Phobos and Deimos on the return to Earth

[redliox]

While I'd love to see exploration of the Martian moons included with journeying to the red planet, the current thinking to reach them or to perform an orbital mission requires a huge amount of buildup for a single expedition...most of which amounts to delivering fuel thanks to physics.  By comparison a surface mission at least offers the advantage of fueling from Mars whereas the moons don't have resource utilization options.  Instead of visiting either moon on the inbound phase of a mission, what about visiting them on the outbound phase, just before Earth return?

I was imagining something akin to Mars Direct or SpaceX's ITS where the mission focuses on the surface and fueling from Martian resources.  The moons come into play on the outbound, Earth-return phase of the mission.  This would be the rough layout:

1) Crew vehicle launches from Mars directly into a transfer orbit
2) Crew vehicle adjusts from transfer to circular orbit (one that near-matches target moon).
3) Crew vehicle rendezvouses with a Moon vehicle.
4) Exploration of moon via Moon vehicle.
5) Crew vehicle departs for Earth.

The Moon vehicle would have been positioned before the mission and possibly reusable for later ones.  Main questions I have lie with how much trouble it'd be to:
1) Launch from Mars surface directly into a transfer orbit matching either Phobos or Deimos orbit
2) Doing a burn that circularizes the transfer to a circular orbit
3) Departing to Earth from said circular orbit
I'd presume a near-Phobos orbit would have a period of 7.5 hours and 30 hours for near-Deimos.

I am aware the disadvantage of visiting the moons is, either on the outbound or inbound, circularizing/matching.  I am simply trying to compute the needs and delta-v of the scenario I explained above.  I'd especially like to compare visiting Phobos to visiting Deimos.  I would only expect to visit one or the other, not both since that'd further increase the fuel load for orbital maneuvering.  We'd be leveraging ISRU from Mars, but the priority would be reserving enough propellant for Earth return after visiting a single moon.

Summon the delta-v masters!  (since mathematician isn't flamboyant enough)

[guckyfan]

Best may be doing Phobos and Deimos from the surface of Mars, then returning to Mars for refuelling again.

[redliox]

Best may be doing Phobos and Deimos from the surface of Mars, then returning to Mars for refuelling again.

Show the numbers then.  'Best' is an opinion, not a fact.

[guckyfan]

Best may be doing Phobos and Deimos from the surface of Mars, then returning to Mars for refuelling again.

Show the numbers then.  'Best' is an opinion, not a fact.

You do know the meaning of "may"?

My reasoning is simple. Getting from Phobos or Deimos to the surface of Mars requires less delta-v than going to earth.

[ThereIWas3]

And returning to Mars lets you top-up the tanks for the trip back to Earth.   Doing the moons on the way to Earth means you have to carry the entire to-Earth fuel load, plus what is needed for the moon arrival and departure.

[redliox]

And returning to Mars lets you top-up the tanks for the trip back to Earth.   Doing the moons on the way to Earth means you have to carry the entire to-Earth fuel load, plus what is needed for the moon arrival and departure.

Heatshields tend to be disposable, heavy things bear in mind.  I'm not debating multiple refueling stops and orbital forays.  I DID spell out in the thread title for specifying return-to-Earth.

[redliox]

My reasoning is simple. Getting from Phobos or Deimos to the surface of Mars requires less delta-v than going to earth.

Getting to Low Mars Orbit or High Mars Orbit from the moons is also easier than getting to either the Martian surface or Earth if the objective is to conserve delta-v, especially Low Orbit with aerobraking.

[redliox]

Reference thanks to Dalhousie: http://old.marssociety.org.au/library/willson-et-al.pdf
In this case they specify Phobos.

[guckyfan]

Heatshields tend to be disposable, heavy things bear in mind.  I'm not debating multiple refueling stops and orbital forays.  I DID spell out in the thread title for specifying return-to-Earth.

