ISS-based cryogenic third stages as expendable Earth-Moon tugs

[Archibald]

As the (horrible, I know) title says.

We suppose an international company. That company buy bulks of cryogenic third stages from countries willing to sell them (list below, from memory)
- USA Centaur and Delta IV second stages
- ESA HM-7 or Vinci (if build one day...)
- Russia / India RD-56
- JAXA LE-5B
- China YF-75
(and perhaps, one day - SpaceX Raptor)
The company evidently also buy a rocket ride to loft these stages into low earth orbit. Evidently no payload is carried; no GEO satellite.

Meanwhile, lunar heavy payloads (15- 20 tons) are launched and dock to the ISS.
The company provide the cryogenic stages with Soyuzautomated rendezvous and docking gear (Kurs + probe-and-drogue).
Then a booster loft the cryogenic third stage near the ISS no fly zone, the payload disengage from the space station, dock to the stage, fire, head to L1 / L2 / LLO.

This a mere expension of CSI Soyuz/ block D scheme that involve the ISS into a lunar flight.

By using the ISS for a lunar program, we don't have to wait 2028 and the end of its useful life to return beyond LEO...

Thoughts ?

[Jason1701]

TRL for cryo storage is too low without extensive development work.

ISS is built for specific and low load paths. A bunch of RL-10s would tear it apart.

[pathfinder_01]

There was a NASA plan that used Delta IV heavy's stage to push a spacecraft that was docked at the ISS to l1. It would aerobreak back to the ISS or be picked up by the shuttle.

[sdsds]

Is it true there has never been a rendezvous conducted with a cryogenic stage?  By any space program?  Ever?

[Downix]

Is it true there has never been a rendezvous conducted with a cryogenic stage?  By any space program?  Ever?

They have with emptied ones I know of.

[Archibald]

There was a NASA plan that used Delta IV heavy's stage to push a spacecraft that was docked at the ISS to l1. It would aerobreak back to the ISS or be picked up by the shuttle.

Ah - here we go. I need a link, or a pdf !  This is what I had in mind.

Only LOX/LH2 stages have enough performance margin on the difficult Earth - Moon segment (3 km/s to Earth escape, 4 km/s to low lunar orbit, with GEO and/or libration points between the two).

The best non-cryogenic stages are Block D and Breez, with a much more limited payload - a mere 8 tons.

Just reminded that Apollo 7 and 9, of course, met their S-IVBs in low Earth orbit. Heck, the Apollo 9 stage still had a big load of propellants - they later expended it into a mock TLI that send it into a heliocentric orbit.

[pathfinder_01]

http://history.nasa.gov/DPT/Architectures/Moon%20-%20L1-Moon%20Exploration%20Architecture%20DPT%20Jun_00.pdf

Here is the link.

[Jorge]

Is it true there has never been a rendezvous conducted with a cryogenic stage?  By any space program?  Ever?

They have with emptied ones I know of.

That's right. To complete the picture:

No cryogenic stage has ever been part of the "active" vehicle in a rendezvous.

No in-space cryogenic propellant transfer has ever been performed.

[Archibald]

Jorge,

can you see any peculiar problem for rendezvous and automated docking with a modified cryogenic stage? mass and size -wise, a Centaur is similar to an ATV, but the "payload" is quite different - a big load of cold propellants that might slosh...

[edkyle99]

Only LOX/LH2 stages have enough performance margin on the difficult Earth - Moon segment (3 km/s to Earth escape, 4 km/s to low lunar orbit, with GEO and/or libration points between the two).

I don't think I agree with that as an absolute statement.  A storable propellant stage could perform a trans-lunar insertion burn just as well as a cryogenic propellant stage.  It would require more propellant mass, but without boiloff the upmass to LEO difference would not be as large as might be expected.  The difference could be further reduced by using a small cryo stage, launched at the last minute, as an augment to the storable stage.

No need for ISS in either case.

 - Ed Kyle

[sdsds]

Only LOX/LH2 stages have enough performance margin on the difficult Earth - Moon segment (3 km/s to Earth escape, 4 km/s to low lunar orbit, with GEO and/or libration points between the two).

I don't think I agree with that as an absolute statement.  A storable propellant stage could perform a trans-lunar insertion burn just as well as a cryogenic propellant stage.  It would require more propellant mass [...]

If the results shown in the attached chart are correct, there's a surprisingly easy to remember sequence of (approximated) payload capabilities.  This uses simplistic dry mass assumptions, with delta-v numbers for getting to the lunar surface from ESAS:
  TLI = 3140 m/s
  Lunar = 855 + 1911 (LOI + Descent) = 2766 m/s

For a cargo spacecraft using hydrolox propulsion for all in-space maneuvers, an ESAS CaLV launcher could put 20 mT on the lunar surface.  For a spacecraft using hypergolic propulsion for all in-space maneuvers, the same launcher puts 10 mT on the surface.  The mix-and-match approach of hydrolox TLI and hypergolic lunar propulsion yields 15 mT.

