sevenperforce

Apparently they were adjusting the mixture ratio mid-ascent -- I'm not sure why -- and ran out of oxidizer 5.4 seconds earlier than planned. Not sure whether it flamed out or if the computer automatically triggered BECO due to a low LOX volume message. I'll see whether the launch videos are available online...if so, comparing this launch to the prior launch should show a noticeable difference at BECO if it was a flameout...probably a hot fuel-rich flare. Does SpaceX use tighter margins? Yes. But with reusable first stages (or at least planned-reuse first stages), this would never happen; they have a comparably high volume of extra propellant left over for their landing attempt, so a launch anomaly can be compensated for with relative ease at the expense of narrower landing margins.

The colors are pink, purple, and pale blue; depending on your computer settings the pale blue may appear to have a greenish tinge. The italized axis labels represent varying apsides. When the apsis on the horizontal axis matches the apsis on the vertical axis, the orbit is circular and has only one speed; when the axes don't match, there is a high apsis (the apoapse/apogee) and a low apsis (the periapse/perigee). Because the axes have the same labels, each possible combination appears twice, once in the pink region and once in the blue region. But since this is now an elliptical orbit without constant speed, the minimum speed (at apoapse) is shown in blue while the maximum speed (at periapse) is shown in pink. For most coplanar orbits, the most efficient way to transfer between circular orbits at different altitudes (for example, to go from a 100 km circular parking orbit to the 400 km circular orbit of the ISS) is to enter an elliptical orbit with a perigee at 100 km and an apogee at 400 km, ride it up to its highest point, then transfer into a circular 400 km orbit with a circularization burn. Virtually all orbital maneuvers involve some sort of elliptic transfer. This table allows you to figure out how much speed you need to enter or to leave any of the given orbits. That's the thing; you don't necessarily have the velocity at a given apsis. If I'm in a 185x185 km circular parking orbit and want to send a satellite to geostationary orbit, I won't know how large of a burn I need. But glancing at the table, I see that the perigee speed for a 185km x 36,786km elliptical orbit is 10,268 m/s. My parking orbit speed is 7,793 m/s (shown in purple). So I need 3,275 m/s of dV to enter the geostationary transfer orbit. Once the satellite completes the transfer, I can look at the apogee speed for that same 185km x 36,786km orbit to figure out how much additional velocity the satellite will need in order to circularize to a 36,786x36,786 km geostationary orbit. So it's not just about finding perigee speed for a given apogee (or vice versa); it's about calculating the energy requirements for every different transfer or maneuver. The dV map I linked to shows all this even more clearly; it was just too large an image to embed here.

I got tired of doing a complex set of calculations from scratch every time I wanted to find the dV of a given Hohmann transfer, so eventually I sat down and made an excel calculator to do it for me. Which led to this: And this: These are reference tables for periapse and apoapse velocities for Hohmann transfers between numerous orbits of interest around the earth and the moon. They should be pretty self-explanatory. These won't give you dV directly; instead, you have to subtract your current velocity from your target velocity. So if you're at a low orbit and want to go up to a higher one, subtract your circular-orbit velocity (in purple) from the periapse velocity (in pink) matching your orbital altitude to the target altitude and execute that burn. Once you reach the apoapse of the Hohmann transfer at your target altitude, subtract your new velocity (in blue, matching the new altitude to your starting altitude) from your target orbit's circular velocity (in purple) and execute that burn to circularize. To drop to a lower orbit, do the same thing in reverse. The EML-1 and EML-2 points are reference for an orbit at that distance; the perigee burn is the same, but the apogee burn needs to match the lunar-circular velocity instead. Actually that's not perfectly correct (since it matches period not speed), but I'm using patched-conic anyway so it's close enough. I've made a correction to the original so that the EML-1 and EML-2 circular velocities are the period-matching velocities rather than the reference velocities for an orbit at that distance. On the lunar side those points are stationary so you don't have to match velocity at all, In the lunar reference table, the circular velocities for EML-1 and EML-2 are for orbits with that distance but at other points; if you are actually reaching one of these points, you just kill your elliptic-orbit apolune velocity. Then I decided to go ahead and create a complete dV map for all major cislunar transfers. I can't attach it here but I posted it at the following link: http://forum.nasaspaceflight.com/index.php?topic=39942.0 Here's a reduced-size version of the dV map; if you want the full-size version, you'll have to go to the link above or click here.

