Author Topic: All EELV VSE architecture  (Read 73659 times)

Offline HIP2BSQRE

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Re: All EELV VSE architecture
« Reply #40 on: 07/02/2009 04:33 AM »
Has anyone read all the proposals--a lot of them like using an existing launcher rather than spending $18b on a new a launcher.  Some even use space depots--like T/Space.  Where have I seen that in another proposal????

Offline libs0n

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Re: All EELV VSE architecture
« Reply #41 on: 07/02/2009 07:11 AM »
My summation of a particular layout to help realize the aims of this thread.

The objective of this white paper architecture is that NASA builds the lunar hardware, the Orion and the Altair, while the major lunar propulsion hardware is based upon an enhanced EELV second stage.  All mission mass is launched on the EELV vehicles.

The EELV vehicles are existing commercial launch vehicles of known reliability, which can be produced in quantity to launch the lunar mission mass.  This commercial utilization will promote the growth and competitive health of the American launch industry in both the domestic and international launch market, while also providing a sound footing for the successful conduct of a new lunar exploration campaign.

The other centerpiece of this white paper should be the use and advantages of the proposed EELV upgrade known as the "Common Upper Stage".  This upgrade appears to be an expansion of the existing Centaur stage, and features variable tank lengths to accommodate different sized propellant loads, and a variable count of RL-10 engines, from 1 to 6 and points in between.

For the purposes of the white paper, further upgrades are made to this Common Upper Stage.  First is the capability to transfer propellant so as to be refueled on orbit.  Second is the capability to automatically rendezvous on orbit, to assist in the transfer of propellant.  Other specific enhancements, like better mission duration extension options, are left to further specific consideration.

This centerpiece stage can serve as the basis for the following:

1. The Trans Lunar Injection stage(or stages).
2. Perform the Lunar Orbit Insertion role for the lunar hardware.
3. Assist in the descent burn of the Altair, thus possibly simplifying the design of such vehicle, and allowing it to fit within the EELV fairing without sacrificing lander functionality.  This propulsive burn can be done with either a discrete EELV stage, or by the element that performs other propulsive duties.  Whatever element performs this assist will be discarded on the moon in a crashing method.
4. The containment vessels for the propellant to be transferred through launch and to on orbit transfer.
5. The base structure of any propellant storage depot to assist in the logistics of refueling.

This upgrade will also boost the payload capability of the EELV launch vehicles that conduct the launch campaign for the lunar mission mass.  As it is applicable to both vehicle lines, it will possibly allow for ease of mission implementation using both vehicles and existing facilities and allow for some measure of assured reliability launch access.

For the conduct of the lunar mission, a particular arrangement is detailed:

Orion + refueled EELV second stage for TLI, LOI
Altair + refueled EELV second stage for TLI, LOI, and possibly assisting in the descent.  Otherwise a discrete crasher stage in addition for the Altair.

The propellant medium to be refueled should be LOX, as it makes up the bulk of the propellant and mission mass, and is presumably the simpler cryogen to transfer.  This propellant can be launched separately in the second stages of launch vehicles dedicated to this task.  The mission components that require this LOX load can be launched without it, and receive such LOX loading on orbit.  This can either take the form of a dedicated LOX depot where LOX is stored until such time as sufficient quantity exists, or in the form of the LOX containers docking and transferring with the propulsive element requiring the loading.  This mission layout requires the development of fuel transfer technology for the second stage container and receiver, as well as automated docking technology.  As such fuel transfer technology must be developed for this scenario, it may make further sense to develop a dedicated LOX storage depot to aid logistics.  Such a depot could make further use of the EELV second stage technology for its basic structure.

The waypoint for such refueling should be LEO, as the bulk of the mass of a lunar mission is in escaping LEO.  What remains is the rendezvous point on the other end: lunar orbit, L1, or L2.  This determination is somewhat ancillary to the white paper as they can all be accommodated depending upon the preferred selection.

In conclusion, such an architecture offers a comprehensive lunar access program equivalent to the current mandate but with a minimum of launcher development, and can take advantage of extensive existing launch facilities and operations while greatly stimulating the use of those vehicle lines.  By allocating the launch and major propulsion duties to an organization already possessing such competency, this EELV architecture will allow for NASA to focus on their own respective capabilities in the construction of lunar mission hardware and the conduct of the lunar missions.  Working together, both can serve in the aim of returning to the moon while building a more expansive tomorrow in space.
« Last Edit: 07/02/2009 07:41 AM by libs0n »

Offline Ben the Space Brit

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Re: All EELV VSE architecture
« Reply #42 on: 07/02/2009 10:43 AM »
FWIW, most EELV-only (and those based on other ~20t-to-LEO launchers) feel a lot like von Braun and Korolev's early moon mission plans before the money was available for super-launchers.  Basically, these were ambitious EOR-LOR missions with 'space stations' assembled in Earth orbit before they set off for the Moon.

