An SSTO as "God and Robert Heinlein intended".

[T.Neo]

I only have this one for winged TSTOs.

Figures, so far all the documents of that sort I've found have had to do with reusable spaceplanes and not more usual launch vehicle stages. :dry:

I think that generally, people really like reusable spaceplanes...

 

[Urwumpe]

Figures, so far all the documents of that sort I've found have had to do with reusable spaceplanes and not more usual launch vehicle stages. :dry:

I think that generally, people really like reusable spaceplanes...

Possible. I think the problem is rather historic: Most active simpler launchers are military technology and many details about their project management like such estimates are either military secret or company secrets.

Space planes are simply of more academic nature.

This one goes into the tank department, which is fairly simple, but good enough to show that the tank structural mass isn't that large.

https://www.princeton.edu/~stengel/MAE342Lecture4.pdf

And this one here goes to teach why the proper design of thrust structures is so important for rockets:

http://er.jsc.nasa.gov/seh/main_EDC_Spacecraft_Structures.pdf

 

[RGClark]

Wasn't this argued not to be the case? I tried to give examples in this post showing that engines could make up a fourth to a third the total stage dry mass, and I didn't include the thrust structure. And the figures here for the Atlas D show a booster mass of over 50% of the vehicle's dry mass! That isn't even including the sustainer engine and its thrust structure.
Also, does anyone have a source describing the mass breakdown percentages of rocket stages? I've been trying to find this information for months! :shrug:

For the Atlas, since that is a booster section that drops away, that makes the comparison more difficult. Also the Atlas had an unusual method to save weight on the propellant tanks. It used what are called "balloon tanks". It's based on the idea that inflating a structure can increase its compressional and bending strength. Consider an inflated tire or basketball for example. Then this fact was used to reduce the thickness of the walls of the Atlas tanks. The reduced strength of the walls was made up for by the increased strength due to the internal pressure.
In fact for the Atlas its walls are so thin the structure can not stand on its own when empty. When in storage its propellant tanks are filled with nitrogen. Otherwise it would collapse under its own weight:

Atlas.
http://www.astronautix.com/lvs/atlas.htm

As for a weight breakdown for the propellant tanks of a proposed SSTO, see the the estimates at the end here:

Mass Estimating Relations
• Review of iterative design approach
• Mass Estimating Relations (MERs)
• Sample vehicle design analysis
http://spacecraft.ssl.umd.edu/academics/483F09/483F09L13.mass_est/483F09L13.MER.pdf

Note the propellant tanks for this hydrogen fueled SSTO weigh more than the engines. This problem of tank weight is especially true for hydrogen fueled vehicles because the tanks for hydrogen have to be so large, therefore heavy, because of hydrogen's low density.


Bob Clark

 

[T.Neo]

For the Atlas, since that is a booster section that drops away, that makes the comparison more difficult.

How so? It would increase the complexity of things, sure, but surely the disconnect/separation system can't mass over a quarter the mass of the entire vehicle!

Also the Atlas had an unusual method to save weight on the propellant tanks. It used what are called "balloon tanks". It's based on the idea that inflating a structure can increase its compressional and bending strength. Consider an inflated tire or basketball for example. Then this fact was used to reduce the thickness of the walls of the Atlas tanks. The reduced strength of the walls was made up for by the increased strength due to the internal pressure.
In fact for the Atlas its walls are so thin the structure can not stand on its own when empty. When in storage its propellant tanks are filled with nitrogen. Otherwise it would collapse under its own weight:

It is not a bad thing, it served Atlas well for decades and continues to be used with the Centaur stage. And the SpaceX Falcon tanks use a similar concept (but are beefed up to be self-supporting, they only require pressurisation to withstand flight loads).

The fact that Centaur needs to be pressurised to support itself is not problematic, with modern technology the process is automated and alerts are sent to engineers via email.

Considering your crusade for the reduction in tank mass, I would have thought you'd have a more positive impression of pressure supported structures.

Note the propellant tanks for this hydrogen fueled SSTO weigh more than the engines. This problem of weight is especially true for hydrogen fueled vehicles because the tanks for hydrogen have to be so heavy because of its low density.

I have not read that pdf very closely (looks very interesting though- thank you), but this page gives the mass of the STS ET LH2 tank as 29 000 "pounds" (approximately 13 200 kilograms) empty. The mass of three SSME is 9531 kilograms (going by figures on Wikipedia) and that is not including thrust structure.

And the DIRECT J-130 core stage dry mass figure, with three SSMEs removed, is still roughly two times more massive than the shuttle ET (the tank is supposedly reinforced to withstand the loads of a heavy upper stage, but still thrust structure mass has to be in there somewhere).

But think of it like this:

Let's say engines make up 25% of a launch vehicle (as shown by the few examples that I gave a page or two back). Let's say that the thrust structure masses the same as the engines, and is also 25% of the dry mass. Let's allocate a further 5% to "other stuff", and leave the remaining 45% for the propellant tank mass.

If we reduce the tank mass by a whole 50%, the vehicle mass is only reduced by 22.5%.

The less the tanks mass as a percentage of dry weight, the less impact lightening them will have. If tanks are only 30% of the vehicle mass, halving the mass of the tanks will only shave 15% of the dry mass off.

