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#120
by
butters
on 16 May, 2009 19:57
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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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#121
by
Kaputnik
on 16 May, 2009 20:05
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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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#122
by
A_M_Swallow
on 17 May, 2009 02:41
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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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#123
by
Kaputnik
on 17 May, 2009 10:18
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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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#124
by
gospacex
on 17 May, 2009 11:17
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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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#125
by
clongton
on 17 May, 2009 13:35
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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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#126
by
Kaputnik
on 19 May, 2009 21:47
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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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#127
by
Archibald
on 20 May, 2009 10:10
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#128
by
Kaputnik
on 20 May, 2009 12:10
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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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#129
by
Archibald
on 20 May, 2009 15:42
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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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#130
by
mmeijeri
on 20 May, 2009 15:52
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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 SpaceIt mentions the possibility of using Sun-Mars L1, and recommends further investigation of that possibility.
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#131
by
Harlan
on 20 May, 2009 17:38
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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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#132
by
Archibald
on 20 May, 2009 19:15
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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
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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#133
by
mmeijeri
on 20 May, 2009 19:33
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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.
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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#134
by
Archibald
on 20 May, 2009 19:43
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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.htmlMore 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=enhttp://ares.jsc.nasa.gov/HumanExplore/Exploration/EXLibrary/DOCS/EIC042.HTML (scroll down to the bottom of the page)
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#135
by
Kaputnik
on 20 May, 2009 21:20
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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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#136
by
mmeijeri
on 21 May, 2009 02:03
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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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#137
by
Kaputnik
on 21 May, 2009 09:14
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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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#138
by
mmeijeri
on 23 May, 2009 19:31
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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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#139
by
mmeijeri
on 23 May, 2009 19:49
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This works out to about 6 km/s, using the numbers found on this NASA
Mars Fact Sheet.
)