Author Topic: Reusable earth departure stages  (Read 20503 times)

Offline Hop_David

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Reusable earth departure stages
« on: 05/26/2014 05:12 pm »
One of my favorite fantasies is staging platforms at EML1 and 2. Outbound spacecraft could go to these places to stock up on propellent, water (for drinking as well as radiation shielding), and air. These staging platforms could be supplied by volatiles from the lunar poles and/or carbonaceous asteroids parked in lunar orbit.

I imagine a spacecraft with an Earth Departure Stage (EDS) departing EML2 for a deep perigee via the 9 day Farquhar route. ~11 km/s at perigee confers a big Oberth benefit for a TMI burn or trans asteroid burn. After TMI, the EDS disengages, turns 180º and then does a braking burn. Slowing down to just below escape velocity would put the EDS on an ~ 9 day orbit. The 3rd perigee would be 27 days later and thus back in the moon's neighborhood. Then it could return to EML2 to get ready to send another spacecraft on its way.

But I was looking at a near parabolic orbit with a 400 km perigee. Precious little time is spent in the neighborhood of perigee. In the attached graphic a region of the ellipse is divided into 16 time increments, each increment about 3.4 minutes. The ship would spend about 54 minutes in the neighborhood portrayed.

Is there enough time to accelerate .6 km/s, disengage, turn 180º and then decelerate .6 km/s?

Offline savuporo

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Re: Reusable earth departure stages
« Reply #1 on: 05/26/2014 05:28 pm »
The ship would spend about 54 minutes in the neighborhood portrayed.
Is there enough time to accelerate .6 km/s, disengage, turn 180º and then decelerate .6 km/s?
An RL-10 powered stage normally has a total burn time of 14-20 minutes. So, yes ?
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Offline e of pi

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Re: Reusable earth departure stages
« Reply #2 on: 05/26/2014 05:45 pm »
At 1G, conducting a 600 m/s burn would take only about a minute. 54 minutes is thus plenty of time for even a relatively low-thrust system (excepting ion drives and such, obviously) to make a 600 m.s burn, drop the payload, and brake back down those 600 m/s. Heck, at even a tenth of a G, 5 minutes would be more than enough. How long would flipping take after dropping the payload? I'm not entirely sure, but I definitely suspect much less than 45 minutes.

Offline jongoff

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Re: Reusable earth departure stages
« Reply #3 on: 05/26/2014 06:01 pm »
But I was looking at a near parabolic orbit with a 400 km perigee. Precious little time is spent in the neighborhood of perigee. In the attached graphic a region of the ellipse is divided into 16 time increments, each increment about 3.4 minutes. The ship would spend about 54 minutes in the neighborhood portrayed.

Is there enough time to accelerate .6 km/s, disengage, turn 180º and then decelerate .6 km/s?

As others have noted, I would think so. It'll depend strongly on your system's T/W ratio though. The T/W ratio of the stage after separation (during the breaking burn) should be much higher. For instance, with a Dual Engine Centaur stage and a 60-ish tonne payload being slung on say a TMI trajectory, your T/W ratio is down around 0.25 for the departure burn (taking roughly 4min for the burn). But once you've staged, your T/W ratio is probably >1 (my super quick BOTE is saying you'd have a burnout mass on the Centaur of <9tonnes, which would give a T/W of ~2.5 for the Centaur, meaning it would only take it 24 seconds for the retro burn.

~Jon

Offline john smith 19

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Re: Reusable earth departure stages
« Reply #4 on: 05/26/2014 06:14 pm »
The Project Troy study by Reaction Engines for a Mars mission utilizes a reusable EDS.

The OP narrows the field to EML1 and 2.

The study can be found here.

http://www.reactionengines.co.uk/tech_docs/mars_troy.pdf

It's a strategy that saves a fair bit of hardware.
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Offline Hop_David

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Re: Reusable earth departure stages
« Reply #5 on: 05/26/2014 11:51 pm »
But I was looking at a near parabolic orbit with a 400 km perigee. Precious little time is spent in the neighborhood of perigee. In the attached graphic a region of the ellipse is divided into 16 time increments, each increment about 3.4 minutes. The ship would spend about 54 minutes in the neighborhood portrayed.

