Design a lunar architecture around the assumption of a F9R reusable first stage, SEP tugs and propellant depots.This, SEP and depots seem to work very well together. You can refuel in LEO and again in LLO with fuel you have moved there efficiently with SEP months in advance.Im not going to propose anything like a complete architecture in the OP, mainly I was interested in the scale that was possible within these limitations. It seems to me to be extremely large. The scheme is to keep the F9R first stage as is but have a modified second stage that could be refueled in LEO and LLO, and act as an EDS and lunar lander.By my working you could land a 50+ ton payload from LLO to the lunar surface, if you could refuel something with the performance of a F9v1.1 second stage there.This is using some figures I grabbed off the internet:F9 v1.1 upper stage:empty: 6tpropellant mass: 93tburn time: 375sISP: 342slunar gravity: 1.622m/s^2LLO 1.87km/s(my guess was gravity loss = burn time * lunar gravity = 608m/s)so total necessary delta v = LLO+gravity loss = 2480m/s (roughly) ?
It's fascinating to look at what could be accomplished with a specialized second stage lofted by the F9R first stage. But I wonder: since the second stage is an integral part of the launch vehicle, is this something that only SpaceX could do? After all, SpaceX doesn't offer suborbital F9R launches....
One should avoid such a large gravity loss in regards to the Moon.Just as guess it should be about 100 m/s.LLO orbit is not 1.87km/s.Spacecraft LRO:"LRO orbit is nominally 50 km, polar, with 2 hour period. Orbital velocity is 1.6 km/sec."http://ilrs.gsfc.nasa.gov/docs/2007/lrolr_mcgarry_0709.pdf
I got the gravity loss by multiplying the published burn time for the upper stage by lunar gravity. That is the right approach right? I was just guessing there. I imagine this is an extreme case where you are landing a fully fueled upper stage one way. I guess burn time would look only half as bad if you were landing a smaller payload but reserving enough for returning to orbit. You would be splitting the burn time over descent and ascent.
Quote from: gbaikie on 07/12/2014 06:37 pmOne should avoid such a large gravity loss in regards to the Moon.Just as guess it should be about 100 m/s.LLO orbit is not 1.87km/s.Spacecraft LRO:"LRO orbit is nominally 50 km, polar, with 2 hour period. Orbital velocity is 1.6 km/sec."http://ilrs.gsfc.nasa.gov/docs/2007/lrolr_mcgarry_0709.pdfI think you are right. I found that 1.87 with a random google. Wikipedia suggested 1.6, In the solar sail thread Jim posted something that I think said 1.72 or 1.73, so I guess my number is very wrong and we can land even more.. but I dont know how to get such objects into LEO in the first place without lots of fiddling, so I will just leave it as we are not limited by what we can transfer from LLO to the lunar surface.(edit: here is where I got the 1.87 figure: http://en.wikipedia.org/wiki/Delta-v_budget )(edit: here is another wiki page suggesting 1.6 )
I got the gravity loss by multiplying the published burn time for the upper stage by lunar gravity. That is the right approach right?
Most sources are probably implicitly or explicitly covering the Apollo lunar ascents. For those, this looks reasonably authoritative:http://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/19790072468_1979072468.pdfIt asserts for Apollo 14 the theoretical minimum would have been 6045.3 fps (1842.6 m/s) over ~430 seconds of flight time.
It's not clearly wrong. Two things -1) Gravity losses are undefined in this number, and are TWR-dependent.2) "Low Lunar Orbit" is a subjective concept. Stability and margin of safety increase at higher altitudes, with the Lunar mass concentrations causing considerable orbital perturbations. It was only with the detailed gravitational mapping of the Moon, and computer-assisted discovery of 'frozen orbits' in the past decades which balanced out masscon influence, that a craft has been able to travel at anywhere near mountain-clipping altitude for long, and even then there's a thin atmosphere to deal with, akin to the Earth's atmosphere at ISS altitude.WP defines LLO as anything under 100km altitude, and notes Lunar mountain ranges reaching up to 6km. Apollo's parking orbit was 110x110, and it reduced periapsis to 15km only briefly for landing.
