Northstar1989

I second that. The game has desperately needed the ability to add more launch sites and runways for some time now- and runways in particular really ought to have been included in this expansion... Better yet, a toolkit to allow players to add multiple different building-types (using the models for the existing buildings in-game, plus making the runway segments individually peaceable), including runway segments, using the Mission Builder. That way players could, for instance, build a longer/wider runway than the KSC one out in a desert somewhere for a Scenario... If players can build missions that essentially become a custom Sandbox setup (i.e. players can keep playing after any objectives are complete, take contracts, etc.) this become a even more useful... If they can't, they should be able to. Adding launch sites is, besides the 5 meter rocket parts, really the only thing in this expansion I consider worth buying. But right now it seems like a half-finished idea that could have been so much more...

You're not trying to get *out* of orbit, just maintain an orbit- which can be done with a few dozen millinewtons of thrust for most satellite designs... Even something the size of the ISS could maintain orbit at 200 km with less than a newton of thrust- which its panels provide enough power for...

Nobody said anything about using this engine to boost a satellite from LEO to GTO. Doing so would just be silly. But if you used these engines to maintain a Propulsive Fluid Accumulator, and transferred collected gasses to a depot in a higher orbit (say 600 or 700 km) you could refuel chemical upper stages with fresh Liquid Oxygen (or Liquid Nitrogen for Thermal Rockets) at 600-700 km before sending them to GTO, the Moon, Mars or wherever...

Most afterburning jet engines *do* expand their exhaust to some degree at certain altitudes and speeds. Just because it's not as obvious as with a rocket doesn't mean it's not happening. In fact, let me clarify something- rocket engines *ARE* jet engines. Rockets are actually just a subtype of jet engine that relies on internal propellant. Don't just take my word for it, though- It's literally in the first paragraph of the Wikipedia page on jet engines: "A jet engine is a reaction engine discharging a fast-moving jet that generates thrust by jet propulsion. This broad definition includes airbreathing jet engines (turbojets, turbofans, ramjets, and pulse jets) and non-airbreathing jet engines (such as rocket engines). In general, jet engines are combustion engines." https://en.m.wikipedia.org/wiki/Jet_engine Internal Combustion Airbreathing Jet Engines (what most people think of when they use the term "jet engine") and rockets can actually be understood with many of the same equations, and share many of the same design-principles... Both generally include a narrow throat (much narrower than the rest of the engine) that compresses the exhaust stream to the speed of sound (Mach 1), and most high-performance jet engines that are designed to operate at supersonic speeds (i.e. most afterburning jets in things like fighter aircraft, but NOT the jets you see on large subsonic passenger aircraft) then expand the exhaust- because that is the only way to accelerate the exhaust beyond Mach 1 (you also got more Thrust that way- right up until you expand the exhaust to ambient atmospheric pressure...) --- So, it's the same principle in an internal combustion airbreathing jet engine designed to fly at supersonic speeds as in a rocket nozzle, really. Compress the exhaust from a combustion chamber until its velocity reaches Mach 1, then expand it to accelerate it further (ideally, enough to equal ambient pressure). The nozzle may look a bit different, but the working principles are largely the same. One of the biggest differences that DOES exist, however, is that many high-performance jets have variable-geometry "petals" that determine the final aperture the exhaust passes through- allowing the jet to produce exhaust at different pressures depending on the altitude as well as throttle/afterburner setting (increasing the throttle or igniting the afterburner increases the Mass Flow Rate through the engine- resulting in higher exhaust pressure unless you increase the expansion-ratio...), such as to better match the ambient pressure. Ideally, the petals should be opened wider at higher altitudes, higher throttle settings, and when using the afterburner... --- I also have a sneaking suspicion you don't fully understand how Expansion Ratio is defined. If you have a 1.25 meter combustion chamber and turbofan on a jet engine, then a 0.05 meter throat, then a variable-geometry nozzle that can be anywhere from 0.600 meters to 1.25 meters in diameter, the Expansion Ratio varies between 12 and 25, not between 0.5 and 1. The Expansion Ratio is determined by the ratio of diameters of the throat to the end of the nozzle, *NOT* by the ratio of diameters of the nozzle-end to the combustion chamber or the rest of the engine. The throat diameter and nozzle diameter are the *only* numbers that matter here, in fact you can increase the Expansion Ratio just by making the throat diameter smaller while keeping the final nozzle diameter the same, although this can create problems with turbulent flow or excessive chamber pressures (higher than the walls of the combustion chamber can handle) if you take this too far...

Well said. I hope you don't think I was saying anything differently than that. Just to be clear, you can *always* increase ISP by increasing chamber pressure or expansion-ratio (keeping one the same and increasing the other), it's just that, as you said, those improvements become progressively smaller and smaller as you approach a mathematical limit. Your math is off, though. It's not nearly as simple as dividing the chamber pressure by the final exhaust pressure to find the percentage of thermal energy you harness. A rocket that expands its exhaust from 100 atmospheres to 1 atmosphere is *a lot* less than 99% efficient. The math involved actually really hard- as in beyond my mathematical abilities hard (my abilities are just advanced single-variable calculus, basic statistics, and very limited amounts of multivariable calculus)- although there are simpler equations that exist that give a reasonable first approximation of the efficiency...