Ok, so maybe that is still advantageous enough to warrant its existence. But it isn't near "halve the vehicle dry mass" or "double mass ratio". And tanks can't just be lightened at a whim, they still have to contain propellant and support payload...

 

[RGClark]

This one goes into the tank department, which is fairly simple, but good enough to show that the tank structural mass isn't that large.

https://www.princeton.edu/~stengel/MAE342Lecture4.pdf

Actually, it doesn't say what the tank mass is for a given propellant load. But it does show on p.9 that you can increase your structural strength by having a pressurized tank, as proven by the early versions of the Atlas rocket that used "balloon tanks".


Bob Clark

 

[Urwumpe]

Actually, it doesn't say what the tank mass is for a given propellant load. But it does show on p.9 that you can increase your structural strength by having a pressurized tank, as proven by the early versions of the Atlas rocket that used "balloon tanks".

yes, but if you did not skip the first 4 pages, you will also be able to put the structural engineering concepts of the various stresses, elasticity and moduli into perspective... even more if you had been involved in structural mechanics projects before.

 

[RGClark]

I have not read that pdf very closely (looks very interesting though- thank you), but this page gives the mass of the STS ET LH2 tank as 29 000 "pounds" (approximately 13 200 kilograms) empty. The mass of three SSME is 9531 kilograms (going by figures on Wikipedia) and that is not including thrust structure.

When I am referring to "propellant tanks" I'm referring to both fuel and oxidizer tanks. It is notable as well that the hydrogen tank is larger and heavier than the oxygen tank on the shuttle ET because of hydrogen's low density even though the oxygen mass is 6 times that of the hydrogen mass.
The mass of the entire external tank also is over 26,000 kg while the the 3 main engines mass about 10,000 kg. But this comparison is misleading because the engines are smaller than they would need to be for a SSTO because most of the launch thrust is provided by the solid rocket boosters.


Bob Clark

---------- Post added at 03:57 PM ---------- Previous post was at 03:50 PM ----------

yes, but if you did not skip the first 4 pages, you will also be able to put the structural engineering concepts of the various stresses, elasticity and moduli into perspective... even more if you had been involved in structural mechanics projects before.

The lecture report by Stengel is very informative on structural loads but it doesn't give the tank mass to propellant mass relationship.


Bob Clark

 

[Urwumpe]

The lecture report by Stengel is very informative on structural loads but it doesn't give the tank mass to propellant mass relationship.

If you know the structural loads, that the tank has to carry as primary structure of the rocket, you can calculate the mass based on material and structural technology.

It sure doesn't say the exact value, since even two slightly different alloys of aluminum are already resulting in completely different masses there.

(But if you for example use the values for calculating the stress of the ET hydrogen tank and then put it into relation to the mass, it would be a very precise estimate of the tank mass for a different tank using the same technology)

 

[RGClark]

The fact that Centaur needs to be pressurised to support itself is not problematic, with modern technology the process is automated and alerts are sent to engineers via email.
Considering your crusade for the reduction in tank mass, I would have thought you'd have a more positive impression of pressure supported structures.

True. I am of two completing minds on the tank pressurization question. On the one hand the X-33 case showed that composite, non-cylindrical tanks under the need to make them lightweight burst when pressurized.
On the other hand the Atlas case showed that you can reduce the structural weight of the tanks by making use of the strength obtained by pressurized structures.
Still I would like to see what would be the mass of tanks if they did not have to be pressurized.


Bob Clark

 

[T.Neo]

The mass of the entire external tank also is over 26,000 kg while the the 3 main engines mass about 10,000 kg.

Sorry, I misread and thought you meant "hydrogen tank of the hydrogen fueled SSTO".

The STS ET is more massive than the engines on STS alone, but there is also the thrust structure to deal with. The mass of the engines and their thrust structure together could be closer to ~25 tons.

But this comparison is misleading because the engines are smaller than they would need to be for a SSTO because most of the launch thrust is provided by the solid rocket boosters.

Good point. You'd likely need six or more SSMEs to get adequate liftoff thrust, which would mean more mass in engines... and a heavier thrust structure as well.

In that case, there is really nothing to gain by taking the extra SSMEs to orbit, it would be better to stage them off in an Atlas-esque manner.

But it would still be a very ungainly launch vehicle. Probably similar performance to the EELV heavies, but more costly.

 

[RGClark]

Space Travel: The Path to Human Immortality?
Space exploration might just be the key to human beings surviving mass genocide, ecocide or omnicide.
July 24, 2009
On December 31st, 1999, National Public Radio interviewed the futurist and science fiction genius Arthur C. Clarke. Since the author had forecast so many of the 20th Century's most fundamental developments, the NPR correspondent asked Clarke if anything had happened in the preceding 100 years that he never could have anticipated. "Yes, absolutely," Clarke replied, without a moment's hesitation. "The one thing I never would have expected is that, after centuries of wonder and imagination and aspiration, we would have gone to the moon ... and then stopped."
http://www.alternet.org/news/141518/space_travel:_the_path_to_human_immortality/