We can see this in the mission profiles earlier in this thread.  You launch a series of modules, which rendezvous in LEO. 

To me, it would seem to need something like this:

* EDS - probably a Wide-Body Centaur;
* Crew Module - Either a Bigelow Sundancer or something derived from the ATV (optional);
* Lander - An Altair is too big, so we would need something that could fit in a 5.4m fairing and fall into the predicted Phase 1 40mt to LEO payload limit; this implies a 2- or 3-man surface crew;
* Crew Return Vehicle - Dragon-D or Orion would both do the job.

With the trans-hab, the CRV isn't automatically required, but subtracting it requires the development of some system from braking into a stable earth orbit from trans-lunar return orbit.

I would say that, dependent on CRV size, the trans-hab definately isn't automatically needed.  You are flying a four-man crew to the Moon (I'm assuming that we do without the cost of developing autonomous loiter software for the orbiter) and dropping them onto the surface for a ~7 day excursion or just waiting in orbit until the previous outpost crew returns to LLO for rendezvous. 

It occurs to me that you could get more lander into orbit if, as Danny suggests, the lander could be fuelled in LEO rather than launched 'wet'.  Mabye a modified 'Dragon-F' or modified OSC Cygnus with the cargo cabin replaced with LCH4 and LOX (or hypergolic fuel) tanks could be used as a IFR vehicle.

Depending on whether you could get the WBC into LEO with enough fuel for TLI, this is a four- or five-launch mission.

Mission Profile

1) EDS to LEO
1a) IFR spacecraft 2 to refuel EDS (if required)
2) Lander to LEO - Telerobotic rendezvous & docking with EDS
3) IFR spacecraft to refuel lander
4) Crew vehicle to LEO

Vehicles rendezvous to form 'train' in LEO.  Standard transfer orbit profile to LLO.  EDS expended after TLI.  LOI handled by lander MPS (third 'orbital manoeuvring' stage on lander expended after LOI?).  Three of four crew descend to surface for excursion or outpost crew rotation.  ROI carried out by CRV.  Re-entry direct from return orbit.

Steps 1-3 can also be used to launch robotic cargo landers.  This can either be used for outpost logistics or for landing extra excursion supplies and equipment ahead of manned landing for extended surface endurance.

NEO Variant
The only significant change is the inclusion of a fourth module (possibly a double-ender, MPLM-derived cargo pod or something derived from the ATV or the Japanese HTV) to increase vehicle endurance for the longer transfer orbit to NEO.  Otherwise, I understand that dV requirements are similar, so there is no reason why the WBC-based EDS cannot be used.

Commercial Synergies
* 'COTS-F' - Commercial refuelling contract for EDS and lander tanking (Initial tanking will have to be by NASA-developed robot tanker);
* 'COTS-L' - Commercial Crew Return Vehicle contract available (to replace/suppliment Orion-Lite) after contractor demonstrates vehicle can handle free-return orbit and direct re-entry from return orbit;
* 'COTS-L1/-L2' - Commercial logistical delivery to outpost using own landers.

Ultimately, one option arising from this archetecture would be to develop aerobraking from return orbit and replacing CRVs as the orbiter with a trans-hab.  Recycling the orbiter and, possibly, the lander several times may reduced overall program costs over the longer-term.  This will allow for COTS-D service providers to launch lunar crews without needing to develop and deploy lunar-rated vehicles.
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Offline mmeijeri

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Re: All EELV VSE architecture
« Reply #43 on: 07/02/2009 10:48 AM »
Phase I only refers to Atlas. There is no comparable "phasing" for Delta IV

There's a combined set of phases in ULA's presentation to the Augustine commission. Phase 1 appears to be a new common upper stage.
« Last Edit: 07/02/2009 11:00 AM by mmeijeri »
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Offline Danny Dot

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Re: All EELV VSE architecture
« Reply #44 on: 07/02/2009 01:36 PM »
@Butters

I am thinking current EELV to lift the "current" Orion to ISS and maybe lunar.  We have data to show this can be done.

