SLS return to the moon

[fredinno]

2 hours ago, Red Fang said:

Trying to approach this from another angle. 

I'd go for as many reusable/permanent items in lunar mission architecture, in order to turn it into a pork project that is difficult to cancel after the first successful flight. STS flew for thirty years before it got dumped, ISS is probably going to satay in LEO for 30 years... Apollo got caned early, and its specialized and expendable nature made it easier. Reusables/permanent structures make governments commit to a long-term project. 

Why would being expendable make that job easier? Making stuff after every flight creates more jobs (assuming it doesn't have Shuttle-level maintenance.)

Apollo was canned as, in an attempt to kill off any future lunar missions, the entirety of AAP minus Skylab (even post-Skylab LEO AAP missions) was killed off- not to mention the decisions of the Nixon Administration, and the Shuttle's promise of low cost and pork was too appealing.

Considering how much Congress defends SLS and try to make sure it doesn't get cancelled (even passing laws to make sure it never happens) (even w/o a proper mission), I highly doubt it'll be killed until a Columbia happens. Even then, the Shuttle was cancelled because it was painfully obviosu it was fundamentally flawed in its design. SLS/Orion is not, for the most part.

They need to keep at least a Lunar Space station to keep SLS running, and if a lunar landing program is international, it's far less likely to be cancelled.

[Rune]

12 hours ago, fredinno said:

Why would it become an ISS 2.0? The proposals for a lunar space station are usually only a few modules (far smaller), and can be based off ISS modules or Cygnus.

A fuel depot might complicate things though.

Logistics requirements, mostly. You are talking about doing a scale model of ISS, yeah, but you are supplying it with SLS and similar superboosters. Meaning a launch cadence of once a year (a mission every two years, BTW) will eat about the same slice of budget pie as the 6+ resupply/crew flights to ISS each year. Rough order of magnitude references here, assuming ISS operations eat ~5 billion and SLS launch costs alone will be something like ~3 billion or so each year, meaning both numbers are between one and ten billion. And yes, I know how rough that sounds, it is that kind of general comment drawing on murky half-remembered figures... but I have a hunch that the contractors will take care to size each effort so it employs the same number of people for the same amount of money, give or take.

12 hours ago, fredinno said:

Why do you need advanced fuels? This report: http://www.sei.aero/eng/papers/uploads/archive/AIAA-2013-5479_Presentation.pdf

shows that the best choice for a lunar lander would be RP-1 (very well understood) or CH4 (less well understood due to new engine, but slightly better performance) I would choose CH4 if you want to refuel using ISRU eventually and do reuse (I would choose this one), or RP-1 if you want to get boots to the ground sooner at a slightly lower cost.

Boil off is not a problem for ~15 day lunar missions, though lunar base missions would complicate things and require better insulation to mitigate boil off.

LH2 might also work, it has the lowest mass of them all, but it's probably bad for the longer term when building a base due to higher boil off rate.

It also shows a single stage lander is infeasible due to ~10T higher launch mass.

Those would be advanced fuels, compared with hydrazine/NTO. Indefinite engine firings are more complicated to engineer, and keeping the propellants happy in their tanks requires much more of an effort. Not to mention the relative mechanical complexity of the engine, of course... I would also choose methane if ISRU was a thing, sure, but it is very much not a thing right now, and won't be for a long time (unless you focus you space program on making it a thing, but then you wouldn't be going to the Moon). Now, if they are the difference between one or more stages? Yeah, maybe, in a reusable architecture. But as I showed with dirty back-of-the-envelope calculations, even then they aren't a clear-cut winner.

12 hours ago, fredinno said:

The amount of fuel needed for HAB reuse is uneconomical unless you use ION drives, and those are pointless for HABs as you need higher thrust engines to make sure the Van Allen doesn't give the astronauts cancer by passing through them too much.

A lunar lander can also be upgraded with more engines, an inflatable heat shield, and a drop tank stage and used as a ISRU using MAV (assuming the lander uses CH4 as fuel), as per se the Boeing proposal. The aerodynamics will be worse, yes, but Mars' atmosphere is very thin. It might not be a huge deal.

