Reusable earth departure stages

[cordwainer]

It might make more sense to have a more long term mission with two different vehicles, one to deliver and take care of the astronauts for several months on the surface and another to pick the astronauts up several months down the line. This would get rid of the complexities of a manned orbiter and allow one to carry sufficient habitat payload for a more extensive period on Mars.

[cordwainer]

As to staging platforms at EML1 and 2 and the use of fuel depoting. I do wonder how feasible it would be to use such techniques to actually fly something like a Skylon SSTO to lunar orbit? I know Skylon is not designed for that and such a long flight regime would require additional robustness against radiation and micro-meteorites for such an SSTO to be reusable, would it even be possible to engineer?

Perhaps, you could rework some Skylons while on orbit into dedicated cislunar shuttles, so they don't have to re-enter the atmosphere as part of their flight regime. Piggy-back lunar landers and cargo delivery modules onto the Skylons to build a Moonbase.

It seems to me that if you want to reduce the cost of space operations then you need to put as many mission functions into one module or vehicle rather than build a bunch of different vehicles.

[Hop_David]

One problem would be that the EDS's engines would be pointed exactly at the departing payload for the braking burn.  So you would need to wait long enough for the stage to move away from the vicinity of the payload so it would be safe to start its engines.  That might take a few minutes depending on the strength of your RCS thrusters.

I hadn't thought of that.

Of course the EDS needs to be separated from it's payload. Is there a separation method that would push the payload forward as well as pushing the EDS backward?

Such a push might give a little distance between payload and EDS in the time it takes for the EDS to turn 180º. The push might also help with the delta V, both for accelerating the payload and decelerating the EDS.

[aero]

One problem would be that the EDS's engines would be pointed exactly at the departing payload for the braking burn.  So you would need to wait long enough for the stage to move away from the vicinity of the payload so it would be safe to start its engines.  That might take a few minutes depending on the strength of your RCS thrusters.

I hadn't thought of that.

Of course the EDS needs to be separated from it's payload. Is there a separation method that would push the payload forward as well as pushing the EDS backward?

Such a push might give a little distance between payload and EDS in the time it takes for the EDS to turn 180º. The push might also help with the delta V, both for accelerating the payload and decelerating the EDS.

Will the return trajectory start at exactly 180 degrees reversal? After all, the Earth has moved a little and firing just 20 degrees off would both slow and move the vehicle laterally. Separation would happen very quickly once the engines fired after which orientation wouldn't be a concern.

[jongoff]

One problem would be that the EDS's engines would be pointed exactly at the departing payload for the braking burn.  So you would need to wait long enough for the stage to move away from the vicinity of the payload so it would be safe to start its engines.  That might take a few minutes depending on the strength of your RCS thrusters.

I hadn't thought of that.

Of course the EDS needs to be separated from it's payload. Is there a separation method that would push the payload forward as well as pushing the EDS backward?

Such a push might give a little distance between payload and EDS in the time it takes for the EDS to turn 180º. The push might also help with the delta V, both for accelerating the payload and decelerating the EDS.

Some possibilities:

1- Use retrorockets on the EDS to push back quickly.

2- It might also be possible for the EDS to fire a little off of the velocity vector (to avoid plume impingement), and then cancel out that off-vector component as soon as there's sufficient space with the payload.

3- Or if the EDS has 2+ engines, you might be able to splay them outward and just take some cosine losses.

4- Or if Magnetoshell Aerocapture works and scales right, maybe you could just turn on the "deflector shield" once the payload has a little distance.

There are probably other options, but those are the first three that come to mind.

~Jon

[savuporo]

Total trip time from first to last perilune is 54 days. Perilune to EML2 is about 3 days -- so add 6 days for a total of 60 days. Is two months too long for oxygen/hydrogen?

As of today, yes, far too much. I may be wrong,  but i seem to remember that Centaur record loiter time was counted in days, not weeks.

EDIT: actually, correction, demonstrated record durations are counted in hours, 8 or so.

