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Last thing I have seen in the news is that there is no firm date for Butch and Suni's return, but not before sometime after the first week in August.
I still have to worry about the thrusters on the entry capsule itself, having seen or heard NOTHING about the them. All the focus has been on the misbehaving service module, which is discarded for destruction before actual entry. I am presuming they've also been tested and checked out, bit I have seen nobody actually say so.
GW
A voice out of the past (me) says once upon a time "go to" statements were essential, at least in the things I used so long ago: Fortran 4 and QuickBASIC. They were most reliably used as part of the branches of an "if-then" test, for accessing different choices of subroutines or different stretches of code instructions.
The "spaghetti code" problems happened when when you used a "go to" to jump outside of a logical repeating loop (however it was set up). That sort of thing was a "no no" even half a century ago.
I do not know, but I get the impression that the more modern languages are organized to the point of not needing "go to"-type statements to set up logical repeating loops and if-then-type choices. Perhaps over-organized. That sort-of idiot-proofs the language, yes, but it also reduces the complexity of the problems you can actually attack with it. It's a tradeoff: there is no such thing as a free lunch.
GW
Tom:
The Boeing problem is a far older problem. "Capitalism" is focused on maximizing revenue, to the exclusion of all else (!!!), unless well-regulated by government (or whoever) for ethics. I've seen it in business school myself firsthand over half a century ago, and in business school grads for decades ever since. Let's just say I have zero respect for the morals and ethics of such business school grads, who almost exclusively comprise corporate management in America (and abroad). Without severe legal consequences, they invariably refuse to comply with society's rules and needs.
Boeing was an oddity, whose corporate management was former product engineers, not business school grads, and who were co-located in Seattle with the current engineers, from WW1 until the late 90's, when they "merged" with McDonnell-Douglas. Right after that acquisition, Boeing's new corporate management moved first to Chicago, to get away from the engineers, and a few years later to DC, the better to lobby Congress. Most of Boeing's actual products are made in Seattle, and Wichita, KS. The troubles with Boeing products routinely threatening the public's lives began then, right after corporate management moved, and publicly announced they were all about shareholder value, NOT the top quality engineering in the world! That timing and that announcement says it all!
I'm entirely unsure what else to point at, as the problem. It seems quite self-evident to me.
And I think the plea bargain over the 737MAX debacle is quite wrong! It les all the guilty parties off the hook! Those corporate managers should instead stand criminal trials. Nothing else will change the way Boeing now does its business.
GW
My point is that there is NO EXCUSE for these problems with NTO-any hydrazine thrusters to be cropping up at all! Not with something that has been flying for over 60 years!
The article all but blames bad material choices for seals. That kind of thruster has been flying since Mercury in 1961. They were on Gemini, they were on Apollo. They were on the warhead busses of all the ICBM's. They have been on everything since. And Boeing has gobbled up almost all the former contractors who made all those things.
There being no excuse, I blame greedy corporate management mandating cheap shit that doesn't work right just so they can have more revenue off everything. If you value money over safety and quality, without regard to laws, regulations, or ethics, that happens EVERY SINGLE TIME! Throughout history!
Besides, Boeing was paid almost twice what SpaceX was paid, and has taken almost twice as long. They were paid NOT to have these problems, and were given the time NOT to have these problems!
It appears to me that Boeing corporate management just took the money, ordered the cheap shit be used, and walked away from owning any of the risks banking on not getting caught. That is NOT the Boeing that built the B-17 or the B-29. Or the B-47. Or the B-52. That is NOT the Boeing that built the B-707, or the -727, or the -747. It is NOT the Boeing that built the early-model -737's.
But is IS the Boeing that built the B-737-MAX and killed two planeloads of people with it, before they got caught being cheap and greedy, and more than a little illegal (!!!), under the FAR's!
In view of all that, why is dangerous trouble with Starliner in the least unexpected?
GW
This illustration in post 150 just above shows the real point to be made here: the launch trajectory has an apogee altitude, where it has gone horizontal before starting downward. At this point, there is some horizontal speed. You make this same point in space your perigee point for a transfer ellipse, with its apogee on the other side of the Earth at your desired circular orbit altitude.
Use the orbits+ course orbit mechanics spreadsheet "orbit basics.xlsx, using the two-endpoints worksheet, set up for Earth. Plug in your perigee altitude and your apogee altitude, and use the average Earth radius value since the inclination is unlikely to be zero. It will give you the transfer ellipse perigee and apogee speeds, and it will also report circular orbit speed at perigee (which you do not need) and at apogee (which you do need).
