Author Topic: MCT Speculation and Discussion Thread 4  (Read 610801 times)

Offline philw1776

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Re: MCT Speculation and Discussion Thread 4
« Reply #460 on: 08/24/2015 05:07 PM »
MCT Mass   180   mT

S2 Mass w/MCT   1025   mT

S2 Mars 25mT Cargo    8.5   Km/sec Rocket Equation

I don't think those three go together.

A stage with 180mt dry mass, 8.5km/s delta v and 25mt payload has a wet mass of 1980mt.

Dry mass return from Mars is 80mT + 25mT payload = 105mT
Total fueled mass is 1025 yielding nearly 8.5 Km/sec with Elon's cited ISP of 380



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Offline nadreck

Re: MCT Speculation and Discussion Thread 4
« Reply #461 on: 08/24/2015 05:12 PM »
So we're saying that the oft stated 200mT MCT landing on Mars is impossible?

We aren't, Oli is, I am sure some agree with him. As opposed, I believe 50t could be the dry mass of a 820t fully loaded and fueled MCT, note that at most there would be 1.3km/s of ΔV or so a re-entry mass around 230t, and a landing mass approaching 150t.

Note my overall mission requirement ΔV is 6km/s.
It is all well and good to quote those things that made it past your confirmation bias that other people wrote, but this is a discussion board damnit! Let us know what you think! And why!

Offline Oli

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Re: MCT Speculation and Discussion Thread 4
« Reply #462 on: 08/24/2015 05:23 PM »
So we're saying that the oft stated 200mT MCT landing on Mars is impossible?

Looking at 5 different 20t+ lander concepts with only supersonic retropropulsion I can find in my pdf collection (:) ), the best payload / structural mass ratio is ~0.7. And that's for an expendable lander with 13m diameter blunt body and a payload of 52t (from austere human mission to Mars, 2009).

Slender bodies are worse, for example the lander from DRM5 has a ratio of 0.41.

100 / 80 is 1.25.
« Last Edit: 08/24/2015 05:23 PM by Oli »

Offline Impaler

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Re: MCT Speculation and Discussion Thread 4
« Reply #463 on: 08/25/2015 01:10 AM »
I think we need to consider entry velocity when looking at any lander, it's the #1 driver of TPS, structural mass and retro-propulsion needs.  Most lander heritage involves direct entry from interplanetary transit and very high entry velocity.  The Viking lander was the exception and is probably the best comparison because it doesn't drop nearly as many parts along the way.  It had a landed mass of 62% of entry mass compared with 10% for the Phoinex lander.

If the requirement is scaled down to ~3.6 km/s as you would have from an entry from Mars orbit then the entry becomes exceedingly gentle, combined with a lifting body shape to further lengthen and distribute the heating load should put it well within the range of radiative metallic systems which would be negligible in additional mass because they would simply BE the outer skin of the vehicle.  G-forces would be reduced to under 1.5 g's meaning the vehicle doesn't need to be anywhere near as strong.  Retro-propulsion needs should be on the order of ~800 m/s.


http://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/20100017668.pdf  (Se slide 14)

http://www.ssdl.gatech.edu/papers/conferencePapers/IAC-2008-D2.3.9.pdf (see page 7)

So I propose a vehicle designed to act as a ferry between the surface and orbit, able to land 100 mt and return to orbit with 25 mt PLUS enough propellant to make another landing on Mars.  At 75 mt dry mass for the vehicle the landing requires 40 mt of propellants.  The launch mass is 400 mt of which 260 is assent (4.1 km/s) propellant, 25 cargo, 40 return propellant (.8 km/s) and 75 the vehicle dry mass.  Total propellant load is just 300 mt meaning propellant production on Mars can be a fraction of that needed for Direct return.  The structural mass fraction at launch is a very conservative 18.75% and payload is 46% of entry mass (including propellant) both very reasonable figures.

