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It's now been over a few weeks since the successful launch of IFT-2 and discussion of the problems encountered following the hot staging.
The first decent analysis of this situation arose on a Scott Manley video where his thoughts paralleled mine: propellant sloshing in the LOX and
the methane tanks creating voids which led to starvation of the turbopumps, resulting in failure of the engines to stay functional and restarting problems.
I haven't heard much more from any official sources (i.e. SpaceX), but Scott also suggested that the ignition of the Starship gimbeled engines possibly exerted enough of a blast effect on the booster that there was a negative acceleration that caused the problem. According to other sources of information, the center 3 engines were all deflected outwards to reduce this effect, but if this did occur as Scott suggested, my suggestion for a "fix" is pretty straightforward and simple: increase the gap between the Starship engines and the defector plate of the booster allowing the gasses from the exhaust to better vent sideways during separation. That would allow the angled exhaust to vent rather than hit the top of the booster. Jus a slight modification to allow the geometry to "do it's thing."
I'd be interested to hear what GW thinks about the issue itself as well as my suggested "fix."
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Howe long were the starships engines heating the BFR tanks to cause it to pressurize fuel in a gaseous state as normally the amount would be small as its typically a vacuum which would be created by normal emptying of the fuel tanks.
But you are right space x has been quite silent for what happened in its launch.
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There's 2 different issues here, about which SpaceX has been very publicly silent: (1) what happened to Superheavy? and (2) what happened to Starship?
Staging took place high enough that there isn't a lot of aerodynamic drag. But the interactions of two objects in close proximity in supersonic flow are not necessarily small, as SpaceX learned long ago trying to figure out how to stage in its Falcon-1 beginnings.
I rather think the jet blast from Starship's engines put a large decelerating force onto Superheavy. That force would be approximately equal to the engines' thrust. It's quite large. This occurs precisely because of the hot staging. It's not so large with staging-in-coast, because with the larger distance, the upper stage plume has time to spread, so that very little of it actually strikes the lower stage.
It's a brief deceleration pulse, but with hot staging, it's a strong deceleration pulse. The unused propellant aboard Superheavy is going to fall toward the forward tank domes as a swarm of globules, plus a thick film flowing rapidly forward along the lateral tank walls. The aft dome suctions will be very nearly devoid of any liquids.
The tanks remain pressurized with evaporated propellant, by design. The venting is such that some design pressure level is maintained, that being a part of the structural strength necessary to survive entry airloads, as well as supporting a heavy upper stage during ascent acceleration. It's definitely not near-vacuum in there.
You have to apply ullage thrust (coming from something) such that the globules and thick films fall back toward the aft tank domes. That process takes time, and the lower the gees induced by the ullage thrust (whatever it is), the longer that process takes.
Any forces acting upon Superheavy after staging but before its flip (using cold gas thrusters I believe), act in the wrong direction, being drag and plume forces from Starship. During the flip, centrifugal force would fling some of the propellants to each end of each tank. Stuff near the forward end gets flung forward, stuff near the rear end gets flung aft. But, if the deceleration pulse was strong, nearly all the free-floating propellants are near the forward end, so the flip's centrifugal forces simply amplify that collection of propellants on the wrong domes.
There's a time required for application of ullage forces to settle the propellants onto the rear tank domes. It depends upon the effective gee level of the applied ullage forces upon the near-empty (but not quite empty) stage. A time constant for this would be the time it takes for a globule to fall from the forward dome to the after dome, at the acceleration level of the applied ullage thrust: s = 0.5 a t^2. Example: ullage thrust acceleration is 0.01 gee, for a = 0.0980667 m/s^2. Presume a typical tank length of 25 meters. The time constant would be t = (2 L /a)^0.5 = 22-23 seconds. This needs to act for 3-4 time constants before it has settled the propellants enough to successfully pull a suction at the feed points with your engine turbopumps.