Is there enough time to accelerate .6 km/s, disengage, turn 180º and then decelerate .6 km/s?

As others have noted, I would think so. It'll depend strongly on your system's T/W ratio though. The T/W ratio of the stage after separation (during the breaking burn) should be much higher. For instance, with a Dual Engine Centaur stage and a 60-ish tonne payload being slung on say a TMI trajectory, your T/W ratio is down around 0.25 for the departure burn (taking roughly 4min for the burn). But once you've staged, your T/W ratio is probably >1 (my super quick BOTE is saying you'd have a burnout mass on the Centaur of <9tonnes, which would give a T/W of ~2.5 for the Centaur, meaning it would only take it 24 seconds for the retro burn.

~Jon

Thanks! Do you know the dry mass and propellent mass a Dual Engine Centaur? Newtons? If I knew those things I believe I could do my own BOTEs.

A few things still unknown:

Time it'd take to flip 180 for the braking burn.

How much delta V it would take to return to EML2. I believe it's doable to time the third apogee to be in the moon's neighborhood. But the route from an apogee in the moon's neighborhood to a halo around EML2 still hasn't coalesced in my imagination.
« Last Edit: 05/26/2014 11:58 pm by Hop_David »

Offline RanulfC

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Re: Reusable earth departure stages
« Reply #6 on: 05/27/2014 02:04 pm »
Thanks! Do you know the dry mass and propellent mass a Dual Engine Centaur? Newtons? If I knew those things I believe I could do my own BOTEs.

I don't think there's been much "official" information on the DEC, these websites might help though:
http://alternatewars.com/BBOW/Boosters/Centaur/Centaur_GIW.htm
http://alternatewars.com/BBOW/Boosters/Centaur/Centaur.htm
http://www.ulalaunch.com/uploads/docs/Published_Papers/Upper_Stages/TheCentaurUpperStageVehicleHistory.pdf

Quote
A few things still unknown:

Time it'd take to flip 180 for the braking burn.

Less than a minute I suspect ;)

Quote
How much delta V it would take to return to EML2. I believe it's doable to time the third apogee to be in the moon's neighborhood. But the route from an apogee in the moon's neighborhood to a halo around EML2 still hasn't coalesced in my imagination.

IIRC some of the early Mars missions (VonBraun? VSI?) called for reusable boosters that would return to Lunar orbit for refueling. I seem to recall that both nuclear and chemical boosters were discussed but my search-fu has failed me in finding a reference at this point.

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Offline metaphor

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Re: Reusable earth departure stages
« Reply #7 on: 05/27/2014 03:04 pm »
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.

Offline john smith 19

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Re: Reusable earth departure stages
« Reply #8 on: 05/27/2014 06:46 pm »
Note the key feature of Project Troy is the use of a resonance orbit in which the "departure" orbit is a sub multiple of the orbit around the planet or the Moon and is slightly below escape velocity, leaving the payload to do the last few m/s.

This in fact requires no braking thrust from the EDS.
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Offline JasonAW3

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Re: Reusable earth departure stages
« Reply #9 on: 05/27/2014 08:54 pm »
Why have the EDS detach and return back to the EML-2 point?

A Mars bound craft is going to need a midcourse correcction burn, and, if you weree using something like a Fusion engine, you'd want to do a continious acceleration out, a midcourse flip, and a continious decelleration to a safe atmospheric or orbital entry speed for the craft going to Mars. Once that speed has been achieved, detach the craft and begin accelletating the EDS back to Earth with a midcourse flip and correction, and decellerate it into EML-2 at that point for refurbishment and refuel.  In the meantime, the next EDS is launched before the first one has returned to Earth.

If an abort is declared prior to the detachement, they simply repeat the same procedure back to Earth.  If they abort to orbit, they wait there until the next EDS arrives and match velocities with it and dock after the EDS has dropped of it's payload.

