Author Topic: IAC 2017 -- BFR v0.2 - DISCUSSION THREAD 3 (Post Speech)  (Read 609700 times)

Offline Nathan2go

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Re: IAC 2017 -- BFR v0.2 - DISCUSSION THREAD 3 (Post Speech)
« Reply #40 on: 10/07/2017 02:32 am »
No need for long duration storage of cryo-propellants?

Did you folks notice the very low speed at which the ship switches from aero-braking to retro-propulsion?  That means that the amount of propellant needed for landing is very low, so storable propellant could be practical.

On Mars, the retro-burn starts at Mach 2.3 = 620 m/s = 1380 mph (which is less than half of what JPL said a 100t, 13' diameter capsule did). Assuming 13% gravity loss (3 gees of thrust and .4 gees of gravity), 10% reserves, and Isp=300s, the landing burn will have a Mass Ratio (Mr) of 1.30.  So the landing propellant is 70t for 150t payload and 85t dry wt.

On return to Earth, aero-braking should get the terminal speed down by 1 order of magnitude, since the air density is 2 orders magnitude more (drag and lift are proportional to density*V^2), perhaps 140 mph.  So the landing propellant requirement is very low.

That means that the ship could use storable propellants for landing, even if all the landing propellant for the round trips is provided at Earth.  Maybe it would be Super-draco derived engines, presumably burning hydrazine and N2O4.

But it would also be possible to make storable propellant on Mars:  propane, which is also storable at room temp, can probably be made in the same chemical reactor that makes methane, if a separator is provided.  N2O4 could be made using Nitrogen from Mar air, but since it is only about 1% N2, maybe it would be easier to bring N2O from Earth (also an easily storable liquid), then react it with more O2 to form the desired N2O4

For on-orbit refilling, a full-tanker boil-off rate of 1% per week is a reasonable goal (for flights every 2 weeks, 10 weeks per Mars departure).  But this can be achieved with a specially insulated propellant depot, and need not impact the ship design.

Of course Musk's initial plan will be the simplest, with more complexity (like propellant depots or storable landing propellant) added when they starting to get deeper into the system design.
« Last Edit: 10/07/2017 02:56 am by Nathan2go »

Offline Robotbeat

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Re: IAC 2017 -- BFR v0.2 - DISCUSSION THREAD 3 (Post Speech)
« Reply #41 on: 10/07/2017 02:42 am »
Methane and oxygen are in a different class from hydrogen when it comes to cryogenics. Both methane and oxygen can be passively stored in space without boiloff. Practically speaking, hydrogen can't.

Hydrogen boiling point: 20K
Oxygen boiling point: 90K
Methane boiling point: 112K

Being MUCH higher above absolute zero means it's also much, MUCH easier and more efficient to actively cool both methane and oxygen than hydrogen. Multiple times less energy required to refrigerate them. That's if it's even necessary to refrigerate them at all.

SpaceX is NOT going to use storables (i.e. hypergols) for BFR. They don't use them for the F9 booster, and they sure as heck won't use them for BFR. The handling and regulatory headaches alone would increase the cost significantly, and they're also about an order of magnitude or two more expensive than methane and oxygen. And they're also WAY harder to make on Mars. Not happening.
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Offline Zed_Noir

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Re: IAC 2017 -- BFR v0.2 - DISCUSSION THREAD 3 (Post Speech)
« Reply #42 on: 10/07/2017 08:22 am »
....
The one thing that is not possible or practical is an escape system at either Mars or the moon.  The reason for this is quite obvious.  All escape systems depend on being rescued after the fact.  This is possible just about any place on earth.  Even the escape system for the B-58 Hustler could sustain somebody floating in the water are in the Arctic for up to 4 days.  There is nobody on the moon or Mars to rescue anybody.  Even if you successfully escaped the BFR you would still inevitably die.  That’s the way it would be for the foreseeable future.  Only after the BFR has made so many trips to those destinations to prove its reliability would rescue in those places be possible but then there would not be a need for a rescue system on the BFR.

That is not quite true. If there is a fueled BFS available on the Moon or Mars. A rescue mission could be mounted with the BFS on a ballistic hop to pick up the survivors.

