Author Topic: Combining the benefits of SEP and Hypergolics for human missions.  (Read 15149 times)

Offline Darkseraph

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Upon seeing the success of solar electrical propulsion on Dawn and a few other missions using this tech I was struck with an odd idea.

My premises are:

1) Solar electric propulsion is impractical for sending humans to the Moon because it could take months in going through radiation belts. But it is extremely mass efficient for prepositioning human equipment like lander, return stages and habitats due to its incredible ISP.

2) We don't have either cryogenic propellant depots or lunar ice mines as of yet. I think both are promising ideas but will cost quite a shiny penny to get off the ground. They have unknown unkowns because the are new tech. We also don't have reusable launch vehicles. But we have performed landings on the moon using hypergolics and we have transferred propellant this way in space on the ISS. In space refueling with hypergolics is a mature technology and doing reusuable single stage on the moon is much easier than earth.

3) Hypergolics have terrible isp, but they can be stored for a very long time without boil off issues and are much more dense than LH/LOX systems.

4) I'm assuming a "what if?" situation were we don't have the SLS, Cryogenic Depots, Reusable Launch Vehicles or Lunar IRSU and have to use mostly available or mature technology to go to the Moon.

So putting all those together, what I am suggesting is that Lunar missions could be done with EELV sized vehicles in a reusable system that combines hypergolics with solar electric propulsion. The hypergolic return stage, lander, habitat is prepostioned (at L1/L2 or LLO) and refueled by an SEP vehicle, which returns to earth orbit to be refueled with more hypergolics and xenon/argon for the SEP itself. A human capsule without such a massive service module is sent to dock with the lander on a regular EELV using conventional chemical propulsion in 3 days. They go to the the surface, explore, the lander goes back up to orbit with enough propellant for station keeping, transfers them to their capsule+hypergollic return stage....rinse repeat until you have built up a bit of base on the moon that may be able to do some water mining, but it's not a requirement.

The return stage (which is basically like an Apollo service module) would be expended in this system, but the moon lander wouldn't. It would be a single stage system that you refuel as well as the SEP tug. The tanker carrying propellant for the SEP and hypergolics for the other systems would be an expendable system similar to Progress or ATV. It meets up in LEO with a mostly empty tug, transfers fluids and de-orbits on its merry way.

There are several possible SEP craft that could be the tug, from what is proposed for the asteroid return mission to the lunar tug suggest by Ad Astra using VASIMR engines.  The SEP modules round trip would be on the order of months, but not that many lunar missions are likely to be done each year. If you want to do more missions or build up a base with cargo, you add more tugs launched EELVs to do that.

So, any particular gaping holes in doing something like this?
« Last Edit: 03/07/2015 06:04 pm by Darkseraph »
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Offline KelvinZero

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Another way of combining benefits of SEP and hydrazine could be to place a series of propellent depots in your path, such that after you expend the propellent of one you are positioned to rendezvous with the next.

I can't find the original thread which had several variations, but I found a summary:
SEP accelerated hydrazine depots

It isn't a serious proposal. If you miss your window or any of the following depots all that effort is wasted. Interestingly though you would not be stranded in space. After each burn you are still on a one year trajectory back to earth. Similarly all your SEP depots could return to aerobrake back into LEO for refueling one year later. (It has some other issues too though.)


Offline Darkseraph

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I've no clue why you would do that other than for Rube Goldbergy reasons. :D


That Jupiter-Exoliner system that was just announced is kinda what I am getting at. I am thinking of how you could achieve lunar missions without SLS, cryogenic depots, ice mining or using so much mass for them to recur. Obviously it is a massive slow boat way of doing it, but I suspect there will be so few missions initially that it won't matter.

If you break it up, only the parts carrying humans have to go fast (either to the moon or away from it). Everything else can actually go pretty slow, esp if it used hypergolics instead of cryogenic fluids.  So that would be the lander's ascent/decent propellant, the capsule return propellant, and station keeping propellant.

Hypergolics are now where as efficient as cryogens, but they are dense, can be stored for easily for years and the enourmous efficiency of SEP systems just dwarfs the relativy poor ISP of hypergolics to Hydrogen or Methane. It would mitigate it.