Or unless you save lugging a reentry capsule all around the solar system, and instead use that mass for extra propellant for a minimal insertion burn, no exotic propulsion system required. Then you can aerobrake over the course of months, while the crew is picked up in the current flavor of crew launcher, or leave the hab parked in a suitable high-energy orbit (DRO, L1/L2, wherever). A habitat for long term flights may weight form around 20mT (salyut) to maybe 40mT (transhab). Oh, and it would count as its own fuel depot and station, but let's not get into that. Bolting a return capsule to that may add 50% to the dry weight (the 20mT of Orion, for example), which is a more than sizeable increment. If that is the case, it is not a clear-cut thing that discarding the habitat even saves mass on total mission architecture, never mind the cost benefits of reusing hardware.

And if you actually do some technology development and refine aerocapture techniques... well, then you are actually saving a lot of mass in order to reuse equipment. I'd call that a win-win. But yeah, I never expected RL government agencies to actually follow the sensible plan because it is sensible. That would be crazy!

About the applicability of lunar landing technology for Martian environments comments ... Yeah, other that having different gravitational, thermal, aerodynamic, ground support, and mission duration environments, they are totally the same mission. Or in other words, about as applicable as shuttle or commsat technology.

 

Rune. To paraphrase Churchill... "the US will always do the right thing, right after it has tried every other possible bad option".

Edited March 21, 2016 by Rune

[fredinno]

6 hours ago, Rune said:

Logistics requirements, mostly. You are talking about doing a scale model of ISS, yeah, but you are supplying it with SLS and similar superboosters. Meaning a launch cadence of once a year (a mission every two years, BTW) will eat about the same slice of budget pie as the 6+ resupply/crew flights to ISS each year. Rough order of magnitude references here, assuming ISS operations eat ~5 billion and SLS launch costs alone will be something like ~3 billion or so each year, meaning both numbers are between one and ten billion. And yes, I know how rough that sounds, it is that kind of general comment drawing on murky half-remembered figures... but I have a hunch that the contractors will take care to size each effort so it employs the same number of people for the same amount of money, give or take.

Those would be advanced fuels, compared with hydrazine/NTO. Indefinite engine firings are more complicated to engineer, and keeping the propellants happy in their tanks requires much more of an effort. Not to mention the relative mechanical complexity of the engine, of course... I would also choose methane if ISRU was a thing, sure, but it is very much not a thing right now, and won't be for a long time (unless you focus you space program on making it a thing, but then you wouldn't be going to the Moon). Now, if they are the difference between one or more stages? Yeah, maybe, in a reusable architecture. But as I showed with dirty back-of-the-envelope calculations, even then they aren't a clear-cut winner.

Or unless you save lugging a reentry capsule all around the solar system, and instead use that mass for extra propellant for a minimal insertion burn, no exotic propulsion system required. Then you can aerobrake over the course of months, while the crew is picked up in the current flavor of crew launcher, or leave the hab parked in a suitable high-energy orbit (DRO, L1/L2, wherever). A habitat for long term flights may weight form around 20mT (salyut) to maybe 40mT (transhab). Oh, and it would count as its own fuel depot and station, but let's not get into that. Bolting a return capsule to that may add 50% to the dry weight (the 20mT of Orion, for example), which is a more than sizeable increment. If that is the case, it is not a clear-cut thing that discarding the habitat even saves mass on total mission architecture, never mind the cost benefits of reusing hardware.

And if you actually do some technology development and refine aerocapture techniques... well, then you are actually saving a lot of mass in order to reuse equipment. I'd call that a win-win. But yeah, I never expected RL government agencies to actually follow the sensible plan because it is sensible. That would be crazy!

About the applicability of lunar landing technology for Martian environments comments ... Yeah, other that having different gravitational, thermal, aerodynamic, ground support, and mission duration environments, they are totally the same mission. Or in other words, about as applicable as shuttle or commsat technology.

 

Rune. To paraphrase Churchill... "the US will always do the right thing, right after it has tried every other possible bad option".

