Moonguy

Consider an unmanned lunar cargo lander that delivers 5,000 kg to the lunar surface. The delta-v for landing from a low lunar orbit is 2,100 m/sec. and the Moon's gravity is .16 Earth's. Now, consider that same lander used to land a payload on Mercury from a low Mercury orbit where the delta-v involved is 3,200 m/sec. Mercury's gravity is .38 Earth's. All of the physical characteristics of the lander must remain the same, only the payload mass can be altered. How much payload must be off-loaded to accomplish the landing? Thank you. . .

The Shuttle experiment ended when the tether snapped. It was one of those embarrassing failures no one at NASA likes to remember.

This is huge news, but I am curious why it was not posted under 'Medical Science'?

Exactly how do we define a 'killer' asteroid? Asteroids currently zipping past Earth range around one to two kilometers in size. Is this large enough to prompt the mass extinction events that have occurred periodically? Or is a larger body required?

Actually, this sounds like the '60's sci-fi movie 'Marooned'. Only it was the American (Apollo) that was stranded and a Russian spacecraft (Voskhod?) came up to rescue. A good yarn if patently unrealistic. . .

Actually, they would need a pressure suit. You can deliver oxygen to the lungs at sufficient pressure without a suit. The problem is that, on Titan, the air pressure is 45% greater than at Earth's surface. Our skin surface, like everything else about us, evolved in Earth's atmospheric pressure and is suited to that. If you increase the pressure 45%, you would get gas diffusion effects into the subcutaneous tissues and eventually the blood stream. Not a good thing when cyanide is one of the chemicals in the 'air'. Other chemicals like ethane permeate into the tissue - in this case at higher pressure - and exert changes to blood cell DNA. It is one of the causes of certain forms of leukemia.

1) The Tanker departs on a solar sail, so we only need the easiest departure point. Most cargo missions will depart from L3, but the Tanker takes on water from the Moon, so Tankers depart from orbits closer to the Moon but with easy departure velocities. That would be either L1 or L2. On reflection, L2 might actually be better as it does not interfere with operations (depots) in the L1 orbit. 2) The Tanker is an expendable unit. There is no need to worry about coordinating or synchronizing the Tanker's orbit for either picking up the crew (for the return to Earth) or for follow-on missions. 3) Also, the Tanker's orbit only needs to intersect Mercury's orbit at one of its nodal crossing points just once. A one-year circular orbit which intersects Earth's orbit and subsequently intersects Mercury's orbit at a node, while Mercury is passing through that node, should be fairly easy to establish. It just may take a couple of years for the sail to spiral into the appropriate orbit. That is why producing propellants from water is preferred over storing cryogens for 'years' before the crew can use them. 4) Water can be stored passively. Cryogens would need liquefaction equipment even with a small boil-off rate to avoid gas build-up. With no crew onboard for most of the time, the chances for problems increases. 5) The MOTS we discussed earlier is deployed before any crews launch to Mercury. The 185-ton mass includes the MOTS' dry mass, the dry mass of two manned Mercury Landers and 100 tons of water. The crew should be able to produce propellant sufficient for return to Earth from that for at least one mission. It depends on the mass of the Crew Module payload and that is still being worked out for this concept . 6) Another advantage to expending the Tanker is that the crew does not have to constrain their stay time on Mercury to the Tanker's orbital period. They can extend their mission stay time indefinitely if needed. Alternately, once the crew arrives at Mercury, they would have at least a year before the Tanker is anywhere near Mercury. Whether it would be near enough for use in a return flight. . .I just do not know for certain at this point. At present I am working on a piece for the Facebook group site concerning Sheltering the crew on Mercury. I hope to have it posted before the weekend. You might want to check into that to get more details. . .

