Iskierka

Not according to multiple independent economic studies - the very worst cases, where they made sure to make it as expensive as reasonable and fly as little as reasonable, it worked out as competitive with Falcon 9. Additionally, they have a lot more resources than you're imagining now, as they've entered into research agreements with USAF development labs, and have multiple clients interested in helping with engine and airframe. They also have much larger margins on the design than you seem to assume - just because they're not publicly quoting it often, doesn't mean they've pushed theory to its limit. Many of the features are specified as significantly heavier than they could be, and the SABRE engine is currently quoted as lower performance than they think they can get. The fuel tanks are specified only as standard aerospace aluminium, whereas they know there are a few likely alternatives that could improve performance - but they know that alloy will work. Yes, the project will fall through if parts underperform - but they've already assumed that -everything- -will- underperform, giving them margins for improvement rather than loss. And they've stated that they did consider an incremental approach, but the trouble is incremental approaches die off. The technology can very definitely get orbital - sub-orbital simply cannot provide enough of a business case to get the funding required to make the final vehicle, and would actually be more expensive as it would require scaling down the full-performance engine. They're aiming straight for orbital because they're confident they can do it, and they know it 100% -will- fail if they try smaller first.

While yes, this rocket design does technically exist, you got some factors slightly wrong. The solid fuel gas-generator section uses a fuel that's actually extremely oxygen-deprived, practically only containing enough to physically burn itself. The remaining solid fuel, now gasified then gets mixed with the air, to provide far more heat for far less mass in the ramjet section, creating a much more efficient ramjet than the design you described. A pure solid booster with air-augmentation has been tested but not used much, as it's minimal Isp improvement for a much more complex design, while the proper design, known as a throttleable ducted rocket, gets much higher Isp, only somewhat less than a liquid fuel ramjet, with the storage advantages of solid fuels.

This is probably because of the variance in its thrust in airbreathing mode - its TWR at static sealevel and transition point are about 9-10, but it peaks at about 14 at M=2.5 or so. I'd suggest using start and end velocitycurve values of 0.67, midpoint of 1, and setting the airbreathing thrust as 50% higher than quoted on wiki. (Unless the velocitycurve can go above 1? In which case it might be easier to handle the numbers using 1, 1.5, 1, and directly quoted value) I'm going to confirm that the body does contribute lift - in the latest D1 specifications, the rear fuselage is even slightly flattened into a lateral ovoid, to promote its lift at high mach and improve trim characteristics. Probably not more than say 15%, but enough that it should be considered. Finally, the runways are almost certainly too short by a large margin. Bearing in mind that a good part is for emergency braking, the runways specified for the actual vehicle are 5.5 km long, whereas I believe the longest KerbinSide runway is a mere 1.8 or so. EDIT: Useful note, the vehicle's own CoM is positioned at around 35% from the front of the cargo bay, if you may need help positioning it. May also be useful for users to know that its quoted payload is only if the payload CoM is in a similar place. The user's manual, section 3.5 shows acceptable positionings for a good reference.

My input: I feel like +100% should mean that control surfaces' zero-input line is at 0 AoA on the control, as that is the most common actual set-up for high AoA or high-manoeuvrability aircraft. This is so that the control surfaces will never stall and retain similar authority over the entire flight envelope. It's notably more common for them to track AoA like that than attempt to fight it, though the latter does exist. But, that's just what my brain instinctively says, I've not yet had chance to test this new feature. (Which will hopefully be useful regardless of whether its numbers make sense, for the aforementioned reason of retaining authority at high AoA.)

As another alternative, try very narrow deltas. Concorde is a good example - with its very low aspect ratio, it retains a very high control margin against the CoM, despite only having one surface. (Plus, the single surface and smooth shaping give it very good L/D, at least for such a low AR) Last Concorde replica I made was one of the most stable aircraft I've made in KSP - granted, it wasn't space-rated due to the strange vertical CoM/thrust offset, and it probably couldn't hold high AoA on aerodynamics alone, but for simply a plane, it's a pretty good design.

