Author Topic: Just F9R, SEP tugs and propellant depots.  (Read 16160 times)

Offline KelvinZero

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Just F9R, SEP tugs and propellant depots.
« on: 07/11/2014 10:49 am »
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: 6t
propellant mass: 93t
burn time: 375s
ISP: 342s
lunar gravity: 1.622m/s^2
LLO 1.87km/s

(my guess was gravity loss = burn time * lunar gravity = 608m/s)
(edit: apparently a huge overestimate, I guess 200m/s in later examples)

so total necessary delta v = LLO+gravity loss = 2480m/s

using deltaV = ISP*g(earth)*ln(m0/m1) I get:

m0/m1 = 2.09

Anyway, assuming that m0 - m1 = 93t (the propellant) and that the 2nd stage mass is 6t I get a value of 57 tons 78 tons of useful payload.

(edit: corrected a basic algebra error above I think)

Given the above it seems totally reasonable to pursue some sort of lunar architecture with just what you can put on top of a F9R, maybe using a F9H for some large indivisible 50 ton cargo that you have to get up to LEO.

------------- Below this line may be edited frequently: -------------



Here is my first thought of how this lander could be layed out. Other people may have better ideas:

(1) The basic lander is shaped like a F9 upper stage, with 10 tons (potentially 11) for a fairly conventional Dragon v2 on top, lunar landing legs etc. (I hope that is sufficient)

(2) This is rather tall and thin, but it has oversized landing gear creating a bit of a pyramid shape when extended.

(3) The oversized landing gear is justified because in general, in addition to the crew and passengers, it can carry a lot of cargo one way. This cargo is attached to the landing legs around the base of the rocket in either 2 or 4 modules, so they are right at the lunar surface when landed. This cargo could be attached at the LEO depot if perishable, or at the LLO/ EML2 depot to exploit slower, more efficient transport for cargo.

(possibly the landing gear stays behind when the vehicle ascends from the moon)



Here is my attempt at a more rounded architecture, ie including crew and cargo and not one way.

Working backwards:

Returning home (moon surface --> earth surface)
I am using direct return from the lunar surface to the earth's surface. What returns is a fairly standard Dragon v2 and the modified second stage.
Im using 6 tons for the second stage, and 10 tons for the dragon v2, which I hope is a healthy over estimate including draco fuel and a bit of returned cargo etc.
So 16 tons are returned to earth,
Im using 2.8 for the deltaV, and 342 for ISP.
I calculate 20.8 tons for the fuel.

(NOTE: no fuel for propulsive landing of upper stage on earth budgeted for. Potentially, being unmanned, it can aerobrake into LEO and be refueled in order to land and be refurbished. The Dragon returns directly to earth.)

EML2 --> Moon surface
So the lander must have that much fuel once it reaches the lunar surface. Im assuming full refueling at EML2, implying 93-20.8 = 72 tons of fuel can be used in landing.
Im using deltaV from EML2 to Moon surface = 2520m/s + 200m/s gravity loss = 2720m/s
I get 57.6 tons landed, including the mass of the stage (6t), the dragon etc (10t), the fuel for return (20.8t).



Conclusion: 20+ tons of one-way cargo can be sent from EML2 to the lunar surface, in addition to what ever cargo the dragon carries and returns.

LEO --> EML2
Im 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 = 3.45
m0/m1 = 2.8
m0 = 93(fuel)+16(2ndStage+ dragon etc) = 109
m1 = 109/2.8 = 39
spare fuel on arrival= 39 - 16 = 23 tons.

(Oops!  probably needed deltaV=4.0 to keep travel time to 5 days, reducing that spare fuel to 17t. This doesn't apply to the unmanned portion, which could perhaps have been even a bit lower.)

Some missions:
(A) All Chemical, A dragonfull of crew and 20 tons cargo to the lunar surface
= 414 IMLEO + lander launch

(B) All Chemical, Dropping the 20 tons cargo, just sending dragonfull of crew.
= 257t IMLEO + lander launch

(C) Same as (A) but exploiting the ARM SEP tug for unmanned portions
= 217t IMLEO + lander launch

(D) Magical Asteroid ISRU fairy
= 20t IMLEO (thats the lunar cargo) + Lander launch.
« Last Edit: 07/15/2014 05:23 am by KelvinZero »

Offline Nilof

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Re: Just F9R, SEP tugs and propellant depots.
« Reply #1 on: 07/12/2014 08:47 am »
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.

- Most low lunar orbit are not stable due to lunar mass concentrations, and will decay very quickly. There are exceptions though, with four "frozen" low lunar orbits at inclinations 27º, 50º, 76º, and 86º which are stable. Unless you went for the 86º one and you are close to the poles, you will have two launch windows per month. You can solve the problem of infrequent launch windows by using L1 or L2 as staging points, at the cost of increased delta-v distance between the lunar surface and the depot. So another tradeoff.
For a variable Isp spacecraft running at constant power and constant acceleration, the mass ratio is linear in delta-v.   Δv = ve0(MR-1). Or equivalently: Δv = vef PMF. Also, this is energy-optimal for a fixed delta-v and mass ratio.

Offline KelvinZero

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Re: Just F9R, SEP tugs and propellant depots.
« Reply #2 on: 07/12/2014 09:59 am »
Yeah my sole reason for using LLO was to get the most impressive sized single object down to the lunar surface, 57 tons, slightly more than you could get to orbit with a FH anyway. There are all sorts of other reasons to avoid aiming for this number. But it is nice to know what chunks you can handle without a bigger launcher.

Hey this is embarrassing.. I think I might have messed up with some basic algebra at the end. Is it more like 78 tons?

