Hypothetical SpaceX SSTO

[Elmar Moelzer]

A stage has much increased surface area, a fair chunk of which will be in the direct path of the airstream, as opposed to a capsule, where quite a lot of it is in the lee of the airstream.

Check the skin temperature maps of Apollo. Peak temperatures were a very small patch on the edge between main heat shield and the lee cone IE the most forward leading edge part of the structure. Outside of this patch temperatures dropped fast.

I've sometimes wondered if you could move that heat to leeward areas is there enough leeward area that you could dump the heat and keep the surface at a reasonable temperature.

If so you could have a semi passive heat pipe TPS.

But, a stage is also very buoyant compared to a capsule and from what I understand will most likely decelerate higher up in the atmosphere. So to the best of my understanding, heat loads could be less than for a capsule and not more. Plus, you could do some active cooling where you turn the engines into the air stream and use residual fuel to cool the engines. Maybe even do some active braking. Am I missing something?

[Zach Swena]

If so you could have a semi passive heat pipe TPS.

Heat pipes would be to heavy.  Risidual fuel or firing the engines should help a lot though.  It can push the shock wave and heat producing high shear air layers out away from the craft.

A SSTO rocket would probably make sense for a rocket the size of BFR.  It has a huge enough payload that it could take a huge payload hit and still lift small type sats.  This only works though once they figure out second stage reuse, and if the extra TPS and other stuff required for orbital re entry work on the BFR first stage also.  This would simply be a way to use the same rocket for both large payloads with a second stage and and small payloads as SSTO.  I don't think it will ever make sense for anything but the smallest payloads flying due to the ultra thin to negative margins involved.  Obviously, negative margins are a no go.

[Norm38]

From a sci-if standpoint I'm always looking for a good, feasible SSTO architecture. Something that lets a small group of characters move around freely without needing a massive infrastructure to get back to orbit.
Even if it only carried a crew of say 5-10, and did nothing more than make low orbit, I can get a lot of use out of that. Especially if the same craft can lift a lot of cargo when and where a 1st stage booster is available.

[Elmar Moelzer]

A SSTO rocket would probably make sense for a rocket the size of BFR.  It has a huge enough payload that it could take a huge payload hit and still lift small type sats.

IIRC, someone did the math based on the information we know and calculated that an SSTO with raptor engines and in the size category of the BFR first stage would have a payload in the tens of tons to LEO. GTO is obviously a different story. I think the biggest advantage of an SSTO is that you don't have to perform an RTLS burn with all the issues involved with that. It also would probably allow for more rapid reuse, since you do not have to restack it with a second stage (unless you need to do a high energy orbit, which would still need one, I guess).

[nadreck]

A SSTO rocket would probably make sense for a rocket the size of BFR.  It has a huge enough payload that it could take a huge payload hit and still lift small type sats.

IIRC, someone did the math based on the information we know and calculated that an SSTO with raptor engines and in the size category of the BFR first stage would have a payload in the tens of tons to LEO. GTO is obviously a different story. I think the biggest advantage of an SSTO is that you don't have to perform an RTLS burn with all the issues involved with that. It also would probably allow for more rapid reuse, since you do not have to restack it with a second stage (unless you need to do a high energy orbit, which would still need one, I guess).

Getting to orbit is feasible but returning it for reuse is not. If we presume that we have any payload left after accounting for TPS, then what we have done is equip something 10 times the size of a reusable S2 with TPS so added 10 times that mass and manufacturing cost while reducing the payload by a significant amount.  If we have a given stage that could do SSTO without a 2nd stage and has a reusable 2nd stage there was no reason to build reusability from orbit into the first stage, so the development effort and manufacturing cost of making the first stage reusable have to be considered as part of the cost when you think of justifying using the first stage for SSTO (or you are not reusing it and throwing away a whole first stage you could have recovered when launching a far higher payload with a 2nd stage).

