sevenperforce

Prompted by a lengthy discussion over on @KerikBalm's "near-future scifi" thread, I created a rather beefy Excel table to model the performance and fuel fraction of a nuclear-thermal turbocharged ramrocket engine, to get an idea of what might be possible for SSTO applications. The spreadsheet worked so well that I expanded it out to provide fuel fractions for a wide range of engine types using various propellant combinations If you want to build an SSTO, you've gotta be able to fit into these fuel fractions. No getting around it. These are, within error, the absolute minimum fuel fractions you'll need to make orbit. If the table says a given fuel fraction is 85.5%, then you've gotta fit your engines, tanks, airframe, and payload (plus margins and recovery hardware) into the remaining 14.5% of GLOW. Note that I didn't even include the values for a "true" airbreather (e.g., turbojet/ramjet/scramjet), because the thrust losses of combusting the airstream while still trying to accelerate make net performance far, far worse than even a pure rocket design. With no further ado, here's the table! Fuel fractions for SSTO Precooled Turbocharged Air-augmented Rocket only NTR (H2O) N/A 72.3% 73.2% 83.1% NTR (LH2) 47.9% 48.9% 49.7% 63.7% Hydrolox 79.8% 80.7% 81.5% 89.0% Methalox 85.5% 85.9% 86.6% 92.3% Kerolox N/A 89.1% 89.7% 94.3% Visually: Each SSTO ascent profile is constructed around a base assumption of 7.8 km/s to orbit, plus 750 m/s of gravity drag and 750 m/s of air drag. The NTRs are Tantalum Halfnium Carbide pebble-bed reactors operating at slightly over 4400 K; basically the best thing we could actually build with modern tech. Assumed specific impulses: NTR (H2O): 469 s at SL, 555 s in vacuum NTR (LH2): 820 s at SL, 971 s in vacuum Hydrolox: 366 s at SL, 452 s in vacuum Methalox: 334 s at SL, 382 s in vacuum Kerolox: 282 s at SL, 348 s in vacuum Additional notes... Rocket only. This is provided mostly for comparison. It is assumed that a rapid climb is used to leave the atmosphere as soon as possible, with peak specific impulse and zero drag being reached around 2 km/s. This design invariably has the worst fuel fraction but allows the highest TWR engines. As mentioned above, altitude compensation is assumed; specific impulse climbs linearly to 2 km/s before plateauing at the vacuum value. Air-augmented. This is an optimally designed intake shroud/duct with area behind the duct for expansion. A 22% static thrust boost is assumed due to pressure entrainment. Thrust augmentation reaches 100% around Mach 1.75, then begins to drop around Mach 6, decreasing to zero at the exhaust velocity of the actual engine. Aerodynamic drag is higher, at 800 m/s, and is spread out over a larger range of airspeeds, with vacuum specific impulse being reached much later. This represents a large fuel fraction gain for a fairly modest decrease in engine TWR. Specific impulse starts at slightly higher than the vacuum isp, then rises rapidly before dropping gradually. Turbocharged. This adds a single-stage compressor fan to the shroud inlet. The increase in static thrust is substantial, allowing a physically smaller engine, but there is only a fractional improvement in net fuel fraction, as the fan is only useful to about Mach 2. Because the fan intake is more demanding than a ram intake, the design and vehicle integration may prove to cost significantly more in dry mass than it saves. However, the turbocharger can (in principle) be used alone as a ducted fan for a hover-light landing, which is a nice and very efficient advantage if you're looking for that (e.g., with an integrated-cabin manned launch vehicle). Aerodynamic drag increases to 850 m/s. Precooled. This is the design employed by SABRE, albeit without combustion of the airstream. Precooling the intake air allows the compressor fan to operate up to Mach 5 with hydrogen and Mach 3 with densified methane, with a corresponding improvement in fuel fraction. Water and kerosene are excluded, for obvious reasons. Dry mass penalty will be hefty, though. For those who REALLY want to nerd out, here are the specific impulse curves, using hydrolox as an example: And here are samples of the drag curves with respect to velocity:

Well I think there is a very good chance that they have a very incompetent CEO... Hint: if you can't decide whether your company manufactures hoverboards, skateboards, or launch vehicles, you might be an incompetent CEO. Naturally. But you'll note that I did address the cost-benefit analysis above; I just don't see any way doing a TSTO could possibly be more expensive if their mass fractions really are as stellar as they claim.

