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About Frank

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  1. Solar Thermal Rocket

    Here are a couple of link to STP that might be of interest: "Reduce tank volume with increased STP specific impulse approaching 1200 seconds - Higher temperatures above 3000K increase Isp - At higher temperatures above 3000K and low pressures, hydrogen starts dissociating and lowers the average molecular weight, which also increases Isp" I guess dissociation of hydrogen will allow more Isp for a given temperature (which is limited by material properties), at the expense of efficiency, requiring more energy input.
  2. Solar Thermal Rocket

    I haven't looked closely at the numbers, but 7x more efficient seems optimistic to me. Here on Earth, it's VERY difficult to make a case for solar thermal based on cost, even where thermal should obviously be better. When it comes to mass as the currency, I don't know yet. The working example of a concentrating solar collector I found is SCARLET It's a fresnel concentrator and thermal management is on page 7 of the report. In 1998, 20 years ago, they managed 45W/kg for the array and DS1 with its ion engine worked. Using a 4 m diameter dish and concentrating on a PV module will generate some heat. Quite a bit of it, because up to 60% of the energy will dissipate in heat if 40% efficient.
  3. Hydrogen has lots of handling issues, maybe they are all solved, but I suspect that the SpaceX team did the cost comparison and decided to go with methane. True, the Lunar ISRU is different, but experience from a Lunar settlement would still be helpful. I don't feel it's an either-or case anymore with private money going into space exploration. Both Lunar outposts AND Mars outposts/colonies are possible concurrently. SpaceX mentioned the Sabatier reaction to convert CO2 + 4H2 → CH4 + 2H2O   ∆H = −165.0 kJ/mol, so Mars CO2 with Earth hydrogen to get methane and water. I'm guessing the boil-off rate is manageable to feed the reactor? Carbon-fibre tank vs hydrogen tank plus insulation.. Energy density of hydrogen 8.491 MJ/L, liquid methane (LNG) 22.2 MJ/L compressed methane (CNG 250 bar) 9 MJ/L. So both tanks are the same size. Mass difference? Ammonia can be handled safely, it's used on crops everywhere, regardless, it has lower energy content than methane, so not ideal as rocket fuel, though it might combine well with thermal thrusters. A space suit is probably enough protection and must be worn anyway. Well, that lander crashed, so there was no recovering. A rover is one solution. A Mars short hop is another if enough fuel is left in the rocket - this would assume a navigation failure. Risky business for sure. A short hop using one of the rockets previously landed and partly refuelled is another.
  4. Solar Thermal Rocket