Yes, it is the equivalent of sending a cruise ship to do the job of a life boat.  :o
« Last Edit: 10/07/2017 09:07 am by Zed_Noir »

Offline octavo

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Re: IAC 2017 -- BFR v0.2 - DISCUSSION THREAD 3 (Post Speech)
« Reply #43 on: 10/07/2017 09:02 am »
....
The one thing that is not possible or practical is an escape system at either Mars or the moon.  The reason for this is quite obvious.  All escape systems depend on being rescued after the fact.  This is possible just about any place on earth.  Even the escape system for the B-58 Hustler could sustain somebody floating in the water are in the Arctic for up to 4 days.  There is nobody on the moon or Mars to rescue anybody.  Even if you successfully escaped the BFR you would still inevitably die.  That’s the way it would be for the foreseeable future.  Only after the BFR has made so many trips to those destinations to prove its reliability would rescue in those places be possible but then there would not be a need for a rescue system on the BFR.

That is not quite true. If there is a fueled BFS available on the Moon or Mars. A rescue mission could be mounted with the BFS on a ballistic hop to pick up the survivors.

Yes, it is the equivalent of sending a cruse ship to do the job of a life boat.  :o
How does the unpiloted bfs plot this hop? I would think this a tall ask in an emergency without some sort of GPS analog around the moon / Mars first?

Offline john smith 19

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Re: IAC 2017 -- BFR v0.2 - DISCUSSION THREAD 3 (Post Speech)
« Reply #44 on: 10/07/2017 09:26 am »
ITS first stage PMF was given as 0.96.  This rocket is going to be smaller, so I don't see how it could have a better ratio. 
Those 31 Raptor engines are going to weigh around 31 tonnes, likely more, all by themselves.  First stage engine mass probably accounts for only 1/4th of the total stage dry mass.  Those assumptions right there gets us close to 0.96.
 - Ed Kyle
That gives a SL Raptor T/W ratio of about 174:1.

That sounds aggressive, until you discover the RD270 (only other complete FFSC actually built) had a T/W of 189:1.
[EDIT. However that was from a Russian site. Astronautix give T/W as 153.25, but running the numbers gives more like 143:1, although I think the Russians sometimes leave off the TVC mass, maximum +/-12deg gimbal) ]

It's also true that Hydrogen (used in the AJR Integrated Powerhead Demonstration) is a special case, both because its a deep cryogen and because it's the only propellant that's compressible at pressures achievable in rocket engines, so difficult to achieve as high a T/W ratio.

But Methane is not H2, it's a hydrocarbon.

It seems hard to believe SX could not achieve the T/W of an engine designed close to 1/2 a century ago, given the advances in design tools (huge) and materials (significant, but perhaps not as dramatic as people imagine given the environment they have to survive)  :( At 189:1 that's a saving of a bit less than 3 tonnes. I'll take a wild stab and guess Musk would target 200:1, because y'know, he's Musk.

[EDIT, and even given the Astronautix values, I'd still say 200:1, as Merlin is already around 180:1 and I'd say Musk is of the "You don't know what the limits are till you exceed them" school  :) ]

The joker in this pack is the weight of the piping to feed those 31 engines at their maximum flow rate. For the same lengths narrow bores should weight less but will cause more fluid hammer (faster fluid flow being stopped) and need higher tank pressures.

Sudden valve closure can give a pressure on the valve 3x driving pressure, enough to snap the pipe off the engine. It should not be underestimated.  :( So 1 tonne/engine (all inclusive) might be accurate.

Modern turbofans are operating at higher and higher temperatures in order to get higher and higher efficiency, and their turbine blades actually have to interact with this hot flow. But in a rocket, only the turbopump's blades have to do that (and it can be designed for lower combustion temperature).

And there's one huge advantage for rocket engines over turbofans when it comes to reliability: turbofans will ingest anything in the air. Birds, insects, sand, volcanic ash, people, etc. That can and does cause catastrophic failure. Rocket engines bring their own air which can be carefully screened for contaminants, with actual screens being put in place to catch anything that might hurt the engine.
There are 2 big issues with rocket engine turbines.

Historically they have run uncooled. In contrast gas turbine blades have typically (except for very small designs) used part of the airflow to do blade cooling, which is why they can run above the melting point of the raw alloys they are made of (and have done so for some time).

The other issue is the very fast startup process.  It turned out a lot of the blade damage in the SSME was caused by a temperature spike of (IIRC) 500F above normal operating temp in the preburner during the few seconds of startup. It, coupled with the LH2 temp  in the cooling channels, lead to "dog kenneling" of the channels (wall thinning leading to leaks) and cracks in the blades. IOW  what happens in those first 3-5 of engine caused damage out of all proportion to the loads during the rest of the time.

