• 

A starship is not an independent entity—no more than a jet plane is independent just because it can leave the ground.

Imagine for a moment, a fully loaded 747 jet airliner flying from Los Angeles to New York. That's several hundred thousand pounds of airplane and three hundred people. First of all, there's the technology to build that airplane. (Not just one factory in the city of Seattle—but aerospace contractors all over the United States supplying components for every part of the plane, from blades for its jet engines to light bulbs over the seats.) Then there's the technology to maintain it; the schools to train the mechanics, the teachers to teach the stewards; the simulators and mockups on which they'll learn; the equipment with which the plane will be serviced, the specialized trucks and tools and devices; and the men and the training to support all of these levels. There's more than the airplane alone. There are airports. Lots of them.

That means the equipment and technology to lay down a perfectly flat runway two miles long. There are ground controllers—that means radar and scopes and computers. And reservation desks and communications systems to service those desks. Luggage-handling systems, porters, allowances for taxis, parking lots, restaurants, rest rooms—and the people to clean, maintain, and work in them.

Then you have to be ready for emergencies—so there will be rescue planes and fire trucks and medical equipment on hand. And detectives to protect the passengers and Federal Marshals to look for hijackers with metal detectors and psychological profile charts.

An airplane burns fuel—a lot of it. A 747 gulps enough petroleum in a single flight to drive an automobile for a year. That requires refineries to crack that oil (men to build and operate those refineries) and trucks to deliver it, tanks to store it. Passengers need meals, that demands another whole service industry. And entertainment—and specialized insurance—and airsickness bags in the back of each seat.

None of these things just happen by chance; they are designed into the system as it grows. The 747 could not exist until most of the support technology has already developed. What did not exist had to be built. All of it was oriented to fit the needs of the passenger as our culture has determined them. The very existence of the 747 as the kind of plane it is, is a direct result of what our culture considers important to the traveler. Imagine an airplane with fourteen bathrooms!

That airplane is a piece of living America. (In fact, I'm told it's the state bird of Hawaii.) It is an active vital symbol of our national technology, but it is no more independent of that technology than is a bird independent of the air in which it flies. The air holds up the bird. Our technology holds up that airplane.

Now. Apply that to a starship.

Extrapolate the needs of the Enterprise. Her fuel requirements, her crew requirements, her maintenance and training needs, her supply needs, her communication and control structures, her relation to the culture that produced her—and why that culture produced her.

Think about it.

Do you think the Enterprise is really an independent entity?

It isn't. It never could be. Her independence is an illusion, just as the independence of the 747 is an illusion. Sooner or later that ship is going to have to return home to have her exhausted energies recharged. Or, if not home, then to a base with an equivalent technology.

From the invited address of Salter Wherry to the United Nations General Assembly, following establishment of Salter Station in a stable six-hour orbit around the Earth, and shortly before Wherry withdrew from contact with the general public:

Nature abhors a vacuum. If there is an open ecological niche, some organism will move to fill it. That's what evolution is all about. Twenty years ago there was a clear emerging crisis in mineral resource supply. Everybody knew that we were heading for shortages of at least twelve key metals. And almost everybody knew that we wouldn't find them in any easily accessible place on Earth. We would be mining fifteen miles down, or at the ocean bottom. I decided it was more logical to mine five thousand miles up. Some of the asteroids are ninety percent metals; what we needed to do was bring them into Earth orbit.

I approached the U.S. Government first with my proposal for asteroid capture and mining. I had full estimates of costs and probable return on investment, and I would have settled for a five percent contract fee.

I was told that it was too controversial, that I would run into questions of international ownership of mineral rights. Other countries would want to be included in the project.

Very well. I came here to the United Nations, and made full disclosure of all my ideas to this group. But after four years of constant debate, and many thousands of hours of my time preparing and presenting additional data, not one line of useful response had been drafted to my proposal. You formed study committees, and committees to study those committees, and that was all you did. You talked.

Life is short. I happened to have one advantage denied to most people. From the 1950s through the 1990s, my father invested his money in computer stocks. I was already very wealthy, and I was frustrated enough to risk it all. You are beginning to see some of the results, in the shape of PSS-One—what the Press seems to prefer to call Salter Station. It will serve as the home for two hundred people, with ease.

But this is no more than a beginning. Although Nature may abhor a vacuum, modern technology loves one; that, and the microgravity environment. I intend to use them to the full. I will construct a succession of large, permanently occupied space stations using asteroidal materials. If any nation here today desires to rent space or facilities from me, or buy my products manufactured in space, I will be happy to consider this—at commercial rates. I also invite people from all nations on Earth to join me in those facilities. We are ready to take all the steps necessary for the human race to begin its exploration of our Universe.

It was past midnight by the time that Jan de Vries had read the full statement twice, then skipped again to the comment with which Salter Wherry had concluded his address. They were words that had become permanently linked to his name, and they had earned him the impotent enmity of every nation on earth: "The conquest of space is too important an enterprise to be entrusted to governments."

For this next depot taxonomy post, we’ll finally be talking about what people usually think about when they hear the term orbital propellant depot — larger, cryogenic refueling facilities, focused on enabling large-scale human spaceflight missions, performed by a diverse variety of users, and going to/from a wide variety of destinations. The idea for such propellant depots for enabling interplanetary human spaceflight dates back to at least 1928 with the writings of Guido von Pirquet¹.

This blog post will be focused on what I call “Low-Orbit” human spaceflight depots. These are depots located near the lowest stable orbit around a planetary body. This is the first place at which you can realistically refuel or switch vehicles on your way from a planetary body, and is the last place at which you can refuel or switch vehicles on your way down. As a firm believer in the idea of refueling early and refueling often, low-orbit depots are an important piece of infrastructure for any planetary system humanity wants to travel to/from regularly. A later blog post in the series will talk about “High-Orbit” depots — depots operating in fixed locations² located out near the edge of a planetary system’s gravitational sphere of influence, and will include a discussion of where those types of depots might make sense.

Before we jump into the weeds about low-orbit human spaceflight depots, I did want to address a recent train of thought I’ve seen that suggests that just using tankers and directly refueling a vehicle is superior to having a depot involved. While this could easily be the topic of a series of its own, I wanted to briefly highlight a few of the biggest advantages I can see of having a depot vs just using direct refueling with tankers:

• Flexibility: A depot, properly designed, with published, standardized grappling, refueling, and power-data interfaces, can be agnostic about who it gets propellant from and who it sells propellant to. Depots can quickly take advantage of whatever the cheapest source of propellant is at a given time (RLVs, ISRU, propellantless launch, buying leftover capacity from other missions going to destinations near a depot, atmospheric gathering, etc), and can easily service both smaller missions and bigger ones. Tanker-based approaches tend to be a lot less adaptable, typically being optimized for one or two specific vehicles that needs refueling.
• Robustness: With a fixed installation, that only has to be launched once for a long mission lifetime, you can afford to throw way more resources (dry mass, volume, and power) at making rendezvous, prox-ops, docking (RPOD), and manipulation as safe and reliable as possible. This could include beacons, larger more capable (and/or redundant) relative navigation sensors and comms, longer reach capture robotics that minimize the dry mass requirements on tanker and client vehicle alike, etc.
  • For tankers, on the other hand, you want to minimize parasitic dry mass that has to be launched every time, and for the departing vehicle you want to minimize mass you have to carry through large, high delta-V in-space maneuvers. You could in theory carry a nicer RPOD kit on your departing vehicle that you jettison before leaving LEO, but now you’re amortizing that mass and cost over a much smaller number of missions, or starting to add more complexity than just doing a depot.
  • Another question to ponder is with vehicles that require large numbers of refueling events per mission, is it best to have the client vehicle handle all of those docking maneuvers with its (by definition) less-capable RPOD capabilities? With a depot, each tanker and each client vehicle only has to perform one mission-critical RPOD/refueling operation per mission, whereas with direct refueling via tankers, the client vehicle would typically have to perform a larger number of mission-critical RPOD events. A depot would also have to handle the same larger number of RPOD/refueling events, but as mentioned before, can throw more resources at making these as reliable and safe as possible.
  • Longer refueling cycles from using a direct-tanker refueling approach also increases the MMOD³ risk to the departing vehicle, which by definition can’t afford to throw as many resources at MMOD protection as a depot can.
  • A corollary to this is that depots make the most sense if you use them in a way that offloads as much of the refueling-unique hardware/software as possible from the delivery and client sides of the system to the depot. Ideally a tanker would be a minimally modified upper stage⁴, and client vehicles would also have similarly minimalistic hardware needed to be grappled and receive propellant. If you’re doing your delivery vehicle or your client vehicle in a way that makes you question the utility of a depot, that may be a hint that you’re doing something wrong.
• Non-Integerality: Yes I may have made up that word, but the point is that tankers tend to come in integer quantities. Unless you always design your departing vehicles to use only integer quantities of tankers, you’ll almost always end up having wasted propellant. This is especially true given how launch vehicles, and in-space vehicles tend to increase performance and get upgraded over time, and don’t necessarily upgrade at the same rates. If you only ever had a monopoly/monopsony situation where you only had one tanker provider, and only one vehicle needing tankers, you might be able to keep tanker size locked to an integer fraction of the amount of propellant needed, but in reality that isn’t going to happen. So a tanker-based system is always going to end up wasting propellants, and this is even more the case when you have diverse customer vehicle and delivery vehicle sizes. The more diversity you have in your space transportation ecosystem, the more a depot makes sense.

There are probably other good arguments I’m glossing over, but long-story short, unless you’re interested in a boring monoculture world where only one type of in-space transportation system exists, depots make a lot of sense. So, without further ado, let’s jump into some of the taxonomical considerations of human spaceflight low-orbit depots.

Human Spaceflight Low Orbit Depots

Application: Refueling large transfer stages or in-space transports for ferrying people and cargo between LEO, the Moon, Mars, Venus, and other destinations of interest.

Location: As discussed earlier, this blog post is focused on depots located near the lowest practical stable orbit around a planetary body. Details vary for different planetary bodies, as described below⁵:

• For Earth, the low-orbit depot location is in LEO, ideally in the lowest inclination that still lets you hit required departure asymptotes and maximize the number of economically useful launch sites to send propellant, people, cargo, and materials to/from the depot⁶. You probably only need one or two such low-orbit depots, though if you get to high-enough earth departure throughput there may eventually make sense to spread more depots out at different RAANs⁷. Likely for a first human spaceflight depot, as with the previously discussed smallsat launcher depots, you’ll want to locate it in LEO near other human-occupied facilities like ISS — far enough away to be safe, but close enough to conveniently move between each other, ideally within one work shift⁸.

• For a low-orbit depot around the Moon, this would likely be a polar or near-polar LLO⁹, though due to the very slow rotation and practically zero J2 perturbation¹⁰, if you have a lot of non-polar surface sites, you eventually may want multiple smaller depots in equally RAAN-spaced near-polar LLO planes, and maybe one in an equatorial orbit. If you’re trying to do lunar surface missions, having your depot in LLO makes way more sense than in a higher orbit like NRHO, for reasons I should probably go into in another blog post.

• For Mars this would also likely be a LMO orbit, with an inclination high enough to be able to access any points of interest on the surface, while still being low enough to minimize delta-V penalties¹¹, and keep the nodal precession rate fast enough to minimize phasing orbit time for three-burn departures. You’ll likely also have to put some thought into perturbations from Phobos and Deimos¹².

