MatterBeam

Gas Core nuclear rockets are pretty awesome. At 2000s Isp, they outperform many electric rockets while providing good mount of thrust in a very high power density package - that doesn't need any radiators! If you give it radiators to absorb the waste heat absorbed by the pressure vessel, Isp in the order of 5000s or more is available. Some preliminary research has been done and physical models built to test some aspects of the reactor, such as handling the fuel particle vortex or how to extract power by bleeding off some of the plasma into an MHD chamber. Look at the 'Nuclear Gas Core rocket' papers on the Nasa Technical Reports Server, or at the RD-600 design. However, I don't see this technology as being available to civilians. They won't benefit much from the high thrust and low mass when just loading up on more propellant is cheaper. There's also the fact that it consumes a lot of expensive fissionable fuels and it has 'explosive containment failure' as a failure mode. Solar thermal rockets so far have focused on the very simple and robust heat exchanger design, where sunlight is used to heat a temperature-resistant metal like tungsten, then run propellant over the metal. 8km/s with hydrogen comes out to a temperature of 2565K, assuming complete dissociation. It is more likely that some hydrogen does not break down and the operating temperature is closer to 3000K. Advanced materials such as tantalum hafnium carbide can survive 4000K temperatures, pushing the maximum solar-thermal heat-exchanger design's exhaust velocity to 10km/s. I am working on a design that can achieve 12km/s using liquid droplets as the heat exchanger. If your departure point is Earth, and you are heading to Mars, you'll have 100% of the thrust at the departure burn and 44% of the thrust at the insertion burn. Thankfully, your spaceship would have expended most of its propellant reserves near Earth so your effective acceleration still goes up. Huge mirrors that reflect solar wavelengths can be made very very lightweight. After all, the Sun is an easy target to track, so you don't need any of the adaptive optics that massively increase the kg/m^2 rating of the mirrors. In fact, if you really want to chase your mass savings, you can use the same designs and structures as solar sails, except that you're not bouncing away the sunlight where it came from but focusing it onto a collector on your spaceship. Solar sails can mass as low as 7g/m^2. A 5 kilogram sail of Mylar at 7g/m^2 can collect as much as a megawatt of solar power.

I see three types of propulsion for three types of ship: -Network: These ships rely on vast infrastructure to handle propulsion externally. Laser beams, kinetic streams and tether chains. The advantage is that the authority operating the infrastructure determines where the spaceship can go and when. Another advantage is that the spaceships are the cheapest possible and probably the easiest to buy. An example is the two-pulse laser ablative rocket. -Hauler: These ships exceed the ability of the existing infrastructure to propel them, or operate outside of its reach. However, their commercial purpose prioritizes efficiency over speed. They will use small quantities of propellant paired with a high Isp reactor-powered electric rocket. An example is the electrodeless plasma rocket. -Fast: These ships can't wait around like haulers. They transport time-sensitive or precious cargo in addition to people on tight schedules. They must be authorized to use high-energy nuclear propulsion. Since cost is less of an issue, engine Isp is sacrificed to increase thrust, resulting in high mass ratios. Nuclear thermal rockets are an example.

I will introduce you to the wonders of 'optimal exhaust velocity': http://www.projectrho.com/public_html/rocket/engines.php#id--Delta-V--Delta-V_Implications and to TRANSIT TIME given distance and desired acceleration: http://www.projectrho.com/public_html/rocket/torchships.php#id--Constant_Acceleration_Equations. You want Jool in 75 days? That's 51.7 billion meters in 6480000 seconds. Travel time = 2 * (Distance/ Acceleration)^0.5 6480000 = 2 * (51.7e9/ Acceleration)^0.5 Acceleration = 51.7e9/(6480000/2)^2/= 4.9mm/s^2 Acceleration = (4 * Distance) / (Travel Time ^2) The minimum acceleration required for your Jool mission is 4.9 millimeters per second squared, which is equivalent to half a milligee. The deltaV expended during this Brachistochrone transit is 31.7km/s. The optimal exhaust velocity is therefore 22.8km/s for a mass ratio 4 rocket. If you want tiny propellant tanks that represent only 20% of the spaceship's mass, so mass ratio 1.25, the the optimal exhaust velocity is 31.7*(1/ln(1.25)): 142km/s. Higher exhaust velocities are wasted. Lower mass ratio gives pitiful practical savings in propellant mass, the cheapest part of a spaceship, while multiplying exponentially the power output of the engines required, the most expensive part of a spaceship. Finally, halving the transit time quadruples the acceleration required. Taking all these into account, high Isp fusion rockets for the tiny distances in the stock Kerbal system are pretty pointless. You want a powerful reactor not to create enough thrust out of engines with Isps of tens of thousands of seconds, but to push massive payloads at the optimal exhaust velocity. Or, for very small transit times. Say I want to reach Jool in an week. 7 days: 604800 Acceleration = (4* 51.7e9)/(604800^2) = 0.56m/s^2 Total deltaV = 0.56 * 604800 = 341.9km/s. Optimal exhaust velocity for mass ratio 2 = 493km/s. If the vessel's dry mass is 1000 tons, its wet mass will be 2000 tons. Average mass is 1500 tons. Using 50% reactor to engines efficiency, I'll need a fusion rocket capable of 414GW. That's how you design a spaceship torchship!

Those numbers are derived from NASA's Transhab project. https://www.nasa.gov/content/life-support-systems

I'm looking forwards to more! A few notes: -Bare minimum human food requirements are 0.8kg of oxygen per day, alongside 0.6kg of food and up to 30kg of water. Both oxygen and water can be recycled to a very high efficiency, so all you really need to carry is 1 day's worth of oxygen, food and water, plus emergency reserves, plus recycling equipment (441kg/1.9kW) and luxury food. The 0.6kg figure is actually nearly pure sugar and/or fat and protein mix. Not very tasty! You'll want to bring along packaged and dried food for your guests. If kerbals only need half the supplies humans need, then I get a 0.3kg per kerbal per day of essential non-recyclable consumables, plus 220kg per kerbal of life support equipment, plus about 100kg for a safety margin against losses. That's a total of 352 tons for your 1000 crew.

@p1t1o: Oh definitely. If two-stage reusable launch vehicles become available without major refurbishment between flights, we could send propellants into orbit by the tons! It would pull the rug out on atmospheric gas scooping concepts.

