Hypothetical SpaceX SSTO

[bioelectromechanic]

Speaking hypothetically, is it theoretically possible for MCT to be an SSTO?

There were quotes from Elon that the F9 first stage can reach orbit with a very minor payload.
There also are some estimates that the first stage of F9 can do close to 9km/s dV.
So an SSTO with current F9 technology seems possible, but not practical.

Let's say it is possible to design MCT as an SSTO that can lift a payload in the F9-F9H range, which can also refuel in orbit. So basically you have a single common system that can launch from Earth, land on Mars, and come back.
The plane-like operation seems appealing because of the possible reduced complexity.

Is such a system even theoretically possible?
How large does it need to be to lift a payload in the F9-FH range?

[hkultala]

Speaking hypothetically, is it theoretically possible for MCT to be an SSTO?

There were quotes from Elon that the F9 first stage can reach orbit with a very minor payload.
There also are some estimates that the first stage of F9 can do close to 9km/s dV.
So an SSTO with current F9 technology seems possible, but not practical.

Let's say it is possible to design MCT as an SSTO that can lift a payload in the F9-F9H range, which can also refuel in orbit. So basically you have a single common system that can launch from Earth, land on Mars, and come back.
The plane-like operation seems appealing because of the possible reduced complexity.

Is such a system even theoretically possible?
How large does it need to be to lift a payload in the F9-FH range?

SSTO is different than SSTO+landing.

And in addition to the mass fraction needed for the delta-v is SSTO, there are many other inefficiencies:

Engines made for takeoff from earth are very unoptimal for doing anything else than taking of from earth;
There is just way too much thrust (unnecessary heavy engines, and maybe also too high acceleration with fuel tanks almost empty), and the nozzle is optimized for wrong wrong air pressure for most part of the flight.

No, it does not make any sense for MCT to have the T/W ratio and sea level nozzles required to take off from earth. On the other hand, same nozzle and same T/W ratio work for both liftoff from mars and upper stage from earth.

[Robotbeat]

Yes it's possible. Essentially Delta Clipper. (paging HMXHMX...)

I'd argue that given what SpaceX are attempting, this would be less efficient. But yes it could be done.

Musk has said it'd basically be a two-stage rocket, with a BFR and BFS. Or think about BFS as the payload doing a lot of the work to get to orbit.

[nadreck]

Yes it's possible. Essentially Delta Clipper. (paging HMXHMX...)

I'd argue that given what SpaceX are attempting, this would be less efficient. But yes it could be done.

Musk has said it'd basically be a two-stage rocket, with a BFR and BFS. Or think about BFS as the payload doing a lot of the work to get to orbit.

Getting to orbit is a possibility, but a functional BFR first stage does not need the TPS of the BFS/MCT/Tanker stage and a perfectly good first stage would be being disposed of by putting it in orbit as it could not be recovered. The limits seem to be about 1000m/s of atmospheric entry speed. Currently the F9 first stage can be travelling anywhere between 2000m/s and 6000m/s when the 2nd stage separates, however at 2000m/s it can expend another 3200-3500 m/s and return to the launch site and still have fuel to slow down and land, if instead it is moving at 4000m/s it probably only has enough ΔV to slow it down to a safe entry speed and land down range. At 6000m/s it has no propellant left at all. In all of these cases except the last one, it needs to perform retropulsion from whatever speed it is going to 1000m/s or less as it hits the thicker atmosphere

[M_Puckett]

Speaking hypothetically, is it theoretically possible for MCT to be an SSTO?

There were quotes from Elon that the F9 first stage can reach orbit with a very minor payload.
There also are some estimates that the first stage of F9 can do close to 9km/s dV.
So an SSTO with current F9 technology seems possible, but not practical.

Let's say it is possible to design MCT as an SSTO that can lift a payload in the F9-F9H range, which can also refuel in orbit. So basically you have a single common system that can launch from Earth, land on Mars, and come back.
The plane-like operation seems appealing because of the possible reduced complexity.

Is such a system even theoretically possible?
How large does it need to be to lift a payload in the F9-FH range?