If the near-empty stage masses, say, 250 metric tons, then to pull 0.01 gees you need to apply 2.5 metric tons-force worth of ullage thrust on it, for a duration in the neighborhood about 90 seconds, before you attempt lighting the engines for the boost-back burn. To shorten that time, you must apply higher ullage thrust.
I did not see any evidence in any of the videos of anything like that happening. So I am unsurprised their scheme did not work right.
BTW, a rocket engine turbopump is a rather fragile thing when it comes to off-design conditions. And sucking in vapor instead of liquid is certainly a VERY off-design condition. The turbopumps subjected to this would likely (at the very least) shed all the blading, and would quite probably explode or otherwise violently self-destruct.
SpaceX has been lighting Raptors at full power. I've been hearing what sounds like an explosion followed by the continuous rumble of thrust. That is diagnostic of lighting at full propellant flow instead of a reduced thrust setting. Lighting at full power runs right at the hairy edge of exploding the engine at ignition. Add turbopumps self-destructing because they could not get suction, followed a split second later by a glob of liquid, and I would be very surprised if those Raptor restart attempts did not blow up some engines!
Using adequate ullage thrust for an adequate interval of time, before you attempt engine restarts, makes all of those risks and failure modes go away. Just running 3 engines at reduced thrust during staging does not. Which is what they did.
The thrust of those 3 engines must be less than the thrust of the engines in the Starship upper stage, or else that upper stage could not pull away from the thrusted lower stage in a hot-staging scenario. That means the upper stage plume strike forces on the lower stage are larger than the thrust of the 3 lower stage engines, resulting in a net deceleration before you flip! The propellants still surge toward the forward tank domes, away from the aft dome suctions. And that WILL kill the 3 running engines, which will suffer fatal damage as they suck vapor instead of liquid!
You are not going to get multiple tons of ullage thrust force from cold gas thrusters, nor from drag in the really thin air, once reversed and clear of the upper stage plume. SpaceX needs real, substantial thrusters for this! They have not yet included them. Given the chronic overtime and turnover, I doubt any of the SpaceX engineers who solved the Falcon-1 staging issues still work there. And no company's managers ever want to pay for writing down ALL the knowledge. The engineers they have will have no experience and little knowledge base, for solving this problem. Especially with the strong "not invented here" attitude that most big companies have. SpaceX is no exception.
My own recommendation would be solid thruster cartridges, similar to what flew on the Saturns long ago, just larger for the larger stages. They are not all that big and heavy, not compared to the size of the stage. Rather trivial, really. They produce a lot of thrust for several seconds. They are not a huge cost item like SRB's! Think of it as "reloading your gun with bullets" each time you refurbish for the next flight. I'd put them on both stages, which is "suspenders and belt". "Armored codpiece".
Anybody hear anything at all about what happened to Starship? I have not.
GW
Last edited by GW Johnson (2023-12-04 10:08:17)
GW Johnson
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"There is nothing as expensive as a dead crew, especially one dead from a bad management decision"
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As usual, GW details things the way they really are, and not in some fanboy manner. He has elaborated what I said and what Scott Manley stated immediately after the launch and hot staging fiasco. I sent a video that an aerospace engineer published showing what was probably happening in the tanks.
https://www.youtube.com/watch?v=j9WFmItGrKk
Last edited by Oldfart1939 (2023-12-04 23:15:18)
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For All.... Dr. Greg Stanley will be giving his monthly report on space activities since the last meeting of the North Houston NSS chapter.
It is highly likely he will attempt to provide the best information available on the recent Starship test.
The meeting is scheduled for this coming Saturday afternoon at 2 PM Eastern time, 3 PM New Hampshire time.
See the North Houston NSS topic for the Zoom link.
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Oldfart's youtube video illustrates almost exactly what I was talking about in post #3 above. Thanks, Oldfart1939.
I have no such computer programs with which to visualize or analyze stuff like that. I just had to think my through, pencil-and-paper style.