If we're allowing German, there's always the good old Rindfleischetikettierungsüberwachungsaufgabenübertragungsgesetz.

One factor to consider, assuming that we are starting at the present time, is that developed nations receive a large amount of entertainment via the internet (as, I presume, we are all doing right now). This, clearly, would change, simply due to lightspeed lag. There would be a local instant internet and some kind of cache for the "full" internet. But that would really affect communication and how information is processed.

Only for a specific speed. While you're still on the ground.

You're correct; the majority of nuclear warheads re-entered pointy-end-down. This was intentional. For one thing, ICBMs re-enter at much lower than orbital velocity, so they don't have nearly as much heating to deal with. But, more to the point, the purpose of a re-entry capsule is to slow down as rapidly as possible. An ICBM, on the other hand, doesn't need to slow down at all, so as long as it survives to impact, everyone is happy. Well, everyone on the sending end. Not so much on the receiving end.

"Kobayashi Maru" should twist everyone's fingers up nicely.

Well, since any power source could simply be pulsed in order to increase its effective power, there's no minimum. If your chosen "dispersal" method requires 100 watts, but your power source is only 1 watt, then you can run your 1-watt power source to a capacitor bank and fire it for one second every 100 seconds. Weight and mass convert easily enough, but your problem is trying to convert between horsepower (power) and thrust (force). Power needs to operate on working mass in order to generate force. In your case, the working mass is the air being pulled through your main helicopter rotor. The conversion is simple enough if you know your rotor efficiency, your rotor length, and the density of air. Keep in mind that a longer rotor will mean a more massive system, which is going to drive up your weight.

Exculpatory. "The failure investigation's probe into materials defects proved to be exculpatory." Phenomenological. "A phenomenological approach to orbital mechanics performed poorly." Quintessential. "The argument over the planetary status of Pluto seemed to many to be the quintessential astronomical debate." Okay, those are all kind of a stretch to call space-related, but........

I have not yet seen a concept I like for an integrated Dragon 3/second stage.

Indeed. In theory, there might be a size threshold where the vehicle is simply too small to stage efficiently. But at that point, square-cube losses would dominate anyway.

Yep. And the bell is actually pressure-stabilized during the burn.

Did a little more digging, and it's all very encouraging. The ARES (Axisymmetric Rocket Ejector Simulation) project found static thrust augmentation of up to 22% with an optimally-designed simple ejector duct. The PR-90, an early test vehicle for the GNOM technology, kicked in at around 200 m/s and went up to 1 km/s, more than doubling the solid-fuel isp at burnout. The GNOM would have kicked into air-augmentation mode at around 600 m/s and gone up to 2 km/s with an average of 100% thrust augmentation. Net thrust augmentation for our pebble-bed water rocket reaches 0% at around 5.4 km/s. The addition of a bypass fan increases net static turbojet thrust by 50-100% depending on bypass ratio, but drops to a fifth of its initial efficiency by Mach 1. So, all things considered, our NTTRR has a specific impulse curve that looks like this: With vertical takeoff and rapid acceleration, fan efficiency drops quickly, but by the time you go supersonic the ram effect is rising fast. By Mach 2, the fan is shut off but the ram effect continues to rise (in combination with decreasing air pressure which increases specific impulse), to a peak of 1,001 seconds at 2 km/s. Ram compression starts to drop off due to increasing airspeed and transitions to pure rocket mode at 5.4 km/s. Integrating the rocket equation across this dV range, allowing for gravity drag and aerodynamic drag distributed appropriately across the velocity profile, provides that a vehicle using a NTTRR will have mass ratio of 2.57:1. In other words, 72% of GLOW is fuel; 28% is dry mass + payload. Of course, the TWR of the NTTRR is not going to be great; probably around 12:1 or maybe 15:1. But even if we say the TWR is 12:1 and the tank/body fraction is as low as 13:1, we're still talking about putting having 14% of our launch mass as payload. Using the same payload as before, we're looking at a GLOW of 106 tonnes and a dry mass of just 14.7 tonnes. Of course, 12.8 tonnes of the "payload" is the integrated crew cabin/capsule, so we're looking at a total vehicle dry mass of 27.5 tonnes. That's about twice the dry mass of a V-22 Osprey and slightly less than the mass of an Mi-26 heavy transport helicopter (the heaviest helicopter in the world). Of course, it would be physically a bit bigger, with enough space for approximately 77 cubic meters of propellant. By comparison, a standard FEU 40' cargo container has a capacity of 67.5 cubic meters.

