propellant boil-off rates well under 0.05% per day can be achieved in LEO.
Quote from: QuantumG on 08/23/2012 06:05 ampropellant boil-off rates well under 0.05% per day can be achieved in LEO.That's 17% per year. Utilization flights would still need to be quite timely.(edit: corrected my maths!)
A new paper "Thermal Optimization of an On-Orbit Long Duration Cryogenic Propellant Depot" co-authored by Bernard Kutter of ULA is attached.They conclude that propellant boil-off rates well under 0.05% per day can be achieved in LEO.
Quote from: mikes on 08/23/2012 06:18 amQuote from: QuantumG on 08/23/2012 06:05 ampropellant boil-off rates well under 0.05% per day can be achieved in LEO.That's 17% per year. Utilization flights would still need to be quite timely.(edit: corrected my maths!)I get 18% per year (perchance your original, uncorrected figure?). Obviously that's not very different, but the important conceptual difference is that I assume a constant absolute boil-off rate (kilograms per day) rather than a constant relative rate (% per day). Heat flux into the tanks should be approximately independent of the remaining propellant load, hence I believe a constant absolute loss rate is applicable.
I wonder how long the campaign has to be before storable propellant wins...This will require me to do math, but before I do: good question?
Don't bother.The conclusions (*cough*) are based on the lie of the average, a very dubious use of the term "propellant", and an implication that the storage life of the "propellant" is related to the average boil-off rates of the propellant's components.
Also note also that the average boil-off rates assume an LO2:LH2 mass ratio of 11:1 (55mT:5mT), which seems a bit odd.
Never mind that the LH2 will boil off over 5-20x faster than the LO2 (based on their numbers), and thus the "propellant depot" will rapidly become an LO2 depot.
Quote from: mikes on 08/23/2012 06:18 amQuote from: QuantumG on 08/23/2012 06:05 ampropellant boil-off rates well under 0.05% per day can be achieved in LEO.That's 17% per year. Utilization flights would still need to be quite timely.(edit: corrected my maths!)I get 18% per year (perchance your original, uncorrected figure?). Obviously that's not very different, but the important conceptual difference is that I assume a constant absolute boil-off rate (kilograms per day) rather than a constant relative rate (% per day). Heat flux into the tanks should be approximately independent of the remaining propellant load, hence I believe a constant absolute loss rate is applicable.Even with a loss rate of 18% per year, the timing of depot-dependent missions could be rather leisurely if the depot is over-sized a bit.
Quote from: QuantumG on 08/23/2012 06:05 amA new paper "Thermal Optimization of an On-Orbit Long Duration Cryogenic Propellant Depot" co-authored by Bernard Kutter of ULA is attached.They conclude that propellant boil-off rates well under 0.05% per day can be achieved in LEO.What about helium for pressurization ? Never really mentioned.
Quote from: Proponent on 08/23/2012 08:23 amQuote from: mikes on 08/23/2012 06:18 amQuote from: QuantumG on 08/23/2012 06:05 ampropellant boil-off rates well under 0.05% per day can be achieved in LEO.That's 17% per year. Utilization flights would still need to be quite timely.(edit: corrected my maths!)I get 18% per year (perchance your original, uncorrected figure?). Obviously that's not very different, but the important conceptual difference is that I assume a constant absolute boil-off rate (kilograms per day) rather than a constant relative rate (% per day). Heat flux into the tanks should be approximately independent of the remaining propellant load, hence I believe a constant absolute loss rate is applicable.Yes, it is nominally a constant absolute rate. Per Equation 1 (pg 7), boil-off is dependent on Lx2 mass, and heat flux is constant for a given Beta and Theta. Thus boil-off rate % increases linearly with decreasing mass (all other factors constant).I ran the numbers using their values for several scenarios, and I'm off by 10x in all cases; not sure where I went wrong, so I'd appreciate a check. E.g., using Eq 1, values in Table 4, Beta 20, and Theta -10 (their avearage boil-off minima) I get boil-off of 0.008% LH2 and 0.001% LO2, average of 0.005%, not 0.05%.In any case, the "average" boil-off rate appears to be the rub. LH2 boil-off is significantly higher than LO2, and appears to dominate in all cases. Which on closer examination isn't surprising (or a cursory examination--compare Fig 5 and Fig 7). Also note also that the average boil-off rates assume an LO2:LH2 mass ratio of 11:1 (55mT:5mT), which seems a bit odd.Presenting the "average" boil-off in this manner and with those numbers seems disingeneous and of questionable utility. What am I missing?
At 17% a year you would loose only about 9.35 tons of mostly hydrogen. Given a single rocket could lift 20+ tons at a time that isn't a big problem.
I wonder how long the campaign has to be before storable propellant wins...
Quote from: joek on 08/23/2012 06:38 pmDon't bother.I take their word (analysis) before yours
Don't bother.
the attached figure illustrates the source of my question [...] if you have a "propellant depot" where one component of the propellant (LH2) is boiling off much faster than another component (LO2), is "total propellant boil-off" a useful metric or figure of merit?
