is going to cost less than it's 70 ton rocket they have already paid for- and they will have actual numbers showing it's actual costs.
« Last Edit: 09/13/2013 05:48 am by gbaikie »

Offline Proponent

  • Senior Member
  • *****
  • Posts: 7276
  • Liked: 2781
  • Likes Given: 1461
Re: Storable Propellant Earth Departure Stages
« Reply #129 on: 09/13/2013 10:16 am »
SLS is costing about 2 and half billion per year in order to develop

That sounds more like the annual cost of SLS and Orion together.

Offline mmeijeri

  • Senior Member
  • *****
  • Posts: 7772
  • Martijn Meijering
  • NL
  • Liked: 397
  • Likes Given: 822
Re: Storable Propellant Earth Departure Stages
« Reply #130 on: 09/13/2013 10:45 am »
Partly for habs as 4.5M is much too narrow and cramped but mostly for Mars EDL systems.

Of course there are good alternatives for EDL as well. In fact, some of these alternatives may even be necessary for large and heavy payloads! But that's off-topic for this thread, because it doesn't so much have to do with storable propellant, but with not having an HLV. We have several threads for that already, including one for EDL.
Pro-tip: you don't have to be a jerk if someone doesn't agree with your theories

Offline gbaikie

  • Full Member
  • ****
  • Posts: 1592
  • Liked: 49
  • Likes Given: 5
Re: Storable Propellant Earth Departure Stages
« Reply #131 on: 09/13/2013 09:04 pm »
SLS is costing about 2 and half billion per year in order to develop

That sounds more like the annual cost of SLS and Orion together.

Believe that's about 3 billion per year.

Offline alexterrell

  • Full Member
  • ****
  • Posts: 1747
  • Germany
  • Liked: 184
  • Likes Given: 107
Re: Storable Propellant Earth Departure Stages
« Reply #132 on: 09/13/2013 09:43 pm »
One issue would be getting a large enough fairing for large BEO payloads but an 8 meter hammerhead fairing on the Delta IV Heavy might do the job.

You don't need 8m fairings for exploration, even though EELVs could perhaps be modified to support them.

I don't understand the concerns about large fairings. Imagine a storable propellant Earth departure stage more-or-less appropriate for SLS  ginned up by stacking three Briz-Ms, each about 4 m in diameter and combined about 8 or 10 m high (depending on interstages). Gross mass ~70 tonnes. Why would one need anything wider?

We've discussed this before and the only real need anyone came up with for large payload fairings is for Mars Heatshields.

I'm sure an inflatable heatshield would be cheaper to develop than SLS. Or bifocal folding, or something new. Or anything for that matter.

An 8m fairing could also be useful for very large (~1MW) solar arrays unless folding technology is improved.

The habitat issue is not an issue assuming Bigelow style modules work. If there's no SLS, the BA-2100 woulod have to become something like the BA-1200.

Offline M129K

  • Full Member
  • ****
  • Posts: 823
    • "a historian too many" blog.
  • Liked: 71
  • Likes Given: 290
Re: Storable Propellant Earth Departure Stages
« Reply #133 on: 09/16/2013 05:42 pm »
Little bit of a nitpick with the opening post: NASA doesn't want a 43 ton lunar lander, the lander in the LSS DRM is brought into orbit by a CPS and launched on Block 1A, and even using the most optimistic numbers don't allow it to get much more than 30 tons into LLO. If Altair wouldn't have to do the LOI burn for itself and Orion and was brought into orbit via CPS it could have a mass of just 22 metric tons.

Sorry for being, what we call an "antf*cker" over here, and enjoy the rest of the thread ;D

Offline Robotbeat

  • Senior Member
  • *****
  • Posts: 39270
  • Minnesota
  • Liked: 25222
  • Likes Given: 12114
Re: Storable Propellant Earth Departure Stages
« Reply #134 on: 09/16/2013 05:48 pm »
So that further strengthens QuantumG's point (as if it needed any more strengthening).


Note: The term we use here in the US of Merka is "nitpicker." ;)
Chris  Whoever loves correction loves knowledge, but he who hates reproof is stupid.

To the maximum extent practicable, the Federal Government shall plan missions to accommodate the space transportation services capabilities of United States commercial providers. US law http://goo.gl/YZYNt0

Offline a_langwich

  • Full Member
  • ****
  • Posts: 735
  • Liked: 212
  • Likes Given: 48
Re: Storable Propellant Earth Departure Stages
« Reply #135 on: 09/16/2013 06:40 pm »
What's the highest ISP currently available from a storable propellant chemical rocket, in the thrust range necessary for an EDS? 

What's the highest ISP available from non-toxic sources?

Are there theoretically higher options out there?

If you excluded LH as the fuel and chose something else, could you make a cryo stage storable for a year plus?  What's the highest ISP alternative there?

Offline M129K

  • Full Member
  • ****
  • Posts: 823
    • "a historian too many" blog.
  • Liked: 71
  • Likes Given: 290
Re: Storable Propellant Earth Departure Stages
« Reply #136 on: 09/16/2013 06:47 pm »
So that further strengthens QuantumG's point (as if it needed any more strengthening).

Eh, I'm not too sure wether it actually strengthens it... The NASA lander would use hydrogen, which isn't storable, and wouldn't do the LOI burn by itself. The lander would be over 30 tons but you'd need to give it 4050 m/s, not 3107. I think that would actually increase the total mass needed to bring into LEO.

Quote from:   a_langwich

What's the highest ISP currently available from a storable propellant chemical rocket, in the thrust range necessary for an EDS? 

