spacex SpaceX's "BFR/BFS/MCT" Thread -- Facts, Conjecture, Math and Hypothesis

[Unobscured Vision]

@DocM is correct about BFR's S1 landing on an uprated ASDS. Falcon-9 isn't a small rocket (3.66m); but at the same time BFR is going to be massive. We're talking probably 15m. The kicker is that once most of its' fuel is spent, it's actually not all that heavy. That's why we calculate Dry Mass and Wet Mass separately, and how the Crawlers can haul big hooey beasts like Apollo/Saturn out to 39A (even at an angle) and they don't fall over so long as you're driving carefully.

 

 

[Joe2mercs]

On 23 May 2016 at 2:56 PM, DocM said:

Further,

 

You say 4% to orbit for F9 without reuse, 2.2% with which vastly overstates the reuse penalty at 45%.

 

Musk estimated the penalty at 30% for a return to launch site landing (RTLS), and 15% for a drone ship (ASDS) landing. RTLS requires 3 burns; braking/boostback (the largest burn of all), reentry and landing. ASDS requires 2 burns; reentry and landing.

 

SpaceX just demonstrated a 3rd variant: suicide-slam, which has a penalty closer to 10% for LEO and an ASDS landing.

 

Now consider a BFS that boosts 5%+ to LEO and has a 10% penalty for a suicide-slam ASDS landing. That's over 4.5% mass to orbit, not 2.2%.

 

 

The higher ISP of the methane engine is marginal over that of RP1 if both are closed cycle designs.  Historically there has been an engine weight penalty associated with using fuels of lighter density mainly due to the need for more powerful fuel pumps.  It remains to be seen whether the extraordinary Merlin 1D thrust to weight ratio (I think about 180:1 with super chilled fuel mix) can be matched by the Raptor design.  Very much a case of comparing apples with oranges but a useful yardstick.  The fact remains that a methane rocket will be a minimum 11% heavier than its super chilled RP1 counterpart just in terms of fuel mix alone.  The dry mass of methane rocket will also be significantly heavier due to a) the extra structure/volume needed to accommodate the less dense methane b) the additional LOX needed for efficient methane ignition c) heavier engines, due to larger fuel pump turbines and the need for extensive reuse with minimal overhaul.  Given the weight penalties I think it unlikely that such a booster will get anywhere near 5% payload to LEO; not that it will ever be flown as an expendable vehicle.  Elon is much more interested in the economics of payload to LEO.  The MCT booster represents a much greater financial investment in terms of development and individual vehicle build cost over that of the Falcon series.  Furthermore, there appears to be minimal likelihood of fee paying launch contracts that would help alleviate the costs associated with the MCT booster development.  

The MCT booster is unlikely to set new records for payloads to LEO as a % of take off mass.  The Falcon heavy is often quoted as 53 tons to LEO in expendable mode and I would be surprised if it ever exceeds 30 tons in reusable mode.  The MCT launcher will never be used in 'suicide slam' mode, it is just far too valuable.  Mark my words the MCT booster will plod up and down to LEO lifting about 2% of its take off mass as payload.  Ironically it is the Falcon series that is ushering in an era whereby rockets will no longer be assessed on their ability to simply lift mass to LEO but on how cheaply they can do it.  The cost of payload to orbit includes fuel, refurbishment (between flights), initial build and amortisation of the development costs over the vehicle lifetime, along with launch complex fees.  The MCT booster will be big and for the reasons outlined above I stand by my figure of 5,100 tons take off mass.