Is this 0.05% going to be hard to improve without a completely new architecture, or could it just be augmented by a second system for reliquifying the boil off?
How likely are we to see this actually flying? So much of the things I was enthusiastic about have been slashed.
It's not quantity but payload requirements of size or delta-v that will bring depots about. Depots at LEO enable GEO sats to be larger than what will be acomodated by existing LV's as well as higher delta-v for BEO missions with heavier payloads.
ULA could be its own customer. GEO missions using a Centaur that can dock and refuel partially increasing the GEO payload capability has an economic savings impact for GEO missions for ULA by using a less number of SRB's to do a larger payload.
Quote from: KelvinZero on 08/26/2012 01:27 pmHow likely are we to see this actually flying? So much of the things I was enthusiastic about have been slashed.We'll see depots when flight rates are large enough to warrant them.
I was really just meaning a demonstrator. Is it still on the cards?
I don't think a demonstrator will be on the cards unless a customer pays for it.
The customer used to be NASA, back when it was still funding technology development (In-orbit propellant transfer and storage was mentioned in Obama's original FY2011 budget under the Flagship Demonstration Program but I realize most of that was slashed. I think I recall some sort of small non-threatening propellant transfer demonstration surviving. )
Quote from: joek on 08/26/2012 02:23 amthe attached figure illustrates the source of my question [...] if you have a "propellant depot" where one component of the propellant (LH2) is boiling off much faster than another component (LO2), is "total propellant boil-off" a useful metric or figure of merit?Great work; thanks for sharing this!Yes, it looks like in this design, as in others, the depot would mainly lose LH2. Perhaps the assumption is that the spacecraft making a withdrawal from the depot will have been launched with excess LH2.
Quote from: KelvinZero on 08/26/2012 12:35 pmIs this 0.05% going to be hard to improve without a completely new architecture, or could it just be augmented by a second system for reliquifying the boil off?Earlier papers illustrated how you could avoid boil-off altogether using cryocoolers. I had thought LH2 cryocoolers would be too large and impractical, but apparently that is not the case. And such cryocoolers have flown in space before, so I don't really understand the focus on passive systems only, but they must have good reasons for it.And bear in mind that 0.05% in LEO is absolutely fine, you don't need any more than that, in fact you could make do with much more boil-off than that.
Quote from: oldAtlas_Eguy on 08/26/2012 04:37 pmULA could be its own customer. GEO missions using a Centaur that can dock and refuel partially increasing the GEO payload capability has an economic savings impact for GEO missions for ULA by using a less number of SRB's to do a larger payload.I see this point, but you seem to be comparing only within the Atlas V family. Could there really be a savings getting a mission to GEO using two Atlas V launches as compared to a single Delta IV-Heavy launch?
(As the Air Force plans to demonstrate starting around 2025... Lockheed has a contract for a subscale demo of the reusable flyback booster for the EELV successor.)
Quote from: sdsds on 08/26/2012 02:44 amQuote from: joek on 08/26/2012 02:23 amthe attached figure illustrates the source of my question [...] if you have a "propellant depot" where one component of the propellant (LH2) is boiling off much faster than another component (LO2), is "total propellant boil-off" a useful metric or figure of merit?Great work; thanks for sharing this!Yes, it looks like in this design, as in others, the depot would mainly lose LH2. Perhaps the assumption is that the spacecraft making a withdrawal from the depot will have been launched with excess LH2.Bernard has made the point several times in the past that it was better to use the boiled off LH2 to suppress all LOX boiloff and then provide for stationkeeping propulsion gas. So long as the total boiloff rate is less than the stationkeeping propellant consumption rate would have been, it doesn't matter. It just means that instead of shipping up RCS propellants (say hydrazine) you're shipping up a potentially lighter mass of LH2.Also, in answer to someone else's question about why the 11:1 O/F ratio, this was due to using the existing Centaur tankage design and fitting it within an existing fairing. Bernard and I discussed in a different paper the idea of doing the depot LH2 tank using a larger tank diameter (possibly integral with the PLF), which would get you much closer to the desired diameter. Ie this was driven by what they could do quickly with off-the-shelf tools (and for tankage that they had good solid thermal data on), not so much what could you do if cost was no object and maximum performance was.~Jon
Now for a tanker vehicle, or a deep-space propulsion stage, lower or zero boiloff might be better.
Quote from: Robotbeat on 08/27/2012 06:27 pm(As the Air Force plans to demonstrate starting around 2025... Lockheed has a contract for a subscale demo of the reusable flyback booster for the EELV successor.)Don't hold your breath. The phase II of RBS, that was supposed to come from a down-select from 3 contractors down to one, has lost all its funding. That puts RBS into an indefinite hold, and puts any date like 2025 in real jeopardy. (This could be the topic for a whole other thread, but I've already said all I can/will say about it.)
1) You will need to keep the depot and Centaur completely inside the faring because of the sunshields.
2) In order to make the ~14' or 4.33m diameter stainless steel tank you will need new tooling, but ULA's ACES plans does include being able to make larger diameter tanks so this effort would be common to use as a larger diameter US ACES.
