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#40
by
Jim
on 17 Mar, 2008 15:56
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Garrett - 17/3/2008 12:40 PM
imagine that the journey times will be optimised not only as a function of the launch vehicle but also of the launch window.
launch period is a function of LV performance. Launch windows occur daily. Launch periods are the days available to launch.
Read "A Case for Mars", it goes into the trades
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#41
by
Podkayne
on 18 Mar, 2008 01:35
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Garrett,
You have asked a complicated question. You can accept a quick off-the-cuff answer or check out some papers in the NASA Technical library, the papers published by the American Astronautical Society , or maybe some online notes from a college orbital mechanics course . Wikipedia is a good place to start, but take what you find on the Internet with a grain of salt.
I'm afraid it is, indeed, rocket science as applied to interplanetary orbital mechanics.
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#42
by
Podkayne
on 18 Mar, 2008 15:23
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Here's a great tutorial on
"The Basics of Spaceflight" from JPL. It includes chapters on trajectories and planetary orbits. Here's another tutorial on the famous
porkchop plots, the "first menu item on a trip to Mars". Yummy!
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#43
by
Garrett
on 19 Mar, 2008 12:44
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Thanks Podkayne for all those links, very useful, though I haven't had time to go through all of them thoroughly. Jim, I'll look into getting that book you've suggested, cheers.
Having read up a little on Hohmann transfer orbits (HTO's), porkchop plots (which are closely correlated to HTO's) and various other stuff, this is my understanding at the moment:
- A HTO must be taken if a trip to Mars is to be made using current technology and financial means.
- Small deviations to a HTO are possible and these are characterised in the porkchop plot.
- The window for an ideal HTO for Earth to Mars transfer takes place every 2.135 years or 780 days.
- For the 2005 Mars launch opportunity, NASA defined a 3 week window for HTO insertion.
- A HTO (or a variation thereof) to Mars takes about 6 months minimum, but can often be up to 8 or 9 months (I think).
What I do not understand yet is:
- When are the windows for a HTO transfer from Mars to Earth? Are they the same as the Earth to Mars windows, i.e. about every 2 years,
and do they occur at the same time or are they somewhat "out of phase"?
If they occur simultaneously then a roundtrip mission to Mars would last about 2 years and 10 months.
- I've heard of many other, shorter, estimations for mission durations so I imagine that my assumptions must be flawed somewhere?
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#44
by
Jim
on 19 Mar, 2008 12:52
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Garrett - 19/3/2008 9:44 AM
- When are the windows for a HTO transfer from Mars to Earth? Are they the same as the Earth to Mars windows, i.e. about every 2 years,
and do they occur at the same time or are they somewhat "out of phase"?
If they occur simultaneously then a roundtrip mission to Mars would last about 2 years and 10 months.
- I've heard of many other, shorter, estimations for mission durations so I imagine that my assumptions must be flawed somewhere?
Read Zubrin's book, he covers all this
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#45
by
Garrett
on 19 Mar, 2008 14:57
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Er, right Jim. Actually, as this is the Mars Q&A I was kinda hoping somebody would cover it here, not just for me but also for others who might come to a Mars Q&A looking for, well, answers. Or maybe I've misunderstood the purpose of this Q&A.
Regards.
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#46
by
Jim
on 19 Mar, 2008 15:23
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Garrett - 19/3/2008 11:57 AM
Er, right Jim. Actually, as this is the Mars Q&A I was kinda hoping somebody would cover it here, not just for me but also for others who might come to a Mars Q&A looking for, well, answers. Or maybe I've misunderstood the purpose of this Q&A.
Regards.
The reason I pointed to the book, is that it too complex with too many variables to summarize here.
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#47
by
Jim
on 19 Mar, 2008 16:13
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from the Mars Direct paper:
"Time is obviously required if the astronauts are to do any
useful exploration, construction, or resource utilization
experimentation on the surface of the destination planet.
This clearly means that opposition class Mars missions
(which involve 1.5 year flight times and 20 day surface
stays) are out of the question. It also means that
architectures involving Lunar or Mars orbital rendezvous
(LOR, MOR) are very undesirable, for the simple reason
that if the surface stay time is long, so is the orbit time."
http://en.wikipedia.org/wiki/Mars_Direct
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#48
by
Podkayne
on 19 Mar, 2008 17:30
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Garrett - 19/3/2008 8:44 AM
What I do not understand yet is:
- When are the windows for a HTO transfer from Mars to Earth? Are they the same as the Earth to Mars windows, i.e. about every 2 years,
and do they occur at the same time or are they somewhat "out of phase"?