I'm pretty sure an airbreathing LANTR pushing liquid hydrogen at launch would have enough T/W to take off vertically. Then you could slowly dial down the propellant flow as airflow took over, use that to get up as high as it will allow, then switch back to pure hydrogen.

Well, it could give us a rough idea of how fast the stage must have been going. Though obviously the stage will splat at lower speeds than the deck will boom, so...

I was thinking more of something like a negative lift-to-drag ratio -- having an aeroshell that would passively re-orient after dropping below a certain speed, but do so at an angle that converted a portion of the drag into negative lift and thus circularize. It wouldn't get you entirely out of the atmosphere, but it might be circular enough that a dedicated space tug could match trajectory near apogee, couple with the payload, and then raise perigee. That would avoid needing any sort of correction on the payload itself. You could use spin-stabilization to keep the payload in a low-drag orientation through the lower atmosphere, designed in such a way that the angular momentum would decay at around 70 km and allow it to tumble into the higher-drag negative-lift orientation.

Hmm. Strictly speaking, the last point where the orbital velocity changes would be approximately where it crosses the Karman line. Until then, it is highly subject to drag. Is there a way to design a projectile which will passively re-orient such that a significant portion of its radial velocity is converted to tangential velocity?

Yeah, the weight of the stage doesn't really matter. This is one of the time when the difference between weight and mass actually makes...a difference. Instead of static weight, we are looking at two things: momentum and kinetic energy. If you want to know (roughly) whether the deck was going to give way, you would take the momentum (mass times velocity) and divide by the "crunch time" (i.e. the time it took the stage to crumple to a stop) to get the impulse force. Then you can divide by average contact area to get the pressure. This depends on some assumptions but can provide a first-order estimate of whether the forces exceeded the load bearing capacity of the deck. Now, to determine how far the stage would penetrate, you need to consider kinetic energy. In cases where there is little cohesion in the target material and the impactor is massively supersonic (e.g. a meteor impact or an antitank missile hitting a ceramic plate), you can use Newton's approximation, but that's not the case here. Instead, divide the kinetic energy of the stage by the compressive strength of the deck, then divide by the impact area. This gives you the distance across which a given amount of force produces work equal to the kinetic energy. Of course some kinetic energy is absorbed by the crumpling stage but this is still just first order.

Well, if he has a forcefield capable of containing quantum neutron degeneracy pressure, it ought to be trivial for him to open up one end of said forcefield and create a rocket with a specific impulse of three million seconds. Simulations suggest that there is a small set of proton+neutron combinations a little larger then unbiseptium which have half-lives of minutes-to-days. The simulations are consistent with what we have currently created, but we haven't gotten any further than unbiseptium (118 IIRC) so far.

The LM massed exactly the same as a standard Dragon V2. And Elon has expressed specific intentions for retrofitting the V2 platform for other worlds, so why wouldn't they be interested? The Dragon V2 has more than twice the available internal pressurized volume than the lunar module ascent stage crew cabin had. That's why I proposed adding an internal tank if the crew is reduced to only two. Of course, I admit that this is the trickiest part. The descent module needs 1.87 km/s minimum...closer to 2.3 km/s for safety's sake. The stock Dragon V2 with nozzle extensions would only have 941 m/s of delta v. So you need to add extra propellant somewhere. But yes, with a double Falcon Heavy launch, the second upper stage ought to have enough remaining fuel to return both capsules to LEO. I'd have to run the math to be sure.

The one NASA veteran launching to the ISS via Soyuz today will be heading up activities related to the BEAM. A barge landing attempt will be made, either because the weight of the BEAM + Dragon makes RTLS too risky, or because they really really want to pull off a successful barge landing and think this mission has enough margin to be sure.

More correcting than competing.... **stops being a jerk** Okay, sorry.