The lander would launch from Earth atop a 2 stage launch vehicle and would be loaded with ~50 mt of propellants providing ~900 m/s DeltaV to be used for emergency separation and propulsive landing in the even of an abort.  This would put the total launch mass at 225 mt.  And the lander could then be fully topped off via another launch of and transfer of 250 mt or propellant from a stretched 2nd stage tanker.  Transit to Mars is then done via a hybrid propulsion system, first a large SEP vehicle would slow push the lander to high Earth orbit where crew would board via a dragon capsule.  Then the SEP would separate and the lander would make a lunar-Earth slingshot burn of ~1 km/s for fast transit to Mars and use propulsive capture and perhaps some airobraking to reach low mars orbit leaving just enough propellant for landing.  On the surface the cargo is unloaded and a 25 mt return cabin is loaded in it's place.  The SEP system would make a slow transit to mars arriving well after the lander and wait in low orbit for rendezvous after which it would bring the lander back to high Earth orbit, crew would disembark again via Dragon capsule and then the unoccupied combined vehicle would do a down-spiral to low Earth orbit where it would be refueled again and use it's propulsive capacity to reduce entry velocity to the 3.6 km/s velocity it can tolerate followed by landing on Earth for reloading and reintegration, SEP remains in orbit ready to be reused.


Total BFR launches would be 6, one for the lander, 2 tankers of chemical propellant, 1 for SEP hardware, 2 estimated for SEP propellant.  In addition 2 trans-lunar Dragon capsule launches likely on Falcon Heavy.  Each subsequent launch would need 5 launches when SEP is reused.  The cycle would be one round trip every 2 synods.  If philw1776 can run the numbers of the size of the BFR needed to do a 225 mt launch that would be most informative to see how it compares.

As the Mars colony is built up the small SEP transit vehicle would be replaced with a large more powerful cargo hauler with extensive habitats aka the mother-ship.  The lander would be fueled in LEO and used to rapidly deliver passengers (100 at a time) to the waiting mother-ship in high orbit and would then travel attached too it as the mother-ships speed is expected to be competitive with the earlier direct flight.  At mars the lander is used to repeatedly ferry between surface and low orbit with containerized cargo being reloaded in orbit.  At Earth the launch of chemical propellant and landers is almost eliminated in favor of launching naked cargo containers and SEP propellant to be loaded onto the mother-ship.

« Last Edit: 08/25/2015 02:09 AM by Impaler »

Offline Vultur

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Re: MCT Speculation and Discussion Thread 4
« Reply #464 on: 08/25/2015 01:26 AM »
Isn't PICA supposed to be significantly better than previous TPS materials though? Would that reduce the required dry mass?

Offline Oli

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Re: MCT Speculation and Discussion Thread 4
« Reply #465 on: 08/25/2015 02:27 AM »
I think we need to consider entry velocity when looking at any lander, it's the #1 driver of TPS, structural mass and retro-propulsion needs.  Most lander heritage involves direct entry from interplanetary transit and very high entry velocity.  The Viking lander was the exception and is probably the best comparison because it doesn't drop nearly as many parts along the way.  It had a landed mass of 62% of entry mass compared with 10% for the Phoinex lander.

The human lander concepts I referred to are all designed for either entry from Mars orbit or aerocapture into orbit + entry from Mars orbit. Granted, gs, heat peak rate and heat load are significantly higher in the aerocapture phase than in the entry phase, although its a lot better than direct entry.

"Parts dropped along the way" are part of the structural mass at entry, which I was refering to. I.e. the entry mass minus the fuel and payload.

The best lander in terms of payload to structural entry mass I cound find is in the pdf attached on page 15, top left. A value of ~1.2. With SIAD, from orbit. It has a total entry mass of only 20t though. The higher the entry mass the worse usually.

With HIAD it may get better, haven't seen the mass break down of such a lander yet. HIAD could be interesting for MCT.

The lander would launch from Earth atop a 2 stage launch vehicle and would be loaded with ~50 mt of propellants providing ~900 m/s DeltaV to be used for emergency separation and propulsive landing in the even of an abort.  This would put the total launch mass at 225 mt.  And the lander could then be fully topped off via another launch of and transfer of 250 mt or propellant from a stretched 2nd stage tanker.  Transit to Mars is then done via a hybrid propulsion system, first a large SEP vehicle would slow push the lander to high Earth orbit where crew would board via a dragon capsule.