3)A 4.33M diameter tank would double the volume for the LH2 tank dropping the ratio to ~5.5:1.
Quote from: jongoff on 08/27/2012 05:56 pmNow for a tanker vehicle, or a deep-space propulsion stage, lower or zero boiloff might be better.Then again, if the deep space stage is actually in deep space, boil-off would likely be considerably lower.
Bernard has made the point several times in the past that it was better to use the boiled off LH2 to suppress all LOX boiloff and then provide for stationkeeping propulsion gas. So long as the total boiloff rate is less than the stationkeeping propellant consumption rate would have been, it doesn't matter. It just means that instead of shipping up RCS propellants (say hydrazine) you're shipping up a potentially lighter mass of LH2.
What's unclear to me in the latest paper is where a visiting spacecraft is supposed to go. In previous concepts they had a central module that does not appear to be present here.
There's cheaper ways to do station keeping... especially in LEO where there are zero-propellant magnetic despin options.
Are you concerned that the presence of a docking/transfer module would significantly change the boil off rate?
Quote from: QuantumG on 08/27/2012 09:54 pmThere's cheaper ways to do station keeping... especially in LEO where there are zero-propellant magnetic despin options.There's attitude control and there's station keeping. Magnetorquers can readily control attitude about two axes, or even three, though that's a little tricker. About any axis, though, the torque delivered is going to be small, absent a large power supply. Small torques are certainly adequate for maintaining attitude on the whole, but I doubt they'd be enough to deal with transients that would arise on docking and undocking.Then there's station keeping....
Quote from: jongoff on 08/27/2012 05:52 pmBernard has made the point several times in the past that it was better to use the boiled off LH2 to suppress all LOX boiloff and then provide for stationkeeping propulsion gas. So long as the total boiloff rate is less than the stationkeeping propellant consumption rate would have been, it doesn't matter. It just means that instead of shipping up RCS propellants (say hydrazine) you're shipping up a potentially lighter mass of LH2.I don't think this is actually true. While having LH2 boil off is certainly more efficient than having LOX boil off, I don't think the station keeping is a net win. It's more like making the best of a bad situation, and it appears to require going to very low altitudes.
What do you mean it "requires going to very lo altitudes"? Do you mean that it's only at very low altitudes where stationkeeping requirements are higher than boiloff?
Mind you, there's also very good orbital dynamics reasons for wanting the depot to be at lower altitudes. The apogee raising maneuver needed to rendezvous with a higher altitude depot is effectively wasted propellant, and can impose a decent mass hit on depot tanker deliveries and visiting vehicle deliveries.
If you can go with a lower altitude depot (which reduces the orbital dynamics penalty for using it), and it enables you to go with a simpler, non-ZBO design, that sounds win-win to me.
That said, in LEO there is the potential of using Electryodynamic tethers for stationkeeping, and if you had that, then yeah, going to a zero boiloff system might make a lot more sense. The problem is that ED tethers are still in the TRL valley of death. If they were an operationally proven technology, it might be a more interesting question.
IVF.http://www.ulalaunch.com/site/docs/publications/Integrated%20Vehicle%20Propulsion%20and%20Power%20System%20for%20Long%20Duration%20Cyrogenic%20Spaceflight%202011.pdf~Jon
To be honest, I'm not entirely sure what their assumptions were about reboost propulsion needs, other propulsion needs, Isp of the warmed GH2 thrusters, etc. I know that depots tend to have fairly low ballistic coefficients, particularly LOX/LH2 depots with sunshades, so they may experience more drag than say the ISS. But I haven't run the numbers well enough to see what altitude a ULA-style depot optimizes out to.
There wasn't any consideration for VV (visiting vehicles) radiative environment from active RCS. So if the depot is busy then the boil off rate will be higher. How much is unknown since it wasn't part of the study.
Rocket exhaust won't be nearly as hot as 6000 K. Even in the combustion chamber, temperatures that high would be extremely problematic, and by the time the exhaust has been fully expanded through the nozzle, it will be much cooler than in the combustion chamber. In addition, the exhaust won't be optically thick, and it will be only very briefly present.
"A Study of CPS Stages for Missions beyond LEO"http://spirit.as.utexas.edu/~fiso/telecon/King_2-15-12/
"A Study of CPS Stages for Missions beyond LEO"
Yeah, for a space RCS nozzle, my guess would be closer to 500-1000K tops. For hydrazine monoprop thrusters, it could potentially be even lower (they start off relatively cool to start with, and you try to expand the crap out of the exhaust to get a good Isp.
Of course, in the not-to-distant future there may be the alternative of using a Sticky Boom...
Quote from: jongoff on 09/03/2012 03:47 pmOf course, in the not-to-distant future there may be the alternative of using a Sticky Boom...Yes, that would eliminate need for burns during the final stages of rendezvous. Looks like Stickey Boom has been getting increased media attention lately.