1) All water/cryogens comes from the Moon. There is no Depot involved. All water/propellant transfers are direct to vehicle. 2) 'Departure Stage' is a misnomer. The stage carrying a solar sail only delivers a payload of 15 tons to the L1 point. The 15 ton payload includes propellant production equipment, a habitat module and the solar sail. At L1, the tanks of the delivery stage are filled with water, the solar sail deploys and the entire assembly departs into a heliocentric orbit with a one-year period. This orbit intersects the orbit of Mercury at one of the nodal crossings. 3) The sail and its payload - now called a Tanker - may have to make one or more orbits before the crewed vehicle is ready for launch. Preserving cryogens for that long is difficult. 4) The Crewed Vehicle launches 'some time later' as Mercury is positioned for encounter. The Crewed Vehicle launches into the same exact orbit as the Tanker when the Tanker encounters Earth. Assuming a successful rendezvous and docking, the crew occupies the Tanker Habitat for the cruise phase of the flight - approximately 78 days. 5) Prior to Mercury encounter, the crew processes water into cryogens and fills the tanks of the Crewed Vehicle's propulsion system. At encounter, the Crewed Vehicle performs the MOI and rendezvous with the MOTS. 6) To return, the Crewed Vehicle performs the same maneuvers as at launch from Earth. The Tanker used for return would not necessarily be the one they arrived on. The Tankers are single-launch items and regarded as expendable. their total mass when filed with water is 185 tons. This demands very minimum mass for the Crewed Vehicle injection mass as all of the maneuvers are budgeted up to 10 km/sec.

The 'Cycler' concept has dropped off the radar. For now, the idea is to focus on the first few ('Alpha') missions based on non-recovered systems. A departure stage is sent to L1. This unit carries a payload of 15 tons. This includes a solar sail (3880) kg, a self-deploying habitat (NOT a 'tin can'), and a docking module. The stage is equipped with water electrolysis units (as in the case of the Cycler), liquefaction and propellant transfer equipment. At L1, the propulsion stage's tanks are filled with up to 100 tons of water. The solar sail and habitat module are deployed and the cluster departs from Earth on a transfer intercept course to Mercury's orbit. This course is elliptical and actually crosses Earth's orbit twice. Much later (1 year+?), a second launch delivers a crewed capsule to the same L1 location where it is refueled with cryogens. As the cluster vehicle approaches, the crewed capsule launches into the same trajectory, where it eventually conducts rendezvous and docks with the cluster. The crew remains on board the cluster for the balance of the transfer, which takes about 78 days. Prior to Mercury encounter, water from the cluster's supply is processed into cryogens sufficient for either an orbit injection at Mercury (with subsequent rendezvous with the afore-mentioned MOTS) or it can be outfitted for a direct descent to Mercury's surface. Such a mission could be supported by existing Atlas, Delta or Falcon systems. the upper stage elements would be derived from the Advanced Centaur systems.

Virtually everything you are saying is why I am confident about the 'transportation problem' (to Mercury) being solved. I particularly agree with you about the need for new ideas to 'percolate'. I grew up in a space advocacy that torpedoed itself with incessant arguing over 'where to go first' or who's idea for how to go was better. . . It made the entire movement a joke. Sadly, I have to admit I engaged in many of these arguments. I like to think I have evolved to where I see the value in 'synergy'. In the present case, this means using Mercury to boost astronomy and planetary exploration efforts and human settlement of Mars and other venues. A Mars settlement effort would provide a needy customer for things Mercury can provide economically. Less clear to me is what a Mars settlement would do to pay for the goods it got from Mercury. Its an interesting question, but I have not had time to work on it.

All of the life-support and agricultural technology issues are already being worked on in the interest of lunar and Mars exploration. The assumption here is that Mercury simply draws on technologies developed with experience from those two venues. The same goes for mining and mineral processing technologies. This is one of the reasons why the Mercury project is so cost effective. There is very little 'Mercury specific' technology needed, at least not in the early stages. Transportation is probably the biggest exception. Solar thermal is my choice for sending crews to Mercury, but has not been considered by NASA for lunar or Mars operations. They still think in terms of the J2-X, lunar water mining and (gulp!) L2 depot/station for departure. Goes a long way top explain why NASA can't sell their program to the public. . .