• For Venus, the extremely low rotation speed and therefor very low J2 pertubation may require you to do multiple smaller depots in similar inclinations but equally RAAN-spaced planes, as you won’t pass over a given point on the surface very frequently, and the very slow nodal precession rate could potentially require very very long phasing orbits for a 3-burn departure. Venus has a deep enough gravity well that you do want to refuel in LVO coming to/from, but it’s not trivial from an orbital dynamics standpoint¹³.

Size: As big as you can practically get away with — ideally you’d want this depot to be at least 2x the propellant capacity as whatever the largest vehicle you’re refueling. So somewhere in the 100-2000mT range, or even bigger¹⁴. Early versions will want to be single-launch if possible, in many cases repurposing at least one of the main propellant tanks from the stage that delivered them to their destination as one of the depot tanks¹⁵. Eventually, it may be possible to do multi-tank depots, but if you can do a single launch depot big enough for refueling two missions, you may be better off making more than one depot instead of trying to make the depot super big.

Propellant Types: For low depots, you’re primarily going to be dealing with large transfer stages (Centaur V, Starship, New Glenn Upper stage), which typically use LOX, and either Methane or LH2 for the fuel. Most of these use autogenous pressurization, and use the main propellants for RCS. So most of the depot will be for LOX, LH2, and/or Methane.

• For Mars or Venus you may eventually also want to store liquid CO for some applications, since it’s an easier ISRU propellant, but that remains TBD.
• A lunar low-orbit depot may also want to stock storable propellants, depending on what lander propellants end up being most popular¹⁶.
• You may eventually also want to store some secondary fluids (Helium or Neon for active cooling loops, life support consumables like water, air, etc), but you may not explicitly need a depot for that function.

Some Other Considerations for Human Spaceflight Low-Orbit Depots

• As mentioned before, human spaceflight depots want to be designed in a way to enable offloading as much of the RPO and docking/or berthing from the vehicles they’re servicing. The less parasitic dry mass that tankers and clients have to lug around on every mission, the better.
• Storing cryogens in low orbits tends to be hard — you have a warm planet blocking half of the sky. So launching the propellant in a subcooled state or even partially frozen (i.e. slushy propellants) can help a lot. Also a lot should be done to minimize heat leaks between the cold part of the depot and any hot sections (habitation, power, etc). If you can’t get to zero boiloff, LH2 is a great thermal sponge, and can be used to chill other propellants, and intercept heat from heat sources before being vented. You may have to vent some hydrogen boiloff, but if you’re smart, you can use that hydrogen boiloff on the way out to eliminate boiloff issues for everything else.
• These depots are also big debris targets¹⁷. Deployable MMOD/MLI¹⁸ solutions could be very helpful to avoid a puncture, which would probably be very hard to patch. Since these depots are fixed, and are typically only performing stationkeeping maneuvers, it may be possible to augment their MLI/MMOD protection over time using in-space assembly/manufacturing techniques¹⁹.
• Especially for low-orbit depots around the Moon/Mars/Venus, there may be a benefit to having some temporary habitation/shelter collocated with the depot, especially if you’re supporting multiple sites, as a search and rescue option during exploration phases, and as a stop-over point to act as a buffer between different sizes of transportation between planets and between the depot and the surface.
• Over time you may want to add in other facilities such as dry docks for assembling, and repairing/maintaining large in-space vehicles/structures, habitation facilities, etc. But they should probably be coorbital, near the depot, not attached (as that will make cryo thermal management all the harder, and the depot is a big hazardous work location, where you should probably minimize the amount of time people spend in close proximity to it). This could be done in two ways — coorbital facilities, spaced where the safe time to travel from one to the other is as short as practical (definitely less than an 8hr work shift if at all possible, and much closer if possible), or by having the two facilities connected by a connecting tether or other structure that includes elevator facilities.
  • If your depot facility starts wanting to have permanent collocated habitation (say for in-space assembly/repair/maintenance of in-space stages), and having the two be coorbital doesn’t work, you’re likely going to want to keep the people separated as far away from propellant tanks as possible, both to minimize heat-leak into the tanks, but also to minimize hazard to the people²⁰.
  • For larger depots, and ones where people will be there more, putting some thought into spatially separating the fuels and oxidizers more could be a good idea. In rockets you often can’t do much to keep the two separated, and many use a common bulkhead, but in a fixed facility it’s more of a possibility. Having fuel and oxidizer that close together for long periods of time is somewhat tempting fate — you kind of have to do it for high efficiency rockets, but there’s something to be said for having your fuel and oxidizer many meters apart for a long-duration facility.
• One area of disagreement I have with other depot advocates is whether propellants should be shipped to a depot as cryogenic propellants (LOX/LH2/LCH4), or if you should ship them as something more storable like water and CO2, and have the depot itself have large-scale electrolyzing, separation, and propellant refrigeration systems. My concern is that while in theory very large solar arrays could be done in space, combining large flexible structures like multi-MW solar arrays and radiators with a facility that sees a lot of docking, propellant slosh, etc seems like a bad idea from a structural dynamics standpoint. Also depots with very large power generation and heat rejection capabilities are likely to come later in the process, since they’ll almost certainly require multiple launches and in-space construction.

Anyhow, I probably could go on, but as with the previous parts of this series, I am only trying to scratch the surface with considerations and operating details, as I introduce each new type of depot. This definitely isn’t the last you’ll hear from me on the topic.

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1. Sykora, Fritz, “Guido von Pirquet-Austrian Pioneer of Astronautics,” History of Rocketry and Astronautics, R. Cargill Hall, ed., AAS Publications, San Diego, 1986, p. 151. And yes, I’d love to see someone actually fly a depot before we hit the 100th anniversary of von Pirquet first discussing the idea
2. Or at least orbits/halo orbits where the only maneuvers being performed are for stationkeeping purposes
3. MicroMeteorite and Orbital Debris
4. With some lightweight grapple fixtures, refueling modified T-0 umblicals, and maybe an upgraded comms/controls system
5. Most of these points could easily justify their own blog post, and knowing me, I’ll probably eventually do said blog posts, but I’ll try to keep things in this post brief.
6. I’m currently noodling doing a paper looking specifically at this optimization — higher inclinations for the depot give you the ability to hit higher departure declinations, and are accessible by more launch sites, but at the cost of slower nodal precession driving longer hang-times for a three-burn departure, and you also have decreased payload mass to higher inclinations. My gut says ISS-like orbit is probably not far from the optimal point, but it would be fun to run the numbers
7. As a reminder, RAAN stands for Right Ascension of the Ascending Node, which is a measure of where a given orbit plane crosses the equator heading northward. Here’s a decent wikipedia explainer. When you see a multi-plane constellation, like say OneWeb or Starlink, in a lot of cases the satellites will be put into multiple planes with each plane having the same inclination and altitude, but a different RAAN. The different RAANs help with more frequent passes over the same point in the ground, and for depots by increasing the frequency with which a depot passes through the plane of a specific departure asymptote. If that makes any sense.
8. I could easily see a scenario where the first human spaceflight depot evolves from smallsat launcher depots.
9. Though how low of a LLO is open to debate — lower orbits have higher stationkeeping requirements, and are harder to do sunshields with because the Moon takes up more of the sky surrounding the depot. I haven’t done an optimization analysis myself, but 250-500km above the Moon might reduce both of these effects noticeably compared to the 100-150km parking orbits used for previous lunar missions
10. The J2 parameter is a measure of how oblate or “round about the middle” a planet is. This is typically tied to the rotation speed of the planet. The J2 perturbation is what causes orbital planes to precess slowly. If a planet had no J2 or very low J2, the planes would stay in a fixed orientation relative to the stars, which is problematic if you’re trying to get the plane to line up with a departure asymptote for an interplanetary departure… Earth, Mars, and most of the gas giants have high J2s, while the Moon, Venus, and Mercury all have very low J2s.
11. The equatorial velocity of Mars is ~half that of earth, so it’s not as big of a deal to launch from or into a higher inclination, but it still is some losses
12. I’m not much of a Mars guy, so am handwaving that part a bit
13. If someone is looking for a PhD dissertation topic, this feels like an area that could use a lot more skullsweat investigating.
14. Depending pretty strongly on how successful Starship ends up being, and how open Elon is to using other people’s depots. In a world where Starship either doesn’t pan out, or Elon insists on doing his own thing, the required size for refueling other missions can be a lot more modest initially. Also a world with depots in low-orbits around destination planets, and roving depots and/or high-orbit fixed depots may not need LEO depots to be quite so big.
15. As shown in Part IV of this series.
16. For long duration lunar landers there are definitely differences of opinion about whether storables or cryogens make more sense.
17. Big LEO facilities that don’t want to dodge things all the time is one of those reasons why moving to a leave-no-trace approach to satellite operations is going to be an important part of growing up as a spacefaring civilization.
18. MLI, or multi-layer insulation is a type of very effective insulation for use in vacuum environments. It is typically made of many layers of thin metalized plastic films separated by nets or spacers so that most of the heat has to transfer via radiation instead of conduction or convection. Having more spacing between layers can help. And in some cases, MLI can be combined with MMOD (which also wants to be multiple thin layers with spacing between them), like what our friends at Quest Thermal have worked on.
19. Most hypervelocity impact shielding benefits a lot from extra spacing between bumper layers, so an in-space assembled/manufactured MMOD solution could be particularly useful
20. Though as my wife pointed out, if your depot facility starts wanting anything, it’s probably a sign of an impending robot apocalypse, so the point might be moot.

      For there to be a race, do both contestants have to know the objective? Or even that a race is, in fact, being raced?
     Most Western studies on the future of human (as opposed to automated) solar system exploration have applied technology developed in the twenty-plus years since Apollo designs were frozen to reduce the costs of more or less traditional goals: a Lunar base, human Mars missions, or asteroid mining. Recently, however, three men—a man space historian Tim Kyger refers to “one of America's astronautical pioneers,” a former astronaut, and a planetary scientist—have all begun to advocate a new post-Space Station goal.
     This goal is one they believe will be much easier to accomplish than any of the older projects, and thus more likely to win political approval in this age of limited budgets. It would combine into a single project the traditional aims of the scientific exploration of the solar system with the newer ones of space industrialization, thereby enlarging the constituencies of both. And it would be a significant step toward that ultimate goal, the beginnings of solar system trade.
     This new goal is the human exploration and exploitation of Mars’ two small moons, Phobos and Deimos. In terms of total velocity change, these bodies are essentially the easiest of solar system objects to reach from Earth orbit. The realization of this single fact could allow the early establishment of a solar system economy based on the low-cost delivery of water and fuel from these moons to industries in Earth orbit and eventually on Luna—perhaps as early as the turn of the century.
     Six years of paper studies by these three scientists appear to have been largely ignored by NASA. They may not have been ignored elsewhere.
     At the 16th annual Lunar and Planetary Science Conference in 1985, the Soviet Union chose to dump a decades-old tradition of absolute prelaunch secrecy on automated solar system exploration. At a conference attended by the cream of Western space scientists and management, Soviet scientists described in unprecedented detail a 1988 mission to none other than the Martian moon Phobos.

