Author Topic: IAC 2017 -- BFR v0.2 - DISCUSSION THREAD 3 (Post Speech)  (Read 375187 times)

Offline envy887

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Re: IAC 2017 -- BFR v0.2 - DISCUSSION THREAD 3 (Post Speech)
« Reply #20 on: 10/06/2017 07:17 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

I agree that 235 tonnes seems a lot more realistic for maximum payload, not that it's ever likely to be used that way.

However, for reusable payload, the 2016 booster was only supposed to need 5% of it's prop load to RTLS and land. And the ship should need only 2-3% (about 800 m/s) to deorbit and land. It should be able to get 150 tonnes to LEO fully reuseable.

Offline RocketmanUS

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Re: IAC 2017 -- BFR v0.2 - DISCUSSION THREAD 3 (Post Speech)
« Reply #21 on: 10/06/2017 07:25 PM »
Is anyone else spooked by all this talk of "no need for an escape system, we'll be safe like an airline?" The parallels with the shuttle program seem almost too obvious.
Yes.
I personally think they need as escape system for lift off and landing for Earth, Lunar and Mars.
To much to risk without one.

New thread for escape system
https://forum.nasaspaceflight.com/index.php?topic=43923.new#new

Offline ncb1397

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Re: IAC 2017 -- BFR v0.2 - DISCUSSION THREAD 3 (Post Speech)
« Reply #22 on: 10/06/2017 07:27 PM »
Another thing that doesn't seem to quite add up is power. Design appears to use maximum - 12 m radius PV fins, which yields something around 90 KW @ earth and 35 KW @ Mars aphelion. Doesn't seem like enough to heat the 853 cubic meters of pressurized volume. Going to need some of that natural gas for heat in a lot of scenarios.
« Last Edit: 10/06/2017 07:28 PM by ncb1397 »

Offline Craig_VG

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Re: IAC 2017 -- BFR v0.2 - DISCUSSION THREAD 3 (Post Speech)
« Reply #23 on: 10/06/2017 07:33 PM »
Another thing that doesn't seem to quite add up is power. Design appears to use maximum - 12 m radius PV fins, which yields something around 90 KW @ earth and 35 KW @ Mars aphelion. Doesn't seem like enough to heat the 853 cubic meters of pressurized volume. Going to need some of that natural gas for heat in a lot of scenarios.

Just for comparison's sake, what PV efficiency did you use for the calculation?

Offline DJPledger

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Re: IAC 2017 -- BFR v0.2 - DISCUSSION THREAD 3 (Post Speech)
« Reply #24 on: 10/06/2017 07:38 PM »
Is anyone else spooked by all this talk of "no need for an escape system, we'll be safe like an airline?" The parallels with the shuttle program seem almost too obvious.
No rocket will ever be as safe as an airliner so EM has made a huge mistake by not designing the BFR system with a LAS. Lets hope EM realizes this and incorporates a LAS in the next BFR design revision which may be announced at next year's IAC.

Also the engine no. on BFR booster needs to come down significantly to simplify the design and make it more reliable. Not to mention excessive maintenance costs that the 31 engines on booster will incur. EM gave BFR 31 engines because he is trying to dev. a nova class rocket with high performance engines as fast as possible on a shoestring budget. This may well bite him in the back in maintenance costs even if BFR never has an engine failure or mission failure.

If 1st FH fails then I will bet on EM reducing the engine no. on BFR in it's next design revision.
« Last Edit: 10/06/2017 07:57 PM by DJPledger »

Offline Kaputnik

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Re: IAC 2017 -- BFR v0.2 - DISCUSSION THREAD 3 (Post Speech)
« Reply #25 on: 10/06/2017 07:40 PM »
Another thing that doesn't seem to quite add up is power. Design appears to use maximum - 12 m radius PV fins, which yields something around 90 KW @ earth and 35 KW @ Mars aphelion. Doesn't seem like enough to heat the 853 cubic meters of pressurized volume. Going to need some of that natural gas for heat in a lot of scenarios.

That's impossible to judge unless you know how much insulation the cabin has, and how much insolation it receives (which depends on vehicle attitude).
Waiting for joy and raptor

Offline ZachF

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Re: IAC 2017 -- BFR v0.2 - DISCUSSION THREAD 3 (Post Speech)
« Reply #26 on: 10/06/2017 07:50 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

I think the 85 tonnes ship mass refers to the Spaceship version.

Recall in ITSv2016, the ship version weighed 150t, and the tanker version 90t. There was no mass given for a cargo version, but it probably would have been closer to the tanker in weight. Also, the payload was 300t for the spaceship variant and 380t for the tanker.