Refigured for 0.1 gee vs 0.01 gee ullage acceleration, the ullage motor thrust level is 7-8 metric tons-force, but for something closer to 30 seconds than ninety seconds. Converted to units I have an intuition for, that's about 16,500 lb of force total of multiple cartridges.
That's about F*t = 495,000 lb-sec of total impulse required of the ullage solid motors collectively. If the vacuum specific impulse of these motors was near 275 secs, the cartridges collectively would require about 1800 lbm of solid propellant, which is only 0.816 metric tons of solid propellant.
If the ratio of propellant mass to loaded motor mass is near 0.65, as it is with a lot of small production solids, the collective mass of the solid ullage motor cartridges would be about 1.25 metric tons. Compared to a launch mass near 5000+ metric tons, that really is trivial.
That sort of thinking, and beneficial outcome, is exactly why solid ullage motors were used on the old Saturns.
GW
Last edited by GW Johnson (2023-12-05 10:50:29)
GW Johnson
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I posted before but I'll repeat. My recommendation to SpaceX is not to use hot staging. Instead use pistons the same as Falcon 9. Pistons pressurized with helium push the stages apart. Because Starship is larger, pistons will either be larger or more of them. If Elon is concerned with mass, the pistons should mass less than the hot staging adapter.
Falcon 9 upper stage uses cold nitrogen thrusters for RCS, and those provide ullage thrust. I believe Starship uses pressurized gas from propellant tanks. Is that enough? I don't know, but that was how the first orbital test was supposed to stage.
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My suggestions are as given here:
(1) Increase the distance between the Starship and Booster stages, by a small but meaningful increment and increase the size of the hot staging vents.
(2) The pressure of the Starship exhaust on the Booster, as calculated by GW, needs to be reduced significantly through the angle of thrust outwards. I believe the gimbel angle is limited to 15 degrees, so the distance I suggested in (1) above could be calculated on this basis.
(3) Include several ullage solid thruster motors to ensure settling of the LOX and Methane before attempting Booster recovery.
I am unsure about whether or not SpaceX is locked into hot staging, but a lot of thought and good engineering should make it work. This was a new approach and mistakes were definitely made, and it may take some expensive "tinkering" to solve the issues tha have arisen; many test flights with less than optimal results.
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We seem to have an opportunity to apply the well known practice of betting ....
SpaceX keeps it's plans close to the vest, so it seems likely no one placing a bet will have an advantage over anyone else.
At this point, I'm putting all my chips on GW's solid rocket ullage motors as the primary solution.
I have evaluated Oldfart1939's suggestion of increasing the flow of hot gas from the Starship through the vents, and see a decent chance that some version of that will show up in the next test. This would be in tandem with trying to reduce the impact of hot gas on the top of the booster stage
I have evaluated RobertDyck's suggestion of trying an entirely different way of moving the stages apart.
There may be someone in the membership willing to put chips on that idea.
I have no experience with bets or managing them, so would appreciate any input from members who are knowledgeable in this specialty.
***
One advantage of using GW's suggestion is that the engines on the booster can be shut off entirely, since they are (apparently) unable to overcome the powerful impulse from Starship during it's stage separation burn. That will save fuel for the recovery process.
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The Russians have been doing hot staging for several decades. So have we, with a variety of missiles. The Russians have not ever tried to do anything with the lower stage after staging (nor did we), which is the only new thing here. That's where lower stage ullage is an issue. Upper stage ullage is ALWAYS an issue.
It looked to me like their hot staging actually worked. The vented interstage piece (the Russians use a truss of bars, same basic idea) seemed adequate, and the forward dome protection for the lower stage seemed adequate, because it did not blow up until considerably after the staging event.
Ullage for the upper stage was the 3 lower stage engines thrusting the vehicle, until the upper stage engines lit and pushed it away from the lower stage. That worked.