Lunar mining is a big project where large installations are installed on the lunar surface to either work autonomously or keep humans alive for years on end. It will cost a lot to set up and operate. Sending propellant up from the lunar surface and down to LEO costs at least 5670m/s. This is less than bringing it up from Earth's surface (9500m/s) or from an asteroid (10km/s+), but it simply cannot compare to the 0m/s cost of collecting propellant gasses from the upper atmosphere. Also, the lunar surface is notably devoid of hydrogen and extracting oxygen from oxidized minerals incurs a hefty cost in the form of melting and electrolysis. It is on the order of ~70MJ per kg of O2 before efficiency losses. Compare this to the ~1MJ/kg energy cost of simply cooling down the oxygen until it becomes liquid. Putting all these elements together, I can safely say that a lunar colony will not be able to produce propellants anywhere as easily as a scoop does, and certainly cannot start producing thousands of tons per year with a single Falcon 9 launch. My last point is that this blog post, like other technology-specific posts before, are all about detailing and analysing a specific concept. Extended comparisons with other technologies using hundreds of price assumptions is far beyond the scope of what I am trying to do, which is familiarize readers with the technology itself. If a reader decides it is very interesting, they are going to tweak the specifics of their own works or settings to push the technology into pro-eminence. I have no control over that, so why close up the number of options a reader has by asserting that one option or another is 'better.' Developmental costs are highly subjective. Operational costs do have historical, political, economical, financial, industrial and cultural variations, but they do correlate more strongly with the science and maths behind the technology. For example, if it costs 70MJ/kg to produce oxygen on the Moon and 1MJ/kg to liquefy it in low orbit, I can at least assert that the equipment and power requirements on the Moon will be much higher. Maybe not exactly 70 times higher, but 'significantly higher'. If the worked example scoop failed immediately after being released from the Falcon 9, it would be a roughly 22 ton mass with a 26604m^2 airbrake creating 252 newtons of drag. If it takes 50m/s for it to fall out of orbit, then it can wait for repairs for about 22000*7800/252*50= 3.4 hours. More likely it won't deploy its massive scoop and just wait with a stowed cross-section area of about 10m^2, so it will survive for 419 days. If the scoop works for a few months and collects 100 tons, then breaks down, it will survive for 19 days. When full at 1250 tons, it will stay in orbit with the scoop deployed for 9 days. With the funnel retracted, it will stay in orbit for 65 years. It becomes apparent that if anything goes wrong with the scoop, all it has to do is retract or eject the funnel and it will stay in orbit for a very long time. During this time, it will wait for repairs just a hop away from any launch site. Also, mechanical intervention can happen every time a cargo hauler docks with it to offload propellant. This can be dozens of times a year. Long range probes are exposed to the full radiation of the sun and space, and multiple passes through the Van Allen belts if they're bringing back propellants using reusable vehicles. The technological and matheatical comparisons are what the ToughSF blog focuses on. Other less technical comparisons are harder to pin down and work out without a fully detailed list of the assumptions an author is using for their setting. An example of this is aversion to nuclear technology. If nuclear thermal rockets are in use, then a high-Isp scoop that collects nitrogen and oxygen is best. If they are not in use, a lower Isp scoop that consumes all of its nitrogen and retains the oxygen is best. That's a massive change, based on something which even in our own reality has oscillated throughout history!

I am comparing the objective merits of several technologies. Prices and costs have an objective basis, but the vast number of assumptions needed to get a directly comparable figure makes them impossible to apply to the vast number of scenarios and settings my readers would use them in. This isn't really something they'll teach you unless you're an aerospace student. I just learnt it from reading up on hypersonic spaceplane designs. I see your point now about storing sunlight. Thankfully it will not be necessary. You can absorb sunlight effectively with a black surface, and discharge the heat through a shadowed surface, or a surface angled edge-on to the sunlight. You can capture the energy across the temperature differential from the heated to cooled sides. The equilibrium temperature in direct sunlight is about 394K. A surface in perfect shadow will cool down to 3K. Across this temperature difference, you can run a heat engine continuously while in sunlight, and on residual heat over the night side of the planet. Radiatively cooled liquefaction of the collected gasses is viable option, but it will require massive surface areas to remove heat as the temperature drops to cryogenic levels. At 394K, you are radiating 1300kW/m^2. At 77K, it drops down to 64W/m^2. To bring nitrogen down below boiling point, you need even lower temperatures. This implies very heavy equipment due to their sheer size, when a relatively lightweight heat pump can do the same thing in a compact structure.

Hi! Friction actually plays a very small role at orbital velocities. This is because gas particles don't flow around objects, they ricochet off surfaces like bullets instead. A funnel that traps them causes compression, which is the main source of heating. Using the heat from the gasses you are ram-scooping is going to be difficult, but possible. The first issue to overcome is the extremely low heat conductivity. It will be hard to move the energy in the gas to whatever generator you'll be using. The second is that you do not want to reduce the pressure of the gasses, and they need to be vented into a holding tank. This means only a fraction of the energy in the gas's temperature is available, and the pressure energy is lost when you collect the gasses. Possible, but hard. It would reduce the size of the solar panels required, but they are already a negligible fraction of the overall mass... I'm not sure how the difference between orbiting in sunlight or shadow can be exploited to generate electricity. Well, one thing I can point out is that the ion engine is roughly 31% efficient and that it would encounter roughly 9.4mN of drag per square meter at those altitudes.

The GOCE's primary design objective was to design a platform devoid of vibrations, oscillations and other disturbances to its ultra-sensitive gravity instruments. This made the usual solutions of chemical thruster and maneuvering thrusters inappropriate. Fins and an electric engine are better. I don't know its surface area or engine Isp, so I can't really make useful data out of it.