Wait, you want to use SEP for transfering a 475t payload from LEO to HEO (e.g. LDHEO)? That would take about 8 years with a 300kw SEP.

« Last Edit: 08/25/2015 02:56 AM by Oli »

Offline philw1776

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Re: MCT Speculation and Discussion Thread 4
« Reply #466 on: 08/25/2015 02:57 AM »
It's late but a quick run on a 225mT to LEO, i.e. >10Km/sec allowing for gravity losses, looks like a 25 engine 12.7 million LBS thrust BFR stage one and a 7 engine stage 2.  Don't like this solution because mass fraction to LEO seems too high, but it's bed time.

S1 avg ISP SL to MECO  330   
S2 vac ISP    380   
1st Stage T/W   1.21   
BFR DIA   12.5   m
MCT Mass   225   mT
1st Stage Tank Length   24.5   m
S1 Propellant Volume   3005   m3
1st Stg Airframe Weight   205   mT
Propellant Weight   3186   mT
S1 Engines Weight   33   mT
S1 Total Weight mT   3424   mT
S1 Total Weight LBS   7.5   Million LBS
S1 empty + S2   1568   mT
DRY Weight   238   mT
%  DRY WEIGHT    7.0%    %
RTLS Propellant   45   mT
RTLS Delta V   0.56   Km/sec
Stage One Km/sec   3.50   Km/sec Rocket Equation
2nd Stage Tank Length   8.5   m
Propellant Volume   1043   m3
Propellant Mass   1105   mT
S2/S1 mass   0.39   
S2 Mass w/MCT   1330   mT
S2 Mass w/MCT   2.9   Million LBS
Calc # Vac Raptors   6.83   1.3 T/W
Stage 2 Thrust   3.8   Million LBS
Stage 2 Km/sec   6.62   Km/sec Rocket Equation
S2 Mars 25mT Cargo    8.1   Km/sec Rocket Equation
TOTAL WT mT   4754   mT
TOTAL WT LBS   10.5   Million LBS
THRUST Needed   12.7   Million LBS
THRUST Needed   56.4   Million Newtons
1st STAGE # ENG    25.0   Eng 16+8+1 = 25
LEO Mass Fract   4.73%    %

I think that even the 1st MCT architecture revealed by SX will be somewhat different than Musk has said so far and by the time it's built, quite different again.  I think the Raptor engine will have 20% or so more thrust than ~ 500 KLBS leading to somewhat fewer engines but still "a lot".
I like the SEP interplanetary haulers myself.  It's just not the "land the whole thing" paradigm that we think Musk was speculating.
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Offline Impaler

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Re: MCT Speculation and Discussion Thread 4
« Reply #467 on: 08/25/2015 04:00 AM »
Isn't PICA supposed to be significantly better than previous TPS materials though? Would that reduce the required dry mass?

PICA is the best ablative, but I intend for the lander to do repeated rapid cargo flights between mars surface and low orbit maybe as many as 100 such round trips before returning to Earth, no ablative can go through that many cycles so I reject it as an option.

The human lander concepts I referred to are all designed for either entry from Mars orbit or aerocapture into orbit + entry from Mars orbit. Granted, gs, heat peak rate and heat load are significantly higher in the aerocapture phase than in the entry phase, although its a lot better than direct entry.

"Parts dropped along the way" are part of the structural mass at entry, which I was refering to. I.e. the entry mass minus the fuel and payload.

The best lander in terms of payload to structural entry mass I cound find is in the pdf attached on page 15, top left. A value of ~1.2. With SIAD, from orbit. It has a total entry mass of only 20t though. The higher the entry mass the worse usually.

With HIAD it may get better, haven't seen the mass break down of such a lander yet. HIAD could be interesting for MCT.

Yes the aerocapture is worse then the subsequent entry, Ideally I would like to avoid it to keep thermal and load requirements to a bare minimum.