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     Water—for life support, rocket fuel, and industrial processes—is the single most important resource required for the colonization and industrialization of the inner solar system. It is also the heaviest. And worse, down here close to the Sun, it is a very precious commodity, either locked up at the bottom of massive gravity wells (Earth, Mars) or entirely absent (Luna). To find an almost infinite supply, at less delta-V (the sum of all the changes in a spacecraft’s velocity required for a given mission; in this case, a measure of its cost) than any other source regularly available from low Earth orbit, is a miracle which could mean the difference between space industrialization in our lifetimes—or never.
     Yet S. Fred Singer, former astronaut Brian O'Leary, and planetary scientist Bruce M. Cordell believe they have found just that.
     If it were not for their circular, equatorial orbits, it would be easy to conclude that the Martian moons are captured asteroids—and many scientists do anyway. Both are small and very dark; both are saturated with craters blasted into ancient regoliths; both maintain the same face toward Mars, with their major axes aligned toward the center of the primary.
     In other ways, these asteroid-moons are very different. On a small scale (hundreds of meters) Deimos appears rougher than Phobos, while on a larger scale, the opposite is true. Phobos is wracked by what appear to be cracks a few tens of meters deep and hundreds wide, possibly resulting from an ancient impact which blasted a crater nearly a third as wide as Phobos’s longest axis. Crater counts suggest the surfaces of both moons are on the order of three billion years old.
     Phobos's orbit is believed to be decaying at a rate which will result in a Martian impact in a few tens of millions of years; Deimos is very close to Mars's geosynchronous orbit and moves ve slowly through the Martian sky—making it potentially very valuable real estate, as we shall see.
     But the main value of these bodies lies in their estimated composition, which is similar to carbonaceous chondrite meteorites or C-type asteroids; that is, rich in hydrated silicates and carbon, as well as limited amounts of metals.
     Water content could be as high as 20 percent.

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     Evidently, the idea was first taken seriously at the first Case For Mars Conference in Boulder, Colo., sponsored by a group of students and held from 29 April through 2 May 1981. (The papers have been published by the American Astronautical Society in Volume 57 of its Science and Technology Series).
     S. Fred Singer, originator of the Minimum Orbital Unmanned Satellite of Earth (MOUSE) proposal of 1953, based on earlier British Interplanetary Society work, presented a paper entitled “The Ph.D. Proposal: A Manned Mission to Phobos and Deimos." Therein he argued that, even for purely scientific missions, it made more sense to send a human crew to Deimos than it did to send either automated or human missions to Mars itself. He felt the high cost of sending a human mission to the planet, involving the development and transportation of landing vehicles, would make near-term approval unlikely, while automated missions are very limited in the science they can obtain—and are not at all that cheap, in any case.
     But what could humans do for science on Deimos? Well, first there's Deimos itself. Thirty years after the beginnings of the space age, Comet Halley is the only one of the solar system’s small bodies humanity is attempting to explore—despite the fact that these bodies could be of the highest long-term importance, both for science and potential industrialization. [As I write, NASA Administrator James Fletcher has declined to request the CRAF Comet Rendezvous and Asteroid. Flyby mission for Fiscal Year 1988, favoring instead international Earth-orbital projects to map yet again (albeit in greater detail) Earth’s radiation belts and ocean surface topography. Both of these are worthy scientific missions, no doubt, but beyond their obvious military utility, I can't help wondering what they will do to advance humanity's long-term future in space.]
     The biggest problem with automated rovers on Mars is the time it takes for a signal to reach Mars and return to Earth, requiring rovers to be highly autonomous and therefore expensive. Singer suggests that a Deimos base would allow essentially real-time control of therefore relatively simple (and therefore inexpensive) rovers and sample return vehicles. Deimos presents a single face to Mars, allowing simple placement of communications antennae, and is in a close-to-geosynchronous orbit, allowing long communication sessions with Mars-based assets. This, in turn, says Singer, allows proliferation of those assets: a Deimos mission would allow “the sequential operation of many rover vehicles at different locations on the surface Mars in a single space mission, thus increasing the probability of discovering scientifically exciting results" at less cost than flying and controlling a similar number of rovers from Earth (emphasis on original). A further advantage would be the effective quarantine of samples and personnel at the Deimos base, reducing the risk of back-contamination of Earth by still barely possible Martian organisms.
     Carl Sagan, meanwhile, has pointed out to Brian O’Leary that telescopes on Phobos and Deimos could easily provide a resolution on Mars ten times superior to that of the Viking cameras, and could define the most interesting locations for sample returns and later human landings. O’Leary also quotes S.J. Adelman and B. Adelman as stating at the second Case For Mars Conference in July 1984 that a Phobos base could also support gravity-wave astronomy, radio astronomy, and astrometry research.
     Singer thus argued that automated Martian landings controlled from Deimos are easier, less costly, safer, and could be done much sooner than a human Mars landing. How costly? “Assumimg that certain development costs can be shared with other programs, the incremental cost … is estimated at less than $10 billion (in 1978 dollars) over a fifteen-year period,” says Singer.
     If Phobos and Deimos can be all that advantageous and inexpensive for purely scientific pursuits, what might they do for the quest to find ever-cheaper resources for proposed Earth-orbital industries? It did not take Brian O’Leary and Bruce Cordell long to find a very favorable answer.
     About this time, aerobraking, the science of using a planet’s atmosphere to take orbital energy from a spacecraft, was coming into vogue as a way of reducing the high fuel costs of Earth-orbit Orbital Transfer Vehicle aperations. Since velocities into the Martian system are always less than those involved in cis-Lunar space, it did not take much imagination to realize that any aerobraking OTV developed for Earth-Lunar or Earth-geosynchronous transportation (it is actually easier to get to Luna than it is to get into geosynchronous orbit) would be applicable to entering Mars orbit from an interplanetary trajectory.
     According to O’Leary, aerobraking at Mars can reduce propulsive delta-V requirements from 1.5 or 2.0 kilometers per second to 590 or 667 meters per second for Phobos and Deimos respectively. However, “even without aeroassist at Mars, the moons of Mars are more accessible to the Earth at biennial opportunities than is the Moon of the Earth. The chief difference is in the requirement to soft land payloads on the Lunar surface… [Further,] the [Phobos/Deimos] missions permit low-impulse propulsion for the entire trip, opening the possibility of using solar electric, mass-drivers, tethers, and solar sails as sources of propulsion. The only advantages the Moon seems to offer are its proximity and launch window frequency: days versus months or years” (National Academy of Sciences/NASA Symposium on Lunar Bases and Space Activities in the 21st Century).
     But Bruce Cordell has gone even further. In a report done for a large U.S. aerospace company, Cordell considers three classes of human missions: a human mission to the Martian surface in the absence of any “significant” Lunar utilization; human missions to Mars after the establishment of a lunar base; and the possibility that “the next major post-Space Station civilian goal in space for the United States (and its collaborators) should be the exploration of the Martian moons. This program should occur prior to, or at least concurrently with, the development of any Lunar industrial capability.”
     Cordell lists many of the same advantages for early Phobos/Deimos development as O’Leary, including regular launch windows, as opposed to asteroids; round-trip delta-V's which “compare favorably with those to any asteroid:” the fact that Phobos and Deimos have “better specified physical and chemical properties than currently known Earth-approaching asteroids;” and potentially of greatest importance, “propellant production plants on Phobos/Deimos … make the manned landings on Mars independent of terrestrial fuel supplies.
     This could be very important because, on Mars, existence of crustal swells, rifting, volcanism, impact cratering, and abundant water … [possible] hydrothermal, dry-magma, and sedimentary mineral concentration processes … [and] tectonic similarities between mineral-rich Africa and portions of Mars, suggest that the potential for mineral wealth on Mars is impressive. ” Any such minerals would reside within Mars's relatively small gravity well, making their use by space-based industries less expensive than Earth-derived materials.
     The main emphasis of Cordell's study was the development of an economy based on the mining and transport of water from these moons. The first human mission of the Martian system, under Cordell’s plan, would take place during the 2001 window, although he can find no reason why that should not happen earlier. In fact, he told me he now considers his entire argument to be far too conservative.
     In order to keep this first interplanetary mission simple (and thus inexpensive), Cordell proposes an unusually small crew of three, one of whom would be a geologist.
     The 2001 mission would use two reusable Orbital Transfer Vehicles stacked one on top of the other into a single spacecraft. All the fuel of one and 30 percent of the fuel of the second would be used to get to Mars, with aerobraking in Mars’ atmosphere used to slow the remaining vehicle. Sixty days would be spent remote sensing, mapping, probing, and sampling both Phobos and Deimos. The spacecraft then would return to Earth via a Venus swingby, and would maneuver to rendezvous with the Space Station. (Altematively, aerobraking in Earth's atmosphere could replace the Venus swingby, resulting in a shorter travel time but at the price of heavier aerobraking equipment.) Total delta-V for this round trip is 7.91 kilometers per second—compared to 9.0 to and from the Lunar surface.
     Using the information obtained on this mission, “intense” research and development on obtaining water and OTV fuel from Phobos/Deimos regolith material would be undertaken. In 2005 a second human mission would continue the exploration begun on the first mission, and would begin initial work on a base on one of the moons, or a rotating orbital station, depending on human zero-gravity endurance. The crew of this OTV would remain in the Mars system, to be relieved at the next Mars window, two years later. Including travel time, each crew would spend approximately four years in space, which, according to James Oberg, would result in sufficient radiation exposure to limit each crew to a single mission.

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     Following construction of what Cordell describes as humanity’s first base outside the Earth-Moon system, a new class of automated OTV optimized for water transport would arrive, to be loaded with water mined from Phobos or Deimos, and then head back for Earth orbit or any Lunar base. Phobos/Deimos-Luna transport loops require about half the total delta-V of Earth surface to Lunar surface loops. “This is due to the large ascent requirement for Earth; from Earth to low Earth orbit it is 9.7 kilometers per second." Keep in mind that the Phobos/Deimos loop was 7.91 kilometers per second. It could be less expensive to fuel cis-Lunar OTVs from the Mars system than it would be to deliver the fuel from Earth, and certainly “all necessary propellants—for both the Mars-bound and Earth-bound portions of [water transport flights]—are produced completely from water extracted from the Martian moons.”

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      Tickets now available: 50 to 107 light-minutes.

     [Heard on live evening newscast 2197-06-01 on Ceres]

     “Twelve kiloseconds ago (~3 hours), Orbital Materials LLC announced their intention to establish an oxyhydrogen propellant depot on Phoebe, a retrograde satellite of Saturn, within the next decade. A spokesman from Orbital said Phoebe was purchased from a private collector. We have colonization analyst Helen Graves here with us on Ceres. Helen?”
     “Right here, Mindy.”
     “Helen, what are your thoughts on the Orbital Materials acquisition?”

     “Well Mindy, as you know, I’ve studied interplanetary colonization for decades, and the Orbital acquisition seems hopelessly long-sighted. Phoebe orbits Saturn, and there are no present plans for colonization that far out. The closest well-frequented base would be Pasiphae Station, in the Jovian system. Frankly, Mindy, they just won’t have any customers.”
     “What do you think is their aim in acquiring such a risky investment, then?”
     “I’m guessing they intend to bootstrap colonization efforts themselves. Phoebe is undeniably well-suited for it. The moon orbits retrograde, which makes it easier to rendezvous with from certain Hohmanns, especially with slingshot capture tethers. It also has the vast wealth in water to make the fuel itself.”

     “Thanks, Helen. Again, if you’re just joining us, Orbital Materials has acquired the moon Phoebe for speculative use as a propellant depot. Construction will start after the first crews arrive; I’m told Orbital will use higher-energy transfers to cut down on the six-year Hohmann from Earth. I’m Mindy Graham, and this is Ceres Evening News.”

Part of my series on countering misconceptions in space journalism.

