Author Topic: Space Elevator for Mars  (Read 21436 times)

Offline Nilof

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Re: Space Elevator for Mars
« Reply #100 on: 01/03/2018 03:15 AM »
Quote
Climber & Rappeller Utility

Yes, a climber would need something other than chemical energy.  Even HVDC power lines are inefficient and unsuitable over MSE distances.  It may be that areosync PV + a superconducting tether power line will be needed, to power an MSE climber system with performance superior to Mars launch.

A cargo descent rappeller is more efficient than a climber:  it uses 0 kW.  :)  And it does so while saving the tremendous energy that would otherwise be devoted to propellant manufacture on Mars, for the cargo ship's surface launch.

There's utility in that, yes?

Mars surface to mars escape is 5.2 km/s. Mars surface to the foot of a 1400 lower phobos tether is 4.1 km/s. Going from Mars surface to Phobos the traditional way takes 5 km/s.

You need the really ambitious versions of the lower Phobos tether to gain a substantial benefit from the rest of the system compared to just burning from LMO with an Oberth benefits. Otherwise, you just spent most of your delta-v alternative cost budget to get to Phobos, and you save a few hundred m/s of delta-b budget at best.

So it really boils down to Mars SSTO vs Mars suborbital vehicle + beamed power climber hybrid vs mass driver.

I think a more likely use for a Rapeller could be with a 1400 km tether to drop cargo to LMO in order to build infrastructure there.
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 LMT

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Re: Space Elevator for Mars
« Reply #101 on: 01/03/2018 04:43 AM »
I think a more likely use for a Rapeller could be with a 1400 km tether to drop cargo to LMO in order to build infrastructure there.

As regards the rappeller, it's used more effectively on the Mars Lift as given.  Dropping off cargo at Arestation and rappelling it to Mars obviates the need for suborbital transfer craft launch and landing, and the associated propellant manufacture on Mars.  It's not hard to make the business case for Omaha Trail cargo, vis-a-vis the upper/lower tethered system. 

--

And yes, the budget required to reach a Phobos tether from Mars, for transit in either direction, is considerable.  It does seem to consume the desired savings of the upper/lower tethered system.  Also some practical issues of Phobos cargo delivery were noted in Dr. Lades' talk.

Offline Hop_David

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Re: Space Elevator for Mars
« Reply #102 on: 01/03/2018 03:20 PM »
The projectiles are guided by the magnetic rings, in your drawing they would be the sections connected to the ground-cables. Between, the projectiles are free-flying. The meteors achieve nothing. See below.

Okay so we've eliminated the enclosing ring continuously accelerating the projectiles

Then the projectiles flying at faster than orbital velocity would not be flying a nice circular orbit from one platform to another. You would have to aim the ricochet to hit the next platform. See illustration below

Holding the platforms aloft by continuously ricocheting machine gun bullets poses some engineering challenges. How far apart are the platforms? What precision is needed in directing the ricochets?

I can imagine some spectacular failure modes with this scheme as well.
« Last Edit: 01/03/2018 03:22 PM by Hop_David »

Offline Paul451

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Re: Space Elevator for Mars
« Reply #103 on: 01/03/2018 06:46 PM »
Okay so we've eliminated the enclosing ring continuously accelerating the projectiles
Then the projectiles flying at faster than orbital velocity would not be flying a nice circular orbit from one platform to another.
You would have to aim the ricochet to hit the next platform.

You might want to read up on orbital rings, you seem to be thinking of it as a passive tube. The concept is that the particles are magnetically accelerated (not "ricocheted") faster than orbital velocity. It's like a launch-loop, but around the entire world. It wouldn't work if the particles were just in a natural circular orbit. The centripetal acceleration of the particles keeps the useful structure aloft.

My point was that if you damage one of the magnetic guides, the others would need to redirect to compensate. But they'll need to do that anyway whenever parts of the system is under maintenance. And I suspect they'll be constantly doing that based on the natural variability in the ring anyway.

But explaining how it works is off-topic for this thread, especially zombied from back in August. If you want to dig around, there might be an existing thread in Advanced Concepts that would be more appropriately re-animated.

Offline Hop_David

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Re: Space Elevator for Mars
« Reply #104 on: 01/03/2018 08:43 PM »
It wouldn't work if the particles were just in a natural circular orbit.