If you mean orbital refueling, then it has already been attempted and is under serious study for large-scale use by NASA and SpaceX. If you mean gas scooping, it is because it forms part of the category of 'unconventional technologies' that are hard to test and must compete for funding with more conventional technologies such as re-usable rockets or nuclear-thermal engines. I don't claim to have thought of it first, I just wanted to describe the concept, expose its merits and put some numbers on it to give interested readers a point to start from. That's... not it at all. I'm comparing the collection of propellant gasses using low orbit scoops to delivering the propellant up from Earth, producing it on a lunar base or from a captured asteroid. All these options have a research and development cost, but it is not what I am focusing the comparison on. The engine determines your thrust per kilowatt rating and the percentage of gasses you can keep. Solar panels integrated behind the scoop ship's main funnel do not contribute to drag. Mass of the ship has no role in the thrust or energy requirements. Nuclear reactors do have a lifetime... but it is measured in years to decades. If the scoop can operate for such a length of time, it would have collected and processed thousands of tons of gas into propellant and would have paid for itself several times over. We do not need anything close to 100% efficiency to make the concept work. Realistic, even pessimistic, efficiency figures have been calculated into all the figures I've posted.

This was proposed in the Titan-specific part of the Saturn Energy network. Here's the mention:

I think a more expressive number is the 48-51% reduction of launch vehicle size for a geostationary satellite, or a 2x increase in payload capability to GEO. With gas-scooped LOX, beyond LEO payloads will become quite cheaper. The difference in launch costs is what you're banking on. For a Mars mission, the difference is even more extreme.

What do you mean by 'path to GTO'? The conventional scoop I worked out the performance for stays in 200km low orbit and does not change altitude. The 'diver' design trades safety and endurance for lower mass, but I only included it as an alternative design at the end. Scoops in low orbit work best by scooping up oxygen, which the store. The rocket fuel the oxidizer burns with comes up from the ground along with the spacecraft that needs it. For hydrogen, you only need to bring up 0.167kg for each kilogram of liquid oxygen received for 'free' from a scoop.

@kerbiloid: I like the Saturn terraformation mega-project idea. Here's a few thoughts I've had: -The cold outer medium you use to get a temperature gradient can always be a set of radiators evacuating heat into the vacuum of space. -I'm assuming these Titan digger-bots are fusion fueled? -There is no metal powder or even rocks on Titan until you reach its core under hundreds of kilometers of hard ice. You'll need to bring it from elsewhere. It might be better to start the project on Enceladus or Dione. -The hydrogen ice on Saturn is at pressures where metallic hydrogen exists. Basically, the pressures are so great that the bonds between hydrogen are crushed and rearranged. Nanobots will not survive! It might be better to have them simply drift on Saturn's winds and collect hydrogen while receiving shipments of carbon and minerals from the moons. -It sounds like it might be much more energy and resource efficient to simply collect De from Saturn and use it to heat Titan instead of trying to warm all of Saturn until it becomes warm like a star or brown dwarf. Maybe you could try terraforming Titan by melting the top layer of ammonia and water ice and using the vapours as a thermal blanket to prevent further heating from leaking into space? Like an oven, Titan would get quite warm very quickly. Over time, you could even lower atmospheric pressure by having hot particles escape into space.

I will try to clarify, but first I must point out that the SABRE engine deals with hundreds of kilograms of air per second and achieves '1GW/m^3' of cooling, while the design I worked out handles micro to nanograms of gas per square meter per second. For the 26604m^2 scoop that a single Falcon 9 FT can put into orbit, about 1.5kW of power needs to be delivered to heat pumps for cooling and liquefying the air. It is a much smaller scale! Heat pumps can perform the same function as open-cycle cooling, admittedly with much heavier equipment and power requirements. I calculated that coefficient of performance for cooling air from 990K to 77K and added another 38% stirling engine efficiency on top of that to work out the 1.5kW figure. The radiators to handle heat loads on the order of a kilowatt are pretty small. Even a lukewarm radiating surface at 300K with a decent emissivity of 0.8 can radiate 367W/m^2. Shipping resources up from the ground costs a lot and requires at least 16 kg of propellants on the launchpad for 1kg of payload in orbit. Add in the mass of engines, propellant tanks, structures and so on and you get the 3-5% payload fractions of modern rockets (20kg for 1kg in orbit). Atmospheric scooping therefore 'saves' 20kg of cost per kilogram it produces. Based on the worked example figures, a single Falcon 9-launched scoop can produce 757 tons of liquid oxygen per year and massively reduce propellant requirements for extra-LEO travel. The biggest market right now for this sort of thing is geostationary satellites. Mars missions would have savings of nearly 200kg on the ground per kilogram produced in orbit. Collecting hydrogen is a bad idea at all altitudes because it is a thousand times rarer than the other elements in terms of particles per cubic meter at most altitudes. If we also consider the molar mass, it is more than 10000 times rarer in kg/m^3. Gas scooping must focus on oxygen and nitrogen. The only advantages to scooping gas at 400km altitude is that the particle count for hydrogen and oxygen is pretty much equal, which facilitates the direct production of hydrolox rocket fuels. Also, the drag forces (and therefore thrust) and cooling requirements are about a 100 times lower, so a scoop that fits within the same Falcon 9 payload budget can grow to 1840m in diameter and collect roughly the same mass of gas as a 184m wide scoop at 200k altitude. It depends on the exhaust velocity of the engines it uses to counter drag. If the exhaust velocity is lower than orbital velocity, the scoop will need to consume more propellant than it could ever collect to stay in space. If the exhaust velocity is equal to orbital velocity, the scoop has to use all of the gasses it collects as propellant and retains none. If the exhaust velocity is higher than orbital velocity, then the scoop will be able to retain some of the gasses in reserve and produce a net gain in propellant collected.

As, so you would be using the heat engines as a waste heat recovery system. Smart! It would be limited in capability, but in an early power-starved colony, every scrap of heat will be used and recycled as much as possible and this sort of thing would help.

Thanks.

