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Everything posted by sevenperforce

  1. Another critical issue is that ICPS, currently, is 13.7 meters high, while Centaur V is 12.3 meters high. The Exploration Upper Stage is planned to be 16.6 meters high. Trying to stack two Centaur Vs on top of each other would raise the height of the stack by 8 meters higher than SLS Block 1B. Adding National Team's lander (without the transfer stage, since we're staging in LLO) would add 12 meters to this. Block 1B Crew stands 111 meters high, so we add 20 meters to that plus the 7-meter-high crawler-transporter, and that brings the entire stack to 138 meters. But you've also doubled the propellant load on the ESM, so it's going to be about twice as long, adding another 4.8 meters and bringing the top of the vehicle to nearly 142 meters, about 3 meters taller than the VAB doors. And that's not even accounting for adapters or spacing between anything. EDIT: Also, merely doubling the propellant load on the ESM wouldn't come anywhere close to what's necessary to brake the National Team's lander into LLO so we're DOA there too.
  2. Because I apparently have too much time on my hands, here are some handy little diagrams showing each of the different modalities. Obviously nothing is to scale.
  3. All of this info can be found online if you know what to search for, but I can also answer these questions (mostly) off the top of my head. The Shuttle's OMS system contained about 300 m/s of total dV depending on the amount of payload it carried, but the overall launch profile dropped it off in a slightly-suborbital trajectory in order to allow the external tank to re-enter passively. It used about a third of this to raise its orbit after jettisoning the external tank and about third of this to lower its orbit back into the atmosphere at the end of the mission. The remaining third was used for on-orbit maneuvering and for attitude control during the first phases of re-entry. The inclination of Mars is less than 2° off from the inclination of Earth. The two orbital planes intersect along a line, and any transfer trajectory between the two bodies will cross this line at one point or node. So typically in a Mars mission, the vehicle perform a very small mid-course correction at whatever place along the transfer orbit crosses that line. It's not the most efficient way to change inclination, but since it is such a small correction it's not that big of a deal and is typically performed with the basic RCS system on the vehicle. Every now and then, the orbital planes and the optimal transfer trajectory will line up such that the node is either at Mars or at Earth. If it's at Mars, then you can just use your atmospheric capture and you never need to correct; if it's at Earth, you perform that correction along with your transfer burn to take advantage of Uncle Oberth. No, it didn't -- Saturn's gravity well is extremely deep and so performing a capture burn as close as possible to Saturn's surface was the more efficient option. Titan is so small compared to Saturn that a gravity assist off of Titan would not have been sufficient to provide a capture. KSP is extremely scaled down. Jool is only about 100 times more massive than Tylo, while in real life Saturn is more than 4,200 times more massive than Titan. It took about 780 m/s of dV, spread over 10 different burns as it got closer and closer. I don't know what your approach is to approach Mercury, but all missions to Mercury -- Mariner 10, MESSENGER, and the ongoing BepiColombo -- have used at least one planet flyby to reduce the propellant needed to reach the innermost planet. Mariner 10 used a single Venus flyby but it never entered Mercurian orbit, instead doing all of its observation during three Mercury flybys. MESSENGER used one Earth flyby, two Venus flybys, and three Mercury flybys before finally performing its Mercurian orbital insertion burn. BepiColombo has already done one Earth flyby, two Venus flybys, and two Mercury flybys, and will do four more Mercury flybys before finally entering a polar orbit around Mercury with minimal propellant use.
  4. So there are three different things here, actually: Vertical ascent to Earth escape A direct ascent launch profile The direct ascent lunar mission architecture The first, which is what the OP describes, has never been planned by any space agency for any mission. It's not feasible at all. You never simply burn straight up to escape velocity; you always perform a gravity turn. It's certainly possible to do, given enough stages, but it would take so much more fuel that it's absolutely never considered. The second, the direct ascent launch profile, simply means that your upper stage burns straight through LEO insertion into the BLEO transfer burn without SECO and without any coast period. This was more often done a long time ago, when we didn't have restartable upper stages. It can technically be slightly more efficient due to the Oberth effect but it requires perfect timing and often is not possible depending on your launch site latitude and the inclination and phasing requirements of the mission. And there's still a brief period when you're technically in LEO, even though you never stop burning. The third describes a mission architecture for going to the moon and coming back. When we were doing our planning for crewed moon missions, there were three options. In a Direct Ascent mission, a single ginormous rocket would take off, go into a LEO parking orbit, burn to the moon, enter lunar orbit, descend to the surface, return to lunar orbit, and then return back to Earth. This never required any orbital rendezvous or docking, but it meant the largest possible launch vehicle. In an Earth Orbit Rendezvous mission, many small launches to LEO would assemble a moon vehicle which, once completed, would burn to the moon, enter lunar orbit, descend to the surface, return to lunar orbit, and then return back to Earth. This could be completed without building any ginormous rocket. In a Lunar Orbit Rendezvous mission, a single ginormous rocket would take off, go into a LEO parking orbit, burn to the moon, and enter lunar orbit, THEN send a small lander down to the surface while the main vehicle remained in lunar orbit, after which the lander would return to the main vehicle and the main vehicle would return home. This meant the original launch vehicle, while still very large, wouldn't have to be nearly as large as in a Direct Ascent approach, and this is what became the Apollo architecture. All three things can be called "direct ascent" but they are all very different. Indeed. It's counterintuitive and it honestly feels a little magical, but Trigonometry Works In Mysterious Ways.