I remember thinking when I first saw 2001 as a teenager and could appreciate it more, I thought it was way too optimistic. We could never have huge rotating space stations and passenger flights to orbit and Moon bases and nuclear-powered interplanetary ships by then.
That's what I thought and probably most people familiar with the space program thought that. And I think I recall Clarke saying once that the year 2001 was selected as more a rhetorical, artistic flourish rather than being a prediction, 2001 being the year of the turn of the millennium (no, it was NOT in the year 2000.)
However, I've now come to the conclusion those could indeed have been possible by 2001. I don't mean the alien monolith or the intelligent computer, but the spaceflights shown in the film.
It all comes down to SSTO's. As I argued above these could have led and WILL lead to the price to orbit coming down to the $100 per kilo range. The required lightweight stages existed since the 60's and 70's for kerosene with the Atlas and Delta stages, and for hydrogen with the Saturn V upper stages. And the high efficiency engines from sea level to vacuum have existed since the 70's with the NK-33 for kerosene, and with the SSME for hydrogen.
The kerosene SSTO's could be smaller and cheaper and would make possible small orbital craft in the price range of business jets, at a few tens of millions of dollars. These would be able to carry a few number of passengers/crew, say of the size of the Dragon capsule. But in analogy with history of aircraft these would soon be followed by large passenger craft.
However, the NK-33 was of Russian design, while the required lightweight stages were of American design. But the 70's was the time of detente, with the Apollo-Soyuz mission. With both sides realizing that collaboration would lead to routine passenger spaceflight, it is conceivable that they could have come together to make possible commercial spaceflight.
There is also the fact that for the hydrogen fueled SSTO's, the Americans had both the required lightweight stages and high efficiency engines, though these SSTO's would have been larger and more expensive. So it would have been advantageous for the Russians to share their engine if the American's shared their lightweight stages.
For the space station, many have soured on the idea because of the ISS with the huge cost overruns. But Bigelow is planning on "space hotels" derived from NASA's Transhab concept. These provide large living space at lightweight. At $100 per kilo launch costs we could form large space stations from the Transhabs linked together in modular fashion, financed purely from the tourism interests. Remember the low price to orbit allows many average citizens to pay for the cost to LEO.
The Transhab was developed in the late 90's so it might be questionable that the space station could be built from them by 2001. But remember in the film the space station was in the process of being built. Also, with large numbers of passengers traveling to space it seems likely that inflatable modules would have been thought of earlier to house the large number of tourists who might want a longer stay.
For the extensive Moon base, judging from the Apollo missions it might be thought any flight to the Moon would be hugely expensive. However, Robert Heinlein once said: once you get to LEO you're half way to anywhere in the Solar System. This is due to the delta-V requirements for getting out of the Earth's gravitational compared to reaching escape velocity.
It is important to note then SSTO's have the capability once refueled in orbit to travel to the Moon, land, and return to Earth on that one fuel load. Because of this there would be a large market for passenger service to the Moon as well. So there would be a commercial justification for Bigelow's Transhab motels to also be transported to the Moon.
Initially the propellant for the fuel depots would have to be lofted from Earth. But we recently found there was water in the permanently shadowed craters on the Moon. Use of this for propellant would reduce the cost to make the flights from LEO to the Moon since the delta-V needed to bring the propellant to LEO from the lunar surface is so much less than that needed to bring it from the Earth's surface to LEO.
This lunar derived propellant could also be placed in depots in lunar orbit and at the Lagrange points. This would make easier flights to the asteroids and the planets. The flights to the asteroids would be especially important for commercial purposes because it is estimated even a small sized asteroid could have trillions of dollars worth of valuable minerals. The availability of such resources would make it financially profitable to develop large bases on the Moon for the sake of the propellant.
Another possible resource was recently discovered on the Moon: uranium. Though further analysis showed the surface abundance to be much less than in Earth mines, it may be that there are localized concentrations just as there are on Earth. Indeed this appears to be the case with some heavy metals such as silver and possibly gold that appear to be concentrated in some polar craters on the Moon.
So even if the uranium is not as abundant as in Earth mines, it may be sufficient to be used for nuclear-powered spacecraft. Then we wouldn't have the problem of large amounts of nuclear material being lofted on rockets on Earth. The physics and engineering of nuclear powered rockets have been understood since the 60's. The main impediment has been the opposition to launching large amounts of radioactive material from Earth into orbit above Earth. Then we very well could have had nuclear-powered spacecraft launching from the Moon for interplanetary missions, especially when you consider the financial incentive provided by minerals in the asteroids of the asteroid belt.

Bob Clark

Just saw this article on The Space Review discussing a recently discovered copy of a 1963 TV interview with Arthur C. Clarke:

The perils of spaceflight prediction.
by Jeff Foust
Monday, December 5, 2011
http://thespacereview.com/article/1981/1

In the interview Clarke gives some predictions of the future of space exploration. From the standpoint of the beginnings of human spaceflight, he suggests a manned Mars mission within 25 years, which would have been by 1988, and Moon bases by the end of the 20th century.
This turned out to be too optimistic. But as I argued above, this could indeed have been technically and even financially feasible: if it had been recognized that reusable SSTO's are possible and in fact aren't even really hard, we would have had routine, private spaceflight by the 1970s.
Such wide spread, frequent launches using reusable spacecraft would have cut the costs to space by two orders of magnitude, at least. This would then have made the costs of lunar bases and manned Mars missions well within the affordability range.
The important point is that the required high efficiency engines and lightweight stages for SSTO's already exist and have for decades. All that is required is to marry the two together. An expendable test SSTO could be produced, like, tomorrow. Just this one simple, cheap test would finally make clear the fact that routine spaceflight is already doable.