In this situation, I'm estimating a round trip from Earth to Mars as 3 to 6 months, depending on orbital dynamics and at least 3 to 4 EDS units.  By using them to drop off supply payloads prior to launching a manned mission, the control and use procedures can be worked out, as well as Man Rating the EDS, prior to a manned mission.
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Offline Nilof

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Re: Reusable earth departure stages
« Reply #10 on: 05/27/2014 10:30 pm »
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.
For a variable Isp spacecraft running at constant power and constant acceleration, the mass ratio is linear in delta-v.   Δv = ve0(MR-1). Or equivalently: Δv = vef PMF. Also, this is energy-optimal for a fixed delta-v and mass ratio.

Offline jongoff

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Re: Reusable earth departure stages
« Reply #11 on: 05/28/2014 10:02 pm »
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

Offline cordwainer

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Re: Reusable earth departure stages
« Reply #12 on: 05/28/2014 11:59 pm »
I am guessing John when you say "resonance" orbit you mean low energy transfer orbit. While optimally you wouldn't need braking for the EDS for Mars you would need some braking for the Orbital Crew Module or payload that you would be sending. Albeit it wouldn't need much retro-thrust and such braking could be down gradually during the cruise phase using a low thrust/high Isp form of propulsion. As for the Moon it would probably be better to build the EDS and payload as one vehicle with two engine modules 180 degrees apart on booms, thus doing away with the need to flip the craft. You would have to have some method other than nose-in docking for refueling from a fuel depot, but if you have sufficient depot infrastructure there is no reason you couldn't build a reusable Cis-lunar shuttle.

Offline cordwainer

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Re: Reusable earth departure stages
« Reply #13 on: 05/29/2014 12:15 am »
Where they discuss resonance orbits is in relation to the OBO to minimize the energy and braking needed by the Skylon's to the departure window to meet up with the OBO.  The braking needed for the EDS to meet up and refuel with the OBO would be minimal, nothing that could not be achieved with simple RCS systems. Then the OBO maneuvers around the returning EDS to allow it to overtake the OBO upon return, which optimally doesn't need any braking. Still it is a somewhat risky maneuver.

Offline john smith 19

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Re: Reusable earth departure stages
« Reply #14 on: 05/29/2014 06:42 am »
Where they discuss resonance orbits is in relation to the OBO to minimize the energy and braking needed by the Skylon's to the departure window to meet up with the OBO.  The braking needed for the EDS to meet up and refuel with the OBO would be minimal, nothing that could not be achieved with simple RCS systems. Then the OBO maneuvers around the returning EDS to allow it to overtake the OBO upon return, which optimally doesn't need any braking. Still it is a somewhat risky maneuver.
Wouldn't that apply to any reusable departure stage?
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Offline MP99

Re: Reusable earth departure stages
« Reply #15 on: 05/29/2014 08:48 am »
After TMI, the EDS disengages, turns 180º and then does a braking burn. Slowing down to just below escape velocity would put the EDS on an ~ 9 day orbit. The 3rd perigee would be 27 days later and thus back in the moon's neighborhood. Then it could return to EML2 to get ready to send another spacecraft on its way.

When you intercept the Moon, you then need to perform further burns to target EML. That could be a problem 30+ days after departure, due to boiloff.

However, if you're prepared to wait, the EDS could follow a weak stability boundary orbit, which takes ~3 months, but requires nothing more than a few m/s of thruster to insert to EML.

I suspect part of the issue with this is whether the vector on the TMI burn lines up anywhere close to that needed for the return orbit.

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Offline Nilof

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Re: Reusable earth departure stages
« Reply #16 on: 05/29/2014 11:43 am »

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

I worked on some numbers on the "SpaceX FX/FXX/BFR Speculation Thread" for Mars transfers for the case where the MCT was an upper stage of an FXX-class rocket. I came up with some numbers for what I called a "Mars toss" mission:

MCT: 100 tons payload to Mars, has a 75 tons dry mass, and contains 775 tons of propellant when fully loaded. Has an Isp of 340 and a Delta-V of ~5.6 km/s with the full payload, 6.5 km/s with a 50 ton payload, 7.2 km/s with a 25 ton payload, and 8 km/s with no payload.

