Author Topic: Ongoing ULA Cryogenic Propellant Work  (Read 48472 times)

Offline QuantumG

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Ongoing ULA Cryogenic Propellant Work
« on: 08/23/2012 06:05 am »
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.
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Offline mikes

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Re: Ongoing ULA Cryogenic Propellant Work
« Reply #1 on: 08/23/2012 06:18 am »
propellant 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!)
« Last Edit: 08/23/2012 06:50 am by mikes »

Offline Proponent

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Re: Ongoing ULA Cryogenic Propellant Work
« Reply #2 on: 08/23/2012 08:23 am »
propellant 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.

Offline QuantumG

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Re: Ongoing ULA Cryogenic Propellant Work
« Reply #3 on: 08/23/2012 09:01 am »
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?

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Offline Proponent

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Re: Ongoing ULA Cryogenic Propellant Work
« Reply #4 on: 08/23/2012 09:06 am »
Two thoughts about that.  The answer will depend on the delta-V needed for station keeping, since boil-off could be used for that purpose and isn't all wasted.

It will also depend on the departure delta-V needed.

Offline Nathan

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Re: Ongoing ULA Cryogenic Propellant Work
« Reply #5 on: 08/23/2012 09:47 am »
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.

What about helium for pressurization ? Never really mentioned.
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Offline Proponent

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Re: Ongoing ULA Cryogenic Propellant Work
« Reply #6 on: 08/23/2012 10:53 am »
Unlike like a rocket stage, a depot does not need to pump propellants at high rates across a large pressure differential.  Perhaps the depot's self-pressurization at 35 PSI is all that's needed?

Offline joek

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Re: Ongoing ULA Cryogenic Propellant Work
« Reply #7 on: 08/23/2012 01:54 pm »
propellant 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? 
« Last Edit: 08/23/2012 02:06 pm by joek »

Online Chris Bergin

Re: Ongoing ULA Cryogenic Propellant Work
« Reply #8 on: 08/23/2012 02:29 pm »
Cool! (no pun intended ;)) Future article in that.

I've moved the thread to this section, as this is the home of all prop depot stuff, including previous ULA content. I understand why it was in the ULA section, but it's already got a lot of threads, so keeping commercial crew and prop depots in the specific sections is best.
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Offline Robotbeat

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Re: Ongoing ULA Cryogenic Propellant Work
« Reply #9 on: 08/23/2012 05:31 pm »
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?


Remember, the claim is that /well under/ 0.05%/day boil-off in LEO is achievable. Not "down to" .05%, but "well under."
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Offline ChileVerde

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Re: Ongoing ULA Cryogenic Propellant Work
« Reply #10 on: 08/23/2012 06:31 pm »
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?

Yes, this is interesting stuff.

Let me ask a question that has probably been answered to death in times long past:  Would it make sense to launch H2 and O2 depots to rendezvous with a water depot + electrolyzer that could make up for all or some of the cryogen boil-off? Long-term storage of water/ice is not, I trust, a huge problem.
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Offline joek

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Re: Ongoing ULA Cryogenic Propellant Work
« Reply #11 on: 08/23/2012 06:38 pm »
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.

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.


Apologies.  That was uncalled for, not to mention really dumb.
« Last Edit: 08/25/2012 07:24 pm by joek »

Offline mmeijeri

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Re: Ongoing ULA Cryogenic Propellant Work
« Reply #12 on: 08/23/2012 06:42 pm »
Don't underestimate the efficiency of liquid hydrogen as a coolant or the utility of a liquid oxygen depot, especially given the high O/F ratio of LOX/LH2.
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Offline Jim

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Re: Ongoing ULA Cryogenic Propellant Work
« Reply #13 on: 08/23/2012 07:10 pm »
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.


I take their word (analysis) before yours


Offline pathfinder_01

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Re: Ongoing ULA Cryogenic Propellant Work
« Reply #14 on: 08/23/2012 10:09 pm »


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.


Rocket engines don't use a 2:1 ratio for LOH/LOX. In fact the shuttle was 7:1.

Offline pathfinder_01

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Re: Ongoing ULA Cryogenic Propellant Work
« Reply #15 on: 08/23/2012 10:13 pm »


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.

The storage of LO2 isn't a problem in orbit. In fact on a trip to mars, you would have to worry about your LOX freezing. LO2 is a very mild cyrogen, in fact liquid nitrogen is colder. On earth it is possible to store LO2 for years.  The question isn't that the LOH boils 5-20X faster than the LOX, but does it keep long enough to mount a mission. With LOX, you have zero boil off esp. if you have cold gasous hydrogen around to rechill the LOX.
« Last Edit: 08/23/2012 10:25 pm by pathfinder_01 »

Offline pathfinder_01

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Re: Ongoing ULA Cryogenic Propellant Work
« Reply #16 on: 08/23/2012 10:30 pm »
propellant 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!)

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.

Offline jongoff

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Re: Ongoing ULA Cryogenic Propellant Work
« Reply #17 on: 08/23/2012 11:53 pm »
propellant 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.

I haven't had the time yet to read the paper (though I have it in my queue). But if they're looking at an ~18%/yr loss in LEO, that means that in L1/L2, you're probably talking 1-2%/yr (based on relative numbers Bernard and Frank have discussed in previous papers).

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Re: Ongoing ULA Cryogenic Propellant Work
« Reply #18 on: 08/23/2012 11:55 pm »
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.

What about helium for pressurization ? Never really mentioned.

IVF.

http://www.ulalaunch.com/site/docs/publications/Integrated%20Vehicle%20Propulsion%20and%20Power%20System%20for%20Long%20Duration%20Cyrogenic%20Spaceflight%202011.pdf

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Offline muomega0

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Re: Ongoing ULA Cryogenic Propellant Work
« Reply #19 on: 08/24/2012 05:01 pm »
propellant 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? 

