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Well, they clearly have some sort of problem going on that they don't yet understand.
The fact that something bad happened in a static test tied down to a thrust stand, suggests a possible resonance in a feed plumbing line, not a vehicle structural mode, which would be strongly damped by the fixity of the thrust stand. Such would explain a lot of the fires and engine outs we have been seeing all along. But it takes a thrust or chamber pressure oscillation to do that, one occurring at one of the organ pipe oscillation modes in that feed line.
I really seriously doubt they have any sort of data acquisition that is not digital. You cannot see an organized oscillation hidden in the noise "hash" with digitally-acquired data: the inherent pixellation completely obscures it. You need a high-speed FM tape recording using a recorder of about 1 MHz capability. You play this back at various speeds and magnifications, either as a paper printout or on a scope display, until you find what you are looking for: a definite periodicity with a nontrivial amplitude. It will be a signal hidden among a nonperiodic "hash" of all sorts of amplitudes and frequencies, so it may not initially "leap off the display" at you. You do this with both thrust and pressure data recordings.
The reason I suspect they have a thrust/pressure oscillation going on is because of the 10-Hz shaking of my doors, windows, and hanging light fixtures when they test Raptors, but NOT Merlins, at the McGregor test facility. You cannot hear this signal, but you can feel and observe its effects. 10 Hz is below the lowest frequency humans can hear (but an elephant could).
This is the sort of problem where the basic approach of "build it, break it, build another" will let you down. This is the sort of thing where you have to stop and investigate carefully just to figure out what went wrong. SpaceX is not used to doing that. SpaceX has no one on their payroll old enough to know about how to investigate anything like that. And they have a bad "not invented here" culture that prevents them from learning from the mistakes of others, especially mistakes made long ago.
They faced a problem like this once before with fatal staging problems on Falcon-1, and lucked-out enough to solve it before it could bankrupt them. This thrust oscillation thing is way more subtle than the supersonic drafting thing that caused their Falcon-1 first stages to get sucked into, and collide with, their Falcon-1 second stages.
SpaceX isn't the only New Space participant having serious problems. Just a couple of days ago, another one dropped a Lockheed satellite into the ocean, for lack of enough second stage thrust to reach orbit. The exit expansion cone broke off their rocket engine. Thrust with only a choked throat, is way lower. This is likely an inadequate heat protection design. But apparently the nozzle loss did not impact any regenerative cooling issues (although I do not know for a fact this was a liquid).
GW
Neither is anything I have seen in the science journals.
GW
By the way, has anybody heard anything about when the next orbital test flight might be? I have seen nothing so far.
GW
"Is it possible for a stage on a test stand to emulate the vibrations that occur during POGO?"
Short answer: NO.
Longer answer: being tied to a thrust stand is what locks out many of the structural vibration modes that thrust oscillations might excite otherwise. The Saturn 1st and 2nd stages, and the Titan-II stages, all suffered POGO. It was NEVER detected on the ground.
Note: I think I feel (I cannot hear) thrust oscillations in Raptor static firings in McGregor, but not in Merlin static firings. The frequency is in the 10 Hz vicinity, below the 20-30 Hz lower limit to human hearing. I feel it, and I see it in the vibrations of my windows, doors, and my chandelier-type dining room light fixture. If there are any (any at all!!!) near-10-Hz structural vibration modes in the vehicle, or near-10-Hz organ pipe modes in any of the propellant transfer piping, those could easily be excited by the thrust oscillations. The vehicle near-10-hz vibration modes will be functions of how much propellant has burned off, the organ pipe modes in plumbing will not be.
There may also be (and have been) a sort of in-tank propellant lateral or radial slosh mode, which might uncover the propellant suction points intermittently, when propellant levels in the tanks are very low. This might be "fixed" with the right anti-slosh baffle designs located on aft tank heads to one extent or another, but if the level is too low, there is no fix. The vortex going into the feed pipe will cavitate. We see it in bathtub and sink drains all the time.