The entire purpose of the project is to establish a permanent base on Mercury to make Mercury useful. The economics strongly favor a base over a brief visit. So the first crew has the challenge of setting up a base that can house an engineering crew and a science/exploration crew. This includes a lab where hundreds of kilograms of rock samples can be examined instead of the minimal amount that could be sent back on a return flight. The construction work can be done at night under lights with the goal of having the base covered (as needed) as soon before sunrise as possible. As a thought experiment, a 10 meter thick shield covering a base that had a 25 m x 25 m footprint (625 m2) would keep a bulldozer with a 2 m3 bucket busy for about 260 hours, assuming an average of 5 minutes to transport/dump each bucket load. That is just over ten days. Obviously using two bulldozers would cut that down to less than a week. . . On Mercury, the average density of regolith poured to make a shield mass 10 meters thick would have an average density of 1.9 metric tons /m3. A 10-meter shield would therefore weigh 19 x .38 or 7.22 metric tons per square meter of surface on the base. This corresponds to roughly 30 kg/in2 or 13.5 lb/in2. More could be added to just exactly equal one full atmospheric pressure inside the base, if it were inflatable. Essentially zero pressure differential increases safety margins. During the daytime, the crew would remain in the base and maintain the base's agricultural system. The 88-day daytime is long enough to see some crop plants all the way through their growth cycle to harvest. Chances are the crew would have the garden up and running long before sunrise, so the second sunset would see the crew with a full harvest in place, all of the critical systems set up and perhaps hundreds of rock & soil samples examined before the next crew arrives.

Radiation dangers are about how much exposure you allow the crew to have. The mass of the crew module for the inter-planet transfer phases is critical in this respect. Right now, the concept of choice is for a laminate radiation shield and use of both potable and waste water (in separate units!) to absorb radiation. It is easy to design shielding mass for any given material or even a series of layered material. The hard part is putting it all together in a mass that can be launched for well below the national debt in cost. The most dangerous exposure time is the period when the crew begins its descent to the surface to rendezvous with the pre-positioned surface elements waiting for them. If the crew descends over the night side, they are blocked from the Sun's heat and most deadly radiations but the spot on the surface where they are to rendezvous would be just a few days (less than a week) away from sunrise. Not enough time to deploy the first base (~ 1 day) AND cover it with enough regolith to secure against radiation AND deploy energy units, So the crew descends over the day side where they are assaulted by the heat from Mercury's surface as well as the Sun. Putting the base near one of the poles and having the crew enter Mercury orbit over the poles eases the problems somewhat. What I do not know at this point for a certainty is what thickness of regolith is needed to reach the NASA baselines for exposure limits. For that reason, I just assume a depth of ten meters regolith all around. That is a lot of regolith to move around. . .

Good clarifications, thanks! Regarding the reflector/concentrator, why are they made of anything rigid at all? Solar sail films already exist with 10 GRAM/m2 density. They ;also have 90% reflectivity. The materials available - at least the samples I have held in my hands - are strong enough to be integrated into an inflatable structure or even a mechanical arrangement like that in an umbrella. Adding scrim or other items to strengthen the reflector material is easily accommodated well with in the 1 kg/m2 mass limit. The system you describe has a lot of flexibility. I have to think a solar thermal stage would be easier to develop than a nuclear thermal stage. The higher Isp (1267 vs ~950) sure looks like ample incentive. Sort of gets you to wondering why. . .

This occurred to me while I was reading the earlier posts: Would it not be possible to channel the concentrated light into the thrust chamber through fiber-optic cables? This would make the orientation of the reflectors independent of the thrust vector. The light from the fiber-optics concentrates inside the chamber as in other concepts, but does not require the chamber to be designed with exotic materials in order to allow light through. I have not had time to check it out, but it would seem this could scale sufficiently for use by a manned vehicle. . . especially one going to Mercury (or Venus).

With upwards of 1250 sec. Isp, I too am interested in Solar-thermal for sending crews to Mercury. Given that it uses Liquid Hydrogen, I conceptualized a stage with a crew cabin integrated into the stage tankage so that the LH2 provides the radiation shielding (most of it anyway) without adding mass for shielding. The STR is only really need for the MOI phase. Even 10 km/sec MOI only require mass ratio around 2.7.

Many apologies for being so long delayed in getting back to this thrill-a-minute subject!! I did get a couple of 'analysis' about cyclers back from some associates. One was incomprehensible to me. the other basically concluded with 'deploy a lot of cyclers and it might work.' Not particularly useful. . . For some historical reference here, I originally worked out a scheme once suggested by Samuel Herrick, the originator of NASA's early navigation systems. The idea was to do away with going into orbit at all and just do a direct landing from a vehicle on a flyby trajectory. The manned lander would have to be pretty beefy, but it was still lighter than a full up MOI stage + lander. The landing velocity would be several times that for a lunar landing, but the payload was small, only about two tons. The big problem, at the time, seemed to be nowhere for the crew to go if the lander could not land. I hope we can discuss this more in the future. . .