If I had to guess, the tanker/cargo version are probably 55,000-60,000 kg dry.


As I go over the numbers, I think the re-usable payload figures are for the spaceship version as well, meaning BFR's payload to LEO/GTO for the cargo variant is likely 25-30 tonnes higher... It's the only way I can reconcile the tanker variant's required ~210 LEO payload (of fuel).

Offline meberbs

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Re: IAC 2017 -- BFR v0.2 - DISCUSSION THREAD 3 (Post Speech)
« Reply #27 on: 10/06/2017 08:01 PM »
Another thing that doesn't seem to quite add up is power. Design appears to use maximum - 12 m radius PV fins, which yields something around 90 KW @ earth and 35 KW @ Mars aphelion. Doesn't seem like enough to heat the 853 cubic meters of pressurized volume. Going to need some of that natural gas for heat in a lot of scenarios.

That's impossible to judge unless you know how much insulation the cabin has, and how much insolation it receives (which depends on vehicle attitude).
In space the only way you lose heat is radiation. I didn't see any radiators, but they could have been hidden on the other side of the solar panels. Also, the heatshield on the bottom should be dark which will serve as a radiator. Waste heat from electronics, the life support system, chemical reactions powering human bodies, and whatever else needs energy on the ship should keep things plenty warm, they just have to keep the radiators sized and positioned correctly to balance the energy needs. If they need more heating, they just need to point the black side of the ship a bit more towards the sun.

Offline Lar

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Re: IAC 2017 -- BFR v0.2 - DISCUSSION THREAD 3 (Post Speech)
« Reply #28 on: 10/06/2017 08:04 PM »

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".

No, it isn't unrealistic. The first stage, despite being bigger, is a lot simpler than the second stage. This PMF doesn't seem unreasonable to me given how large the booster is, and that it uses engines likely to be very efficient weightwise and composite tanks.

Also, putting "ship" in quotes reads as if it is intended to cast aspersion. Ship is the correct term and scare quotes are not helpful.
« Last Edit: 10/06/2017 08:05 PM by Lar »
"I think it would be great to be born on Earth and to die on Mars. Just hopefully not at the point of impact." -Elon Musk
"We're a little bit like the dog who caught the bus" - Musk after CRS-8 S1 successfully landed on ASDS OCISLY

Offline fthomassy

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Re: IAC 2017 -- BFR v0.2 - DISCUSSION THREAD 3 (Post Speech)
« Reply #29 on: 10/06/2017 08:10 PM »
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?
gyatm . . . Fern

Offline DJPledger

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Re: IAC 2017 -- BFR v0.2 - DISCUSSION THREAD 3 (Post Speech)
« Reply #30 on: 10/06/2017 08:18 PM »
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. The Raptor engines on BFR will be running at even closer to chemistry and materials limits than many other rocket engines due to it's high Pc. Will take many decades before rocket engines approach the reliability level of commercial turbofans.

Offline fthomassy

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Re: IAC 2017 -- BFR v0.2 - DISCUSSION THREAD 3 (Post Speech)
« Reply #31 on: 10/06/2017 08:47 PM »
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. The Raptor engines on BFR will be running at even closer to chemistry and materials limits than many other rocket engines due to it's high Pc. Will take many decades before rocket engines approach the reliability level of commercial turbofans.
Emphasis mine.  My point is never say never.  Much of engineering is about balancing limits to achieve both performance and reliability.  Having thinner margins and greater reliability is not an oxymoron.
gyatm . . . Fern

Offline Lar

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Re: IAC 2017 -- BFR v0.2 - DISCUSSION THREAD 3 (Post Speech)
« Reply #32 on: 10/06/2017 08:51 PM »

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".

No, it isn't unrealistic. The first stage, despite being bigger, is a lot simpler than the second stage. This PMF doesn't seem unreasonable to me given how large the booster is, and that it uses engines likely to be very efficient weightwise and composite tanks.

Also, putting "ship" in quotes reads as if it is intended to cast aspersion. Ship is the correct term and scare quotes are not helpful.
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.

"Aspersion"?  Don't be silly.  I was using quotes to identify the name "Ship" as given by Mr. Musk.