[Hop_David]

One problem would be that the EDS's engines would be pointed exactly at the departing payload for the braking burn.  So you would need to wait long enough for the stage to move away from the vicinity of the payload so it would be safe to start its engines.  That might take a few minutes depending on the strength of your RCS thrusters.

I hadn't thought of that.

Of course the EDS needs to be separated from it's payload. Is there a separation method that would push the payload forward as well as pushing the EDS backward?

Such a push might give a little distance between payload and EDS in the time it takes for the EDS to turn 180º. The push might also help with the delta V, both for accelerating the payload and decelerating the EDS.

Will the return trajectory start at exactly 180 degrees reversal? After all, the Earth has moved a little and firing just 20 degrees off would both slow and move the vehicle laterally. Separation would happen very quickly once the engines fired after which orientation wouldn't be a concern.

Good point. Jon Goff mentioned Centaurs and I've found dry mass, propellent mass and newtons thrust of a Centaur. Knowing the newtons gives me a handle on how long accelerations would take for different masses. However elsewhere in this thread someone mentioned a long duration round trip might have hydrogen boil off problems and the shortest round trip I've found so far is 60 days.

If hydrogen is out I'll have to plug in a different exhaust velocity to the rocket equation although my thrust might be improved. When I get a better handle on burn times, I'll have a better idea at what longitude the EDS starts its braking burn. I believe you're right, the needed rotation might be less than 180.

[DarkenedOne]

The thing is that reusable chemical EDS do not make any sense so long as it would have to be refueled by a expendable rocket.  When you compare an architecture that utilizes a reusable chemical EDS that is refueled by an expendable rocket to an architecture that utilizes an expendable chemical EDS that is launched by an expendable rocket the expendable system comes out superior. 

Chemical reusable EDS stages would only make sense if the cost of fuel in LEO is rather low.  There are a number of systems that could theoretically make that happen.  One would be gun launch.  Another would be some form of space based ISRU.

Now for EDS that use high ISP propulsion systems like nuclear thermal, nuclear electric, solar thermal, and solar electric it would make no sense for them not to be reusable.  The cost of refueling them even with an expendable launch system would be lower than the cost of replacing them. 

[Hop_David]

[quote author=Hop_David link=topic=34822.msg1206963#msg1206963 date=14014671
Some possibilities:

1- Use retrorockets on the EDS to push back quickly.

2- It might also be possible for the EDS to fire a little off of the velocity vector (to avoid plume impingement), and then cancel out that off-vector component as soon as there's sufficient space with the payload.

3- Or if the EDS has 2+ engines, you might be able to splay them outward and just take some cosine losses.

4- Or if Magnetoshell Aerocapture works and scales right, maybe you could just turn on the "deflector shield" once the payload has a little distance.

There are probably other options, but those are the first three that come to mind.

~Jon

The EDS would do all it's work in vacuum so no need to make it aerodynamic.

Could the attitude jets be put on arms and pointing an opposite direction to the main engine?



Putting them on arms would give them more torque.

[Hop_David]

The thing is that reusable chemical EDS do not make any sense so long as it would have to be refueled by a expendable rocket.

See the OP. I imagine the EML2 staging platform being supplied by either carbonaceous asteroids in lunar orbit or volatiles from the lunar poles.

Delta V from moon to EML2 is about 2.5 km/s. Delta V from a rock in lunar orbit would be even less.

With this sort of delta V budget, reusable tankers to supply the platform is plausible.

[Chris Bergin]

Please do not embed massive images. If they break the forum width the post will be removed as we can't spend all day editing posts.

Attach.

[muomega0]

One problem would be that the EDS's engines would be pointed exactly at the departing payload for the braking burn.  So you would need to wait long enough for the stage to move away from the vicinity of the payload so it would be safe to start its engines.  That might take a few minutes depending on the strength of your RCS thrusters.

I hadn't thought of that.

Of course the EDS needs to be separated from it's payload. Is there a separation method that would push the payload forward as well as pushing the EDS backward?

Such a push might give a little distance between payload and EDS in the time it takes for the EDS to turn 180º. The push might also help with the delta V, both for accelerating the payload and decelerating the EDS.

