Starship is Go...

kbd512 · 2019-02-26 17:00:13

GW,

The doc I read said the charred material needs to be removed and then the remaining PICA-X can be reused as long as there's enough remaining material to contend with the next reentry. I'm sure that NASA could be wrong, given what they thought about Columbia's heat shield at the time that incident happened.

Is it better to refurbish or replace, as required, or roll the dice and hope for the best?

I only ask because it didn't work out very well for Challenger and Columbia. You're the one that's always telling us to test assertions before betting lives on them. How well has this been tested?

Are there any examples of ablative heat shield reuse on winged reentry vehicles without minimal refurbishment?

Is removing the charred material to leave a clean surface too difficult to do with a robotically-controlled sander?

What happens if a chunk of that charred material breaks off and hits one of Starship's giant fins when it blasts off from Mars?

I dunno, but it seems like we've been here before.

SpaceNut · 2019-02-26 18:03:46

The foam incidents were due to removing the sealing paint and formular changes for the external tank as well as the cold during launch which has formed ice and cracked foam chunks in addition to the O-Ring bleed by... since this was falling away strikes that hit the shuttle the BFR will not have that issue as there is nothing to fall into since its moving away from it.

kbd512 · 2019-02-26 21:27:53

SpaceNut,

The hubris is stunning. Lest we forget:

How We Nearly Lost Discovery by Wayne Hale

Read the article. Mr. Hale publicly apologized to the crew responsible for installation of the ET's TPS. There were no defects in workmanship. The act of filling and draining the tank cracked the foam. A piece of foam hit Discovery on the way up, too. Luckily, it struck a glancing blow and did no major damage.

From the article:

So were we stupid? Yes.
Can you learn from our mistake? I hope so.
So when you go to the Smithsonian and see Discovery there, think how lucky you are to see her whole, intact, and with her crews safely on the ground.
You see, this is how I found out that we were never really as smart as we thought we were.
Maybe that is a lesson that applies to you, too.

The Starship has bow planes or fins. It also has tail fins that protrude quite a ways away from the body of the vehicle. If a chunk of material breaks off there, where is it likely to go?

Let's say that never happens. That still doesn't matter. PICA-X is installed in blocks. On Mars, the propellant tanks won't be instantly filled to the top with propellant. It'll take lots of time. There will be lots of thermal cycles. Differential thermal expansion will happen as a result.

Fracture in Phenolic Impregnated Carbon Ablator

From the document:

C. Charred PICA Tests
Several specimens with ~2.3 mm notch size were tested inside the electron microscope. The crack initiation and growth was rapid once the critical stress was achieved, showing very similar behavior as virgin PICA samples. Videos and micrographs were obtained for the notched specimens. The crack initiated in the porous matrix phase and the crack path followed the porous matrix. Carbon particle shedding in the porous matrix was also observed in some videos. Fiber pull-out and bridging phenomenon, similar to virgin PICA samples, were observed at some locations as shown in Figure14. The stress-strain data and tensile strength for representative samples are shown in Figure 15. Most specimens failed between 600 kPa - 900 kPa. Again, based on test data from 8 samples, the introduction of a notch did not make any difference in maximum tensile stress at which failure occurred. The maximum tensile stress achieved for these specimens was slightly higher than FiberForm samples but significantly lower than virgin PICA. This suggests that fracture toughness of PICA ablator is driven by the extent of energy that can be absorbed by the porous matrix phase. The charred PICA specimens showed a similar non-linear stress-strain relation as FiberForm and virgin PICA samples. However, the strain-to-failure in charred PICA samples was lower than virgin PICA. To date, based on our knowledge; no other tensile testing of charred PICA material has been conducted. This was the first test series where stress-strain plots and tensile strength data for the in-plane tension loading configuration were obtained.

Let's get some hard data from the bubbas at ERC Corp who did the testing and Dr. Dan Rasky.

PICA-X is supposed to be made of sterner stuff, but how tough?

Maybe TUFROC could be used in critical areas?

