totm nov 2023 SpaceX Discussion Thread

[magnemoe]

22 hours ago, tater said:

BFS (alone) as an SSTO with even 10t cargo is the USAF's wish come true. Doesn't even need launch clamps, just a blast trench underneath (just to increase lifespan by decreasing wear and tear, it has no such thing on the Moon or Mars).

Seriously. Fuel up, and go. Land at any AFB you will overfly. By the time anyone has the orbital elements, you can be back at base.

does it have enough trust? 3 atmospheric and 4 vacuum engines, yes you could replace two vacuum for atmospheric, you could add more engines but it would cut into payload and make it an separate version. 
Yes the point to point on earth also make far more sense with an single stage. 

[tater]

All I know is that Musk said the BFS was an SSTO with a small payload. Assume cargo version.

[magnemoe]

5 hours ago, tater said:

All I know is that Musk said the BFS was an SSTO with a small payload. Assume cargo version.

Think this was pretty theoretical as in above 9km/s dV with no cargo but full fuel load. 
The two atmospheric engines would not be capable of above 1 g trust in any case, yes its up to 3 engines now but think this is more for redundancy and more return cargo. 
 

[MatterBeam]

8 hours ago, tater said:

All I know is that Musk said the BFS was an SSTO with a small payload. Assume cargo version.

 

3 hours ago, magnemoe said:

Think this was pretty theoretical as in above 9km/s dV with no cargo but full fuel load. 
The two atmospheric engines would not be capable of above 1 g trust in any case, yes its up to 3 engines now but think this is more for redundancy and more return cargo. 
 

I did the calculations. 
The 85 ton BFS with 1100 tons of propellant has a mass ratio of 13.94. With its 375s Isp Raptor engines, it has 9693m/s of deltaV. RSS players know you can reach orbit with about 9400m/s of deltaV, so you can 'waste' 293m/s of deltaV pushing off on the sea level Raptor engines at 330s to 356s Isp. Estimating payload capacity has proven extremely difficult with the tools I have, but I estimate it at about 1.5 to 2.5 tons. 

If you want a TWR of 1.3 on liftoff with a fully loaded BFS, you'd need nine 1700kN sea-level Raptor engines though. This will certainly eat into whatever payload capacity you should have left over.

The picture changes if SpaceX manages to stick closely to the design figure for dry mass of 75 tons. This gives us 10 tons extra margin to work with. You can increase the payload capacity to 12 tons, or launch 2.5 tons into space and land with 300m/s deltaV for a retro-burn. 

[tater]

The 75-85 ton vehicle is the pictured pressurized version. I think the cargo/tanker version is lighter as it was in the 2016 presentation.

[SuperFastJellyfish]

nm  

Edited October 18, 2017 by SuperFastJellyfish

[magnemoe]

1 hour ago, tater said:

The 75-85 ton vehicle is the pictured pressurized version. I think the cargo/tanker version is lighter as it was in the 2016 presentation.

1 hour ago, MatterBeam said:

 

I did the calculations. 
The 85 ton BFS with 1100 tons of propellant has a mass ratio of 13.94. With its 375s Isp Raptor engines, it has 9693m/s of deltaV. RSS players know you can reach orbit with about 9400m/s of deltaV, so you can 'waste' 293m/s of deltaV pushing off on the sea level Raptor engines at 330s to 356s Isp. Estimating payload capacity has proven extremely difficult with the tools I have, but I estimate it at about 1.5 to 2.5 tons. 

If you want a TWR of 1.3 on liftoff with a fully loaded BFS, you'd need nine 1700kN sea-level Raptor engines though. This will certainly eat into whatever payload capacity you should have left over.

The picture changes if SpaceX manages to stick closely to the design figure for dry mass of 75 tons. This gives us 10 tons extra margin to work with. You can increase the payload capacity to 12 tons, or launch 2.5 tons into space and land with 300m/s deltaV for a retro-burn. 