Given m0/m1 = R and..
m0 = Mc+Mp, (mass landed, mass of propellant) and..
m1 = Mc, (just mass landed)
then Mc = Mp/(R-1) = 93/(2.1-1) = 84 tons landed
..giving 78 tons of cargo, assuming 6 tons for the F9v1.1 upper stage.

Someone really has to check my math! Not just this bit either.

Offline sdsds

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Re: Just F9R, SEP tugs and propellant depots.
« Reply #3 on: 07/12/2014 03:32 pm »
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....
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Offline gbaikie

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Re: Just F9R, SEP tugs and propellant depots.
« Reply #4 on: 07/12/2014 06:37 pm »
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: 6t
propellant mass: 93t
burn time: 375s
ISP: 342s
lunar gravity: 1.622m/s^2
LLO 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) ?

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 suppose one could have perigee velocity at 1.87km/s if it's highly elliptical.
In Apollo the delta-v used to go from LLO to surface was about 1.8 km [that includes gravity loss].

Generally to have lowest gravity loss leaving or landing on the Moon, the zero to 1 km/sec velocity must adequate acceleration. Or with leaving earth the zero to say 4 km/sec must have adequate acceleration.
Or this in the range where rockets generally have most the their gravity loss.
Or if landing on the Moon, if lands with low and constant thrust one will have excessive gravity loss.
So generally, at LLO, you do burn that puts the perigee intersecting surface, and as get near to crashing, one applies thrust so as to avoid hitting the surface a high velocity.  So ideally or excessive in terms of lowering gravity loss [which isn't a big deal] one have very high thrust- say 20 meter/sec/sec.
So say 1600 m/s divided by 20 is 80 seconds and after the 80 second you end up less than 1000 meter above the surface, and this in less than 1000 meter above the surface one makes up the remaining velocity one must lose.
So that is path have less than 100 m/s of gravity loss- but it would be fairly scary the more one wanted to reduce gravity loss. And leaving the Moon means one has start at least 5 m/s/s acceleration. I believe with Apollo was around 1 gee acceleration [9.8 m/s/s].
Or if you only accelerate at 5 or less meter per second per second, one will have a significant amount of gravity loss.
Now what you doing is using a second stage, and thrust needed on any second stage does not need to be high, to prevent gravity loss- one does have much gravity loss leaving earth after one gotten to around 4 km/sec. Plus you putting a large payload on it.
So one could modify it, by increasing engines- say 3 engines instead of one.


« Last Edit: 07/12/2014 06:52 pm by gbaikie »

Offline Burninate

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Re: Just F9R, SEP tugs and propellant depots.
« Reply #5 on: 07/12/2014 08:12 pm »
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....
A second stage with no payload would possess 0.9 TWR.  This is prohibitive for, say, launching straight up from the pad.  The second stage's thrust would be outweighed by gravity for a while, though I think it would be able to get to positive TWR before it actually reached apex, after it's drained some fraction of the tank.

However - Lop the second stage off and attach a payload directly to the first stage*, running it until it's dry, and suborbital flights work just fine for extremely large payloads.

This is a configuration that *might* make some sense for very heavy F9H flights to LEO.  If the second stage dry mass is not so much of an issue due to the enormous bulk of the payload, and thrust is scarce, then the central core could act as the stage that gets the payload to orbit.

*Obviously this would require some amount of additional engineering.

EDIT: An SLS-class payload, 70T, would get 7658m/s out of this arrangement at 300s Isp.  A 60T payload, 7986.  A 50T payload, 8361m/s.  This is close enough that if the payload already has a sizable engine for other purposes (EDS?), it could provide the last kick without requiring a high TWR.

EDIT2: You could also simply fly the second stage half-full.
« Last Edit: 07/14/2014 07:19 am by Burninate »

Offline KelvinZero

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Re: Just F9R, SEP tugs and propellant depots.
« Reply #6 on: 07/12/2014 11:31 pm »
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....
Hi sdsds, I guess what is meant is we have already looked at shoestring approaches with modified dragon etc. I was interested in what we could do with a properly designed architecture BUT limited to what we can fit on top of a F9R. It is not so much about getting to the moon faster but imagining a world where the F9R first stage is a proven article, along with SEP and depots. So we are not limited to the upper stage in any way. I feel safer sticking to the performance of the upper stage since it obviously does fit on top of the first stage and I can find numbers.

It looks like in principle you can land extremely large objects, assuming one way and unmanned. I imagine assembling entire 70 ton lunar outposts in LEO, perhaps a donut shape around the base of the upper stage, then tugging the whole thing to lunar obit, refueling and depositing in one go.

A more reasonable goal would probably be to consider modules we can launch in one go. Since the upper stage may never return to earth, or could be refueled in orbit by reusable upperstages, I think I heard somewhere this could be only 15% smaller than current F9v1.1 payloads, so around 11 tons? That could pretty much all be payload so it can far exceed the apollo ascent vehicle that was around 4.5 tons including ascent rockets and fuel.

Offline Alf Fass

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Re: Just F9R, SEP tugs and propellant depots.
« Reply #7 on: 07/12/2014 11:58 pm »
That's a thumbs up from me.
When my information changes, I alter my conclusions. What do you do, sir?
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Offline KelvinZero

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Re: Just F9R, SEP tugs and propellant depots.
« Reply #8 on: 07/13/2014 12:33 am »

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 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? 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.
« Last Edit: 07/13/2014 01:27 am by KelvinZero »

Offline Alf Fass

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Re: Just F9R, SEP tugs and propellant depots.
« Reply #9 on: 07/13/2014 01:53 am »

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.

