MatterBeam

I mentioned importing Nitrogen from elsewhere. It costs less in terms of kJ/kg to extract oxygen from Mercury's crust than to ship it in from other planets. If you've got a plan to create a magnetic field stronger than Earth, and move thousands of tons of material and equipment from other planets to Mars, then digging a bit deeper to get oxidizer metals is trivial.How does sunlight ionize metals underneath the surface anyway, let alone 100m deep anyway? It is possible to terraform Mercury. You just need a lot of energy. There is vastly more oxygen than necessary, locked up in metal oxides on Mercury, than is needed to create a thick atmosphere on Mercury. A rough estimate puts 300km/2440km:~12% of Mercury's mass in the crust. That's 4e22kg of crust. Earth's atmosphere is contains 1.2e18kg. We need 2.6 times more for Mercury's atmosphere, so 3.12e18kg. Therefore, you only need the crust to contain 0.078% oxygen by mass to fulfil your needs. You'll find the actual fraction to be closer to 40%+.

@PB666: You are a very engaging read, but in many of your posts, you get lost in a storm of assumptions and successive conclusions. Above is one of them. Here's what I have to note: -I presented the trip calculator (an excel file) on that link. The maths works out, whatever the set of numbers you put in. And I'm not defending the author's numbers. -Your argument about the potential of fusion vs the weight of the components.... is just another way of saying kW/kg. If your fusion reactor has 0.1kW/kg, it loses out to existing solar panels at 0.3kW/kg, before other complications are included (fusion fuel availability, sunlight intensity, ect). -Propulsion power density should include the mass of all the systems that create, convert and use power, as well as the waste heat management systems. Using further subdivisions of the components (reactor, radiators, engines) is much less useful, and rather pointless as they will all be mostly proportional, so they can all be brought into the same kg sum per kW of output. -Area and volume has nearly zero relevance in space. Space is free, mass is not. The volume of the engines will have nearly zero effect on the mission performance, only their mass does. -Solar panel power density is usually rated by output, not input. -1MW of input corresponds to 8.55x85.5m with 1367W/m^2 of sunlight. 10x100feet corresponds to 127kW of sunlight. This is why everyone hates imperial. -At 32% efficiency, you's produce 320kW, so I think that's a typo. -Where did you get '1 ton' from? If we look at current developments, the triple-junction solar cells that can get 32% efficiency mass about 0.85kg/m^2. That 1MW input array would mass 621kg in total. The structural mass you'd need on top of that figure would be about 1kg/m^2 for 731kg, but it can also be ~3000 times lower than that for any solar panel lifted up from Earth. If it is meant to only stay in space and handle the gentle accelerations of a milligee craft, it will never need to structural support to survive the 3g+ peak accelerations and vibrations of a launch vehicle. -If we work only with the power density figure, we can easily work out that an advanced multi-junction solar panel would produce roughly 320/(621+731): ~236W/kg. A bit better than what is being achieved today (150W/kg). -Why would the panels on a solar electric craft have to be retractable? -The ISS solar panels are a terrible example. They're old, use old technology, are built for sturdiness and reliability and to be easily replaced by an astronaut in EVA, and only produce 27W/kg. -If the ion drive matches the power density of the solar panels, then you're halving the system power density. If they provide a better power density, then you're approaching the power density of the solar panels. If they're lower, then you're losing system power density. The relationship is 1/x = 1/y + 1/z. X System Power Density, Y Panel Power Density, Z Engine Power Density. VASIMR by Ad Astra published 200kW/620kg including radiators and other sub-systems, or 0.32kW/kg. Put together, we get 1/236+1/320: 1/136. The overall system power density would be 136W/kg. From here, we can calculate all the mission parameters. It is a much simpler approach than calculating truss diameters and the N/m^2 rating of ion engines. -The acceleration figures you ended up with are purely the result of the assumptions you came up with yourself. -Extremely high Isp is not a good thing. The savings on propellant are inversely proportional. If you halve your propellant load by increasing Isp from 1000s to 2000s, you'll get less savings by doubling again from 2000s to 4000s and so on. Propellant is the cheapest part of a spaceship. Lowering the Isp until a decent balance between mass ratio and average acceleration is what actual rocket design should do. It is an iterative process though, and real world designs are subject to constraints outside of the scope of this discussion (what the year's NASA budget is, what the diameter of the fairing is, what are the current president's preferred space objectives and so on). -Getting to Mars in 89 days requires about 25km/s of deltaV. With a 3500s Isp engine, that's a mass ratio of 2.04. I don't see where the physically/practically impossible part is.

Phase-change cooling seems to me to be more and more effective the closer I look into it. Expect some interesting designs to come from this! Quite right. The distinction is trivial though, when you can only detect these ships at distances of a few km, while combat ranges are measured in thousands of km. In fact, if the engineering constraints make the hydrogen ship a lousy stealth ship than can be detected at 1000km, then it's actually fine as long as your weapon range is 2000km.

Oh it's certainly possible. You'd need tremendous amounts of energy. The first thing to do is unlock the nitrogen and oxygen locked in the planet's minerals. Earth's crust should not be terribly different from that of Mercury, so you can expect nearly half of every kg of rock to be oxygen. You'd then import nitrogen from other sources in the Solar System, such as Titan. You'd need a 2.6 times thicker atmosphere to create 1atm on the ground. You will have to do this continuously too. Earth loses about 90 tons of atmosphere every day. At the same temperature but with a 0.38g gravity, you'd lose 236 tons per day. To protect Mercury, you'd need a strong magnetic field. Mercury is a 17.8 times smaller volume to cover with a magnetic field than Earth, but it is 13 times closer to the Sun. So, you need to generate a magnetic field about 38% stronger than that of Earth. Then, you'd need to handle the sunlight. On one side, you need to reflect away 92% of the incoming sunlight. The hard way to do it is to create a very transparent filter with a high refractive index, that you stack until 92% of the sunlight bounces away and 8% goes through, while nearly nothing is absorbed by the filter itself. It might look like a Bragg (dielectric) reflector and made of quartz. The easy way is to create billions of small rotating shutters that open up 8% of the time. It'll look terrible and give you a headache though. That should be enough. A barren, dry world where you slowly die of microgravity sickness while the light blinks like a badly tuned fluorescent line above, but hey, at least you can take your helmet off!