Will the return trajectory start at exactly 180 degrees reversal? After all, the Earth has moved a little and firing just 20 degrees off would both slow and move the vehicle laterally. Separation would happen very quickly once the engines fired after which orientation wouldn't be a concern.

Good point. Jon Goff mentioned Centaurs and I've found dry mass, propellent mass and newtons thrust of a Centaur. Knowing the newtons gives me a handle on how long accelerations would take for different masses. However elsewhere in this thread someone mentioned a long duration round trip might have hydrogen boil off problems and the shortest round trip I've found so far is 60 days.

If hydrogen is out I'll have to plug in a different exhaust velocity to the rocket equation although my thrust might be improved. When I get a better handle on burn times, I'll have a better idea at what longitude the EDS starts its braking burn. I believe you're right, the needed rotation might be less than 180.

Centaur Upper Stage Applicability for Several Day Mission Durations with Minor Insulation Modifications
The LEO boiloff rates for vary by MLI layers and type of propellant.

 3 layer MLI     LOX    2  %/day    LH2: 4-5 %/day    based on 1 hour hold  Table 2 and Table 4 :
20 layer MLI    LOX   0.8%/day     LH2  2.5%/day      + 100lbs 

The ULA depot adds a conical sunshield to the transfer stage, which brings these rates down an order of magnitude, perhaps 0.1%/day for LH2 away from LEO.
30 to 60 days * .25%/day is roughly  6 to 12% of initial LH2, or 3 to 6% at 0.1%/day.
Longer durations eventually require solar arrays and cryocoolers in the trade.

[A_M_Swallow]

Good point. Jon Goff mentioned Centaurs and I've found dry mass, propellent mass and newtons thrust of a Centaur. Knowing the newtons gives me a handle on how long accelerations would take for different masses. However elsewhere in this thread someone mentioned a long duration round trip might have hydrogen boil off problems and the shortest round trip I've found so far is 60 days.

If hydrogen is out I'll have to plug in a different exhaust velocity to the rocket equation although my thrust might be improved. When I get a better handle on burn times, I'll have a better idea at what longitude the EDS starts its braking burn. I believe you're right, the needed rotation might be less than 180.

If you are not using hydrogen due to boiloff problems then the EDS can use the same engines as the lander.  Same propellant and a common pool of replacement parts will simplify the logistics.  Possibilities include Super Draco (NTO/MMH, Isp 235) and Morpheus HD5 (methane/LOX, Isp 321).

[savuporo]

The ULA depot adds a conical sunshield to the transfer stage, which brings these rates down an order of magnitude, perhaps 0.1%/day for LH2 away from LEO....

Its not just about boil off, long duration loiter for Centaur involves multiple other adjustments that need to be made, including things like batteries.

[jongoff]

The ULA depot adds a conical sunshield to the transfer stage, which brings these rates down an order of magnitude, perhaps 0.1%/day for LH2 away from LEO....

Its not just about boil off, long duration loiter for Centaur involves multiple other adjustments that need to be made, including things like batteries.

That's part of what the whole Integrated Vehicle Fluids project is about. It replaces the batteries, the hydrazine thrusters (used for settling and ACS), and the helium pressurization (for repressurizing the tanks prior to a burn) with their IVF system. It taps boiled-off GOX/GH2 from the tanks to run a small internal combustion engine, which recharges the batteries, and runs a compressor for boosting the pressure of the GOX/GH2 prior to warming it for autogenous pressurization. With IVF you can run the stage as long as there is LOX and LH2 left in the tank. They had a lot of the prototype hardware for it at the Space Symposium. It's looking like they were going to fly part of it in 2015, and the rest of it in I think 2016. Once it's there, the duration of Centaur goes way up, the dry weight goes down quite a bit, and refuelability becomes easier since you're just dealing with two fluids.

Combine that with the improved insulation (MLI or a sun-shield), and using the rotational settling, and there's no reason you couldn't handle months-long missions.

