sevenperforce

The landing engine would be recessed in the octaweb with the landing engine igniter glued into place. Landing leg deployment will really mess up aerodynamics; that's why I'd time it as late as possible. Perhaps I could use an automated servo with an ultrasonic rangefinder to deploy at 3 meters or so. With a landing motor TWR < 1, the rocket should still descend shuttlecock-style on the grid fin airbrakes even at full throttle.

My kids love watching rocket launches, so I got them some model rockets and we were launching them. I had also made my oldest a 1:80 scale model of the Falcon 9 with pop-out grid fins and landing legs. He kept asking me when we could launch his Falcon 9 and make it land on a boat. So I got to thinking...how hard would it be to build a hobby rocket with propulsive landing capability? Proper timing for a true suicide burn is hard enough with a liquid-fueled rocket; it would be even more difficult with a solid-fueled rocket, even if you had a laser rangefinder. Instead of trying to do a true suicide burn, then, you could build your model rocket with over-engineered shock-absorbing landing legs. Then, with slightly oversized grid-fin-style airbrakes, the terminal velocity would be fairly low. You could get away with a TWR < 1, since you would only need to decrease terminal velocity, not zero it out entirely. Your landing legs would catch you regardless of whether the motor burned out a few feet above the ground or was still firing on landing. Here's a 1:40 scale model, reproduced in SketchUp. Two stage rocket. Eight motors for ascent; central motor is reserved for landing. The octaweb at the base holds the launch and landing motors, mounts the landing legs, and holds the leg deployment channels (they are deployed by the landing-motor ejection charge). It is either aluminum or dense resin. The legs are likely either aluminum or PVC. They snap into place at the top and are deployed by springs (not shown). Main body (C) is light plastic to allow it to be launched without a license. Fairing halves (A) and (B) are plastic. Upper stage (C) is dense 3D-printed plastic; interstage (D) and "grid fin" air brakes (E) are either aluminum or dense resin. The landing ignition wire guide (F) is also light plastic. Detail view of interstage and upper stage: On ascent, the landing legs serve as a sort of shuttlecock to maintain guidance, with the upper stage providing sufficient counterweight. Return airbrakes on the lower stage and guidance brakes on the upper stage both pop out on springs at stage separation, which also closes a circuit to ignite the upper stage motor. On upper stage burnout, fairing halves pop out to deploy parachutes; these are tethered to the second stage (not shown) and allow for recovery. Lower stage descent and landing:

Technically, the launch vehicle does the insertion, while the spacecraft does the circularization.

That's pretty typical. Just bits of ice and such that were stuck to the stage by acceleration during the burn and are now free to float away.

Everything on Mars is pretty darn old.

Not a lot to say until more research is done. We weren't really looking for another particle in this area, so it's odd to find it, but at the same time the existence of a noticeable signal where we weren't looking is significant. If real, this particle is a gauge boson, but a massive one like the Higgs rather than a massless one like the photon. Its energy level suggests a region of action on the order of the diameter of an atomic nucleus. This could provide an explanation for how dark matter interacts with itself.

Yeah, unless I misremember something, each GTO mission will have essentially the exact same window (with variation in coast time) because you're launching to the same geostationary position relative to your launch point.

Rumor was of a boost in the range, but now there is talk of propellant loading issues. Could be the boat made the pause the prop loading, resulting in a temperature spike and a need to rechill. If they load all the propellant and then have to hold, dumping and rechilling the entire load might take too long.

Dunno who on Reddit saw it, but here's a zoomed-in image. They are faint, but they are definitely there.

Someone on Reddit pointed out that there are very thin barely-visible red stripes on the landed JCSAT-14 booster, so this may be old news.

That's debatable. I actually misspoke; there aren't any materials that can handle 5000C and remain solid. Tantalum hafnium carbide melts at 3900C. So that's your upper limit for any thermal rocket of any kind. For liquid hydrogen fuel I'm guessing you're looking at around 1300 seconds of specific impulse, max. Maybe 2000 seconds if you can get full disassociation of the hydrogen. If your propellant was somehow contained and was opaque to certain laser frequencies, I suppose you could make a solar electric laser-pumped thermal rocket with a higher temperature and better exhaust velocity. But you'd need some sort of nozzle, which is problematic. You'd want a magnetic nozzle, at which point it would seriously be better to just go SEP because your dry mass is entirely ungodly. A novel idea is to use solid lithium hydride as your propellant and use a solar-electric cyclotron or neutronistor to induce a neutron spray that fissions the lithium directly. Poor thrust but you end up with relativistic exhaust velocities. Scales down pretty well, and could be used with a secondary propellant for high thrust.

However, there are materials that can handle 5000C without melting. A solid-core nuclear reactor will melt down at around 3000C. So the ISP can be higher than NTR even though it isn't as high as with SEP.

Gravity is not very strong. Just saying.

I can't think of any other reason why one of the engine bells would be bedazzled with stripes. Then again, I can't think of any reason why stripes would be used. Maybe they want a visual indicator? I suppose that in the worst-case scenario if the rocket ripped itself apart near maxQ, you would want to know whether the reused engine was the culprit, and so you would want to be able to identify it visually in the footage.