I suppose that would be the 324 second Isp on the German Aestus II engine. I've seen higher though, including 326 for Briz-M (though that probably isn't powerful enough) and 327.5 claimed for the Orion main engine. I suppose that's about the theoretical limit.

Quote from:  a_langwich
What's the highest ISP available from non-toxic sources?
RL-10B2 with 461.5 seconds, soon to be overtaken by Vinci at 465 seconds  ;D

Offline mmeijeri

  • Senior Member
  • *****
  • Posts: 7772
  • Martijn Meijering
  • NL
  • Liked: 397
  • Likes Given: 822
Re: Storable Propellant Earth Departure Stages
« Reply #137 on: 09/16/2013 07:28 pm »
I suppose that would be the 324 second Isp on the German Aestus II engine.

Aestus 2 would be 340s.
Pro-tip: you don't have to be a jerk if someone doesn't agree with your theories

Offline M129K

  • Full Member
  • ****
  • Posts: 823
    • "a historian too many" blog.
  • Liked: 71
  • Likes Given: 290
Re: Storable Propellant Earth Departure Stages
« Reply #138 on: 09/16/2013 07:33 pm »
I suppose that would be the 324 second Isp on the German Aestus II engine.

Aestus 2 would be 340s.

It is? Whoa. Didn't think that was possible. I guess that means Aestus 2 is the most efficient storable propellant engine available?

Offline mmeijeri

  • Senior Member
  • *****
  • Posts: 7772
  • Martijn Meijering
  • NL
  • Liked: 397
  • Likes Given: 822
Re: Storable Propellant Earth Departure Stages
« Reply #139 on: 09/16/2013 07:41 pm »
It is? Whoa. Didn't think that was possible. I guess that means Aestus 2 is the most efficient storable propellant engine available?

I believe the Russians have staged combustion storable engines, and those might do even better, but I can't find a reference now. Note that Aestus 2 has not been finished, though there have apparently been hot firings.
Pro-tip: you don't have to be a jerk if someone doesn't agree with your theories

Offline a_langwich

  • Full Member
  • ****
  • Posts: 735
  • Liked: 212
  • Likes Given: 48
Re: Storable Propellant Earth Departure Stages
« Reply #140 on: 09/17/2013 05:46 am »

Quote from:  a_langwich
What's the highest ISP available from non-toxic sources?
RL-10B2 with 461.5 seconds, soon to be overtaken by Vinci at 465 seconds  ;D


Non-toxic storable. 

Offline M129K

  • Full Member
  • ****
  • Posts: 823
    • "a historian too many" blog.
  • Liked: 71
  • Likes Given: 290
Re: Storable Propellant Earth Departure Stages
« Reply #141 on: 09/17/2013 06:47 pm »
Eh, not so sure in that case. The highest I can remember from the top of my head is 320 seconds for Kerosene/Hydrogen Peroxide.

Offline Robotbeat

  • Senior Member
  • *****
  • Posts: 39270
  • Minnesota
  • Liked: 25222
  • Likes Given: 12114
Re: Storable Propellant Earth Departure Stages
« Reply #142 on: 09/17/2013 08:22 pm »
Eh, not so sure in that case. The highest I can remember from the top of my head is 320 seconds for Kerosene/Hydrogen Peroxide.
Methane/LOx are storable in this environment. Isp over 350s, perhaps 375s.
Chris  Whoever loves correction loves knowledge, but he who hates reproof is stupid.

To the maximum extent practicable, the Federal Government shall plan missions to accommodate the space transportation services capabilities of United States commercial providers. US law http://goo.gl/YZYNt0

Offline Patchouli

  • Senior Member
  • *****
  • Posts: 4490
  • Liked: 253
  • Likes Given: 457
Re: Storable Propellant Earth Departure Stages
« Reply #143 on: 09/17/2013 08:43 pm »
Eh, not so sure in that case. The highest I can remember from the top of my head is 320 seconds for Kerosene/Hydrogen Peroxide.
Methane/LOx are storable in this environment. Isp over 350s, perhaps 375s.

Methane engines large enough for use in an EDS have been tested.

This example is in the RL-10's thrust class a cluster of four to six would be up to the task of propelling a S-IVB sized EDS.
A small EDS such as one needed to send Orion to EML 1 would only need two of them.

Offline M129K

  • Full Member
  • ****
  • Posts: 823
    • "a historian too many" blog.
  • Liked: 71
  • Likes Given: 290
Re: Storable Propellant Earth Departure Stages
« Reply #144 on: 09/17/2013 08:45 pm »
Eh, not so sure in that case. The highest I can remember from the top of my head is 320 seconds for Kerosene/Hydrogen Peroxide.
Methane/LOx are storable in this environment. Isp over 350s, perhaps 375s.

I thought that didn't count. They still require some light boil off control, though for lunar trips it's pretty much storable. If light cryogenics are allowed, methane wins hands down.

Maybe butane is easier to store though. Not sure.

Offline a_langwich

  • Full Member
  • ****
  • Posts: 735
  • Liked: 212
  • Likes Given: 48
Re: Storable Propellant Earth Departure Stages
« Reply #145 on: 09/17/2013 09:33 pm »
Eh, not so sure in that case. The highest I can remember from the top of my head is 320 seconds for Kerosene/Hydrogen Peroxide.
Methane/LOx are storable in this environment. Isp over 350s, perhaps 375s.

