PB666

Maybe an NGO will pick it up.

Lowering the temperature of CO2 and increasing the pressure gives dry ice, separates from O2. Exactly at 520 kPa (5.15 ATM) and -56.3'C Dry Ice will separate from O2. Carbon monoxide has a triple point of 68.10 K (−205.05 °C)15.37 kPa Oxygen has triple point of 54.361 K (−218.79 °C), 0.1463 kPa. Therefore you can precipitate out CO2, the CO, leaving O2 The remaining CO in O2 can be catalyzed to CO2 and recrystallize by passing oxygen through passing through an electrified glass tube creating ozone, this will lower the combustion temperature (I believe this requires a paladium catalyst and some pressure). Or in the presence of NO can be converted with UV light to CO2 and 03. NO2 can be hydrated to nitric acid and removed by reacting it with the urine waste stream and used as a fertilizer.

Not to worry, you'll be hiding from Galactic cosmic rays.

The short answer is that in a perfectly spiral exit orbit to the planets SOI uses twice the dV as a infinitely short pulse at the parabolic origin (the developing Pe). The equation is roughly a parabola, neither elliptical or hyperbolic. I am working or a program to solve the La grangian for the best orbit when [escape velocity - orbital velocity >> 0.5 * specific thrust * orbital period]. I can tell you that if the Θ of your escape vector then don't burn at all between Θ - 45' to Θ + 45', it only wastes propellant and it does not save time. (Θ ± π/4) or using the Pe as the standard reference, don't burn at all between π ± π/4. You can achieve most of your gains between for thrust between ± π/2. There are problems however, Assumption 1, you don't have nuclear power. (too heavy) Assumption 2, power density of your batteries are not infinite. Assumption 3, your target heliocentric orbit is something other than Earths orbit. Assumption 4 is you are using solar power. Assumption 5 is that your starting orbit is minimal (LEO) and that the Earth is opaque. The earth blocks the sun for the lowest orbits, as a result your escape vector needs to be such that the Pe needs to be place on the sunlit side of the planet. The oberth effect in the last pass of the Pe before exits is the point in which, if you have additional thrust, then the thrust should be applied, you do not have to reach the target heliocentric orbit, this can be done by burning along the prograde as you are exiting Many ION drives can reduce the ISP in favor of thrust, if you can reduce the ISP by two fold then applying it at Pe roughly saves fuel because of the oberth effect. You can change the ISP in KSP by altering aspects in the permanent file. You can use batteries to extend the length of the burn. You can also use hyperglolics to assist with the Oberth effect during the last pass of Pe. You can vector down a few degrees approaching Pe and a few degrees leaving Pe to keep the Pe as low as possible. You can retro at Apo to reduce Pe using highest ISP setting on ION drive. You can bring a chemical rocket to assist with your escape burn. If you account for the oberth effect in exit burns, the effect of spiraling (performance reduction) increases with the absolute value of the difference of specific potential energy between the source heliocentric orbit and the target specific potential energy orbit. So that spiraling out and pushing out the Pe makes hohmann transfers less efficient than if you have a short pulse burn. An example a burn to Pluto or mercury might use 3 times as much propellant a burn to mars might only use 2.3 times as much propellant compared to an infinitely powerful ION drive with same ISP. SME = SKE - SPE (frequently called the Hamiltonian equation) when SME is such that SKE>SPE any positive SME is multiplied by 2 and SQRT of which is velocity at escape, that is the transfer velocity. S = specific, M = mechanical, E = energy, K = kinetic, P = potential. SPE = µ/r and r is to be calculated at the semimajor axis a. Since KE = v2/2 the change of KE, SME is greatest when dV is applied at the highest possible velocity in an orbit . . .this is the core basis of the oberth effect. And any SKE you gain in excess of SPE at Pe is yours to keep when you escape. As you approache Pe your velocity can be calculated with the Vis-visa equation. This can be compared with the PE of a circular orbit of same radius (r), from this you can calculate the SKE you have, that you need and the SPE that you need to overcome. For that reason Spiral orbits and at the last pass of the planet power inefficient burns are both counterproductive.