[bioelectromechanic]

Looks like if it were possible to do 1/10 - 1/20 of the two stage payload mass to LEO while being reusable, a BFR with 100-200 tonnes to LEO would be able to lift payloads in the F9 range.

Now the first problem is reuse from orbital velocities, possibly a SSTO is not even possible within these constraints.
Second problem is you're going to need a spacecraft to transfer people or equipment.
Third is you're going to need to bring a Mars a ascent/decent stage (>4.1km/s dV) with a heat shield.

So now the whole thing starts to looks impractical, because you need to have what's essentially a reusable second stage / space tug anyway.

This seems to suggest the Mars infrastructure is probably structurally similar to current one - first stage, reusable second stage / tug/ ascent vehicle / MCT, a crew module with escape capabilities.

[Zach Swena]

Basically, the infrastructure I see being useful would be a single vehicle that could function either as a first stage for an ultra heavy lift rocket with boost back RTLS or barge landing profiles, or a SSTO vehicle for F9 or smaller class missions.  This would mean more dry weight then required for the two stage launches, but the same is true of using a RTLS capable first stage for expendable missions like Spacex has been doing.  It matters less when both options include reuse.

The logical stepping stone between the current architecture and a craft that can be used either as a first stage, or SSTO vehicle is the development of a reusable second stage.  Only then will they know how feasible it would be to combine the capabilities for a reusable SSTO capable craft.

[bioelectromechanic]

A stage which is sometimes a SSTO and sometimes not has to be orbital-reusable, yet probably offers no advantage over a TSTO stack with a multiple payload adapter.

Say a BFR SSTO is 10 tonnes to LEO.
Let's look at some common cases:
Case 1: customer wants 10 tonnes to GTO.
Case 2: customer wants 15 tonnes to LEO.
Case 3: customer wants 4 payloads, each 2.5 tonnes, each in a different orbit.
You can't do any of those.

Now say a BFR TSTO is 100 tonnes to LEO with the second stage being Mars capable.
You can accommodate multiple large payloads with very high orbital flexebility.

Take into account that Elon (others too) expect most future business to not come from traditional payloads (dare say proven).

Sadly it looks like a BFR SSTO would be an over designed system that can only serve a fraction of expected launches.

[Zach Swena]

Sadly it looks like a BFR SSTO would be an over designed system that can only serve a fraction of expected launches.

I agree, yet this still seems like the most useful SSTO configuration, especially so if the reentry TPS on the first stage was removable.  I think this highlights why we don't have any SSTO designs being seriously considered these days.

[sevenperforce]

I think the primary use nowadays for an SSTO would be as a crew ferry to LEO.

But Challenger taught us that you need a launch abort system, and you need your crew to be on top of the rocket rather than slung alongside. Columbia taught us that your abort system needs to be capable of re-entry. Dragon V2 teaches us that a launch abort system can double as a propulsive landing system. Falcon 9R teaches us that landing a tall orbital-class rocket on its tail, while possible, has a tipover probability far too high to be acceptable for manned landings.

So unless you have a gliding landing, then you need to come down vertically...in a horizontal attitude. One reason this is nice is that in normal launches, your crew capsule's abort/landing engines can double as the front thrusters for a horizontal-attitude vertical landing. Then you only need something on the back end to lower it to the ground.

Under what circumstances would you have perpendicular thrusters on the back end of your rocket?

One possibility: if you had partial air-augmentation shrouds with monopropellant ejector jets to induce and compress airflow during launch, several of these could be reserved for propulsive landing.

[Rabidpanda]

Falcon 9R teaches us that landing a tall orbital-class rocket on its tail, while possible, has a tipover probability far too high to be acceptable for manned landings.

How can you possibility make this assertion without more data?

[Mader Levap]

Falcon 9R teaches us that landing a tall orbital-class rocket on its tail, while possible, has a tipover probability far too high to be acceptable for manned landings.

Capsule lands completely seperately and independently from rest of rocket. It has also completely different construction. They are pretty much as different as possible, while still having propulsive landings.