That was my point. You had altered my illustration of a continuous circular ring. You removed the ring  but left the projectiles still sailing nicely along circular paths from one platform to another.

But the paths would look more like straight lines.

So rather than constant acceleration along a circular ring, you'd have discreet, abrupt accelerations at the corners of a large polygon.

Below I drew the projectiles being magnetically redirected from one platform to another.

My objections remain.

It would take very precise aiming to send a projectile from one platform to another thousands of kilometers away. The projectile would have to hit the exact center of the platform or it would set the platform spinning instead of halting and reversing it's downward fall.

In the case of earth based platforms 60 degrees apart and approximating the hyperbolic paths as straight lines, the change in vertical velocity component would be about the same the projectile speed. How fast are the projectiles moving? 10 km/s? 16 km/s? That's a very drastic velocity change in a very short time. How many g's would the projectile endure? How many g's to the platform? The pusher plate in the boom boom Orion vehicle comes to mind.

But explaining how it works is off-topic for this thread, especially zombied from back in August. If you want to dig around, there might be an existing thread in Advanced Concepts that would be more appropriately re-animated.

Isaac Arthur's scheme was off topic back in August as well (in my opinion). But I address the August arguments because Nilof just responded to them within the past few days. If you like, you can open a new thread on Isaac Arthur's orbital rings.

Offline LMT

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Re: Space Elevator for Mars
« Reply #105 on: 01/05/2018 06:27 AM »
No Tankers

Notably, one of the benefits of the Omaha Trail proposal is that cargo flights can be launched without dedicated tanker ships. 

No tankers at all.  Just returning cargo ships.

This could be especially beneficial to the construction phase of SpaceX's Mars City, which could take decades.  The work might go 10x faster at Omaha Crater, but still, it's a benefit.



Cargo flight staging. Deimos propellant. Mars Lift space elevator in gold.

Depot

That slide didn't garner comment, but it shows on-orbit depot use, an ambition at SpaceX and elsewhere.

Quote from: Elon Musk
By establishing a propellant depot in the asteroid belt or one of the moons of Jupiter, you can make flights from Mars to Jupiter, no problem.

Quote from: Charles Miller
...produce propellant [from water in the lunar soil]. You’d put that in a propellant depot in lunar orbit to allow humans to go anywhere in the solar system.

Here the depot is at Deimos, with propellant shuttled to LEO.  Typical #s, allowing notional 55 t propellant reserves:

(1) Burn 397 t of 1008 t propellant after DRL, for transit and LEO circularization.

(2) Launch cargo to LEO.

(3) Transfer 540 t in LEO.

(4) Burn 16 t for EDL.

(5) Burn 540 t for transit and Mars Lift Arestation circularization.

Option:  To double payload, (1) is performed with 1066 t propellant load, twice, filling tanks in (3).  A second Earth launch transfers the second cargo.  With full tanks (5) delivers 300 t to Arestation.

All is done without an Earth tanker fleet, tanker booster fleet, and their propellant.  Cost savings would follow efficiency. 
« Last Edit: 01/08/2018 01:21 AM by LMT »

Offline meberbs

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Re: Space Elevator for Mars
« Reply #106 on: 01/05/2018 02:49 PM »
That slide didn't garner comment, but it shows in-orbit depot use, an ambition at SpaceX and elsewhere.
False. My previous post was directly discussing that slide. Your claims in it are wrong, and the claimed numbers you just provided make this easy to show.

Here the depot is at Deimos, with propellant shuttled to LEO.  Typical #s, allowing notional 55 t propellant reserves:

(1) Burn 397 t of 1008 t propellant after DRL, for transit and LEO circularization.
Not sure why you are limiting yourself to 1008 t instead of the 1100 that constitutes a full load of the tanks.
Also, I get that to just do the 1.55 km/s burn to get on an Earth trajectory it would take 399 t with your claimed propellant load and the 50 t cargo needed for this architecture to be compared to SpaceX's.

This does not include losses in transit or LEO circularization.

(2) Launch cargo to LEO.

(3) Transfer 540 t in LEO.