Those are just illustrations to show how the various components are arranged in relation to each other and the direction of travel. Don't take everything so literally, don't read too much into pictures! Mirrors, radiators? Please clarify your second point. Just to be sure, the mirrors do not have anything to do with radiators. Ions can definitely be used for fast interplanetary travel. In fact, if you can travel several times faster than the biggest chemical rockets even if your acceleration is measured in milligees (1/1000 of a g). This is because although the departure from Earth takes weeks instead of minus, you can keep accelerating to higher velocities. Use the trip calculator provided here: http://www.rocketpunk-manifesto.com/2011/08/mission-to-mars.html You will notice that even if you input modest figures, such as 1kW/kg, 25% of dry mass dedicated to propulsion, 3500s Isp and so on, you'll end up with a 289 ton spacecraft that sends 100 tons of payload to Mars in just 85 days. This is departure from Earth and insertion in low Mars orbit. It has a mass ratio of 1. A 450s Isp chemical rocket with a similar mass would put only 80 tons around Mars, and will take 8.6 months to do so.

The whole point of the 'hydrogen steamer' is that it uses hydrogen boil off to cool the hull down to temperatures where it emits so little radiation that it is undetectable. Hence, no infrared signature.

Solar pressure on a 1000m^2 reflector would be at most 0.018N. If the reflector massed just 7kg, an acceleration of 0.01g would impose a force of 0.67N, which is about 38 times greater. As the reflectors get lighter and lighter and acceleration lower, solar pressure plays a bigger role. You can create a sunshade to cover the wires, at the cost of blocking out part of their radiative surfaces. Or just accept the inefficiency and increase the cooling capacity. Either way, why would you have superconducting wires dangling in space?! In both designs I described, the photovoltaics are embedded deep within a structure inside the spacecraft. Only the reflectors and the optics are exposed to space. I promise you all, in Part 3, I will propose a detailed design of a solar-powered spacecraft meant to travel from Earth to Jupiter. The 'telescope-like' design I have in mind is closer to the Newtonian telescope. Using an non-imaging surface, I can get even get light to be focused coming in from the sun at an angle, so that the secondary mirror does not block any of the sunlight. Look at type (h). We definitely need to build these things in space. As I said in the blog post, the key to the performance of future solar systems is to maximize on the thinnest reflectors possible and minimize the solar panel area required. The engine just needs to point through your center of gravity in the direction of acceleration you need. If you put your propellant tanks and engine in the middle, solar reflector on one side and payload on the other, you can create a spaceship where aligning these things doesn't matter. Examples of sideways-on-a-stick designs: Exactly this. It all boils down to the power density.

I've been looking around for nozzle calculators I got the following results: Hydrogen, gas constant 4184, input temperature 2330K, ratio between nozzle throat area and exit area: 30000, exit temperature 22K, exit velocity 8.2km/s. So, a nozzle fitted onto some sort of nuclear thermal would allows for a near undetectable exhaust, while propelling the rocket with an Isp of 835s. If we want a thrust of roughly 100kN, we'd use a nozzle throat 7.6cm wide at a pressure of 1atm. The resultant nozzle exit diameter is 13.1 meters. Higher input temperatures are acceptable for moderate increases in exit temperature. 3000K input, with the same parameters, allows for an exit temperature of 28K and 9.39km/s exhaust velocity. Increase the expansion ratio to a ridiculous 100000, and you can even use high temperature 4000K hydrogen for 10.8km/s exhaust velocity and yet only release it at 23K. At that point, you're absorbing over 80MJ of heat per kg of hydrogen. I've done the calculations above. Expansion ratios, measured in nozzle exit area over nozzle throat area, will be at least 30000. To convert into a ratio of diameters, divide by 3.14 and calculate the square root - in this case, it'll be (30000/3.14)^0.5: 173 The coilgun/mass driver propulsion option is a good idea. The only downsides are the extremely poor thrust to weight ratio (compare the mass of the magnets, cooling systems, reactor, energy storage and so on, to a simple Nuclear Thermal Rocket) and a much lower amount of heat absorbed per kilogram of hydrogen onboard.

Hi! Rigidity through centripetal force is definitely an option, but you will have difficulty preventing the sails from bending under the slightest acceleration. This is because the sails are so thin that they are not very strong. That means they cannot support a large centripetal force which would come from spinning the sails more rapidly to prevent bending. Static electricity is another option, but remember that space is not empty. Solar winds and the interplanetary medium would interact with this charge, causing electric arcing and a lot of damage to the sails. If you reduce the charge, the repulsion between same-charged-surfaces becomes weaker and less effective. So yeah, these options are doable for a solar sail, but for a solar reflector that accelerates much more quickly than a solar sail, you'll have trouble.

The thermal conductivity of gasses is about five to ten times lower than for liquids, but that only increases the size of the heat exchanger, not the overall efficiency. Peltier junctions have 10-15% the maximum efficiency of a heat pump, while the Rankine systems you mention can achieve 40-60%... You can always run the hot exhaust over a heat exchanger and carry only the heat inside.

Sorry, the part balancing and mod implementation is up to you!