Check of calculations.   Using eqn 1 with 24 in the numerator for your conditions, the %Boil-off for the first day results in 0.01% and 0.05%  for LOX and LH2 for the values provided. Multiplying these two % times the initial mass and dividing by the total mass of both,  the total % mass loss for the first day is 0.0124%. If one averages .05 and .01, its 0.03, so Figure 7 is likely just adding the two together ~(0.05 + 0.01=0.06)--which would only be meaningful if the two tanks had the same mass.   So with 24 in the numerator, one obtains the values in Figure 7 by adding the two %s.

Proponent is correct:  boiloff is governed by heat leakage, and the rate in kilograms per hour does not depend on the amount of propellant in the tanks (first order, with many other factors that could be considered).

Is there any rationale for a 11:1 ratio?

The CPD goal of 0.05%/day appears to be an *arbitrary* one.

Many things are "missing".    For example, cut the mass of the LOX by two to achieve a more reasonable ratio.  Destination/payload with the available fuel?  Is this a robotic program?  Recall that one needs ~ 90 mT of prop for a 6 day lunar sortie.  Two  6-day lunar sorties would need  ~2x 13000 kg of LH2.   Too many others missing things to list.

----
Quote
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.

At 170M/launch?  At 350M/launch? Every year?

For 5000 lb of LH2 to start, how does one boil-off 9.35 tons of hydrogen?

---------------
Quote
I wonder how long the campaign has to be before storable propellant wins...

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. ::)

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Re: Ongoing ULA Cryogenic Propellant Work
« Reply #20 on: 08/25/2012 05:58 pm »
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.

Offline joek

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Re: Ongoing ULA Cryogenic Propellant Work
« Reply #21 on: 08/25/2012 07:44 pm »
Don't bother.
I take their word (analysis) before yours

Yes, I was overly-caustic, and stupid to suggest the author's don't know their business.  Apologies (previous post updated).  Question remains as to whether the boil-off rates presented are representative of a useful LO2-LH2 propellant depot.

Offline joek

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Re: Ongoing ULA Cryogenic Propellant Work
« Reply #22 on: 08/26/2012 02:23 am »
Backing up a bit, the attached figure illustrates the source of my question about whether the data in this paper represents a usable propellant depot--that is, if one considers "propellant" to be a useful combination of LH2 and LO2.*

The attached figure was generated using the values from the paper.  The values for Q (heat leak rate) might be off by a bit as they had to be eyeball'd from Fig. 5 and Fig. 6.  Nevertheless, the result tracks closely with the results shown in Fig. 7.

Of note, the paper shows only the total propellant boil-off.  As the paper suggests (but does not explicitly state), and the attached figure shows, the constraint on "propellant" is the LH2 boil-off rate, which appears to be significantly higher than the LO2 boil-off rate.

Again, 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?  Methinks not.


* edit: And in particular, if the claimed "< 0.05%" boil-off rate is something to be celebrated or viewed with skepticism.
« Last Edit: 08/26/2012 03:12 am by joek »

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Re: Ongoing ULA Cryogenic Propellant Work
« Reply #23 on: 08/26/2012 02:44 am »
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?

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.
« Last Edit: 08/26/2012 02:45 am by sdsds »
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Offline mmeijeri

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Re: Ongoing ULA Cryogenic Propellant Work
« Reply #24 on: 08/26/2012 09:17 am »
Earlier papers did talk about significantly higher hydrogen losses and that makes sense. I imagine that's part of the reason for the design they chose in this latest paper. Still, it's total boil-off that matters, even if all hydrogen boils off a LOX depot is still very useful.
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Offline spectre9

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Re: Ongoing ULA Cryogenic Propellant Work
« Reply #25 on: 08/26/2012 11:43 am »
Interesting paper but I'm really not advanced enough in my knowledge to be able to really understand much of it.

I'm thinking that trying to use this method is a way of using existing technology to test cryogenic depots in the near term.

They're not going to get a full blown ACES overnight so now they are looking to make smaller steps. That's just my guess.

Found this nice presentation on the Centaur sun shield. Would this be the technology used here?

Might more than the one RL-10 engine be good if this was to be useful?

Offline KelvinZero

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Re: Ongoing ULA Cryogenic Propellant Work
« Reply #26 on: 08/26/2012 12:35 pm »
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?

Im imagining something fairly separate from this portion, but still casting a shadow on it. Im guessing whatever heat it radiated would not be as much problem as the direct sunlight it removed, so at least you are not worse off.

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Re: Ongoing ULA Cryogenic Propellant Work
« Reply #27 on: 08/26/2012 12:40 pm »
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?

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.
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Offline KelvinZero

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Re: Ongoing ULA Cryogenic Propellant Work
« Reply #28 on: 08/26/2012 01:27 pm »
cool ;)

Maybe they were only testing one aspect rather than spend more money and possibly complicating the results.

How likely are we to see this actually flying? So much of the things I was enthusiastic about have been slashed.

Offline douglas100

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Re: Ongoing ULA Cryogenic Propellant Work
« Reply #29 on: 08/26/2012 03:40 pm »

How likely are we to see this actually flying? So much of the things I was enthusiastic about have been slashed.

Difficult to say. I'd like to see it too, but there would have to be a customer for such a depot. But ULA's continuing work to improve their cryogenic stages and their smart engineering innovations are very impressive. And I agree that the paper QG cites is very interesting and encouraging.
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Re: Ongoing ULA Cryogenic Propellant Work
« Reply #30 on: 08/26/2012 04:37 pm »
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. The first version of ACES that can dock and refuel, followed by a more advanced veriosion with a long drift time between burns for use on Lunar missions (a crasher stage).

A 551 has the cheapest LEO rate for depot resupply which translates to a savings when using a depot with the 0, 1 or 2 SRB versions with much larger GEO sats that would have required 3, 4 or 5 SRB's or even not capable without a depot being larger than what a 551 is capable of similar or larger than possible with a FH to GTO.

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Re: Ongoing ULA Cryogenic Propellant Work
« Reply #31 on: 08/26/2012 04:39 pm »
How 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.
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Re: Ongoing ULA Cryogenic Propellant Work
« Reply #32 on: 08/26/2012 06:03 pm »
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.