Then Phase I for the rest of the stuff.  I found an earlier study that the Delta can get up to about 50t without having to increase core size.  I haven't found Phase I for the Atlas.  I am leaning to this 50t to LEO for lunar missions.

In flight refueling of hypergolics is also on the table, but very skeptical of cryo inflight refueling.

The theme is to keep development cost down.

Direct/EELV is a good option, but it belongs in with the Direct team in the Direct thread.

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Offline Danny Dot

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Re: All EELV VSE architecture
« Reply #45 on: 07/02/2009 01:39 PM »
Phase I only refers to Atlas. There is no comparable "phasing" for Delta IV

There's a combined set of phases in ULA's presentation to the Augustine commission. Phase 1 appears to be a new common upper stage.

Come to think about it, I didn't see the Delta data use the term Phase I?  But they did have a nice chart of future growth.  50t looks like a reasonable number for an upgraded Delta.

Do you know what the mass to LEO is for a Phase I Atlas Heavy?

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Offline Danny Dot

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Re: All EELV VSE architecture
« Reply #46 on: 07/02/2009 01:48 PM »

snip

What am I missing that makes cryo fuel transfer so scary?

For the reasons you state  :D

Spinning will not work, because the fuel is on both sides of the c.g.

A settling rocket would, but I don't think it has ever been done.  The smaller the rocket, the smaller the flow rate.  Rockets take a lot of design work and use prop. 

Cryo can not use diaphragms.  Cryo transfer takes lots of work on how the lines react to the cold.  Liquid Hydrogen has a history of being a pain in the butt to handle.

It is a high risk development.  Could it be done with enough time and money?  Yes. 

The whole point of this effort is to minimize risk and development dollars.

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Offline mmeijeri

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Re: All EELV VSE architecture
« Reply #47 on: 07/02/2009 02:16 PM »
Come to think about it, I didn't see the Delta data use the term Phase I?  But they did have a nice chart of future growth.  50t looks like a reasonable number for an upgraded Delta.

The pre-ESAS and pre-ULA proposals for EELV upgrades were divided into phases for Atlas whereas Delta had a long list of incremental upgrades. Page 9 of the ULA presentation has combined these plans. It looks as if the old Atlas Phase 1 together with all the incremental upgrades to Delta have been combined into a new Phase 1. It would make sense if this phase included a new common upper stage, which is a possibility that did not exist before ULA was created.

Quote
Do you know what the mass to LEO is for a Phase I Atlas Heavy?

 According to the presentation it is 40mT.
« Last Edit: 07/02/2009 06:29 PM by mmeijeri »
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Offline Ben the Space Brit

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Re: All EELV VSE architecture
« Reply #48 on: 07/02/2009 02:17 PM »
I haven't found Phase I for the Atlas.  I am leaning to this 50t to LEO for lunar missions.

From what I have seen, Atlas-V Phase 1 is the current Atlas-V CCB with a WBC in place of the current version of Centaur.  The estimated payload weights go from 9t for the 'vanilla' 501 model to 40t for the 'heavy'.

IIRC, 50t for Delta-IVH is RS-68A and WBC. 

Jim will correct me if I'm wrong, I have no doubt.
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Offline Danny Dot

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Re: All EELV VSE architecture
« Reply #49 on: 07/02/2009 02:25 PM »
I am thinking 50t to LEO.  Anybody object.  This will avoid going to larger diameter cores.

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Re: All EELV VSE architecture
« Reply #50 on: 07/02/2009 02:30 PM »
Danny,
I'm not trying to trivialize the effort needed in any way, but I wanted to provide some insight into this stuff (since I'm probably closer to some of the key players in the propellant depot field than most).



Spinning will not work, because the fuel is on both sides of the c.g.

Ok, before I start, what our friend is talking about here is what ULA and LM called "settled" cryo handling--ie finding some way to provide a sufficiently unbalancing force to cause the propellant to assume a desired orientation to enable either long duration storage (ie by keeping the liquid away from the vents when you need to relieve pressure, or by keeping the liquid in one place and the gas in another for doing one of the active cooling ideas), or transfer (if the gas and liquid are both where they should be, you can do rapid vented transfers, or you can do no-vent transfers using the two-pipe system someone here explained--though that requires some sort of pump or compressor).