I agree about Cyclers being 'less straightforward'. What is at issue, aside from crew safety, is the on-going cost of using Cyclers as opposed to other alternatives. A system of Cyclers designed to provide at least one flight opportunity per year seems like it would be hideously expensive. In fact, if we stay away from using the SLS units (at $2 Billion/per launch) it is possible to develop a Cycler system well below the costs quoted for Mars cyclers. That is a particular question I am developing as it is the whole center of the issue: 'What is it going to cost?' Once I have a good picture of just know complex the Cyclers are, I can be more detailed about that. Thermal management for the MOTS is not solely the strategy of flying over the terminator to minimize heating. It has thermal rejection built into both the propellant manufacturing system and the crew habitat. Your idea of using radially-mounted radiators (or are they inert thermal 'vanes'?) is very practical.

As noted earlier, this idea (Cyclers) is complex enough to warrant a computer-based analysis of the orbital configurations over a long period of time.. I have contacted a number of sources and will see who raises there hand first... I think it fair to table further discussion on Cyclers for a week or two that info is in hand. I'll address the matter at that time. When the project started, the first concept for crew transport was 'Mercury Direct' where the crew left Earth on a combined Earth Return/Crew Habitat/Mercury Lander Module, but with no MOI stage. The crew just landed on Mercury directly from a flyby orbit. A neat concept that did not require Cyclers, the MOTS, the Mercury Orbit Insertion stage or orbital refueling anywhere except for the Departure from Earth. Still, it has issues of its own and is a very high-risk approach. I'm ok with the risk level, but that is a difficult hurdle to get most folks over, so I have not pushed it. Manned spaceflight is always about making trade-offs between conflicting requirements. Nature never gives you everything you want or need. I'm not particularly concerned about what method is used. All I'm concerned with is choosing an approach that is sustainable in terms of cost and complexity. So, moving on. . .

At least you had orbital mechanics! I learned mine on the street. . .and it probably shows. The ideal rendezvous is one where the Cycler and Mercury are at exactly the same point on Mercury's orbit at exactly the same time. Practically speaking, this never really happens. Both the Cycler and the Crew Vehicle need to have some margin in their ability to self-modify their orbits. the trick is to not have to modify it by very much. What I gather you are saying is that the Cycler will tend to rendezvous with Mercury on successive orbits, but at progressively greater distances. This may very well be the case, but I'm thinking it will need a computer analysis to say for sure and by how much. There is a very interesting article in the current issue Astronomy Magazine, by Robert Zimmerman, that brings into question the availability of water-ice on the Moon. It does not make a conclusion about it either way, but it really gives you a reason to have a good think before planning anything.

It was my understanding that to encounter Earth from L2 required a Delta-V of about 1.45 km/sec. If so, a propulsion system designed for 9.5 km/sec. total Delta-V would still have 8.0 km/sec. potential Delta-V at Earth encounter. Stay times at Mercury are multiples of Mercury's orbital period. The Mercury-Earth transfer opportunities (ie, launch windows) come at 115.9 day intervals. The Cycler's orbital period is 351 days. This is almost exactly three (actually 3.05) synodic periods, which defines the launch windows. Keep in mind that launch windows from Mercury are open for approximately 20 days, so I'm thinking the '.05' noted here is not a problem even though it is not a perfect, match-up mathematically. Mercury's orbital period is 87.97 days. Multiplying this by 4 gives 351.88 days. The orbital characteristics of the MOTS defines the times when arriving and departing spacecraft can do their thing. Recall the idea of designing the spacecraft to handle Delta-V up to 9.5 km/sec.? Mercury Orbit Insertions can be made as low as 6.1 km/sec. Every calendar year has at least one, often two, flight opportunities where the MOI is we4ll below 8.0 km/sec. For the latter example, that would leave a 1.5 km/sec. Delta-V 'surplus' available to effect a slight plane change if the MOTS was a tad off in alignment with the incoming/outgoing trajectories. Since the 0-180 degree meridians are perpendicular to the Sun-Mercury line every other perihelion (ie. every 176 days) arrivals and departures can occur every third synodic period, or 347.7 days. If I am correct (sound of nails being bitten in background. . .) the departure flights would have to be off-optimum by about 5 days, but still within the 20-day launch window limit. Functionally, I have always preferred an equatorial orbit for the MOTS and an equatorial location for the surface Base. The problem is all the water for propellant is at the poles. If it is there at all. the logistical issues for supporting an aggressive launch schedule are disquieting. . .

