An important article...

louis · 2020-03-10 19:24:03

And here's the other way of doing it! Progress on New Glenn:

https://www.youtube.com/watch?v=avg0XZU2OBo

I prefer Space X's pioneer spirit!

Last edited by louis (2020-03-10 19:24:28)

Mark Friedenbach · 2020-03-10 20:02:12

SpaceNut wrote:

Musk tweeted that the tiles are hexagonal-shaped because that provides "no straight path for hot gas to accelerate through the gaps." The tiles will be installed on the windward side, towards the direction of re-entry, "with no shielded need on the leeward side." The hottest sections will have a "transpiration cooling" system, with microscope pores on the exterior that allow water or methane to ooze out and cool the exterior. That would minimize damage on the heat shields and allow the Starship to return to service shortly after a flight merely by refilling the heatshield reservoir. "Transpiration cooling will be added wherever we see erosion of the shield,".
Which means he does not know whether he needs it or not.....

I don't know how you got that idea from the quote. He knows that transpiration cooling will be required--the physics demands that. What's not clear is which parts of the exterior will need it the most. Which is reasonable since reentry dynamics are not something which simulations are 100% correct on. So put extra tiles on the first test model, and add the active cooling where the outer tiles burn through.

This is just old-school iterative engineering design.

louis · 2020-03-10 20:11:27

I agree Mark. Elon rightly sees it as a trade-off between the tiles and the main structure of the rocket. He expressed it as "how thick do the oven mitts need to be when you are handling something hot" ("something hot" being the tiles).

Mark Friedenbach wrote:

SpaceNut wrote:
Musk tweeted that the tiles are hexagonal-shaped because that provides "no straight path for hot gas to accelerate through the gaps." The tiles will be installed on the windward side, towards the direction of re-entry, "with no shielded need on the leeward side." The hottest sections will have a "transpiration cooling" system, with microscope pores on the exterior that allow water or methane to ooze out and cool the exterior. That would minimize damage on the heat shields and allow the Starship to return to service shortly after a flight merely by refilling the heatshield reservoir. "Transpiration cooling will be added wherever we see erosion of the shield,".
Which means he does not know whether he needs it or not.....
I don't know how you got that idea from the quote. He knows that transpiration cooling will be required--the physics demands that. What's not clear is which parts of the exterior will need it the most. Which is reasonable since reentry dynamics are not something which simulations are 100% correct on. So put extra tiles on the first test model, and add the active cooling where the outer tiles burn through.
This is just old-school iterative engineering design.

SpaceNut · 2020-03-10 20:25:54

The cone shaped nose says that the distance from the tip back several meters will be the burn through area and if the ridge lines are raised along the length they will as well where it changes from heat paint to tiles..The length of these tiles that will need cooling will be dependant on the angle of attack. Next is the fin edges that are for the landing legs will also need protection until its done with the portion of the down ward slope such that it spins for a tail down landing.
I got all of that from shuttle fuselage that required the RCC tiles on those locations....

Here is the issue with transpiring is the holes for the liquid must hold back the fluid until it heats and builds pressure. That said its going to push the tile away from the air frame that its sucured to....the larger the tile unless its made thicker will crack under that pressure build up. The tiles get weaker as a function of heating..The hole size will be the limiter for the cooling and pressure build up as a function of the heat passing through the tile. As the heat increase so will the pressure and the venting of that pressure can not happen as the hole is the restrictor to it doing the cooling....

kbd512 · 2020-03-13 16:10:48

If SpaceX is planning on a thermal soak and transpiration cooling design to contend with reentry heating, there's going to be a significant payload performance penalty associated with using transpiration cooling- such as extra propellant and pressurized structures. Incidentally, ESA was planning on using this technology to cool the leading edges of hypersonic space planes and reentry vehicles. I think they gave up on this when they figured out that the mass associated with transpiration cooling exceeded the mass of traditional heat shield materials and better UHTC materials were tested and found to be adequate for the purpose through ongoing experimentation.

Anyway, we do know that this works because ESA carried out a series of extended duration experiments proving that it works (they ran the experiments for as long as 120 minutes at Mach 4 and as long as 30 minutes at Mach 10), but any mass savings over a passive heat shield will be minimal to nonexistent and the structure is significantly more maintenance intensive because, as SpaceNut indicated, cracking occurs if the structure isn't appropriately reinforced to contend with heating and cooling and cooling is dependent upon small laser-drilled pores that aren't clogged with debris. If the pores become clogged, you get local heating and failure of the structure. The fine Martian dust could pose a significant problem for this cooling scheme.