So 9 engines, you probably want some vacuum engine too, on the other hand it will not only be the cargo version but you can also cut some other parts like smaller cargo hold with an smaller door. 
I also understood 75 ton was for the crewed version but might not be the one who goes to Mars (storm cellar and extended life support and mission length) 

[MatterBeam]

3 hours ago, magnemoe said:

So 9 engines, you probably want some vacuum engine too, on the other hand it will not only be the cargo version but you can also cut some other parts like smaller cargo hold with an smaller door. 
I also understood 75 ton was for the crewed version but might not be the one who goes to Mars (storm cellar and extended life support and mission length) 

I must remind you that my calculations ignored the fact that seven more Raptor engines will add up to 6.7 tons to the dry mass and reduce the payload capacity from razor thin to negative.

[Rune]

21 hours ago, tater said:

The 75-85 ton vehicle is the pictured pressurized version. I think the cargo/tanker version is lighter as it was in the 2016 presentation.

Per Elon's AMA the other day, initially there will be no tanker version, just a BFR with an empty payload bay. And reading a bit between the lines, if they eventually build a dedicated tanker, it would be a BFR without payload section, with the main tank stretched a bit and nothing on top.

 

Rune. A pretty phallic picture, the dome-like nose.

[sevenperforce]

If the cargo BFS can be a zero-payload SSTO, I really don't see why you'd need the booster for point-to-point.

Then again, the lowest-energy suborbital trajectories intersecting any two points on a sphere sometimes have very very high apoapses, so higher dV might be required to depress trajectories and decrease both travel time and radiation exposure, not to mention harshness of re-entry.

I wonder if SpaceX would consider making a methane tanker of the BFS, so that you could do LOX-only ISRU on Mars.

[IncongruousGoat]

1 minute ago, sevenperforce said:

If the cargo BFS can be a zero-payload SSTO, I really don't see why you'd need the booster for point-to-point.

Because people have mass? If you've got 100 people, and (rough estimate time) approximate everyone's weight to 65 kg, that's 6.5t  of cargo, not counting the mass of chairs, and ECLSS stuff. That's definitely not a zero payload, and Elon will probably want to have more than just 100 people on any given point-to-point flight.

[Racescort666]

49 minutes ago, IncongruousGoat said:

approximate everyone's weight to 65 kg

Even that is generous, 50th percentile male is 78 kg (not that all of the passengers would be male) but from a human factors point of view, 65 kg is low.

[Rakaydos]

23 hours ago, MatterBeam said:

I must remind you that my calculations ignored the fact that seven more Raptor engines will add up to 6.7 tons to the dry mass and reduce the payload capacity from razor thin to negative.

It's only 2 more. 4 are actually lighter because the large vacum bell is cut down to sea level nozzels.

[insert_name]

On an unrelated note, iridium 4 is now using a flight proven booster, probably one used in a previous iridium launch. However this makes RLTS impossible as the block 3 does not have enough fuel

https://www.nasaspaceflight.com/2017/10/iridium-4-flight-proven-falcon-9-rtls-vandenberg-delayed/

[Exoscientist]

 I'm getting surprisingly high values for the thermal protection of the BFR upper stage, either spaceship and tanker versions. I'm using the fact as indicated in the wiki page on the BFR that it will use the Pica-X thermal protection material. Several references give its density as about 0.25 gm/cc = 250 kg/m³, and the thickness as on the Dragon 2 as 7.5 cm, 0.075 m, about 3 inches. The BFR upper stage has a length of 48 meters and a width of 9 meters. The top part of the stage is conical, so the bottom surface is not rectangular but for simplicity I'll approximate the bottom area to be covered by thermal protection as a rectangle. So the area that needs to be covered is approx. 48*9 = 432 m².

 Then the volume of the thermal protection material is 432 * 0.075 = 32.4 m³. At a density of 250 kg/m³, that amounts to a mass for the thermal protection of 32.4 * 250 = 8,100 kg, which is a surprisingly high addition to the dry mass of 85 tons for the spaceship upper stage or to the 50 tons for the tanker upper stage. One possibility, is the thickness of the PICA-X for the Dragon 2 is coming from the fact it is doing a ballistic reentry, thus generating high heat. However, the BFR upper stage will be doing a more gentle gliding reentry. So perhaps the thermal protection will only need to be half as thick.