  5. Yes, we could design a rocket that performs a direct burn to the Earth-Moon Lagrange Point 1 (L-1) without first achieving low earth orbit (LEO). It would be wildly, wildly more Δv-expensive than performing a pitchover and entering orbit, whether that orbit was a parking orbit with a subsequent trans-lunar injection (TLI) burn or a so-called "direct ascent" using a continuous burn to avoid any coast period. The Δv required to go from Earth's surface to LEO varies based on ascent profile, but is typically considered to be between 9.3 and 10.0 km/s. The Δv from LEO to C3=0 (that's Earth escape) is 3.22 km/s, while the Δv from LEO to trans-lunar injection is 3.20 km/s. The escape velocity from Earth's surface is 11.86 km/s, so I will assume for the purpose of this calculation that the Δv required for a direct burn from the Earth's surface to L-1 is approximately 11.84 km/s. Now if you look at the above numbers, then you'd think that the "just burn straight up" approach seems better, because a direct burn is 11.84 km/s while going to LEO first will cost you between 12.5-13.2 km/s. However, you're forgetting that the direct burn to L-1 is not instantaneous; it takes time. If your launch vehicle can accelerate at somewhere between 1.5 gees and 3.0 gees, then it's going to take between 400 and 800 seconds of time to do that. Let's say you have a LOT of thrust on your launch vehicle and so you can do all of this in just 500 seconds. Well, that means you're going to be pushing against Earth's gravity for 500 seconds, and Earth's gravity is going to pull you back down at 9.81 m/s2. Accordingly, your required Δv is going to go up by 4.9 km/s, bringing total dV to 16.74 km/s. So you're WAY better off going to orbit first. The reason for this, as many here probably already know, is that gravity drag drops off as you ascend to an orbit, thanks to centripetal acceleration. What China did with Tianwen-1 is the same thing we did with Pioneer-4 when we launched it on the Juno II: a direct ascent into low earth orbit that continues directly into the injection burn without SECO or a coast period. No launch beyond Earth has ever gone straight up without at least momentarily entering low earth orbit.
  6. It doesn't even close, either, because it contemplates magically adding 10 tonnes of propellants to the ESM without any dry mass growth.
  7. If I was designing an ejection system for a Mach 10+ crewed vehicle, I would certainly do an ejectable cockpit. Make the cockpit basically a miniature space capsule that blasts itself free and then tumbles to aerodynamically stabilize. A chute system for such a large capsule would probably be prohibitively heavy and require secondary cushioning like airbags, which make things much more tricky. So I would give the chute system a drogue that automatically deploys at low altitude and then a lower-energy ejection seat inside it so that the pilot can land with a personal parachute. But that's not exactly what is reflected in the film. Maverick's suit is all sooty; he clearly had quite a toasty time. It's possible that the flight suit would be designed to ablate some to absorb heat at that point. Another possibility is that the ejection system actually flips the pilot's seat backward and out of the cockpit in order to position the back of the seat windward and keep the (now-inverted) pilot leeward.
  8. Putting an abort system under it accomplishes the off the stack bit, and could be used with a crew X-37 derivative in a similar fashion, I think the issue is what happens aerodynamically during that event. It has to stay in controlled flight at transonic/supersonic speeds while also being pulled away (lateral as well, I assume) from the launch vehicle. In theory you could have a ring of hypergolic abort engines plumbed to the vehicle OMS propellants but mounted on the PAF, and attach the PAF to the upper stage with pyro bolts. Then, in a nominal launch, the fairings are jettisoned at the normal time and the vehicle separates (cutting the plumbing lines) after SECO. In an abort, the pyro bolts all fire to separate the PAF from the upper stage and the hypergolic abort engines push the entire PAF and fairing assembly free of the upper stage.
  9. Hard to come up with a workable abort system for a spaceplane, though. Clearly helps with shock point attachment.
  10. Doesn't merit inclusion elsewhere but I had to roll my eyes. For a smart person, Neil is really dumb sometimes. He thinks that because he's got degrees and this mini cult of personality that he can just shoot from the hip without actually thinking about physics. Or, I don't know, looking up what dynamic pressure is.
  11. Depends on whether you count sample-return missions: The Genesis mission spent a little over three years in space collecting solar wind particles before re-entering Earth's atmosphere. It crash-landed due to a drogue failure but some of the samples were usable. The Stardust mission collected particles from the coma of the Wild 2 comet; its mission duration was over 6 years before successfully returning to Earth. Hayabusa took a little over 7 years to return from its sampling mission of 25143 Itokawa, and Hayabusa2 took 6 years and 3 days to do the same thing with 162173 Ryugu. OSIRIS-REx is expected to take just over 7 years to complete its sample return from 101955 Bennu. Of course none of those can be reflown like X-37B or the Shuttle.
  12. Looks like they have started LOX fill tests on the fully stacked booster 7 + ship 24.
  13. Oh, absolutely a Russian video. I was just wondering whether this is a video from an active aerial military engagement ongoing in the Russo-Ukrainian War, or something else like a training exercise. I don't speak Russian so I have no idea what the comms are discussing. If it's a training exercise, then the plane streaking past is clearly a training partner. If it's an active engagement, then the other plane might be Ukrainian.
  14. Looks like only a partially successful relight on the second test.
  15. 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.
  16. 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."
  17. 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.
  18. 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.
  19. 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.
  20. 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.
  21. 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.
  22. 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.
  23. 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.
  24. 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.
  25. There's some speculation/hope/indication that they may have some integrated mechanical joints.
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