Right, that gets filed under cryo non-LH alternatives. 
I'm semi-arbitrarily forcing everthing into four categories:  toxic storable, non-toxic storable, cryo storable, and cryo non-storable.  The middle two are preferred, but have the least heritage of flown designs, although maybe you could toss solid US into the non-toxic storable category?

Were you assuming a sunshade design for those LOX/methane tanks?  What sorts of data do we have on LOX boiloff in these situations?  Centaur stage tests?

Offline Robotbeat

  • Senior Member
  • *****
  • Posts: 39270
  • Minnesota
  • Liked: 25222
  • Likes Given: 12114
Re: Storable Propellant Earth Departure Stages
« Reply #146 on: 09/17/2013 09:42 pm »
Eh, not so sure in that case. The highest I can remember from the top of my head is 320 seconds for Kerosene/Hydrogen Peroxide.
Methane/LOx are storable in this environment. Isp over 350s, perhaps 375s.

I thought that didn't count. They still require some light boil off control, though for lunar trips it's pretty much storable. If light cryogenics are allowed, methane wins hands down.

Maybe butane is easier to store though. Not sure.
No boiloff is needed at all. Passive thermal management (and without a giant sunshield) is sufficient to keep the propellants liquid indefinitely. XCOR calls methane/LOx "space storable."
Chris  Whoever loves correction loves knowledge, but he who hates reproof is stupid.

To the maximum extent practicable, the Federal Government shall plan missions to accommodate the space transportation services capabilities of United States commercial providers. US law http://goo.gl/YZYNt0

Offline QuantumG

  • Senior Member
  • *****
  • Posts: 9238
  • Australia
  • Liked: 4477
  • Likes Given: 1108
Re: Storable Propellant Earth Departure Stages
« Reply #147 on: 09/17/2013 10:59 pm »
There's space storable and then there's space storable. You're not doing a multiyear launch campaign to fill a lox/methane EDS.
Human spaceflight is basically just LARPing now.

Offline gbaikie

  • Full Member
  • ****
  • Posts: 1592
  • Liked: 49
  • Likes Given: 5
Re: Storable Propellant Earth Departure Stages
« Reply #148 on: 09/18/2013 12:20 am »
If we had market for rocket fuel in space, than the issue would be the price difference of rocket fuels.

Therefore one might pay more for the best rocket fuel in space environment- LOX & LH.
Then again, if water was mined in space, you might pay less for LOX & LH, and more for more easily storable
Methane & LOX or other kinds of rarer but more suitable for some purpose rocket fuel.

In modern world, we don't store toasters for a long time- and it has little to do with storability of toasters.
If we had market for rocket fuel we wouldn't have much of issue with storing LOX & LH in Cislunar space- one could have weekly or monthly deliveries of rocket fuel.
If interested in buying delivery of rocket fuel, you don't care to much about whether delivered rocket which is in theory cheaper, you care whether it is actually cheaper. Though probably would how much to would be to deliver by next tuesday- and if cheaper rocket not available you choose that which is available at time you need it- on time delivery becomes a more important way to get customers.

Storability in regards to Mars orbit, would slightly different than a destination which is less than a week travel distance. But still it's a matter of price. A solution might be to ship water to Mars and make rocket fuel in Mars orbit. The biggest factor is how much rocket fuel is needed in Mars orbit per month and the largest amount needed in any given week or month.
If made rocket fuel in Mars orbit and you had low monthly/yearly demand, than you wouldn't have gas station which made a lot of rocket fuel. If one only needed a large amount at one time per two or 3 year period, than rather making a lot of rocket, it would be mostly about storing a lot of rocket fuel, but one would still tend to not want to store large quality over long time. Because in retail you avoid having large inventories, but you want large inventory available at busy time like say Christmas.

But however it's done, it's a matter different prices, so if at Mars orbit if you had choice LH&LOX which was $5000 per lb, how would pay for something else, whether methane, kerosene or whatever.
Or suppose the choice actually more like, LOX at $3000 per lb, Liquid Hydrogen at 12,000 per lb. So 6 parts Oxygen to 1 part Hydrogen- 7 lb of LH&LOX for 30,000 [$4285 per lb].
So if there something else, would want it if it was cheaper at say $4000 per lb. Or is worth more, and therefore you could be willing to pay more for it?
« Last Edit: 09/18/2013 12:29 am by gbaikie »

Offline Proponent

  • Senior Member
  • *****
  • Posts: 7276
  • Liked: 2781
  • Likes Given: 1461
Re: Storable Propellant Earth Departure Stages
« Reply #149 on: 09/18/2013 07:59 pm »
Eh, not so sure in that case. The highest I can remember from the top of my head is 320 seconds for Kerosene/Hydrogen Peroxide.
Methane/LOx are storable in this environment. Isp over 350s, perhaps 375s.

I thought that didn't count. They still require some light boil off control, though for lunar trips it's pretty much storable. If light cryogenics are allowed, methane wins hands down.

Maybe butane is easier to store though. Not sure.

I'll ask my usual question here: why not, say, propane or propylene?  Higher boiling points, greater density and comparable (propane) or slightly superior (propylene) Isp.

Offline M129K

  • Full Member
  • ****
  • Posts: 823
    • "a historian too many" blog.
  • Liked: 71
  • Likes Given: 290
Re: Storable Propellant Earth Departure Stages
« Reply #150 on: 09/18/2013 08:01 pm »
I'll ask my usual question here: why not, say, propane or propylene?  Higher boiling points, greater density and comparable (propane) or slightly superior (propylene) Isp.