 

  Bob Clark

[sevenperforce]

44 minutes ago, Exoscientist said:

 I'm getting surprisingly high values for the thermal protection of the BFR upper stage, either spaceship and tanker versions. I'm using the fact as indicated in the wiki page on the BFR that it will use the Pica-X thermal protection material. Several references give its density as about 0.25 gm/cc = 250 kg/m³, and the thickness as on the Dragon 2 as 7.5 cm, 0.075 m, about 3 inches. The BFR upper stage has a length of 48 meters and a width of 9 meters. The top part of the stage is conical, so the bottom surface is not rectangular but for simplicity I'll approximate the bottom area to be covered by thermal protection as a rectangle. So the area that needs to be covered is approx. 48*9 = 432 m².

 Then the volume of the thermal protection material is 432 * 0.075 = 32.4 m³. At a density of 250 kg/m³, that amounts to a mass for the thermal protection of 32.4 * 250 = 8,100 kg, which is a surprisingly high addition to the dry mass of 85 tons for the spaceship upper stage or to the 50 tons for the tanker upper stage. One possibility, is the thickness of the PICA-X for the Dragon 2 is coming from the fact it is doing a ballistic reentry, thus generating high heat. However, the BFR upper stage will be doing a more gentle gliding reentry. So perhaps the thermal protection will only need to be half as thick.

 

  Bob Clark

The Dragon 2 is designed to be able to do a partly lift-generating entry but can survive a ballistic entry.

Bigger is better for re-entry, though, because an empty stage will have a TON of surface area but not much mass at all. IIRC, the Space Shuttle's ET always broke up due to aerodynamic stresses, then burned up, rather than the other way around. Fluffy is great. The BFR upper stage won't need nearly as much shielding as the Dragon 2, comparatively.

Plus, wouldn't the 85 tons already include shielding mass?

2 hours ago, insert_name said:

On an unrelated note, iridium 4 is now using a flight proven booster, probably one used in a previous iridium launch. However this makes RLTS impossible as the block 3 does not have enough fuel

https://www.nasaspaceflight.com/2017/10/iridium-4-flight-proven-falcon-9-rtls-vandenberg-delayed/

This would seem to indicate that the only reason they haven't reused more boosters so far is because customers simply haven't been entirely comfortable with the idea.

In other words, they need buy-in before they can reuse.

Which, they're getting. So, good!

[Rune]

3 hours ago, Exoscientist said:

 I'm getting surprisingly high values for the thermal protection of the BFR upper stage, either spaceship and tanker versions. I'm using the fact as indicated in the wiki page on the BFR that it will use the Pica-X thermal protection material. Several references give its density as about 0.25 gm/cc = 250 kg/m³, and the thickness as on the Dragon 2 as 7.5 cm, 0.075 m, about 3 inches. The BFR upper stage has a length of 48 meters and a width of 9 meters. The top part of the stage is conical, so the bottom surface is not rectangular but for simplicity I'll approximate the bottom area to be covered by thermal protection as a rectangle. So the area that needs to be covered is approx. 48*9 = 432 m².

 Then the volume of the thermal protection material is 432 * 0.075 = 32.4 m³. At a density of 250 kg/m³, that amounts to a mass for the thermal protection of 32.4 * 250 = 8,100 kg, which is a surprisingly high addition to the dry mass of 85 tons for the spaceship upper stage or to the 50 tons for the tanker upper stage. One possibility, is the thickness of the PICA-X for the Dragon 2 is coming from the fact it is doing a ballistic reentry, thus generating high heat. However, the BFR upper stage will be doing a more gentle gliding reentry. So perhaps the thermal protection will only need to be half as thick.

 

  Bob Clark

That is not surprising at all. The thermal loading is, at the bare minimum, proportional to the ballistic coefficient, and an empty stage will always beat a capsule at that. Which is why Shuttle could reenter with non-ablative silica tiles, it had a low-ish ballistic coefficient at high angles of attack. A humongous empty upper stage (BFS), doubly so.

 

Rune. Figuring out the thermal protection requirements requires so much more math, you'd need propietary software.

[Exoscientist]

10 hours ago, sevenperforce said:

The Dragon 2 is designed to be able to do a partly lift-generating entry but can survive a ballistic entry.

Bigger is better for re-entry, though, because an empty stage will have a TON of surface area but not much mass at all. IIRC, the Space Shuttle's ET always broke up due to aerodynamic stresses, then burned up, rather than the other way around. Fluffy is great. The BFR upper stage won't need nearly as much shielding as the Dragon 2, comparatively.