If they occur simultaneously then a roundtrip mission to Mars would last about 2 years and 10 months.
- I've heard of many other, shorter, estimations for mission durations so I imagine that my assumptions must be flawed somewhere?
I'm not sure if this will answer your questions, but here is a discussion of short (planetary surface) stay missions versus long stay missions . Here's a large (10 Mb) Powerpoint file that discusses www.marsinstitute.info/epo/docs/jan05/human_studies.ppt ">human Mars mission exploration strategies , including mission mode.
Interplanetary mission planning isn't my specialty, but I recall that all the studies of human Mars missions that we did while I was involved in them had a total mission duration of about 1000 days (almost 3 years). So-called short stay missions (with a surface mission duration of about 30 - 45 days) were studied in the early 90's as part of the Space Exploration Initiative, but for reasons quoted above from Dr. Zubrin's book, longer surface stays are considered to be less risky for the crew and have been the focus of human Mars mission studies since the mid-1990's.
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#49
by
Seer
on 19 Mar, 2008 18:04
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About the mars aerocapture/entry heatshield - could one use a watercooled shield? I.e, spray water through holes in the shield. That would mean the spacecraft would get progressively lighter throughout. Another thing, one could use ISRU to replenish that water, in addition to the propellant for the trip back.
This scheme does assume that you can use the same shape shield for both aerocapture/entry at mars, and aerobraking back at earth. Does anyone know whether that is possible?
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#50
by
Kaputnik
on 19 Mar, 2008 20:55
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Seer - 19/3/2008 8:04 PM
About the mars aerocapture/entry heatshield - could one use a watercooled shield? I.e, spray water through holes in the shield. That would mean the spacecraft would get progressively lighter throughout. Another thing, one could use ISRU to replenish that water, in addition to the propellant for the trip back.
This scheme does assume that you can use the same shape shield for both aerocapture/entry at mars, and aerobraking back at earth. Does anyone know whether that is possible?
Flawed idea, on several grounds.
1- we already have materials that make perfectly good Mars entry heatshields, and can do the job at lower mass than your proposed active cooling system.
2- Current ISRU plans would not create water, they would only draw on CO2 as a resource because it is more easily extracted.
3- No, a Mars entry craft and an Earth entry craft are not interchangeable. In any case you wouldn't want to do this because it means dragging the spacecraft all the wya down to Mars and then launching it back up again- very, very wasteful in energy.
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#51
by
neviden
on 20 Mar, 2008 10:20
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#52
by
Kaputnik
on 20 Mar, 2008 10:41
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AFAIK nobody has yet built or used such a device successfully. It would undoubtedly weigh more than a rigid device; the same would go for an inflatable shield.
I think that the real problem to date is the low payload mass ratio of Mars entry vehicles. The MERs, with their airbags and a separate lander, had a very low 22.5% payload fraction- ouch! Whilst the 'lander' itself was quite heavy, the backshell/parachute system was also surprisingly heavy- much more so than the heatshield. I wonder how high we could get the payload fraction in the context of manned vehicles?
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#53
by
neviden
on 20 Mar, 2008 10:57
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22.5% is indeed very low. Maybe it would actually make sense to go with propulsive breaking in combination with aeroshell. It's only 4,1 km/s delta-v from Mars Low orbit. DC-X style.
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#54
by
Kaputnik
on 20 Mar, 2008 11:19
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That's no better. To achieve 4100m/s delta-v (assuming storable propellants giving 320s isp) you'd need to have about 75% of the vehicle as propellant.
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#55
by
neviden
on 20 Mar, 2008 11:27
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Kaputnik - 20/3/2008 2:19 PM
That's no better. To achieve 4100m/s delta-v (assuming storable propellants giving 320s isp) you'd need to have about 75% of the vehicle as propellant.
That is true, but that would also mean that if you refuel it with propellant you could return back to low orbit the same way. 4.1 km/s is quite doable for SSTO vehicle.