Bob Clark

 

[T.Neo]

RGClark, each time I read one of your posts on SSTO, it is almost like I want to curl away in a small cave somewhere and cry.

The issue I have here, is what from here to Mu Arae do you base your "SSTOs can cut costs by two orders of magnitude" asssertion on. I really want to see the calculation you performed that reached this conclusion.

An expendable test SSTO could be produced, like, tomorrow.

And it would achieve, like, nothing.

Seriously: such a vehicle, would do so much more for what a TSTO could do for so much less.

It would be by no means cheap, and if a NASA project or similar it would take many years at least.

The required mass-optimised stages are almost nowhere to be seen (balloon tank Atlas and S-II don't exist, Space Shuttle ET would need to be redesigned for the task), except perhaps for the SpaceX Falcon tanks. Marrying high performance engines with tanks built for completely different systems would be a non-trivial task.

And all you would get for it in the end is a launch vehicle too expensive for its payload capability.

Just this one simple, cheap test would finally make clear the fact that routine spaceflight is already doable.

I can guarantee you, that if you stick an SSME on anything, it will not be routine at all.

The point is, you want to avoid high performance engines, high performance structures, stuff like that. They drive costs up.

By staging, you complicate the recovery process... but you also reduce your GLOW, and the technological intensity of the hardware required.

And even then, you need some pretty extraordinary evidence for an extraordinary claim of reducing launch costs by 1-2 orders of magnitude.

 

[RGClark]

RGClark, each time I read one of your posts on SSTO, it is almost like I want to curl away in a small cave somewhere and cry.

The issue I have here, is what from here to Mu Arae do you base your "SSTOs can cut costs by two orders of magnitude" asssertion on. I really want to see the calculation you performed that reached this conclusion.

?????

You ask me this me every time, I tell you every time and you ignore the fact that I told you like a half dozen times before.
It's coming from reusability. The low cost, expendable test version is important because it will show not only that SSTO's are possible, but in fact they can be done at high payload fraction.
Once this is recognized then the more expensive reusable ones can be undertaken since the expendable version showed you had so much payload that you could add reentry/landing systems and still have significant payload left over.


Bob Clark

 

[PhantomCruiser]

My $0.02...

I'm not sure if this has been previosly mentioned in this thread or not. I've tried going back to read it all, and well... It's almost as hard to read as some of the proceedures I have to work through here at a nuclear plant.

I think it's been shown that reusability (in present form of the words) is still somewhat mythological. The SSME are reuseable, but only after a long and lengthy (and expensive) rebuild process to again certify them flight ready.

SpaceX wants Dragon to be reuseable, but there will still be some in-depth non-destructive examination (NDI), along with other tests to recert the Dragon for spaceflight, correct?

 

[T.Neo]

You ask me this me every time, I tell you every time and you ignore the fact that I told you like a half dozen times before.
It's coming from reusability.

Yes, but here is what annoys me about your claims: you treat reusability as if it is some magic phenomena that cuts costs by two orders of magnitude, magically.

This is emphatically not the case, and STS is the concrete evidence of this fact. The expendable parts of STS weren't the problem here, it was rather how the reusable parts needed to be treated.

To have any benefit from reusability, you need to do it correctly. There are many ways you can fail and many ways you could succeed. This is what makes your "the cut of two orders of magnitude comes from reusability" claim even more annoying.

This is why I ask you time, and time, and time again, for the calculation(s) that you performed that led to the "two orders of magnitude" statement. "It comes from reusability" is no calculation, it is just an empty claim. The more extraordinary something is, the more extraordinary evidence it needs behind it.

The low cost, expendable test version is important because it will show not only that SSTO's are possible

The expendable test version would be a total waste of time, better to test out subscale models/construction techniques to verify predictive calculations and estimations of mass.

in fact they can be done at high payload fraction.

This is your claim, and as I understand it, a pretty dubious one.

you had so much payload that you could add reentry/landing systems and still have significant payload left over.

Again, your claim, and one that I have trouble believing. Have you ever calculated the mass of these systems? Conservatively, and not with "modern materials magically cut this mass in half"?

I think it's been shown that reusability (in present form of the words) is still somewhat mythological.

It isn't mythological just because it hasn't been done right before (unless you ask Mike Griffin). The SSME took a high performance, bleeding-edge-of-technology engine and tried to make it reusable. And on top of that, they increased the throttle rating to increase payload.

So there's no wonder the engine required such intense refurbishment. It doesn't necessarily have to correlate with other engines, but it can correlate with high performance ones, which is generally why you don't want to stick to those.

There are other options: for example, you could run the SSME at a 70% throttle setting throughout the flight. This will reduce T/W and ISP (a bit), and make the engine grossly overdesigned for this role, but this could help both with reusability and reliability.

The Common Evolved Cryogenic Engine (CECE, an RL-10 derivative) is listed on the PWR website as having a service life of 10 000 seconds, and 50 in-space starts.

This is enough for perhaps 20 flights, including a two-burn trajectory.