So, for this first one, spinning, it's actually useful for at least part of the cryo-fluid management problem--long term storage.  In fact, ULA is going to be flight testing the concept next month (August).  They have a, I think, DMSP satellite going up that's only going to use about 3/4 of the Centaur propellant load (leaving over 11klb of useful propellant in the stage after the mission).  Instead of venting this all overboard immediately, they're going to do some experiments (as they have been doing in the past for several decades).  They want to demonstrate transition from propulsive settling to spin settling and back and forth several times.  So, this isn't a crazy idea, and we should have interesting data back on this in the very near future.

Quote
A settling rocket would, but I don't think it has ever been done.  The smaller the rocket, the smaller the flow rate.  Rockets take a lot of design work and use prop.

Actually, this is by far the most mature settling technique.  Just about every cryo upper stage in history uses this technique (Centaur, SIV-B, DIV-US, etc).  We literally have hundreds of flights worth of experience using this, and tweaking it.  Read: http://www.ulalaunch.com/docs/publications/SettledCryogenicPropellantTransfer.pdf

Quote
Cryo can not use diaphragms.

The research I read was that they had options that could work for a limited number of tries, but I agree this isn't the best approach.

You guys left out my personal favorite approach: magnetic propellant positioning.  LOX is paramagnetic.  Hydrogen, Methane, and Propane are all mildly diamagnetic.  LOX is sufficiently paramagnetic that you can pick a little of it up with a strong permanent magnet on earth, in a 1G field.  The diamagnetic side is trickier, but there should be some data on that this year (based on analysis calibrated against some previous flight experiments).  This one is a lot less mature than the propulsive settling, but it gives you the benefit that it simplifies passive storage, transfer, and active cooling.  And once you've figured out how to do it for hydrogen, it actually works even better for methane and propane (which are good reusable lander fuels).

Quote
Cryo transfer takes lots of work on how the lines react to the cold.  Liquid Hydrogen has a history of being a pain in the butt to handle.

True, but a lot of that is stuff that upper stage guys had to figure out to get their stages to work--we're not starting from scratch here.  We have over 150 centaur flights, about a dozen saturn flights, and several cryo delta flights that all are contributing their experience base.

Quote
It is a high risk development.  Could it be done with enough time and money?  Yes. 

The whole point of this effort is to minimize risk and development dollars.

With the technologies we have now, there are ways to flight test out the last few remaining questions inexpensively.  ULA has been working on several cryo testbed options over the years, for instance.  The last one that they've publicly talked about was the Centaur Test Bed:

http://www.ulalaunch.com/docs/publications/CentaurTestBedCTBforCryogenicFluidManagement20064603.pdf

Stuff like this allows you to fly those CFM experiments as secondary payloads on Atlas flights (and if I have my way, I'm trying to make sure our suborbital vehicle is designed to handle their testbed concepts as well). 

If this was a priority, I think you could be to the point where you could do a demo 1st gen depot (wouldn't be as fancy as what we'd want in the long run, but enough to demonstrate the concept and to be useful for early operations) within a couple of years and for about what DoD spent on Orbital Express (ie about half a billion including launch costs).

I'll repeat it again--we have far more experience on cryo depots today than we did on orbital rendezvous and docking when Apollo had to make their mission mode decision.  They could've taken the "low technical risk" option, and gone for Direct Ascent, but it would've killed their budget trying to build bigger launchers. They were smart enough to take a calculated risk, and systematically beat that risk into the ground.  If NASA isn't capable of managing risk like that anymore, they're not really ever going to make it beyond LEO in a sustainable fashion.

~Jon

Offline mmeijeri

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Re: All EELV VSE architecture
« Reply #51 on: 07/02/2009 02:32 PM »
50mT could be just enough to get an Orion to L1 in one launch, provided it doesn't get much heavier. That's a potentially interesting mission scenario in itself, provided you have something to do there. Radiation experiments are my standby example.

Topping up Orion propellant from a depot would help even if performance turns out to be less or if Orion turns out to be heavier, but then you lose some of the benefit of single-launch. Staging at L2 and using 12 day trajectories should be possible even without refueling, though without an Altair I'd be wary of doing that. It would likely also make for a harder sell, especially since some will no doubt seize on this.
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Offline mmeijeri

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Re: All EELV VSE architecture
« Reply #52 on: 07/02/2009 02:37 PM »
A little over 30mT is probably enough if you stage at the ISS, which would be my preference. It reduces schedule pressure and risk to crew through boil-off associated with the cryogenic EDS. The crew could launch early and wait at the ISS. It also gets the ISS into the lunar program, which is something ESA has asked for and something that helps demonstrate to Congress and the public that all that money was spent wisely.