[Joe2mercs]

On 29 May 2016 at 6:08 AM, Joe2mercs said:

The higher ISP of the methane engine is marginal over that of RP1 if both are closed cycle designs.  Historically there has been an engine weight penalty associated with using fuels of lighter density mainly due to the need for more powerful fuel pumps.  It remains to be seen whether the extraordinary Merlin 1D thrust to weight ratio (I think about 180:1 with super chilled fuel mix) can be matched by the Raptor design.  Very much a case of comparing apples with oranges but a useful yardstick.  The fact remains that a methane rocket will be a minimum 11% heavier than its super chilled RP1 counterpart just in terms of fuel mix alone.  The dry mass of methane rocket will also be significantly heavier due to a) the extra structure/volume needed to accommodate the less dense methane b) the additional LOX needed for efficient methane ignition c) heavier engines, due to larger fuel pump turbines and the need for extensive reuse with minimal overhaul.  Given the weight penalties I think it unlikely that such a booster will get anywhere near 5% payload to LEO; not that it will ever be flown as an expendable vehicle.  Elon is much more interested in the economics of payload to LEO.  The MCT booster represents a much greater financial investment in terms of development and individual vehicle build cost over that of the Falcon series.  Furthermore, there appears to be minimal likelihood of fee paying launch contracts that would help alleviate the costs associated with the MCT booster development.  

The MCT booster is unlikely to set new records for payloads to LEO as a % of take off mass.  The Falcon heavy is often quoted as 53 tons to LEO in expendable mode and I would be surprised if it ever exceeds 30 tons in reusable mode.  The MCT launcher will never be used in 'suicide slam' mode, it is just far too valuable.  Mark my words the MCT booster will plod up and down to LEO lifting about 2% of its take off mass as payload.  Ironically it is the Falcon series that is ushering in an era whereby rockets will no longer be assessed on their ability to simply lift mass to LEO but on how cheaply they can do it.  The cost of payload to orbit includes fuel, refurbishment (between flights), initial build and amortisation of the development costs over the vehicle lifetime, along with launch complex fees.  The MCT booster will be big and for the reasons outlined above I stand by my figure of 5,100 tons take off mass.

The numbers speak for themselves. JCSAT 14 (mass of about 4,700Kg) and THAICOM 8 (mass of about 3,800kg) were both recently placed into geostationary transfer orbit using the latest Falcon F9FT and the first stage was successfully recovered on the barge after a 'suicide slam' landings in both cases.  The use of such landing profiles indicates that these landings were at the very limit of fuel margin.  These payload figures for reusable launches compare to the stated GTO specification of 8,300Kg for expendable F9FT launches.  Based on this real data it would appear that the payload penalty for recovery of the first stage is roughly half.  When reuse becomes the norm and the Falcon heavy enters service I think that to post payloads to LEO and GTO reflecting expendable launches will become completely irrelevant as nobody, but the insane, will contemplate paying a massive premium of throwing the rocket away. 

[DocM]

"Based on this real data it would appear that the payload penalty for recovery of the first stage is roughly half."

 

Incorrect. ~25-30% for RTLS, 8-10% for ASDS landings.

 

Super-Slam drastically cuts propellant use by totally eliminating the boostback burn, minimizing the reentry burn, and using a high-G 3 engine landing burn to minimize gravity losses. 

[DocM]

Good NSF article in the run-up to September's Mars architecture presentation at IAC in Guadalajara, Mexico. 

 

https://www.nasaspaceflight.com/2016/06/reusability-successes-earth-proving-ground-mars/

 

SpaceX's Lars Hoffman, Senior Director of Government Sales

 

Quote

>
>
"In 2018 we're looking to send a Dragon to the surface of Mars. There are no runways and theres no water (to allow for) splashdowns on Mars, so we've decided to perfect the vertical landing capability on Earth to allow us to build that transportation network on Mars to be able to land our capsules and other vehicles on Mars later on."

 

L2 image based on direct data. Leg count guessed vs booster size.

[Unobscured Vision]

Makes sense that they're going the "N1 Route". Concentrically-arranged engines on an outer ring, some gappage and then more of the business inside. They're going with whatever the physics and thermodynamics simulations say will work the best.

[DocM]

IF they stick with 15 meters as the core diameter, that's 2 meters wider than a US Ohio-class ballistic missile submarine.

 

IF its aspect ratio (fineness) is in the 8:1-10:1 range for rockets, which has advantages WRT stage landings, it'll be 15m wide x 120-150m tall. Saturn V: 10.1m wide x 110.6m tall (10.95:1)

 

By comparison; Atlas V is 15:1 and Falcon 9 is 18.7:1, depending on the fairing length. Flying pencils.

