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SpaceX Discussion Thread

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13 minutes ago, MaverickSawyer said:

Y'know, just had a bit of a brainwave...

Ablative heat shields outgas to carry away the heat, right? I can't remember if the outgassing  has an impact on the shock front, but if it does... Maybe there's an advantage to the venting of methane on the windward side... Something like gas film cooling in a rocket engine, perhaps?

A gas film that can combust? Sounds nice.

Why dont they use LOX instead?

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15 minutes ago, MaverickSawyer said:

Y'know, just had a bit of a brainwave...

Ablative heat shields outgas to carry away the heat, right? I can't remember if the outgassing  has an impact on the shock front, but if it does... Maybe there's an advantage to the venting of methane on the windward side... Something like gas film cooling in a rocket engine, perhaps?

That's the point.

The bow shock where the majority of reentry heating occurs is not at the hull, it's some distance from that (air compresses as much as possible, then heats). Outgassing basically moves the bow shock a little further from the hull. It will also carry some heat away as it disassociates.

Edited by tater

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18 minutes ago, MaverickSawyer said:

Y'know, just had a bit of a brainwave...

Ablative heat shields outgas to carry away the heat, right? I can't remember if the outgassing  has an impact on the shock front, but if it does... Maybe there's an advantage to the venting of methane on the windward side... Something like gas film cooling in a rocket engine, perhaps?

It does have an impact, in fact, it’s one of the main mechanisms of heat dissipation in ablative heat shields. Scott Manley talked about it in his recent vid about heat shields.

Edited by sh1pman

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In the Ablator vs Active Cooling debate, I think there are two factors to remember here.  The first is that a stainless steel skin that sneers at 1,000oF doesn't need nearly as much protection as lesser materials, and requires no extra protection on the back side. The second is that the double-skin cooling channel approach should also provide a lot of the strengthening and stiffening of the windward side that would be required to deal with the aerodynamic forces anyways. All that should save a lot of mass compared to ablator.

Edited by StrandedonEarth
tweaking
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3 minutes ago, Xd the great said:

A gas film that can combust? Sounds nice.

Why dont they use LOX instead?

Wait, lemme get this straight...

You're worried about a combustible gas being used for cooling, and turn around to suggest that instead, they should create a high temperature, pure oxygen environment... One so hot that, if memory serves, can actually turnm molecular oxygen into atomic oxygen.

rofl.gif

Disregarding the fact that oxygen has terrible thermal absorption properties, you'd be exposing the entire vehicle to a HIGHLY oxidizing environment, and even stainless steel has its limits... Just look at the exhaust stacks of a Cessna.

Flaming methane is actually preferable, as that would keep it from becoming trapped in the atmosphere as a greenhouse gas.

 

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4 minutes ago, MaverickSawyer said:

You're worried about a combustible gas being used for cooling, and turn around to suggest that instead, they should create a high temperature, pure oxygen environment... One so hot that, if memory serves, can actually turnm molecular oxygen into atomic oxygen.

If the windward side is already at 1750K, then methane will just instantly turn into a soup of ionized particles and dissociation products. It won’t really burn.

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2 hours ago, MaverickSawyer said:

I strongly suspect after. Think about it... Rocket engines use actively cooled walls in the hottest areas, right? It'd be easier from a production standpoint to use ablative cooling, but they use regenerative cooling instead because it's more reliable.

Regenerative cooling is effectively "free" from a mass standpoint (other than the plumbing involved). The cooling material is propellant that's going to the combustion chamber anyways, and is therefore wonderful propellant mass, rather than hateful not-propellant mass.

Also, it's basically free heat, letting you ramp up combustion chamber temperature a bit without needing additional combustion. For a hydrolox engine, that's definitely a neat perk; the preferred exhaust gas is unburned hydrogen, so if you gain additional heat from regenerative cooling, you can run a little bit more fuel-rich while maintaining chamber temperature.

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41 minutes ago, Starman4308 said:

Regenerative cooling is effectively "free" from a mass standpoint (other than the plumbing involved). The cooling material is propellant that's going to the combustion chamber anyways, and is therefore wonderful propellant mass, rather than hateful not-propellant mass.

For this active, transpirational cooling, the propellant is of course lost. Still, even at the high end of 20 tons for Starship (some calcs I have seen people do say 5 tonnes), that's a small price to pay.