Plus, wouldn't the 85 tons already include shielding mass?

 

 Very good point. In Elon's description of the Interplanetary Transport System (ITS) from last year both tanker and spaceship version of the upper stage were intended to be reusable multiple times, so likely the heat shield mass was already included in the quoted vehicle mass values:

Making Humans a Multi-Planetary Species.
Musk Elon. 
New Space. June 2017, 5(2): 46-61.
https://doi.org/10.1089/space.2017.29009.emu

 Likewise, the descriptions of the tanker and spaceship upper stages in this years BFR version were also already described as being reusable, so the heat shield mass was also included there.

 In any case, you get a high mass for the payload of the BFR tanker as an expendable SSTO. Estimating the BFR tanker dry mass as approx. half the dry mass of the ITS version, as Elon confirmed with the spaceship case, the tanker dry mass would be in the range of 45 to 50 metric tons, and the payload as expendable SSTO would be in the range of 55 to 50 tons.

 But this puts it as an expendable SSTO in the payload range of the Falcon Heavy while being in the same size range of the expendable Falcon Heavy. So this SSTO would get the same payload fraction as a 2 and 1/2 stage vehicle. Moreover, judging from the fact the ITS tanker upper stage was to cost $130 million production cost, the half size BFR tanker might only be $65 million, so it would be half the cost of the Falcon Heavy. But the Falcon Heavy as an expendable launcher already would be a significant cut in the cost to orbit. So the BFR tanker as an expendable SSTO could be a great reduction in the cost to space, compared to current values.

 But Elon wants to go beyond expendables and has implied the reusable version of the BFR upper stage would only get perhaps in the range of 10 to 15 metric tons payload (by saying it's an order of magnitude less than the full BFR 150 ton reusable payload.) That loss in payload seems high, 40 tons, nearly the size of the entire vehicle dry mass, presumably because of the size of the propellant that needs to be kept on reserve for landing on return.

 I'd like to see a trade study of the payload of instead going with wings for a horizontal landing. Wings typically take up only 10% of an aircraft dry mass. Then with carbon composites, that would be cut to less than 5% of the landed (dry) mass. Keep in mind the loss in payload with vertical, propulsive landing is nearly 100% of the vehicle dry mass. Also, going with short, stubby wings as with the X-37B, you can make the wing weight even less:

  The areal size of the wings in that case would also be less than that of bottom area of the BFR tanker, perhaps only 1/3rd to 1/2 the areal size. So the increase in heat shield mass would only be at most 1/2 that of the approx. 8,100 kg mass of the current heat shield, so perhaps an extra 4,000 kg. But actually the addition of wings gives a gentler glide slope so probably the heat shield thickness could be reduced. The result might even be the total heat shield mass would be reduced by adding wings. 

  Bob Clark

Edited October 20, 2017 by Exoscientist

[magnemoe]

15 hours ago, Rakaydos said:

It's only 2 more. 4 are actually lighter because the large vacum bell is cut down to sea level nozzels.

The vacuum engines give 20 higher ISP, don't think the large bell adds a lot to mass or they had not used them and just used 4 sea level ones, but most uses vacuum engines for upper stage, only reason for not is not having room for it. 

[sevenperforce]

8 hours ago, Exoscientist said:

Very good point. In Elon's description of the Interplanetary Transport System (ITS) from last year both tanker and spaceship version of the upper stage were intended to be reusable multiple times, so likely the heat shield mass was already included in the quoted vehicle mass values:

Making Humans a Multi-Planetary Species.
Musk Elon. 
New Space. June 2017, 5(2): 46-61.
https://doi.org/10.1089/space.2017.29009.emu

 Likewise, the descriptions of the tanker and spaceship upper stages in this years BFR version were also already described as being reusable, so the heat shield mass was also included there.

 In any case, you get a high mass for the payload of the BFR tanker as an expendable SSTO. Estimating the BFR tanker dry mass as approx. half the dry mass of the ITS version, as Elon confirmed with the spaceship case, the tanker dry mass would be in the range of 45 to 50 metric tons, and the payload as expendable SSTO would be in the range of 55 to 50 tons.

