sevenperforce
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Posts posted by sevenperforce
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22 hours ago, tater said:
I'm not changing my forum theme to read the post (can see it in your quote), but is that the booster, or the ship? In the case of the ship—doesn't work on Mars. Off the table. For the booster, you are of course exactly right.
You can always Ctrl+A to select all text on the page and read it quickly without changing your forum theme.
On a related note, thanks for pointing out that there is a dark theme. Much easier on the eyes.
Our polymath fellow is proposing wings for Superheavy, which is a nonstarter in just so many ways.
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15 minutes ago, Exoscientist said:
If you see out of nominal state for one or more engines, you can shut the test down, examine the engines not operating properly and compare to the ones that are.
This is very helpful if you do not have a well-instrumented engine with modern realtime telemetry and must physically inspect parts to see what went wrong.
This is much less helpful when you have so much advanced instrumentation and realtime telemetry that you typically don't perform physical inspections of your parts even after a static test because you already know what went wrong before the vehicle can even be detanked.
17 minutes ago, Exoscientist said:SpaceX should lower the Raptor chamber pressure and thrust level.https://exoscientist.blogspot.com/2024/01/towards-advancing-spacex-starship-to.html
[snip]
I advise first start with reducing the dry masses by optimally lightweighting the expendable versions of both stages. Surprisingly this gives a greater expendable payload than the expendable payload of the current version.
It does not.
You cannot take the mass ratio of an aspirational upper stage and imagine that it will simply scale to the same mass ratio for a lower stage. Lower stages have to be stronger than upper stages. That's basic.
20 minutes ago, Exoscientist said:I suggest using winged, horizontal approach to reusability gives a much reduced payload loss due to reusability.
If you're removing mass left and right for an expendable version, it's not going to be able to support itself horizontally.
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On 1/6/2024 at 3:15 PM, RCgothic said:
Even if it was a knife-edge of mostly exactly 120s tests with a few falling short (which this data isn't), that would *still* not say anything about raptor's reliability because 3rd party observers have no idea what's being tested or what the abort criteria are.
It could be GSE faults. It could be test aborts more conservative than flight. It could be testing above 100% throttle. It could be deliberate tests to failure.
To expand on this point, let's say that @Exoscientist's speculation was 100% true -- the majority of these tests are aiming for exactly 120 seconds, and some significant number of those tests fall short of 120 seconds because the engines spontaneously fail somewhere between 110 and 115 seconds (or something like that). Raptor is clearly unreliable, right?
Nope, this still doesn't provide any meaningful evidence that Raptor is too unreliable for flights, let alone flight tests. For example, let us imagine that the Raptor manufacturing process has some chance (let's say 3%) of introducing a fatal defect in the turbopump exhaust injection manifold, and that defect is undetectable except through static testing. Let us further suppose that engines with the defect have a 99.999% chance of failing before 120 seconds, and that engines without the defect cannot develop the defect by static fire testing. Because SpaceX is hardware-rich, they could simply elect to static test each of their engines a few times and throw out the bad ones. This would be a completely reasonable way to eliminate the defect, if that's the way they chose to do it.
Absolutely nothing can be inferred from these observations, other than the fact that SpaceX has an active, healthy, aggressive test-firing program.
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1 hour ago, magnemoe said:
In part this is why Rocketlab want to reuse electron but here it might be more about the composite stage.
I think you're correct. If I remember accurately, Rocketlab is fairly flush with engines but not so flush with stages, the opposite of ULA (which can churn out stages rather quickly but is depending on BO for the engines). Might explain why Rocketlab is less worried about saltwater incursion in its engines, because it can just replace them if they don't perform well in test fires. The Rutherford engines are probably more accepting of harsh conditions because they lack combustion-based turbopumps.
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12 minutes ago, StrandedonEarth said:
Congrats to ULA for a successful debut of Vulcan! Now, how long until S.M.A.R.T. (better than nothing) reuse becomes a thing? And how long after that will they decide S.M.A.R.T. is as economical as STS SRB re-use (as in, not really)?
One of the primary reasons to pursue SMART is to increase launch cadence, not decrease launch costs. Blue Origin is not exactly flush with engines and their factory isn't moving very fast, so if ULA (or its successor) wants to get a reasonably high launch cadence for launching the Kuiper satellite constellation, they might need to recover the engines simply to avoid schedule delays.
If they can recover and refurbish the engines faster than Blue Origin can build and ship them, then SMART makes a lot of sense for that reason alone, even if the cost savings are marginal.
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6 hours ago, Ultimate Steve said:
If this goes right, a small wafer with my name engraved onto it (and an SD card with a picture I made on it) will be going to the Moon.
It's got a photo of me and my kids, too.
2 hours ago, JoeSchmuckatelli said:I'm just glad to see the US have two domestic produced orbital class rocket engine manufacturers.
Not a fan of monopolies.
Competition is gud!
All Hail Sherman
6 hours ago, Ultimate Steve said:Man, we've had so much Falcon 9 saturation for the past several years that I forgot just how much of a sporty vehicle Falcon 9 is. Vulcan has like a 5 minute first stage burn, a 20-25 second stage separation sequence from cutoff to startup, and a 10 minute second stage burn.