Logistics for a Lunar space station would be less than that for the ISS-Shuttle used through most of its lifetime, as the Shuttle cost around the same amount as the SLS per marginal launch cost ($450 million vs $500 million). Yes, it'll be more expensive than Commercial cargo, but as I said earlier, the cost of the ISS was bloated significantly via the SLS. Using 4-5 launches to build a space station in Lunar Orbit. And you can launch up to 3 times with SLS using existing infrastructure- as long as you have the money. The ISS will be retired around the same time SLS will start picking up and building its lunar space station (2024) if we start right after EM-2.

You know, I would rather trust a published report that clearly states a lunar landing mission with only a single stage is unavailable due to the additional mass. Seriously dude, look at page 16, which shows the vehicle masses for different propellants and lander types. Single stage always loses in terms of mass. Even the report itself concludes what I've been telling you.

We can use ULA's IVF tech for the 'infinite burns' necessary- Ch4 simplifies this though as it does not need to be pressurized (it self pressurizes)- only the Lox needs to be. RP-1 is very well understood- keeping the propellants happy is something the Atlas ICBM had to face all those years ago. The only part that needs real new development in the descent stage is IVF tech, which is already partially developed.

Bull. You need a 0.7 km/s burn to get to Earth C3 a 'minimal insertion burn'. That doesn't sound like much, but that's quite a bit of propellant, since this is the very last burn. This map shows aerobrake is possible to C3, but really, that's only if you have a heat shield, otherwise, you're just going to be flying back off into space. Rendezvous with that would require one SLS launch, and add more R+D money to the costs, as noone has ever done rendezvous in such an elliptical orbit. Noone wants to be stranded with 0 food supplies left. Also, reuse is pointless if you only have 1 launch every 2 years, which is what the Mars Launch windows limit you to, assuming one mission per window.

And Salyut as a HAB is laughable. Do you have any idea how crowded it would be after you put in all the food supplies? BA-330 is pretty much the minimum size for a 4 man crew on a Mars mission as a HAB. And where are you getting the 50% dry mass for Orion? BA-330 is 20T by itself (not to mention the exact number for an Orion is 18T, removing a lot of unnecessary propellant. Here is Mars Direct. http://www.astronautix.com/craft/marirect.htm Keep in mind you'd need more launches, as Mars direct uses a 140T LV instead of the 100T of SLS Block IB, and that NASA already rejected it, and added a in space HAB/ propulsion module, and another Mars Direct LV launch.

How is reuse supposed to shave mass? I would expect a final burn to cost more in mass with propellant than 18T. Mars Direct in fact rejected what you are proposing, and made the system expendable- unless you had inflatable HABs, reuse added too much more mass to be possible with a 2 launch infrastructure.

[Rune]

41 minutes ago, fredinno said:

Logistics for a Lunar space station would be less than that for the ISS-Shuttle used through most of its lifetime, as the Shuttle cost around the same amount as the SLS per marginal launch cost ($450 million vs $500 million). Yes, it'll be more expensive than Commercial cargo, but as I said earlier, the cost of the ISS was bloated significantly via the SLS. Using 4-5 launches to build a space station in Lunar Orbit. And you can launch up to 3 times with SLS using existing infrastructure- as long as you have the money. The ISS will be retired around the same time SLS will start picking up and building its lunar space station (2024) if we start right after EM-2.

You know, I would rather trust a published report that clearly states a lunar landing mission with only a single stage is unavailable due to the additional mass. Seriously dude, look at page 16, which shows the vehicle masses for different propellants and lander types. Single stage always loses in terms of mass. Even the report itself concludes what I've been telling you.

Sure, it is doable. I would also wager it is way more likely than what I would do. And sure, single-staging it is less efficient that using more than one stage. I actually started saying I would use a two-stage lander in this very same thread! That way it can do cargo missions with a lot of payload. But the thing that started me commenting here was a dual-launch SLS lunar mission architecture. With such a monster dedicated to a lander flight (or fuel to refuel it as Nibb proposed), being too subtle about the lander's efficiency is not really such a strong requirement.

41 minutes ago, fredinno said:

We can use ULA's IVF tech for the 'infinite burns' necessary- Ch4 simplifies this though as it does not need to be pressurized (it self pressurizes)- only the Lox needs to be. RP-1 is very well understood- keeping the propellants happy is something the Atlas ICBM had to face all those years ago. The only part that needs real new development in the descent stage is IVF tech, which is already partially developed.