This is the full blog post on the topic of Atmospheric Gas Scooping as a source of propellant for Orbital Refuelling. It focuses on currently available technologies and compares it to the existing situation, but the maths and equations form the basis for discussions on futuristic situations or the use of Gas Scooping on other planets. You can read it in the original format here. Low Earth Orbit Atmospheric Scoops The rocket equation is often described as tyrannical. Low exhaust velocity chemical rockets that are used today need payloads to be mounted on towers of propellant to reach extraterrestrial destinations. Re-fueling a rocket in orbit has been floated as a solution to drastically reduce rocket sizes. The best source of propellant for this purpose is much closer than is generally thought... Orbital refuelling Orbital propellant depot Orbital refueling is a concept where a spaceship's propellant tanks are refilled after it achieves orbit. Without refuelling, a spaceship must carry the propellant it needs to reach its extraterrestrial destination up from the ground. With refuelling, it can launch empty or even reuse a propellant tank once in orbit. The propellant can be launched separately or manufactured from extra-terrestrial sources such as the Moon or an asteroid. In this post, we will be looking at a way to simply scoop up propellant from the atmosphere. Beating the rocket equation A quick calculation reveals how advantageous orbital refuelling is. SpaceX's Mars mission architecture relies on orbital refueling for good reason. Imagine a 10 ton spaceship that we want to send to Mars. It needs 9.5km/s of deltaV to enter orbit, and another 5.6km/s to go to Mars. Let us imagine that it uses engines with an average specific impulse (Isp) of 350 seconds, which corresponds to kerosene-liquid oxygen rockets. We can also write this as an exhaust velocity of 350*9.81: 3433m/s. The rocket equation is as follows: DeltaV = Exhaust velocity * ln(Mass Ratio) DeltaV, in m/s, is the velocity a rocket can reach if it expends all of its propellant reserves. Exhaust velocity, in m/s, is how fast a spaceship's engines expel propellant. Mass ratio is the fully fuelled or 'wet' mass divided by the empty or 'dry' mass. We can re-write that equation to work out the mass ratio required to obtain a certain deltaV. Mass ratio = e^(DeltaV/Exhaust Velocity) e is the exponent function, roughly equal to 2.718. If we want to reach orbit, we need 9500m/s of deltaV. Using an average exhaust velocity of 3433m/s, we calculate that the mass ratio required is e^(9500/3433): 15.9 This means that the wet mass is 15.9 times greater than the dry mass on the launchpad. Going from orbit to Mars requires a mass ratio equal to e^(5600/3433): 5.1 DeltaV requirements for travel between Earth, Moon and Mars. The total mass ratio on the ground is the product of both mass ratios. 15.9*5.1: 81. If our spaceship is 10 tons, then the rocket on the launchpad weighs at least 810 tons. A realistic spaceship with propellant tanks, structural support and engines will cut into the available payload mass. We will end up with a towering rocket that is mostly propellant, with a tiny cargo bay at its tip. What if we use orbital refuelling? How reusability and orbital refueling can make a Moon mission drastically smaller The mass ratio for launch remains the same. However, we do not need to carry up propellant from the ground to go to Mars with it - instead, we refill the same tanks with more propellant in orbit. The mass ratio on the ground is therefore 15.9. The rocket will mass only 159 tons on the launchpad. This is a rocket more than five times smaller than the previous example. Using the same calculations with the Moon as destination (3300m/s) still means the rocket is e^(3300/3343): 2.61 times smaller than without refuelling. Current challenges and solutions ITS and its tanker docking in LEO to transfer propellants. The reason orbital refuelling has not been used to drastically reduce the size of rockets on the launchpad is that currently, the only way of delivering propellant to a spaceship in orbit is with another launch. This negates any direct advantage refuelling provides. Building a lunar base. The other option discussed as a source of propellants is in-situ resource utilization. This involves setting up an industrial base on the Moon, an asteroid or elsewhere to produce propellant from locally available materials. The products are shipped back to Low Earth Orbit for refuelling spaceships. It makes sense from a deltaV perspective: it is easier to send propellant back from Mars (5.6km/s) than it is to lift it up from Earth (9.5km/s)! However, it would require a massive investment in extraterrestrial infrastructure that would not help anyone for several decades, requiring automated technology that can work independently or a permanent manned settlement. China Aerospace Science and Technology Corp.'s 30km/s Hall-effect ion thruster. NASA has been working on alternatives to orbital refuelling for high-deltaV missions. The best solution so far is the use of high exhaust velocity solar-electric propulsion. Dawn approaching Ceres. As noted above, mass ratio depends on exhaust velocity. A higher exhaust velocity drastically reduces mass ratio. Solar electric rockets such as Dawn's ion engine have an Isp of 3100 seconds. Other electric designs are capable of even greater performance, powered possibly by nuclear reactors. High Isp rockets could reduce the mass ratio required for a mission to mars down to e^(5600/30411):1.2 or less. It would be more than four times smaller than a chemically-fuelled rocket. Atmospheric Gas Scooping This concept consists simply of running a gas scoop through the upper atmosphere and collecting the atmospheric gasses to be used as propellant. Some of the propellant is consumed by the scoop itself, the rest of made available for orbital refuelling of other craft. How a gas scoop might approximately look like We have to work out some orbital mechanics to find out how much gas the scoop-equipped spacecraft gets to keep. In essence, it is ramming a scoop through immobile air and accelerating it to orbital velocity. The scoop loses momentum - it must expel some of that gas as propellant through a rocket to recover that momentum. The momentum a scoop loses by collecting 1 kilogram of air at low altitude is 1*7800: 7800 Newton-seconds of momentum. The momentum gained by using a rocket is equal to exhaust velocity*propellant flow. The exhaust velocity*propellant flow product must be greater than 7800 Newton-seconds. If the exhaust velocity is lower than orbital velocity, the scoop will need to consume more propellant than it could ever collect to stay in space. If the exhaust velocity is equal to orbital velocity, the scoop has to use all of the gasses it collects as propellant and retains none. This is the case for solid-core nuclear thermal rockets. If the exhaust velocity is higher than orbital velocity, then the scoop will be able to retain some of the gasses in reserve and produce a net gain in propellant collected. The ratio between propellant retained and propellant consumed is simple. A rocket like Dawn's ion engine will produce 1*30411: 30411 Newton-seconds of momentum per kilogram of propellant. It can therefore retain up to (30411-7800)/30411: 0.743 or 74.3% of the gasses it collects. The gas retention percentage is simply calculated as: Retention % = [(Exhaust Velocity - Orbital Velocity)/Exhaust Velocity]* 100 We can quickly work out that propulsion systems with an exhaust velocity just above orbital velocity will have a very hard time collecting propellant. An ideal engine has an exhaust velocity several times the orbital velocity. However, as we will now see, getting enough power to these engines is problematic. Powering the scoops against drag Running an electric rocket requires a good amount of power. In space, this is hard to come by. The power requirements of a gas scoop are determined by the thrust it needs to produce, which depends in turn on the drag it experiences. Aerodynamics play no role at orbital velocities, drag is only a function of cross-section and atmospheric density. This equation will give you an estimate within +/-10% of reality. Drag: 0.5 * Orbital velocity^2 * Cross-section area * Gas density Drag will be measured in Newtons. Orbital velocity will be in m/s. Gas density chart up to 300km altitude from here. Cross-section is the frontal surface area (m^2) of the gas scoop ship that intersects with the gasses it is traversing. Some of it will be the scoop opening, into which gas is funnelled, some of it will be closed surfaces that gas bounced off of. A gas scoop ship will want its funnel to cover the entire frontal area to minimize wasteful drag (drag not contributing towards collecting gasses). Therefore, the cross-section area is the funnel opening area. Gas density is determined by altitude and noted in kg/m^3. If the scoop operates at a single altitude, we use a single value. If it changes altitude, we use an average value based on the time spent at each altitude. At very high altitudes, some sources will note a value in particles per cubic meter. This has to be converted into kg/m^3. Particle counts for different elements per cubic meter for altitudes up to 1000km. Notice how mass has nothing to do with drag or the thrust requirements. Gas scoops can be as massive as they want so long as they are able to handle drag. Thrust requirements will not change as it collects gas either, nor will they be lower when the scoop's tanks are empty. So how much power do we need? We first must estimate the drag generated and give the scoop ship enough power to produce sufficient thrust to counter the drag, plus a safety margin. The ISS reboosts its orbit using a Zvezda module's engines. It could be performed by electric rockets instead. Here is a table of drag values listing the drag force per square meter of cross-section area at different altitudes, for 7800m/s orbital velocity, using data from this NASA website: We can see that above 200km altitude, the drag force is in millinewtons, descending down to tenths of micronewtons per square meter at the edge of the atmosphere. A more accurate reading would take into account the small differences in orbital velocity as the altitude changes. It is 7800m/s at 150km altitude, but reduces slightly to 7350m/s at 1000km altitude. The equation for engine