Not that it's bad, there are advantages to a sustainer architecture, but I just kind of haven't seen a sustainer rocket go up in a while (and had kind of assumed Vulcan was a tad sportier than it is).
Vulcan Centaur 0 (the one with no SRBs) can't quite get off the ground on its own. The BE-4s are underpowered for launching it single-stick, so they have to launch it partially detanked. With two SRBs (like in this launch), they can fully load the first-stage tanks but it's still rather slow getting up.
Sustainer architectures are weird.
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10 hours ago, Exoscientist said:
Such tests are more challenging for the engines and the stages. But that is the point.
I am pretty damn sure that all-up, full-mission-duration static fire tests are NOT more challenging for engines and stages than all-up, full-mission-duration launch tests.
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2 hours ago, Exoscientist said:
But going by counting the number of tests for the Raptor that fail to reach that 115 to 120 second mark, it may be 1 in 5 to 1 in 6 fail to reach it. Note as the author of the video observes some tests are planned to be shorter. For some for instance they were intended to be about 47 seconds long. But there are a block of tests I marked off in the attached image that appear to be aiming for that 115 to 120 second mark, and several of them don’t make it. I estimate 5 or 6 out of the 30 I marked off failed to reach that planned burn length.
Another questionable issue of these static tests is the planned lengths. The largest portion them were of a planned length of about 120 seconds, 2 minutes. But judging by the two test flights the actual burn time for the booster is in the range of 2 minutes 39 seconds to 2 minutes 49 seconds range. Only very few of the test stand burns went this long or longer.
The video gives a link where you can watch the test stand burns NSF.live/McGregor. Another useful aspect here is you may be able to judge the power level of the burns. There is a graphic that shows the sound level of the burns. From that you may be able to judge whether or not the engines were firing at or close to full thrust.
In the image below, the burns in white are those shorter burns of about 47 second lengths the author of the video made note of. They may be tests of the boost back or landing burns. The ones I’m commenting on are under the yellow bar, which I estimate to be at about the120 burn time. There 5 or 6 out of 30 don’t reach the planned burned time.
Dude, this is hogwash. You have absolutely no idea what the planned burn time was for any of these tests, what was being tested, whether these were acceptance tests or tests to failure or outlier tests...nothing.
You're looking for patterns that don't exist. You might as well throw in your lot with the day-trading dopes arguing about which candles predict a new stock market trend.
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8 minutes ago, mikegarrison said:
But they had already flown the Saturn I, which used the third stage of the Saturn V as its second stage.
Sort of, but not entirely. The Saturn I used an RL10-based second stage, while the Saturn IB used the same basic upper stage with a J-2 engine, which became the S-IVB-200. It flew a few times before Apollo 4. However, the third stage that flew on the Saturn V in the Apollo 4 test was a different configuration, the S-IVB-500, with a flared interstage, a different helium pressurization system, a new auxiliary propulsion system, and a different separation system.
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1 minute ago, zolotiyeruki said:
During the space race, the concepts of rapid iteration and "fail fast" weren't even conceivable. Because each iteration took so much time and money to create, everything had to be as perfect and complete as possible before it could even be tested. That in turn made things even more expensive and slow. SpaceX have recognized that the ability to rapidly and cheaply iterate means that failure is an option. Unlike any space program before, it's ok if it doesn't work perfectly the first time, or the second, or the fifth.
I'll even go one step further and point out that the development of the Saturn V actually deviated significantly from prior US launch vehicle development by doing an all-up test on the first actual launch. Prior to the Apollo program, virtually all rockets were tested one stage at a time. The first stage would be ground-tested, then test-launched with a dummy upper stage and payload. The second stage would then be ground-tested, and only after all of this would it be stacked onto the first stage with a dummy payload for an integrated flight test. For three-stage rockets, this would proceed even slower (first stage with two dummy stages, then first two stages with one dummy stage, then an all-up test with a dummy payload, then a true integrated test launch).
The Apollo program deviated dramatically from this approach by doing an integrated flight test of all three stages AND a functional CSM on Apollo 4. They focused on validating all of the systems independently (and in parallel) so that they would be able to put everything together on the first go. And of course the Apollo program was wildly successful.
So if there is a "lesson learned" from Apollo, perhaps it is the lesson that deviating from past practices can be a really good idea if you have a consistent vision and the resources to make it work.
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11 hours ago, AckSed said:
It might be a series of requirements: IF hydrogen AND low-density second stage EQUALS low consumption of fuel for cooling heatshield PLUS lighter heatshield AND high vacuum specific impulse... THEN hydrogen.
Also, keep in mind: a lot of start ups gravitate toward kerosene for their first stage (or even solids). Because of the high density of kerosene, you tend to have a much thinner rocket. Not only are they going for full reusability, but they want a methane first stage, so it makes sense to use a wider stage due to methane’s low density, which in turn gives them more volume to work with for their hydrogen upper stage.
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14 minutes ago, Errol said:
First, thank you for the detailed reply.