Far too pessimistic, gravity loss leaving Earth is usually less that 1km/s, even though burn times around 500 seconds or more.

The sooner you can switch away from vertical accent the lower gravity loss. Cant say for sure but I'd expect loss of much less than 200m/s
When my information changes, I alter my conclusions. What do you do, sir?
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Offline Nilof

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Re: Just F9R, SEP tugs and propellant depots.
« Reply #10 on: 07/13/2014 02:31 am »
Yes, since the moon has no atmosphere you can skip the vertical ascent phase entirely. You just need enough altitude to turn the lander and at that point a lander capable of 1g acceleration will only need to be pointed 9.5 degrees (arcsin(1/6) ) above the horizon to cancel out gravity. Cos(9.5 deg) = 0.986 so you would only need lose 1.5% of your delta-v to gravity drag at most, and that is an upper estimate since you can point your spacecraft closer to the horizon as you gain speed.
For a variable Isp spacecraft running at constant power and constant acceleration, the mass ratio is linear in delta-v.   Δv = ve0(MR-1). Or equivalently: Δv = vef PMF. Also, this is energy-optimal for a fixed delta-v and mass ratio.

Offline KelvinZero

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Re: Just F9R, SEP tugs and propellant depots.
« Reply #11 on: 07/13/2014 03:28 am »
Haha.. with my bad algebra (57 should have been 78) high lunar gravity (1.87 could be as low as 1.6) and gravity loss overestimate we are probably getting closer to a round 100 tons one way. Two Falcon heavy launches of just payload.

Roll on the day we are landing two 50-ish ton diggers with each mission, but not today.

Here is my first thought of how this lander could be layed out. Other people may have better ideas:

(1) The basic lander is shaped like a F9 upper stage, with up to 11 tons for crew and two-way cargo on top. Possibly this is in the form of a Dragon and trunk, or otherwise crew could board only when it reaches the LEO depot.

(2) This is rather tall and thin, but it has oversized landing gear creating a bit of a pyramid shape when extended.

(3) The oversized landing gear is justified because in general, in addition to the crew and passengers, it can carry a lot of cargo one way. This cargo is attached to the landing legs around the base of the rocket in either 2 or 4 modules, so they are right at the lunar surface when landed. This cargo could be attached at the LEO depot if perishable, or at the LLO/ EML2 depot to exploit slower, more efficient transport for cargo.

(possibly the landing gear stays behind when the vehicle ascends from the moon)

Offline gbaikie

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Re: Just F9R, SEP tugs and propellant depots.
« Reply #12 on: 07/13/2014 03:44 am »

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 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 would say, it can't be much less than 1.6 km. 
Or 1.6 km/sec is low end of that which is possible.
The wiki number of 1.87 is simply wrong- or I don't know how or where they are getting it- as wild guess maybe it from a lunar trajectory from Earth and including some braking. Or I mean, a sort of powered capture type orbit which would have is lowest point at low lunar orbit distance which then immediately descent to surface [not done before, as far as I know].

Here they are suggesting it's about 1.6 km/sec:
http://forum.nasaspaceflight.com/index.php?topic=29195.0
Here's a more detailed analysis, indicating an aspect of landing one moon require a period of time to allow pilots to see the terrain they going to land on:
http://web.mit.edu/digitalapollo/Documents/Chapter8/lunarlandingsymposium.pdf
E.g:
Summary of Braking Phase -
"The braking phase, lasting about 450 seconds, covers
some 243 nautical miles during which the velocity is reduced from 5500 ftlsec.
to approximately 600 ft/sec., and the altitude from 50,000 feet to about 9,000 feet.
The attitude during the phase is normally such that the thrust vector is close to being aligned
with the flight path angle. In this attitude, the pilot is not able to look in the direction of the
intended landing area. "

So in meters:  5500 ft is 1676.4 meter and 4900 feet is 1493.52 meters per second.
So from orbit, 1493.52 m/s delta-v gets to about 3 km above surface and going at velocity 600 ft/sec [182.88 m/s]. So that's so pilots can see where they are going.

I also recalled they changed the trajectory somewhere around Apollo 14 so it was more efficient [or at least, better in some way].
And if one going to location one had previously gone to before and had things like landing beacons, one could use less delta-v [one would know exactly where one was going to and not need to look where to land].
Quote
I got the gravity loss by multiplying the published burn time for the upper stage by lunar gravity. That is the right approach right?
Well, it gives the most it could be, but basically, no.
Though it does sort of indicate the problem with using Falcon upper stage with one engine.

Now, perhaps, if you wanted to use falcon upper stage with one engine, one could probably use as a stage. It could use low thrust to get about 1/2 way to the surface, and with the final stage having higher thrust to weight ratio. So final stage only needed about 1 km/sec of delta-v. And falcon-9 handling the .6 km + delta-v needed, and then falcon 9 upper stage separates and returns to be re-fuel at LLO.

So with it's total burn time of 375 second, one use around 300 second to descent and 75 second to return to orbit- just as SWAG. And/or have shorter tank or more rockets.
Though it could be too complicated.

Offline KelvinZero

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Re: Just F9R, SEP tugs and propellant depots.
« Reply #13 on: 07/13/2014 03:51 am »
One question is if the the upper stage is sufficiently throttle-able. I could imagine some super dracos being exploited at the last moment, even what is essentially a dragon v2 (possibly minus heat shield) on top. This might provide additional thrust at end plus interesting abort scenarios if the main engine fails.

---

Ok here is my attempt at a more rounded architecture, ie including crew and cargo and not one way.