Centrifuges generate alot of heat. Particularly during the initial startup. Even at martian pressures centrifuges generate alot of internal heat. In general a centrifuge would benefit from lower gravity because the stress along the vector of angular momentum is less, but having said that you can have frictionless bearings and take that out. The radius of bearings for centrifuges are intentionally made small for that reason. Within the centrifuge you have a manifold that at some point has a maximum radius (in which is a mass which conducts heat but if the bearing is small net conduction is tiny) it is at the manifold maximum radius that you have collisions at speeds from 10s of meters per second to 100s of meters per second. The centrifuges we had in our group that were refridgerated required a magnitude more expenditure on issues dealing with vacuum pumps and refrigeration (essentially vacuum pump is part of the cooling system as it lowers the mass flow of gas around rotors) There are two or more ways to centrifuge a livable object. You could spin the entire object (this includes things that are inductively tied to another object but do not allow passive exchange between inhabitants). The other way is to use a manifold within or extended along another object. The sealing mechanism for end-attached systems is complicated, likely to loose gas mass over time and provide friction. Another way is to put the vessel to be turned inside another vessel, then just collect lost gas and return it to the first. This results in less friction an no net-losses. For example, with a minimal neoprene/teflon bearing you could use a light weight oil (such as found in a torque converter) to hold the gas in the system. On Mars or the moon its not too much of a problem because you can torque against the moon. In space its a problem if you want the shell to have zero angular momentum because as friction accumulates the shell will turn and you will constantly be needing to add thrust to stop it, so ignoring that you have a navigation system that compensates for the gyration. Also rotating space craft do not steer as non-rotating craft, so there are times when you might want friction just to kill the rotation and neutralize rotation of the shell. While I doubt we need to apply such technology on Mars, you could test it on the Lunar surface. But, IMO, if you are continually going to be hauling people back and forth to Mars, then its better to have such a system in a interplanetary cycler that has protection from GCR and provide artificial gravity. In this case BFR is not going to be a useful thing for travel to Mars, since at most it just going to transfer passengers between 450km LEO and some equally high Martain orbit. And then they transfer again for a lander (one-way) to Mars. One option to deal with the increased mass is to send them to L1 or L2, then refuel the ships and redirect to LMO. A third option is to have a tiny device with two airlocks that only fit 2 people, they stay in the device for a few hours each day doing say 20 minutes of stress exercises. Then the device can be mounted on a spindle and spun up with ION drives or whatever. Alternatively NASA could find better solutions for dealing with space-flight, like prosthetic devices that simulate some effects of standing on the cardiovascular system. Assuming 100 days in transit, I don't see why this is a terrible problem, particularly when they get to Mars g force is only 0.39. Bottom line is if the system is going to be used in IP crew transfers, then it needs to be tested in orbit.

Once again moving posts to this thread. The problem is more specific to flight. Basic assumption is that Kids will not be playing on the surface of Mars. Logic. Atmosphere is much, much riskier than cosmic radiation. Since they are not going to be playing on the surface then all facilities for childrearing can be buried, which means if you get Atmosphere control issue taken care of, then burial issue, and the gravity issue . . . . . But the child bearers are going to have to pass from MEO to Mars. What is the profile, males and females. Well educated, late 20s to late 30s in age, extensive trained technicians (post graduate + 4-8 additional years of additional training and 2 years in flight related. So here is the problem, the rate of SNP mutation is something like 3.5 time higher in the male gonad that the female. These are largely unrepaired errors (DNA polymerase errors and uncorrected damage repair). In males risk of damage increases linearly from age 13 to Death. In females the SNP risk is largely from birth to 10 years of age (females have a higher risk of autosomal recombination with advanced age). If a male passes through a zone of 150 days in which is gonadal exposure is 1000 times higher than the general population, its as if his guys aged 300 years in a year. So what you would expect in his offspring is a higher than average occurrence of leukemia, lymphoma, schizophrenia and autism in a naive population of males (if you select non-naive males-see reference below). In addition within a few generations, because population is small, a much higher occurrence of rare autosomal recessive mutations. These mutations would increase during the exposure but would not be evident for at least 2 generations and more likely around 4 generations. We can compare this with acute radiation exposure (e.g. a blast) the risk is immediate, only the cells involved in a replication phase are at high risk. Some cells are permanently transformed and others are not. In the case of chronic exposure all cells will be damaged because all cells are exposed in all phases and risk accumulates with time. There will be some individuals, unknown apriori, who are resilient to radiation damage of this type, and there will be individuals who are more susceptible, and so there is an expected evolutionary outcome. If you are talking about static populations with a sustainable resource base (e.g. some village on the s. coast of the Caspian Sea) then overtime the rate of damage will decrease and all individuals will become resilient to mutation due to natural selection. https://www.ncbi.nlm.nih.gov/pmc/articles/PMC4030667/ If you are sending elderly couples to Mars, you hardly care. This has been recommended, don't send reproductive age adults to Mars. Next issue, is there a way to defend against the radiation. 4 cm of water in a shell around the spacecraft will suffice. So if the craft has an aluminum hull, just split the thickness of the wall and fill it with 4 cm of water. That's it. Well, not so, you probably will have to aim at doing a hohmann from E-L2 to M-L1 (taking longer) and then taking a different craft to re-enter Mars, carrying a big bucket of water to martian surface is not wise.

