sevenperforce

While SpaceX certainly won't be doing thirty launches this year, I don't see why they won't continue to recover stages. Not to fanboy, but landing a rocket on its tail is an engineering/guidance challenge, not a physics challenge, and SpaceX has hit all the high points there. The Orbcomm landing couldn't have been more perfect, and they've gotten the kinks out of water landings as well; the only reason Jason-3 failed was because it was using the v1.1 booster with the older, weaker landing legs. There's no reason to think that the next Falcon 9 booster to launch will also be the first orbital-class booster to be reused. All SpaceX has to do is offer reused boosters at a vastly discounted price (say, $40 million) and there will be numerous companies eager to jump on it; I'll predict we see a booster relaunch before July. Either SES-9 (if they can stick the water landing) or CRS-8 would be refurbished and ready for reuse by the time JCSAT-14 or Amos-6 launch, and both JSAT and Spacecom would probably be willing to volunteer. As far as the rest of the list, the political climate is really going to have to cool off before we get any kind of meaningful cooperation between the superpowers.

It probably will, unless you do an Apollo-style thing with only a few landings, which is sadly the most likely scenario. Then it only makes sense to reuse the propulsion system, maybe for a reusable GEO satellite tug. If the political climate improves by the time we're setting out to do this, then building an overengineered interplanetary transport capable of going to Mars or to Ceres or to Venus would probably be the best way of ensuring that we actually go to Mars, to Ceres, and to Venus. Once they've invested heavily it doesn't make sense not to keep going. You don't want to over-overengineer, though, because then your cost goes too high and they're loath to build a second or a third or a fourth one. You kinda want a fleet of these things if you're going to be able to sustain a manned Martian base. So you want to design a craft which is robust enough to last for a long time and be adapted for a lot of different missions (i.e., more spaceship than spacecraft), but is still cheap enough that we could conceivably build several of them. To that end, I'd suggest a combined hab/command center with some autonomous function, probably assembled on Earth and lifted to orbit by a single super-heavy launch. Lots of open space to keep your crew from going stir-crazy and to allow for expansion/adaptation. It would need to have internal airlocks, of course, to prevent catastrophic depressurization in the event of a hull breach, but you'd still want to try and manage a fairly open design. It would need its own engines (ideally VASMIR, but alternatively some sort of hybrid engine that can use either ion prop or chemical prop through the same nozzle) and power supply (fold-out or body-fixed solar panels with a decent battery system) so that it could still limp along if a meteoroid impact or other problem took out the main propulsive engines. Probably want to throw on an electrolysis system for converting water to fuel using solar power if you needed to. For the sake of expanded options in future missions, you'd want it to be able to be used as an emergency lander with low-gravity worlds without atmospheres, since those are the worlds where an aerobraking capsule is useless. So...Luna, Ceres, etc. ought to be realizable. This spaceship would mount to a sturdy frame which would be used for interplanetary transfers. The frame would include a spartan pressurized command capsule and feature attachment points for engines, cargo, fuel, and power supplies. The frame ought to be strong enough that it could be used for brachistrochrone transfers once our engine technology advances. This allows for modular replacement/upgrading without modification to the main spaceship, and the frame could be left in orbit if an emergency landing was needed.

How about a cage match in a hemispherical pressurized chamber on the surface of an asteroid, so there's gravity, but it's very, very low gravity and you can jump to the top of the chamber easily?

And then someone's wifi goes out momentarily and...grey goo.