The cold gas thrusters flipped it around. Could not have been the 3 engines, because I think they died with the propellants surging forward, away from the aft dome engine suctions. Whatever, the flip worked.
What did NOT work was lighting lower stage engines after the flip, precisely because no effective ullage solution was installed. What they assumed would work (keeping the 3 engines lit) did NOT work.
Solution: install the solid ullage motors in the base of the lower stage. Light the upper stage engines for hot staging, but almost simultaneously shut down the 3 engines in the lower stage to prevent fatal damage from the propellants surging forward, as they ALWAYS will. Flip it with the cold gas thrusters, then fire the solid ullage motors. Then light up whatever engines you think you need for the boost-back burn.
Alternatively, if you go back to cold staging, you must install solid ullage motors in the base of BOTH stages. You shut down your lower stage engines, then do your gas/piston thing to separate them, almost simultaneously firing your upper stage solid ullage motors, which push the upper stage away significantly while settling its propellants. Then fire your upper stage engines before the ullage solids burn out.
Meanwhile, flip your lower stage with the cold has thrusters. Then fire its ullage motors to settle its propellants, and light whatever engines you need for a boost-back burn before those solids burn out.
Sorry. That's simply what you have to do with this thing, no matter which way you eventually decide to accomplish the staging.
The booster needs no ullage motors for its entry burn. Its attitude and the drag of the grid fins as the entry point is approached, will settle the propellants. The same is true of the landing burn. Both those events take place deeper in the atmosphere where drag forces are significant at 1-2 km/s speeds. Getting away with that all along, has made them unjustifiably overconfident with regard to staging this new vehicle, hot or otherwise.
And BTW, the first flight ended up attempting its staging much deeper in the atmosphere (because of the multiple lost engines), where aerodynamic forces were significant enough to threaten the integrity of the entire tumbling vehicle. To this day, I do not believe their flight termination system blew it up. I think the air loads tore it apart, causing the same-appearing explosion effects. The 41 sec delay from signal-sent to the explosion is your clue! While they haven't said in public, I think the FAA agreed with me on that issue. That why a more reliable flight termination system is one of the fixes they demanded.
GW
Last edited by GW Johnson (2023-12-06 13:00:19)
GW Johnson
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"There is nothing as expensive as a dead crew, especially one dead from a bad management decision"
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Alternatively, if you go back to cold staging, you must install solid ullage motors in the base of BOTH stages. You shut down your lower stage engines, then do your gas/piston thing to separate them, almost simultaneously firing your upper stage solid ullage motors, which push the upper stage away significantly while settling its propellants. Then fire your upper stage engines before the ullage solids burn out.
That's one way, but not the only way. The S-IVB stage was the 3rd stage of Saturn V. It used solid rocket motors for ullage thrust. It had 2 sets, so allowed 2 engine starts. The first was immediately after separating the 3rd stage from the 2nd, to ignite the 3rd stage for insertion into Earth orbit. The second was to depart from Earth orbit, for Trans-Lunar Injection (TLI).
However, the Apollo Service Module (SM) thruster quads used monomethyl hydrazine (MMH) and N2O4 aka NTO as oxidizer. Space Shuttle used the same fuel and oxidizer for both RCS thrusters and Orbital Manoeuvring System (OMS). Apollo SM main engine used aerozine-50 for fuel, and N2O4 as oxidizer. Each thruster quad had its own fuel and oxidizer tank. For thruster quads, fuel and oxidizer was in a silicone bladder, and that was within a metal tank. Between the bladder and tank was pressurized helium. This meant both fuel and oxidizer were squeezed out like toothpaste from a tube, so no settling issue. The Apollo Service Module would use the aft-pointing thruster for each thruster quad for ullage thrust. That means provide a enough thrust to settle propellant in the main thank, then ignite the main engine. At least two thruster quads had to be operational: one on each side.