With ISS staging you could even consider launching the Orion unmanned so you could fly more aggressive trajectories if needed. It would also allow commercial or international players to launch astronauts to LEO even for lunar missions. I imagine this might help with national prestige for the Russians. They might prefer to launch astronauts to an ISS staging point with Soyuz and then to contribute propellant to a joint Orion flight beyond LEO since that would make their own role bigger.
« Last Edit: 07/02/2009 03:01 PM by mmeijeri »
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Re: All EELV VSE architecture
« Reply #53 on: 07/02/2009 02:56 PM »
As you will know I have a strong preference for immediate work on hypergolic depots and use of a hypergolic lander, initially in the form of in-flight refueling (or even one time fueling if that's easier) as suggested by Danny. This would get (makeshift) depots and commercial launchers into the picture immediately.

Having said that, it would be wise to pick an architecture that could make immediate use of cryogenic depots once they became available, without putting them on the critical path and while stimulating their commercial development straight away. Topping up cryogenic upper stages and using them as EDSs or even reusable transfer stages seems like the most logical initial application. Cryogenic landers could be a later application, although I'd like to see reusable landers first, which would likely be easier with hypergolics. And with reusable hypergolic landers / cis-lunar shuttles you could have a sizeable market for hypergolic propellant both in LEO and at L1/L2 long before there was a moon base. This by itself might spur ULA to develop a refuelable cryogenic upper stage if that makes propellant transfer to L1/L2 more efficient.

To stimulate rapid development of cryogenic depots more than just lip service would be needed. Things like the Centaur testbed flight should be fully funded. Other than that, I would like to see NASA work with industry to define standardised, open interfaces for depots, fund some research, commit to buying propellant or orbit and then get out of the way as was suggested by Jon. Interestingly, some work on standardised interfaces has already been done by Moog. I wouldn't want to see NASA in the cryogenic depot development business any more than I would want to see them in the launcher business. ULA/Boeing/LM are the experts when it comes to cryogenic fluid management. On the other hand, I would have no problem with NASA involving itself with more experimental technologies such as magnetic propellant positioning.
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Re: All EELV VSE architecture
« Reply #54 on: 07/02/2009 03:12 PM »
50mT could be just enough to get an Orion to L1 in one launch, provided it doesn't get much heavier. That's a potentially interesting mission scenario in itself, provided you have something to do there. Radiation experiments are my standby example.

Topping up Orion propellant from a depot would help even if performance turns out to be less or if Orion turns out to be heavier, but then you lose some of the benefit of single-launch. Staging at L2 and using 12 day trajectories should be possible even without refueling, though without an Altair I'd be wary of doing that. It would likely also make for a harder sell, especially since some will no doubt seize on this.

I guess my big concern here is that EELVs + depots are fundamentally different than HLVs.  Sure, you can do a "black aluminum" design (to borrow a composites phrase), but it's not going to look very good compared to actually taking advantage of commercial launch + depots' strong points.

For instance, take Orion.  Orion's CM is actually only a little heavier than the Apollo CM (18klb vs 11klb), in spite of being much bigger.  Some of that's bloated a bit, but it's not too crappy compared to previous work.  The reason Orion is so big is that its SM is still huge.  doing drylaunch like you suggest, or even going to a different mission mode (say staging some things out of L1/L2) can completely change the dV requirements on Orion, precipitating a lot of good positive changes.  If Orion is going to L2, it needs less than half the dV capacity...

Anyhow, we're flight testing today, and I need to head down to the shop.  Just saying though, if your EELV/depot architecture ends up looking very similar to the ESAS architecture, there's a good chance you've still nowhere close to an optimal design.

~Jon

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Re: All EELV VSE architecture
« Reply #55 on: 07/02/2009 03:30 PM »
Spinning will not work, because the fuel is on both sides of the c.g.

Liquid Hydrogen has a history of being a pain in the butt to handle.

Most of the propellant mass is the LOX. LH2 is extra painful to transfer.  For these reasons LOX is the low-hanging fruit so the first cryo propellant transfer should transfer LOX only. Don't discount LOX transfer just because LH2 transfer is hard! Launch the EDS with empty LOX tank but full LH2 tank.

