Hypothetical SpaceX SSTO

[nadreck]

Elmar, I consider that the mass of the Dragon 2 at re-entry to be something on the order of 6 to 8t. But it is not well documented. The only places I have seen specs for the Dragon 2 for fuel capacity are in this document on page 2-5.  That puts the fuel mass at 1.6t or 20% of the 8t I am expecting as re-entry mass.  I never specified the landing propulsion system for the reusable Raptor upper stage, but I feel that the 5t of propellant for landing was a reasonable number as it scales linearly with the size of the vehicle presuming relatively similar ISPs.  A deep throttling Raptor-Vac would have much lower ISP than a hypergolic fueled superDraco I presume. I expect either smaller methalox landing engines, or superDracos and hypergolic fuels, but either way 5t seems reasonable. Which scales to 50t when your reentry mass is 10 times.

As for TPS, on a Dragon the TPS covers something with an area of 10m² for 8t of mass being slowed from orbital velocity. That implies dissipating 25gJ per m² if you took a 5 meter cylinder and presented it with the smallest cross section as you are suggesting you have cover an area of 20m² this puts the heat dissipation at about 35gJ per m² needing thicker TPS over about twice the area as the Dragon. I don't have any source on the TPS mass on the Dragon, but I understood that the general trade off between covering a larger area over a smaller one was that for the same given mass re-entering you could have somewhat thinner coverage due to the lower total energy dissipation but a fully proportional decrease because you increased the rate of dissipation more than proportional to the reduction of time elapsed as the deceleration rate increased.  I also note that you had suggested that there was some advantage to this when you wrote about how it was easier to slow down an SSTO and the heating load would be lighter.  In your most recent post you seem to be suggesting that you want the craft to slow down more quickly exposing less area on re-entry. I will note that with the big squat BFR sized thing you have less variation in area whether it surfs down like a cylindrical lifting body with an 1/6th to 1/4 sphere engine shield in a nose high orientation or whether it puts a more traditional butt first heat shield between the re-entry stream and the engines.

What I would suggest is that if you think butt first with a heatshield works better, that will be complicated and more weighty than what I had envisaged as this heatshield has to somehow come out of storage at the sides of the engine area and form around the engines.  I will also note that this method, with its much smaller cross section area, will give the craft a much higher terminal velocity and impact the amount of propellant required for a landing negatively.

[Elmar Moelzer]

Elmar, I consider that the mass of the Dragon 2 at re-entry to be something on the order of 6 to 8t. But it is not well documented. The only places I have seen specs for the Dragon 2 for fuel capacity are in this document on page 2-5.  That puts the fuel mass at 1.6t or 20% of the 8t I am expecting as re-entry mass. 

I calculated 1200 kg (400 gallons of hydrazine) for the 400 gallons of fuel that Dragon will carry for the full propulsive hop! It only needs 300 gallons for the propulsive landing tests that might involve hovering. In any case a rocket stage can do a much riskier hover slam landing than a manned capsule. So you can expect additional savings there. All in all, I would estimate the total fuel weight to be at most 900 kg for the Dragon and I would therefore say that even in a worst case scenario, the second stage would need at most a ton of fuel (and I am aiming high here).

I never specified the landing propulsion system for the reusable Raptor upper stage, but I feel that the 5t of propellant for landing was a reasonable number as it scales linearly with the size of the vehicle presuming relatively similar ISPs.  A deep throttling Raptor-Vac would have much lower ISP than a hypergolic fueled superDraco I presume. I expect either smaller methalox landing engines, or superDracos and hypergolic fuels, but either way 5t seems reasonable. Which scales to 50t when your reentry mass is 10 times.

Raptors wont have to throttle down that far on a SSTO with say 9 engines. They probably still have a better Isp than the Superdracos. I think that 10 tons is the worst case scenario for a ten times as heavy vehicle.

As for TPS, on a Dragon the TPS covers something with an area of 10m² for 8t of mass being slowed from orbital velocity. That implies dissipating 25gJ per m² if you took a 5 meter cylinder and presented it with the smallest cross section as you are suggesting you have cover an area of 20m² this puts the heat dissipation at about 35gJ per m² needing thicker TPS over about twice the area as the Dragon. I don't have any source on the TPS mass on the Dragon, but I understood that the general trade off between covering a larger area over a smaller one was that for the same given mass re-entering you could have somewhat thinner coverage due to the lower total energy dissipation but a fully proportional decrease because you increased the rate of dissipation more than proportional to the reduction of time elapsed as the deceleration rate increased.