No problem.
14 minutes ago, Errol said:Initially I had tried this (not by hand, my math is shaky at best, so I used a TWR calculator script I found online). I wasn't sure if you are supposed to compensate for lunar gravity by adjusting just the weight or both the thrust and the weight.
I would suggest making a spreadsheet, carefully laying out the dry mass and propellant capacity of each stage, and so forth.
How you compensate for lunar gravity is up to you; just keep it consistent. The TWR is the total thrust of all firing engines divided by the weight of the vehicle; the weight of the vehicle is its mass times the gravitational acceleration. Personally, I just imagine it's Earth gravity for everything and then if I am dealing with lunar touchdown or takeoff I apply the appropriate transformation afterward.
14 minutes ago, Errol said:I had heard about the ascent module not ever being the active translational actor during docking, but wasn't the ascent stage responsible for all of the phasing/catch-up/rendezvous maneuvers, only switching to station keeping once close enough for final approach to be initiated by the CSM?
Yes, that's correct.
The lunar module ascent had two phases: a ten-second burn straight up to clear terrain, and then a pitchover and burn to orbit. The precise moment and orientation of liftoff was chosen to minimize phasing and plane changes. If the ascent propulsion engine had failed during the last thirty seconds of the orbital insertion, the four little aft-facing RCS thrusters had sufficient umph to complete orbital insertion. That was one of the reasons for the valve that would allow the main tank propellants to flow directly into the RCS thrusters.
14 minutes ago, Errol said:[snip]
Also, if the ascent stage was responsible for rendezvous maneuvers like I asked about above, does this mean that they only used the RCS translation for those burns?
Yes; following successful orbital insertion, the rendezvous maneuvers were performed entirely by the RCS system. Although the RCS system had three-plane translational capability, all of these maneuvers were being done pretty much manually, often using slide rules to compute the proper heading, time, and duration of each burn. As a result, the aft-facing RCS thrusters were the only ones used for these maneuvers; the other thrusters only provided attitude control.
For Apollo 11, the initial launch reached an 87.6x17.6 km elliptical orbit at main propulsion system burnout. RCS was used about an hour later, at apoapsis, to circularize. They had planned for a plane change burn as well but didn't need it. By this point, they were nearly on a collision course with the CSM so they did a Constant Delta Height burn (still with RCS) which ensured that the two vehicles would be constantly 28 km apart and they would have time to plan the remaining rendezvous burns.
As an example -- the RCS circulation burn for Apollo 11 required 15.7 m/s of dV or around 13.7 kg of propellant. The burn took just over two minutes, starting at MET 125:19:34.70 and ending at around MET 125:21:36. Had the burn been performed with the actual ascent propulsion engine, it would have taken less than three seconds.
14 minutes ago, Errol said:Do you know if there were any other procedures required if this valve was going to be used? Was there some sort of additional pressurization hardware (like a bladder or something) to enable this, otherwise how would the main tanks deal with ullage?
I don't know the specific procedures but it definitely all would have happened pretty quickly, so it was likely automated.
The RCS propellant tanks used teflon bladders to hold the propellant inside a pressurized tank. As helium pressurant was vented into the tank, the teflon bladders were compressed and pushed the RCS propellant out, obviating the need for any separate RCS ullage burn. If all of the RCS tanks had COMPLETELY failed and the main tank was being used exclusively, then ullage burns would become a problem. There was a chance that the surface tension of the propellant in the feed lines would be enough for the initial ullage puff but it would have been tricky. But that was an unlikely contingency (lots of other stuff would have to fail which would probably be LOCV anyway).
14 minutes ago, Errol said:On 1/3/2024 at 5:08 PM, sevenperforce said:Although the ascent propulsion system was technically restartable, it was not ordinarily restarted during ascents because the helium pressurization for restarts was a little tricky: it was a single burn from the lunar surface to lunar orbit. It was restarted for disposal burns, and required an RCS ullage burn for those restarts.
What was tricky about it?
It used multiple helium tanks with burst discs (for simplicity) so it could only be started twice.
14 minutes ago, Errol said:Does this mean that the CSM RCS was used for the TLI ullage burn?
No, definitely not. The CSM never did anything until separation from the third stage.
The S-IVB third stage of the Saturn V was equipped with a pair of Auxiliary Propulsion System modules using hypergolic propellants. Each module carried around 120 kg of propellant and boasted a trio of RCS engines and a single ullage engine. They were used to provide ullage for third-stage restarts, roll control (and backup pitch/yaw control) during the third stage burns, and general attitude control during the transposition and docking maneuver.
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On 1/3/2024 at 12:06 PM, Exoscientist said:
As long as Raptors fail in flight tests, there will be questions about its reliability.
Why use "flight tests" here? Why not just say "as long as Raptors fail, there will be questions about its reliability" period? How is a flight test magically different from a static test?
On 1/3/2024 at 12:06 PM, Exoscientist said:My opinion a moon rocket should be A moon rocket(singular). We did this 50 years ago. There is no reason why we can’t do that now.