[sevenperforce]

Falcon 9R teaches us that landing a tall orbital-class rocket on its tail, while possible, has a tipover probability far too high to be acceptable for manned landings.

Capsule lands completely seperately and independently from rest of rocket. It has also completely different construction. They are pretty much as different as possible, while still having propulsive landings.

But that is by definition multiple stages, no?

If the LES thrusters can be used for propulsive landing....

[QuantumG]

I think the more sensible conclusion is that Falcon 9R is a prototype and it'd take a lot of effort to not get better.

[RobLynn]

Attempting to put 1st stages into orbit is a sensible use for it at end of safe life because as scrap on ground it has value little above the materials cost or as a collectable, probably ~$10 per kg, but in orbit those same mostly metallic materials could (eventually) be worth $1000/kg or more, so perhaps $20million?.  Remove the legs and grid fins of course.

[Jim]

Attempting to put 1st stages into orbit is a sensible use for it at end of safe life because as scrap on ground it has value little above the materials cost or as a collectable, probably ~$10 per kg, but in orbit those same mostly metallic materials could (eventually) be worth $1000/kg or more, so perhaps $20million?.  Remove the legs and grid fins of course.

Not really.  Falcon 9 won't still be around when there is such an infrastructure in place that can make use of the hardware.

Also, still better to put an upperstage on the vehicle and put something really useful in orbit. 
Additionally, booster stages don't make good spacecraft.

[john smith 19]

It could be that what makes sense for ULA and Arianespace isn't the case for SpaceX. Because SpaceX is vertically integrated, it might be easier for them to reuse their first stage and continue to keep their production line open.

Exactly. If you are a "rockets are LEGO elements" company like OrbitalATK that assembles things and each stage is dissimilar, if you start reusing S1 a lot, you (or your S1 vendor) have a vacant S1 line, and repurposing it for, say, lunar landers, or in space tugs, is a lot harder.  Your S2 is from a different vendor so you can't repurpose the S1 line to make them...

The (theoretical) aerospace model is that is the suppliers problem.  The user is decoupled from the supplier.

IRL such stages have been highly coupled between the suppliers and the customers.

Whether it was dumb luck forced on them due to limited resources, or shrewd thinking (I think the latter but I'm biased), SpaceX does not have this problem. S2 is made on the SAME line as S1... same tankage, a lot of the same internal fixtures, etc, just shorter. yes, it's different, but reconfiguring the line to make 3x ... and then 5X... and then 10X (as reuse fraction goes up)  S2 as you do S1 isn't nearly as hard.

True.  This is a cheaper way to make a rocket. Industry SOP says it's it gives a less "optimal" system but the approach produced the Titan II which worked pretty well. OTOH it also produced the Delta IV. This suggests the devils in the details. 

(and this is what kind of bugs me about the talk of a Raptor upper stage for F9... all of a sudden you're eroding a lot of commonality. ESPECIALLY if you go to a different tank size like so many people here like)

People forget to this you have to re-plumb every launch pad that supports launches.

SX have shown they have a distinct preference for commonality.  I suspect they now have quite good data on the value of that commonality.

If so you could have a semi passive heat pipe TPS.

Heat pipes would be to heavy.  Risidual fuel or firing the engines should help a lot though.  It can push the shock wave and heat producing high shear air layers out away from the craft.

TBH I thinking more in terms of a capsule, although they were tested for the Shuttle leading edge. The temperature map for Apollo showed the most intense heating on the skin was in fact localised to a  very small patch on the leading edge.  Max Faget famously kept a piece of the Mylar protecting the ablative layer on the rear cone of an Apollo capsule during transport and recovered from it after it landed. With a inconel skin if you could halve the high temperature areas temperature (and spread the rest out a bit, still keeping the temperature high enough to give high emittance). With a heat pipe rated at 30Kw/sq cm of cross sectional area and 100w/cm for the rating of the Apollo heat shield that's roughly 46 sq inches/pipe.

