I was looking at H2 for propellant not just for the ozone, which I'd not known about, but makes sense, but also for isotopes produced by the neutron flux. Running air will make a variety I'm sure.
Not so much, I think.
(A) The neutron flux from a p-11B reactor is about one hundred millionth of the neutron flux from a neutronic reactor of comparable power. I wouldn't expect significant activation even of the reactor structure, never mind the propellant.
(B) If the reactor is fully shielded, the airflow won't even see that
neutron flux. If it's only shadow-shielded, how does the propellant airflow see any more neutrons than the aerodynamic flow around the vehicle?
Also, the nozzles and other hardware can be optimized better for a single fluid, I'd imagine.
That depends on the engine design. De Laval nozzles are pretty universal. The rest of the engine does get substantially heavier and more complicated if you try to make it airbreathing, but even with chemical fuels it's possible to design an engine that makes it worthwhile
. The potential improvement in capability from running airbreathing is not small, as you'll see below.
I've been told efficiency relies in part on how fast your exhaust stream is flowing, versus the airflow around the form. This is why turbofans are so much more efficient than tubojets--you have a supersonic exhaust stream from the jet, whereas the fan has it closer to the cruising speed of the airliner.
Not quite. Thrust-to-power ratio (which is a completely different efficiency measure from Isp, by the way) depends on the difference between initial and final propellant velocities (initial propellant velocity being the velocity of the ambient air for an airbreather, or the velocity of the vehicle for a rocket). So what you've said is true only for airbreathing engines.
The underlying reason is simple Newtonian mechanics - momentum is proportional to mass and velocity, but energy is proportional to mass and the square
of the velocity. If you do the math, you find that an increase in mass flow rate combined with a proportionate decrease in propellant velocity change gives you the same thrust for less power.
Unfortunately, this means airbreathing engines tend to get less efficient at high speeds, all else being equal. Since the airflow initial velocity is higher, the energy required to induce a given change in velocity is also higher. And since the power requirement curve flattens as you approach final velocity=initial velocity, once you pass a certain airspeed you find that a straight rocket design would get more thrust than an airbreathing engine at the same power level no matter how high the airbreather's mass flow rate gets (the airspeed in question obviously depends on the Isp of the rocket being compared). Of course, the airbreather will still have higher Isp than the rocket because most of the propellant is still free...
With chemical engines all of this is heavily constrained by the coupling of power and internal propellant expenditure, in addition to various associated practical difficulties. Fusion power frees things up a bit, but you do eventually still run into practical difficulties (ozone, magnetic shielding requirements, minimum coolant mass flow rates, etc.).
For Bussard's idea, you wouldn't necessarily need traditional turbojets, you'd just need to run your engine so that the exhaust velocity is closer to the airstream velocity. If he decided to use kerosene for this reason, it's because he didn't think that you could run the engine at the different flow rates. I find this hard to believe. throttling a pump isn't hard, and the fact that you'd be throttling at the upper end of the flight path anyway, as the high thrust isn't as necessary, means you might as well start out hard, and lean out the propellant feed as needed.
There are two measures of engine performance. One is thrust-to-power ratio, which is a measure of how fast your vehicle can go with a given energy input. The other is specific impulse, which is a measure of how fast your vehicle can go by expending a given fraction of its internal mass as propellant.
For a jet engine, thrust-to-power ratio is directly proportional to specific impulse (roughly, since you're probably still expending some
For a rocket engine, thrust-to-power ratio is INVERSELY proportional to specific impulse.
You can't just run a lot of propellant out the back of a rocket and expect high "efficiency". You'll get a lot of thrust, yes, but you'll run out of propellant in no time flat. Modern multistage chemical rockets actually do this - that's why they're multistage. The Saturn V used almost a tenth of its first stage propellant just clearing the launch umbilical tower.
Check out Figure 7 on page 5 of this pdf
. Bussard has apparently done the analysis, and turbojets are just better than the ARC-QED rocket at low Mach numbers.
Even easier if you're not limited to just onboard propellant, too.
easier. But engine design gets... interesting, which may be why Bussard didn't go that route right away. More advanced SSTO designs are probably possible...
I agree that the runway shuttle is a better design, but if you look at page eleven of the pdf I linked to, and check out the size of those rockets, you'll see what I saw--that kind of thing is ridiculous. that let to the thought of what you'd need to get the same specific impulse with QED rockets. Confusion of a variety of things led to the slightly easier process of duplicating the Saturn V performance, as an easier problem that can be checked easier.
The payload mass fraction of that thing is below 5%. Even with relatively casual design, Bussard's spaceplane gets 14%, and I suspect it could do better with heavy optimization by a couple of competing teams of geniuses. It could also be scaled up - the Skylon D series is already bigger and heavier than Bussard's SSTO. Of course, all this only buys you maybe ~100 tonnes to orbit, unless you want to get really
special with the runways and ground handling...
Also remember that Polywell would be useful for in-space propulsion as well, which would result in a substantial reduction in IMLEO requirements for missions anywhere in the solar system. Not necessarily two orders of magnitude substantial, but still...
VTOL SSTO could prove to be feasible with Polywell, if the power-to-weight ratio of the reactor and supporting equipment can get high enough. You could then make a heavy lifter as big as you wanted (subject to ground infrastructure and acoustic concerns). The enormous specific energy advantage from fusion power, combined with the large penalty in dry weight, might make airbreathing a good idea even for a VTOL - probably ram air in an RBCC configuration, accepting the penalty associated with a rocket liftoff and offsetting it with the advantage of much lower propellant usage in flight. Back in the optimistic '60s they tossed around ideas for air-augmented super heavy SSTOs
- fusion might make something like that worthwhile. But I haven't looked into it in detail; the lack of hard numbers on all-up Polywell system masses makes it difficult...
One thing you'd have to watch is the power requirements for the rocket liftoff. If you could get the Polywell cluster to produce MUCH more power than would be developed by a kerolox cluster capable of lifting the vehicle, it might be possible to just do a Liberty-ship
-type design with high-thrust rocket propulsion in the multi-thousand-second range for Isp (the chart in Bussard's paper is right around there, you'll recall), allowing partially propulsive descent and landing after a launch. Naturally this possibility is dependent on a sufficiently large achievable power-to-weight ratio for the Polywell and associated systems...