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Does anyone here know the maximum amount of mass that one could bring to Mars and still use aerobraking?
Does anyone here know the maximum amount of mass that one could bring to Mars and still use propellant to achieve Mars orbit?
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scienceguy - 8/3/2008 1:07 PM
1. Does anyone here know the maximum amount of mass that one could bring to Mars and still use aerobraking?
2. Does anyone here know the maximum amount of mass that one could bring to Mars and still use propellant to achieve Mars orbit?
It could be the size of one of the moons. Just have a large enough aeroshell or big propellant tank with big engines.
The size is limited by the mass you can send to Mars, not how you brake into orbit
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Thanks for the quick response.
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Why are their limits on the mass/area of the Mars entry vehicle?
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So, during the trip to Mars, how are they going to solve a Problem with the Astronauts being so lonley from away from family?
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The mass is limited by the area of the heatshield. The area of the heatshield is limited by the diameter of the payload fairing, which is itself related to the choice of LV.
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So even if a massive ship with AG was constructed in LEO, it would still be able to aerobrake at Mars if it had a wide enough heat shield?
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Well... yeah, of course it would. There's no inherent limit on these things. It comes down to the density of the entry vehicle and the practical limitations of what can be assembled.
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scienceguy - 9/3/2008 4:37 PM
So even if a massive ship with AG was constructed in LEO, it would still be able to aerobrake at Mars if it had a wide enough heat shield?
Yes
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Orbiter - 9/3/2008 11:27 AM
So, during the trip to Mars, how are they going to solve a Problem with the Astronauts being so lonley from away from family?
I know of two ways. First, the plans I've glanced at have multiple astronauts (4 seems the favored number). Second, the astronauts would remain in contact with mission control and their families.
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The second method may not work owing to speed-of-light communication delays (up to 20 minutes+)
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tnphysics - 9/3/2008 8:11 PM
The second method may not work owing to speed-of-light communication delays (up to 20 minutes+)
It will work, it is called email
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I ment telephone.
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khallow - 10/3/2008 9:21 AM
Orbiter - 9/3/2008 11:27 AM
So, during the trip to Mars, how are they going to solve a Problem with the Astronauts being so lonley from away from family?
I know of two ways. First, the plans I've glanced at have multiple astronauts (4 seems the favored number). Second, the astronauts would remain in contact with mission control and their families.
By 2035, the computers will be able to hold a decent conversation with them. Until it goes mad, kills the crew and you have to pull the modules for its higher functions... additionally, they could have all sorts of robots as well.
Plants also have a calming effect. Those will almost certainly be carried, albeit probably in pot plant form.
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Lampyridae - 10/3/2008 1:31 AM
Plants also have a calming effect. Those will almost certainly be carried, albeit probably in pot plant form.
Time to develop plants that are:
a) edible
b) like weak sun light
c) and thrive in high levels of CO2.
We can heat the greenhouse artificially and recycle the water.
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A_M_Swallow - 9/3/2008 9:54 PM
Lampyridae - 10/3/2008 1:31 AM
Plants also have a calming effect. Those will almost certainly be carried, albeit probably in pot plant form.
Time to develop plants that are:
a) edible
b) like weak sun light
c) and thrive in high levels of CO2.
We can heat the greenhouse artificially and recycle the water.
Spinach
a) is edible
b) does fine in weak sunlight
c) thrives in high levels of CO2 (actually most plants if not all do better in high CO2)
Also, spinach survives well under salty conditions (Zapata et. al. 2004), and water found on Mars would probably be salty. As a final kicker, spinach absorbs a lot of iron from soil and we know how important that would be in the rusty dunes of Mars.
Reference
Pedro J. Zapata, María Serrano, M. Teresa Pretel, Asunción Amorós and M. Ángeles Botella (2004) Polyamines and ethylene changes during germination of different plant species under salinity. Plant Science, Volume 167, Issue 4, Pages 781-788
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Perhaps with the correct sort of "pot" plants, they wouldn't care...
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Will it be possible to return to Mars orbit on the same spacecraft that was used for entry.
I have read some proposed architectures, many propose the ascent and descent vehicles are seperate spacecraft.
Does anyone know why?
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Lampyridae - 10/3/2008 1:31 AM
khallow - 10/3/2008 9:21 AM
Orbiter - 9/3/2008 11:27 AM
So, during the trip to Mars, how are they going to solve a Problem with the Astronauts being so lonley from away from family?
I know of two ways. First, the plans I've glanced at have multiple astronauts (4 seems the favored number). Second, the astronauts would remain in contact with mission control and their families.
By 2035, the computers will be able to hold a decent conversation with them. Until it goes mad, kills the crew and you have to pull the modules for its higher functions... additionally, they could have all sorts of robots as well.
Plants also have a calming effect. Those will almost certainly be carried, albeit probably in pot plant form.
I think the 'out of contact' 'problem' is over-stated. Given how long explorers used to have to go without contact, a 20-minute delay is absolutely nothing. When Franklin disappeared it was two whole years before anybody could be sure that something had gone wrong. Anyway, there was discussion of this on the 'would you go to Mars now' thread/poll in the Ares-V section.
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MB123 - 10/3/2008 9:20 AM
Will it be possible to return to Mars orbit on the same spacecraft that was used for entry.
I have read some proposed architectures, many propose the ascent and descent vehicles are seperate spacecraft.
Does anyone know why?
There are also architectures proposed which use a dual purpose ascent-descent vehicle.
I can think of two reasons for having separate craft, though:
1- landed mass bottleneck forces you to strip down a craft ot the bare minimum functionality
2- you want the crew to land in their habitat module, a la Zubrin, for whatever reason- ascent vehicle needs to have refuelled itself, better chance for crew survivial if they go off course, etc.
Note that in the first option, you wouldn't save all that much mass by having two dedicated and separate ascent and descent vehicles. The ascent vehicle has to descend to Mars safely at some point anyway, so you might as well stick a crew on it- if that's compatible with the rest of the mission plan.
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These proposed architectures, are they available online? Will I be able to find many of them on the NASA techhnical reports server?
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Look for Mars Direct, Mars For Less, NASA's 'Design Referece Mission version 1 to 3, amongst others.
Another example was published recently and details are discussed on the Ares-V part of the forum.
Yet another is JPL's NEP-AG study, although that didn't look at operations outside of Mars transit itself.
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scienceguy - 8/3/2008 10:07 AM
Does anyone here know the maximum amount of mass that one could bring to Mars and still use aerobraking?
Does anyone here know the maximum amount of mass that one could bring to Mars and still use propellant to achieve Mars orbit?
The basic question has already been answered, but I thought I'd point out that aerobraking and achieving orbit are two separate actions. Aerobraking is used to circularize an orbit, such as the Mars Reconnaissance Orbiter did after its orbital insertion burn. This involves repeated dips through the upper atmosphere to lower the apogee of the highly elliptical insertion orbit. The final step is another engine burn to raise the perigee so the orbit is nearly circular.
Rather than a burn, aero-capture is another potential option for entering orbit. This is involves passing even lower into the atmosphere to bleed off a large amount of velocity in a single pass, entering an elliptical orbit that way. If desired, aerobraking or engine firing can then be used to circularize the orbit.
Aero-capture has never been done before. As I understand, it requires greater navigational precision, and a heat shield approximately as capable as one designed for direct entry.
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iamlucky13 - 13/3/2008 12:35 AM
Aero-capture has never been done before. As I understand, it requires greater navigational precision, and a heat shield approximately as capable as one designed for direct entry.
To be viable the aero-capture heat shied only has to save its own weight in fuel, larger savings are even better. Is the Mars breaking delta-v of approx 3.3 km/s sufficiently low that a reusable heat sheild made out of say titanium possible?
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Is there an elliptical/ballistic Earth orbit that can be used to test aero-capture?
To develop the techniques and heat shields we only need a small payload.
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Is the MOI delta-v really as much as 3.3km/s? I thought it was more like a third of that. Oh well.
I believe that ESA are talking about an aerocapture demonstrator. I don't know the flight profile but you'd think an elliptical orbit or a lunar flyby would do the job.
Until we have a reliable network of guidance beacons at Mars (similar to GPS) I doubt that a manned spacecraft or even an unmanned one if it were mission critical would be allowed to use aerocapture, but that's just IMHO.
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Depends how many passes you can do. You're going to need some propulsion maneuvers anyway too. So it's not strictly an either-or-problem.
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Kaputnik - 13/3/2008 1:55 PM
Is the MOI delta-v really as much as 3.3km/s? I thought it was more like a third of that. Oh well.
No. Rechecking that figure includes visiting both of Mars's moons.
Earth C3 to Mars transfer is 0.6 km/s
Mars transfer to Mars C3 is 0.9 km/s
Mars C3 to low Mars orbit = 0.2 + 0.3 + 0.9 = 1.4 km/s
So Mars transfer to low Mars orbit = 0.9 + 1.4 = 2.3 km/s
http://en.wikipedia.org/wiki/Delta-v_budget
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A_M_Swallow - 12/3/2008 5:01 PM
iamlucky13 - 13/3/2008 12:35 AM
Aero-capture has never been done before. As I understand, it requires greater navigational precision, and a heat shield approximately as capable as one designed for direct entry.
To be viable the aero-capture heat shied only has to save its own weight in fuel, larger savings are even better. Is the Mars breaking delta-v of approx 3.3 km/s sufficiently low that a reusable heat sheild made out of say titanium possible?
A metal probably isn't a great material choice, and I'm not aware of any reason a ceramic or similar shield couldn't be used for both capture and entry. Still, some of the things I've read have suggested a two stage heat shield to reduce entry mass (less fuel to de-orbit, better mass/surface area ratio) and to reduce risk of heat soak from the capture from damaging anything. Ceramic heat shields have very low thermal conductivity, so over the relatively short time of the capture they may heat up quite a bit, but you'd have a fair amount of time to jettison it before the heat soaked into the rest of the craft.
The heating from the capture manuever is actually greater than that for the following entry from low orbit.
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tnphysics - 10/3/2008 10:11 AMThe second method may not work owing to speed-of-light communication delays (up to 20 minutes+)
It's a 40-minute delay (20 minutes each way) when Mars & Earth are furthest apart
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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.
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For published reports on human missions to Mars, search David Portree's excellent compendium "Romance to Reality"
You could also try The Mars Society and NASA's technical reports library .
Many of the answers to questions on this thread can be found in scholarly, researched, and peer-reviewed papers. There is no need for speculative, incomplete, or shoot-from-the-hip answers - unless that's what you are looking for.
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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.
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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.
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As you probably know, it takes a while for the results of NASA studies to make their way into publication. You might check the proceedings from the AIAA Space 2007 Conference or the 2008 Space Technologies and Applications International Forum for the most recent results that have made it past the technical and management reviews.
I haven't seen much on Constellation studies of human Mars missions since the House threatened to forbid NASA to fund any studies related to the subject. (This is not the first time Congress has forbid NASA to even think about human missions to Mars.)
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How well would a book like Red Mars hold up to reality? Do the ideas in it have any basis in the scientific community or is it misleading?
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Here's and extensive bibliography on the subject of terraforming Mars. The Astrobiology web has a page on the subject also.
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Where can we get reliable (scholarly) information on launch windows, journey times and total mission durations for human Mars missions? From Wikipedia it appears that the minimum energy launch window is about every 2.135 years (780 days)
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Garrett - 17/3/2008 12:06 PM
Where can we get reliable (scholarly) information on launch windows, journey times and total mission durations for human Mars missions? From Wikipedia it appears that the minimum energy launch window is about every 2.135 years (780 days)
They aren't fixed numbers. They are interrelated variables
Launch vehicle performance drives many of this numbers (launch windows, journey times, etc )
mission duration is a function of the mission architecture, which will or be affected Launch vehicle performance
the only thing fixed is the muinimum energy launch window of 26 months
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Cheers for the quick reply. I actually hadn't finished my post - I pressed the "Submit" button by accident!
I understand that journey times and mission durations are not very closely correlated to launch windows, but I imagine that actual launches will be preferably close to the minimum energy launch window? An author on Wikipedia also suggested another, longer trajectory going into Venus.
I would still have believed that the mission duration will not be completely independent of these launch windows. Likewise, I imagine that the journey times will be optimised not only as a function of the launch vehicle but also of the launch window. What I'm wondering is whether the variables are as interrelated as you say or whether the situation is actually closer to a few "eigenvalues" so to speak?
I suppose the best answer is that I should probably start reading a "rocket science" book and work out the equations myself! :-p
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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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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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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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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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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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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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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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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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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 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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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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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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To be more on topic.. what is wrong with deployable heat shield that would look something like this:
http://www.russianspaceweb.com/mars_lander_200.jpg
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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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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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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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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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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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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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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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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
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Seer - 22/3/2008 7:07 PM
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.
There are three or four different types of hab, erv, and ascent vehicle etc for a very good reason- eahc one is optimised for a different purpose. A vehicle suitable for Mars ascent hasn't got enough living space to work as a hab or erv, for example.
Martian water can, of course, be obtained, but it's not much use for your active cooling system until you get it into orbit and pumped aboard an entry vehicle. That means that the first lot of spacecraft would have to be either stupidly heavy, with water brought from Earth, or more sensibly be conventional in design.
Honestly the entire concept of active water cooling is a solution looking for a problem. A normal heatshield can do the job, and at lower mass, complexity, and probably cost too.
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I know NASA's plan for the moon is to set up an outpost there. Does the VSE include an outpost on Mars as well? Have they planned that far in advance?
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Jason - 17/3/2008 10:02 AM
How well would a book like Red Mars hold up to reality? Do the ideas in it have any basis in the scientific community or is it misleading?
Other than being wildly optimistic, there's no bad science. In one scene in the book, the first man on Mars recalled a solar storm that badly irradiated the crew on the first expedition. Leukaemia kills a few of the colonists later on in the books. Dust was a major pain in the butt for the colonists; it got into everything.
The massive colony ship was built out of spent ETs, which is a dubious proposition given their fragility and work required to get them habitable. Additionally, inflatables gave much better volume to mass ratios. The colonists also seemed to have a huge amount of elbow room for only an 8 or 9 month voyage. Finally, the colony ship, which must have massed at least 1000mT, was pushed out of orbit and on to Mars with LOX/LH2 engines... and also braked with them.
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scienceguy - 23/4/2008 10:50 AM
I know NASA's plan for the moon is to set up an outpost there. Does the VSE include an outpost on Mars as well? Have they planned that far in advance?
It's not NASA's plan, it's the President's... but NASA have made it their own. NASA has only a vague plan for Mars, which are as about as firm as their plans for manned missions to Jupiter and Venus, which is to say not at all. Congress has forbidden NASA (again) from spending any on a mission to Mars this year.
NASA is so dominated by the desire to build Ares V "to go to Mars" it's making a mess of everything. 6 Ares V launches are needed to send 6 people to Mars IIRC under the current plan. At the moment, there is not even enough money for the required number of lunar missions. With current funding (with Ares V), there will probably be a manned lunar mission by 2025, that is the reality which is implied but not stated by just about everybody including Griffin.
IMHO, if NASA goes with Direct and an aggressive program on the moon, a lunar COTS program and decent propulsion technologies, then a manned Mars mission is possible by 2035 or so.
Ironically, though, Deimos is closer to Earth than the Moon in terms of delta-Vee. I think it would actually be a good first exploration target, maybe even offering some kind of ISRU possibilities.
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Does anyone know if NASA TV will carry live coverage of the the Mars Phoenix lander Entry, Descent and Landing? The event is not listed on the NASA TV schedule. I'm looking at this page: http://www.nasa.gov/multimedia/nasatv/MM_NTV_Breaking.html. Should I be looking somewhere else?
Thanks in advance. I'm planning a watching party, as long as I know that the event will be carried live. I watched both the MER-A and MER-B landing feeds live from JPL, and it made for engrossing television.
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It will be shown live
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Can anyone describe the mission trajectory for a direct ascent option to the apollo missions?s
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Can anyone describe the mission trajectory for a direct ascent option to the apollo missions?s
No parking orbit. The launch vehicle performs all the stage burns in succession with no second burn for the 3rd stage.
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Also to my surprise a methane fuel option though I'd hate to see how many an ERV would need to get off Mars then again it can't be any worse then Falcon 9.
Not many people still cling to a Zubrin-esque ERV capable of supporting a crew from Mars ascent right through to landing on Earth.
A far more realistic concept is to have rendezvous in Mars orbit with the ERV. This reduces the launch mass of the Mars Ascent Vehicle to something like 15-25,000kg. Given the much lower gravity on Mars, one or two RL10s would be perfect for this application.