I gave you the calculation for the TPS mass earlier. PICA-X has 0.25 grams per cm3.
At 6 cm thickness that was 1.8 tons, even for the 125 m2 that you were suggesting.
I admit that I might be wrong about base first re entry but from all I remember it was always the preferred method for re- entry for VTOL SSTOs that did not have a large cross range as a requirement.
But even if we take your 125 m2 that you suggested, the heat shield mass would still be a lot lower than what you suggested and I am convinced that we will need much less than that even with side re entry as an empty SSTO is much more buoyant than a Dragon capsule. But for the sake of the argument, lets stick with 1.8 tons.

[nadreck]

Elmar, hydrazine has a higher specific gravity than water, 400 gallons = 1600 litres > 1600 kg.  The dragon, and as I pointed out a higher weight to cross section area ratio craft will be coming much faster than a "hop test" or a helicopter drop and will require more retropulsion than a DragonFly not less. I felt I was generous keeping it at 20%. Then we go from an 8t craft to a 25t craft and you only at 11% for braking 900kg to 1t?  It scales linearly there is no savings because of scale. If you need 1600kg to give a craft that masses 8t  X deltaV then a 16t craft needs 4800kg.

Where do you get 10 times as heavy a vehicle? If the dragon is 8t and the BFR sized SSTO is 250t that is 32 times the mass, requiring 32 x 1.6t 

I can't accept 1.8t and 6cm without some further background on PICA, do you (or does anyone) have something to support this. My understanding is that the mass of heatshield on the Dragon is more substantial than that. Also my understanding is that as you spread the heatshield out over a larger relative area decreasing the mass to area that you have to use proportionally more heatshield per unit mass of the vehicle because as the rate of heating the material increases the need for the shielding more than the reduction of heating time reduces the need.

[Elmar Moelzer]

Elmar, hydrazine has a higher specific gravity than water, 400 gallons = 1600 litres > 1600 kg. 

According to google 1 cm3 of hydrazine is 0.795g. Therefore a 1 liter of hydrazine is 0.795 kg. I was generous and made it 0.8kg for my calculation (temperature differences could admittedly make it more dense, so I rounded up).

The dragon, and as I pointed out a higher weight to cross section area ratio craft will be coming much faster than a "hop test" or a helicopter drop and will require more retropulsion than a DragonFly not less.

I would assume that with a drop test from 10,000 feet, it will come in at terminal velocity either way, no?
The full propulsive landing from a 10,000 foot drop test was 300 gallons of fuel.

Where do you get 10 times as heavy a vehicle? If the dragon is 8t and the BFR sized SSTO is 250t that is 32 times the mass, requiring 32 x 1.6t

You mentioned 10 times as heavy in your earlier post. I even quoted it. That was what I responded to.

I can't accept 1.8t and 6cm without some further background on PICA, do you (or does anyone) have something to support this. My understanding is that the mass of heatshield on the Dragon is more substantial than that.

I gave you the mass earlier. It is 0.25g/ cm3 from every information that I could find. There was an article that claimed that each of the heat shield tiles of the Dragon capsule was only 2 pounds (there are about 50 that I counted). But that seemed too low, even to me. So I went with the higher number which is 0.25g/cm3.
https://linuxacademy.com/blog/space/comparing-heat-shields-mars-science-lab-vs-spacex-dragon/
I cant find the article quoting the 0.25g/cm3 anymore but this one here puts it at 0.27g/cm3. Close enough, I think:
http://136.142.82.187/eng12/history/spring2013/pdf/3131.pdf
Allegedly Space-X has since then further improved the material (it is now called "version 3", if I remember correctly). So it is plausible that it would be as low as 0.25g/cm3, maybe even lower. Link to article mentioning 3rd version of heat shield technology:
http://www.fastcocreate.com/3031641/inside-the-dragon-with-elon-musk

Hope that sets things right now.

[nadreck]

Elmar, hydrazine has a higher specific gravity than water, 400 gallons = 1600 litres > 1600 kg. 

According to google 1 cm3 of hydrazine is 0.795g. Therefore a 1 liter of hydrazine is 0.795 kg. I was generous and made it 0.8kg for my calculation (temperature differences could admittedly make it more dense, so I rounded up).