The first question to ask is what they did. They had a orbit at ~50,000 km from the surface (for reference the moon is ~300,000 km from Earth) and they need an orbit much smaller. They started with an orbit that was (surface-radius + 426 km) X (surface radius + 44,500km) to understand what they did and why you need to study Keplar's laws of Plantary motion. The equations that you want to use are the definition of a (semi-major axis) and the vis-visa equation. To avoid confusing radius (e.g. a circle) with radial distance, imagine that Pe as a vector point from the planets center of mass to the point of lowest orbit, this is called a position vector, the length is the magnitude of that vector, so we are not referring to a radius of r but the magnitude of the position vector, the orbital maneuvers change the position of the vectors. Step one: convert altitudes to radial distances. Pe = 3,806,000 meters, Apo = 47,972,000 two: a = Average the two . . . . 25,889,000 three: calculate the velocity at apoapsis. SQRT( µ * (2/a - 1/r)) = 362.28 m/s four: the desired orbit (braking orbit); although they did not say what this was, lets assume that orbits Pe was altitude = 50 km (r = 3430000). the velocity at apo = 345.14. five: at apoapsis burn retrograde (362.28 - 345.14) = 17.034 m/s This manuever now puts the spacecrafts motion into the upper atmosphere of Mars. It would be equivalent approximately to placing a spacecraft periapsis about 90km from the surface of Earth, such an orbit would decay quickly. six: Lets say that during each pass at the periapsis the spacecraft looses 100 m/s of velocity, how many passes would it need to place the apoapsis at altitude = 450 km. This is not what they did, what they did was more compllicated, they started at a higher altitude to bleed off peak velocity and let the orbit decay, what I will do is just simplify. The r-vector is 3,430,000, this then needs a velocity of 3250 at the Apogee. But this is not what we want we want the difference of the velocity at the perigee. So for that we have to apply the vis-visa equation to the perigee. The original velocity of Pe (That is the speed of the craft at Pe) = 4,827 m/s and we want to drop that speed to 3,629 which means we a dV of 1129 m/s seven: After a dozen passes in the atmosphere, prior to the last pass we raise the perigee slightly and then at Apo we burn prograde to put the perigee at 450 km. by doing this they reduced an amount of fuel that results in 1130 m/s of velocity, and the cost was only 17.034 and the cost of correcting Pe by 400 km (3,430,000 to 3,830,000) So this is not a re-entry aerobraking procedure, this is an orbital correction aerobraking procedure. Its a form of a hohmann transfer. https://en.wikipedia.org/wiki/Hohmann_transfer_orbit but instead of correcting at Pe, the first burn (a little longer) is a lower than desired Pe, and aerobraking replaces the second burn, and finally a correction is applied. For a lander you will be coming in upwards of 5,000 m/s . . . and you will not have a delicate satellite. You will have some sort of ablative shield and your perigee will be much lower, if you pass as MRO passed you will fly through the martian atmosphere and back into heliocentric orbit. Because your craft is 10t (very close to the limit of what Earth can launch to mars currently) you will not likely have the fuel on board to slow down 5,500 m/s of velocity to land (which would triple the size of your craft). In this case you need to use the Martian atmosphere to bleed off alot of speed and do it rather quickly. If you have unlimited resources from Earth place a refueling depot in orbit around Mars. In this case you could use the procedure above to place your craft in a parking orbit, refuel, and then proceed to land on Mars. Here is the problem, you need about 4400 dV to land. So the choice fuel outside of Earth