Recently it seems NTR and NEP missions such as BNTR have gained the upper hand over ISRU centered Mars missions.
Not sure I follow you here. ISRU is not and cannot be a factor until you get there. NTR is most likely to be used for TMI so is compatible with any type of mission.
The Mars ascent vehicle may still use ISRU and it's launch mass might be closer to 50T vs 25T of course the weight would be 1/3 on Mars.
How do you figure that? Do some maths. All you need is a cabin suitable for the four to six people most commonly assumed to make up the crew, and then a rocket stage to go under it. You can use scaled up numbers from the LEM ascent stage, or from the Soyuz BO, or whatever, to come up with the 'cabin' mass, and the main propulsion system ends up being a bit like an EELV upper stage. 50t is way too much.
I figure the most bare bones ascent vehicle would be an Orion CM and a methane ascent stage.
No it wouldn't. The Orion is pretty substantial, far heavier than necessary. The bare bones would be some sort of pressure-stabilised Al-Li sphere housing the crew. Or even an unpressurised Langley type design.
Mars's gravity well though less then Earth's is still much deeper then the Moon's.
I figure the Mars ascent vehicle also would need enough delta V to play an active role in seeking the MTV and docking with it in just a few days.
The MAV needs something like 4km/s delta-v to reach orbit (IIRC), vs. about 2.6km/s for the moon. IMHO the simplest design would have a single RL10-derived engine in the CH4/LOX main ascent stage, and then the crew capsule itself would have a small hypergolic RCS. The ascent stage would be jettisoned before attmepting to dock with the orbiting vehicle.
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Also to my surprise a methane fuel option though I'd hate to see how many an ERV would need to get off Mars then again it can't be any worse then Falcon 9.
Not many people still cling to a Zubrin-esque ERV capable of supporting a crew from Mars ascent right through to landing on Earth.
A far more realistic concept is to have rendezvous in Mars orbit with the ERV. This reduces the launch mass of the Mars Ascent Vehicle to something like 15-25,000kg. Given the much lower gravity on Mars, one or two RL10s would be perfect for this application.
Recently it seems NTR and NEP missions such as BNTR have gained the upper hand over ISRU centered Mars missions.
Not sure I follow you here. ISRU is not and cannot be a factor until you get there. NTR is most likely to be used for TMI so is compatible with any type of mission.
The Mars ascent vehicle may still use ISRU and it's launch mass might be closer to 50T vs 25T of course the weight would be 1/3 on Mars.
How do you figure that? Do some maths. All you need is a cabin suitable for the four to six people most commonly assumed to make up the crew, and then a rocket stage to go under it. You can use scaled up numbers from the LEM ascent stage, or from the Soyuz BO, or whatever, to come up with the 'cabin' mass, and the main propulsion system ends up being a bit like an EELV upper stage. 50t is way too much.
I figure the most bare bones ascent vehicle would be an Orion CM and a methane ascent stage.
No it wouldn't. The Orion is pretty substantial, far heavier than necessary. The bare bones would be some sort of pressure-stabilised Al-Li sphere housing the crew. Or even an unpressurised Langley type design.
Mars's gravity well though less then Earth's is still much deeper then the Moon's.
I figure the Mars ascent vehicle also would need enough delta V to play an active role in seeking the MTV and docking with it in just a few days.
The MAV needs something like 4km/s delta-v to reach orbit (IIRC), vs. about 2.6km/s for the moon. IMHO the simplest design would have a single RL10-derived engine in the CH4/LOX main ascent stage, and then the crew capsule itself would have a small hypergolic RCS. The ascent stage would be jettisoned before attmepting to dock with the orbiting vehicle.
I was thinking the Orion CM because this is NASA we're talking about and they tend to prefer to error on the side of caution.
Plus NASA would want a few days of loiter time just in case the MAV is forced to launch with the MTV out of position.
Though there might be one advantage to using an Orion CM for the ascent vehicle the crew can leave the MTV for home once it's captured.
This also would allow the MTV to take a more leisurely trip back to LEO after capturing at one of the Earth/Moon liberation points.
It also could be used if something goes wrong and the MTV can't perform the Earth capture burn though the reentry speeds might be pretty close to Orion's limits but it's better then getting stuck up there.
You'd get the crew back and your biggest worry would be congress complaining they lost an expensive asset and the environmentalists complaining about a big nuclear space craft being in an Earth crossing orbit though the flyby can be designed in such a way it'll never return to earth for centuries.
Which leaves one question what to do with EOL MTVs that have flown their last trip to Mars?
Refuel the drive section one last time after stripping the vehicles and send them on a one way trip as a booster for a large probe to the outer solar system?
Jupiter seems like a good place to send them I think it's less delta V then reaching the sun.
As for an open ascent vehicle Mars does still have an atmosphere wouldn't it still be trouble?
Plus you'd be cutting your time from leaving the surface hab to rendezvous to under 9 hours the max limit a spacesuit can scrub CO2 and provide O2.
Though I guess the MTV could be moved into low Mars orbit ahead of time and the MAV launch in such a way it can rendezvouses quickly.
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I was thinking the Orion CM because this is NASA we're talking about and they tend to prefer to error on the side of caution.
Plus NASA would want a few days of loiter time just in case the MAV is forced to launch with the MTV out of position.
The CM would be heavier than necessary. End of story. It offers no particular advantage over, say, a lightweight cylinder with Orion systems installed. Why use a conical shape with a useless heatshield?
Though there might be one advantage to using an Orion CM for the ascent vehicle the crew can leave the MTV for home once it's captured.
This also would allow the MTV to take a more leisurely trip back to LEO after capturing at one of the Earth/Moon liberation points.
It also could be used if something goes wrong and the MTV can't perform the Earth capture burn though the reentry speeds might be pretty close to Orion's limits but it's better then getting stuck up there.
You'd get the crew back and your biggest worry would be congress complaining they lost an expensive asset and the environmentalists complaining about a big nuclear space craft being in an Earth crossing orbit though the flyby can be designed in such a way it'll never return to earth for centuries.
FYI it's 'libration' points although they could have a 'liberating' effect on exploration (sorry, bad pun).
There's no reason not to take an Orion to Mars orbit and back, just dragging it all the way up from the surface is a waste of mass/energy.
With a thicker heathshield and a skipping entry, I believe that Orion could easily handle a direct return from Mars.
As for an open ascent vehicle Mars does still have an atmosphere wouldn't it still be trouble?
Plus you'd be cutting your time from leaving the surface hab to rendezvous to under 9 hours the max limit a spacesuit can scrub CO2 and provide O2.
Well I was half-joking. Perhaps the max-q would indeed be too much. Btw, I'm not sure where you get the nine hours from, Apollo suits were supposed to last up to 110hrs in a dire emergency. Consumables can be supplied externally to extend the mission time of a suit.[/quote]
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Question:
Do the orbital mechanics allow you to choose to return to Earth early after you get to Mars? i.e. you intend a conjunction class mission and fly a normal outward leg, but carry enough propellant to make an opposition return leg with a Venus flyby?
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Theres the problem of entering and landing a heavy load thru the thin Mars atmosphere, but given the what you can throw at the problem - its workable - the first thing that pops into my mind are a serries of large, potentially bulky parachutes optimized for different velocities and altitudes to slow it down enough for a powered landing. Maybe it would make a Hab look like a 5 story tuna-can instead of a 3 story one. The point is with enough throw capability you could make the last parachute in the serries as large as it has to be in order to slow the vehicle down to the point where rockets can be fired for the landing.
Parachutes aren't the answer. There's a really good paper on Mars EDLS that you can probably dig up off google.
Essentially, different methods of decelleration are useful at certain airspeed ranges. So, at the very top end, you need a robust heat-proof drag device (a heatshield) to withstand the very harshest point of entry. If that shield can provide sufficient drag in relation to the vehicle's mass, you have a chance of slowing sufficiently before hitting the surface, and you can deploy another drag device.
Historically, a supersonic parachute has been employed as the second drag device, and, again, this needs to impart sufficient decelleration to the vehicle in a short time so that the next system can be used- this is propulsive descent, which should only be used once you have reached subsonic speeds.
Once you are travelling subsonic you can fire retro-rockets which whatever thrust/impulse is necessary to do the job.
So, each drag device needs to slow the vehicle into the region where another type of system can take over, before it hits the surface.
Ballutes have been suggested as a bridge between parachutes and heatshields; inflatable heatshields have also been suggested. Lifting entry can help a bit too. It is likely that some combination of such technologies will be needed to enable manned missions.
Using current technology limits, the entry vehicle cannot have a 'density' (mass/heatshield area) of more than about 150kg/m2, of else it hits the surface before it has slowed enough to deploy a parachute. So for something using a 10m PLF this would mean an entry mass of less than 12t.
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Theres the problem of entering and landing a heavy load thru the thin Mars atmosphere, but given the what you can throw at the problem - its workable - the first thing that pops into my mind are a serries of large, potentially bulky parachutes optimized for different velocities and altitudes to slow it down enough for a powered landing. Maybe it would make a Hab look like a 5 story tuna-can instead of a 3 story one. The point is with enough throw capability you could make the last parachute in the serries as large as it has to be in order to slow the vehicle down to the point where rockets can be fired for the landing.
Parachutes aren't the answer. There's a really good paper on Mars EDLS that you can probably dig up off google.
Essentially, different methods of decelleration are useful at certain airspeed ranges. So, at the very top end, you need a robust heat-proof drag device (a heatshield) to withstand the very harshest point of entry. If that shield can provide sufficient drag in relation to the vehicle's mass, you have a chance of slowing sufficiently before hitting the surface, and you can deploy another drag device.
Historically, a supersonic parachute has been employed as the second drag device, and, again, this needs to impart sufficient decelleration to the vehicle in a short time so that the next system can be used- this is propulsive descent, which should only be used once you have reached subsonic speeds.
Once you are travelling subsonic you can fire retro-rockets which whatever thrust/impulse is necessary to do the job.
So, each drag device needs to slow the vehicle into the region where another type of system can take over, before it hits the surface.
Ballutes have been suggested as a bridge between parachutes and heatshields; inflatable heatshields have also been suggested. Lifting entry can help a bit too. It is likely that some combination of such technologies will be needed to enable manned missions.
Using current technology limits, the entry vehicle cannot have a 'density' (mass/heatshield area) of more than about 150kg/m2, of else it hits the surface before it has slowed enough to deploy a parachute. So for something using a 10m PLF this would mean an entry mass of less than 12t.
Well it would appear current technology is just not good enough for a manned mission then.
I kinda skipped to the end with the above and assumed an aeroshield for orbital capture and entry - maybe an inflatable ring can extend the surface area of the initial aeroshield. The serries of parachutes afterthat would of course have to be hypersonic to deccel between the point where the aeroshield was effective to where you can have a powered landing.
From what I've read the problem is in this greyzone where the size of the chutes have to be large enough to catch enough of the thin Martian atmosphere.
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Well it would appear current technology is just not good enough for a manned mission then.
Well, actually it is (as far as entry and landing), we just don't have the cost-effective lift capacity to get enough mass to Mars to make it feasible.
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Well it would appear current technology is just not good enough for a manned mission then.
Well, actually it is (as far as entry and landing), we just don't have the cost-effective lift capacity to get enough mass to Mars to make it feasible.
Sure there is 2 J246s per mars direct component
If 10m payload fairing on a j246 limits a landable mass to 12mT and Zubrins idea of a Hab was 25mT then the current state of the art may not be good enough. A mars sample return mission should work though.
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Sure there is 2 J246s per mars direct component
If 10m payload fairing on a j246 limits a landable mass to 12mT and Zubrins idea of a Hab was 25mT then the current state of the art may not be good enough.
Was 25mT landable? or total?
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Using current technology limits, the entry vehicle cannot have a 'density' (mass/heatshield area) of more than about 150kg/m2, of else it hits the surface before it has slowed enough to deploy a parachute. So for something using a 10m PLF this would mean an entry mass of less than 12t.
Unless your heat shield is launched to LEO in 2 or more pieces. The it can be arbitarily large.
IIRC, Mars Drive were suggesting a hinged heat shield. 10m PLF would allow almost 20m diameter. 4 times the area, 48t? 12m PLF gives almost 70t
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tnphysics - 9/3/2008 8:11 PM
The second method may not work owing to speed-of-light communication delays (up to 20 minutes+)
It will work, it is called email
I ment telephone.
Trailing a cable all the way to Mars is a bitch, though.
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Well it would appear current technology is just not good enough for a manned mission then.
Well, actually it is (as far as entry and landing), we just don't have the cost-effective lift capacity to get enough mass to Mars to make it feasible.
Mass? Volume.
Current state of the art is, what, 5m PLF? That allows about 3t entry mass. IIRC that's pretty much what MSL will be, so the current system is maxed out.
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Mass? Volume.
Current state of the art is, what, 5m PLF? That allows about 3t entry mass. IIRC that's pretty much what MSL will be, so the current system is maxed out.
So is that the real reason for the size of Ares V?
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Sure there is 2 J246s per mars direct component
If 10m payload fairing on a j246 limits a landable mass to 12mT and Zubrins idea of a Hab was 25mT then the current state of the art may not be good enough.
Was 25mT landable? or total?
It was the landed payload. The EDLS is extra.
Interestingly, historical payload fractions for Mars payloads have been less than a third of entry mass! Ouch.
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Using current technology limits, the entry vehicle cannot have a 'density' (mass/heatshield area) of more than about 150kg/m2, of else it hits the surface before it has slowed enough to deploy a parachute. So for something using a 10m PLF this would mean an entry mass of less than 12t.
Unless your heat shield is launched to LEO in 2 or more pieces. The it can be arbitarily large.
IIRC, Mars Drive were suggesting a hinged heat shield. 10m PLF would allow almost 20m diameter. 4 times the area, 48t? 12m PLF gives almost 70t
A segmented rigid heatshield is certainly one possibility. It would have to weighed up against an inflatable or a biconic system though. IMHO an inflatable system has the most promise. I could envisage a rigid folding system being quite heavy because you'd need to build the shield more heavily around the joint, with overlaps to ensure no plasma leakage. I also don't like the idea of having to choose between a) having the shield exposed all the way to Mars or b) not finding out there's a deployment problem until you get there.
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Who said we are going to land on arachutes anyway, it might easyly be lighter to drop the remaining ground speed and altitude with rockets. If the Russians really go the no parachutes route with PPTS here on earth that would really push this design idea.
Another possibility I could imagine is to use a rocket propelled rotor for landing.
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Another possibility I could imagine is to use a rocket propelled rotor for landing.
In the thin air of mars? Pucker factor as you wait for the rotors to deploy aside.
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Who said we are going to land on arachutes anyway, it might easyly be lighter to drop the remaining ground speed and altitude with rockets. If the Russians really go the no parachutes route with PPTS here on earth that would really push this design idea.
Another possibility I could imagine is to use a rocket propelled rotor for landing.
The 'remaining' speed is hypersonic- unless you have a gigantic heatshield which will be a far higher mass solution than a parachute of the same area.
You cannot fire up your landing rockets whilst travelling hypersonic. The interaction of the plume with the airstream is a bit of a mystery, apparently, and the real-life tests needed to determine just what happens would stack up to a massive research effort. This money would all be spent in the knowledge that propulsive descent will be less mass-efficient than parachutes, so in at least one aspect, inferior. Couple that with uncertainty over whether hypersonic retro-propulsion would even work, and it seems like a bit of a dead end to go down, IMHO.
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I don't see why we want to put the bulk of the mission payload down on the surface at the bottom of the gravity well. Stage everything as a primary mission base in low mars orbit. Then send robots, cargo, and crew down to the surface in a reusable surface shuttle in pieces that are manageable given the physical limitations of the entry vehicle.
This is what worries me about mars mission planning. If we haven't figured out how to use LEO as an effective space base, then we're likely to get wildly off course in planning for mars exploration. And the Mars Direct proposals seems to actively badmouth LEO development as a distraction from the self-proclaimed objective of putting man on mars.
It's no easier to live in a life support vessel on the martian surface than it is to live in a life support vessel in mars orbit, but it's a lot harder to get back from the surface than it is to get back from orbit.
These days, space exploration usually only makes the evening news when something goes wrong. One of the few exceptions was the pair of mars rovers, which substantially exceeded expectations and whose stunning images touched the particular demographic that needs to be engaged in order for space exploration to succeed: those who are mindful of the need to fund open-ended scientific discovery but skeptical of the payoff from space exploration in particular.
The decisive scientific and political success of the mars rover program should be in the back our minds as we plan further mars missions. The vocal minority in favor of colonizing mars is blinded to a variety of other mission profiles that are more likely to generate positive results and popular support.