Haha! Totally impossible. No way to have a stable nucleon matrix of that size. Nucleus stability is a little like the variety of regular convex 3-symmetric polyhedrons. Tetrahedrons are simple. Cubes are simple. Icosidodecahedrons are a little more complicated. The more faces you add, the more types of faces you'll have to add, and eventually it becomes less and less possible to make it work. Same thing with nucleons; as you try to pack them into a tight sphere, the ones at the surface keep getting less and less firmly affixed. They're not competing theories....

Yeah, if you do EOR and LOR then you have dV to spare. The "safest" approach is to do a double series LOR. Send the Lunar Dragon unmanned on a Falcon Heavy to LLO, then do a second launch with a manned, unmodified Dragon V2 to LLO only after the Lunar Dragon has been safely circularized. The second launch performs a rendezvous with the first launch and the crew transfers to verify systems. The first Falcon upper stage has more than enough dV to act as a crasher stage, setting the Lunar Dragon down gently. They conduct their mission, then ascend and conduct a second rendezvous to the unmodified Dragon V2 and they transfer, along with samples and so forth. The second Falcon upper stage will have more than enough dV for the transfer back. In fact, they might even be able to bring the Lunar Dragon back for reuse, or at least for ISS docking and investigation. Definitely a much more flexible mission profile. If the Lunar Dragon's systems won't come online, they can just make it an orbit-and-return instead. The unpleasant bit is being unable to do it all in one go. But the double LOR is much better than an EOR to LOR. The goal would be for SpaceX to show that their platforms are capable enough to be used for a wide range of mission profiles without requiring ground-up designs.

The "punching through" aspect is what makes me think it hit hard. It may have still been supersonic. I know we appear to see a plume hovering, but that's probably poor frame rate more than anything else.

A LOR mission profile is challenging. The primary issue is that you have to bring along two Dragons rather than just one, which really wrecks your mass fraction. The other thing is that you can't use the Falcon upper stage for your return trip. The crew has to be in the Earth re-entry vehicle for launch due to abort safety requirements, so you have to stack the modified Lunar Dragon on top of the Falcon upper stage, with the unmodified (or loosely modified) Dragon V2 on top of it: Because they have to be in the unmodified Dragon V2 (1) for launch, but then break away and dock nose-to-nose with the Lunar Dragon (4), which then must also break away for descent, the Falcon upper stage (5) can't be used for the Earth return transfer. So you need drop tanks (2) on the unmodified Dragon V2 inside its service module trunk (3) so it can serve as the Earth return transfer stage. As you pointed out, though, the Dragon V2 is very mass-efficient. So I don't know how much mass can be gained by removing the parachute and heat shields from the Lunar Dragon and replacing them with a larger tank. The underside would need to be completely rebuilt...probably using the same landing legs, but with a larger fuel tank in place of the heat shield and a stripped-down aeroshell. It needs its own power supply as well, and it needs an inflatable airlock in place of the egress hatch. Recall that the dV needed for ascent-to-transfer is more than the dV needed for descent-to-ascent. I don't think it's possible to expand the fuel tank capacity enough to have both ascent and descent. So you'd still need to use the Falcon upper stage as a crasher stage. If we suppose that the Lunar Dragon can have all those modifications made without increasing dry mass beyond 4,200 kg, and we estimate a crewed payload of 1,103 kg as in my original estimate, then it is going to need 2.07 km/s (allowing 100 m/s of descent plus a little over two minutes of hover time), which corresponds to 4,640 kg of fuel and a total mass of 9,943 kg. To deliver this from LLO to just above the lunar surface, the Falcon upper stage will need roughly 1.87 km/s, or 10 tonnes of fuel. But if the return Dragon V2 has a dry mass of 4,200 kg, a loaded mass of 5,303 kg, and needs 1.31 km/s of dV to return, it will need 2600 kg of fuel. Even if we ignore the mass of the drop tanks and consumables, that's a total launch mass of 6,800 kg which, combined with the 9,943 kg mass of the Lunar Dragon, means the Falcon upper stage only has 2.87 tonnes of fuel remaining after LLO injection. Far more development on the LOR approach. Internal tanks won't fill up the room; I explained how much available space there is in the very first post. As it turns out I hadn't ever played KSP until after I started this thread.