Attach the donor and recipient spacecraft, which are cylindrical, top to top with a cable and a LOX hose and spin them around their common center of gravity. The axis of rotation should be perpendicular to their common axis of symmetry, so the centrifugal (pseudo-)force pushes roughly towards the bottoms of the tanks. Put the LOX outlet at the bottom of the donor tank. The majority of the mass of the system with the full donor LOX tank, empty recipient LOX tank and already full "recipient" LH2 tank is the LOX, so initially the c.g. would probably be inside the LOX tank. This looks like a problem since part of the LOX doesn't look like it will flow into the outlet, but this problem goes away over time. As you transfer the LOX that you have access to, which is at least half of the total, the c.g. will shift into the recipient vehicle and the previously inaccessible LOX will become accessible.

There's a potential risk that a lot of LOX might shift from one end of the tank to the other at once and damage something. Here's why I don't expect that to happen. The surface of the LOX will always be a portion of a cylinder with axis equal to the axis of rotation (passing through the c.g.). Since the tanks are also cylindrical (with a perpendicular axis) the only LOX that will fall is from the two points on the surface of the tank on opposing sides of the c.g. There isn't much LOX near those points because of curvature of the tank and LOX surface, so even a shock shouldn't transfer much LOX. If a shock causes too much LOX to fall from the top of the tank to the bottom the c.g. will shift towards the bottom the the tank, which reduces the tendency of additional LOX to fall. So the amount of LOX on top of the tank seems stable. Of course my ability to do fluid dynamics in rotating frames of reference in my head is rather limited so this could easily be wrong.

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Re: All EELV VSE architecture
« Reply #56 on: 07/02/2009 03:37 PM »
I've been mostly thinking about L1 (and more recently MEO and GEO) as staging points, not L2. And even L1 requires less delta-v than LLO. On the other hand I would want to have fast any time return, since I'm mostly interested in using Orion as a transporter to/from the nearest staging point and in using it as an escape pod. I'm not sure the scenario described in your most recent blog post allows for a quick return. Also, if you want to do operations in MEO and GEO, you really need the delta-v because deorbit burns from there are so expensive. Even the current 1.4 km/s is a bit tight. With this capability Orion (in combination with Altair) could safely operate anywhere from LEO to the Sun Earth Lagrange points and even somewhat beyond (NEOs) with a six person crew.

On the other hand, I'm all for commercial Dragon-sized or Soyuz-sized crew capsules in an L1/L2 COTS-D/CRS.
« Last Edit: 07/02/2009 04:30 PM by mmeijeri »
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Re: All EELV VSE architecture
« Reply #57 on: 07/02/2009 03:41 PM »
For these reasons LOX is the low-hanging fruit so the first cryo propellant transfer should transfer LOX only. Don't discount LOX transfer just because LH2 transfer is hard! Launch the EDS with empty LOX tank but full LH2 tank.

Or, as Jon suggested the other day, LOX and methane, or even LOX and kerosene. Jon also suggested using L1/L2 because of its better thermal environment. The low delta-v to most interesting places also helps compensate for any loss in efficiency due to lower Isp.
We will be vic-toooooo-ri-ous!!!

Offline Danny Dot

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Re: All EELV VSE architecture
« Reply #58 on: 07/02/2009 03:54 PM »
We are looking for numbers and details.

How many pounds of settling prop to transer X tons of cryo?

Show me the report of a test of a diaphram that works at cryo temps.

Show me the design of an on-orbit hook up of  LH2 or LO2.

Why did NASA reject LO2 methane for Orion and why is it OK now?

Delta Vs and trip times for different lunar trajectories?

Etc.

Danny Deger
Danny Deger

Offline mmeijeri

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Re: All EELV VSE architecture
« Reply #59 on: 07/02/2009 04:07 PM »
Is there anything that argues against a multi-track approach of both cryogenic depots and noncryogenic ones? I'm getting the feeling that Jon would be worried about being stuck with MMH/NTO depots even when cryogenic ones turn out not to be all that more difficult and Danny would be worried about cost and risk of cryogenic depots/in-flight refueling.

With NASA being only peripherally involved in the development of cryogenic depots but well-positioned to take advantage of it, a secondary track would not need to bust the budget. And note that cryogenic depots only matter for getting commercial players beyond LEO. Having depots at all and needing a high flight-rate to fill them would still help with the development of cheaper launchers and thus reducing the cost of getting into LEO. For commercial development of space that would be the more important strategic consideration.
We will be vic-toooooo-ri-ous!!!

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