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Very well said.
Analyst
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It's no easier to live in a life support vessel on the martian surface than it is to live in a life support vessel in mars orbit,
Incorrect. ISRU is feasible on the surface
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It's no easier to live in a life support vessel on the martian surface than it is to live in a life support vessel in mars orbit,
Incorrect. ISRU is feasible on the surface
Perhaps, but locating the main base in orbit doesn't preclude landing a Sabatier reactor module on the surface and lifting the propellant back up to the orbiting base in several trips with a reusable surface shuttle.
We don't want to land a massive return vehicle. Even if the propellant is produced on the surface, we still don't want to send the return vehicle down the to surface to fuel up. We want to bring the propellant up to the vehicle stationed in orbit.
Also, conditions on the surface can be challenging and variable, so it would be nice for the surface crew to be able to retreat to the orbiting base in any number of contingencies without committing to earth return.
The fact that mars has an atmosphere is convenient for propellant production but decidedly inconvenient in most other ways, from aerodynamic loading to dust storms.
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The "reusable" shuttle is the hole in your plan.
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I don't see why we want to put the bulk of the mission payload down on the surface at the bottom of the gravity well. Stage everything as a primary mission base in low mars orbit. Then send robots, cargo, and crew down to the surface in a reusable surface shuttle in pieces that are manageable given the physical limitations of the entry vehicle.
It depends which DRM you look at. A 'semi direct' plan leaves the MTV and TEI stage in Mars orbit and only sends the surface habitat, rover, and ascent vehicle to the surface.
It's no easier to live in a life support vessel on the martian surface than it is to live in a life support vessel in mars orbit, but it's a lot harder to get back from the surface than it is to get back from orbit.
Absolutely incorrect. On the surface, mars can provide you with gravity, materials fort radiation shielding, water, oxygen, propellant, even a medium for growing food in the long term. If you are in space, you must be 100% self-reliant.
The decisive scientific and political success of the mars rover program should be in the back our minds as we plan further mars missions. The vocal minority in favor of colonizing mars is blinded to a variety of other mission profiles that are more likely to generate positive results and popular support.
I think you have some preconceived notions at work here. Some people do advocate 'colonising' by building up surface resources at a single location from the first mission onwards. But most people, myself included, would opt for a series of initial missions visiting different locations.
One idea which I have looked into personally is a 'traverse' model where the crew would deliberately land thousands of km from their ascent vehicle, using mobile habitats to slowly wander across the surface over the course of eighteen months. There are some major technical hurdles with this idea though.
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Perhaps, but locating the main base in orbit doesn't preclude landing a Sabatier reactor module on the surface and lifting the propellant back up to the orbiting base in several trips with a reusable surface shuttle.
Your 'reusable shuttle' would need a collosal development effort to make it a reality. We're talking a leap in technology akin to STS. You're also proposing a model which has yet to be demonstrated as viable here on Earth- RLVs supplying a propellant depot.
We don't want to land a massive return vehicle. Even if the propellant is produced on the surface, we still don't want to send the return vehicle down the to surface to fuel up. We want to bring the propellant up to the vehicle stationed in orbit.
This was the flaw in the original Mars Direct. Having the ERV ascend from the surface meant that it was unfeasibly small. More recent proposals leave the MTV in orbit as a return vehicle which allows the ascent vehicle to be a smaller and simpler affair.
Also, conditions on the surface can be challenging and variable, so it would be nice for the surface crew to be able to retreat to the orbiting base in any number of contingencies without committing to earth return.
The fact that mars has an atmosphere is convenient for propellant production but decidedly inconvenient in most other ways, from aerodynamic loading to dust storms.
If I understand you correctly, though, you are proposing an orbit-based model where the crew make multiple trips to and from the surface in a reusable vehicle. So they pass through atmospheric entry and landing several times, not just once. You are over-exposing them to the two most hazardous phases of a Mars mission.
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Is the top of Olympus Mons still in the atmosphere?
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Yes, the surface shuttle is certainly the long pole in my proposed mission profile, and yes, we should absolutely gain experience with propellant depots in LEO before we attempt this in mars orbit.
That's my point: we have to substantially expand development of LEO as an orbiting spaceport and substantially refine the performance, economics, and capacity of mass transport between surface and orbit.
We have to get it right on earth before we can do it right on mars. COTS-C/D is a step in the right direction, but we have a ways to go. Low earth orbit is the gateway to interplanetary space travel.
We put men on the moon, and then we left. We squandered a whole lot of developmental momentum, and we haven't been back since. Wherever man sets foot in the universe, we should establish and maintain an orbiting spaceport. That way, we never regress, and we gain footholds from which to stage missions to more distant worlds.
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The first mission to mars should be a preparatory unmanned mission that carries the initial module of the mars orbiting habitat and a surface shuttle loaded with a Sabatier reactor. After the habitat is orbited, the shuttle would descend, land, deploy the reactor, refuel, ascend, and re-dock with the habitat.
This would be the proof of concept mission to qualify the shuttle and reactor for a manned mission. Because there's no crew to support and nothing to return to earth, smaller or fewer launch vehicles would be required and/or a larger habitat can be delivered.
I think that demonstrating the production of propellant from the martian atmosphere would go a long way to impress the kinds of people we need to engage in order to protect and expand funding for space exploration, and this is a short path to that milestone.
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Is the top of Olympus Mons still in the atmosphere?
No, even at 27 km, the atmospheric pressure at the top is at least 5% of the surface pressure. The lower gravity on Mars allows its atmosphere to extend much higher. Pretty thin but enough for serious heating at orbital velocity.
Anyway, if you were going to slow to sub-sonic velocity relative to the surface above the atmosphere, then land, would 27 km less make that much difference? Maybe 400 m/s delta v for a powered descent.
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Is the top of Olympus Mons still in the atmosphere?
No, even at 27 km, the atmospheric pressure at the top is at least 5% of the surface pressure. The lower gravity on Mars allows its atmosphere to extend much higher. Pretty thin but enough for serious heating at orbital velocity.
Anyway, if you were going to slow to sub-sonic velocity relative to the surface above the atmosphere, then land, would 27 km less make that much difference? Maybe 400 m/s delta v for a powered descent.
I'm thinking the summit of the mountain should be looked at as a possible landing place
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Landing at higher altitude is a difficulty, not something that makes the process easier. The paper on EDLS explains this.
If you aim for the lowest points on Mars, e.g. Hellas Basin, you can land the largest payloads. This is because your aerodynamic decellerators have the best chance to work.
The only way landing on Olympus Mons could be 'beneficial' would be in an all-powered descent, which would be horribly costly in delta-v.
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The first mission to mars should be a preparatory unmanned mission that carries the initial module of the mars orbiting habitat and a surface shuttle loaded with a Sabatier reactor. After the habitat is orbited, the shuttle would descend, land, deploy the reactor, refuel, ascend, and re-dock with the habitat.
This would be the proof of concept mission to qualify the shuttle and reactor for a manned mission. Because there's no crew to support and nothing to return to earth, smaller or fewer launch vehicles would be required and/or a larger habitat can be delivered.
I think that demonstrating the production of propellant from the martian atmosphere would go a long way to impress the kinds of people we need to engage in order to protect and expand funding for space exploration, and this is a short path to that milestone.
For a long term plan, perhaps your model makes sense. But I think that we accomplish exploration of Mars using more proven systems first. I want to see a Mars landing in my lifetime.
I do think, though, that your reasoning may be off. Mars orbit is not really a safer place than the surface. You aren't really any closer to getting home since you still have to wait for a launch window.
Let's look at worst case scenarios- some major malfuction to a propulsion system strands you at Mars.
If you're on the surface, you can continue to grow food, generate oxygen, mine water ice, even manufacture new propellant in case you suffered a leak and lost the first batch. You might have to sit tight till the next launch window but you'll basically be alright.
If you're in orbit, and a failure loses you the launch window, then you need either a 100% closed ECLSS or a huge margin of reserve consumables to tide you over while you wait on a rescue mission at the next window.
I know where I'd rather be.
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The first mission to mars should be a preparatory unmanned mission that carries the initial module of the mars orbiting habitat and a surface shuttle loaded with a Sabatier reactor. After the habitat is orbited, the shuttle would descend, land, deploy the reactor, refuel, ascend, and re-dock with the habitat.
This would be the proof of concept mission to qualify the shuttle and reactor for a manned mission. Because there's no crew to support and nothing to return to earth, smaller or fewer launch vehicles would be required and/or a larger habitat can be delivered.
I think that demonstrating the production of propellant from the martian atmosphere would go a long way to impress the kinds of people we need to engage in order to protect and expand funding for space exploration, and this is a short path to that milestone.
For a long term plan, perhaps your model makes sense. But I think that we accomplish exploration of Mars using more proven systems first. I want to see a Mars landing in my lifetime.
I do think, though, that your reasoning may be off. Mars orbit is not really a safer place than the surface. You aren't really any closer to getting home since you still have to wait for a launch window.
Let's look at worst case scenarios- some major malfuction to a propulsion system strands you at Mars.
If you're on the surface, you can continue to grow food, generate oxygen, mine water ice, even manufacture new propellant in case you suffered a leak and lost the first batch. You might have to sit tight till the next launch window but you'll basically be alright.
If you're in orbit, and a failure loses you the launch window, then you need either a 100% closed ECLSS or a huge margin of reserve consumables to tide you over while you wait on a rescue mission at the next window.
I know where I'd rather be.
That scenario assumes an ISRU-based architecture in the first place, and I will be surprised if it gets done that way. If it was, some parts of ISRU (like fuel manufacture, since it may be a chondritic body) may not be out of the question at a Phobos base, which would do just fine for an orbital site at Mars. It may not have much gravity, but it has enough components left lying on the ground aren't going to simply drift away. One other benefit of an LMO or Phobos based expeditionary architeture is, it makes resupply from Earth easier. Assuming you had HLLV available (doesn't matter which one, could be Ares V, JS-246, or evolved EELV), you can ship fresh supplies via something like ATV, HTV, or even Dragon or Orion (the last two might allow you to consider aerobraking at Mars). People tend to forget the opposition class trajectory is to maximize payload and minimize flight time simultaneously. With a smaller payload, you can shorten flight times for non-opposition (non-Hohmann) trajectories. If I were planning an orbital/Phobos-based expeditionary architecture to explore Mars, I'd want that unmanned almost-anytime resupply capability as part of the architecture from day one. Planning for survival by ISRU on the ground means your architecture is essentially a Mars-base architecture.
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For a long term plan, perhaps your model makes sense. But I think that we accomplish exploration of Mars using more proven systems first. I want to see a Mars landing in my lifetime.
The unfortunate aspect of the surface-oriented mission profile is that launching two separate long-term life support vessels for transit and surface, although incurring a substantial mass penalty, is still a better option than launching and landing a rocket stage big enough to lift the habitat for the return voyage.
I say, put off the long-term surface base for now and do a 2-day flag-planting photo-op using a small lander departing from an orbiting long-term habitat and ERV. Landing men on mars is easier than supporting men on mars for any extended stay, so shouldn't we accomplish that first?
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That scenario assumes an ISRU-based architecture in the first place, and I will be surprised if it gets done that way.
I've actually come to the conclusion that an ISRU ascent stage is virtually a necessity.
Run the numbers on a ISRU MAV- for a crew of four to six you're talking anywhere from about 20-45t entry mass depending on how you calculate it. Big, but within the potential capabilities of systems similar to those used already and scaled up to fit on an SDLV.
The mass for a pre-fuelled MAV is just scary, by comparison. The extra mass of the ascent propellant drives up the mass of the rest of the design, so you're talking 60-100t at entry. You're trying to land a fully fuelled rocket on the surface, and it will not be easy, cheap, or safe. It will absolutely require a fundamental change in entry/descent technology- it will make STS look easy by comparison.
If it was, some parts of ISRU (like fuel manufacture, since it may be a chondritic body) may not be out of the question at a Phobos base, which would do just fine for an orbital site at Mars. It may not have much gravity, but it has enough components left lying on the ground aren't going to simply drift away.
If the MAV is ISRU, then it should be a mission rule that it is ready to launch by the time the crew get there. This means robotic deployment and operations. All near-term proposals have used atmospheric ISRU only, because it is a globally available, homogenous resource, that can be sucked in by a far no matter where you are on the planet. It can be processed easily by cryogenic distillation. Regolith processing, whether on Mars or Phobos, requires a whole different set of tools- a dose of good luck to land in the right place, machinery to grind up the material, and massive amounts of heat to drive off the substances that you want. It just doesn't sound like something you could achieve without a human presence, and certianly is out of the question for the first missions.
One other benefit of an LMO or Phobos based expeditionary architeture is, it makes resupply from Earth easier. Assuming you had HLLV available (doesn't matter which one, could be Ares V, JS-246, or evolved EELV), you can ship fresh supplies via something like ATV, HTV, or even Dragon or Orion (the last two might allow you to consider aerobraking at Mars). People tend to forget the opposition class trajectory is to maximize payload and minimize flight time simultaneously. With a smaller payload, you can shorten flight times for non-opposition (non-Hohmann) trajectories. If I were planning an orbital/Phobos-based expeditionary architecture to explore Mars, I'd want that unmanned almost-anytime resupply capability as part of the architecture from day one. Planning for survival by ISRU on the ground means your architecture is essentially a Mars-base architecture.
Your fast flight time requires a lot of braking at the other end. For unmanned cargo flights, why not just employ direct entry and landing on Mars? This will be a lot lower mass than propulsive capture and rendezvous with Phobos. It will also make more sense than aerocapture followed by Phobos rendezvous.
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For a long term plan, perhaps your model makes sense. But I think that we accomplish exploration of Mars using more proven systems first. I want to see a Mars landing in my lifetime.
The unfortunate aspect of the surface-oriented mission profile is that launching two separate long-term life support vessels for transit and surface, although incurring a substantial mass penalty, is still a better option than launching and landing a rocket stage big enough to lift the habitat for the return voyage.
I say, put off the long-term surface base for now and do a 2-day flag-planting photo-op using a small lander departing from an orbiting long-term habitat and ERV. Landing men on mars is easier than supporting men on mars for any extended stay, so shouldn't we accomplish that first?
A Mars mission MUST last two to three years. From the moment you complete TMI, you are commited to that length of mission.
So your suggestion is that crew spend two days on Mars and then return to orbit to wait for the remaining eighteen months for their launch window? And the benefit is what, exactly?
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Some people do advocate 'colonising' by building up surface resources at a single location from the first mission onwards. But most people, myself included, would opt for a series of initial missions visiting different locations.
IMO, it is a very reasonable compromise to begin clustering hardware together from the first missions. This is especially true with the distances and durations involved with a Mars mission. The cost and safety advantages are too great to ignore. Scavenging a part from a previous mission could save a science objective if an instrument failed or perhaps save a crew if something critical failed.
Regarding exploration activities, it's hard to imagine that a single expedition would exhaust all the research opportunities within the range of a given landing site. A pressurized rover with a few hundred kilometers of range would keep the first few expeditions plenty busy. By the later missions, perhaps a suborbital "hopper" could open even greater distances.
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If it was, some parts of ISRU (like fuel manufacture, since it may be a chondritic body) may not be out of the question at a Phobos base, which would do just fine for an orbital site at Mars. It may not have much gravity, but it has enough components left lying on the ground aren't going to simply drift away.
If the MAV is ISRU, then it should be a mission rule that it is ready to launch by the time the crew get there. This means robotic deployment and operations. All near-term proposals have used atmospheric ISRU only, because it is a globally available, homogenous resource, that can be sucked in by a far no matter where you are on the planet. It can be processed easily by cryogenic distillation. Regolith processing, whether on Mars or Phobos, requires a whole different set of tools- a dose of good luck to land in the right place, machinery to grind up the material, and massive amounts of heat to drive off the substances that you want. It just doesn't sound like something you could achieve without a human presence, and certianly is out of the question for the first missions.
The gas ISRU lander and the solid ISRU lander can be different machine at different locations, providing there is a way of moving the refined resources to the Mars outpost. This will reduce the mass of each lander at the cost of an extra launch vehicle.
On Mars the hydrogen appears to be in the form of the solid called ice. Before the big ISRU equipment arrives mini rovers may be needed to prospect for surface ice. If they move at 2 miles per hour during day light they should be able to cover about
2 * 24.6 / 2 = 24.6 miles.
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If they move at 2 miles per hour during day light they should be able to cover about
2 * 24.6 / 2 = 24.6 miles.
meaningless numbers. What determines the 2 mph? Also not all the daylight hours are usable.