[Elmar Moelzer]

Getting to orbit is feasible but returning it for reuse is not.

Not what I remember.
You lose payload for RTLS of the first stage as well. It is not free and bringing a second stage back is something SpaceX has not even attempted yet. It will cost payload as well. You do not have to do an RTLS burn with an SSTO. Yes, you lose some weight to TPS, but the TPS weight is not that high. PICA is very lightweight and there are other ways to reduce the TPS weight. You could do active cooling with water and base first re-entry in a biconic launch vehicle. Also, as I mentioned before, it is my understanding that a very buoyant (almost empty) launch vehicle should decelerate much further up in the atmosphere and thus would be subjected to less heating than a more compact capsule.

[nadreck]

Getting to orbit is feasible but returning it for reuse is not.

Not what I remember.
You lose payload for RTLS of the first stage as well. It is not free and bringing a second stage back is something SpaceX has not even attempted yet. It will cost payload as well. You do not have to do an RTLS burn with an SSTO. Yes, you lose some weight to TPS, but the TPS weight is not that high. PICA is very lightweight and there are other ways to reduce the TPS weight. You could do active cooling with water and base first re-entry in a biconic launch vehicle. Also, as I mentioned before, it is my understanding that a very buoyant (almost empty) launch vehicle should decelerate much further up in the atmosphere and thus would be subjected to less heating than a more compact capsule.

I am not sure if you misunderstood me, what I was saying was that building an SSTO is practical when you both scale up the size and use a higher performance engine (and it is presumed that the Raptor will be higher performance overall even if it loses a little in the thrust to weight ratio). It is the fact that you have to create something much more expensive and potentially beyond the envelop of what is possible without going to a negative payload number.  The TPS system needs to deal with atmospheric entry of an SSTO must handle dissipating the energy of far more mass than just a reusable 2nd stage so the hit on payload which is already a much smaller fraction of the dry weight of an SSTO vs a TSTO is much more pronounced.

So lets compare 3 rough cases:

FH class TSTO with Kerolox first stage and reuseable raptor upper stage.
FH class SSTO with raptor engines
BFR class SSTO

So the TSTO case has a GLOW of ~1600t for a 55t to LEO payload with booster reuse (2 side cores RTLS, centre core  ASDS which gives the Raptor upper stage 2,500 m/s towards orbital velocity with a dry weight 10t, 385 second ISP, and initial propellant load of 210t).  If we presume that the upper stage is structurally a cylinder with a 5 meter diameter it will be 13 meters long without the engine to have enough volume (250m³) to contain the propellant. Including a disk shaped common bulkhead the area of material needed to make the structure of the tank would be about 285m². If we presume that we need to re-enter the atmosphere with 5t of propellant for maneuvering and landing, that we need 2t of landing gear, 3t of structure and mechanicals to supprt the TPS around the engine and that 5t of TPS would adequately protect the 125m² of surface of half the cylinder and engine shield for re-entry.  So we have a re-entry mass of 25t (15t came from our payload capability to allow for reuse) and that 25t represents 780gigajoules of energy that has to be dealt with on re-entry or 6.25gJ per m².