Quote from: Ronsmytheiii on 09/04/2012 11:26 pm"A Study of CPS Stages for Missions beyond LEO"Interesting. For me the big takeaway point is driven by the parametric sensitivity analysis that shows the extreme importance of the propellant mass fraction for all the CPS-enabled missions they analyzed. That, plus the observation that the CPS can be less massive if it carries less propellant during launch, drives the conclusion that the CPS would ideally take on its propellant once in LEO.Presumably the tanker that comes to deliver that propellant can be built plenty-sturdy (i.e. with a poor PMF) because it never has to go beyond LEO.
Yeah, for a space RCS nozzle, my guess would be closer to 500-1000K tops. For hydrazine monoprop thrusters, it could potentially be even lower (they start off relatively cool to start with, and you try to expand the crap out of the exhaust to get a good Isp.Of course, in the not-to-distant future there may be the alternative of using a Sticky Boom...~Jon
Quote from: jongoff on 09/03/2012 03:47 pmYeah, for a space RCS nozzle, my guess would be closer to 500-1000K tops. For hydrazine monoprop thrusters, it could potentially be even lower (they start off relatively cool to start with, and you try to expand the crap out of the exhaust to get a good Isp.Of course, in the not-to-distant future there may be the alternative of using a Sticky Boom...~JonAny N2H4 mono thruster with an expansion ratio above about 40 (most of them), is going to have pretty close to room temp exhaust.
Liquid methane or liquid natural gas (LNG) has about the same boil off temperature as oxygen. The RL-10 has already been tested and can use methane. Why not use LOX and LNG? Boil off would be less than hydrogen. It can therefore be stored for much longer periods. I think that is why they were thinking about using LNG made from Martian atmosphere and water. I know hydrogen is better with ISP, but methane isn't that bad.
Take 5 minutes and compare the IMLEO required to transfer from LEO the same payload to the moon using a 320 or 345 Isp stage versus a 450 s stage. Account for boil-off, or assume zero . Do not forget stage mass also increases as the ISP is lowered (ironic huh?) in the mass fractions. The result may be that you stop asking the question.
Yet another area where depots (or at least cryo transfer) makes a ton of sense if we're going to be serious about exploration/settlement.
Have internal combustion engines ever been operated in space before? The notion of an old inline Chevy straight-6 on Centaur is amusing, but an ICE/alternator might be cheaper than a fuel cell(s) for a given power requirement.Pictures I've seen of the Delta 4's upper stage always impressed/dismayed me with the sheer clutter of hardware and cabling/tubing and I don't doubt Centaur is any less cluttered with similar systems. Anything to clean up and simplify systems has got to be A Good Thing.
I've attached an image taken from the presentation showing their nominal LEO departure stack for an NEA mission. Making each of those (3!) CPS stages sufficiently lightweight (i.e. high PMF) means loading them with propellant only once they are in orbit.
Using a LEO departure stack is a needless complication. Propellant transfer at a Lagrange point is all you need. High mass fractions and cryogens are no more than a nice to have.
High mass fractions and cryogens are very nice to have at departure, where boil-off is much less of an issue.
... but obviously controlling (or making use of) boil off is very desirable.
Quote from: kkattula on 09/07/2012 04:03 amHigh mass fractions and cryogens are very nice to have at departure, where boil-off is much less of an issue. Sure, very nice, but if you depart straight from LEO, then you start in the worst possible thermal environment, making boil-off much more of an issue. You also have a large delta-v for a single hop, which makes mass fractions more important.With a refueling point at L1/L2, things become much easier. For departure from LEO to L1/L2 little more than a Centaur or DCSS capable of staying in space for no more than a couple of days without excessive losses is needed....
Actually, I was thinking of an L1/L2 departure. Assemble the payload and hypergolic propulsion there, then send a cryo stage to do the departure burn. Saves about 30% on departure propulsion stage mass for a Mars mission.
I'm thinking of campaign, involving staging the storable propulsion and spacecraft to L2, say with 5 launches over 1 or 2 years, then a single launch of the cryo departure stage.
At that 0.05% per day boil-off rate, it's break even on whether to use cryo for Mars arrival. The last three launches would have to be within a few months of each other. Cryo's seems to be a disadvantage(for long stay missions), for Mars departure, and L1/L2 return.
If you're using boil off for station keeping, seems to me something with low molecular mass gives more bang for the buck. This might be another argument for H2.
QuoteI'm thinking of campaign, involving staging the storable propulsion and spacecraft to L2, say with 5 launches over 1 or 2 years, then a single launch of the cryo departure stage.Are you imagining using Atlas Phase 2? Otherwise 5 launches sounds low. You don't really need a huge EDS with Lagrange point staging, and using an HLV to launch propellant takes away much of the benefit of using depots.
Quote from: Hop_David on 09/07/2012 04:46 pmIf you're using boil off for station keeping, seems to me something with low molecular mass gives more bang for the buck. This might be another argument for H2.Well, yes and no. Using hydrogen boil-off gives higher Isp than hydrazine monopropellant,
I said station keeping by boil off. Boil off removes heat from the cryogens.
And I'm not sure hydrogen boil off would have a better ISP than hydrazine. I would expect the hydrogen boil off to be only slightly warmer than the liquid hydrogen. If the hydrogen gas is in the double digits Kelvin, it's exhaust velocity probably isn't high.