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Offline sdsds

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This idea seems to presuppose a large advantage can be gained by moving mass up Earth's gravity well using SEP. I suggest that this step can be much more easily accomplished using hydrolox upper stages of launch vehicles to get to an orbit like super-sync GTO, and then a small hypergolic (or even solid) kick stage to circularize there, or maybe even push out into a larger ellipse. At that point, sure, SEP is your "slow boat" friend to move onto trajectories which use three-body (Earth, Moon, spacecraft) dynamics to get you where you want to go.

That much is all possible with today's technology. You then "only" need to design your mission architecture so the pieces rendezvous and aggregate in one of the places in the Earth-Moon system where trajectories lead to and from ... anywhere. The frequently discussed possibilities include the colinear Lagrange points and distant retrograde lunar orbits.

Was this the kind of architecture you had in mind, or were you planning on the long spiral out from LEO using SEP?
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Offline gbaikie

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Another way of combining benefits of SEP and hydrazine could be to place a series of propellent depots in your path, such that after you expend the propellent of one you are positioned to rendezvous with the next.

I can't find the original thread which had several variations, but I found a summary:
SEP accelerated hydrazine depots

It isn't a serious proposal. If you miss your window or any of the following depots all that effort is wasted. Interestingly though you would not be stranded in space. After each burn you are still on a one year trajectory back to earth. Similarly all your SEP depots could return to aerobrake back into LEO for refueling one year later. (It has some other issues too though.)

Well, instead of solar system one could also do this for Earth/Moon system.
So start with so 400 by 20,000 km orbit. Next step is a 20,000 circular orbit. Next a 20,000 by 100,000 km.

let's say SEP begin or stays at 20,000 km circular orbit. And Xenon and storable rocket fuel delivered to 400 by 20,000 and which the SEP with no payload, goes to the 400 by 20,000 and gets payload, and lifts it to 20,000 km circular. Leaves some payload at 20,000 km and lifts the the rest up to 20K to 100K, then returns with no payload to 20,000 circular.
And spacecraft which uses the rocket fuel gets a launch to 400 by 20,000, and it uses onboard fuel to raise up to 20,000 by 20,000, then refuels, then goes to 20,000 by 100,000 and refuels again.
So SEP job is to deliver the fuel and have fuel in right part of orbit so the spacecraft which needs it will have in the right spot of the orbit for where ever that spacecraft is going. And if spacecraft fails to show up move fuel again to right location for the spacecraft and this could include altering the inclination of these orbits. And the assumption is one knows when and which spacecraft will use it.
Of course the 20K circular orbit is less than 24 hours [shorter than GEO]. And fuel can be in location that once spacecraft has circularize the orbit it's minutes away for beginning it's docking and refueling. So less than day to fire the engine after refueling to reach the 20k by 100K in order to refuel the second time.
And by time refueled at the 20k by 100k orbit, it might immediate burn it's rockets to go where ever it's going or it might have spend entire orbit period before doing this and/or it could dock with SEP which uses it's power to send it somewhere [and that would also take a lot of time] but it could retain it's fuel for a lunar or Mars capture burn. And the spacecraft only goes once thru Van Allen- could be crew or cargo.
Also SEP does not need to spend much time repetitively going thru the Van Allen belts [over it's life maybe dozens of times but not hundreds of times].
Have little clue if economical, but one could also mix in the use cryogenic fuels- such as LOX and/or whatever. And one could collect a bunch empty fuel tanks for some kind of future use. And could one ship Xenon to places where other SEP spacecraft need it.

Offline KelvinZero

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Yeah sorry didnt want to pull it off the moon OP topic.

I also like the idea of a lunar cycler which is sort of like a single instance of these non-stationary depots, and means that abandoning free return and stopping at the moon could be done with SEP-prepositioned hydrogolic propellent.

Offline mmeijeri

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3) Hypergolics have terrible isp, but they can be stored for a very long time without boil off issues and are much more dense than LH/LOX systems.

Actually, it's not terrible at all, it's similar to LOX/kerosene. It's just that LOX/LH2 has spectacular performance for chemical propulsion, at the cost of additional complexity.

Combining hypergolics and SEP seems like the obvious choice to me, and you don't even have to rule out some use of LOX/LH2 even at the beginning. If you use L1/L2 as a staging point (which has a long list of additional advantages), then you can still use LOX/LH2 to get stuff to L1/L2. You could do that either in a single launch for smaller payloads, say with an Atlas V, or in two launches, with the hypergolics payload going first and a transfer stage to take it to L1/L2 second. That requires no new technology either.