are hypoglolics, they have ISPs of around 250 (2500 m/s exhaust velocity). Why am i mentioning this, many rockets you see launched from Earth are using cryogenic or semi-cryogenic engines (sometimes with sloppy boosters of SRB). These can produce ISP up to 466, twice that of a hypoglolic. In fact, your LEO>Mars injection was probably the result of a Hydrolox or Metholox fuel system. The rocket equation specifies dV = ISP*g*ln(Starting mass/Final Mass). We can calcuate a starting mass for a final mass of 10t. The starting mass of rocket in LMO required to retrograde burn down to the planet is 58t. If we assume that 1/10th of that is engines and fuel tanks, that leaves your lander with 42% of its original mass. To correct we need to increase starting mass to 138t. Right now we cannot even place 138t of mass in LEO, let alone send to Mars. So lets imagine we can send 20t of ship to LMO at a time, each ship having 5t of fuel. That would take 20 deliveries. A very slow ION drive type tug could 'handwaving away the power-production issue' deliver 20t to orbit, but that tug would take 20 years to deliver the fuel required, you would need 5 @ 200t tugs to reduce this to a 2 years, so even that is a non-starter. You could have a solar paneled cryogenic refueling station, metholox (ISP 375), you probably could get away with a 35t lander at the LMO station resulting in a 12t lander on Mars (less if it uses drop tanks). In the most speculative mid-far future range is hydrolox refueling station. Potentially ISP of 475. This is a spacecraft at LMO of 29t (again you can lower with drop tanks). Left without choices you end up mulling over the aerodynamics of re-entry. The aerodynamics of re-entry is about as difficult as the rocket science required to get into orbit. This is complex math. The density, temperature and composition of the gas changes with altitude. PV=nRt, but there are particle behaviors, supersonic, transsonic and subsonic dynamics that need to be dealt with. But the ideal situation is that you aim for a Pe in which your KE at Pe (drag losses included) is slightly higher than µ/2r which gives a slight vertical velocity allowing you increase breaking time. This is very difficult to nail down. At this point velocity erodes quickly and then radial velocity starts to decrease rapidly in the direction of -3.8 m/s2. Believe it or not, the higher the initial reentry velocity the easier this is. The reason is that circular orbits degrade by spiraling through the atmosphere, and by definition increase the loss of altitude as a function of time. So really fast entries are a balance between escape velocities, capture braking (MRO-like) and reentry braking procedure and what is idea (provided a great shield) is something between reentry breaking (mostly) and capture braking (a little). At the point that velocity is below Mach 5.0 the shields need to come off and the craft needs to start burning to land. Anything that the craft can provide to keep its radial/verticle velocity closer to zero means more time spent dragging down the speed. But at some point wings will not provide significant lift because of the thinness of the martian atmosphere, at that point they only add weight and should be discarded. This is the point to burn off horizontal and some vertical velocity. If you are doing this right at the end of your re-entry procedure you will have about 400 m/s of residual velocity. This means you have saved 4400-500 = 3900 m/s of DV . . . . . .in the best case scenario you save 17t, in the worst case 100t. So prepare yourself to do some hairy math. If you really want to be a rocket guru, figure out a way to get your ship down to the surface and back to a refueling station (or Earths atmosphere) without using ISRU or the like. I think if you can do that and prove it can be done you will have alot of job offers coming in.

