Quote from: mmeijeri on 08/08/2009 11:12 amAt Mars, instead of rendez-vousing with Deimos, you could also brake into a much higher orbit and end up at Sun Mars L1/L2. This should be a lot cheaper. I need to update my spreadsheet for this, but in the mean-time I'd be really interested in what numbers David comes up with.Earth departure remains the same: .7 to drop from EML1 and a .5 burn at perigee for TMI for about 1.2Putting Mars apogee at 1.08 million kilometers does make the Hohmann exit burn smaller: .6888 Mars perigee burn to park the ship in an orbit with that high apoapsis.My spread sheet givesCircle V at apoapsis .1990Ellipse V at apoapsis .0146Apoapsis circulize burn .1825But this is wrong since my spreadsheet is old school 2-body patched conics and doesn't consider the sun's influence on this Mars orbit. The velocity of Sun-Mars L1 wrt Mars isn't .1990 but 0 km/sec.So I would guess the "circulize burn" at Mars apoapsis would be .0146.Totalling all these, my guess would be 1.84 km/sec for EML1 to SunMarsL1
At Mars, instead of rendez-vousing with Deimos, you could also brake into a much higher orbit and end up at Sun Mars L1/L2. This should be a lot cheaper. I need to update my spreadsheet for this, but in the mean-time I'd be really interested in what numbers David comes up with.
Quote from: Hop_David on 08/08/2009 05:33 pmQuote from: mmeijeri on 08/08/2009 11:12 amAt Mars, instead of rendez-vousing with Deimos, you could also brake into a much higher orbit and end up at Sun Mars L1/L2. This should be a lot cheaper. I need to update my spreadsheet for this, but in the mean-time I'd be really interested in what numbers David comes up with.Earth departure remains the same: .7 to drop from EML1 and a .5 burn at perigee for TMI for about 1.2Putting Mars apogee at 1.08 million kilometers does make the Hohmann exit burn smaller: .6888 Mars perigee burn to park the ship in an orbit with that high apoapsis.My spread sheet givesCircle V at apoapsis .1990Ellipse V at apoapsis .0146Apoapsis circulize burn .1825But this is wrong since my spreadsheet is old school 2-body patched conics and doesn't consider the sun's influence on this Mars orbit. The velocity of Sun-Mars L1 wrt Mars isn't .1990 but 0 km/sec.So I would guess the "circulize burn" at Mars apoapsis would be .0146.Totalling all these, my guess would be 1.84 km/sec for EML1 to SunMarsL1Thread necromancy... Is the 1.84 km/s one way, thus 3.68 km/s if return included ? (EML-1 > SML-1 > EML-1)
A fuel depot at L1 has no real advantage over LEO except enabling you to use different launch configurations.... unless, you have electric propulsion to transport fuel from LEO to L1.Consider a 1 MW VASIMR tug, weighing 10 tons, which uses 10 tons of Argon to lift a 60 ton payload to L1 over the course of 12 months. Two such tugs deliver 120 tons per year. If most of this is water, it could be electrolysed "on demand" for a variety of missions.
A fuel depot at L1 has no real advantage over LEO
So here is an off the wall thought: An EML-1 / 2 Gateway station may be the first practical location in the solar system for the use of a space tether all the way to a planetary surface. I don't know how practical the capital investment / material science is, but if you could drop a tether from EML-1 to the lunar surface you could achieve a very practical means of potentially getting a lot of mass into an orbital location at very little cost (solar power arrays). Does anyone know if this has been studied at all?Thanks, - OsaEdit: Answered my own question via wikipedia:http://en.wikipedia.org/wiki/Lunar_space_elevator
<snip>See papers linked belowhttp://ccar.colorado.edu/nag/papers/AAS%2006-132.pdf
Before we go much further, let's get one thing straight:It is VASIMR, not VASIMIR.Quote from: iontyre on 08/07/2009 12:32 pmBefore we go much further, let's get one thing straight:It is VASIMR, not VASIMIR.Variable Specific Impulse Magnetoplasma RocketI think they also prefer the case sensitive nomenclature: VaSIMR.Variable Specific Impulse Magnetoplasma Rocket
Before we go much further, let's get one thing straight:It is VASIMR, not VASIMIR.Variable Specific Impulse Magnetoplasma Rocket
Your satellite could "orbiting" the libration point at a distance of ~50,000 km.
@robotjunkieIt is my non-technical understanding that attaining a low lunar orbit (LLO) is "easy" when starting from EML-1 or EML-2. Thereafter, a surface landing would occur in the same fashion as if LLO were achieved direct from LEO.Have you attempted to model EML to LLO?
Have you tried using:http://astrojava.com/lunar-trajectory-simulation?Edit to add:This simulator can be a bit frustrating at first, but it shows:If you depart EML-1 with a delta-v of 125 m/s in a direction coplanar with the orbit of the Moon and at an angle of 63.2 degrees (where zero degrees is along the line from the Earth to the Moon) you will reach a perilune of 110.1 km 68.5 hours later. When there you can insert into an essentially circular LLO with -630 m/s of delta-v.Sadly that orbit is equatorial. I can't coerce the simulator to show non-coplanar EML-1 departures, but intuition says there would be a non-coplanar departure with similar delta-v that led to a polar LLO.
I like the conclusions of the paper; using halo orbits for constellations of com sats, placed in orbit with low energy trajectories. They could also be power sats for the ISRU facilities as well.