Author Topic: Deimos Aligned - Conservative Architecture for the Martian Moons  (Read 6522 times)

Offline redliox

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The Martian moons may be the second step on the path humans take to Mars, the first being either the Moon or the asteroids.  Personally I favor them, especially Deimos, for their strategic proximity that offers a science bonus.  There's the chance of ice can ISRU, but this architecture would not rely on that - instead it would rely almost exclusively on modern and conservative methods, even more so than my previous thread 'Mars Aligned.'

I've previously stated I reference Mars Semi-Direct as a first basis for reasonable plans, but in the case of the Martian moons nothing is a perfect match.  The best analogs would be the old Apollo lunar flights and the current ARM plans.  The largest disadvantage is all the fuel must be carried with, which will result in heftier vehicles.  The greatest advantage is the low delta-v, especially at Deimos, resulting in less need for fuel.  The only aerodynamic maneuvering happens strictly at Earth, at launch and for crew return.

Compared to Martian schemes, this would have a far simpler setup:
-2 Launches
-2 Spacecraft
-Docking only near Earth

The two vehicles involved are derived from my previous Mars Aligned scheme, but may prove to be far more capable for a Deimos/Phobos expedition:
-Transit Vehicle - TV - enhanced version of Orion.
-Deimos Excursion Vehicle - DEV - essentially a Mars hab lander but with larger fuel tanks and no heat shields.

I will elaborate on the vehicles, especially DEV, in a second posting, but the flight plan itself is simple enough that I'll lay it out straightaway:

An SLS launches DEV to an Earth-Moon-Lagrange point (most likely EM-1).  Shortly after via a 2nd SLS, TV joins with a crew.  At the Lagrange staging point, the TV and DEV dock, and remain linked until Earth return.  On arrival, TV performs the capture burn.  DEV performs all further orbital maneuvers, matching orbit with either satellite, rendezvousing, and a slow landing.  TV-DEV nominally operate at the moon for a month; or if supplies allow a +200 day mission with visits to multiple sites at both Phobos and Deimos.  Leaving Mars, either DEV or TV provide the depature burn.  Earth return would use a combination of burns and Lunar Gravity assist, as TV-DEV is aiming for a Lagrange point.  DEV remains at the EML-point while the TV returns the crew to Earth itself.

More to come...
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Offline Russel

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As you know I favor an architecture that uses a lander/ascent vehicle that is also capable of visiting the moons. And has a capable transit vehicle that also doubles as a mars orbiter. I can foresee the possibility that such an architecture might be tested incrementally with a visit to the moons of Mars (and a robotic trip for the lander/ascent vehicle perhaps.)

Offline redliox

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As you know I favor an architecture that uses a lander/ascent vehicle that is also capable of visiting the moons. And has a capable transit vehicle that also doubles as a mars orbiter. I can foresee the possibility that such an architecture might be tested incrementally with a visit to the moons of Mars (and a robotic trip for the lander/ascent vehicle perhaps.)

Right, the moons could be used as a testing ground.  A Martian Hab, especially if meant as a base module, is typically imagined as a fat tuna can vaguely resembling an ISS module with legs.  If you can build a module, you can build a lander.  In the micro-gravity of Phobos, Deimos, and any asteroid things can be treated as if for the ISS.

Specifically, the element on the lander that needs the most attention would be the legs and rocketry.  Best to gear them for Mars but the low-g makes testing easier plus, especially if the initial mission needs all the fuel it can get, the moon-lander can assist on MOI and departure.

Further bonus, presuming the delta-v needs of a round Martian trip are met, the Deimos/Phobos lander could be reused.  A Martian lander will go through so much hell in its use...but not so much the Deimos/Phobos route.
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Offline redliox

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Now for an overall review of what's required to operate at the Martian moons...

We are going to presume the vehicles launched to Mars escape from Earth via SLS and its EDS; given that each element (the TV and DEV) launches seperately to an EML-point rendezvous both would have an EDS with a partial fuel load it's not an unreasonable assumption.  The mission vehicles will need to accomplish the following: capture into Mars orbit, transfer to Deimos, land on Deimos, transfer to Phobos, land on Phobos, escape from Mars, and capture into an EML-point.  Plenty of work to do...