But its not like they would be launching over the S.Indian ocean or some other remote body of water. If you launch over the South China Sea, you also have to clear sea traffic out of the way some of which lacks radio communication. I saw a statistic on this once. Something like a more than a billion people live within 200 miles of the east coast of Asia. Of course there is barge launching, I don't think they would go for that.

This posting had nothing to do with SLS launch system. Finally moved here others have commented about how the Mars stuff. Zubrin is not an expert in risk analysis and his comments should be taken with a grain of salt. Short and sweet. Envirnomental risk analysis is extremely difficult to parse out even when the total risk is known, its not that the risk is absent, it because the methods and numbers of participants required to determine risk are much larger than it is feasible to study. I can only tell you this, at high radiation levels there are associated neuropathies, for example sub-lethal exposure of gamma radiation causes tingling sensations is the radiated areas. People who undergo radiation therapy sometimes lose or alter their taste of food. Risk is not about lab animals, risk spreads out widely a few individuals are at high risk at low exposures, and others can survive superlethal doses. Thats the way naturally breeding populations are. The public sector agencies are not looking at risk from the perspective of how to keep that last man alive, they look at it from the perspective of how to keep the first man healthy and free of lifelong risk. In a vacuum of information they guess at what that level is. I should point out that Zubrin did cherry pick his data, the more recent data is the most accurate, and he refutes it.

Note that Musk was responding to a comment about the ULA Vulcan rocket, that's why I made the point. Its not that he likes hats, its he's very dubious about the progress of the Vulcan . . .So am I.

If I am a very knowledgeable customer, Just let them drop the craft at LEO-2/3rd GTO and use ION drive to get me to GTO and I will keep my payload and thumb my nose at the centaur.

As reported here earlier today the Vulcan Rocket design has its problems. The RL10C engines (smaller nozzle) they want to replace the RL10b-2 engine with have not yet been developed. It has to hurt ULA. FH is here, now, and Vulcan is at least 3 years away.

Whoever made that chart misread the data, thats 500 days on Mars, not in transit. 150 mSv per year for ISS = 0.42 mSv / day. . . . . .Average 742 day stay would be 311.64 mSv 16.34 months transiting to Mars would be 960 mSv 16.34 month on Mars would be 330 mSv. Kerr, Richard (31 May 2013). "Radiation Will Make Astronauts' Trip to Mars Even Riskier". Science. 340 (6136): 1031. Zeitlin, C.; et al. (31 May 2013). "Measurements of Energetic Particle Radiation in Transit to Mars on the Mars Science Laboratory". Science. 340 (6136): 1080–1084. Chang, Kenneth (30 May 2013). "Data Point to Radiation Risk for Travelers to Mars". New York Times. Retrieved 31 May 2013. 660 mSv is twice the average highest exposure on ISS. Where are Robert Zubrin's references for his values?

I should note that while all the other numbers are dT the Mars gives no numbers. Number of days they are using, nor do they give time spent on Mars. IOW, its an invalid comparison.

I would draw your attention to the first post in the thread, because this rely's on a performances of RL10s at such a thrust that cosine losses are minimal. How much are the customers willing to pay for an a high efficiency engine(s) that is only useful a few 1000 dV short of Vorbital that are hard to produce, expensive and currently in short supply. The replacement that would have been useful for such S2 staging to GTO for large payloads . . . .the F-2X was mothballed . . . .so . . . . . . . . . Proof of the pudding is in the eating, while they list 19,200 PL to GTO, their prices (551) are $153,000,000 for 8900 to GTO. SpaceX offers 14,200 to GTO (fully expendable) but 8,000 kT to GT0, they charge $90,000,000 M To arrive at 62KT in LEO you would need 4@ATLAs V 551 = $600 million. So they are only 60% more expensive to GTO and 700% more expensive to LEO . . . . . . . .