SuperDracos, probably.

Mass flow difference is roughly 1-2%. The engines are significantly different -- the combustion chamber is a different shape and the throat leading to the nozzle is wider -- but SL thrust would be about the same. In fact, the Merlin 1D Vacuum might not actually be able to throttle as low at SL as it can in vacuum without choking in the nozzle, since the throat is wider. Yeah, I wondered this as well, then got pilloried over at NSF when I asked in one of the threads.

Household aluminum foil usually comes in at around 0.016 millimeters. So the niobium alloy nozzle is almost twenty times thicker. Yes, it's very thin. You can cut it with kitchen shears, although it would ruin the shears. It's so thin that there is a stabilization ring around it, to keep it from being warped while it is in the interstage. Once stage separation and startup happens, the heat causes the stabilization ring to expand and fly off. You can see it in most of the SpaceX webcasts. Great idea, but it still wouldn't work. The first stage is 22.2 tonnes dry. The second stage is 4 tonnes dry. The first stage is already coming in on a suicide burn with a TWR > 1 at minimum throttle; the thrust-to-weight ratio would be more than five times greater with the smaller second stage.

Hey, no one is saying it wouldn't work, per se. Eight Merlins arranged around a central plug nozzle with an air-augmentation shroud to kick up the thrust would SSTO easily enough.

It's a common proposal, and it's not a terribly bad one. Now, towering up to 36,000' is a bit overly optimistic. The infrastructure investment is just so high that we'd never be able to get it set up when you can just put your orbital stage on top of a first stage and launch from anywhere. Not to mention that you're limited as to inclination.

The interstage is between the first stage and the second stage; the trunk is between the second stage and the Dragon capsule. The trunk stays with the capsule on orbit and is jettisoned shortly before re-entry.

It won't be going at 200 km/s until it's way, way below Mercury's orbit. A gravitational slingshot per se won't work unless you're already on a hyperbolic trajectory with respect to your target body. Think of a gravitational slingshot like riding a skateboard past a speeding car. If you grab hold of the speeding car for a moment as it passes, you can use its momentum to help slingshot you forward much faster. But if you're already traveling alongside of the speeding car, holding onto it, then you can't slingshot off of it. However, you can certainly boost your final velocity by using the Oberth effect to burn hard at perihelion. Back to our skateboard analogy; if you are holding onto the speeding car and push off of it with your foot, you'll get more of a kick than if you tried to simply push off the ground.

A mass driver on the moon or on an asteroid is a fine idea, sure.

Conversion would be virtually impossible in this particular case. Aerospike/plug nozzles are designed in one of two ways. Either they have a toroidal combustion chamber which opens in a ring around the nozzle, or they have many small combustion chambers arranged around the perimeter. For linear aerospike engines, it's always the latter. Toroidal combustion chamber: Ring of small combustion chambers: Linear aerospike, many small chambers: The Merlin engine has a very strong, lightweight combustion chamber enabling high combustion efficiency and high chamber pressure, leading to good sea-level specific impulse. The combustion chamber would have to be completely redesigned in order to be converted into a toroidal one, as combustion and flow in a toroidal chamber is completely different. Now, it might be possible, in theory, to use the Merlin engines themselves as the smaller thrust chambers arranged around a single aerospike nozzle. But that would be....just ridiculously massive.

"Spleen access port here."

I'm not sure if density is the key factor. IIRC the efficacy of an ablative heat shield is a combination of its heat capacity, its heat conductivity, and its total mass. You want something with a high heat capacity so it takes a lot of heat to ablate a given mass away, and you want something with very low heat conductivity so that each microlayer must ablate away completely before the next layer feels any significant heating.

Yeah, I've been curious as to whether the new titanium grid fins will be square, like the current ones, or switch to something like the diamond-shaped ones depicted for the ITS Booster. Simply placing a cork lip around the base of the second stage should be enough to keep it on track. The TPS panels currently used for the Falcon 9 first stage are cork.