Silavite
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Posts posted by Silavite
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3 hours ago, Hannu2 said:
301 is so called austenitic steel. It means it has face centered cubic lattice, in which iron is not ferromagnetic. It is ferromagnetic only if it is in body centered cubic lattice (ferritic iron). However, there may be small ferritic areas due to for example lattice defects and sometimes austenitic steels may have mild ferromagnetic properties. It is far too weak effect to hold anything but you can barely feel it with strong magnet. I also think that such defects are not wanted in rocket hull because they affect structural and chemical properties too.
This is getting a little beyond the scope of the topic, but since we're on a materials science tangent, I may as well add one more detail. The statement that 301 (or, more generally, 300 series austenitic stainless steels) are nonmagnetic is only true for the material that is in the as-cast or annealed form. Cold working can cause austenite to transition to martensite, which is magnetic.
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Is it just me, or does the vertical stabilizer on Boom's redesign look really undersized?
Spoiler
I'm wondering how lateral stability (particularly CN,β) is achieved...
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AIAA has a huge job board: https://careercenter.aiaa.org/
AIAA also has networking events and local chapters.
Aerospace (though more on the "space" side) job board: https://rocketcrew.space/
For a fun twist, here's the scoreboard for the LA Aerospace Games this year. (With a list of companies/organizations attending.) https://www.facebook.com/photo?fbid=10222010097137180&set=gm.10160034282721100
I'm in the same boat (T-1 year to graduation) so I can relate to your search.
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7 hours ago, Pthigrivi said:
So, what do we think about all this?
Using a company wide email to distribute something like this without approval definitely sounds like a bad idea. I don't know if that alone is justification to fire these employees, but they definitely screwed up.
I think that the main tension here is the fact that there is not a bright line between Elon's personal statements and SpaceX's official statements. Elon's Tweets are often official statements on SpaceX's doings. Sometimes this is really cool (such as when we get to see glimpses of technical details or decisions through his Tweets), but other times this doesn't work out so well. If any other employee said some of the things Elon regularly posts on his Twitter through an official channel, then they would be fired on the spot.
In my opinion, reserving the SpaceX Twitter for official statements while letting Elon do whatever he wants (but making it explicitly clear that none of the Tweets from his account are official company statements) would be the best solution. Let Elon be Elon (even if plenty of folks may disagree with him), but let SpaceX be SpaceX. That's my $0.02, at any rate.
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22 minutes ago, tater said:
Make it slightly bigger, a lot cheaper—and send a few, just in case.
That said, the, "send a few," part of this has been done before. Viking 1/2, Pioneer 10/11, Voyager 1/2, Spirit/Opportunity were all probes which were launched in pairs. Going beyond 2 would increase margins, however.
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1 hour ago, tater said:
For WDR June best case is June1. Looks like estimates of launch after earlier WDR dates (that were not met) show a min 1 month period between (4-7 weeks). If they can hit the beginning of that, then they might have a chance at that 1 week window in July, else, off to August. Note that summer in FL means a substantial chance of weather violations, then once they hit August—hurricane season. Of the manned Apollo missions, 2 flew in July, the rest between October and May.
Most of the issues during summer in Florida are due to lightning criteria violations. NASA-STD-4010 has more details on the specific criteria for those who are curious. Anyway, the lightning during that period is mostly caused by sea breeze thunderstorms, which means that lightning almost always occurs during the afternoon and evening hours, so it's something that can be accommodated for. (Though it could make delays on day-of-launch more dicey... I assume a daylight launch is desired, so the earliest window would be about 7 AM. Delay for 4 hours or more and the chances of nasty weather rolling in shoot up.)
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If you had to pick one equation as fundamental for a given field, which would you pick and why? For example, I think that it would be reasonable to say that F = ma is the fundamental equation for dynamics and ΣF = 0 is the fundamental equation for statics.
The field can be as broad or as specific as is desired. Families of equations (such as the Navier-Stokes equations for fluid mechanics) are acceptable but discouraged.