Humans like to reason by analogy. It’s a powerful problem solving technique because so much of our experience generalizes, while first-principles thinking is computationally costly. In space, however, thinking by analogy is almost always wrong, because our terrestrial experience shares so little with the realities of space travel.

One concrete example of this is our intuition around fuel. Humans need to eat to stay alive, and similarly we charge our devices and fuel our cars. All of them can perform their essential duties with a well-contained, discrete battery or fuel tank.

Rockets, on the other hand, are little BUT fuel tank, and it’s important to understand why. Going to orbit involves more than flying beyond the atmosphere, which while difficult is comparatively easy. Going to orbit involves going fast! Roughly 7.8 km/s, or 17,000 mph. These numbers are so huge it’s difficult to imagine how to go that fast.

Rockets work by throwing mass out the back, as fast as possible. A really good rocket engine can eject hypersonic exhaust gas at more than 10 times the speed of sound, which seems fast enough for anything. On the other hand, orbital velocity is more like 25 times the speed of sound. This means that a rocket entering orbit is throwing exhaust products behind it literally as fast as chemistry and physics allows, and yet that gas is still travelling forwards faster than its ejection speed.

The exhaust results from burning fuel, and the fuel that’s ejected has to also be accelerated to these great speeds. This is similar to the effect that fuel mass has on the efficiency of long haul jets, but much much worse. In a jet, the aircraft has to carry passengers, cargo, and the fuel it needs to land for the whole journey. In a rocket, the vehicle also has to carry the oxidizer and the speeds involved are much, much greater.

As a result, the final velocity of the rocket increases only logarithmically with the ratio of fuel mass to everything else, a brain-melting problem often called the “Tyranny of the rocket equation.”

Δv = v_e log(Mi/Mf),

where Δv is the change in velocity, v_e is the exhaust velocity, Mi is the initial mass, and Mf is the final mass. Astronaut Don Pettit has a nice summary of these issues.

In space travel, Δv is everything, and it determines how much fuel is needed to go from place to place, as summarized in this handy chart:

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In addition to the absurd 9.3km/s necessary to reach Earth orbit, most other destinations are a reasonable fraction of any achievable exhaust velocity. As a result, mission design is primarily about figuring out how big the fuel tank is and where to put it.

This horrible state of affairs means that it’s basically impossible to get to orbit, let alone deep space, using conventional engineering. In other words, all rocket scientists need to employ at least one crazy idea if they want to get there. The problem with crazy ideas is that it’s hard to tell which ones are almost practical.

As an example, many rocket scientists will reach for hydrogen and oxygen as a high performance fuel, even though hydrogen’s low density and hard cryo temperature mean that the mass fraction of the rocket takes significant penalties. The Space Shuttle is another example of where following wild ideas can lead. SpaceX instead employed lithium/aluminium alloys, materials that are extremely light and almost impossible to weld. There are no easy options – all options involve a great deal of difficulty, and many turn out to be impossible.

At the end of all this, a really good rocket is able to deliver about 4% of its launch mass to orbit. Everything else is structure, engines, and fuel. This is why rockets are really nothing like cars and bicycles and planes.

So it is that, faced with the impossibility of the problem, creative scientists and engineers will face temptation to veer off into wild hypotheticals. Many of the subjects of this blog series deal with the relative impracticality of some of these ideas.

The subject of this blog, after a this fairly sizable preamble, is refueling depots.

The idea behind refueling depots can begin with an analogy to gas stations. Most launch vehicles arrive in LEO with both cargo, such as a satellite, and an empty upper stage. If that stage could be refueled like a car, it would have enough Δv to go almost anywhere in the solar system.

Yes, you read that right. A close examination of the chart above shows that while getting from the surface of the Earth to LEO requires about 9.3km/s of Δv, LEO to the Moon, Mars, or Jupiter requires only about another 3-5km/s of Δv. Since the upper stage of most rockets delivers just over half of the orbital Δv, if refueled they would be able to send missions all over the place.

This fact has been appreciated for a very long time. Indeed, the science fiction author Robert Heinlein said it best when he said “If you can get your ship into orbit, you’re halfway to anywhere.”

Under the status quo, these perfectly good rocket stages are discarded, wasted, and either left to rot in orbit or burned up in Earth’s atmosphere. If a refueling capability existed, existing rockets could also launch larger payloads, or existing deep space payloads could be launched on smaller rockets.

As it happens, I believe that miniaturization and modularity are not sustainable ways to save money on large-scale space projects, a concept embodied in SpaceX’s outsized Starship vehicle. But it turns out that there are bigger weaknesses with the fuel depot concept.

The fundamental problem with refueling from a depot is that, in space, the cost of fuel is not determined by volume or weight, but by location. It’s not possible to economize on fuel by making a hybrid Prius rocket and driving it more slowly, at least not in any conventional sense. Fuel needs to be in the right tank at the right time in the right quantity, and never for very long before it’s used.

The obvious place for a fuel depot is in LEO. Fuel depots further out, such as near or on the moon, have much less benefit. Launch to LEO is the hardest leg, so it makes sense to refuel at each end.

In more detail, when we think of space we think of the night sky overhead, with satellites zooming around. What this disguises, and was ignored in the film Gravity, is that even a set of orbits as restricted as LEO is not one place. In practice, getting from one orbit to another can require more fuel than launching from Earth to that orbit in the first place.

Most orbital vehicles, such as Soyuz, Dragon, or the Shuttle, have 200-500m/s of Δv for orbital maneuvering. Typically, this fuel is used to correct imperfections of the launch, approach and dock with the space station, and then de-orbit.

As players of Kerbal Space Program will know, 500m/s is plenty of Δv to tweak the relative orbital phase, eccentricity, and semi-major axis. It is, however, vastly inadequate to change the inclination or the longitude of the ascending node, which are two other Keplerian orbital elements. Indeed, if we break up all LEO orbits into inaccessible adjacent slots accessible within a 500m/s window, there are about 200 distinct inclinations, and 200 distinct longitudes of ascending node, for about 40,000 discrete orbits.

If orbital refueling was standard practice for every launch, then we’d need to build (and keep supplied) on order 40,000 separate depots. This is obviously impractical, since the cost of building and operating even one is probably more expensive than continuing to manage without them.

In practice, we have to select just a handful of these orbits, just as the ISS needed a discrete orbit. Various considerations can be employed to select these orbits. For example, the ISS, which can be thought of as an orbital depot of people and fatigued aluminium rather than fuel, was placed in an orbit that was accessible for the launch sites and vehicles of the contributing countries.

An ISS-like orbit is a natural choice for a fuel depot, but it would only be useful for a subset of deep space launches. If the depot’s orbital plane doesn’t align with the destination during the relevant launch window, that’s just tough. For the moon, this would mean being able to launch only two or three days per month. For Mars, a dedicated depot would be required for each launch window. There are orbits whose orbital plane precesses in as little as two months, enabling alignment during any launch window, but for only a few days in that time. It is also possible to reduce fuel requirements by changing planes during an extended series of departure burns, but even this strategy requires around 1km/s of Δv.

Additionally, a high inclination orbit like the ISS is designed to be accessible for launches from Baikonur,. This increases the inefficiency of launches from any other launch site, which inflates the operational cost of keeping it fueled.

Some proponents believe that the depot could be topped up by salvaging the dregs of conventional launches. All rockets carry slightly more fuel than they need to to ensure margin for orbital insertion. What little is left over, if it happened to be in the same orbit, could be transferred to a depot for storage and later dispensing. Although such fuel would be “free” as salvage, it’s unclear how enough fuel could be salvaged from a huge variety of launch orbits to meet any kind of demand. If fuel margin on launch is 2%, then the depot would need at least 50 successful salvage operations in close succession to be able to refuel even one additional stage.

Operationally, depot management is complicated by a boom-bust cycle of use. For regular gas stations, usage is relatively steady and predictable, while the cost of unsold fuel lurking in an underground tank for a week or two is relatively negligible. Deep space launches occur infrequently and have differing quantities of use and even propellants. In a future where human flight to LEO or the Moon is ramping up, a depot would have to constantly grow to meet demand, which is a nontrivial requirement. Storing cryogenic fuels for long periods in space is difficult because they heat up and boil off, not least because the Earth is radiating significant heat into the LEO space region.

In practice, an impending deep space launch that required depot refueling would require the depot to be refueled “just in time” by a series of accessory launches, all of which would involve a partly-empty upper stage docking and transferring fuel, only to transfer that fuel back to the probe-carrying upper stage. Given the inefficiencies and losses of doing so, it is much easier to dock the fuel-carrying upper stage(s) with the payload, and perform the requisite injection burn directly. Same outcome, reduced complexity, greatly reduced overhead.

There is a broad exception to the above considerations; a case in which propellant transfer in LEO does make sense. In the SpaceX Starship concept, a fully reusable Starship is refilled by a series of launches while in LEO, before continuing its journey. This is a different model to an orbital fuel depot, though conceivably a Starship could be permanently parked in some orbit as a depot if there was a good enough reason. Orbital refilling is more like the in-flight refueling of a fighter jet than the establishment of a chain of gas stations.

The Starship concept reflects a different line of reasoning. Rather than compensate for a vehicle with a small Δv with a series of gas stations, build a vehicle with a huge Δv and, by refilling in LEO, enable it to fly enormous distances with no further support. Indeed, unlike conventional three stage Lunar lander schemes capable of transporting at most a couple of tonnes of cargo, the Starship is able to refuel in LEO, deliver hundreds of tonnes of cargo to the Moon, and fly all the way back to Earth in a single stage, reusable format. This is more like cargo flights to remote bases and islands that do not refuel after landing.

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This graph shows the cargo capacity of Starship to deliver or return cargo from the Moon, depending on where it’s refilled and how low the dry mass ends up being. It can fly similar quantities of cargo to Mars, but must be refueled there by a local propellant factory.

In summary, reasoning by analogy in space almost never works. Be wary of glib assertions that future cis-Lunar industry will employ orbital fuel depots filled by Lunar water and staffed by self-replicating robots. The natural place to fuel rockets is at their launch pad!

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     Introduction: On February 19th 2019 the world entered an era of impending Space-based military asset proliferation with U.S. Space Policy Directive-4 (SPD-4), which codifies legislative efforts in support of the establishment of the U.S. Space Force. There are currently approximately 81 U.S. military assets in Space, based on information that is publicly available. This study provides a preliminary quantitative assessment of potential U.S. military demand for in-Space water-based fuel in support of future U.S. military assets in Earth’s orbit. Our study concludes that if the U.S. military transitioned its future assets to utilization of water-based propulsion, there could be a military market demand for in-Space water-based fuel of 25 metric tons per year (baseline demand). In addition, and more importantly for the evolution of military demand for in-Space water fuel, we determined a parametric relation-ship between the supply of in-Space water on the Moon, Near-Earth Asteroids (NEAs) and the Asteroid Belt and the potentiality of increased U.S. military assets and increased propulsive capability per asset, thus leading to even further increased military demand for in-Space water in the future.

     Note: We make no assertions as to the merit of the establishment of a U.S. Space Force and the militarization of Space. However, increased market demand from civil, commercial and military use cases will all contribute to increased economic viability of Space Resources Utilization (SRU) and investability of associated technological capabilities and SRU value-chain public and private enterprises. The purpose of this preliminary assessment is to provoke holistic thinking on how U.S. military needs for strategic deterrence could drive significant demand for an in-Space water-based fuel supply-chain.