Now if we imagine a similar take off mass of SSTO that is a 5 meter diameter you can squeeze 1500t of methalox propellant in a cylinder 92 meters long.  The structure of this cylinder with the common bulkhead would have an area of 1520m² or 5.33 times the structure of the Raptor upper stage.  While the engines have to scale up more than 8 times they are a smaller proportion of structural weight than the tank walls. So I will scale the dry weight to 55t. Now the vacuum ISP achievable with a Raptor that works efficiently at sea level is estimated at 363 and the sea level ISP is estimated at 321 a higher proportion of the propellant is burned at lower altitude but if we generously estimate an average ISP of 350 you are going to need GLOW to be limited to 15 times dry weight to make LEO. With 1500t of Methalox and 55t of structure (just under 2 times the mass of an F9 core) you could have 45t of payload in an expendable SSTO of FH class. The cross section area of the cylinder reentering with the engine protection area would work out to 750m² and even if we could keep the landing gear and engine protection structure to the same mass as the much smaller raptor upper stage we still have to scale up the propellant to at least 15t  and the mass re-entering will be at least 75t requiring 2.4tJoules of energy absorption on re-entry or about 3.125gJ per m². TPS might be able to be thinner but has to cover 6 times the area so if TPS fits into 15t (which I doubt) making the FH sized SSTO go from 45t of payload as an expendable to 10t if the same landing gear and protective structure for the engines on re-entry works. If you need even 75% of the thickness of TPS on the surface as compared to the Raptor upper stage re-entry you are down to 2.5t of payload to LEO on the SSTO.

Finally what if you upscale from 1600t GLOW and 5m diameter to 6,400t GLOW and 15m diameter? You increase the structural walls of the tanks from 1520m² to 2400m² because it is a more efficient volumetric shape, but you have increased the required thrust by 4 times so I suggest that you have to at least double the dry weight of the disposable SSTO to 110t but the good news is as an expendable SSTO it should be able to put 290t in LEO. However when it comes to re-entering if we presume a re-entry mass of 250t we need to include 50t of propellant in that, and with the landing mass 10 times the mass of the re-entering raptor upper stage upgrading the engine protection structure and landing gear by 4 times from the raptor is in order and that gets us 20t more there. The cross section area re-entering is half the cylinder area and the engine protection at about 1100m² and with the 250t mass coming back the energy dissipated at re-entry is up to 7.8tJoule or 7.1gJ per m² so the TPS has to be thicker over nearly ten times the area as the Raptor upper stage (so 50t TPS).  At this assumption then we have 150t left for payload in a reusable BFR sized SSTO.

But what if we underestimated the TPS by 50% in the Raptor upper stage? Then doubling the TPS requirement in the Raptor upper stage took away 12.5% (5t) more from the payload leaving us at 35t. On the FH class 5 meter SSTO we would have lost more than the total payload mass. And on the BFR class SSTO you would have lost 33% of your payload (another 50t). If it turned out that the Raptor upper stage TPS requirement was 20t then there would be NO payload on the BFR class SSTO but there still would have been 25t payload on the reusable Raptor upper stage. Note that there is virtually no payload on the BFR class reusable SSTO to GTO with the most generous assumption above.

[Elmar Moelzer]

I am not sure if you misunderstood me, what I was saying was that building an SSTO is practical when you both scale up the size and use a higher performance engine (and it is presumed that the Raptor will be higher performance overall even if it loses a little in the thrust to weight ratio). It is the fact that you have to create something much more expensive and potentially beyond the envelop of what is possible without going to a negative payload number. 

I was always assuming a higher performance engine like the raptor. The Phoenix studies by Garry Hudson demonstrated that relatively small SSTOs would be possible with a meaningful payload to LEO, even with 1980ies technology. Generally, structural mass is more important than engine performance. We have had the ability to build light weight tank structures since the 1960ies. SpaceX has demonstrated extremely high mass fractions. The Merlin 1D has a thrust to weight ratio of 180 and it is very likely that the Raptor engine which is also optimized for thrust to weight to have an only slightly lower T/W.

The TPS system needs to deal with atmospheric entry of an SSTO must handle dissipating the energy of far more mass than just a reusable 2nd stage so the hit on payload which is already a much smaller fraction of the dry weight of an SSTO vs a TSTO is much more pronounced.

The SSTO is mostly empty tank mass, which is much more buoyant than a more compact second stage or a Dragon capsule. This means lower TPS requirements. With base first re- entry and a biconic shape, you could reduce the need of TPS on sidewalls. It is possible that active cooling with water could further reduce the weight of the TPS (Gary Hudson did some studies on this).