Ceramic Matrix Composite (CMC) Thermal Protection Systems (TPS) and Hot Structures for Hypersonic Vehicles

louis · 2019-02-27 08:55:41

No one seems to have dealt with the specifics of what Everyday Astronaut says is being proposed for the Starship:

https://www.youtube.com/watch?v=LogE40_wR9k (from about 08:30).

1. A special stainless steel shell (with higher heat resistance) covering a regenerative cooling system (lots of pipes) through which liquid methane will flow and so cool the outer shell. It's the sort of system that has been used on the Space Shuttle's nozzles.

2. The shell will also be porous and will thus be able to sweat cold methane to help protect the shell.

These technologies seem to be in place, they just haven't been brought to bear on the issue of heat on re-entry.

Are people saying this concept is fatally flawed? If so, how?

I said no one: GW did address the issues. He had criticisms of the video but concluded: "However, a stainless design with liquid cooling probably could be made to work as a reusable item". GW states the body should be black. But it seems such a design is feasible. People certainly should not dismiss the design.

Last edited by louis (2019-02-27 09:01:54)

elderflower · 2019-02-27 11:15:53

That's what my fish scale proposal was about. The overall porosity that has been assumed will be subject to coking as the methane is pyrolised in the pores. I wanted bigger vents which would not so easily be blocked.
Maybe ammonia or demineralised water would be a better coolant.

kbd512 · 2019-02-27 14:34:02

Louis,

I think the concept can be made to work for an acceptable mass penalty, but any little defect could render that active thermal protection system inoperable. If that happens during a high velocity interplanetary reentry, the rocket and anything inside it will get BBQ'd. It's not as mass-efficient as passive heat shielding, but it does work.

1. DLR has proven that this works. Google "SHEFEX II". It was launched on a Brazilian VS-40 sounding rocket. SHEFEX II survived peak heating of 2,500C upon reentry using CC/SiC composites and tested the viability of the AKTIV transpiration cooling experiment.

2. DLR is in the process of working on "SHEFEX III", which is a miniature spaceplane / lifting body, sort of like Boeing's much larger X37, except that it has sharp leading edges. SHEFEX III is a 500kg reentry vehicle. The "SHEFEX III" reentry vehicle is expected to reach Mach 17 upon reentry and is the next step in flight validation of the transpiration cooling concept.

3. Incidentally, X37 has already successfully tested TUFROC on its leading edges during an orbital velocity reentry after an extended duration space flight. That solution is just .4g/cm^3, vs .21g/cm^3 for PICA-X, and should be lighter than any transpiration cooling scheme would be, presuming I did my math correctly using the equations that DLR provided from their arc jet ground test experiments. DLR tested air, H2O, and Ar during their ground test experiments, but not CH4, so far as I recall.

4. The "SHEFEX IV" vehicle is intended to be a full scale reusable spaceplane, like SNC's DreamChaser.

5. The only active heat shielding concept I've seen with the potential to be lighter than any of the present competing heat shielding technologies is the plasma heat shield technology from MSNW LLC. This has been successfully demonstrated in an arc jet heating facility, but was never developed into a flight test demonstrator for lack of funding from NASA.

Here's a brief regarding where DLR wants to take this technology:

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

Interesting conglomeration of info from NASA showing the lay of the land for cutting edge small vehicle tech, to include heat shielding:

State of the Art of Small Spacecraft Technology

RobertDyck · 2019-02-27 14:50:17

These ideas for re-entry cooling are all old. I remember hearing about them as NASA developed Apollo and did the first steps of Shuttle in the 1960s. Yes, NASA seriously tendered to their contractors request proposals for Shuttle in 1968. Ended up different. The problem with regenerative cooling during atmospheric entry has always been the mass of coolant is greater than an ablative heat shield. And plumbing adds complication, things that can fail. An ablative heat shield is a hunk of material on the bottom of your capsule, there's very little that can go wrong.