 

Yeah, that's still much more complex than four valves and a couple helium bottles in my book. And all non-storables also need pretty specialized equipment in the tank to endure multi-day stays in space because of the thermal environment, and have losses BTW. Otherwise they would be called storables. Sure, you can minimize losses with tons of thermal protection, and actually use those losses for stationkeeping and producing power and such. Neat trick, but it is not free. If you really stop all leaks, your tank will blow up from overpressure as heats builds up pressure, or have a sophisticated radiator and heat exchanger system capable of cryocooling. And of course, all the fancy systems have a given failure rate...

41 minutes ago, fredinno said:

Bull. You need a 0.7 km/s burn to get to Earth C3 a 'minimal insertion burn'. That doesn't sound like much, but that's quite a bit of propellant, since this is the very last burn. This map shows aerobrake is possible to C3, but really, that's only if you have a heat shield, otherwise, you're just going to be flying back off into space. Rendezvous with that would require one SLS launch, and add more R+D money to the costs, as noone has ever done rendezvous in such an elliptical orbit. Noone wants to be stranded with 0 food supplies left. Also, reuse is pointless if you only have 1 launch every 2 years, which is what the Mars Launch windows limit you to, assuming one mission per window.

And Salyut as a HAB is laughable. Do you have any idea how crowded it would be after you put in all the food supplies? BA-330 is pretty much the minimum size for a 4 man crew on a Mars mission as a HAB. And where are you getting the 50% dry mass for Orion? BA-330 is 20T by itself (not to mention the exact number for an Orion is 18T, removing a lot of unnecessary propellant. Here is Mars Direct. http://www.astronautix.com/craft/marirect.htm Keep in mind you'd need more launches, as Mars direct uses a 140T LV instead of the 100T of SLS Block IB, and that NASA already rejected it, and added a in space HAB/ propulsion module, and another Mars Direct LV launch.

How is reuse supposed to shave mass? I would expect a final burn to cost more in mass with propellant than 18T. Mars Direct in fact rejected what you are proposing, and made the system expendable- unless you had inflatable HABs, reuse added too much more mass to be possible with a 2 launch infrastructure.

Actually, 0.7km/s doesn't sound too bad. With 320s, that's mass ratio a bit under 1.25, meaning 25% of the payload mass. If we are talking about a ~40mT Transhab, that's about 10mT of propellant (storables have such awesome tankage fractions, you can almost discount the stage dry weight), which is actually way less than what Orion weights. After that, no heatshield, just successive passes through the upper atmosphere like all Martian probes do over the space of a couple of months. And of course you would go back into space. Repeatedly. But you are in a habitat that survived a trip to Mars, so why not? Or, you know, get fancy the way JPL people like to do, and capture using the Moon as a force multiplier for your burns.

Salyut as habitat made some impressive record for long duration stays, BTW. Salyut 6 had a single mission spanning 185 days. Not quite Mars-and-back, and they did launch Progresses to it, but it already had all the important things (attitude control, ECLSS, airlock and the like), it just needs more room and supplies to handle longer missions. Taking the venerable DOS modules as a base for a deep-space ship is a really obvious move, I think... they've been flying with extraordinary success for a really long time.

Reuse of course, saves no mass, it saves cost on subsequent missions and (potentially) adds confidence in the systems. But not lugging an Orion all the way to Mars and back, for example, saves you quite a few tons, as I have just shown.

 

Rune. We are actually quite in agreement on most points, I think.

 

Edited March 21, 2016 by Rune

[fredinno]

40 minutes ago, Rune said:

Sure, it is doable. I would also wager it is way more likely than what I would do. And sure, single-staging it is less efficient that using more than one stage. I actually started saying I would use a two-stage lander in this very same thread! That way it can do cargo missions with a lot of payload. But the thing that started me commenting here was a dual-launch SLS lunar mission architecture. With such a monster dedicated to a lander flight (or fuel to refuel it as Nibb proposed), being too subtle about the lander's efficiency is not really such a strong requirement.