GW
The picture in post 6 just above should make clear how to reach Vanguard-1 for retrieval. Whatever craft might be used for this, needs to have (1) a way to get the satellite aboard and stow it safely for the ride home, and (2) radiation protection for the crew if manned, since this orbit penetrates deeply into the lower Van Allen belt. A person would get a pretty fair dose during a time measured in minutes (not hours), especially if that person went EVA to grab the satellite, any time the craft is above 1600 km altitude, which is crudely half the 2-hour+ orbit.
I'd recommend going to 650 km circular LEO, and waiting to synchronize: when Vanguard-1 is there at its perigee with you there also, you accelerate by 0.66 km/s to get onto its orbit adjacent to it. That makes rendezvous easy. I'd recommend a typical rendezvous budget of 0.2 km/s dV "just in case", plus a deorbit dV budget of about 0.1 km/s to come home. I'd deorbit directly from the Vanguard ellipse, preferably at its apogee, that's the lower requirement, but you can do it at perigee if you must. It just takes a much bigger burn (about 0.76 km/s). The whole dV budget for the craft is then just under 1 km/s, if done right. Done wrong, it's closer to 1.8 km/s.
I did this orbit work with the orbits spreadsheet supplied as part of the course materials for the "orbits+" course linked from the "orbital mechanics class traditional" thread. Any of you can use it to do just what I did. I just read the results and sketched the orbits in Windows 2-D "Paintbrush" (the old one). Any of y'all can do something like that too. Most of you are far better with computers than I am.
GW
The study linked in post 524 just above is a follow-up to my earlier computations of entry versus a very wide range of ballistic coefficients. That earlier study document is linked in post 2049 of the "starship is go" thread. In it, you can see the plots made from the entry studies using the entry spreadsheet that is supplied among the "orbits+" course materials. I ran this at constant entry speed, altitude, and angle. Nothing about it surprised me except a constant 6 peak deceleration gees.
In the follow-up study just above, I took those heating and deceleration data across that very wide range of ballistic coefficients, and worked out the distributions of heating round the vehicle, and from those the surface temperature distributions around the vehicle, under the assumption that only re-radiation cooling was going on. From the ballistic coefficients at constant CD, I worked out the mass loadings W/A on the blockage areas (which change, mass does not), and from those the pressure distributions on the windward side of the heat shields.
Note in the data tables of the second figure the mass loadings in US customary. Those fill the same basic role for entry as does wing loading in an airplane. There are no planes (other than models) with wing loadings under 5 lb/sq.ft. The space shuttle and a lot of supersonic military planes have wing loadings at or over 100 lb/sq.ft. None of these ever survived turning broadside to the oncoming wind in flight. Lower wing loading is structurally more fragile.
The temperatures tell the real tale. None of the stagnation temperatures are survivable without either refractive or ablative heat shielding, down to impossibly fragile mass loadings. You might get by at the mass loading corresponding to a light aircraft's wing loading, with bare metal afterbody in the wake zone, but not with aluminum, and certainly not with any sorts of elastomers or organic-matrix composites. You will need stainless steel, or something even better, like Rene 41. And structural strength to go dead broadside to the wind is in very serious question! If you don't go dead broadside, you don't decelerate!
I did not address the structural question. I merely point out that (1) it is there, and (2) it is very serious.
GW
The link to the document in post 226 works just fine.
GW
That link in the quote in post 2060 just above does not take me to anything I ever sketched. It takes me to the imgur webpage, not to any particular image.
Meanwhile, just to let everybody know, SpaceX has been testing a lot of Raptor engines lately, quite a few up on the tower stand where they are quite loud. There's about a 10-Hertz inaudible pressure wave from these tests that really rattles the windows and doors and chandelier-type light fixtures. I feel it with Raptors, but not Merlins. I suspect the soot cloud in the flame with kerosene has an oscillation damping effect that is missing with the methane.