I'm just thinking of alternatives to Stirling heat pumps for moving the heat from a cold gas coolant into a heat exchanger that leads into a radiator. Currently, it is very difficult to design a system that wants to run cold and handle a lot of power in space. Examples include particle accelerators, laser generators, ion engines, ect. They run below room temperature, down to 100K or less when superconductors are involved. And yet, kilowatts to megawatts of waste heat are generated by their operation. Dealing with that waste heat passively means using a set of radiators at even cooler temperatures. Due to the low radiation per square meter, the radiators would have to be massive. Not something a spacecraft can afford! An ideal option would be increase the radiator's operating temperature. For this to happen, you need heat pumps to move cool heat up a temperature gradient - you need heat pumps. The problem is, current heat pumps have poor efficiency and horrendous power density. The mass you save on radiators, you'll lose it with the pumps! Brayton-cycle turbines can combine high power density and high efficiency. If they can be run in reverse as heat pumps, then you can solve some of the problems involved with running high-energy cold systems in space. Here's a 1MW example: An VASIMR engine needs to be kept at 273K temperature for its magnets to operate at peak efficiency. It is 60% efficient and has a power density of 1kW/kg. To handle 1MW, it needs to mass 1 ton. This engine produces 400kW of waste heat. A passive heating solution would be use radiators that run at 273K. The Stefan-Boltzman equation says that even very black radiators (emissivity 0.9) would emit only 283W/m^2. You'd need an area of at least 1413m^2. Even if the radiators are radiators are relatively lightweight and double-sided, you'd need to dedicate 3.5 to 7 tons just for radiators. Let's introduce an existing heat pump. A stirling-cycle heat pump has an efficiency of up to 35% and a power density of 0.1kW/kg. We will use it to increase the temperature to 600K, so its thermal performance will be 83%. Combined with the pump's own efficiency, you need 3.4W of input to move heat up that temperature gradient. It would allows for the radiators to be 23.3 times smaller and mass only 150 to 300kg. However, to handle all of the waste heat, you need heat pumps that consume 1360kW of power and mass 13.6 tons! Not only are you losing all the gains from the smaller radiator, but you are also more than doubling your energy budget! How does it go differently if we used a heat pump based on the Brayton cycle and built like a modern turbine? It could be up to 80% efficient and have a power density of 10kW/kg. It only needs 1.5W of input per 1W of heat moved, and it will mass only 60kg. Quite a difference! That is the solution that this 'reverse ramjet' is for.

Ramjets. In reverse. We all know that ramjets work by compressing incoming air, which heats it up, then injecting fuel and burning it, so it heats up even more, and then expanding it through a nozzle. The net energy gain from the fuel burning is translated into warmer gasses exiting the nozzle at a velocity faster than at the intake, at roughly the same pressure as the exterior. The ramjet is therefore a heat engine. Could we run a ramjet in reverse, turning it into a heat pump? A heat pump takes heat from a cold source, and moves into a warm sink. It expends energy to move heat against the temperature gradient. Let us consider a ramjet that is meant to move heat from a cold source (cold gas) into a heat sink. The reverse-ramjet would accept the cold gas at the intake. It compresses it, raising the temperature. The temperature gradient between the gas and heat sink is reversed, allowing heat to flow out of the gas. Then, the gas re-expands in the nozzle. If it is expanded to the same pressure as the inlet, it ends up being colder, as it has lost heat in the compressor. The energy to run the ramjet comes from accelerating the cold gas, probably from a fan.

For superconducting wires, there is no internal resistance, so no heat generated when current runs through the wire. All heat will come from external sources, primarily sunlight. Insulation reduces the amount of sunlight absorbed and reduces the rate at which that heat reaches the wire itself. Even for non-superconductors, like a thick length of aluminium, heating should be minimal. What helps is that many solar cells in series produce a high voltage, while the distance to be traversed from the solar cell to the engine doesn't have to be very long (a dozen or so meters at most) because as I've said, the solar collectors are not directly connected to the solar cells. Let's say we have a 100MW to deal with. We'll set the voltage to a modest 100V, stepped up to 10kV. The current is therefore 10kA. What combined thickness of aluminium wires should we use to carry this current over a distance of ten meters without suffering more than a certain rate of heating, matched by the cooling capacity? Let's imagine passive cooling fins attached to the wires. Four fins of 20cm width would have an area of 8m^2, but only 5.6m^2 effective after interreflection. At 300K, it radiates 2.44kW. This means that with a current of 10kA allows for a resistance of 24.4microOhms. This means wires as thin as 0.01m^2 are possible. Using better cooling allows for thinner wires. The liquid helium/hydrogen comment was with aluminium superconductivity in mind, which requires 1.2K temperatures. I should have clarified.

This should be in Mod development, but here's my suggestion: Use the USI-LS mod files. Go to the converter that consumes Supplies and produces Mulch, and add heat generation to it, just like the ISRU part does. Then, add to all command pods a generator that consumes electricity, produces heat but has no 'output'. If you want to use your own heating system instead of the stock heat generation mechanics, add a weightless resource called HEAT and create generators that produce it as an output resource.

The forces involved are what matters, and they're quite low when you're accelerating at 0.01g. The trusses might end up being decently massive, but they would still contribute little to the overall mass per m^2 of reflectors, and even less to the kW/kg rating of the solar electric system.

Those are just... visual aids. Not actual examples... I must insist, the wiring we have today is not relevant for a multi-megawatts solar electric craft. The wires will not be bare, so the emissive constant you'll use is that of the insulating layers. They can be actively cooled through jackets of circulating fluid, and even kept at cryogenic, down to superconducting temperatures, by keeping them in shadow and a liquid helium/hydrogen cooling system. High-temp superconductors can even be cooled by liquid nitrogen (70K) and would allow for massive currents at low voltage. Also, the shapes you are referring to are for the reflectors. The reflectors don't have wires running out of them, as they don't generate electricity. Because the wires are so lightweight, it is likely that the spacecraft can afford to keep several kilometers of wiring as a backup.

It is an issue! But not a big one. The micrometeorites will make very small holes. The area of the hole, divided by the total area of the reflective surfaces, would be insignificantly small. Decades of hits would be required to make an impact on the amount of light you can focus onto the solar cells.