This allows you 4.385 km/s delta-v in total. (Cargo is 150 tons for direct comparison of course) Typically LEO to MTO takes 4.3 km/s on its own, and that is not for fast transfers. If you want to get straight back to Deimos like you show, you would need to burn the same 0.94 km/s to circularize at Deimos that the rail launcher gave you on the way out, and that is assuming you accurately hit the aerocapature, getting into the exact desired capture orbit and have very accurate timing of the transfer as well.

(4) Burn 16 t for EDL.
I'm curious where you got this number, I don't think anyone outside SpaceX has a good idea how much fuel is needed forth this on BFR, there are too many unknowns.

Option:  To double payload, (1) is performed with 1066 t propellant load, twice, filling tanks in (3).  A second Earth launch transfers the second cargo.  With full tanks (5) delivers 300 t to Arestation.
Not going to bother with this since it  obviously just has larger problems than the first plan.

Note that the extra 92 tons I mentioned that you could start with would also cause an extra 25 t to be lost in the initial burn from Earth, and then with the losses in transit and LEO circularization, this will not come close to rescuing your plan.

All is done without an Earth tanker fleet, tanker booster fleet, and their propellant.  Cost savings would follow efficiency.
SpaceX is planning to fully fill up the ship's fuel before departing for Mars. They aren't doing this for no reason, so it should be obvious that your plan to only half fuel the ship will not work. None of the claimed benefits would emerge.
« Last Edit: 01/05/2018 02:50 PM by meberbs »

Offline LMT

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Re: Space Elevator for Mars
« Reply #107 on: 01/06/2018 09:01 PM »
Cargo Depot Redo

False. My previous post was directly discussing that slide.

Your architecture expects you to then fully fuel that other ship for it to transfer to Mars

compare with SpaceX's architecture  apples to apples, 50 tons return to Earth

No, you didn't mention the slide, and your text is not consistent with that flight plan.

There's no need to "fully fuel" cargo ships in LEO.  That's required only for crewed missions, where delta-v must be maximized.  Different flight plan.

Also there's no cargo on the return flight.  These cargo ships never land on Mars.  So unless you assume a booming market for Deimos cinder block at Home Depot, return cargo is naturally 0 t.

To discuss a particular flight plan, paste the image and reference the numbered steps; that'll keep your text clear.

Not sure why you are limiting yourself to 1008 t

Because that's the minimum needed.  We add more for options, such as doubled payload.

to just do the 1.55 km/s burn to get on an Earth trajectory it would take 399 t with your claimed propellant load and the 50 t cargo needed for this architecture to be compared to SpaceX's.

This does not include losses in transit or LEO circularization.

As we noted previously, perihelion requires ~1.55 km/s, aphelion requires ~0.8 km/s.  "Typical" #s are intermediate, not the extremes.  For example, 1.16 km/s.   

Then we tacked on 0.5 km/s for circularization.  (Your 0.94 km/s reasoning isn't right, and typical #s are smaller.)

As for losses, long-duration tankage scenarios expect some ZBOT method.  Any further loss would get some of the 92 t left at Deimos.

I'm curious where you got this number, [Burn 16 t for EDL.]

ITS drops 99%+ of KE aerodynamically, leaving only a landing burn.  We used a 40 s hover.

I don't think anyone outside SpaceX has a good idea how much fuel is needed forth this on BFR

"Fuel"?  "Propellant".  :)  Redo the cargo flight plan to see how the numbers improve.

Offline meberbs

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Re: Space Elevator for Mars
« Reply #108 on: 01/06/2018 11:32 PM »
No, you didn't mention the slide, and your text is not consistent with that flight plan.
I shouldn't have to explicitly mention the slide, it should be obvious that is what I was discussing, and it follows your mission plan exactly, other than pointing out that you have entirely insufficient fuel. 

There's no need to "fully fuel" cargo ships in LEO.  That's required only for crewed missions, where delta-v must be maximized.  Different flight plan.
Wrong, please read the numbers in the post I just made again. You do not have anywhere near enough fuel for what you are proposing. I suggest you actually think about the fact that SpaceX is not planning to use fewer tankers for cargo vs. crew flights and why this is their plan.

Also there's no cargo on the return flight.
Wrong, look at SpaceX's architecture again.

These cargo ships never land on Mars.  So unless you assume a booming market for Deimos cinder block at Home Depot, return cargo is naturally 0 t.
Your architecture still involves transport capability from Mars surface, no reason cargo could only come from Deimos.