For an LEO shuttle, they came up with tiles. The tile material expands when hot, so it's manufactured as small tiles to prevent cracking/breaking, and applied with felt between the tile and aluminum alloy skin so expansion doesn't cause it to sheer off. Silica foam tiles are light weight (low mass) and do the job very well. They were manufactured with a densified under side where it adheres to the felt, so it won't break off. Black tiles had a black glaze, as thin as a coat of paint, that would emit (radiate) heat. White tiles did not. White tiles were lighter due to lack of glaze, but didn't radiate as much heat, so could only be applied where temperature was lower. STS-1 had a problem: vibration from SRBs caused a lot of the white tiles to shake off. The military used a telescope in Hawaii to image the Shuttle on-orbit, identify which tiles were missing. NASA engineers determined Columbia was able to safely return, even with the tiles missing. However, they very quickly developed AFRSI. That's a quilt; top layer is fabric made of fibres with the same high purity silica as the white tiles. Back fabric is normal fibreglass. Batting in the quilt is also fibres made of the same high purity silica. A water-proof polymer coating was applied on top to prevent the quilt from getting soaked by rain. This quilt actually provided better thermal protection than white tiles, but the reason they changed to it was vibration would not shake any off, and even strikes by ice-filled foam from ET would not significantly damage the quilt. There were "cuts" made by foam strikes, but a fabric flap 1 inch wide proved to be not a problem.

A craft that doesn't have foam over an ET does not have the problems that Shuttle had. Foam with nothing to protect it from humidity of Florida's coast? Foam applied to a tank of cryogenic liquid oxygen (the top tank)? Expect that will fill the foam with ice. The 1968 design included a piloted fly-back booster with wings, and an aircraft skin over the foam. STS-1 painted the foam. Anything to protect it from Florida humidity. But SLS and BFR will not have any ET with exposed foam.

louis · 2019-02-27 16:20:04

I've never liked tiles and I believe I'm right in saying they would need detailed inspection, maintenance and repair after landing...in my view not practical for Mission One to Mars.

RobertDyck wrote:

These ideas for re-entry cooling are all old. I remember hearing about them as NASA developed Apollo and did the first steps of Shuttle in the 1960s. Yes, NASA seriously tendered to their contractors request proposals for Shuttle in 1968. Ended up different. The problem with regenerative cooling during atmospheric entry has always been the mass of coolant is greater than an ablative heat shield. And plumbing adds complication, things that can fail. An ablative heat shield is a hunk of material on the bottom of your capsule, there's very little that can go wrong.
For an LEO shuttle, they came up with tiles. The tile material expands when hot, so it's manufactured as small tiles to prevent cracking/breaking, and applied with felt between the tile and aluminum alloy skin so expansion doesn't cause it to sheer off. Silica foam tiles are light weight (low mass) and do the job very well. They were manufactured with a densified under side where it adheres to the felt, so it won't break off. Black tiles had a black glaze, as thin as a coat of paint, that would emit (radiate) heat. White tiles did not. White tiles were lighter due to lack of glaze, but didn't radiate as much heat, so could only be applied where temperature was lower. STS-1 had a problem: vibration from SRBs caused a lot of the white tiles to shake off. The military used a telescope in Hawaii to image the Shuttle on-orbit, identify which tiles were missing. NASA engineers determined Columbia was able to safely return, even with the tiles missing. However, they very quickly developed AFRSI. That's a quilt; top layer is fabric made of fibres with the same high purity silica as the white tiles. Back fabric is normal fibreglass. Batting in the quilt is also fibres made of the same high purity silica. A water-proof polymer coating was applied on top to prevent the quilt from getting soaked by rain. This quilt actually provided better thermal protection than white tiles, but the reason they changed to it was vibration would not shake any off, and even strikes by ice-filled foam from ET would not significantly damage the quilt. There were "cuts" made by foam strikes, but a fabric flap 1 inch wide proved to be not a problem.
A craft that doesn't have foam over an ET does not have the problems that Shuttle had. Foam with nothing to protect it from humidity of Florida's coast? Foam applied to a tank of cryogenic liquid oxygen (the top tank)? Expect that will fill the foam with ice. The 1968 design included a piloted fly-back booster with wings, and an aircraft skin over the foam. STS-1 painted the foam. Anything to protect it from Florida humidity. But SLS and BFR will not have any ET with exposed foam.

louis · 2019-02-27 16:23:42

It's surely the case with any entry system - whether parachutes, ablative shields or tiles - that any defect could cause a disaster.