No need for bridge trusses. I was thinking of tension structures like these: A minimal number of 'struts' takes on the load of wires pulling the fabric (reflective surfaces) into curved shapes. The actual arrangement of the solar concentrators will look like the larger circles on this: At the 0.01g accelerations I am aiming for, the forces are relatively moderate. Consider a 200m wide parabolic reflector dish on a 100m 'stick' jutting from the side of the spaceship. The dish's surface area is 31400m^2. At 1g/m^2, it would weigh 31.4kg. If structural support doubles this figure, it is still only about 35kg. If the stick had the support the entire weight of the dish on an attachment point at its tip, it would need to handle 0.34N at the tip. After spending a stupid amount of time on this calculator: http://www.amesweb.info/StructuralBeamDeflection/CantileverBeamConcentratedLoad.aspx, I can say that a N moment force can be handled by a triangular diamond-like carbon stick 10cm wide and 0.1cm thick that masses 15kg. It will bend by 17 degrees under maximum acceleration. Total mass per m^2 with structure would be 1.59g/m^2. If we add a 'guy wire' to the tip of the stick, we can support most of the bending forces (0.28N) by a tension wire such as Zylon only 0.25mm thick. The 'stick' can now be considerably lighter, about 5.3kg and allowing for a deflection of 13.6 degrees. Total mass per m^2 to 1.27g/m^2. Did the calculations above. They would increase the mass per area by 30 to 50% for the solar collectors. Parachute was only a description of the shape. Nothing to do with drag or wing.

Higher efficiency is not self defeating. It means you reduce the waste heat production, the mass dedicated to radiators, the size of the cooling systems and the ratio of solar collector area to solar panel area. All good things. These advanced solar electric concepts are meant for use in space. The priority is increasing the kW/kg rating. Home use is on a completely irrelevant scale with with very different concerns, such a $/kW and payback times. The wires and equipment that will be designed to handle this current will not be made to hardware store specifications, but tailored for use by the spacecraft's engines. The wires, electrical equipment and solar panels themselves will all be in enclosed, protected and charge-neutralized environments within the spaceships. The only thing exposed to space for both of the designs I described is the solar collectors, and all they have to do is stay relatively reflective and generally straight. Non-imaging optics can handle light coming from many directions and can therefore correct for some wobble and bending, so even those requirements are less strict than, say, on a solar sail. I seriously doubt that the wires taking the current from the solar cells to the electrical regulators/transformers/modulators ect. will be exposed to hard vacuum without protection. They will likely be protected by many layers of insulators and insulation and looped through environments filled with a safety gas. Your designs seems to use solar power to ionize a gas and turn it into a plasma, but you still need electrical power to accelerate it further? I'm not sure what 'angle' refers to. The details of a high powered electrical thruster are very complex and involve electromagnetic theory I am not familiar with. The departure burn scheme you describe does not seem to be a good option. For one, it limits the burn times to a very narrow portion of the orbit. Waiting for the next periapsis is measured in hours in low orbit, but quickly becomes days or weeks as the apoapsis is pushed past lunar distances. An orbit with an apoapsis at 400000km and a periapsis of 200km would have an orbital period of 10.8 days, and you're still only 85% done with the departure deltaV, but still well within Earth's sphere of influence. You'd do the short burn and then have to loop all the way around to do another short burn which will raise the apoapsis even more. A spiralling orbit allows for continuous acceleration and gradual increase in velocity. It is what NASA and Ad Astra propose for their solar-electric craft and the best way to make the most of their low acceleration. I agree with your first point, hence the blog post introducing the kW/kg rating as an important figure in the introduction. For your second point, I must insist that it is simply a matter of reaching a thermal equilibrium. If heat in matches heat out, the temperature does not rise. That's all there is to it. As for 'a thousand suns', solar cells are regularly tested in concentrated photovoltaics research to x100, x500 and x1000 solar intensities. Today's laboratories already have designs at these intensities, and they can go higher. The cells remain at room temperature (300K). The data I quote in the blog post includes link to those experimental studies. According to wikipedia, liquid droplet radiators were tested of Shuttle missions STS-77 back in 1996 (https://en.wikipedia.org/wiki/Liquid_droplet_radiator#cite_note-Dickinson1996-6).

Thanks! There's more to come. A tension-wire structure should be decent. The wires can pass in front of the parabolic disk without a problem, so you can form a 3D pyramid-like structure where the reflective films are pulled by each corner of the pyramid. Another option is inflatable rims. I think you should take a more step by step approach. Some of the limitations you mention are self-imposed. Due to the second law of thermodynamics, the heat exchanger in a solar thermal rocket cannot be hotter than the Sun. Heat flows from a hotter source to a colder sink. 1. Current frames are built out of aluminium and designed to survive the 3g+ accelerations and vibrations during a launch from Earth. Much lower masses can be achieved by in-space construction. 50% efficiency with 0.1kg/m^2 means a power density of 6.8kW/kg using sunlight at Earth orbit. Impressive when compared to today's 300W/kg, but still lower than what is potentially possible with concentrated sunlight. Without concentrated sunlight, you would find it very difficult to reach that efficiency anyway. 2. Voltage/Amperage are not things I mentioned because they are mostly engineering problems. Superconductivity, for example, makes ultra-high amperages a non-issue for example. 3. Why would you use an ion thruster with such an extreme Isp? Such Isp becomes wasteful for interplanetary trajectories because your propellant savings are inversely proportional to your deltaV requirement. Each increase in exhaust velocity saves less and less propellant, until you're negotiating kilograms on a multi-ton ship. Suppose we use an electrodeless plasma thruster. Let's set the Isp to a more reasonable 2000s. Per megawatt, the engine produces 102kN before efficiency losses. If you have a 50 ton dry mass rocket and want to go to Mars, you need a deltaV of 6km/s. That's a mass ratio of 1.36, or 18 tons of propellant. Wet mass is 68 tons. If we have access to just 10MW of power, the initial acceleration is 0.015m/s^2, with the average acceleration at 0.017m/s^2. The departure burn to Mars will require a departure burn duration of 56 hours, which is tiny compared to the 8.6 month travel time to Mars. Let's go big. Let's dedicate an entire 50 tons to power and propulsion, averaging 10kW/kg. That gives us 500MW of propulsive power to push a 100 ton dry mass rocket. With an engine Isp bumped up to 4000s, thrust is 25.5kN. We want to cut the Earth-Mars trip down to 2 months. This is an impulse-2 trajectory that requires 52.9km/s of deltaV. The mass ratio requires is 3.85, so now our fully loaded rocket masses 385 tons. It starts accelerating at a low rate of 0.066m/s^2, but it averages 0.1m/s^2. This 2-months to mars rocket spends roughly 3 days accelerating away from Earth. Decent, no? This is without even using the 'advanced' designs mentioned in the blog, at 1MW to 2MW/kg. Payload should be 30 to 40 tons. For the designs I proposed, there are no large films. There are large solar collectors, but they're just reflective dishes that only bounce sunlight. The actual solar panels are a much smaller surface embedded in the rocket, surrounded by cooling equipment and wires as thick as necessary to carry away the current.