power is the following: Engine Power: Thrust * Exhaust Velocity /2 If we combine it with the drag equation, we obtain this: Engine Power: Drag force per m^2 * Cross-section * Exhaust Velocity /2 Engine power is in Watts. Thrust and Drag force per m^2 are in Newtons. The Cross-section area is in m^2 and the Exhaust velocity is measured in m/s. Example: 10m wide funnel at 200km Cross-section area: (10/2)^2 * 3.14 = 78.5m^2 Drag force per m^2 at 200km: 9.51 * 10^-3 Newtons Exhaust velocity: 30411m/s Engine power: 11.35kW 100m wide funnel at 400km Cross-section area: (100/2)^2 * 3.14 = 7850m^2 Drag force per m^2 at 400km: 1.03 * 10^-4 Newtons Exhaust velocity: 30411m/s Engine power: 12.3kW The engine power here is the effective power output going out of the nozzle. Rocket engines are not 100% efficient, and there are further losses in the systems that generate, transport and convert electricity going into the engine. These might double the actual power consumption. There are many options for producing the required energy to power the engines, but only a few are practical or achievable with the technologies available in the near future. Apollo's fuel cells. Chemical energy, such as in a fuel cell, is not a viable option. The system masses too much and requires constant refuelling. Radio-isotopes are compact and reliable, but their output is very, very low at only about 2 to 5kW per ton. The two options remaining are solar panels and nuclear reactors. Solar panels are cheap and lightweight, able to produce over 580W/m^2 at peak efficiency. However, they can become very large if we need kilowatts of power. They must be packed behind the main funnel or they would increase the cross-section area. A 10kW array of solar panels might be divided into eight segments 2m wide and 1.07m long. Thin-film solar panels might reduce mass down to a few dozen kilograms. Nuclear reactors are very powerful are unlikely to need much mass or volume to produce the power output required. Experiments have been conducted and prototypes have been flown of space-rated nuclear reactors, but historical and political reasons have prevented their widespread use so far. We can send a 100kW reactor today into space massing only 512kg. Gas collection and processing A turbomolecular pump A funnel scoops up gasses by ramming through the air and compressing them into a tube. In that tube, pressures rise until it reaches stagnation pressure. This is the pressure when the gasses have been completely stopped relative to the scoop. Because the scoop is travelling so much faster than the particle velocity inside the gasses being collected, it can act as a giant piston like in a turbomolecular pump. The scoop naturally compresses gasses while they cannot escape back out of the opening. The first step is to move the pressurized gas into an empty gas tank. It can be vented without any pumping equipment and only a valve to maintain the flow going one way. As long as the collection tanks are kept at a lower pressure, no active pumping is needed. To keep collection tanks at a lower pressure, the gasses must be cooled until they become liquid, then they are transferred to an insulated long-term storage tank. This is the second step. Air intake and compression are tasks performed by the scoop during its operation. Cooling the collected air until it liquefies allows for its fractional distillation into different gasses. The gasses have different boiling points and separate into different fluids as the temperature drops. The relative proportions of each gas this fourth step produces is determined by the atmospheric composition at the collection altitude. Number of N2 particles per cubic centimeter at altitudes 100 to 1000km Number of O2 particles per cubic centimeter at altitudes 100 to 1000km Number of atomic nitrogen N particles per cubic centimeter at altitudes 100 to 1000km Number of atomic oxygen O particles per cubic centimeter at altitudes 100 to 1000km Number of atomic hydrogen H particles per cubic centimeter at altitudes 100 to 1000km As noted before, we must convert particle counts into kg/m^3 densities. For example, at 200km altitude, we note that we get 3.8*10^9 N2 particles, 3.25*10^8 O2 particles, 4.39*10^9 O particles, 9.62*10^6 N particles and 2.15*10^5 H particles per cubic centimetre. We multiply the values by a million to get the cubic meter values, then convert the particle numbers into masses by multiplying by the element's molar mass in g/mol then dividing by Avogadros' constant (6.022 * 10^23). We can work out the following figures: N2 particle mass: 28.0134 / 6.022 * 10^23 = 4.65 * 10^-23 grams N2 mass per m^3 at 200km: 1.78 * 10^-16 kg O2 particle mass: 31.9988 / 6.022 * 10^23 = 5.313 * 10^-23 grams O2 mass per m^3 at 200km: 1.72 * 10^-17 kg O particle mass: 15.9994 / 6.022 * 10^23 = 2.65 * 10^-23 grams O mass per m^3 at 200km: 1.16 * 10^-16 kg N particle mass: 14.0067 / 6.022 * 10^23 = 2.325 * 10^-23 grams N mass per m^3 at 200km: 2.23 * 10^-19 kg H particle mass: 1.008 / 6.022 * 10^23 = 1.67 * 10^-24 grams H mass per m^3 at 200km: 3.59 * 10^-21 kg What can we understand from these figures? The atmospheric composition at 200km is dominated by nitrogen and monoatomic oxygen. Density when all gasses are considered is 3.127*10^-16kg/m^3... nitrogen therefore represents 56.9% of the mass of air collected at this altitude, with monoatomic oxygen second at 37%. The relative proportions of the gasses allows us to estimate the average heat capacity; it should be about 1.2kJ/kg/K. Stirling engine animation It is important as it tells us how much energy will be needed to reduce the temperature of the gasses down to their liquefaction values. A Stirling-engine cryocooler is suited to the task, with a pump efficiency of roughly 40%. Here is a table of the temperatures, energies and power consumption levels involved in liquefying air collected at 7800m/s for altitudes between 100 and 1000km: Note that the power ratings are in milliWatts per square meter of scoop cross-section area. It is the electrical consumption for a heat pump cooling down the gasses from the collection temperature to 77K, where nitrogen liquefies. For 200km, we read that 19.48W per square meter is required. A 10m wide scoop would need to devote only 1.5kW to cryocooling at this altitude. Gas scoop design example We will attempt to work out a ballpark estimate of the mass, size and performance of a conventional gas scoop that uses currently available technology. This simple design operates at one altitude and continuously uses its engines counter the drag forces of ramming through the air. We will stay within the limits of a payload that could be launched by a single Falcon 9 FT rocket to LEO, so 22.8 tons to 200km altitude. Let's start with the engine. We want it to use the gasses it collects as propellant. This excludes current electric rockets propelled by xenon or argon. Instead, we need to find examples of designs that are happy to run on nitrogen. Oxygen is viciously corrosive when hot, so should be avoided as a propellant despite its relative abundance. Diagram of the electrodeless RF plasma thruster. Looking at this engine list, we find the electrodeless RF plasma thruster. It has the ability to accelerate any propellant to extreme velocities at high efficiency, without the temperature or endurance limits of other electric rockets. 23km/s exhaust velocity with nitrogen as propellant can be achieved by a plasma with an electron temperature as low as 5eV, and an ion temperature of 25eV. The VASIMR rocket engine by AdAstra. The most similar engine we have mass and efficiency data on is the VASIMR, with 580W/kg in low gear and 60% overall efficiency. We can work out that each kW of electrical power supplied to the RF plasma thruster produces 0.052N of thrust out of 1.72kg of engine. As mentioned before, solar electric power is the most realistic option. Supplying 1kW of electricity using thin-film technology requires 2.02m^2 of solar panels and about a kilogram of electrical equipment after including 15% losses during electrical conversion and transport to the engines. Putting all these figures together gives us a figure of 34kg and 2.02m^2 of power and propulsion per Newton of thrust. At 200km altitude, the orbital velocity is 7790m/s. Air density is 3.127*10^-10 kg/m^3. The drag force per square meter of scoop cross-section area is 9.487*10^-3 N/m^2. Using the power and propulsion figures, we determine that each square meter of scoop area requires roughly 0.326kg of equipment to counter drag forces. A drag force in the range of milliNewtons allows for the use of very lightweight scoop materials. The closest analogy would be a hemispherical balloon holding up to a pressure differential, with structural support lines transmitting load or tension forces to a payload. Stratospheric balloons survive 200Pa pressure differentials and carry several tons using membranes massing 55 grams per square meter. We might be able to say that despite the hundred-thousand-fold decrease the in the forces involved we have decided to include massive redundancies against micrometeorite impacts and friction ablation so the scoop will mass 0.05kg/m^2, similar to a stratospheric balloon. Processing the gasses requires some more equipment. 2.44 micrograms per square meter per second. are collected at this altitude and velocity. 94% of that is N2 and O, reducing the collected mass to 2.29 * 10^-9 kg/m^2/s. It requires a heat pump and gas handling system able to handle 18W of power. SABRE reaction engines claims to be able to handle 400MW of cooling with a one ton device, but we will use a much more conservative 400W/kg. 18W of cooling would require 0.045kg of equipment. The total so far is 0.426kg/m^2. We want our gas scoop to collect gasses for one year, then offload them. We therefore need storage tanks able to handle liquid oxygen and nitrogen for one year. They store at similar temperatures, so we will use 5% of the propellant mass as tank mass, something based on this NASA report for the Jupiter launch vehicle. It is based on Aluminium-Lithium materials, but even more lightweight carbon fibre is possible. World's largest carbon-fibre tank for holding liquid oxygen. Liquid oxygen has a density of 1141kg/m^3. Liquid nitrogen has a density of 808kg/m^3. Over a year, a gas scoop orbiting at 200km altitude produces 4.37*10^-2kg of liquid nitrogen and 2.84*10^-2kg of liquid oxygen, per square meter of scoop area. This requires a combined storage volume of 7.89 * 10^-5 m^3, massing a negligible 3.65 grams. Due to the 23000 to 7800 ratio between the engine's exhaust velocity and orbital velocity, 33.9% of the gasses collected must be expelled as propellant. This will reduce the amount of nitrogen collected from 56.9% to 22.98% of the total mass of gasses collected. This reduces the nitrogen storage volume required by a factor 2.47 and the yearly tank mass to 4.67 * 10^-5 m^3, and the mass to 2.3 grams. With all functions accounted for, the scoop will require 0.428kg of equipment per m^2 of cross section area. If we double this figure to account for everything including attitude thrusters, docking structures, thrust frames, communications, propellant handling, safety margins and so on, it is still about 0.857kg/m^2. Now we can derive the maximum size of this scoop and its yearly performance. 