Thanks for the info.

kbd512 wrote:

Louis,
I think the concept can be made to work for an acceptable mass penalty, but any little defect could render that active thermal protection system inoperable. If that happens during a high velocity interplanetary reentry, the rocket and anything inside it will get BBQ'd. It's not as mass-efficient as passive heat shielding, but it does work.
1. DLR has proven that this works. Google "SHEFEX II". It was launched on a Brazilian VS-40 sounding rocket. SHEFEX II survived peak heating of 2,500C upon reentry using CC/SiC composites and tested the viability of the AKTIV transpiration cooling experiment.
2. DLR is in the process of working on "SHEFEX III", which is a miniature spaceplane / lifting body, sort of like Boeing's much larger X37, except that it has sharp leading edges. SHEFEX III is a 500kg reentry vehicle. The "SHEFEX III" reentry vehicle is expected to reach Mach 17 upon reentry and is the next step in flight validation of the transpiration cooling concept.
3. Incidentally, X37 has already successfully tested TUFROC on its leading edges during an orbital velocity reentry after an extended duration space flight. That solution is just .4g/cm^3, vs .21g/cm^3 for PICA-X, and should be lighter than any transpiration cooling scheme would be, presuming I did my math correctly using the equations that DLR provided from their arc jet ground test experiments. DLR tested air, H2O, and Ar during their ground test experiments, but not CH4, so far as I recall.
4. The "SHEFEX IV" vehicle is intended to be a full scale reusable spaceplane, like SNC's DreamChaser.
5. The only active heat shielding concept I've seen with the potential to be lighter than any of the present competing heat shielding technologies is the plasma heat shield technology from MSNW LLC. This has been successfully demonstrated in an arc jet heating facility, but was never developed into a flight test demonstrator for lack of funding from NASA.
Here's a brief regarding where DLR wants to take this technology:
Hypersonic Flight and (Re)-Entry in Germany – Overview and Selected Projects
Interesting conglomeration of info from NASA showing the lay of the land for cutting edge small vehicle tech, to include heat shielding:
State of the Art of Small Spacecraft Technology

GW Johnson · 2019-02-27 16:50:18

Spacex may be looking at a liquid-cooled metal surface, but I'd bet they will decide against it for weight and cost. I hope they can make it work, but nobody prior ever has.

To use rocket fuel (a hydrocarbon) as coolant, particularly a cryogenic (liquid methane), they will have to use very high pressures indeed inside the cooling passages to keep the methane from boiling, at a very significant metal temperature (or it won't transfer heat). It will tend to coke up on those passages and plug them. Kerosene is worse, but ALL hydrocarbons coke up.

The high temperature metal is weak precisely because it is hot. That inherently means very thick walls to contain the high coolant pressure. THAT is HEAVY! You cannot use titanium to save weight, because it is VERY MOST DEFINITELY NOT a high temperature material! Structurally, it is crap at 700-800 F.

You cannot get much advantage to re-radiation cooling from the hot metal at only 1000-1200 F, this effect is far more effective in the 1600-1800 F range, and higher still will re-radiate even better. Stainless is structurally junk at 1600 F, it still has a little strength at 1000-1200 F.

There are metals with better strength at somewhat higher temperatures than stainless, but they are expensive as all hell, and at least as heavy as stainless if not heavier.

Now you know why there are no liquid cooled skin designs actually flying, despite all the lab R&D contracts for the last half century. You can afford the weight and expense in a rocket engine, but not the whole damned airframe!

Guys, that puts you back to ablatives. You are right back in the shuttle tile problem, just with PICA-X panels instead of the more fragile silicate tiles.