[...]
Mars toss transfer missions: gets fully refueled at LEO, places a 100 ton payload in a Mars transfer orbit with a 4 km/s burn, separates from the payload with ~200 tons of propellant(so ~4km/s delta-v) left. Slows down and lands at the launch site or docks with the depot within a day or two for a second mission.

The basic idea is that you can get back to the depot rather quickly if it's in LEO, using a high T/W transfer stage that can place a payload in a hyperbolic orbit and then quickly brake into an orbit that will put it back at the depot within a small integer multiple of the depot's orbital period.

With that said, Mars transfers do chug through propellant, as in roughly 7-8 tons of propellant per ton of payload for Kerolox or Hypergolics. However, stage reuse doesn't change that much if your stage has a decent mass ratio. For multiple launches, the large depots needed would be the bottleneck either way.
« Last Edit: 05/29/2014 11:56 am by Nilof »
For a variable Isp spacecraft running at constant power and constant acceleration, the mass ratio is linear in delta-v.   Δv = ve0(MR-1). Or equivalently: Δv = vef PMF. Also, this is energy-optimal for a fixed delta-v and mass ratio.

Offline jongoff

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Re: Reusable earth departure stages
« Reply #17 on: 05/29/2014 12:48 pm »

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

I worked on some numbers on the "SpaceX FX/FXX/BFR Speculation Thread" for Mars transfers for the case where the MCT was an upper stage of an FXX-class rocket. I came up with some numbers for what I called a "Mars toss" mission:

MCT: 100 tons payload to Mars, has a 75 tons dry mass, and contains 775 tons of propellant when fully loaded. Has an Isp of 340 and a Delta-V of ~5.6 km/s with the full payload, 6.5 km/s with a 50 ton payload, 7.2 km/s with a 25 ton payload, and 8 km/s with no payload.

[...]
Mars toss transfer missions: gets fully refueled at LEO, places a 100 ton payload in a Mars transfer orbit with a 4 km/s burn, separates from the payload with ~200 tons of propellant(so ~4km/s delta-v) left. Slows down and lands at the launch site or docks with the depot within a day or two for a second mission.

The basic idea is that you can get back to the depot rather quickly if it's in LEO, using a high T/W transfer stage that can place a payload in a hyperbolic orbit and then quickly brake into an orbit that will put it back at the depot within a small integer multiple of the depot's orbital period.

With that said, Mars transfers do chug through propellant, as in roughly 7-8 tons of propellant per ton of payload for Kerolox or Hypergolics. However, stage reuse doesn't change that much if your stage has a decent mass ratio. For multiple launches, the large depots needed would be the bottleneck either way.

Nilof,

One challenge that I think Hop is hitting on is that if you do a direct from LEO departure burn to Mars, by the time you're done with the toss, you're going to be pretty far from periapsis, and may thus pay a pretty heft propellant "fine" to keep the stage. I would think that if you wanted to recover the stage, your best bet would be first to boost to a highly elliptical earth orbit (or come in from L1/L2), and only do the last couple hundred m/s of the departure burn when you're back approaching periapsis.

I'm doing a paper right now that's vaguely related (it's looking at a departure burn method to enable departures to arbitrary BEO destinations using a LEO depot), if we actually get it written this year, it might be good to do something like this as a follow-on.

~Jon

Offline Nilof

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Re: Reusable earth departure stages
« Reply #18 on: 05/29/2014 07:37 pm »
Well, when I ran the numbers, my main conclusion was that the ~4 km/s Mars burn is small enough that there's lots of margin in the delta-v budget for slowing down.