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Re: Ongoing ULA Cryogenic Propellant Work
« Reply #33 on: 08/26/2012 06:13 pm »
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.

Well, traffic, or launch tonnage then. And it will be a long time before ULA could recoup the investment required for deploying a LEO depot from increased launch business for larger payloads. Especially since launching the spacecraft mostly dry and refueling it in GEO would be easier. But that's fine, depots are a means to an end, not an end in themselves.
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Re: Ongoing ULA Cryogenic Propellant Work
« Reply #34 on: 08/26/2012 07:57 pm »
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.

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?
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Offline KelvinZero

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Re: Ongoing ULA Cryogenic Propellant Work
« Reply #35 on: 08/27/2012 06:15 am »
How 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?

Offline douglas100

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Re: Ongoing ULA Cryogenic Propellant Work
« Reply #36 on: 08/27/2012 07:59 am »

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 paper is about techniques and technology for reducing cryogenic boil off. The depot presented is described as a "concept." But ULA are working steadily on their technology improvements and some will fly on future cryogenic stages. If a program were funded they could probably produce a demonstrator depot quite quickly.

There is the other issue of ULA not being allowed to produce payloads. I don't know enough to say whether this would affect building a demonstrator.
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Offline KelvinZero

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Re: Ongoing ULA Cryogenic Propellant Work
« Reply #37 on: 08/27/2012 11:15 am »
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  :P

(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. )
« Last Edit: 08/27/2012 11:16 am by KelvinZero »

Offline douglas100

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Re: Ongoing ULA Cryogenic Propellant Work
« Reply #38 on: 08/27/2012 02:23 pm »

The customer used to be NASA, back when it was still funding technology development  :P

(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. )

That's encouraging. NASA was really what I had in mind when I used the word "customer."
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Re: Ongoing ULA Cryogenic Propellant Work
« Reply #39 on: 08/27/2012 04:24 pm »
I was really just meaning a demonstrator. Is it still on the cards?

CRYOTE hasn't been cancelled, has it? And ULA has been talking about an upper stage testbed.
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Re: Ongoing ULA Cryogenic Propellant Work
« Reply #40 on: 08/27/2012 05:52 pm »
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?

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.

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Re: Ongoing ULA Cryogenic Propellant Work
« Reply #41 on: 08/27/2012 05:56 pm »
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?

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.

Yeah, the .05%/day number in LEO allows you to get rid of hydrazine for propulsive station keeping. It only makes sense for a LEO depot to go any lower if you've got a propellantless (or higher Isp) method of stationkeeping. I guess if you could feed the GH2 into some sort of solar electric or solar thermal system (without dumping more heat leak into the system) it might make sense, or if you used an ED tether for reboost (and settling). But if you're just using chemical propulsion for settling and reboost/stationkeeping, a purely passive boiloff system makes sense.

Now for a tanker vehicle, or a deep-space propulsion stage, lower or zero boiloff might be better. But for a LEO depot that has to keep itself in orbit, ZBO is gilding the lily.

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Re: Ongoing ULA Cryogenic Propellant Work
« Reply #42 on: 08/27/2012 06:27 pm »
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.

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?
If the Atlas V first stages are reusable... (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.)
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Re: Ongoing ULA Cryogenic Propellant Work
« Reply #43 on: 08/27/2012 07:05 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.)

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Re: Ongoing ULA Cryogenic Propellant Work
« Reply #44 on: 08/27/2012 07:07 pm »
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?

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

Several items here:

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.

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Re: Ongoing ULA Cryogenic Propellant Work
« Reply #45 on: 08/27/2012 07:11 pm »
Now 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.
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Re: Ongoing ULA Cryogenic Propellant Work
« Reply #46 on: 08/27/2012 08:16 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.)
Good to know. Sucks. In the name of cost-saving, we'll be stuck with higher cost launch vehicles. So. Short. Sighted.
« Last Edit: 08/27/2012 08:17 pm by Robotbeat »
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Offline jongoff

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Re: Ongoing ULA Cryogenic Propellant Work
« Reply #47 on: 08/27/2012 08:24 pm »
1) You will need to keep the depot and Centaur completely inside the faring because of the sunshields.

Not necessarily. Part of the reason for developing the pneumatically deployable sunshields in the first place was that they might not be able to have it deployed before launch. There were also some intriguing possibilities under development by Quest Products before NASA's SBIR program screwed things up last year that could've been great (I talked about them here: http://selenianboondocks.com/2011/11/tooting-someone-elses-horn-quest-product-development-corps-advanced-mli-technologies/

[NASA SBIR screwed things up by poor implementation of their change in how the fund Phase II SBIRs, which unnecessarily resulted in them giving out half as many Phase I and Phase II awards last year in spite of getting a budget increase for the SBIR program last year. Quest was one of the many companies with great technologies that had solid Phase I efforts that got completely hosed by how NASA SBIR implemented that policy change.]

Quote
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.

Yeah, as I was saying it would require some new tank work. Might be able to leverage some of the DCSS tank fixtures, or there may be some leftover tank tooling from the bigger-diameter Centaur tanks I think they did for Shuttle Centaur/Centaur G.

Quote
3)A 4.33M diameter tank would double the volume for the LH2 tank dropping the ratio to ~5.5:1.

Yeah, I'd have to dig up my old numbers but I found the performance optimizing solution for a single-launch Centaur-derived depot involved using a 5m diameter LH2 tank for the depot, launched empty, with the deployable sunshield, and a slight tank stretch on the Centaur LH2 tank to give you more LOX capacity while keeping the MR reasonable. IIRC it was close to 75mT propellant capacity at a reasonable MR...but I'd have to dig the numbers up to confirm.

~Jon

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Re: Ongoing ULA Cryogenic Propellant Work
« Reply #48 on: 08/27/2012 08:25 pm »
Now 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.

Oh definitely. Even at L1/L2, the boiloff rate drops by ~10X compared to LEO. At least according to some of ULA's previous papers.

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Re: Ongoing ULA Cryogenic Propellant Work
« Reply #49 on: 08/27/2012 08:34 pm »
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.