Working backwards:

I am using direct return from the lunar surface to the earth's surface. What returns is a fairly standard Dragon v2 and the modified second stage.
Im using 6 tons for the second stage, and 10 tons for the dragon v2, which I hope is a healthy over estimate including draco fuel and a bit of returned cargo etc.
So 16 tons are returned to earth,
Im using 2.8 for the deltaV, and 342 for ISP.
I calculate 20.8 tons for the fuel.

So the lander must have that much fuel once it reaches the lunar surface. Im assuming full refueling at EML2, implying 93-20.8 = 72 tons of fuel can be used in landing.
Im using deltaV from EML2 to Moon surface = 2520m/s + 200m/s gravity loss = 2720m/s
I get 57.6 tons landed, including the mass of the stage (6t), the dragon etc (10t), the fuel for return (20.8t).

Conclusion: 20+ tons of one-way cargo can be sent from EML2 to the lunar surface, in addition to what ever cargo the dragon carries and returns.

(rats, just noticed an error: I did not include fuel for the 2nd stage to propulsively land on earth. It is possible it refuels in orbit I suppose. It is also likely that recovering the 2nd stage from the moon just won't be attempted initially)
« Last Edit: 07/13/2014 11:34 am by KelvinZero »

Offline Burninate

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Re: Just F9R, SEP tugs and propellant depots.
« Reply #14 on: 07/13/2014 06:53 pm »
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.

Offline KelvinZero

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Re: Just F9R, SEP tugs and propellant depots.
« Reply #15 on: 07/13/2014 08:41 pm »
It probably came from Apollo. I decided to go with EML2 to be conservative. In any case I wouldn't put my depot in the lowest possible orbit. That would probably only be relevant if I had another stage that handled to and from the moon to squeeze the largest possible payload size out of the actual lander.

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.pdf

It asserts for Apollo 14 the theoretical minimum would have been 6045.3 fps (1842.6 m/s) over ~430 seconds of flight time.

gbaikie also had a similar number for Apollo on this thread.
« Last Edit: 07/13/2014 08:45 pm by KelvinZero »

Offline gbaikie

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Re: Just F9R, SEP tugs and propellant depots.
« Reply #16 on: 07/14/2014 03:26 am »
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.

If was instead lunar orbit rather than lunar low orbit, 1.8 km/sec would be fine. But saying LLO and 1.87 km/sec is suggesting a certain degree of accuracy.
1.8 km/sec would be more accurate than 1.87 km/sec- or in other words, if one carries something to lower decimal point one is implying a degree of precision.

Offline jongoff

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Re: Just F9R, SEP tugs and propellant depots.
« Reply #17 on: 07/14/2014 05:03 am »
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.

I'm not sure if this is correct. My understanding was that with most LEO depot options you get windows on average every 6-9 days, which is more like 3-5x per month. Even twice a month is plenty frequent for the near term... If it ever cramps your style too much, you can add another depot in the same inclination but with its RAAN offset by say 180 degrees.

~Jon

Offline KelvinZero

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Re: Just F9R, SEP tugs and propellant depots.
« Reply #18 on: 07/14/2014 05:19 am »
I have a concern that my delta-v values for EML2 etc may assume unreasonably long travel times. Should I use different numbers for reasonable travel times? (This might also become a reason to use lunar orbit possibly)

I would also like some numbers to use for unmanned cargo. I know there are slower, more efficient trajectories to the moon.. were they perhaps about 25% cheaper? Im also really hazy. Im Hazy if SEP would help, or what numbers to use if SEP is assumed.

Currently we are looking at getting 93t of propellant, 20 tons of cargo, and the vehicle+crew (16t) to EML2, the last being sent without exploiting low energy trajectories or SEP.

Offline KelvinZero

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Re: Just F9R, SEP tugs and propellant depots.
« Reply #19 on: 07/14/2014 06:35 am »
I decided to paste the evolving plan into the OP.

Just added this step:

LEO --> EML2
Im 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 = 3.45
m0/m1 = 2.8
m0 = 93(fuel)+16(2ndStage+ dragon etc) = 109
m1 = 109/2.8 = 39
spare fuel on arrival= 39 - 16 = 23 tons.

Assuming we send propellant by sending a full upper stage and delivering the remaining fuel,
100t in LEO --> 28t at EML2, and assuming the same ratio for cargo

So, so far we need
* 93 tons of fuel in LEO,
* 70 tons of fuel delivered to EML2 --> 250t in LEO
* 20 tons of optional lunar cargo delivered to EML2 -->71t in LEO

so thats about 414 tons in leo, or about 42 F9R launches (plus the lander+crew launch)
to deliver a dragonfull of crew and 20 tons cargo to the lunar surface, all elements potentially reusable.

(edit: dropping the 20 tons of cargo, that 70t becomes 46t and total IMLEO becomes 257t or 26 F9R launches, (plus the lander+crew launch))

The depots are pretty much equivalent to the upper stages so I think they are already included. In fact, since everything is chemical you probably dont need EML2 and could launch both the full lunar vehicle and the "depot" at once.. Or you could investigate SEP and possibly bring IMLEO down.

rats: should probably rework this with gbaikie's numbers:
deltaV of 4km/s for fast (5 day) to EML2
deltaV of 3.2 for slow (180 day) to EML2.
« Last Edit: 07/14/2014 10:57 am by KelvinZero »

Offline gbaikie

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Re: Just F9R, SEP tugs and propellant depots.
« Reply #20 on: 07/14/2014 06:51 am »
LEO to L-2:
http://www.nasaspaceflight.com/_docs/NASA-TN-D-6365.pdf