My personal favorite is a lithium-hydride NTR/NSWR fission fragment rocket. When exposed to a sufficient neutron flux (like, thermonuclear bomb or gas core NTR level), lithium-6 hydride fissions cleanly to helium and tritium. One of the highest specific impulses possible from nuclear rockets. The primary advantage over SEP is mass; you can get a pretty decent TWR because your engine mass can be very very low. You can actually get a higher specific impulse for a given propellant than an NTR, because your heat exchanger can be made out of a material much more heat-resistant than a solid-core nuclear reactor. As others have said, SEP has the advantage of higher specific impulse, denser propellants, and a more controllable deployment arrangement. You can always have mirrors that fold out, of course, but that really drives up dry weight because you need fine control of alignment, whereas solar panels can still function even when they aren't all perfectly aligned. Reactor-based nuclear electric propulsion is fairly nice though, and not nearly as controversial. Actually, a solar thermal rocket can have a better TWR than an NTR, because it is basically an NTR with the shielding replaced by a heat exchanger and the reactor replaced by a gigantic mirror. A mirror can be much, much more lightweight than a reactor. It does have to be scaled up pretty dramatically to get to that point, though. One way to "cheat" with solar thermal propulsion would be to use a dual-layer heat exchanger: a solid ceramic core with some kind of ablative or otherwise sacrificial secondary layer. With a much larger surface area, the secondary layer would allow for high thrust during the initial escape while ablating away, adding its mass to the thrust of the rocket. Then, once it ablated/burned away, you would be left with a much more lightweight core for sustained propulsion or course correction maneuvers. Rocket equation likes that.

Disadvantage is just that it's pretty much useless beyond Earth, and almost completely useless beyond Mars. You also have to deal with an inflatable mirror design that will hold up against continual solar radiation. Heat rejection is an issue too, as with any thermal rocket. Google "Solar moth" for examples. It has a very futuristic feel.

Solar thermal is not too bad, particularly if you're doing stuff at less than 1 AU, but it's limited by propellant choice. With hydrogen, you can get up to 1000 seconds of impulse, but with something like water you're looking at 190 seconds. Dry mass can be quite low, though; you can use hydrogen or helium to inflate balloon mirrors to get pretty good performance.

The really great thing about chemical fuels is that your power source and your propellant are one and the same, so you get to shed it as you no longer need it. The rocket equation loves that. Your propellant is also the coolant for your power source, which saves additional mass. If there was a more energy-dense chemical fuel, you could use your chemical fuel to heat a separate propellant...but you really still can't beat hydrolox for raw energy to weight ratio. If you don't go chemical and beamed power isn't an option, you carry your power supply (and, often, your coolant) with you as dry mass to orbit. One possibility would be something like a shielded-loop propellant-cooled nuclear-electric pulsed inductive thruster. A pulsed inductive thruster is basically the fluid version of a railgun. You get to use ammonia as your reaction mass, which has great density and can double as your coolant due to its convenient catalyzed endothermic decomposition, and it is quite amicable to ISRU as well. It requires no electrodes, accelerates a charge-neutral plasma with a specific impulse of up to 10,000 seconds, and varies its power consumption by changing pulse frequency without a drop in efficiency. You just need a way to get it to operate in a non-vacuum...which, if you were pumping it as nuclear reactor coolant, shouldn't be hard. Then you set up with a duct and a supercharger to use the lower atmosphere as reaction mass, and you're set.

Naturally.

Devil's brew is putting it lightly. I wanna see an ion engine that can use the atmosphere as reaction mass, so you can get to orbit with it.

I can't even find the original.

Late, but the static fire completed last night with all systems go for launch tomorrow. I've seen speculation that in the future, SpaceX may begin doing static fires the same day as launch. The launch ignition is provided by ground sources of TEA/TEB, so all they would have to do is refuel the stage and refill the suppression tanks. Kinda makes sense. Any major problems would take more than 40 hours to fix anyway, and this way you save having to set up for launch twice. Does anyone know the source of the liquid oxygen that SpaceX is using? Is it provided by chemical industrial process and trucked in, or is it fractionally distilled from the atmosphere offsite and trucked in, or is it fractionally distilled onsite?

Technically? Yes. Measurably? Not even close.

Ah, I misread the formula for HAN. Sucrose (sugar) can burn because it has carbons linking the hydrogen and oxygen together. You add oxygen, which bonds to the carbon to form CO2 and leaves the oxygen and hydrogen to bond to each other. With HAN, however, you actually end up with two free oxygen atoms per decomposition. So HAN is an oxidizer, not a fuel, in biprop mix. Guess it could burn with just about any fuel then. HAN/hydrazine, anyone? **shudder**

Less on KSP and more on propellant choices... The Wikipedia article says that HAN can be used as a bipropellant, but it doesn't say how, and I can't figure out how that would work. It has both hydrogen and oxygen attached, so what would it be burned with? Pretty cool monoprop though. It is still toxic, volatile, and likely carcinogenic, but it isn't quite as bad as hydrazine. For NERVA options beyond pure LH2, I rather like anhydrous ammonia. It is about as dense as kerosene, disassociates easily, and the nitrogen doesn't have the residue problem that CH4 would suffer from. In an NTR with ammonia you get a specific impulse significantly above the best possible hydrolox engines with a TWR that is much better than standard nuclear rockets. Ammonia decomposes endothermically with catalysis, making it a superb coolant for an expander cycle NTR, and it is just cryogenic enough to be self-pressurizing without requiring insulation or particularly careful handling. Finally, it would be extremely amenable to either LOX-afterburning or a supercharged ramrocket cycle, which could mitigate its TWR disadvantage. Finally, it is cheap and available practically everywhere.