The diving technique works (and is often more efficient, it was a common technique for breaking Mach 1 in both Concorde and the SR-71) because of something known as specific excess power, which is how much power the aircraft can deliver versus power required to maintain steady level flight. Around Mach 1 due to increased drag, there's effectively a throat of reduced SEP that gives a narrow and difficult gap to pass through. However, the aircraft can still get much excess power just a bit slower just before this throat - so it's often better to climb high while subsonic with plenty of power, then enter a dive, which will convert altitude to speed while your engines keep thrusting, pass through the throat quickly, then climb again after you are no longer restricted by the increased M=1 drag.

A nuclear precooled rocket would be impossible, because the precooler requires a cold sink to cool the incoming air. In SABRE, this is the fuel it is burning in the engine, so there is no issue - but an NTR equivalent would only have a hot core to operate with, making precooling impossible. A SABRE type engine must operate with cryogenic fuel, and very likely would only work with hydrogen, due to the low temperatures and very high specific heat capacity of LH2. False. The SABRE engines operate airbreathing from a standstill, and are even capable of self-ignition with no ground connection, which drastically simplifies launch operation. Having to burn in inefficient rocket mode for the initial acceleration phase would significantly reduce efficiency, so this is not done. Doubtful, as the expelled gasses will be passed near a radioactive source, they will always be at least partially radioactive. The people living there care about the radiation, as I'm presuming this far in the future we would have far-reaching manned missions, and any scientists collecting data from it would care, as it will interfere with results from samples. And if you're going to be flying a single-use vehicle, there's no benefit to making it single-staged, or carrying extra mass for components required to do so. Skylon only needs wings for return, a simpler vehicle based on the engine could carry a reduced payload on the engine, though there would be little point to going to all that cost. Running organic chemicals through a nuclear reactor sounds like a bad idea for Titan, and Venus should probably just be forgotten for any long-endurance missions. On Mars, if you were to bring a nuclear reactor, its weight would be very impractical for a transport vehicle - it would be smarter to use the reactor to electrolyse CO2 to CO+O2, which can act as a rocket fuel of 90% Isp as compared to kerolox. Build a simpler hopper-rocket, that returns to the reactor base to refuel, and that is a much more practical planet exploring vehicle.

This is often taught but not entirely accurate. Two kerosene stages of 4500 m/s requires a mass ratio of ~3.7 each, with a little spare fuel, or 13.5 overall. If you adjust the split slightly, say 5000/4000, you find that the overall ratio is still 13.5. This doesn't account for certain factors, which when included, results in the upper stage being given more dv, due to factors like a lower minimum TWR requirement, resulting in a lighter engine. Thus making it preferable to have the smaller stage ratio on the lower stage, to avoid burning more fuel to lift the upper stage's. Further upper stages will be optimised for greater expansion, giving a higher Isp and making it preferable to burn with this engine. As a result splits more like 6000/3000 are common. Higher is seen regularly when cryogenic upper stages are involved, as the kerosene then becomes a boost stage to get the efficient hydrogen stage moving, and they want to use as little weight as possible on the less efficient fuel. Kerosene lower stages with cryogenic upper stages are commonly only 2000-2500 m/s, leaving the rest to the upper stage.

SpaceX are noticing issues with full reusability, such as if they actually reserve enough fuel to return both stages, they can't even launch an empty Dragon. The electromagnetic launch tubes are quite infeasible, due to the sheer scale of the engineering required, where every part needs to withstand the counterforce of some silly-high acceleration on the payload. If you want a similar alternative that actually could be done with today's technology, look at launch loops - they require one very long cable (a few hundred to thousand km) and two base stations, but in terms of designing and making, that's a much more reasonable request. In terms of power requirements, they're very modest, as they accelerate small payloads at much lower rates and longer distances, even to the point that passengers would be feasible. Stopping and starting the cable would be an impressive challenge, though there are some solutions, such as pylons that can support the cables electromagnetically at speed, and pulling the cable taut by moving the propulsive nodes in the base stations further apart. In the case of cable failure there might be some major concerns - but this is why you built it away from civilization, and didn't under-design the cable to allow that to happen, right?