From post:
"Hollister David says:   
October 2, 2012 at 10:28 am   

I may have misremembered the 3.2 km/s 8 day trip. I did find this paper by Farquhar http://www.nasaspaceflight.com/_docs/NASA-TN-D-6365.pdf where he describes an 8.8 day trip to EML2 taking 3.47 km/s (page 27)...."
http://www.spudislunarresources.com/blog/gateway-l-2-mission-opening-cislunar-space-or-dead-end/
So, from PDF:
Pg: 23:
"A typical two-impulse trajectory between the earth parking orbit and L, is depicted
in Figure 20. The variation in 4 V as a function of the transfer time for this mode is
given in Figure 21. Notice that the AV curve for the impulse at L, is starting to flatten
out at t = 108 hours. Thus, it appears that only very small savings can be realized by
extending the transfer time beyond 108 hours."
And:
Pg 24:
"Fortunately, it is possible to reduce the AV costs below those for the two-impulse
transfer by using a powered lunar swingby (Reference 3). This three-impulse transfer
mode is illustrated in Figure 22. The total transfer time for this mode is almost 9
days, but the reduction in A V is appreciable."

Hmm, I thought I recall something like 9 days and about 3.8 km to L-2 from LEO. But anyhow,
here is another ref with some longer travel time [and it seems they advocating low thrust as most efficient]:
http://ccar.colorado.edu/asen5050/projects/projects_2012/wolma/
« Last Edit: 07/14/2014 06:54 am by gbaikie »

Offline KelvinZero

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Re: Just F9R, SEP tugs and propellant depots.
« Reply #21 on: 07/14/2014 01:20 pm »
I had a go at estimating the difference of the SEP tug could make.

Found these numbers (they may not be latest)
SEP tug
Total: 15.5t
Propellant: 11.8t
ISP: ~3000
(I remember something about the mission taking 6 years, so Im assuming that is how long to exhaust fuel.)

Low thrust deltaV to EML2.. dunno. Im using 7km/s (ie EML1 value)

m0/m1 = e^(deltav/ISP*g) = 1.27

From which I deduce the SEP tug could move 40 tons of cargo to EML2 (taking 6 years)
55 tons delivered to LEO becomes 40 tons at EML2. Two SEP tugs would deliver slightly less than desired.
Im just going to assume something with proportional performance.

93 tons at LEO ( fuel for tanking up lander in LEO )
90 tons at EML2 --> 124 at LEO (20 lunar cargo + 70 refuel at EML2)
IMLEO is now 217 tons, or about 22 F9R, plus the lander+crew launch.

Using something very similar to two ARRM SEP tugs almost halves IMLEO but takes 6 years to set up. That timeframe could be acceptable but I haven't considered weight of active cooling to solve boiloff.
« Last Edit: 07/14/2014 01:24 pm by KelvinZero »

Offline A_M_Swallow

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Re: Just F9R, SEP tugs and propellant depots.
« Reply #22 on: 07/14/2014 01:41 pm »
To calculate the time you need either the force (thrust) or the energy used by the SEP.  The ARM SEP is aiming for 40 kW.

Offline KelvinZero

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Re: Just F9R, SEP tugs and propellant depots.
« Reply #23 on: 07/14/2014 10:39 pm »
Um.. thanks a little?  8)

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?

Offline A_M_Swallow

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Re: Just F9R, SEP tugs and propellant depots.
« Reply #24 on: 07/15/2014 08:24 am »
Um.. thanks a little?  8)

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?

Yes.  Ion thrusters come with an energy efficiency.  I do not know what it is for the chosen thruster.

Offline gbaikie

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Re: Just F9R, SEP tugs and propellant depots.
« Reply #25 on: 07/15/2014 09:59 am »
I had a go at estimating the difference of the SEP tug could make.

Found these numbers (they may not be latest)
SEP tug
Total: 15.5t
Propellant: 11.8t
ISP: ~3000
(I remember something about the mission taking 6 years, so Im assuming that is how long to exhaust fuel.)
"The NEXT ion thruster is one of NASA’s latest generation of engines. With a power output of seven kilowatts, it’s over twice as powerful as the ones used aboard the unmanned Dawn space probe. Yet it is simpler in design, lighter and more efficient, and is also designed for very high endurance.

Its current record of 43,000 hours is the equivalent of nearly five years of continuous operation while consuming only 770 kg (1697.5 lbs) of xenon propellant. "
http://www.gizmag.com/next-ion-record/25570/

So per 7 Kw, it's 770 kgs in 43,000 hours.
Or if it there were 5 of these engines using total of 35 Kw, in nearly 5 year it would use up 3.85 tons.

But I would suppose that this is at a thrust level which provides highest exhaust velocity and it is most efficient. And to get more thrust it would be less efficient and the 7 Kw engine could consume much more than 770 kg in 43,000 hours.

« Last Edit: 07/15/2014 10:04 am by gbaikie »

Offline KelvinZero

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Re: Just F9R, SEP tugs and propellant depots.
« Reply #26 on: 07/15/2014 11:05 am »
the article also mentions "30 million-newton-seconds of total impulse", which might be enough to work out whether its performance is comparable.. if you trust them to have put meaningful numbers next to each other.

I think I saw 1.5N quoted for the SEP tug. Using T = v*dm/dt I get around 10 tons in 6 years. It is still my best guess.

What about the general architecture? Is ten tons enough for the dragon+extended life support+landing legs and so on? Is landing the whole dragon+lander(+cargo) and returning directly to earth reasonable? Is the tall lander configuration with 20 tons of cargo (or up to 80 tons in extreme, one way cargo configurations) around the base reasonable?