It's not stupid to suggest cracking the water into hydrogen and oxygen; it's just a different approach for different missions. Your propellant choice will depend on what kind of craft you have, what goal you're trying to achieve, what energy source you're using, and the availability of fuel. Liquid water is fairly easy to store as long as you have at least some mild source of waste heat to keep it from freezing. It's also quite dense. It offers pretty good thrust if you have a powerful heat source like a nuclear reactor. Its thrust is a little less impressive if you're using something like a solar reflector (e.g. solar moth) but still respectable. Isp sucks, though, so you'll need a lot of it...but depending on your mission and the availability of water, that might not be a problem. Cracking water into hydrogen and oxygen can allow you to skip carrying an energy source, if you want. Might be useful if you have a deep-space mission far from the Sun but don't want to use a nuclear reactor (for whatever reason). Unfortunately, neither hydrogen nor oxygen are particularly dense. If you are going to carry your own energy generator and merely use the hydrogen as propellant, you need it to be fairly dense when it's injected into the reaction chamber, which either means compressing it mechanically (requiring energy for the compressor, a heat sink for the waste heat from the mechanical compressor, and resulting in preheated hydrogen which decreases the effectiveness of the engine) or keeping it liquid in the first place. But liquid hydrogen has very low density and is difficult to handle or store for long periods of time. So you may be better off just using water unless you need a high Isp. The autonomous space tugs of the future may well be giant water balloons with nuclear-thermal reactors on one end...

Earth offers plenty of options which are much better than steam for rocket propulsion. However, using a tank of liquid water with a nuclear reactor or a solar-thermal rocket is a nice simple design if you can collect the water in space.

Would an NTR then be a better choice for the interplanetary space station transfer ship thing? This is only tangentially related...but in analogy to a hybrid rocket that uses a solid fuel and liquid oxidizer, I wonder if it would be feasible to build a hybrid nuclear rocket, using lithium-6 hydride saltwater and natural uranium. Lithium-6 could be kept in solid pellets and dissolved in ordinary water as needed for each burst of thrust. Dissolving lithium-6 into water produces gaseous hydrogen (used to provide pressure to run the pump) and lithium-6 hydride saltwater, which would pass into the nuclear reactor, absorb low-energy neutrons from radioactive decay of the uranium, and decay into hydrogen, tritium, and fast neutrons, greatly increasing the energy released.

Eh, CO2 scrubbers are simple enough. If you're going to be going into space you can afford to bring your oxygen with you. But this leads to a secondary question: if you have a fuel depot with access to water, capable of performing electrolysis, is it more efficient (from an Earth-launch perspective) to provide tugs with hydrogen and oxygen to burn, or with hydrogen to put into an NTR? Or would the simplicity of just giving them water to put into their NTR work as well?

You'd likely need to use a solar sail approach to get the asteroid into place, which would then double as your station-keeping mechanism. I'm curious whether it would be possible to have a large enough body to allow for successful assists, but still be small enough and in the right spot to avoid destabilizing existing orbits.

On a different, but vaguely related topic... Would it be possible to place a largeish asteroid in any sort of stable Earth orbit where it could be used as a gravitational tractor for LEO insertion assist, without destabilizing GTO or LEO? I'm thinking about an orbit with a perigee a little above LEO and an apogee a little below GTO. Instead of needing to carry almost 9 km/s of dv to reach LEO, launch vehicles could merely execute a nearly-vertical ascent timed to slingshot around the asteroid at its perigee into a medium-low Earth orbit. After all, it's a lot easier to get to space than it is to stay in space. This is probably hella dangerous, but I'm just wondering if it's even vaguely possible.

Also how the captain insisted on using the MMU herself in place of the person who actually had the most training...

One major reason to mine asteroids and the moon is to get water for rocket fuel in space, rather than having to lift it out of Earth's gravity well. Suppose that you've got to build a generic solar-powered water-fueled rocket for something like an unmanned space tug. You have to budget tankage, solar collector, engine, and associated structure/modules. Assuming that you have access to mined water in orbit at each possible destination, what makes more sense: building a solar-thermal water rocket, that simply uses giant mirrors to focus solar radiation, boil water, and eject it out the back of your rocket (low Isp but very simple), or solar panels that convert solar energy to electricity that is used for electrolysis to split water into hydrogen and oxygen to burn in a thruster?

Making an accelerated trip to Mars will be too much for chemical propulsion, but making an accelerated trip back might work a little better. If you have a 53-tonne budget for an self-contained ion thruster to use as a Mars return accelerator, what's the best choice? It will be a fine balance of thrust-to-weight ratio, power-to-weight ratio, and Isp... If you can't accelerate at least one leg of the trip, then you have to build a very large and complicated hab, in which case it only makes sense to do this. If you aren't going to build a dedicated, persistent transfer hab, then it makes a lot more sense to jettison the hab at Mars, accelerate on your return trip, and aerobrake in.