In this thread on the NSF forums, it is said that the RS-68 has a flight rating of 1200 seconds, a maximum burn time certification of 1800 seconds, and has been tested to a maximum of 2400 seconds and 8 starts (unfortunately the link provided as a source no longer seems available). The limiting factor in the RS-68's maximum burn time is most probably the ablative nozzle, and even at a maximum of 1200 seconds, you can fit in four whole 250 second burns.

The limiting factor in actually making the RS-68 reusable is probably a lack of access panels, etc.

So if you have a $10-30 million engine (and RL-10s are relatively cheap), that you could reuse 20 times or more for only a bit of inspection between each flight, you could achieve pretty significant cost savings. Still not down to two orders of magnitude less- for that, you'd probably have to amortise the cost of the hardware over perhaps something like 1000 flights.

I'm not saying that is impossible- maybe an engine could exist someday that could be robust enough and with good enough monitoring systems to withstand and that have a low refurbishment cost, but if you look at the RL-10, it is a pretty low performance engine- in terms of ISP it is quite good, but it has a far lower chamber pressure than say the SSME (640 psi on the RL-10-B2, compared to 3280 psi on the SSME), and makes up for it with a high expansion ratio (the RL-10 is a vacuum engine so this doesn't matter, the flip side is that it cannot be used from takeoff and is thus unsuitable for an SSTO unless it has a thrust augmented nozzle).

RGClark keeps saying that with the highest performance engines around, SSTO can be created and costs can be reduced by two orders of magnitude. But "highest performance" and "cost reduction" just don't jive. You don't want to have to select for the highest performance around, but rather the most reliable, practical system.

The thing is, SSTO only makes economic sense for reusability, and the reason it is attractive for reusability is because it simplifies the recovery and re-launch process. However, if you have to deal with the burden of extra costs from the propulsion side for example, this advantage could quickly diminish and won't be worth it at all.

 

[T.Neo]

SSTO discussion continued on from here in the Falcon 9 Flight 3 updates thread.

The only reason SSTO was even considered is because people thought it would make for a more completely and rapidly reusable launcher. If SpaceX can pull that off with two stages (or 2.75 stages, with Falcon Heavy), SSTO is unnecessary.

It doesn't mean SSTO is 'unecessary'. The existence of fixed-wing aircraft doesn't negate the purpose of helicopters for example.

The advantages to SSTO exist and will exist if and when SpaceX or whoever else achieves successful reusability with a TSTO.

The problem is that these advantages are intrinsically linked to the set of disadvantages that make SSTOs unattractive, which have been discussed at length here and elsewhere.

 

[RGClark]

Altitude compensation on the Falcon 9 first stage and applications.

Copied below is an argument for how two copies of the Falcon 1 combined together using a single Merlin engine with an altitude compensating nozzle could be SSTO. I had forgotten though that the Falcon 1e, which SpaceX wants to move to for small launches anyway, has about twice the propellant load of the Falcon so it could form the SSTO.

Aerospike nozzles.
So how to get the altitude compensation? The aerospike[1],[2] is the most extensively studied method of altitude compensation so to get to this quickly it would be nice if this could be used. However, the aerospike (or plug nozzle for the shortened version) requires a toroidal combustion chamber. This would require significant modification of the engine.

However, one concept for getting an aerospike engine uses multiple small chambers arranged around a central spike. This was the idea for the X-33/VentureStar[3]. It was also used earlier on the Beta SSTO concept of Koelle[4],[5]. This could be conveniently used on the Falcon 9 first stage because of its multiple engines. The idea would be to greatly shorten the nozzles on the Merlins and arrange them around a central spike.

We need an estimate of the mass of the Falcon 9 first stage. This environmental assess- ment report on the SpaceX Grasshopper VTVL test vehicle[6] gives on p. 7 (page 17 according to the numbering as a PDF file) the Falcon 9 first stage kerosene load as 24,900 gallons and 38,900 gallons of LOX. Using a density of .820 gm/cc for kerosene
and 1.14 gm/cc for LOX, this amounts to about 245,000 kg propellant for the first stage.

Falcon 9 with aerospike nozzle becomes SSTO.
SpaceX has said the Falcon 9 first stage has a better than 20 to 1 mass ratio. This would give it a first stage dry mass of 13,000 kg. But this is using the Merlin 1C engine. From the thrust level and thrust/weight ratio[7] for this engine we can estimate its mass as about 650 kg. The Merlin 1D is supposed to be lighter, estimated as 440 kg. Using these brings the dry mass down to 11,000 kg.

The question is what would be the weight with the truncated nozzles and aerospike? A complaint against the aerospike used on the X-33 is that it had a worse T/W ratio than other LH2/LOX engines. However it had been planned for the full VentureStar version to use lightweight ceramics for the aerospike, expected to double the T/W ratio to about 80 to 1, better than the SSME's. With the advances in ceramics necessitated by the research and test flights with the hypersonic vehicles such high temperature lightweight ceramics should be further along now than they were with the VentureStar. For instance, the method of transpiration cooling using ceramics should make rocket engine combustion chambers and nozzles lighter and more reusable[8]. So I'll assume the total engine weight remains the same with the aerospike.