Edited June 3, 2016 by DocM

[Unobscured Vision]

Woof.  

 

Let's see ... optimum AR for 15m ... need some factors here:

 

- Let's consider that we don't use those sharp angles like Saturn did (because they're aerodynamically not good, and will impair performance), so it'll be 15m all the way up -

- Less mass + materials/m^2, because this is not the 60's and we actually have better materials and processes now -

- Fuel mass at liftoff will be less because Methane is lighter than RP by a factor of almost two. LO2 capacity needs to be increased but it's not going to substantially impact weight (10%) -

 

-- Our net weight savings is around 35% relative to Saturn V at the same height from ground (348'). That's gonna help a lot. (And yes, I know, Saturn V used LH/LO2 in the S2 and SIV-B.)

-- Our net aerodynamic efficiency is improved by 12%~15% versus Saturn V simply by using modern aerodynamic modeling and practices. This translates into a net gain of 3%~5% performance improvement of the platform as a whole. And yes, it does make a difference at those levels of power. It's an extra 15 tons of payload.

 

Soooo ... optimum Aspect Ratio? Aerodynamically, that'll be 10.4~11.6:1. The "sweet spot" is 11.15:1.  

[DocM]

Now, what happens if its tanks are not made of friction stir welded aluminum-lithium alloy like Falcon 9 or Falcon Heavy?

 

NASA's been doing a lot of work on large scale composite cryogenic tanks, and we know how strongly they've been supporting commercial outfits via Space Act Agreements. NASA says these can be 30% lighter than Al-Li alloy, and IIRC they qualified one 2 years ago.

 

Just sayin'

 

Edited June 3, 2016 by DocM

[Unobscured Vision]

53 minutes ago, DocM said:

Now, what happens if its tanks are not made of friction stir welded aluminum-lithium alloy like Falcon 9 or Falcon Heavy?

 

NASA's been doing a lot of work on large scale composite cryogenic tanks, and we know how strongly they've been supporting commercial outfits via Space Act Agreements. NASA says these can be 30% lighter than Al-Li alloy, and IIRC they qualified one 2 years ago.

 

Just sayin'

 

Any weight saved is more weight that can be added to the payload -- and it's less weight that the S1 has to contend with when it's landing the stage.  Every bit helps.

[Unobscured Vision]

Too bad the Carbon composites didn't hold up to extremes of temperature or pressure. Or much of any other kind of demand that was put on them ... even the CF-Impregnated stuff didn't hold up when eeeeeveryone was proclaiming it to be "the future of structural materials". Pressurize it past 10 atmospheres and watch it pop like a balloon ... and 480psi really isn't much.

 

Ahhh, the all-mighty hype machine ... how we love to hate ye. 

[Unobscured Vision]

Thinking about the Aspect Ratio some more, and the weight reduction from using lighter tanks as well as the aerodynamics of the BFR needing to be a slightly longer rocket due to weight reduction compared to Apollo/Saturn.

 

They'd almost have to deliberately push it to the long edge of the AR's "sweet spot". They're gonna be running 12~18 mlbf (mission-dictated) on one end, and they need to balance the other end accordingly. Naturally, they'll want to use aerodynamics to stabilize as much as possible while heading uphill to prevent it somersaulting/frontflipping/cantering. A longer chassis will provide that negative/positive pressure to even things out along the hull ... and a curious idea comes to mind too.

 

Ducting. 

 

Since there's all of that force being exerted (possibly more than desired), why not build in the capability to channel that force where it's needed or not needed? It's already there ... use it!  Take some cues from the VTOL Jets around the world, and design some wind channels that can be opened and closed as needed to help balance those aerodynamic forces on the chassis as it's heading uphill and downhill. So instead of pushing its' way through the atmosphere, it slices its' way through. This could add another 15% performance improvement by itself if it's done right.  And really, it wouldn't impact the fuel capacity at all; in fact, I bet it would help stretch it out even further. Win-win!