Either I am REALLY missing something, or that's not actually a thing. There's nothing to keep an electron orbiting a neutron. An electron in orbit around a neutron is a bound state but there are no conservative fields to enable additional mass to be contained within that bound state. Neutronium is sort of a slang term for any state or phase or arrangement of matter consisting primarily of neutrons. It's mainly used in reference to neutron-degenerate matter under gravitational compression and Pauli exclusion pressure, but it can also be used for metastable neutron pairs, triplets, and so forth in temporary strong-force attraction. A piece of a neutron star would explode due to nuclear forces repelling the neutrons apart, with the sudden removal of gravity that was confiding them. But a lone neutron has a half life of 15 minutes. Compared to most of the transactinide elements humans have synthesised, that's AGES. Strictly speaking, a sample of neutron-degenerate matter would explode due to quantum forces. The stuff at the surface of the neutron star is held in place by the strong force, which attracts; gravity also ensures that it remains on the surface. As you go deeper, though, the weight of the stuff above you increases until the nuclear strong force no longer matters, because the weight of gravity is holding the nucleons together even more tightly than the nuclear strong force can. The density increases as the nucleons are squeezed more and more tightly together, until finally quantum spin characteristics (Pauli exclusion principle) prevent them from merging entirely. This is an incredibly high amount of force; gravity squeezes them together against this outward force and thus stores tremendous potential energy in that quantum field. If you removed a sample of NDM from the gravitational field holding it together, the energy stored in the quantum field would be released instantly, and that is what would make it explode.

According to this article, SpaceX will attempt another barge landing after the April 8 ISS resupply run. I'm not sure if this is because the weight of the BEAM plus the Dragon 1 is too great for the first stage to RTLS. It might be. More likely, though, the stage would be capable of RTLS but they want to try and actually stick a barge landing, both for publicity reasons and for more practice. They've already gotten a RTLS perfectly on the first attempt; failing a RTLS landing would be very bad for publicity and give them very little useful data, whereas failing a barge landing would be less damaging and give them more data. Plus, if the first stage has enough fuel for RTLS, then landing on a barge instead would give them more practice landing under a range of weights. It might even be easier because T/W will be lower. Even though they have very good estimates of remaining fuel and so forth, the inherent variability means that the computer onboard the stage have to assess the exact conditions and velocity and altitude and use that to precisely time the suicide burn, and those equations are going to continue to need tweaking. My experience in metallurgical engineering is very limited, but I am extremely doubtful that the engine plume would significantly weaken the strength of the steel. There's just not enough time. It was probably energetic enough to blast off particles (like a sandblaster that is actually a flamethrower) but weakening the tensile strength of the metal is just not likely. Yeah, I think they will be able to stick landings like this in the future, either immediately or once they have additional data.

Yeah, if JFK hadn't been assassinated we probably wouldn't have gone to the moon.

Dunno. How are they getting funded now?

Then again there is always the Bigelow commercial space station.

I know that Elon has a tendency to make exaggerated or overly-optimistic statements, but I don't see anything particularly exaggerated about his claim that Dragon V2 on Falcon Heavy can deliver a 2-4 tonne payload to Mars, ostensibly without modification. Terminal velocity would be a few times higher, I suppose, but this could be decreased by using chutes in combination with propulsive landing. Landing on Earth will be part of operations, yes. You would absolutely need an internal auxiliary tank; no question about that. But I outlined how much dV you would have. With a crasher stage, direct ascent to Earth return has lower dV requirements than LOR.