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59 minutes ago, mikegarrison said:
Combustion is generally understood to be a "black art", although modern computational tools are changing that somewhat.
Anyway, if it were easy to do, I can safely say my career path would have been very different.
You need to get the flame to start burning, get it to keep burning, manage the heat transfer to the combustor walls, avoid acoustic instability ("screech"), minimize the amount of cooling you can get away with (because that is always parasitic on the ultimate power output), etc.
Typically the problem is going to be that there are "hot spots". This is because the fuel and oxygen are injected in discrete locations, not magically teleported into the combustion chamber in a perfect premix.
Also, while the main engine does not have this worry, the upstream turbopumps have to not only keep the combustor from melting down but also have to worry about keeping the turbine alive.
54 minutes ago, JoeSchmuckatelli said:I was going to suggest metal foam, but then Mike said that you need to minimize cooling - so I guess it really is a dark art.
Ogun, show us the way!
To clarify this point about minimizing cooling, it is specifically film/dump cooling that should be minimized. Film cooling acts to prevent heat transfer to the chamber walls by injecting a layer of (usually) fuel around the chamber periphery. This creates a region of very rich, off-nominal mixture ratio fluid near the wall that does not burn and serves to protect the chamber wall from the extremely hot "core" combustion. This has a negative effect on performance because you're intentionally creating an area of maldistributed (from a combustion efficiency perspective) fuel which does not completely burn. This paper partially discusses the deleterious effects of film cooling and how designers might try to use less of it.
I'm unsure about exactly what effects ablative cooling have on performance. The only things I can say with much certainty are that the engine's thrust will increase (due to the increased mass flow rate from the increase in throat diameter as the throat material ablates away), and that the engine's vacuum exhaust velocity will decrease (as a consequence of decreased expansion ratio, which is in turn a consequence of increasing throat diameter).
Regenerative cooling of the combustion chamber has basically no effect on performance. The energy lost through the chamber liner into the coolant ends up going back into the chamber as somewhat-warmer-than-ambient propellant (indeed, expander cycle engines take this concept to the extreme), so very little heat actually escapes the engine system in regenerative cooling. Regenerative cooling of the supersonic flow in the nozzle extension can actually have a positive effect on performance (due to the decrease in entropy), though the effect is quite small (on the order of 10 m/s exhaust velocity) for all but the highest of expansion ratios. See here for more reading about the effects of regenerative cooling on performance.
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9 hours ago, mikegarrison said:
The way it is done is to use PD design tools to basically create the design yourself. See how close you came to the data you know, and redo it in a loop until it converges.
PD design tools are not really "back of the envelope", but in a sense they are a more sophisticated, calibrated version of "back of the envelope". They still are not a detailed design, which is the step you do once a PD design gets the go-ahead for further development.
Sometimes you come up with the answer that the claimed performance is unlikely to be achieved, and that's also an educational outcome.
It's different with a design that is actually in service, because in that case if you come up with the answer that the performance can't be achieved and yet it is clearly being achieved in service, then you know your tools need fixing. Real world data trumps all.
Apologies for the slight derailment, but what is meant by the term, "PD design tools?" Google is not helpful here and I assume you're not talking about proportional-derivative controllers or physical design in the electronic context. Are they things like Roskam Class 1 / Class 2 methods?
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52 minutes ago, caecilliusinhorto said:
I thought the reason they made the launch mount so high was so that they wouldn't have to use water deluge systems?
The reason that the launch mount is high up is so a flame trench/diverter isn't necessary. A flame trench exists to prevent the exhaust from physically damaging the launch site (and by extension protect the rocket from debris kicked up by the exhaust), whereas a water deluge system exists to damp the sound of the rocket launch.
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2 hours ago, Beccab said:
Yeah, if B4 and S20 are the ones to fly then 100% will be ready by the time they get the launch license. If it's B8 and S20 then probably it will be at the beginning of its testing campaign
Edit: forgot to add, rumour is that it's the FWS mainly holding back the final EA. Coincidentally though having this delay instead of directly ruling for an EIS makes a mitigated FONSI more likely to happen. Whatever the current problem may be it doesn't seem to be impossible to overcome, or there would be no reason to wait two more months
What is FWS in this context?