Obviously Raptor startup will have been modeled extensively (unlike SSME, which had no CFD models of its preburner built until much later) and AIUI the full flow SC means no inter-propellant seals between an X rich turbine driving a Y rich pump (where X & Y could be either the Oxidizer or the Fuel), which eliminates a major failure mode and also eliminates a shed load of purge gas.

The problem is that start transient is likely to remain short, so is likely to remain (by gas turbine standards) very stressful.
« Last Edit: 10/07/2017 10:03 am by john smith 19 »
MCT ITS BFR SS. The worlds first Methane fueled FFSC engined CFRP SS structure A380 sized aerospaceplane tail sitter capable of Earth & Mars atmospheric flight.First flight to Mars by end of 2022 TBC. T&C apply. Trust nothing. Run your own #s "Extraordinary claims require extraordinary proof" R. Simberg."Competitve" means cheaper ¬cheap SCramjet proposed 1956. First +ve thrust 2004. US R&D spend to date > $10Bn. #deployed designs. Zero.

Offline Zed_Noir

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Re: IAC 2017 -- BFR v0.2 - DISCUSSION THREAD 3 (Post Speech)
« Reply #45 on: 10/07/2017 09:29 am »
....
The one thing that is not possible or practical is an escape system at either Mars or the moon.  The reason for this is quite obvious.  All escape systems depend on being rescued after the fact.  This is possible just about any place on earth.  Even the escape system for the B-58 Hustler could sustain somebody floating in the water are in the Arctic for up to 4 days.  There is nobody on the moon or Mars to rescue anybody.  Even if you successfully escaped the BFR you would still inevitably die.  That’s the way it would be for the foreseeable future.  Only after the BFR has made so many trips to those destinations to prove its reliability would rescue in those places be possible but then there would not be a need for a rescue system on the BFR.

That is not quite true. If there is a fueled BFS available on the Moon or Mars. A rescue mission could be mounted with the BFS on a ballistic hop to pick up the survivors.

Yes, it is the equivalent of sending a cruse ship to do the job of a life boat.  :o
How does the unpiloted bfs plot this hop? I would think this a tall ask in an emergency without some sort of GPS analog around the moon / Mars first?

There should be SX Starlink satellites around if the BFS is there at the Moon or Mars.

Distress beacons and reflective panels  from the survivors's means of escape from a doomed BFS to tracked the trajectory to the landing site from orbital and static observation asserts.

Since both the Moon and Mars are extensively survey by orbital satellites. The BFS doing the rescue can use a terrain matching navigation system to go to the survivors once they are located by orbital asserts.


Offline DJPledger

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Re: IAC 2017 -- BFR v0.2 - DISCUSSION THREAD 3 (Post Speech)
« Reply #46 on: 10/07/2017 11:37 am »
No rocket will ever be as safe as an airliner so ...
I understand the anxiety and lack of trust.  But what is your basis for this being permanent state?
Because rocket engines are running much closer to the limits of chemistry and materials than commercial turbofans. ...

...like some of your other statements, that isn't actually true.

Modern turbofans are operating at higher and higher temperatures in order to get higher and higher efficiency, and their turbine blades actually have to interact with this hot flow. But in a rocket, only the turbopump's blades have to do that (and it can be designed for lower combustion temperature).

And there's one huge advantage for rocket engines over turbofans when it comes to reliability: turbofans will ingest anything in the air. Birds, insects, sand, volcanic ash, people, etc. That can and does cause catastrophic failure. Rocket engines bring their own air which can be carefully screened for contaminants, with actual screens being put in place to catch anything that might hurt the engine.
Are rocket engines designed to withstand FOD damage? I think not. Takes just a 1mm size speck of metal entering a TP to destroy an RD-171 engine. EM should look closely at this and design Raptor to withstand FOD such as stray small pieces of metal etc. Giving Raptor FOD tolerance like modern turbofans have should make it and BFR much more reliable.

Offline jpo234

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Re: IAC 2017 -- BFR v0.2 - DISCUSSION THREAD 3 (Post Speech)
« Reply #47 on: 10/07/2017 12:59 pm »


Are rocket engines designed to withstand FOD damage? I think not. Takes just a 1mm size speck of metal entering a TP to destroy an RD-171 engine. EM should look closely at this and design Raptor to withstand FOD such as stray small pieces of metal etc. Giving Raptor FOD tolerance like modern turbofans have should make it and BFR much more reliable.

This is a rocket engine. A foreign object would have to come from one of the tanks. This different from an air breathing engine in a plane.