The difference between transfer ellipse perigee speed and your launch trajectory apogee speed is the ideal dV you need, to get onto the transfer ellipse from your launch trajectory. This might need a little bit of factor-up to cover some small gravity and/or drag losses, more if you do not burn quickly enough to be "impulsive".
You will return to perigee if you do not circularize at apogee. If you do, it is below the entry interface altitude effective for orbital class speeds. That means you will re-enter, unless you raise your transfer ellipse perigee altitude above the entry interface altitude. You do that with a small burn at your transfer orbit apogee.
You avoid that fate and get where you really want to go by burning at apogee to raise your perigee all the way up to that apogee altitude. That IS circularization, exactly per the 1st 3 lessons in the orbits+ course set! The dV there is the difference between circular orbit speed and transfer ellipse apogee speed. It needs no factor-up as long as you do it quickly enough to be "impulsive".
GW
The 8013 perigee speed for a 60x300 transfer ellipse looks to be about right. My spreadsheet is still set up for a 59.5 x 300 transfer ellipse, and shows 7949 m/s at perigee and 7662 m/s at apogee about 44 minutes later. It also shows the circular orbit speed at 300 km to be 7734 m/s. That's an error of only 64 m/s perigee speed, which is likely round-off error somewhere.
For a sling launch apogee speed of 1000 m/s, the difference between transfer ellipse perigee speed and launch trajectory apogee speed is 7949-1000 = 6949 m/s, something that must be gained fairly quickly in order to be considered "impulsive". The circularization dV is the difference between circular speed and apogee speed, both at 300 km, which is 7734-7662 = 72 m/s. It's not much, but it needs to be done quickly in order to be "impulsive".
The transfer perigee point is down in the atmosphere well below the entry interface altitude of 140 km, effective for orbital class speeds. That has two implications: (1) there may be a little bit of gravity and/or drag loss associated with the dV to get onto the transfer ellipse. That 6949 m/s dV for it should be factored up a little bit. (2) If you fail to circularize, you will hit the atmosphere and start entry as you return to the transfer perigee.
For the rocket calculations, I rounded velocities to 3 significant figures, and I applied a 2% factor-up for losses at the perigee end. There are no losses at the apogee end for circularization, unless the burn is not "impulsive".
As for the mass figures, the inert mass fraction as I have been using it applies to the ignition mass, not the payload mass. I think the 222 kg is incorrect. The weight statements organize this way for a single stage:
item..........mass..................mass/Wig
payload......Wpay.................Wpay/Wig
inert..........Winert................Winert/Wig
burnout......Wbo=pay+inert..."allowance"= Wpay/Wig+Winert/Wig = 1 - Wp/Wig
propellant Wp....................Wp/Wig = 1-1/MR
ignition.......Wig=prop+bo.....1 = "allowance"+Wp/Wig
Note that mass ratio MR = Wig/Wbo as usual.
GW
It's unclear to me, but it sounds like he's using hot hydrogen at pressure to push the projectile, not combusting. He does evacuate the barrel, because you only want to push the projectile out of the barrel, not it plus a mass of air, too.
I think his estimate of half a km/s velocity loss is too low. My slinger launch drag loss numbers run much higher than that. More than 2/3 of the muzzle velocity gets lost in the first 10 seconds. Then you're in the thin air and it loses very little from there. It's why the velocity-time plots are shaped the way they are.
His estimates of SpaceX first stage entry speeds at Mach 9 or 10 are quite wrong. The staging velocity is actually quite low: ~2 km/s, which is only about Mach 6 or 7, and it takes place essentially exoatmospheric. They do an entry burn to slow speeds down to around Mach 2.5 to 3.5 for hitting the atmosphere, heat-sinking their way through the brief heating pulse with their aluminum Falcon cores. Steady state, aluminum is only good for Mach 2.5 speeds. The boost-back burn is a slow arcing trajectory with an entry speed Mach 3 to 4. That's too high for aluminum.
You'll note that Starship/Superheavy has a similar staging speed and altitude, but does no entry burn, because stainless steel can take the heat steady-state at Mach 3.5. The boost-back burn puts it on a slower arcing trajectory, specifically to limit entry speed to something the bare steel can take.
He's looking at a gun trajectory that apogees-out at orbital altitude, something not seen with the slinger. Note also how he talks about the extreme heating and the ablative required to survive it.
Yes, my 2-D Cartesian trajectory spreadsheet could model such projectiles, given a muzzle velocity and a launch angle. It is the same as what I was doing for the slinger, just more extreme.