Stretch it (the BFR first stage) and make it a tanker to fill a depot.

That is your most structurally efficient way to carry a meaningful payload.  Plus if ups your launch cadence using the same facilities as the MCT.

The residual prop minus what you need to deorbit and land are your payload.

[sanman]

Didn't Musk comment that F9 booster would be capable of SSTO if it didn't have to carry an upper stage on top.

[john smith 19]

Speaking hypothetically, is it theoretically possible for MCT to be an SSTO?

There were quotes from Elon that the F9 first stage can reach orbit with a very minor payload.
There also are some estimates that the first stage of F9 can do close to 9km/s dV.
So an SSTO with current F9 technology seems possible, but not practical.

Let's say it is possible to design MCT as an SSTO that can lift a payload in the F9-F9H range, which can also refuel in orbit. So basically you have a single common system that can launch from Earth, land on Mars, and come back.
The plane-like operation seems appealing because of the possible reduced complexity.

Is such a system even theoretically possible?
How large does it need to be to lift a payload in the F9-FH range?

Your missing a few points.

Vertical take off Single Stage to orbit has been possible for decades.

The issues have been a)You loose 2/3 to 3/4 of the TSTO payload b)The stage (not a capsule) has to handle the full -7900ms delta v to land again c) This is the Mars Colony/i] transporter, which IIRC Musk has said  is in the 100 passenger (IE small airliner) category.

For the SSTO question b) is probably the killer. The amount of KE you need to dump rises at the square of the velocity. What's X at 2000 m/s is 16x that at about 8000 m/s in round numbers.

the notion the engine bells would survive that temperature becomes very doubtful. Likewise the amount of TPS you're going to ablate away on every single landing is likely to be considerable. 

All multiplied together there is no just not a good reason to do this.

[nadreck]

Didn't Musk comment that F9 booster would be capable of SSTO if it didn't have to carry an upper stage on top.

and didn't cary a payload

[Robotbeat]

Didn't Musk comment that F9 booster would be capable of SSTO if it didn't have to carry an upper stage on top.

and didn't cary a payload

No, it definitely could carry a payload, but it'd be significantly smaller than F9 can do with an upper. Plus it'd be expended.

[HMXHMX]

Didn't Musk comment that F9 booster would be capable of SSTO if it didn't have to carry an upper stage on top.

and didn't cary a payload

The payload would be between 0 and 10K lbm depending on assumptions.

[nadreck]

Didn't Musk comment that F9 booster would be capable of SSTO if it didn't have to carry an upper stage on top.

and didn't carry a payload

The payload would be between 0 and 10K lbm depending on assumptions.

The assumptions I am making bring it to zero *wink*

Seriously to get between 0 and 10k lbm you are throwing away the stage as compared to putting up more than 4 times the maximum there if recover the stage and only throw away a 2nd stage. That 0 to 10k lbm you mention would have to include all margins for any other purpose than getting to low orbit as well as a hypothetical payload.

[Exastro]

Suppose you were free to tweak the stage design to increase the payload? 

Without the US, interstage, payload, or fairing, the T/W should be a fair bit higher.  So you might gain by stretching the stage.  Or replacing the center engine with an MVac, which might be started only after the stage reaches near-vacuum. 

I realized it's unlikely that the stage would be able either to land or carry enough payload to compete with a normal F9 for cost/kg.  But suppose the stage itself had some value -- refuel it with several FH flights and it could push a big payload to TLI, TMI, or elsewhere, say.  Of course you'd need to tweak it some more to be able to do that. 

[HMXHMX]

Didn't Musk comment that F9 booster would be capable of SSTO if it didn't have to carry an upper stage on top.

and didn't carry a payload

The payload would be between 0 and 10K lbm depending on assumptions.

The assumptions I am making bring it to zero *wink*

Seriously to get between 0 and 10k lbm you are throwing away the stage as compared to putting up more than 4 times the maximum there if recover the stage and only throw away a 2nd stage. That 0 to 10k lbm you mention would have to include all margins for any other purpose than getting to low orbit as well as a hypothetical payload.