The solar collectors are made of the same materials, structures and general configurations as solar sails. Solar sails today can be made of Mylar, averaging 7 grams per square meter. In the future, thinner sheets of aluminum and lighter structural materials, such as graphene foam, can allow for reflective surfaces of only 0.1 grams per square meter. If you really get down to it, nanometers-thick metals such as silver with a silica coating can mass as little as a few milligrams per square meter. I will work out a full example spaceship in Part III to demonstrate the level of performance possible with Advanced Solar-Electric power and propulsion schemes.

Technically, all the technologies I mention have already been deployed in space or have real-life working demonstrators. Concentrator solar panels have been used on Deep Space 1: https://en.wikipedia.org/wiki/Deep_Space_1 and thermophotovoltaics are a well-understood technology. The collector equipment is usually the component with the lightest mass per square meter: only a few grams or less. The more collector area we use (and higher solar concentration), the lower the average mass per square meter of the entire system. Around Jupiter, we would be receiving 10% of the sunlight around Earth, around Saturn 1% and 0.1%. 1kW/kg is generally accepted as the minimum needed for rapid interplanetary travel. It is the performance required of a generator to power a VASIMR rocket to Mars in 39 days. This means a power and propulsion system can travel quickly around Jupiter if it had 10kW/kg, around Saturn at 100kW/kg and Neptune at 1MW/kg. Currently, solar electric systems are unable to produce this level of power density, so they would become incredibly underpowered in the outer solar system. With the designs I suggest, travel around Saturn and even Neptune under solar power is possible using the advanced designs I described. The reason solar-electric power is considered weak today is because it struggles to reach 100 to 300W/kg. Quite right! Making the solar collectors larger should not be a big problem, as they mass only a few grams per square meter while the rest of the system (solar cells, cooling ect) does not have to get more massive. One application I will talk about in Part 2 is beamed power. Towering structures holding together several km^2 of parabolic reflectors can collect a lot of sunlight even in the outer solar system. It is converted into electricity by the solar-electric systems I am describing. This electricity can be used to power lasers to transmit that energy to other spacecraft. This means that you only need one big collector while all the rest of the spacecraft only uses a tiny laser dish to power its rockets.

What is TANSIS?

