In many discussions of rocket technology, a skeptic will often make some comment about how things would be so much better if we had Warp Drive. But the reality is that we don’t really need Warp Drive for things to be interesting. We just need Sufficiently Advanced Propulsion Technologyâ„¢ (name derived from Clarke’s Third Law).
While all sorts of arbitrary definitions for Sufficiently Advanced Propulsion Technologyâ„¢ can likely be suggested, I would suggest the following high-level requirements as a minimum for a candidate technology:
- Can operate safely in a biosphere.
- Has an Isp high enough to enable an SSTO vehicle with a Mass Ratio comparable to existing jumbo jets
- Has an engine thrust/weight sufficient to enable a high enough vehicle thrust/weight for minimized gravity losses while still keeping total engine mass comparable to the engine mass fractions of existing jumbo-jets
My logic for this is simply that if you had a propulsion system like this, the airframe and systems requirements for the vehicle would end up being not that much higher than that of an airliner, and most likely the service life of the system would be measured in the low tens of thousands of flights. While I think that it’s possible to design a fully-reusable TSTO and maybe even SSTO rockets using existing rocket propulsion technology, in order to hit the mass fractions necessary, you either get really miniscule payload fractions, or you end up getting very very challenging mass ratios, or most likely both. But with SAPTâ„¢ you get very high payload fractions, with very unchallenging mass ratios. Now realistically, you would want to spread the pain a little bit more between the propulsion and the vehicle structures/systems, but I think this is a good first-pass that’s easy to analyze–i.e. this is what a propulsion system would need to do in order to enable easy highly-reusable SSTO vehicles without having to push hard on any of the other technical areas.
So what would this mean as far as performance specifications? Let’s start with stats from a representative jumbo-jet. I’ll use the 777 Freighter as an example of an existing, long-range jumbo jet.
Payload Fraction: .29
Propellant Mass Fraction: .43
Engine Mass Fraction: .05
Let’s start first with a “dense” propellant SAPTâ„¢ system (say something like subcooled ammonia or water), since for that you could assume that the propellant storage and handling dry mass is similar to the 777’s propellant handling and storage dry mass. If you assume a required 9500m/s of delta-V to orbit (a reasonable SWAG including gravity losses and drag losses with a high-Isp propulsion system), and a .43 pmf (which works out to a MR=1.75), the required Isp is approximately 1750s. If you assume a vehicle launch T/W of 1.4 (to keep gravity losses modest in spite of the very slow change in vehicle mass with respect to time), that comes out to an engine T/W of about 29. By comparison, most solid-core NTRs have Isps with subcooled ammonia of only about 450-500s, and T/W ratios in the 10-30 range (with the better numbers being for more theoretical designs and the worse numbers for ones that actually were closer to operations back when the US did NTRs).
For a LH2 propellant SAPTâ„¢ system, you’d need to do something to factor in the much lower density of LH2 compared to Jet A, liquid ammonia, or water. If you want to keep the payload fraction the same, and assume that the propellant-density-scaling parts of mass ratio (tanks and tank volume-driven structures) make up 10% of the non-engine dry mass ratio, that gives you ~2.3%. With LH2 being ~10x less dense than Jet A, and assuming you keep the payload fraction, engine fraction, and the non-propellant-density-scaling dry fraction constant (ie you take all the extra tank mass out of the propellant fraction), that drops the propellant fraction to about 22%. Using a slightly higher required delta-V of 10km/s (to factor in higher drag losses due to lower system density, and higher gravity losses due to the higher Isp engines not accelerating the vehicle as fast due to slower mass change), you get a required engine Isp of about 4100m/s.
Are there any even semi-sane propulsion technologies that come anywhere close to these numbers? The only ones I can think of that meet all three criteria might be Bussard’s QED ARC engines or variants that run off of something like Winterberg’s micro-chemical fusion bomblets…both of which involve not-exactly-proven-out fusion technology. Pretty much you’re stuck assuming some sort of advanced fusion or anti-matter system. Solid-core nuclear thermal or laser or microwave thermal are all too low of Isp. Gas Core Nuclear thermal could be high enough Isp, maybe, for an open cycle design. But finding a way to do that without flagrantly violating requirement #1 would be tough. Some of the pulsed nuclear propulsion ideas get up into the right Isp and T/W ratio range, but all involve EMP levels that would fry anything in LEO. There may be something else that none of us are thinking of, but nothing that looks even remotely near-term.