Assume the high number, 20t. Apollo TPS was ~116kg/m2. If Starship had to use that much TPS (half the hull, plus fins, etc, we're approaching 200 tonnes of TPS. PICA-X is much less dense, but even if it is 10X lighter, the high end prop cost for active cooling via transpiration is a wash. If PICA isn't 10X better than Apollo's TPS, then they have more room for higher entry velocities without sacrificing anything at all. If the steel actually masses less than CFC, they also get that gain. If the leeward side need less protection as well, they gain even more.

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If the methane explodes after being vented...

*Kerbal hat on*

...it's like an engine on the entire side of the ship to slow it down faster!

*Kerbal hat off*

Am I wrong, or am I very wrong?

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58 minutes ago, tater said:

Assume the high number, 20t. Apollo TPS was ~116kg/m2. If Starship had to use that much TPS (half the hull, plus fins, etc, we're approaching 200 tonnes of TPS. PICA-X is much less dense, but even if it is 10X lighter, the high end prop cost for active cooling via transpiration is a wash. If PICA isn't 10X better than Apollo's TPS, then they have more room for higher entry velocities without sacrificing anything at all. If the steel actually masses less than CFC, they also get that gain. If the leeward side need less protection as well, they gain even more.

I think Starship will have an easier time re-entering than Apollo, because it has more surface area per mass, and can also generate more lift during re-entry, so it should need a thinner TPS than Apollo needed, although it probably won't be 10x thinner. Especially if they want to re-use it.

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Here are some numbers for the re-entry cooling energy budget :

 

from this thread https://www.reddit.com/r/spacex/comments/a9y9r0/an_energy_budget_for_starship_reentry/

The entry speed they were using : 8km/s (so LEO reentry, not mars(12km/s?) or lunar (11km/s?))

Mass : 110t

Kinetic Energy = 3500 GJ (LEO re-entry)......

Heat energy reaching the craft = ~ 1% or 35GJ for the cooling system to deal with. 

This is the key number here and i'm not sure how accurate or precise it is.

I'm a thermodynamics rookie, so i have to rely on other folks to refine it.

The other critical number is the transpiration cooling : 3GJ of cooling per tonne of methane.... high?, low? about right?

The spacecraft itself might end up sinking 5, 10 or 20GJ? for a couple of minutes before it starts to cool in the lower atmosphere.

I will be keen to see if Elon gives us some actual numbers.

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I'd been hum-drumming about a post about this over the weekend, and now I'm late to the party! Oh well, maybe some of this will still be interesting to consider. Please keep in mind, these are just the thoughts floating in the head of some engineer! No claims for gospel or magic here.

How to pump the fluid? First thing's first, a big advantage of cooling by perspiration rather than just circulating coolant along the inner wall is that the pumping will be very much simplified. For cooling with circulation you'd need to keep very high flow rates of coolant across the inner surface to keep the heat transfer rate high. High circulation rates means constant pumping, which is very difficult when your fluid has evaporating gas bubbles in it. Turbopumps can't handle it, so you'd either need to separate out the gas in a vapor-liquid separator (heavy, and generally hate shaking / changing gravity) or pump the fluid with a positive displacement pump (really heavy). However, by instead 'perspiring' the coolant through microporous channels, we make sure all the pesky bubbles are leaving the rocket right as they form! The coolant could just be pumped slowly without circulation, like water irrigating a field. This lets us use a nice, 'small', 'low-power' turbo pump.

Would the coolant evaporate within the pores, or be sprayed out to evaporate 'in the wind'? I'd guess we'd evaporate in the pores, hopefully reasonably close to the surface. I'm thinking since most of the heat is apparently absorbed from radiative emissions (rather than conductively/convectively from the air) increasing the thickness of the boundary layer will be helpful but not the primary effect. I could be convinced otherwise though, it's interesting to consider!