However if station keeping is achieved by boil-off, it seems to me hydrogen has the best ISP.
...And I'm not sure hydrogen boil off would have a better ISP than hydrazine. I would expect the hydrogen boil off to be only slightly warmer than the liquid hydrogen. If the hydrogen gas is in the double digits Kelvin, it's exhaust velocity probably isn't high....
The boiled off H2 could be put through a small, solar powered resistojet, for higher Isp.
Quote from: jongoff on 08/23/2012 11:55 pmIVF.http://www.ulalaunch.com/site/docs/publications/Integrated%20Vehicle%20Propulsion%20and%20Power%20System%20for%20Long%20Duration%20Cyrogenic%20Spaceflight%202011.pdf~JonThanks for the link. This is a superb paper. It's such a neat idea that it's hard to believe it's taken this long to get this far (lateral thrusters look like something that should have been retro-fitted a long time ago). ULA seem *very* determined to make it happen. As a way to *radically* improve a Centaur's flexibility (however many engines it's carrying) this looks like a real game changer. Increasing payload to the point where you could eliminate an SRB, lowering boil off by a *minimum* of 50%, *unlimited* main engine starts are all very impressive. Obviously these improvements are aimed at ULA's existing customers but I think it's biggest impact could be changing people's mindset about integrated versus separate systems.It's probably too late but it might be interesting to see what happens to a Moon architecture when you have a LH2/LO2 stage that can trade boil off for duration IE pulsed Vs continuous propellant settling, or that can do unlimited restarts Vs the current 2 (3 at a pinch) Centaur, set (I was surprised to find) by the GHe capacity.
RELEASE: 12-333NASA SELECTS SPACE LAUNCH SYSTEM ADVANCED DEVELOPMENT PROPOSALSWASHINGTON -- NASA has selected 26 proposals from academia and industry for advanced development activities for the nation's next heavy lift rocket, the Space Launch System (SLS). Proposals selected under this NASA Research Announcement (NRA) seek innovative and affordable solutions to evolve the launch vehicle from its initial configuration to its full lift capacity capable of sending humans farther into deep space than ever before.
Industry proposals selected for contract negotiations are: -- "Development of a Fluid-Structure Interaction Methodology for Predicting Engine Loads," ATA Engineering, Inc., San Diego -- "Space Launch System (SLS) Advanced Development Affordable Composite Structures," ATK Space Systems, Inc., Clearfield, Utah -- "Ball Reliable Advanced Integrated Network," Ball Aerospace & Technologies Corp., Huntsville, Ala. -- "Affordable Structural Weight Reduction for SLS Block 1A," Collier Research and Development Corp., Newport News, Va. -- "DESLA Systems Engineering and Risk Reduction for AUSEP," Exquadrum, Inc., Adelanto, Calif. -- "Space Launch System Program AUSEP LOX Flow Control Valve," MOOG, Inc. Space and Defense Group, Aurora, N.Y. -- "Affordable Upper Stage Engine Advanced Development," Northrop Grumman Systems Corp., Redondo Beach, Calif. -- "Hybrid Precision Casting for Regeneratively-Cooled Thrust Chamber Components," Orbital Technologies Corp., Madison, Wis. -- "NASA Space Launch System (SLS) Advanced Development, Affordable Upper Stage Engine Program (AUSE) Study," Pratt & Whitney Rocketdyne, Inc., Jupiter, Fla. -- "Advanced Ordnance Systems Demonstration," Reynolds Systems, Inc., Middletown, Calif. -- "Cryo-Tracker Mass Gauging System," Sierra Lobo, Inc., Freemont, Ohio -- "Efficient High-Fidelity Design and Analysis Tool for Unsteady Flow Physics in Space Propulsion Geometries," Streamline Numerics, Inc. Gainesville, Fla. -- "Robust Distributed Sensor Interface Modules (DSIM) for SLS," The Boeing Company, Huntington Beach, Calif. -- "Integrated Vehicle Fluids (IVF)," United Launch Alliance, Centennial, Colo.
Surprised this wasn't cross referencedSnip-- "Integrated Vehicle Fluids (IVF)," United Launch Alliance, Centennial, Colo.
Quote from: Ronsmytheiii on 09/25/2012 02:55 amSurprised this wasn't cross referencedQuoteSnip-- "Integrated Vehicle Fluids (IVF)," United Launch Alliance, Centennial, Colo. http://forum.nasaspaceflight.com/index.php?topic=26853.450Thanks Ron. Yeah this is a good break for ULA. I have no idea how much IRAD budget they have, but from what I've seen it's probably a lot smaller than most people here would suspect (for instance I think their yearly IRAD budget is probably smaller than Project Morpheus' yearly budget), and getting a $1-3M contract to keep pushing IVF forward, not just for Centaur, but for SLS, is definitely welcome news.~Jon
Surprised this wasn't cross referencedQuoteSnip-- "Integrated Vehicle Fluids (IVF)," United Launch Alliance, Centennial, Colo. http://forum.nasaspaceflight.com/index.php?topic=26853.450
Snip-- "Integrated Vehicle Fluids (IVF)," United Launch Alliance, Centennial, Colo.