Upper stage of Falcon 9 uses cold nitrogen thrusters. Those thrusters are used for manoeuvring, but aft-pointing thrusters are used for ullage thrust. Just like Apollo SM thruster quads. I'm saying Starship upper stage could do the same thing. Would residual gas pressure in the main tanks be enough? Or is a separate thruster system necessary? I'll let engineers argue that detail, but I'm saying some sort of ullage thrust is necessary.
As for SuperHeavy first stage after separation: just use solid rockets for ullage like GW said. If Elon really really doesn't want to do that, an alternative is to keep the centre three engines operating and burning long enough for propellant to settle before igniting the middle 10 engines. SuperHeavy is much larger than Falcon 9 first stage, and has 33 suction intakes instead of 9. Sloshing is a much bigger issue. Expect anti-slosh baffles to be more complex, and settling time to be significantly longer.
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How do the header tanks work? Both Starship and SuperHeavy have header tanks. Does that alleviate the ullage settling problem? When are header tanks used?
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https://www.yahoo.com/news/spacex-may-t … 00086.html
Projections for 3rd Starship/Superheavy test flight ...
Gizmodo
SpaceX May Test Key Technique for Moon Mission on Starship’s Third FlightGeorge Dvorsky
Wed, December 6, 2023 at 12:15 PM EST·3 min read
36Starship during its second test flight, November 18, 2023.
SpaceX is actively preparing for the third test of its Starship launch system, a mission that may feature the inaugural attempt at in-flight propellant transfer. This capability is essential for enabling deep space missions and is a crucial component of NASA’s Artemis Moon plans.
<snip>
Apache Internal Server Error objected to something in this article
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I quite agree with Rob about the thrusters and main engine on the Apollo service module. What made that work as well as it did (and it worked very well indeed) was the pressurized bladder feed of room temperature storables. You cannot do that with cryogenics, there are no materials available with the enormous cold elongation capability.
The S-IVB used the solid ullage motors, precisely because it was fueled with cryogens. (The second stage of Saturn-5 used them, too, because of its liquid oxygen, which is also why they did not need bladders for the kerosene.) That is necessarily tankage without bladders, and propellants can slosh about in zero-gee, forming both floating globules and thick films on the interior tank walls.
If you draw very little propellant, you can use the thick film effect to provide liquid at your suction point. The propellant will settle after about 3 time constants under thrust, at which point you can pull a lot more and generate a lot more thrust. But if you try to pull a lot of propellant with the thick film effect, you will fail. You will suck vapor instead of liquid into your turbopump, destroying it.
The problem is the difference between suddenly just being un-thrusted in zero gee (free fall), and having forces impact you while un-thrusted, including swapping end for end (centrifugal effects). Impacting forces could be aerodynamic drag, and plume impact from an upper stage. They really tend to scatter the propellant into globules, with big "slosh" motions about the tanks.
Be aware that at close range, an upper stage plume has not had time to spread laterally into near-vacuum, and thus most of it hits the front end of a lower stage. When that obtains, the impact force is of the same magnitude as the engine thrust that produced said plume. This effect has been known for many decades. It lies at the heart of structural design for jet blast barriers, since the end of WW2 with the very first jets.
Header tanks reduce the magnitude of the sloshing and cavitation effects, but they do NOT eliminate it! There is always free volume within any tank that is not a bladder tank. If that free volume is at your suction, you destroy your turbopump sucking vapor. The header tanks that SpaceX has been using are more about thermal "insulation" with the empty outer tank's shell than anything else. The heat goes into the vapor in the empty tank, instead of evaporating the liquid still in the header. Not perfect, but it works very well. These contain the propellant for subsequent burns.
GW
Last edited by GW Johnson (2023-12-06 19:00:54)
GW Johnson
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"There is nothing as expensive as a dead crew, especially one dead from a bad management decision"
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Since starship and the BFR are based on the Falcon and dragon then making the single large tank into a cluster would with valves would stop all of the issues that are being seen.