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Sort of related, I found a paper written by this company's current chief designer (Igor N. Nikischenko, who formerly worked as the Deputy Chief Designer in the Liquid Propulsion Department at Yuzhnoye in Ukraine) a few years back. It talks about the rationale for using LOX as a regenerative coolant and also discusses some novel combustion cycles. It looks as though the RD-58MF also uses LOX as a regenerative coolant if what this paper says is true.
QuoteDue to its properties, LOX is more suitable for regenerative cooling than kerosene. Additionally, LOX flow rates are normally several times higher than kerosene flow rates (according to MR showed in Table 1 it is at least in 2.4 times higher). It is tacitly believed that oxidizers at all and LOX specifically are not suitable for thrust chamber cooling. This belief is based on the principal concern that metal alloys are vulnerable to heavy oxidation and ignition in oxidizing environments. However, LOX regenerative cooling of thrust chamber was successfully confirmed in the 1950-60s. [...] The high effectiveness of LOX cooling was experimentally proved, as well as its feasibility. In the course of RKK Energia investigations the small leakages in the thrust chamber inner wall were simulated during the firing tests and it was proved that such defects do not cause thrust chamber destruction. [...] Intensive investigations of LOX cooling were also conducted by NASA. In particular, successful firing tests were conducted using experimental thrust chambers. Similar to RKK Energia, NASA experimentally confirmed that small leakages in thrust chamber inner wall did not cause destruction of the thrust chamber.
It is intended that LOX regenerative cooling of 11D58MF thrust chamber will exclude the internal [film] cooling, thus avoiding the associated specific impulse losses.
In hindsight, the big advantages to using LOX over RP-1 for cooling are clear: LOX mass flow rate is much higher than RP-1 mass flow rate and you also get the benefit of latent heat release from LOX (whereas RP-1 must be kept cool enough to prevent formation of waxes in the cooling channels).
The whole paper is fascinating and really worth a read. (Maybe semi-expander semi-gas-generator cycles will be the trendy new thing for upper stage engines, eh?)
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Also... one other thought. There was a lot of talk about simplifying operations, but Neutron is planning to balloon tanks in its upper stage (which are not exactly easy to handle). Admittedly having the launch site right next to the production facility should simplify operations in regards to using such a structure (no transportation), but I'm still wondering if they have any other special procedures for working with balloon tanks.
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Continuing with the engine stuff, the Scott Manley and NSF interviews with Beck reveal that the chamber pressure is 1,500 psi. Plugging in the known data:
- Vacuum thrust: 1075 kN *
- Sea level thrust: 851.4 kN *
- Chamber pressure: 10.34 MPa (1,500 psi)
Along with some educated guesses:
- Contraction ratio: 1.67 (same as the H-1, which is in a similar thrust class)
- Combustion efficiency: 0.99 (Beck talked a bit about the injector in Tim Dodd's interview, and I'm guessing that they're using a coaxial swirl configuration from his somewhat oblique comments. That type supports excellent mixing/atomization characteristics and also has some throttling ability which would be needed for the upper stage.)
- Mixture ratio: 3.2 (guess which is near the ISP maxima for this chamber pressure)
- Relative gas generator flow rate: 0.02 of the main thrust chamber mass flow rate (guess based on averages in Sutton and SP-125)
And a not-so-educated guess:
- Freezing area ratio: 2.5 (flow is assumed to be in chemical equilibrium until it reaches the point in the nozzle at which the area is 2.5 times the throat area; if somebody has a better guess for this, I would appreciate it!)