Your math is wrong. If you are traveling at 7.7 km/s and you need to expel exhaust at 4 times that velocity to maintain orbit, then this only amounts to an exhaust velocity of (7700 * 4=30800 m/s) 30.8 km/s, which is an ISP of (30800 / 9.8066 = 3140.7 seconds) 3140.7 s, not 3300 s. Any Gridded Ion Thruster (a type closely related to this airbreathing design) can easily beat this ISP- indeed the Dual Stage 4 Grid design can achieve exhaust velocities of 210 km/s (21,414 seconds ISP). https://en.m.wikipedia.org/wiki/Gridded_ion_thruster KSP isn't real life. Real electric thrusters have much higher ISP, and much lower Thrust, than the ion thrusters in KSP. Or maybe your figures came from the Dawn/DS1 probes, which both used the NSTAR design, which maxed out its ISP around 3100 seconds... https://en.m.wikipedia.org/wiki/Dawn_(spacecraft) Problem is, that's not even close to the best ISP thst can be achieved. Even the NEXT thruster, developed as a successor to the Dawn/DS1 thrusters and using similar ion thruster technology, can achieve an ISP of 4190 seconds: https://en.m.wikipedia.org/wiki/NEXT_(ion_thruster) Once again, though, the ESA airbreathing thruster is a dual-stage design (one that, interestingly, thermalizes the flow before ionizing it). Which means it's going to have higher ISP but lower Thrust than any design similar to NSTAR or NEXT... https://m.esa.int/Our_Activities/Space_Engineering_Technology/World-first_firing_of_air-breathing_electric_thruster Regards, Northstar

--- Also interesting: I bet this ESA-developed air electric thruster would be darn useful for the JP Aerospace "Airship-to-Orbit" project... https://en.m.wikipedia.org/wiki/JP_Aerospace Instead of the Ascender stage needing to carry considerable supplies of onboard propellant for ion thrusters, an airbreathing electric thruster would allow the use of the surrounding upper atmosphere as propellant: potentially significantly reducing the mass requirements for propulsion... --- Some interesting background on airships: Airships grow more efficient as they become larger. For instance, an airship that grows 8-fold in volume (by doubling its length), assuming an isometric relationship between structural mass and surface area, cuts in half the structural mass per square meter of volume (as Surface Area grows with the 2nd power of length for a given shape, but volume with the third power of length, in yet another example of the Square-Cube Law), and frees up mass for purposes like propulsion and payload... So, for instance, the current JP Aerospace design is a 1.8 km long airship... (by the way, interesting article discussing it below) http://www.science20.com/robert_walker/can_giant_airships_accelerate_to_orbit_jp_aerospaces_idea-225058 Let's make up some numbers to illustrate how its performance might change with scaling... If the 1.8 km design weighed 16 metric tons in structural mass and required 10 tons of propellant and engines to bring its 30 ton payload to orbit, what would happen to a half-sized design if you cut the propulsion/fuel system mass in half? Well, a design with half the dimensions (900 meters long) and isometric scaling of the structural mass based on Surface Area would weigh 4 tons in structural mass and require around 1.25 tons of propellant and engines (assuming the same proportion of total vehicle mass for propulsion as with the 1800 meter design) to bring a maximum payload of 1.75 tons to orbit (1/8th the volume but 1/4th the structural mass leads to a much lower payload fraction...) On the other hand, assuming isometric scaling of structural mass based on Surface Area, a double-sized (3.6 km long) design would have a structural mass of 64 tons and require 80 tons of propellant and engines to bring a 304 ton payload to orbit... Therefore, and this is the interesting bit relevant to Electric Air Thrusters, smaller airships operate on much thinner margins than larger ones, and if Air-Electric Thruster saved the 900 meter design just 500 kg on propulsion system mass, it would increase the payload mass by more than 28% to 2250 kg... Consequently, Air-Electric Thrusters might bring down the minimum size of an Airship-to-Orbit design significantly- perhaps allowing for much earlier development of a functional prototype carrying a tiny payload as proof-of-concept...

Not seen that thread before, no. But I'm not surprised nobody there brought up Propulsive Fluid Accumulators. The only other person I know of who seems to be as obsessed as I am with the idea of low-cost access to space is Elon Musk, and well, he doesn't exactly haunt these forums... Most people don't even know what a Propulsive Fluid Accumulator is (or what a Mass Driver assisted launch would look like, or how a Microwave Beamed Power system might work, or how an Airship-to-Orbit system might use electric thrusters to reach orbital velocities in the upper atmosphere, for that matter...) Speaking of which, I bet this system could really benefit the Airship-to-Orbit concept as well... https://en.m.wikipedia.org/wiki/JP_Aerospace Instead of ion thrusters relying on internal propellant reserves, this ESA electric air thruster would allow the Airship to operate off the surrounding atmosphere- potentially drastically reducing the size of airship that would be required to attain neutral buoyancy with a given sized payload...