http://en.wikipedia.org/wiki/Boeing_747#Development_and_testing However, this is a slight understatement. The engine had passed all airworthiness requirements, such as maximum thrust, sustained thrust, endurance, etc. However, when Boeing attempted to actually use the engine in the aircraft, they found that certain routine engine movements involving changing the thrust would repeatedly cause the engines to explode, particularly with the distorted casings. P&W insisted the engine was fine, as it had passed all tests, so the senior engineers grabbed the main guy and dragged him out to a test stand, where he couldn't blame aircraft systems. And then they started performing that failure condition, again, and again, and again, blowing up many engines in front of this guy, until he eventually conceded that yes, the certification wasn't everything, and the engine did need improving before it could be allowed to fly. This of course ignores a few problems the 747 itself had during development, such as landing gear being torn off during taxiing, but the whole engine debacle did occur as described. Nowadays manufacturers are smart enough to not try pass off an easily broken engine as working - back then, things were more lax.

In pretty much all two-stage or two-plus-boost-stage rockets, by the time the second stage is burning they're well beyond the need of overcoming gravity. The thrust ratio comes down to economics - in the Atlas, politics requires them to have lots of separate manufacturing, so two very different engines will get ordered. As such, with no restrictions, the second stage engine is very low thrust, intended to build up velocity slowly rather than fight gravity, and save fuel by increasing mass ratio. In Falcon, and Russian rockets, the engine manufacturers are generally the same. As such, since they already have to make a kind of engine, they'd like to use a similar engine on the next stage, so that parts can be duplicated to save costs. In rocket launches, what you find is that fuel cost is negligible, consistently less than about half a percent. When you have to use separate manufacturing anyway so you can't save cost by sharing engines, it's preferable to make the rest of the rocket slightly smaller and save a little fuel, as with the Atlas. When you don't, what you find is that most of the cost is in the rocket, and a large fraction of that is in the engines. So, rather than save on fuel or rocket size, it turns out more profitable to save money by duplicating engine components, as SpaceX and the Russians have been doing. If you look at the stage stats, you'll notice that despite the very different thrust ratios, the two stages will cover about the same ranges in terms of dv ratio, and thus contribute about the same getting into orbit regardless of second stage size.

Using stock aerodynamics, this was calculated long ago and is shown in mechjeb. For how to do it from basics in code, Fd = maximum_drag * mass * flightglobals.drag_multiplier * atmospheric density * v^2, therefore v^2 = g / (maximum_drag * flightglobals.drag_multiplier * atmospheric density) If you're not aiming to do this in code, just use mechjeb, or flight engineer may also measure terminal velocity. There isn't a straightforward rule of thumb.

The trouble with air-augmented rockets is they get a marginal advantage only. They might achieve something around double Isp, but have quite a large mass penalty due to large ductwork and thus, while they make an SSTO easier, it is not significantly so. Additionally their airbreathing thrust achievable can be quite limited due to the large ductwork eventually needed - at very large sizes, getting the through-flow and the exhaust to mix becomes very difficult, meaning either lots of small heavy engines, or a single large but lighter engine, that achieves much less Isp advantage. This is compared to a SABRE, which weighs around four to five times as much, but achieves ten times the Isp, up to very high altitude and velocity, with plenty of thrust capability for the kind of vehicle targeted. Rescaling the SABRE may be difficult due to the variety of unique parts, but it is perfectly reasonable to do so, as it's a fairly straightforward engine overall, and doesn't have air-exhaust mixing issues. As such the combustion chamber is quite happy to enlarge or shrink.