     Results and conclusions: Our estimate of U.S. military baseline demand for in-Space water, assuming today’s number of U.S. military assets in Earth’s orbit, is 25 metric tons, which using ULA’s $3,000/kg price point at LEO, deduces a total market opportunity of approximately $75M USD per year. This market opportunity is not significant and probably does not warrant in itself significant investment in the realization of in-Space water supply-chain for U.S. military needs. However, given the U.S. military’s desire to grow its presence in Earth’s orbit, and the proliferation of low-cost, highly maneuverable, and shorter lifetime military assets in LEO (with higher de-orbiting fuel needs), we foresee the $75M market opportunity growing by at least an order of magnitude in the foreseeable future. In addition, on the supply side, the estimated Lunar, NEA and asteroid belt deposits posit a new paradigm for how the U.S. military could deploy, operate, utilize, refuel, repair and retire or repurpose its assets in Earth’s orbit (and beyond). The U.S.’s notional apportionment of Lunar water deposits alone could, for example, enable the deployment of 33,000(!) equivalent assets (by mass) in Earth’s orbit, with 100X the maneuverability (i.e. station keeping reserves) for 500 years(!). Even though it is not reasonable to assume all of the U.S.’s theoretical Lunar water apportionment would be used for military needs, even a fraction of it would drive significant demand for in-Space water fuel, with a market opportunity in the $Bs/year.

     As noted in the introduction, we make no assertions as to the merit of the establishment of a U.S. Space Force and the militarization of Space. The sole purpose of this preliminary assessment is to provoke holistic thinking on how military needs for strategic deterrence could drive significant demand for an in-Space water-based fuel supply-chain, that enhances the value-chain of SRU for civil and commercial needs, and thus ena-bles increased investability of SRU technologies and enterprises at a global scale.

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Water Supply and Demand

Supply

	Moon	Near Earth Asteroids	Asteroid Belt
Pros	Close	Potential low delta-V	Enormous supply
	Diverse resources & applications	Significant supply	Diverse resources
Cons	Relatively high delta-V	Highly variable	High delta-V
	Finite supply	Significant unknowns	Far away

Demand

• Demand estimates based on current U.S. military assets in orbit
• Future trends for satellite development are unknown
• Future assets may be smaller & decentralized, but advancements could dramatically change this

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Calculating Demand

Demand Assumptions and Methodology

Assumptions

Deriving water needs:

• All fuel is replaced with water (Isp~180)
• Fueling occurs in LEO for deployment
• Dry mass of satellites does not change
• CubeSats not considered

Future assets:

• Follow same orbital distribution
• Lifetime of 20 years
• 10X current maneuverability
• 10% share of all ISRU water

Methodology

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USG Assets Results

Based on ~130 military assets, it was estimated:

• Current demand is ~45 tons of water per year
• 333 kg of water per asset per year

A future water propelled U.S. military asset would require:

• 3,000 kg for deployment
• 130 kg for disposal
• 610 kg per year for station keeping

Supply chain demand:

• ~40% Fuel for LEO
• ~20% Fuel for MEO
• ~40% Fuel for GEO

Transportation to destination orbit needs 50% for LH2/LOX and 10% for ion/plasma

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Supply Chain Road Map

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Calculating Supply 1: The Moon

Supply Chain Road Map

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Lunar Assumptions and Methods

Assumptions

• 1.2 x 10¹² kg of water located at the Lunar Poles
• U.S. Military has 10% share of total
• About 2/3 of the dry mass of Lunar escape vehicles is payload
• LH2/LOX fuel for Lunar escape
• Ion/plasma propulsion used for final fuel delivery

Methods

• 10% multiplier for 10% share of total water
• ~50% multiplier for water loses due to Lunar escape
• ~90% multiplier for water loses due to destination orbit insertion

Leaving 5.2 x 10¹⁰ kg for military assets

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Lunar Results

• Accessible with conventional technology
• Relatively high magnitudes of water
• Important stepping stone to future supply chains
• Easier to disrupt
• More Competition

Independent Lunar Supply Chain
Source Water Mass	1.2 x 1012 kg
Delivered Water Mass	5.2 x 1010 kg
Theoretical Number of Deployable Assets	~3.4 Million Assets
Time Until Depletion	~510 years

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Calculating Supply 2: Near Earth Asteroids

Supply Chain Road Map

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Near Term NEA Assumptions and Methods

Assumptions

• Only 5-30m asteroids are minable
• C types make up 20% of NEAs
• Even distribution of asteroid size and type
• U.S. Military has a 10% share
• Solar baking method is scalable
• Ratio of asteroid to spacecraft is 60
• Min of 50% of water reserved for return
• 75% of the water is mined out and the remainder of the asteroid is ditched
• Fuel to despin asteroids is negligible
• LH2/LOX is used for outbound
• Derived LH2/LOX fuel for Earth return
• Ion/plasma propulsion is used for fuel delivery

Methods

• Use assumptions to find delta V max
• Use delta V in broken plane delta V function to estimate mass in range
• Adjust based on ratio in size range
• 20% multiplier for C types
• Density adjustment multiplier
• 10% multiplier for percent water
• 75% multiplier for water taken
• 50% multiplier for water saved
• Subtract water inbound for net
• Adjust using Reiman sums to account for greater recovery at all lower delta Vs
• 10% share of total water
• ~90% multiplier for water loses

Leaving 5.3 x 10⁷ kg for military assets

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Near Term NEA Model

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Near Term NEA Results

• Comparatively low water mass
• Extremely efficient for early ISRU cascade all within Lunar delta V
• Water return ratio here is 5 times the investment
• Provides value as a redundant supply chain
• Hard to disrupt

Independent Near Term NEA Supply Chain
Source Water Mass	1.6 x 109 kg
Delivered Water Mass	5.3 x 107 kg
Theoretical Number of Deployable Assets	~3400 Assets
Time Until Depletion	~170 years

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Long Term NEA Assumptions and Methods

Assumptions

• All NEAs accessible with minimal in-bound loss
• C types make up 40% of NEAs
• Even distribution of asteroid size and type across delta-V ranges
• All C type asteroids can be mined
• 100% of the water is mined and the remainder of the asteroid is ditched
• Fuel to despin asteroids is negligible
• U.S. Military has 10% share of water
• Derived LH2/LOX fuel for Earth return
• Ion/plasma propulsion used for final fuel delivery

Methods

• 20% multiplier for C types
• Density adjustment multiplier
• 10% multiplier for percent Water
• 75% multiplier for water taken
• ~16% multiplier for loses due to Earth return
• 10% multiplier for military share
• ~90% multiplier for water loses due to destination orbit insertion

Leaving 9.8 x 10¹² kg for military assets

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Long Term NEA Results

• Tremendous resources available in the NEA population
• Much of it is currently inaccessible
• Requires significant R&D but it’s worthwhile
• Hard to disrupt

Independent Long Term NEA Supply Chain
Source Water Mass	6.7 x 1014 kg
Delivered Water Mass	9.8 x 1012 kg
Theoretical Number of Deployable Assets	1 Million Assets*
Time Until Depletion	12,800 years*

*Capped at one million with demand becoming linear thereafter, as # of assets was far too great to be realistic

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Calculating Supply 3: The Asteroid Belt

Supply Chain Road Map

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Asteroid Belt Assumptions and Methods

Assumptions

• Asteroid belt mass is 3x10²¹ kg
• C type make up 40% of the population
• C types are 15% water by mass
• Asteroid intercept requires average delta-V of 8000 km/s
• Fuel losses due to asteroid escape are negligible
• Derived LH2/LOX fuel for Earth return
• About 2/3 of the dry mass of miner is payload
• U.S. Military has 10% share
• Ion/plasma propulsion is used for final fuel delivery

Methods

• 40% multiplier for percent C types
• 15% multiplier for percent water
• ~13% multiplier loses due to Earth return
• ~90% multiplier for water loses due to destination orbit insertion
• 10% multiplier for the 10% share of total water

Leaving 2.1 x 10¹⁸ kg for military assets

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Asteroid Belt Results

• Provides inexhaustible resources
• Much of this water is centralized to Ceres
• Could support planned Moon like manufacturing
• Ceres will likely be an important strategic & economic hot spot for humanity

Independent Asteroid Belt Supply Chain
Source Water Mass	1.8 x 1020 kg
Delivered Water Mass	2.1 x 1018 kg
Theoretical Number of Deployable Assets	1 Million Assets*
Time Until Depletion	2.8 Billion Years**

*Capped at one million with demand becoming linear thereafter, as # of assets was far too great to be realistic

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All Independent Supply Chain Outcomes

Given a 10% share of the total resources and estimated requirements for resource return, based on the average military asset with 10X maneuverability and 2% exponential growth, each supply chain could independently support:

	Lunar Supply Chain	Near Term NEA Supply Chain	Long Term NEA Supply Chain	Asteroid Belt Supply Chain
Source Water Mass	1.2 x 1012 kg	1.6 x 109 kg	6.7 x 1014 kg	1.8 x 1020 kg
Delivered Water Mass	5.2 x 1010 kg	5.3 x 107 kg	9.8 x 1012 kg	2.1 x 1018 kg
Theoretical Number of Deployable Assets	3.4 Million	3,400	1 Million*	1 Million*
Time until Depleition	510 Years	170 Years	12,800 Years*	Unlimited*

*Capped at one million w/linear demand after for asset replacement, as the exponential # of assets was unrealistic

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Controversy and confusion continue to swirl around the issue of a cislunar space base, no matter where it is proposed to be built, what it is named, or what it is supposed to be for. Assuming that one of the primary purposes of such a base should be to supply lunar-derived propellant to vehicles in or departing from cislunar space, there are several immediate problems. First, there is a massive conflict in the potential schedule, since human missions to Mars that could be fueled from the Moon could take place about the time we find out if there is any actually accessible lunar ice available. We know that there are massive ice deposits on Mars, while most of the Moon is more than bone dry and the critical polar water deposit surface characteristics are still hidden from us.

Another problem area is the related but unknown costs of building the lunar mining base, the cislunar base, and the transport system to move the propellant from the lunar surface to the cislunar base, and how much that would add to the cost of lunar propellant. It is very hard to estimate the cost of developing the mining base. Some claim that the cost would be so high as to make the lunar propellant more expensive than propellant brought from the Earth, even though the cost of moving propellant from the Moon is usually quoted at about 15–20 times less that bringing it from the Earth’s surface.

Finally, there is the NASA plan, dormant for a while and now seemingly moving ahead, to create a way to use its obsolete and expendable SLS rocket to support what it still refers to as a lunar “gateway.” This is the project I have referred to in the past as the “Gateway without a Gate.” Very recently, the name and orientation of this project has changed again, to the Lunar Orbital Platform-Gateway, or LOP-G. It may seem as if a station in lunar orbit is more closely associated with actual lunar development, but placing a station in any actual lunar orbit, since it is then in an orbital plane, restricts the number of lunar surface locations that are easily accessible, and some of these so-called “lunar” orbits spend a lot of time far from the Moon.

Since most plans and sources do not mention refueling and logistics facilities as integral and initial parts of this project, such a “gateway” would neither be able to support a lunar base nor support dispensing lunar propellant produced by such a base, and are unlikely to be added later. A NASA request for information (RFI) in late 2017 covered many aspects of gateway science, but not a single transport issue. Most people would agree that an actual gateway provides a pathway to some physical location. So where is the physical path for this gateway? Where does it lead from and what does it lead to? In spite of the acronym, the planners have lopped off the critical gateway (transport-related) features—if they were ever there to begin with.