As before I'll use the Merlin Vacuum Isp in calculating the delta-V and take the required delta-V to orbit as 9,150 m/s for kerosene engines. Then we can get 6,000 kg payload:

342*9.8ln(1 + 245/(11 + 6)) = 9,167 m/s.

An SSTO is best utilized as a reusable though. Estimates of the added weight of reentry/landing systems are in the range of 28% [9]. However, with modern materials this probably can be cut to half that. Then the payload will be reduced to about 4.5 metric tons.

"An increase of 10% in Isp corresponds to an increase in 100% in payload."
This example illustrates well the importance of altitude compensation. Using it we are able to increase our engine Isp by 10% or more, to the extent we can achieve an SSTO with significant payload. Note that because of the rocket equation just being able to increase the Isp by 10% is no trivial feat. A rule of thumb among propulsion engineers is that "an increase of 10% in Isp corresponds to an increase in 100% in payload"[10]. I'll illustrate this with a single stage vehicle. A common estimate is that a kerosene-fueled SSTO needs a mass ratio of 20 to carry significant payload. Let's say a Falcon 9 sized rocket had instead high efficiency engines such as the NK-33[11] with an Isp of 331 s. You would need three of these. These have better thrust/weight than the Merlin 1C. So the dry weight is reduced to about 11 mT. Then it could carry a payload of 4.5 mT:

331*9.8ln(1 + 245/(11 + 4.5)) = 9,150 m/s.

With vacuum optimized nozzles high efficiency engines such as the NK-33 can get an Isp at 360+ s. Ten percent higher Isp than 331 s is at 364 s. Using this as the Isp allows a payload of 9.4 mT: 364*9.8ln(1 + 245/(11 + 9.4)) = 9,150 m/s.

Falcon Heavy with aerospike matches phase 1 SLS a hundred times cheaper.
Interestingly an improvement in payload using altitude compensation also applies to staged vehicles. A preliminary calculation showed that with a two-stage vehicle you can increase your payload in the range of 25% to 30%. The improvement can be even better if the vehicle uses parallel staging, perhaps up to 50%. This is because with parallel staging the second stage still has to use the same nozzles as a lower stage because they still fire from the ground.

I'll illustrate this with the Falcon Heavy. SpaceX has said its side boosters will achieve a 30 to 1 mass ratio[12]. Reportedly the lower stages also will be stretched to hold more propellant. A posting on the NASASpaceFlight forum suggests a stretched version of the Falcon 9 will have 480,000 kg gross mass[13]. I'll estimate the stretched Falcon 9 based boosters as having a 435 mT propellant load and 15 mt dry mass. The central core stage has to be stronger since it holds the upper stage as well as the heavy payload. I'll take the dry mass for this lower core stage as 20 mT with the same 435 mT propellant load.

For the upper stage, I'll take the propellant load as 30 mT and the dry mass + fairing as 6 mT. For the vacuum Isp of the Merlin 1D I'll take the announced 310 s, and for the Isp of the upper stage, that of the Merlin Vacuum, 342 s. Then we can lift a payload of 51 mT:

310*9.8ln(1 + 2*435/(2*15 + 435 + 20 + 36 + 51)) + 310*9.8ln(1 + 435/(20 + 36 + 51)) + 342*9.8ln(1 + 30/(6 + 51)) = 9,155 m/s.

Now suppose we use altitude compensation to be able to get 342 s vacuum Isp for all the engines. Then we could lift 70 mT:

342*9.8ln(1 + 2*435/(2*15 + 435 + 20 + 36 + 70)) + 342*9.8ln(1 + 435/(20 + 36 + 70)) + 342*9.8ln(1 + 30/(6 + 70)) = 9,150 m/s

This means we could get a 70 mT launcher from essentially the same vehicle as the Falcon Heavy by using altitude compensation on the engines. SpaceX has said they intend to sell the Falcon Heavy for the range of $80 to $125 million per launch. The modifications to the aerospike nozzle should be relatively low cost compared to designing a whole new engine so the price should still be in this range.

Compare this to the estimates of the costs of the SLS program:

Space Launch System.
"Program costs.
During the joint Senate-NASA presentation in September 2011, it was
stated that the SLS program has a projected development cost of $18
billion through 2017, with $10B for the SLS rocket, $6B for the Orion
Multi-Purpose Crew Vehicle and $2B for upgrades to the launch pad and
other facilities at Kennedy Space Center. An unofficial NASA
document estimates the cost of the program through 2025 will total at
least $41B for four 70 metric ton launches (1 unmanned in 2017, 3
manned starting in 2021). The 130 metric ton version should not be
ready earlier than 2030."
http://en.wikipedia.org/wiki/Space_Launch_System#Program_costs

If that estimate for the total costs of the SLS is correct then that's $10 billion per launch for the interim 70 mT payload vehicle. That's two orders of magnitude higher than the Falcon Heavy with altitude compensation.

Reusable first stage Falcon 9 with aerospike also serves as a reusable booster.
This is for the expendable version of the Falcon stages. However, another benefit of the reusable version is that it could serve as the first stage of the reusable booster program (RBS) of the Air Force[14].