You want to be inspired by things. You want to wake up in the morning and think the future is going to be great. That's what being a spacefaring civilization is all about. It's about believing in the future and believing the future will be better than the past. And I can't think of anything more exciting than being out there among the stars.

Offline Robotbeat

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Re: IAC 2017 -- BFR v0.2 - DISCUSSION THREAD 3 (Post Speech)
« Reply #48 on: 10/07/2017 01:19 pm »
Merlin was designed to ingest a nut.

"Fleck of paint" is, as usual, false.
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Offline Kaputnik

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Re: IAC 2017 -- BFR v0.2 - DISCUSSION THREAD 3 (Post Speech)
« Reply #49 on: 10/07/2017 01:21 pm »


Are rocket engines designed to withstand FOD damage? I think not. Takes just a 1mm size speck of metal entering a TP to destroy an RD-171 engine. EM should look closely at this and design Raptor to withstand FOD such as stray small pieces of metal etc. Giving Raptor FOD tolerance like modern turbofans have should make it and BFR much more reliable.

This is a rocket engine. A foreign object would have to come from one of the tanks. This different from an air breathing engine in a plane.



I'm unable to find the quote but I'm sure I remember reading about an engine where part of the design brief was to be able to ingest a loose nut into the turbopump without failure. Anyone else recall this?
(Sorry for slight thread drift...)
"I don't care what anything was DESIGNED to do, I care about what it CAN do"- Gene Kranz

Offline Kaputnik

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Re: IAC 2017 -- BFR v0.2 - DISCUSSION THREAD 3 (Post Speech)
« Reply #50 on: 10/07/2017 01:30 pm »


Are rocket engines designed to withstand FOD damage? I think not. Takes just a 1mm size speck of metal entering a TP to destroy an RD-171 engine. EM should look closely at this and design Raptor to withstand FOD such as stray small pieces of metal etc. Giving Raptor FOD tolerance like modern turbofans have should make it and BFR much more reliable.

This is a rocket engine. A foreign object would have to come from one of the tanks. This different from an air breathing engine in a plane.



I'm unable to find the quote but I'm sure I remember reading about an engine where part of the design brief was to be able to ingest a loose nut into the turbopump without failure. Anyone else recall this?
(Sorry for slight thread drift...)

Ok found it- it was the Merlin!
https://www.airspacemag.com/space/is-spacex-changing-the-rocket-equation-132285884/?no-ist=&page=2

Although IMHO I'm wary of the accuracy of this account and would prefer a better source. It just sounds very unlikely to me that part of qualification testing involves chucking nuts and bolts into the fuel tanks...
"I don't care what anything was DESIGNED to do, I care about what it CAN do"- Gene Kranz

Offline John Alan

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Re: IAC 2017 -- BFR v0.2 - DISCUSSION THREAD 3 (Post Speech)
« Reply #51 on: 10/07/2017 02:15 pm »
Shifting subtopic gears here a minute away from FOD in the prop tanks...  ;)

Quote
Meanwhile, I've found a solution for the bounded problem (150 t LEO/20 t GTO for reuse, 250 t LEO for expendable version).  The solution requires that second stage dry mass be roughly 45 tonnes, much less than the 85 tonnes mentioned in the presentation.  With PMF ~ 0.96 for both stages, the numbers work out if something like 6-7% propellant fraction is assumed to be required for RTLS, landing, etc.  I have S1 at 3278 t/131 t GLOW/Dry and S2 at 1122 t/45 t.

 - Ed Kyle

Using Ed's cargo solution numbers...
Where do we end up on the Tanker version?
How many tonnes of off loadable prop to LEO... the 220 tonnes (1/5 full) hinted in the 2017 presentation?...
And are the tanks likely a stretched 1250 tonnes prop volume as some have opinion'd?...Or something else?
With no need to support a payload in front of them... could the tanks be a lighter, less beefy version?
It's also thought the nose section is as light as possible with no openings beyond maybe a maintenance access hatch...
 ???
I have been looking at Ed's solution and pondering it's various nuances...
Trying to Wrap my head around what SpX is doing here...  ???