Also, you wouldn't even need full hypergolics depots initially, a spacecraft could be its own depot, directly supplied by tankers, just like the ISS today and Soviet stations before it.
« Last Edit: 03/17/2015 06:59 pm by mmeijeri »
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Offline KelvinZero

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This idea seems to presuppose a large advantage can be gained by moving mass up Earth's gravity well using SEP. I suggest that this step can be much more easily accomplished using hydrolox upper stages of launch vehicles to get to an orbit like super-sync GTO, and then a small hypergolic (or even solid) kick stage to circularize there..
I would have assumed that. Any numbers to put it into perspective for a layman like me?

Offline Darkseraph

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I've read a paper on this using VASIMR for a lunar tug that delivers cargo landers. VASIMR is of course just one form of SEP and there are others that are very similar. But for the same initial mass to LEO, this system could deliver twice the payload of a LH/LOX chemical system to the Moon with a roundtrip time of 7 months. The next payload delivered can be even bigger since you're not putting up the mass of the SEP tug again, just cargo/fuel.


That is actually a massive advantage over LH/LOX, it allows you to downsize your vehicles from something the size of SLS to something the size of EELVs/Falcon Heavy or their international equivalents. You can make the whole system mostly reusable too without needing cryogenic depots or lunar ice mining up front. You can't just send the whole mission on one super-rocket Saturn style, you have to chop it into pieces, but most of those pieces are unmanned equipment prepositioned and tested, so a failure is not a program ending tragedy, you just postpone the human rendezvous launch.


http://www.nasaspaceflight.com/2012/01/aerojet-solar-electric-propulsion-enabler-exploration-gateway/

Forget about this, but they're talking about something similar.
« Last Edit: 03/15/2015 11:10 pm by Darkseraph »
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Offline sdsds

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Any numbers to put it into perspective for a layman like me?

I think a fair "apples to apples" comparison would start with two identical launch vehicles. As an example let's use Atlas V 431. One is tasked with delivering a SEP tug and its payload to 400 km circular LEO. It delivers about 15,130 kg of payload systems weight to that target. The other is tasked with delivering a kick stage and it payload to a highly elliptical 185 x 150,000 km orbit. It delivers about 6,293 kg of payload systems weight to that target.

Is the question, "Which approach delivers more to e.g. a 150,000 x 150,000 circular orbit?"
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Offline Nilof

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The Jupiter/Exoliner system is a good example of the benefits of a reusable tug. You can use the entire LV payload to send up "dumb" cans of propellant, with far less packaging needed. With chemical, you lose a lot more payload by making things reusable, and the expendable systems will also mean more wrapping on top of the payload hit from the rocket equation.
« Last Edit: 03/17/2015 11:22 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 dkovacic

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This idea seems to presuppose a large advantage can be gained by moving mass up Earth's gravity well using SEP. I suggest that this step can be much more easily accomplished using hydrolox upper stages of launch vehicles to get to an orbit like super-sync GTO, and then a small hypergolic (or even solid) kick stage to circularize there, or maybe even push out into a larger ellipse. At that point, sure, SEP is your "slow boat" friend to move onto trajectories which use three-body (Earth, Moon, spacecraft) dynamics to get you where you want to go.

That much is all possible with today's technology. You then "only" need to design your mission architecture so the pieces rendezvous and aggregate in one of the places in the Earth-Moon system where trajectories lead to and from ... anywhere. The frequently discussed possibilities include the colinear Lagrange points and distant retrograde lunar orbits.

Was this the kind of architecture you had in mind, or were you planning on the long spiral out from LEO using SEP?
Here are some arguments against it:

Hydrolox stage is not reusable without major redesign.
Hydrolox stage has very short mission span (hours or days at best?) without active cryocooling.
Hydrolox stage has significantly lower ISP than SEP.
Hydrolox stage cannot be refueled in space (without major development effort).
Hydrolox stage has to be launched together with payload (cannot do EOR and assembly in orbit), and is thus mass and volume limited by the LV (essentially it needs SLS).
Hydrolox stage is as expensive as SEP (ok, this is just my estimate, it might not be true). RL-10 alone is 38 million USD.

By hydrolox stage I assume Centaur derived stage, not completely new stage.