Not sure, I thought propane had a lower Isp than methane and I never thought about propylene. If propylene indeed has the potential for an Isp higher than 370 it's probably the superior (storable) fuel.
« Last Edit: 09/18/2013 08:11 pm by M129K »

Offline mmeijeri

  • Senior Member
  • *****
  • Posts: 7772
  • Martijn Meijering
  • NL
  • Liked: 397
  • Likes Given: 822
Re: Storable Propellant Earth Departure Stages
« Reply #151 on: 09/18/2013 08:25 pm »
While it's fun to speculate, it doesn't really matter if there might be other storable propellants, because the main attraction of storable propellants in this context is that they are available now, so are engines using them and so are on-orbit transfer and storage. The same is not true for exotic new combinations any more than it is for LOX/LH2. Who knows what kinds of propellant combinations might be used in the far future. By that time LOX/LH2 will be considered a storable propellant too, at least for a month to a couple of months in LEO and for a year or more at L1/L2, which is all that would be needed to make it more attractive for TMI than noncryogenic propellant. The storable propellants that matter in this context are the ones we have right now, mainly traditional hypergolics and perhaps hydrogen peroxide.
« Last Edit: 09/19/2013 05:33 pm by mmeijeri »
Pro-tip: you don't have to be a jerk if someone doesn't agree with your theories

Offline A_M_Swallow

  • Elite Veteran
  • Senior Member
  • *****
  • Posts: 8906
  • South coast of England
  • Liked: 500
  • Likes Given: 223
Re: Storable Propellant Earth Departure Stages
« Reply #152 on: 09/20/2013 02:55 am »
One thing to watch out for, make sure that your propellant is not pressurised using helium.  Helium, unlike nitrogen, is not storable.  Some engines permit the pumping of the propellant.

Offline Robotbeat

  • Senior Member
  • *****
  • Posts: 39270
  • Minnesota
  • Liked: 25222
  • Likes Given: 12114
Re: Storable Propellant Earth Departure Stages
« Reply #153 on: 09/20/2013 03:25 am »
One thing to watch out for, make sure that your propellant is not pressurised using helium.  Helium, unlike nitrogen, is not storable.  Some engines permit the pumping of the propellant.
Helium is perfectly storable. It's stored in carbon fiber tanks as a gas.
Chris  Whoever loves correction loves knowledge, but he who hates reproof is stupid.

To the maximum extent practicable, the Federal Government shall plan missions to accommodate the space transportation services capabilities of United States commercial providers. US law http://goo.gl/YZYNt0

Online sdsds

  • Senior Member
  • *****
  • Posts: 7194
  • “With peace and hope for all mankind.”
  • Seattle
  • Liked: 2039
  • Likes Given: 1962
Re: Storable Propellant Earth Departure Stages
« Reply #154 on: 06/16/2014 04:56 am »
Recently there has been discussion in other threads of proposals. Dalhousie wrote:
Quote
It is very hard to design a Mars mission using launchers of less than 70 tonnes.  There have only been a handful.  People who keeping banging on about doing Mars with small launches need to come up with the architectures.

I agree with that. In response I wrote
Quote
It would take on-orbit propellant storage and transfer. A lot of it. Reliably. Let's in general terms describe it as "a zillion" launches of propellant.

My favorite design would include two stages just for LEO departure. The first stage would be powered by a cluster of nine Aestus II (RS 72) engines; the second by a cluster of four.
http://forum.nasaspaceflight.com/index.php?topic=34843.msg1213978#msg1213978

Attached I provide an almost farcically simplistic sketch of this LEO departure system, which I call "TSSP." It sends 33 t of payload from LEO to Earth-escape using 62 t of MMH/NTO propellant.
— 𝐬𝐝𝐒𝐝𝐬 —

Offline mmeijeri

  • Senior Member
  • *****
  • Posts: 7772
  • Martijn Meijering
  • NL
  • Liked: 397
  • Likes Given: 822
Re: Storable Propellant Earth Departure Stages
« Reply #155 on: 06/16/2014 11:05 am »
Recently there has been discussion in other threads of proposals. Dalhousie wrote:
Quote
It is very hard to design a Mars mission using launchers of less than 70 tonnes.  There have only been a handful.  People who keeping banging on about doing Mars with small launches need to come up with the architectures.

I disagree, I don't think it's hard at all. Huntress described an architecture that could easily be adapted to the use of smaller launchers. With refueling at EML1/2, SML1/2 and LMO, propulsive orbit insertion and heavily propulsive Mars landing and prepositioning of propellant by means of SEP we have all the technologies we need. It would still be expensive of course, but it is going to be expensive any way you do it.
« Last Edit: 06/16/2014 11:11 am by mmeijeri »
Pro-tip: you don't have to be a jerk if someone doesn't agree with your theories

Offline spacenut

  • Senior Member
  • *****
  • Posts: 5180
  • East Alabama
  • Liked: 2587
  • Likes Given: 2895
Re: Storable Propellant Earth Departure Stages
« Reply #156 on: 06/16/2014 09:50 pm »
No mater what fuel is use, it seems like you must have lox.  Lox and liquid methane are very close in boiling point so they would use the same technology for storage, valves, etc.  Unlike liquid hydrogen which is much colder.  Methane to me wins hands down as it is nothing but pure natural gas, and natural gas is already stored in liquid form around the country during the summer for winter release.  I know, I was a natural gas utility engineer for 39 years, recently retired.  Also, if ULA or SpaceX develops an upper stage using liquid natural gas, the isp is better.