Oxidiser-rich is not pure. And further, quickly doing research on Encyclopaedia Astronautica, I found only one Russian engine that exceeded the shuttle's LOx ratio, achieving 6.2. The majority matched, at 6, and maybe a third were lower, most around 5.8, some as low as 4.2. Running pure for hydrogen would require a ratio of 8, which would have to be exceeded to run oxygen-rich, meaning no Russian engines did run lean as you suggest. Additionally, I never said it had to create low chamber temperatures - it just has to lower them from the pure temperatures, which can relatively easily reach 4500 K and beyond, over 1000 beyond the temperatures actually used. And, in theory, this is still a low temperature - if you could get 100% combustion of hydrogen, the peak temperature would theoretically be around 6700 K, but in practice this has never yet been observed, as at these temperatures hydrogen and oxygen begin refusing to bond to each other. As a result, what should be H2O + lots of energy, becomes some mixture of H2O + HO- + H+ + O(2-) + a lot of the energy, but considerably less. For more fun, if anyone can design a rocket that can somehow overcome this and create a rocket with 100% hydrogen usage, then even a pure rocket SSTO would require a mass ratio of just 5.5, as the exhaust velocity would reach around 5600 m/s. Hydrogen -should- be the best fuel by far, instead, it's just the best fuel by an appreciable margin.

The space junk problem is currently self-feeding. Because it's so expensive to go into orbit, no-one's going to do it for anything they don't profit from reasonably quickly - that includes cleaning up orbit. As soon as we can get to orbit cheaply, the sunk cost is relatively low, so investors are less concerned about getting paid back, and can be happy with knowing it will allow more profit in future. In short, cheap space access will increase Kessler syndrome, but it will remove the elements of this system that cannot actively fight the increase. Active satellites can avoid each other, the concerns are with the large regions of junk around 1000-2000 km that are out of control and just building up more debris. Similarly, the large costs of satellites currently are because of large launch costs. When a launch costs $100m, you don't want to spend that and then not have your satellite either turn on in the first place, or last for its intended lifetime. As such, the company will probably spend $100m on the satellite for a $100m launch, to make sure things don't go wrong on their end. Make the launch cost $5m, and the satellites will start costing $5m, because it's considered to not be so great a loss, and a failed satellite can be replaced at that cost. This, however, is ignoring that the hydrogen mass becomes a tiny fraction of the overall fuel mass, so the 10% mass figure is negligible overall, and likely slightly exaggerated, particularly for modern tanks. The oxygen tank, which is a more similar percentage to kerosene, makes up the majority of the overall tank mass, leading to, ex, the 3.6% figure for the shuttle's external tank. Which isn't even purely tank, it includes piping and such that is not counted as the tank's mass when designing the container. Plus the shuttle's engines run very hydrogen-rich, meaning that 10% tank mass figure becomes much more significant - for a reusable SSTO, more development would be put on the engine performance, helping it run at higher temperature and thus leaner than the shuttle, reducing average tank mass further. Aside: Skylon wouldn't benefit from this despite running much purer, but this is because of the small amount of additional fuel to make first ascent. Tank mass would likely still be 3-4% 30% is a very high estimate of the difference, 50% would only be seen on very poor LH2 designs versus very good RP-1 designs. More likely is between 20-30% of the TWR, and this results in the overall dv still being much higher. That same link you quoted specifically points out the overall mass of LH2 SSTOs is significantly lower than for RP-1 for the same payload, despite any drawbacks to LH2.