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
Your landing propellant fraction is too high, I think, and so is your raptor weight.
"I think it would be great to be born on Earth and to die on Mars. Just hopefully not at the point of impact." -Elon Musk
"We're a little bit like the dog who caught the bus" - Musk after CRS-8 S1 successfully landed on ASDS OCISLY

Offline John Alan

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Re: IAC 2017 -- BFR v0.2 - DISCUSSION THREAD 3 (Post Speech)
« Reply #33 on: 10/06/2017 09:06 PM »
Ed... glad to hear the numbers now match up to what was 2017 presented...  8)

I agree with your thought the tanker and cargo are in the 45 Tonnes dry weight range... (45~50)

I also agree with you on Raptor weight... I'm at 980kg each SL these days personal estimate...
« Last Edit: 10/06/2017 09:11 PM by John Alan »

Offline ZachF

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Re: IAC 2017 -- BFR v0.2 - DISCUSSION THREAD 3 (Post Speech)
« Reply #34 on: 10/06/2017 09:09 PM »
ITS2016 required 7% of fuel to land its first stage per the slideshow, BFR1`17 stages at a lower velocity.

Offline Semmel

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Re: IAC 2017 -- BFR v0.2 - DISCUSSION THREAD 3 (Post Speech)
« Reply #35 on: 10/06/2017 09:13 PM »
I just realized, for BFR/BFS dont need a TEL. If I understand correctly, its a large part of the current launch pad structure of F9. I think it will be eliminated for BFR/BFS. because it doesnt need it. The first stage is put onto the launch mount by the tower crane seen in the animation. The bottom umbilicals and launch mounts are connected then. The second stage is vertically set on top by the tower crane again. All umbilicals for the BFS run through BFR. The rocket is structurally stable enough that it does not need a support from a TEL. I would even take the guess that all connections between first and second stage are actuated and connect automatically without manual interaction.

After the first flight, BFR will stay on the launch mount after each flight. It will not go back to the hanger. And if it needs to go back, it is craned to a transporter. This will need a lot of manual work and they will avoid that like the plaque.

It seems BFR and BFS are designed to minimize operation costs. The way to minimize operation costs and reduce turn around time is automation. Seeing the way the fuel pipes are designed in the presentation convinced me of that. I would guess that they will make all checkouts after integration of BFR and BFS automatically and only have an operator on a console. This will save a lot of costs compared to F9 launch processing.

Also, due to the mating procedure, vertical integration comes with it. There is no other way of doing it anyway. This will make a lot of people happy. Turns out, for BFR, vertical integration is simpler and far cheaper than horizontal. Its just differently done than anyone else does vertical integration right now.

Offline cppetrie

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Re: IAC 2017 -- BFR v0.2 - DISCUSSION THREAD 3 (Post Speech)
« Reply #36 on: 10/06/2017 09:48 PM »
The BFS canít really be expended unless they donít bother to fill the header tanks. Since those are specifically for landing fuel and the only engines for landing are the two sea-level engines, I would guess that the header tanks will only be plumbed to feed fuel to the sea-level raptors and nowhere else. I suppose there could be plumbing to move fuel between tanks, but that would add weight and complexity. Do we have a sense of how much fuel the header tanks can hold? How much is mass to orbit reduced to have full header tanks so you can recover the ship. Itíll never be used expendable so it hardly matters but just curious.

Offline John Alan

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Re: IAC 2017 -- BFR v0.2 - DISCUSSION THREAD 3 (Post Speech)
« Reply #37 on: 10/06/2017 10:40 PM »
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...
 ???
« Last Edit: 10/06/2017 10:59 PM by John Alan »

Offline DAZ

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Re: IAC 2017 -- BFR v0.2 - DISCUSSION THREAD 3 (Post Speech)
« Reply #38 on: 10/06/2017 11:28 PM »
Is anyone else spooked by all this talk of "no need for an escape system, we'll be safe like an airline?" The parallels with the shuttle program seem almost too obvious.
Yes.
I personally think they need as escape system for lift off and landing for Earth, Lunar and Mars.
To much to risk without one.

New thread for escape system
https://forum.nasaspaceflight.com/index.php?topic=43923.new#new

The following should not be considered an argument either for or against the need for an emergency escape system.

Considering the work that was done on both the B-58 Hustler and the XB-70 Valkyrie escape pod systems, with the additional work that was accomplished under project MOOSE, it would seem that an escape system would both be possible/practical.  This system could be usable from the pad all the way into orbit and during return from the moon or Mars reentry all the way to touchdown.

I am not stating that designing/building/testing such a system would be simple nor cheap just that it would be possible and practical.  As the earlier flights of the BFR would be limited in crew size to possibly 10 – 20 and the BFR has so much performance to spare that the performance hit would be relatively minor.  Even when the crew size gets up to the 100-200 the BFR has so much performance that it could still absorb the performance hit.

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.
« Last Edit: 10/07/2017 01:55 AM by DAZ »

Online Robotbeat

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Re: IAC 2017 -- BFR v0.2 - DISCUSSION THREAD 3 (Post Speech)
« Reply #39 on: 10/07/2017 01:18 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.
« Last Edit: 10/07/2017 01:40 AM by Robotbeat »
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