Shedding bits of char shouldn't be much of a problem during ascent, because the thing will leave the sensible atmosphere at only about Mach 2 to 3, just like the shuttle (and X-37) did. Upon descent from LEO, entry interface is Mach 25-ish at about 90 km (60 miles). The air pressures are lower. Peak heating and peak deceleration gees are not at precisely the same time, but both are near the halfway point down: around Mach 12-ish at about 45 km (30 miles).

THAT is where the airloads would rip loose any old char, which is nothing but a very porous charcoal-like material. It will have no ice content at all. If your heat shield is tougher than a shuttle silicate tile (and PICA-X is), and if it is tougher than a carbon-carbon shuttle leading edge or nose cap (and PICA-X is), then you should not have much of a problem or risk. I really don't see why we have to remove char from a BFS heat shield after landing on Mars. After landing on Earth returning from Mars, you replace the whole thing anyway.

We'll get a chance to confirm this when BFR/BFS actually flies to LEO. It should be capable of about 4 round trips before they have to replace it. Let's see what sort of processing they do before they send it up for flights 2 - 4 on the same heat shield.

GW

Last edited by GW Johnson (2019-02-27 16:54:24)

SpaceNut · 2019-02-27 17:27:43

Pica is also a molded block that is shaped and adhered to the surface.
Its that way even on the current probes sent to mars and even on the Orion capsules.
BFR will be no different for how it will be shaped to fit on the structure.

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Oldfart1939 · 2019-02-27 20:25:18

Based upon GW's comments regarding ablatives, and after a single Earth reentry from Mars. this would argue for multiple vehicles--Starship type, and fewer BFR booster stages. Boosters should be capable of rapid reuse, but replacement of the PICA-X will take time. There is also the unmentioned need for the tanker stages, as well. Reentry of these from LEO should give a requirement of 2 tankers per Starship, with each flying 3 refueling missions before refurbishment. Or maybe I'm missing something here?

kbd512 · 2019-02-27 21:06:04

Louis and Oldfart1939,

I actually think UHTC's (ZrB2 or HfB2) and TUFROC are the way forward. Neither are ablative materials, so protected surfaces remain serviceable for some number of flights before oxidation inevitably exacts its toll on the heat shield materials. The leading edges of the fins can be protected by the UHTC's, the CH4 can be utilized as a thermal sink for the hot structure, the gaseous CH4 can either be dumped into the much larger primary tanks if sidewall secondary tanks are used (instead of that weird tank within a tank concept) or be consumed by Starship's RCS for attitude control during reentry. The TUFROC will passively protect the majority of the vehicle. In any event, those two technologies, both of which are space flight qualified at orbital reentry velocity, negates the requirement for PICA replacement. Think of it as a combination passive and semi-active system.

Ultimately, plasma aeroshells created with rotating electromagnetic fields are the future. It's also a convenient way to propel a spacecraft through interplanetary space without fuel and protect it from SPE's and GCR's through collisions with the sheath of charged particles it creates.

A Plasma Aerocapture and Entry System for Manned Missions and Planetary Deep Space Orbiters

GW Johnson · 2019-02-28 13:06:33

It'll be interesting indeed to see if any of the ultra-high-temperature ceramics (and similar) make it to ready-to-apply status instead of the gravy-train and exploratory research status programs they've been in.

Sure the material itself may be good up to 5000-7000 F, but how do you hang onto something that hot? And with what? All these materials are high-density, and therefore very high thermal conductivity (closer to values we associate with metals than with insulation materials). Such parts will run much closer to isothermal than anyone is used to. That means there will be no cool end to glom onto.

When that "trivial little difficulty" is finally solved (and it hasn't been solved yet), then these new "miracle" materials may actually see general application. Until then, only certain very restricted applications are even possible at all.

GW

Last edited by GW Johnson (2019-02-28 13:06:56)

kbd512 · 2019-02-28 14:26:36

GW,

How did the Space Shuttle engineers manage to hold onto the RCC and HRSI tiles?

