cryogen

The equilibrium temperature depends on the radiative properties of the spacecraft. Your temperatures are reasonable for planet surfaces, but for an artificial material with very low solar-emissivity, Teq can be much colder than -100 °C. Moreso, there's the trick that you can stack several layers of radiating surfaces, to get a combined, effective emissivity much lower than any physical material's. For a famous example, the James Webb telescope has a sunshield with five layers, εeff = 0.00038 (!!), and Teq = 40 K (-233 °C) on the cold side. That temperature is achieved solely by passive, radiative cooling. http://jwst.nasa.gov/sunshield.html In theory, you can get Teq so cold, that it's below the boiling point of liquid hydrogen (20 K at 1 atm). Alternatively, it can be slightly warmer, and the tank can be kept artificially cold by active refrigeration. Either way, no hydrogen is vented into space: it's "zero boil-off" (ZBO). http://develop.nttc.edu/sbipp/technologyportfolios/portfolios/ISS-Propellant_T-S/Archive%5C20060042833.pdf - - - Updated - - - This sounds very interesting, if it can work. For reasons I don't fully understand, the Caltech study calls this a "prohibitive" mass cost. Asteroid Retrieval Feasibility Study http://www.kiss.caltech.edu/study/asteroid/asteroid_final_report.pdf I'd note that's definitely outside the C3 ability of existing rockets; this would need an SLS launch at least, or in-orbit fueling. Should also note that, if you're picking up an asteroid whole, the mass uncertainty means you can't feasibly exclude a 1,000 ton asteroid. (So your 500 ton figure understates it). But if you're going for a boulder option, then you could arrange that. Also, don't forget there's a pretty large delta-v cost to match orbits with an NEA, even if it's extremely low in C3. That adds another fraction to the propellant ratio. https://www.nasa.gov/sites/default/files/files/Creech_SLS_Deep_Space.pdf

I think so, Tse, J. S., and D. D. Klug. "Molecular Dynamics Studies of High Pressure Transformations and Structures." In High Pressure Molecular Science, pp. 59-85. Springer Netherlands, 1999. https://books.google.com/books?id=FYztCAAAQBAJ&pg=PA73&lpg=PA73 Zha, Chang-sheng, Ronald E. Cohen, Ho-kwang Mao, and Russell J. Hemley. "Raman measurements of phase transitions in dense solid hydrogen and deuterium to 325 GPa." Proceedings of the National Academy of Sciences 111, no. 13 (2014): 4792-4797. http://www.pnas.org/content/111/13/4792.long

It means ordinary, molecular hydrogen (H2) below its 14 Kelvin melting point -- not the exotic metallic hydrogen.

What would be the additional cost of building and launching a 2nd JWST, having already built one?

It's worse than that, we don't even have a human-rated rocket that can reach SSO (retrograde polar orbit). Wikipedia says Soyuz-FG gets 4,500 kg to 800 km SSO, which is a lot less than the 7,150 kg Soyuz TMA-M spacecraft.

Solid hydrogen, actually. https://en.wikipedia.org/wiki/Wide_Field_Infrared_Explorer Wide Field Infrared Explorer (WIRE)

The asteroids are in sun orbit, so it takes far less delta-v to move them to a high earth orbit (like EML1), than a low one. In the reference concepts, the propulsive delta-v's are on the order of 300 m/s -- i.e. 1,000 tonnes asteroid mass, 10 tons Xe propellant, and 3,000s Isp. (The NEA v_inf's are higher, but most of the delta-v comes from a lunar gravity assist). A low-thrust spiral to LEO would add about 6 km/s, increasing the total delta-v by a factor of 20. That's probably not realistic. http://www.kiss.caltech.edu/study/asteroid/papers/near.pdf

The best HLV is one that shares stages and cores with a smaller rocket, one that launches frequently. For example, tri-cores like Falcon Heavy and Vulcan Heavy. This is how you can get reliable, tested components. All of the shuttle-derived concepts have a single, giant core that's only used for super-heavy payloads, so it gets launched very rarely. This is dodgy. If you share cores with a commercial sat launcher, everything's tested dozens of times a year: you get lots of engineering data, opportunities to learn and iterate.