That "massive" Skylon is actually very narrow, so while it's very long, its volume is quite low, comparable to rockets in its payload category. Which is actually 15 tonnes, though it can do 33 if the vehicle itself does not reach orbit and launches a second stage at the peak of its parabola. Going to the rocket equation around earth, we find that for a traditional vertical-launch LH2/LOx rocket, the required mass ratio for the delta-v is 8.3 or so, leaving no spare fuel. (A rare occurrence) Thanks to the fancy-pants engine, we find that Skylon drastically reduces its dv to launch, and the mass ratio, including spare fuel, is just under 5.1. Additionally, because of its fancy engines and use of the wings on ascent, its initial TWR can be less than one, drastically reducing empty weight of the vehicle. Going to a traditional vertical-launch RP-1/LOx rocket, which is the design you are suggesting, minus extra weight for wings, we find that a mass ratio of 15.1 is required. Again, with no spare fuel, you can probably make that 16 or 17 accounting for that. Now, for the RP-1/LOx rocket, we have certain requirements. Firstly, its TWR needs to be greater than 1. Secondly, we are looking at at least three times the mass, due to that ratio. So the engines need to be likely 5-6 times as powerful versus the Skylon designed there, at a modest estimate. Skylon's engines are quite heavy, with a TWR of only 14 in airbreathing mode, but 5-6 times more and even RP-1 rockets struggle to match that overall weight. So not only do we need a lot more fuel, but our engines are heavier. Meaning now we need more fuel. Plus the wings need to be bigger to carry those big engines. Which means more fuel to get it into orbit. And because of all this fuel weight, the structure needs to be much stronger - especially when all this weight has to survive multiple G's during ascent! So now we add more fuel. At this point, you're using four times as much fuel as Skylon by mass, easy. At this size, Skylon will still have slightly larger LH2 tanks than your proposed RP-1 tanks, but Skylon's LOx tanks have nothing on yours. Which have to be built really strong and heavy and cause all sorts of difficulties. While Skylon's can sacrifice much strength, become very light, and then make it really easy to reenter, as the vehicle is practically light as a feather for its size, meaning its deceleration on reentry starts long before things heat up. Meanwhile you have a smaller, very, very heavy vehicle to try protect from heat ... Good luck. This is also ignoring that overall rocket mass ratios are quite similar for LH2 and RP-1. RP-1 does get the better overall in general, but not by as great a margin as LH2's required mass ratio, meaning LH2 has the greater payload capacity, if you care to build the rocket large enough. Most people don't due to the cost of building something that big repeatedly, which would be no issue on a reusable vehicle like Skylon. Yes, you should. Then you might realise this design won't work A pure-rocket SSTO was tried with the much more likely LH2 fuel, and it's been pointed out earlier in the thread how well THAT went. It was so difficult to manufacture to LH2's less demanding spec that they lost funding, they never would have gotten anywhere with RP-1.

While true that the main page listing for the Buran is technically wrong, the number is correct, since the Energia can (and did, at least for test launching) operate independently, meaning the Energia payload is 100+ tonnes, while the STS system reaches an upper limit of 24. (Plus human payload and supplies, which tend to be discounted when they're relatively consistent between missions and the interesting payload is a separate item)

I'm sorry, but -thin- margins? Skylon is not working with thin margins. A thin margin is when 15% increase in mass eats up the entirety of the payload, and that's only thin because many projects might see a 10% increase over design point, or on rare cases around 20%. Go look at the numbers behind Skylon. Payload is a good 25%+ of operating empty weight with the C2 design variant and set to increase with the in-progress D1. Even if the final design sees that outlier weight penalty of 20%, that's still a payload of 5%+ of the OEW, and at the target operating cost, that's still cheaper than anything else on the market. And this is assuming REL have been quoting numbers the way other entities have, showing the most promising figures they can get away with in a desperate effort to acquire the most funding. REL may be in need of funding, but everything we've seen so far, seems to be pointing at their number quoting being much more sensible. It's probably not off the cards that the final weight may be less than preliminary quotes. I also see lots of people very seriously underestimating what the work they've done so far represents. A partially-proven precooler that's 10% of the new engine design? Try more like a mostly proven precooler, which all other heat exchangers in the system will be based off. The only major differences between the precooling HE and the helium circuit HEs are scale and the precooler's requirement of an anti-frost system, which, interestingly enough, is basically totally proven now. Other components from the helium loop can be demonstrated via other technologies - maglev electromagnets that use helium for superconductors, MRI scanners, experimental fusion reactors*. Granted, these are not facing the same stresses, except perhaps the reactor, but the precooler's size demonstrates a willingness to use large amounts of inconel, which tends to help. * Related fact: Alan Bond did actually work on the Joint European Torus for a couple decades, so there should be relevant knowledge of how to make high-capacity coolant systems when facing very high temperature stresses. All this heat exchanger work has basically proven half of the engine technology which is particularly exotic, if not 3/4. Throw on top the nozzle work done by the University of Bristol for them and research into advanced combustors in pursuit of the SCIMITAR engine and they've done the majority of the work required for individual pieces, they just need to assemble it. The most unproven component of it all is probably the assisting ramjet that has to work with all this thermodynamics-cheating tomfoolery next-door and disturbing its incoming flow. And specifically, Nibb31, they've already done that kind of ruggedness testing to the heat exchanger. They strapped one of around 20-40% scale to a RR Viper jet, and ran both up to capacity - if that through-flow, vibration, and other general violence caused by doing such doesn't prove its durability, I'm not sure what does until they're funded to build the full thing. They also did all this testing in the rain - condensation from the precooler actually meant every test was in a small patch of it that they made. The airframe is not an unconventional design, only derivative of current designs to adapt to specific needs of a large entering empty structure. There is also no convention of TPS methods - capsules use a weighty, but small ablative shield. The shuttle only used tiles because it was the only option of the day, and we found out they were a bad idea. Skylon is admittedly trying a new method, but with its low ballistic coefficient, there's no reason to anticipate significant issue yet. Watch that space, but don't go saying it's a major issue to panic over. Also don't go on about reinventing the wheel because every aircraft manufacturer is doing that with every design, what with the move towards composites and integrating them more and more in different components - it's less of a severe obstacle than it might seem. As might be gathered from the long post, I'm certainly an advocate of Skylon, and have looked in good detail at the work they've done and publicised. I'm certainly aware that it's a risk and there is plenty of work to be done, but the work with the most reason to be doubted has been, and finishing it is more a function of time and money, rather than having to solve many huge remaining issues. And to respond to the obvious reply to that - if there weren't minor details to tidy, it wouldn't take time or money to finish the work. And that applies to any project, yet it still costs several billion to develop an airliner that's just a generational improvement. Everything has small issues. The issue is, trying to get some people on board, first to fund, and then to manufacture the other components, since their intention is to manufacture the heat exchangers and associated piping/control systems, then to have a real engine manufacturer provide the rest for SABRE, and someone else make the vehicle itself. As mentioned RR would be smart to get in on the engine, EADS/Astrium is likely for the vehicle. Possibly EADS/Airbus if they want to adjust the branding for any reason.