(edit) I also think the scale from using a second stage is just too big. I think I will come up with a less massive but wider (5.2 meter) stage. If you want scale, get a fully fueled 1st stage to the moon!

Considering only volume, going from 3.7 width to 5.2 lets you halve the height of everything between the dragon capsule and the main engine. It would look a lot more like a moon lander. On the other hand it is not clear to me that the landing gear that has to reach past the main engine would be much lighter.

Since the Dragon remains attached to the second stage until reentry at earth, we could look at distributing the dragon trunk's functionality around the second stage. This can get us wider and squatter (and with trunk cargo closer to lunar surface) while still leaving clear channels for the dragon superdracos to contribute to controlled lunar landing. (using that main engine so near the regolith is a bit scary and perhaps can be avoided?)
« Last Edit: 07/15/2014 11:38 pm by KelvinZero »

Offline gbaikie

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Re: Just F9R, SEP tugs and propellant depots.
« Reply #27 on: 07/16/2014 11:32 pm »
If NASA has LOX depot in LEO, it seems to me that NASA can do a gemini type:
http://www.astronautix.com/articles/bygemoon.htm
Lunar exploration program.
Or NASA doesn't need to put large payloads on the Moon.
So a NASA lunar program is robotic which leads to more detailed exploration
which includes manned landings.

And result of such exploration would be which pole of the Moon has better places
which could be mined. And one would probably convene government panel which evaluated the entire
program and issued a report to Congress with it's findings.

And far as NASA concerned it then go forward with major program to explore Mars. And like Lunar exploration a key element of Mars exploration would be focused on finding Martian water, which would used
by Mars bases [and possibly making rocket fuel to get back to Earth].
For first Mars bases one could use Mars water at surface, but it seems with crew on Mars the focus should shift to find water underground. A underground liquid water source would be also related to find any life on Mars. Underground water could also to related to finding caves which also related to finding present or ancient life on Mars.
It seems in first decade of exploring Mars, one have a limited number of crew on Mars surface [5-10 or less]
and over years one would exploring with purpose of increasing amount people living on Mars, and more permanent bases. Budget restraints will limit future crew size, and lower costs would allow large number on Mars [more than 20 crew at any time period]. With the larger amount crew allowing more exploration of Mars in shorter time period.
 And during period of time NASA is exploring Mars, commercial lunar and/or asteroid mining could be occurring.
With commercial operation, there will effort to increase size of payload [and with tourism. number of passengers] with go to the Moon.
« Last Edit: 07/16/2014 11:33 pm by gbaikie »

Offline KelvinZero

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Re: Just F9R, SEP tugs and propellant depots.
« Reply #28 on: 07/17/2014 12:10 am »
Yes I was thinking about the other direction, how small can we go while not being unrealistic.

One approach could be to consider Apollo-like payloads. Then I dont need to worry about how realistic my lander is, you just need a Dragon trip to low lunar orbit and home. The Lander is 15 tons that could perhaps be moved there earlier with SEP or at least a more efficient transfer.

We could use an oxygen depot and deliver up to 10 tons of hydrogen in a final F9R payload to push the dragon to the moon and perhaps back. with a 6:1 ratio thats about 70 tons of LOX/LH propellant (neglecting tank mass). I think that is way more than you need.

Its fun playing with these numbers but I know I wont impress anyone with my abilities there. What is clear is the mind boggling scale you can achieve: a lander that you would have difficulty finding enough cargo for.

Im also becoming more interested in the dragon as perhaps the ideal shape for depositing your cargo trunk right on the surface and then lifting off again. I dont mean necessarily using a dragon, but there was a mars proposal recently that used that configuration with multiple rockets higher up and cargo right at the bottom. (And you could use an "uncrasher stage" to get nearer to the surface and still have all reusable components)

Offline Burninate

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Re: Just F9R, SEP tugs and propellant depots.
« Reply #29 on: 07/17/2014 06:16 am »
Yes I was thinking about the other direction, how small can we go while not being unrealistic.
Falcon Heavy will be cheaper per kilogram to LEO than Falcon 9, or any other LV, by a substantial margin.  Even if crossfeed never works, and no boosters are recovered, it beats the leader of the 20T class, the Proton, by launching more than twice as much mass for about the same price.  It will never be enough to launch a substantial surface return mission all on its own, but it's the best option on the horizon for a depot/departure stage.

In terms of mission chunk size, the missions that require a high Earth orbit will benefit substantially from 20T blocks to GTO in one piece, with ~10-15T (the initial size class for a small modular station demonstrated on the ISS) getting through to C0 or lunar resonant orbit or EML2.  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.

---

For SEP missions, the simplified approximation I've been throwing around for a unit from which to build a SEP tug is a ~7kW thruster - solar panel - battery package that weighs ~200kg, and burns ~200kg of propellant per year at 4000s Isp for 0.25N

Use those in groups of five and you have a very convenient shorthand unit: 1 ton of engine/power burns 1 ton of propellant per year.  1 ton of propellant at 4000s Isp aboard a 53T IMLEO mission nets you ~750m/s.  Use 2T of power and 1T of fuel and you can go through it in 6 months for 750m/s.  Use 12T of power and 1T of fuel and you can get through that 750m/s in 1 month.

Use 10T (~20% IMLEO) of power and 10T (~20% IMLEO) of fuel for a 1 year burn aboard a 50T IMLEO mission, and you net around 9km/s, enough to get 30T of payload to any useful orbit in the Earth-Moon system in 1 year of spiralling.