That too-low-to-hear but significant amplitude signal suggests a nontrivial thrust oscillation with a fairly-well-defined 10 Hertz frequency in the Raptor that the Merlins just do not have. If there are any vehicle or plumbing modes near that 10-Hertz frequency, resonance could easily cause POGO problems.
Those tests have been running pretty close to 6-7 minutes each, essentially a full duration ascent burn for either stage. I cannot tell a vacuum Raptor from a sea level Raptor, they all sound about the same, from 6 miles away. I suspect the vacuum tests are running with something approximating a normal shock, essentially just barely aft of the exit plane. They would have to be at near full pressure, and would be separated between ignition at some reduced pressure, and throttle-up to test conditions. I usually don't hear the "explosions" that were full-power starts. Von Braun found reduced-throttle ignition to be far more survivable, about 80 years ago.
As for Bob's question, the object in the wake zone behind an extended heat shield sees an exposure to exactly the same gas/plasma effective temperature as exists at the stagnation point, of value in degrees K numerically equal to flight speed in m/s, to about 10% accuracy. At 8 km/s you are looking at about 8000 K effective plasma temperature. At 4 km/s, 4000 K. Etc.
This high effective temperature persists well into the flow alongside the wake zone. The gas in the wake zone is still at that same temperature, just not moving very fast at all relative to the surface. The convective heating rates in the wake are much lower than stagnation, because there is less or almost no fluid scrubbing effects, and the pressure (density) is nil. A crude estimate is stagnation convection/10. The plasma radiation heating rates are also lower, because the radiating plasma shock layer is more distant from the surfaces at risk, yet it still "shines" on everything. A crude estimate in the separated wake zone would stagnation radiation heating/2.
On the windward side of your extended heat shield, there is attached flow at significant pressure (density) all the way out to the rim, and the radiating shocked plasma layer is very close to all of the windward surface. Crude estimates near the rim might be stagnation convection/3, and the same radiation heating rate as the stagnation radiation heating rate, just all over the heat shield, right out to the rim.
The "nose radius" of the heat shield affects the stagnation heating values, but in opposite ways for convection and radiation. Sharper nose radius rapidly increases convective stagnation heating. Flatter nose radii greatly increase stagnation plasma radiation heating. Exactly how those balance out is both atmosphere model-dependent, and entry trajectory details-dependent.
The net result is quite different at Mars than Earth. Earth heat shields do better as more-or-less spherical segments with large "nose" radii. Mars heat shields do better as conical shapes with slightly-blunted tips. It's not one-size-fits-all, not by a long shot! Every situation is different!
GW
Actually, I quite agree with you.
GW
Bob:
The final results are excruciatingly-sensitive to the speed and angle at entry interface, something I did not vary in this study. Just a tad faster or steeper raises gees and peak heating. Just a tad slower or shallower lowers peak gees and peak heating.
Dig out the entry spreadsheet posted in the "interplanetary transportation" topic, "orbit mechanics class traditional" thread., and run it for yourself. It has a user's manual. Just be sure to use realistic inputs. It's a "garbage-in, garbage-out" thing.
Kbd512:
Nextel now makes several ceramic cloths, all pretty much intended for aircraft engine nacelle fire curtain application. I am very familiar with the oldest one, an AF-19 fabric made from Nextel-312 fibers, which are aluminosilicate minerals. All these have a solid phase change at about 2300 F, leading to ~3% volume shrinkage, and catastrophic embrittlement. Which upon cooldown renders them extremely fragile, crumbling away to dust at the touch of the slightest breath of air. Those minerals melt at about 3300 F, but you cannot reuse those cloths (or anything else you make from them) if you exceed about 2200-2300 F.
GW
RE: post 2049 just above.
The image number labels are at the tops of the images. They are out of order in the posting.
