RCgothic

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Everything posted by RCgothic

  1. Falcon Heavy is something like 8.2 tonnes per second. Assuming Saturn V dumps negligible propellant into the flame trench after it clears the tower (exhaust slowed by distance, "aim" of exhaust wanders) then with a linear reduction it would need to take longer than 14.5s to exceed the Falcon Heavy's static fire. It doesn't take that long, so Falcon Heavy probably wins. But it's moving slowly to begin with. Assuming again zero as Saturn clears the tower and, 10s to clear the tower, and reduction proportional to distance ascended.... I'll do the maths later if nobody else beats me to it. I think it's close.
  2. It takes about 6-10 seconds for a Saturn V to clear the tower, so it's probably reasonable to say that a 12 second Falcon Heavy hold down burn is probably comparable in terms of energy delivered to the flame trench.
  3. I am also with Slashy. Yes, some risks you just have to suck up. There are failure modes you just can't prepare against. If the LM ascent stage springs a leak on the lunar surface there's just no way back from that. You can't pack additional fuel tanks when the margins are that fine. You design the LM so it's as reliable as possible and that's that. But some risks you don't have to just suck up. Was there an unavoidable need for cargo and crew to be on the craft following maturation of rendezvous technology? No. Was putting an orbiter on the side of a fuel tank that had to be insulated with foam the only way to put crew and or cargo in space? No. Were solid rocket boosters the only choice? No. Would a different architecture have allowed full abort capability for the duration of ascent? Yes. Would the airforce have bought in to a conventional rocket family without crippling it with barely-acheivable design requirements? Most likely. Shuttle was a bad design forced to work until it didn't. This isn't just in hindsight - there were articles at the time criticising the choices that were being made. There's no reason except for flawed design goals that NASA's post-Saturn vehicle(s) couldn't have had the reliability record of ATLAS at a fraction of the cost of shuttle.
  4. Fair points both. Capsule and Liquid boosters is still safest though.
  5. But it could have torn free completely, at which point it overtakes the stack and even a capsule with a rapid abort response is at risk. Liquid boosters can be throttled down. SRBs have no place on a manned launch.
  6. Yes, the SRBs could tear free and overtake the stack, LOCV even for a capsule. SRBs have no business on a manned rocket. LFBs could be throttled down to prevent this. SLS is not free from this failure mode. The other failure mode is being on the side of the stack in the path of falling debris. The orbiter was on the side of the stack to enable recovery and reuse of the main engines in a gliding recovery. That wasn't as cheap and effective as advertised, and procluded abort modes for substantial portions of the flight. The exposure of the orbiter to these risks was in hindsight entirely pointless, and it should have been with foresight following STS-1 if not sooner. Manned vehicles must have abort modes from 0-0.
  7. I believe the reason for the long delay this time is that the range needs to be clear for the prep for an Atlas launch today and tomorrow.
  8. Do you mean shear/extrusion? An o-ring can't have tension in a cross-sectional plane unless it's glued to the sealing surfaces, it would simply lose contact (and not seal).
  9. Given all its design compromises, its impressive that STS even worked in the first place. Manned launches have no business being on a rocket with no abort modes. This was entirely foreseeable and my big beef with the shuttle. At some point somewhere someone went "we don't need those" and that's unacceptable when an alternative design with capsule on top has no issues. Possibly rapid reusability wasn't possible without compromising on the abort modes, but in that case the goal shouldn't have been pursued until it could be done in a way that preserved the abort modes, and yes, I apply this reservation to BFR as well. The second fundamental problem with shuttle is cargo and crew on the same vehicle. You just don't have to risk a crew and pay for crew rating to put up a comsat. If what you need to put up is a crew, there's no reason why you can't rendezvous with whatever they need to be doing in orbit. The only ability unique to shuttle was significant downmass, and there's no reason a low-g return module could not be designed and deployed at a fraction of the price of a couple of shuttle missions had that ability ever been really needed. The third big problem is that shuttle became a pork project. It went on so long subcontractors and senators feel entitled to the work shuttle provided. The program has never really been allowed to end despite not flying anymore, and that stops NASA from doing any better in subsequent designs. Finally, a replacement should have been designed in parallel to shuttle operation. That budget for doing so was not available was a congressional issue, not an issue with NASA. On a side note (because not shuttle specific) the comparison of what happened to Challenger to wind shear detectors at airports. It's not the same thing. The engineers knew the o-ring wasn't good for those temperatures and said so. The risk was known at the time and they were ignored, it's not an unknown phenomena suddenly tripping us up. The failure could happen to any craft abused beyond its limits. If you want all-weather launch capability it needs to be designed in from the start. That shuttle didn't have it isn't really shuttle's fault as it's lack of a design requirement.