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.

And if you were to choose cryocoolers, the trade-off would change again. It is instructive to compare the old papers about long time storability of DCSS and Centaur. Centaur's common bulkhead helps keep the LOX cold, but it makes things worse for the LH2 and DCSS's large diameter helps limit boil-off. And note that the latest paper does not feature a common bulkhead between the LOX and the 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. Adding radiators as well as solar panels for powering cryocoolers would seem to make the problem worse.
« Last Edit: 08/27/2012 08:35 pm by mmeijeri »
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Re: Ongoing ULA Cryogenic Propellant Work
« Reply #50 on: 08/27/2012 09:54 pm »
There's cheaper ways to do station keeping... especially in LEO where there are zero-propellant magnetic despin options.

However, if you're using an upper stage with RL-10 performance (say, 465s of ISP) then you can afford to have boiloff.

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Re: Ongoing ULA Cryogenic Propellant Work
« Reply #51 on: 08/28/2012 04:26 am »
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.

Good point. I think the absence of that module points to this design being the basis for a thermodynamic model, but not the basis for an actual depot. Are you concerned that the presence of a docking/transfer module would significantly change the boil off rate?
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Re: Ongoing ULA Cryogenic Propellant Work
« Reply #52 on: 08/28/2012 12:28 pm »
There'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....
« Last Edit: 08/28/2012 12:52 pm by Proponent »

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Re: Ongoing ULA Cryogenic Propellant Work
« Reply #53 on: 08/28/2012 01:41 pm »
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. 

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Re: Ongoing ULA Cryogenic Propellant Work
« Reply #54 on: 08/28/2012 04:29 pm »
Are you concerned that the presence of a docking/transfer module would significantly change the boil off rate?

I wouldn't say concerned, but certainly curious. It does look like a complication.
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Re: Ongoing ULA Cryogenic Propellant Work
« Reply #55 on: 08/28/2012 10:12 pm »
There'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....

Proponent beat me to the punch on this one. Attitude control and stationkeeping (maintaining orbital altitude, inclination, and RAAN as necessary for operations and in the face of atmospheric drag and other perturbations) are totally different things, and magnetic torquers don't do anything for stationkeeping.

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.

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Re: Ongoing ULA Cryogenic Propellant Work
« Reply #56 on: 08/28/2012 10:15 pm »
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.

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.

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Re: Ongoing ULA Cryogenic Propellant Work
« Reply #57 on: 08/29/2012 12:20 am »
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?

Yeah, that's what I was thinking. Otherwise you might prefer something like ISS altitude, and fortunately ISS doesn't require anywhere near 0.05 % / day in reboost, and actually increased its altitude after the Shuttle.

After all, atmospheric density is roughly negatively exponential in altitude. Orbital velocity squared varies less over the range concerned, but also diminishes with altitude. I don't know if D = 1/2 rho v^2 Cd A is still accurate at the speeds and densities involved, but I believe it is still at least a rough approximation, certainly at low altitudes. Or perhaps I should say "if Cd can be treated as roughly constant over the range involved", since IIRC the drag equation is basically the definition of Cd.

And wouldn't atomic oxygen be more of a problem at lower altitudes, for roughly the same reasons?

I can believe that at 0.05% / day you could afford the higher drag losses, but not that you'll end up with lower total losses than a ZBO depot at ISS altitude.

Do you know if exponential density and constant Cd are accurate enough approximations? If so, it would be fun and possibly enlightening to do some calculations.

Quote
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.

Only 20 m/s or so IIRC, not by itself worth 0.05% / day. It may have been more of a problem for the Shuttle than it would be for launchers with a less extreme delta-v split between their stages, just as with the payload hit for the higher inclination, and for similar reasons.

Quote
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.

It would if it were a net win, but I don't think it is. The 0.05% / day is certainly good enough for a LEO depot, which after all only needs to export propellant to L1/L2, but the low altitudes don't look like something you'd do if you had ZBO or something close to it.
« Last Edit: 08/29/2012 12:28 am by mmeijeri »
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Re: Ongoing ULA Cryogenic Propellant Work
« Reply #58 on: 08/29/2012 12:23 am »
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.

So are depots of course...
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Re: Ongoing ULA Cryogenic Propellant Work
« Reply #59 on: 08/29/2012 02:25 am »
Martijn,

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.

~Jon

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Re: Ongoing ULA Cryogenic Propellant Work
« Reply #60 on: 08/29/2012 07:50 am »
IVF.

http://www.ulalaunch.com/site/docs/publications/Integrated%20Vehicle%20Propulsion%20and%20Power%20System%20for%20Long%20Duration%20Cyrogenic%20Spaceflight%202011.pdf

~Jon
Thanks 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.
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Offline mmeijeri

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Re: Ongoing ULA Cryogenic Propellant Work
« Reply #61 on: 08/29/2012 09:50 am »
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.

It would be interesting if you could find out more. In the mean time doing some calculations might still be fun.

Phasing could be another complication for very low orbits. If your tanker can't go above the altitude of the depot because that would negate the delta-v advantage, then it may have to stay in a very low orbit indeed, and take a long time to line up precisely, during which time it would suffer from increased boil-off and drag. Or you would be limited in your launch windows, which could lead to less frequent deliveries from a LEO depot to an L1/L2 depot, which in turn also increases boil-off and reduces utilisation.

Possibly not enough of a complication to make you not want to use low orbits given the boil-off, but another reason not to make it a net selling point and another reason not to use low orbits on ZBO depots.
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Re: Ongoing ULA Cryogenic Propellant Work
« Reply #62 on: 09/01/2012 05:54 pm »
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.

Interesting question.

The two heat sources considered are the earth and sun. Earth is about 300 degrees K and subtends almost 2 pi steradians as seen from a depot in LEO. The sun is 5778 degrees K and subtends about 5e-5 steradians (if my arithmetic's right).

Quick Googling seems to indicate rocket exhaust is 6000 degrees K, in the neighborhood of the sun's surface temperature. I don't know at what distance the rendezvous burns would be made so I don't know if they'd subtend a solid angle comparable to the sun's.

