Author Topic: "Mission to Mars Using Six 'Not So Easy' Pieces" • Mike Raftery , Boeing  (Read 30723 times)

Offline metaphor

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I don't really understand why the cargo is launched directly to EML-2 to meet an SEP transfer stage there. Launching to EML-2 takes only a few hundred m/s less than launching directly to Mars for aerocapture or direct entry. A single SLS launch should be sufficient to place a habitat for four crew on the surface of Mars.

That seems like a response to the proposal that voices meaningful concerns, rather than flying off into discussions of how SpaceX will get there first. So it seems like it deserves to be considered! Could you start by putting some crispness into the numbers? I use 3,188 m/s for LEO to Earth escape, and 4,400 m/s for TMI from LEO. The difference is quite a bit more than "a few hundred m/s"!

The lowest transfer windows to Mars are about 3.54 km/s from LEO.  The more difficult launch windows might require up to 3.8 km/s.  Lowering transit time increases that delta-v, but that shouldn't be a concern for cargo launches.  For crew launches, if you want to lower the transit time to less than 7 months, 4.00 km/s is enough for any launch window, while a less than 6-month transit time takes less than 4.15 km/s.  Trajectory Browser

Offline sdsds

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The lowest transfer windows to Mars are about 3.54 km/s from LEO.  The more difficult launch windows might require up to 3.8 km/s.[...] Trajectory Browser

Those clearly exist. I suppose the question is, "Can you count on launching into them?" Attached is a chart from "Cryogenic Propulsive Stage (CPS) Mission Sensitivity Studies - Low Earth Orbit Departure Results - Revision D" dated 4 October 2012 by Mark Schaffer of SpaceWorks Engineering. To get a 30 day launch window on this opportunity required 4400 m/s.
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Offline metaphor

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The lowest transfer windows to Mars are about 3.54 km/s from LEO.  The more difficult launch windows might require up to 3.8 km/s.[...] Trajectory Browser

Those clearly exist. I suppose the question is, "Can you count on launching into them?" Attached is a chart from "Cryogenic Propulsive Stage (CPS) Mission Sensitivity Studies - Low Earth Orbit Departure Results - Revision D" dated 4 October 2012 by Mark Schaffer of SpaceWorks Engineering. To get a 30 day launch window on this opportunity required 4400 m/s.

I'm not sure why the lowest point on that blue line is still higher than 4 km/s.  The 2022 launch window is actually one of the hardest but it should still be less than that.

I was assuming cargo launches land directly on Mars and don't go into orbit first.  In that case, the Mars arrival delta-v does not matter so there's no reason to optimize for it.

Offline ThereIWas3

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After Apollo 11, each mission was to a different location.  They were exploring, not building.  So none of the equipment was designed to be reused.  And they launched  one or two missions per year.  Maximum stay was 3 days, so they did not need a lot of equipment.

Mars being a lot further away, stays will be longer and more equipment will be required; more than a single launch of any of the proposed vehicles could manage.   So an integration location needs to be chosen.

LEO has advocates, as does EML.  I would like to see both Mars orbit and Mars surface considered as well for rendezvous/integration locations.   It might allow for first launches a few cycles earlier than building up everything at the Earth end first.
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Offline Robotbeat

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Many rendezvous points probably makes sense. EML2 (or EML1 or high Earth orbit... Practically all the same from an energy perspective), in addition to Mars orbit and Mars surface. By having the SEP stage and transit hab be reusable, you have the two biggest pieces (which are also most flexible... SEP could be used as a tug for, say, a depot for lunar propellant and both pieces could be used for a mission to, say, Ceres) which could be used again and again even after we go beyond SLS and Orion. Also, it'd be nice to have a better performance, single-stage reusable Mars lander that could be refueled in orbit or on the surface.
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Offline Rocket Science

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It would be interesting if Habs and Rovers could be repositioned by being refueled and making short “hops” to a new location for the next crew...
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Offline sdsds

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Comparing the two trajectories below seems useful. Both are for the same (2035) opportunity. Raftery's uses SEP; the trajbrowser assumes impulsive (i.e. chemical) TMI. 256 days in transit, versus 112.
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Offline Robotbeat