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The gas ISRU lander and the solid ISRU lander can be different machine at different locations, providing there is a way of moving the refined resources to the Mars outpost. This will reduce the mass of each lander at the cost of an extra launch vehicle.
read the post, this is not for the early missions. Again, you bring up things not appropriate for the early missions. None of these missions, both lunar and martian are settlements, they are outposts. There won't be mining or excavators for many years, they are way outside the timeframe of discussions on these threads. When are you going realize this and stop inserting them needlessly into the discussions.
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If they move at 2 miles per hour during day light they should be able to cover about
2 * 24.6 / 2 = 24.6 miles.
meaningless numbers. What determines the 2 mph? Also not all the daylight hours are usable.
E = 0.5 * m v2
The current Mars rovers show that energy is restricted by the size of the solar panels and mass restrictions mean we can only have small panels.
2 mph is walking speed. Slower than that and only small areas can be surveyed. Even with restricted day light hours the rover should be able to travel 80 to 90 miles in 4 days. Hopefully the probe can be landed that close to its target location.
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The gas ISRU lander and the solid ISRU lander can be different machine at different locations, providing there is a way of moving the refined resources to the Mars outpost. This will reduce the mass of each lander at the cost of an extra launch vehicle.
read the post, this is not for the early missions. Again, you bring up things not appropriate for the early missions. None of these missions, both lunar and martian are settlements, they are outposts. There won't be mining or excavators for many years, they are way outside the timeframe of discussions on these threads. When are you going realize this and stop inserting them needlessly into the discussions.
I am aiming this at the initial boots and flag mission.
The other posters are planning to use ISRU propellant for the return fuel. See quotes below.
The ascent propellant could be pure hydrogen or its compound methane. Since there are only trace quantities of hydrogen in Mars's atmosphere usable quantities of hydrogen will have to be extracted from ice in the soil.
See
Posted on: 29 April 2009, 13:23:42
Posted by: Kaputnik
Quote from: William Barton on 29 April 2009, 12:14:06
That scenario assumes an ISRU-based architecture in the first place, and I will be surprised if it gets done that way.
I've actually come to the conclusion that an ISRU ascent stage is virtually a necessity.
Run the numbers on a ISRU MAV- for a crew of four to six you're talking anywhere from about 20-45t entry mass depending on how you calculate it. Big, but within the potential capabilities of systems similar to those used already and scaled up to fit on an SDLV.
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If they move at 2 miles per hour during day light they should be able to cover about
2 * 24.6 / 2 = 24.6 miles.
meaningless numbers. What determines the 2 mph? Also not all the daylight hours are usable.
E = 0.5 * m v2
The current Mars rovers show that energy is restricted by the size of the solar panels and mass restrictions mean we can only have small panels.
...
Oh please.
That's the formula for the kinetic energy of a moving mass. Not for how much power you need to compensate for friction at a particular speed, in a particular environment.
It's completely irrelevant in this context.
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E = 0.5 * m v2
The current Mars rovers show that energy is restricted by the size of the solar panels and mass restrictions mean we can only have small panels.
2 mph is walking speed. Slower than that and only small areas can be surveyed. Even with restricted day light hours the rover should be able to travel 80 to 90 miles in 4 days. Hopefully the probe can be landed that close to its target location.
Another meaningless and clueless post
The current mars rovers did nothing of the sort. The fact is common sense which is something absent in the post
What mass restrictions?
A rover wouldn't be able to travel that far in four days with current collision avoidance systems
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Given how long it's taking NASA to design a 'simple' lunar lander ...
For a manned program there is no such thing as a simple lander.
Agreee, hence my use of inverted commas.
I was comparing the Altair design effort with the work required to build the capability to manufacture lunar-derived propellants for a Mars mission. i.e. if it takes many years and billions of dollars to do Altair, what will it take to make machines that work in harsh lunar conditions grinding up regolith and processing it into useful propellant? Not to mention spacecraft to deliver that propellant to where it's needed.
Perhaps it will be a reality one day. But why do people want to put it on the criticla path to Mars?
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What mass restrictions?
The manned Mars transfer vehicle may be constructed in space but the ISRU lander may have to go up on a single launch. Any associated rover will have to be restricted to only some of that mass.
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What mass restrictions?
The manned Mars transfer vehicle may be constructed in space but the ISRU lander may have to go up on a single launch. Any associated rover will have to be restricted to only some of that mass.
Where does it say there are rovers with the ISRU lander?
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What mass restrictions?
The manned Mars transfer vehicle may be constructed in space but the ISRU lander may have to go up on a single launch. Any associated rover will have to be restricted to only some of that mass.
Where does it say there are rovers with the ISRU lander?
That is my suggestion.
The rover(s) would collect the water for the ISRU chemical processor.
I will not preclude a trade study finding separate landers for the processor and rovers to be cheaper but I am not assuming it.
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Who says a future rover needs to avoid small rocks? there is a project ongoing on earth that uses a inflatable ball as the basis for a roving system, and there are many other options.
If you look at current robotics technology on earthyou we see the fastet progress in history currently ongoing, this is because there are now competitive events focusing on autnomous robotics.
Look at these soccer playing robots http://www.youtube.com/watch?v=DS0xuXZ6r8g
it's worth noting that both teams play without any human interaction (except for starting/stopping the game), they have all sensors mounted on board (notice the mirror on the top that is used for 360 degree vision with a single camera) and they also carry all their computing power on board.
That will put the possibilities of automated collision avoidance into perspective.
I think the biggest obstacle for really autonomous rovers is risk aversity. one needs to trust the AI to do it's thing. We are seeing this on the current rovers as well as they have completed their primary mission more and more of the navigation needs is done directly on the rovers and even though they are really old now they travel faster then ever before.
I think the best thing for Rover technologies would be to do a price, e.g. the team that sends a rover through desert X within the timeframe Y wins a million dollars. I think this concept could work very well for rovers because the entry barrier is low enough (those soccer playing robots have a building cost below 10 000 $)
The hardware we'd need for efficient rovers is available what we need to do is make it radiation hardened (e.g. the Robots from above run on standard laptops (they are no high end models though a Netbook would be enough)) and we need to write the software of which many problems will be solved by projects like those soccer playing robots.
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Zubrin suggested packaging a rover with the ISRU lander, so as to carry the reactor to a safe distance.
However it all depends on what you cna actualy land in one go. There's no sense in needlessly splitting things up, but many payloads could be divided without a lot of difficulty.
The ascent vehicle could be separate from the ISRU plant. The reactor could be separate from both of these, and the rover separate again. But it doesn make life difficult and probably increases total landed mass because you have to factor in the connections between these units. It also introduces new failure modes, e.g. if the connections are damaged.
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I presume that everything needed to land on mars, refuel, and return to earth must be tested prior to sending a crew.
So at the risk of potentially delaying a manned mars mission, why don't we pursue a flagship robotic mars mission that combines heavy exploration rovers with heavy sample return based on ISRU?
This would exercise most of the capabilities we will need for a manned mission, on a sufficiently large but somewhat smaller scale.
The general public perceives the MER program far more positively than anything else that NASA has done recently. So now we send robots to find martian water and bring it back to earth using propellants manufactured from martian atmosphere.
That's how we write the missing chapter in the future history of space exploration that connects what we've already done with where we want to go. We've proven that we can send robots to mars. Now we send robots to prove that we can send humans to mars.
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Some people, me included, suggest a 'mission zero' be flown prior to a manned mission. This would use only slightly modified hardware and go through all crucial mission phases, returning a cargo of rock samples. Depending on the mission design, you might need such a mission to pre-place some elements anyway.
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The robotic mission would leave the reactor, which might be reused for the first manned mission. The lander could carry other useful payload. It could even be a habitat.
Another, perhaps riskier option is to send pressurized rovers that can be operated autonomously, remotely, and then later manually.
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What I was thinking was that if the mission design relied upon a SEP Tug, or similar, this could be put in place for 'mission zero'. Any surface elements like power sources or rovers could be left in place as backups for the first manned mission.
One mission design which I've been drawing up myself would fly out two MAVs (ascent vehicles) in mission zero, one being used for the sample-return; subsequent manned missions would use the oldest available MAV, leaving a new one so that there is always a backup.
However when talking of missions at this level of detail it must be remembered that the overall mission design is dependent on the LVs, entry systems, propulsion technology, trajectories, and crew size.
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MAV and rovers able to take people are big ones, possibly more than an initial sample return would need.
A Mars SEP could be a Lunar SEP with bigger fuel tanks and solar arrays.
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MAV and rovers able to take people are big ones, possibly more than an initial sample return would need.
Indeed. But IMHO it will be be better and cheaper to use the manned hardware designs for the MSR. It avoids having to build a different set of spacecraft, and it provides a thorough test of the designs in operation. If you plan on doing an all-up test of the the manned equipment later, then the MSR mission cannot share costs to the same extent.
A Mars SEP could be a Lunar SEP with bigger fuel tanks and solar arrays.
All depends on mission plan. If a SEP Tug is only used to get components up to L2, for example, it needn't really be any bigger at all.
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So, during the trip to Mars, how are they going to solve a Problem with the Astronauts being so lonley from away from family?
Send non-married astronauts :)
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So, during the trip to Mars, how are they going to solve a Problem with the Astronauts being so lonley from away from family?
Shorten the trip by starting from and returning to EML-2. Using SEP (or NEP), accelerate out for half the deltaV budget and decelerate in for the other half. Perform the surface mission. On the return leg, do the same back to EML-2. By the time they arrive back at EML-2, they should be able to enter the halo orbit with little more than docking thrusters.
Powered flights like this solve several problems, not the least of which is a very reduced transit time. Also, while the g-load will be very light (*almost* zero-g), its very presence, no matter how slight, will make equipment and hardware easier to design.
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Question about low-thrust TMI.
A standard 1-impulse manoeuvre only needs about 4km/s to achieve TMI. However a slow spiral under continuous thrusting is far less efficient.
Aside from the issue about crossing the VA belts, does anybody know what sort of delta-v is required to achieve TMI by spiralling?
Secondly, are there any clever strategies to this which can balance thrust, propellant, VA exposure, and time taken, by thrusting only at certain points in the manoeuvre and/or supplementing with high-thrust propulsion?
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Here's another question.
Zubrin "Mars Direct" and electric-propulsion proposals have something in common = low IMLEO - for different reasons.
So the question is = would it be possible to go to Mars using SEP, then going back using Mars Direct sheme - ISRU at Mars ?
I mean, having a SEP-ferry on the way up, and a chemical LOX/Methane engine on the way back.
In this sheme the crew could return Earth two ways
- going back with the SEP-tug (low-thrust, commuting with an Orion at EML2)
or
- going back using LOX/Methane (the usual high-thrust trajectory)
Or
- using the chemical engine to "boost" solar-electric propulsion and cut the trip time when returning.
In other words, I want to use the chemical Earth Return Vehicle to boost the SEP-Mars-ferry on the way back.
Hybrid chemical/electric proopulsion tug looks feasible. Despite differences in trajectories it looks like low- and high- thrusts may be combined.
Here's some links.
http://pdf.aiaa.org/GetFileGoogle.cfm?gID=3791&gTable=japaperimport
http://www.stormingmedia.us/29/2948/A294804.html
http://pdf.aiaa.org/GetFileGoogle.cfm?gID=10191&gTable=mtgpaper
http://pdf.aiaa.org/GetFileGoogle.cfm?gID=26623&gTable=japaperimportPre97
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Yes, you can combine different propulsion systems.
LEO to higher rendezvous point (high circular or elliptical orbit, or a lagrange point) can be done chemically or by high-isp.
From there through TMI can likewise be done in a variety of ways- or combined with the first manoeuvre.
Arrival at Mars can be by direct entry and landing, or aerocapture, or a variety of propulsive methods.
Ascent from Mars surface obviously needs to be chemical propulsion, IMO it is virtually essential that this uses ISRU though, for a number of reasons.
From Mars orbit through TEI can be done chemically, either with Earth-sourced or Mars-sourced propellants, or by high-isp propulsion.
On arrival at Earth the options are similar to Mars arrival- direct entry, aerocapture, or chemical or high-isp propulsion.
FWIW, the original 'Mars Direct' scheme seems to be out of favour these days. It forces the return habitat vehicle to be far too small- you cannot live in an Orion for six months. More up to date plans involve rendezvous in Mars orbit with a return hab. This negates some of the advantages of using ISRU for the TEI burn because leaving it in Mars orbit becomes quite attractive.
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So, during the trip to Mars, how are they going to solve a Problem with the Astronauts being so lonley from away from family?
Shorten the trip by starting from and returning to EML-2. Using SEP (or NEP), accelerate out for half the deltaV budget and decelerate in for the other half. Perform the surface mission. On the return leg, do the same back to EML-2. By the time they arrive back at EML-2, they should be able to enter the halo orbit with little more than docking thrusters.
Powered flights like this solve several problems, not the least of which is a very reduced transit time. Also, while the g-load will be very light (*almost* zero-g), its very presence, no matter how slight, will make equipment and hardware easier to design.
I like this concept.
I feel that Sun-Mars L1 should be included in such shemes. Maybe incremental steps could be of interest,, considering limitations in budget ?
something like
Expedition 1- Sun-Mars L1
Expedition 2- Phobos landing
Expedition 3- Mars Orbit
Expedition 4- Mars landing.
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Expedition 1- Sun-Mars L1
Expedition 2- Phobos landing
Expedition 3- Mars Orbit
Expedition 4- Mars landing.
This is very similar to what's proposed in the following study, though starting from SEL-2 as that's even more efficient:
Next Steps In Exploring Deep Space (http://iaaweb.org/iaa/Studies/nextsteps.pdf)
It mentions the possibility of using Sun-Mars L1, and recommends further investigation of that possibility.
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Expedition 1- Sun-Mars L1
Expedition 2- Phobos landing
Expedition 3- Mars Orbit
Expedition 4- Mars landing.
Just a note that Expeditions 2 and 3 provide a really important scientific tool, which is the proximity and bandwidth required to do real-time telerobotic control of probes on the Mars surface (or in the atmosphere). So they're valuable in their own right, not just practice for the "real thing."
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Expedition 1- Sun-Mars L1
Expedition 2- Phobos landing
Expedition 3- Mars Orbit
Expedition 4- Mars landing.
This is very similar to what's proposed in the following study, though starting from SEL-2 as that's even more efficient:
Next Steps In Exploring Deep Space (http://iaaweb.org/iaa/Studies/nextsteps.pdf)
It mentions the possibility of using Sun-Mars L1, and recommends further investigation of that possibility.
I knew the document but I had missed the "sun mars L1" part.
Going to SEL-2 takes more time, up to 90 days. I would prefer telescope servicing missions there.
Just a note that Expeditions 2 and 3 provide a really important scientific tool, which is the proximity and bandwidth required to do real-time telerobotic control of probes on the Mars surface (or in the atmosphere). So they're valuable in their own right, not just practice for the "real thing."
And there come Zubrin's Athena. :)
http://pdf.aiaa.org/getfile.cfm?urlX=85%26%5D0%3BU%2BDN%26S7R%20CMU%24CBQ%3A%2B64K8%26%5FOGJ%0A&urla=%25%2ARH%27%21P%2C%20%0A&urlb=%21%2A%20%20%20%0A&urlc=%21%2A0%20%20%0A&urle=%27%2B%22D%22%23PJCU0%20%20%0A
(sorry for the disastrously long URL !)
I know that Zubrin tends to be way too optimistic on his mass budget estimations (he took many flak on the subject with Mars Direct)
As a result Athena was to be a 20 tons spacecraft only (!). Kind of Salyut-to-Mars.
What is interesting in the proposal is Zubrin little calculations.
According to the paper, the Athena ship could be sent to Mars
within a single Energia launch
or, as an alternative...
Four Protons are then used to lift to orbit and
mate with the hab four storable propulsion stages,
each with a propellant mass of 18 tonnes and a dry
mass of 2 tonnes, and an Isp of 326.5 (i.e. Russian
RD-0210 N2O4/UDMH engines). This combination
can throw 26 tonnes onto TMI with a 03 of 18
km2/s2.
Let's replace the four Proton by two Jupiter 120... more power, more margins. Keep the hypergolics for the sake of simplicity.
Athena was not intented to Sun-Mars L1 nor Phobos nor Mars orbit; it was to be a "double flyby". Maybe Athena could be "retargeted" to Sun-Mars L1 (feedback welcome). Or even to Phobos, which delta-vee is said to be lower than going to the Moon...
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Going to SEL-2 takes more time, up to 90 days.
I think it's only 14 days if you use chemical propulsion.