No one wants a sortie lander. To achieve the goals of Artemis, we need substantial downmass -- something closer to the 46 tonne (launch) mass of the Altair lander. Add a reasonably-sized crew capsule, and doing this in a single launch means a vehicle capable of throwing upwards of 70 tonnes to TLI.
We also know you dislike solids generally, so let's look at an all-liquid architecture. Three stages are going to be required, obviously. Rule of thumb splits delta-v among stages, and you need 9.4 km/s to reach orbit and 3.2 km/s to reach TLI, so that's a total of 12.6 km/s, or 4.2 km/s on each stage. Notionally, let's imagine a kerolox first stage, a methalox second stage, and a hydrolox third stage.
The EUS packs 126 tonnes of hydrolox, so let's start there. With a 75-tonne payload, it develops just 3.5 km/s. If we swap out its four RL10C-3s for a pair of J-2Xs, it will get even less. But the EUS is the largest hydrolox upper stage in development, so let's stick with that. Now our first two stages need to somehow deliver 4.5 km/s each.
Now for our notional methalox second stage. We know you dislike Raptor, so let's imagine a BE-4U with improbably comparable specific impulse to RVac. Let's imagine a 45-tonne stripped-down Starship as our main tank, pushed by a quincux of BE-4Us. The BE-4 produces 2.4 MN at an estimated sea level specific impulse of 315 seconds, so if we bump that up to 380 seconds then five of them will produce a whopping 14.5 MN together. With the required nozzle extension they're going to weigh in around 4 tonnes each. The third stage and payload together come in at 232 tonnes so stage dry mass is 297 tonnes, which gets us a delightful 5.5 km/s from the 1000 tonnes of methalox onboard (reduced from 1200 tonnes because BE-4 can't take densified methalox). We've got a T/W ratio of 1.1:1, which is low but acceptable. ETwo stages deep, and we're ahead of schedule! Our first stage will only need to deliver 3.6 km/s (or roughly 4.1 km/s if you want to treat sea level specific impulse as pressure drag and just use vacuum specific impulse).
But here we have a problem. The most powerful kerolox engine in the world, the RD-171, gets 337.2 seconds of vacuum specific impulse. You need a wet:dry mass ratio of 3.5:1 to achieve 4.1 km/s, putting us at a liftoff mass of at least 4,500 tonnes, almost double the launch mass of the Saturn V. You'd need at least nine RD-171s or at least 17 RD-180s to get off the ground. Or you can use a cluster of 77 Merlin 1Ds.
Good luck with that.
On 1/3/2024 at 12:19 PM, Exoscientist said:I major irritation of mine is that SpaceX dismisses the lessons of Apollo. It dismissed the importance of a flame trench and dismissed the importance of having powered stage separation.
How were these "lessons" from Apollo?
The Saturn V used a flame trench and powered stage separation, but these weren't lessons. NASA didn't attempt to go without powered stage separation and then fail and correct it; they just chose to use solid separation motors from the start. You seem to be confusing "lessons learned" (e.g., "we tried it one way and realized it didn't work and so we found a better way") with standard operating procedure ("this is the way we decided to do it and it worked").
Besides, Superheavy has always had a flame "trench". It's actually six very big trenches that go in every direction.
On 1/3/2024 at 12:19 PM, Exoscientist said:But the most egregious of these is dismissing the importance of having full thrust, full up(all engines), full mission duration(actual minutes long flight length) static burns:
[snip]
The "T-Bird" S-IC-T was a Saturn V first stage designed specifically for static fire tests; it never flew and never could have flown. It completed its test firings prior to the construction of the actual flight articles.
AFAIK, neither the three S-IC stages launched in the Apollo 4-6 flight tests nor subsequent S-IC stages used for actual crewed missions ever received full-thrust, full-up, full-mission-duration static burns.
It continues to puzzle me why you insist that building a new facility for full-duration static test fires of Superheavy would somehow be categorically and qualitatively better than conducting full-duration test fires simultaneous with the test launches themselves.
On 1/3/2024 at 3:17 PM, Exoscientist said:SpaceX needs to be open about how many of these static fires are done at full power.
We have no evidence that SpaceX fails to share thrust level and telemetry data with its government partners.
12 hours ago, CatastrophicFailure said:Apollo used a significant percent of GDP at the time to put two dudes in a bedroom closet on the moon for a couple of days. I’d rather see all parties involved did not simply do that again, and actually expanded our scope and capabilities, for a fraction of the relative cost.
Let’s not merely repeat the past, let’s actually build the future. I’m ok with a heavily-regulated player with massive oversight blowing up a few pre-prototype concept demonstration rockets to accomplish that.
Agreed. As outlined above, building an actual single-launch architecture for meaningful moon landings would require a rocket almost double the liftoff mass of the Saturn V (and likely more).
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4 hours ago, Errol said:First major question I have is the thrust to weight ratio for each stage. [snip] I've been able to find references to the burn times for each stage fairly easily, but I can't seem to pin down the thrust to weight ratios. So to list what I am missing, I need the TWR for S-II, SIV-B, the CSM, LEM Descent stage (lunar TWR, to be clear) and also for the Launch Escape System (with the CM attached).