Actually, a system like this would use more electrical power the more reaction mass passes through it. That's because an engine like this is essentially designed to ionize ALL the molecules entering the intake (or as close to 100% as it can get, anyways) and then accelerate all of them. So, thrust and power consumption remain directly proportional until the point where the engine "chokes" on excess pressure... The practical lower limit for altitude a satellite could maintain using this system would be determined by the availability of electrical power from the solar panels, anyways. The lower your altitude the more solar panels you need to power the thruster, but also the more drag those solar panels generate... This system has the widest margins at *higher* altitudes, where the power to drag ratio of the solar panels is better (the same area of panels generates the same power, but less drag, at higher altitudes...) --- This is also my opportunity to sneak in the subject of Propulsive Fluid Accumulators... https://en.m.wikipedia.org/wiki/Propulsive_fluid_accumulator As part of an effort to figure out whether Propulsive Fluid Accumulators would be feasible, the math for air-electric thrusters was already pioneered DECADES ago, long before the first functional prototype air-electric thruster was ever built this year, in 2018 (the first that I know of, anyways). Originally, when the idea was first looked at in the late 1960's through the 1970's, every proposal required a nuclear reactor to achieve a positive power:drag ratio for the electric thrusters systems at ANY altitude, nevertheless have any power left over to actually collect atmospheric gases (early solar panel designs existed, but their power output was pathetic). But when the idea was revisited in the 2000's, solar technology had improved far enough that solar-powered systems were feasible- but only at HIGHER altitudes (nuclear designs were suggested to operate at 100 km, for reasons having to do with economics rather than efficiency- whereas solar designs were suggested operate at altitudes of 200 km or more). Since that time, solar technology has improved further- and the Air-Electric Thruster developed by this ESA researcher could probably operate at lower altitudes (say, 180 km) than past solar air-electric thruster concepts (which were never actually built) were designed for due to power constraints and drag from the solar panels needed to power the whole thing... But I still doubt it could maintain a 90, or even 100 km orbit with today's solar panel efficiencies... http://www.esa.int/Our_Activities/Space_Engineering_Technology/World-first_firing_of_air-breathing_electric_thruster Also, if a satellite can MAINTAIN orbit at 200 km using this Air-Electric Thruster (as advertised in the ESA post), then it's perfectly reasonable to think a satellite operating at a higher altitude (say, 300 km) could operate with a surplus of electrical power: enough to potentially run a Propulsive Fluid Accumulator system and generate enough extra Thrust to counteract the extra Drag it and some extra solar panels would generate...

The ISP is only higher, and Thrust lower, for a given amount of energy imparted to each molecule. If you impart less energy to a larger number of lighter molecules, you can still get the same ISP and Thrust as with a heavier propellant... The only reason this isn't done with electric thrusters that work off internal propellant is that it kind of defeats the purpose of using a lighter propellant- the mass and volume of fuel tanks required is determined by the number of molecules stored and their pressure, not their molecular mass. So storing 1 ton of Hydrogen requires more mass in fuel tanks than storing 1 ton of Argon than storing 1 ton of Xenon. Lighter propellants also tend to be harder to store long-term. Which means the only reason to ever choose a lighter propellant (besides cost/availability) is if you're going to operate your electric thruster at a higher ISP- which lighter propellants are better suited for...

The explanations given here about Expansion Ratios do an excellent job of summarizing why vacuum rockets have larger nozzles, but I would just like to expand the discussion a step further... The mass-effectiveness of an engine nozzle is determined not only by ambient pressure but also, in large part, the [Combustion] Chamber Pressure and [Combustion] Chamber Temperature of the rocket... The hotter and more pressurized the exhaust is before it reaches the throat of the rocket nozzle, the more additional Thrust you can obtain by expanding the exhaust flow further. This is because what a rocket nozzle essentially does is convert thermal energy and pressure into kinetic energy- which is why the degree to which the exhaust cools and decreases in pressure is directly related to how much it increases in velocity. As a result of this, advances in the Chamber Pressures of current rocket designs make it worthwhile to have a larger nozzle than in past eras. For instance the SpaceX Merlin engine set new records for Chamber Pressure for conventional rocket engines- and its vacuum version thus derives relatively more benefit from the Expansion Ratio of its nozzle than an engine with a lower Chamber Pressure and the same nozzle would... Note that Chamber Pressure is tied up with Mass Flow Rate, but the two are not the same- an engine with a higher Chamber Pressure will often tend to have a higher Mass Flow Rate than an engine with a lower Chamber Pressure and, but if its Expansion Ratio is higher, it will also tend to have a higher ISP...