22800kg / 0.857kg = 26604m^2 A single Falcon 9 FT can put a scoop with a cross-section area of 26604m^2 into orbit. It would represent a disk 184 meters wide, although multiple smaller scoops would be more sensible. They would be equipped with thrusters producing a total 252 newtons of thrust and use 509m^2 of solar panels. These scoops would collect 2046 tons of gas, of which 471 tons of liquid nitrogen and 757tons of liquid oxygen is retained, for a total of 1228 tons. 'diver' establishes an elliptic orbit, with the lowest point dipping into the atmosphere. For most of the orbit, a small, efficient electric rocket slowly accelerates the scoop. It then descends rapidly until it hits the atmosphere, using its excess velocity to compensate for the increasing drag. This allows it to 'dive' into the denser parts of the atmosphere and quickly collect a lot of gas. Momentum carries the scoop back up and out of the atmosphere, giving it a lot of time to regain the lost velocity and process the collected gasses. Divers are the simplest and most robust design, with the lowest energy and mass requirements. However, they have the lowest endurance. What's the point of all this? Remember the mass ratio equations at the beginning of this post? A mass ratio of 15.9 to reach orbit can be understood as each kilogram of propellant in orbit being worth 15.9 times more than the same kilogram on the ground. 1128 tons in orbit would be worth 19525 tons on the ground. These thousands of tons could be saved without having to capture an asteroid, mine the moon or invent a radically new launch system. They alone would make the entire concept of atmospheric gas collection worthwhile. Understandably, the gasses collected cannot be simply loaded into rocket propellant tanks and expected to burn. Liquid nitrogen is inert and cannot be 'burned'. Liquid oxygen needs a fuel, which can be liquid oxygen, methane or kerosene lifted up from the ground or already on-board the spaceship expecting to be refuelled. In a typical LH2/LOX rocket engine, six kilograms of oxygen are consumed for every kilogram of hydrogen. Looking at it another way, for each kilogram of liquid oxygen a spaceship receives through orbital refuelling, it needs to carry along 0.167kg of liquid hydrogen from the ground or separate launches. A 10 ton dry mass spaceship being sent from low earth orbit to geostationary orbit would require a deltaV of 3931m/s. If it uses a LH2/LOX 450s Isp rocket engine, it would require 14.3 tons of propellant. It would need to bring up 2.39 tons of liquid hydrogen from the ground, but would save on 11.96 tons of liquid oxygen propellant. The lifter on the launchpad would be 49% lighter. A single Falcon 9 FT-launched scoop could provide enough liquid oxygen to push 633 tons of payload from LEO to geostationary orbit per year. For comparison, a 353s Isp kerosene/liquid oxygen rocket with an oxidizer/fuel ratio of 2.56 would reduce lifter mass on the launchpad by 52% and is enough for 498 tons to GEO. There are further savings if we consider that that launcher would need smaller propellant tanks, structures and fewer engines when it is nearly two times smaller. LEO to GEO is a common requirement for satellite launches. These savings might allow for much more realizable Lunar or Martian missions. Other secondary benefits exist. The name 'Orbital Transfer Vehicle' or OTV is given to the rocket stage that moves a payload from low earth orbit to geostationary or lunar orbits. EDS stands for 'Earth Departure Stage' for payloads being sent to Mars or elsewhere. An electric or nuclear-thermal OTV or EDS can use all of the gasses collected (1228 tons) and can essentially launch with empty tanks, meaning that the full advantages of orbital refuelling can be reaped. The availability of hundreds to thousands of tons of propellant waiting in orbit will incentivise the development of non-chemical propulsion technologies. Alternative designs and improvements The two heaviest components of an atmospheric scoop are power and propulsion equipment. Here are some solutions to reducing their mass requirements so that even bigger scoops which can collect even more gas per year can be launched on the same launcher: -Use all the nitrogen: If atmospheric gas scooping is going to provide oxidizer to spaceships in low orbit, then only liquid oxygen is of interest. All the liquid nitrogen can be converted into propellant on-board the scoop. At 200km, this allows for up to 57% of the gasses collected to be consumed, meaning Isp can be as low as 18km/s and thrust per kg will increase by a third compared to our example. Furthermore, if the nitrogen is being consumed as it is collected, then there is no need to liquefy it and less cryocooling equipment will be needed. -Use better power density and space-grade components: In our example, the power density of the electric rocket engine was 580W/kg, cryocoolers were 400W/kg and we doubled the mass per square meter in the end. Specialized technology using space-grade materials could greatly reduce the mass per square meter. -Trawler design Instead of fish, it would be collecting gasses from a lower altitude. To fulfill all the energy needs on-board our scoop design, we needed to equip it with 509m^2 of solar panels. Keeping these panels and other equipment hidden behind the scoop's funnel is challenging, as it elongates the design and increases structural requirements to prevent bending and flexing. If the main body of the design and its scoop funnel were physically separated, the overall design could be lighter. An arrangement where the funnel is dragged at a lower altitude by tethers and pumps connecting it to the main body is called a 'trawler', as it resembles boats dragging along nets under the sea. For one, solar panels can be arranged any way desired. Second, a scoop traversing lower altitude air can collect more gas per square meter, allowing for a smaller cross-section for the same performance and lower mass per square meter. Third, the scoop ship does not have to consume a lot of propellant to lift itself out of a lower altitude if it wants to stop scooping: it can simply winch up the funnel. The disadvantages are greater complexity and the need to pump gasses from where they are collected to where they are processed and stored. -Diver design A conventional or trawler gas scoop design needs engines to counter-act drag with an equal amount of thrust at all times. Solar panels must provide sufficient energy and store it to keep propulsion running around the night side of Earth. There is a design that allows a scoop ship to get away with much smaller engines and power sources. A 'diver' operates at an elliptical orbit, dipping into the atmosphere only at its lowest point. At that point, it rapidly scoops up gas and rams its way up and out with its momentum. For the rest of the orbit, it fires up its small engine to recover the lost momentum over a long period of time. Much smaller solar panels would be needed to feed propulsion and gas processing equipment. If the diver goes into the lower atmosphere with excess velocity, like an aerocapture maneuver, it could use a comparatively tiny scoop to collect a lot of gas, so overall the diver scoop will be much smaller and lightweight compared to other designs. If it needs more time, the scoop can perform one diver every several orbits. The downside is that ramming through the lower atmosphere at higher than orbital velocity involves significant heating which lowers the ship's useful life. Collecting gasses quickly means that some intermediary store must be available before they are processed and liquefied: holding large volumes of hot gas would require voluminous tanks that add drag and weight to the scoop ship. -Electrodynamic tethers Scoop ships spend a lot of time in space, have a lot of electrical power available and only need tiny thrust output to maintain their altitude. Electrodynamic tethers are perfectly suited to atmospheric gas scooping operations. They push against the Earth's magnetic field to provide a form of propulsion that does not require any on-board propellant. Electrodynamics tethers can supplement or replace electric rocket engines. -Electric scoop Between 60 and 1000km altitude, Earth's atmosphere stops being a continuous medium and becomes a loose plasma dominated by charged particles. This is called the ionosphere. For example, monoatomic oxygen is a negatively charged particle that represents a large fraction of gasses at 200k altitude and above. These ions can be collected by an electromagnetic/electrostatic scoop. The electromagnetic section is composed of large magnetic fields that direct ions towards the ship from a large volume of space. It would be similar in operation to a Bussard Ramjet. The electrostatic section is composed of extremely thin rings with a charged interior. Ions of the unwanted charge are deflected, ions of the desired charge are pulled to the center, where a physical scoop collects the gas. The main advantage of this design is that it allows a tiny scoop and a set of magnets to do the work of a much larger funnel. An interstellar voyage by the Bussard Ramjet An electric scoop would be the only practical way to collect significant amounts of hydrogen from altitudes above 400km. Scoop cross-sections several kilometers in diameter would have the same performance as a scoop a hundred meters wide at lower altitudes. With an electric scoop, such diameters are possible without scaling the mass to cross-section area.