The one labeled “image 8” was intended to be the first one, showing exactly what and how I was trying to do in the study. What I varied, how I varied it, and more importantly, what did not vary. I kept the heat shield nose radius ratioed to its diameter constant, as that nose radius is a major influence on the heating. Larger is lower heating. Had it not increased, the heating results would have been very much higher.
There were 4 images that were entry trajectory analysis results illustrations, to show where the numbers came from. The one labeled “image 5” was the first, with heat shield diameter 4 m equal to the 4 m base diameter of the protected object. This corresponds to a ballistic coefficient of 300 kg/sq.m, comparable to the Apollo command module’s 313.
The one labeled “image 3” was intended to be next, showing an 8 m dia heat shield on that same 4 m dia object, for a ballistic coefficient of 75 kg/sq.m.
I doubled the heat shield dia again to 16 m, for that same 4 m dia object, in the image labeled “image 6”. That produced a ballistic coefficient of 18.75 kg/sq.m.
I doubled the heat shield diameter yet again to 32 m for the same 4 m dia object, for a ballistic coefficient of only 4.68+ kg/sq.m. That one is labeled as “image 1”.
These 4 analyses produced peak convective heating rates at stagnation, and peak plasma radiation heating rates at stagnation, that were relatively trivial. I summed these to a total heating rate at stagnation, and computed the stagnation point surface temperatures for which thermal re-radiation equaled the combined convective-radiation input, without ablation, without conduction into the interior, and without any other cooling at all. Those results are summarized in plot form in the image labeled “image 2”. The temperatures decrease but stay rather high until the ballistic coefficient falls below about 50 kg./sq.m.
I also used the peak deceleration gees to estimate the peak average pressure exerted across the heat shield. Those results are also plotted in “image 2”. I was surprised to see the peak deceleration gee being the same for all cases. I expected higher gee with more heat shield area. But the thinner air higher up apparently just offsets that effect.
I went back and updated these results by estimating temperatures away from the stagnation point with the results plotted in “image 4”. The stagnation point plot is there, plus a plot representing attached flow on the windward side of the heat shield near its rim, and another plot representing any of the leeward heat shield surfaces, or the lateral surfaces of the protected object. All of these are surfaces within the separated wake zone.
Looking at those results in “image 4”, at ballistic coefficient near 50 kg/sq.m, the stagnation point temperature is at least 1000 C. Higher if reflective. Near the rim of the heat shield, that is still at least 800 C, higher if reflective. Only for the separated wake zone surfaces is it 500 C, yet higher if reflective. There are lower temperatures shown in the plots at lower ballistic coefficients, but the structures may well be getting too fragile to fly, despite the lower average pressure below the 4 KPa at 50 kg/sq.m.
I put together a little table of max service temperatures for several materials, which got posted as “image 7”. It’s not in any way comprehensive. With a rim temperature of 800 C or higher, there is simply no way in hell that any sort of silicone elastomer on any sort of fabric material, is going to be in reusable shape after a single entry from low Earth orbit! That is the real takeaway here!
There is a very good reason that practical spacecraft heat shields have all been ablatives up to now, the only exceptions being the refractory ceramics on the Space Shuttle and the X-37B. I think I just showed you exactly why that has been true. By doing the same kind of things H. Julian Allen and A. J. Eggers were doing for warhead entry scenarios, back in the early 1950’s. The physics has not changed. Some of the materials have.
GW
I looked up the Stoke Space "Nova" design. It's more-or-less clear how the first stage is to work, rather similar to the Falcon cores and Superheavy, just using landing legs of some sort, unlike Superheavy, but like the Falcon cores.
It's the design of the "Andromeda" second stage that still seems to be in flux as testing proceeds. It would have some sort of metal base that is the heat shield, cooled by the hydrogen that fuels the circumference-mounted second stage thruster engines. I cannot really tell if it is simply to be cooled by backside hydrogen flow, or if it is transpiration-cooled by leaking hydrogen through the heat shield surface. I'm not sure Stoke knows yet, either.