This story has been making headlines lately: https://www.theguardian.com/science/2018/mar/08/air-fuelled-engine-development-low-flying-satellites http://www.esa.int/Our_Activities/Space_Engineering_Technology/World-first_firing_of_air-breathing_electric_thruster Basically, a European Space Agency researcher developed a working prototype of an electric thruster that can operate off of the residual atmosphere found at 200 kilometers altitude. The Guardian's article correctly identifies the utility of this technology for maintaining the orbits of ultra-low altitude satellites flying at 200 km or so (the idea of such satellites isn't new- the ESA recently kept one in Ultra Low Earth Orbit at as low as 250 km for 4.5 years using nothing but a Xenon-Electric thruster, and NASA has designs on the books for similar satellites around Mars...) but nobody seems to have yet realized there is a much more useful type of Ultra Low Earth Orbit satellite we can build than scientific or comms satellites... --- Propulsive Fluid Accumulators. https://en.m.wikipedia.org/wiki/Propulsive_fluid_accumulator Satellites which by some means (such as a Xenon-Electric Thruster, an electromagnetic tether that pushes off the Earth's magnetic field, or an Air-Electric Thruster) maintain a very low orbit (200 km is the sweet spot for Earth- which this thruster just happens to be designed for...) and collect residual atmospheric gases with a specialized intake- some of which are then compressed and cooled to cryogenic temperatures, to store supplies of Liquid Oxygen, Liquid Nitrogen, or even Liquid CO2 (if you're willing to pay the mass penalty to compress it to 6-7 atmospheres) locally collected in Low Earth Orbit... The cryogenic liquids would then, in theory, be delivered to a fuel depot in a higher orbit- either by the Propulsive Fluid Accumulator itself (which would have to ascend to a higher orbit) or by a smaller tug/tanker/ferry of some sort. Either way, the ultimate destination for these liquids is spacecraft destined for missions beyond Low Earth Orbit- anything from Mars or Moon missions, to transfers of comm satellites designed for Geostationary Orbit... LOX can be burned with Kerosene, CH4, or H2; whereas N2 can be relatively easily reacted with O2 to form N2O4- a common oxidizer for Hydrazine-based hypergolic rocket propellants (intriguingly, Hydrazine is just N2H4, and it's also possible to manufacture hydrazine-derived hypergolic in LEO using the locally collected N2 as a feedstock along with CH4 and H2 launched from Earth- just significantly more complicated...) or just heated and used as propellant directly, as with compressed-nitrogen RCS systems (already use, require storing the N2 as a compressed gas though, which requires much stronger/heavier pressure vessels than cryogenic liquid...) or Nuclear/Microwave/Solar Thermal Rockets... --- The point is this: the development of a propulsion system that doesn't require regular fuel launches to LEO (although even with such launches, it's still possible to leverage a few dozen kg of Xenon into *hundreds* of kg of LOX with a Propulsive Fluid Accumulator) massively improves the economics of such a technology. As such, this Air-Electric Thruster which could allow satellites virtually unlimited loiter-times in 200 km orbit (well, as long as the solar panels last before they degrade...) could be the key breakthrough that finally makes Propulsive Fluid Accumulators a reality... I can't be the only one to have noticed this (although, due to how few people even know what a Propulsive Fluid Accumulator is, and don't scoff at it as an impossible idea because they don't understand what's actually possible, the number of other people who realized this probably number in the dozens, at best) but I would certainly like to draw some attention to this idea. If this Air-Electric Thruster really works at 200 km Low Earth Orbits like it is designed to, then there's no reason to think somebody couldn't design and build a Propulsive Fluid Accumulator (though likely only a tiny, test system) within the next decade as proof-of-concept. If it worked, that would also really light a fire on developing functional orbital cryogenic fuel-transfer system to take full advantage of the nearly-unlimited supplies of cryogenic gases this would make available in Low Earth Orbit... Regards, Northstar