The warm areas to live in on Titan will be heated by human metabolism and heaters. Humans consume food, which takes electricity to produce, heaters, use electricity directly. This electricity must be generated from another source. Stirling engines used as generators must work with a temperature gradient going from hot to cold. There are no good heat sources on Titan!

Accelerating a spaceship so that it reaches Titan requires energy, either in the form of propellant or a reactor. That energy is best used on Titan itself. Methane fuel requires oxygen oxidizer, which is not available on Titan. Plants are a method of converting solar energy into chemical energy. There is very little sunlight on Titan, and plant efficiency is extremely low (about a 3% percent or less). If fusion energy is available, then it would solve energy needs everywhere.

Hi Kosmonaut! I believe deuterium is plentiful in our oceans as an isotope of hydrogen contained in 'heavy water', and being able to produce tritium out of lithium industrially does favour D-T over D-He3 until we start mining the Moon. 20% of the speed of light is for interstellar travel! It would also need extremely high exhaust velocities, since you'd need a minimum deltaV of 120000km/s. Even with a mass ratio of 1000, the exhaust velocity is 17300km/s! D-He3 cannot produce something that fast! You'd have to be running a particle accelerator as an engine with terrible efficiency and absolutely pitiful thrust. Aerobraking, which can maybe shave off 10km/s from the trip if you use massive heat shields, shouldn't even be in the same book at 0.2C. Maybe if you greatly reduced the speeds the spaceship reaches? At 500km/s, you're still cutting down the trip from Earth to Jupiter to two weeks... The difference between the lift produced by a vacuum balloon and a hot hydrogen balloon is very very small, at least in the rarified atmosphere of a gas giant. Saturn has a gas density of 0.19kg/m^3. A vacuum balloon can lift 0.19kg per m^3 of balloon. A hot hydrogen balloon might be able to lift 0.15kg/m^3, approaching 0.19kg/m^3 as you increase the temperature. Very small difference! A lofstrom loop has to cover thousands of kilometers of land with a 'danger, risk of falling accelerator' signs. People won't want to live under it or near it and it would cost a lot to build. A space fountain has a much smaller footprint... but that's not a big advantage in the long run, I agree. I doubt my stream-rider post will appear this month! I've got a lot of my plate and other subjects queued, so don't hold your breath D-T fusion is the easiest and using a neutron-absorbing shell around the fusion products, either on the fuel pellets themselves or on the reactor walls, allows you to recover most of the fusion energy. Proton-Boron11 is aneutronic and powerful, and you could deal with current ignition issues by using massive proton accelerators to bulldoze your way past any obstacle, but you'll have to deal with X-rays leaking and wasting a lot of the fusion energy. Tokamak is a torus-shaped, closed reactor. Its a glorified coal boiler to provide heat for electrical generators. You could use the electricity to power and electric rocket. A fusion rocket is an open reaction chamber. Its a rocket engine with a very hot core and lots of propellant coming in through the top and out the bottom. The designs are incompatible! One or the other.