GW
With the smaller vehicles, I was showing 1 to 1.08 km/s horizontal velocity at apogee, altitude 37 statue miles (59.5 km). That was for 45 degree launch at 5000 to 5500 mph. Most of the sling launcher exit kinetic energy is lost to extreme hypersonic drag (low altitude!) in the first 10 seconds. Larger size vehicles gain mass faster than they do frontal area, at the same apparent overall density and the same L/D and hypersonic drag coefficient. The frontal area mass loading (analog of ballistic coefficient) increases in proportion to diameter. The larger it is, the less initial KE you lose in the first 10 seconds, but the effect is not at all dramatic!
So I looked at larger-mass vehicles of larger diameter, and I reduced the launch angle to just maintain the apogee altitude, hoping for an increase in apogee horizontal velocity. I got it, but it is not dramatic. Even so, it lowers the dV required to reach orbit (theoretically 7.95 km/s minus the slinger horizontal apogee speed), to which you add a tad of loss, and about 0.08 km/s to circularize at 300 km circular. That raises the overall payload/ignition mass fractions. A little bit!
The problem is these larger slung vehicles are getting to be quite large, implying a ridiculously-large slinger, and the deliverable payloads are still quite low compared to Tom's desired 200 kg.
I cannot work on this continuously, there are too many other things around here demanding my attention. But I have made some progress, and I sent it to Tom's personal email.
I don't know enough about hybrid rockets to investigate them properly. But I have been looking at storable liquids, and solids, for the propulsion from slinger apogee to low circular Earth orbit. That's in the stuff I have sent Tom so far. And none of that is a converged sizing.
GW
Just goes to prove that other properties than simple tensile strength get into real-world engineering design.
GW
Does the ChatGPT Python program vary CD with Mach number, or is that a constant? It should vary with Mach, because the subsonic and hypersonic values are not the same, and the transonic values are about twice either the subsonic or hypersonic values.
I could not quite tell what the program was really trying to do. The slinger trajectory is one problem, while making the burns to get onto a transfer ellipse is quite a different problem.
GW
The 3rd and 4th studies I did presumed a 3000 lb object 22.45 inches diameter. The 3rd study showed apogee at 56.5 km altitude and 1.0 km/s speed (all horizontal at apogee, by definition!!!). It launched from the slinger at 45 degrees and 5000 mph. The 4th study threw exactly the same object from the slinger at 45 degrees and 5500 mph (10% faster). It achieved apogee at 65.5 km altitude and 1.08 km/s speed.
The two altitudes bracket your 37 stature miles = 59.5 km, and the two speeds are almost the same at 1.0 to 1.08 km/s. Altitudes in that range are high enough to neglect air drag for the ascent to orbit, and the radial change is low enough to reduce gravity loss for the ascent from that apogee to orbit. These data were in the plots that I sent you in the 3rd and 4th documents.
The notion is to eject the payload and rockets from the carrier vehicle and fire the main rocket stages, all right at apogee conditions. The transfer ellipse is the apogee altitude as the ellipse min altitude 59.5 km (give or take a smidge), and ellipse max altitude 300 km (give or take a bit). We have to get from from 1.0-1.08 km/s to the transfer perigee speed with the main rocket burns, then wait for it to reach transfer apogee to circularize with a final small burn.
I am assuming a tug goes and retrieves it, relieving it of needing to burn for rendezvous and docking. I also do not assume a deorbit burn for disposal, which presumes "something ese" deals with that issue.
This scenario is in fact what I have been using for the initial bounding calculations attempting to size NTO-MMH and solid rockets to do this job. The ideal dV without losses is 6.95 km/s to get onto the transfer orbit, and 0.08 km/s to circularize at 300 km altitude. The loss I figured as 2% of surface circular orbit speed. I tried limiting them to 3000 kg = 6600 lbm, knowing the larger size might give me a bit more altitude, or alternatively let me lower the slinger launch speed and gees. Maybe I should lower that to 3000 lbm, or maybe I should look at 6600 lbm slung vehicles. I do not yet know.
The liquid did fairly well as a two-stage item making 3 burns (the second stage relights to circularize. The solid I first tried as a 3-stager, with two main stages for the burns onto the transfer orbit, plus a very small solid motor to circularize. It did not do very well, being lower Isp. I got better results with a 4-stager, that being 3 solid stages to get onto the transfer ellipse, plus a very small motor to circularize. The results I got at 3000 kg just scale up or down with payload size. The velocity requirements do not scale.
I still don't believe any of this, not having done any design layouts and checking the requirements to survive the sling, as well as do the propulsion job.