[DocM]

1 hour ago, Unobscured Vision said:

Too bad the Carbon composites didn't hold up to extremes of temperature or pressure. Or much of any other kind of demand that was put on them ... even the CF-Impregnated stuff didn't hold up when eeeeeveryone was proclaiming it to be "the future of structural materials". Pressurize it past 10 atmospheres and watch it pop like a balloon ... and 480psi really isn't much.

 

Ahhh, the all-mighty hype machine ... how we love to hate ye. 

Already solved. 3 years ago NASA hit -423°F LH2 at 135 psi, rocket propellant tanks only need 50+ psi, and they could do 20 full cycles then. An outfit in WA figured out how to do the joints.

[Joe2mercs]

On 3 June 2016 at 7:09 AM, Unobscured Vision said:

Woof.  

 

Let's see ... optimum AR for 15m ... need some factors here:

 

- Let's consider that we don't use those sharp angles like Saturn did (because they're aerodynamically not good, and will impair performance), so it'll be 15m all the way up -

- Less mass + materials/m^2, because this is not the 60's and we actually have better materials and processes now -

- Fuel mass at liftoff will be less because Methane is lighter than RP by a factor of almost two. LO2 capacity needs to be increased but it's not going to substantially impact weight (10%) -

 

-- Our net weight savings is around 35% relative to Saturn V at the same height from ground (348'). That's gonna help a lot. (And yes, I know, Saturn V used LH/LO2 in the S2 and SIV-B.)

-- Our net aerodynamic efficiency is improved by 12%~15% versus Saturn V simply by using modern aerodynamic modeling and practices. This translates into a net gain of 3%~5% performance improvement of the platform as a whole. And yes, it does make a difference at those levels of power. It's an extra 15 tons of payload.

 

Soooo ... optimum Aspect Ratio? Aerodynamically, that'll be 10.4~11.6:1. The "sweet spot" is 11.15:1.  

The Arodynamic efficiency of rocket is not nearly as important as it's dry mass to volume ratio.  A rocket can lift quite slowly through the thick lower layers of atmosphere just so long as its engines are of high ISP.  Note that the point of maximum aerodynamic pressure (MaxQ) varies according to rocket design (how much additional strain can the structure sustain over and above its mass and maximum vibration) and mission profile (MaxQ is not he same for all flights but is a function of altitude and speed). A fast ascent hits a MaxQ (a) at low altitude and uses lots of energy pushing through thick air whereas a slower ascent hits a MaxQ (b) at high altitude.

 

Some have said that because methane is lighter then the rocket will be lighter.  This is a wrong assumption.  When burning 1kg of RP1 a rocket must also use 2.6kgs of LOX.  Likewise when burning 1kg of Methane 3.8kgs of LOX is burnt.  So for engines of the same efficiency (ISP) the methane rocket will be 33% heavier I terms of fuel mix alone (3.6kgs RP1/LOX versus 4.8kgs Methane/LOX). To this must be added the weight of the greater sized tanks as methane is less dense; 422g per litre for liquid Methane versus about 1kg per litre for super chilled RP1.  Kg for Kg the fuel tanks needed for liquid Methane are about 2.4 times bigger than those for super chilled RP1! The increased ISP (efficiency) of closed cycle engines means that the rocket can be slightly reduced in size for a given payload requirement but payload for payload the Methane rocket (raptor) will always be heavier than it's equivalent RP1 rocket (RD170 for example) due to the LOX burden alone.  Add in the mass of extra materials for the larger tank sizes with attendant support structures and a methane rocket is going to be that bit heavier again contriving to reduce the net payload to LEO as a percentage of take off mass.