One key thing that blew me away...
(on edit... I'm thinking the 1.25:1 take off thrust to weight working on Gravity Losses is key in this working)

250 tonnes payload (expendable) only has 850 tonnes of prop sloshing around in BFS's 1100 tonnes capacity tanks...
And between the BFR and BFS both running to empty...
It gains 9200+m/s (LEO) velocity...
...295 tonnes...
250 tonnes payload (plus the 45 tonnes empty BFS)...
...going around and around in orbit...
And the prop tanks were not full...  :o

Profound... and here is why I think so...
As the Raptor matures and chamber pressure inches toward the 300 bar goal.
They can instantly take advantage to add more delta/V to the stack...
As take off thrust from those 31 Raptors increases... add more prop to the BFS...
Someday... 250 tonnes RTLS may be possible... (wild guess not confirmed)

The other concept that I think helps understand it, is this example...
Lets say the Payload is a 250 tonnes water tank...
You drain 15 tonnes of water out prelaunch and put 15 tons more prop in the BFS...
You get to LEO with the 235 tonnes of payload... 9200+m/s and SECO it...
But now you have 15 tonnes of prop left in the BFS tanks...
You can eject the payload and recover the BFS (just over 1000m/s delta/v avalible + aerobraking)

So ironically... You have to expend stage 1... but can save stage 2...LMAO...  ;D

Now then... Continuing down the payload to prop weight redistribution path...
In between the 235 and 150 tonnes payload range...
Still adding more and more Prop to BFS to keep GLOW at 4400 tons spec...
You end up with enough delta/v margin at SECO, to add a MECO and start doing ASDS style BFR recoveries...
In other words... LOAN spare delta/v from the BFS back to the BFR...
At first the landings are "HOT" and very ballistic...
But with less and less payload... they get cooler and easier on the hardware with bigger delta/V loans...

At 150 tonnes payload...
You have enough delta/v to add a flip/boost back burn and start landing back on the launchpad..
And ironically... the BFS still only has 950 tonnes in the tanks at launch and still could take another 150 if the Raptor 300 bar upgrade pans out...

From 150 tonnes down to 0 tonnes payload..
If needed be... Adding more prop to BFS = more delta/v to do plane or orbit changes after reaching LEO...
You are gaining delta/v, because GLOW is now falling below 4400 tonnes...

Now to the tanker...
I'm thinking... but I'm not sure on this part...
IF the payload is reduced to zero...
And then you go a bit further on a tanker BFS by cutting weight to a bare minimum on the airframe...
I mean after all, in the big picture sense...
There is no need to have a load path thru the tank to hold a payload...
A flying gas tank could be well under 45 tonnes empty weight...
Every tonnes saved in EW helps in what is needed on the landing burn also...

It's implied by the 2017 slides and presentation...
...that something like 217 tonnes of prop can be hauled up on each load...
217x5 loads would be 1085 tonnes of prop...
...plus the 15 tonnes assumed on board that could have got it home...
Equals full tanks... 1100 tons...

But it's still not clear to me how they got from 150 tons of spare prop to a 217 or so number...
There is a 67 tonnes gap there that needs explained somehow...  :P

Anyway... My thoughts on BFS/BFR and how as a system it may work... :)

(On edit... now back to the ongoing FOD discussion)
« Last Edit: 10/07/2017 03:26 pm by John Alan »

Offline RobLynn

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Re: IAC 2017 -- BFR v0.2 - DISCUSSION THREAD 3 (Post Speech)
« Reply #52 on: 10/07/2017 02:32 pm »
That gives a SL Raptor T/W ratio of about 174:1.
That sounds aggressive, until you discover the RD270 (only other complete FFSC actually built) had a T/W of 189:1.
[EDIT. However that was from a Russian site. Astronautix give T/W as 153.25, but running the numbers gives more like 143:1, although I think the Russians sometimes leave off the TVC mass, maximum +/-12deg gimbal) ]
similar sized 50 year old NK33/AJ26 staged combustion engine with ~1500kN and ~15MPa chamber pressure has T/W ~140.  have a look at the size of its huge turbopump https://en.wikipedia.org/wiki/NK-33#/media/File:Aerojet_AJ26_in_the_Stennis_E-1_Test_Stand_-_cropped.jpg and it is pretty clear that the Raptor (if the released CAD is to be believed) is going to be much much higher T/W. Raptor T/W of 200-250 is likely.
The glass is neither half full nor half empty, it's just twice as big as it needs to be.

Offline oldAtlas_Eguy

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Re: IAC 2017 -- BFR v0.2 - DISCUSSION THREAD 3 (Post Speech)
« Reply #53 on: 10/07/2017 02:52 pm »


Are rocket engines designed to withstand FOD damage? I think not. Takes just a 1mm size speck of metal entering a TP to destroy an RD-171 engine. EM should look closely at this and design Raptor to withstand FOD such as stray small pieces of metal etc. Giving Raptor FOD tolerance like modern turbofans have should make it and BFR much more reliable.