Came back to this thread after a long hiatus, and man do you guys love completely derailing threads... Putting aside the silly Groupthink that they can't possibly be worthwhile, Cyclers make sense at a basic level because you only have to launch them ONCE at a high cost, but can re-use them many times. Stop comparing them to one-off missions, that never was a Cycler's intended purpose. If an Aldrin Cycler gets re-used, say, 10 times, and it masses twice as much as a "conventional" Orbit-to-Orbit spacecraft you transfer to Mars, capture, and transfer back (that is, a dedicated orbital habitat that doesn't land on Mars, but rather remains in orbit, and gets re-used many times), here's what the propellant budget looks like for the main vehicles: 2 Cyclers (1 each way): Year 0: 2A+2A (2A per cycler) Years 2-18: minor course corrections Conventional Orbit-to-Orbit Habitat: Year 0: A+B Years 2-18: A+B (9A and 9B in all) Where A is the amount of fuel for a "conventional" Orbit-to-Orbit habitat to make an Earth to Mars transfer (with a transfer-time equal to a Cycler's short leg), and B is the fuel for a Mars to Earth transfer. Aerocapture is assumed, but like the course-corrections is not considered to require enough fuel consumption to bother counting. If B is 60% the value of A, then your fuel-consumption is 16A- exactly 4x greater over the course of 20 years of Mars missions (during the last 2 years you're just bringing the final crew home) using a conventional habitat vs. accelerating the Cyclers itself (which is twice as massive- and can include mass-saving devices like a greenhouse to grow food for mass-savings, mental health and recreation). Of course, the Interceptor Ship needs to make two burns each trip. We can reasonably assume it would need space and life-support capacity for about 4-5 days (not "weeks"- although the fuel consumption is higher, shorter ferry trips save you on mass for habitation space and extended life support systems). So it might reasonably be expected to mass in about 20% as much as a habitat that sustains crews for 5 or more months (people can reasonably be expected to tolerate very cramped conditions for 120 hours they'd go nuts in over 150 days...) So, the new fuel budget: 2 Cyclers: Year 0: 2A + 2A + 0.2A + 0.2B Years 2-18: 0.2A + O.2B (1.8 A and 1.8 B in all) This bumps the fuel consumption to the Cyclers to exactly 43% that of the conventional architecture over 20 years, with the break-even point in year 8 of the mission (prior to the 5th misdion, fuel-consumption for the Cycler architecture is higher). It's worth noting that a separate lander is needed with both the Cycler and "Conventional" architecture to get from Mars Orbit to Mars, but that lander either travels the same trajectory as the Interceptor Ship or the Orbit-to-Orbit Habitat (or, alternatively, "lives" on Mars with the surface base and only launches to meet with the orbital vessels a couple days every 2 years, receiving refueling and repairs the rest of the year...) and adds exactly the same amount to the fuel consumption of both architectures (though it does make the ratio between them more even). In order to get a 4-5 day ferry flight and extraplanetary docking, you might need more than 20% the fuel-consumption of a dedicated Orbit-to-Orbit Habitat making an interplanetary transfer, but there's a LOT of room for error with the Cycler architecture still requiring a lot less fuel-consumption than a "Conventional" architecture. And the 2-Cycler architecture puts a lot less wear-and-tear on the Cyclers than the dedicated Orbit-to-Orbit "conventional" vessel receives, as each Cycler only has to make the equivalent of a high-speed Mars-transfer once in its lifetime, and then spends the rest of its lifetime essentially in microgravity (course-corrections would likely be carried out with ion thrusters to save fuel, under insignificant acceleration). By contrast, a "conventional" ship makes a 5-month Mars transfer (almost *exactly* the same Delta-V as entering a Cycler Orbit) and a return-journey to Earth under significant thrust both ways every 2 years, and might not last 20, or even 10 years of active service... (since you could build at least two for the cost of 2 Cycler ships, a shorter service-life would be acceptable) I do agree with the point others have made that Cycler architectures require a larger upfront cost for a smaller marginal cost- but that's the case with literally ANY plan to get to Mars cheaper. Even technology-driven approaches (new propulsion methods, new materials etc.) require an upfront R&D cost in exchange for a lower marginal cost for future missions... That's the nature of any type of affordable access to space- you have to be willing to INVEST lots of money upfront in order to get an acceptable marginal amd lower long term cost- there are no free lunches, any low-hanging fruit that reduces both marginal and upfront cost for lottle effort has already been picked... It's not something to be cynical about- investing money now to save more money later is the very cornrstone of good economic stewardship... For not quite 3 times the initial fuel consumption in year 0, you more than cut in half your long-term fuel requirements with the Cycler architecture. And while it's R&D costs that actually dominate the costs for a Mars mission, the Cycler need not be anything more than just a larger version of the Orbit-to-Orbit Habitat with extra radiation-shielding and better docking capabilities- its required capabilities (including the ability to sustain crew for a minimum of 5 months between resupply) are otherwise essentially the same... (note the Orbit-to-Orbit Habitat could resupply at Mars before making the return to Earth- otherwise it needs to be able to support crew for at least *10* months) Anything else, like greenhouses, more crew space, or heavier comm-systems, is just added gravy for the Cycler architecture... The Interceptor Ship could quite literally be an Orion Capsule with the proposed Cygnus Exploration Augmentation Module and a heavier transfer stage- capabilities we'd already want to have before going to Mars anyways. In fact, an Orion+Cygnus EAM would be overkill- able to support a crew of 4 for 60 days (whereas a ferry-orbit shouldn't take more than a week) for a bit less than 30 (29.3) tons of launch mass (including the 15.5 metric ton Orion Service Module, 10.4 metric ton Orion Capsule, and 3.4 ton 4-segment Cygnus EAM) plus also a transfer-stage (or simply an extended Service Module with extra fuel onboard) http://www.spaceflightinsider.com/missions/commercial/orbital-proposes-future-deep-space-applications-cygnus/ https://en.m.wikipedia.org/wiki/Orion_(spacecraft) 29.3 metric tons for an Interceptor Ship may sound like a lot, but any Mars vessel calable of sustaining 4 crew on both a 5-month journey to Mars and a 5-month journey back (let's forget resupply at Mars for a moment) could easily end up massing 160 metric tons or more, assuming plenty of radiation-shielding... A Cycler ship would be about twice the size- so about 320 tons going by this estimate... (A lot of that in heavy radiation-shielding to protect the crew and onboard electronics from as much cosmic radiation as possible) It would literally mass more than 10 times the Interceptor Ship- which is why any argument that having a Cycler isn't worthwhile just because the Interceptor would need to have life-support for a week or two is utterly hollow... An Orion+Cygnus Exploration Augmentation Module might not be light, at 30 tons (you'd definitely need a dedicated Falcon Heavy launch to lift it and its transfer-stage in one launch) but it's certainly a lot lighter than a 320-ton Cycler Ship or a 160-ton "conventional" Orbit-to-Orbit transfer ship/habitat (either one of which would need to be launched in segments, as even a Falcon Heavy couldn't lift either in one go...) and would require a lot less fuel for a ferry-orbit to a Cycler ship (a considerable bit more expensive in Delta-V than a 5-month Mars transfer) All my masses assume very generous allowances for radiation-shielding, thermal insulation, and redundancy systems. I might be way off on how much/little space I assume the crew can get by on for a Mars transfer however- but nobody will really know the limits of human tolerance for overcrowding on an interplanetary mission until they try it with trained astronauts (civilian experiences with overcrowding don't cut it, as astronauts are specifically selected for their psychological resilience, and civvies are not. Russian experiences with Mir, however, seem to indicate highly-trained, psychologically-screened humans are much more capable of withstanding cramped conditions than most people assume...)

The ESA is apparently feeling kind of left behind by all this (read to the bottom of this article) https://arstechnica.com/science/2018/02/three-years-of-sls-development-could-buy-86-falcon-heavy-launches/ One wonders if this might drive them to invest in the SKYLON spaceplane (the design that would use the SABRE engine). Or to try something really crazy like magnetic launch-assists (like a slightly toned-down StarTram, where the rocket still fires it's upper-stage engines on the way out of the atmosphere) or even Microwave Thermal Rocketry... https://www.space.com/38384-could-startram-revolutionize-space-travel.html http://parkinresearch.com/microwave-thermal-rockets/#TheBottomLine Of course, they'll probably just sit on their hands and get left in the dust of groundbreakers like SpaceX if my suspicions are correct. Government space agencies have proven far too unwilling to gamble on risky new launch technologies in recent decades- which is why SpaceX beat them all to the next big breakthrough with reusable launch stages. ------------ But if the ESA *were* willing to take a big risk, like building a magnetic launch-assist system (the smart thing to do would be to build a reduced-length system: a 5-mile track should be able to accelerate payloads to 1/4th the speed of the 81-mile track proposed for StarTram for only 1/16th the construction cost- 2.2 km/s for $1.25 - $2.5 billion... Still a quite substantial boost on the way to orbit, and still a long enough tube to reach up the side of a mountain for ejection at 18,000 feet...) or investing in Microwave Thermal Rocketry (possible launch costs of under $1000/kg, beating even SpaceX reusables) the potential benefits would be *enormous*- at least assuming SpaceX didn't catch on and find a way to take the same approach faster, better, cheaper...