We can arrive at the following ISP figures for sea level and vacuum:
- Sea level: 272.0 sec
- Vacuum: 344.4 sec
* These ISP numbers seem to be unusually far apart for a first stage engine, but you have to consider that the ratio of vacuum thrust to sea level thrust (and thus vacuum ISP to sea level ISP, since mass flow rate is constant/choked at the throat) is really high for a first stage engine (1.26, whereas the Merlin 1D is 1.08, the RD-180 is also 1.08, and the H-1 is 1.13). This implies a high expansion ratio, but I was unprepared for how high; 37 according to RPA.
According to Sutton, an expansion ratio of 37 (for γ ~ 1.2 and pressure ratio of 100 at sea level) is within the incipient flow separation region, so I'm curious as to what kind of tricks Rocket Lab may be employing in their nozzle design:
Spoiler
Of course, trying to predict thrust/ISP levels in an incipient flow-separation regime may not be the most reliable for a (relatively) simple program like RPA, so I'd take the exact results with a grain of salt, but the general idea still stands.
Edit: I forgot to include ISP loss due to the gas generator in my original post (d'oh!)
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6 hours ago, DDE said:
Now George Sutton is the world's foremost engine-eer.
Another interesting story from the development of that same engine:
QuoteA humorous (although it was not thought so at the time) anecdote involving a materials selection for the Redstone engine occurred during its first test. In front of Army generals and corporate executives, the first engine test resulted in an explosion which destroyed the engine. It seems the designer had selected alloy steel for the LOX dome which, at an operating temperature of -290F (-179C), was well below the ductile-to-brittle transition of the steel. Startup shock fractured the dome and resulted in the explosion. From then on, LOX domes were made from aluminum alloys.
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1 hour ago, sevenperforce said:
Curiouser and curiouser. The Rutherford has a big whopping advantage over the RD-180, of course, because 100% of its propellant goes directly to thrust, while the ORSC cycle of the RD-180 uses up a significant amount of chemical energy in the preburner. That is why it has a fairly low TWR: the electric motor and batteries are hella heavy. But the RD-180 has a chamber pressure more than double that of the Rutherford, too.
Generally, what is the relationship between chamber pressure and specific impulse for a given propellant? Is it quadratic? Exponential?
That said, staged combustion engines have gas-liquid (or in the case of FFSC, gas-gas) injectors, which affords higher combustion efficiency in the main combustion chamber. It's not big enough to make up the difference, but the factor is there.
The relevant (approximate) relationship is this one from Sutton:
Wherein,
- k - Heat capacity ratio
- R - Specific gas constant
- R' - Universal gas constant
- M - Molecular weight
- T1 - Chamber static temperature
- T0 - Chamber stagnation temperature (assumed to be approximately equal to T1)
- P1 - Chamber pressure
- P2 - Exit pressure (equal to the environment if optimally expanded)
The biggest assumption here is that of a constant heat capacity ratio, but there are others (such as no friction or heat transfer to the walls, no Rayleigh losses (all of these go together in the assumption of isentropic flow), single phase homogeneous ideal gas, velocity at the exit is purely uniaxial).
Plotting performance curves for RP-1/LOX (O/F = 2.7) and CH4/LOX (O/F = 3.2) via RPA gives the following, respectively:
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5 minutes ago, sevenperforce said:
I know he just means that CF is 4X lighter than SS in terms of weight per unit area, not that the mass ratio will be 4X better than Centaur, but still…wow.
I’m sure the upper stage engine will have to have decent throttling. You need that for precision orbital insertion, anyway. Unthrottleable or limited-throttle liquid engines are mostly the regime of disposable first-stage engines. And throttling isn’t too hard to do with a GG engine.
The additional density of methane as opposed to LH2 should also help with the mass fraction.
On another note, after looking at Archimedes claimed sea level(?) Isp, I decided to look at Rutherford's sea level Isp. The Wikipedia page claims 311 seconds at sea level, but that number is equivalent to the sea level Isp of the RD-180! As good as Rutherford is, I have a very hard time believing that it matches the RD-180 in that regard. Electron's user guide says that the specific impulse for the sea level Rutherford engine is 311 seconds, but it does not explicitly say that 311 seconds is the value for the sea level Rutherfords at sea level. It seems likely that the Isp numbers for the Rutherford and Archimedes are both some kind of average between sea level and vacuum.