Online sdsds

  • Senior Member
  • *****
  • Posts: 7194
  • “With peace and hope for all mankind.”
  • Seattle
  • Liked: 2039
  • Likes Given: 1962
Re: Storable Propellant Earth Departure Stages
« Reply #157 on: 06/17/2014 03:31 am »
No mater what fuel is use, it seems like you must have lox

Not really. MMH and NTO are both liquids at room temperature, and the NTO (nitrogen tetroxide, N2O4) serves as a really good oxidizer. No LOX required. And the MMH ignites on contact with NTO, making the ignition system rather foolproof!

www.astronautix.com/props/n2o4mmh.htm
— 𝐬𝐝𝐒𝐝𝐬 —

Offline RanulfC

  • Senior Member
  • *****
  • Posts: 4595
  • Heus tu Omnis! Vigilate Hoc!
  • Liked: 900
  • Likes Given: 32
Re: Storable Propellant Earth Departure Stages
« Reply #158 on: 06/18/2014 09:17 pm »
No mater what fuel is use, it seems like you must have lox

Not really. MMH and NTO are both liquids at room temperature, and the NTO (nitrogen tetroxide, N2O4) serves as a really good oxidizer. No LOX required. And the MMH ignites on contact with NTO, making the ignition system rather foolproof!

www.astronautix.com/props/n2o4mmh.htm

Hmm, yes.. But if you use H2O2 instead it gives you auxilary power and life support :)

Its also liquid at "room-temperature" and indefinatly storable as long as its "cooled" below room temperature. While not strictly "hypergolic" as MMH/NTO it has a wide available fuel list and by using a solid screen catalyst it ACTS hypergolic with most fuels. Fun part is you can make it just about anywhere you can make other propellants :) Lastly it's ISP isn't that much lower than MMH/NTO so...

Randy
From The Amazing Catstronaut on the Black Arrow LV:
British physics, old chap. It's undignified to belch flames and effluvia all over the pad, what. A true gentlemen's orbital conveyance lifts itself into the air unostentatiously, with the minimum of spectacle and a modicum of grace. Not like our American cousins' launch vehicles, eh?

Offline savuporo

  • Senior Member
  • *****
  • Posts: 5152
  • Liked: 1002
  • Likes Given: 342
Re: Storable Propellant Earth Departure Stages
« Reply #159 on: 06/18/2014 09:58 pm »
Its also liquid at "room-temperature" and indefinatly storable as long as its "cooled" below room temperature. ..

And compared to MMH/NTO peroxide has negligible flight history and track record in deep space. That includes engines and storage.  Storability of hydrazine and ability to restart engines has been demonstrated after decades in space.
It will take decades before any proposed peroxide system will gain a similar reliability record.
Orion - the first and only manned not-too-deep-space craft

Offline Burninate

  • Full Member
  • ****
  • Posts: 1145
  • Liked: 360
  • Likes Given: 74
Re: Storable Propellant Earth Departure Stages
« Reply #160 on: 06/18/2014 10:02 pm »
Out of curiosity: What is the liquid temperature range of MMH?  of NTO?

Methane has an advantage over most of the less volatile hydrocarbons in that its liquid temperature range at standard pressure very nearly overlaps with oxygen's, allowing for a common thermal environment at a manageable pressure.  ~90K is also high enough that a passive (paint / controlled sunshield-based) cryocooler is practical.  I've seen comments to the effect that this is only workable BLEO due to the blackbody of Earth, but I have not seen any engineering studies.
« Last Edit: 06/18/2014 10:06 pm by Burninate »

Online sdsds

  • Senior Member
  • *****
  • Posts: 7194
  • “With peace and hope for all mankind.”
  • Seattle
  • Liked: 2039
  • Likes Given: 1962
Re: Storable Propellant Earth Departure Stages
« Reply #161 on: 06/19/2014 03:32 am »
Out of curiosity: What is the liquid temperature range of MMH?  of NTO?

The attached table was snipped from PHYSICAL & THERMODYNAMIC PROPERTIES OF HYPERGOLIC PROPELLANTS: A REVIEW
AND UPDATE
, S.L ARNOLD
http://ns.gentoogeek.org/steves_world/hypergol_properties.pdf

Not being a chemist I don't understand most of what's in that paper, so I may be taking this bit out of context. Corrections appreciated!
— 𝐬𝐝𝐒𝐝𝐬 —

Online sdsds

  • Senior Member
  • *****
  • Posts: 7194
  • “With peace and hope for all mankind.”
  • Seattle
  • Liked: 2039
  • Likes Given: 1962
Re: Storable Propellant Earth Departure Stages
« Reply #162 on: 06/21/2014 07:07 pm »
I'm slowly developing a list of reasons why a two-stage Earth-orbit departure system makes sense, particularly when using space-storable propellants. I hope to present that reasoning at some point, but at the moment I'm struggling a bit with the extra "degree of freedom" a two-stage system provides. In particular, there's the question of how to choose the sizes for the upper and lower stages.

The attached images show three scenarios. The first shows stage sizes optimized (I hope!) for sending a 33 t payload to Earth-escape velocity. The second shows the payload mass (17.7 t) which stages of those sizes could carry through a (4400 m/s) trans-Mars departure burn. The third shows the stage sizes if they were designed to optimally send 17.7 tons through TMI.