If thrust is applied with a duty cycle substantially lower than 100%, then highly elliptical orbits and the Oberth effect become practical to use, reducing total delta V requirements by up to half, but mission duration grows very rapidly to obscene levels.  The other efficient alternative is to settle for a ~5 year mission (the lifetime of the ion thrusters) using low-thrust spiral orbits, but a ratio of power to propellant of 5:1.  This last trade is the route you might take for routinized, zero-boiloff mild-cryogenic or non-cryogenic bulk fuel deliveries using highly shielded rad-hardened electronics: use a smaller tug (2T = 5 year mission but 8T extra payload mass, 4T = 2.5 year mission but 6T extra payload mass), that still burns through 10 tons of fuel to deliver 30-40T of payload through a ~9km/s trajectory.

Low-thrust propulsion is very good for raising and lowering periapsis, very bad for raising and lowering apoapsis, and the default near-circular spiral-out mode both increases radiation exposure significantly and takes up to twice as much dV as highly-elliptical Hohmann transfers using high-thrust propulsion.  Very mild aerobraking may be a good complement to its capabilities depending on how ugly the stresses are on the solar panels.  Feathering the solar panels and using Magnetoshell AeroCapture for braking would be great, but like the delicate solar panels, it probably won't be able to be retracted + reused.

The object of the above is just to give a quick way to calculate the very rough capacities of SEP systems based on the NEXT thruster, for the Earth-Moon system, with conservative assumptions.
« Last Edit: 07/17/2014 08:47 am by Burninate »

Offline KelvinZero

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Re: Just F9R, SEP tugs and propellant depots.
« Reply #30 on: 07/17/2014 07:03 am »
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.
I had really only been considering lunar missions and no modular construction, apart from adding the lunar cargo which has to be in two parts less than 10t each.

I had been assuming the F9R could deliver a bit less than 10 tons to LEO plus the 2nd stage itself of course. That was based on the current payload reduced by 30%

If falcon heavy ends up cheaper that is great. One of the nice things about depots is that you are not tied into your original architecture for most of your mass. FH, asteroid ISRU, lunar ISRU, and Skylon are all welcome.


Offline Hop_David

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Re: Just F9R, SEP tugs and propellant depots.
« Reply #31 on: 07/17/2014 05:20 pm »
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.

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.
« Last Edit: 07/17/2014 06:52 pm by Hop_David »

Offline Robotbeat

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Re: Just F9R, SEP tugs and propellant depots.
« Reply #32 on: 07/17/2014 11:18 pm »
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.
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Offline gbaikie

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Re: Just F9R, SEP tugs and propellant depots.
« Reply #33 on: 07/18/2014 01:12 am »
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.

Yeah, Gemini-style missions were a possible alternative to the Saturn V Apollo program- and without using depots.
If one were private entity, I would not suggest developing depots if one wanted to go to the Moon, but as public policy to it appears a necessity to develop depots.
Or a need of having depots could be argument for needing a space agency. Or the general useless of NASA is that it has not develop depots in the 60 years of it's existence.
And if imagine that NASA should explore the Moon and Mars, then what are if you thinking if you are against developing depots in space?
It certainty can't be we are in some kind of rush to go the Moon.
It seems if we are ever going to mine the Moon and/or Asteroids then a prerequisite is depots in microgravity. If we aren't going to do this, and maybe One imagines one should go to Mars instead [or whatever nonsensical approach] then one still needs depots to do manned exploration of Mars.
Or the moon is the easiest destination as compared to Mars, rocks, Venus, or Mercury.
I assume one wants to determine if there is minable water on the Moon. So that requires depots. If going to harder place to get to, then depots would be more useful.
« Last Edit: 07/18/2014 01:17 am by gbaikie »

Offline KelvinZero

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Re: Just F9R, SEP tugs and propellant depots.
« Reply #34 on: 07/18/2014 01:34 am »
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.

Offline Robotbeat

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Re: Just F9R, SEP tugs and propellant depots.
« Reply #35 on: 07/18/2014 02:14 am »
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.
You'd need a little bigger lander if you wanted to do EML2 rendezvous, though.

I really like SEP, but SEP may not even make sense if you have cheap propellant in LEO and higher. $7 million for, say, 7 tons of propellnt in LEO? Cheaper to use a chemical tug!

That story changes if you've got truly powerful SEP tugs, though (especially if combined with a sort of aerobraking). A 2 Megawatt tug (probably 2.5-3 MW with margin) could shuttle 20 tons to EML2 and go back to LEO in less than 2 weeks or so (at lower Isp). Or perhaps push about 80 tons to EML2 and return to LEO about 3 times a year. In both cases consuming just 7-8 tons of propellant. With just 100 kilowatts, round-trip missions are harder to justify, it's more useful as a kind of disposable kick-stage. Still definitely enabling in our current expendable launch paradigm, but 100kW isn't terribly relevant for a largely reusable architecture. Luckily, solar power is improving all the time. And 1kW/kg is possible with enough effort (and was demonstrated already to some extent on the Japanese IKAROS), so the solar array for a 3MW tug may only weigh 3 tons! But even a huge 1MW array at 150W/kg (roughly current best-of-breed) would only weigh 7 tons. But a lot of work would be needed to actually build it.
« Last Edit: 07/18/2014 02:25 am by Robotbeat »
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Offline KelvinZero

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Re: Just F9R, SEP tugs and propellant depots.
« Reply #36 on: 07/18/2014 03:21 am »
Sorry I missed you were referring to EML2, RobotBeat.