What size of pore and how distantly spaced? Let's just ballpark some numbers. I was thinking an outer steel skin of 2-20mm would be sane. Assuming the temperature drop is still mostly across this outer skin (since it's coming apparently predominantly from radiative heat) We'd want a pore aspect ratio (length/diameter) of maybe 3:1 - 20:1 with closely packed pores in order to prevent there from being hot spots in the steel where the heat sneaks through and causes bubbles to evaporate inside the vessel. So, perhaps 250-750um pores? Just taking some easy values, a 4mm skin with 10:1 aspect ratio pores would have 400um diameter pores, hexagonally arranged at maybe a 1.6mm center-to-center spacing. This makes the total material about 80% dense (20% air gaps in the steel), and is arranged so any point on the surface is atleast 5 times closer to a pore than it is to the inner surface. This should keep most of the original strength and also have minimal conduction to the interior. Nothing's magic in these numbers, they just sound like nice guesses for an initial picture. Perhaps for practical reasons, the coolant may evaporate .5-1mm into the pore rather than right at the surface. In this case, my example would have any point on the surface being 4 times closer to the evaporation in a pore than it is to the inner surface. This was the number I hunted for to get the 1.6mm spacing.

How many pores on the craft? Starship is planned to be 9m Diameter x 54m length, and only half of it (I assume) will have pores. Given it has a tapered shape and fins, I'm seeing about a 750m^2 surface area needing pores. Dividing through I'm getting a little over 300 million pores. Yeah, that sounds about right!

How do we make 300 million tiny pores? There's a dizzying array of options, but laser ablation is the runner to beat. 400um holes with a 10:1 aspect ratio made by the billion in stainless steel is right up its ally. We could talk a lot about the particulars for a case like this, but I think it's by far the simplest option.

My main issue with everything is corrosion. Stainless steel is corrosion _resistant_ not corrosion proof in sea-spray conditions, like after a barge landing. These super-high-surface-area pores will easily be able to catch and hold water by surface tension. It's kinda asking for trouble, so maybe.... Just Maybe... They'll prevent this by applying a positive pressure of clean dry air to the inner jacket, to constantly blow air out the pores.

And just maybe... If they do... I'll get to play air hockey on it! 

Looking forward to hearing what people think.

 

Edited by Cunjo Carl
Oops, off by an order, then a factor. Hopefully fixed now.
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1 hour ago, cubinator said:

If the methane explodes after being vented...

*Kerbal hat on*

...it's like an engine on the entire side of the ship to slow it down faster!

*Kerbal hat off*

Am I wrong, or am I very wrong?

an engine that thrust the spacecraft into more heat? If the methane combust behind?

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10 minutes ago, Cunjo Carl said:

<snipped great stuff>

Looking forward to hearing what people think.

Sounds pretty reasonable, as does the positive pressure to mitigate corrosion at a launch site. That could presumably work for dust as well (Mars, for example, looking down the line).

Edited by tater
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1 minute ago, tater said:

Sounds pretty reasonable, as does the positive pressure to mitigate corrosion at a launch site. That could presumably work for dust as well (Mars, for example, looking down the line).

Awh, jeeze. Good call! The statically charged and super-fine martian dust would have a field day with those pores.

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PS--I guestimated the surface area at 5000 square meters as well in my TPS calc just up thread.

1 minute ago, Cunjo Carl said:

Awh, jeeze. Good call! The statically charged and super-fine martian dust would have a field day with those pores.

I am not an "Occupy Mars!" person in the least, but SpaceX certainly likes to think ahead to this goal, so I bet that idea would make total sense to them (both for South TX as you suggest, or Mars).

(also the thing has to sit on Mars a LONG time to refill the tanks with ISRU (assuming that TRL ever gets anywhere near doing for real).

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2 minutes ago, tater said:

PS--I guestimated the surface area at 5000 square meters as well in my TPS calc just up thread.

I am not an "Occupy Mars!" person in the least, but SpaceX certainly likes to think ahead to this goal, so I bet that idea would make total sense to them (both for South TX as you suggest, or Mars).

Huh, good to know. It's pi*D*H for surface area of a cylinder, and I'd wound up squaring the D by accident in my first calc! Oops.

My latest looks like: ~500m^2 = ((54m/1.5)*3.14*9m)/2

where the 1/1.5 factor on the length accounts for the pointed shape, and the 1/2 factor on the whole thing counts for the pores on just one side... But you know what, my estimation is ignoring the fins which would totally need cooling! Now I'm getting ~750m^2, but still no where close to the 5000. I'm pretty shot though. I'll re-edit my post for the 750m^2 value. It doesn't change any of the bottom lines, but it'll make me feel better.

 

I'd gotten into space flight as a kid when my grandpa gave me Zubrin's book on going to mars... I won't pretend it's smart or practical, but dang if it's not cool! To be honest though, for me, actually going there is an after thought. I just like the tech!