So hydrolox could be used for initial GTO/HEO insertion, but for the "similar" amount of money SEP tug (which is far-reaching technology enabler) could be developed and used.

Offline Oli

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IMO the biggest advantage of SEP/(pressure-fed)Hypergolics is that we know that it works reliably for years or even decades. For human missions BEO reliability is key for cost-effectiveness.

Offline redliox

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Combining SEP & hypergolics seems like a 2 sided coin:
1) Combining the best non-cryogenic systems that are space-proven.
2) Combining the least practical systems that come with their own monkey wrenches.

I'll explain each side reasonably...

1)  SEP is very efficient.  It's been used several times on probes and even commercial satellites, and if scaled up properly could certainly move (at least cargo) elements for human missions with less fuel.  Hypergolics, while "old school" alongside SEP, are steadfast and easily manageable compared to crygenics; and are used on almost every spacecraft for either simple attitude control or orbital insertion.  A combination of a simple setup and a complex one.

2) SEP...is slower than molasses...literally.  Moving a space probe that is, typically, at worst 2 tonnes in mass is manageable but manned vessels will be on the order of 20-50, including unmanned ones.  If your flight to Mars requires more than a month to break into orbit in addition to ~6 months of cruise...it certainly won't be on the crewed leg.  As for hypergolics, it's heavy, low specific impulse, and explosive (then again what rocket fuel isn't).

In the end, it would make an interesting combination.  The trick would be cobbling together something that stays within mass parameters and delta-v budgets, not to mention accommodating human needs.  I doubt it would work well for humans but, outside of aerocapture (which has its own drawbacks), SEP supplemented by hypergolics would be the best way to get cargo into Mars orbit.  Plenty of potential but you should remain aware of the drawbacks.
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Offline Darkseraph

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That's why I put the thread in the Moon section. I don't think this would be a good way to go to Mars, except for some cargo.

Unless you use SEP to put heavy radiation shielded closed loop life support artificial gravity cyclers into their correct orbits...this solves some problems, but you couldn't avoid needing a large chemical stage to boost the crew to meet the cycler. I won't touch how you solve that problem with 50 light year clown pole!
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Offline mmeijeri

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That's why I put the thread in the Moon section. I don't think this would be a good way to go to Mars, except for some cargo.

It would be excellent for prepositioning propellant to SM-1/2, LMO and perhaps L1/L2, which would be very practical for Mars.

Quote
Unless you use SEP to put heavy radiation shielded closed loop life support artificial gravity cyclers into their correct orbits...this solves some problems, but you couldn't avoid needing a large chemical stage to boost the crew to meet the cycler. I won't touch how you solve that problem with 50 light year clown pole!

You still need a chemical stage, but it doesn't have to be huge. At least not so huge as to require a very large launch vehicle. The trick is to use the transfer stage from L1/L2 onwards, and to use modified LV upper stages to get from LEO to L1/L2.
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Offline redliox

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That's why I put the thread in the Moon section. I don't think this would be a good way to go to Mars, except for some cargo.

It would be excellent for prepositioning propellant to SM-1/2, LMO and perhaps L1/L2, which would be very practical for Mars.

I think I agree with both of you guys on those points.  The scheme NASA's currently following, using SLS, Orion, and a SEP tug would be lousy to use for an expedition to Mars (mainly due to the later 2 elements)...but they would serve well for a Lunar one.

The Orion can send humans to the Moon swiftly enough without excessive support.  An SEP tug could haul a lander, fuel tanks, and perhaps even a space station to EML-1, EML-2, or Lunar orbit beforehand...and even be reusable with an adequate xenon supply.  SEP and crew wouldn't be tied together, but they could work independently of each other well.
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Offline redliox

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I'd like to see an architecture made from this.  I suppose the first starting point is establishing the delta-v needs along with the number of flights.

Let's assume there are two key staging points: LEO and L1 (L2 could be another possibility, but most of its benefits are for interplanetary flights).  The delta-v for a one-way trip from one to the other is 3.8 km/sec, 7.6 for a round trip.  From L1 to the lunar surface, the d-v is 2.5, 5 for a complete trip to-from L1 and the Moon.  We have two groups of numbers to work with: 7.6 km/sec for Earthly vehicles and 5 km/sec for Lunar vehicles.