I think the docking vehicle would be in the neighborhood of 300 degrees K. And it would subtend a large angle during the time it's docked.

These thoughts have given me a new appreciation for hydrogen boiloff station keeping, though. It seems to me conventional station keeping burns would be heat sources subtending a large solid angle and within the O2 MLI skirt.
« Last Edit: 09/01/2012 06:20 pm by Hop_David »

Offline Proponent

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Re: Ongoing ULA Cryogenic Propellant Work
« Reply #63 on: 09/02/2012 04:12 pm »
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.

Offline jongoff

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Re: Ongoing ULA Cryogenic Propellant Work
« Reply #64 on: 09/03/2012 03:47 pm »
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.

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
« Last Edit: 09/03/2012 03:47 pm by jongoff »

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Re: Ongoing ULA Cryogenic Propellant Work
« Reply #65 on: 09/04/2012 11:26 pm »
"A Study of CPS Stages for Missions beyond LEO"

http://spirit.as.utexas.edu/~fiso/telecon/King_2-15-12/

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Re: Ongoing ULA Cryogenic Propellant Work
« Reply #67 on: 09/05/2012 04:32 am »
"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.
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Re: Ongoing ULA Cryogenic Propellant Work
« Reply #68 on: 09/05/2012 05:38 pm »
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.

Good to know.

Of 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.

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Re: Ongoing ULA Cryogenic Propellant Work
« Reply #69 on: 09/05/2012 06:55 pm »
Of 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.

Yeah, with the SAA we just signed with LaRC, the second part of that would be for long-reach manipulators that would be useful for space facilities like ISS or propellant depots. We should be getting into that more next year (this year's focus is on smaller arms for crew/cargo vehicles and the like).

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Re: Ongoing ULA Cryogenic Propellant Work
« Reply #70 on: 09/05/2012 06:58 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.

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.

~Jon

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Re: Ongoing ULA Cryogenic Propellant Work
« Reply #71 on: 09/05/2012 07:25 pm »
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
Any N2H4 mono thruster with an expansion ratio above about 40 (most of them), is going to have pretty close to room temp exhaust.

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Re: Ongoing ULA Cryogenic Propellant Work
« Reply #72 on: 09/05/2012 07:35 pm »
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
Any N2H4 mono thruster with an expansion ratio above about 40 (most of them), is going to have pretty close to room temp exhaust.

Ok, that's about what I figured, but I was being intentionally wishy-washy on the low-side numbers, since I was too lazy to run the analysis. I knew the high-side temp for them was ~1200K, but wasn't sure how fast they cooled off. I know that for some space thrusters with condensable exhaust products (like LOX/LH2), that one of the limits on expansion ratio is condensing or freezing in the exhaust...

~Jon
« Last Edit: 09/05/2012 07:36 pm by jongoff »

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Re: Ongoing ULA Cryogenic Propellant Work
« Reply #73 on: 09/05/2012 08:17 pm »
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. 

I've been asking the same thing about liquid methane.
Go back to page 2 of the thread and read muomega0's post:
Quote
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.

I believe this means liquid methane loses, even when considering boil-off, but obviously controlling (or making use of) boil off is very desirable.

Of course, using fuel available in-situ at a destination is best for the return flight (shout out to Titan)!

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Re: Ongoing ULA Cryogenic Propellant Work
« Reply #74 on: 09/06/2012 01:38 am »
"A Study of CPS Stages for Missions beyond LEO"

A final version of the presentation, along with the full paper presented at the Global Space Exploration Conference, is available at:
http://www.sei.aero/news/newsindex.php?id=516

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.

Agreed. Even for relatively easy first steps along the path, like a crewed visit to an NEA, the amount of propellant required for an architecture that calls for all-hydrolox propulsion is staggering. It deserves some rethinking of the Apollo S-IVB or CxP EDS concept.

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. Doing it the other way is kind of like trying to do an Apollo lunar mission without lunar orbit rendezvous....
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Offline Damon Hill

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Re: Ongoing ULA Cryogenic Propellant Work
« Reply #75 on: 09/06/2012 02:54 am »
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.

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Re: Ongoing ULA Cryogenic Propellant Work
« Reply #76 on: 09/06/2012 03:41 am »
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.

ah in space power is usually provided by batteries, solar cells and occasionally APU(which mix hyrdrozine and nitrogentetoxide in a combustion camber of sorts). Fuel cells were only used to provide power for Gemini, Apollo and the Shuttle. Rockets usually don't use fuel cells.  Upper stages are usually battery powered.

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Re: Ongoing ULA Cryogenic Propellant Work
« Reply #77 on: 09/06/2012 05:52 pm »
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.
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Re: Ongoing ULA Cryogenic Propellant Work
« Reply #78 on: 09/07/2012 04:03 am »
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.

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Re: Ongoing ULA Cryogenic Propellant Work
« Reply #79 on: 09/07/2012 08:08 am »
High 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.

If you think about it, starting from LEO towards your destination in a single hop is to refueling and multiple hops as using an HLV to launch EDS + propellant in one launch is to using multiple smaller launchers and depots. A large EDS and a single hop is the in-space equivalent of an HLV and a single launch, the in-space equivalent of the Saturn / Apollo approach to launch.
« Last Edit: 09/07/2012 09:15 am by mmeijeri »
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Re: Ongoing ULA Cryogenic Propellant Work
« Reply #80 on: 09/07/2012 04:46 pm »
... but obviously controlling (or making use of) boil off is very desirable.

Especially when you consider that conventional station keeping would make heat in the propellant depot's rocket chambers. And making heat is the last thing you want.

But using boil off for station keeping expels heat.

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.

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Re: Ongoing ULA Cryogenic Propellant Work
« Reply #81 on: 09/08/2012 11:17 am »
High 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.