To answer these questions reliably, your best approach is to look at the actual thrust of each engine, the number of engines on each stage, and the weight of the stack at each staging event. Then just do the math. All of that is going to be much more accurate than guessing at whether publicly-posted numbers were posted by people who did the math correctly. The wet and dry masses of each stage is all public from NASA documents.
4 hours ago, Errol said:The other major area of concern I have that I can't quite figure out is the fuel systems routing and ullage requirements for the LEM.
I happen to know a good bit about this so you're in luck!
4 hours ago, Errol said:I've read that the same hypergolic fuel was used for the descent engine, ascent engine, AND the RCS thrusters, is this correct?
The same fuel type was used, yes, but it was not sourced from the same tanks.
The RCS thrusters on the ascent stage provided 100% of the reaction control for both the descent and ascent; there were no RCS thrusters on the descent stage. The tanks on the descent and ascent stages were completely separate without any interconnections. The descent propulsion engine was a throttleable, restartable, gimballed hypergolic engine fed exclusively from tanks housed within the descent stage. Because this engine could be gimballed, the RCS on the ascent stage was used only for roll control while the descent stage was firing and ullage while the descent stage was starting up. While I know that the RCS thruster controls were designed to permit translational burns for docking, I am not sure whether translational firing was active during hover and landing. I should also note that although the lunar module was capable of acting as the active translational actor during docking, it never did so in practice; it would just hold orientation and allow the CSM to come to it.
The ascent stage had three sets of propellant tanks: one main set of propellant tanks which fed the ascent propulsion system and a pair of redundant propellant tank sets for the RCS system. Each of the redundant RCS propellant tank systems contained half the propellant needed for descent RCS plus all of the propellant needed for ascent RCS, so that if one of the tanks experienced a problem they could still perform all necessary ascent burns. There was an additional (closed) valve linking the main propulsion propellant tanks to the RCS system so that the RCS system could be powered directly from the main prop tanks if both sets of RCS system tanks failed, although this would reduce the amount of propellant available for the ascent propulsion system.
The ascent propulsion system was constant-thrust and fixed, so the RCS system had to provide pitch and yaw as well as roll during ascent. Although the ascent propulsion system was technically restartable, it was not ordinarily restarted during ascents because the helium pressurization for restarts was a little tricky: it was a single burn from the lunar surface to lunar orbit. It was restarted for disposal burns, and required an RCS ullage burn for those restarts. There was no ullage burn off the lunar surface because lunar gravity provided sufficient propellant settling.
4 hours ago, Errol said:Last question is more broadly speaking about all the missions that went to the moon at all. It's one detail about the flight plan that I seem to have found evidence for two different possibilities. I'm wondering about the trans-lunar injection burn. Was the burn one long continuous burn from stage one on the ground, all the way to the moon, with the correct time of day used to ensure the correct phase angle for departure from earth OR did they cut the engines after circularizing in LEO, and then re-light the SIV-B for the Hohmann transfer?
@StrandedonEarth is correct: the SIV-B did a single burn from staging to parking orbit, circled Earth a few times, then restarted for the trans-lunar injection burn. The parking orbit was both to simplify phasing and to allow time for systems check-out before committing to the rest of the mission. Oberth, drag, and boiloff losses were accounted for but minimal.
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6 hours ago, Exoscientist said:On 12/7/2023 at 3:14 PM, sevenperforce said:
Look at the RL10A-4-2. It has an expansion ratio of 84:1, lower than the expansion ratio of the J-2X, but it achieves 451 seconds Isp because it is an intrinsically higher-performing engine. Gas generator engines simply do not have the same efficiency potential as expander cycle engines.
The comparison between the RL10A-4 and the J-2X is illustrative. The RL10A-4 has an expansion ratio 84 to 1 and gets an Isp of 451 s. The J-2X has an expansion ratio of 80 to 1 and gets an Isp of 448 s: http://www.astronautix.com/j/j-2x.html . So the J-2X Isp is over 99% of the RL10A-4 value.
I'm not sure where astronautix got the 80:1 figure, but the actual manufacturer said the expansion ratio was 92:1. So I'm inclined to believe them.
Which goes to my point. You can't squeeze expander cycle performance out of a gas generator engine. It's a different beast.
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7 hours ago, Spaceception said:
Thanks, I'm on my phone.
According to Lapsa, it's not the full length first stage tank.
Ah, yep, that makes sense. I was going to be pretty shocked if the fineness was THAT low. I could see it for a hydrolox rocket but not for methalox.
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3 hours ago, RCgothic said:
SN8 had a sudden loss of pressure in the methane header tank
IIRC that was basically the same failure mechanism as (presumptively) IFT-2 experienced: slosh in the tanks during a rapid flip that dumped vapor into the turbopumps.
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2 hours ago, Exoscientist said:3 hours ago, Kartoffelkuchen said:
Not again.
If your cars fuel pump breaks down and the engine stops working, would you say that the engine is unreliable?
I understand the point you’re making but perhaps it’s not the best example to give in regards to a rocket engine. For rocket engines the turbo pumps are part of the engine.