This is a rocket engine. A foreign object would have to come from one of the tanks. This different from an air breathing engine in a plane.



I'm unable to find the quote but I'm sure I remember reading about an engine where part of the design brief was to be able to ingest a loose nut into the turbopump without failure. Anyone else recall this?
(Sorry for slight thread drift...)

Ok found it- it was the Merlin!
https://www.airspacemag.com/space/is-spacex-changing-the-rocket-equation-132285884/?no-ist=&page=2

Although IMHO I'm wary of the accuracy of this account and would prefer a better source. It just sounds very unlikely to me that part of qualification testing involves chucking nuts and bolts into the fuel tanks...
The reason is that inside the LOX tank were COPV Helium tanks fastened by Nuts and Bolts. If one of those came loose(unlikely because such as shown before results in a rather spectatular tank failure) or a stray got left in the tank during build (likely).

For Raptor a more likely scenario is frozen LOX or Methane in the super cooled prop.
« Last Edit: 10/07/2017 02:57 pm by oldAtlas_Eguy »

Offline rakaydos

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Re: IAC 2017 -- BFR v0.2 - DISCUSSION THREAD 3 (Post Speech)
« Reply #54 on: 10/07/2017 02:57 pm »
Merlin was designed to ingest a nut.
Merlins have proven to be excellent engines, but we have to remember that Raptor will operate at higher pressures and will have full-flow preburners, different propellants, etc.  No guarantee that Raptor will end up as reliable as Merlin.

 - Ed Kyle
They didnt design the merlin to injest a nut because they expected it to ingest a nut. They designed it that way so that merlin would have suffucet margins for reuse. Raptor has the same design requirment, though perhaps not so vividly described as "eating a nut."

Offline guckyfan

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Re: IAC 2017 -- BFR v0.2 - DISCUSSION THREAD 3 (Post Speech)
« Reply #55 on: 10/07/2017 03:05 pm »
They didnt design the merlin to injest a nut because they expected it to ingest a nut. They designed it that way so that merlin would have suffucet margins for reuse. Raptor has the same design requirment, though perhaps not so vividly described as "eating a nut."

I think it is so even worst case engine out does not become a RUD for the whole vehicle. With the nut the engine is no longer operational but won't destroy neighbouring engines or the tank.

Offline oldAtlas_Eguy

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Re: IAC 2017 -- BFR v0.2 - DISCUSSION THREAD 3 (Post Speech)
« Reply #56 on: 10/07/2017 03:48 pm »
They didnt design the merlin to injest a nut because they expected it to ingest a nut. They designed it that way so that merlin would have suffucet margins for reuse. Raptor has the same design requirment, though perhaps not so vividly described as "eating a nut."

I think it is so even worst case engine out does not become a RUD for the whole vehicle. With the nut the engine is no longer operational but won't destroy neighbouring engines or the tank.
These scenarios in qualification testing speak to the culture of the engine design team wanting to be able to test to destruction to find out if the engine controller software and the sensors that is being used can act fast enough to forestall a RUD on a failing engine. Starve the engine quickly enough of prop and a potential RUD will fail to occur but the engine is definitely out but when you have 31 of them or even just 4 or 2 and only need n-1 to fulfill mission with the Raptor VACs probably can do the mission with as few as 2. And for the booster probably can loose as many as 6 out of the 31.

Offline Robotbeat

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Re: IAC 2017 -- BFR v0.2 - DISCUSSION THREAD 3 (Post Speech)
« Reply #57 on: 10/07/2017 03:55 pm »
By the way, this payload corresponds to a payload mass fraction of 5.68% (250/4400t).  Saturn V was 3.88%; Energia was 3.96%; F9 FT is 4.15% IIRC.  (!)
I've been trying to rocket-equation this, with little success. 

Here are the "knowns".
GLOW 4400 t
Thrust Liftoff 5400 t, ISP = 330/356 sec
Ship dry mass 85 t
Ship Mp 1100 t
Ship Thrust 775 t (4 engines) ISP 375 sec
Ship Thrust 347 t (2 SL engines) ISP 330/356 sec

These imply a first stage mass = 4400 t - 1185 t = 3215 t
Unknown is first stage propellant mass fraction. 
When I plug the known numbers into the rocket equation, I get a first stage PMF required to be 0.97938 to get 250 tonnes to 9,200 m/s ideal delta-v (LEO).  That's unrealistic because the first stage ends up with 20 tonnes lighter dry mass than the second stage "Ship".  With PMF1 a more "reasonable" 0.96, I get total ideal delta-v = 9061 m/s, not usually good enough for LEO, but it depends on the details of the ascent.  To get 9200 m/s with PMF1 = 0.96, payload maximum is 235 tonnes.