In the example above I assumed a rather mediocre mass ratio(~11) for the upper stage because I assumed that MCT would have a heat shield. If it is a reusable transfer stage and nothing else, you can assume a mass ratio of ~16 for a Kerolox stage, and you get significantly more than 4 km/s available for a braking burn. So even if the timing on the first braking burn isn't ideal, braking down the stage is still quite practical.
« Last Edit: 05/29/2014 07:42 pm by Nilof »
For a variable Isp spacecraft running at constant power and constant acceleration, the mass ratio is linear in delta-v.   Δv = ve0(MR-1). Or equivalently: Δv = vef PMF. Also, this is energy-optimal for a fixed delta-v and mass ratio.

Offline cordwainer

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Re: Reusable earth departure stages
« Reply #19 on: 05/29/2014 08:36 pm »
True, any EVA docking would be risky.  If that is what your referring to, then yes. I think it comes down to designing appropriate docking, guidance and EDS capture technology with sufficient safeguards and redundancy to the OBO. I think you would want two OBO's in trailing orbit of one another in case you fail to capture on the first attempt or if something catastrophically goes wrong with one of the OBO's.

Offline cordwainer

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Re: Reusable earth departure stages
« Reply #20 on: 05/29/2014 08:51 pm »
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.

Offline cordwainer

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Re: Reusable earth departure stages
« Reply #21 on: 05/29/2014 09:08 pm »
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.

Offline Hop_David

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Re: Reusable earth departure stages
« Reply #22 on: 05/30/2014 04:25 pm »
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.

Offline aero

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Re: Reusable earth departure stages
« Reply #23 on: 05/30/2014 05:01 pm »
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.
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Offline jongoff

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Re: Reusable earth departure stages
« Reply #24 on: 05/30/2014 05:02 pm »
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

Offline savuporo

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Re: Reusable earth departure stages
« Reply #25 on: 05/30/2014 05:41 pm »
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.
« Last Edit: 05/30/2014 06:03 pm by savuporo »
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Offline Hop_David

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Re: Reusable earth departure stages
« Reply #26 on: 05/30/2014 05:50 pm »
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.

Offline DarkenedOne

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Re: Reusable earth departure stages
« Reply #27 on: 05/30/2014 06:05 pm »
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. 

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Re: Reusable earth departure stages
« Reply #28 on: 05/30/2014 06:07 pm »
[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.

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Re: Reusable earth departure stages
« Reply #29 on: 05/30/2014 06:12 pm »
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.
« Last Edit: 05/30/2014 06:13 pm by Hop_David »

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Re: Reusable earth departure stages
« Reply #30 on: 05/30/2014 07:06 pm »
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Offline muomega0

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Re: Reusable earth departure stages
« Reply #31 on: 05/30/2014 07:41 pm »
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.

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Re: Reusable earth departure stages
« Reply #32 on: 05/30/2014 08:00 pm »

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).

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Re: Reusable earth departure stages
« Reply #33 on: 05/30/2014 08:12 pm »
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.
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Re: Reusable earth departure stages
« Reply #34 on: 06/01/2014 02:57 am »
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

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Re: Reusable earth departure stages
« Reply #35 on: 06/01/2014 02:59 am »
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

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Re: Reusable earth departure stages
« Reply #36 on: 06/01/2014 03:40 am »
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.
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Re: Reusable earth departure stages
« Reply #37 on: 06/01/2014 06:41 am »
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.
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Re: Reusable earth departure stages
« Reply #38 on: 06/01/2014 04:10 pm »
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

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Re: Reusable earth departure stages
« Reply #39 on: 06/01/2014 07:14 pm »
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.
« Last Edit: 06/01/2014 07:15 pm by Nilof »
For a variable Isp spacecraft running at constant power and constant acceleration, the mass ratio is linear in delta-v.   Δv = ve0(MR-1). Or equivalently: Δv = vef PMF. Also, this is energy-optimal for a fixed delta-v and mass ratio.

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Re: Reusable earth departure stages
« Reply #40 on: 06/01/2014 08:15 pm »
Reposting Farquhar's Route, this time as attached rather than embedded image.