You don't need an ice-covered asteroid; you can extract water from ordinary rocks. You can find them in similar heliocentric orbits as the earth. http://www.kiss.caltech.edu/study/asteroid/asteroid_final_report.pdf The delta-v distances involved can be extremely low. EML-1 is only ~150 m/s from heliocentric orbits, and some near-earth asteroids are only slightly further away.

True, but that's only a flyby mission; it uses very little delta-v in deep space. From a quick Google: MRO and MAVEN used monopropellants; Magellan used a solid rocket; and all of Galileo, Cassini, MESSENGER, Juno, Mangalyaan, Mars Express, and the future Europa Clipper and JUICE use bipropellants. Still, they all use bipropellants for the main engines, not monoprops. In very large amounts too! e.g., [TABLE=width: 600] [TR] [TD]Shuttle OMS[/TD] [TD]MMH/NTO, 21,660 kg[/TD] [/TR] [TR] [TD]Soyuz TMA-M (current)[/TD] [TD]UDMH/NTO, 800 kg[/TD] [/TR] [TR] [TD]SpaceX Dragon v2[/TD] [TD]MMH/NTO, 1,290 kg[/TD] [/TR] [TR] [TD]Boeing CST-100[/TD] [TD]MMH/NTO[/TD] [/TR] [TR] [TD]Orion Service Module (ESM)[/TD] [TD]MMH/MON-3, 9,200 kg[/TD] [/TR] [/TABLE]

I don't think this is right. The only deep-space missions that used monoprops were prehistoric ones, that were flyby-only. (To my knowledge). I don't know of any orbiter (high deep-space delta-v) that used monoprops. The tank fraction isn't meaningfully higher for biprops than monoprop. Biprops are very reliable; they have lots of space experience in GEO satellites. If you need more reliability, you can add redundant spares, since their weight is trivial (<4 kg for 400 N) compared to the fuel mass. Cassini did this:

I dug up some data. The specific impulse is 266s, compared to 242s for monopropellant hydrazine. (They don't specify the assumptions for this figure). It's still much worse than storable bipropellants, like hydrazine / N2O4 (344s?). So I'm not sure how useful this really is. It could replace small RCS thrusters, but it doesn't look viable for main propulsion (i.e., apogee thrusters on GEO comsats), which is where most of the hydrazine actually goes. http://www.nature.com/news/green-fuels-blast-off-1.13603 https://en.wikipedia.org/wiki/Green_Propellant_Infusion_Mission

67th. https://en.wikipedia.org/wiki/Ariane_5#Launch_history Launches #15 - #81, inclusive. I noticed that too. Apparently gaining speed quickly is more important than holding altitude.It's probably more efficient with an altitude drop. Tilting the trajectory upwards would spend more delta-v on gravity losses.

It's not possible for a Falcon first stage to do that, both because the delta-v is too low (kerosene's Isp is too low), and because it's too fragile to survive reentry. (The reusable stages only reenter at a gentle 2.0 km/s). You can look up the stage masses here.

ESA made the Herschel telescope! It's built world-leading space telescopes, in several types of astronomy. Planck is the world's best telescope at CMBR wavelengths; Gaia is the highest-precision astrorometry telescope; XMM-Newton is one of the two giant x-ray telescopes (and the upcoming ATHENA telescope will be even bigger). Another European intergovermental, ESO, is building the largest ground-based telescope (E-ELT, 39 meters). It's not quite NASA-size, but it's certainly competent, especially when you consider how small a budget it works with. That couldn't work for Kepler, because it's in a heliocentric orbit. Actually a lot of space observatories are in very distant orbits, the biggest reason being to avoid earthshine.