I wasn't making an economic case.  Only pointing out what the payload of an expendable SSTO based on  the SpaceX technology might be, in contravention of the comment that it couldn't carry any payload.

[llanitedave]

Why?  What purpose would it serve?

[bioelectromechanic]

It seemed counter inutitive for a fully reusable design to have 2 sets of landing legs, heat shields, avionics, engines, tanks, etc.

Interesting that it actually might be possible, even though only to LEO.

[su27k]

Why?  What purpose would it serve?

So that they can retire F9/FH after MCT came online.

A lot of people didn't read OP's question carefully, this is not about using F9 1st stage or any 1st stage to SSTO, the question is specifically asked about MCT, i.e. the 2nd stage of BFR, which should already have the TPS for reentry built-in.

[john smith 19]

It seemed counter inutitive for a fully reusable design to have 2 sets of landing legs, heat shields, avionics, engines, tanks, etc.

Interesting that it actually might be possible, even though only to LEO.

What's counter intuitive?

That you need 2 stages or that you need landing legs, avionics and heat shields for both?

[Hotblack Desiato]

From what I have gathered, MCT is supposed to be an SSTO.

Just not on Earth, where it needs to have a second stage, and in an environment where it is not supposed to carry a large payload (on Mars, MCT has to bring cargo down to the surface, not up into orbit).

[Elmar Moelzer]

b)The stage (not a capsule) has to handle the full -7900ms delta v to land again

Not sure what you mean with this point. Can you please elaborate on that?

[john smith 19]

b)The stage (not a capsule) has to handle the full -7900ms delta v to land again

Not sure what you mean with this point. Can you please elaborate on that?

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.

[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!

[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.

[hkultala]

SSTO doesn't work for Earth.

.. because you say so, without any calculations to prove your point?

SSTLEO does work fine and could be done on todays or even tens of years old technology, but it just not has been economical to make, because
1) without air-breathing engines the payload fraction is so bad that the rocket is much bigger and more expensive for same payload, and suitable air-breathing engines have not yet been developed(though SABRE development is underway)
2) Not very many commercial launches are to LEO


I think you are messing up SSTO and reusable SSTO. Reusable SSTO is much harder than SSTO, but even that can be done, but with a very high development cost.

[Impaler]

I think that is exactly what Cat meant, it won't work for SpaceX Earth launches because they care about little things like having their launch vehicle be cost effective.

BTW some discussion earlier about heat-shielding on entry, my understanding is that you always save mass by making the entry object more compact and the surface area of shielding smaller even if it must be thicker.  The reason is that your dissipated heat dose not scale linearly with vehicle mass divided by deceleration energy, a more compact area experiences a shorter higher temperature pulse but less total heat actually directed into the vehicle due to the detached shock-layer effects which put more of the energy into the atmosphere as radiation.

The g-forces that humans can survive is the limiting factor in a manned capsule, if we could arbitrarily compress the crew and make them invulnerable to g-forces then we would design the vehicle like a sample return capsule with very small area of thick shielding.

So when returning a rocket stage you would always orient it vertically to present the smallest area, also it's the only direction that the tank structure has any strength and can survive in.  In addition the engine will have to point forward because it dominates the remaining mass and trying to go nose first would be unstable with the vehicle wanting to flip over.

The heat-shield for a 2nd stage is either going to be some kind of clam-shell inside the inter-stage which opens after separation but before ignition and then closes again for entry (maybe opening again to act as landing legs), or just a quantity of propellant for SRP to create shock layer standoff from the engine.

[rsdavis9]

So when returning a rocket stage you would always orient it vertically to present the smallest area, also it's the only direction that the tank structure has any strength and can survive in.  In addition the engine will have to point forward because it dominates the remaining mass and trying to go nose first would be unstable with the vehicle wanting to flip over.

so what was the CG of the space shuttle? It had 3 engines in the back. Have all reentry vehicles had the CG near the front? It sounds like a hard balancing act to have the CG in the back but if it could be done it would simplify the heat shield placement.