I haven't read the book. Is it radiothermally powered, or powered by an actual nuclear-thermal rocket? Ostensibly, the Hermes only carries the fuel for the transfer from LEO to LMO and back. The MAVs and Martian habs were apparently placed via unmanned missions some time earlier, while the crew is ferried to the Hermes in an ordinary chemical rocket and uses a capsule for descent. Refueling the Hermes between missions would probably happen along with the crew-ferrying trips. In the movie (and I assume the book as well), Mark Watney uses a buried radiothermal generator as a heat source and mentions they had brought it with them...does it ever say where this came from? The curved trusses are definitely unlikely; even if you must have toroidal rotating hab, there'd be no reason not to join them with straight trusses. If you're going to have a rotating hab for gravity simulation, wouldn't it make more sense to just do a vertically stacked system, rather than a toroidal one? I suppose the toroidal arrangement allows for more floor space and a uniform artificial gravity, but a stacked system would be so much easier to construct.

Yeah, ran the numbers above and figured that out. Adjusted mission profile is here.

A handful of thoughts... To start with, "the world's space agencies" basically means the US and Russia, and maybe China. Nobody else has anywhere near the budget for this sort of thing. And since the US and Russia aren't really getting along right now...yuck. This would be more of an interplanetary space station than a transfer vehicle, so if we want it to last for a while, we're going to need to think long-term. Chemical propulsion probably isn't the best option; we'd need to wait until VASMIR engines had been integrated into the ISS and all the kinks had been worked out. Micrometeoroid shielding is another problem if it's going to be spending really long periods of time coasting around the solar system. Another possible propulsion mode would be to have an entire engine module which received liquid water from asteroids or the moon, so that the fuel supply would be separate from the propellant supply. You'd need fixed docking points for spacecraft and attachment points to strap on ascent/descent vehicles, plus a full-size robotic arm attachment system. Depending on the size we're dealing with, it might also make sense to throw some retractable solar sails on there.