Some people fear that this project would eat up any and all funds for an actual lunar mining base and thus are now insisting that any lunar base be supported only from low Earth orbit, which is less efficient. If the platform is supported by the SLS, it would eat up even more funds, leaving little for any other human space projects, and could delay the establishment of an actual lunar base by a decade. With time, bureaucrats could decide that it is too dangerous to have propellant depots docked at the “gateway.” When the powers that be finally realize that crews stationed beyond LEO actually do need a significant amount of shielding mass to protect them from cosmic radiation, the whole project could be thrown into redesign disarray and last another decade before even being launched.

Obviously, a vehicle about to depart for Mars will not want to land on the Moon to get its propellant. Even getting into lunar orbit from a cislunar location would waste fuel. Thus, most experts believe that a location like Earth-Moon L1 or L2 is the best place to accumulate a large store of propellant, since it is always in the same position relative to the Moon and Earth, and thus is not subject to orbital plane limitations. Since vehicles ready for either departures to Mars or to cyclers going to Mars would need to be positioned at locations other than L1 (but at about the same distance from Earth), the main propellant depot would probably be at L1, with temporary depots positioned at other locations during Mars transit windows every 26 months.

So why is having a lunar-derived propellant supply in a near-lunar location so favorable? It’s the propellant cost, along with some other good reasons. If you want to go to any location outside the Earth-Moon system, whether it is Mars or an asteroid, a departure location high above Earth is best since that allows a very efficient Oberth maneuver, which uses a departure burn at L1 and another during a close pass of the Earth. This saves more than half of the departure propellant compared to departure from LEO, and for Mars missions, this means the Mars transit propellant weighs less than the mission’s dry mass. (If you are not using any lunar propellant, the advantage of the high departure point is much less.) From L1, lunar propellant can be delivered to LEO for only about 0.85 kilometers per second of velocity change, and even directly from the lunar surface for about 2.74 kilometers per second, making delivery much cheaper there than Earth propellant, which needs about 9.5 kilometers per second for delivery.

Having a base in a location like L1 makes initial support of a new lunar base much easier, since it is possible to reach L1 from anywhere on the Moon’s surface (or the reverse) in about 12 hours without worrying about the orbital planes. It also breaks the trip from LEO to the lunar surface into two smaller steps in terms of velocity change, thus decreasing the dry mass and fuel mass fraction of each vehicle, and allowing each one to carry relatively more cargo or propellant. This also improves the safety factor, since smaller rocket engines can be used, and they do not fire as long.

The big factor, however, is still the vast cost difference in moving the fuel to L1. The difference is primarily caused by the fact that you need a huge amount of propellant to move the Mars transit propellant from Earth, but only a small amount to move it from the Moon. Let us assume that we have vehicles ready for a Mars mission, either three large 85–100 ton dry mass vehicles, similar in size to the SpaceX BFS stage, able to carry a small crew plus a lot of cargo, or a set of ten smaller 30-ton-range dry mass vehicles, some for crew and cargo and some just for cargo. Assume that both fleets have about 4,000 tons of dry mass and need about 2,100 tons of propellant at L1 to depart via an Oberth maneuver. To show the huge numeric difference in cost between Earth-based and lunar based propellant at L1, we do need to do some simple calculations.

Note that in these calculations, I distinguish between the mass of the rocket propellant needed to move the Mars transit propellant to L1, and the mass of the Mars transit propellant (the payload) itself. To avoid confusion, I will refer to the transit propellant as the transit propellant payload, the surface to LEO propellant (a) carrying the transit payload to LEO, and the surface to LEO propellant (b) carrying up the LEO to L1 propellant needed in LEO to move the transit payload from LEO to L1. On the opposite side of the scale, I will refer to fuel for the cislunar tanker which carries the lunar-derived transit propellant payload as the lunar to L1 propellant. The multiple kinds of propellant uses may be confusing, but these distinctions are crucial to understanding the huge cost difference. All named transport propellant loads include the return to base propellant, as all vehicles are reusable.

Let’s assume you have moved your Mars expedition fleet dry mass (less the transit and bootstrapping propellant) to either a position at L1 or into a high orbit that is at a similar distance from Earth as L1 is, ready for its Oberth-style Mars departure maneuver. The transit propellant payload will need to include all of the propellant (a minimum of 400 tons) needed for bootstrapping the initial landings on Mars before the surface propellant plant there can be set up. During transit, this propellant can be kept in vehicles with cryo-coolers or in one or more propellant depots equipped with cryo-coolers. All in-space propellant discussed here is cryogenic liquid oxygen (LOX)-liquid hydrogen with an assumed specific impulse (Isp) of 460 seconds. All loads of the transit and bootstrap propellant to L1 are 150 tons, and thus a single expedition needs 14 such loads.

What is the amount of propellant mass to move the needed approximately 2,100 tons of transit propellant to the fleet at L1 or a similarly high orbit? Propellant created on Earth for use by this fleet would be moved to L1 in two steps, supported by multiple launches from the surface. If the BFR tanker version is used, 150 tons at a time can be delivered by one BFR tanker to a LEO logistics base. Reaching LEO takes about 9.5 kilometers per second of velocity change. There it is transferred to a small reusable tanker with sunshade and cryocoolers, which will take it from LEO to L1. This takes another 3.77 kilometers per second, for a total delta-V of about 13.27 kilometers per second. To provide propellant for the small tanker, almost two more BFR tanker loads of LEO to L1 propellant need to be delivered to the LEO base per 150-ton load of transit propellant payload. All of the BFR tankers would reenter and land back on Earth. The small tanker would offload the cryogenic propellant to shaded and refrigerated depots at the L1 base, and then, minus its payload, would drop back toward Earth, where it would use a single pass aero-capture maneuver to get into LEO again for a tiny amount of fuel.

For transport of Earth-based propellant, I will use current numbers for the SpaceX BFR (tanker version), as the launcher, which uses LOX-methane propellant. At 4,400 tons liftoff mass, we can subtract the dry mass to get the total propellant mass. The payload (wet or dry) is 150 tons, the upper stage is 85 tons, and the first stage is probably at about 125 tons, giving a total dry mass of 360 tons and thus a nominal surface to LEO propellant (a) mass of 4,040 tons, with almost 1,000 tons of propellant in the second stage. The propellant to payload ratio for this description of a BFR would be 4,040/150 or 26.93: 27 tons of propellant is needed for every ton delivered to an LEO base. The payload mass ratio is an impressive 0.0341. The structural mass ratio for the second stage is also an impressively low 6.9 percent (85 tons/1,225 tons). These numbers will probably change some as the extremely efficient BFR designs are refined further by SpaceX.

We will assume that, at LEO, the 150-ton transit propellant payload is then transferred to the smaller tanker of the same cargo capacity but with smaller engines which use only LOX-hydrogen. Note that the launches from the Earth’s surface need to supply the small cislunar tanker with both the LOX-hydrogen LEO to L1 propellant plus the LOX-Hydrogen transit propellant payload itself. The LEO to L1 tanker is already in orbit. Assuming that the smaller tanker is about 25 tons and has its own cryo-coolers and sunshade, its structure and payload would weigh 175 tons. The delta-V from LEO to L1 is 3.77 kilometers per second, and about 1.0 kilometers per second for the return trip where aerocapture and a small orbit trim is all that is needed. This means the small tanker needs to carry only 7 tons of descent propellant with it to L1, with a margin. This then means the mass that reaches L 1 must be 182 tons, so the ascent (LEO to L1) propellant mass, also with a margin, is 250 tons. The total mass departing LEO for L1 is now 432 tons. Note that the structural dry mass for this small tanker at departure from LEO is 5.8 percent.

However, the 257 tons of LEO to L1 propellant needed to move the single load of transit propellant payload propellant to L1 and get the empty tanker back to LEO also needs to be moved up to LEO first, and requires the use of 6,921 tons of LOX-methane surface to LEO propellant (b) on more than one ride to LEO via the BFR. It costs about $200,000 to launch a Falcon 9 with about 500 tons of fuel (LOX and RP1) on board, thus that fuel combination costs about $400 dollars per ton. LOX-methane might cost about the same. So, in order to get each 150-ton batch of Mars departure propellant to L1 from Earth, it takes the following components:

Propellant masses for delivery of 1 & 14 loads of 150 tons of LOX-hydrogen to L1

Payload	payl mass	prop mass	vehicle	from	to	propel. cost
Mars Transit prop (payload)	150	4,040	1 BFR tanker	Earth	LEO	$1,616,000
LEO to L1 tanker propellant	257	6,921	2 BFR tankers	Earth	LEO	$2,768,400
Mars transit prop (payload)	150	257	1 cisln tanker	LEO	L1	102,800
Total propellant mass	-	11,218	both	Earth	L1	$4,487,200
(Multiply by 14 BFR and tanker loads to L1)
Total propellant mass to L1	2100	157,052	both	Earth	L1	$62,800,000

This is a 74.79-to-1 ratio of propellant to payload delivered to L1 (11,218/150). The transport propellant for one load of Earth propellant thus costs $4.49 million at $400 per ton so that just one ton of Earth propellant delivered to L1 would be worth about $30,000, based only on delivery propellant costs. The total propellant mass needed to deliver 2,100 tons, or 14 BFR loads, of transit propellant payload to L1 is 157,052 tons and costs $63 million, making the propellant cost a major part of a single Mars mission cost when conducted with reusable vehicles. If there were 100 civilian passengers on the trip, the fuel would cost each person $628,000.

Now let’s see how much lunar to L1 propellant is needed to get the transit propellant payload propellant from the Moon to L1. In this case, we are using a lunar to L1 tanker with LOX-hydrogen fuel. It takes 2.6 kilometers per second to go from the lunar surface to L1 and the same to return, but fuel use on return is minimal since there is no payload. I assume that this tanker is 30 tons, since it needs landing legs and it leaves the lunar surface with the same 150 tons of cargo, for a dry and payload mass of 180 tons. The empty 30-ton tanker will need just 25 tons of fuel (with a margin) to return to and land at its lunar base without its payload, so the structural mass, payload, and return fuel mass is 205 tons. The vehicle leaving the lunar base with its payload of fuel needs 170 tons of propellant, with a “wet” mass of 375 tons, so for the whole round trip (where the payload is left at L1), the fuel needed is just 195 tons of propellant.

Reusable Lunar surface to L1 & return tanker masses

Trip Component	1 load mass	mass fraction	mass for 14 loads
Dry structural mass	0 tons	0.08	420
Transit propellant payload	150 tons	0.40	210
subtotal	180	0.48	2520
return to base propellant	25 tons	0.07	350
subtotal	205		2870
surface to L1 propellant	170 tons	0.45	2380
total mass at liftoff	375	1.00	250
total round trip propellant	195 tons	0.52	2730

Notice that this means for a well-designed light tanker ferry, for every 1.0 tons of payload propellant it only takes 1.3 tons of fuel propellant to move the payload from the surface base to L1. At lunar takeoff, this tanker would mass 375 tons, but since it is taking off in lunar gravity, it would effectively weigh only about 62 tons. The composite propellant tanks, similar to what will be used for the BFR, would probably mass only 20 tons, leaving 2 tons for the engines, 2 tons for the landing legs, and 6 tons for the rest of the structure, including the cryo-coolers, sunshade and power system. This provides a structural mass fraction of 8.0 percent.