Those things get so hot that they greatly resemble incandescent lightbulbs in action. UHTC's and TUFROC have already reentered and generated temperatures comparable to the leading edges of the Space Shuttle. The vehicles they were attached to were still in one pretty piece when they landed. If that hadn't happened, then I'd think that between AFRL, NASA, ESA, and DLR, someone would've put the brakes on those research projects.

TUFROC has already been applied to the leading edges of the X37 and will be applied to the leading edges of DreamChaser. It's .4g/cm^3 vs .21g/cm^3. That would mean that a TUFROC heat shield is roughly double the mass of PICA-X, but TUFROC isn't ablating away during reentry so only inspection is required before re-flight and there would be less removal and replacement. It's not a tile technology. It's fabricated to net-shape during fabrication. It's 4t vs 2t for PICA. If you have to replace the PICA after every round trip between Mars and Earth, how long does it take to become cost-ineffective and a significant component of flight costs?

I'm only suggesting that we use LOX-filled heat pipes (bow fins and nose) CH4-filled heat pipes (aft fins / landing gear) to transfer the heat from the UHTC's and supporting hot structure for the leading edges into the propellant and then consuming the pressurized gaseous propellant from the main tanks for RCS control of Starship during reentry maneuvering and by a micro gas turbine APU that provides electrical and hydraulic power during reentry and landing and subsequent deployment of the solar panels after Starship has landed. Combined with side wall tanks instead of tanks-within-tanks to prevent sloshing, that does away with Helium COPV's, landing tanks that don't provide structural reinforcement of the main tank sidewall, and minimizes the requirement for batteries optimized for power density over energy density specifically to handle high current draws from control surface actuators for very brief periods of time in comparison to the overall flight regime.

GW Johnson · 2019-02-28 17:29:47

Kbd512:

The shuttle tiles were very low density objects (Sp.gr. somewhere near 0.01 to 0.02). That means quite inherently that their thermal conductivities were also very low (also on the order of 0.01 to 0.02 if measured in BTU/hr-ft-R). Those things were 98%+ void space internally.

I made some stuff like that, and held a 12,000 F/inch gradient through a 1/4-inch thickness in a steady state test. That's the experimental stuff in my ceramic heat shield paper at the Boulder CO Mars Society convention a few years ago. Stuff like that is very fragile structurally, although mine was tougher than NASA'a, since mine was a fabric-reinforced composite, and theirs was not.

Used as a very low density heat shield, there would be a huge thermal gradient between the hot surface that sees the hypersonic flow and the colder surface attached by bonding to the airframe structure. If instead the density were high (near sp.gr. 1), the thermal conductivity would fall in the 0.1 to 1 range (again BTU/hr-ft-R). You simply would NOT get a large thermal gradient, no matter what the heat flow per unit surface really was.

Metals fall in the 1-10+ range when hot (again, BTU/hr-ft-R).

The shuttle stuff soaked out in a furnace glows incandescently. Removed from the oven, the corners and edges cool quickly, precisely because of those huge thermal gradients inside the material. That is how a bare hand can be used to pick up a shuttle tile sample that glows, as long as those edges are dark. You and I both have seen the photos or videos of that stunt.

You cannot do that with a dense ceramic. And the new UHTC materials are at least as dense as crucible ceramics. No one ever picked such a thing up with anything but tongs, because the high conductivity means a near-isothermal part. There are NO cool edges or corners.

And that's the problem: how can you take advantage of strength at incandescent temperatures, if you have nothing by which to hang onto a white-hot part? Might as well just use ordinary crucible ceramics. They already go hotter than you can hang onto. Cheaper, too.

GW

Last edited by GW Johnson (2019-02-28 17:33:38)

louis · 2019-02-28 18:19:55

Am I right in remembering that the Space Shuttles would always lose a large number of tiles on re-entry that would need to be replaced before the Shuttles could be re-used?