Hello! I share with you the recent post from here: http://toughsf.blogspot.com/2017/11/advanced-solar-energy-in-space-part-i.html I hope you will find it interesting and will have something to discuss! Advanced Solar Energy in Space: Part I Solar Thermal Rockets can be efficient and have high performance. However, they remain temperature-limited to an exhaust velocity of 12km/s. How do we surpass this limit? The limits NASA's Suntower concept. Solar Thermal Rockets have been shown to have great potential if we use modern materials technology - they can be as performant as Nuclear Thermal Rockets in the inner Solar System. With Liquid Rhenium Solar Thermal Rocket, we demonstrated that it could be possible to increase the maximum operating temperature from the 4500K of the most advanced solid heat exchangers to the 5900K of a liquid rhenium-based heat exchanger. However, despite these high temperatures, a Solar Thermal Rocket can never exceed an exhaust velocity of 12km/s. Due to the second law of thermodynamics (the heat exchanger cannot get hotter than the source), the heat exchanger in a Solar Thermal Rocket cannot get hotter than the surface of the sun. As exhaust velocity depends on temperature, we cannot obtain more than 12km/s exhaust velocity out of a rocket that uses the Sun as a heat source. Electric rockets Solar Electric rockets are well known for powering propulsion systems with much higher exhaust velocities. The ion thruster on the Dawn probe was powered by 38m^2 of solar panels and achieved an exhaust velocity of 31.3km/s (3200s Isp). A high exhaust velocity allows for the amount of propellant carried to be drastically reduced for any deltaV requirement. For example, if we wanted to go from Earth to Mars, the deltaV requirement is 6000m/s. A chemical-fuel rocket with an exhaust velocity of 3678m/s (375s Isp, like the Raptors on SpaceX's BFR) would need to consume 4.1kg of propellant for each 1kg of dry mass to fulfil the deltaV requirement. An ion thruster like Dawn's would only need 0.21kg for each 1kg of dry mass: a twenty-fold decrease. In short, the main advantage of electric rocket is that they allow for very small spaceships that don't need a lot of propellant to go to further and faster. So what's the catch? Electric rockets have two major downsides. The first is the power source. So far, we have used solar energy in the form of solar panels, or nuclear energy in the form of RTGs, to power electric rockets. Solar panels have a fundamental efficiency limit called the Shockley–Queisser limit. It states that no more than 33.7% of the energy of sunlight can be extracted by a single solar cell. Most common solar cells have the potential to extract 32% of the Sun's energy using silicon band-gaps, while commercial versions manage an efficiency of only 24%. The world record for silicon solar panel efficiency is 26.3%. An early Space-Based Solar Power concept. Solar panels are rather heavy for their area and performance. They are usually several kilograms per square meter. Research into making solar panels lighter has produced designs such as thin-film solar cells with 0.2 or even 0.1kg/m^2. Combining the efficiency of the most advanced solar panels with the sectional density of thin-films solar cells makes for a system with a power density of 1.5kW/kg at most. Modern advanced solar cells for spacecraft aim for 0.3kW/kg. Other power options rely on nuclear energy. The current form of nuclear energy of spacecraft, RTGs, has woefully poor performance. A power density of 1 to 10W/kg is to be expected. The SP100 nuclear reactor. Nuclear reactors with a carnot heat cycle (Stirling heat engine, steam turbines and so on) have great potential, but currently they are limited by the low temperature difference between the reactor core and the radiators. This is the result of having to keep the fissile fuels inside the reactor core solid and safe (so a low maximum temperature) and the low performance of thermal radiators currently employed (so a high minimum temperature). Despite these limitations, nuclear reactors producing over 10kW/kg have been designed. They have a potential of over 100kW/kg or more. Nuclear power has problems not related to its performance as well. For the foreseeable future, fissile fuels are expensive, dangerous to handle and a hot-button political and environmental topic. The radioactivity continuously degrades the power generating equipment and makes refurbishment or repairs a complicated affair. A lot of effort will have to be put into finding sources of fuels if we intend to exploit the Solar System, otherwise we'd have to wait for alternative nuclear technologies such as fusion reactors to mature. Low power density means low acceleration. Spiralling trajectories are the result. The second problem with electric rockets is their low propulsive power. One aspect is that the electrical power divided by the exhaust velocity leads to a very low thrust output. Another aspect is that the cryogenic magnets, the superconducting coils, the electrostatic chambers... are simply quite heavy. Current laboratory-tested concepts such as VASIMR are only expected to have about 1kW/kg, with most other designs struggling to reach ten times less specific power. Between the two, electric rockets end up having extremely low thrust. This prolongs the burns that chemical rockets can perform in minutes into weeks-long affairs that spend an inordinate amount of time in Earth's Van Allen belts. The benefits of the Oberth effect are completely lost and gravity losses that come from accelerating away from the optimal angle become significant. Reducing the duration of these burns requires dedicating most of the spacecraft's dry mass to power generation and propulsion, so as to increase the rate of accelerations. The propellant to payload ratio quickly drops to levels comparable to chemical propulsion. Advanced Solar Electric Energy We need to improve the performance of electric rockets. It is possible to do this without relying on problematic nuclear propulsion, limp solar panels or the massive amounts of chemical fuels needed for interplanetary travel. What we need is advanced solar energy concepts with much higher power densities. We will now look at designs that allow the efficient use of sunlight to produce electricity out of compact and lightweight generators They common key to these designs' performance is the use of solar collectors made of extremely lightweight materials, based on the technology developed for solar sails. These reflective surfaces of only a few grams per square meter can focus huge amounts of sunlight onto a small surface. Intense sunlight allows for higher performance and greater temperatures. Concentrated photovoltaics Photovoltaics convert the light they absorb into electricity. By increasing the intensity of this light, more electricity can be produced from the same solar cell. This increases power density. Peak gains are obtained from x1000 -x3000 concentration. Concentrated photovoltaics attempt to multiply the intensity of sunlight collected by a solar panel by adding a concentrator. A concentrator focuses sunlight onto the solar panel's surface. Since the concentrator only needs to be reflective, it can be much lighter than an equivalent surface area of solar panels. By maximizing the collector area (lightweight) and minimizing the solar cell area (heavy), a better power density can be achieved. On top of simple mass optimization, efficiency can be improved. Conventional solar cells use a single silicon p-n junction. The efficiency therefore capped by the the Shockley-Queisser limit. This is sufficient as the design is cheap and relatively lightweight, so higher output is achieved by adding more solar panel area. Doped silicon solar cell. Concentrated photovoltaics leads to a very small solar panel surface area. It becomes more reasonable to use more complex solar cells that improve efficiency. Multijunction solar cells use multiple p-n junctions on top of each