FH class TSTO with Kerolox first stage and reuseable raptor upper stage.
FH class SSTO with raptor engines
BFR class SSTO

So the TSTO case has a GLOW of ~1600t for a 55t to LEO payload with booster reuse (2 side cores RTLS, centre core  ASDS which gives the Raptor upper stage 2,500 m/s towards orbital velocity with a dry weight 10t, 385 second ISP, and initial propellant load of 210t).  If we presume that the upper stage is structurally a cylinder with a 5 meter diameter it will be 13 meters long without the engine to have enough volume (250m³) to contain the propellant. Including a disk shaped common bulkhead the area of material needed to make the structure of the tank would be about 285m². If we presume that we need to re-enter the atmosphere with 5t of propellant for maneuvering and landing, that we need 2t of landing gear, 3t of structure and mechanicals to supprt the TPS around the engine and that 5t of TPS would adequately protect the 125m² of surface of half the cylinder and engine shield for re-entry.  So we have a re-entry mass of 25t (15t came from our payload capability to allow for reuse) and that 25t represents 780gigajoules of energy that has to be dealt with on re-entry or 6.25gJ per m².

Missing a few things here.
1. Your reusable upper stage also needs extra engines (superdracos probably) and fuel/tanks for landing. The single Raptor can not throttle low enough for landing. You completely disregarded that here.
2. You are assuming ASDS landing for the center stage, which I still believe to be the exception rather than the rule since ASDS landing will probably always be a pain and less cost effective than RTLS landing. RTLS landing of all 3 cores will reduce the payload to LEO a lot more. An SSTO does not need to spend the fuel for RTLS.
3. 5 tons of propellant seems extremely generous. The Dragon 2 capsule does the same thing, is less buoyant and still does not need 5 tons for powered landing, not even close to that, nor does it need 5 tons of TPS. It has maybe 600 kg of TPS (and could be as low 80 kg, depending on the source, I use). Even by using a worst case calculation and your side entry surface, I am getting 1.8 tons for the weight of the tps (6 cm thick, 0.25g/ cm3, 125 m2) and I think that is way off compared to the Dragon.

The cross section area of the cylinder reentering with the engine protection area would work out to 750m² and even if we could keep the landing gear and engine protection structure to the same mass as the much smaller raptor upper stage we still have to scale up the propellant to at least 15t  and the mass re-entering will be at least 75t requiring 2.4tJoules of energy absorption on re-entry or about 3.125gJ per m².

I disagree about the propellant mass needed. Where does that number come from?

TPS might be able to be thinner but has to cover 6 times the area so if TPS

No, it does not. You are assuming side re- entry. If you assume base first re- entry, the surface area is not going to increase with a longer cylinder.

fits into 15t (which I doubt)

Again, you are way off with your TPS mass.

However when it comes to re-entering if we presume a re-entry mass of 250t we need to include 50t of propellant in that

20% of the dry weight for re- entry and landing propellant seems excessive compared to a much more compact Dragon capsule that uses less for its propulsive landing (and needs a lot of safety margins and cant do a much riskier hoverslam).

, and with the landing mass 10 times the mass of the re-entering raptor upper stage upgrading the engine protection structure and landing gear by 4 times from the raptor is in order and that gets us 20t more there.

This seems excessively pessimistic. You might be able to get away with a much lighter landing gar, if you are doing a biconic LV with a larger base and less height (Dragon has a minimal landing gear).

The cross section area re-entering is half the cylinder area and the engine protection at about 1100m² and with the 250t mass coming back the energy dissipated at re-entry is up to 7.8tJoule or 7.1gJ per m² so the TPS has to be thicker over nearly ten times the area as the Raptor upper stage (so 50t TPS).  At this assumption then we have 150t left for payload in a reusable BFR sized SSTO.

Again, you are assuming a side re- entry instead of a base first re-entry. You are also overestimating the weight of the TPS. If you multiplied the weight of the Dragon 2 by 25, you would not get 50 tons for its TPS, no way!