Quote from: Ronsmytheiii on 09/25/2012 02:55 amSurprised this wasn't cross referencedSnip-- "Integrated Vehicle Fluids (IVF)," United Launch Alliance, Centennial, Colo. http://forum.nasaspaceflight.com/index.php?topic=26853.450
ULA’s work will develop a heat exchanger and cooling system to support the utilization of an integrated vehicle fluids system as an auxiliary power unit. A noteworthy consideration was the extensive experience and effort that was previously put into this system between ULA and its partner, which has been proven to be a very effective teaming arrangement. Of particular importance in this effort is the utilization of the constant boil-off of gaseous hydrogen and oxygen for power generation, while at the same time maintaining low peak pressures for the system, which greatly reduces complexity through the elimination of numerous subsystems and components. This will also reduce operating costs and greatly improve operating safety, which is always a critical area of focus when developing launch vehicle systems.
Quote from: jongoff on 09/25/2012 04:27 pmQuote from: Ronsmytheiii on 09/25/2012 02:55 amSurprised this wasn't cross referencedSnip-- "Integrated Vehicle Fluids (IVF)," United Launch Alliance, Centennial, Colo. http://forum.nasaspaceflight.com/index.php?topic=26853.450Thanks Ron. Yeah this is a good break for ULA. I have no idea how much IRAD budget they have, but from what I've seen it's probably a lot smaller than most people here would suspect (for instance I think their yearly IRAD budget is probably smaller than Project Morpheus' yearly budget), and getting a $1-3M contract to keep pushing IVF forward, not just for Centaur, but for SLS, is definitely welcome news.~JonBtw, here's the relevant part of the selection statement, posted in the SLS thread:QuoteULA’s work will develop a heat exchanger and cooling system to support the utilization of an integrated vehicle fluids system as an auxiliary power unit. A noteworthy consideration was the extensive experience and effort that was previously put into this system between ULA and its partner, which has been proven to be a very effective teaming arrangement. Of particular importance in this effort is the utilization of the constant boil-off of gaseous hydrogen and oxygen for power generation, while at the same time maintaining low peak pressures for the system, which greatly reduces complexity through the elimination of numerous subsystems and components. This will also reduce operating costs and greatly improve operating safety, which is always a critical area of focus when developing launch vehicle systems.
Quote from: Ronsmytheiii on 09/25/2012 02:55 amSurprised this wasn't cross referencedSnip-- "Integrated Vehicle Fluids (IVF)," United Launch Alliance, Centennial, Colo. http://forum.nasaspaceflight.com/index.php?topic=26853.450Thanks Ron. Yeah this is a good break for ULA. I have no idea how much IRAD budget they have, but from what I've seen it's probably a lot smaller than most people here would suspect (for instance I think their yearly IRAD budget is probably smaller than Project Morpheus' yearly budget), and getting a $1-3M contract to keep pushing IVF forward, not just for Centaur, but for SLS, is definitely welcome news.~Jon
A new paper "Thermal Optimization of an On-Orbit Long Duration Cryogenic Propellant Depot" co-authored by Bernard Kutter of ULA is attached.They conclude that propellant boil-off rates well under of 0.05% per day for LH2 can be achieved in LEO, for a certain set of conditions.
Quote from: QuantumG on 08/23/2012 06:05 amA new paper "Thermal Optimization of an On-Orbit Long Duration Cryogenic Propellant Depot" co-authored by Bernard Kutter of ULA is attached.They conclude that propellant boil-off rates well under of 0.05% per day for LH2 can be achieved in LEO, for a certain set of conditions.As mentioned earlier, many things are missing from the report, hence the modifications to the conclusion above.Recall that in earlier ULA papers, the boiloff rate for LOX (not LH2) *may* reach 0.1% (the altitude was 1300 km, about double the altitude of this report).Also recall that *with upgrades (25 layers of MLI)* to the 3m diam tank Centaur upper stage, LH2 boil-off can be reduced to 2.4%. Hence the depot with many assumptions would have to reduce the boiloff passively *50 times more* to achieve 0.05% for LH2.
Further, there are many scenarios when the 0.05%/day passively cannot not be met over a 180 day cycle for a lunar sortie fuel capacity, not the robotic mission sized tanks of this analysis with a 11:1 prop ratio. It is easy to estimate that the passive depot design loss of LH2 at 40% or more over 180 days based on the design presented as it takes time to assemble 100mT of propellant at a 6:1 ratio. Contingencies are not included in the 40% estimate.The good news from the report is that under certain conditions, the passive rate for LH2 (0.05%) is approaching the 0.01%/day or less required for economics. The IMLEO mass savings of using LH2/LO2 is substantial.IOW: the conclusion of the report has not been substantiated by any means for all of the LEO conditions, operations, and contingencies, but progress is being made on the passive design.The great news is that by simply adding few pumps and power to the LEO depot to make it ZBO for LH2, not just LO2, substantial dollars can be saved filling up LEO before departing to L2. SLS need not apply in this depot centric architecture.