Apparently repurposing the stage sizes chosen for an escape burn requires 89,124 kg of mass in LEO, whereas stages optimal for TMI would require only 87,624 kg in LEO. Is that reasonable?
— 𝐬𝐝𝐒𝐝𝐬 —

Offline mmeijeri

  • Senior Member
  • *****
  • Posts: 7772
  • Martijn Meijering
  • NL
  • Liked: 397
  • Likes Given: 822
Re: Storable Propellant Earth Departure Stages
« Reply #163 on: 06/21/2014 07:27 pm »
I'm slowly developing a list of reasons why a two-stage Earth-orbit departure system makes sense, particularly when using space-storable propellants. I hope to present that reasoning at some point, but at the moment I'm struggling a bit with the extra "degree of freedom" a two-stage system provides. In particular, there's the question of how to choose the sizes for the upper and lower stages.

I see no reason to insist on a storable first stage. Are you assuming the entire stack departs from LEO directly, in one piece? If you assume departure from L1/L2, you can leave open the mode of transport to L1/L2, and only a single storable stage might be needed from L1/L2 onward. Modules can travel to L1/L2 individually.
Pro-tip: you don't have to be a jerk if someone doesn't agree with your theories

Online sdsds

  • Senior Member
  • *****
  • Posts: 7194
  • “With peace and hope for all mankind.”
  • Seattle
  • Liked: 2039
  • Likes Given: 1962
Re: Storable Propellant Earth Departure Stages
« Reply #164 on: 06/21/2014 09:09 pm »
If you assume departure from L1/L2 [...] only a single storable stage might be needed from L1/L2 onward.

I agree. I'm just not convinced that should be considered an "Earth departure stage." Putting the rendezvous (and thus departure) point is somewhere in cis-lunar space -- whether that's a Lagrange point, DRO, or whatever -- considerably redefines the departure task! I also believe LEO departure using a "high energy" stage integrated with the launch vehicle (and e.g. loaded with propellant as part of the launch vehicle count down) is certainly going to be more efficient by many measures. That's what makes something like the Exploration Upper Stage proposed for SLS so exciting.

But LEO rendezvous and departure is a meme that seems quite resilient!
— 𝐬𝐝𝐒𝐝𝐬 —

Offline mmeijeri

  • Senior Member
  • *****
  • Posts: 7772
  • Martijn Meijering
  • NL
  • Liked: 397
  • Likes Given: 822
Re: Storable Propellant Earth Departure Stages
« Reply #165 on: 06/21/2014 09:20 pm »
I agree. I'm just not convinced that should be considered an "Earth departure stage." Putting the rendezvous (and thus departure) point is somewhere in cis-lunar space -- whether that's a Lagrange point, DRO, or whatever -- considerably redefines the departure task!

It's what would take you outside the gravity well, but yeah, you'd have to get there first.

Quote
I also believe LEO departure using a "high energy" stage integrated with the launch vehicle (and e.g. loaded with propellant as part of the launch vehicle count down) is certainly going to be more efficient by many measures. That's what makes something like the Exploration Upper Stage proposed for SLS so exciting.

Well, as you know I'm totally against SLS anyway, but whatever launch vehicle you use, I don't see how LEO departure is more efficient. If you are using a storable stage from LEO, then 3.2km/s of your delta-v is provided by storable propulsion, which is less efficient than LOX/LH2.

Quote
But LEO rendezvous and departure is a meme that seems quite resilient!

Persistent is the word I'd use, as I think it's a big mistake...
Pro-tip: you don't have to be a jerk if someone doesn't agree with your theories

Online sdsds

  • Senior Member
  • *****
  • Posts: 7194
  • “With peace and hope for all mankind.”
  • Seattle
  • Liked: 2039
  • Likes Given: 1962
Re: Storable Propellant Earth Departure Stages
« Reply #166 on: 06/22/2014 02:37 am »
whatever launch vehicle you use, I don't see how LEO departure is more efficient. If you are using a storable stage from LEO, then 3.2km/s of your delta-v is provided by storable propulsion, which is less efficient than LOX/LH2.

Yes, when measured by mass of propellant required, LH2 is a winner. But that's not the only factor. There is also cost.

I have trouble imagining a low cost launch provider that uses a high energy upper stage. Delta, Atlas, Ariane, H-IIA; all are high cost. Further I doubt there is a low cost means of developing a launch system that includes high energy upper stage propulsion. In contrast I hope there may someday be at least one low cost launch provider using a moderate energy upper stage propulsion system: SpaceX, using a fully recovered F9R. And there might be others!
— 𝐬𝐝𝐒𝐝𝐬 —

Offline savuporo

  • Senior Member
  • *****
  • Posts: 5152
  • Liked: 1002
  • Likes Given: 342
Re: Storable Propellant Earth Departure Stages
« Reply #167 on: 06/22/2014 08:15 am »
..I hope to present that reasoning at some point, but at the moment I'm struggling a bit with the extra "degree of freedom" a two-stage system provides. In particular, there's the question of how to choose the sizes for the upper and lower stages.
Gives you another freedom : you can use flight proven solid fuel kick stages as first stage, with high mass fraction. Or zeroth stage ..

http://cms.atk.com/SiteCollectionDocuments/ProductsAndServices/ATK-Motor-Catalog-2012.pdf
Orion - the first and only manned not-too-deep-space craft

Offline mmeijeri

  • Senior Member
  • *****
  • Posts: 7772
  • Martijn Meijering
  • NL
  • Liked: 397
  • Likes Given: 822
Re: Storable Propellant Earth Departure Stages
« Reply #168 on: 06/22/2014 10:54 am »
I have trouble imagining a low cost launch provider that uses a high energy upper stage. Delta, Atlas, Ariane, H-IIA; all are high cost. Further I doubt there is a low cost means of developing a launch system that includes high energy upper stage propulsion. In contrast I hope there may someday be at least one low cost launch provider using a moderate energy upper stage propulsion system: SpaceX, using a fully recovered F9R. And there might be others!