Hmm your internet has a different value than mine (which comes in at 1.02gm per cm³ see my google foo).

The dragon, and as I pointed out a higher weight to cross section area ratio craft will be coming much faster than a "hop test" or a helicopter drop and will require more retropulsion than a DragonFly not less.

I would assume that with a drop test from 10,000 feet, it will come in at terminal velocity either way, no?
The full propulsive landing from a 10,000 foot drop test was 300 gallons of fuel.

As it is descending into thicker and thicker air and decelerating at the same time it will always be going faster than terminal velocity for the altitude it is at, so that when braking starts it is over terminal velocity for that altitude. I doubt it will get down to terminal velocity for 10,000 feet even, and a helicopter drop or hop to 10,000 feet will have it at terminal velocity somewhere below 7,000 feet.

Where do you get 10 times as heavy a vehicle? If the dragon is 8t and the BFR sized SSTO is 250t that is 32 times the mass, requiring 32 x 1.6t

You mentioned 10 times as heavy in your earlier post. I even quoted it. That was what I responded to.

That was 10 times as heavy as my example of a Raptor based reusable upper stage which I said had 5t of propellant reserve. The dragon is just under 1/3^rd that mass.

I can't accept 1.8t and 6cm without some further background on PICA, do you (or does anyone) have something to support this. My understanding is that the mass of heatshield on the Dragon is more substantial than that.

I gave you the mass earlier. It is 0.25g/ cm3 from every information that I could find. There was an article that claimed that each of the heat shield tiles of the Dragon capsule was only 2 pounds (there are about 50 that I counted). But that seemed too low, even to me. So I went with the higher number which is 0.25g/cm3.
https://linuxacademy.com/blog/space/comparing-heat-shields-mars-science-lab-vs-spacex-dragon/
I cant find the article quoting the 0.25g/cm3 anymore but this one here puts it at 0.27g/cm3. Close enough, I think:
http://136.142.82.187/eng12/history/spring2013/pdf/3131.pdf
Allegedly Space-X has since then further improved the material (it is now called "version 3", if I remember correctly). So it is plausible that it would be as low as 0.25g/cm3, maybe even lower. Link to article mentioning 3rd version of heat shield technology:
http://www.fastcocreate.com/3031641/inside-the-dragon-with-elon-musk

Hope that sets things right now.

So the first article you linked me to reports the PicaX heat shield as 8 CM thick, if I use .27gm per cm³ then I get a mass of 216kg for a disk shaped heatshield that is 3.7m in diameter but in actual fact it is curved and has a slightly larger area. Since the area I was suggesting on the proposed reusable raptor upper stage was 125m² which is 12.5 times the area of the dragon heatshield 216kg * 12.5 is 2700kg.  This is still a much more significant mass than you are claiming.

[Elmar Moelzer]

Hmm your internet has a different value than mine (which comes in at 1.02gm per cm³ see my google foo).

http://www.endmemo.com/cconvert/lbsgalus.php

As it is descending into thicker and thicker air and decelerating at the same time it will always be going faster than terminal velocity for the altitude it is at, so that when braking starts it is over terminal velocity for that altitude. I doubt it will get down to terminal velocity for 10,000 feet even, and a helicopter drop or hop to 10,000 feet will have it at terminal velocity somewhere below 7,000 feet.

I doubt that SpaceX would bother with this sort of propulsive landing test if it did not resemble a real life situation relatively closely.
Either way, you can go and bend the numbers as much as you want, I don't think that you will get even close to the 5 tons that you projected for the fuel requirements.

That was 10 times as heavy as my example of a Raptor based reusable upper stage which I said had 5t of propellant reserve. The dragon is just under 1/3^rd that mass.

You gave a dry weight for your upper stage of 10 tons. The dragon has an empty weight of almost 7 tons if I am not mistaken. So we are at 2/3rd not 1/3rd. Add some cargo, etc and you are probably closer to 8.

So the first article you linked me to reports the PicaX heat shield as 8 CM thick, if I use .27gm per cm³ then I get a mass of 216kg for a disk shaped heatshield that is 3.7m in diameter but in actual fact it is curved and has a slightly larger area. Since the area I was suggesting on the proposed reusable raptor upper stage was 125m² which is 12.5 times the area of the dragon heatshield 216kg * 12.5 is 2700kg.  This is still a much more significant mass than you are claiming.