 

The aspect ratio is function of aerodynamics, dry mass and engineering (as the efficiency of Merlin engines improved it was cheaper, from an engineering point of view, for SpaceX to lengthen the Falcon rocket to incorporate the additional fuel load than it was to keep its height and make it fatter). The Saturn V's shape was arrived at by incorporating  tried and tested stages from other rockets.   At this point it is useful to note that it's first stage constituted nearly 80% of the Saturn V take off mass (about 3000 tons).  The second stage was in turn 80% of the remaining mass, with the last 4% being the 140 ton payload (third stage) in LEO.   If we assume a diameter of 15m for the MCT booster we can easily determine its height as a function of its take off mass and vice versa.  We know that the mass ratio of methane to LOX is 1:3.8 which translates into volume terms of a ratio of about 2:3 (LOX at 1.141kg/litre is so much denser than liquid methane 0.462kg/litre).  I think that SpaceX can achieve a methane rocket structure of 5% dry mass.  If I calculate the height of my own preferred speculative MCT booster with a take off mass of 5,100 tons:-  I make the assumption that the first 5 meters will be engines, plumbing and associated thrust structure;  the next 30 meters contain a conjoined tank (sharing a common partition) of about about 808 tons methane (~40% by volume) and a LOX tank of about 3068 tons (~60% by volume). Such a first stage would have mass of 4080 tons and be 35 meters tall including the engines.  So when assessing rockets that are 100+ meters tall use the rule of thumb of 100 tons per meter to approximate the entire lift off mass.  Talk of an aerodynamic 'sweet spot' rocket of an aspect ratio of about 11 (of a 15 meter diameter rocket) implies a rocket 165 meters tall and perhaps some 16,000 tons;  perhaps a tad unrealistic.

[Draggendrop Veteran]

^  @Joe2mercs     You have put a lot of effort into your posts, well done.....and welcome to Neowin.   

[Joe2mercs]

On 07/06/2016 at 5:01 AM, Draggendrop said:

^  @Joe2mercs     You have put a lot of effort into your posts, well done.....and welcome to Neowin.   

Thanks for the approbation and welcome message.

[Unobscured Vision]

Welcome indeed.

 

I'd like to reserve my response until SpaceX officially announces the particulars of BFR/MCT this summer (possibly at the big conference this month or next month?) ... there's still a lot of unknowns about Raptor's performance right now. The only really solid source here (Wikipedia) has a lot of conflicting and outdated information, and I'd like some updated (and public) numbers about performance before saying much more about that.

 

As for the next part of BFR/MCT, "how tall/wide/etc will it actually be?" ... well, that's up for debate too. We're all fairly certain it'll be a 15-meter hull. Following how SpaceX "does things", and generally those of us who are aware of Engineering practices (as you obviously are, J2M), we can certainly put it in the ballpark. I stated my criteria for the conclusions I came up with in my post, and I stand by it. Does it make me some kind of "authority"? Of course not. Engineers (and even Engineers-in-training like myself) all have different ideas and opinions about how best to go about getting hardware designed and built within certain parameters. The information that I myself have at that particular time says that it could be done such-and-such a way, and I presented a solution as I saw it. I'm open to correction and peer review, as any Engineer should be; but I stand by my results.  An Aspect Ratio of 10.4~11.6 is completely legitimate for 15m. I called the "sweet spot" at 11.15 to allow BFR's S1 to carry extra fuel and to accommodate Stability Ducting during atmospheric travel during launch; but SpaceX will likely go conservative with that number to further avoid somersaulting/flipping without needing active control.

 

I of course don't expect SpaceX to read anything I've typed and simply show up on my doorstep as if I were the "Golden Child" or any such nonsense; it's an intellectual pursuit that I enjoy taking part in whilst I'm in College and I'm happy to participate in the discussion. I'm studying to be an Engineer. I want to work for SpaceX.  