~Jon

[jongoff]

The thing is that reusable chemical EDS do not make any sense so long as it would have to be refueled by a expendable rocket.  When you compare an architecture that utilizes a reusable chemical EDS that is refueled by an expendable rocket to an architecture that utilizes an expendable chemical EDS that is launched by an expendable rocket the expendable system comes out superior. 

Chemical reusable EDS stages would only make sense if the cost of fuel in LEO is rather low.  There are a number of systems that could theoretically make that happen.  One would be gun launch.  Another would be some form of space based ISRU.

Now for EDS that use high ISP propulsion systems like nuclear thermal, nuclear electric, solar thermal, and solar electric it would make no sense for them not to be reusable.  The cost of refueling them even with an expendable launch system would be lower than the cost of replacing them. 

Darkened One,

Have you actually run numbers on this, or are you stating your opinion? Not trying to be rude, just curious.

~Jon

[Robotbeat]

Well, if you reuse the EDS you can have it accelerate several payloads into Mars transfer orbit in one window. You can use it for sending stuff to the moon or to near earth asteroids as well.

My inner manufacturing engineer is a big fan of getting more "inventory turns" on your expensive hardware than once every two years. It would be interesting to see if you could find a way to enable multiple Mars departures in a single launch window with a single reusable EDS...

~Jon

This. And I don't see why you couldn't. The Mars window is somewhat flexible, especially if you're planning this ahead of time. If you can refuel quickly, I don't see why you couldn't get, say, a dozen launched in a single window with a single EDS.


...By the way, a reusable EDS (or a reusable first stage for that matter) is very, VERY sensitive to mass ratio. If you have a good enough mass ratio, then the retro burn consumes very little propellant. I'm sure Jon's quite aware of this, just making a general point. Isp can stay the same, for all we care.

Here is an example: Suppose you need 2km/s, you have a 4km/s exhaust velocity and have a stage full/empty ratio of 5:1. That means you can handle 5 tons of payload for every 5 tons of fueled stage mass. But if you can decrease the empty mass of your stage such that you have a stage full/empty ratio of, say, 10:1, you can actually carry plenty of propellant to do a full retro burn with more than 700m/s left over for docking, etc. For the same payload and the same initial mass.

So dry mass reductions can make a huge difference in many (especially reusable) architectures. Balloon tanks and turbopumps with a dense propellant can beat a much more expensive hydrogen nuclear thermal rocket stage in many metrics.

[john smith 19]

If you are not using hydrogen due to boiloff problems then the EDS can use the same engines as the lander.  Same propellant and a common pool of replacement parts will simplify the logistics.  Possibilities include Super Draco (NTO/MMH, Isp 235) and Morpheus HD5 (methane/LOX, Isp 321).

You might like to revise that Isp for NTO/MMH for Super Draco as Musk said the the chamber pressure is about 1000psi.

[A_M_Swallow]

If you are not using hydrogen due to boiloff problems then the EDS can use the same engines as the lander.  Same propellant and a common pool of replacement parts will simplify the logistics.  Possibilities include Super Draco (NTO/MMH, Isp 235) and Morpheus HD5 (methane/LOX, Isp 321).

You might like to revise that Isp for NTO/MMH for Super Draco as Musk said the the chamber pressure is about 1000psi.

The Appendix A Noise to Draft Environmental Assessment for Issue an Experimental Permit to SpaceX for Operation of the Dragon Vehicle at the McGregor Test Site, Texas - published May 2014 gives the Super Draco exhaust velocity as 2,300 m/s (7,546 ft/s) on page 12.

Checking the maths 2300 / 9.81 = 234.45

If you think the figure in the report is err ... out of date I will leave you to get it changed.

http://www.faa.gov/about/office_org/headquarters_offices/ast/media/20140513_DragonFly_DraftEA_Appendices%28reduced%29.pdf

[Nilof]

Yes, but that is most likely the sea level Isp, and even there the superdraco looked underexpanded during testing, most likely to keep it compact.

If you gave it a vaccum-optimized nozzle to allow the exhaust to expand properly in vaccum you could probably squeeze out at least 270s.