The majority of my math can be referred to here, when I initially conceived the figures the TV will be dealing with: http://forum.nasaspaceflight.com/index.php?topic=36713.80

Capturing into Mars orbit is the heaviest chore, in part because all the mission mass is being delivered.  Prior delving into the minimum requirements to enter a high but stable Martian, aiming for an orbit of approximately 48 hours, yielded the need for a full kilometer/second of effort.  Aerocapture would require far less fuel, but applying it to a crewed vehicle would be a dicey matter...best applied to the Mars lander when its turn comes.

As for moving between Capture/High orbit and Deimos and finally Phobos, the lump sum in delta-V is approximately 1.95 km/sec.  Suffice to say, budgeting for 2 km/sec of delta-V between capture and Phobos landing works well.

Escape from Mars is would happen from Phobos' orbit, since the mission would be working inward.  Escaping from Phobos' orbit totals to 1.8 km/sec; by comparison an escape from Deimos would be 1.4 km/sec; this difference is inconsequential for a conjunction-class-long-stay mission that visits bot moons, but for an opposition-class-short-stay mission it could merit focusing on Deimos.  Again, budgeting for 2 km/sec should cover the needs.

Finally, the return to Earth...or more specifically the EML-point where DEV could park and the TV returns the crew from.  The best figures I can find place this number between 1.1 and 2.  I am going to assume 1.5 km/sec (but if someone can crunch the numbers from a capture from Mars into the EML-points please inform us all here).

So the numbers the mission will deal with sequentially: 1+2+2+1.5 for a total of 6.5 km/sec of delta-v.
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Offline redliox

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I've done recalculations and found both Deimos and Phobos would be impossible to accomplish purely propulsively with only 105 tonnes.  However, it may remain possible to venture to Deimos and still accomplish a viable orbital mission of Mars and its outermost moon.

Referring back to math I did for 'Mars Aligned' (here at: http://forum.nasaspaceflight.com/index.php?topic=36713.80 ), a 105 tonnes mass breaking into High Mars Orbit to perform the first 1 km/sec the mission requires 25 tonnes of that mass to be fuel (in this case a methane/oxygen mix).  This burn would come from the TV, as, during a Mars mission, it would provide breaking for both itself and the Mars lander, so with an identical mass it performs the same role and function.  Out of its total mass of 55 tonnes 35.75 is propellant, so post burn it is left with just over 10 tonnes of fuel.

The next two sets of challenges will draw from the DEV vehicle, which will be a 50 tonne lander charged with handling the majority of maneuvers within Mars' sphere of influence.  A potential logistics nightmare...however, DEV will be dropping mass substantially, translating to better results in velocity changes.  Presuming DEV's mass is 80% propellant, it has 40 tonnes to draw upon.  I will divide this up incrementally, first with Deimos ops, then Martian escape....

Beginning from High Orbit, the remaining TV-DEV mass begins at 80 tonnes.  The complete transfer and landing to Deimos comes down to just a hair less than 1km/sec of effort.  Calculating via stout.net's Delta-V, this will require 19 tonness of propellant, 6 tonnes less than capture did.  The mass of DEV reduces from 50 to 31, down to a mission total of 61 tonnes at this point.

With TV-DEV still high in the Martian gravity well escape from Deimos requires precisely 1.67 km/sec, 2 km/sec to give margin.  DEV's remaining fuel at this point is 21 tonnes, translating to 1.57 km/sec - over 90% of the total escape from Deimos/Mars.  The final 0.42 km/sec would draw from the TV, which would be pushing the remaining 40 tonnes with 5 tonnes of its propellant.

The final leg of the journey would be difficult in delta-v, yet still hopeful.  The remaining mass of the 105 initially sent to Mars is 35 tonnes at this point, only 5 of which is propellant.  This amounts to just under 0.6 km/sec.  Were the TV-DEV aiming for LEO or Earth itself, returning the crew would be impossible.  However, including trajectory corrections en route to Earth, a minimum capture into an orbit nominally meant for transferring to the Moon is around 0.48 km/sec, within the narrow budget of the Transfer Vehicle.