Making a lunar ferry and lander from an already flight-proven upper stage and using depots is a very viable strategy. Masten Space System has been doing some work on a centaur-derived lander called XEUS.I thought I'd add a couple notes on orbits:-Nodal precession makes it impossible to permanently align the orbit of a LEO depot with the orbit of the moon. So using a permanent LEO depot as a staging point for cis-lunar missions means you will have to eat a delta-v penalty for the inclination change. Since you can do it near apoapse, i.e. include it in the lunar capture burn, the delta-v cost isn't too high but doing that will limit the number of launch windows to two per month. This can be circumvented entirely by using temporary staging points where the transfer stage is being refilled directly by tankers launched from the ground.
LEO --> EML2Im choosing to send the vehicle with a full tank and no additional cargo, and arrive with fuel to spare.(we could also have sent it with the 20 tons of cargo attached and still have a bit of fuel to spare)deltaV = 3.45m0/m1 = 2.8m0 = 93(fuel)+16(2ndStage+ dragon etc) = 109m1 = 109/2.8 = 39spare fuel on arrival= 39 - 16 = 23 tons.
Um.. thanks a little? Plugging those numbers in gave me 4.0 years, but that is relying on my algebra and zero inefficiency. Isn't 40kW just the energy that goes in?
I had a go at estimating the difference of the SEP tug could make.Found these numbers (they may not be latest)SEP tugTotal: 15.5tPropellant: 11.8tISP: ~3000(I remember something about the mission taking 6 years, so Im assuming that is how long to exhaust fuel.)
Yes I was thinking about the other direction, how small can we go while not being unrealistic.
Building a human-scale station out of 3-5T modules launchable to anywhere in the Earth system via F9R is much harder, for the class of missions which don't do any assembly in LEO and rely on purely chemical propulsion. It's also more expensive at the outset to launch three F9R missions than one FH mission, without taking into account the complexity of assembly.
NASA can already do Gemini-style missions without a depot. Orion on D4H, 28-45t stage on Falcon heavy (can be small enough for a D4H, too, if hydrolox), rendezvous with a pre-launched lander at EML2.
I don't think RobotBeat is arguing against depots. You could just keep the human part nice and simple: go directly to the moon and meet your lander (which is effectively your depot) there. If you don't meet it, you still have everything to get home.For Apollo I think the entire lander (ascent + descent) was under 15 tons, and with space storable propellant.I think we could probably do a really nice mission with two FH if you exploited SEP to get your lander there ahead of time. I bet that has already been looked at in another thread somewhere.
Quote from: Nilof on 07/12/2014 08:47 amMaking a lunar ferry and lander from an already flight-proven upper stage and using depots is a very viable strategy. Masten Space System has been doing some work on a centaur-derived lander called XEUS.I thought I'd add a couple notes on orbits:-Nodal precession makes it impossible to permanently align the orbit of a LEO depot with the orbit of the moon. So using a permanent LEO depot as a staging point for cis-lunar missions means you will have to eat a delta-v penalty for the inclination change. Since you can do it near apoapse, i.e. include it in the lunar capture burn, the delta-v cost isn't too high but doing that will limit the number of launch windows to two per month. This can be circumvented entirely by using temporary staging points where the transfer stage is being refilled directly by tankers launched from the ground.If the goal is the poles, you want a transfer orbit inclined with regard to the moon's orbit. I imagine transfer orbits coplanar with the earth's equator.. No nodal precession for an equatorial LEO depot. And equatorial orbits are easier to reach from earth's surface as well.As you mention, a lunar transfer orbit's apogee is quite slow. So an ~20º direction change at apogee isn't a big hit on delta V. The big disadvantage is loss of anytime return. Lunar polar orbits would have two per month launch opportunities, as you say.
I don't think RobotBeat is arguing against depots. You could just keep the human part nice and simple: go directly to the moon and meet your lander (which is effectively your depot) there. If you don't meet it, you still have everything to get home.
For Apollo I think the entire lander (ascent + descent) was under 15 tons, and with space storable propellant.
"Start in an equatorial low earth orbit at 300 kilometer altitude."Hmm, that's sort of a "spherical cow in vacuum" sort of thing. How about from an orbit that's reachable from KSC (or Boca Chica :-) )? Cheers, Martin
Quote from: MP99 on 07/19/2014 05:48 pm"Start in an equatorial low earth orbit at 300 kilometer altitude."Hmm, that's sort of a "spherical cow in vacuum" sort of thing. How about from an orbit that's reachable from KSC (or Boca Chica :-) )? Cheers, MartinANY orbit is reachable from ANY launch site with enough doglegging and plane change maneuvers. I think what you meant was "practically" reachable...