So in conclusion, this was a fun and somewhat silly exercise, but it does show you that in order for a propulsion system to be advanced enough to make the rest of designing a high-flight-rate SSTO-class vehicle “easy”. We aren’t talking about warp drive per se, but we are talking about technologies that are sufficiently advanced compared to the state of the art to still seem somewhat magic.
Until we have something like SAPTâ„¢, it looks like we’re probably ought to have focus on technologies that allow us better T/W ratios for our engines, much better mass ratios for our vehicles (due to better materials), keep living with much crappier payload mass fractions, and live with more Rube Goldberg launch methods (like TSTOs, boostback, airlaunch, tether assisted launch, etc). It isn’t the end of the world, and I think there’s a ton of room left to be squeezed out of existing, boring, chemical propulsion. But yeah, the inner sci-fi nerd in me hopes that some genius wunderkinden out there are working on propulsion technologies that are indistinguishable from magic.
Jonathan Goff
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Coincidentally, I got into a discussion yesterday about how NTRs were the tech we’ve all been waiting for over on Americaspace. I somehow think there will be more to it than a single breakthrough. I think the writer did think NTRs were SAPT.
NTRs are the exact opposite of what you want for airline-like operations. If you want evidence for that, just look at the nuclear-powered-bomber work done by both the US and the USSR during the Cold War. They abandoned the efforts pretty quickly. And an NTR would get all manner of neutron activation after thousands of flights, limiting its service life to much lower than airliner operations. And again, the Isp for dense NTR propellants isn’t better than hydrolox. And although I like the concept of fusion (and think power production from something like inertial confinement fusion is possible and even within reach of our current capabilities–see NIF), I don’t think fusion is likely to be better than fission in this regard. I think much of the reason fusion looks so much better for propulsion is simply because fusion systems are much further in fantasyland than fission systems are, thus their flaws are less obviously apparent to someone just plugging numbers into engineering models.
But what are we optimizing for? Airliners are huge, clumsy aluminum vehicles with high landing stresses and large bending moments (from the wings), etc. We can do much better with rockets, I think.
I also think that structure mass depends more on propellant density than one might first think, since it sizes the other structures as well. And if we keep vehicle size constant, dense propellants work very well, as explored here: http://www.dunnspace.com/alternate_ssto_propellants.htm
For TSTO vehicles and in-space reusable stages, where fly back is important, mass ratio makes an even bigger difference than it would normally, since the only payload the stage has (for fly back) is its own dry mass. I still don’t think we’ve explored the whole envelope for high mass fraction tanks and high T/W rocket engines. Merlin 1D has a very high T/W (around 150), and it doesn’t use a terribly exotic engine cycle or anything like a thrust-augmented nozzle.
Breaking your launch vehicle into two stages just makes a LOT of things easier for an Earth launch vehicle. And if stage integration is just /so/ expensive that it makes TSTO unattractive compared to SSTO, then I think the right step would be to address what causes stage integration to be such a pain, to automate the process and provide just as many resources into solving that not-as-sexy problem as we do on the sexier propulsion problem. Because we still need to integrate payloads (even with a magic SSTO), and the knowledge gained in integrating stages would do well for integrating payloads as well.
But really, development cost is a huge issue, largely because the demand is so low now. If I had $20 billion to invest in a super-duper SSTO but I only had a market for 20 launches a year, I would be far better served just investing the money, letting it accumulate for half a decade and then buying a bunch of Falcon 9 (or whatever) expendable launches every year with the interest earned.
Great post summarizing why NTR isn’t necessarily the end-all, be-all (although focusing on in-space stages, not launch vehicles): http://forum.nasaspaceflight.com/index.php?topic=30812.msg1006138#msg1006138
If we could secure propellant from in-space sources then we could afford to continue to live with current propulsion because, for BEO missions, we could stop lifting propellant and lift only the most valuable payloads (e.g. people and high-tech equipment). In-orbit servicing would also reduce the launches from Earth’s surface.
Doug:
Sure, with just better ISRU (but no propulsion advancements) we could /afford/ to do NASA-type human spaceflight, but that still means no vast expansion into space.
We need reusable rockets, both launch vehicles and in-space stages.
We don’t need fewer launches from Earth, we need more.
Let’s flip this around a bit… given a reaction mass molecular weight, what temperature would we need to heat it to in order to achieve the required Isp? If you assume density is proportional to molecular weight, you can add the scaling relations for vehicle size and come up with a minimum temperature and corresponding molecular weight. That gives you a better idea of where and how you ought to be hunting for SAPT.