A further source of delta-v is Earth's own moon.  As is, Luna has been used for gravity assists: the STEREO twins, Nozomi, WIND, GEOTAIL, and even a handful of communication satellites stranded in off-orbits.  Furthermore, given the attention ARM has with lunar orbit, NASA will devote research into lunar operations years before Mars gets put on paper (even if ARM is ultimately canceled).  Only actual delta-v numbers I lack, but I see 2 angles in which the arriving TV-DEV utilizes a lunar gravity assist:
1) Directly from Mars, TV-DEV pass directly in front of Luna's path.  Luna's gravity naturally pulls on the paired spacecraft, the Moon absorbing delta-v much as Venus did for both Mercury-bound Mariner 10 and MESSENGER.
2) In a loose High Earth Orbit, TV-DEV make a pass of Luna, possibly multiple ones, that shape its trajectory about Earth.

The end result, at least hypothetically, would put TV-DEV at one of the Lunar Lagrange points.  Given the Moon obviously influences these points, it is within reason drawing on Luna's gravity can slide the pair into one.  From there, the Orion-based TV finally separates and returns in the style of Orion and Apollo to Earth.  DEV can be refurnished and reunited with a new TV for a further mission to Deimos.

The math to do a purely Earth-fueled, chemically-propelled expedition to Deimos is definitely brutal.  Out of the 105 tonnes sent to Mars & Deimos, roughly 30 is dry weight (10 for DEV and 20 for TV) with 75 solidly propellant.  Near the end fuel becomes crucial, but can complete the trip and even reuse a landing craft.

In light of TakeOff's "What would orbiting Mars teach us?" thread (here for due credit: http://forum.nasaspaceflight.com/index.php?topic=37218.0 ), I come to the following conclusions in regards to my Deimos mission here:
1) Orbiting Mars is difficult, specifically for human spaceflight, with current technology.
2) Deimos is the threshold for current human spacecraft, but just
3) Resources (especially propellant) from Earth alone aren't sufficient for exploring Mars or Phobos.
4) Ways for reducing velocity without propellant are required (be it ion drive or aerobraking).

Specifically in my Deimos Aligned architecture (which is a modification from Mars Semi-Direct), we would be able to ensure a safe way of entering and exiting Martian space without anything more exotic than methane/oxygen rocketry.  However, Mars' son of terror marks the boundary line for anything directly tied to Earth.  A Martian lander, after safe delivery to high orbit, would need to worm its way downward by unconventional means while a return vehicle could await by Deimos.
« Last Edit: 04/08/2015 11:51 AM by redliox »
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Offline Nilof

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You could significantly reduce the mass needed with a near-term SEP/Chemical hybrid system(SEP for spiral-out and interplanetary transfers, chemical for Mars capture and orbital maneuvers). This study is a good place to start, and also considers the detailed logistics of the habitat: spirit.as.utexas.edu/~fiso/telecon/Oleson_3-6-13/Oleson_3-6-13.pdf

They get 190 tonnes for a rather spacious six-man vehicle vs 415 tonnes with a 450s Isp hydrolox departure stage and 330s Isp storable thrusters. You also get the advantage of full reusabillity so that the same spacecraft can be used for a later Mars landing mission if you use it for a Mars moon mission at first.

EDIT: attached the pdf to prevent link rot.

« Last Edit: 04/08/2015 02:47 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 redliox

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You could significantly reduce the mass needed with a near-term SEP/Chemical hybrid system(SEP for spiral-out and interplanetary transfers, chemical for Mars capture and orbital maneuvers). This study is a good place to start, and also considers the detailed logistics of the habitat: spirit.as.utexas.edu/~fiso/telecon/Oleson_3-6-13/Oleson_3-6-13.pdf

They get 190 tonnes for a rather spacious six-man vehicle vs 415 tonnes with a 450s Isp hydrolox departure stage and 330s Isp storable thrusters. You also get the advantage of full reusabillity so that the same spacecraft can be used for a later Mars landing mission if you use it for a Mars moon mission at first.