The numbers as best we know them say that either regenerative cooling or transpiration cooling should work, but neither concept has ever actually flown and been tested in a real entry. Those and some other ideas were supposed to have been tested on the old X-20 "Dyna-Soar" spaceplane, but it was cancelled in 1963 before it could ever fly. X-20 had bare-metal exotic-alloy wings fuselage, and tail fins at the wingtips. It had leading edges and a nosetip made of graphite ablative, modified with rods (metal, ceramic, not sure) to conduct heat inward better. Graphite is both a poor insulator and a poor conductor of heat. It starts burning in air at about 5000 F, though. We got away with it as a slow-eroding throat insert in one-use solid rocket nozzles up to 6000 F gas temperatures, though.
GW
Quite right, Tom.
As for the topic, I finished and posted the ballistic coefficient study that Bob Clark was interested in, over at my "exrocketman" site. It is titled "Ballistic Coefficient Study for Earth Entry", and dated 4-12-2025. Top of the stack right now. It has 6 figures in it, which precluded my posting it here. The one with the surface temperature and pressure results that Bob needs is the last one. I emailed him to let him know it was posted.
GW
The question of whether any given material can survive entry without adding extra heat protection is what I am trying to get to, with the entry study I am doing. The usual criteria for heat shield (or other) material selection are (1) the equilibrium surface temperature, and (2) the applied pressure, both of which it must withstand.
The question of whether a super-large lifting (or drag) surface can survive entry is a question of aerodynamic loads vs material strengths, as much as it is surviving temperatures and pressures. I am NOT addressing those issues! I leave that to others. Be aware that any given design must survive in both scenarios.
Also be aware that anything riding unprotected in the wake zone behind a heat shield of any kind, is exposed to plasma at the same effective temperature as that seen by the stagnation zone (numerically the speed in m/s equals the effective temperature in degrees K, to about 10% accuracy). The difference is a far lower heat transfer coefficient leading to far lower heating rates because the scrubbing action is far less away from the stagnation point.
This effect allowed Mercury and Gemini to have bare metal afterbodies of high-temperature exotic alloys, separated from the pressure shell by mineral wool insulation, coming back at roughly 7.74 km/s. Apollo returning from the moon at about 10.9 km/s could not; its afterbody was coated with cell-gunned Avcoat insulating ablative, same material as the windward-side heat shield, just thinner.
I will post my entry study on my "exrocketman" blog site once it is done. That will be soon. It will have a handful of figures in it, which makes it impossible to post directly here on the forums. Without those figures, the words will seem meaningless.
And by the way, the entry spreadsheet I am using to do this is the same one supplied with the "orbits+" course materials offered through these forums. That spreadsheet and its user's manual are supplied as course materials for lesson 7/7B. Using that tool, anyone one of you can do what I am doing! There's enough information in the user's manual to enable you to do that!
GW
The Tesla dealerships and owners did nothing to deserve what I consider to be domestic terrorism. Musk's atrocious behavior is the cause of those radicals acting out.
I would add that there have been a lot of Tesla owners who are appalled at Musk, and selling their vehicles. They clearly do not want to be associated in any way with him.
This is not "the left" as you termed it, but a small radical faction, similar to the right wing radicals that also killed people and wrought damage at various rallies, too. It is radicals of either stripe that are the "bad guys".
GW
I predict massive changes in NASA's space agenda under the new guy. I rather suspect that SLS, Starliner, and maybe even Orion, are going to go away, and sooner than anyone suspects. Artemis as we know it will also go away soon, but it will not vanish. It will continue as something embodied very differently, but more than "flags-and-footprints", to counter the Chinese. The name may change.