  10. I mean there may be volatility issues with sealants or materials or such that you'd only discover in a true vacuum test. For pressure vessel integrity I completely agree that 3bar should be good enough.
  11. In terms of pressure differential I'm sure three atmospheres would be fine, but there may be other factors that can only be tested in a near vacuum environment.
  12. An asparagus has leaves overlapping in a spiral pattern. For an asparagus staged rocket with a large number of stages, the stages usual end of dropping off in a spiral pattern.
  13. Honestly I doubt there's an analytic solution for Pn engines. I envisage a program where you step through every possible combination of engine thrusts in 10% increments for combinations of values close to zero on all but the 1 degree of freedom you're interested in. Then iterate at a finer resolution around those values. Repeat for each degree of freedom. It's a brute force approach but should give you a solution of there is one. May miss 'spiky' solutions where equilibrium is reached very quickly from an unlikely combination.
  14. Not an easy problem! You need to work out what combination of engine thrusts provides rotational equilibrium. As others have mentioned there may be more than one unique solution. Then you either want the axis (solution) in which you get maximum thrust or maximum ISP depending on whether or not you are fuel limited or time limited. Over any reasonable time period the answer will always be to first rotate the axis of maximum engine thrust or maximum isp onto your custom vector.
  15. I have no doubt we could manage it with current tech. The problem is budgets and timescales. On current budget the timescale is not getting any closer to the present. On budget max you'd still need to design build and test the interplanetary hardware before the next Mars window, as SLS is not nearly sufficient. That's not happening, as there's too much to do. Even the window for a first interplanetary unmanned test after that is only 2020, which is really pushing it. Finally even if that goes perfectly you need to wait again for the next window in 2022. That's April 2023 manned arrival at Mars at earliest.
  16. Yes. Falcon 9 almost gets away with it because second stage does most of the work of getting to orbital velocity. Falcon Heavy first stages lift more payload to a little faster than Falcon 9, but the 2nd stage really struggles. Half as much DV or less.
  17. From what I've heard on Reddit, it was a ridiculously close call and insiders say they're not going to release it for PR reasons. Firstly, look at the scorch trail across the barge. Also the last frame shows spray being kicked up way off the mark. The rocket clearly pulled some crazy last minute manoeuvres. Secondly, insiders describe the booster balancing on one leg and nearly toppling. There is also a reference to CRS6 and the little thruster that couldn't. Apparently this time it could. So the video sounds ridiculously awesome, but don't hold out any hope for a release!
  18. New Glenn is projected at 45t to LEO and 16t to GTO reusable. It will never fly expendable. Falcon Heavy is projected at 8t to GTO reusable. Payload to LEO reusable is speculative. To match New Glenn 16t to GTO at least the center booster need to be expended (possibility). Falcon Heavy has the potential to beat New Glenn flying expendable, but that's not really comparing like to like.
  19. If they can stick the landing with accuracy, maybe they'll be able to build a freshwater /de-ionised landing pool?
  20. Of course the poor dwell time of the hydrogen EUS means you won't be assembling anything in LEO, which limits SLS to small high energy payloads. It's basically useless for anything else.
  21. I think there's also a slight plane change manoeuvre required, which is cheaper at higher apoapsis.
  22. Can we back calculate what the maximum payload to the minimum GTO is and compare it to Spacex's stated values?