sevenperforce

Looks like only a partially successful relight on the second test.

Engagement in Ukraine? There's a second plane visible streaking past as his chute opens but it's too far to see what it is. Probably another Su-25, but since the Su-25 is flown by both Russia and Ukraine that wouldn't make much of a difference.

What a cool job. "So, what do you do?" "I shoot tiny objects at experimental human-sized spaceships at ridiculous speeds in order to estimate how bulletproof we can make our astronauts."

Direct Air Capture has been studied for a long time, but I believe the cost remains somewhat prohibitive. In particular, the energy requirements (cycles of heating and cooling in the sorbent used to capture the CO2) make any system carbon-positive unless they can use purely zero-emission energy sources, which in turn reduce the availability (and drive up the cost) of those zero-emission energy sources in the market. That's one way of doing it, but I believe that is vastly more expensive (from an energy standpoint) than using sodium hydroxide or some other sorbent to do the trick.

Yes, the hypersonic glide ratio was 1:1 which was "insane" in the sense that it was about double the maximum achievable hypersonic glide ratio of a capsule-based design. The Shuttle could also do a 2:1 glide ratio in supersonic flight and a 4.5:1 glide ratio in subsonic flight, although the subsonic regime was too short to get appreciable crossrange during that period. A capsule-based design would not have been able to re-enter and glide back to the launch site in a once-around polar orbit.

An appropriately sized vacuum nozzle for an F-1 engine would need a nozzle exit area of approximately 431 square meters. Diameter of 11.7 meters, 40% larger than the diameter of the SLS. I dunno what insane rocket you're going to have as a first stage if you have that ginormous thing as a second stage.

The difference in impactor mass and target mass are so disparate that it's almost a pure energy analysis, I think. Even with a DU penetrator, the actual penetration depth is going to be significantly smaller than the crater depth. Given that momentum exchange is not the primary driver, I'd think that relative velocity is more important than impactor mass. Momentum is linear to velocity; energy is quadratic.

You could make a thread. Looks like they flew something, too, which is farther than many startups have gotten. The video talks about an issue with the "oxidizer flow" so I'm guessing we're looking at a hybrid rocket with a non-toxic solid fuel? EDIT: Yes, looks that way. This article has a video of a successful test and specifically identifies it as a hybrid rocket.

Indeed. However, it's my understanding that even without servo assist, mechanically constrained joints can solve the ballooning problem all on their own. The possibility to later add servos is surely not lost on Musk.

One of the wildest things ever done with an SR-71 (which brings this thread back from its brief derailment) was to flight-test one of NASA's aerospike engine designs. Here's the hydrolox linear aerospike on the ground: And here's what it looked like when it was mounted to the back of the SR-71: Sadly there are no images of it being fired in flight.

Here's a fantastic video showing late night prep and takeoff. Note you can see fuel visibly leaking at 2:25 and following. Spectacular afterburner ignition at 11:10.

There's some speculation/hope/indication that they may have some integrated mechanical joints.

Yes, in order to keep the tanks from buckling under thermal expansion during flight, there were big gaps that left it leaking on the ground. It also had to refuel immediately after takeoff, but this was not because of leaks; the volume of leaked fuel was negligible. Rather, it could not take off with a full fuel load at all, even with both afterburners. So it would take off with a partial fuel load and rendezvous with a specially modified tanker for top-off. Even then, it would eventually become too heavy to sustain level flight and would need to fire up a single off-axis afterburner to hold altitude and complete the fuel transfer. Most tanker refueling runs are performed at or around cruising speed but the tanker had to get all the way up to its max speed of ~500 knots to keep pace with the Blackbird.

I don't think anything other than the chopsticks can utilize the hardpoints. You end up with too many degrees of freedom.

Was literally typing this out. The SR-71 could not achieve supersonic flight on its engines alone without afterburners, but once through the transonic regime, it could shut off the afterburners and cruise just fine at around Mach 1.2 with military power. However it was actually more efficient in ramjet mode with full afterburner at Mach 2++. The ramjet inlets were designed for Mach 3.24 so that's where they flew the "cleanest" and that was usually the target "top speed" for sorties. Purportedly it got to over Mach 3.5 on a few occasions.

I think they keep the crane attachments in place until they are ready to actually launch, and then they remove them and put on the tile cover sections before final stacking. Until they are ready to actually launch, those crane hooks remain the only way of lifting other than the chopsticks.

My oldest son is just under 10 so I'll have to do it all myself.

That's just a shockingly huge change in period. Prior to the impact, the period was 11.9216 hours with an eccentricity of 0.03. The periapsis velocity and distance would have been 0.1777 m/s and 1.1427 km while the apoapsis velocity and distance would have been 0.1674 m/s and 1.2134 km. Assuming an impact near apoapsis, the dV change required to decrease the period to 11.38 hours would have been 0.00285 m/s or 2.85 cm/s. Assuming an impact near periapsis, the dV change required to decrease the period to 11.38 hours would have been 0.00229 m/s or 2.29 cm/s. Using the 5e9 kg mass estimate for Dimorphous, that's a change in kinetic energy of at least 13.11 kJ and a change in momentum of at least 11.45 million kg*m/s. DART itself, at 500 kg and a relative speed of 6.6 km/s, had a kinetic energy of 10,890,000 kJ and a momentum of 3.3 million kg*m/s. So we are looking at a momentum gain factor of about 347% for an impact with a rubble pile. Extraordinarily good news.

That's beyond amazing to see. Especially in the way it picks up the entry burn. And that's even more amazing to see. Come on, let's light this candle!

Obligatory.

I don't know. Possible that since it is an expander cycle, they want to maximize the surface area to maximize the power output of the cycle. The more heat they can acquire, the higher chamber pressure they can achieve. That's particularly important with methalox, after all, since methane has a lower heat capacity than hydrogen. It's a 3D-printed engine, I believe, so that makes designing the cooling channels for maximum heat takeup a little easier than it otherwise would be. I'm guessing it's a pure dual closed expander cycle without a split. But who can tell for sure? Interestingly enough one of their patents contemplates first-stage recovery using ducted lift fans located perpendicularly up the sides of the first stage.

Dilithium crystals don't provide enough lithiums to counteract flux capacitor depression. The only true solution is the trilithium crystal. 50% more lithium and 100% as much crystal.

I didn't see a thread yet for this, so I figured why not go ahead and talk about it? Spanish smallsat launch startup with a dual-regeneratively-cooled methalox aerospike engine. Or, more accurate, a "spike" engine; there's no truncation of the nozzle. Here's one of the hot fires: Some of their corporate material suggests a two-stage reusable rocket: This cutaway seems to confirm an annular chamber: I do love a dual-expander cycle methalox design, if that's what they're doing here. Won't have a TWR ratio as good as staged combustion but there's still something very pure about it.

One of the issues with a prepared landing surface is that you can end up making the problem worse if you’re not careful. As long as your landing site is sufficiently flat, your landing engine plume is **probably** only going to spew dust everywhere; there’s a risk of excavating larger pieces of debris but there’s also no guarantee there are any such pieces of debris to be excavated. In preparing a pad, you’ve gotta be 100% certain that your preparation will result in a surface that’s truly impermeable to an exhaust plume, because otherwise you have basically just created a giant layer of potential shrapnel.