Quote from: muomega0 on 09/27/2012 01:18 pmQuote from: QuantumG on 08/23/2012 06:05 amA new paper "Thermal Optimization of an On-Orbit Long Duration Cryogenic Propellant Depot" co-authored by Bernard Kutter of ULA is attached.They conclude that propellant boil-off rates well under of 0.05% per day for LH2 can be achieved in LEO, for a certain set of conditions.As mentioned earlier, many things are missing from the report, hence the modifications to the conclusion above.Recall that in earlier ULA papers, the boiloff rate for LOX (not LH2) *may* reach 0.1% (the altitude was 1300 km, about double the altitude of this report).Also recall that *with upgrades (25 layers of MLI)* to the 3m diam tank Centaur upper stage, LH2 boil-off can be reduced to 2.4%. Hence the depot with many assumptions would have to reduce the boiloff passively *50 times more* to achieve 0.05% for LH2.What was the sunshield on that configuration (can you link it, or give the paper's title)? I would take this paper at face value, and assume it supersedes earlier work.
The trade-off for putting the depot at 1300 km is that tankers deliver less prop per launch, and the recipient has to reduce it's dry mass to reach where it can be refuelled.
Quote from: muomega0 on 09/27/2012 01:18 pmFurther, there are many scenarios when the 0.05%/day passively cannot not be met over a 180 day cycle for a lunar sortie fuel capacity, not the robotic mission sized tanks of this analysis with a 11:1 prop ratio. It is easy to estimate that the passive depot design loss of LH2 at 40% or more over 180 days based on the design presented as it takes time to assemble 100mT of propellant at a 6:1 ratio. Contingencies are not included in the 40% estimate.The good news from the report is that under certain conditions, the passive rate for LH2 (0.05%) is approaching the 0.01%/day or less required for economics. The IMLEO mass savings of using LH2/LO2 is substantial.IOW: the conclusion of the report has not been substantiated by any means for all of the LEO conditions, operations, and contingencies, but progress is being made on the passive design.The great news is that by simply adding few pumps and power to the LEO depot to make it ZBO for LH2, not just LO2, substantial dollars can be saved filling up LEO before departing to L2. SLS need not apply in this depot centric architecture.For thetas of 0-10o total boiloff is shown as 0.04% or less, which would be 7.2% or less over 180 days. For the 60mT prop load stated, that would be 4.3mt boiled off.
The classic question which is not answered by ZBO hardware is how much prop needs to be sent overboard in station-keeping? I understand the figure is around 0.05% per day (at 300s Isp ULA intends on using LH2/LOX for station keeping TMK, so the ISP would be ~428s). With ZBO you'd need another RCS system (presumably hypergolic) that would kill the CPD once it's prop ran out (or complicates the number of fluids required when re-fuelling).
Quote from: MP99 on 09/28/2012 09:36 amQuote from: muomega0 on 09/27/2012 01:18 pmQuote from: QuantumG on 08/23/2012 06:05 amA new paper "Thermal Optimization of an On-Orbit Long Duration Cryogenic Propellant Depot" co-authored by Bernard Kutter of ULA is attached.They conclude that propellant boil-off rates well under of 0.05% per day for LH2 can be achieved in LEO, for a certain set of conditions.As mentioned earlier, many things are missing from the report, hence the modifications to the conclusion above.Recall that in earlier ULA papers, the boiloff rate for LOX (not LH2) *may* reach 0.1% (the altitude was 1300 km, about double the altitude of this report).Also recall that *with upgrades (25 layers of MLI)* to the 3m diam tank Centaur upper stage, LH2 boil-off can be reduced to 2.4%. Hence the depot with many assumptions would have to reduce the boiloff passively *50 times more* to achieve 0.05% for LH2.What was the sunshield on that configuration (can you link it, or give the paper's title)? I would take this paper at face value, and assume it supersedes earlier work.Practical Depot by Kutter was the original work cited.Face value? can you state how the rate was reduced from the original cite to the paper?
Analysis shows that LO2 equivalent side-wall absorbed heat fluxes of approximately 0.5 BTU/hr/ft² can be obtained for a tank with no surface MLI. Note that this is calculated by taking all heat loads, inclusive of conducted heat, into the tank and dividing by the total surface area of the tank. This is roughly equivalent to a boil-off rate of less than 0.1% of full tank volume per day.Further design and analysis optimization to minimize parasitic heat loads can provide significant further improvement in the cryogenic fluid storage. These improvements include improved thermal isolation of the tank from the warm avionics structure, refined deployable sun shield geometry, and use of LO2 boil-off gas for cooling the sun shield.
Notice how long the sunshield is in the paper to achieve the 0.05%, and this only carries a 11:1 ratio for tank sizes that cannot conduct a lunar sortie, for example.
So how did one easily conclude 40% or more? The 40% does *not* include prop for reboost or station keeping. Now note the attitude, and how it must change continuously over time in the thermal analysis. Do you know what the torque equilibrium attitude would be for the depot and its relation to the attitudes shown?IOW: This station keeping propellant consumption will dwarf reboost, at least for the passive depot.But even the passive depot with superior ISP beats out all other fuels.
