Quote from: AncientU on 10/06/2017 05:13 PMBy 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 tThrust Liftoff 5400 t, ISP = 330/356 secShip dry mass 85 tShip Mp 1100 tShip Thrust 775 t (4 engines) ISP 375 secShip Thrust 347 t (2 SL engines) ISP 330/356 secThese imply a first stage mass = 4400 t - 1185 t = 3215 tUnknown 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/sS2: 1185 t > 85 t, ISP 375 sec, delta-v = 5479 m/sPL: 235 t, delta-v total = 9217 m/sWhen 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/sS2: 1185 t > 151 t, ISP 375 sec, delta-v (ascent) = 5446 m/sPL: 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

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. (!)

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.

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.

Quote from: ncb1397 on 10/06/2017 07:27 PMAnother 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).

I've been trying to rocket-equation this, with little success. Here are the "knowns".GLOW 4400 tThrust Liftoff 5400 t, ISP = 330/356 secShip dry mass 85 tShip Mp 1100 tShip Thrust 775 t (4 engines) ISP 375 secShip Thrust 347 t (2 SL engines) ISP 330/356 secThese imply a first stage mass = 4400 t - 1185 t = 3215 tUnknown 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 rocket will ever be as safe as an airliner so ...

Quote from: DJPledger on 10/06/2017 07:38 PMNo 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?

Quote from: fthomassy on 10/06/2017 08:10 PMQuote from: DJPledger on 10/06/2017 07:38 PMNo 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.

Quote from: Lar on 10/06/2017 08:04 PMQuote from: edkyle99 on 10/06/2017 06:31 PMI've been trying to rocket-equation this, with little success. Here are the "knowns".GLOW 4400 tThrust Liftoff 5400 t, ISP = 330/356 secShip dry mass 85 tShip Mp 1100 tShip Thrust 775 t (4 engines) ISP 375 secShip Thrust 347 t (2 SL engines) ISP 330/356 secThese imply a first stage mass = 4400 t - 1185 t = 3215 tUnknown 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

Quote from: edkyle99 on 10/06/2017 06:31 PMI've been trying to rocket-equation this, with little success. Here are the "knowns".GLOW 4400 tThrust Liftoff 5400 t, ISP = 330/356 secShip dry mass 85 tShip Mp 1100 tShip Thrust 775 t (4 engines) ISP 375 secShip Thrust 347 t (2 SL engines) ISP 330/356 secThese imply a first stage mass = 4400 t - 1185 t = 3215 tUnknown 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.

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

Quote from: QuadmasterXLII on 10/06/2017 04:23 PMIs 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 systemhttps://forum.nasaspaceflight.com/index.php?topic=43923.new#new

Quote from: fthomassy on 10/06/2017 08:10 PMQuote from: DJPledger on 10/06/2017 07:38 PMNo 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. ...