I posted a series of links to DLR's work on transpiration cooling and use of UHTC's to permit sharp leading edges for hypersonic vehicles, but here's another link to one of the papers:

Transpiration Cooling Using Liquid Water

Here's another paper on using UHTC's for hypersonic vehicles (UK research efforts):

UHTC Composites for Hypersonic Applications

A PowerPoint presentation on hypersonic vehicle research in Germany (from 2011, so a bit dated now):

Hypersonic Flight and (Re)-Entry in Germany – Overview and Selected Projects

GW Johnson · 2020-03-14 10:22:18

I think it is clear that Spacex would only implement transpiration cooling in specific areas, presumed to be small, that are identified in prototype flight tests. That seems to be the intent expressed by Musk. Which means ablative tiles on windward steel substructure.

Personally, I think that once they are forced to address entry speeds above 8 km/s LEO entry speed, they will find out they need something on the leeward surfaces, too. That is precisely why the lateral sides of Apollo were covered with ablative, while Gemini and Mercury were bare metal coated very dark to radiate efficiently.

Apollo's lateral ablative was white, not dark, in order to reflect the plasma radiative heating that dominates the picture above about 10 km/s entry speeds. Speed returning from the moon was 10.9 km/s. Surprise, surprise.

The paint-on ablative used on Apollo was inadequate for free returns from Mars. That's why NASA developed PICA. PICA was expensive and difficult; Spacex developed its own version that was easier and less expensive. That is where PICA-X came from. Free returns from Mars vary with both average travel speed on the interplanetary orbit, and on the orbital positions of the planets. They vary from 12-to-17 km/s at entry interface. You take what you get, you don't get to choose.

Stagnation point convective heating varies proportional to velocity cubed. Stagnation point plasma radiation heating varies proportional to the sixth power of velocity.

These ugly little facts of life are EXACTLY why I say that bare steel leeward surfaces on Starship will never fly outside LEO. No transpiration system is ever going to change that picture; they are so heavy that they barely make any sense for just LEO entry at 8 km/s.

GW

Last edited by GW Johnson (2020-03-14 10:31:47)

louis · 2020-03-14 11:58:25

I think Paul Wooster, Chief Engineer for the Mars Mission at Space X, has stated explicitly they will go for orbital capture (at both ends) - two or three orbits, before you descend...doesn't that change the picture somewhat?

GW Johnson wrote:

I think it is clear that Spacex would only implement transpiration cooling in specific areas, presumed to be small, that are identified in prototype flight tests. That seems to be the intent expressed by Musk. Which means ablative tiles on windward steel substructure.
Personally, I think that once they are forced to address entry speeds above 8 km/s LEO entry speed, they will find out they need something on the leeward surfaces, too. That is precisely why the lateral sides of Apollo were covered with ablative, while Gemini and Mercury were bare metal coated very dark to radiate efficiently.
Apollo's lateral ablative was white, not dark, in order to reflect the plasma radiative heating that dominates the picture above about 10 km/s entry speeds. Speed returning from the moon was 10.9 km/s. Surprise, surprise.
The paint-on ablative used on Apollo was inadequate for free returns from Mars. That's why NASA developed PICA. PICA was expensive and difficult; Spacex developed its own version that was easier and less expensive. That is where PICA-X came from. Free returns from Mars vary with both average travel speed on the interplanetary orbit, and on the orbital positions of the planets. They vary from 12-to-17 km/s at entry interface. You take what you get, you don't get to choose.
Stagnation point convective heating varies proportional to velocity cubed. Stagnation point plasma radiation heating varies proportional to the sixth power of velocity.
These ugly little facts of life are EXACTLY why I say that bare steel leeward surfaces on Starship will never fly outside LEO. No transpiration system is ever going to change that picture; they are so heavy that they barely make any sense for just LEO entry at 8 km/s.
GW

kbd512 · 2020-03-14 14:47:25

Louis,

If you fail to aero-capture using an atmosphere with a density that varies wildly in comparison to Earth's upper atmosphere, then you either become another crater on the surface of the planet or you get ejected into deep space with no possibility of return. Mars' atmospheric density varies with the time of the day and season of the year. Since JPL still doesn't have that entirely figured out, neither does SpaceX. We'd need another one of those "useless" robotic missions to figure that out, or we can just experiment with Starships until SpaceX runs out of money or gets it right.

GW Johnson · 2020-03-14 14:49:59

"I think Paul Wooster, Chief Engineer for the Mars Mission at Space X, has stated explicitly they will go for orbital capture (at both ends) - two or three orbits, before you descend...doesn't that change the picture somewhat?"

It would if they could. I myself have never heard that from anyone at Spacex. Where will they get the extra delta-vee to do that, even from a Hohmann trajectory? It's even higher if you fly a faster trajectory.

Mars arrival is delta-vee free except for touchdown, if you do a direct entry. It's about 6 km/s entry speed from Hohmann, 7.5 km/s from a 6-month trajectory. Touchdown is likely more than they've been allowing; closer to 1 km/s than 0.3 to 0.5.

If you go into Mars orbit first, that's 1.8 to 2.1 km/s more from Hohmann, more yet from a faster trajectory. Then it's only 3.6 km/s entry from orbit, and the same touchdown burn as direct entry.