Quote from: MP99 on 09/28/2012 09:36 amThe trade-off for putting the depot at 1300 km is that tankers deliver less prop per launch, and the recipient has to reduce it's dry mass to reach where it can be refuelled.Reboost is a very minor trade.
ISS is constrained to 300 to 400 km due to the shuttle mass fraction. However, with the existing fleet, the performance penalty of reaching 300km vs 1200 km may be 500kg of capacity. However, ISS basically cuts it reboost needs in half from 300 to 400 km as drag decreases with altitude.IOW. The depot with smaller LVs can fly higher with minimal reboost needs and capacity hits, and its drag for the same altitude is *substantially* less than ISS due to a substantially smaller footprint.So when ULA states that "a LEO depot....will require the expenditure of propellant at the rate of tons per year", it is not reboost, which leaves station keeping. This is from the Evolving the Depot....paper. Do you take the "tons per year" at face value?
Quote from: MP99 on 09/28/2012 09:36 amQuote from: muomega0 on 09/27/2012 01:18 pmFurther, there are many scenarios when the 0.05%/day passively cannot not be met over a 180 day cycle for a lunar sortie fuel capacity, not the robotic mission sized tanks of this analysis with a 11:1 prop ratio. It is easy to estimate that the passive depot design loss of LH2 at 40% or more over 180 days based on the design presented as it takes time to assemble 100mT of propellant at a 6:1 ratio. Contingencies are not included in the 40% estimate.The good news from the report is that under certain conditions, the passive rate for LH2 (0.05%) is approaching the 0.01%/day or less required for economics. The IMLEO mass savings of using LH2/LO2 is substantial.IOW: the conclusion of the report has not been substantiated by any means for all of the LEO conditions, operations, and contingencies, but progress is being made on the passive design.The great news is that by simply adding few pumps and power to the LEO depot to make it ZBO for LH2, not just LO2, substantial dollars can be saved filling up LEO before departing to L2. SLS need not apply in this depot centric architecture.For thetas of 0-10o total boiloff is shown as 0.04% or less, which would be 7.2% or less over 180 days. For the 60mT prop load stated, that would be 4.3mt boiled off.Again, this is a mission specific lasting a few days or maybe a week or so. It actually states this in the paper! The "beta angle", or the angle between the orbit plane and the sun as known in the thermal community, or the theta angle cited here, varies for ISS +/- 75.1 degrees over a year many times. Here is an example in Figure 2 of beta angle changing over time, the max depends on the inclination.
A long duration CPD is a conceptual space vehicle that can store large quantities of cryogenic propellant for extended periods.
Quote from: MP99 on 09/28/2012 09:36 amThe classic question which is not answered by ZBO hardware is how much prop needs to be sent overboard in station-keeping? I understand the figure is around 0.05% per day (at 300s Isp ULA intends on using LH2/LOX for station keeping TMK, so the ISP would be ~428s). With ZBO you'd need another RCS system (presumably hypergolic) that would kill the CPD once it's prop ran out (or complicates the number of fluids required when re-fuelling).What a fabulous classic question! Why send any propellant overboard? LH2 is volume limited so the cost to launch will only be 3 to 4 times higher vs LO2 due to the 6:1 ratio. By density, its ~ 16 times lighter.So while the passive depot consumes tons of prop for year for station keeping......The Zero Boiloff depot not only has ZBO for LH2 and LO2 the ZBO depot station keeping prop needs are also ZERO[/i].
Frank provided an IVF update at Space 2012http://www.ulalaunch.com/site/docs/publications/IVF-Space-2012.pdf
Practical Depot by Kutter was the original work cited.
Quote from: CommercialSpaceFan on 09/24/2012 12:50 amFrank provided an IVF update at Space 2012http://www.ulalaunch.com/site/docs/publications/IVF-Space-2012.pdfThanks for the link.This also explains *why* they went piston rather than Wankel, which was not clear from the paper Jon referenced. It's not the first time Wankel's have had high temperature seal issues. Id like to note the author has a nice line in understatement.
One item for all those reusable space tug fans is the concept of no more start cartridges.
Quote from: muomega0 on 09/28/2012 01:15 pmPractical Depot by Kutter was the original work cited.Having read this paper is anyone bothered by a couple of things...AFAIK Centaur tanks are the last survivors of the Atlas tank technology. They use stainless steel grade 301. IIRC When I checked this Jim confirmed they still do so. 301 has had 1/10 the thermal conductivity of Aluminum alloys. Can you trust the numbers and graphs and treat the tank description as a typo? Or not?2) IIRC 1300km altitude puts it right in the van allan radiation belts.Does no one think this *might* be an issue?
BTW ULA's partner is XCOR which is doing most of the work.
The new paper (the OP) updates the altitude to 365 nm (676 km).cheers, Martin
One item for all those reusable space tug fans is the concept of no more start cartridges. The ICE is started then the cryo pumps are engaged feeding the GG of the main engine which is ignited and once running the pumps disengaged and the ICE throttled back to a RCS and pressurization role.