The Orion- which was supposed to be capable of going to Mars with an extended habitation module- is supposed to slate in at 10.4 metric tons for the capsule alone. https://en.m.wikipedia.org/wiki/Orion_(spacecraft) The EAM proposed by Orbital AKT is basically just a modified Cygnus. The 2-segment one is slated to be able to support 4 crew for 60 days on its own, the 4-segment one would probably be at least twice as capable... https://en.m.wikipedia.org/wiki/Cygnus_(spacecraft) http://www.spaceflightinsider.com/missions/commercial/orbital-proposes-future-deep-space-applications-cygnus/ The "standard Cygnus" in cargo configuration with a 2 ton payload capacity only masses in at 1.5 metric tons. The "enhanced" version with 3.2 to 3.5 tons of payload capacity only masses in at 1.8 tons (so there is a less-than-linear relation between size and dry mass). A modified version for crew probably would weigh 5-6 tons for the 4-segment manned version at most... (that's likely more than a 100% increase in mass vs. a cargo version with the same dimensions) A Falcon Heavy can lift 63.8 metric tons to LEO in a single launch. https://en.m.wikipedia.org/wiki/Falcon_Heavy So, you could easily piece together a "flag and footprints" Mars mission by launching something like an Orion Capsule (10.4 tons), a few 4-segment EAM's based on the Cygnus (5-6 tons each), and a Mars lander (probably not more than 40 tons dry mass) on 4 or 5 Falcon Heavy launches- with plenty of extra mass budget to work with for the transfer-stages... With a little orbital-docking, and perhaps staging the components in a highly-elliptical orbit before the astronauts even arrive, it should be perfectly possible to carry out a flag-and-footprints style Mars mission with the Falcon Heavy as the launch vehicle. It's capable of launching any individual component necessary for a Mars mission- it's just a matter of finding ways to hook them all together in Earth Orbit (propellant-transfer between the payloads also would help here...) BFR is nice, but it's only really necessary for Mars colonization or long-term stays- not for simple "flag-and-footprints" missions similar to Apollo but on Mars... And at $450 million for five Falcon Heavy launches (going by SpaceX's estimated price tag, of $90 million per Falcon Heavy launch) you could launch the components for TWO Mars missions for less than the cost of a single Saturn V... (which would also help with redundancy- send two missions at once to different parts of Mars and they can potentially provide assistance to each other if one gets in trouble...) Each Falcon Heavy has a bit more than half the payload capacity of the Saturn V, but only costs less than 10% as much...

Just found time and energy to watch the launch (working as an EMT is demanding!) Very cool! Now they've just gotta start looking at more things they can do with all that payload. Developing a system of Propulsive Fluid Accumulators and orbital fuel depots would make it possible to refuel the upper stage in orbit (by launching it with surplus kerosene and getting LOX from the PFA's) to enable upper stage recovery, and go a long way towards bringing down the cost of interplanetary space travel by allowing mission designers to leverage the fuel mass they bring to Low Earth Orbit (so 1 ton of Kerosene becomes as good as more than 3 tons of Kero/LOX in Low Earth Orbit, as Kero/LOX is more than 2/3rd's Oxygen by mass...) for instance... https://en.m.wikipedia.org/wiki/Propulsive_fluid_accumulator http://www.alnaspaceprogram.org/blog/?p=25 Maybe somebody could also talk Musk into seriously looking into use of Microwave Thermal Rockets for use in Earth Departure Stages (ideally, a reusable one that could be added to the BFR as an optional second stage between the booster and final stage to increase its payload capacity, and would seperate from the BFR before leaving Earth's SOI and do a boostback/recapture burn+aerobraking to eventually land back on Earth for refurbishing...) while they're at it. Because Musk's original field of study was Physics, he should be well aware of the potential benefits of using energy beamed from outside a spacecraft, rather than just onboard fuel, for propulsion- particularly as this would allow you to propel a spacecraft with just pure Hydrogen (for great ISP) or Nitrogen (which would be available in nearly-unlimited quantities from Propulsive Fluid Accumulators), and the same beamed power technology could be used to supercharge a fleet of Propulsive Fluid Accumulators (which otherwise would have to run off solar panels and only manage a small propellant surplus each month...) either with power from Earth's surface or from solar collection satellites in higher (and therefore lower-drag and more stable) orbits... And, of course, such technology could be used to provide a Moon Base or orbiting space station with extra electrical power as well... Just some ideas on breakthrough technologies that Musk could work on next if Congress/Senate still isn't biting on a manned Mars mission at Falcon Heavy prices and the BFR doesn't work out. Because Musk's background *was* in Physics and the main thing holding Microwave Thermal Rocketry back is actually the development of cheaper and more robust Gyrotrons... (the aiming and relay-systems to shoot those microwaves at a rocket during launch or while it's in orbit overhead, or to bounce and refocus those microwaves off a relay satellite in low orbit if the target isn't in direct Line of Sight are actually relatively easily achieved- it's the Gyrotrons and power-sources of batteries or capacitors which drive up the costs and prevent research into this field... Interestingly when Musk started his PhD in Physics, he initially wanted to work on developing better high-energy capacitors- so this would give him the opportunity to revisit that old love with a much larger team...) this isn't so far-fetched as it seems. And, by providing mass-leveraging (PFA's powered by microwaves), or much better ISP (Microwave Thermal Rockets using LH2) for the Earth Departure Burn they could really bring down the cost-per-kg of BFR by a lot... (by reducing relative mission mass-requirements to orbit. For example a detachable second stage could allow for a much larger portion of the BFR final stage to be dedicated to payload- since less Delta-V would be required of it to get to Mars- and improve overall mass efficiency by increasing the number of stages. Getting back to Earth from Mars would then require a little more effort with a reduced final stage fuel capacity though: probably a longer/slower return-trip at least...) Anyhow, I look forward to seeing what Musk does next. Sadly, I fear the Falcon Heavy, while a big step in the right direction, won't be quite enough to really push politicians to return to the Moon or go to Mars (flag-and-footprints style) on its own. For that, more work on orbital-refueling (needed for the BFR anyways) and technologies like orbital fuel-depots and Propulsive Fluid Accumulators (to mass-leverage the fuel we launch to LEO) to bring down the costs even further will probably be needed... Falcon Heavy will probably prove useful for bigger satellites, probes, and space stations though...