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I just redid my calc, I thought 5000 was what I used, but it was in fact 1500. I did 2Pirh (no need for the end caps), then checked 50 vs 55m, and rounded down to ~1500.

1500*116 (the apollo tps density/m^2)= 174,000kg (which I rounded up to 200 tonnes (fins, etc)) as an upper bound.

So I was off by a factor of 2, I meant to divide by 2 but forgot.

So upper limit on the order of 100 tonnes (Apollo TPS), and realistically PICA-X for this application is some fraction of that. 20% would be 20 tonnes.

PS--like your EV Nova avatar image. ;)

 

Edited by tater

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9 hours ago, MaverickSawyer said:

It'd be easier from a production standpoint to use ablative cooling

Nope; ablative coating has fallen out of favour on smaller thrusters, replaced by various forms of pure radiative cooling. On larger engines, it’s supposed to be simpler (e.g. the “low-cost by design” RS-68 uses it), but it has enough wrinkles to make regenerative cooling simpler.

Plus, as you may notice, regenerative cooling dates back to the 1930s, while ablatives would arrive from large solid-propellant missiles.

12 hours ago, MaverickSawyer said:

If they tried that with Raptor at full power, I personally guarantee they'd have nothing left of the test stand in under a second. Staged combustion is EXTREMELY intolerant of errors or flaws... just ask the teams behind the NK-33 and Aerojet Rocketdyne's conversion of them for use on Antares...

Shouldn’t happen. NK’s teams were well-burned by random debris ingestion on the N-1, and RD-170 (171) had similar problems on the Zenit (but not on Energiya). There are pretty fine wire mesh filters on all intakes now.

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2 minutes ago, DDE said:

Nope; ablative coating has fallen out of favour on smaller thrusters, replaced by various forms of pure radiative cooling. On larger engines, it’s supposed to be simpler (e.g. the “low-cost by design” RS-68 uses it), but it has enough wrinkles to make regenerative cooling simpler.

Plus, as you may notice, regenerative cooling dates back to the 1930s, while ablatives would arrive from large solid-propellant missiles.

Hmm. I guess I misunderstood ablative cooling, then, as I thought it was more prevalent in older engines from the '50s and '60s. I know from second-hand experience that building modern regeneratively-cooled chambers and nozzles can be an absolute hunchfuster, depending on the process used. It either needed extremely precise machining and advanced bonding techniques/facilities, or it's manual labor intensive. (Additive manufacturing could seriously upend that equation, but that's not germane to the discussion at hand.)

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33 minutes ago, MaverickSawyer said:

I know from second-hand experience that building modern regeneratively-cooled chambers and nozzles can be an absolute hunchfuster, depending on the process used. It either needed extremely precise machining and advanced bonding techniques/facilities, or it's manual labor intensive.

Yes, but the esoteric techniques have been mastered to the point where even IRFNA and (I suspect) trifluoride/pentafluoride could be used fairly reliably. Ablatives remain rather unpredictable, however, and that’s probably inherent.

Edit: it appears that Energomash fired an Isayev 4D75 staged combustion engine with pentafluoride instead of NTO as RD-503. Zis is not nuts, siz is super-nuts, but I’m not sure which propellant served as coolant.

Edited by DDE

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4 hours ago, DDE said:

Yes, but the esoteric techniques have been mastered to the point where even IRFNA and (I suspect) trifluoride/pentafluoride could be used fairly reliably. Ablatives remain rather unpredictable, however, and that’s probably inherent.

Edit: it appears that Energomash fired an Isayev 4D75 staged combustion engine with pentafluoride instead of NTO as RD-503. Zis is not nuts, siz is super-nuts, but I’m not sure which propellant served as coolant.

pentafluoride is not the real problem, but the real problem would be pressure.

How would a series of tubes welded  together hold up at 300atm, as in a raptor engine?

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9 minutes ago, Xd the great said:

How would a series of tubes welded  together hold up at 300atm, as in a raptor engine?

A series of vacuum-brazed tubes are doing just fine in the 230 atm RD-170, although the throat region is machined instead.

Glushko developed initial vacuum-brazing techniques back in the late 1940s, along with welding techniques that allowed leak-proof welds in the propellant injector manifolds, something US designers avoided for decades.

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Uh oh.

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