Silavite

I sat in the FAA public comment hearing earlier tonight for four hours, only to have them end the session before I got a chance to speak. Admittedly I was able to do other stuff while I waited (and some people's comments were... entertaining in and of themselves), so it wasn't too deadening.

In addition to the points Kerbiloid mentioned, gaseous nitrogen also has the annoying tendency to dissolve in LOX. This dilutes your fuel and reduces performance. Also, if I can speak to one particular aspect of rocket engines: Heat transfer in the thrust chamber. Particularly, heat flux at the throat (where it is maximized due to the low boundary layer thickness and high mass flux). A somewhat extreme example of this is the RS-25. As figure 6 demonstrates, the RS-25 has a peak wall heat flux of 163 MW/m2 (or about 100 BTU/s/in2 in freedom units). To put that in perspective, let's consider the hot side wall temperature at the wall that would result from such a heat flux. Fourier's law of conduction for one dimension can be written as: q = κ * ΔT / L where q is the heat flux κ is the wall material's thermal conductivity, ΔT is the temperature rise, and L is the wall thickness. Let's approximate the cold side temperature as 20 K, the boiling point of LH2 at atmospheric pressure. Say that we want to make our thrust chamber out of Inconel, an alloy known for its resilience to high temperatures. κ = 15 W/m/K for Inconel. Supposing that the chamber wall is 3 mm thick (about 1/8"), that results in a wall side temperature of 32,600 K. Obviously, this won't work. Let's try copper, which has an excellent thermal conductivity of κ = 390 W/m/K. This results in a wall side temperature of 1,250 K, which is still too high. Suppose that we want to limit the hot side temperature to 600 K (about 327 deg C or 620 deg F) to prevent the copper wall from losing too much of its yield strength (the relevant line on this chart is for Cu-DHP, which is very nearly pure copper). With a hot side temperature of 600 K, we can say that the chamber wall must be no thicker than 1.4 mm (slightly thinner than 1/16"). This is an approximate upper limit on wall thickness from thermal considerations. There will also be a lower limit on wall thickness from structural and manufacturing considerations. For the RS-25, I'd imagine that the margin between those two limits is very slim indeed. Addendum: The problem is both worse and better than this analysis predicts. Worse because this neglects the fact that the cooling channels have a finite thickness, so there will be some fraction of the wall which doesn't have coolant behind it. Better because the structural load between the thrust chamber and cooling channels is relatively small; the LH2 in the channels will be at a slightly higher pressure than the LH2 in the thrust chamber, thus pressure differential between the thrust chamber and cooling channels is much smaller than the pressure differential between the thrust chamber and outside.

I must admit that this whole catching mechanism is making me say something I never thought I'd say... it's got me interested in a structures problem! (No offense to any structures people here .)

Has Bill Gerstenmaier talked about it at all? He's with SpaceX now, but he's the first person that comes to mind linking senior NASA folks and SpaceX.

If nothing else, at least there will be no risk of foam insulation projectiles.

The title of this thread is so close to being accurate...

The chance of a hurricane striking French Guiana is quite low due to the fact that it's almost at the equator. Tropical cyclones need to be some distance from the equator to form because the rotation associated with their structure initially derives from the Coriolis force, and the Coriolis force is zero at the equator. In addition, it would be unlikely for an extant tropical system to move south towards French Guiana due to the beta effect (on slides 11 - 18). Starbase, on the other hand... well, it's only a matter of time, and those tents don't look very windproof.

There is an old legal adage that says, "If you have the facts on your side, pound the facts; if you have the law on your side, pound the law; if you have neither the facts nor the law, pound the table." Unfortunately, I think that BO is pounding the table at this point.

This logic seems sound, but it seems that BO was granted an injunction anyway. Any idea as to how?

https://spacenews.com/starliner-investigation-finds-numerous-problems-in-boeing-software-development-process/

You beat me to it! (Though... I must admit I am extremely tempted to drive down there. It would "only" be a ~6.5 hour drive, so doing a daytrip is theoretically possible, and if it happens on Saturday when I don't have classes... well, we'll see.)

This is still a long way off, and odds are likely that nothing comes of it... but how much protection does Starbase have against high winds, surge, and rain?

I suppose the rationale is something along these lines. Seriously though, this means that the booster will need to be caught and locked in place on the pad quickly enough such that mechzilla is free to catch Starship. You also need to put notches for holding SS somewhere that won't get destroyed by reentry heating.

The first objection I can think of is that of joints. How would you make these double-suit joints without the outer layer pinching the pressurized sublayer?

The STS proved that SRMs can be reused, but the refurbishment required was immense. A big part of this was likely due to the fact that the motors were dropped in seawater, but I find myself wondering if a different landing system would have made practical reusability (in the vein of the Falcon 9 first stage) possible. Even assuming that SRMs are able to magically soft land near the launch site, the refueling process involves recasting and installing multiple segments of solid rocket fuel. Even handwaving the issue of landing (say that a winged flyback system with horizontal landing is used) and inspection, the refueling process sounds insurmountably labor intensive. Could practical reusability (on the level of the Falcon 9 first stage, which pushes costs down significantly but isn't a paradigm changer) or rapid reusability (on the level of Starship's aspirations with almost no inspection) be achieved with solids? If so, how?

Tim Dodd's interview with Elon talks about this almost explicitly. Dragon requires extreme conservatism, Falcon is a bit less conservative, and Starship requires rapid testing/failure/iteration.

I can't speak for the Raptor specifically, but the RS-25's and RL10's exterior definitely get downright chilly when firing.

Maybe they'll be able to get the dent out via pressurization? It worked for Falcon 1 flight 4.

I am something of a neophyte when it comes to making sugar (or similar) rockets, but I was under the impression that when using sucrose (table sugar), you're not supposed to let it caramelize. In the words of Richard Nakka, Sorbitol, by contrast, does not caramelize. I'll quote the final section of the sorbitol/KNO3 page here, since it seems relevant. It's worth noting that Nakka refers to sorbitol/KNO3 as KNSB and sucrose/KNO3 as KNSU.

The study I quoted there has this to say in brief: https://energy.mit.edu/news/building-nuclear-power-plants/ More detail is available in the section, "Sources of Increasing Cost."

The IPCC discusses the role of anthropogenic (human) forcings versus natural forcings. See: Summary for Policymakers (SPM), pages 5 and 6.

https://businessdesk.co.nz/article/technology/soul-crushing-inside-rocket-labs-toxic-workplace-culture The financials shown in this Twitter thread are particularly worrisome.

I was reading about materials when this statement caught my eye: Source (page 283/284 of the PDF, section 12.5.7) How does the crystal structure influence the emergence of a ductile to brittle transition?

I am cautiously optimistic about the catching idea since Superheavy will be able to hover. If they were doing a hoverslam (like F9), then I'd be much more doubtful.