10 / 15 / 20.  Curiously easy to remember!

[Archibald]

I have my own little empiric rule - that a cryogenic stage has a TLI payload roughly equal to its fully-fueled mass (in EOR mode of course).
A 15 tons Centaur has a 15 ton payload, and so forth.
Non-cryogenic (hypergolic or kerosene stages) with staged combustion and ISP around 350 - 360 seconds can place a bit less than half of their mass through TLI. Breez or Block D are around 15 - 20 tons, with a TLI payload of 8 tons+.
It's a pity there's no equivalent to Block D or Breez in other space agencies.
It is all a matter of payload, EOR style. Powerful cryogenic stages for heavy unmanned payloads, hypergolics / kerosen stages for a small manned taxi.
Soyuz and Gemini should be the models for the manned taxi. A 3500 kg Gemini housed two men for 14 days in 1966 - not bad. A lunar variant would be heavier of course, but it would also benefit from 45 years of progresses... 

[Patchouli]

TRL for cryo storage is too low without extensive development work.

ISS is built for specific and low load paths. A bunch of RL-10s would tear it apart.

ISS stays in LEO it's only used as a staging point.

If you wanted to move the ISS to L1 it would be best to simply upgrade the solar cells and fit it with several VASIMR engines as it's too heavy for even an S-IVB to push to L1.
But really it would be best to simply send a Bigelow type station to L1 if you need a L1 station.

http://history.nasa.gov/DPT/Architectures/Moon%20-%20L1-Moon%20Exploration%20Architecture%20DPT%20Jun_00.pdf

Here is the link.

Looking at this concept and how much sense it makes why did they even consider the expensive architecture they used with Constellation.

It seems we can defiantly do this and use Dragon and Orion as transfer vehicle.

I wonder how much work would it take to make the Orion SM refuelable in LEO so the Atlas 552 can be used to launch it.

[Robotbeat]

Only LOX/LH2 stages have enough performance margin on the difficult Earth - Moon segment (3 km/s to Earth escape, 4 km/s to low lunar orbit, with GEO and/or libration points between the two).

I don't think I agree with that as an absolute statement.  A storable propellant stage could perform a trans-lunar insertion burn just as well as a cryogenic propellant stage.  It would require more propellant mass [...]

If the results shown in the attached chart are correct, there's a surprisingly easy to remember sequence of (approximated) payload capabilities.  This uses simplistic dry mass assumptions, with delta-v numbers for getting to the lunar surface from ESAS:
  TLI = 3140 m/s
  Lunar = 855 + 1911 (LOI + Descent) = 2766 m/s

For a cargo spacecraft using hydrolox propulsion for all in-space maneuvers, an ESAS CaLV launcher could put 20 mT on the lunar surface.  For a spacecraft using hypergolic propulsion for all in-space maneuvers, the same launcher puts 10 mT on the surface.  The mix-and-match approach of hydrolox TLI and hypergolic lunar propulsion yields 15 mT.

10 / 15 / 20.  Curiously easy to remember!

There are problems with your assumptions, there... A hydrolox stage would weigh more than a hypergolic stage of the same thrust, since hydrogen is a lot less dense. That's not the end of the story, but your assumptions there are a little simplistic.

[libs0n]

Only LOX/LH2 stages have enough performance margin on the difficult Earth - Moon segment (3 km/s to Earth escape, 4 km/s to low lunar orbit, with GEO and/or libration points between the two).

I don't think I agree with that as an absolute statement.  A storable propellant stage could perform a trans-lunar insertion burn just as well as a cryogenic propellant stage.  It would require more propellant mass, but without boiloff the upmass to LEO difference would not be as large as might be expected.  The difference could be further reduced by using a small cryo stage, launched at the last minute, as an augment to the storable stage.

No need for ISS in either case.

 - Ed Kyle

Ed, some thoughts for you,

1. Hydrolox injection can also be taken advantage of by injecting fuel directly into a post TLI staging area using a launch vehicle capable of TLI, like the Delta 4 Heavy, and launching the lander dry, to be fueled at the staging area, thus reducing EDS requirements.

2. Forget the EDS.  Launch a dry lunar lander to L2/L1, and then send fuel to it to, or the other way around.  I reckon a dry lunar lander can fit into the TLI of a Delta 4 Heavy or 2, and storable propellant can be shaped to whatever capability.  A modest upgrade srb boosted Delta 4 Heavy can probably send a Dragon to L2/L1, or, given the timeframe of an actual mission, the Falcon Heavy could be an option. 

3. And to support your post, mass efficient hydrolox TLI insertion is irrelevant if the heavier propellant alternative can offer a better price.  Delta 4 Heavy vs Falcon Heavy.