Here's the current observational limits (click to embiggen). The optical magnitude falls off very steeply, as R-4 (because both the amount of sunlight, and the distance from earth observers, fall off as R-2). Mid-infrared magnitudes (thermal radiation) only fall off as R-2, but you need an intrinsic heat source -- a planet large enough to still have primordial heat. Left graph is IR + optical on a log scale; right graph is zoomed in on small objects. [TABLE=width: 800] [TR] [TD][/TD] [TD][/TD] [/TR] [/TABLE] Sources (open-access): Luhman, K. L. (2014). A search for a distant companion to the sun with the Wide-Field Infrared Survey Explorer. The Astrophysical Journal, 781(1), 4. https://iopscience.iop.org/0004-637X/781/1/4/pdf/0004-637X_781_1_4.pdf Sheppard, S. S., et al. (2011). A southern sky and galactic plane survey for bright Kuiper belt objects. The Astronomical Journal, 142(4), 98. http://arxiv.org/abs/1107.5309 Summary: [TABLE=width: 960] [TR] [TD][/TD] [TD]limit[/TD] [TD]method[/TD] [TD]limiting telescope[/TD] [/TR] [TR] [TD]Sedna-size, albedo = 0.15 (assume r=500 km)[/TD] [TD]~70 AU[/TD] [TD]optical, R band[/TD] [TD]Palomar 1.2m + Las Campanas 1.3m[/TD] [/TR] [TR] [TD]Pluto-size, albedo = 0.15[/TD] [TD]~110 AU[/TD] [TD]optical, R band[/TD] [TD]Palomar 1.2m + Las Campanas 1.3m[/TD] [/TR] [TR] [TD]Earth-size, albedo = 0.15[/TD] [TD]~270 AU[/TD] [TD]optical, R band[/TD] [TD]Palomar 1.2m + Las Campanas 1.3m[/TD] [/TR] [TR] [TD]Saturn-size (Fortney et al. brown dwarf model)[/TD] [TD]~28,000 AU (0.4 l.y.)[/TD] [TD]IR, W2 band (4.6 μm)[/TD] [TD]WISE[/TD] [/TR] [TR] [TD]Jupiter-size (Fortney model)[/TD] [TD]~82,000 AU (1.3 l.y.)[/TD] [TD]IR, W2 band (4.6 μm)[/TD] [TD]WISE[/TD] [/TR] [/TABLE]

One of the lunar Lagrange points, perhaps EML1. Zero-g High solar constant One of the cheapest destinations from earth, after LEO (In particular, no delta-v penalty for landing on a surface. This also allows low-thrust (SEP) cargo shipments from LEO) Reasonably fast evacuation to earth possible (~1-3 days) Very cheap to near-earth heliocentric, to bring in NEO asteroid material (minimum of just 140 m/s to C3=0!) Very cheap to SEL2, to bring in space telescopes for upgrades Very cheap to low lunar orbit, and the lunar surface Reasonably cheap to bring in satellites from GEO (probably low-thrust / SEP) Cheap to high-C3 interplanetary (descent to low earth altitude + Oberth burn) Generally strikes good delta-v balance between a lot of places Compared against similar locations, Solar Lagrange points need more delta-v from earth Near-earth asteroids need more delta-v, and have very long transit times Lunar orbit needs more delta-v, and has high station-keeping costs Various tradeoffs between EML1/EML2 (unstable) and EML4/EML5 (stable halo orbits) LEO is cheaper from earth, but very expensive from heliocentric. If you want to use lots of asteroid material (water ice, shielding mass), you want to assemble things in a high orbit like EML or SEL GEO is more expensive from both earth and heliocentric (VERY suboptimal) Lunar surface is in a deep gravity well, has day/night cycle, and is covered in dust (Of course, this isn't an original idea at all!) edit: Should emphasize how surprisingly cheap EML could be. The low-thrust delta-v for LEO -> EML1 is 7 km/s. With a solar-electric tug, you could probably get 1.2 : 1 mass ratios or better (e.g. for a 1-year spiral, at 8,000s Isp and with 200 W/kg solar panels). I.e., take 10 Falcon Heavy launches to LEO (500 tons). With a 4 MW solar array (~50 tons), and 50 tons of xenon propellant, you could lift 400 tons of payload to EML1 (on paper). That's <$2 billion of launch vehicles, for the entire mass of the ISS. (The ion engines would probably cost more). For bringing in mass from a low-C3 NEO, you could get LEO multipliers as high as 100:1. I.e. one Falcon Heavy launch (50 tons) gets 40 tons of SEP propulsion to a NEO, which then pushes 4,000 tons of NEO mass by 300 m/s, into EML1. For bulk shielding, that's insanely cheap! 4,000 tons of iron-nickel, for example, would surround a 10-meter TransHab with >1 meter thickness. This is just silly back-of-the-envelope stuff, but it's provocative.