Reusable Falcon 9 with aerospike also as next-generation shuttle.
Another interesting possibility is suggested by the recent report of investigations of bringing back the shuttle as a commercial satellite launcher[15],[16]. My view is that the shuttle orbiter is too heavy for that role, ca. 80 mT in dry mass. This cuts greatly into the payload capacity. According to the reports the investigations also considered building their own shuttle but using the advancements made since the shuttle was designed. Then the reusable Falcon SSTO could be used to launch small payloads or even crew capsules. It could be modeled on the aerodynamic design of the shuttle since the aerodynamics for that are so well studied. Since kerolox has an overall density about 1,000 kg/m^3, the propellant could fit within the 300 m^3 shuttle-sized payload bay[17]with 55 m^3 left over. For a cargo only version we could also use the 75 m^3 sized volume of the crew cabin, for a total of 130 m^3.

The stretched version of the Falcon 9 first stage to be used for the side boosters of the Falcon Heavy are expected by SpaceX to have 30 to 1 mass ratio. The improved mass ratio over the current Falcon 9 first stage is probably coming from the fact that making your rocket larger in general improves your mass ratio, the fact that the new Merlin 1D is lighter, and also the fact the boosters do not have to support the weight of the upper stage and payload of the full vehicle. Using a 435 mT propellant load and a 15 mT dry mass, this could launch 15.3 mT: 342*9.8ln(1 + 435/(15 + 15.3)) = 9,155 m/s. If you take the reentry/landing systems mass with modern materials as 14% of the dry mass this takes up 2.1 mT from the payload, so to 13.2 mT. To scale up the space shuttle design for a 435 mT propellant tank compared to a 245 mT tank, the linear dimensions would only have to be scaled up by 20%.



Bob Clark


1.)Aerospike Engine.
http://www.aerospaceweb.org/design/aerospike/main.shtml

2.)Nozzle Design.
by R.A. O'Leary and J. E. Beck, Spring 1992
http://www.rocketdynetech.com/articles/nozzledesign.htm

3.)Aerospike engine.
[ame="http://en.wikipedia.org/wiki/Aerospike_engine"]Aerospike engine - Wikipedia, the free encyclopedia[/ame]

4.)Beta, A Single Stage Reusable Ballistic Space Shuttle Concept.
Based on a study contract of the German Federal Ministery for Education and Science, Bonn (RFT 1017).
May 1970
Space Division, Messerschmitt-Bolkow-Blohm ( MBB), Munich, Germany.
www.spacefuture.com/archive/beta_a_single_stage_reusable_ballistic_space_shuttle_concept.shtml

5.)A Cost Engineered Launch Vehicle for Space Tourism.
D E Koelle
TCS-TransCostSystems, Ottobrunn, Germany.
IAA-98-IAA.1.5.07
http://www.spacefuture.com/archive/a_cost_engineered_launch_vehicle_for_space_tourism.shtml

6.)Draft Environmental Assessment for Issuing an Experimental Permit to SpaceX
for Operation of the Grasshopper Vehicle at the McGregor Test Site,Texas.
September, 2011
http://www.faa.gov/about/office_org...0110922 SpaceX Grasshopper Draft EA.Final.pdf

7.)Merlin(rocket engine).
http://en.wikipedia.org/wiki/Merlin_(rocket_engine)

8.)Ceramic Materials for Reusable Liquid Fueled Rocket Engine Combustion Devices.
http://ammtiac.alionscience.com/pdf/AMPQ8_1ART06.pdf

9.)Newsgroups: sci.space.policy, sci.space.history, sci.astro, sci.physics
From: Robert Clark <[email protected]>
Date: Tue, 4 Oct 2011 10:04:42 -0700 (PDT)
Subject: Re: Elon Musk's SpaceX to build 'Grasshopper' hover-rocket
http://groups.google.com/group/sci.physics/msg/f3b4c27da1f13027?hl=en

10.)Discovery of New Molecule Could Lead to More Efficient Rocket Fuel.
ScienceDaily (Dec. 22, 2010)
http://www.sciencedaily.com/releases/2010/12/101222071831.htm

11.)NK-33.
http://www.astronautix.com/engines/nk33.htm

12.)SPACEX ANNOUNCES LAUNCH DATE FOR THE WORLD'S MOST POWERFUL ROCKET.
http://www.spacex.com/press.php?page=20110405

13.)Re: Falcon Heavy Master Update Thread.
http://forum.nasaspaceflight.com/index.php?topic=24711.msg723410#msg723410

14.)Doing a 180 - AFRL's Rocket-back Pathfinder.
Posted by Graham Warwick at 4/7/2010 7:52 AM CDT
http://www.aviationweek.com/aw/blog...79a7Post:1553afc7-cc1e-4b5e-9a6c-ced39704d348

15.)Next Gen Shuttle-Capable vehicle interest as secret effort to save orbiters ends.
December 19th, 2011 by Chris Bergin
http://www.nasaspaceflight.com/2011/12/next-gen-shuttle-vehicle-secret-effort-save-orbiters-ends/

16.)Atlantis Journal – Epilogue.
by MLD on Dec.19, 2011, under Commercial Space, Space Exploration, Space Policy, Space Shuttle Program
http://www.marylynnedittmar.com/?p=1303

17.)Space Shuttle orbiter.
Shuttle Orbiter Specifications.
http://en.wikipedia.org/wiki/Space_Shuttle_orbiter#Shuttle_Orbiter_Specifications