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For SpaceNut re #15
Your post is so short, I'm wondering if you might have been suggesting that the solution to the sloshing problem might be to use two sets of tanks in the booster?
One set would be larger than the other, and it would contain the propellant for the boost phase of flight.
The second smaller set would be full to the brim at stage separation, ready to go to work to accomplish the landing.
The trade off would be additional metal for tankage, and additional piping, but no ullage motors would be needed.
I will count this as an additional option in the new topic for bets on what happens for Starship/Booster flight #3.
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I took the header tank issue on, in the final paragraph of my post #14 just above. The second sentence in that paragraph is the key to understanding how this works. The tank is NEVER totally "brim full": that way lies a bursting or explosion danger, because liquids are incompressible. The advantage you get with a header is a shorter settling time, simply because the distance a globule can fall is shorter. You do NOT get positive coverage by liquid at the suction.
GW
GW Johnson
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For GW Johnson re #17 and #14
Thanks for explaining that in your use of the term "header" you were thinking about a separate set of tanks to be used for recovery from staging.
As you have explained many times, and for many forum members, a membrane cannot be used to hold cryogenic liquids, because no membrane is known to exist that can retain it's integrity at cryogenic temperatures.
However, in post #17, it seems to me you have raised a red herring.
I didn't know what that expression means, so here is a Google take on it:
red her·ring
/ˌred ˈheriNG/
noun
noun: red herring; plural noun: red herrings1.
a dried smoked herring, which is turned red by the smoke.
2.
a clue or piece of information that is, or is intended to be, misleading or distracting.
"the book is fast-paced, exciting, and full of red herrings"
h
OK ... I withdraw the expression "red herring". I did not know the etymology and recognize it is inappropriate for this situation.
It seems to me you are missing an opportunity to solve a really difficult technical problem.
It seems to me you took the easy (and understandable) way out by suggesting solid fuel rockets, but I am only putting half my chips on that solution, because I think SpaceNut's idea has merit.
You have pointed out a significant operational issue ... the tank must have some gas space at the top during the ascent to staging.
However, there is absolutely NOTHING to prevent topping off the small tanks with liquid just before staging, and expelling ALL the gas that is present for safety reasons. The small tanks for use at staging could even be OPEN during ascent, and filled with liquid in the main tanks. At the moment just before staging, the valves to those two tanks could be closed, so that liquid cannot escape when Starship shoves HARD on the top of the booster.
After the booster has revolved to the attitude needed for the first burn, the tanks would both be filled with liquid and ready to supply the engines. The valves to the main tanks could then be opened, to allow access to the remaining propellant, which by that time will have settled in the bottom of the respective main tanks.
I am leaving 50 chips on SpaceNut's option.
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... pressurized bladder ... You cannot do that with cryogenics, there are no materials available with the enormous cold elongation capability.
There is one: PolyChloroTriFluoroEthylene (PCTFE), sold by Honeywell under the brand name Clarus. Also sold by Honeywell as Aclar, but that's marketed for pharaceutical packaging, and comes in standard sizes. Clarus and Aclar are made by the same company, in the same factory, on the same machines. However, Clarus is marketed for military and aeorspace applications, and customization is available. Don't get cheap, just buy Clarus. A company in Japan called Daikin makes it under the name Neoflon, however they used that name for all their fluoropolymers so you must specify Neoflon PCTFE. This material will NOT handle hard cryogenic cold of liquid hydrogen, but will handle densified liquid methane and densified liquid oxygen. It's service temperature goes down to -240°C. Liquid hydrogen boiling temperature is -252.78°C at 1 atmosphere pressure. At 13.13 bar pressure absolute (14.50 psia) LH2 boils at -239.96°C, it's critical point. SpaceX chills LCH4 to -180°C, and LOX to -207°C.
PCTFE is the most impermeable to moisture of any known polymer. There are other polymers more impermeable to oxygen, but they become brittle and break in the cold of LOX. A polymer film can be made even more impermeable by vacuum deposition of aluminum. Aluminum clogs the pores of the polymer. Aluminizing the film would definitely be a custom order, so again Clarus.
