It only has an isp of about 312 s?
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.
...Um, that sounds too complex.
Quote from: Robotbeat on 08/29/2013 04:28 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.
From my calculations you will need 158 metric tons of storable propellant to make this work from LEO.
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.
Quote from: Downix on 08/29/2013 05:30 amFrom my calculations you will need 158 metric tons of storable propellant to make this work from LEO.Please show your work.. I did.
Quote5 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.
I checked yours. You neglected the weight of the lander itself.
Quote from: QuantumGOnly in pixie land is that the cost of an SLS launch.That's the cost on-record.
Only in pixie land is that the cost of an SLS launch.
Sorry to hurt your bash-SLS fest.
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.
Quote from: QuantumG on 08/29/2013 01:49 amAccording 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...
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.
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.
Quote from: Downix on 08/29/2013 05:51 amI checked yours. You neglected the weight of the lander itself. No I didn't.Dry mass: 10458Payload mass: 43000Total mass at burnout: 53458Isp: 312Delta-v: 3107Apply the rocket equation, you get: 147579 in LEO.Subtract the payload, you get: 104579Subtract the dry mass, you get: 94121
QuoteQuote from: QuantumGOnly 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 from: Warren Platts on 08/29/2013 02:39 amIt only has an isp of about 312 s? For the sake of argument.
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.
QuoteReason 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 ?
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.
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?
Quote from: Ben the Space Brit on 08/29/2013 09:48 pmYou 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.
Quote from: mmeijeri on 08/29/2013 09:50 pmQuote from: Ben the Space Brit on 08/29/2013 09:48 pmYou 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.
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.
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?
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.
Quote from: QuantumG on 08/29/2013 01:49 amAccording 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?
Large commercial satellites are launched with tons of hypergolic propellants, especially large military birds. I don't see this as a serious concern.
Quote from: mmeijeri on 08/29/2013 09:43 pmUnless 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?
If you think toxicity is an issue, you could use kerosene / peroxide too. Or even peroxide as a monopropellant.
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.
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?
Quote from: joek on 08/29/2013 11:15 pmTrue, 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.
Quote from: mmeijeri on 08/29/2013 11:19 pmQuote from: joek on 08/29/2013 11:15 pmTrue, 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).
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.
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.
Quote from: Robotbeat on 08/29/2013 11:54 pmAlso, 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.
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.
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?
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...
SuperDraco probably has very low Isp.
Quote from: mmeijeri on 08/30/2013 01:00 amSuperDraco probably has very low Isp....and would have heavy tanks.
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.
Quote from: QuantumG on 08/30/2013 01:08 amBecause 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?
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
Quote from: Ben the Space Brit on 08/29/2013 09:52 pmDifferent 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.
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.
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.
Quote from: RocketmanUS on 08/30/2013 11:11 pmHypergolics 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.
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.
Quote from: QuantumG on 08/29/2013 02:54 amQuote from: Warren Platts on 08/29/2013 02:39 amIt 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.
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.
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.
{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...
Quote from: Robotbeat on 08/31/2013 11:51 pm{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.
Falcon Heavy is still on the drawing board although it does share components with the currently produced F9 v1.1.
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.ShuttleProtonSoyuz-FGSoyuz-UAriane 5H-IIBFalcon 9Antares 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.
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.
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.
And to a large extent, the low launch rate of Delta IV Heavy is due to the low demand.
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.
Quote from: Patchouli on 09/01/2013 03:41 amOne 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.
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.
Quote from: QuantumG on 08/29/2013 01:49 amAccording 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?
Quote from: mmeijeri on 09/01/2013 11:21 amQuote from: Patchouli on 09/01/2013 03:41 amOne 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.
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.
Quote from: mmeijeri on 09/04/2013 04:39 pmGoing 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.
Quote from: clongton on 09/06/2013 01:10 amQuote from: mmeijeri on 09/04/2013 04:39 pmGoing 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.
Quote from: Robotbeat on 09/06/2013 01:18 amQuote from: clongton on 09/06/2013 01:10 amQuote from: mmeijeri on 09/04/2013 04:39 pmGoing 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.
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.
Quote from: newpylong on 09/04/2013 04:19 pmQuote from: mmeijeri on 09/01/2013 11:21 amQuote from: Patchouli on 09/01/2013 03:41 amOne 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.pdfAnd 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).
Quote from: Robotbeat on 09/04/2013 04:31 pmQuote from: newpylong on 09/04/2013 04:19 pmQuote from: mmeijeri on 09/01/2013 11:21 amQuote from: Patchouli on 09/01/2013 03:41 amOne 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.pdfAnd 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.
Quote from: muomega0 on 09/06/2013 01:22 pmQuote from: Robotbeat on 09/04/2013 04:31 pmQuote from: newpylong on 09/04/2013 04:19 pmQuote from: mmeijeri on 09/01/2013 11:21 amQuote from: Patchouli on 09/01/2013 03:41 amOne 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.pdfAnd 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.
...40 to 50 kW power bus with EP-tanks is quite a high IMLEO....
Quote from: muomega0 on 09/06/2013 05:06 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.
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.
Solar arrays can do about 150kW/kg
Quote from: Robotbeat on 09/06/2013 05:31 pmSolar arrays can do about 150kW/kgTypo! (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
Why would one need anything wider?
Quote from: mmeijeri on 09/01/2013 11:21 amQuote from: Patchouli on 09/01/2013 03:41 amOne 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?
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
Partly for habs as 4.5M is much too narrow and cramped but mostly for Mars EDL systems.
Quote from: gbaikie on 09/13/2013 05:40 amSLS is costing about 2 and half billion per year in order to developThat sounds more like the annual cost of SLS and Orion together.
So that further strengthens QuantumG's point (as if it needed any more strengthening).
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?
I suppose that would be the 324 second Isp on the German Aestus II engine.
Quote from: M129K on 09/16/2013 06:47 pmI 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?
Quote from: a_langwichWhat's the highest ISP available from non-toxic sources?RL-10B2 with 461.5 seconds, soon to be overtaken by Vinci at 465 seconds
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.
Quote from: M129K on 09/17/2013 06:47 pmEh, 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.
Quote from: Robotbeat on 09/17/2013 08:22 pmQuote from: M129K on 09/17/2013 06:47 pmEh, 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.
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.
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.
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.
Recently there has been discussion in other threads of proposals. Dalhousie wrote:QuoteIt 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.
No mater what fuel is use, it seems like you must have lox
Quote from: spacenut on 06/16/2014 09:50 pmNo mater what fuel is use, it seems like you must have loxNot 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
Its also liquid at "room-temperature" and indefinatly storable as long as its "cooled" below room temperature. ..
Out of curiosity: What is the liquid temperature range of MMH? of NTO?
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.
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!
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.
..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 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!
Quote from: sdsds on 06/22/2014 02:37 amI 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?
Quote from: sdsds on 06/22/2014 02:37 amthere 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?
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!
It's about the upper stage recovery of a "fully recovered F9R."
Quote from: RanulfC on 06/18/2014 09:17 pmIts 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 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).
Quote from: Proponent on 06/26/2014 07:40 pmPeroxide 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....
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.