If you understand the point being made, take the point being made instead of taking some other point that isn't being made.
Suppose a truck manufacturer builds a semi-truck, and early tests of the semi by the manufacturer show that the rear wheel assemblies have a tendency to seize up, stop spinning, and frag the axles at highway speed. The truck manufacturer promptly redesigns the wheel assemblies, and they no longer exhibit this problem. Much later, in acceptance testing, the truck manufacturer learns that the transmission is not downshifting properly and is transmitting ten times too much torque to the axles, which causes the axles to fail.
It would be very silly to suggest that this transmission failure in acceptance testing is the same wheel assembly failure exhibited in early testing, simply because both involve the axles.
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1 hour ago, tater said:
That is a CHONKY boi.
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13 hours ago, cubinator said:
I watched the whole thing lamenting that there wasn't any audio included, and then realized my computer speakers were muted.
Time to enjoy again!
It's so cool to watch with the audio on, because the plasma stream starts to heat up visibly before you get any significant wind noise. It's really something.
All those burning bits from the ablatives are crazy to watch too.
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1 hour ago, tater said:
Yeah. The issue is that the difference between some of them is likely small for the 2 core configs, and the target orbit as well. F9 now does 18.6t to LEO with ASDS recovery. Expended, SpaceX claims 22.8t. Wiki says FH with all 3 recovered is ~30t. Presumably a 2 core variant would be somewhere in between for 18.6t and 30t for 2 core recovery operations. As usual, the payloads are sort of volume limited (the 18.6t Starlink v2minis are probably the densest payload possible since they stack tightly), so this is very likely then more driven by target orbit since the added payload will be propellant residuals in S2.
I went and glanced through the last couple years of launches to try and get some benchmarks for the largest launches to given destinations.
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Falcon Heavy
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All Cores Expended
- ViaSat-3, 6.722 tonnes to GEO (5/1/2023)
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Center Expended, Boosters RTLS
- Psyche, 2.608 tonnes to heliocentric orbit (8/4/2022)
- Echostar-24 (Jupiter-2), 9.2 tonnes to GTO (7/29/2023)
- USSF-67, 3.75 tonnes to GEO (1/15/2023)
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Center ASDS, Boosters RTLS
- ArabSat-6A, 6.465 tonnes to supersynch GTO at 90,000 km apogee and 23° (4/11/2019)
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All Cores Expended
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Falcon 9
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Expended
- Galaxy 31 & 32, 6.6 tonnes to supersynch GTO at 283x58,433 and 24.2° (11/12/2022)
- Eutelsat 10B, 5.5 tonnes to supersynch GTO at 261x59,831 and 22.8° (11/23/2022)
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ASDS
- SES-18 & 19, 7 tonnes to GTO (3/17/2023)
- Intelsat 40e TEMPO, 5.59 tonnes to GTO (4/7/2023)
- Inmarsat-6 F2, 5.47 tonnes to supersynch GTO at 387x41,592 and 27° (2/18/2023)
- Starlink, 18.4 tonnes to LEO (12/7/2023)
- USA-343 GPS III-06, 4.352 tonnes to MEO (1/18/2023)
- Galaxy 33 & 34, 7.35 tonnes to subsynch GTO with apogee 19,800 km (10/8/2022)
- Danuri, 0.679 tonnes to ballistic lunar transfer (8/4/2022)
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RTLS
- OneWeb #17, 6 tonnes to LEO (3/9/2023)
- SARah 1, 4 tonnes to SSO (6/18/2022)
- Hakuto-R, 1 tonne to ballistic lunar transfer (12/11/2022)
- EROS-C3, 0.4 tonnes to retrograde LEO (12/30/2022)
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Expended
I feel like the supersynch launches are particularly helpful because they represent a maximization of the capability of the stage with a given mission profile. May try to math around with this a little.
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Falcon Heavy
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18 hours ago, CatastrophicFailure said:On 12/9/2023 at 9:25 PM, SunlitZelkova said:
One blog called it “too big for LEO missions, too small for lunar missions”
So, basically, SLS?
OOOOOOH that burns like an pad without a deluge.
19 hours ago, tater said:Testing new, smaller tiles.
I dunno why but that gives me the ick.
On 12/9/2023 at 11:32 PM, darthgently said:Someone asked Jared Isaacman what was taking Polaris Dawn so long...
Kind of figured something along these lines. A real solid workable EVA suit -- one they actually want to build on rather than use as a tethered one-off -- is tough. But doable.
On 12/9/2023 at 10:36 PM, tater said:Properly done, a 2 core F9 only saves the cost of props and refurb for 1 booster in a use case where 3 cores (all reused) is overkill. We'd have to look at the payload numbers for F9 with booster reuse (RTLS, then ASDS), compare it to FH with 3 cores recovered, vs 1 expended—and the difference between whichever cores are ASDS vs RTLS). That number is not $0, but it's not a huge amount, either, and will be for some narrow payload mass regime I bet.
There are a lot of unknowns about internal operational cost of ASDS vs RTLS and so forth.