Four Protons are then used to lift to orbit and
mate with the hab four storable propulsion stages,
each with a propellant mass of 18 tonnes and a dry
mass of 2 tonnes, and an Isp of 326.5 (i.e. Russian
RD-0210 N2O4/UDMH engines). This combination
can throw 26 tonnes onto TMI with a 03 of 18
km2/s2.
Talk about coincidence. As you may know I have been thinking a lot about architectures that involve hypergolic depots. The other day I read that von Braun's initial "Marsprojekt" involved hypergolic propellants. A quick look at a delta-v chart (from the non-authoritative Wikipedia) shows that with properly prepositioned propellant, you could make a journey to Mars orbit with individual hops of no more than 2 km/s.
(http://upload.wikimedia.org/wikipedia/commons/thumb/c/c9/Deltavs.svg/590px-Deltavs.svg.png)
With such small delta-v's the inefficiency of hypergolics is only about 20%. And the good thing is, you could use SEP to preposition the propellant - meaning storables might actually be more efficient than LOX/LH2 that is not prepositioned! If anyone has a good source for the delta-v's involved, we could do the sums.
In other words, it looks as if we could really go to Mars soon. Of course, the propellant transfer is risky, but we could do an unmanned precursor soon. Once you have a hypergolic propellant depot and a SEP tug that would not have to be capable of crossing the van-Allens multiple times, you could send an unmanned Orion on a trip to Mars orbit and back to Earth. You don't even need a bigger launcher. Two EELVs will get an Orion and Centaur to LEO. The Centaur would need some kind of mission kit for limited boil-off mitigation, and this has already been investigated by ULA. The Orion would need slightly bigger propellant tanks. The Centaur is able to get an unmanned Orion to L1 on a very slow trajectory (>100 days). Once you're at L1, you have the two biggest delta-v hurdles (Earth->LEO, LEO->L1) behind you. From there on it would be plain sailing for the unmanned Orion.
How's that for a legacy for Obama? ;)
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Going to SEL-2 takes more time, up to 90 days.
I think it's only 14 days if you use chemical propulsion.
Four Protons are then used to lift to orbit and
mate with the hab four storable propulsion stages,
each with a propellant mass of 18 tonnes and a dry
mass of 2 tonnes, and an Isp of 326.5 (i.e. Russian
RD-0210 N2O4/UDMH engines). This combination
can throw 26 tonnes onto TMI with a 03 of 18
km2/s2.
Talk about coincidence. As you may know I have been thinking a lot about architectures that involve hypergolic depots. The other day I read that von Braun's initial "Marsprojekt" involved hypergolic propellants. A quick look at a delta-v chart (from the non-authoritative Wikipedia) shows that with properly prepositioned propellant, you could make a journey to Mars with individual hops of no more than 2 km/s.
----
With such small delta-v's the inefficiency of hypergolics is only about 20%. And the good thing is, you could use SEP to preposition the propellant - meaning storables might actually be more efficient than LOX/LH2 that is not prepositioned! If anyone has a good source for the delta-v's involved, we could do the sums.
In other words, it looks as if we could really go to Mars soon. Of course, the propellant transfer is risky, but we could do an unmanned precursor soon. Once you have a hypergolic propellant depot and a SEP tug that would not have to be capable of crossing the van-Allens multiple times, you could send an unmanned Orion on a trip to Mars orbit and back to Earth. You don't even need a bigger launcher. Two EELVs will get an Orion and Centaur to LEO. The Centaur would need some kind of mission kit for limited boil-off mitigation, and this has already been investigated by ULA. The Orion would need slightly bigger propellant tanks. The Centaur is able to get an unmanned Orion to L1 on a very slow trajectory (>100 days). Once you're at L1, you have the two biggest delta-v hurdles (Earth->LEO, LEO->L1) behind you. From there on it would be plain sailing for the unmanned Orion.
How's that for a legacy for Obama? ;)
Much better (and colourful!) than wikipedia
http://www.clowder.net/hop/railroad/deltaveemap.html
More delta-vee numbers for various locations
(note : It is beyond my understanding that GEO takes 3.8 km/s while Earth escape is 3.2 km/s only ??? )
http://www.lr.tudelft.nl/live/pagina.jsp?id=f62334be-f957-48e2-9646-88edf39eb738&lang=en
http://ares.jsc.nasa.gov/HumanExplore/Exploration/EXLibrary/DOCS/EIC042.HTML (scroll down to the bottom of the page)
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(note : It is beyond my understanding that GEO takes 3.8 km/s while Earth escape is 3.2 km/s only ??? )
Just a guess, but does the 3.8km/s include the circularisation burn?
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Just a guess, but does the 3.8km/s include the circularisation burn?
According to wikipedia it does, but it doesn't include a plane change. So this would be the number from an equatorial launch site. Wikipedia gives 3.9 km/s to be precise.
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So in that case it's pretty clear how it takes more energy to reach GEO than to escape from Earth orbit entirely.
Back to the question I asked, I'm a bit surprised nobody's been able to help out on the low-thrust trajectories issue. I'll keep googling, but a lot of it goes over my head...
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Back to the question I asked, I'm a bit surprised nobody's been able to help out on the low-thrust trajectories issue. I'll keep googling, but a lot of it goes over my head...
The answer to your question is beyond my limited orbital mechanics fu, but while googling I found something called Edelbaum's equation.
Dv=sqrt(v_0^2 + v_f^2 - 2v_0*v_f*cos(i*pi/2))
v_0 would be the Earth's velocity around the sun, v_f the velocity of Mars around the sun, and i any inclination change.
This should give you the delta-v from Earth C3 to Mars C3.
Can any experts comment?
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This works out to about 6 km/s, using the numbers found on this NASA Mars Fact Sheet (http://nssdc.gsfc.nasa.gov/planetary/factsheet/marsfact.html).
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Back to the question I asked, I'm a bit surprised nobody's been able to help out on the low-thrust trajectories issue. I'll keep googling, but a lot of it goes over my head...
The answer to your question is beyond my limited orbital mechanics fu, but while googling I found something called Edelbaum's equation.
Dv=sqrt(v_0^2 + v_f^2 - 2v_0*v_f*cos(i*pi/2))
v_0 would be the Earth's velocity around the sun, v_f the velocity of Mars around the sun, and i any inclination change.
This should give you the delta-v from Earth C3 to Mars C3.
Can any experts comment?
If this is the formula I think it is, its only actually half the manouver. It looks like the formula to find the first deltaV in a Hohmann transfer orbit, although given the day I'm having I might be wrong. Wikipedia has more information on Hohmann tranfer orbits than I can easily explain.
With this sort of problem though, the transfer orbit is only half of it; to leave the Earths orbit around the Sun with a velocity greater than the Earth's you need to use a hyperbolic escape orbit with an excess velocity equal to what you need to leave the Earth's orbit with. Then you have a similar problem when you get to Mars.
I might have misunderstood something and so all of this is probably wrong (its been one of those days).
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The document I took this from was about low-thrust trajectories. The 6 km/s should only refer to the delta-v to go from Earth C3 to Mars C3. To that should be added LEO->Earth C3 and Mars C3->LMO, but I think Kaputnik was considering doing that with chemical propulsion.
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The document I took this from was about low-thrust trajectories. The 6 km/s should only refer to the delta-v to go from Earth C3 to Mars C3. To that should be added LEO->Earth C3 and Mars C3->LMO, but I think Kaputnik was considering doing that with chemical propulsion.
Hmmm, this is complicated stuff.
Are you sure you need to find 6km/s? The wiki delta-v chart shows 0.6km/s for that manoeuvre, but presumably using a single impulse.
Actually, what I'd originally been looking for was more about the LEO-C3 aspect. I think this generally requires about 4km/s on a single impulse trajectory, and is a hefty mass penalty for any interplanetary mission. So I was looking for information on how we can use low-thrust high-isp systems to spiral away from LEO instead- whilst not getting fried at the VA belts, or taking years to do it.
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Hmmm, this is complicated stuff.
I agree. :) I'm still wrestling with the basics.
Are you sure you need to find 6km/s? The wiki delta-v chart shows 0.6km/s for that manoeuvre, but presumably using a single impulse.
The 0.6 km/s is for Earth C3->Mars transfer. Mars transfer->Mars C3 is 0.9 another km/s. I take it these are the two impulses for a Hohmann transfer.
Actually, what I'd originally been looking for was more about the LEO-C3 aspect. I think this generally requires about 4km/s on a single impulse trajectory, and is a hefty mass penalty for any interplanetary mission. So I was looking for information on how we can use low-thrust high-isp systems to spiral away from LEO instead- whilst not getting fried at the VA belts, or taking years to do it.
The TU Delft link given by Archibald says LEO->C3 takes 5.8-7.4 km/s for SEP, depending on thrust and 3.2 km/s for an impulsive trajectory. Are you thinking about cargo or crew? Because crew could easily stage at L1.
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You'd keep the crew on the ground up until everything was ready at the rendezvous point- whether this is a high orbit, or a lagrange point, is up to cleverer people than me. The send them in something like an Orion+DHCSS so they don't linger in the VA belts. This 'taxi' concept is well laid out in JPL's study of an AG-NEP Mars architecture.
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Yeah, that was what I was thinking of.
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Much better (and colourful!) than wikipedia
http://www.clowder.net/hop/railroad/deltaveemap.html
More delta-vee numbers for various locations
(note : It is beyond my understanding that GEO takes 3.8 km/s while Earth escape is 3.2 km/s only ??? )
Thank you for linking to my delta vee map! In my figures the delta vee for for reaching GEO includes a circulization burn at apogee. Some caveats: My figures assume circular coplanar orbits and so are a rough approximation. Also LEO is an ambiguous term. It can be any orbit from earth's surface to 2000 km up, with any inclination. In my models it is an equatorial, circular orbit 300 kilometers above the earth's surface.
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Where can we get reliable (scholarly) information on launch windows, journey times and total mission durations for human Mars missions? From Wikipedia it appears that the minimum energy launch window is about every 2.135 years (780 days)
I have made a spread sheet for Hohmann launch windows:
http://www.clowder.net/hop/railroad/Hohmann.xls
You type departure and destination planets into colored cells. You can also enter apoapsis and periapsis of planetary parking orbits you're departing from and arriving at. The sheet returns launch windows, trip times, and delta vees.
The spreadsheet assumes circular, coplanar orbits. So it's an apprximation.
Is it reliable and scholarly? My most impressive credential is graduating in the top 60% of my high school class. I do make misteaks. If anyone finds wrong assumptions or errors in my spread sheet, I'd appreciate a heads up.
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I want to compare Zubrin's Athena "double flyby" with Sun-Mars L1/ L2 alternatives.
http://forum.nasaspaceflight.com/index.php?topic=1337.0
http://forum.nasaspaceflight.com/index.php?topic=13532.0
[…] to get to L2 you are initially targeting a close pass of the Moon. At perilune you do a small DV to reduce your orbital energy relative to the Moon, then you coast out to L2 and finally do an insertion burn at L2. It is this "powered lunar swingby" that is the whole secret to the DV advantage of L2 over L1 (3.3 km/s Vs 3.8 km/s)
Do you think we can apply such manoeuvre to a Mars (or Venus!) mission ?
Something like
1- approach on a hohman orbit
2- powered martian swingby
3- enter halo orbit around Sun-Mars L2 with minimal delta-V
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I want to compare Zubrin's Athena "double flyby" with Sun-Mars L1/ L2 alternatives.
http://forum.nasaspaceflight.com/index.php?topic=1337.0
http://forum.nasaspaceflight.com/index.php?topic=13532.0
[…] to get to L2 you are initially targeting a close pass of the Moon. At perilune you do a small DV to reduce your orbital energy relative to the Moon, then you coast out to L2 and finally do an insertion burn at L2. It is this "powered lunar swingby" that is the whole secret to the DV advantage of L2 over L1 (3.3 km/s Vs 3.8 km/s)
Do you think we can apply such manoeuvre to a Mars (or Venus!) mission ?
Something like
1- approach on a hohman orbit
2- powered martian swingby
3- enter halo orbit around Sun-Mars L2 with minimal delta-V
The Earth Moon L1 and L2 points are about 84% and 116% of a lunar distance. A normal orbit at L1 altitude would travel about 1.1 km/sec wrt earth. But since L1 moves at the same angular velocity as the moon, it's only going about .85 km/sec. The difference between ordinary orbits and these L points are dramatic!
The Sun-Mars L 1 & 2 points lie about 1.08 million kilometers from Mars. These are about .5% of the Mars distance from the sun. Since these orbits are going the same angular velocity as Mars, they are moving 99.5 and 100.5% of Mars velocity wrt the Sun.
But I will plug in 1.08 million kilometers into Mars apoapsis with 300 km as Mars periapsis.
Coming in from an earth Hohmann, a .7 km/sec burn at Mars periapsis would drop the apoapsis to 1.08 million km apoapsis. Perhaps some of that .7 km/sec could be accomplished with aerobraking.
The natural circular orbit at this altitude would be about .2 km/sec (if there were no sun to perturb it). At periapsis the payload orbit is moving about .016 km/sec. A .184 km/sec circulization burn would be needed if this were an ordinary orbit.
But the L1 point is moving 0 km/sec wrt Mars. So a .016 km/sec burn would suffice to park it there.
So leaving Earth Mars Hohmann to Sun Mars L1 would take (.7 + .16) km/sec. Or about .9 km/sec.
To launch a payload from Mars to the Mars Sun L1 would take about 6 km/sec.
To launch a payload from Deimos to Mars Sun L1 would take about about .6 km/sec. About .9 km/sec for Phobos.
I think. Take what I say with a grain of salt.
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Thank you very much !
I think. Take what I say with a grain of salt.
Understood. Others posters there will critic these numbers better than I...
In the athena paper Zubrin tell us that
The crew will approach and depart Mars twice with
a hyperbolic velocity of about 4 km/s.
How much delta-V is this ? How does this compare to SML-1 / SML-2 mission profiles ?
I had this idea of two identical Athena spaceships flying to Mars in convoy (a la Von Braun Mars 1969 expedition).
They would ride in space above a Jupiter 232 / 246 or an Ares V. A single launch should be enough - Athena fit into a single Energia booster.
When arriving at Mars one would go into a "double flyby" path, the other into a halo orbit around SML-2.
Both would be +/- 1 million kilometer away from Mars.
Just a random thought.
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What is the current status of
a) VASIMR
b) Nuclear Reactor Rockets?
c) Antimatter Catalyzed Micro Fission/Fusion?
B&C have no work being done.
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1. What should be done in order to put Man on Mars within the next 10 years (before 2020)?
2. Should NASA and ESA co-operate?
3. Should we still use chemical propulsion?
...
1. Convince more than 50% of the population in Europe, US, Russia, China, ... that it's necessary for all of us to either/or survive as a species and/or do well economically as a world.
Currently, according to CBO research grants for universities the mantra is "we will make you wealthy". Back in the 60's, it was "we will make you safe" ala the Cold War, which is how Apollo was funded as a proxy war with the Soviets.
2. With the election of Obama, there's been a sea change in the desire for international partnership. However, it takes more than desire - the prior administration dismantled most international partnerships, so the US could be a more effective bully - er, able to "act unilaterally". There is still too much sentiment to wind down the ISS as an unsuccessful experiment in international cooperation.
It's incredibility subjective the hundreds of different perspectives I've heard on international cooperation including the ISS. Not everybody likes it. What you have to remember is that when government gets involved, the elements of political dealmaking (like sausage making) is a bit awkward/ugly/cumbersome/inexplicable/inefficient/... I could go on, but its just the way it works.
Now multiply the governments involved ... add more for situational instability and additional geopolitical agenda "pay offs" ... and you get the idea.
So it has little to do with the rational need your comment implies.
And at the moment there are unpredictable aspects to international cooperation that make it difficult.
Still, I think Obama will bring off some big things this way ... eventually. The timing at this moment doesn't allow it. First the US needs to get its "NASA house" in order - we're a long way from that.
3. No. Realistically, radiation exposure and logistics costs require nuclear.
You can do it with chemical but because you can do something doesn't mean it does get done that way.
I personally don't think there's any serious interest in getting to Mars in a decade unless you have funded, active nuclear exo atmospheric propulsion.
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3. No. Realistically, radiation exposure and logistics costs require nuclear.
Even with things like SEL-2 staging, Earth swingby, reuse of the MTV and use of the interplanetary transport network for cargo and propellant?
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3. No. Realistically, radiation exposure and logistics costs require nuclear.
Even with things like SEL-2 staging, Earth swingby, reuse of the MTV and use of the interplanetary transport network for cargo and propellant?