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The traj isn't optimizing for Mars capture delta-v, not an appropriate comparison. That delta-v isn't free. Appropriate for a flyby, not for an actual Mars mission with Mars-orbit rendezvous (chemical OR SEP-chemical hybrid).
« Last Edit: 07/08/2014 08:21 PM by Robotbeat »
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Offline sdsds

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Right, or if one had sufficient faith in one's direct Mars entry/descent/landing technology, perhaps because it had been tested on prior cargo missions. That would lead to a Mars surface rendezvous approach....
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Offline Robotbeat

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Right, or if one had sufficient faith in one's direct Mars entry/descent/landing technology, perhaps because it had been tested on prior cargo missions. That would lead to a Mars surface rendezvous approach....
That'd be fine except now your lander must be much bigger because it has to be big enough to support the whole transit. The lander/ascent-vehicle is just about the hardest part so you'd want to minimize it.

All in all, the approach by Raftery is more conducive to sustainable, flexible operations.
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Offline sdsds

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All in all, the approach by Raftery is more conducive to sustainable, flexible operations.

This being the Raftery thread, I'll have to agree with that. "Have SEP, will travel!"
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Offline Halidon

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Maybe its artistic license, but the SLS EUS appears to be using different engines for the caro and crew launches. Certainly looks like RL-10s for the crew version, RL-60/RL-X for cargo?

Offline Archibald

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The pictures on that pdf, they were so beautiful, I made a 5 minutes powerpoint animation with them. I added Coldplay "a sky full of stars" as soundtrack. I should post that on YouTube  :o
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Offline Robotbeat

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The pictures on that pdf, they were so beautiful, I made a 5 minutes powerpoint animation with them. I added Coldplay "a sky full of stars" as soundtrack. I should post that on YouTube  :o
That'd be cool.
Chris  Whoever loves correction loves knowledge, but he who hates reproof is stupid.

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Offline metaphor

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This is an ingenious proposal.

But one thing I feel might be a problem is the high orbit rendezvous.  That means that the MAV has to use much more fuel than a low orbit rendezvous, and also an additional engine start.  Compared to the ~3.8 km/s of delta-v to low Mars orbit from the surface, going to areostationary orbit requires an additional two burns of 1.1 km/s and 0.7 km/s.  If the spacecraft has a dry mass of 10 tons, a MAV using LOX/methane would need 20 tons of fuel to get to low orbit, but 40 tons of fuel to get to areostationary orbit.  That might be a problem, especially if not using ISRU and that fuel has to be carried from Earth.

Offline ThereIWas3

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This is an ingenious proposal.

But one thing I feel might be a problem is the high orbit rendezvous.  That means that the MAV has to use much more fuel than a low orbit rendezvous, and also an additional engine start.  Compared to the ~3.8 km/s of delta-v to low Mars orbit from the surface, going to areostationary orbit requires an additional two burns of 1.1 km/s and 0.7 km/s.  If the spacecraft has a dry mass of 10 tons, a MAV using LOX/methane would need 20 tons of fuel to get to low orbit, but 40 tons of fuel to get to areostationary orbit.  That might be a problem, especially if not using ISRU and that fuel has to be carried from Earth.

A cynic might say that Boeing would prefer the higher mass solution, as it reinforces the need for SLS.
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Offline Robotbeat

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Cynical: Doesn't make that much sense, really. With refueling, the 30-40mT lander is likely to weigh ~10mT dry.

A higher rendezvous orbit means a smaller SEP tug, and the SEP tug has more need of the volume than the lander.
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Offline yg1968

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All of this stuff is very easily transferable to a more commercially supplied architecture.

In Boeing's proposal, Orion could easily be replaced by an upgraded Dragon since the spacecraft doesn't go further than the L2 Gateway.
« Last Edit: 07/10/2014 08:01 PM by yg1968 »

Offline Robotbeat

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All of this stuff is very easily transferable to a more commercially supplied architecture.

In Boeing's proposal, Orion could easily replaced by an upgraded Dragon since the spacecraft doesn't go further than the L2 Gateway.
Indeed. Raftery also explicitly mentions commercial propellant delivery for the SEP, which could logically be extended to the lander, too (especially if you could manage a single-stage reusable lander).
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Offline Oli

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Awesome. SEP tugs are the real deal.

I have a feeling this might actually happen.
« Last Edit: 07/10/2014 02:58 AM by Oli »

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