>

In post #136 I argued that small, low cost SSTO's are doable
now using lightweight design and high efficiency engines. However,
I was surprised to find after doing the calculation you don't even need
the high efficiency engines to get the SSTO. The low efficiency SpaceX
Merlin engines would be sufficient for example, IF you have altitude compensation.
The impetus for trying the calculation was from a report by SpaceX
that you could get the same performance from a planned heavy lift
first stage using a lower performance Merlin 2 engine compared to the
high performance RS-84 engine. The reason was the lower Isp of the
Merlin was made up for by its lower weight:

SpaceX Propulsion.
http://images.spaceref.com/news/2010/SpaceX_Propulsion.pdf


Now note that the biggest single contributor to the vacuum Isp of an
engine is not the chamber pressure, but the nozzle length. For
example, the Merlin Vacuum raises its vacuum Isp to 342 s from the 304
s Isp of the Merlin 1C by having a longer nozzle, even though the
chamber pressure remains the same, ca. 100 bar.
So I'll redo the calculation for the SSTO using the SpaceX Falcon 1
first stage but using Merlin engines this time. We'll assume that
using altitude compensation we are able to get an engine with the same
vacuum Isp as the Merlin Vacuum but able to launch from ground.
We'll use the soon to be introduced Merlin 1D:


SpaceX Plans To Be Top World Rocket Maker.
Aug 11, 2011
By Guy Norris
San Diego
http://www.aviationweek.com/aw/generic/story.jsp?id=news/awst/2011/08/08/AW_08_08_2011_p27-354586.xml&headline=SpaceX%20Plans%20To%20Be%20Top%20World%20Rocket%20Maker&channel=defense


Using the 160 to 1 thrust/weight ratio and 155,000 lbs. vacuum thrust
given, it has a mass of 970 lbs., 440 kg. However, this would make it
overpowered for the Falcon 1 first stage only. So we'll use two copies
of this stage powered by a single Merlin 1D.
The original Falcon 1 first stage with the Merlin 1C engine has a dry mass
of 1,360 kg. I estimated the mass of the Merlin 1C in the prior post to
be 650 kg. So without the engine, the stage weighs 710 kg. So two of
them will be 1,420 kg without engines, and adding on the Merlin 1D
engine gives this a mass of 1,860 kg.
The propellant mass of the two copies of the first stage is 43,080
kg. Then to calculate the payload that can be carried I'll again just
use the vacuum Isp and take the required delta-V as 9,150 m/s. We
conclude a payload of 1,140 kg can be lofted:

342*9.8ln(1 + 43,080/(1,860 + 1,140)) = 9,160.

Now we'll estimate how much the payload can be if we use a higher
energy density fuel such as methylacetylene and use lightweight
composites for the stage. I'll get a rough idea how high the Isp can
be for this case by assuming it is increased proportionally to the
same degree as for the high efficiency engine case. That is, using
methylacetylene in the high efficiency case resulted in increasing the
vacuum Isp to 384 s from the 360 s vacuum Isp for the kerosene.
Assuming the vacuum Isp will be increased to the same proportion here
gives us a vacuum Isp of 365 s for methylacetylene and the Merlin 1D
engine.
For the reduced stage weight using composites, assume again it will
be reduced by 40% aside from the engines. Then the stage weight with
the Merlin 1D engine will be .6*1,420 + 440 kg = 1,290 kg. Then will
be able to loft a payload of 2,320 kg:


365*9.8ln(1 + 43,080/(1,290 + 2,320)) = 9,160 m/s.


Also, quite likely SpaceX could make a half-size version of the
Merlin 1D engine. So you could use a single copy of the Falcon 1 first
stage. Then the payload would be approximately cut in half, 570 kg for
the kerosene/standard stage version and 1,160 kg for the
methylacetylene/composite stage version.

Note that low chamber pressure, low performance engines can also be
used to power the SSTO's is extremely important. Such engines have
less complicated combustion cycles and have to withstand much less
strenuous operating regimes. This makes them cheaper, simpler, easier
to maintain, and easier to make reusable. So the most costly component
of any rocket, the engines, become markedly cheaper for the proposed
SSTO.

What is key though is to come up with ways to get the needed altitude
compensation without adding on too much to the engine weight. In a
following post I'll discuss some methods this might be accomplished.



Bob Clark

 

[Urwumpe]

TLDR

 

[Hlynkacg]

There are too many questionable (borderline nonsensical) assumptions in that post to even begin parsing them all.

 

[Sky Captain]

The problem is if you want to build reusable SSTO you have to use the most bleeding edge hardware and that's not cheap. Consider expandable SSTO based on shutle ET. it would need 6 - 7 SSME's to lift off the pad and 6 - 7 SSME and cheap don't mix together. Now take the same ET put 2 RS 68 under it strap on 2 srb's and you have cheaper and much more capable rocket.
With reusable SSTO you would need bleeding edge engines, structure, landing system, thermal protection, bleeding edge everything made by using most advanced production methods. That would not be cheap. And after every flight ecerything would have to be taken apart and inspected by army of technicians.

IMHO only way how to make reusable space plane is to use airbreathing engines or some sort of launch assist that would cut deltaV requirements by 1 km/s or more. That way there would be much more margin for weight growth and lower performance more reliable componenets could be used.