GW
I see SpaceX got a contract from NASA to build a Dragon with a way-oversize service module, modified for viability after many months in space, to be the deorbit propulsion device for the ISS. This will get used along about 2031 or 2032 when ISS is decommissioned and disposed of. There is some question remaining to be solved about how to launch such a hugely-overweight Dragon to ISS. My guess is Falcon-Heavy. But I know nothing.
GW
Sounds like he knows something about compressible aerodynamics, then. He'd probably like what I did for the rocket engine estimator spreadsheet. There's nozzle compressible flow underlying it.
GW
From AIAA's "Daily Launch" for Fri 7-19-2024:
AP NEWS
Boeing is closer to understanding thruster failures on its first astronaut flight with latest test
Boeing is closer to understanding what went wrong with its astronaut capsule in orbit, now that testing is complete on a spare thruster here on Earth. Officials said Thursday there’s still no return date for astronauts Butch Wilmore and Suni Williams. Engineers will first disassemble the thruster that was test-fired in New Mexico over the past couple of weeks. Then they’ll analyze the data before clearing Starliner for the trip home.
My take:
They haven't been testing a whole system, just one thruster. I'm not sure I like the sound of that. But I know zero about what the nature of the thruster problem really was on Starliner going to ISS. Note also that there is not one single word about helium leaks.
Look closely at the wording of the last sentence in that press release: clearing them to ride Starliner back to Earth is apparently a foregone conclusion! So, either things are not as bad with the thrusters as has been ballyhooed, or else a really bad management decision has already been made. If the crew survives coming home, we'll never know which!
GW
Composite solid propellants today pretty much use a polymer binder in which you disperse the solid ingredients during mixing, then cure in place by both chemistry and cooking (essentially vulcanizing the rubber binder). The solids are usually AP (a class 1.1 mass detonable material as a powder that is also friction-sensitive), aluminum powder (a fire hazard), iron oxide powder, and carbon lack powder. The older ones used AN oxidizer, but AP performs a lot better.
The composite binders are CTPB or HTPB if you need to cold soak (these go down to -65F), or PBAN (if you never go below about 20 F, it cracks easily in the cold). Processing these things requires remote-operated equipment located in explosion-proof revetments. There is also GAP, but it is a liquid monopropellant explosive. Usually only a double-base plant can handle that.
The double base is even worse, you need liquid explosive precautions (impermeable surfaces everywhere) and you must handle raw nitroglycerin, which is extremely sensitive to all sorts of stimuli! The propellant is pelletized nitrocellulose flooded with nitroglycerin, which reacts to form a sort of plastic material. The same oxidizer and ballistic-aid solids can be dispersed among the NC particles, to form what is called a composite-modified double base. Those usually used AN as the added oxidizer, but AP is being used more, now.
As you can see, the facility investment and operating expenses of a solid propellant plant are enormous. Not many smaller businesses can afford to get started doing this anymore. "Aerojet Rocketdyne"/"Northrup Grumman" possibly operates a strategic big-motor plant in Utah (one of the two that were in Salt Lake, I don't know which one), and a small-motor tactical plant back east that was once Atlantic Research. Or they might be running the other other big-motor plant that was UTC-CSD's big plant on the Mississippi. I doubt they are operating two big-motor plants. I don't think the other small-motor plant that was Allegheny Ballistic Laboratory in West Virginia is operating any more.
The big-motor plants at least used to be mostly silo-based missiles, or space business stuff, that never soaks out very cold. They really like PBAN for such applications, and so have no experience using CTPB or HTPB. The small-motor guys had to meet severe cold soak to -65 F, and so stayed completely away from PBAN. CTPB was the older choice, HTPB is the newer one, and slightly better.
GW
edit update 7-24-2024: this posting is about the Falcon-9 second-stage failure trying to orbit several Starlink satellites:
I saw one small, obscure item that talked about an oxygen leak. A cryo liquid spewing into the engine bay in significant quantities from a plumbing leak would cause freezing of the high coastal humidity into lots of ice, accumulated into the engine compartment. Falling chunks of ice can do serious damage, as it did to Shuttle Columbia's wing. Such falling chunks of ice were seen in the video coverage.
The engine explosion (or whatever it was) could have been either damage from being struck by falling chunks of ice, or something to do with leaking oxygen in the presence of almost-inevitable leaking methane, or both. Hopefully, there enough of a data record to figure this out fairly quickly, for the FAA investigation.
GW
Edit update 7-27-2024: they seem to be cleared to fly Falcons again. Although, I have seen nothing about what the failure turned out to be.