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If this upper stage really does have such an incredible mass ratio, then I'd imagine you're looking at some pretty serious accelerations on lighter, higher energy payloads near the end of the 2nd stage burn. Has there been any information released about Archimedes (assuming that they're using a vacuum version of Archimedes for the second stage) having throttling capabilities?
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2 hours ago, sevenperforce said:
The Merlin 1D clocks in at just under 10 MPa; Raptor is 30+ MPa. What is typical thermodynamic efficiency in gas generators?
I honestly couldn't tell you for the gas generators themselves. Gas generators are basically small rocket engines, so I'd assume that the combustion efficiency isn't that different? That said, the highly off-stoichiometric mixture ratios may muck that assumption up somewhat...
For supersonic impulse turbines, nozzle efficiency into the first rotor is on the order of 80% - 96% according to Huzel & Huang.
Turbine efficiency seems to be on the order of 60% - 80% as indicated by Sutton as well as Huzel & Huang. (This can be increased if a gearbox is used, but that adds mass.)
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Also, I'm guessing that the Isp of 320 seconds is an average between sea level and vacuum. If my fiddlings with RPA are correct, you'd need around a ~20 MPa chamber pressure to make that happen at sea level (accounting for a combustion efficiency of 99%, a nozzle efficiency of 98%, and sending 2% of the propellants to the gas generator) which seems a bit silly for an open cycle.
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30 minutes ago, JoeSchmuckatelli said:
That's happening in Coal country. Thing is, I keep hoping that some bright chemist will discover a way to make coal less polluting. Would love to be able to squeeze the juice out of it and see it become a viable resource... but I suspect our continued failure to do so means it's not feasible (or at least not economical).
Thing is, I don't have a lot of sympathy for dirty ol' coal. Like an ex girlfriend; you appreciate the history with them, and how they helped you develop... but at some point, it's time to move on.
Speaking of coal; ran across this and found it interesting.
Mapped: The world’s coal power plants in 2020 (carbonbrief.org)
In addition to the local air pollution and CO2 emissions from coal, there's also the fact that it just isn't economically viable when compared to the rising tide of renewables (and natural gas, due to the lack of maturity in power storage solutions). It's true virtually everywhere that, for building a new power station, a renewable source will be cheaper than coal. It's true in a sizable minority of places that building new renewables is cheaper than running existing coal power stations.
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4 hours ago, StrandedonEarth said:
Yes, but that had oxygen and fuel (wiring/insulation) to burn, and it was probably reasonably full happening on the outbound leg. But I didn't know about inerting the tanks on commercial aircraft; that's sensible.
So yeah, rocket fuel tanks shouldn't have to worry about vapors igniting. Oxygen tanks, on the other hand, well, best not to have anything flammable in there. although pure oxygen will burn darn near anything if it gets hot enough...
On the gripping hand, that damaged tank should have never been cleared for flight...
Its worth noting that LOX can react explosively with titanium, but I don't think that would be an issue here. (The reaction also requires substantial impact to occur, so it's even less relevant. For example, the Titan I used titanium pressure vessels located inside the LOX tank to hold helium pressurant.)
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3 hours ago, mikegarrison said:
On the other hand, the skin is also the fuel tank. So any heating (or worse, burn-through) is directly impacting the fuel and oxidizer.
The CH4 header tank is ensconced in the center the larger LOX/CH4 bulkhead, so it is relatively well-insulated. The LOX header tank, on the other hand, is right at the nose (where the highest head loads will likely be). I'd imagine that losing a tile covering the LOX header tank could be very dangerous indeed.
Desmos/Equation question
in Science & Spaceflight
Posted
Desmos can do piecewise functions with the following format: f(x) = {condition : output, condition : output, condition : output, etc...}
From the description, I believe that this is what you're looking for: https://www.desmos.com/calculator/n0rx0hi3nj