sevenperforce

That's possible. The F9 landings actually have a degree of risk associated with them because the stage has to scrub non-negligible horizontal velocity at engine relight. The original impact zone is offshore, but the grid fins steer the impact zone beyond the pad, so that the engine relight will scrub both horizontal and vertical velocity to complete the hoverslam. If the grid fins operated properly but the engines failed to light, it would actually impact farther inland than the pad, not between the pad and the shore. In contrast, the Starship actively controls its fall with its flaperons but then actually gains significant momentum toward anterior at relight because the engines are lit while horizontal. They're gimbaled as far back as possible, but still it pushes forward toward the landing site, and then the flip must actually tip the whole thing back beyond radial to scrub velocity. So if Starship boosts out over the water and then uses its flaperons to glide back, it could actually be safer than the F9 landings.

Nah, they won't sacrifice SN8 just for excrementss and giggles. But they may use a different landing pad so that there is no collateral damage if it pancakes.

As long as the header tanks are full at engine relight, they're fine. Once the engines are lit, you have ullage and so fuel reserves in the main tanks will come available. It's definitely a lot to have only simulated.

To be clear, the Venera landers WERE sterilized prior to encapsulation and launch. It's not like the USSR didn't even try. I don't, however, know how good of a job the USSR did sterilizing them. There's data from Venera, which did measure atmosphere composition. It's certainly worth a second look, though I suspect that someone would have noticed it if we had this evidence earlier. That said, with Mars getting all the attention, it's possible that it was in the data all along, and it was overlooked. The measurements would be from spectra, not atmospheric composition; a digital sniffer on Venera wouldn't have picked up anything. I don't know what spectrometric resolution you need in order to resolve the signature of phosphine. There were published spectra measurements in 1968, 1975, 1985, 2013, and 2017 but I don't know anything about how good the resolution was. Now that we know where to look, it would be interesting to hunt through that data. Of course we would also need to know what concentration was measured recently in order to know whether finding or not-finding phosphine would be significant.

I assume they're talking about overall gross volume.

The lander has lots of dV -- it's Orion that is lacking.

Equatorial landings have lower dV requirements, which is increasingly important because LOP-G seems less and less likely to be ready in time for A3.

Yes, I can make the estimate. It's 50-50. Error bars on that estimate are rather large, though.

I remember one of my first actual upper-level physics lab projects was rediscovering the absorption lines in the stellar spectrum.

It was leaked.

The cool thing, I think, is that this is the most likely way of finding life on another planet. We would see novel chemistry first.

The altitude-compensating nozzle with the highest vacuum efficiency is a dual-bell nozzle, and that is what the RVac has. It is more efficient in vacuum than the adjusted-exit-angle nozzle of the RS-25, but less efficient at sea level due to parasitic plume recirculation within the extension. That's ok, though, because the RVac would only fire at sea level in the event of an abort, which would be a time for prioritizing thrust, not efficiency. Precisely. An altitude-compensating sustainer engine will be more efficient than a sea-level-optimized over the course of the entire ascent, but it has lower thrust than a smaller engine at sea level AND takes up more space.

I agree...which is why I wouldn't ask Elon stupidly technical questions like, "From a materials integrity standpoint, what are the primarily failure modes if Raptor's turbopumps were pushed too far in an emergency abort?" or "What are the pros and cons of using a thorium salt reactor to operate the Martian ISRU plant?" But I am confident he could speak to Starship's broad failure modes, demand for hypersonic passenger transport, and RCS thruster design.

Eh, I don't know about that. Elon's position as chief engineer is not just window dressing. He might not be THE injector expert or THE turbopump expert or THE combustion efficiency expert or THE metallurgy expert, but he probably knows as much diverse information about his rockets as anyone at SpaceX. Just a few examples of the many questions about Starship I'd ask if I could: In which case air augmentation is wasted mass

Fair point. I should have said "realistically" not "by definition". I don't want to ask Elon tough questions; I want to ask Elon interesting questions that I can't learn the answer to by googling for twelve seconds. Sigh. Not everyone can nerd out on this stuff as much as we do.