    If we get 40% conversion efficiency from a concentrating solar array, is pumping this energy into an ion engine more efficient when its high Isp is considered? All this technology is also known and working so the risk of development isn't there (or as high). Also, electricity is a generally useful form of energy (compared to heat) and may have use along the coast phase or even when parked waiting for the next mission. I believe there is some cooling needed for those concentrating collectors, but no mirrors needed, which may end up adding to the mass of the array in the balance - hard to tell.
  5. One thing about moon based Oxygen production is that it would be nice to know that ISRU works before we send stuff to a far-away planet. In other words, if someone starts putting moon hotels in place (Bigelow) or decides that it's about time we have a moon base, then the Oxygen creation infrastructure must work. Then sending fuel to EML-1 or Earth orbit or even Mars is possible and may even be economical. I didn't begin to calculate the cost difference. I'm just leaving the door open for that possibility, same with the mass drivers. If these things ever get built and work as advertised, that's probably an order of magnitude cheaper at least. " last person putting time in it? " - I don't follow. In any case, I was under the impression the Solar Thermal Propulsion threads were about long stay and short trips, so I wondered if short stays and long trips were possible and how short can the trips can be given Solar Thermal Propulsion. Well, methane can be stored in pressure tanks, but would need to be liquefied before use, as can oxygen. Freezing is probably avoided by burying the tanks. Hydrogen is tricky to get any kind of density as is evident from our lack of hydrogen cars, efficient storage is a problem. Active cooling is a possibility - sure. I mostly went with methane based on the Raptor engine and the idea that fuel can be synthesized on Mars (someday). I'm not a chemist, so I assume smart people are working for Musk. Aerobraking I took from the Musk plan. Assuming it would work for a smaller craft. Now that I think about it, it isn't clear how the dragon capsule and the glider would work together. I guess they don't which is too bad, because I liked the escape feature. Maybe a fairing style of glider skin with a spare one on the main ship for the Earth reentry. Needs more work... As for re-use, returning upper stages to Earth seems too expensive still, so my thought there was to re-use the upper stages in space. Even though the upper stage used for the Mars mission would return to Earth it seems that the 5 refills and two re-entries would be enough re-use . Maybe burn up on Earth re-entry with a capsule landing instead of a glider landing? Accurate landings on Earth have been going very well for SpaceX hitting the bull's eye 18 times in a row now. I had though both orbit and land drops, no ISRU at first, though setting up the equipment for future missions might be a part of the short stay mission. The main idea for me is to re-fuel at Mars and make the manned trips as short as possible. The important thing about the Raptor vs the Merlin was the fuel type as RP-1 isn't going to help anyone. 3 MN would sustain gravity acceleration for a 300 ton ship with a single engine! Artificial gravity anyone? It is also more efficient. Another thought is that having at least two engines would be nice redundancy should one fail on landing for example. They did develop a scaled version of the Raptor that might work better for this application. Hydrogen vs Methane - still not convinced either way. Ammonia or Ammonia hydrogen-capable might be good too. Don't know. If you're saying that chemical rockets are more efficient than solar thermal for LEO to staging area (EML-1 or high orbit), then why bother with staging at LEO at all? Just go straight there, robotically assemble/refuel at EML-1.
  6. "Spiral" was not the right word, still learning the nomenclature, I just meant to say get to EML-1 efficiently and so, slowly, using solar engines. I guess I'll try and lay out a first draft scenario... Main ship assembled and fuelled at LEO, efficiently, slowly, using solar energy, reach EML-1. Mate the upper stage with crew capsule of a Falcon Heavy fitted with a Raptor engine to the main ship which has all the habitat module, fuel tanks and solar thermal engines/concentrators. Accelerate the ship with the Raptor engine. Use the solar thermal engine to accelerate and decelerate to capture orbit at Mars. Separate the upper stage and crew capsule for a Mars aerobraking landing. Previous drops have fuel and life-support , habitat etc.. Orbital refuelling of the joined ship and similar return journey to Earth. Separated main ship parks at EML-1 until the next mission. Following the Musk plan for colonization, more ships would be added for each launch window. No numbers yet or even a feasibility inkling... I agree about hydrogen, I was thinking methane if long term storage is needed or hydrogen if usable quickly. Boil-off seems to be an issue. SpaceX numbers? Don't know either way. They may have a secret/alternate/improved/yet unknown plan that differs from the original, would not surprise me. I'm not keen on the drop 100 people off for 2+ years on the first try without even knowing if 38% gravity is enough to sustain life that long... Why EML-1? The idea of getting LOX from the lunar surface is appealing, if not in the near term, in the long term since getting Oxygen from the Lunar surface is less difficult than from Earth and can reuse upper stages of rockets as vehicles. Lunar LOX could also be dropped at some Mars orbit and on Mars for refuelling. This all assumes it is cheaper to get Oxygen from the moon than the extra cost of lifting it from Earth to space. Also, a mass driver might ballistically launch fuel from Earth or Luna to EML-1 without much rocket Delta-V, basically just manoeuvring thrusters. Raptor Upper stage? That means 2..3.5 MN of thrust, perhaps convertible/augmentable with hydrogen instead of methane (it can be done in IC engines): “Nevertheless, Raptor itself is clearly well on the way to full production, partly thanks to the US Air Force. Yup, SpaceX isn't the only organisation interested in this technology. From 2009 to 2015 Raptor development was funded solely by SpaceX, but in January 2016 SpaceX pocketed $33.6 million from the US Defense Department to develop a prototype version of an upper-stage variant of the Raptor designed to be used on the upper stage of a Falcon 9 and a Falcon Heavy.”
  7. Enthalpy's threads say it can be done: Solar Thermal Rocket Manned Mars Mission Non-Hohmann to Mars Here's a Delta-V map to get an idea of the scale of things: I like Enthalpy's idea of using STR (Solar Thermal Rocket) to spiral up from LEO. Maybe it was implied, but the manned capsule would not be aboard for the slow trip, it would rendez-vous with it at, say EML-1 where the ships dock and maybe refuel. Doing this would save a lot of launches compared to sending everything directly to EML-1. It seems high delta-V is key for fast transit which is required to keep Radiation exposure minimal since shielding so far requires very large mass which is difficult to move. This is where a STR or Nuclear Rocket or Electric Drive is important, though Musk's ITS also promises 80 to 150 days, but for a 2+ year stay using a Methane-Oxygen Rocket with 382s Isp. . "The transport capacity of the 2016 spaceship from low Earth orbit to a Mars trajectory—with a trans-Mars trajectory insertion energy gain of 6 km/s (3.7 mi/s) and full propellant tanks—was projected to be 450 tonnes (500 tons) to Mars orbit, or 300 tonnes (330 tons) landed on the surface with retropropulsive landing.[32] SpaceX estimated Earth-Mars transit times to vary between 80–150 days, depending on particular planetary alignments during the nine discrete 2020–2037 mission opportunities, assuming 6 km/s delta-v added at trans-Mars injection.[32]" From reference 32 above, SpaceX expects to burn off 8.5 km/s in atmosphere at Mars and 12.5 km/s at Earth, so aerobraking instead of deceleration. Given a hydrogen rocket Isp of 450s, higher speeds could be reached, and would require some deceleration, perhaps offset by the reduced mass of hydrogen? Maybe ~15% faster? Compare to the Nuclear mission in the previous post, SpaceX would take 90 days instead of 124 there and 206 back, though the nuclear short-stay mission is a greater distance... Hmmm.
  8. One year roundtrip, 30 day stay to Mars. Is it possible with solar thermal engines instead of nuclear thermal? We can assume Earth and Mars capture orbits for the solar thermal part and chemical rocket to and from capture orbits. Perhaps with slower unmanned propellant drop missions ahead of mission if necessary.