So how does that argue in favour of LEO departure?
Pro-tip: you don't have to be a jerk if someone doesn't agree with your theories

Offline A_M_Swallow

  • Elite Veteran
  • Senior Member
  • *****
  • Posts: 8906
  • South coast of England
  • Liked: 500
  • Likes Given: 223
Re: Storable Propellant Earth Departure Stages
« Reply #169 on: 06/22/2014 12:22 pm »
I have trouble imagining a low cost launch provider that uses a high energy upper stage. Delta, Atlas, Ariane, H-IIA; all are high cost. Further I doubt there is a low cost means of developing a launch system that includes high energy upper stage propulsion. In contrast I hope there may someday be at least one low cost launch provider using a moderate energy upper stage propulsion system: SpaceX, using a fully recovered F9R. And there might be others!

So how does that argue in favour of LEO departure?

Restrictions on upper stages tend to also apply to in space stages (including space tugs).

A space tug that uses liquid hydrogen as a fuel has to 'launch' within minutes of being fuelled to minimise boil-off.

Whereas a space tug that uses say methane/LOX can orbit the Earth a few times to get to the right launch location.

Offline mmeijeri

  • Senior Member
  • *****
  • Posts: 7772
  • Martijn Meijering
  • NL
  • Liked: 397
  • Likes Given: 822
Re: Storable Propellant Earth Departure Stages
« Reply #170 on: 06/22/2014 12:51 pm »
That doesn't matter for a storable stage.
Pro-tip: you don't have to be a jerk if someone doesn't agree with your theories

Online sdsds

  • Senior Member
  • *****
  • Posts: 7194
  • “With peace and hope for all mankind.”
  • Seattle
  • Liked: 2039
  • Likes Given: 1962
Re: Storable Propellant Earth Departure Stages
« Reply #171 on: 06/23/2014 03:41 am »
there may someday be at least one low cost launch provider using a moderate energy upper stage propulsion system: SpaceX, using a fully recovered F9R. And there might be others!

So how does that argue in favour of LEO departure?

It's about the upper stage recovery of a "fully recovered F9R." Personally I doubt a launch vehicle upper stage which provides propulsion all the way to a cis-lunar rendezvous and departure point is going to be able to return to Earth for recovery and reuse. But an upper stage delivering its payload to 350 x 350 km? I hope that stage could get home again intact!

Discussion of reusability leads into what could be a major advantage of a two stage storable propellant LEO departure vehicle. It seems quite possible the lower stage could return from the intermediate orbit it reaches, without expending extraordinary amounts of propellant and without a major hit to the payload mass. In particular, a vehicle which in expendable mode takes 33 t from LEO through Earth-escape might instead take 30.7 t (93%) while returning its lower stage to LEO. For TMI, the payload goes from 17.7 t to 16 t (90%).
— 𝐬𝐝𝐒𝐝𝐬 —

Offline mmeijeri

  • Senior Member
  • *****
  • Posts: 7772
  • Martijn Meijering
  • NL
  • Liked: 397
  • Likes Given: 822
Re: Storable Propellant Earth Departure Stages
« Reply #172 on: 06/25/2014 03:37 pm »
It's about the upper stage recovery of a "fully recovered F9R."

Ah, OK, but in that case what you need is the ability to drop off propellant in LEO. It's not required that the whole spacecraft stack should depart from LEO in one piece.
Pro-tip: you don't have to be a jerk if someone doesn't agree with your theories

Offline RanulfC

  • Senior Member
  • *****
  • Posts: 4595
  • Heus tu Omnis! Vigilate Hoc!
  • Liked: 900
  • Likes Given: 32
Re: Storable Propellant Earth Departure Stages
« Reply #173 on: 06/26/2014 03:52 pm »
Its also liquid at "room-temperature" and indefinatly storable as long as its "cooled" below room temperature. ..

And compared to MMH/NTO peroxide has negligible flight history and track record in deep space. That includes engines and storage. Storability of hydrazine and ability to restart engines has been demonstrated after decades in space.
It will take decades before any proposed peroxide system will gain a similar reliability record.

Peroxide has had a lot of work done on its stability and storage possibilites even though it lacks an actual "history" of use in space. Preoxide has the same qualities and usability as hydrazine but MMH has the "history" of use behind it. Everyone "knows" peroxide would be just as good and in many cases better than hydrazine the main issue is no one wants to make the change. Given the depth of "background" that MMH/NOT has even I can see "why" there is the reluctance to do so but thats no reason to assume it never will. It probably won't, but... :)

Randy
From The Amazing Catstronaut on the Black Arrow LV:
British physics, old chap. It's undignified to belch flames and effluvia all over the pad, what. A true gentlemen's orbital conveyance lifts itself into the air unostentatiously, with the minimum of spectacle and a modicum of grace. Not like our American cousins' launch vehicles, eh?

Offline Proponent

  • Senior Member
  • *****
  • Posts: 7276
  • Liked: 2781
  • Likes Given: 1461
Re: Storable Propellant Earth Departure Stages
« Reply #174 on: 06/26/2014 07:40 pm »
Peroxide has an extensive history, it's just that it was so long ago everybody's forgotten about it.  In the 50s and early 60s, is was commonly used.  Mercury's thrusters, for example, were peroxide-powered.  Other applications included attitude control on the upper stages of the Scout, Titan I and Titan II, as well as the Syncom satellites (the first geosynch comsats).