I just realized that all those calculations I just performed -- determining that Falcon Heavy's upper stage would be left with 56 tonnes of fuel if it launched without a payload -- could have been avoided if I had merely looked up Falcon Heavy's launch specs and noticed that it is capable of putting a 53-tonne payload in LEO. Oops. In any case, this means that if SpaceX were to launch Falcon 9 LEO payloads (10-13 tonnes) using the Falcon Heavy instead, it would cost them around $30 million extra per launch, and leave them with 40+ tonnes of fuel in LEO. The brachistrochrone is far more advantageous on the return trip than it is on an outgoing trip. Dragon V2's integrated ballast sled allows for precision attitude and lift control on re-entry, which (if I understand correctly) gives it unprecedented aerobraking capabilities; it would be able to execute multiple passes through the atmosphere to bleed off a high transfer speed. Thus, the cost of a return brachistrochrone is only 50-60% the cost of an outgoing brachistrochrone. Plus, while the outgoing trip requires a Mars Ascent vehicle, the return trip does not, allowing the same amount of delta-V to be achieved with far less fuel. The Dragon V2 has a manned persistence of something like two weeks, IIRC.... As mentioned above, I think the cheapest manned Martian landing would involve two separate descents: first the unmanned ascent vehicle with a propulsive landing, followed by the manned Dragon V2 on an aerobraking trajectory once the ascent vehicle had made a confirmed landing. With the Apollo missions, it didn't make sense to do two separate landings because there was no way to aerobrake and the ascent delta-V is modest. But splitting them up for a Martian landing would be a significant delta-V savings. The ascent vehicle would carry the Earth-return Dragon V2, but it would only need enough dv to get up to Low Mars Orbit. At liftoff, the manned Dragon V2 would mass 8,898 kg, plus a standard Falcon-9 upperstage dry mass of 3.9 tonnes, meaning about 30 tonnes of fuel is needed to reach LMO. By this time, propulsive Falcon-9 stage 1 landings will likely have been mastered, so an upper stage retrofit with landings legs and grid fins should be able to make the descent with the remaining 719 m/s of dV. The full-thrust Falcon 9 upper stage is capable of 934 kN at max thrust, allowing nearly 4.8 martian-gees of acceleration: definitely enough. So the overall mission would look like this: Accumulate upper stages in LEO using Falcon Heavy for Falcon 9 missions with about 40 tonnes of remaining fuel ($30 million each). Launch self-contained high-thrust ion engine with attached power supply and fuel reserves with Falcon Heavy ($90 million for launch + cost of ion engine assembly and fuel). Launch Mars Ascent Vehicle in the form of a Falcon 9 upper stage retrofit with landing legs and attached Dragon V2 (for return transfer and Earth descent) with Falcon Heavy ($90 million for launch + cost of Dragon V2). Launch inflatable transfer hab, Transfer Vehicle Frame, and manned Dragon V2 (for Martian descent) with Falcon Heavy ($90 million for launch + cost of hab, frame, and Dragon V2). Inflate hab, Earth Orbit Rendezvous, assemble transfer vehicle via EVA. (Abort mode: rendezvous with ISS or use Dragon to re-enter.) Use chemical rockets for Hohmann transfer to Mars Capture Orbit; loop around Deimos to enter elliptical orbit around Mars. Mars Ascent Vehicle breaks off along with single Falcon 9 upperstage; upper stage ferries it to a minimum-dV aerobraking trajectory and burns up; Mars Ascent Vehicle descends to a propulsive landing. (Abort mode: skip to step 11 using Dragon V2 used for launch.) Once MAV landing is confirmed, crew enters the Dragon V2 used for launch, uses small burn to cross Deimos/Phobos and enter aerobraking trajectory, use repeated aerobraking passes to descend to a propulsive landing less than 1 km from MAV. Jettison hab, ion engine ferries remaining transfer stages to low martian orbit. After manned EVA on Mars, crew moves to MAV for ascent; rendezvous with ion engine and remaining transfer stages in LMO. Transfer stages fire to place craft on Earth return trajectory; ion engine kicks in to start pushing on a low-thrust one-way brachistochrone to Earth. On Earth approach, jettison ion engine and use repeated aerobraking passes in Dragon V2 to land. Doable? Maybe.

Related question about docking... In real life, how long does the actual process of maneuvering into position and making the docking connection take? Assuming, for the sake of argument, that you're already on a "collision course" to align orbits and positions to an arbitrary distance (e.g., crossing within <10 m at relative speeds of <5 m/s), how much time does it take to RCS into alignment and go through the docking process? Two or three minutes? Ten or fifteen minutes?

Hmm, let's see. The Falcon 9v1.1 first stage has a dry mass of roughly 25.6 tonnes in its reusable configuration and carries 395.7 tonnes of fuel. I don't know exactly how much fuel needs to be reserved for landing, but Elon said that landing on a barge allows staging at 2.5 km/s. Assuming first-stage gravity losses are on the order of 0.5 km/s, and the typical launch mass is 541.3 tonnes with specific impulse of 282 s, the rocket equation says that total mass at staging will be 183 tonnes, for a fuel consumption of 358 tonnes. This suggests 37.7 tonnes of fuel must be reserved for landing. So if three of these are launched together without any upper stage and full crossfeed (not exactly right, but good enough for an estimate), the launch mass is (395.7t + 25.6t)*3 or 1263.9 tonnes. Staging takes place after 716 tonnes of fuel have been consumed, corresponding to a dV of 2.3 km/s. Since there are three Falcon 9 cores firing simultaneously, gravity drag will be lower -- about 0.2 km/s -- so the true dV is 2.1 km/s, leaving 5.7 km/s of dV to get up to a parking orbit. At this point, gravity drag and aerodynamic drag will be negligible, and the higher vacuum Isp of 311 s can be used, so you'll need to burn... ...356.4 tonnes of propellant to get to LEO. Leaving you with less than 40 tonnes of propellant. That's WAY less than I expected. Well, that's disappointing. Looks like you were right. What if the upper stage is launched into LEO without payload using a 2/3-reusable Falcon Heavy? The upper stage has an inert mass of 3.9 tonnes and carries 92.7 tonnes of fuel, for a total mass of 96.6 tonnes. Launch mass will be 1361 tonnes; the same math as before gives a booster separation at 1.7 km/s (allowing 0.3 km/s of gravity/aero drag losses). At this point, the vehicle has a mass of 518 tonnes and 395.7 tonnes of fuel in the core stage, with a vacuum Isp of 311 s, for a core-stage dV of 4.4 km/s. At core MECO, the craft will be traveling at 6.1 km/s, so the upper stage will need to burn 36.7 tonnes of propellant for final orbital insertion, leaving it with 56 tonnes of propellant. That's an improvement, I suppose, but not much of one. It would have to be a partial brachistochrone -- burn for a while, then coast, then turn around and burn in the opposite direction. Still potentially an order of magnitude faster than a Hohmann transfer.