You assume that mature electric plane designs would look anything like modern airliners. I've already made the case there would be several major design differences- biplane designs instead of monoplanes (to improve range at the expense of energy-efficiency), thinner wings, and possibly dual-fuselage designs (one possibility might be to have the fuselages merge somewhere along their length, creating a "V" shaped fuselage bridged into an "A" shape by the wings. Passengers would simply have to walk further to reach seats on the opposite side of the plane...) And such ideas only skim the surface. Flying-wing designs (like the B-2 Bomber) might be easier to manage with a Center of Mass that doesn't shift due to fuel-consumption, for instance...

If you have multiple pairs of engines (remember I said add *additional* outboard engines) you can just shut down the opposing engine, or throttle down multiple engines on the opposite side to remove any yaw tendency... It's already something pilots of multi-engine aircraft are trained how to do...

I never said it was, and these are several additional factors that are also important in engine design. But the basic point stands- high exhaust velocity engines work better at high speeds than do low exhaust velocity engines, and vise-versa. Which is why electric turbofans are likely to see more widespread use than electric propellers in passenger aircraft... We're talking about how much Thrust you can generate with a given amount of energy/fuel- which is pretty much the definition of thrust efficiency. So yes, that is *exactly* what I'm talking about, and I *do* get it...

Yes and No. The wings would be *relatively* thinner. This does not necessarily mean thin in absolute terms. Just a bit thinner than they are now. Not thin enough to create issues with transonic flight. And, although removing the fuel mass from the wings *would* exasperate the root bending-moment issue, there *are* workarounds for this problem- like switching to a dual-fuselage design (this would work best with smaller aircraft, where passengers tend to walk out on the runway to load anyways...) with a section of wing between the two fuselages, or adding extra engines further outboard on the main wing (which would allow each engine to operate at an exhaust-velocity closer to cruise-speed: each engine would generate less thrust, but consume even less energy- and the total airflow accelerated would be greater...) I didn't mean matches as in "equals". I meant it as in "is based on". Nor did I use the term "Ve=V0" in my original phrasing- you introduced that. YNM accurately expressed my idea when he said: "Thrust generation comes best when the exhaust comes close to the airspeed." Note the phrase *close to*. In the real world, you want Ve to be *close to* V0, but some definite amount larger than it...

This is a limit problem. The closer you get to Ve=V0 the more air is required for maximum thrust. You reach a point where it takes more air than there is mass in the universe to reach maximum thrust long before you reach Ve=V0, and obviously you can't examine the two when they're actually equal. But, within practical limits, the closer your exhaust velocity is to your airspeed the more Thrust you can generate within the frame of reference of that V0. E= 1/2 mv^2 but Thrust = mv. The less kinetic energy you give each molecule of exhaust the more momentum you can generate by spreading your energy over more molecules... Again, all of this within practical limits. You don't want an air intake larger than your plane! (And the Drag produced would be enormous!)

You're mis-applying formulae. Thrust per kg of air goes down, but due to the higher propulsive efficiency and lower exhaust velocity you can accelerate a lot more air with the same energy. Thrust, after accounting for these factors and properly increasing airflow, increases when exhaust velocity approaches airspeed.

Oh the irony! (to say people know nothing from the armchair- then make such an inaccurate claim...) The labor costs would be *trivial* compared to all the other labor costs already incurred in operating an airline industry. More than made up for by the fuel-savings costs. It's not that it would be cheap. It's that all the other costs of running an airline- including fuel- are absolutely massive. A chore, yes. Difficult, ABSOLUTELY. But far from impossible. Far greater hurdles are overcome in aerospace design on a daily basis. Most likely you'd slot the batteries in the bottom of the fuselage, similar to where checked luggage is currently stored. This would also allow you to make the wings thinner (reducing their drag), since fuel wouldn't be stored in them anymore... So you'd trade off thinner wings for a longer fuselage, with the extra internal volume dedicated to batteries rather than passengers... The fuel-savings are also *MUCH* more substantial than you think. Consider that on a Sydney to Los Angeles A380 flight, fuel-costs amount for $244,539 of the $305,735 cost of a 484-passenger flight. http://www.opshots.net/2015/04/aircraft-operating-series-aircraft-operating-expenses/ So, if batteries costed 1/3rd what jet fuel costs to power such a flight (amortizing out the cost of purchasing the batteries over all the flights they'll be used on in an average service life), you could comfortably cut the passenger-capacity in HALF to make room for all the batteries, and still have a lower cost-per-passenger than the traditional 484-seat A380 flight... And, of course, it's easier to fill out all the seats on a 242-seat aircraft than a 484-seat one. You could probably eventually (once the technology is established) even charge passengers *slightly* more for the electric flight, as it would be quieter... And you could DEFINITELY make an argument for tax-incentives from the government as you aren't creating as much air pollution...