A heatshield for any vehicle flying between Earth and Mars should be of the advanced types in the class of PicaX and be able to do a number of reentries. Especially a less stressing one like returning to Mars from orbit.

But you did specify on the return to Earth, so my suggestion was in some way OT, though I believe it is the better option. I won't continue arguing it here.

[redliox]

Heatshields tend to be disposable, heavy things bear in mind.  I'm not debating multiple refueling stops and orbital forays.  I DID spell out in the thread title for specifying return-to-Earth.

A heatshield for any vehicle flying between Earth and Mars should be of the advanced types in the class of PicaX and be able to do a number of reentries. Especially a less stressing one like returning to Mars from orbit.

But you did specify on the return to Earth, so my suggestion was in some way OT, though I believe it is the better option. I won't continue arguing it here.

Correct; there are many scenarios.  Refueling and reusability are great ideas.  I was simply trying to keep things on topic with no offense intended.

[dror]

IF some ISRU can be done on phobos and you can get some H2+O2 fuel on it, this could save some significant ISRU on Mars

[Hop_David]

While I'd love to see exploration of the Martian moons included with journeying to the red planet, the current thinking to reach them or to perform an orbital mission requires a huge amount of buildup for a single expedition...most of which amounts to delivering fuel thanks to physics.  By comparison a surface mission at least offers the advantage of fueling from Mars whereas the moons don't have resource utilization options.  Instead of visiting either moon on the inbound phase of a mission, what about visiting them on the outbound phase, just before Earth return?

I was imagining something akin to Mars Direct or SpaceX's ITS where the mission focuses on the surface and fueling from Martian resources.  The moons come into play on the outbound, Earth-return phase of the mission.  This would be the rough layout:

1) Crew vehicle launches from Mars directly into a transfer orbit
2) Crew vehicle adjusts from transfer to circular orbit (one that near-matches target moon).
3) Crew vehicle rendezvouses with a Moon vehicle.
4) Exploration of moon via Moon vehicle.
5) Crew vehicle departs for Earth.

The Moon vehicle would have been positioned before the mission and possibly reusable for later ones.  Main questions I have lie with how much trouble it'd be to:
1) Launch from Mars surface directly into a transfer orbit matching either Phobos or Deimos orbit
2) Doing a burn that circularizes the transfer to a circular orbit
3) Departing to Earth from said circular orbit
I'd presume a near-Phobos orbit would have a period of 7.5 hours and 30 hours for near-Deimos.

I am aware the disadvantage of visiting the moons is, either on the outbound or inbound, circularizing/matching.  I am simply trying to compute the needs and delta-v of the scenario I explained above.  I'd especially like to compare visiting Phobos to visiting Deimos.  I would only expect to visit one or the other, not both since that'd further increase the fuel load for orbital maneuvering.  We'd be leveraging ISRU from Mars, but the priority would be reserving enough propellant for Earth return after visiting a single moon.

Summon the delta-v masters!  (since mathematician isn't flamboyant enough)

Phobos and Mars are near equatorial. If launching from the equator, there's little plane change expense.

Here's a few numbers just based on the vis viva equation, not looking at gravity loss during vertical ascent to get above Mars' thickest atmosphere:

Mars to 300 km LMO: 3.7 km/s
Mars to Phobos: 4.9 km/s
Mars to Deimos: 5.4 km/s

Deimos would be more interesting if we were using ion engine interplanetary transfer vehicles such as the Hermes in The Martian.

A rough rule of thumb for delta V with low thrust engines is subtracting speed of destination orbit from speed of departure orbit. An ion craft would enter Mars sphere of influence with a C3=0 traveling close to 0 km/s wrt to Mars. Deimos travels 1.4 km/s. A 300 km LMO is about 3.4 km/s. So rendezvous with Deimos would save about 2 km/s delta V over parking in LMO.