Not sure why you are limiting yourself to 1008 t

Because that's the minimum needed.  We add more for options, such as doubled payload.
Except as I showed it is nowhere near enough. Doubled payload would fail even with full tanks.

to just do the 1.55 km/s burn to get on an Earth trajectory it would take 399 t with your claimed propellant load and the 50 t cargo needed for this architecture to be compared to SpaceX's.

This does not include losses in transit or LEO circularization.

As we noted previously, perihelion requires ~1.55 km/s, aphelion requires ~0.8 km/s.  "Typical" #s are intermediate, not the extremes.  For example, 1.16 km/s.
Your architecture is basically useless unless it works every window, not just when you get an easy window. You need to plan for the worst case.

Then we tacked on 0.5 km/s for circularization.  (Your 0.94 km/s reasoning isn't right, and typical #s are smaller.)
The 0.5 km/s should be conservative for circularization at Earth after aerocapture, however as my calculation showed, you actually allowed less than 0 km/s for this.

The 0.94 number has nothing to do with the LEO circularization, and everything to do with the fact that you used 0.94 km/s to get from Deimos to an elliptical orbit with a periapsis close to Mars. In the best case scenario, that is also the orbit you would end up in after an aerocapture maneuver (actually you would need a lower periapsis to be deep enough in the atmosphere, but that just makes things worse for you). Circularizing that elliptical orbit to match Deimos takes exactly the same delta-v as entering that orbit from Deimos took in the other direction.

As for losses, long-duration tankage scenarios expect some ZBOT method.  Any further loss would get some of the 92 t left at Deimos.
Unfortunately for you, SpaceX is not using ZBOT, and their method of reducing boil off is to vent the outer tanks for greatly improved insulation. You can't do this, since you need the fuel that is in the outer tanks.

I'm curious where you got this number, [Burn 16 t for EDL.]

ITS drops 99%+ of KE aerodynamically, leaving only a landing burn.  We used a 40 s hover.
99% KE is for direct landing on Mars from MTO. Your use of it in this case is simply ignoring the deorbit burn entirely. It also isn't a number that is necessarily very exact. The 40s hover that you guessed doesn't mean much on its own. It depends on how many engines at what throttle.

Redo the cargo flight plan to see how the numbers improve.
No, you redo the numbers without using assumptions that contradict your architecture, and without assuming best case transfer windows, and without ignoring the fuel needed to circularize at Deimos.

Offline LMT

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Re: Space Elevator for Mars
« Reply #109 on: 01/10/2018 06:41 AM »
No, you didn't mention the slide, and your text is not consistent with that flight plan.
I shouldn't have to explicitly mention the slide, it should be obvious that is what I was discussing, and it follows your mission plan exactly, other than pointing out that you have entirely insufficient fuel.

Not at all.  The given, specific cargo flight plan has no requirements or assumptions for full LEO tanks or 50 t return cargo -- and of course in this flight plan there's no Mars landing for a 50-t cargo pickup anyway. 

So our mass and delta-v numbers work as given, with corresponding efficiency improvements over SpaceX baseline.  The numbers shouldn't be very surprising, especially given the familiarity of the route and spacecraft.

Then we tacked on 0.5 km/s for circularization.  (Your 0.94 km/s reasoning isn't right, and typical #s are smaller.)
The 0.94 number has nothing to do with the LEO circularization, and everything to do with the fact that you used 0.94 km/s to get from Deimos to an elliptical orbit with a periapsis close to Mars. In the best case scenario, that is also the orbit you would end up in after an aerocapture maneuver (actually you would need a lower periapsis to be deep enough in the atmosphere, but that just makes things worse for you). Circularizing that elliptical orbit to match Deimos takes exactly the same delta-v as entering that orbit from Deimos took in the other direction.

That reasoning is quite wrong.  The delta-v coming off the DRL by no means sets the delta-v for aerocapture circularization. 

For one thing, the DRL launch vector is certainly not the most efficient for lowering periapsis; so it's not the delta-v that one must "take".  That's easy to see with the example GMAT script: tweak the DRL VNB to reach the same periapsis using much less than 0.94 km/s delta-v.