10t = 10,000 kg. Atmosphere density on Mars is like 0.6/101 what it is on earth g-forces are about a 3rd. If terminal velocity is say 40-50 meters per second on Earth then on Mars it will be over the speed of sound for an object of density 1, a length of about 2 meters and however wide it is (say 1.2 meters in diameter). 400 m/s and if you account for Mach shock say 0.95 whatever the speed of sound is. Its a lower speed if you choose a higher elevation, higher speed if you choose a lower elevation. If you set the landing velocity maximum to 10 m/s then you would need a parachute roughly 500 m in radius to slow you down (you wont be able to do this), at minumum (without parachute) you would have to burn 500 dV of fuel. Another method is to aerodynamically load surfaces for example a wing, this then allows the craft to fight gravity at the same time conserving lift and providing a horizontal direction of motion which to create drag in, BFR. A particle traveling to its perigee achieves a flat trajectory at PE and there after rises, it momentum is horizontal, and the drag it creates is in the horizontal direction, if lift or thrust is provided in the verticle direction the craft can stay alot long enough to dissipate much of its speed, while eventually letting g-v2/r to craft intercept the ground. With a properly designed entry and lift surface this can get you optimistically down to 200 m/s after which you will need dV to land. Dropping vertically from space is a bad idea, because integral of drag forces with respect to distance is dominated by proximity to the ground, the craft will never reach terminal velocity before engines would need to fire to burn off speed and that negates most of the drag loses from energy. Download RSS (real solar systems). There is a Mars there and you can practice making structures and see which structures need the least dV to land. Be aware that Mars have no ocean, elevation is measured from the lowest point on Mars, which would be the equivilent of measure altitude on the Earth from the bottom of the Mariannas trench. As a consequence your target altitudes are going to be higher than you think. I made a WWI styles triplane that was very effective of dissipating drag forces continuously (sort of), but KSP is not good at modeling mach forces at high altitude. And basically once you get close to M=1 <350 m/s you need to abandon the wing and retrofire the rockets. Remember that the wing that provides lift below Mach 1 does not work the same above Mach 1. Below Mach 1 a wing tilt above 14' to AoA is generally a stall that provided no lift. In space Wings act as a particle deflector if you assume elastic collisions, it means that a tilte between 30 and 45 may be more effective until atmosphere thickens. Guidelines Fuel enough to burn off 500 dV Some sort of Engine capable of producing a multiple of the crafts weight 10kt * g: e.g. 3 x 10,000 x 3.8 m/s = 114,000 N of thrust (RL10b-2 has 110,000 but you cannot carry the fuel to mars, too volatile). Some sort of Carriage that provides lift and drag (expendable) Some sort of technology to avert the effects of re-entry (an ablatable surface) Some sort of landing struts. (on some of my craft I just us piping, the piping can be assembled as the lunar model, the more bends in the piping from the verticle access the more cushion it can give). Some sort of steering engines (landing generally works best with main engine pointing down) Some sort of guidance system (knows direction of ground and space).

They have an F1 F-2 class engine (using a SpaceX 9-engine layout) or RS68 either of which can replace SSME. The reason there will be no build-up not going to happen is the 2 ton elephant in the room. Demand is being treated as inflexible, if you cut the cost on a function 5-fold, demand is very flexible. If FH does not, by the end of the year, have 3 times as many customers for FH that it has right now I would be surprise. That means 3 customers to 9 customers from now to the end of the year . . .lets see. https://in.mobile.reuters.com/article/amp/idINKBN1FQ38W http://spacenews.com/38331spacex-challenge-has-arianespace-rethinking-pricing-policies/ ArianeSpace is also reconsidering the viability of their launch vehicles. The SLS program is so expensively structured that only public sector can support it. If you run a cost up to 1 billion$ per launch with 2 billion$ of gov't support then who in the private sector wants to buy into that. In fact there's lies a very important game-theory metric for a competitor. If a competitor can come in, basically take all of the companies future customers, make it be known that they have acquired these customers and drives the public sector out of the base-line support, the that would drive down the competitor's competition and make what he is doing more profitable.