The HIAD dose look like it saves a lot of propellant, but I'm concerned that it's basically a disposable system and I'm looking for a repeatable surface-2-orbit shuttle for mars which rules out any disposable systems.  One alternative I'm considering is if an HIAD like equivalent can be produced via a circle of body flaps at the rear of a biconic vehicle.  I'm looking at 13 m diameter with flaps 2 -4 m long effectively creating a diameter of 15-17 m, the flaps would also provide control authority possibly saving more propellant.

Wait, you want to use SEP for transfering a 475t payload from LEO to HEO (e.g. LDHEO)? That would take about 8 years with a 300kw SEP.

Who said 300 kw?  I said large and that it would have DRY mass of up to 250 mt aka an entire BFR launch load, that's a lot more then 300 kw, though I have not done the full number crunch on this vehicle it is by no means ARM or anything that small.  Both the spiral to HEO and the transit to mars could be more then a year.

Offline Owlon

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Re: MCT Speculation and Discussion Thread 4
« Reply #468 on: 08/25/2015 04:44 AM »
Isn't PICA supposed to be significantly better than previous TPS materials though? Would that reduce the required dry mass?

PICA is the best ablative, but I intend for the lander to do repeated rapid cargo flights between mars surface and low orbit maybe as many as 100 such round trips before returning to Earth, no ablative can go through that many cycles so I reject it as an option.

I'm actually pretty sure Elon Musk said at one point (probably the Dragon 2 unveiling) that version 3 of PICA-X would be good for at least 10 Dragon flights, and that they were shooting for as many as 100 in the long term (presumably a version 4 of something). PICA-X version 3 is debuting with Dragon 2, and they were planning on switching to the then-new PICA-X version 2 on Dragon 1 whenever this was said. If they can achieve those sorts of numbers, they can probably get something that works for the necessary 30ish Earth+Mars reentries for MCT or many hundreds from LMO.

Additionally, I'm pretty sure it each version of PICA developed by SpaceX has been lighter than the previous.

Most of that comes from the Dragon 2 unveiling, and I'm thinking the 10/100 flights bit came from a video at the same event with some Q&A recorded by a forum member here.

Offline Oli

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Re: MCT Speculation and Discussion Thread 4
« Reply #469 on: 08/25/2015 05:51 AM »
As the Mars colony is built up the small SEP transit vehicle would be replaced with a large more powerful cargo hauler with extensive habitats aka the mother-ship.  The lander would be fueled in LEO and used to rapidly deliver passengers (100 at a time) to the waiting mother-ship in high orbit and would then travel attached too it as the mother-ships speed is expected to be competitive with the earlier direct flight.  At mars the lander is used to repeatedly ferry between surface and low orbit with containerized cargo being reloaded in orbit.  At Earth the launch of chemical propellant and landers is almost eliminated in favor of launching naked cargo containers and SEP propellant to be loaded onto the mother-ship.

Why not start this way?

Following elements:

- 1.5MW SEP tug capable of transporting 100mt from HEO to Mars orbit (e.g. 1 sol). The SEP tug in the Raftery concept can do that.
- 100mt reusable Mars lander.
- 100mt Habitat.
- 100mt Cargo "container".

A colonial fleet starts in HEO with:

- Cargo containers.
- 1 Habitat.
- 1 brand new Mars lander.

Each with its own SEP tug. All elements are being transfered to Mars. Then:

- Crew lands with brand new lander.
- People from the surface come up with the lander.
- Habitat plus all SEP tugs head home to HEO.
- Lander continues to "ship" Cargo containers from Mars orbit to the surface.

Requires some sort of Lander maintenance on the surface, but the idea is to fly it until its broken. Better start with the high value Cargo :).

Obvious advantages:

- The Lander is far better utilized.
- Far simpler. Only entry from Mars orbit and back.
- Doesn't spend years in deep space.
- No huge volume (living space, space for supplies) required for the crew, as the crew only lands in it.

Anyway, I guess that wouldn't be MCT anymore so its OT  ;).

Offline Vultur

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Re: MCT Speculation and Discussion Thread 4
« Reply #470 on: 08/25/2015 05:59 AM »
PICA is the best ablative, but I intend for the lander to do repeated rapid cargo flights between mars surface and low orbit maybe as many as 100 such round trips before returning to Earth

Ah, I was assuming one landing, one launch direct to Earth.