"Wait a minute!" someone might say. "You can use aerobraking to park in LMO. So LMO isn't 3.4 km/s." To achieve a decent alpha, Hermes would need either acres of solar panels or acres of light weight radiator are to dump a nuke power source's waste heat. In either case the power source would be large and fragile and not able to endure aerobraking.

Weir gave Hermes a 2 mm/s^2 acceleration. That assumes an implausibly good alpha (in my opinion) but we'll go with it. An extra 2 km/s at 2 mm/s^2 acceleration would take a million seconds or about 12 days. I believe 1 mm/s^2 is perhaps plausible in which case it'd take an extra 24 days to spiral down to LMO.

An interplanetary ship harbored in a planet's orbit must depart as well as arrive. Harboring at Deimos vs LMO would save 4 km/s of propellent and 24 to 48 days of time depending on what acceleration an ion craft could achieve.

For a similar reason I advocate harboring something like the Hermes at EML2 rather than LEO. Parking at the edge of earth's gravity rather than LEO would save Hermes 14 km/s climbing in and out of earth's gravity well. And skipping the climb to near bottom and back would save 80 days if acceleration is 2 mm/s^2. 160 days if we use a more plausible 1 mm/s^2.

I'm attaching a graphic illustrating my day dream where ion ships are harbored as close as possible to the edge of planetary gravity wells.

[redliox]

Phobos and Mars are near equatorial. If launching from the equator, there's little plane change expense.

Here's a few numbers just based on the vis viva equation, not looking at gravity loss during vertical ascent to get above Mars' thickest atmosphere:

Mars to 300 km LMO: 3.7 km/s
Mars to Phobos: 4.9 km/s
Mars to Deimos: 5.4 km/s

Deimos would be more interesting if we were using ion engine interplanetary transfer vehicles such as the Hermes in The Martian.

A rough rule of thumb for delta V with low thrust engines is subtracting speed of destination orbit from speed of departure orbit. An ion craft would enter Mars sphere of influence with a C3=0 traveling close to 0 km/s wrt to Mars. Deimos travels 1.4 km/s. A 300 km LMO is about 3.4 km/s. So rendezvous with Deimos would save about 2 km/s delta V over parking in LMO.

"Wait a minute!" someone might say. "You can use aerobraking to park in LMO. So LMO isn't 3.4 km/s." To achieve a decent alpha, Hermes would need either acres of solar panels or acres of light weight radiator are to dump a nuke power source's waste heat. In either case the power source would be large and fragile and not able to endure aerobraking.

Weir gave Hermes a 2 mm/s^2 acceleration. That assumes an implausibly good alpha (in my opinion) but we'll go with it. An extra 2 km/s at 2 mm/s^2 acceleration would take a million seconds or about 12 days. I believe 1 mm/s^2 is perhaps plausible in which case it'd take an extra 24 days to spiral down to LMO.

An interplanetary ship harbored in a planet's orbit must depart as well as arrive. Harboring at Deimos vs LMO would save 4 km/s of propellent and 24 to 48 days of time depending on what acceleration an ion craft could achieve.

For a similar reason I advocate harboring something like the Hermes at EML2 rather than LEO. Parking at the edge of earth's gravity rather than LEO would save Hermes 14 km/s climbing in and out of earth's gravity well. And skipping the climb to near bottom and back would save 80 days if acceleration is 2 mm/s^2. 160 days if we use a more plausible 1 mm/s^2.

I'm attaching a graphic illustrating my day dream where ion ships are harbored as close as possible to the edge of planetary gravity wells.

Clearly you favor ion propulsion; the only thing I have against it is it's too slow for human transport although suitable for most probes.

A critical question I have is do those numbers, 4.9 and 5.4 for Phobos and Deimos respectively, include the burns to circularize an orbit?  I assume your ion engines would spiral outward from a low orbit but, assuming, purely chemical rockets do the orbital maneuvers, basically periapsis raises, alter those numbers significantly?