[Impaler]

Center of Gravity on the Shuttle during re-entry was DANGEROUSLY far to the rear and presented a constant danger that the vehicle would tumble during re-entry because the shape was unstable, which is what ultimately happened to Columbia.  This is one of the reasons Buran was considered superior to the Shuttle, not having the main engine in the orbiter made the entry hugely safer because center of mass was basically in the center of the vehicle.

The X-33 project was abandoned in large part because the center of gravity of the empty re-entry vehicle could not be moved far enough forward due to the mass of the rear-engine, the vehicle would have again been dangerously unstable on re-entry.

Even the Skylon concept only works because the engines are placed at the center of the vehicle on stub-wings, the earlier HTOL concept that it evolved from had the same fatal rear-engine flaw. 

Basically every attempt to re-enter a tubular object with the center of mass in the rear has been an deemed an engineering failure, if even regular air-space engineers have learned their lesson then Musk will likely avoid that mistake when he tries to re-enter a 2nd stage.

[Elmar Moelzer]

Center of Gravity on the Shuttle during re-entry was DANGEROUSLY far to the rear and presented a constant danger that the vehicle would tumble during re-entry because the shape was unstable, which is what ultimately happened to Columbia.  This is one of the reasons Buran was considered superior to the Shuttle, not having the main engine in the orbiter made the entry hugely safer because center of mass was basically in the center of the vehicle.

The X-33 project was abandoned in large part because the center of gravity of the empty re-entry vehicle could not be moved far enough forward due to the mass of the rear-engine, the vehicle would have again been dangerously unstable on re-entry.

Even the Skylon concept only works because the engines are placed at the center of the vehicle on stub-wings, the earlier HTOL concept that it evolved from had the same fatal rear-engine flaw. 

Basically every attempt to re-enter a tubular object with the center of mass in the rear has been an deemed an engineering failure, if even regular air-space engineers have learned their lesson then Musk will likely avoid that mistake when he tries to re-enter a 2nd stage.

Not if you reenter bottom first.

[ChrisWilson68]

Center of Gravity on the Shuttle during re-entry was DANGEROUSLY far to the rear and presented a constant danger that the vehicle would tumble during re-entry because the shape was unstable, which is what ultimately happened to Columbia.  This is one of the reasons Buran was considered superior to the Shuttle, not having the main engine in the orbiter made the entry hugely safer because center of mass was basically in the center of the vehicle.

The shuttle had various issues, and having the CoG toward the rear was not the greatest of them.  It was an entirely manageable issue.

Engineering is about trade-offs.  Every choice has positives and negatives.  Stating some negatives and then just entirely dismissing a choice is not justified.

The X-33 project was abandoned in large part because the center of gravity of the empty re-entry vehicle could not be moved far enough forward due to the mass of the rear-engine, the vehicle would have again been dangerously unstable on re-entry.

Most people think there were several reasons X-33/VentureStar was abandoned, and CoG was again not the biggest of them.  The hydrogen tanks with complex shapes and linear aerospike engine were looking like they could not be produced with low enough mass to come even close to the necessary mass fraction.

[The Amazing Catstronaut]

.. because you say so, without any calculations to prove your point?

The calculations are old, readily available, and the same calculations done by one more amateur isn't going to shift anyone's reasoning.

SSTLEO does work fine and could be done on todays or even tens of years old technology, but it just not has been economical to make, because
1) without air-breathing engines the payload fraction is so bad that the rocket is much bigger and more expensive for same payload, and suitable air-breathing engines have not yet been developed(though SABRE development is underway)
2) Not very many commercial launches are to LEO

SSTO is fine if you want to lift the mass of the SSTO into LEO. If you want to carry payload to anywhere other than a few rigid trajectories, you need a massive SSTO. An expendable TSTO is objectively more sound on a bang-for-buck basis than an equivalently voluminous expendable SSTO using the same engines and fuel types. Fiddling with the structural mass doesn't bias performance towards the SSTO either, as both architectures benefit from structural improvements.