Duna's atmosphere starts at over 40 km, and Kerbin's ends at ~69.1 km. Besides that, your logic is flawed: it declares that the mun has no size whatsoever.

Since it's being mentioned I'll just throw in that the SR-71 is incapable of supersonic level flight. To break the sound barrier it has to enter a dive under full power, then re-ascend once it has less drag and more thrust from being significantly above mach 1.

I have both laptop and desktop, and use both interchangably. While the desktop is more powerful, the laptop is close enough that in most cases, the difference isn't noticeable, and the laptop has a higher-quality screen, so even with marginally less clutter or shadow resolution, the game still tends to look better. It's not all in the settings - sometimes, it just needs better colours to be a better experience.

That depends on how something is designed. Certainly with aircraft, if you increase size, skin thickness of all components stays constant. Thus, mass is only a square proportion with any given dimension. Presumably, a similar thing is true of spacecraft, given their similar requirements, and so Nova is correct, mass is 4x when all given dimensions are 2x.

Kerbin's equivalent of the von Karman line is somewhere between 40-45 km. Pretty good, I didn't think it was possible to get so high on them.

Not even orbital mechanics, it's basic physics. If it didn't work, something would be majorly wrong.

Which he was actually using (1.2e7 m = 12000 km), but he missed one significant aspect in that such orbital disturbance is only usually a significant issue when you have orbital resonance. Unless the orbits are timed just right so that the satellite being considered passes the mun at the same point in the absolute reference frame, the effect of the mun\'s gravity will tend to cancel itself, as it finds itself pulling oppositely at other passing points. The mun may provide ~0.07% of the orbital speed in dv per hour, but per orbit, Kerbin provides a whole 628% of the orbital speed in dv, and in LKO, ~1570% per hour. Perhaps it\'s a non-negligible effect, but it\'s a long, long way from being dominant. (For a reference of what\'s generally considered 'stable', having the SoI about three times more expansive than the current orbit is fairly widely accepted, and without orbital resonance effects LKO is well within this with respect to the mun.) I suspect any such orbiter experiments probably suffered from time warp inaccuracies, which, surprise surprise, KSP\'s patched conics is intended to prevent. Orbiter already has concerningly large numerical precision issues in LEO at high warp rate, the 1/11th size of Kerbin isn\'t going to help that.