If I am, I think that argues against using anything like ceramic tiles as a heat shield for a Mars Mission.

kbd512 · 2019-02-28 19:22:07

GW,

The Space Shuttle's RCC parts ranged from 1.6g/cm^3 to 1.98g/cm^3. How did they get that to work without active cooling?

The UHTC's are affixed with ceramic fasteners and some are actively cooled by heat pipes using a gas or liquid to carry away the heat. That was what the AKTIV experiment set out to prove. SHEFEX II survived reentry intact as a result. The basic process works. If it didn't, how else do you imagine it is that SHEFEX II came back in one piece? DLR has pictures of it being fished out of the ocean on their website. UHTC's have also been ground tested with single and multi-wall CNT's as grain boundary reinforcement to inhibit superheated O2 from penetrating into the microstructure of the UHTC.

Here's how it works, according to the people who built SHEFEX II:

The SHEFEX 2 Campaign

Then there's this possibility:

Controlling the thermal contact resistance of a carbon nanotube heat spreader

SpaceNut · 2019-02-28 20:41:47

Not all tiles were equal in surface area or thickness also they did change in material content as well for a variety of locations for heat shielding. Its the thickening of the materials that reduces the heat to the vehicle skin.

https://en.wikipedia.org/wiki/Space_Shu … ion_system

All the tile must do is get you through the majority of the heat build up such that underlying materials can withstand the heat that is still being generated for a short period of time before its starts to drop as the air speed drops.

https://www.nytimes.com/2003/02/03/us/l … start.html

https://www.geek.com/news/2200f-space-s … s-1559855/

RobertDyck · 2019-03-01 07:32:02

Shuttle RCC was the nose cap and leading edges of wings. When it experienced maximum heating, it had maximum air flow. And the hypersonic shock wave increased density of air on those parts. That increases heating, but also increases air flow across the part. So it increases cooling as it increases heating at the same time. And there wasn't anything behind it, the RCC was bolted to the front of the wing as a whole piece.

kbd512 · 2019-03-01 08:55:52

Robert,

Could a combination of graphite felt insulation under flexible RCC sewn or bolted into / over 310 stainless work?

Overwrap with SiSiC (Silicon / Silicon Carbide; a plasma deposition coating for oxidation resistance, actually tested in an arc jet facility) coating for oxidation resistance:

Carbon Composites Inc - Flex-Shield

Insulation layer between the 310 and flexible RCC:

Carbon Composites Inc - Soft Carbon and Graphite Felt

It's not a purpose-built and tested-to-perfection solution, but it's comparatively cheap, field-replaceable using hand cutting tools, and available in rolls with a variety of dimensions. NASA had pretty good luck with COTS fabrics for HIAD and ADEPT, so why not try using this stuff since we want a light / cheap / easily replaceable fabric heat shield / insulator combination?

CCI recommends / uses carbon composite bolts / screws to fasten it to furnaces, which is exactly what SHEFEX II used to bolt the UHTC panels to the vehicle structure. All I want to know is if we can use a hybrid stainless / composites / fabrics protection approach in a way that makes economic and operational sense. NASA is also looking at using UHTC's in leading edges with RCC backing panels to thermal soak the rest of the structure.

GW Johnson · 2019-03-01 13:15:00

Kbd512:

Shuttle tiles were not 1.6 to 2-something g/cc, the rock fibers from which they were made had that density. The tiles were around 98-99% void space internally. That's 50-100+ times more void volume than rock fiber volume.

That is why the as-built tiles were lighter than a piece of Styrofoam, at somewhere in the 0.01 to 0.02 g/cc range. It's not an aerogel, or a conventional foam. It's like a very open-structured felt. The fibers were quite far apart, and sort of randomly connected.

The material I made from commercial supplies was a water-based pipe insulation paste of flakes and fibers, same alumino-silicate minerals as the fibers in shuttle tile. I alternated layers of this with layers of alumino-silicate fire curtain cloth. It would have been denser, except that I cured it above 212 F, at about 215 F. When you do that, the steam wormholes its ways out, reducing density by increasing void space.