A Center of Gravity that moves during flight as fuel is consumed has the potential to destabilize an aircraft (if the Center of Mass of the fuel is fore of the CoM of the aircraft as a whole). It also necessarily limits design-choices. A CoM that stays fixed in place liberates the designer... As for the 2nd bit- you hit the nail on the head. Thrust generation for an airbreathing vessel is maximized when exhaust velocity matches airspeed. Which is why jets produce more thrust than propellers at high speeds, but less at low speeds... (jets have a higher exhaust velocity than propellers) Since people want to get to their destination FAST, and higher flight speeds allow higher (and therefore less turbulent) cruising altitudes, it's likely any *successful* electric aircraft will use electric turbofans for their higher exhaust velocities...

Electric engines are better-suited for contra-rotating propeller designs as they are easier to miniaturize and you can have one engine for each direction. The batteries stay on as "deadweight", but this is also an advantage of sorts- the Center of Gravity doesn't shift due to fuel-consumption, which allows for some interesting new design choices... As for replacing jet engines- that's what the Rolls Royce "E-Fan" is designed to accomplish. It's essentially an electric turbofan capable of operating up to higher exhaust velocities than propellers, allowing for higher speeds. However It's also heavier and less energy-efficient than propellers at low speeds: much like jet engines are heavier and less fuel-efficient than ICE propellers at low speeds... I suspect that due to the limitations of battery technology, propeller planes might dominate the electric aircraft markets for some time before electric turbofans eventually take over. That's because they can travel farther on a kilogram of batteries than can electric turbofans: at the expense of longer flight-time. The designs of the past may be the designs of the future with electric aircraft. Biplanes, for instance, are useful for increasing your range without exceeding airport size-limitations: a biplane with twice the wing-area and the same wing-chord as a monoplane produces 20% more Lift for 2x the subsonic Drag from the wings (or about a 60% increase in overall drag on the entire airframe), a better ratio (maybe 10%/40% or 5%/32% instead of 20%/100%) if the wing-chord is reduced on each wing for a smaller overall increase in wing-area but a better aspect-ratio... At these ratios, the increased Lift is useful for increasing your range by allowing you to pack in more batteries: for instance in a plane that is 12% batteries by mass to start with, a 20% increase in Lift might allow for a 130% increase in battery-mass, allowing for a 43.75% increase in maximum flight-time assuming constant throttle (or a roughly 50% increase in range, assuming substantial time spent to reach Cruise Altitude and speed, and higher Takeoff and Climb Thrust than Cruise Thrust in the shorter-range monoplane...) In fact, some of the Rolls Royce/Airbus designs for a Distributed Electric Propulsion plane even incorporate a short biplane span beneath the electric turbofans- though presumably this is for structural reasons to support 3-4 electric turbofans on each wing than for any aerodynamic reason... With electric aircraft, raw performance is far more important than energy-efficiency. Batteries and electricity are substantially cheaper than jet fuel when amortized over thousands of flights (assuming they can be made sufficiently reliable- and designed such that individual batteries in massive arrays can be replaced when they fail, rather than the entire array), whereas the range-limitations impose a massive economic cost. So one would expect that modern biplanes would be initially desirable to increase the range- even if the extra batteries and wings drove up initial construction costs of the planes significantly... So, in short, I predict a progression of electric planes similar to ICE planes: Propeller Biplanes --> Propeller Monoplanes--> Turbofan Monoplanes The advancement will be limited largely by the energy-density of batteries: as it improves, designs will follow a similar progression to ICE designs to chase shorter flight-times and higher cruise-speeds... Right now we're still in the Wright Brothers-like era of experimental electric aircraft. But as electrics progress, I predict we'll see the first commercial electric passenger planes in a generation or two. Once again, the first ones will be biplanes- because they can either carry more passengers or stay in the air longer for the same wingspan, and achieve potentially longer ranges. These initial flights will be slow and relatively shorter-range (the best designs *might* manage a Transatlantic flight-time but probably won't be used this way due to the very long flight-times) but cheaper than ICE flights due to their much lower fuel-costs (automation will also help reduce the higher labor-costs of slower flights: the days of robotic flight-attendants and pilotless airplanes are just around the corner, in my opinion- and will follow in the wake of driverless trucks and buses). They will mainly be used for shorter-range flights between nearby cities, while ICE jets continue to service long-distance flights. However as battery technology progresses, faster, longer-range electrics will become a reality... Alternatively, electric blimps (filled with Helium) with heavier-than-air lifting-body designs might also become the first generation of commercial electric aircraft. Blimps require far less energy expenditure per ton-mile of cargo, due to their use of buoyancy to support most of their weight, and heavier-than-air lifting-body designs can achieve some respectable flight speeds. Thus blimps have the capability to provide an alternative form of long-distance passenger transport (*especially* if the piloting of the blimps is automated) which may be more comfortable for passengers due to the inclusion of gambling areas, restaurants, etc. on board to take advantage of the much lower cost per ton-mile than ICE planes. These blimps might in turn give way to the first generation of electric passenger planes- likely biplanes for some of the reasons above...