Meanwhile, Mars will become a minimal "flags-and-footprints" mission, because Trump is uninterested in anything but gaining Earthly wealth and power. That plus Musk's deserved public unpopularity will drive the wedge between him and Musk rather deeper. I also predict Musk will no longer be part of Trump's administration in a year or so.
As near as I can tell, Musk's public image is so bad (and deservedly so) that Tesla's prospects are pretty much finished. Although it takes time for the huge dead dinosaur to fall over. SpaceX may survive this, if Gwynn Shotwell can keep doing the right things without being over-ruled by Musk. Tesla has no Shotwell, so they have no chance. SpaceX does have a chance, but it's not a "sure thing". Musk is Trump's "sacrificial lamb" to avoid taking all the blame himself for all the chaos he is causing.
GW
Bear in mind that lifting bodies have a substantially-lower lift curve slope that delta wings, in turn less than swept wings, and that in turn less than straight wings. There is more to available lift than just wing area! By around a factor of 2 to 3! Planform shape is a big influence, as is wing section shape (lifting body vs airfoil). Especially considering the differences between subsonic and supersonic/hypersonic designs!
That being said, I have been looking closely at Bob's suggestion that low ballistic coefficient might eliminate the need for heat shielding. I am running a generic study across a very wide range of ballistic coefficients, to see what the peak stagnation heating rates and peak deceleration gees look like, as well as end-of-hypersonics altitudes, for fixed entry speed and entry angle below horizontal, in Earth's atmosphere. These are at fixed mass and hypersonic drag coefficient, with a fixed "nose radius"/diameter ratio. I vary diameter. Just generic, but well within the ballpark.
I will not take this through any realistic vehicle designs, but it will provide design constraints in terms of surface stagnation zone temperatures, and average pressures across the heat shield at peak deceleration. Both are crucial heat shield parameters.
GW
I do not know what that heating system might really look like. I have yet to even bound the heat transfer problem, which will be the "long pole in the tent".
You have to understand, heat transfer, especially via gaseous media, is about the slowest physical process that we know. The proof of that thesis is the vast difference in time it takes 32 F air to kill you, versus 32 F water. And that transfer process is termed "convection", not "conduction". Conduction occurs within solid media that are in intimate contact.
GW
Retrieving that satellite is an interesting idea. Worthy of pursuit at some level!
Not the least would be an evaluation of those materials, particularly the solar cells, after so many years exposed to space and Van Allen belt radiation. 67 years is a long time.
Does anyone have current data for the current values of its orbit? The most import would be min and max distances and the inclination. I could run the intercept orbit dV's with the orbits spreadsheet from the "orbits+" course materials.
This thing is a sphere about 2 feet to a cubit in diameter. It actually resembled Sputnik 1 in shape, although smaller in size, if memory serves. It's not that heavy, a single person could lift it.
GW
The first time I know of for point-to-point rocket delivery was Von Braun attempting rocket-delivered mail in the 1930's, before he came to the attention of the Wehrmacht.
GW
The nozzle needs cooling because the most practical materials to build it are metals, and none of those survive 3000 K temperatures. The carbon-based chamber material can survive 3000 K temperatures, and if it has the strength to contain 5-atm-class pressures at such temperatures, then it needs no cooling.
Thermal expansion is going to be a real bugaboo with such a design, if you try to affix the engine to the vehicle at more than one axial station. I wouldn't do that myself, but I would make the mounting point and hardware very strong and stiff, so as to maintain firm control over the thrust vector direction line passing through the vehicle cg, no matter what thermal expansion occurs.
There is a downside and a work-around: firm mounting precludes easy gimbal-type thrust vectoring! You will need some other form of attitude control.
I would suggest attitude control thrusters. You can tap some hydrogen off the main chamber to run them, up nearer the cooler front end,
so the not-so-hot gas is easier to handle with metal tubing and valves.