On the discussion concerning the orbit for MOTS being in a terminator orbit, I may have had a useful re-think. The MOTS only really needs to be in a terminator orbit to receive incoming Crew Vehicles or launch Crew Vehicles into the proper Trans-Earth trajectory. Any other time the MOYS can be reached from the polar base no matter what the MOTS' orbital orientation to the Sun is. I should think that reduces propulsive requirements for orbital maintenance. The Oberth effect for Earth departures also has another important advantage. The flight time from L2 to Earth encounter would be more than long enough to determine if all the systems on the spacecraft had performed properly during the injection burn. If not, there would likely be time to effect a direct entry abort.

I'm not at all discouraged by anything you have contributed to date. I knew going into the study that this would be a complex undertaking. There are 1001 details I have yet to think about, much less resolve. That is why I brought the matter to these forums. Alas, this is the only forum where I have gotten any kind of credible, useful feedback. Pretty sad when you consider some the minds that dwell here. . . On the up side, I have yet to see anything that has 'SHOWSTOPPER' stamped on it. I only wish I had more time to work through these issues. . .

One of the objectives of the MESSENGER mission's to characterize mercury's interior make-up due to mass distribution irregularities. The Bepi-Colombo mission will also have an interest in this. All I can say for now is that the real situation is 'To Be Determined'. My guess at this point is Mercury's interior, while definitely fractured and differentiated, is actually fairly evenly mass distributed. I say this because Mercury has a molten outer core that would tend to distribute mass (in a molten state') evenly. This core is very large compared to the planet's volume. I'm not discounting mascon effects over formations like the Caloris basin, but they might not be as pronounced as lunar mascons. More data is needed. . . The 'Sun-synchronous' orbit of the MOTS derives from the initial orbit being established over the terminator, then maintained in that orientation by the solar sail and the occasional higher-impulse maneuvers. The latter are done when the MOTS is lowest in mass due to propellant consumption. As you noted the velocity requirements for Cycler-based Earth departures and Mercury arrival are higher than the lowest possible case for standard Hohmann missions. The Crew Vehicle's propellant tankage must be designed for highest propellant mass required for any phase of the mission. that would be the 9.5 km/sec. you noted. That is a scary number for payloads of 20 tons, which is about what interplanetary crew modules (usually Mars bound) tend to be. A crew module for Cycler missions could reasonably be a third or a fourth of that. This is where propellant mass is saved. Eventually, propellant water comes from both Mercury and the Moon. Initially, actual propellants are brought up from the Moon to an orbit around the Earth-Moon L2 point. The propellants are loaded directly onto the Crew Vehicle. There is no 'depot' required because the delivery vehicles are sized to provide the full propellant loads in one or two flights of the lunar-based tankers (that is, near simultaneous flights of two tankers). A Mars mission, by comparison, might need three or four tankers per mission, making a depot more of a necessity. The departure maneuver is timed for less than 48 hours after propellant loading. I do have calculations for vehicle mass ratios, but, alas, they are based on the assumed use of the SLS Block II large upper stage and its J2-X engines. For a 9.5 km/sec. delta-V at Mercury, the mass ratio would have to be 8.1. Not a happy number, but not hideous for a vehicle starting with full tanks in an Earth orbit only 140 m/sec. to C3=0. With full tanks, the vehicle would have a 2.0 km/sec. delta-V margin to effect an Abort-To-Earth maneuver if needed. . .assuming the problem causing the abort did not disable the engines altogether. So I have homework to do to pin down the engine and propellant requirements for the Cycler option.

The MOTS' orbit is Sun-synchronous. Mercury rotates beneath the orbit so that over the course of 56 days - Mercury's period of rotation - the MOTS will view Mercury's entire surface. The sail that delivers MOTS to this orbit remains attached to perform station-keeping maneuvers.