[Joe2mercs]

Just had another speculative thought.  We know that a sphere is most efficient structure to contain pressurized fluids/gases.  From this it is likely that the diameter of the MCT booster will be dictated by the diameter of it's second stage which in turn will be dictated by the diameter of the spherical tank needed to contain the methane in that stage.  Since Space X is all about the economy of modular structures my guess is that the two domes needed for the second stage methane spherical tank would be also used to cap the ends of the LOX tank along with both the methane and LOX tanks of stage one.  Effectively the diameter of the stage two tank (and the rocket itself) will be dictated by the lift mass of the rocket.  Here are some speculative figures:-

Take off mass     Rocket diameter  Stage 1 + Stage 2

(metric tons)        (meters)                  (meters) 

2000                            7.4                          69.2

5000                          10                             90

8600                          12                           106

16,600                       15                           130

The total height of the rockets will depend on the size of the third stage/LEO payload (assume 2% that of take off mass)

[DocM]

SpaceX's MO is to use common bulkhead tankage, and common stage diameters. That's old news.

 

Also: only 2 stages, with the BFS  spacecraft doing duty as the upper/departure/return stages and they'll "land the whole thing" on Mars. BFS will be pushed just past MAXQ by S1 which stages early for easier reuse (see F9), doing the rest of orbital insertion itself and being refuelled there by tankers. Musk confirmed this on the public side earlier this year.

 

The only way to interpret this is that BFS is a very large  "battlestar" vehicle. NSF's editor was given data by SpaceX confirming this, and partly shared it on the L2 private forum. On the public side he reported literally falling off his couch.

 

So, if they deliver 5-6% by now dry mass to orbit, by the time it refuels and departs it's back up near launch wet mass with a huge deltaV. Now, how high is that departure orbit? LEO? L1/2?

Edited June 8, 2016 by DocM

[Unobscured Vision]

I'm gonna go with LEO for the TLI/TMI, like Saturn/Apollo did. At least until the L1/L2 facilities are up and running. Then afterwards, yeah, SpaceX will switch to a L1/L2 TMI model.

[DocM]

Summary of known and implied info

 

Stated

 

BFR/BFS: huge, makes Saturn V look "small"

 

2 stages; BFR booster and BFS spaceship/upper stage

 

Raptor engines: methane/LOX, full flow staged combustion, largely 3D printed

 

Liftoff thrust: ~15,000,000 lbf
GLOW: public estimates are 5,000 to 6,000 tonnes

 

Landed cargo on Mars: ~100 tonnes
Return cargo from Mars: ~25 tonnes
Crew: up to 100

 

Conops: BFR stages low, lands for reuse; BFS pushes to orbit (?how high?), is refuelled then departs to Mars. "...land the whole thing"

 

Implied publicly: 

 

15 meter core for BFR and BFS.

 

BFS is a battlestar, a massive manned spacecraft which would  dwarf the Shuttle.

 

Musk mentioned an easily deployable cargo module. Interpreted as the lower portion of BFS is a utility module which remains on Mars. Similar to the (former SpaceX intern) Max Fagin supersonic retro-propulsion thesis defense. See attached.

 

Info from new Washington Post interview,

 

Washington Post....

 

NET dates

 

2018: Red Dragon 1
2020: Red Dragon 2 & 3 (minimum)
2022: BFS cargo/test mission to Mars
2024: crew BFS launched to Mars
2025: BFS & crew land on Mars

 

Musk: 

Quote

I'm so tempted to talk more about the details of it. But I have to restrain myself. 
>
...the first mission wouldn't have a huge number of people on it because if something goes wrong, we want to risk the fewest number of lives as possible.
>
But I do want to emphasize this is not about sending a few people to Mars.... It's about having an architecture that would enable the creation of a self-sustaining city on Mars with the objective of being a multi-planet species and a true space-faring civilization and one day being out there among the stars.

Max Fagin concept

 

[DocM]

Also notable: 

 

Musk speaks of a "fleet" leaving for Mars at each synod after the test and setup missions, likely a mix of passenger and cargo BFS's. 

[DocM]

IAC 2016 PLENARIES AND HIGHLIGHT LECTURES

 

Date: Tuesday September 27

 

Time: 13:30-14:30 CDT (14:30-15:30 EDT)

 

Colonizing Mars -- A deep technical presentation on the space transport architecture needed to colonize Mars (SpaceX late breaking)
 

[Joe2mercs]

On 08/06/2016 at 0:41 PM, Unobscured Vision said:

Welcome indeed.