I sealed the surface against gas intrusion with a paint coating of an alumino-silicate cement, and cured that hard at 215 F. I used it as a low density / low conductivity liner inside a miniature ram combustor at gas temperatures up to about 3500 F, and surface temperatures to a max of 3300 F, the alumino-silicate's meltpoint.

Solid phase-change cracking was not an issue for me because my liner was well-contained within the pressure shell, and my outer layers of insulation never got that hot. I could see tiny shrinkage cracks on the ID surface, but they posed no problems for my application.

As an external heat shield, you cannot tolerate such surface cracking, because retention is not positive, and because wind shear stress tears at the cracks. They do not occur if you stay below the phase change temperature of 2300 F. And that is why the same mineral as shuttle tile was restricted to 2000 F on the windward surface black tiles. It is why they could NOT serve as leading edge or nosecap insulation. The slow ablative carbon-carbon-composite had to serve there, where surface temperatures ran ~ 3000 F.

GW

kbd512 · 2019-03-01 14:40:37

GW,

I'm asking about whether or not the flexible RCC (not tiles) material developed by Carbon Composites Inc (with a Si/SiC layer deposited onto its surface to inhibit oxidation), in conjunction with one or more layers of graphite felt backer, could be used. Think of it as RCC over tile, like TUFROC, but without the added weight and flexible (min bend radius 4 to 12 inches, dependent on the thickness of the RCC) instead of tiles or rigid RCC. Imagine that we had a series of bolts or attachment points on the vehicle's exterior on the leeward side of the craft (in line with the tail fin that doesn't move). I want to wrap flexible RCC over the top, graphite felt as insulation underneath the RCC, and I want to connect it to the vehicle on the leeward side by bolting through the material with another RCC cap on top to inhibit movement.

I took Elderflower's "scales" idea and translated that into layers of flexible RCC with insulator backing material. It doesn't have to be graphite felt, could be anything flexible or spray on or otherwise formable to the desired shapes. The vehicle is "shrink wrapped" with rolls of flexible insulation that can be removed and field replaced using hand cutting tools (scissors or shears) and RCC bolts (like SHEFEX II). It's like tennis racket taping, but requires no adhesives because it's secured using pins / rods / bolts to the vehicle's spine.

RobertDyck · 2019-03-01 17:38:06

Oh. I can look up stuff, study stuff, but don't consider myself an expert with heat shields. Gary Johnson has actual hands-on experience. I would love to actually participate in designing a spacecraft. I watched the space program live as it happened from my parent's TV, from the last two missions of Mercury through Apollo 11 and beyond. When I was a child, my dream was to be an aerospace engineer. I could go on, but... Ok, let's see how far we can get.

Shuttle reusable heat works because atmospheric entry heating is less from LEO. Apollo used AVCOAT ablative to return from the Moon. NASA determined returning from Mars would hit the atmosphere at significantly higher speed, AVCOAT wouldn't cut it. In 1970 NASA announced PICA for that purpose. A PICA heat shield designed for Apollo CM and designed to safely return from Mars would have 3 times the mass of AVCOAT for lunar mission. NASA used PICA for robotic sample-return missions in 1995, and SpaceX designed Dragon with PICA-X. Presumably the new version of PICA is lower mass, but still greater than AVCOAT. By the way, Orion has an AVCOAT heat shield designed to return from the Moon. Orion would require a new heat shield for Mars.

So BFR/BFS/Starship. To answer your question I will have to research entry vehicle details. Realize one reason HIAD and ADEPT work is because they're designed to enter Mars atmosphere or Earth from LEO, they aren't intended to enter Earth atmosphere at trans-planetary velocity. Gary already posted speed differences. And posted a formula for heating due to speed. Notice the exponent! Do you want a heat shield to return Starship to Earth from Mars? That's a lot harder.

Last edited by RobertDyck (2019-03-01 18:00:06)

SpaceNut · 2019-03-01 18:35:57

Knowing that interplanetary speed is the issue for a ship having not enough fuel to slow, then what else can we do?

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