That first article also gave 2 pounds per tile. There are no more than 50 tiles at the bottom of the Dragon capsule. I therefore rather believe the other articles that mention 6 cm.
A returning SSTO vehicle would experience less heating during re- entry than a Dragon capsule because it is A LOT less dense.
Either way, even 2700kg is a lot less than the 5 tons you projected earlier.
And IMHO all of these are worst case scenarios, here.

[Robotbeat]

Heatshield mass for something like PICA-X is likely more proportional to entry mass than it is to area. If it's proportional to area, that would have you reenter the stage straight down to minimize heatshield area, which insults the aerospace intuition and goes against many other entry vehicle designs.

A friend of mine who worked at NASA on thermal protection system (and related materials) development told me that a rule of thumb for an achievable number for an RLV is that TPS mass is 10% of your entry mass.

[nadreck]

Fine Elmar, I gave you, and anyone else who cared to read, my logic, math and sources. Believe what you want. I believe that reusable SSTOs on chemical rocket ISP's on Earth are just not worth it financially when compared to TSTO reusables. My logic dictates that you get the same payload to orbit for less dollars no matter how you approach it with TSTO reuse than SSTO reuse.

[Elmar Moelzer]

Heatshield mass for something like PICA-X is likely more proportional to entry mass than it is to area. If it's proportional to area, that would have you reenter the stage straight down to minimize heatshield area, which insults the aerospace intuition and goes against many other entry vehicle designs.

A friend of mine who worked at NASA on thermal protection system (and related materials) development told me that a rule of thumb for an achievable number for an RLV is that TPS mass is 10% of your entry mass.

Which would put TPS masses much more in line with my predictions, even assuming that the newer versions of PICA-X developed by SpaceX have not brought any improvement over that number.

[Elmar Moelzer]

Fine Elmar, I gave you, and anyone else who cared to read, my logic, math and sources. Believe what you want. I believe that reusable SSTOs on chemical rocket ISP's on Earth are just not worth it financially when compared to TSTO reusables. My logic dictates that you get the same payload to orbit for less dollars no matter how you approach it with TSTO reuse than SSTO reuse.

nadreck, don't be annoyed by us defending our positions, please. This is all meant in good sport and we can all (me, you, everyone who reads this thread) learn something from this exchange. I certainly did learn a few things. Nothing wrong with that, hmm? I mean, you made good and fair points. I hope, I did so as well.
Whether SSTO RLVs make more sense than TSTOs probably depends on the market more than the technology. My point was that you can have an SSTO RLV with a meaningful payload. Whether it is economic depends on the market and that is something that is much more difficult than the engineering details, we discussed earlier (and could be almost put into the realm of believe).
So, please continue your posts. I liked our discussion a lot.

[Robotbeat]

Fine Elmar, I gave you, and anyone else who cared to read, my logic, math and sources. Believe what you want. I believe that reusable SSTOs on chemical rocket ISP's on Earth are just not worth it financially when compared to TSTO reusables. My logic dictates that you get the same payload to orbit for less dollars no matter how you approach it with TSTO reuse than SSTO reuse.

I agree with you, although I do think we should try to develop an SSTO RLV anyway, since it'd be more convenient for some applications (though not bulk lift).

One big reason I like TSTO particularly with a RTLS first stage is that you can theoretically relaunch the first stage (with another reusable upper stage on it) before the initial upper stage even completes a single orbit. That's 90% of your rocket that you can reuse in less than an hour (!). And you can also afford healthy margins for the reusable upper stage (compared to a SSTO RLV), allowing fast turnaround there as well.

If you can get the performance high enough and the reliability high enough and the operations automated and streamlined enough, you could use such a capability to drive the cost to orbit down to a small multiple of the propellant cost. That's like down near $10/kg in LEO, if you use methane!

[nadreck]

nadreck, don't be annoyed by us defending our positions, please.

Elmar, I am sorry if you felt I was being curt or was annoyed with you. I enjoyed the exchange, when people challenge me where I do my logic I need to flesh it out and put some of it down in characters, spread sheet formula and pixels, but I think there is nothing further to be gained with continued debate on this.

