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#2051 2025-04-13 11:05:24
- GW Johnson
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- From: McGregor, Texas USA
- Registered: 2011-12-04
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Re: Starship is Go...
RE: post 2049 just above.
The image number labels are at the tops of the images. They are out of order in the posting.
The one labeled “image 8” was intended to be the first one, showing exactly what and how I was trying to do in the study. What I varied, how I varied it, and more importantly, what did not vary. I kept the heat shield nose radius ratioed to its diameter constant, as that nose radius is a major influence on the heating. Larger is lower heating. Had it not increased, the heating results would have been very much higher.
There were 4 images that were entry trajectory analysis results illustrations, to show where the numbers came from. The one labeled “image 5” was the first, with heat shield diameter 4 m equal to the 4 m base diameter of the protected object. This corresponds to a ballistic coefficient of 300 kg/sq.m, comparable to the Apollo command module’s 313.
The one labeled “image 3” was intended to be next, showing an 8 m dia heat shield on that same 4 m dia object, for a ballistic coefficient of 75 kg/sq.m.
I doubled the heat shield dia again to 16 m, for that same 4 m dia object, in the image labeled “image 6”. That produced a ballistic coefficient of 18.75 kg/sq.m.
I doubled the heat shield diameter yet again to 32 m for the same 4 m dia object, for a ballistic coefficient of only 4.68+ kg/sq.m. That one is labeled as “image 1”.
These 4 analyses produced peak convective heating rates at stagnation, and peak plasma radiation heating rates at stagnation, that were relatively trivial. I summed these to a total heating rate at stagnation, and computed the stagnation point surface temperatures for which thermal re-radiation equaled the combined convective-radiation input, without ablation, without conduction into the interior, and without any other cooling at all. Those results are summarized in plot form in the image labeled “image 2”. The temperatures decrease but stay rather high until the ballistic coefficient falls below about 50 kg./sq.m.
I also used the peak deceleration gees to estimate the peak average pressure exerted across the heat shield. Those results are also plotted in “image 2”. I was surprised to see the peak deceleration gee being the same for all cases. I expected higher gee with more heat shield area. But the thinner air higher up apparently just offsets that effect.
I went back and updated these results by estimating temperatures away from the stagnation point with the results plotted in “image 4”. The stagnation point plot is there, plus a plot representing attached flow on the windward side of the heat shield near its rim, and another plot representing any of the leeward heat shield surfaces, or the lateral surfaces of the protected object. All of these are surfaces within the separated wake zone.
Looking at those results in “image 4”, at ballistic coefficient near 50 kg/sq.m, the stagnation point temperature is at least 1000 C. Higher if reflective. Near the rim of the heat shield, that is still at least 800 C, higher if reflective. Only for the separated wake zone surfaces is it 500 C, yet higher if reflective. There are lower temperatures shown in the plots at lower ballistic coefficients, but the structures may well be getting too fragile to fly, despite the lower average pressure below the 4 KPa at 50 kg/sq.m.
I put together a little table of max service temperatures for several materials, which got posted as “image 7”. It’s not in any way comprehensive. With a rim temperature of 800 C or higher, there is simply no way in hell that any sort of silicone elastomer on any sort of fabric material, is going to be in reusable shape after a single entry from low Earth orbit! That is the real takeaway here!
There is a very good reason that practical spacecraft heat shields have all been ablatives up to now, the only exceptions being the refractory ceramics on the Space Shuttle and the X-37B. I think I just showed you exactly why that has been true. By doing the same kind of things H. Julian Allen and A. J. Eggers were doing for warhead entry scenarios, back in the early 1950’s. The physics has not changed. Some of the materials have.
GW
Last edited by GW Johnson (2025-04-13 11:08:51)
GW Johnson
McGregor, Texas
"There is nothing as expensive as a dead crew, especially one dead from a bad management decision"
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#2052 2025-04-14 11:29:56
Re: Starship is Go...
Thanks for that, GW. Your updated version of Fig 6. is closer to the Dr. Akin result of approx. stagnation temperature 800°C for a ballistic coefficient of ca. 20 kg/sq.m:
Your updated Fig 6:
This is in the range of the max. service temperature for some steel alloys:
Bob Clark
Last edited by RGClark (2025-04-14 11:32:35)
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“Anything worth doing is worth doing for a billion dollars.”