My question for SEP, while decidedly useful for unmanned craft, is how much time?  It's so damn slow for crew timelines.  Unless it can do at least 1 km/sec of change within a month for at least a 50 tonne vehicle it isn't worth the time versus chemical/aerobrake.
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Offline Nilof

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The exact time figures depend on the particular launch window used. The stay time quoted in the study is 300 days with near term SEP/chemical hybrid vs 500 days for chemical(if you allow a faster trajectory than the minimum energy one used in the mass comparison for fairness), with a similar total mission time. In other words, you still get lots of time for science work with both. For Mars moon missions there is almost no difference between the stay time and the transit time in term of the effect on crew health.

With that said, if instead of looking at equal crew facilities but lower cost SEP you look at equal mass architectures, you can give the crew much better facilities. Such as better exercising equipment or a Nautilus-type centrifuge. Or you could significantly expand the science-related payload and bring a full lab with you.

In general, I think more payload > shorter transit time on the short term. On the longer term, you have the advantage that SEP scales better and better leading to shorter transit times. This has the effect of making chemical architectures increasingly obsolete as time passes.
« Last Edit: 04/08/2015 01:59 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 redliox

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Finally a few numbers on Lunar gravity assist!  http://www.permanent.com/space-transportation-lunar-gravity-assist.html

Quoting from PERMANENT specifically on the delta-v of an LGA:
Quote
The maximum braking the Moon can provide is about 2.2 km/sec, using a "double lunar gravity assist", whereby the asteroid passes by the Moon coming in, then past the Earth, then past the Moon again going back out. This would divert the asteroid by almost 90 degrees from its original path, and capture it into a highly elliptical Earth orbit. Subsequent gravity assists would insert it into a more circular orbit around Earth after which it would perform final small thrusting maneuvers to achieve its desired destination orbit.

Many asteroids require a delta-v of much less than 2.2 km/sec, and require only a single lunar gravity assist (not an Earth gravity assist) to be captured, and optionally additional lunar gravity assists to divert the asteroid into a more circular orbit.

So flying by Luna could reasonably yield 2 km/sec to brake the incoming TV-DEV.  The initial 1.5 km/sec for reaching a Lagrange point could be met with the remaining ~0.6 km/sec worth of fuel offering ample margin.  Hypothetically this could also work for a departing mission, using fly-bys to boost the craft to a classic Hoffman transfer orbit.

If anyone can crunch Lunar ga number further do post the results here.
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Offline redliox

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Delving into a Deimos-centered architecture once again.  Posting it in here and placing the numbers I've crunched.  In this version I am presuming 3 elements (4 if you count Orion shuttling up crew) all meant to put a crew around Mars and to explore Deimos and areosynchronous orbit before a return to Earth.  I'll begin by naming these elements and their base masses:

Mars Transit Habitat
Hab-30 mt/Kick Stage (Dry) 6.3 mt
Propellant 9.6 metric tons (N204/MMH - ISP 336)
Total Mass: 45.9 metric tons


Methalox Stage
Total Dry Mass: 15 metric tons
Total Wet Mass: 115 metric tons
Total propellant: 100 metric tons (MH4/O2 - ISP 380)


Interplanetary Ion Tug - ITIT
Total Dry Mass: 30 metric tons
Total Wet Mass: 80 metric tons
Total Propellant: 50 metric tons (Xenon - ISP 1100)


Next, I will show the numbers these elements will deal with:

Earth Departure Operations
LEO to EML 1 -  4 km/s
TMI from EML 1 - 1.5 km/s


Mars Orbital Operations
MOI - 1.5 km/s
Deimos - 1.1 km/s
Reserve - 0.5 km/s
ETI - 1.9 km/s
Total: 5 km/s


Earth Return Operations
EML 1 Capture - 2 km/s
Earth return - 0.7 km/s
Total: 2.7 km/s


The MTH and Methalox stages would unite at EML-1 whereas the ITIT would be delivered to areosynchronous orbit.  In short, the mission would start chemically from Lunar space and return via SEP to it.  From the MTH's POV, it would undergo 9 km/s of delta-v to complete a round trip in the following sequence:

1) TMI from EML 1 - Mass Before 275.9 mt / Mass After 175.9 mt
Propellant consumed: 100 mt
Delta-v: 1.6 km/s
2) 1st Methalox Stage Jettisoned
3) MOI - Mass Before 160.9 mt / Mass After 105.9 mt
Propellant consumed: 55 mt
Delta-v: 1.5 km/s
4) Deimos - Mass Before 105.9 mt / Mass After 77.9 mt
Propellant consumed: 28 mt
Delta-v: 1.1 km/s
5) Reserve - Mass Before 77.9 mt / Mass After 60.9 mt
Propellant consumed: 17 mt
Delta-v: 0.9 km/s
6) Jettison 2nd Methalox Stage
7) Dock with ITIT

8) ETI - Mass Before 125.9 mt / Mass After 104.4 mt
Propellant consumed: 21.5 mt
Delta-v: 2 km/s
9) EM 1 Capture - Mass Before 104.4 mt / Mass After 84.4 mt
Propellant consumed: 20 mt
Delta-v: 2.2 km/s

Total d-v from all this: 9.3 km/s.  ITIT would still have 8.5 metric tons of xenon and the MTH would have 9.6 to further draw upon, the later particularly for docking maneuvering.
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Offline kris

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Out of curiosity how do phobos and deimos compare in delta v needed to get there and back?

I sure hope the http://en.wikipedia.org/wiki/Phobos_And_Deimos_%26_Mars_Environment get's selected.

Offline redliox

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Out of curiosity how do phobos and deimos compare in delta v needed to get there and back?

I sure hope the http://en.wikipedia.org/wiki/Phobos_And_Deimos_%26_Mars_Environment get's selected.

There's a section in Space Science about Phobos/Deimos spacecraft - check there.

As for Phobos and Deimos, flying around each one is child's play.  As in about 1.5 km/s of dv each is sufficient to handle.  If anything's trick it's getting enough dv to handle both simultaneously.
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Offline Oli

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Thoughts:

- Some of your delta-v numbers seem to be significantly higher than the ones that can be found on wiki for example. Not saying yours are wrong, but I wonder why the difference.

- 30t for the Habitat incl. supplies is probably not enough. I don't know the crew size and mission duration you assume, but I think for a crew of 4 for ~1000 days it should definitely be more.

- 6.3t dry mass for a kick stage with 9.6t propellant? Or is that the excursion vehicle?

- 30t dry mass for the SEP tug is also rather high, unless it's very powerful. How many kw do you assume?

- SEP isp of only 1100s?

- Not sure why you don't use the SEP tug on the inbound trajectory, in combination with chemical.
« Last Edit: 05/14/2015 08:06 PM by Oli »

Offline redliox

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- Some of your delta-v numbers seem to be significantly higher than the ones that can be found on wiki for example. Not saying yours are wrong, but I wonder why the difference.

I've had many people criticize on 'not enough', and warned that specifically Martian launch windows can vary in their dv needs.  I figured to take things more conservatively.  I've seen charts with the numbers vary a little too, so I've tried to go by the mean numbers in general.

An example: the alleged "minimum" dv to capture into Mars orbit is 0.6 km.  More than likely this would be a >50 hour orbit easily influenced by the Sun or even Jupiter.  When I referenced space probes like the MGS and MRO, when each performed MOI into an elliptical orbit they both targeted 1km/s or more, and I confirmed it through their websites and printed sources; more to the point it establishes a working range to aim for. 

- 30t for the Habitat incl. supplies is probably not enough. I don't know the crew size and mission duration you assume, but I think for a crew of 4 for ~1000 days it should definitely be more.

- 6.3t dry mass for a kick stage with 9.6t propellant? Or is that the excursion vehicle?

Referencing numbers from the Deep Space Habitat thread (with it's documents) and Mars Direct.  Both respective architectures seem confident large habitats could include consumables for years within a 25 mt mass; me boosting it to 30 is my way to ensure more can be stowed away.