For that case I was assuming a LLO rendezvous with something practically identical to the Apollo capsule, but that is a bit more delta-v for the human side. I figured the SEP tug would move the lander as low as practical just before the human crew arrived.

You mention very high power tugs, but for just propellant we could just have more smaller ones right?  I like the idea of starting smaller and having faster evolution.

Another case where size and speed matters less is if we have a constant ongoing mission, the only type I consider valuable for HSF.

I was also interested in seeing what we could do with specifically the ARM SEP tug, since that is the hardware closest to possibly existing at the moment.

Offline MP99

Re: Just F9R, SEP tugs and propellant depots.
« Reply #37 on: 07/19/2014 05:48 pm »


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.

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.

"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

Offline gbaikie

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Re: Just F9R, SEP tugs and propellant depots.
« Reply #38 on: 07/19/2014 09:02 pm »
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.

Sure. One could divide task of human and robotic operations so that crew spend a higher portion of total mission time at lunar surface as compared to the time needed for journey to and back from the Moon.

Quote
For Apollo I think the entire lander (ascent + descent) was under 15 tons, and with space storable propellant.
With ascent being about 5 tons and descent being about 10 tons. If don't need the crew landing with it's return fuel, one can have the lander be 1/3rd the mass.
The elements which will require time: rendezvous, docking, and refuel could be shifted so it's done robotically- allowing crew to spend less time involved with this kind of thing.
And so for crew one wants to maximize time on Moon and lower time it takes to get there and back.

According wiki, in regards to Apollo 16:
"John Young and Charles Duke spent 71 hours—just under three days—on the lunar surface, during which they conducted three extra-vehicular activities or moonwalks, totaling 20 hours and 14 minutes."
And total "Mission duration    11 days, 1 hours, 51 minutes, 5 seconds "
http://en.wikipedia.org/wiki/Apollo_16
That's pretty good, 3 days on surface and 8 days coming and going.
As starting point, you don't want to do much worst than that, though having everything on one rocket, helps a lot in terms of doing this.
An simple way to improve it, is have crew spend 4 or more days rather 3 days on surface, but a focus should be reducing this time period of the 8 days coming and going. And if crew is spending days in LEO and/or lunar orbit, it will extend this time period.
So if crew first goes to LEO [which Apollo spent a brief period of time in] it shouldn't take the 2 days that is required to get to ISS:
"The next crew to launch toward the International Space Station will make the trip faster than any astronauts before them, thanks to a new docking plan being tested this month...
While it normally takes Soyuz vehicles two days to reach the orbiting laboratory after launch, Cassidy, Vinogradov and Misurkin will make the trip in just six hours."
http://www.space.com/20099-space-station-crew-fast-spaceflight.html

One advantage over ISS, is that the docking could occur at lower orbit. And another potential advantage is the depot could dock to launch vehicle, rather than launch vehicle docking to a station. Meaning a separate space tug can be part of depot.
A critical aspect is hitting the lunar trajectory window, and delays of launches [technical/weather] would require a depot to have flexibility to allow for this [have lower orbit when depot is filled [if not in immediate use, it could be parked higher- to reduce atmospheric drag] and eccentric orbits for tugs to adjust timing for planned launches.
Or with crew, one could simply repeat Apollo but with it's payload being is just crew to lunar orbit. With earth reentry and lander vehicle already sent [unmanned] to lunar orbit.
Or moon direct- with no time in LEO or lunar orbit, crew launch from earth in a lunar lander which directly lands on lunar surface and having return ascent vehicle already at lunar surface [and earth return vehicle in lunar orbit]. And this would need lunar lander would need more delta-v than Apollo's LEM [the ascent or descend stages- but not combined], since the vehicle is not carrying an ascent vehicle, it could be around 10 tons.
In either of these cases most of payload mass which involve in manned mission would be unmanned launches which could spend weeks in space and could be launched from various launch sites around the world.


Offline Lar

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Re: Just F9R, SEP tugs and propellant depots.
« Reply #39 on: 07/20/2014 01:44 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, Martin

ANY orbit is reachable from ANY launch site with enough doglegging and plane change maneuvers. I think what you meant was "practically" reachable...   
"I think it would be great to be born on Earth and to die on Mars. Just hopefully not at the point of impact." -Elon Musk
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Offline Burninate

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Re: Just F9R, SEP tugs and propellant depots.
« Reply #40 on: 07/23/2014 09:43 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, Martin

ANY orbit is reachable from ANY launch site with enough doglegging and plane change maneuvers. I think what you meant was "practically" reachable...
Also - Boca Chica is near the Tropic of Cancer.  That would correspond to only a few degrees off the Lunar orbital plane - which is *better* than an equatorial low Earth orbit for the purposes of lunar missions.

Equatorial LEO is actually rather difficult to reach if you don't have a handy dandy equatorial launch site.  Any launchsite can efficiently launch into a plane of inclination equal to their latitude or higher - the benefit of equatorial launchsite going into equatorial-prograde orbit due to the rotation of Earth's surface is only around 465m/s dV over going from equatorial launchsite into polar orbit or polar launchsite into polar orbit.  This is vastly overstated in the popular media, as it's about 5% of launch dV.  Any major plane change maneuver in LEO swamps this effect.  The major benefit for LEO is that an equatorial launchsite can launch into any orbit/inclination, while a polar launchsite is stuck with polar ones unless it wants to spend tons of dV.  Adding onto that modest 465m/s in commercial commsat practice is a comparable dV apogee kick plane change maneuver necessary to reach the equatorial plane and transition from inclined-GTO to eq-GSO from an inclined launchsite.
« Last Edit: 07/23/2014 09:55 pm by Burninate »

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