Nice write-up as always. Loving the artwork! A few notes: -Floating in gas giants doesn't necessarily require vacuum balloons. Hot hydrogen works well enough and would allow for equalizing the pressure between the inside and outside of the balloon, allowing for vastly thinner envelope materials. -Personally, I am skeptical of Lofstrom loops. They are massive moving structures requiring constant upkeep and expensive hardware running at high temperatures throughout. Every single aspect is risky, costly or both. A space fountain at least is much smaller, and Robert L. Forward's suggested modification is even better: shoot the mass stream straight upwards and have spaceships ride them. Vacuum tubes at the bottom protect them from too much drag, and the stream is deflected downwards at the top. My own modification would be not to bother at all with a physical structure at the top. Just angle the stream to 45 degrees through a vacuum tube 10km tall and 10 long and you can have spaceships just ride the stream to a large fraction of orbital velocity. I'll write up a 'stream rider launch system' blog post soon. -Orbital rings can have a spinning inner section and a stationary external hull. The hull can be connected to the ground with towers. More importantly, you can portray the immobile exterior of the ring around Kerbin. For the interplanetary rocket: A tokamak is a closed 'fusion bottle'. The heat of the fusion reaction must escape somehow. Tokamak designs have the heat be absorbed by a high-temperature coolant loop, run through a heat exchanger, and into a coolant loop of containing a pressurized fluid just above boiling point. The latter fluid is allowed to boil and expand through a turbine for electricity. It is the electricity which is then used to power the rocket engine. So basically, the tokamak is you big electricity generator and the rocket engine is how you use the electricity. Alternatives include the use of thermocouples moving heat from the tokamak wall to the radiators, or a magnetohydrodynamic generator that bleeds off plasma from the reactor.

No support for RSS: is this a final decision?

Hard to say. The limits of the KSP engine are currently your biggest obstacle to presenting fascinating technologies such as space elevators and aerovators. Perhaps you could start on that turn-based model-representation game?