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#2053 2025-04-14 12:19:20
- kbd512
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- Registered: 2015-01-02
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Re: Starship is Go...
I'd like to point out how ridiculously low 18.75kg/m^2 (3.84lbs/ft^2) truly is. That is the wing loading equivalent of the average ultra-light aircraft, none of which are made from materials that will withstand 800C. For a 150,000kg dry mass vehicle, the heat shield surface area is 8,000m^2. A regulation American football field is 57,600ft^2, or 5,351m^2, which means the heat shield would need to be 1.5X the size of a football field. For all practical purposes, we don't build any flying vehicles of that size.
The Nextel fabric that ADEPT was made from costs around $20/ft^2. It would survive 850C without issue. However, the fabric for this heat shield would cost more than all 6 Raptor engines (now less than $200K per copy), and I'm guessing that this deployable fabric heat shield would need to be jettisoned to allow Starship to land. How do we do that?
This seems like a highly impractical solution without some kind of radical redesign of Starship. Starship might be able to deploy such an enormous heat shield, and it should protect Starship from reentry heating without any real issue, but how does one then land a vehicle ensconced in this giant fabric "lifting body surf board"?
Starship on a football field:
Where the heck are we stowing a heat shield that large and how are we either getting rid of it to land vertically, or somehow using the heat shield to soft land on the ocean?
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#2054 2025-04-14 14:22:10
- GW Johnson
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- From: McGregor, Texas USA
- Registered: 2011-12-04
- Posts: 5,918
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Re: Starship is Go...
Bob:
The final results are excruciatingly-sensitive to the speed and angle at entry interface, something I did not vary in this study. Just a tad faster or steeper raises gees and peak heating. Just a tad slower or shallower lowers peak gees and peak heating.
Dig out the entry spreadsheet posted in the "interplanetary transportation" topic, "orbit mechanics class traditional" thread., and run it for yourself. It has a user's manual. Just be sure to use realistic inputs. It's a "garbage-in, garbage-out" thing.
Kbd512:
Nextel now makes several ceramic cloths, all pretty much intended for aircraft engine nacelle fire curtain application. I am very familiar with the oldest one, an AF-19 fabric made from Nextel-312 fibers, which are aluminosilicate minerals. All these have a solid phase change at about 2300 F, leading to ~3% volume shrinkage, and catastrophic embrittlement. Which upon cooldown renders them extremely fragile, crumbling away to dust at the touch of the slightest breath of air. Those minerals melt at about 3300 F, but you cannot reuse those cloths (or anything else you make from them) if you exceed about 2200-2300 F.
GW
GW Johnson
McGregor, Texas
"There is nothing as expensive as a dead crew, especially one dead from a bad management decision"
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#2055 2025-04-14 14:45:55
- kbd512
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- Registered: 2015-01-02
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Re: Starship is Go...
GW,
My underlying point is that SpaceX requires RCC to make their fully exposed body fins and hinge lines survive reentry heating. We're not going to achieve ultra-light-like wing loading for a super heavy lift launch vehicle because it's not practical. We have our "go-to" materials for surviving reentry heating and they happen to work.
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#2056 2025-04-14 16:49:16
- GW Johnson
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- From: McGregor, Texas USA
- Registered: 2011-12-04
- Posts: 5,918
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Re: Starship is Go...
Actually, I quite agree with you.
GW
GW Johnson
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#2057 2025-04-14 17:22:14
- tahanson43206
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- Registered: 2018-04-27
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Re: Starship is Go...
For RGClark...
Here is an image GW asked me to post for you:
(th)
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#2058 2025-04-15 08:03:21
Re: Starship is Go...
I'd like to point out how ridiculously low 18.75kg/m^2 (3.84lbs/ft^2) truly is. That is the wing loading equivalent of the average ultra-light aircraft, none of which are made from materials that will withstand 800C. For a 150,000kg dry mass vehicle, the heat shield surface area is 8,000m^2. A regulation American football field is 57,600ft^2, or 5,351m^2, which means the heat shield would need to be 1.5X the size of a football field. For all practical purposes, we don't build any flying vehicles of that size.