Good luck convincing public opinion you want to fly nuclear reactors in space... ::)
(not that I share these fears: Molten Salt Reactors look like a fine concept, sure enough to fly in space)
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Good luck convincing public opinion you want to fly nuclear reactors in space... ::)
Maybe it would be acceptable in sufficiently high orbits. No nukes below GEO sounds like a plausible if arbitrary rule. But the things I mentioned above were proposed by Huntress et al to avoid the use of nukes.
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Well, this is one of the reasons we're not going to Mars.
And I suspect we won't until we become more realistic as a culture.
Its not the technology that limits us, or the budget. Its human nature and motivations.
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3. No. Realistically, radiation exposure and logistics costs require nuclear.
Even with things like SEL-2 staging, Earth swingby, reuse of the MTV and use of the interplanetary transport network for cargo and propellant?
Good luck convincing public opinion you want to fly nuclear reactors in space... ::)
(not that I share these fears: Molten Salt Reactors look like a fine concept, sure enough to fly in space)
I feel the anti nuclear crowd probably are the biggest obstacle outside of politics and the Earth's gravity well to human space exploration beyond LEO.
It's just such an enabling technology you have to be a complete fool to not take advantage of it.
Some engines such as the molten saltwater rocket have such great performance they make deep space missions almost easy.
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So what are nasa's plans on figuring out microgravity effects on biological creatures prior to moon/mars/neo missions, if they threw away the "artificial gravity" maker (centrifugal module)
Wont that want to be known prior to sending dudes up for 3 years?
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I'm not aware of any active attempts to gather this data. ESA and the Russians are performing bed-rest tests.
Since the current baseline is very much Moon-first, it would seem sensible to gather partial-g data there. Remember, we already send people into zero g for six months at a time, so we can surely perform the first lunar missions in confidence.
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One of the big questions is if the transit to Mars is in microgravity, how hard will it be for the crew to adapt quickly to Mars grav once on the surface and start to work. You wouldn't want them just sitting in the descent capsule trying to adapt for several days. One way to test might be a long (6 month) duration stay at ISS followed by a trip to the Moon as an analogue, though of course gravity is less, would probably be useful.
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Yes, you can combine different propulsion systems.
LEO to higher rendezvous point (high circular or elliptical orbit, or a lagrange point) can be done chemically or by high-isp.
From there through TMI can likewise be done in a variety of ways- or combined with the first manoeuvre.
Arrival at Mars can be by direct entry and landing, or aerocapture, or a variety of propulsive methods.
Ascent from Mars surface obviously needs to be chemical propulsion, IMO it is virtually essential that this uses ISRU though, for a number of reasons.
From Mars orbit through TEI can be done chemically, either with Earth-sourced or Mars-sourced propellants, or by high-isp propulsion.
On arrival at Earth the options are similar to Mars arrival- direct entry, aerocapture, or chemical or high-isp propulsion.
FWIW, the original 'Mars Direct' scheme seems to be out of favour these days. It forces the return habitat vehicle to be far too small- you cannot live in an Orion for six months. More up to date plans involve rendezvous in Mars orbit with a return hab. This negates some of the advantages of using ISRU for the TEI burn because leaving it in Mars orbit becomes quite attractive.
Bringing the thread back to life...
I've dugged further electric/ chemical combinations. Found this
http://adsabs.harvard.edu/abs/2005AcAau..57..829M
This bring a question.
Electric propulsion trajectories are complex spirals (because thrust is too low to use the Oberth effect).
High-thrust trajectories are Hohmans.
Is there a way of combining the two into a electric-chemical trajectory ? For example, between Earth and MArs ?
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SEP cannot do a Hohmann transfer because thrust is miniscule, but then again it doesn't have to because Isp is so high. The higher delta-v is more than compensated for by the higher Isp, it's the ratio delta-v/Isp that counts.
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Yes, you can combine different propulsion systems.
LEO to higher rendezvous point (high circular or elliptical orbit, or a lagrange point) can be done chemically or by high-isp.
From there through TMI can likewise be done in a variety of ways- or combined with the first manoeuvre.
Arrival at Mars can be by direct entry and landing, or aerocapture, or a variety of propulsive methods.
Ascent from Mars surface obviously needs to be chemical propulsion, IMO it is virtually essential that this uses ISRU though, for a number of reasons.
From Mars orbit through TEI can be done chemically, either with Earth-sourced or Mars-sourced propellants, or by high-isp propulsion.
On arrival at Earth the options are similar to Mars arrival- direct entry, aerocapture, or chemical or high-isp propulsion.
FWIW, the original 'Mars Direct' scheme seems to be out of favour these days. It forces the return habitat vehicle to be far too small- you cannot live in an Orion for six months. More up to date plans involve rendezvous in Mars orbit with a return hab. This negates some of the advantages of using ISRU for the TEI burn because leaving it in Mars orbit becomes quite attractive.
Bringing the thread back to life...
I've dugged further electric/ chemical combinations. Found this
http://adsabs.harvard.edu/abs/2005AcAau..57..829M
This bring a question.
Electric propulsion trajectories are complex spirals (because thrust is too low to use the Oberth effect).
High-thrust trajectories are Hohmans.
Is there a way of combining the two into a electric-chemical trajectory ? For example, between Earth and MArs ?
I can think of two ways to send an ion engine on a Hohmann trip: Tethers and railguns.
However exiting the Hohmann to achieve a capture orbit about the destination planet might be a problem with ion engines. I think it would be more doable for Jupiter, Saturn and beyond as the ship would hang around the destination for awhile after reaching apohelion.
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hi people i am new in the forum and i registered because i love space and investigation for life and i was watching history chanel and all the misions to mars and they mentioned the mount olimpo and i wanted to know if there are any plans or idea coming or already done to send a ship to investigate what is inside like to send a shuttle inside the crater to see what is inside it and see if it can go to the nuk of the planet it would be 2 achievements
investigate mars underground and if lucky see for first time a planets nuk directly maybe at least reach some lava or any sign material inside it to take it and send it back to earth and investigate it and if mount olimpos crater is covered maybe the ship can have a drill or throw a bomb and take of that piece of rock covering it.
other idea i have is to throw a shuttle on its north pole on the white ``ice´´ and a hughe explosion blow its cover and lets the shuttle go in the ice as deep as posible
and the question is...
have they done any of those yet or is it planed to be done???
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no and no
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is it posible to do any of the above with todays tech??
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is it posible to do any of the above with todays tech??
Nuking or bombing is a bad idea
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one of my best friends father is president of a club of astronomy and he told me is posible i just called him
but to go inside the mount olimpo or the ice inside the poles can be done with todays tech and why havnt they done it??
p.s. my friends father said they did that to titan in saturn to go inside the ice
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Anything is possible (for the most part) if given enough money. There are plenty of companies that will do as you ask if you or friend's father want to pay for it ;-)
Ask your friend's Dad to give you the references of the Titan thing, please and review what it took to get there, etc (if such did occur - I am not sure of it myself) Returning samples from Mars, or any other planetary body, is currently in 'debate' stage as it is difficult to return things to Earth for a number of technical reasons (as well as $$).
I highly suggest 'researching' what you are asking about as you show a very low level of knowledge about such (not an insult, I promise you). Most of what you ask is all within this site and is very well explained as to possible or not, etc. Welcome to the site - and be glad that Jim was as nice as he was, LOL. You will find this site VERY educational as well as accessible, but you *will* need to check stuff out at least minimally before asking and then depending upon things you hear from friends that are not 'in the business' ;-) Your answers will likely come from folks that do such things professionally and know *exactly* what they are talking about - be wary about debating them if you have not done your homework (!!!!!!!!!)
Do a bit of googling about Mars and the active 'studies' happening..bet you will be surprised how much is actually happening right now and in the near future. Possibly some of what interests you is available already or will be soon (?) Heck, there is even an image of a 'lander' landing under parachute from an overhead satellite. Not real science in that image, but might help show just how well various things are being covered, IMO.
Again, welcome to NSF - the real Spaceflight site (at least as far as I am concerned),
Alex
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ty!!!!
the next question will be how long it would take a suicide satelite or ship or wut ever to trhow it into a queasar or dark whole or a a worm tunel
btw sos for my english i am learning and i am a 13 yr old latin very interested on this =)
love dark wholes (FROM SPACE XD) and ty again
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Using current technology limits, the entry vehicle cannot have a 'density' (mass/heatshield area) of more than about 150kg/m2, of else it hits the surface before it has slowed enough to deploy a parachute. So for something using a 10m PLF this would mean an entry mass of less than 12t.
Which is completely useable, if you use a little creativity ;-) (See the link in my sig.)
Hey guys, I was just looking for some good back-of-the-envelope figures to help me with my mental models.
1. So far I have been assuming 1/3.5 ratio of IMLEO to TMI (not including the mass of the spent TMI stage). For ~200 day trajectories at the reasonably opportunistic time periods, is this a reasonable figure?
2. Related to that, I have been trying to find the TMI payload for a J246. I tried to use the search but to no avail. Does anybody know it?
3. Also, what is a reasonable PMF (propellant mass fraction for those that are visiting Q&A) for a methane/LOX powered ascent vehicle to low-mars orbit, and also (3b) to an elliptical orbit suitable for rendezvous with an awaiting ERV?
4. What is a suitable PMF for a methane/LOX powered ERV for a nominal ~200 day return trajectory, from both a low-mars orbit and (4b) a higher-energy elliptical orbit?
I hope this is the right thread for these questions.
Thanks guys!
- Mike
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Using current technology limits, the entry vehicle cannot have a 'density' (mass/heatshield area) of more than about 150kg/m2, of else it hits the surface before it has slowed enough to deploy a parachute. So for something using a 10m PLF this would mean an entry mass of less than 12t.
Which is completely useable, if you use a little creativity ;-) (See the link in my sig.)
Hey guys, I was just looking for some good back-of-the-envelope figures to help me with my mental models.
1. So far I have been assuming 1/3.5 ratio of IMLEO to TMI (not including the mass of the spent TMI stage). For ~200 day trajectories at the reasonably opportunistic time periods, is this a reasonable figure?
2. Related to that, I have been trying to find the TMI payload for a J246. I tried to use the search but to no avail. Does anybody know it?
3. Also, what is a reasonable PMF (propellant mass fraction for those that are visiting Q&A) for a methane/LOX powered ascent vehicle to low-mars orbit, and also (3b) to an elliptical orbit suitable for rendezvous with an awaiting ERV?
4. What is a suitable PMF for a methane/LOX powered ERV for a nominal ~200 day return trajectory, from both a low-mars orbit and (4b) a higher-energy elliptical orbit?
I hope this is the right thread for these questions.
Thanks guys!
- Mike
Mike.
I just discovered this thread and after reading it all I think it IS the right thread for these questions. I had been thinking that there should be one thread dedicated to ALL things MARS. It sometimes seems things get hashed and re-hashed over on multiple threads on different forums. Would it not be better to keep it all in one place ? Anyone ?
Mick.
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1. So far I have been assuming 1/3.5 ratio of IMLEO to TMI (not including the mass of the spent TMI stage). For ~200 day trajectories at the reasonably opportunistic time periods, is this a reasonable figure?
Off-hand, I think the TMI burn is about 4km/s. Assuming an isp of 450s, that gives a basic ratio of 1:2.5. It would be reasonable (conservative?) to assume a dry stage mass of 10%, so the actual payload to IMLEO ratio becomes about 1:2.9
On some windows, and trajectories, TMI could be as little as 3.6km/s... using RL10B for 462s isp... and assume a lower dry mass stage... and you are optimistically looking at less than 1:2.5
2. Related to that, I have been trying to find the TMI payload for a J246. I tried to use the search but to no avail. Does anybody know it?
The exact figure would depend upon how fast a trajectory to Mars you wanted (e.g. minimum energy Hohmann transfer vs. a two year heliocentric free return to Earth). Plus the usual variation per launch year.
3. Also, what is a reasonable PMF (propellant mass fraction for those that are visiting Q&A) for a methane/LOX powered ascent vehicle to low-mars orbit, and also (3b) to an elliptical orbit suitable for rendezvous with an awaiting ERV?
As nobody has built one yet, cannot say definitively. I think you need the ascent vehicle to be capable of about 4.3km/s to reach low Mars orbit, and about another 1km/s got high orbit... but these are just the numbers I recall so don't take them as gospel.
4. What is a suitable PMF for a methane/LOX powered ERV for a nominal ~200 day return trajectory, from both a low-mars orbit and (4b) a higher-energy elliptical orbit?
Again, nobody has built one of these yet.
Best bet is probably to research the delta-v requirements for the two trajectories and then work out the numbers based on real life hardware. Bear in mind you may be asking your hypothetical vehicles to do more (e.g. longer storage of propellants etc) than existing stages and would need generous mass margins to cover than capability.
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Will the MAVEN spacecraft carry a relay radio system for landers as to do the current Mars orbiters?
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Will the MAVEN spacecraft carry a relay radio system for landers as to do the current Mars orbiters?
yes
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Does anyone know of any studies analyzing the minimum amount of mass per person needed to get to and from Mars (assume ISRU)?
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Reading DRM5 and similar, I see that as in Apollo only thing that returns to Earth is the crew capsule. Six or seven monstrous Ares V launches; over thousand tons IMLEO and you recover and reuse nothing. Ok the hab stays on Mars; if you want your eventual second mission to go to very same place the first one did.
How much would departure vehicle mass IMLEO grow if you wanted to
return entire vehicle to say L1?
I understand there is very extreme and very widespread hatred for anything that can be mocked with "Battlestar" tag; but is it really economic to build thousand ton vehicles that cost tens of billions and then throw away every single part of them during one mission.
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BINGO! WE HAVE A WINNER!!!
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It's simply a fact that the majority of IMLEO is fuel.. you can't get around it (yet).
As for leaving the base on Mars.. I believe the idea with fixed-habitat-and-rover designs is to land your second habitat rover-range distance from the first habitat. Then land your third habitat rover-range distance from the first two, and so on. The result is an ever increasing base of operations.
Of course, to believe in that you have to think there will actually be more than one mission :)
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Reading DRM5 and similar, I see that as in Apollo only thing that returns to Earth is the crew capsule. Six or seven monstrous Ares V launches; over thousand tons IMLEO and you recover and reuse nothing. Ok the hab stays on Mars; if you want your eventual second mission to go to very same place the first one did.
How much would departure vehicle mass IMLEO grow if you wanted to
return entire vehicle to say L1?
I understand there is very extreme and very widespread hatred for anything that can be mocked with "Battlestar" tag; but is it really economic to build thousand ton vehicles that cost tens of billions and then throw away every single part of them during one mission.
It should be possible with aerocapture, but I'm not sure if this has ever been done - though it should work.
My Mars reference plan (and I'm beyond version 5), now has SEP tugs carrying everything to L1, then depart from L1 (chemical) before aerocapture into Mars orbit. Then depart from Mars orbit and Aerocapture in Earth Orbit with rendez-vous again at L1.
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I suppose the empty Mars Transfer Vehicle (in a heliocentric orbit) could be captured by a SEP tug and brought back to EML1. You'd probably want to launch the SEP tug on a rendezvous trajectory before the crew even left the MTV (perhaps even before the crew left Mars).
For a MTV of ~100 tons and with ~4km/s to get it from its heliocentric trajectory to EML1, you'd need a SEP tug with a dry mass of ~10 tons, ~20 tons of fuel, given state-of-the-art solar and electric propulsion technology (and about half a year of thrusting with a 500kW solar array). Probably just as doable with a larger delta-v, it'd just take longer (it's going to take a while for the MTV to circle back to Earth, anyways).
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I've been thinking about Mars mission architectures recently.
If you launch a bunch of cargo landers to Mars how hard would it be to get them to land relatively close to one another (say < 5 km). So astronauts on the surface can get to them. Assume heat shield, parachute, rocket propulsion for entry, descent, and landing.
What kind of navigation system would this entail? INS, of course but what else? Something like a homing beacon (akin to ILS in aviation) maybe?
Is this an open problem? If not, I'd like to read up on what ideas people have about doing precision landings on another world with no nav aids.
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I've been thinking about Mars mission architectures recently.
If you launch a bunch of cargo landers to Mars how hard would it be to get them to land relatively close to one another (say < 5 km). So astronauts on the surface can get to them. Assume heat shield, parachute, rocket propulsion for entry, descent, and landing.
What kind of navigation system would this entail? INS, of course but what else? Something like a homing beacon (akin to ILS in aviation) maybe?
Is this an open problem? If not, I'd like to read up on what ideas people have about doing precision landings on another world with no nav aids.