Yeah, that's still much more complex than four valves and a couple helium bottles in my book. And all non-storables also need pretty specialized equipment in the tank to endure multi-day stays in space because of the thermal environment, and have losses BTW. Otherwise they would be called storables. Sure, you can minimize losses with tons of thermal protection, and actually use those losses for stationkeeping and producing power and such. Neat trick, but it is not free. If you really stop all leaks, your tank will blow up from overpressure as heats builds up pressure, or have a sophisticated radiator and heat exchanger system capable of cryocooling. And of course, all the fancy systems have a given failure rate...

Actually, 0.7km/s doesn't sound too bad. With 320s, that's mass ratio a bit under 1.25, meaning 25% of the payload mass. If we are talking about a ~40mT Transhab, that's about 10mT of propellant (storables have such awesome tankage fractions, you can almost discount the stage dry weight), which is actually way less than what Orion weights. After that, no heatshield, just successive passes through the upper atmosphere like all Martian probes do over the space of a couple of months. And of course you would go back into space. Repeatedly. But you are in a habitat that survived a trip to Mars, so why not? Or, you know, get fancy the way JPL people like to do, and capture using the Moon as a force multiplier for your burns.

Salyut as habitat made some impressive record for long duration stays, BTW. Salyut 6 had a single mission spanning 185 days. Not quite Mars-and-back, and they did launch Progresses to it, but it already had all the important things (attitude control, ECLSS, airlock and the like), it just needs more room and supplies to handle longer missions. Taking the venerable DOS modules as a base for a deep-space ship is a really obvious move, I think... they've been flying with extraordinary success for a really long time.

Reuse of course, saves no mass, it saves cost on subsequent missions and (potentially) adds confidence in the systems. But not lugging an Orion all the way to Mars and back, for example, saves you quite a few tons, as I have just shown.

 

Rune. We are actually quite in agreement on most points, I think.

 

Yeah, but Salyut is too small without additional modules to hold the extra supplies required. Also, we now have 4 crew members in it instead of three if we use a modified version to go to Mars, making it even more crowded

And the Transhab is actually way less than 40T. http://spaceflight.nasa.gov/history/station/transhab/

The mass pretty much depends on how much you put in there- if you want the astronauts to do science experiments on the way to and fro Mars, then you likely need ~40T (even more if you add teleoperating probes or something like the Apollo Telescope Mount). If not, it whould be closer to ~20T. Here's a calculator:http://www.projectrho.com/public_html/rocket/transhabcalc.php

IVF is technically less complex and massive- it's more that the tech hasn't matured enough yet (it also improves overall performance) Lunar missions will only need 15 days of life support (until you start building bases, then you should probably land landers in shadowed craters), but this may increase if you have the LV capacity (like SLS has) and want to do experiments through the Lunar night. But you're definitely not going to keep the lander around in space/the moon for more than a month and a half. Boil off is much less of a problem in such short durations. RP-1 does not boil off at all, and O2 does relatively slowly, making this problem even less problematic. Mars is a different story, or course.

You had stated in your previous post single stage landers were more efficient, and NOW you're backtracking? Either way, though single stage landers allow for more cargo, the 10T higher mass of a lander is pretty darn huge. That's literally the difference between upgrading from Block I to Block IB. A bad idea in my book, and a mass that should definitely be used for more experiments and surface duration instead of commonality with a farther-out lunar base that may not happen (or will not happen for a long time)

Edited March 21, 2016 by fredinno

[Rune]

21 minutes ago, fredinno said:

Yeah, but Salyut is too small without additional modules to hold the extra supplies required. Also, we now have 4 crew members in it instead of three if we use a modified version to go to Mars, making it even more crowded

And the Transhab is actually way less than 40T. http://spaceflight.nasa.gov/history/station/transhab/

The mass pretty much depends on how much you put in there- if you want the astronauts to do science experiments on the way to and fro Mars, then you likely need ~40T (even more if you add teleoperating probes or something like the Apollo Telescope Mount). If not, it whould be closer to ~20T. Here's a calculator:http://www.projectrho.com/public_html/rocket/transhabcalc.php

So somewhere between 20 and 40mT is a good estimate for a long duration habitat. Good! Then all the estimates I have done about that are correct, in your own words.