I will point out one other thing though I said that a Raptor US was 10t in expendable mode, I pointed out that to make it reusable it needed 15t more of fuel, TPS, landing gear, and engine protection support for engine TPS. Even if I concede 2.3t of TPS to you the Raptor Upper Stage it still weighs in at 22.7t in reuseable mode.  While all of this discussion is the modern equivalent of back of the envelope rocket science, which I have practised since I was about 9 years old, the intrinsic issue to me is that until you get north of 400 to 500 seconds ISP at sea level with water density of propellant and scale up from there for lower density, TSTO is really cheaper than SSTO with all the technologies that can be brought to bear, especially when you are developing these technologies step wise as SpaceX is doing. Give me water density 2000 second ISP at high thrust (say water or some other reaction mass heated by a compact fusion source) and I will spec out Star Trek like shuttle craft. But until then, using chemical, we need to think more complex to be economically simple.

[Elmar Moelzer]

I would love to see better engines as well (and I am myself very into nuclear fusion, to put it mildly).
That said, Gary Hudson will tell you that lighter structures are more important than engine performance for SSTOs.

[rsdavis9]

since this thread seems to be about reentry and heat shields...

Has anybody done the calculations to see if the falcon 9 s1 had a heat shield in the inter stage and reentered inter stage first and then did a flip would it be worth it. I.E. no reentry burn but extra mass of heat shield.

Problems:
1. doing a flip while in the atmosphere.
2. Designing a heat shield that is inside the inter stage at first but deploys out enough to protect the stage.

[karanfildavut]

I really don't understand the fascination with SSTO architecture. If you think about it, you're lugging a whole lot of unnecessary dry mass up to whatever orbital+ speed that is desired, resulting in a significant PMF hit even in expendable mode. There's a reason why almost every modern rocket architecture uses multiple stages to orbit. You want to lose the extra mass for your fuel container as soon as that fuel is depleted.

In fact, I don't think SSTO with chemical rockets offers any major advantages over TSTO. Worse PMF, problems returning stage from orbital velocity, increased complexity due to additional shielding necessary for higher orbital return velocity, worse ISP due to lack of engine optimization, the list goes on. Tsiolkovsky says that your PMF increases the more stages there are, i.e. with TSTO you get better performance for a given rocket size, any size. From Falcon 1 to BFR.

For those who dream of Star Wars shuttles, the "spacecraft" were never created by people with a good understanding of orbital mechanics. There is absolutely no reason to move to SSTO with chemical rockets. Now if you have some futuristic technology such as a mass driver or orbital tether, then sure, by all means, design your shuttle craft around that. But until then, the first stage is here to stay, and second stage recovery ops will be infinitely more important towards building out space architecture quickly than designing an oversize rocket with worse performance metrics than anything flying today.

[rsdavis9]

The only appeal I get is the theoretical one vehicle simplicity like an airplane that just needs to land and be refueled.  None of this staging and 2 vehicles that need to land separately then need to be rejoined. At some point maybe through material advances it could be possible.

[sevenperforce]

Given that most commercial payloads require GTO capability, it is hard to see a business case for SSTO. You are straining to get any payload at all to orbit. Of course, if you have a space station with refueling depots and space tugs, that's a different story...but we don't.

Now I can absolutely see a case for PSTO, parallel stage to orbit. That's essentially what the Shuttle was, except that it had two parallel "stage" drops instead of just one. Three, if you count the SRBs as two stages. If executed properly, PSTO has all the advantages of TSTO with most of the simplicity of SSTO.

If you took a Shuttle, made it a bit smaller (say one SSME instead of three), replaced the thermal tiles with something a bit simpler, put an internal tank where the payload bay was, and replaced the SRBs and external tank with a single kerolox booster carrying a hydrolox tank for crossfeed to the orbiter, you would have a fantastic launch system.

[rsdavis9]

As somebody said (musk) you just build a bigger rocket so you have the margins for easy reusability.

[guckyfan]

As somebody said (musk) you just build a bigger rocket so you have the margins for easy reusability.

If the margin is near zero, a bigger rocket does not help much. The margin will remain near zero. Unless you introduce another advantage, like Skylon does, SSTO just does not look promising, even with aerospike or similar improvements.

[The Amazing Catstronaut]

SSTO doesn't work for Earth. Any hypothetical SpaceX SSTO would be launching from a BEO gravity well. Technically BFS is a SSTO for this reason. The Lunar Module ascent stage was an SSTO.


So yes, SpaceX will make SSTOs. But finagling them to work for Earth is irrelevant until we have exotic propulsion systems.