The kick stage is a reference to a stage derived from the Orion service module.  The DSH papers included some numbers for station keeping fuel, but considering this would be a vehicle actively involved in exploration I thought giving it a propulsion module in addition to the ITIT would be prudent.  At the habitat's heft, the same system that gives Orion roughly 1.5 km/s of dv would only give about 0.7 km/s.  However, it isn't meant for significant maneuvers; its main function would be delicate docking maneuvers, the landing sequence on Deimos (possibly Phobos too), EML-point adjustments, and "emergency" maneuvers that need speed the ion drive couldn't give.  Hypothetically a methalox stage could be used in its place, but if we're planning for the near term hypergollics is the more practical choice (rockets already available and lasts years without any cryogenic wrapping paper needs).

- 30t dry mass for the SEP tug is also rather high, unless it's very powerful. How many kw do you assume?

- SEP isp of only 1100s?

Was giving a large margin of error presuming a heavy system with less-than-ideal engines and power supply.  Whenever I look at the Boeing Mars proposal, for example, you can't tell me that giraffe-crossed-with-a-TIE-fighter will weigh less than 10 tons even without fuel.  What I want is something more compact and focused on moving just the crew in their hab, treating the setup like a shuttle visiting a space station but with more supplies.

1100 isp is a lower number I've seen mentioned for older ion engines and better practiced with.  Reading into Hall thrusters, they actually have a throttle range from 1000 to 4100.  Again, for argument sake, I worked with the lower "conservative" end.  Ideally, I would hope for a setup that needs 20 kW or less, is 30 tons or less in equipment, and uses 50 tons of propellant (xeonon, argon, anything minimally toxic) with an average isp of 3000.  I suspect while 4000 is the max...when at Mars, they would have to throttle back.

- Not sure why you don't use the SEP tug on the inbound trajectory, in combination with chemical.

Simple: too much mass.  Elaborating further, the first flight to Mars is looking to be orbital, which means options may have to be kept non-cryogenic unless you're using it within say 2 months time; furthermore the first mission probably won't have a fuel supply on Mars established just yet.  That leaves SEP and hypergollics, and hypergollics are damnably heavy in quantity (believe me, I tried the calculations for N204/MMH, H2/O2, and CH4/O2).

After this, going to consolidate my 'Aligned architectures into a single thread.
« Last Edit: 05/15/2015 08:10 AM by redliox »
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Offline Oli

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An example: the alleged "minimum" dv to capture into Mars orbit is 0.6 km.  More than likely this would be a >50 hour orbit easily influenced by the Sun or even Jupiter.  When I referenced space probes like the MGS and MRO, when each performed MOI into an elliptical orbit they both targeted 1km/s or more, and I confirmed it through their websites and printed sources; more to the point it establishes a working range to aim for. 

Referencing numbers from the Deep Space Habitat thread (with it's documents) and Mars Direct.  Both respective architectures seem confident large habitats could include consumables for years within a 25 mt mass; me boosting it to 30 is my way to ensure more can be stowed away.

1100 isp is a lower number I've seen mentioned for older ion engines and better practiced with.  Reading into Hall thrusters, they actually have a throttle range from 1000 to 4100.  Again, for argument sake, I worked with the lower "conservative" end.  Ideally, I would hope for a setup that needs 20 kW or less, is 30 tons or less in equipment, and uses 50 tons of propellant (xeonon, argon, anything minimally toxic) with an average isp of 3000.  I suspect while 4000 is the max...when at Mars, they would have to throttle back.

- Yes, but assume 1.5 km/s for capture. For EML to Mars transfer you assume 1.6 km/s, which is like twice the value of numbers I find elsewhere. I guess the values for the way back are for low-thrust propulsion, that's why they're higher.

- I have yet to see a Mars mission design with a Habitat of less than 40t. And that's only for the transfer, not for the stay on the surface. A Deimos mission is likely going to need significantly more unless you do ISRU on Deimos. I think Mars direct is not realistic, not only in that aspect.

- Well, a Martian moons mission is not going to happen for another 15 years at least, so there's that. It's just that in the papers I've read SEP performance is usually assumed to be lot better. For example in the one attached. A tug with 300kw, 10-12t dry mass, 40t xenon and 2000s isp.

« Last Edit: 05/15/2015 09:30 AM by Oli »