21 minutes ago, fredinno said:

IVF is technically less complex and massive- it's more that the tech hasn't matured enough yet (it also improves overall performance) Lunar missions will only need 15 days of life support (until you start building bases, then you should probably land landers in shadowed craters), but this may increase if you have the LV capacity (like SLS has) and want to do experiments through the Lunar night. But you're definitely not going to keep the lander around in space/the moon for more than a month and a half. Boil off is much less of a problem in such short durations. RP-1 does not boil off at all, and O2 does relatively slowly, making this problem even less problematic. Mars is a different story, or course.

If you wait between launches of SLS (as you would in a dual launch architecture), an ACES stage would have to endure at the very least a multi-month period of inactivity, perhaps a multi-year one between missions. Without serious thermal protection, fluid loss would be enough to deplete a very appreciable fraction of the fuel it held. And in whatever case, a storable propellant stage would be less risky and faster to develop, as well as having way less possible failure points. Since those were the points I was making, not that we shouldn't develop ACES or that ACES is technically flawed in any way, I don't think you are telling me anything I didn't already know (and stated) here. BTW, as far as I know, IFV is only being developed for H2/LOX stages, by ULA/Boeing, for ACES, so it has little to no applicability to hypothetical methane stages using hypothetical engines that don't exist. Also, don't forget that my comments in this thread mostly stem from Nibb's comments about reusable lunar landers!

21 minutes ago, fredinno said:

You had stated in your previous post single stage landers were more efficient, and NOW you're backtracking? Either way, though single stage landers allow for more cargo, the 10T higher mass of a lander is pretty darn huge. That's literally the difference between upgrading from Block I to Block IB. A bad idea in my book, and a mass that should definitely be used for more experiments and surface duration instead of commonality with a farther-out lunar base that may not happen (or will not happen for a long time)

On 20/3/2016 at 11:54 AM, Rune said:

So the lander is a single stage? Then what propulsion were you thinking about? Because I can tell you right now, H2/LOX won't do, on account of the propellant transfer and surface boiloff. So you are probably looking at storables, and thus a pretty massive Mass ratio, so a lot of lander fuel to be moved around.

And how do you take the fuel to the lander, exactly, and insert that into EML? What does the stationkeeping between missions for the lander? What re-certifies the lander after each flight, and how many flights can it be used? And what, exactly, handles the fuel transfer? Equipment on the lander that you take to the surface and back, or you envision some kind of tanker bolted to an EUS?

Also, note that an expendable two-stage lander can be used for unmanned cargo flights, landing big items on the surface for, say, extended stays or the construction of lunar bases. I don't know how payload delivery would work in your scenario...

 

Rune. I'd stick to the expendable for the first missions, at least.

That is my second post in this thread, in the first one I hadn't mentioned anything about stages, just propellants, because I thought it was pretty much a given that an expendable lander would be as stage-optimized as possible. Not only that, I actually calculated later that using two stages is slightly more mass efficient for the mission I was considering. Don't put words in my mouth, please.

 

Rune. I think something is being lost in translation here.

[fredinno]

9 minutes ago, Rune said:

So somewhere between 20 and 40mT is a good estimate for a long duration habitat. Good! Then all the estimates I have done about that are correct, in your own words.

If you wait between launches of SLS (as you would in a dual launch architecture), an ACES stage would have to endure at the very least a multi-month period of inactivity, perhaps a multi-year one between missions. Without serious thermal protection, fluid loss would be enough to deplete a very appreciable fraction of the fuel it held. And in whatever case, a storable propellant stage would be less risky and faster to develop, as well as having way less possible failure points. Since those were the points I was making, not that we shouldn't develop ACES or that ACES is technically flawed in any way, I don't think you are telling me anything I didn't already know (and stated) here. BTW, as far as I know, IFV is only being developed for H2/LOX stages, by ULA/Boeing, for ACES, so it has little to no applicability to hypothetical methane stages using hypothetical engines that don't exist. Also, don't forget that my comments in this thread mostly stem from Nibb's comments about reusable lunar landers!