These are Falcon Heavy components, so they're already designed to be strapped together. And sure, that's a lot of engines, but you needed those engines to lift the fuel into space, so they're kind of already there. Plus, the strap-on boosters are themselves only good for a limited number of launches, so this would be a way of "recycling" boosters that won't last beyond a couple more launches. I suppose you could lift full second-stage engines into space even more cheaply, though. I don't know how the cost works out...whether the cost per m/s of dV is lower if you're launching nearly-full First Stages with 2/3 booster reuse or completely full Second Stages with 3/3 booster reuse. In the former case you do get vacuum-specific engines, which means a higher Isp. Using a modular approach where each component of the transfer vehicle stack has its own fuel supply and engine offers a lot more redundancy than trying to launch fuel and engines separately. Finally, increasing dV saves more than consumables; it saves development time. This is something we could do now, with current tech. No need to design a multi-month hab. No, but with propellant crossfeed, the core booster from a Falcon Heavy can have enough remaining fuel at staging to reach orbit in a nearly-full state.

Well, "simple" is relative. I wasn't suggesting that executing this design would be simple, only that the motion would be translation, which is a "simple" motion in comparison to swiveling, rotating, pivoting, or any other number of more complex transformations. Also, it should be noted that this doesn't eliminate the need for oxidizer. An air-augmented rocket carries all of its own fuel and oxidizer, though it has the option of burning fuel-rich during the limited range of airspeeds where secondary combustion is mechanically efficient. Well, people have certainly thought of it before, since they've built air-augmented rockets on several occasions. I don't think anyone has thought of a central-bypass AAR; at least, I wasn't able to find it in any patent literature. But it wouldn't necessarily have been possible before now, simply because we lacked the right materials. For a central-bypass AAR to enable SSTO RLVs, the air intake will need to be lightweight, strong, heat-resistant, and reusable. In the past, we didn't have any materials with more than two or three of those qualities, but materials science has advanced quite a bit in the past couple of decades.

Elon Musk has said SpaceX's likely target date for a manned Mars mission is 2025. I just like speculating about how you could manage it with current vehicles. You could manage the landing by bringing two descent vehicles along: a full Falcon 9 Stage 1 booster with an empty Dragon V2 already mounted on top plus a separate Dragon V2. After reaching low Martian orbit, the booster would break off first and execute an autonomous powered entry and descent to a tail-first landing. Once it could be confirmed that the booster had landed successfully and had enough fuel for a powered ascent, the astronauts would make an aerobraking-to-powered descent in the separate Dragon V2. They'd hang out, collect soil samples, plant flags, and then transfer to the booster, leaving the descent vehicle behind. I'm not quite sure how this differs from what I'm proposing...can you elaborate?