Have enough delta V to place the lander in orbit before PDI instead of a direct entry fromthe transfer orbit and use a lifting reentry like the MSL will. You could also land a pathfinder payload with a homing beacon first and use that as a nav aid.
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Would the Space X Dragon be a good mars lander, and only a lander.
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I ment telephone.
Polar explorers managed without telephones until quite recently, and often spent years away from their families. Submariners manage without telephones, even today.
People on a Mars mission will have email, video messages, voice messages, and probably something like twitter and SMS. Compared with previous explorers they will live in luxury.
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... If not, I'd like to read up on what ideas people have about doing precision landings on another world with no nav aids.
I suggest landing a rover first, preferably tele-operated from orbit (Phobos?), that would survey a landing site, place several radio beacons around the site to allow precision landing, and even clear small rocks.
Give the rover a bulldozer attachment (or land a separate ROV bulldozer) and clear a nearby base area. Give the landers wheels, and the rover could tow each one to the base area. Maybe even deploy some equipment, hook up comms & power cables etc.
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How can I calculate the pressure of the Marsian wind at a given altitude? I'm wondering what sort of strain it would put on a vertical solar panel, and if it would work as a sail. I've seen that most if not all solar powered devices on Mars have used mostly horizontal solar panels. So I'm wondering if wind pressure is an issue, or is simply avoiding to actually point the panel in the Sun's direction.
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How many cycler stations in Hohmann orbits would be required to allow for reliable available at all windows(every 780 days, I think)?
Would it be reasonable to put up cycler stations on more eccentric/faster orbits, such that one would be able to facilitate Earth/Mars transit on similarly or more frequent time windows?
Thanks.
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How many cycler stations in Hohmann orbits would be required to allow for reliable available at all windows(every 780 days, I think)?
Would it be reasonable to put up cycler stations on more eccentric/faster orbits, such that one would be able to facilitate Earth/Mars transit on similarly or more frequent time windows?
Thanks.
Depends on what cycler you use. There's an Aldrin cycler that flies by earth and Mars each synodic period (roughly 2 1/7 years). But the line of apsides must be substantially rotated each orbit. And it zooms by Mars orbit at a pretty good angle, so the taxis moving between Mars and the cycler would have a steep delta V budget. So in terms of delta V, this cycler's not so good.
More Hohmann like are the Niehoff VISIT 1 and VISIT 2 cyclers. Their periods are 1.25 years and 1.5 years. (Period of an Earth Mars Hohmann ellipse is about 1.4 years). Less delta V for orbit maintenance as well as less for taxi rendezvous. Planetary fly bys more infrequent, though.
With a synodic period of about 2 1/7 years, you can see 7 synodic periods are about 15 years. But 7 periods isn't exactly 15 years, more like 14.95 years, so constant tweaking would be needed.
Venus is much more amenable to cyclers. Earth Venus synodic period is 1.5987 years, quite close to 1 3/5. 5 synodic periods is 7.993 years, very close to an 8 year cycle. What's more is an earth Venus Hohmann ellipse has a period of .7998 years, very close to 4/5 of a year. This makes for a 5 pointed star where the cyclers are traveling nearly Hohmann orbits (http://hop41.deviantart.com/art/Venus-Cycler-Movie-124141316).
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Cyclers would be a great way to supply a base. The transfer points could be EML-2 and Phobos for Mars and EML-2 and HVO for Venus.
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Cyclers would be a great way to supply a base. The transfer points could be EML-2 and Phobos for Mars and EML-2 and HVO for Venus.
Actually they are not at all suited for cargo. They don't give any delta-v for free. Getting something to the cycler and from there to the destination will require more delta-v than flying direct. It is worth it only if you save on mass which you cannot do with cargo.
What they do is give astronauts a habitat for the transfer so you save the mass of the habitat. Even that may not be enough to justify a cycler unless it has equipment for water and air recycling and maybe even food production which would save a lot of mass in supplys to launch with the astronauts.
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I am trying to understand different possible mission profiles.
Frequently missions are proposed using L-points as staging points for missions to Mars. It seemed plausible to me. But then I found out about the Oberth-Effect and found some interesting delta-v charts for missions to the Moon, to Mars, and to Venus.
So getting to any L-point requires as much delta-v as direct transfer to Mars on a Homann-Trajectory. It looks to me as if using L-points for staging missions could be worth it only if using both SEP-tugs from LEO to the L-point and lunar fuel assuming quite low cost for both. And even then the advantage would not be very big compared to launching directly from LEO.
Am I missing something here?
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I am trying to understand different possible mission profiles.
Frequently missions are proposed using L-points as staging points for missions to Mars. It seemed plausible to me. But then I found out about the Oberth-Effect and found some interesting delta-v charts for missions to the Moon, to Mars, and to Venus.
You can still use the Oberth effect together with Lagrange points. The general idea is to drop from a Lagrange point to LEO altitude (perhaps with a lunar flyby) and then do a powered flyby to Mars.
So getting to any L-point requires as much delta-v as direct transfer to Mars on a Homann-Trajectory.
Considerably less than Mars, 3.2km/s - 3.8km/s depending on how long your trip is allowed to be. In addition, since EML1/2 is close to Earth you can use it as an assembly point, which means you can launch individual modules that will fit on a Centaur or DCSS instead of needing a huge EDS. You couldn't do that in Mars orbit, because the trip to Mars orbit takes far too long for that to be practical.
It looks to me as if using L-points for staging missions could be worth it only if using both SEP-tugs from LEO to the L-point and lunar fuel assuming quite low cost for both. And even then the advantage would not be very big compared to launching directly from LEO.
SEP and ISRU would add to the utility of Lagrange points, which is large to begin with.
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You can still use the Oberth effect together with Lagrange points. The general idea is to drop from a Lagrange point to LEO altitude (perhaps with a lunar flyby) and then do a powered flyby to Mars.
That would involve two more passes of the VanAllen Belt. I don't think that would be a good idea especially for manned flights. Also it would require a lot of delta-v eating up much of the Oberth-Effect advantage. I can't really calulate that though so I cannot be positive on this.
So getting to any L-point requires as much delta-v as direct transfer to Mars on a Homann-Trajectory.
Considerably less than Mars, 3.2km/s - 3.8km/s depending on how long your trip is allowed to be. In addition, since EML1/2 is close to Earth you can use it as an assenbly point, which means you can launch individual modules that will fit on a Centaur or DCSS instead of needing a huge EDS. You couldn't do that in Mars orbit, because the trip to Mars orbit takes far too long for that to be practical.
The two charts I have seen both showed more delta-v for L-points than for a Homann-trajectory to Mars. Even a geostationary Orbit needs more.
You can assemble the modules in LEO which may be a good idea anyway even if you then lift them with SEP tugs afterwards.
Edited to correct quote nesting
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That would involve two more passes of the VanAllen Belt. I don't think that would be a good idea especially for manned flights.
That's not a problem in an MTV, which needs good radiation shielding anyway.
Also it would require a lot of delta-v eating up much of the Oberth-Effect advantage. I can't really calulate that though so I cannot be positive on this.
It wouldn't. Traveling between EML1/2 and high Mars orbit would be on the order of 2.5km/s, which is the same as to the lunar surface. This cuts down enormously on the required size of a transfer stage, as well as on the thrust and Isp requirements.
The two charts I have seen both showed more delta-v for L-points than for a Homann-trajectory to Mars. Even a geostationary Orbit needs more.
You'd need slightly more delta-v, but much smaller launch vehicles and transfer stages, and you could greatly increase effective Isp.
You can assemble the modules in LEO which may be a good idea anyway even if you then lift them with SEP tugs afterwards.
You can, but that's more difficult technologically. And even if you did, there would still be many advantages to using Lagrange point rendez-vous in addition to LEO rendez-vous.
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That would involve two more passes of the VanAllen Belt. I don't think that would be a good idea especially for manned flights.
That's not a problem in an MTV, which needs good radiation shielding anyway.
The radiation shielding cannot be that good. The radiation in the belt is a lot higher than outside.
Also it would require a lot of delta-v eating up much of the Oberth-Effect advantage. I can't really calulate that though so I cannot be positive on this.
It wouldn't. Traveling between EML1/2 and high Mars orbit would be on the order of 2.5km/s, which is the same as to the lunar surface. This cuts down enormously on the required size of a transfer stage, as well as on the thrust and Isp requirements.
How does increased requirement of acceleration translate into enormously reduced transfer stage?
But you have one point. If you can get very cheap fuel from Moon the mass from LEO becomes less. However the requirement of delta-v from L-points is at least half that from LEO so this becomes an advantage only if fuel from the moon to L-point is much cheaper than from Earth to LEO. I don't see that any time soon.
Edit: OK if you assume SEP-Tug to L-Point the transfer stage can be smaller.
You can assemble the modules in LEO which may be a good idea anyway even if you then lift them with SEP tugs afterwards.
You can, but that's more difficult technologically. And even if you did, there would still be many advantages to using Lagrange point rendez-vous in addition to LEO rendez-vous.
How is assembly in LEO more difficult than in an L-point?
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The radiation shielding cannot be that good. The radiation in the belt is a lot higher than outside.
It has to be that good, because it has to shield the crew against solar particle events. The idea is to have a storm shelter where the crew can sit things out for a couple of hours to days. It would be perfect for crossing the van Allens too.
How does increased requirement of acceleration translate into enormously reduced transfer stage?
By the rocket equation (http://en.wikipedia.org/wiki/Tsiolkovsky_rocket_equation) propellant usage is an exponential function of delta-v.
But you have one point.
I have many, not invented by me but by experts in the field.
If you can get very cheap fuel from Moon the mass from LEO becomes less. However the requirement of delta-v from L-points is at least half that from LEO so this becomes an advantage only if fuel from the moon to L-point is much cheaper than from Earth to LEO. I don't see that any time soon.
Your reasoning doesn't make sense to me.
Edit: OK if you assume SEP-Tug to L-Point the transfer stage can be smaller.
Not just if you use a SEP tug. There are two things that can significantly affect the size of the transfer stage: transporting modules individually instead of together and much lower delta-v.
How is assembly in LEO more difficult than in an L-point?
Assembly isn't more difficult in LEO, it may actually be easier, though not by much. Transporting large payloads to Mars orbit or even just L1/L2 by SEP is much more challenging than transporting individual modules by chemical propulsion. You'd need very large SEP tugs.
You seem to be trying to prove Lagrange points are a bad idea instead of trying to learn why they might be a good idea. Why is that?
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How does increased requirement of acceleration translate into enormously reduced transfer stage?
By the rocket equation (http://en.wikipedia.org/wiki/Tsiolkovsky_rocket_equation) propellant usage is an exponential function of delta-v.
So the rocket equation tells you you need less propellant for more acceleration? Somehow I don't believe that.
If you can get very cheap fuel from Moon the mass from LEO becomes less. However the requirement of delta-v from L-points is at least half that from LEO so this becomes an advantage only if fuel from the moon to L-point is much cheaper than from Earth to LEO. I don't see that any time soon.
Your reasoning doesn't make sense to me.
And yours does not to me, see the Rocket equation.
Edit: OK if you assume SEP-Tug to L-Point the transfer stage can be smaller.
Not just if you use a SEP tug. There are two things that can significantly affect the size of the transfer stage: transporting modules individually instead of together and much lower delta-v.
Again your reasoning does not make any sense to me at all. But maybe that is because we are talking about different things. I assumed you were talking about the transferstage that sends the completed stack off to Mars. I read your latest statement however as you are talking about the SEP-tugs that can be smaller if they shift the modules individually.
How is assembly in LEO more difficult than in an L-point?
Assembly isn't more difficult in LEO, it may actually be easier, though not by much.
But you said so in your last post.
Transporting large payloads to Mars orbit or even just L1/L2 by SEP is much more challenging than transporting individual modules by chemical propulsion. You'd need very large SEP tugs.
Agree, I was only talking about SEP-tug LEO L1/L2, to make that point clear. Using chemical propulsion for that part however would get you nothing. For the same or very similar delta-v you can get directly to Mars so why bother with L1/L2.
The delta-v maps I am using give me 3.6km/s from LEO to Hohmann transfer.
They give me 3.8 to L1, that is actually more.
Are they so extremely off?
You seem to be trying to prove Lagrange points are a bad idea instead of trying to learn why they might be a good idea. Why is that?
Not at all. I am trying hard to understand any advantage of L1/L2.
Maybe we should start all over from zero. We may have accumulated misunderstandings due to the nested quotes.
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So the rocket equation tells you you need less propellant for more acceleration? Somehow I don't believe that.
No it doesn't tell you that. Let me try to explain it better than I did before. The point is that if you use various staging orbits along the way and refuel there, then you can effectively reset the rocket equation at each node. So instead of having to send everything through TMI and MOI from LEO, say 5.7 km/s, you only have to push individual modules from LEO to L1/L2, say 3.2 km/s and the whole assembly from L1/L2 to a Mars Lagrange point, roughly 2.5km/s. The individual hops are now much smaller, and you need a much smaller transfer stage.
Again your reasoning does not make any sense to me at all. But maybe that is because we are talking about different things. I assumed you were talking about the transferstage that sends the completed stack off to Mars.
I was, but with a Lagrange point you only have to move the completed stack through 2.5km/s instead of the 5.7 you might need for a direct appoach.
I read your latest statement however as you are talking about the SEP-tugs that can be smaller if they shift the modules individually.
Not just the SEP tugs, the same would be true for chemical transfer stages.
But you said so in your last post.
I didn't phrase it as clearly as I might have. I meant that moving the completed stack to L1/L2 with SEP would be much more technologically challenging. It would require enormous scaling up compared to the state of the art.
Agree, I was only talking about SEP-tug LEO L1/L2, to make that point clear. Using chemical propulsion for that part however would get you nothing. For the same or very similar delta-v you can get directly to Mars so why bother with L1/L2.
Many reasons, smaller size of the individual hops, easier phasing, longer launch windows, easier reuse of the MTV, easier to increase effective Isp by using different forms of propulsion for payload / segment combinations where that makes sense, better thermal environment. By contrast, why should we care about a small difference in total delta-v, one that is easily swamped by higher effective Isp?
The delta-v maps I am using give me 3.6km/s from LEO to Hohmann transfer.
They give me 3.8 to L1, that is actually more.
Are they so extremely off?
3.8km/s is correct for a fast transfer to L1. You might want to use that for crew.
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And yours does not to me, see the Rocket equation.
The rocket equation is exponential. To illustrate exponential growth, the story of a chess wager is often used. Terms of wager: 1st square of chessboard 1 grain of rice, 2nd square 2 grains, third square 4 grains, doubling each subsequent square.
If using oxygen and hydrogen, each 3 km/s added to your delta V budget doubles initial mass.
As Martijn says, you get to start over at each propellant source. Here's a graphic of the chess wager vs a chess wager with a propellant source every 4th square:
(http://clowder.net/hop/TMI/DepotVsNoDepot.jpg)
With lots of aerobraking, round trip from LEO to Low Mars Orbit and back can be around 6 km/s.
Round trip from L2 to Low Mars Orbit and back can be around 3 km/s (using much less aerobraking than the LEO round trip).
You might say that from LEO it takes 3.5 km/s to reach L2 so total delta V is 6.5. And 6.5 is more than six.
But e3/4.4 + e3.5/4.4 < e6/4.4
If you're math-phobic here's an analogy that might help: You will save miles on a transcontinental trip if you stay on the interstate. But turning off the interstate occasionally to get gas is worthwhile even if it adds to the miles you have to drive.
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Using chemical rockets:
LEO to C3 = 3.22 km/s.
vs
LEO to EML2 = 3.47 km/s (9 day transit).
LEO to EML2 = 4.31 km/s (4 day transit).
EML2 to C3 = 0.14 km/s.
It may seem obvious that going directly to C3 is the better option than going via EML2, however SEP changes the equation. As can be seen above, if you have the option, it's better to use slow transits. That is especially the case with SEP, which have a very high specific impulse as well.
SEP allows you to build a large vehicle at EML2. You can use either SEP or chemical propulsion to do the small delta-v to C3 and onto Mars transit.
The crew will still go on fast transits, to avoid spending too much time in the Van Allen radiation belts. The 9 day vs 4 day options above are probably not all that important, but they give you the idea.
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A big save is if you can reuse the habitat for transit, for example. Plus the reliability of a very thorough checkout.
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SEP allows you to build a large vehicle at EML2.
True. So does chemical propulsion by the way.
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SEP allows you to build a large vehicle at EML2.
True. So does chemical propulsion by the way.
Yep. I think you get more bang for your buck if you can do a SEP tug.. but that said, we don't actually have one yet.