At this point, we do not know the actual cost of the lunar propellants as delivered to the tanker at the lunar base, including the vehicle development costs, but finally we can compare the relative mass of the delivery propellants. Assuming that we have not underestimated the tanker mass, the delivery propellant mass from Earth to LEO to L1 is 57.5 times more (74.8/1.3) than the delivery propellant mass from the Moon to L1 for the same 1 ton of propellant as payload. (Note that this mass difference is about three to four times larger than the typical cost difference of 15–20.) Cost of lunar fuel would include fuel production costs, transport propellant costs, and vehicle operation costs. The wear and tear on lunar-to-L1 vehicles is much less, since no vehicles need to take off from the Earth or re-enter the Earth’s atmosphere. Even the cislunar tanker must undergo an aerocapture to return to LEO from L1, while the lunar-to-L1 tanker would encounter only the stresses of thrust and landings. In addition, propellant at L1 is at least twice as valuable as propellant in LEO, due to the gravitational potential and additional velocity from an Oberth maneuver performed starting at L1.

So if the lunar propellant cost proportionally as much as the delivered Earth propellant before the lunar propellant is delivered to L1, it might cost about $23,012 per ton (57.5 times $400), or ($29,915/1.3) on the lunar surface. If future lunar entrepreneurs can beat that price, and if lunar water ice does exist in minable quantities, the lunar fuel production enterprise seems assured. Sales of large amounts of propellant for Mars expeditions would assure a robust human presence on the Moon. However, to assure that the logistics capability of any cislunar or L1 base will exist, the propellant depot and cargo handling capabilities must be a part of the design of the base from the beginning. Plans for the NASA cislunar base must accept that large propellant depots can be attached to it.

We are getting significant indications of international interest in a lunar base. At the same time, NASA says that there may not be enough room on the LOP-G station to accommodate international science participation. Why does the cislunar base have to be so small if there is support for it? The solution to this problem is a package deal agreement where there is a combined cislunar, lunar, and fully reusable transport development effort taking place simultaneously, with the international partners providing some of the transport and lunar surface infrastructure, with heavy reliance on commercial launches. Development would take place so that the first human landings at a lunar base site would take place within a year of the cislunar base completion. Without such a “package deal,” the NASA cislunar base would probably become the “gateway with no gate” or a “space station junior” copy with very limited utility, as many of us have feared.

Next up: in-space refueling. The first and most obvious customer is NASA; an orbital fuel depot would allow them to launch satellites on smaller LVs or launch larger satellites, allowing a choice between savings on the LV and increased capabilities on the spacecraft. That could mean buying an Atlas 401, Zenit, Soyuz, H-IIA or Falcon 9 instead of an Atlas 551, Ariane 5, H-IIB or Delta 4(5,4). Perhaps less obvious, Russia would see a significant benefit from a LEO depot in the plane of the Baikonur launch site. Vehicles would refuel in order to plane-change to an equatorial orbit for GEO deployment. Further into the future a fuel depot would be essential for the smooth operation of tugs and satellite tenders, serving as a buffer between fuel launches and fuel used in missions.

I think the current leader is Boeing with their in-development ACES vehicle using integrated fluids management. However, they are focused on LOX/LH2 propellant; few customers today have cryogenic upper stages. Hypergolic fuels require a different set of technologies and would most likely require shipping expendable supplies of a pressurant, either nitrogen or helium, but they have a larger potential market right now as hypergolics are typically used for satellite stationkeeping and orbit changes. The third fuel category is inert gases (argon, xenon) for ion engines; these can be stored as compressed gases or cryogenic liquids.

I think a near-term possibility is simply to ship water. It is dense, relatively inert and can be used as a life support consumable or as a propellant after electrolysis. It has a high surface tension and can be wicked out of a bulk tank in microgravity without pressurants or membranes. There are cubesat-scale thrusters available today that separate water over time, accumulating a charge of gaseous O2 and H2 using small amounts of power, then ignite that fuel in a high-efficiency engine. If future satellites were to adopt this technology for RCS and stationkeeping then they could nearly double their Isp while eliminating toxic fuels and cutting down to a single storage tank. Beyond the near-term possibilities, a water depot operator would be able to buy water from any LEO cargo provider as well as any asteroid mining company, relying on the proven launch capabilities today while safely and cheaply allowing for a riskier but cheaper future supply.

     This entry covers cis-lunar space. The topic of lunar mining and fuel supply has a rich field of information available and I cannot claim to know all of it, but hopefully this will show how we can begin to harvest most of our propellant instead of shipping it from Earth.

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     First let's establish some basics:

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     I assume that we will develop zero boiloff cryogenic storage, reliable cryogenic fluid transfer and reliable cryogenic engine restart. Vehicles will be designed to last 20-30 years, but generally are planned to be replaced every 10 years. Obsolete systems will keep operating until they fail, providing some bonus production capacity.

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     The Moon's surface gravity is 1.62m/s², about 16.5% of Earth gravity.
     It takes about 1.87km/s of dV to land or take off from the equator. Only a little more is needed for the poles.
     Another 0.64km/s will take you from low Lunar orbit to EML1 (Earth-Moon Lagrange Point One).
     Low orbits around the Moon are not stable so you should not park anything important there.

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     Large amounts of water ice are available at the south pole and most likely at the north pole as well.
     Plenty of metal oxides are available (including iron, aluminum and titanium), but carbon is rare.
     In some locations the mantle interface material (KREEP) is accessible at the surface; this rock is rich in incompatible elements like phosphorus, potassium, rare earths and radioactives.

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      Low Earth orbit takes anywhere from 9.4 to 10 km/s of dV to reach.
     Each launch site has a specific, most-efficient inclination. Changing inclinations is very expensive.
     From any low-Earth orbit, EML1 is 3.77km/s away.
     From Kennedy Space Center LEO, geosynchronous orbit is 4.33km/s away. It's only 3.9km/s from an equatorial orbit.
     Earth escape is 3.22km/s away.
     EML1 is the balance point between Earth and the Moon, the place where the gravity of each body cancels out.
     From EML1, a craft with a heatshield can get to any low Earth orbit for 0.77km/s.
     Without a heatshield, 3.77km/s is required.
     A similar maneuver can use the Earth as a slingshot, departing from EML1 with a little nudge into any inclination and then burning at closest approach for best use of the Oberth effect.
     From EML1, geosynchronous orbit is only 1.38km/s away.

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     Putting all of that together:

• The best place for a fuel depot is at Earth-Moon Lagrange Point One (EML1). Fuel can be harvested at the harvester's pace, shipped to the depot at the tanker's pace and accumulated for later use. Fuel is shipped to LEO only as needed and into the correct inclination.
  
  
• The surface to EML1 tanker needs to use chemical engines to overcome the Moon's gravity. It must carry about 4.6km/s of fuel, though the second half of the trip is empty and requires much less fuel. My reference tanker is 5.5 tons, cryogenic with zero boiloff and can deliver 29 tons of propellant from the Lunar surface to EML1, then land at the harvester. Each trip burns just under 34 tons, so the harvester must produce about 2.2kg of propellant for each kg in EML1.
  
  
• The EML1 to LEO tanker can use either chemical or ion engines. The low thrust option requires a very different trajectory and a dV of about 7km/s, plus a lot of onboard power. This can be a competitive option but initially it would be simpler to use the same tanker for each leg. Because the vehicle needs to aerobrake, it requires a reusable heatshield. The same tanker design is used, but it is launched with a dual-use shroud of about 6.3 tons that is kept with the vehicle. This trip delivers 44 tons of propellant at a cost of 19 tons. So, for each kg in LEO the harvester has to produce 3.1 kg. That lines up fairly well with other sources suggesting 75 tons to be harvested for 25 tons in LEO.
  
  

     The harvester is an unknown quantity. There are several concepts (including one of mine), but I will use a more generic NASA number of 10kg propellant per kg of harvester per year, with 10% spares. That means if we want to harvest 100 tons of fuel per year we need to send 10 tons of equipment to start and 1 ton of spares for each year. This was for a Mars ISRU system with rover/excavator and integrated power systems. A Lunar system would be much simpler since it would only be melting and filtering water ice then electrolyzing it, so it is possible the equipment will produce far more propellant than this estimate.
     Cheap propellant in LEO is useful for missions to other planets or moons. It is also useful for supplying a LEO station with water and stationkeeping fuel. Satellite servicing tugs could base out of EML1 with a ready supply of fuel and easy access to all inclinations. If commercial satellites were modified to use a water electrolysis thruster system for stationkeeping then they could be launched empty to EML1 for less dV than a direct launch to GEO. A fill of fuel at EML1 and a 1.4km/s nudge would place the satellite in GEO with decades of stationkeeping thrust available. As a bonus, the upper stage of the launch system could be refueled and repurposed. Lastly, water is extremely useful for manned operations as radiation shielding, drinking water and a reserve source of oxygen.

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     An example of why the interorbital exchange would be useful is NASA's Mars Design Reference Mission (DRM). Each manned trip requires three flights (one crew and two cargo), each using around 190 tons of fuel. Just the fuel would require 9 Block 1 SLS launches or 6 of the 105-ton launches; being generous we're talking about at least $4.5 billion. This is 570 tons of propellant over 2.1 years, or about 267 tons per year at around $7.9 million per ton. A full campaign of three flights to Mars would cost $13.5 billion in fuel. With lunar ISRU we can save at least $4.7 billion as described below (potentially over $9 billion) and gain a sustainable supply of propellant in EML1/LEO with minimal Earth mass.
     An ISRU plant with this capacity would mass at least 83 tons and consume about 18 tons of spares per Mars opportunity. A modified version of my tanker could land perhaps 17 tons of cargo on the Moon for about 62 tons of fuel; this would require two Falcon Heavy flights (~$300 million) or an SLS block 1 flight with two additional Falcon 9 refueling flights (~$850 million). This initial 15-ton plant (with 1 year of spares and a 500-kg dextrous robot for remote repair operations) would provide the first ISRU fuel as a proof of concept. The cargo lander would be refueled and returned to LEO two months later with plenty of lunar samples. (3 tons of samples, 25 tons of fuel.) At this point the whole architecture can be validated and the final ISRU design can be settled. Until this step happens we have to use some pessimistic numbers (as below), but there is the possibility that the pilot plant will be many times more productive than expected as it will be running a much simpler refining process.
     A standard tanker with dual-use shroud (11.8t empty) would be launched with a Falcon 9. Another standard tanker with 43.5 tons of fuel would be launched on a Falcon Heavy, transfer 27.8 tons of fuel to the shrouded tanker and then head to the Moon and land by the ISRU plant. The shrouded tanker would travel to EML1. At this point fuel can be transferred all the way to LEO; without a proper depot there are some inefficiencies, so only about 40 tons are delivered in each shrouded tanker (requiring two standard tanker flights to fuel up). Fortunately this is enough for the outbound trip of a cargo tug.
     The first shipment of 40 tons would take the rest of the year to produce and would be delivered to the cargo tug in LEO, allowing it to deliver a 22-ton payload to the Lunar surface. (Payload would be a second 15-ton ISRU plant, two years of spares for both plants and 1 ton of other cargo.) This could be a Falcon Heavy carrying two payloads or one of several competing options. After 30 days of surface operation the cargo tug returns to LEO (again, 3 tons of samples and 25 tons of fuel).
     The next 40 tons of fuel take a bit over five months to produce and allow the cargo tug to bring the next 22-ton package. Return fuel for the tug would take three weeks.
     At this point the infrastructure on the moon is producing 1.25 tons of fuel per day. Each cargo trip takes less and less time to refuel (3.4 months, then 2.6 months, then 2 months). A proper depot at EML1 and three more ISRU packages round out the system. The depot is 300-ton capacity, about 20 tons of hardware and delivered to EML1 by a Falcon Heavy. In total the buildup phase takes two years and two months, almost exactly one synodic period. 90 tons of ISRU hardware is on the surface with an annual production of 900 tons and spares requirements of 9 tons. This system can deliver 289 tons of fuel to LEO every year, or 617 tons per Mars synodic period. The target of 570 tons is met with an extra 47 tons of fuel for delivering spares.
     A total of seven Falcon Heavy flights are required, plus one Falcon 9. Let's call this $1.1 billion in launch costs over 26 months. Let's also assume that spares are supplied as 11-ton packages by Falcon 9, two per synodic period. A third assumption is that we will need to replace about 20% of the system every period (at a conservative 10-year lifespan for components), so we budget a Falcon Heavy flight for $150 million. That's an operational cost of about $250 million per period. In the first synodic period the fuel in LEO will cost $2.2 million per ton in launch costs. Each additional period will cost only $405,200 per ton in launch costs.