If SpaceX gets its booster return down to a science, and Falcon Heavy performs as expected, then an interesting possibility emerges. The Falcon 9 v1.1 FT Stage 1 booster is capable of SSTO on its own, though without payload or capacity for return. If a Falcon Heavy was launched without any second stage, however, you'd end up with a nearly-full first stage in orbit and two empty strap-on boosters returned safely to the ground, ready to refuel and relaunch. A single Falcon 9 launch costs $61 million, with fuel accounting for roughly $200,000 of that. Thus, Falcon Heavy would allow SpaceX to put a nearly-full Falcon 9 first stage into LEO for marginally more than the cost of a single Falcon 9 launch. With a $1 billion investment, that would be no less than fifteen nearly-full Falcon 9 first stages in LEO. Strap them together and you've got a launch stack capable of a Brachistochrone transfer to Mars for a manned mission in a minute fraction of the Hohmann transfer time. A short transfer time means your consumables budget can be much smaller, enabling an even-faster transfer. Can't think of a cheaper way of doing it.

Hmm...we may be conceptualizing the design a little differently, because I'm not sure those things apply. First of all, the individual nozzles should have peak efficiency; rocket engines scale quite well within reasonable limits such as these, and using many smaller engines/nozzles may even allow pressure-feeding instead of turbopumping, for a net reduction in weight. I'm a little confused by your mention of a "single aerospike type combustor"; in most of the aerospike designs I've seen, numerous small nozzles end up being the rule. Consider this and this cylindrical design and this linear design. Granted, there are also designs which appear to have a single toroidal combustion chamber, but these seem to be the exception, and I can't imagine that it would be easy to control or adjust. Gimballing on a single axis is pretty simple compared to the multiaxial gimballing on most rocket engines. When they are angled all the way "out", there won't be any plume recirculation issues because they are functioning like an inside-out aerospike: That's also how they'd be oriented for vacuum operation (though in vacuum you'd obviously see more expansion away from the wall). In this configuration, they operate identically to conventional thrusters without any loss in efficiency; the only drawback is the added weight of the ducting. There are no heating issues here that conventional rockets don't already have to deal with. Gimballing the nozzles inward would be done only so far as was necessary to take maximal advantage of the airflow. You're absolutely correct that airflow is a function of speed and altitude, but the purpose of inward gimballing is to direct the rocket exhaust toward the point of maximum ram compression, so that the airstream is mixed with the exhaust most efficiently. At launch, the thrusters would be angled similarly to how they are angled in vacuum operation, but as airspeed increases, they would be gradually angled further and further in. Only once the forward airspeed becomes a substantial percentage of the exhaust velocity would the thrusters start to be angled back out; this would coincide with the transition to exoatmospheric orbital insertion. If there was a point in the launch sequence at which reducing the oxidizer supply would allow partial ramjet operation with net increase in efficiency, this could be done easily, but it wouldn't depend on that. That might also depend on the load size and overall mission profile. I wasn't planning on airflow cooling the nozzles. It's possible that they'd be large enough to have 3D-printed flow channels for regenerative cooling; if not, making them out of a niobium alloy and using radiative cooling would also work. Using many small nozzles allows for a high surface-area-to-volume ratio. And sure, the downstream portion of the shroud is effectively a secondary rocket bell and would need regenerative cooling, but that's no surprise, and hardly outside of what conventional rockets already deal with. The intake can be made of the same PICA-X type material used for the heatshield on the Dragon V2. Heat dissipation is a major problem with ramjet and scramjet designs because they have very low thrust-to-weight ratios, meaning they must spend a long time in-atmo at high speeds in order to build up velocity. In practice, this is what dooms them more than any other issue. Air augmentation is different; while ramjets and scramjets focus on using atmospheric oxygen to reduce fuel consumption, air augmentation focuses on using the atmosphere to augment thrust for the same amount of fuel. Because of this, air-augmentation designs can have T/W ratios approaching those for bare rocket engines, and would be able to accelerate rapidly enough that cooling is much less of an issue. As somebody pointed out earlier, the Russians did do substantial development of a truck-mounted air-augmented ICBM that would have been vastly smaller than anything the US was able to field at the time, but they scrapped the project because the inventor kicked the bucket. Adding an air-augmentation shroud to existing ELV engine designs is not really cost-effective for the fuel savings. Fuel is cheap. But there's nothing about air-augmentation-design that requires exotic or extremely expensive tech; designing an optimized AA engine from the ground up is just a lot of testing and development and fine-tuning. Once development is done, the manufacturing and materials costs would not be significantly different than existing liquid-fueled rocket engines. At that point, there would be no reason not to use it on SSTO ELVs, but the huge cost savings would be SSTO RLVs. I need to crunch the math to figure out the upper limit on delta-V for thrust augmentation for a high-bypass-ratio central AA engine. As I pointed out above, there's no need for a second set of engines. A central-bypass engine can point its thrusters straight down to "ignore" the central airstream altogether; there's no loss of efficiency in vacuum operation. The central airstream is used exactly to the degree that you need to use it.