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Re: Ongoing ULA Cryogenic Propellant Work
« Reply #82 on: 09/08/2012 11:35 am »
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 see. Depending on the time frame I'm not opposed to this at all. Bear in mind though that if you were to use SEP to preposition propellant used to transfer from HMO to LMO, transfer back to HMO, transfer to Earth and insert into HEO the trade could still favour hypergolics + SEP vs all LOX/LH2. Once you have long term storage of LOX/LH2 (at least in deep space), the balance could tip in favour of SEP + LOX/LH2. In any event, it seems undesirable to add an artificial dependency of SEP tugs on cryo depots, with no progress being made on the former until deployment of the latter.

Quote
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.

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
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.

Agreed. LOX/LH2 is clearly superior for LEO to L1/L2, and could overtake hypergolics in the medium term for L1/L2 to Mars transfer orbit. And manned missions to Mars or Mars orbit are not likely to happen before the long term anyway. For unmanned missions however, hypergolics would be superior, but there single launch is still preferable to propellant transfer from the end user's perspective.

ULA's work on actual depots is interesting, but modifying Centaur or DCSS for EOR and EDS service seems more immediately important. And it would still be a significant step in the direction of cryogenic depots, in several respects. You'd have to find near term advantages to something you want for long term reasons. IVF does that, but the near term advantages are only small. They have a whole list of minor advantages, but not a single big one.

If we had a market for hypergolics at L1/L2, then all that could change. A large market can support greater investment in infrastructure, which is what depots are. It would mean a less prominent role for ULA however, which is probably not attractive to them, even though it would be better for commercial development of space.
« Last Edit: 09/08/2012 01:31 pm by mmeijeri »
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Re: Ongoing ULA Cryogenic Propellant Work
« Reply #83 on: 09/08/2012 12:03 pm »
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.

Well, yes and no. Using hydrogen boil-off gives higher Isp than hydrazine monopropellant, but the boil-off is much higher than what you need for station keeping, unless you go to really low altitudes. So what you gain in Isp, you lose in additional boil-off, and going to lower altitudes only recovers a bit of that.

The strongest argument for LOX/LH2 still appears to be its unrivalled Isp.
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Re: Ongoing ULA Cryogenic Propellant Work
« Reply #84 on: 09/08/2012 04:51 pm »
Quote
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.

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.


Originally, I was thinking of a fairly large manned Mars mission using SLS and no SEP tug, but it works out for smaller LVs too.  Whatever mass you want to put through TMI from L2, requires about 20% more mass of cryo propulsion stage, or 30% more of hypergolic propulsion stage.

So if you want your final launch to be the one and only cryo stage, and all cryo stage, you can have five previous launches. With SEP tug, I expect that number drops to 2 or 3, for the same mass through TMI, but you have the added cost of developing & deploying multiple tugs and extra rad hardening of all payloads. But in that case you'd be better off sending a hypergolic propulsion stage via SEP anyway. :)

So if you have neither ZBO nor SEP tugs, and must assemble the mission before TMI, I suggest this is a quite efficient way to do it. Although we're getting a little OT.
« Last Edit: 09/08/2012 04:53 pm by kkattula »

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Re: Ongoing ULA Cryogenic Propellant Work
« Reply #85 on: 09/09/2012 01:11 am »
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.

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.

Although the hydrazine exothermic reactions aren't nearly as horrible as the 6000 degree K rocket chambers I initially imagined, these burns still seem like you'd want to avoid if there were an alternative.

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.

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Re: Ongoing ULA Cryogenic Propellant Work
« Reply #86 on: 09/09/2012 08:57 am »
I said station keeping by boil off. Boil off removes heat from the cryogens.

Sure, that's how I understood what you said. I don't think a brief hydrazine station keeping burn will add much heat, but maybe I'm all wrong about that.

Quote
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.

I vaguely recall reading it was higher than 300s Isp, which is almost biprop level, and better than hydrazine.

Quote
However if station keeping is achieved by boil-off, it seems to me hydrogen has the best ISP.

I don't understand. How would the source of the gas impact Isp?
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Re: Ongoing ULA Cryogenic Propellant Work
« Reply #87 on: 09/09/2012 04:28 pm »
...
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.

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Re: Ongoing ULA Cryogenic Propellant Work
« Reply #88 on: 09/09/2012 04:32 pm »
The boiled off H2 could be put through a small, solar powered resistojet, for higher Isp.

They did talk about something like that (solar thermal), but it does you no good if your boil-off rate is higher than your station keeping needs. You may need a lower amount of hydrogen for your thrusters, but you're still venting the rest. That's also why the efficiency of the IVF ICE isn't all that important.
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Offline mmeijeri

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Re: Ongoing ULA Cryogenic Propellant Work
« Reply #89 on: 09/09/2012 04:37 pm »
@Hop David:

Maybe we were talking past each other. Did you mean that using boiled off hydrogen is preferable to using hydrazine on a non-zero boil-off LOX/LH2 depot? Or did you mean that boil-off is actually an advantage? If the former, we agree.
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Offline CommercialSpaceFan

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Re: Ongoing ULA Cryogenic Propellant Work
« Reply #90 on: 09/24/2012 12:50 am »
IVF.

http://www.ulalaunch.com/site/docs/publications/Integrated%20Vehicle%20Propulsion%20and%20Power%20System%20for%20Long%20Duration%20Cyrogenic%20Spaceflight%202011.pdf

~Jon
Thanks 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.

Frank provided an IVF update at Space 2012
http://www.ulalaunch.com/site/docs/publications/IVF-Space-2012.pdf

Offline Ronsmytheiii

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Re: Ongoing ULA Cryogenic Propellant Work
« Reply #91 on: 09/25/2012 02:55 am »
Surprised this wasn't cross referenced

RELEASE: 12-333

NASA SELECTS SPACE LAUNCH SYSTEM ADVANCED DEVELOPMENT PROPOSALS

WASHINGTON -- 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.