 But this puts it as an expendable SSTO in the payload range of the Falcon Heavy while being in the same size range of the expendable Falcon Heavy. So this SSTO would get the same payload fraction as a 2 and 1/2 stage vehicle. Moreover, judging from the fact the ITS tanker upper stage was to cost $130 million production cost, the half size BFR tanker might only be $65 million, so it would be half the cost of the Falcon Heavy. But the Falcon Heavy as an expendable launcher already would be a significant cut in the cost to orbit. So the BFR tanker as an expendable SSTO could be a great reduction in the cost to space, compared to current values.

 But Elon wants to go beyond expendables and has implied the reusable version of the BFR upper stage would only get perhaps in the range of 10 to 15 metric tons payload (by saying it's an order of magnitude less than the full BFR 150 ton reusable payload.) That loss in payload seems high, 40 tons, nearly the size of the entire vehicle dry mass, presumably because of the size of the propellant that needs to be kept on reserve for landing on return.

 I'd like to see a trade study of the payload of instead going with wings for a horizontal landing. Wings typically take up only 10% of an aircraft dry mass. Then with carbon composites, that would be cut to less than 5% of the landed (dry) mass. Keep in mind the loss in payload with vertical, propulsive landing is nearly 100% of the vehicle dry mass. Also, going with short, stubby wings as with the X-37B, you can make the wing weight even less:

  The areal size of the wings in that case would also be less than that of bottom area of the BFR tanker, perhaps only 1/3rd to 1/2 the areal size. So the increase in heat shield mass would only be at most 1/2 that of the approx. 8,100 kg mass of the current heat shield, so perhaps an extra 4,000 kg. But actually the addition of wings gives a gentler glide slope so probably the heat shield thickness could be reduced. The result might even be the total heat shield mass would be reduced by adding wings. 

The problem with the Shuttle was that glider-wings gave far, far too much lift during re-entry, to the point that the vehicle would start ascending while trying to enter, eventually lowering its airspeed to the point that it was no longer aerodynamically controllable. Thus the need for the complex banked S-curves for re-entry, and a VERY long re-entry profile with large portions of the body exposed to rather high heat fluxes. The thermal tiles themselves never failed, but obviously Columbia had a Big Problem in this area.

Landing legs are also going to be a good deal lower in mass than landing gear, adding another element.

[cubinator]

23 minutes ago, sevenperforce said:

The problem with the Shuttle was that glider-wings gave far, far too much lift during re-entry, to the point that the vehicle would start ascending while trying to enter, eventually lowering its airspeed to the point that it was no longer aerodynamically controllable. Thus the need for the complex banked S-curves for re-entry, and a VERY long re-entry profile with large portions of the body exposed to rather high heat fluxes. The thermal tiles themselves never failed, but obviously Columbia had a Big Problem in this area.

Landing legs are also going to be a good deal lower in mass than landing gear, adding another element.

You also can't land with wings on Mars, which SpaceX needs to be able to do if they want to have a one-size-fits-all planetary lander, which they do.

[sevenperforce]

24 minutes ago, cubinator said:

You also can't land with wings on Mars, which SpaceX needs to be able to do if they want to have a one-size-fits-all planetary lander, which they do.

Let alone the Moon. 

Of course I want to see a dual-thrust-axis lander, but baby that's just meeeeee....

[Rune]

17 hours ago, sevenperforce said:

Let alone the Moon. 

Of course I want to see a dual-thrust-axis lander, but baby that's just meeeeee....

Hey, I also want to the see that, 'cause sci-fi. But sadly, it is simpler to design stuff with a single axis of thrust.

 

Rune. KISS.

[StrandedonEarth]

22 hours ago, sevenperforce said:

Of course I want to see a dual-thrust-axis lander, but baby that's just meeeeee....

We already have the design...

C'mon, Elon, Make it so!

[Exoscientist]

  Perhaps we will see such lander with this new program:

NASA preparing call for proposals for commercial lunar landers

by Jeff Foust — September 7, 2017

http://spacenews.com/nasa-preparing-call-for-proposals-for-commercial-lunar-landers/

  Bob Clark

Edited October 21, 2017 by Exoscientist