Air breathing engines make your mass problem worse because you end up carting them with you. Combined cycle engines of the SABRE ilk have terrible thrust-to-weight margins when compared with equivalent conventional rockets. There's a vast margin of the ascent phase where your wings are more of a hinderance than a help, and once you're up in orbit all that mass dedicated to atmospheric flight is going to be carted around with you to wherever in the solar system you happen to be going to.

I can see a winged SSTO being used for orbital space tourism, delivery of passengers to an LEO space station, max, and there might be an economic case for one then. Hypothetical SSTOs work for LEO servicing, would doubtless be aesthetically sexy, but beyond that, what market is there? If we're trying to force SSTOs to fit mission parameters that vehicles with staging events have been able to conduct for decades - that's not progress. That's retrogression.

SSTOs seem like a great way to spend a lot of dollars for a vehicle tailored to very specific mission parameters.

So yes, my statement that SSTO doesn't work on Earth, was an economic statement. I think it's a sexy sci-fi idea we keep toying with because it appeals to that part of us that believes that LVs need to be aircraft-esque to become aircraft-practical, but it's not a logical connection. SpaceX is not going to build one unless the tech paradigm massively evolves. 

[QuantumG]

SSTO is fine if you want to lift the mass of the SSTO into LEO. If you want to carry payload to anywhere other than a few rigid trajectories, you need a massive SSTO. An expendable TSTO is objectively more sound on a bang-for-buck basis than an equivalently voluminous expendable SSTO using the same engines and fuel types. Fiddling with the structural mass doesn't bias performance towards the SSTO either, as both architectures benefit from structural improvements.

I don't know where you get this from. An SSTO carries a payload of about 2% of GLOW to orbit. An expendable TSTO carries a payload of about 4% of GLOW to orbit. So far we "know" that a reusable first stage eats about 2% of the TSTO performance. Minimum gauge analysis tells us that the lower GLOW for a fully reusable system would be with the SSTO. Reusability economics suggests that the fully reusable SSTO would beat the fully reusable TSTO. There's still a long way to go before we know enough about reusable systems to make any more definitive statements.

[Impaler]

Center of Gravity on the Shuttle during re-entry was DANGEROUSLY far to the rear and presented a constant danger that the vehicle would tumble during re-entry because the shape was unstable, which is what ultimately happened to Columbia.  This is one of the reasons Buran was considered superior to the Shuttle, not having the main engine in the orbiter made the entry hugely safer because center of mass was basically in the center of the vehicle.

The X-33 project was abandoned in large part because the center of gravity of the empty re-entry vehicle could not be moved far enough forward due to the mass of the rear-engine, the vehicle would have again been dangerously unstable on re-entry.

Even the Skylon concept only works because the engines are placed at the center of the vehicle on stub-wings, the earlier HTOL concept that it evolved from had the same fatal rear-engine flaw. 

Basically every attempt to re-enter a tubular object with the center of mass in the rear has been an deemed an engineering failure, if even regular air-space engineers have learned their lesson then Musk will likely avoid that mistake when he tries to re-enter a 2nd stage.

Not if you reenter bottom first.

I think your confusing my response with someone else's, I'm saying that bottom first aka the heavy engine and Center of Mass being first is the only stable and viable configuration, so your agreeing with me but your wording clearly implies disagreement, can you clarify.

[Elmar Moelzer]

I think your confusing my response with someone else's, I'm saying that bottom first aka the heavy engine and Center of Mass being first is the only stable and viable configuration, so your agreeing with me but your wording clearly implies disagreement, can you clarify.

I apologize. Was a misunderstanding on my side.

[ChrisWilson68]

SSTO is fine if you want to lift the mass of the SSTO into LEO. If you want to carry payload to anywhere other than a few rigid trajectories, you need a massive SSTO. An expendable TSTO is objectively more sound on a bang-for-buck basis than an equivalently voluminous expendable SSTO using the same engines and fuel types. Fiddling with the structural mass doesn't bias performance towards the SSTO either, as both architectures benefit from structural improvements.