A few important numbers: 72 hours from EML2 to perilune and then 140 hours from perilune to perigee.

A small EML2 burn and a small perilune burn suffice to send the EDS and payload towards a deep perigee. These two burns total about .4 km/s.

At perigee the whole shebang's moving about 10.8 km/s. Since 11.4 km/s is enough for TMI, a mere .6 km/s burn suffices for Trans Mars Insertion.

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Re: Reusable earth departure stages
« Reply #41 on: 06/01/2014 09:01 pm »
Am reposting the ~50 day route back to the moon but as attached image.

During 140 hours the moon advances 76 degrees and the EDS and payload advance 180º. So I start the sim with the perigee 104 degrees ahead of the moon. I found a 10.85 km/s perigee speed will give the EDS an orbit with a 396,000 km apogee and an orbit whose period is about 2/5 of the moon's period.

From perilune at the beginning to the perilne at the end is about 54 days. Add in two 3 day trips between perilune and EML2 and the total trip takes 60 days.

Reading earlier comments I was coming to the conclusion that 60 day round trip was a show stopper. Hydrogen boil off seemed to be too high over two months.

I liked A. M. Swallow's notion of using methane and thus the EDS have propellent and parts in common with the lander. That would simplify a number of things. However my notion rests on propellent sources either from an carbonaceous asteroid parked in lunar orbit or volatiles from the lunar cold traps. LCROSS gave evidence of some carbon compounds, so maybe methane is a possibility.

Then I read Jon Goff's comments that boil of could be mitigated -- he seemed to be talking about the same techniques that Goff, Zegler,  Kutter and Bar have written about in their propellent depot papers.

I Had been thinking of modeling my EDS after a Centaur stage with 99.2 K newton thrust, 20,830 kg fuel and oxidizer mass and 2,247 dry mass. Maybe I should up the dry mass to accommodate some extra MLI and apparatus for exploiting boil off.
« Last Edit: 06/01/2014 09:02 pm by Hop_David »

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Re: Reusable earth departure stages
« Reply #42 on: 06/01/2014 09:44 pm »
Am reposting the ~50 day route back to the moon but as attached image.

I love images like these; thanks for reposting them!

Does your software have the ability to show the pellet trajectories in a rotating frame of reference where both the Earth and Moon appear fixed? I understand a rotating frame is strange, and I won't try here to justify its use, but note it is used in e.g. the Farquhar route shown in your previous post....
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Re: Reusable earth departure stages
« Reply #43 on: 06/02/2014 02:51 pm »
Am reposting the ~50 day route back to the moon but as attached image.

I love images like these; thanks for reposting them!

Does your software have the ability to show the pellet trajectories in a rotating frame of reference where both the Earth and Moon appear fixed? I understand a rotating frame is strange, and I won't try here to justify its use, but note it is used in e.g. the Farquhar route shown in your previous post....

Not my software, I use Bob Jenkins' Java orbital sim. I wish I could get it to hold motionless the central and orbital body as well as the line between them. But I don't know how.

Offline muomega0

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Re: Reusable earth departure stages
« Reply #44 on: 06/02/2014 03:33 pm »
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
Its in the sun once out of LEO...use solar arrays.
Adding two more technologies to the above (advanced batteries and coolers) reduces the LH2 boiloff to zero and adds additional contingency operational modes including settling.  Using propellant for power is quite expensive and allows more efficient (longer) trajectories.

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Re: Reusable earth departure stages
« Reply #45 on: 06/04/2014 03:23 pm »

Then I read Jon Goff's comments that boil of could be mitigated -- he seemed to be talking about the same techniques that Goff, Zegler,  Kutter and Bar have written about in their propellent depot papers.
Depots and IVF share many of the same techniques and approaches. Current Centaur stages have a life measured in hours due to consumables depletion and propellant boil off. Boil off reduction is a critical technology to increase mission duration. I'd say 1st generation upgrades are good to a week or two. 60 days will probably need to go to the full sun shield and active cryo cooler upgrades.
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