Attitude thrusters are how we had positive flight control on the 4-stage solid-propellant "Scout" launcher without trying to do flexible-structure vectorable nozzles on the solid motors! Those were very simple, very lightweight hydrogen peroxide monopropellant decomposition thrusters. Electrically-heated catalyst beds in the thruster chambers. This required high-test peroxide, but the storage time before use was short enough to avoid that danger.
Trying to do vectorable nozzles on solid motors was very expensive, usually way too heavy in small sizes, and added multiple possible leak failure modes to the design. Only the strategic-size motors had them. They are usually hydraulically actuated, requiring both substantial motive power and some very heavy equipment to run them.
GW
The water column height that makes this work is not the top to bottom height of the J-tube. Is is the difference in elevations between the outside entrance surface and the habitat floor level surface inside. There does not need to be a deep bend.
Put the habitat instead just under the crater floor, and the entrance instead 13.5 m (less the hab floor burial depth) higher up the adjacent crater wall, and you do not even need the ladder!
Biggest problem will be the water freezing, or the water evaporating at a high rate from the exposed water surface at the entrance. I'm not sure which will dominate.
But the local atmospheric pressure on Mars is 6 to 7 mbar generally. That's not enough to stop violent boiling away of the water. Even if it is a bit higher in some areas so that the violent boiling didn't occur, you would still have a very high evaporation rate!
The only thing that stops the (non-violent-boiling) high evaporation rate is when the vapor pressure of the water in the local air, not just the local air pressure, exceeds the equilibrium vapor pressure of just-at-freezepoint water. Warm the water and that vapor pressure requirement is even higher.
What you really need for such an application is a liquid with almost no vapor pressure at all to slow its evaporation rate, a much lower freezepoint to limit freezing effects, and preferably a much higher density. The higher density would lower the height difference between the exposed liquid surfaces, making the design far easier to implement.
Mercury immediately comes to mind, but with all the toxicity problems it brings.
GW
Rob:
I don't know what "Starship siphon" refers to. But I think your suggestion of propellant slosh may be right. That surface is variably sloshing around, which may transiently uncover the suction intakes for the turbopumps when the propellant level is low enough. That would be a separate problem from the possible POGO structural instability driven by thrust oscillations.
Bob:
My old silicate composite material has a limited temperature range and while stronger than NASA's shuttle tiles, limited strength. It was something that worked by chance for me 4 decades ago, but was never developed since.
Ed Pope has several modern unique materials for very harsh temperature and structural environments, all of which are better than my old stuff. His zirconium carbide oxide material in particular caught my eye.
GW
Max efficiency by itself is a very misleading measure! It also has to survive extended operation at full soak-out temperatures, and for a duration that suggests a decent lifetime. The ones that meet those criteria have yet to exceed about 27% conversion efficiency. So says AAAS's "Science" Journal, which is peer-reviewed to a fault.
GW
Today I finally did make the video work of Flight Test 8 on the SpaceX website. I DO NOT do X or any of that other social media BS. I never will.
I saw no surprises, except that the Starship upper stage loss occurred right before the end of the ascent burn, with little propellant left aboard. I did see first stage engine losses on both ascent and descent, excepting the 3 gimballed engines dead center. That makes me wonder if we didn't see POGO in both stages.
If the control software is formulated correctly, ascent oscillations risking engine damage might cause shutdowns that do not recur during descent. Except that there were 2 engines that failed to restart during the 13-engine initial boost-back burn. 11 were enough to slow it so that the "design 3" could recover it on the tower.
The upper stage Starship did not seem to suffer problems until right before the end of its ascent burn onto the ascent ellipse. Then it lost 1 vac engine, then all 3 SL engines (with the thrust vector gimbal capability), then another vac engine, then it obviously tumbled totally out of control until self-destruct ended this.
Bear in mind that this ascent ellipse has apogee at LEO altitude, but a perigee at, or very near, the surface. THAT is what ensures automatic de-orbit, if you do not do a circularization burn at apogee. The ascent ellipse perigee is well down in the atmosphere.
GW