[Hop_David]

A critical question I have is do those numbers, 4.9 and 5.4 for Phobos and Deimos respectively, include the burns to circularize an orbit?

Yes. I'm hoping people will use my Hohmann spreadsheet. I'll walk through the steps with a transfer orbit from Mars surface to Phobos.

Altitude periapsis transfer orbit is 0. Altitude apoapsis is same as altitude of Phobos orbit, 5981 km.

Elliptical transfer orbit is moving 4.3033 km/s at periapsis (Mars surface). To circularize at apoapis takes .5787 km/s. I add these together to get 4.88206 km/s. Which I call 4.9 km/s since my model makes simplifying assumptions that reduce precision.

Attached is a screen capture of what I'm talking about. I locked most the spreadsheet to keep myself from accidentally overwriting important cells. The blue area is left unlocked so I can use it as scratch paper.

[Rei]

Phobos and Mars are near equatorial. If launching from the equator, there's little plane change expense.

Of course, Mars's axial tilt is ~25 degrees, so the orbits are not aligned with the ecliptic.

"Wait a minute!" someone might say. "You can use aerobraking to park in LMO. So LMO isn't 3.4 km/s." To achieve a decent alpha, Hermes would need either acres of solar panels or acres of light weight radiator are to dump a nuke power source's waste heat. In either case the power source would be large and fragile and not able to endure aerobraking.

How did the modules that make up Hermes survive launch from Earth if they're that fragile? They were launched retracted, right? Nowadays, large solar power systems are generally deployed on retractable booms or fans. If you needed to, there's no reason you couldn't engineer the deployment to be reversible, so you can retract for aerobraking.

Not that aerobraking is by any stretch a simple process.  But...

An interplanetary ship harbored in a planet's orbit must depart as well as arrive. Harboring at Deimos vs LMO would save 4 km/s of propellent and 24 to 48 days of time depending on what acceleration an ion craft could achieve.

On the other hand, if your objective is to have a large heavy transfer stage, and a small but high dV ascent/descent stage , it raises the possibility question, why not go all the way and put the heavy craft on a cycler trajectory?

[Hop_David]

Clearly you favor ion propulsion; the only thing I have against it is it's too slow for human transport although suitable for most probes.

If a ship's acceleration is a tiny fraction of local gravity field, path will be gradually expanding spiral. If ship's acceleration is a significant fraction of local gravity field, burns can be modeled as impulsive burns allowing ship to enter and exit an elliptical transfer path.

In LEO local acceleration is is around 9 km/s^2. Obviously 1 mm/s^2 is a tiny fraction of that.

But heliocentric orbits are a different story. At 1 A.U. sun's acceleration is about 6 mm/s^2. At Mars, less than 3 mm/s. So a ship with 1 mm/s^2 acceleration would do something closer to an elliptical transfer orbit rather than the slow ion spirals we leaving LEO.

tl;dr Ion engines suck at climbing in and out of planetary gravity wells. But make good time for moving between heliocentric orbits.

[redliox]

A critical question I have is do those numbers, 4.9 and 5.4 for Phobos and Deimos respectively, include the burns to circularize an orbit?

Yes. I'm hoping people will use my Hohmann spreadsheet. I'll walk through the steps with a transfer orbit from Mars surface to Phobos.

Altitude periapsis transfer orbit is 0. Altitude apoapsis is same as altitude of Phobos orbit, 5981 km.

Elliptical transfer orbit is moving 4.3033 km/s at periapsis (Mars surface). To circularize at apoapis takes .5787 km/s. I add these together to get 4.88206 km/s. Which I call 4.9 km/s since my model makes simplifying assumptions that reduce precision.

Attached is a screen capture of what I'm talking about. I locked most the spreadsheet to keep myself from accidentally overwriting important cells. The blue area is left unlocked so I can use it as scratch paper.

Impressive.  I see that the periapsis burn for a Phobos orbit just under 600 m/s.  I will take your word that you included the same burn for Deimos in its number, but how does that one likewise compare?