The problems are much deeper than that, in a choice between any group of parts that can roughly accomplish the same goal as a different group of parts, they choose the most expensive most time-consuming to make options. SLS's PL>LEO is not that great, and everything is based on cryogenics, which have (as noted above in this thread) a limited shelf life. Block 1 will be wiped by FH. All SpaceX would have to do is to create a refueling depot, since the majority of heaviest payloads in or passing through LEO is fuel, you get the fuel to a depot and in two flights you have satisfied block 2 (130 kT>LEO). SpaceX could launch 2 FH a month, it would take them >6 months to do the same. If the price was double you could justify SLS, the price now is already in the 10s of billions. It doesn't matter how you get fuel to orbit, the fuel doesn't care it doesn't have feelings.....its just a mass to be bound within a volume, you could do it on FH that blows up every second launch (Just saying) but the fuel that gets to space doesn't care how it got there. You could launch a habitation phase, dock, fuel the depot, get the RSA to soyuz in a crew, and bring the fuel up and you don't have to worry about the crew rating for FH at all. 2020-2021 for first unmanned mission 2022-2025 to get Europa Clipper (unmanned) 2022-2025 for EM2 (first manned mission) Getting a makeshift hub and multiple dockings would be a logistics problem that 3 to 7 years could not solve? Someone please tell me, how is SLS superior to the functionality and the availability we had with STS. With STS you could have a docking port in the shuttle bay, you could bring the crew, and you could use individual fuel tanks to load up a rocket. Yep, we have that now, its called FH.

I think short term ISRU (As in BFR > Mars scheme) is idealized and ill conceived. Find a water honey hole first and then talk about harvesting and production. As long as Space X has not disclosed a spot they intend to land and gather water from we can pretty much assume that all of this part of their future is vapor-ware. My idea was to build space tugs now and haul stable fuels into LMO and stock pile them. Then you only have to worry about having enough dV to gain LMO, for Mars I think its 4800dV or so to orbit. Keep your transfer ship in Orbit and drop a habitation ship. Again its a good idea to bury them if you can. But honestly, if you are going to do that don't even worry about the return flight, just send a pair of mature couples to Mars. The cost of resourcing MARS orbit IIRC about dV 1500 for orbital insertion on an Efficient-Hohmann transfer. This compares to the 5500 to 6000 dv required to land and get back to Mars orbit. However absolute cost is a matter of Payload, and even with chemical rockets you can get dV in the 11,000 range if the PL is small. So basically create a coffin sized PL with enough room for 1 crewman in a space suit. Don't even worry about docking ports, just let him space walk to the return vessel. You could even refuel the vessel from LMO if you had a small docking port and send it back for a second pickup. The priority is on sample return, not people return, if you are going to colonize, people deciding to go back home is not something you want. While I appreciate the effort that Musk is making to go to Mars, it is in his best interest to support the discovery missions first, discovery is about resource analysis, that is why you want geologist and chemist on the ground, not breeding pairs. The issues regarding Mars is where is 'where is it feasible to land and develop (ISRU) and after you have supplied enough resources to maintain the colony then do you have the resources for return flight. This means you need many missions to different sites and autonomous drilling vessels, etc. My opinion is that return flight ISRU exploitation would kill the overwhelming proportion of emergent colonies. This has nothing to do with their plan to develop BFR and send a ship to Mars, its about what you do on Mars once you get there.

Though I certainly hope they change, because it would be of benefit to both SpaceX and NASA if the cooperated on the Manned mission more than just feeding them publicly available information. This is just orion deployement and its first functional missions are going to be to EM https://en.wikipedia.org/wiki/Orion_Service_Module#ATV-based_module By mission they mean they are going to have 13 cubesat satellites delivered to orbits near the moon. My guess is 2020 or 2021. This will get no-one anywhere by 2024, basically.

I see the slide presentation was from 6 years ago, I see that dragon is not going to Mars or Moon, not even sure if its going to be crew rated, I see that FH is not being directed at Mars exploration. Basically he chose a set of vaporware and speculated about it. The question is why you are presenting this now as if it could happen when Space X says its not going to happen with F9 or FH?

I don't understand how this ended up in the SpaceX thread, the Rocket Zubrin defines is more or less a ULA based design (4 SSME), it gives no details on the second stage of the rocket and shows a PL which is larger in diameter than any rocket can currently launch. Then in the midst of the slides he tosses in a Dragon with almost no details on how this would work in his tethered Mars scheme.

How many launch pads are they going to have at boca chica? I should add . . . . . how do they intend to cycling between F9B5 and FH at 39A, even that could present problems in the logistics. The correct point to judge this is when they can demonstrate the promised turn-around times on launches for F9.