Quote
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
.

http://forum.nasaspaceflight.com/index.php?topic=26853.450
« Last Edit: 09/25/2012 02:55 am by Ronsmytheiii »

Offline jongoff

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Re: Ongoing ULA Cryogenic Propellant Work
« Reply #92 on: 09/25/2012 04:27 pm »
Surprised this wasn't cross referenced

Snip

-- "Integrated Vehicle Fluids (IVF)," United Launch Alliance,
Centennial, Colo
.

http://forum.nasaspaceflight.com/index.php?topic=26853.450[/quote]

Thanks 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

Offline Robotbeat

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Re: Ongoing ULA Cryogenic Propellant Work
« Reply #93 on: 09/25/2012 07:30 pm »
Surprised this wasn't cross referenced
Quote
Snip

-- "Integrated Vehicle Fluids (IVF)," United Launch Alliance,
Centennial, Colo
.

http://forum.nasaspaceflight.com/index.php?topic=26853.450

Thanks 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
Yes, definitely cool.

And really, also cool that it's on SLS. May not be the first choice for you and I to have such a launch vehicle (when there are other, cheaper options), but if we MUST have SLS, it's best to make the very best of it!  Glad to hear SLS may get IVF as well.

I do not subscribe to the viewpoint that just because (insert unfavored spaceflight architectural approach) is counterproductive, anything good for (insert unfavored spaceflight architectural approach) is bad for spaceflight...
« Last Edit: 09/26/2012 12:32 am by Robotbeat »
Chris  Whoever loves correction loves knowledge, but he who hates reproof is stupid.

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

Offline neilh

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Re: Ongoing ULA Cryogenic Propellant Work
« Reply #94 on: 09/25/2012 10:06 pm »
Surprised this wasn't cross referenced

Snip

-- "Integrated Vehicle Fluids (IVF)," United Launch Alliance,
Centennial, Colo
.

http://forum.nasaspaceflight.com/index.php?topic=26853.450

Thanks 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
[/quote]

Btw, here's the relevant part of the selection statement, posted in the SLS thread:

Quote
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.
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Online oldAtlas_Eguy

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Re: Ongoing ULA Cryogenic Propellant Work
« Reply #95 on: 09/26/2012 04:46 pm »
Surprised this wasn't cross referenced

Snip

-- "Integrated Vehicle Fluids (IVF)," United Launch Alliance,
Centennial, Colo
.

http://forum.nasaspaceflight.com/index.php?topic=26853.450

Thanks 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

Btw, here's the relevant part of the selection statement, posted in the SLS thread:

Quote
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.

BTW ULA's partner is XCOR which is doing most of the work.
« Last Edit: 09/26/2012 04:48 pm by oldAtlas_Eguy »

Offline muomega0

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Re: Ongoing ULA Cryogenic Propellant Work
« Reply #96 on: 09/27/2012 01:18 pm »
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.


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.

Offline MP99

Re: Ongoing ULA Cryogenic Propellant Work
« Reply #97 on: 09/28/2012 09:36 am »
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.

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.



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.

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). 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).

IVF (turning the boiloff "lemon" into RCS "lemonade") allows the depot to keep operating until the schedule catches up again, though it may require a small top-up delivery

cheers, Martin

Offline muomega0

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Re: Ongoing ULA Cryogenic Propellant Work
« Reply #98 on: 09/28/2012 01:15 pm »
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.

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?

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.  Worse, it appears this depot will spin about the axis to settle the fluid at the "bottom" of the tanks, rather than about their centerline.  :o

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
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.

Reboost is a very minor trade.  Let's compare to ISS Reboost  and ATV-2 331 kg + 35 kg prop raised ISS from 345 km to 388 km.   ISS basically cuts it reboost needs in half from 300 to 400 km as drag decreases with altitude, and its likely 4X the drag area than the depot.

ISS is constrained to 300 to 400 km due to the shuttle mass fraction.

Ares V Mars DRM Figure 4-4 pg 24
  300 km  125 mT    400 km 120 mT   650 km  108 mT
thats a 17mT hit at 650 km, still below the van allen belts.

However, with the existing fleet, the performance penalty of reaching 300km vs 1200 km may be 500kg (0.5 mT) of capacity.

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?


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.

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 

Quote
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).


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].
« Last Edit: 12/19/2014 01:24 pm by muomega0 »

Offline MP99

Re: Ongoing ULA Cryogenic Propellant Work
« Reply #99 on: 09/28/2012 11:36 pm »
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.

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?

The new design includes MLI which they tried to avoid in the earlier design.
Quote
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.

Given this is a later paper, it seems reasonable to me that this is the result of "further design and analysis optimization".



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. 

If you just half-fill the LO2 tank, you'd have a 5.5:1 ratio (or adjust as needed).



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. :)

The attitude relative to the Sun (theta) is held steady, it doesn't "change continuously over time". The paper discusses how the albedo & IR from Earth impinges on the H2 & O2 ends at various points throughout the orbit, and how this in turn changes as the beta angle varies (which I presume will happen naturally as with ISS).



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.

Reboost is a very minor trade.

I didn't mention reboost, just the additional prop that tankers must use to reach a depot at a higher orbit. (And equivalent for client spacecraft that might receive the contents of the depot. For instance, I assume no-upper-stage SLS would suffer quite an impact to payload if having to go to 1300 km instead of 675 km.)



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?

I said 4.3 mT boiloff for 6 months, which is 8.6mT per year. So yes, "at face value". Don't understand where your 40% figure comes from.

The paper [table 1] also states the altitude as 676 km (365 nm), not the 300 km you use in your example. Reboost shouldn't be a major concern at this altitude?



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.

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.

From the first sentence of the introduction of the new paper:-
Quote
A long duration CPD is a conceptual space vehicle that can store large quantities of cryogenic propellant for extended periods.

I also don't understand your comment "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". What sort of scenarios?

Also, I think the ULA paper uses Beta in exactly the same sense as in your quoted paper (not "or the theta angle cited here"). Theta is the attitude the CPD holds itself relative to the Sun.



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).

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].

I don't understand how you perform RCS functions without sending prop mass overboard?

cheers, Martin

Offline john smith 19

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Re: Ongoing ULA Cryogenic Propellant Work
« Reply #100 on: 09/29/2012 11:10 am »
Frank provided an IVF update at Space 2012
http://www.ulalaunch.com/site/docs/publications/IVF-Space-2012.pdf

Thanks 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.  :)
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Offline john smith 19

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Re: Ongoing ULA Cryogenic Propellant Work
« Reply #101 on: 09/29/2012 11:47 am »

Practical Depot by Kutter was the original work cited.