Getting from Earth orbit to a Hohmann trajectory to Mars is about 4 km/s delta vee. More for a faster trajectory, call it 5. It's the same in reverse getting from the trajectory into Earth orbit.

For the outbound trip: call it 4-5 km/s to depart, then 6 to 7.5 km/s for the heat shield to dissipate upon Mars direct entry, then as much as 1 km/s to land. That's a total of about 5-6 km/s required to do the one-way mission.

If instead you stop in orbit, it's still 4-5 km/s to depart, about 2 km/s to arrive, dissipate only 3.6 km/s on entry, then use up to 1 km/s to land. Total = 7-8 km/s. If all they have available is 5-6 at the desired payload, where do the other 2 km/s come from?

For the return trip: it takes 6 km/s to get off the surface direct to the interplanetary Hohmann trajectory, even more for a faster trip, like 7.5. Dissipate 12-17 km/s for direct entry at Earth with the heat shield, then use about 1 km/s to touch down. That's 7-8.5 km/s min to return, which at the same inert and a full propellant load means a LOT less payload returning.

If instead you decelerate into Earth orbit, it's still 6-7.5 to depart Mars, plus 4 to 5 to enter Earth orbit. Dissipate only 8 km/s with the heat shield for an orbital entry, then use about 1 km/s to land. That's something in the neighborhood of 10-13.5 km/s delta vee for the trip home. That's around 3-5.5 km/s more delta-vee than the direct entry trip. So where is it going to come from?

The mass ratio and the Isp does NOT support such in a single-stage vehicle. I used the earlier projections of inert mass, propellant mass, and payload mass to run the mass ratio for this vehicle. It does NOT support such high delta-vees. Try 385 sec Isp (effective exhaust velocity 3.776 km/s), inert 100 tons, propellant 1100 tons, and payload from 0 up to 200 tons. See what you get with the rocket equation.

Here's what I get:

0 payload: MR = (1100 prop +100 inert)/(100 inert) = 12. dV = Vex ln(MR) = 9.38 km/s. Looks attractive, but that's ZERO payload!

100 ton payload, MR = (1100 prop + 100 inert + 100 payload)/(100 inert + 100 payload) = 6.5 for dV = 7.07 km/s. We need 5-6 direct, 7-8 to stop in orbit. Well, we can certainly get there on a direct entry flight using an 8.5 month Hohmann, and it still works faster, but not enough to reliably stop in orbit. We need 7-8.5 to return direct, and 10-13.5 to stop in Earth orbit. Not really enough to reliably get home on a direct entry flight, even on a min energy Hohmann. Return is therefore under 100 tons payload, and "slowboat" direct at that.

200 ton payload, MR = (1100+100+200)/(100+200) = 4.667 for dV = 5.82 km/s. We need 5-6 km/s for the direct-entry trip there, and 7-8 to stop in Mars orbit. You need 7-8.5 km/s to return direct, and 10-13.5 to stop in Earth orbit. You could get there with a min-energy Hohmann direct-entry trajectory. Forget coming home at that payload. And forget stopping in orbit at either end.

I'm sorry Louis, that's just the numbers of this. Spacex has been doing great things, but NO ONE can violate physics and math.

They have a very long way to go to get to a vehicle that performs so highly as was projected in 2019. That would be 85 tons inert mass with 1100 tons propellant plus 100-200 tons payload. So I used 100 tons here, and I really don't believe that. Not yet. None of the prototypes is anywhere near that lightweight at 120-ton class, and that's still without many square meters of 4+ cm thick heat shield.

When you hear ridiculous-claim stuff like that coming from various figures at Spacex, even from Musk himself, it shows their ignorance, really. Quite honestly, they themselves don't know what this thing will wind up becoming. Nor what troubles it will face in order to fly. They have never done this before, and (most importantly) they don't bother to ask any of us who have. They have people on their staff who have been through the same numbers I crunched, but manager types rarely listen well enough to their engineers. Gets in the way of what they want to sell.

Technical arrogance like that equates (generally and in all practicality) to technical ignorance, which I have observed multiple times throughout most of my many years. It kinda shows, Louis. And it also most certainly does show recently at Boeing, who has been killing people with their technical arrogance. (So has Airbus, but that hasn't been in the news of late.)

Nevertheless, I think Spacex will eventually fly something that could work, at least as a big transport to LEO. They will have to take huge losses losing some of these vehicles in flight testing. They need to be prepared for that, and I'm not at all sure they really are. But they MUST get prepared, because it WILL happen! And because those losses cost a whole lot more if you lose a crew. Which they well could.

GW

Last edited by GW Johnson (2020-03-14 15:01:14)

GW Johnson · 2020-03-14 15:03:53

Aerocapture, meaning multiple-pass aerobraking without an actual descent and landing, has been done once or twice with unmanned probes. It is one hell of a long way off from being something reliabe enough to risk a crew. Especially at Mars with its atmospheric density variable by a factor of 2 to 3, or even more.

GW

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