 

I'd like to reserve my response until SpaceX officially announces the particulars of BFR/MCT this summer (possibly at the big conference this month or next month?) ... there's still a lot of unknowns about Raptor's performance right now. The only really solid source here (Wikipedia) has a lot of conflicting and outdated information, and I'd like some updated (and public) numbers about performance before saying much more about that.

 

As for the next part of BFR/MCT, "how tall/wide/etc will it actually be?" ... well, that's up for debate too. We're all fairly certain it'll be a 15-meter hull. Following how SpaceX "does things", and generally those of us who are aware of Engineering practices (as you obviously are, J2M), we can certainly put it in the ballpark. I stated my criteria for the conclusions I came up with in my post, and I stand by it. Does it make me some kind of "authority"? Of course not. Engineers (and even Engineers-in-training like myself) all have different ideas and opinions about how best to go about getting hardware designed and built within certain parameters. The information that I myself have at that particular time says that it could be done such-and-such a way, and I presented a solution as I saw it. I'm open to correction and peer review, as any Engineer should be; but I stand by my results.  An Aspect Ratio of 10.4~11.6 is completely legitimate for 15m. I called the "sweet spot" at 11.15 to allow BFR's S1 to carry extra fuel and to accommodate Stability Ducting during atmospheric travel during launch; but SpaceX will likely go conservative with that number to further avoid somersaulting/flipping without needing active control.

 

I of course don't expect SpaceX to read anything I've typed and simply show up on my doorstep as if I were the "Golden Child" or any such nonsense; it's an intellectual pursuit that I enjoy taking part in whilst I'm in College and I'm happy to participate in the discussion. I'm studying to be an Engineer. I want to work for SpaceX.  

Sage words indeed, an old head on young shoulders!   It is fun to speculate and the chances are good that we are all going to be wrong :-)  However the only person to never put a foot wrong never put a foot out.  My own speculations are based on extrapolations of real numbers and experience.  Here is good one....take a 330ml can of Coke and weigh it (~344g). Drink the contents and weigh the can again (~14g).  This ratio of the empty weight (4%) to the former full weight is what Space X engineers are aiming to get close to when they design their rockets; includes rocket engines, thrust frame, plumbing, pressure tanks and fairings.  I wish you well in your pursuit of your goal at Space X.  Speculation grounded in reality helps to sort out the considered ideas from the fantasies.  One of my heroes in Aerospace was John Houbolt,  the chap credited with the concept of Lunar orbit rendezvous.  His speculation made the Moon landing using just one Saturn V possible.  Werner von Braun, to his great credit, recognized Houbolt's unique insight even though many of his own team had dismissed Houbolt's ideas as unworkable fantasy.

[DocM]

The 4% mass fraction may make sense of you're using the same construction material as current launchers; aluminum-lithium alloy, and fineness ratio.  That's not a given if they go clean sheet, and NASA has recently built large tanks made of cryo-capable composites which were 30%+ lighter and stronger. 

 

Another item:

 

did some crunching WRT the potential explosive yield of a failed BFR and what kind of blast radius, overpressures etc. would be a result.

 

Less than a 5 kt tactical nuke airburst, a lot less. Perhaps 1.5 - 2.0 kt.  Saturn V was estimated at ~0.5 kt by NASA.

 

A worst case 5 kt nuclear airburst, an order of magnitude greater than Saturn V, at an optimised distance over the pad would deliver a 1.5 psi overpressure (broken windows) at about 2.75 km from the pad. Boca Chica village is 3.22 km and downtown Brownsville is 35 km.

 

This is for a detonation, not the conflagration which is the usual liquid launcher failure mode. Buy shatter-proof windows if you're in the village.

 

Also, as soon as it clears the tower it'll make a gravity turn East, increasing the distance between it, the village and Brownsville - further reducing the effects. By MaxQ it's so far away as to be inconsequential.

 

It could launch from Boca Chica.

Edited June 20, 2016 by DocM