Example vectors:

GMAT TOI.Element1 = 0.5;
GMAT TOI.Element2 = 0;
GMAT TOI.Element3 = 0.5; 

GMAT TOI.Element1 = 0.64;
GMAT TOI.Element2 = 0;
GMAT TOI.Element3 = 0;

Second, the mission literature shows you what's actually needed for circularization.  In one Lockheed Martin flight plan the spacecraft changes plane and circularizes to Deimos with a pair of burns totaling 607 m/s delta-v.



In an ITS cargo plan MOI is aerocapture, with aerodynamic plane change.  This removes propulsive plane change, lowering the burns for Arestation down toward 0.5 km/s.

But the rocket equation and orbit changes are, by themselves, OT.  Focus should be on possibilities for the Mars Space Elevator and related tether systems.

Offline meberbs

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Re: Space Elevator for Mars
« Reply #110 on: 01/10/2018 08:41 AM »
Not at all.  The given, specific cargo flight plan has no requirements or assumptions for full LEO tanks or 50 t return cargo -- and of course in this flight plan there's no Mars landing for a 50-t cargo pickup anyway. 
50 tons return cargo is part of the SpaceX baseline you claim to be comparing to, you can't just throw out capability and claim your system is equivalent but more efficient.

Full LEO tanks isn't an assumption, it is a fact of what SpaceX has presented as their plan.

So our mass and delta-v numbers work as given, with corresponding efficiency improvements over SpaceX baseline.  The numbers shouldn't be very surprising, especially given the familiarity of the route and spacecraft.
No, your numbers don't work. You have simply ignored most of what I wrote, and have made no attempt to do things like update your numbers to account for worst case transfer windows.

Then we tacked on 0.5 km/s for circularization.  (Your 0.94 km/s reasoning isn't right, and typical #s are smaller.)
The 0.94 number has nothing to do with the LEO circularization, and everything to do with the fact that you used 0.94 km/s to get from Deimos to an elliptical orbit with a periapsis close to Mars. In the best case scenario, that is also the orbit you would end up in after an aerocapture maneuver (actually you would need a lower periapsis to be deep enough in the atmosphere, but that just makes things worse for you). Circularizing that elliptical orbit to match Deimos takes exactly the same delta-v as entering that orbit from Deimos took in the other direction.

That reasoning is quite wrong.  The delta-v coming off the DRL by no means sets the delta-v for aerocapture circularization. 

For one thing, the DRL launch vector is certainly not the most efficient for lowering periapsis; so it's not the delta-v that one must "take".  That's easy to see with the example GMAT script: tweak the DRL VNB to reach the same periapsis using much less than 0.94 km/s delta-v.

...

Second, the mission literature shows you what's actually needed for circularization.  In one Lockheed Martin flight plan the spacecraft changes plane and circularizes to Deimos with a pair of burns totaling 607 m/s delta-v.
I just used the number you provided for simplicity, assuming that you had bothered to orient the vector in close to optimal direction. I figured that a maneuver like in the picture you posted could reduce the delta-V some, but the order of magnitude is still the same. As I showed you barely have enough fuel transferred to get to MTO from LEO (using your number for transferred fuel rather than a corrected value.) You would have less than 100 m/s remaining in a typical case, so it doesn't matter whether you need 600 or 900 m/s for the last step, my point remains the same.

But the rocket equation and orbit changes are, by themselves, OT.  Focus should be on possibilities for the Mars Space Elevator and related tether systems.
You brought up your plan for supposedly improving on the BFR system using your tether architecture for Deimos-Mars transport. Pointing out that your numbers don't add up is on topic if your architecture is. (And if not your architecture can be split off into a separate topic.)

You have simply ignored multiple of the points I made about the flaws in your plan and analysis. I have stuck to ones provable with straightforward numbers, no point in bringing up the others when you can't recognize these.

Offline stefan r

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Re: Space Elevator for Mars
« Reply #111 on: 01/10/2018 08:41 PM »

(1) Burn 397 t of 1008 t propellant after DRL, for transit and LEO circularization.


Why LEO circularized?  The shuttle already has a tether attachment.  Seems like a waste of momentum. 

Offline LMT

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Re: Space Elevator for Mars
« Reply #112 on: 01/12/2018 04:25 AM »
I just used the [DRL] number you provided for simplicity, assuming that you had bothered to orient the vector in close to optimal direction.