The Earthly vehicles consist of the SEP tug and Orion.  The Orion's true strength comes from the SLS and returning directly to Earth via reentry.  The SEP has a hard job ahead, pushing cargo under its own power from LEO all the way to L1, then slowly returning back to Earth.  The SEP would rendezvous with objects largely in LEO, whereas the Orion rendezvous at L1.

The Lunar vehicles would be either cargo or crew landers.  The cargo variant would have the least amount of fuel of any vehicle, as it would only journey down to the Moon once whereas the crew version needs to return to L1.  Neither would deal with Earth's gravity well, as the SEP tug would shove both to L1 and do the hard work for them.  The crew lander, between missions, would wait at L1, get refueled by another SEP visit, and then another Orion on its second mission.
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Offline A_M_Swallow

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I'd like to see an architecture made from this.  I suppose the first starting point is establishing the delta-v needs along with the number of flights.

Let's assume there are two key staging points: LEO and L1 (L2 could be another possibility, but most of its benefits are for interplanetary flights).  The delta-v for a one-way trip from one to the other is 3.8 km/sec, 7.6 for a round trip.  From L1 to the lunar surface, the d-v is 2.5, 5 for a complete trip to-from L1 and the Moon.  We have two groups of numbers to work with: 7.6 km/sec for Earthly vehicles and 5 km/sec for Lunar vehicles.

The Earthly vehicles consist of the SEP tug and Orion.  The Orion's true strength comes from the SLS and returning directly to Earth via reentry.  The SEP has a hard job ahead, pushing cargo under its own power from LEO all the way to L1, then slowly returning back to Earth.  The SEP would rendezvous with objects largely in LEO, whereas the Orion rendezvous at L1.

The Lunar vehicles would be either cargo or crew landers.  The cargo variant would have the least amount of fuel of any vehicle, as it would only journey down to the Moon once whereas the crew version needs to return to L1.  Neither would deal with Earth's gravity well, as the SEP tug would shove both to L1 and do the hard work for them.  The crew lander, between missions, would wait at L1, get refueled by another SEP visit, and then another Orion on its second mission.

The SEP cannot use the Oberth effect so the LEO to EML-1 delta-v is 7.0 km/s, round trip 14.0 km/s. SEP can still be used but it takes longer and needs much more propellant.

Direct reentry from EML-1 is 0.77 km/s.

The burn between LEO and EML-1 can be performed by the Orion's service module. Larger fuel tanks will be needed than the service module currently being developed or Orion could pick up drop tanks at the LEO spacestation.

Offline redliox

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The SEP cannot use the Oberth effect so the LEO to EML-1 delta-v is 7.0 km/s, round trip 14.0 km/s. SEP can still be used but it takes longer and needs much more propellant.

Direct reentry from EML-1 is 0.77 km/s.

The burn between LEO and EML-1 can be performed by the Orion's service module. Larger fuel tanks will be needed than the service module currently being developed or Orion could pick up drop tanks at the LEO spacestation.

I figured the Orion and it's service module could easily handle EML-1, but I know it's delta-v capacity is slightly less than 1.5 km/sec, so reach low lunar orbit is just beyond it's capacity.  The SEP would have the harder task of the two vehicles.  While the Oberth effect wouldn't be available, there is still the chance a regular gravity assist from the Moon could be applied - most of the flights for Orion seem to involve those to reach retrograde orbit, and reaching any of the Lagrange points often involves using the Moon's gravity.  Either way, I would presume a SEP tug would be given the capacity to reach the Lagrange point directly.
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Offline TrevorMonty

Every time a SEP tug goes fro LEO to L1 its solar panels a degradated by approx 20%,  so multiple round trips are not practical. What is needed is alternative uses for SEP tug once it gets to cislunar space.
Providing power to lunar base is one option. Use 30kw (1200kg dry) to deliver a few tons to L1 then spiral down to LLO and use a small  hypergolic lander to place it on surface. The flexiwings can track sun during lunar day maximizing  power output.
Other SEP tugs can stay at L1 and operate as solar power stations using lasers to transmit power to lunar SEP. Efficiency is about 10%, so 3kw received on lunar surface for every L1 30kw SPS. Every additional L1 SPS adds another 3kw or more if their efficiency increases.

3kw is not enough to support a manned lunar base through a lunar night. With enough L1 SPSs, eventually the base can be manned continuously especially if stored power is also used.

Offline MP99

What height orbit are you thinking of to avoid issues with Van Allan belt radiation?