Online sdsds

  • Senior Member
  • *****
  • Posts: 7194
  • “With peace and hope for all mankind.”
  • Seattle
  • Liked: 2039
  • Likes Given: 1962
Re: Storable Propellant Earth Departure Stages
« Reply #175 on: 06/27/2014 01:47 am »
Peroxide has an extensive history, it's just that it was so long ago everybody's forgotten about it.  In the 50s and early 60s, is was commonly used.  Mercury's thrusters, for example, were peroxide-powered.  Other applications included attitude control on the upper stages of the Scout, Titan I and Titan II, as well as the Syncom satellites (the first geosynch comsats).

Historically has peroxide always been pressure-fed? I know e.g. the V2 used peroxide for its gas generator, but I don't think the pump it drove pressurized the peroxide, only the main engine propellants.

I've been rather assuming that the "winner" for a storable propellant EDS was going to be a pump-fed engine....
— 𝐬𝐝𝐒𝐝𝐬 —

Offline Proponent

  • Senior Member
  • *****
  • Posts: 7276
  • Liked: 2781
  • Likes Given: 1461
Re: Storable Propellant Earth Departure Stages
« Reply #176 on: 06/27/2014 09:18 am »
I'd be surprised if the monoprop peroxide thrusters were anything but pressure-fed.

The largest application of peroxide has been in Britain's Gamma engine family, where it served as an oxidizer for kerosene.  These engines had turbopumps driven, I believe, by peroxide decomposition.  Among other applications, they powered the first two stages of Britain's Black Arrow launch vehicle.

Offline RanulfC

  • Senior Member
  • *****
  • Posts: 4595
  • Heus tu Omnis! Vigilate Hoc!
  • Liked: 900
  • Likes Given: 32
Re: Storable Propellant Earth Departure Stages
« Reply #177 on: 06/27/2014 04:16 pm »
Peroxide has an extensive history, it's just that it was so long ago everybody's forgotten about it.  In the 50s and early 60s, is was commonly used.  Mercury's thrusters, for example, were peroxide-powered.  Other applications included attitude control on the upper stages of the Scout, Titan I and Titan II, as well as the Syncom satellites (the first geosynch comsats).

Historically has peroxide always been pressure-fed? I know e.g. the V2 used peroxide for its gas generator, but I don't think the pump it drove pressurized the peroxide, only the main engine propellants.

I've been rather assuming that the "winner" for a storable propellant EDS was going to be a pump-fed engine....

No there were peroxide pump fed engines as well, the most efficent ones actually decomposed the peroxide to drive the fuel/oxidizer pumps themselves. Made for a very compact, powerful engine. (http://www.hydrogen-peroxide.us/history-US-Reaction-Motors/AIAA-2001-3838_History_of_RMI_Super_Performance_90_Percent_H2O2-Kerosene_LR-40_RE-pitch.pdf) There has been ongoing work on various peroxide motors by various people/companies but since "official" interest in peroxide motors waxes and wanes with alarming irregularity its rather hard for anyone to get "serious" work done.

I'd be surprised if the monoprop peroxide thrusters were anything but pressure-fed.

The largest application of peroxide has been in Britain's Gamma engine family, where it served as an oxidizer for kerosene.  These engines had turbopumps driven, I believe, by peroxide decomposition.  Among other applications, they powered the first two stages of Britain's Black Arrow launch vehicle.

Actually the "largest" peroxide engine WAS actually pressure fed. The Beal BA-2 was a TSTO designed to use one pressure fed H2O2/Kerosene engine per stage. (http://en.wikipedia.org/wiki/Beal_Aerospace) Beal test fired BA0810 engine for the second stage in 2000 at their (then :) ) McGregor rocket test site in Texas. At 810,000lbs of thrust it was the largest engine tested at the time since the Apollo program. Beal planned on using pressure-fed engines and composite tanks to reduce construction costs and then launch from an island in the Caribbean. Testing went fine, (composite tanks and pressure fed engines of large size were compatible with peroxide) but the business plan failed and negotiations on the launch site went south as well.

My favorite "fantasy" booster has pump-fed engines on the order of the BA810 as the booster engines for a VTVL, boost-back 1st stage.

On a historical note, A Vertical Empire (http://www.spaceuk.org/) points out that had British space plans been given the go-ahead instead of turning to American launchers the plans called for the replacement of the Gamma series motors with the more powerful and more efficent Stentor motors designed for the Blue Streak missile program. The Stentor engine cluster (4-chamber) version of the Black Arrow would have had about double the capacity of the BA itself but at a higher cost, though it should be noted that the "costing" of the Black Arrow was highly inaccurate as going from the Black Knight to Black Arrow was more difficult and expensive than anticipated.

In addition as the above page notes the Black Arrow ended up being the "ulitmate" development for the Gamma series engines and applications. Meanwhile the Stentor engines and design could be built upon for larger and more capable launch vehicles. As the page also notes though, Britan at the time could AFFORD the Black Arrow and could NOT afford the alternative design. Unfortunatly.

Randy
From The Amazing Catstronaut on the Black Arrow LV:
British physics, old chap. It's undignified to belch flames and effluvia all over the pad, what. A true gentlemen's orbital conveyance lifts itself into the air unostentatiously, with the minimum of spectacle and a modicum of grace. Not like our American cousins' launch vehicles, eh?

Tags:
 

Advertisement NovaTech
Advertisement Northrop Grumman
Advertisement
Advertisement Margaritaville Beach Resort South Padre Island
Advertisement Brady Kenniston
Advertisement NextSpaceflight
Advertisement Nathan Barker Photography
0