?  The DRL vector isn't remotely "close to optimal direction", obviously.  You don't seem to understand VNB impulse.  You might "look it up," as you put it.  OT here.

The given, specific cargo flight plan has no requirements or assumptions for full LEO tanks or 50 t return cargo -- and of course in this flight plan there's no Mars landing for a 50-t cargo pickup anyway. 
50 tons return cargo is part of the SpaceX baseline you claim to be comparing to, you can't just throw out capability and claim your system is equivalent but more efficient.

Full LEO tanks isn't an assumption, it is a fact of what SpaceX has presented as their plan.

Oh, it's fair comparison because equivalent and properly calculated.  Same ship, same cargo, out and back, with and without Omaha Trail infrastructure.  "150 out / 0 back" is naturally the baseline ITS requirement.   

Each variation on cargo and crewed flights is calculated and compared separately.  We don't mix willy-nilly.

--

SpaceX hasn't proposed an Omaha Trail themselves.  But if tomorrow SpaceX proposed it under their own logo, I'm pretty sure no one would say, "But this isn't the plan!  They can't claim this is more efficient!"  More likely folks would brainstorm ways to deliver tunnel-boring machines and Tesla Roadsters by elevator.  ;)

Quote from: Rob Manning, Universe Today
"Mars is really begging for a space elevator," said Manning. "I think it has great potential. That would solve a lot of problems, and Mars would be an excellent platform to try it."

Offline meberbs

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Re: Space Elevator for Mars
« Reply #113 on: 01/12/2018 06:45 AM »
I just used the [DRL] number you provided for simplicity, assuming that you had bothered to orient the vector in close to optimal direction.

?  The DRL vector isn't remotely "close to optimal direction", obviously.  You don't seem to understand VNB impulse.  You might "look it up," as you put it.  OT here.

I already said why I was wrong and why it doesn't matter, but let me rephrase since you seem to have not read my post.

I didn't bother looking at what vector you used in what frame, so the issue has nothing to do with my understanding of the frame. The 940 m/s seemed to be (and is, your next post showed that in a practical example, the answer is around 600 m/s) the correct order of magnitude, so I took the path of not doing more math, since my experience told me that the number was in the correct ballpark.

The real issue that you are failing to address is that you don't have the available delta V for any of these trajectories.

The given, specific cargo flight plan has no requirements or assumptions for full LEO tanks or 50 t return cargo -- and of course in this flight plan there's no Mars landing for a 50-t cargo pickup anyway. 
50 tons return cargo is part of the SpaceX baseline you claim to be comparing to, you can't just throw out capability and claim your system is equivalent but more efficient.

Full LEO tanks isn't an assumption, it is a fact of what SpaceX has presented as their plan.

Oh, it's fair comparison because equivalent and properly calculated.  Same ship, same cargo, out and back, with and without Omaha Trail infrastructure.  "150 out / 0 back" is naturally the baseline ITS requirement.   
Only if you ignore SpaceX's stated cargo return capability, so no that is not the baseline. When you are comparing to SpaceX the baseline is what SpaceX stated, 150 out and 50 back, not something you made up.

Maybe you didn't understand their presentation, they are using basically the same ship for cargo or crew the capability is the same either way.

SpaceX hasn't proposed an Omaha Trail themselves.  But if tomorrow SpaceX proposed it under their own logo, I'm pretty sure no one would say, "But this isn't the plan!  They can't claim this is more efficient!"  More likely folks would brainstorm ways to deliver tunnel-boring machines and Tesla Roadsters by elevator.
SpaceX wouldn't propose what you have themselves because the numbers don't work out. If they did people would add up the numbers and point out that fact. (This has already happened on this site, notably in L2 when people added up some numbers and determined the BFR rocket dimensions did not make sense. It turns out one of the numbers was out of date or miscommunicated, and their prediction came close to the actual dimension)

Quote from: Rob Manning, Universe Today
"Mars is really begging for a space elevator," said Manning. "I think it has great potential. That would solve a lot of problems, and Mars would be an excellent platform to try it."
Mars could be a great location for a space elevator, but your proposed use of it in combination with BFR simply does not work the way you claim it does.

Please stop ignoring all of the problems I have pointed out.

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