MCT is supposed to be "land the whole thing".

So only one Earth launch, one Mars entry, one Mars launch, and one Earth entry between servicings (on Earth).

I think PICA-X will be used for MCT.

Offline llanitedave

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Re: MCT Speculation and Discussion Thread 4
« Reply #471 on: 08/25/2015 06:06 AM »

PICA is the best ablative, but I intend for the lander to do repeated rapid cargo flights between mars surface and low orbit maybe as many as 100 such round trips before returning to Earth, no ablative can go through that many cycles so I reject it as an option.



The same flights that bring up fuel and cargo to MCTs in orbit could also bring up replacement heat shields as necessary.
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Offline Vultur

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Re: MCT Speculation and Discussion Thread 4
« Reply #472 on: 08/25/2015 06:24 AM »
As opposed, I believe 50t could be the dry mass of a 820t fully loaded and fueled MCT, note that at most there would be 1.3km/s of ΔV or so a re-entry mass around 230t, and a landing mass approaching 150t.

Does it really need that much delta-v to go from terminal velocity to landing?

Offline Rocket Surgeon

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Re: MCT Speculation and Discussion Thread 4
« Reply #473 on: 08/25/2015 06:26 AM »
With regards to testing the BFR before flight, is it possible for the Launch Pad to double as the full first stage testing site? It could possibly save on costs rather than having to develop new sites. The main hangar could also be used as the final production facility ala the N1 (except this time, with pre-launch testing).

Depending on costs you could build the BFR pieces off site, ship them to the launch pad, and assemble/test them there.

Offline Impaler

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Re: MCT Speculation and Discussion Thread 4
« Reply #474 on: 08/25/2015 06:36 AM »
PICA is the best ablative, but I intend for the lander to do repeated rapid cargo flights between mars surface and low orbit maybe as many as 100 such round trips before returning to Earth

Ah, I was assuming one landing, one launch direct to Earth.

MCT is supposed to be "land the whole thing".

So only one Earth launch, one Mars entry, one Mars launch, and one Earth entry between servicings (on Earth).

I think PICA-X will be used for MCT.

It is supposed to get the mission done as efficiently as possible, I'm making the case that this hybrid architecture is 1) Actually achievable without using pixie-dust and 2) more efficient to boot.  Compared to that any 'supposed to be' is moot.

Online guckyfan

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Re: MCT Speculation and Discussion Thread 4
« Reply #475 on: 08/25/2015 06:49 AM »
It is supposed to get the mission done as efficiently as possible, I'm making the case that this hybrid architecture is 1) Actually achievable without using pixie-dust and 2) more efficient to boot.  Compared to that any 'supposed to be' is moot.

To be clear on this. You are arguing that you know better than SpaceX and Elon Musk?

Online MikeAtkinson

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Re: MCT Speculation and Discussion Thread 4
« Reply #476 on: 08/25/2015 07:15 AM »
So I propose a vehicle designed to act as a ferry between the surface and orbit, able to land 100 mt and return to orbit with 25 mt PLUS enough propellant to make another landing on Mars.  At 75 mt dry mass for the vehicle the landing requires 40 mt of propellants.  The launch mass is 400 mt of which 260 is assent (4.1 km/s) propellant, 25 cargo, 40 return propellant (.8 km/s) and 75 the vehicle dry mass.  Total propellant load is just 300 mt meaning propellant production on Mars can be a fraction of that needed for Direct return.  The structural mass fraction at launch is a very conservative 18.75% and payload is 46% of entry mass (including propellant) both very reasonable figures.