Ozone as a rocket oxidizer was studied in depth in the 1950's. In theory you could get 20-30 seconds increase in Isp from it, over plain oxygen. It's thought too unstable to be practical. Here's an excerpt from John D. Clark's Ignition!: An informal history of liquid rocket propellants, which is out of print (and currently sells for $2,904 on Amazon ) Here's what pure liquid ozone looks like, http://labphoto.tumblr.com/post/32406799158/liquid-ozone-just-made-some-with-an-ozonisator

Just to be clear, there's exactly two numbers going into that calculation: the diameter and the orbit period. That's it; that's all the data we have. It's easy to overinterpret these two numbers, when they're hyped up in this crazy way. The other numbers are pure guesses. The density is a generic figure for rocky planets; the mass and surface gravity are derived from that (+ measured diameter) The temperature is a WAG. It's a simple toy radiative balance, like this one: https://en.wikipedia.org/wiki/Stefan%E2%80%93Boltzmann_law#Temperature_of_the_Earth No atmosphere is guessed at. We have no knowledge about the surface albedo, so that's a WAG as well (arbitrarily set equal to earth's albedo, 0.3). Fun fact: if you calculate the ESI of Venus, using only the data available for 452b, the figure's about 0.95 (since all the knowledge about the greenhouse gets discarded). You can easily calculate this, http://phl.upr.edu/projects/earth-similarity-index-esi

Look at the design of Solar Probe Plus. It's meant to go within 0.044 AU of the sun: http://solarprobe.jhuapl.edu/mission/docs/SolarProbeME.pdf At its closest approach, only 34 cm2 of solar cells will be exposed to the sun, with 2 m2 of radiators rejecting heat -- a ratio of 600:1. So it doesn't look like you can save mass this way: you're limited fundamentally by the temperature limit of the solar cells (about 100° C here), and the radiator area needed to keep the equilibrium temperature that cold.

Quick question -- anyone have a link to a description of KSP's thermal model? I'm curious how it works and what its limits are. Can you make a multi-layer heat shield in KSP? update: I guess not.. Ok this is confusing...

Is it actually feasible to have SEP that close to the sun at all? I'm wondering how much mass is needed for thermal design, and what that does to the power / mass ratio.

Here's a NASA study on this idea: "Venus Surface Sample Return: A Weighty High-Pressure Challenge" http://ntrs.nasa.gov/archive/nasa/ca...0010003946.pdf Since Venus is a large planet (Earth-size), you'd need quite big an ascent rocket to get off of it. Even bigger than a Mars sample return. In the concept I linked, they try an Apollo-type mission, with a separate Venus orbiter/return vehicle, and Venus lander/ascent vehicle. The ascent uses three stages of solid rocket, and gets launched from a balloon, inflated with helium from tanks. I guess, given the very thick atmosphere, a giant balloon would be cheaper than a rocket powerful enough to push through the soup. Getting a sample-drilling machine to work at the surface would be a major challenge. Silicon IC's won't work at the high temperatures of the Venus surface. You'd need either a very heavy, inefficient refrigeration system, or a very fast surface mission packed with ice, or a new / exotic type of high-temperature electronics. The overall mission size isn't bad -- they think it could be launched with a single Delta IV M+ (the single-core with SRB's -- smaller than the three-core 'heavy'. That's less than 13 tons in LEO).

Only two stages, because the asparagus wasn't installed correctly! The burn was about 3 hours of simulated time. - - - Updated - - - Ah, that's an idea!

Here's my amateur-ish attempt. One surviving Kerbal in an EVA suit, at v∞= 43.69 km/s. Here's a screenshot with precise numbers. Disclosure -- my album mixes screenshots from several attempts. After a few failures, I stopped taking screenshots from the steps I was repeating. Sorry about this probable rule violation! KSP version is stock 1.0.2, no mods, "normal" difficulty in sandbox. No debugs, except for the debug setting that shows thermal information (to figure out why things were exploding ).