 

That is my second post in this thread, in the first one I hadn't mentioned anything about stages, just propellants, because I thought it was pretty much a given that an expendable lander would be as stage-optimized as possible. Not only that, I actually calculated later that using two stages is slightly more mass efficient for the mission I was considering. Don't put words in my mouth, please.

 

Rune. I think something is being lost in translation here.

A HAB needs 35T, per the calculator, and a Lab, 45T. So you were closer.

I was using RP-1, not CH4, as a baseline. H2/O2 is possible, and has advantages (it's actually less massive overall), but doing that also means you probably would waive off reuse (if you can't mitigate hydrogen embrittlement enough. However, this has experience via New Shepard and the DC-X, so maybe H2 O2 is better?) I'm worried about using H2 due to boil off concerns due to the 30+ day missions I would like a SLS lunar lander to have (the LV capacity is more than enough)

I wanted a RP-1 engine (existing) using IVF, modified for use with that fuel- since RP-1 is storable, only the O2 needs to have IVF hooked up to it to catch gases to, making the system simpler overall. Since it would be the H2 IVF with the fuel boil off capture removed, little modification to the existing IVF system is necessary.

Again, reuse is pointless if you only have max. 3 missions a year.

 

OK, sorry on that one.

[Rune]

On 21/3/2016 at 11:27 PM, fredinno said:

I was using RP-1, not CH4, as a baseline. H2/O2 is possible, and has advantages (it's actually less massive overall), but doing that also means you probably would waive off reuse (if you can't mitigate hydrogen embrittlement enough. However, this has experience via New Shepard and the DC-X, so maybe H2 O2 is better?) I'm worried about using H2 due to boil off concerns due to the 30+ day missions I would like a SLS lunar lander to have (the LV capacity is more than enough)

I wanted a RP-1 engine (existing) using IVF, modified for use with that fuel- since RP-1 is storable, only the O2 needs to have IVF hooked up to it to catch gases to, making the system simpler overall. Since it would be the H2 IVF with the fuel boil off capture removed, little modification to the existing IVF system is necessary.

Hum, IFV. I think you have an incomplete understanding of the technology, I'm afraid. RP-1/LOX could never use IFV. For a variety of reasons, the most paradoxical of all being that you wouldn't be able to pressurize the RP-1. Or start the turbopumps.

See, IFV works great because it uses H2/LOX. Let's explore why a bit. In any H2/LOX stage, as the tanks heat, they produce gaseous H2 and O2. Those are great pressurizers in their own tanks (obviously you wouldn't pressurize the O2 tanks with hydrogen, or vice-versa, unless you are feeling very suicidal), so much so that they would make the tanks blow up if left unchecked. To handle that, you usually have relief valves, plus helium bottles to kick-start the flow to the turbopumps when you want to fire. With RP1/LOX, it's pretty much the same, only instead of keeping the RP-1 tank cold and venting, you have to keep heating it to prevent the fuel from freezing (and stirring, too), plus provide the pressure to move it around with the usual He bottles.

Now, IFV. IFV is amazing, frankly, because it uses all that vented propellant to do other stuff. First, it runs the gaseous hydrogen and oxygen through a combustion engine, getting energy and pressure to run everything else. That is, like, the key, energy and pressure (by heating more of the gaseous losses and using the combustion products). Because most things in a rocket are started with pressure, starting with the pumps. Also, it grabs all the excess gaseous propellants, and it uses tiny thrusters to do station keeping with them. Still wondering how they get tiny RCS to work on H2/LOX, even in their gaseous form, BTW. Must spend a ton of juice on electric igniters and be one clunky thruster. Awesome Isp for RCS, though.

And those are the reasons why it can't work on RP-1, basically. No gaseous fuel to pressurize the tank, no easy way to pump it into an engine, completely different thermal requirements, and you can't get rid of the He lines. IFV is meant to be used on cryogenic fuels, for good reasons. Might be able to be adapted to methane, might be methane poses some issue I'm unaware of (carbon deposits?). But even there, it doesn't get rid of the boiloff, it just uses it to do other stuff. ACES endurance is advertised as weeks, and that is because ACES is very well insulated for an expendable stage, and they say nothing about how much dV you lose if you wait. Fuel depots running on this technology could work, yes (the more thermal protection, the less you lose), but they would still waste a percentage of whatever you put in there.