In theory, there are a total of 17 different flight options for the Falcon family:
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One core
- RTLS
- ASDS
- Expended
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Two core
- Both RTLS
- Booster RTLS and center ASDS
- Both ASDS
- Booster RTLS and center expended
- Booster ASDS and center expended
- Both expended
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Three core
- All RTLS
- Boosters RTLS, center ASDS
- Boosters RTLS, center expended
- One booster RTLS, one booster ASDS, center ASDS
- One booster RTLS, one booster ASDS, center expended
- Both boosters ASDS, center ASDS
- Both boosters ASDS, center expended
- All expended
I really have no idea what the correct ordering is in terms of payload capability.
1 minute ago, darthgently said:19 hours ago, tater said:ExpandTesting new, smaller tiles.
I suppose the idea is that if a tile comes loose, a smaller area will be exposed. Am I missing other possible advantages?
That's one advantage, but another one probably comes before that: smaller tiles have a lower torque arm moment and thus are less sensitive to flexion of the underlying substrate surface. This would make sense if this is a problem area where they keep losing tiles no matter what.
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One core
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52 minutes ago, CBase said:
I doubt the header tanks contain 10 metric tons of liquid hydrogen. According to multiple sources that is the amount to transfer. More likely something is installed inside payload area.
The header tanks contain 6.25 tonnes of liquid methane and 23.75 tonnes of liquid oxygen.
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7 hours ago, Exoscientist said:
the payload would still be quite high if we added a nozzle extension to the J-2X to bring the ISP to the range of 465.5s already reached by the RL-10 with nozzle extensions:
I don't know how else to tell you, man -- if you put garbage in, you're going to get garbage out. If we could get magic, weightless nozzle extensions for free, that don't take up any volume or mass, then rocket engineers everywhere would be adding them to every engine everywhere.
7 hours ago, Exoscientist said:Still over 130 tons to LEO.
And no, it's still not over 130 tonnes to LEO, because not only are your numbers all wrong, but you're also not applying the overestimation factor.
7 hours ago, Exoscientist said:for that 482.5 ISP, it is a matter of a nozzle extension. Those are already used on the highest performing RL-10 versions that bring their ISP’s to the 465.5 range.
It's a matter of adding a nozzle extension to an already intrinsically-high-performing engine cycle. An expander cycle engine is fundamentally different than a gas generator engine.
Look at the RL10A-4-2. It has an expansion ratio of 84:1, lower than the expansion ratio of the J-2X, but it achieves 451 seconds Isp because it is an intrinsically higher-performing engine. Gas generator engines simply do not have the same efficiency potential as expander cycle engines. The highest-efficiency GG engine flying today, the HM-7B, has an expansion ratio matching the RL10A-4-2 but it only pulls 446 seconds. The record-holding LE-5 gas generator engine had an expansion ratio of 140:1, more than 50% higher than the J-2X, and it was only able to reach 450 seconds Isp.
And this is before you factor in issues like changes in O:F ratios, chamber pressure, chamber temperature, and so forth. Rocket science is a complex tradeoff balancing many different factors to decrease weight and size while increasing thrust, efficiency, and reliability. You can't just slap an expander-cycle nozzle extension onto a gas generator engine and expect to get expander-cycle efficiency any more than you can slap the straight pipes from a Mustang onto a Prius and expect the Prius to go 0-60 in 4 seconds.
7 hours ago, Exoscientist said:You can’t get much higher than this on current rockets that use them such as the Delta IV Heavy and Atlas V because the expansion ratio is limited by the rocket diameter.
This just isn't true. The largest-diameter nozzle extension for an RL10 comes in at 2.15 meters, while the interstage of the Delta IV has a diameter of 5 meters. A 4.66-meter nozzle extension for the RL10 (the same size, relative to the Delta IV diameter, as your notional J-2Z) would bring its expansion ratio to 1,315:1. So why not do that? It should be obvious: the efficiency savings wouldn't outweigh the mass bloat.
As I discussed upthread, impulse scales approximately with the square root of (1 - RP(k-1)/k), where RP is the inverse of the expansion ratio and k is a specific heat ratio. Let's pause and take a closer look. If the expansion ratio is infinite, the RP term drops away, and the specific impulse scales to the square root of 1, which is just . . . 1. That's because every engine has an inherent maximum specific impulse set by the chemistry of the propellants and the thermodynamic constraints of the universe. Increasing the expansion ratio decreases the loss associated with underexpansion at the nozzle outlet, but it can never exceed that theoretical maximum.
However, that equation for impulse is only part of the picture.