FH is a commercial transport system as it is designed, its not a grocery store for a future Mars mission, at least as it stands BFR is being designed to do that. This discussion is now dragged over several threads and I should remind you that, it could be, but as of yet the most massive and bulky commodity is fuel, and the BFR refueler is not ready anytime soon, so that pretty much ends that discussion. FH will be improved, thats inevitable but only at a timeframe that the marketplace dictates. Lets at least try to keep the two sets of eggs separated. SpaceX has defined their first attempt. Their best Mars windows are in the 2028 to 2036 for Mars, which should give them enough time to test the BFR system.

But building a 'grasshopper'- like BFR, given their current state of technology is not revolutionary. Their learning curve was 2012 to 2014, they are now the world's leading expert is doing it, note:they are still testing so. I think when they say engineering they are talking about New Product engineering, which means FH becomes current product refinement which despite what they say never ends until you discontinue the product deployment; and even after that you probably have a couple of design and safety engineers go in and write a complete report on the evolution, problems, solutions and remaining problems. BFR breaks into to three aspects, the heavy lifter/tanker, the interplanetary transport module and the booster. FH is a heavy lifter, F9 series went from a PL>LEO 8500 F9v1.0 to 22,500 F9FT (probably higher now) to 63,000 w/FH to whatever it will be with B5 cores and boosters. The booster platform is not a spectacular jump, its just an expensive one, but given how much they already know probably not that expense is not so great. How actually expensive it really turns out to be depends on the number of customers for FH, if there are not many customers for FH, then its not probably a profitable venture . . . .its more or less Musk's hobby. Where as the interplanetary transport module is completely (as or right now) without any public or commercial support and will likely not have commercial support for a long while. I would be surprised if we see anything on the transport module anytime soon. Atlas Rocket System- Started in 1950s, Atlas II, 1991. Since 2000 there have only been 4275 launches with 1 failure, that is roughly 5 2.27 launches per year and a failure rate higher than STS. Delta IV Has only launched 35 times. over the last decade (3.5 times per year). 1/36 losses (greater than STS) But for the RS68A it is not clear to me what the commercial value of this system is. Vulcan (under development) Yeah, RL-10 is not a high performance engine, its basically a high dV engine you want to use once required dV < 20% of insertion requirement of any heavy lift to orbit, any attempt to multiply these engines either means the PL diameter has to increase markedly or the reduce the nozzle length and diameter. This problem was easily predictable, if fact I brought the issue here months ago. From what I understand the bell of the RL10b-2 rocket is the hardest part to design and the most expensive part of the engine to make, an RL10b-2 is already very expensive and redesigning it (RL10-C) as a S2--->Orbit multiblock engine is going to be really expensive. https://en.wikipedia.org/wiki/RL10#Variants Variant RL10C RL10B2 Mass - 190 227 kg Thrus - 102 110 kN <----- note the loss in thrust Vexhau - 4400 4532 m/s <----- note the loss in ISP Lengt - 2.2 4.14 m DNozzle 1.44 2.2 m <----- note the reduction in nozzle diamater Why would there futurespace rocket rely on the most expensive version of a deep space engine available, modify the engine (increasing the cost and reducing performance) and then add 4 of them and then add that to a new Centuar V rocket? Make sense to you? https://en.wikipedia.org/wiki/Centaur_(rocket_stage)#Vulcan-Centaur Then replace that with this. https://en.wikipedia.org/wiki/Advanced_Cryogenic_Evolved_Stage By the time they get this baby into orbit (2024.5) , BFR will be a thing. What is their primary goal? Anyone heard of ION drives (HiPEP) and solar panels. Falcon heavy could launch 100 of these for the price of one Vulcan launched vehicle and 1 ACES. One of the problem with ULA is that all their proprietary technologies have patents that have mostly already expired. That would be a competitive problem except for the fact that no-one except them really wants to use them, they are inefficient. The fact that you see none of their competitors, either here are abroad just stealing these technologies is basically a description development problem. Bash, no, tell me where the heck they think they are going?