Having read this paper is anyone bothered by a couple of things.

1) It states the depot builds on Centaur tank technology with the tanks being spun Aluminum alloy friction stir welded to monocoque cylinders.

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?
MCT ITS BFR SS. The worlds first Methane fueled FFSC engined CFRP SS structure A380 sized aerospaceplane tail sitter capable of Earth & Mars atmospheric flight.First flight to Mars by end of 2022 TBC. T&C apply. Trust nothing. Run your own #s "Extraordinary claims require extraordinary proof" R. Simberg."Competitve" means cheaper ¬cheap SCramjet proposed 1956. First +ve thrust 2004. US R&D spend to date > $10Bn. #deployed designs. Zero.

Online oldAtlas_Eguy

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Re: Ongoing ULA Cryogenic Propellant Work
« Reply #102 on: 09/29/2012 11:34 pm »
Frank provided an IVF update at Space 2012
http://www.ulalaunch.com/site/docs/publications/IVF-Space-2012.pdf

Thanks 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. 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.

Offline mmeijeri

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Re: Ongoing ULA Cryogenic Propellant Work
« Reply #103 on: 09/30/2012 08:49 am »
One item for all those reusable space tug fans is the concept of no more start cartridges.

I thought RL-10 didn't need start cartridges to begin with. And, at least in the previous paper, IVF had sort of self-replenishing oxygen and hydrogen bottles. Of course, you still need to purge the engine.
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Offline MP99

Re: Ongoing ULA Cryogenic Propellant Work
« Reply #104 on: 09/30/2012 10:41 am »

Practical 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?

The new paper (the OP) updates the altitude to 365 nm (676 km).

cheers, Martin
« Last Edit: 09/30/2012 10:42 am by MP99 »

Offline mmeijeri

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Re: Ongoing ULA Cryogenic Propellant Work
« Reply #105 on: 09/30/2012 12:29 pm »
Would it be useful to have a larger version of IVF on the Delta IV first stage to pressurise the LOX instead of by using stored helium? I've read that the RS-68 has a simpler pressurisation system than the SSME, is that because unlike the SSME the RS-68 doesn't have an oxidiser heat exchanger? Would IVF be preferable to using a heat exchanger extracting heat from the main engine if you did want to replace helium?
« Last Edit: 09/30/2012 12:33 pm by mmeijeri »
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Offline john smith 19

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Re: Ongoing ULA Cryogenic Propellant Work
« Reply #106 on: 09/30/2012 04:29 pm »

BTW ULA's partner is XCOR which is doing most of the work.

The report states the pump design is based on Xcorp designs but I'm not sure they are partners on this project. IIRC Xcorp pump designs are patented so even if they are not they should be getting something for their previous efforts.

However the *key* contractor seems to be "Roush Industries" for the core of the I6 engine, integrated starter generator, H2/O2 pump integration etc. They appear to be a fast turnaround prototyping and engineering operation with a strong automotive background (Their head office is in Michigan).
MCT ITS BFR SS. The worlds first Methane fueled FFSC engined CFRP SS structure A380 sized aerospaceplane tail sitter capable of Earth & Mars atmospheric flight.First flight to Mars by end of 2022 TBC. T&C apply. Trust nothing. Run your own #s "Extraordinary claims require extraordinary proof" R. Simberg."Competitve" means cheaper ¬cheap SCramjet proposed 1956. First +ve thrust 2004. US R&D spend to date > $10Bn. #deployed designs. Zero.

Offline john smith 19

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Re: Ongoing ULA Cryogenic Propellant Work
« Reply #107 on: 09/30/2012 04:44 pm »
The new paper (the OP) updates the altitude to 365 nm (676 km).

cheers, Martin

This seems more viable in the short term, although still somewhat higher than the ISS orbit.

I note the heat transfer coefficient is listed 0.75BTU/hr-ft^2-R. Given the explicit comment that the whole design is based on Centaur stage technology I'd want to confirm that figure is for a steel tank or an aluminum tank.  It *should* be for SS301 but (given the mix up with the other paper) I'm not sure it is.
MCT ITS BFR SS. The worlds first Methane fueled FFSC engined CFRP SS structure A380 sized aerospaceplane tail sitter capable of Earth & Mars atmospheric flight.First flight to Mars by end of 2022 TBC. T&C apply. Trust nothing. Run your own #s "Extraordinary claims require extraordinary proof" R. Simberg."Competitve" means cheaper ¬cheap SCramjet proposed 1956. First +ve thrust 2004. US R&D spend to date > $10Bn. #deployed designs. Zero.

Offline john smith 19

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Re: Ongoing ULA Cryogenic Propellant Work
« Reply #108 on: 10/16/2012 08:50 am »
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.
I'm not sure how well know it is but toward the end of the original J2-X development programme they reported they had found a way to eliminate start cartridges. The key issue seemed to be the back pressure from the combustion chamber and nozzle cooling pipes being too high to allow the turbine to spin up from the boiling hydrogen in the inlet ducts. 

It seemed it was a question of the proper valve scheduling, which would (had the programme not been terminated) given NASA a simple high thrust, high performance 2nd stage engine with *unlimited* starts.

AFAIK RL10 does not use start cartridges but IMHO eliminating consumables (He tank pressurization) and limits on starting is always a good idea.
MCT ITS BFR SS. The worlds first Methane fueled FFSC engined CFRP SS structure A380 sized aerospaceplane tail sitter capable of Earth & Mars atmospheric flight.First flight to Mars by end of 2022 TBC. T&C apply. Trust nothing. Run your own #s "Extraordinary claims require extraordinary proof" R. Simberg."Competitve" means cheaper ¬cheap SCramjet proposed 1956. First +ve thrust 2004. US R&D spend to date > $10Bn. #deployed designs. Zero.

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