…
I discussed before I think SpaceX is not taking the best approach to developing the Superheavy/Starship. They were spectacularly successful with the Falcon 9 by first getting the expendable version, then proceeding to reusability. If they had taken that approach with the SH/SS they would already be flying the expendable version and perhaps even also the partially reusable one, i.e., reusing the booster only, a la the Falcon 9.
Note then for the expendable version the dry mass of the Starship might have been as low as 40 tons:
Elon Musk @ElonMusk
Probably no fairing either & just 3 Raptor Vacuum engines. Mass ratio of ~30 (1200 tons full, 40 tons empty) with Isp of 380. Then drop a few dozen modified Starlink satellites from empty engine bays with ~1600 Isp, MR 2. Spread out, see what’s there. Not impossible.
https://x.com/elonmusk/status/1111798912141017089?s=61
Then with the ca. 20 ton fairing, the needed ‘wing’ area need to be added might be ca. 1,800 sq.m. But this wouldn’t be as heavy as regular wings with their thickness to generate aerodynamic lift. It would only have the character of a thin plate since it is meant only to be a drag decelerator.
Bob Clark
Last edited by RGClark (2025-04-15 08:53:00)
Old Space rule of acquisition (with a nod to Star Trek - the Next Generation):
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#2059 2025-04-15 09:40:41
Re: Starship is Go...
I discussed in the blog post the inflatable conical shield being investigated to allow the Cygnus cargo capsule to be reusable had the same ballistic coefficient as the Starship of ca. 60 kg/sq.m IF you take the dry mass of the Starship at the expendable 40 tons.
The problem is this conical shield was sized for a returning craft of mass of ca. 5 tons and it’s not certain how the conical shield would scale to higher mass, such as the Starship.
But there might be an example that would give us a reusable thermal shield for a vehicle the size of Starship. I’m thinking of the X-33/Venturestar.
The length in meters was 38.7m and width 39m. For the dry mass, the total gross weight was 2,186,000 lbs, propellant weight 1,929,000 lbs, and payload weight 45,000 lbs; giving a dry weight of 212,000 lbs, or 96,400 kg.
Using a hypersonic drag coefficient of 2, and considering the triangular planform requires multiplying by 1/2 the length*width to get the area, the ballistic coefficient calculates out to be 96,400/(2*1/2*38.7*39) = 64 kg/sq.m.
Remarkably close to the ballistic coefficient of the Starship at the 60,000 kg mass of the expendable’s dry mass + fairing mass.
But the added weight of the metallic shingle TPS of the X-33/Venturestar can’t be too high to allow the ballistic coefficient to remain close to this value.
The areal density of the metallic shingle TPS was about 10 kg/sq.m:
REUSABLE METALLIC THERMAL PROTECTION SYSTEMS DEVELOPMENT
Max L. Blosser*, Carl J. Martin*, Kamran Daryabeigi*, Carl C. Poteet **
*NASA Langley Research Center, Hampton, VA, USA
** JIAFS, The George Washington University, Hampton, VA, USA
https://ntrs.nasa.gov/api/citations/200 … 095922.pdf
The metallic tiles had better resistance to impact and rain than the ceramics at about the same weight.
Fig.3 Layered metallic sheeting separated by insulation.
Fig.21 Metallic TPS at same weight of ceramic tiles, ~10kg/sq.m.
At a 10 kg/sq.m. areal density, the added weight covering just the lower half of the Starship would be (1/2)*Pi*9*50*(10 kg/sq.m.) = 7,060 kg, proportionally small enough that the ballistic coefficient would still be ca. 60 kg/sq.m.
This would be advantageous in that you don’t need added wings and you don’t need an additional conical shield.
BUT for this to work SpaceX would have to go back to the smaller, expendable mass of the Starship. SpaceX had tested the X-33 metallic shingles and concluded they were inadequate. But that was with temperatures developed with the higher 150+ ton Starship. With a lighter dry mass, much reduced temperatures result.
Bob Clark
Last edited by RGClark (2025-04-15 09:46:40)
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#2060 Yesterday 09:44:37
Re: Starship is Go...
Thanks for that, GW. I wanted to ask in regards to the “hot metal” TPS, the heat would rapidly propagate to the upper side of the vehicle. This means the radiative surface area would double, thus doubling the heat emission.
Shouldn’t this improve the survivability?
Bob Clark
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