Simplest solution seems to be to have the upper stage of the launcher deliver the payload to just that height, instead of a minimal LEO orbit. Then just have SEP take over from there.

Cheers, Martin

Offline Burninate

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Every time a SEP tug goes fro LEO to L1 its solar panels a degradated by approx 20%
Can you provide a citation for that?

Offline TrevorMonty

Every time a SEP tug goes fro LEO to L1 its solar panels a degradated by approx 20%
Can you provide a citation for that?

http://www.iepc2013.org/get?id=045

With each one-way trip from earth orbit to a distant retrograde orbit or a destination such as EML2, the
spacecraft must pass through the Van-Allen radiation belts. Each pass through the belts reduces the solar array
output by approximately 12% (possibly more, depending on the cover-glass thickness and solar array type).
Increasing the solar array size is an obvious work-around to this issue, but since a single re-use of a SEP spacecraft
would require an additional two trips through the Van-Allen belts, the requisite increase in solar array size can
become a major design driver.

LEO - L2 =88%
L2- LEO = 77%
LEO - L2= 68%



Offline redliox

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Every time a SEP tug goes fro LEO to L1 its solar panels a degradated by approx 20%
Can you provide a citation for that?

http://www.iepc2013.org/get?id=045

With each one-way trip from earth orbit to a distant retrograde orbit or a destination such as EML2, the
spacecraft must pass through the Van-Allen radiation belts. Each pass through the belts reduces the solar array
output by approximately 12% (possibly more, depending on the cover-glass thickness and solar array type).
Increasing the solar array size is an obvious work-around to this issue, but since a single re-use of a SEP spacecraft
would require an additional two trips through the Van-Allen belts, the requisite increase in solar array size can
become a major design driver.

LEO - L2 =88%
L2- LEO = 77%
LEO - L2= 68%

So a reusable spacecraft would need to get the arrays replaced once every 2 round-trips at best?  Not ideal but not impossible.   Radiation degradation is wise to factor in all setups.
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Offline mmeijeri

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There are apparently panels that somehow anneal back to the 80% state. In other words, they do degrade, but not below 80%. I think Robotbeat has links handy.
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Offline MATTBLAK

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What about keeping the arrays furled up until chemical thrust sends them past the Van Allen belts, then unfurling them after that?
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Offline Oli

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LEO - L2 =88%
L2- LEO = 77%
LEO - L2= 68%

That seems a bit excessive. I've come across the graphs below in a paper from 2007.

Note that according to this even a planar 300kw array with 30mils shielding would weight "only" 6t.

(SLA stands for "stretched lens array")
« Last Edit: 08/10/2015 11:26 am by Oli »

Offline TrevorMonty

LEO - L2 =88%
L2- LEO = 77%
LEO - L2= 68%

That seems a bit excessive. I've come across the graphs below in a paper from 2007.

Note that according to this even a planar 300kw array with 30mils shielding would weight "only" 6t.

(SLA stands for "stretched lens array")

Offline TrevorMonty



LEO - L2 =88%
L2- LEO = 77%
LEO - L2= 68%

That seems a bit excessive. I've come across the graphs below in a paper from 2007.

I was applying 12% degradation each transit, but it may work differently than that. I'm guessing longer transit time higher degradation, so it may not be so bad on L2-LEO trip when SEP tug is empty.
There is also issue of the cargo's exposure, to be considered.


Offline savuporo

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But we have performed landings on the moon using hypergolics and we have transferred propellant this way in space on the ISS. In space refueling with hypergolics is a mature technology and doing reusuable single stage on the moon is much easier than earth.
A small but obvious nitpick : its 'mature technology' in Russia.  On US side, still developing
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Offline redliox

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But we have performed landings on the moon using hypergolics and we have transferred propellant this way in space on the ISS. In space refueling with hypergolics is a mature technology and doing reusuable single stage on the moon is much easier than earth.
A small but obvious nitpick : its 'mature technology' in Russia.  On US side, still developing

*coughs Russianengineimportsrememberthatmess coughs*

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Offline mmeijeri

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A small but obvious nitpick : its 'mature technology' in Russia.  On US side, still developing

That's for noncooperative satellites. Orbital refueling for hypergolics is simply mature technology. Orbital Express demonstrated it years ago.
Pro-tip: you don't have to be a jerk if someone doesn't agree with your theories

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