The lander would launch from Earth atop a 2 stage launch vehicle and would be loaded with ~50 mt of propellants providing ~900 m/s DeltaV to be used for emergency separation and propulsive landing in the even of an abort.  This would put the total launch mass at 225 mt.  And the lander could then be fully topped off via another launch of and transfer of 250 mt or propellant from a stretched 2nd stage tanker.  Transit to Mars is then done via a hybrid propulsion system, first a large SEP vehicle would slow push the lander to high Earth orbit where crew would board via a dragon capsule.  Then the SEP would separate and the lander would make a lunar-Earth slingshot burn of ~1 km/s for fast transit to Mars and use propulsive capture and perhaps some airobraking to reach low mars orbit leaving just enough propellant for landing.  On the surface the cargo is unloaded and a 25 mt return cabin is loaded in it's place.  The SEP system would make a slow transit to mars arriving well after the lander and wait in low orbit for rendezvous after which it would bring the lander back to high Earth orbit, crew would disembark again via Dragon capsule and then the unoccupied combined vehicle would do a down-spiral to low Earth orbit where it would be refueled again and use it's propulsive capacity to reduce entry velocity to the 3.6 km/s velocity it can tolerate followed by landing on Earth for reloading and reintegration, SEP remains in orbit ready to be reused.


Total BFR launches would be 6, one for the lander, 2 tankers of chemical propellant, 1 for SEP hardware, 2 estimated for SEP propellant.  In addition 2 trans-lunar Dragon capsule launches likely on Falcon Heavy.  Each subsequent launch would need 5 launches when SEP is reused.  The cycle would be one round trip every 2 synods.  If philw1776 can run the numbers of the size of the BFR needed to do a 225 mt launch that would be most informative to see how it compares.

I really like this architecture. It has high reusability and great flexibility, obvious changes for lunar, asteroid and Mars moons missions.

Perhaps you should write it up as a paper.

Offline Impaler

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Re: MCT Speculation and Discussion Thread 4
« Reply #477 on: 08/25/2015 08:59 AM »
It is supposed to get the mission done as efficiently as possible, I'm making the case that this hybrid architecture is 1) Actually achievable without using pixie-dust and 2) more efficient to boot.  Compared to that any 'supposed to be' is moot.

To be clear on this. You are arguing that you know better than SpaceX and Elon Musk?

I'm confident they will reach the same conclusion I have after figuring out (possibly the hard way) that Direct Earth return is impossible.  People are being way to fast to grasp at nebulous ideas and speculations even if they come from Musk as THE ONE AND ONLY way it will be done,  Musk like any good programmer tries to think of the simplest possible system he thinks could possibly work, we saw that with F9 reuse plans, they are now WAY more complex then originally planned, MCT will be the same.




Perhaps you should write it up as a paper.

I might do that, though I still need to work out the SEP transit vehicle.

Offline JamesH

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Re: MCT Speculation and Discussion Thread 4
« Reply #478 on: 08/25/2015 09:17 AM »
It is supposed to get the mission done as efficiently as possible, I'm making the case that this hybrid architecture is 1) Actually achievable without using pixie-dust and 2) more efficient to boot.  Compared to that any 'supposed to be' is moot.

To be clear on this. You are arguing that you know better than SpaceX and Elon Musk?

Since the amount of information known about MCT outside of SpaceX can fit on a postage stamp, written with a crayon, and that SpaceX plans change considerable over time, how do we know this isn't the plan? No point in saying "MUSK SAID THIS", because aforesaid plans change.

Online guckyfan

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Re: MCT Speculation and Discussion Thread 4
« Reply #479 on: 08/25/2015 09:39 AM »
To be clear on this. You are arguing that you know better than SpaceX and Elon Musk?

Since the amount of information known about MCT outside of SpaceX can fit on a postage stamp, written with a crayon, and that SpaceX plans change considerable over time, how do we know this isn't the plan? No point in saying "MUSK SAID THIS", because aforesaid plans change.

I hear that argument a lot. But it really is not true. There is a lot of info that is consistent over time.

My base argument is that complex architectures are not likely to get anywhere near the aimed for cost level. At least not before hundreds if not thousands of MCT go to Mars at every synod. SpaceX is aiming to build at least a permanently manned base and they could not afford designing and building such a complex architecture.

I don't deny that your architecture is an interesting one. I only say it is not compatible with what we know about the SpaceX architecture and this is the MCT thread. We have a whole section on this forum for Mars. Your Mars architecture should be discussed there IMO.

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