A "sea level" version of Raptor for Mars would be just 0.7% smaller in diameter than the RVac. I could see it being used to great effect by solid-fueled multistage smallsat launchers. They develop high atmospheric speeds due to their high thrust and they stage rapidly, so you can shed the extra weight of the augmentation system. Something like the SS-520 could probably be cut nearly in half with air augmentation. If you want to go reusable, it's tougher. John Bucknell is in love with air augmentation for ascent but honestly EDL is always the killer. I can't think of any way to make reusability and air augmentation play together due to EDL issues, unless you were to go with a HOTOL design. Smart people can still ask dumb questions. Just saiyan.

I always laugh whenever I go back and watch Elon trying to explain why aerospikes on Starship are a bad idea and how he has to basically explain it like Tim is 5: "Um, so uh, you've gotta get your combustion efficiency, uh...so you know, there's really two parts to...like, when you have a rocket engine, what are you trying to do? You want to shoot things out, uh, as fast as possible, in a straight line. So you've got your combustion efficiency. What percentage of max theoretical combustion efficiency are you, um, and then what sort of nozzle efficiency? Are you straightening the flow? You have to shoot it out in a straight line so then you go in the other direction." And then later... "I would love it if someone could show us how an aerospike is a smart move, because then we would just do an aerospike! You just have to show that your combustion efficiency is not affected, and that you're straightening the flow sufficiently. And this is the other thing: if you've got a two-stage rocket, where your boost stage is primarily in atmosphere and your upper stage is primarily in vacuum, then you can specialize for a vacuum nozzle and for a sea level nozzle, and then why do you need the aerospike?" Any altitude-compensating nozzle is by definition less efficient at sea level than a specialized sea level nozzle and less efficient in vacuum then a specialized vacuum nozzle.

I don't mean to crap on Tim. Really, I don't. But asking "why don't more sea level engines use tapered vacuum nozzles" suggests he does not have nearly as deep an understanding of rocket science as he seems to hold out. It is kind of a waste of time to ask Elon dumb questions. If I had 1/10th the access to Elon that Tim has, we could REALLY learn some good stuff.

What Tim may or may not understand is that the RS-25 was designed in order to have maximal efficiency in vacuum without risking flow separation at sea level. If the RS-25 had been built as a pure sea level engine, it would have had an expansion ratio of roughly 25:1, making its nozzle only 1.4 meters across instead of 2.4 meters across. The expansion bell was far larger than it needed to be at sea level and so it had to be tapered in order to prevent flow separation. The RS-25's vacuum specific impulse is almost 10% greater than the RS-68 because it is a staged-combustion engine (which is intrinsically more efficient than a GG engine) and has a chamber pressure more than twice as high as the RS-68, yet its sea level specific impulse was only 0.3% greater. A dedicated sea-level version of the RS-25 would have a higher sea level specific impulse than the RS-25 actually has, but a markedly lower vacuum specific impulse. So that is why. Expansion ratio is not everything. Because the RS-25 was a sustainer engine which needed to provide the bulk of the Shuttle's dV in vacuum, Rocketdyne built an altitude-compensating nozzle which prioritized vacuum specific impulse over sea level specific impulse. The Superheavy booster does not need to provide very much vacuum dV and so it is a much better idea to optimize for sea level specific impulse (and it's much easier because of the ridiculous chamber pressure). Elon obviously knows all this, but he is giving Tim the simple answer -- booster too small, nozzles too big, can't fit more.