Lunar orbit and lunar flybys would be a good step, yes, but Dragon V2 is a landing vehicle, expressly intended to have capabilities for landing on Earth and Mars as well as airless worlds. And a manned landing demonstration on the moon, realizeable with existing platforms, is far cheaper than a manned landing demonstration anywhere else. I mean, delivering an unmanned payload to Mars would be a good idea too. But a manned landing isn't going to happen anywhere other than the moon, not without decades of development. What makes you think he's running it now? Yeah, I was going to say "Gwenne Shotwell, just like now". But it had contingency by going to Lunar Orbit first in case the landing site turned out to be not-so-great after all, could use the orbiter for extra science, and had more capability overall (3 people to Lunar Orbit instead of 2). And any direct ascent profile is automatically less efficient than a comparable LOR mission- that's why all serious mission proposals used LOR- it just offered that much more performance. The Dragon V2 is very mass-efficient due to its interior basically being empty aside from seats and a small control center. My profile still has contingency; as long as the Falcon upper stage engine ignites successfully after passing EML-1, you have more than enough dV for abort to orbit/flyby and return at any time. Plus, landing sites aren't as big of a deal now, since we have far better lunar surface mapping than we did during Apollo. Abort to flyby or orbit still has the same "extra science" capability that the CM had (they were mass-restricted too, after all). And the only reason Apollo had three people in lunar orbit was because they needed a warm body flying the CM for the LOR and return; crew payload to the surface would be the same. If LOR is automatically more efficient than direct ascent, then show me a mission profile which can successfully pull off a two-man landing in a Dragon V2 using an expendable Falcon Heavy FT as your launch vehicle. By my calculations, expandable FH FT can deliver a payload of 19.6 tonnes to LLO, not including the 3.9 tonne Falcon upper stage, so that's your mass budget (and remember, you need 1.31 km/s of dV to get back to an aerobraking trajectory).

Cheap? Sure. Lightweight? Not so much. Shock-absorbing landing legs work much better and are lighter. Here is the basic problem with parachutes. F = c*A*v2 This is the (simplified) drag equation. F is force, A is area, v is velocity, and c is the applicable drag constant. Setting force equal to m*g and rearranging, we find the terminal velocity equation: v = [(m*g)/(c*A)]1/2 This is the problem. Terminal velocity is proportional to the inverse square root of your parachute's area, so each time you want to cut your velocity in half, you have to quadruple the area of your parachute. Plus, it's actually worse, because you have to factor in the mass of your parachute: v = [((m+d*A)*g)/(c*A)]1/2 In this equation, d is the cross-sectional density of your parachute. So, the curve ends up looking like this: You'll have a parachute that weighs more than your payload before you get down to a safe landing velocity. And because of money that comes from comsat operators. The unmanned Dragon V2, mounted on Falcon Heavy, is intended to deliver payloads to the surface of Mars without modification. We've also been waiting fifty years for a way to get people into space and back for less than $60 million+ per seat. Dragon V2 cuts that in half, or better. Dragon V2 is about ten times cheaper than the Shuttle flights were. If the price comes down enough, previously unprofitable reasons to send people to space become profitable. Not that I know what those reasons are. Here's a idea, though. Many companies use subsidiaries or shell corporations to purchase small slices of land in small countries with financial secrecy, allowing them to use those countries as tax havens. Because most countries will offer tax exemption if you can show that you have already paid taxes to another entity (to prevent double taxation), this can be hugely profitable when large assets need to be transferred. In Germany, for example, the amount of German-owned assets in international tax havens is roughly a fifth of the country's GDP. Of course, the local government controlling the jurisdiction is able to control what goes on and what the local tax rates are. It can also set the bar for entry in such a way that only large corporations can afford to utilize the tax benefits. I wonder whether it would be profitable for an independent company to set up an orbital server to conduct currency exchange, banking operations, asset transfer, investment brokering, and a variety of other operations outside of the jurisdiction of any government. The operating company wouldn't be doing the operations themselves, of course; they would lease server space/time to individual banks and other companies. An orbital server would be able to accommodate a ton of traffic without incurring additional overhead, enabling it to offer space to numerous smaller entities which wouldn't be able to get space in a tax haven. Such a server would likely need regular servicing/updating for a variety of reasons, beyond what would be possible to do automatically. So they would need some kind of short-term hab. Assuming that the servers and an inflatable hab could be lofted with a single Falcon Heavy launch and imagining a ten-year lifetime with manned servicing missions every six-eight months, we're looking at a total cost of roughly $2.5 billion, plus the cost of the server/hab itself. I wonder if an orbital server like that could earn more than $250 million per year.

I need a heat shield if I am going to land propulsively. Or just RTLS because I know the ground around the launch pad is roughly 70 m above SL.