I ask these kinds of questions because, whenever I get commentary on mission setups, I get numbers like these flung in my face among technical concerns.  The biggest challenge for the moons, in this case at least, is ensuring there's reserve fuel for the return to Earth afterwards.

Does leaving Mars from a circling orbit effect departure any differently from the Martian surface or an elliptical orbit?

[redliox]

Gave the spreadsheet a try, in this case for Deimos.

Assuming I used it correctly, a periapsis raise results in 680 m/s required to put the Martian vehicle into a Deimos orbit.  Also reading your spreadsheet, I further see that the transfer orbit's period would be 13.12 hours, meaning after launch from Mars the vehicle (in this case for Deimos of course) the raise burn happens just over 6 hours later.

Is this correct so far?

[Hop_David]

How did the modules that make up Hermes survive launch from Earth if they're that fragile?

If we had dexterous robotic avatars, something like Hermes could be launched in several payloads and assembled in orbit. I regard this as plausible. Launching the whole shebang from earth's surface is less plausible.

They were launched retracted, right? Nowadays, large solar power systems are generally deployed on retractable booms or fans. If you needed to, there's no reason you couldn't engineer the deployment to be reversible, so you can retract for aerobraking.

We're taking about acres of something like cellophane here. Retracting and deploying them each trip would be a nightmare.

On the other hand, if your objective is to have a large heavy transfer stage, and a small but high dV ascent/descent stage , it raises the possibility question, why not go all the way and put the heavy craft on a cycler trajectory?

I don't think an Aldrin cycler is plausible. All Mars cycler schemes I've seen rely on the notion that earth Mars synodic period is 2 1/7 years giving a cycler geometry that repeats every 15 years. The fact that it isn't quite 2 1/7 complicates things. Moreover Mars eccentricity and inclination also requires course corrections. And the big kicker is you need to rotate the Aldrin cycler line of apsides 50º every orbit. It would take a lot of propellent to maintain the cycler orbit, especially if the cycler is massive.

Also the Aldrin cycler has an aphelion clear out in the Main Belt. During fly by the cycler's velocity vector is almost 70º from Mars' velocity vector. This makes for a big delta V budget for the taxis traveling between Mars and cycler. Around 9 km/s, if memory serves.

The VISIT 1 and 2 cyclers are lower maintenance and the Taxis don't need whopping big delta V budgets. But these cyclers don't pass by each planet each synodic period. If I remember right, sometimes passengers would need to sit tight for as long as 7.5 years.

[Hop_David]

Gave the spreadsheet a try, in this case for Deimos.

Assuming I used it correctly, a periapsis raise results in 680 m/s required to put the Martian vehicle into a Deimos orbit.  Also reading your spreadsheet, I further see that the transfer orbit's period would be 13.12 hours, meaning after launch from Mars the vehicle (in this case for Deimos of course) the raise burn happens just over 6 hours later.

Is this correct so far?

You have the right peri and apo apsis for a transfer from Mars surface (the 0 km altitude) to Deimos (the 200062 km altitude). Your'e reading the period right. And you're correct that transfer is half that.

Not sure if you're interpreting the delta V burns correctly. To get off Mars look at Ellipse periapsis velocity (in this case it's  4.692 km/s). And then add to the the circularize burn at apoapsis. In this case it's .6717. I think you're getting the 680 meters per second from ellipse V at apoapsis (.6797 Km/s)? That's the speed the ship is moving at the height of it's orbit. To circularize it needs to be moving 1.3515 km/s. 1.35 - .68 is more like .67 km/s.

So circularizing at the top of transfer orbit is apoapsis circular minus apoapsis ellipse speed. And dV need to get off Mars is just ellipse speed at periapsis.

Finding relevant numbers is like looking for needles in a haystack so I greyed out some of the irrelevant stuff. The dV from Mars surface to Deimos would be about 4.7 + .67 km/s.