S1:  3215 t > 128.6 t, ISP 347.4 sec, delta-v = 3734 m/s
S2:  1185 t > 85 t, ISP 375 sec, delta-v = 5479 m/s
PL:  235 t, delta-v total = 9217 m/s

When I try to model the reusable alternative, assuming 10% propellant saved for first stage flyback landing and 6% for second stage retro and landing, I get only 105 tonnes of LEO payload, as follows.

S1:  3215 t > 437 t, ISP 347.4 sec, delta-v (ascent) = 3265 m/s
S2:  1185 t > 151 t, ISP 375 sec, delta-v (ascent) = 5446 m/s
PL:  105 t, delta-v total = 9211 m/s

Rough guesses, obviously, but I've yet to match the SpaceX charts.  When I try to model the 20 tonne GTO mass, the numbers don't converge at all.  I get no payload to GTO.

 - Ed Kyle
A big incorrect assumption here: that 9.2km/s is required. Technically the absolute minimum energy for a 200km orbit (i.e. Including the potential energy from 200km altitude) above the equator (so using Earth's spin to maximum effect) is just 7.5-7.6km/s.

BFR doesn't use hydrogen, so it should get lower aero losses and higher averaged thrust to weight ratio (since your tanks empty sooner). Therefore the benchmark 9.2km/s need not apply, and I am nearly certain I've seen a legitimate estimate for BFR that shows a trajectory of 8.9something km/s to orbit.

If you use the exact same mass fraction as last year's ITS booster (slightly better than 96%) and your numbers for Isp and stage mass, then you get 250 tons expendable to LEO at about 8.99km/s. That fits perfectly with all the rest of the info we have.

No mystery. If you try to mess with SpaceX's numbers to fit more conservative assumptions, then of course you'll not be able to recreate their figures. That's not a mystery, either.
« Last Edit: 10/07/2017 03:58 pm by Robotbeat »
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Offline DAZ

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Re: IAC 2017 -- BFR v0.2 - DISCUSSION THREAD 3 (Post Speech)
« Reply #58 on: 10/07/2017 05:34 pm »
....
The one thing that is not possible or practical is an escape system at either Mars or the moon.  The reason for this is quite obvious.  All escape systems depend on being rescued after the fact.  This is possible just about any place on earth.  Even the escape system for the B-58 Hustler could sustain somebody floating in the water are in the Arctic for up to 4 days.  There is nobody on the moon or Mars to rescue anybody.  Even if you successfully escaped the BFR you would still inevitably die.  That’s the way it would be for the foreseeable future.  Only after the BFR has made so many trips to those destinations to prove its reliability would rescue in those places be possible but then there would not be a need for a rescue system on the BFR.

That is not quite true. If there is a fueled BFS available on the Moon or Mars. A rescue mission could be mounted with the BFS on a ballistic hop to pick up the survivors.

Yes, it is the equivalent of sending a cruse ship to do the job of a life boat.  :o
How does the unpiloted bfs plot this hop? I would think this a tall ask in an emergency without some sort of GPS analog around the moon / Mars first?

There should be SX Starlink satellites around if the BFS is there at the Moon or Mars.

Distress beacons and reflective panels  from the survivors's means of escape from a doomed BFS to tracked the trajectory to the landing site from orbital and static observation asserts.

Since both the Moon and Mars are extensively survey by orbital satellites. The BFS doing the rescue can use a terrain matching navigation system to go to the survivors once they are located by orbital asserts.

SAR requires an extensive amount of infrastructure.  This infrastructure is available almost everywhere to some extent or another on earth.  It would definitely take a significant amount of infrastructure to pull off the multiple SAR missions to recover the possible 1 – 2 dozen rescue pods scattered across hundreds of square miles.  It is definitely not as simple as sending out a taxi to pick up your overly indulgent girlfriend/boyfriend.  If the BFR is successful as Musk envisions than something like the Starlink constellation will probably happen on the Moon/Mars.  This would only be part of the required infrastructure that would be needed for a successful SAR infrastructure.  The basic premise though is the same.  This extensive infrastructure would not be available until very much later when most likely the need for an emergency bailout system on the BFR most likely will no longer be needed.