Edit update 7-28-2024: saw one report saying it was some sort of sensing line that cracked in fatigue because its supporting bracket was busted somehow. That broken line let a bunch of liquid oxygen out. The report said NOT ONE WORD about the big falling ice chunks, but instead had some not-very-understandable words about the cold slowing part of the propellant flow, making the engine inoperable. Accordingly, I think (1) we are not being told the truth about the actual problem the oxygen leak caused, and (2) whatever that truth really is, they apparently fixed it, and successfully flew 3 Falcons in one day. The news article said they simply removed the sensor line from the engine design as no longer needed (which then begs the question of why it was still there).
Kbd512:
Is your system set up to do net metering? If so, it is essentially on the utility side of the isolation switch. It cannot power you during an outage; that requires it be hooked up on your side of the isolation switch. Them's just the rules in Texas for doing net metering. Any battery or generator needs to be on your side of the isolation switch, to cover you for outages.
We are doing a solar installation out here on the farm near McGregor; both the ranch house and the farm shop, which have two separate electric meters. We are doing it as net metering, because that's what makes the econ0mics look good. So I will have to install a generator on my side of the ranch house isolation switch, to cover outages, which are far less frequent with the co-op than the big utility that serves the cities around here. I already have a small generator for the shop, and will have to hook it in, the same way.
GW
Wasn't sure where to post this.
I saw a report that Ariane-6 flew, but there was a problem getting the whole mission done. The upper stage has some sort of APU (auxiliary power unit) whose exhaust supplies the ullage thrust necessary to light the engine for subsequent burns,. They got up there and did one relight, but after that the APU failed, and so the other relights could not happen. Partial success.
GW
The last Starlink launch on a Falcon-9 had a second-stage problem. It blew up its engine trying to ignite for the second burn that raises orbit perigee. The stage itself survived this, and they deployed the satellites. However, if the satellite ion thrusters cannot keep up with drag at too low a perigee, they may fall back and enter.
A photo taken by a camera on the stage shows a very unusual and heavy ice accumulation near the plumbing at the engine power head, shedding big chunks of ice. Whatever the problem was, somehow that ice is very likely related to it.
GW
I would only point out that heavy-lift rockets are a technology already-in-use and well-proven. Neither spin launch, nor light gas guns, nor any sort of electromagnetic catapult, are ready-to-use launch technologies. All would require considerable development and shakedown to become operational and reliable, if any of them actually prove feasible for launch at all.
GW
News reports indicate that NASA and Boeing on the ground still do not not understand where the problems came from. The plan (such as it is) now goes past the 45 day battery life, justified by charging at the ISS.
News reports still lie about "stabilized helium leaks". They shut off the shut-off valves. Open them back up to operate the systems, and the leaks resume. They are just counting on getting away from ISS, doing the deorbit burn, and jettisoning the service module before the helium can run out.
And I have not seen one single word about the status of the attitude control thrusters on the capsule itself, critical for entry attitude control! If the service module has problems, why not the capsule, too? Both were made by the same Boeing! They just haven't surfaced yet because those systems have not yet been activated in flight.
Other than that, I stand by what I said in post #90 just above.
GW
I see in the news reports the next booster is being sent to the launch pad.
GW
Re: pictures in post 330 just above.
The first one is what a conventional solid rocket design in a cylindrical case would look like with a propellant grain shaped exactly like a pool of liquid for sideways gee. At 1000's of gees, the solid propellant would indeed flow viscously into such a shape. It cannot stand up by itself at such extreme gee. I don't like this design because it has very low cross-sectional loading of propellant, leading to a very poor low ratio of propellant to total loaded motor weight, and thus a high inert fraction, compounded by the structural robustness required for the case itself to resist extreme lateral gee.
The second one is a design that I did for a thrusted entry decoy for warheads about 1988. It is entirely unsuitable for extreme-gee conditions in any direction, but shows that very unique and challenging problems can really be solved with out-of-the-box design thinking. Almost nobody in the industry would have attempted this, but I showed in subscale lab motor tests that the round-the-corner grain design actually works, and overcast technology was already well-known to work.
The third one is a sketch representing a very simple production small motor used for ullage on the old Saturns during Apollo. With the case beefed up, it could serve as an extreme-gee motor of high cross-sectional loading, for axially-directed extreme gee, because the propellant charge is the same shape as a liquid pool. But this is oriented the wrong way for spin launch as we understand it.
GW
That's an interesting idea. Never thought about it in those terms.
You would be somewhat restricted with a 15-day mission length. The old WW2 fleet boats were min 90-day missions, and some went past 120 days.
GW
The physics is favorable, but there is a cost: extreme launch gees.
GW