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SEP allows you to build a large vehicle at EML2.
True. So does chemical propulsion by the way.
Yep. I think you get more bang for your buck if you can do a SEP tug.. but that said, we don't actually have one yet.
Ok, thanks to all first. :)
The picture is not all clear yet to me. But at least it starts forming.
I am still not convinced that EML2 all chemical from earth is a better solution than directly from LEO. Especially for a one off flight. But as soon as other methods like SEP-tug come into the picture that would change.
About SEP-tugs, yes we don't have them. But the required technology is now there, if we have a real need for them. But to need them we would have to have some sustained beyond LEO program.
I thougt trajectories would be quite clearcut and dependant on the depth of the involved gravity wells and available delta-v only. Learning about the Oberth effect and understanding how it works made the first deep crack in that.
Those delta-v charts also make assumtions, sometimes unspoken assumptions.
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SEP allows you to build a large vehicle at EML2.
True. So does chemical propulsion by the way.
Yep. I think you get more bang for your buck if you can do a SEP tug.. but that said, we don't actually have one yet.
Propellant transfer and storage are also easier because you're handled compressed gases. And we do have them. We're just in the bad habit of putting transponders on them and leaving them in GEO.
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Using chemical rockets:
LEO to C3 = 3.22 km/s.
vs
LEO to EML2 = 3.47 km/s (9 day transit).
LEO to EML2 = 4.31 km/s (4 day transit).
EML2 to C3 = 0.14 km/s.
Someone might look at .14 km/s from EML2 to C3 and then .5 km/s from C3 to TMI and conclude a TMI burn from EML2 would be .64 km/s. Doesn't work that way.
Kinetic energy is 1/2 m * v2. If v is large, a small boost to v can make a big energy boost.
So 1/2 m * (v + h)2 can be a lot bigger than 1/2 m * v2 even if h is small.
For example lets compare a boost of velocity boost 1 for 10 vs 100:
(10 + 1)2 = 121, a boost of 21.
(100 + 1)2 = 10201, a boost of 201.
Deep in a gravity well, orbital velocity is high so a small velocity change yields big increase in energy. This is the Oberth benefit.
Someone like Guckyfan will look at orbital speed at EML2 and note 1.19 km/s is pretty slow. He correctly notes there is very little Oberth benefit at this altitude. Indeed, TMI directly from EML2 is a little more than 2 km/s.
But when orbital speed is low, it takes only a small dv to brake the orbit and plummet to the earth. For example with Farquhar's 9 day route, it takes only .33 km/s to drop from EML2 to a very low perigee:
(http://clowder.net/hop/TMI/LEO-lunar-L2.jpg)
Once at the an altitude of 185.2 km, it's moving 3.14 km faster than orbital speed at that altitude (7.79 km/s). 3.14 + 7.79 = 10.93 km/s. 10.93 km/s is a big v! When you're moving almost 11 km/s, a mere .5 km/s burn suffices for TMI.
So this route takes 3 burns: .15 km/s to drop from EML2, a .18 km/s burn at perilune and then a .5 km/s burn at perigee for TMI. This totals between .8 and .9 km/s.
A fully fueled and stocked ship departing from EML2 has a huge advantage over a fully fueled and stocked ship departing from LEO.
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There's some of your nomenclature that's confusing me. When we talk about C3, don't we talk about km²/s² of inherent energy, while when we talk about delta-v we talk about the integration of an impulse maneuver?
For example, I could change the inclination of an orbit, which would consume delta-v and have zero C3 change. Or, if I have used a supersynchronous insertion and then circularize, I'd need a lot of delta-v (and two burns) for a very small C3 change. Am I right?
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Propellant transfer and storage are also easier because you're handled compressed gases. And we do have them. We're just in the bad habit of putting transponders on them and leaving them in GEO.
It looks as if Alphabus could easily double its electrical power if you put extra solar panels on it instead of transponders. Can you confirm or deny this is the case? If so, it would mean 40kW tugs are more or less the state of the art.
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It looks as if Alphabus could easily double its electrical power if you put extra solar panels on it instead of transponders. Can you confirm or deny this is the case? If so, it would mean 40kW tugs are more or less the state of the art.
Something around 50kW is state of the art with minimal changes to existing architectures. So, that sounds pretty reasonable. There are a good handful of ~20kW HETs that are qualified, or nearly so, and a 2 thruster setup is pretty common.
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Something around 50kW is state of the art with minimal changes to existing architectures. So, that sounds pretty reasonable. There are a good handful of ~20kW HETs that are qualified, or nearly so, and a 2 thruster setup is pretty common.
This could also be useful with arcjets instead of ion propulsion. I remember reading about high power arcjets a while ago. I think it may have been ESEX. Is that a state of the art thruster, or has there been work on high power arcjets since then? You could power two on a 50kW tug.
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There's some of your nomenclature that's confusing me. When we talk about C3, don't we talk about km²/s² of inherent energy, while when we talk about delta-v we talk about the integration of an impulse maneuver?
(Googling...) Looks like C3 means characteristic energy (http://en.wikipedia.org/wiki/Characteristic_energy).
Up until now I thought it meant energy of a parabolic trajectory, in other words where C3 = 0.
A lot of people seem to use C3 as a synonym for escape or parabolic orbit. For example in a post above QuantumG says it's .14 km/s from EML2 to C3. From EML2, it takes about .14 km/s to achieve escape.
For example, I could change the inclination of an orbit, which would consume delta-v and have zero C3 change.
With my new understanding of C3, I see that you're right.
I'll reword some of my arguments hoping I'm using the term C3 correctly.
For a high altitude orbit with C3 close to zero, kinetic energy as well as potential energy are close to zero. With kinetic energy close to zero, velocity is close to zero and there is very little Oberth benefit. However with low velocity, it's less expensive to do plane changes. It seems to me a high altitude is the best place to change inclination.
Given a high altitude, low velocity orbit, it is also less expensive to change the flight path angle so this low energy orbit can have a low perigee.
Potential energy is mass * - μ/r. So shrinking r increases potential energy.
Since v2/2 - μ/r is close to zero, increasing potential energy means increasing kinetic energy. Greater kinetic energy means greater velocity.
At a low perigee, a near parabolic orbit has nearly enough velocity for Trans Mars Injection, a .5 km/s perigee burn suffices for TMI.
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The Post by Hop_David made a lot of things much clearer to me. Thanks for that.
http://forum.nasaspaceflight.com/index.php?topic=12196.msg965593#msg965593 (http://forum.nasaspaceflight.com/index.php?topic=12196.msg965593#msg965593)
In this trajectory the passes through the VanAllen Belt are quite fast so the radiation would not be too bad as well. Also I now understand how you can take advantage of the Oberth-Effect from EML2.
This would have been my next question. How exactly does such a trajectory look like. That is now answered.
BTW I am not allergic to math. :) However over many decades my math exposure has been reduced a lot.
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SEP allows you to build a large vehicle at EML2.
True. So does chemical propulsion by the way.
Yep. I think you get more bang for your buck if you can do a SEP tug.. but that said, we don't actually have one yet.
Propellant transfer and storage are also easier because you're handled compressed gases. And we do have them. We're just in the bad habit of putting transponders on them and leaving them in GEO.
+1 :D
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Looking for an answer in all the wrong places.
I need an answer in regards to the dust problems that previous rovers have encountered. I know Curiosity is powered internally without the need for solar power, but with a slim chance that her lenses might get dumped on with dust dose she have any way to deal with it?
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Looking for an answer in all the wrong places.
I need an answer in regards to the dust problems that previous rovers have encountered. I know Curiosity is powered internally without the need for solar power, but with a slim chance that her lenses might get dumped on with dust dose she have any way to deal with it?
No
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Opportunity has been on Mars fOr over eight years and still can see just fine.
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This could also be useful with arcjets instead of ion propulsion. I remember reading about high power arcjets a while ago. I think it may have been ESEX. Is that a state of the art thruster, or has there been work on high power arcjets since then? You could power two on a 50kW tug.
ESEX is state of the art. There hasn't been much interest in high-power arcjets. Most of the focus for cislunar is on Hall Effect thrusters.
Fun fact on arcjets too. All modern thrusters use an exotic alloy for the thruster chamber/nozzle that is no longer produced. The entire world supply is currently sitting in a warehouse in Redmond, WA.
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Ha, guarded by top men I bet! Top. Men.
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I've been thinking of landing cargo. There is the problem of getting multiple landers to drop in a relatively close proximity to each other. Apart from the usual "All eggs in one basket" argument , why not join all together in Mars orbit and use some serious thrust ( RD-180 style ) to land it all together ?
Hab, lab, rover/s, power, ISRU plant, supplies, tools everything together in the one selected LZ. Why not ???
Mick.
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I've been thinking of landing cargo. There is the problem of getting multiple landers to drop in a relatively close proximity to each other. Apart from the usual "All eggs in one basket" argument , why not join all together in Mars orbit and use some serious thrust ( RD-180 style ) to land it all together ?
Hab, lab, rover/s, power, ISRU plant, supplies, tools everything together in the one selected LZ. Why not ???
Mick.
meeting in orbit takes alot of DV.
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meeting in orbit takes alot of DV.
It also makes the landing problem much more serious, eg. if you're using aerodynamic landing you need a much, much larger aerobrake (which imposes all sorts of packaging problems on launch, nevermind robotically managing it in Mars orbit...) and if you're using propulsive landing...well, the mass ratios *should* be the only important thing there, so I suppose it wouldn't actually affect things that much supposing you have the same mass ratio for the same payload as independent and combined pieces.
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Ok. Say all items to be landed are combined together in one big lander before Earth departure and then sent to LMO as one unit by whatever means is most suitable at the time. Could it be propulsively landed with current chemical engines as a complete base ? If so, what engines would be best ?
Mick.
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'Best' can be interpreted differently for different purposes.
FWIW I would go with a pump-fed hypergolic engine of some sort. Arguments can be made for cryogenic propellants instead.
The real issue is mass. An entirely propulsive landing will need a fair bit of propellant- about three quarters of the total mass of the vehicle. In theory there would be no difference in mass whatsoever if you landed one big lander, or several smaller ones.
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I don't know much about engine and propellant tech so could you use, as an example, RD-180's with kero/lox brought from Earth for a one off powered landing of a single base unit containing everything needed for the ground phase of the mission ??? Once safely down the crew would land separately in the ML/MAV and occupy the Base for the duration.
Mick.
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I don't know much about engine and propellant tech so could you use, as an example, RD-180's with kero/lox brought from Earth for a one off powered landing of a single base unit containing everything needed for the ground phase of the mission ??? Once safely down the crew would land separately in the ML/MAV and occupy the Base for the duration.
Mick.
Suggest you do a little research :)
The RD-180 would be a really lousy choice.
- it is ground lit
- is is optimised for Earth sea-level operation
- the LOX would need to be kept from boiling off during months of storage
As I said, pump-fed hypergolics would be more sensible. Similar isp but much better reliability and operability.
Once you have decided what mass you want to land on Mars, look up the delta-v required and plug some numbers into http://www.strout.net/info/science/delta-v/ (http://www.strout.net/info/science/delta-v/) then see how big your mission is when it leaves Earth...
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Anyone know/can do simulations to find how much payload can be landed at - 7km MOLA? The Red Dragon is said to be able to land 1000 kg at -1.3km MOLA. So basically, I'm wondering how much payload can be landed in the Hellas Basin (Hellas Planitia).
In addition, attached is a document that may help and provide more insight into the Red Dragon Mars Mission and give specifications needed to calculate, if anyone tries/decides to.:D
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Does any one think that terraforming Mars through greenhouse effect is boring? I wonder of possibility of igniting the sleeping volcanos in Mars,
Mars is chilling, it's atmosphere is too cold, can't sustain plants that could transform it's CO2 into O2, so it needs heat, what if NASA detonate some dinamites to give the planet the needed CPR? reactivate it's volcanos ring , temperature rise, ice will melt? plants will tan and drink no? or hurl an asteroid to Mars to expose the core a little, very easy no?
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Anyone know/can do simulations to find how much payload can be landed at - 7km MOLA? The Red Dragon is said to be able to land 1000 kg at -1.3km MOLA. So basically, I'm wondering how much payload can be landed in the Hellas Basin (Hellas Planitia).
In addition, attached is a document that may help and provide more insight into the Red Dragon Mars Mission and give specifications needed to calculate, if anyone tries/decides to.:D
Looking at Figure 17 of
http://www.ssdl.gatech.edu/papers/conferencePapers/IEEE-2006-0076.pdf
and
http://www.lpi.usra.edu/meetings/marsconcepts2012/pdf/4216.pdf
I would guess that the lower altiutude would save on the order of 150 m/s of delta vee from additional drag, which should improve payload on the order of (e^(150/2800)-1) * 5180 = ~300 kg. Be wary of my guess however as IANARS.
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Anyone know/can do simulations to find how much payload can be landed at - 7km MOLA? The Red Dragon is said to be able to land 1000 kg at -1.3km MOLA. So basically, I'm wondering how much payload can be landed in the Hellas Basin (Hellas Planitia).
In addition, attached is a document that may help and provide more insight into the Red Dragon Mars Mission and give specifications needed to calculate, if anyone tries/decides to.:D
Looking at Figure 17 of
http://www.ssdl.gatech.edu/papers/conferencePapers/IEEE-2006-0076.pdf
and
http://www.lpi.usra.edu/meetings/marsconcepts2012/pdf/4216.pdf
I would guess that the lower altiutude would save on the order of 150 m/s of delta vee from additional drag, which should improve payload on the order of (e^(150/2800)-1) * 5180 = ~300 kg. Be wary of my guess however as IANARS.
Thanks a lot. I was also wondering the same for MSL if it were to land at around -7km MOLA. So what would the payload be if MSL were to land at an area that was -7km MOLA?
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Strange question really.
I assume you could increase the weight of the payload and still slow it down to around mach 2 where the parachute deploys but I speculate the parachute would not be strong enough to deal with the extra mass without a disreef.
I'm no expert but I think MSL is right on the edge of what's possible with parachutes.
For heavier payloads retropropulsion and drag surfaces like flaps on the capsule are required.
Thought I'd have a crack since the question hasn't been answered yet. If anybody has any further insight I'd welcome it.
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I don't know much about engine and propellant tech so could you use, as an example, RD-180's with kero/lox brought from Earth for a one off powered landing of a single base unit containing everything needed for the ground phase of the mission ??? Once safely down the crew would land separately in the ML/MAV and occupy the Base for the duration.
Mick.
A couple of NK43s would be a better choice if you want to use a big kerolox engine for landing.
Though a cluster of super Dracos or CECEs might be your best bet for a lander engine.
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The talk lately about capturing asteroids and carting them about the galactic neighborhood led me to a question.
Could a waylayed asteroid be realistically brought along to Mars as part of a human visit and if so, would positioning it between the sun and the crew act as shield from radiation?
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The talk lately about capturing asteroids and carting them about the galactic neighborhood led me to a question.
Could a waylayed asteroid be realistically brought along to Mars as part of a human visit and if so, would positioning it between the sun and the crew act as shield from radiation?
Maybe part of the asteroid, but maneuvering the asteroid to and from Mars would cost so much because of all the propellant, that it would be in-feasible. With regolith from the asteroid put into sort of sandbags, maybe it could be used as radiation shielding.
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Is Mars more or less spherical than the Earth?
How about in terms of gravity, does it have a "bulge" like Earth?
Or, is it the case that Mars orbiting spacecraft have to fly so high above the surface that gravitational perturbations don't impact the orbit?
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Is Mars more or less spherical than the Earth?
How about in terms of gravity, does it have a "bulge" like Earth?
Or, is it the case that Mars orbiting spacecraft have to fly so high above the surface that gravitational perturbations don't impact the orbit?
In terms of shape:
The average geometrical deviation from ideal sphere is higher for Mars;
The max geometrical deviation from ideal sphere is higher for Mars (~4x compared to Earth);
Mars is substantially less symmetrical than Earth;
In terms of gravity field:
Asymmetry of the Earth's lithosphere is partially compensated by oceans.
Asymmetry of the Mars' lithosphere is higher and there is nothing to compensate it.
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How does that asymmetry impact orbital flight of Mars probes? Is it such that spacecraft in low Mars orbits, even those beyond the tangible atmosphere, have lifetime problems, like lunar orbiting craft?
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I am looking at the Elektra radio relay system, which is a de facto standard for Mars missions, and was wondering if Elektra represents a hardware implementation around a frequency, or whether Elektra is also a digital communications protocol?
My end issue is wondering if navigational positioning data could be streamed from a Mars navsat via Elektra transmissions to surface rovers.