Also, you can probably get away with transforming at least some of the 777’s impressive cargo fraction to reaction mass without knocking yourself out of the airliner-like operations bucket. If you assume that half or two-thirds of the cargo is propellant instead, you make the requirements for SAPT much less severe without changing the underlying validity of your analysis.
Pierce,
Interesting way of looking at the problem. You want to run the numbers? 🙂
I also agree with you that 777 cargo capacity (about 29% of GLOW) isn’t necessarily required to be interesting. Even 10% of GLOW would still be very interesting. Just running the numbers that would mean needing an Isp of about 2000-2500 with LH2, and about 1000s with denser propellants.
Less magicky at that point.
~Jon
Sure, in my copious free time… 🙂 I realized, based on looking it up, that density and molecular weight aren’t strongly related, so you have to start with a database of light molecules and do them one by one. That makes it substantially more time consuming.
What really drives the design expense is the mass fraction, not the payload fraction. An airliner mass fraction SSTO with even a 5% payload fraction would qualify as SAPT to my mind.
The beginning of the periodic table is pretty short. You’ve got hydrogen, helium, and then metal(loid)s and then carbon. Hydrogen is basically the best we’re likely to ever get.
We can do advanced chemical rockets like hydrogen, lithium, fluoride tripropellant, but just like nuclear, I’m pretty sure it’s more trouble than it’s worth.
There’s a lot we can do with two stages. We could get good Isp with the first stage using some sort of air-breathing engine, and the second stage can just use hydrogen, slush hydrogen, or methane or propylene or something (I doubt using lithium or fluoride is going to be cheap… actually, helium, lithium, beryllium, and boron are probably too expensive to use as rocket propellant in a cheap RLV). It can use a vacuum-optimized nozzle for main propulsion (unlike something that starts at sea level) and a lot of work can be be done to make the TPS lightweight, robust, and very low maintenance.
And I think work on things like TPS are just as important, maybe more so, than work on propulsion. Remember, TPS is involved with perhaps 7-8km/s of the return delta-v for a reusable launch vehicle!
What a great post and comments. Everything begins with the idea and SAPT is a great one. What if you get rid of the first stage entirely and just fly a reusable second stage? (I think weird.) How?
A [very high speed] rail launch system (up a suitable mountainside) would bring maxQ earlier, closer to the release point, but would that be too big a problem?
Heavy lift balloons may not be sexy, but would they do the trick?
Pournelle thought laser launch made sense.
Me, I still like the original Orion. But I think radiation in those small amounts are healthy. Most people think anything like Orion could only be used in space. I think if we were serious about space, getting mass to orbit (the first half of anywhere) is more important. We can certainly mitigate the radiation issue (if not the political issue.) While Orion has been described as Rube Goldberg by some, it’s actually simpler that other designs (Zubrin’s NSWR for example) and could easily be built (they pretty much had the design completed in the 70s.)
But all of this has to do with bringing costs down without realizing that in capitalism, cost is only half the equation… profit motivates because ROI exists.
They do exist! [cue horror movie music]
If you are just looking to get mass to orbit and not looking at in-space propulsion here, there is one idea that gives you railroad-like operations: a space elevator. Carbon nanotube production is getting better and better, with a 10x increase in length every two years or so. All your propellant is laser light beamed from a power plant on the surface (or eventually from a solar power satellite), so screw mass fractions. Instead of rare launches you get a continuous stream of materiel into orbit.
A space elevator is some mega project and even if we have the structural material in sufficient quantity the earth’s dynamics provide a lot of risk factors. By comparison, balloon are cheaper, simpler, faster (crawling that ribbon takes a long time.) Even if the elevator turns out to be a really great idea, we have steps we should take first. It’s time we started using other planets to test risky ideas (for example: terraform some other planet, before trying anything on our own home. Not mars, it’s already a close enough analogy. I say we practice on venus where we can’t hurt anything and the reward is much greater on success.)
Oops should be for other neighborhoods.
Chris,
Of course with ISRU + SAPT you don’t need TPS.
Ed,
A full space elevator isn’t necessary. The same benefits can come from a rotovator type skyhook; with vastly shorter length, easier material requirements, and the ability to scale up as materials tech improves and demand increases.
What about the Skylon concept? The Alan Bond from UK with his ReactionEngines limited…
Granted, mass fraction is at about 4 to 5%
Fuel fraction about 80%
And engine T/W from 8 to 14 depending on speed, ISP from 3000 at mach 2, to 1500 at mach 6.