The use of individual escape pods has many advantages over a single escape lifeboat type system.  They directly scale to the number of crewmen that you're trying to rescue.  They require relatively little modification to the overall BFR system.  They can be placed on the BFR system at early implementation and removed later when no longer needed with little impact to the overall system.  They could greatly reduce the need for the number of launch/reentry pressure suits.  This last alone could almost make it worthwhile to use escape pods even if to only protect against depressurization events.  The use of ejection seats over the decades has shown us two very important things.  One that they generally save more lives than they cost and 2 they don't always work.  If you have a system where everybody is on the same system it becomes an all or nothing proposition.  Using individual escape pods it is very conceivable that you could have a situation where you could eject 20 of them but only 10 of them survive.  This obviously would not be the best situation but it would be a better situation than losing all 20.

Offline lamontagne

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Re: IAC 2017 -- BFR v0.2 - DISCUSSION THREAD 3 (Post Speech)
« Reply #59 on: 10/07/2017 06:06 pm »
By the way, this payload corresponds to a payload mass fraction of 5.68% (250/4400t).  Saturn V was 3.88%; Energia was 3.96%; F9 FT is 4.15% IIRC.  (!)
I've been trying to rocket-equation this, with little success. 

Here are the "knowns".
GLOW 4400 t
Thrust Liftoff 5400 t, ISP = 330/356 sec
Ship dry mass 85 t
Ship Mp 1100 t
Ship Thrust 775 t (4 engines) ISP 375 sec
Ship Thrust 347 t (2 SL engines) ISP 330/356 sec

These imply a first stage mass = 4400 t - 1185 t = 3215 t
Unknown is first stage propellant mass fraction. 
When I plug the known numbers into the rocket equation, I get a first stage PMF required to be 0.97938 to get 250 tonnes to 9,200 m/s ideal delta-v (LEO).  That's unrealistic because the first stage ends up with 20 tonnes lighter dry mass than the second stage "Ship".  With PMF1 a more "reasonable" 0.96, I get total ideal delta-v = 9061 m/s, not usually good enough for LEO, but it depends on the details of the ascent.  To get 9200 m/s with PMF1 = 0.96, payload maximum is 235 tonnes.

S1:  3215 t > 128.6 t, ISP 347.4 sec, delta-v = 3734 m/s
S2:  1185 t > 85 t, ISP 375 sec, delta-v = 5479 m/s
PL:  235 t, delta-v total = 9217 m/s

When I try to model the reusable alternative, assuming 10% propellant saved for first stage flyback landing and 6% for second stage retro and landing, I get only 105 tonnes of LEO payload, as follows.

S1:  3215 t > 437 t, ISP 347.4 sec, delta-v (ascent) = 3265 m/s
S2:  1185 t > 151 t, ISP 375 sec, delta-v (ascent) = 5446 m/s
PL:  105 t, delta-v total = 9211 m/s

Rough guesses, obviously, but I've yet to match the SpaceX charts.  When I try to model the 20 tonne GTO mass, the numbers don't converge at all.  I get no payload to GTO.

 - Ed Kyle
A big incorrect assumption here: that 9.2km/s is required. Technically the absolute minimum energy for a 200km orbit (i.e. Including the potential energy from 200km altitude) above the equator (so using Earth's spin to maximum effect) is just 7.5-7.6km/s.

BFR doesn't use hydrogen, so it should get lower aero losses and higher averaged thrust to weight ratio (since your tanks empty sooner). Therefore the benchmark 9.2km/s need not apply, and I am nearly certain I've seen a legitimate estimate for BFR that shows a trajectory of 8.9something km/s to orbit.

If you use the exact same mass fraction as last year's ITS booster (slightly better than 96%) and your numbers for Isp and stage mass, then you get 250 tons expendable to LEO at about 8.99km/s. That fits perfectly with all the rest of the info we have.

No mystery. If you try to mess with SpaceX's numbers to fit more conservative assumptions, then of course you'll not be able to recreate their figures. That's not a mystery, either.
Perhaps your stage 2 is not heavy enough?  I find that stage 2 is 1335 tonnes, 1100 propellant, 40 return propellant, 150 cargo, 85 structure.  This has better ISP than the first stage, hence better performance overall.  I don't get more than 9 km/s either though.  Not as good as last years vehicle though, at 9600 m/s

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