Logic indicates you are correct with higher pressure you can have an larger nozzle before pressure become to low at the end. However you might want to go with smaller nozzles than optimum because of space constrains. Not exactly. The higher your chamber pressure is, the greater expansion ratio (and therefore the greater nozzle exit area for a given chamber/throat size) you need. This is true. The expansion ratio is the ratio of the nozzle exit to the area of the throat. The throat is the point of highest dynamic pressure; the exit is the point of lowest dynamic pressure, and the larger your pressure drop from throat to nozzle exit, the more efficiently you're converting thermal energy into kinetic energy. RaptorSL, for example, has a nozzle expansion ratio of 40:1, meaning that the area of the exit is forty times the area of the throat. This is dramatically higher than the MerlinSL, with its expansion ratio of just 16. This makes intuitive sense; the Raptor's chamber pressure is 3.4x greater than the Merlin's and so you need a bigger expansion ratio in order to translate as much of that pressure differential as possible into a higher exhaust velocity. The throat of the Merlin is actually 2.8 centimeters larger in diameter than the throat of the Raptor! However, when we are talking about nozzle extensions, we are talking about the difference between the expansion ratio at sea level and the expansion ratio in vacuum. And so then it actually goes in the opposite direction. Remember that the efficiency with which you convert thermal energy into kinetic energy is a function of the pressure drop between any two points. An infinitely long, perfectly efficient nozzle would convert 100% of the thermal energy into kinetic energy. However, at sea level you must truncate the nozzle to avoid overexpansion and resulting flow separation. The MerlinSL cannot expand from 1410 psi to 0 psi, but from 1410 psi to a minimum of 14.7 psi. Likewise, the RaptorSL cannot expand from 4800 psi to 0 psi, but from 4800 psi to a minimum of 14.7 psi. But if you think in terms of efficiency, what does that tell you? Well, the MerlinSL is quite efficient -- it converts 98.96% of its chamber pressure into exhaust velocity at sea level; only 1.04% of its pressure is wasted (14.7/1410 = 1.04%). RaptorSL, on the other hand, is even more efficient -- it converts 99.69% of its chamber pressure into exhaust velocity at sea level, with only 0.31% of its pressure being wasted (14.7/4800 = 0.31%). This is why higher chamber pressure is so important. It's not just a matter of thrust per unit area (although that's important), but of specific impulse. The higher your chamber pressure, the more work you can extract from your propellant in the pressure drop from chamber to sea level pressure. What does all this mean? Well, it means that the MerlinSL has a significant amount of "work" left to extract, and so the MVac needs a large expansion nozzle (the MVac's expansion ratio is 10.3x greater than the MerlinSL's). The RaptorSL, on the other hand, doesn't have as much leftover "work" in its exhaust at nozzle exit, and so the RVac doesn't need as large of an expansion nozzle relative to its SL nozzle. The RVac's expansion ratio is only 3.2x greater than the RaptorSL's! 10.3/3.2 = 3.22, which is remarkably close to how much more chamber pressure the Raptor has in comparison to the Merlin. tl;dr -- the higher your chamber pressure, the more efficiently you can extract work from your propellant while expanding to sea level, and so the less remaining work there is to extract while expanding to vacuum, requiring a proportionally smaller vacuum expansion bell.

Pixel counting of that side-by-side puts the bell diameters ratio at 1.78 (with some uncertainty due to foreshortening). Pixel counting of official SpaceX renders puts the ratio at 1.76 (lunar Starship landing shot) to 1.77 (Mars entry interface) to 1.8 (standard rear view). So I don't think we're going to see any larger extension.

I am not so sure. The chamber pressure of the Merlin is just 1410 psi vs the Raptor at up to 4800 psi. The higher the chamber pressure, the less nozzle extension you need. I'm inclined to think that is in fact the full-size engine bell.

A dual bell nozzle is an ordinary nozzle with a circumferential kink/inflection point. At sea level, flow separation occurs at the inflection point. In vacuum it expands to fill the entire bell. Basically it is controlled flow separation. The RS-25 avoided flow separation by angling the nozzle in at the base to increased pressure at the expense of some underexpansion efficiency loss in vacuum. A dual bell accepts flow separation by essentially designating where it will occur. Info on a dual bell nozzle design

Am I right? Is that or isn’t that the most powerful vacuum engine ever manufactured? It’s more than twice the thrust of the J-2, and even the SSME’s vacuum thrust is slightly less. The F-1 had more vacuum thrust, of course, but it wasn’t a vacuum engine. I wonder if it is dual-bell. It doesn’t look dual bell but it is hard to tell. I do not think we have a sea level compensating bell like the RS-25. As others have said, it won’t — both because the nozzle is not as thin, and because it will not gimbal. They could not use a radiatively cooled nozzle extension like the MVac because the engine is not directly exposed during the burn, due to Starship’s skirt/heat shield. The Raptor will be fixed in place and likely be in contact with the skirt. I assumed 15% gimbal range because that’s what Elon said. Cosine loss can go up to 3.41%, but that’s from the minimum throttle setting already which does not help as much. In other words, that’s 3.41%, not 3.41 percentage points. Not quite enough to get below one gee.

I know. It reminded me of animation renders like these: Except now it's actually real.