It would likely be possible to design an intake/nozzle shroud that can be simply slid forward or backward relative to the central rocket so that it is perfectly matched to airspeed, air density, and ambient pressure. But I'm primarily looking at the central-bypass approach anyway, simply because it offers a higher bypass ratio and intrinsic altitude/pressure compensation. Precisely. With an open central bypass, it's a lot easier to use shockwaves to your advantage. If necessary, there could be an internal asymmetry so that the position of the peak shockwave was independent of airspeed. On re-entry, you can pick your attitude to produce the best possible airflow characteristics. During vertical propulsive landing, backwards airflow will be a source of additional aerodynamic drag; I don't think it would hurt the engine performance. The small rocket engine bells underneath the lip would be independently throttleable and be capable of gimbaling radially about 100 degrees with respect to the engine bypass, further allowing complete control over the airflow: Depending on the angle of each of the engine bells, compensation for an extremely wide range of conditions would be possible. Looking at that wide expanse of intake surface, I almost want to contemplate adding some kind of exoskeletal turbine to enhance low-speed airflow (basically a drum-shaped inside-out turbine, where the blades are mounted on the inner surface rather than on a central spine) which can be shuttered at high speeds. Might be an unacceptable weight cost, though. Oh, the inlet wouldn't be sealed. Since the thrusters can be angled straight down (and out, actually), vacuum operation would be totally normal. The angling would have an effect comparable to an expansion-deflection spike. And yes, cooling issues are going to be present...but that's why you pick a launch trajectory that optimizes airflow to give the best possible performance. Drag, compression heating, and mass flow are nonlinear; there will be an ideal trajectory which balances all of these things to get the lowest possible gravity drag and fuel consumption numbers. Moreover, while this engine need not depend on atmospheric oxygen for combustion, it can use atmospheric air for combustion. There will likely only be a very narrow speed range where this offers any advantage at all, but because it requires no modification to do it, that's an added bonus. A prograde re-entry will allow the same heat shielding used for launch to be used for re-entry. You'd need to flip around to do a retrograde burn, of course, but that's to be expected. Not enough time for a heat exchanger to function. As the airspeed increases, the compressive heating of the airflow is going to increase, until you can no longer heat up the airflow enough to get meaningful expansion or net thrust (especially with a nozzle which has become very short in comparison to how fast you're moving forward). You can also think about it in terms of momentum; if you slow down air as it comes into your engine, you have to speed it back up to the same speed to even break even, and eject it out the back at a far greater speed to get net thrust. That's the problem that scramjets have encountered; they can get combustion to work just fine, but the body of the craft is absorbing so much kinetic energy from the onrushing air that they can't get net thrust above a certain point. The whole reason for having a very large central bypass is that you can slow down the incoming air as little or as much as you want. At really high speeds, you can let the air flow through with as little drag as possible and only get a modest augmentation. With the right design shape, this can happen automatically as a consequence of the airflow speed itself. At a standstill, you have some induced flow due to the low pressure adjacent to the rocket exhaust flows; this generally causes air to be pulled through. Of course, as you pick up speed, forward momentum forces the air inside, first due to pitot pressure and then due to a ram effect. The nice thing about an air-augmented rocket is that you don't have to worry about having enough time for combustion to take place; you just need to make sure you have enough time for expansion of the exhaust gases within the compressed airflow. Another great thing is that the drag equation actually works for you. Drag is proportional to cross-sectional area, which for a cylinder of constant thickness is linear with respect to the radius. At the same time, mass flow through the center of a cylinder is proportional to the square of the radius. So doubling the radius of your bypass will double your drag but quadruple your potential thrust augmentation.