Propulsive breaking would be much more attractive if you would have large ISRU process going on the Moon/Asteroids/Phobos/Deimos. Oxygen is plentiful practically everywhere and itself represents a great deal of mass in the propellant. Even the fuel (H2? methane?) itself could probably be found and made in space.
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#56
by
meiza
on 20 Mar, 2008 12:53
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DC-X style works where there is enough atmosphere that your terminal velocity is low. It's no better in parachutes in that regard.
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#57
by
Kaputnik
on 21 Mar, 2008 17:00
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neviden - 20/3/2008 1:27 PM
Kaputnik - 20/3/2008 2:19 PM
That's no better. To achieve 4100m/s delta-v (assuming storable propellants giving 320s isp) you'd need to have about 75% of the vehicle as propellant.
That is true, but that would also mean that if you refuel it with propellant you could return back to low orbit the same way. 4.1 km/s is quite doable for SSTO vehicle.
Propulsive breaking would be much more attractive if you would have large ISRU process going on the Moon/Asteroids/Phobos/Deimos. Oxygen is plentiful practically everywhere and itself represents a great deal of mass in the propellant. Even the fuel (H2? methane?) itself could probably be found and made in space.
Jumping the gun a bit, no? There's clearly a real issue with getting a good payload fraction on entry vehicles, and there'll be plenty of one-way cargo going to Mars, they won't all be reusable SSTO ascent/descent vehicles.
Personally I think more research on hypersonic parachutes, or even ballutes if they can be made to work, could wides the bottleneck a bit, at what should be a relatively low mass penalty.
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#58
by
Seer
on 22 Mar, 2008 17:07
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Kaputnik - 19/3/2008 4:55 PM
Seer - 19/3/2008 8:04 PM
About the mars aerocapture/entry heatshield - could one use a watercooled shield? I.e, spray water through holes in the shield. That would mean the spacecraft would get progressively lighter throughout. Another thing, one could use ISRU to replenish that water, in addition to the propellant for the trip back.
This scheme does assume that you can use the same shape shield for both aerocapture/entry at mars, and aerobraking back at earth. Does anyone know whether that is possible?
Flawed idea, on several grounds.
1- we already have materials that make perfectly good Mars entry heatshields, and can do the job at lower mass than your proposed active cooling system.
2- Current ISRU plans would not create water, they would only draw on CO2 as a resource because it is more easily extracted.
3- No, a Mars entry craft and an Earth entry craft are not interchangeable. In any case you wouldn't want to do this because it means dragging the spacecraft all the wya down to Mars and then launching it back up again- very, very wasteful in energy.
Regarding your last point: in my scenario, one would only have to build one type of spacecraft, which would be refueled in earth orbit, instead of the normal 4 different habs/erv/mars ascent vehicles in other plans.
As for deriving water on mars - that would either come from underground ice or from manufacturing water from co2 and h2 feedstock brought from earth.
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#59
by
jcopella
on 06 Apr, 2008 14:13
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jcopella - 16/3/2008 1:35 PM
Podkayne - 16/3/2008 1:03 PM
jcopella - 16/3/2008 10:51 AM
What is the current thinking (plans, designs, revised risk assessments) regarding radiation shielding? I had a casual association with someone in the Navy (somehow he got involved thru knowing Bill Readdy, IIRC) who worked on a Mars study group in the mid-90s. At the time he told me the radiation problem was a (the) show-stopper.
For a scholarly discussion of the risks of radiation exposure and other hazards to humans on Mars, see " Safe on Mars: Precursor Measurements Necessary to Support Human Operations on the Martian Surface (2002) " from the National Research Council.
LOL
I've actually read that paper. It doesn't answer the question I posed (which I should clarify), but thanks -- it does deal with radiation impacts but the focus is on risks to Mars surface operations.
To clarify:
My question related to the shielding-related design impacts for a Mars Transfer Vehicle. The last I heard from someone actively working on a precursor project to VSE/Cx, the shielding requirements were a show-stopper.
I'm just wondering if there's been any evolution in that position since VSE/Cx, and if so, what the current thinking is.
Looks like it's still a show-stopper:
http://forum.nasaspaceflight.com/forums/thread-view.asp?tid=12532&posts=39&start=1