Imagine a rocket engine with a nozzle of zero length. If the expansion ratio is simply 1, then 1 raised to the power of anything is still 1, and so impulse would scale with the square root of (1-1) which is zero. Does a rocket engine with a zero-length nozzle have no thrust? Of course not -- intuitively we know that it would still have thrust:
The thrust produced by this engine is simple: it's the chamber pressure times the throat area. Of course this is highly inefficient, because all of the heat from the exhaust is being lost to vacuum and all of the gas molecules are flying in every conceivable direction instead of lending their momentum properly. But it's still thrust. If you imagine chopping the nozzle off of an engine (ignoring that many engines would stop working because they use regenerative cooling in their cycle), this is the pressure thrust (FPressure) you would get from various engines:
Engine Chamber Pressure (kPa) Throat Area (m2) FPressure (kN) RL10B-2 4412 0.01297 57.2 RL10A-4-2 4200 0.01280 53.8 RL10C-1 4400 0.01270 55.9 J-2X 9515 0.07931 754.6 J-2 5260 0.10994 578.3 RS-25D 20640 0.05327 1099.4 RS-68A 10260 0.21571 2213.2 Note that because even similar engines have slightly different mass flows, FPressure is not going to line up exactly the same. In a pressure-thrust regime, the only way to increase thrust is to either increase chamber pressure or use a wider nozzle. Using a wider nozzle will of course require you to have higher mass flow to maintain chamber pressure. In order to better utilize some of that heat and pressure, we add a nozzle to our engine:
We still have some hot, underexpanded gases escaping around the edges, but now a bunch of our exhaust has cooled and is now pointing downrange, which helps us out tremendously. Now, our total thrust becomes the sum of two components: the original pressure thrust component and a new "momentum thrust" (or "dynamic pressure") component:
FTotal = FPressure + FMomentum
The difficulty is that the contributing proportions of these two components change nonlinearly with chamber pressure, expansion ratio, combustion temperature, mixture ratio, and more -- certainly beyond the scope of this thread. But this should illustrate why you can't just keep imagining a longer and longer nozzle (and apparently a weightless one at that?) and conclude that arbitrarily high specific impulses are possible.
12 hours ago, Exoscientist said:if we added a nozzle extension to the J-2X
The J-2X already has a nozzle extension. A J-2X with a different nozzle extension would be a different engine.
Even if we imagined that the J-2X was not a gas generator and was actually just a really big uprated RL10, getting up to 465.5 seconds of specific impulse would make it prohibitively heavy.
Since we keep running into the same issues with Silverbird, I went ahead and did a little tweaking. As I said in my last post, setting up the actual values for SLS with a TLI trajectory gives about 11.2% more than what Block 1B can actually send to TLI:
Spoiler
This is primarily from overestimation due to how Silverbird handles altitude-compensating engines like the RS-25.
To fix this, just downgrade the efficiency of the RS-25s to 95.95%, setting total core thrust to 8,747 kN and vacuum specific impulse to 434 seconds:
Spoiler
This allows Silverbird to better characterize the performance of the RS-25s and yields 42.2 tonnes to TLI, right about where you want it to be.
Note that setting this to a 185 x 185 km nominal parking orbit with direct ascent yields 111.1 tonnes to LEO, again right about where NASA says the max is for Block 1B.
Time to enjoy again!
Planetary orbit shifts?
in Science & Spaceflight
Posted
I saw the original post, started typing a reply, scrolled up a bit, and then saw you had said this, which was 95% of what I was going to say.
Milankovitch cycles are particularly interesting (for me) because they leave directly observable traces on Earth's surface...some of which are even visible with the naked eye.
The full Milankovitch cycle is the fusion of periodic changes in axial tilt, orbital eccentricity, perihelion, and precession rate. Of these, eccentricity has the largest forcing function on global temperature, and it operates on a roughly-100,000-year cycle created primarily by Jupiter. I believe that the amplitude of the changes are forced primarily by Saturn, though I'm not sure. The function looks like this (going back and extrapolated forward by just under a million years):
Whenever eccentricity reaches a maximum, the increase in solar insolation at perihelion reaches a maximum, and so global temperature starts to rise. This increase in temperature releases carbon dioxide stored by the oceans, which further drives up global temperature in a rapid spike. As glaciers melt and ocean levels rise, the dissolved gas carrying capacity of the ocean increases, allowing the oceans to slowly scrub the atmosphere of excess CO2, allowing global temperature to gradually trickle back down until the next peak in eccentricity. These temperature cycles are recorded on Earth in a number of ways, both organic and inorganic. Foraminifera, or forams, are single-celled organisms with calcium carbonate shells, and since calcium carbonate (CaCO3) contains oxygen atoms, foram shells trapped in benthic seafloor sediment create a record of the isotopic concentration of oxygen in the atmosphere at the time. Because 18O evaporates more readily at higher temperatures, higher ocean temperatures lead to a higher 18O/16O ratio in the atmosphere. Thus, benthic forams record global temperature. Similarly, air bubbles trapped in ice cores also preserve samples of the global atmosphere. Sure enough, both benthic forams and ancient ice cores reflect this sudden temperature spike and gradual decline on the exact same period as Earth's variation in eccentricity:
And it's not just these two records. Because global temperature impacts sedimentation rates, and sedimentation rates impact the density of sedimentary rock, we can literally see Milankovitch cycles recorded on the sides of cliffs in certain areas, where the rapidly-deposited sediment has weathered away faster than the slowly-deposited sediment:
I'm a fan of this because it's one of the bodies of evidence that helped me to leave the science-denial cult I grew up in.