other. Each p-n junction is tuned to a portion of the electromagnetic spectrum. Sunlight ranges from radiations in the infrared to X-rays. Only 47% of sunlight's energy is contained in the narrow portion of the electromagnetic spectrum called the visual spectrum, that corresponds to radiations of wavelengths 400 nanometers to 700 nanometers. Conventional solar cells only absorb a fraction of this small segment. Red, green and blue p-n junctions capture energy from wavelengths that correspond to the low (infrared to red), middle (yellow and green) and high energy (blue to UV) sunlight. Triple junction solar cell. Up to 50% efficiency. Quadruple junction solar cell extending into the deep IR wavelengths. Better suited for space use. Up to 56% efficiency. By increasing the number of junction layers and dividing light received into small slices corresponding to each layer, a theoretical maximal efficiency of 86.8% is possible. The number of layers for each efficiency improvement increases exponentially. Reaching 86.8% efficiency requires an infinite number of layers. By three layers, the efficiency cap is raised to 63%, which we will deem sufficient. Another concern is heating. The sunlight that is not converted into electricity becomes waste heat instead. Increasing the temperature of a solar cell lowers its efficiency. When concentrators are focusing sunlight to tens to hundreds of times its normal intensity, the heating can quickly become problematic. Active cooling is therefore required to keep concentrated photovoltaics cool and efficient for space applications. In the above graph, a solar cell under 1000x sunlight intensity is tested at a range of temperatures. At 400K, it is estimated that there is an efficiency loss of 11% of the value at 280K. Lower temperatures improve the rate at which photons are converted in the semiconductors, prevent losses from re-radiated energy and lessen the effect of other inefficiencies such as recombination. Another study cooled down solar cells down to 50 Kelvins. The graph above is the theoretical maximal efficiency achieved at each temperature. We notice that at 50K, the peak efficiency (41%) is 37% higher than the peak efficiency at 300K (30%). Three-junction solar cells can have a mass of 0.85kg per square meter or less. In this book, 0.1kg/m^2 is cited for a concentrated solar cell array, although it cannot be determined if it is multi-junction. This recent NASA proposal for concentrator quadruple-junction solar cells designed to survive extreme environments cites 0.24kg/m^2 as the figure for solar cells without their concentrators. Let us now consider two designs for modern or advanced multi-junction concentrator solar energy systems. We note that waste heat management is a determinant factor in the potential performance of these designs. Refer to All the Radiators for more details. Modern concentrated photovoltaic example: Different solar concentrator configurations. We will use conservative figures. 0.85kg/m^2 for the solar cells, with quadruple junctions operating at room temperature (300K) to provide an efficiency of 40%. The concentrator is a parabolic dish of mass 7g/m^2 and a reflectivity 95%. The solar concentration is x1000. 1.29MW of solar energy is focused by 1000m^2 of concentrator onto each 1m^2 of solar panel. It is converted into 519kW of electricity and 779kW of waste heat. The waste heat must be dealt with using a lightweight, low temperature radiator. A liquid droplet radiator is ideal for this task. 0.1mm wide droplets of mercury coated in black paint, spaced by 1mm and released at a velocity of 20m/s across a 1m wide gap would cool down from 300K to 5K before being captured as solid mercury balls. The radiators would remove 27kW of heat for every kilogram of droplets. Using multiple sheets of droplets multiplies this figure at the cost of a slight efficiency loss due to interreflection. The component masses for 1m^2 are 0.85kg of solar panels, 7kg of collectors and 28.9kg of radiators. It produces 519kW. The system power density becomes 14.1kW/kg. A more realistic figure would include the mass of additional systems such as power converters, droplet radiator booms and pumps, solar tracking mechanisms for the collectors and so on, probably bringing down the system power density to 10kW/kg. Advanced concentrated photovoltaic example: Graphene foam can form the basis for ultra-lightweight solar collectors. We will use optimistic figures. 0.25kg/m^2 for quadruple junction solar cells operating at a 275K temperature. The efficiency is 60%. Micron-thick aluminium concentrators resting on graphene foam or tensed by Zylon wires have a mass of only 1 gram per square meter. Reflectivity is 95% and solar concentration is 10000x. 13MW of sunlight reach the solar cells. 7.79MW becomes electricity while 5.19MW becomes waste heat. We will use a hybrid wire/droplet radiator. The wires can have alternating hydrophobic and hydrophilic surfaces. Small droplets of water are held by charged surfaces, alternating hydrophobic/hydrophobic patches or simple surface tension, on a thin wire. The wire drags the droplets along like a conveyor belt. Ink turns the water black and improves emissivity. Each 1m^2 of radiator area is composed of 1000 parallel wires holding 1 million droplets of 1mm diameter each, moving along at 10m/s. The passage through the vacuum cools the droplets from 275K to 64K. The high heat capacity of water and its high heat of fusion (energy needed to freeze it) means that more than 1.5MW of heat can be removed by 1kg of water and wires. The component masses are 0.25kg of solar panels, 10kg of concentrators and 3.5kg of radiators. System power density can approach 500kW/kg. Even more advanced designs that use thinner reflectors, faster wire/droplet radiators and lower temperature (higher efficiency) solar cells can probably reach 1MW/kg. High temperature thermophotovoltaics Photovoltaics are most efficient when converting light composed of wavelength exactly matching the band-gap of the n-p junction materials they are composed of. This is what allows laser-to-electrical conversion to achieve very high efficiencies. The wavelength of the laser exactly matches the band-gap of the converter. We cannot expect such efficiencies using the Sun's broad spectrum of radiations. Even multi-junction solar cells can only cover part of the spectrum, at the cost of greater complexity, cost and mass per square meter. Using different receiver/emitter materials helps specify which wavelengths reach the PV cell There is a solution in tuned thermo-photovoltaics. In this design, the Sun's rays are focused on a heat exchanger. The heat exchanger absorbs the entire solar spectrum and radiates it back in a narrower range of wavelengths. Selective filters reflect anything outside of an even narrower selection of wavelengths back to the heat exchanger so that the energy is not wasted. A heat exchanger combined with efficient filters allows us to reduce the solar spectrum to emissions exactly matching the band-gap of a solar cell. Efficiencies of 98% of the maximum thermodynamic efficiency are possible using a single n-p junction. TPV system with optical concentrator. Currently, thermophotovoltaics are limited by the stability of the heat exchanger, the loss of re-radiated energy, the lack of cooling systems and simple lack of development when compared to traditional photovoltaics. For example, most designs tested today focus on Silicon Carbide or Tungsten heated to 1500K, while the thermophotovoltaic cells reach 350K or more. The maximal thermal efficiency becomes 76%. From this amount, 20 to 50% of the radiations from the heat exchanger are lost or simply go in the wrong direction in typical flat heat exchanger designs. Solar cells operating at high temperature, losing efficiency compared to their counterparts at 300 or 270K. The radiations outside the band gap range of wavelengths are not always efficiently recycled back into the heat exchanger too. Finally, the low temperature