I don't know where you get this from. An SSTO carries a payload of about 2% of GLOW to orbit. An expendable TSTO carries a payload of about 4% of GLOW to orbit. So far we "know" that a reusable first stage eats about 2% of the TSTO performance. Minimum gauge analysis tells us that the lower GLOW for a fully reusable system would be with the SSTO. Reusability economics suggests that the fully reusable SSTO would beat the fully reusable TSTO. There's still a long way to go before we know enough about reusable systems to make any more definitive statements.

What SSTO gets a payload of 2% of GLOW to orbit?  Is that an expendable SSTO or reusable SSTO?  And is it projections for a proposed SSTO or an actual proven number?

[Impaler]

And most importantly at what ISP do you expect that kind of payload.  SpaceX doesn't use Hydro-Lox which has long been considered the only possible propellant that could give SSTO.  For the expected 380s ISP of Raptor the propellant mass alone would need to be 93% for a 9800 m/s launch.  That leaves 7% for all vehicle mass, payload and landing propellant.  A fantastically light rocket would be 3.3% dry-mass, landing propellant is likely to be more then the remaining ~4% because the estimated retained propellant for F9 to do a down-range landing are around 5%, so their is really nothing left for payload.

[QuantumG]

SpaceX doesn't use Hydro-Lox which has long been considered the only possible propellant that could give SSTO.

Whitehead. PDF

[Elmar Moelzer]

And most importantly at what ISP do you expect that kind of payload.  SpaceX doesn't use Hydro-Lox which has long been considered the only possible propellant that could give SSTO.  For the expected 380s ISP of Raptor the propellant mass alone would need to be 93% for a 9800 m/s launch.  That leaves 7% for all vehicle mass, payload and landing propellant.  A fantastically light rocket would be 3.3% dry-mass, landing propellant is likely to be more then the remaining ~4% because the estimated retained propellant for F9 to do a down-range landing are around 5%, so their is really nothing left for payload.

That is obviously not correct. The Falcon9 rocket has an amazing mass fraction to orbit and it only uses RP1.
Isp is not everything.

[sevenperforce]

I'm inclined to agree. For SSTO, impulse density and TWR vastly outweigh specific impulse. If your impulse density is high enough, it really doesn't matter how low your ISP is as long as your TWR can keep pace.

[Impaler]

And most importantly at what ISP do you expect that kind of payload.  SpaceX doesn't use Hydro-Lox which has long been considered the only possible propellant that could give SSTO.  For the expected 380s ISP of Raptor the propellant mass alone would need to be 93% for a 9800 m/s launch.  That leaves 7% for all vehicle mass, payload and landing propellant.  A fantastically light rocket would be 3.3% dry-mass, landing propellant is likely to be more then the remaining ~4% because the estimated retained propellant for F9 to do a down-range landing are around 5%, so their is really nothing left for payload.

That is obviously not correct. The Falcon9 rocket has an amazing mass fraction to orbit and it only uses RP1.
Isp is not everything.

I applied the F9 Heavy side boosters anticipated 30:1 wet:dry ratio, as I was giving SpaceX the most generous numbers possible least someone accuse me of sandbagging them (alas to no avail), but it still doesn't work.  What is obviously not correct is to try to present personal incredulity as an argument in an engineering discussion.

Try http://www.silverbirdastronautics.com/LVperform.html and see if you can make a SSTO rocket with a positive payload.

SpaceX doesn't use Hydro-Lox which has long been considered the only possible propellant that could give SSTO.

Whitehead. PDF

I wasn't trying to imply that I agreed with the consensus around Hydrogen fuel superiority, simply that it was the dominant view and you have to factor in the increased difficulty that results from lower ISP.  The paper in question was clearly against the mainstream thought of it's day.

The study is comparing the portion of final orbit reaching mass that is composed of known propellant density driven components.  While the fraction of orbital mass that was the dry hardware of the propulsion system is comparable, the low ISP based vehicle needs to have a higher gross take off weight to do the same job.

The biggest factor being ignored in a SpaceX SSTO vehicle is that SpaceX will certainly be using a retro-propulsive landing method which means the vehicles DeltaV needs are well in excess of just going to orbit.  It's more like going to GTO.