And of course, all the points about mechanical complexity still stand. There are orders of magnitude of difference here. For example, Hydrazine/N2O4 tanks are thermally stable in a passive way, with zero losses for decades, just sitting there with their ridiculously low weight fraction. Their only design requirement is to handle pressure and be painted with the right thermal coating!

 

Rune. KSP has taught me to respect low Isps, Ignition! made me fall in love with Hydrazine.

Edited March 23, 2016 by Rune

[fredinno]

3 hours ago, Rune said:

Hum, IFV. I think you have an incomplete understanding of the technology, I'm afraid. RP-1/LOX could never use IFV. For a variety of reasons, the most paradoxical of all being that you wouldn't be able to pressurize the RP-1. Or start the turbopumps.

See, IFV works great because it uses H2/LOX. Let's explore why a bit. In any H2/LOX stage, as the tanks heat, they produce gaseous H2 and O2. Those are great pressurizers in their own tanks (obviously you wouldn't pressurize the O2 tanks with hydrogen, or vice-versa, unless you are feeling very suicidal), so much so that they would make the tanks blow up if left unchecked. To handle that, you usually have relief valves, plus helium bottles to kick-start the flow to the turbopumps when you want to fire. With RP1/LOX, it's pretty much the same, only instead of keeping the RP-1 tank cold and venting, you have to keep heating it to prevent the fuel from freezing (and stirring, too), plus provide the pressure to move it around with the usual He bottles.

Now, IFV. IFV is amazing, frankly, because it uses all that vented propellant to do other stuff. First, it runs the gaseous hydrogen and oxygen through a combustion engine, getting energy and pressure to run everything else. That is, like, the key, energy and pressure (by heating more of the gaseous losses and using the combustion products). Because most things in a rocket are started with pressure, starting with the pumps. Also, it grabs all the excess gaseous propellants, and it uses tiny thrusters to do station keeping with them. Still wondering how they get tiny RCS to work on H2/LOX, even in their gaseous form, BTW. Must spend a ton of juice on electric igniters and be one clunky thruster. Awesome Isp for RCS, though.

And those are the reasons why it can't work on RP-1, basically. No gaseous fuel to pressurize the tank, no easy way to pump it into an engine, completely different thermal requirements, and you can't get rid of the He lines. IFV is meant to be used on cryogenic fuels, for good reasons. Might be able to be adapted to methane, might be methane poses some issue I'm unaware of (carbon deposits?). But even there, it doesn't get rid of the boiloff, it just uses it to do other stuff. ACES endurance is advertised as weeks, and that is because ACES is very well insulated for an expendable stage, and they say nothing about how much dV you lose if you wait. Fuel depots running on this technology could work, yes (the more thermal protection, the less you lose), but they would still waste a percentage of whatever you put in there.

And of course, all the points about mechanical complexity still stand. There are orders of magnitude of difference here. For example, Hydrazine/N2O4 tanks are thermally stable in a passive way, with zero losses for decades, just sitting there with their ridiculously low weight fraction. Their only design requirement is to handle pressure and be painted with the right thermal coating!

 

Rune. KSP has taught me to respect low Isps, Ignition! made me fall in love with Hydrazine.

Then I guess H2 Lox is the way to go. Only problem is long-term boil-off for bases (or in Earth orbit storage- but you should carry the lander and Orion on the last launch together to minimize boil-off), but that's a problem for all cryogens. It also likely waives off reuse and increases lander cost, but does that really matter so much for 2 launches a year?

You know what else is orders of magnitude higher with hypergol? Fuel mass. I'd agree with you if we were talking about Mars landers, but not far shorter duration Lunar Landers. http://www.sei.aero/eng/papers/uploads/archive/AIAA-2013-5479_Presentation.pdf

Again, referring to that lander report, back on page 16, landers with H2O2 propellants are ~15 T lighter than hypergol equivalents. You need a LOT of boil off to mitigate that performance advantage.