emitters cannot achieve the high emission intensities (MW/m^2) that have helped the efficiency of concentrated solar photovoltaics. A solution to these problems can be found. Hot tungsten coils. Using higher-temperature heat exchanger materials is a start. Carbon is ideal, with a melting temperature of over 4000K. Tungsten can reach 3000K temperatures without a problem. Instead of using small band-gap p-n junctions such as Gallium Arsenide, Silicon with a band gap of 1.1eV, which corresponds to wavelengths of 1100nm, can be selected. This corresponds to the peak emissions of an emitter at 2660K. Active cooling can handle the heat load to reduce the cold end temperature to 270K. The thermal efficiency of system with a 2660K hot end and 270K cold end is 89%. This is an immediate 13% improvement over current systems. A photonic crystal composed of layers of gallium arsenide and air with an otherwise impossible band gap. Even more recent research focusing on artificial band-gaps created by using photonic crystals permits efficiencies greater than 40%. The configuration of a cylindrical heat exchanger absorbing sunlight from surfaces on the ends. The heat exchanger can be shaped into a very long cylinder inserted into a closed chamber. The chamber's inner walls are coated with TPV cells. Sunlight concentrators focus sunlight onto the exposed top and bottom of the cylinder. The cylinder then re-radiates this heat from the enclosed lateral walls onto the TPV cells. It is called a 'thermal well'. The large exposed to enclosed surface area ratio of the cylindrical heat exchanger means that very little of the heat exchanger's radiations are lost. This is another 20-50% improvement over current designs. A good optical filter is needed. For silicon, a filter would need to only allow wavelengths longer than about 800nm or shorter than 1100nm. This is possible today with developments in nano-structured metamaterials and photonic crystals. Finally, as demonstrated in the previous section, it is possible to gain a 20% increase in efficiency or more by cooling the solar cells. Together, these optimizations can achieve the 50% efficiency purported in recent research, or approach 70 to 80% efficiency as dictated by theoretical limits. We will now look at a modern, then advanced, design for a solar TPV system. Modern thermophotovoltaic example: We will only use figures available in today's research. High temperature thermo-photovoltaic device. We use a tungsten heat exchanger, heated to 3000K. The thermophotovoltaics are indium-gallium-arsenide-phosphide cells with a bandgap tuned exactly to the peak emissions of the heat exchanger, for an efficiency of 61%. The n-p junction is only 50 nanometers thick and backed by a silver plate to reflect unabsorbed wavelengths back to the heat exchanger. Blackbody emission spectrums for various temperatures. The heat exchanger will be a carbon-coated cylinder. Sunlight is focused onto its flat ends. The light is absorbed and heats up the tungsten. Heat conducts down the cylinder and is re-radiated out of the lateral surfaces. A shell of solar cells intercepts these radiations. At 3000K, the tungsten is emitting at 4.1MW/m^2 but only 1.15MW/m^2 reaches the solar cells. This means that for each square meter of tungsten, there will be 3.56 square meters of solar cells. The thinner the heat exchanger, the lighter it can be. A 10cm wide, 1m long tungsten cylinder would mass 153.8kg and have a lateral surface area of 0.31m^2. It will be able to illuminate 1.1m^2 of solar cells. The mass of tungsten per square meter of solar cell becomes 139.8kg. If we use a 1cm wide cylinder instead, we calculate a tungsten mass of just 14 kg per square meter of solar cells. We will use the latter figure. Another benefit is that the radiations lost from the cylinder ends amount to only 0.25% of the total radiations. The full radiation of the heat exchanger's lateral surfaces will have be matched by an equivalent input from its top and bottom surfaces. For a 1cm wide rod, this means that 0.031m^2 of lateral surface area are supplied by 1.57cm^2 of illuminated area; a ratio of 395:1. The solar concentrators are therefore heating the tungsten at an intensity of 1632MW/m^2. 95% reflective solar concentrators would be achieving a concentration of x628399. If the tungsten rod is I-shaped, with larger absorbing surfaces at the top and bottom, the radiative efficiency is lessened but the concentration factor becomes more manageable. If a 10% radiative loss is acceptable, the a concentration factor of 'only' x15710 is needed. At 7g/m^2, about 0.68kg of reflective concentrator surfaces would be needed to heat a 1cm and 1m long wide tungsten rod. This represents about 0.61kg per square meter of solar cells. Plasmonic light-trapping nanoparticles help very thin solar cells capture as much light as thick solar cells. As this is a single-layer solar cell, we can use the masses of thin-film solar cell arrays as a basis for our power density calculation. This means 0.2kg/m^2 or less. A silver backing plate and a denser semiconductor mix might mean an area density closer to 1kg/m^2. They convert the radiations into 631.3kW of electricity and 403.7kW of heat. The solar cells must be kept at 300K. Cooling will rely on the mercury droplet radiator mentioned above. 14.9kg of droplets are required to remove the waste heat. Component masses come out as 1kg for the solar cells, 14kg for the heat exchanger, 0.61kg for the solar concentrator and 14.9kg for the droplets. System power density is 20.6kW/kg, although other components we have not considered might lower this somewhat. Advanced thermophotovoltaic example: Advanced materials technology and miniaturization techniques can drastically increase the system power density of thermo-photovoltaics. Photoluminescent emitters allow short wavelength radiations without the need for high temperatures. Indium gallium phosphide, typically used as the 'blue cell' in a multi-junction solar panel, will be our only layer. It operates best when it receives 545nm wavelength light. This light will be supplied by a 5317 Kelvin blackbody, with metamaterials filtering out wavelengths too long to be efficiently converted by the solar cell. The solar cells have an 80% conversion efficiency of light at 10MW/m^2 intensity, while massing only 0.1kg/m^2. The lightbulb structure needed to contain the liquid rhenium. The heat exchanger will be liquid rhenium, held inside a transparent tube of fused quartz. The tube walls are actively cooled by circulating hydrogen gas in a manner similar to what was proposed for Closed-Cycle Gaseous-Core Nuclear Thermal Rockets, or 'nuclear lightbulbs'. An alternative would be electromagnetically contained plasma, such as cesium ions. Non-imaging optics allow for extreme solar concentration ratios with sub-1% radiative losses. Parabolic compound troughs can also reduce the sun-tracking requirements of the solar collectors. The liquid rhenium heat exchanging cylinder will be 2mm wide. It masses only 66 grams per meter length. Heat is absorbed from end-caps 6.3mm wide, giving it an I-shape. The tube radiates at 40.78MW/m^2 and receives 4078MW/m^2 through the end caps. Each tube shines on 0.0256m^2 of solar cells, so it adds 2.57kg per square meter of solar cells. Sunlight is supplied by 197m^2 of 95% reflective solar collectors. At 1g/m^2, the solar collectors would mass about 0.2kg. Each square meter of solar cells produces 8MW of electricity but also 2MW of waste heat. We use the water and wire arrangement of radiators from the advanced version of the concentrated sunlight generator to handle this. It would require 1.33kg. Component masses are 0.1kg for the solar cells, 2.57kg for the heat exchanger, 0.2kg for the solar collectors and 1.33kg for the radiators. System power density should be close to 1.9MW/kg, though realistically it will be lower. Part II In the second part of this series, we will look at power generation options that do not rely on photovoltaics.