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For RobertDyck re #19
Thank you for providing the name of a bladder material that might work for Starship, based upon the capabilities you described.
SearchTerm:Bladder cryogenic conditions (methane)
SearchTerm:cryogenic bladder candidate for liquid methane
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If anyone is thinking about what SpaceX will actually ** do ** in preparing for the next launch, I think that GW Johnson's recommendation of solid rockets would be the most feasible to add to the existing stack in the time available.
For that reason, I will shift my bet to 80/20.
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I have already checked on the cold elongation capability of PCTFE and some others. These fall in the 0.5-2% at cryogenic temperatures. For an everting bladder, or for the localized wrinkles in a crumple-type bladder, the elongation at the outside of the wrinkle is nearer 100+%, for an elongation capability about two orders of magnitude short of what is required, if not more.
These materials will crack at cryogenic temperatures, in the type of inherent-wrinkle service that a expulsion-type fuel tank bladder must have. It has to crumple or wrinkle or evert-fold, in order to have only liquid inside it, as the external applied gas pressure drives that liquid out. If it does not serve in that way, then it is only a tank wall inner surface coating, and you still have the same ullage problem that you had without it.
The inside of the fold is in severe compression at an incredibly tiny radius, but the percentage negative elongation is modest. The outer side of the fold is in severe tension at a radius that is about the thickness of the material. The shorter that radius, the higher the required tensile elongation, which at about 2-3 mm is already in the 100-200% range. Thinner is worse, and also weaker against bursting or tearing from localized application of pressure difference upon those wrinkles or folds.
Sorry.
GW
Last edited by GW Johnson (2023-12-10 14:18:07)
GW Johnson
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"There is nothing as expensive as a dead crew, especially one dead from a bad management decision"
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I have not seen anything about that. Technical documents from Honeywell and other just say service temperature goes down to -240°C. Where did you get the elongation stuff? And the electronics industry found soft electronics can break circuits if bent too sharply. Solution was to ensure it cannot bend sharper than a certain safe radius. That means an everting bladder would have a rolling edge that just doesn't bend so sharply that it cracks. The 8.5"x11" samples of PCTFE that I have do not appear to have any problems. But my house is room temperature. As I said, the only technical documents I have been able to find do not have any mention of the issues you raise. Could you post a link?
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For all working this problem ...
The problem to be solved is sloshing of propellant due to G force imposed by the departing Starship.
SpaceX ** will ** solve the problem.
The challenge for members of NewMars is to figure out ** how ** SpaceX will solve the problem, before they do.
Time is short ... we have a bladder concept in play, and we have a solid rocket ullage system in play.
SpaceNut suggested separate tanks for the staging problem. These would be full and open to the main tank prior to staging. Just before staging, the ports to these smaller tanks would be closed, so that propellant cannot slosh around inside them during the negative G excursion. The ports could be opened again after the positive G forces are restored.
This solution would require mass for the smaller tanks, but the mass for the solid rocket motors would not be needed.
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The idea of a flexible bladder, that will remain flexible at cryogenic temperatures and not fatigue over multiple uses, is a tall order. I consider it unlikely.
Some kind of internal subdivision within the tank, preventing slothing, is possible without too much weight penalty. The tank could be divided into cells using thin aluminium spacers. Maybe a combination of this and cold gas ullage thrusters will do the job. The internal aluminium cells would have holes drilled through them, allowing them to drain under thrust. But under staging, enough propellant would remain at the bottom of the tank to allow engine restart, which would restore gravity and drain the cells.
Last edited by Calliban (2023-12-11 16:06:51)
"Plan and prepare for every possibility, and you will never act. It is nobler to have courage as we stumble into half the things we fear than to analyse every possible obstacle and begin nothing. Great things are achieved by embracing great dangers."
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