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     The cost of all that hardware is difficult to estimate. Let's look at two very different systems and get a ballpark figure. The BA-330 expandable space station module from Bigelow is estimated to mass about 22 tons and cost about $250 million, or about $12 million per ton. By contrast, the Iridium NEXT constellation of 72 satellites is expected to mass 57.6 tons in total and cost $2.4 billion (not counting launch costs), or about $42 million per ton. The comsats include costs like bandwidth leases and operations, but let's use that figure anyway as a first approximation. The startup phase requires 131 tons of hardware, or about $5.5 billion. Extended operation will replace about 10% of that hardware per year at a cost of about $550 million. This ignores the significant commonalities between systems; all the propulsion units are identical, all tankers identical other than the cargo adapter and all six 15-ton ISRU plants are identical.
     This gives us the second part of fuel costs: capital expense. Each ton of fuel in LEO in the first period costs $9.8 million, with additional periods costing $890,000. These numbers assume we do not spread the capital costs over five periods, but front-load the entire bill into the first launch season. The first Mars caravan would cost $6 billion instead of $4.5 billion in fuel, but the next one would pay only $1.4 billion. A campaign of three flights to Mars would save about $4.7 billion in fuel costs. As a budget line item, the cis-lunar fuel network would cost NASA or a private operator $375 million per year once established.
     What if, instead, a cost-effective hardware program is used that is closer to the BA-330 in terms of cost per kg? Now our 131 tons of hardware only costs about $1.6 billion with annual replacement costs of $160 million (or per-period costs of $342 million). Under this assumption, the startup phase would cost $2.7 billion with operational costs of $592 million per period ($278 million per year). Propellant would cost $5.34 million per ton in the first period and $960,000 per ton thereafter. NASA would save about nine billion dollars on three Mars flights, and would even save about $1.2 billion on the first flight.

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The first Earth-Moon Lagrange point, or EML-1, offers a number of key advantages that make it an ideal destination for activities in cislunar space. Over the near-term, however, its utility is constrained by a lack of physical infrastructure. This can change if our approach to space moves away from destinations and towards a strategy of enabling capabilities.

Talk abounds of going beyond Earth orbit, although once beyond low Earth orbit (LEO) what happens next becomes a little fuzzy in most discussions. This need not be the case, as capabilities can be built from the very first test of a trans-LEO vehicle. Some sample test runs:

1. Out to GEO: Given launch locations of likely US crewed vehicles, the mission could involve a plane change to geostationary (perhaps through a bi-elliptical transfer for extra-credit), close-approach to a “zombiesat”, and perhaps even retrieval of some old hardware for forensic analysis.
2. Free-return trajectory: A loop around the Moon to give the heat shield a workout on the return. Perhaps some maneuvering out around the Moon. How close can they shave the rear-end of the Moon at perilune?
3. EML-1 visit: The main purpose would be to establish a halo orbit. Once there, it would make sense to drop off a package of instruments that could serve a number of purposes.

Instrumental to the use of EML-1 is the concept of the halo orbit. It is a technique that allows an object to orbit an empty space, typically a Lagrange point, as with SOHO, WMAP, Genesis, the future JWST, and others. The best, though inapt, way to think about it is as a sort of gyroscopic effect, as with a bicycle wheel, where the hub is on the line connecting the centers of gravity of two objects (like the Earth and Moon), and the satellite is on the wheel. The action of the orbit helps to keep the object in place; some station keeping is required, but at a level orders of magnitude less than the ISS.

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The EML-1 neighborhood is good for looking not only outward, but also back in towards Earth. EML-1 is a naturally clutter-free environment, since space junk doesn’t normally have station-keeping capability, as well as a “high ground” that allows observation not only of the Earth but also everything that’s orbiting around the Earth out to GEO. Instruments could be placed there for the specific purpose of keeping an eye on the traffic in cis-GEO space. By the same token, it is also ideal for looking at the Moon, and serves as a natural gateway to the entirety of the Moon’s surface. Bigelow Aerospace has proposed using EML-1 as an aggregation point for modules to be emplaced on the Moon’s surface using Armadillo Aerospace rockets for the descent. As far back as 1986 a Lunar Spaceport was envisioned as a kind of motel/gas station/warehouse/restaurant/garage for space travelers.

The Moon is still a bit in the future, though, as there are still a lot of things to consider on the Earth side of EML-1. One of the key advantages of staging at EML-1 is, as Brad Blair notes, “its ability to fall into various inclinations without a major [delta-V] penalty, thus increasing the number of customers that could be reached by a small set of vehicles and systems elements.”. What this means is that the inclination you end up in in LEO is set by how you depart from EML-1.

If you picture the Earth-Moon system, and the line connecting their centers of gravity, EML-1 is about 85 percent of the way to the Moon. Trace an ellipse from the EML-1 point down around the Earth and back up to EML-1. Put the perigee close enough to Earth and you can even get some aerobraking. This ellipse can be assumed to be in the orbital plane of the Moon. Now, rotate that ellipse around the Earth-Moon axis. This means that every inclination of LEO orbit is accessible, from the 28.5° of Kennedy Space Center, to the 51.6° of ISS, to even polar orbits, although the latter suffer from more delta-V penalty by virtue of the Earth’s oblateness.

If all of the LEO inclinations are available from EML-1, then it is also true that any LEO inclination can get to EML-1. This means that the ISS can serve as a staging point for missions to EML-1 in the nearer term, and later stations in different inclinations can also reach EML-1 as they come online. There is no need to wait to get started. As soon as a crew vehicle comes online it can start staging from ISS to EML-1, first as test-runs, then as missions to emplace as well as service, upgrade, and refuel assets. By the time crewed facilities are emplaced there will already be regular traffic to the location.

One question often raised is “What would a crew do at EML-1?” There are a myriad of answers:

1) In terms of propellant, it is cheaper to go from EML-1 to GEO and back to EML-1 than it is merely to go from LEO to GEO. Over the long term, it makes sense to stage GEO operations from EML-1. What kind of operations? The easiest answer is salvage, given the hundreds of tonnes of scrap circulating in GEO. Crews could fall down to GEO, spend a few days retrieving defunct equipment like failed satellites, and then return to EML-1 to process it. Whatever could be reused in some way is unknown, but the real value is in the forensic analysis of how the satellite weathered in the GEO environment over a known period of time. That kind of information allows for better satellite design.

2) EML-1 is an on-ramp to what are known as the InterPlanetary Superhighways (IPS). These are a network of ridges and ripples in space created by the gravitational effects of the planets and Sun. A satellite pushed onto the IPS will travel very, very slowly along this network to its destination, where it can kick itself into a halo orbit around a Lagrange point and collect data. Locations of interest would include the Sun-Mars L-2 and Sun-Jupiter L-1, to observe the Asteroid Belt; the Sun-Venus equilaterals at L-4 and L-5 to provide communications relay when Mars is on the other side of the Sun from Earth; Sun-Saturn L-2 to look at the Kuiper Belt; Sun-Neptune L-2 to look at the Oort Cloud; Sun-Mars L-1 as a waypoint on the way to Mars and the Asteroid Belt; Sun-Earth L-1 to watch the Sun; Sun-Earth L-2 to watch the stars. The key is that all of these instruments would also be able to return via the IPS to EML-1 for regular maintenance and servicing. As more probes are added to the network, instead of thrown into the void, there will be an increasing stream of probes in need of work.

3) The time lag from EML-1 to the Moon and back is much less than that for Earth-Moon. As a result, it is a better location for safe teleoperation of robots on the lunar surface.

4) As lunar activities ramp up, there will be an increasing need for freight handling of goods destined for the Moon, as well as those from the Moon. Early lunar exports are likely to be low-value-added goods such as oxygen, water, raw regolith, and some metals, but as more capabilities are established the exports will start creeping up the value-added chain: foodstuffs from lunar greenhouses, crafts created by the locals from local materials, increasingly sophisticated entertainment like dance and music, and so forth.

5) EML-1 is an ideal location to aggregate mission components for a trip to an asteroid. Fuel can come from the Moon, while spacecraft come from Earth. Someone’s going to have to put all of that together.

6) A facility at EML-1 can serve as a communications node for lunar operations to overcome the line-of-sight issue. Additionally, with solar sails “pole-sitting” above the north and south lunar poles, communications with the far side can be established.

7) Port services. While probes returning on the IPS will end up in the neighborhood of EML-1, they will need to be picked up. Same thing with freeflyer platforms sent on low-energy trajectories around the Moon for production runs. A space tug would be a good tool to have, and someone has to fly it.

These are just a few ideas, which can easily be expanded. If a crew can fall down to GEO for a servicing run, it can also fall down to HEO, MEO, and LEO for servicing missions. If an asteroid mission can be assembled at EML-1, so too can a Mars mission. If crews are salvaging dead satellites and kick stages from GEO, they may be able to cobble together the parts for other missions, selling the result to whomever wants to buy a space probe.

      Earth Moon Lagrange 2 or EML2 is one of 5 locations where earth's gravity, moon's gravity and so called centrifugal force all cancel out. It lies beyond the far side of the moon at about 7/6 of a lunar distance from earth.

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     Infrastructure at any of these 5 locations could be kept in place with a small station keeping expense. Other high earth orbits would be destabilized by the influences of the earth, moon or sun.

     Of these 5 locations, EML2 is the closest to escape. How close?

Specific orbital energy is given by

v²/2 - GM/r

v: velocity with regards to the earth
G: gravitational constant
M: mass earth
r: distance from earth center

     Here's specific orbital energies for a few orbits:

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     For EML1 and EML2 I'm looking at resulting earth orbits for payloads nudged away from Luna's Hill Sphere.

     Most the energy is getting from earth's surface to Low Earth Orbit (LEO). Then another huge chunk is getting from LEO to escape.

     EML2 is right next door to escape (aka C3=0). If the goal line is Trans Mars Injection, EML2 is on the 9 yard line.

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     EML2's orbital energy is about -180,000 joules per kilogram. How much is that? Well, Kattie is standing next to the small generator which provides electricity for our business during power outages during the summer monsoons. It would take this 20 kilo-watt generator 9 seconds to crank out 180,000 joules.

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     An EML2 payload nudged away from Luna would rise to an 1.8 million km apogee. An ordinary earth orbit at 450,000 kilometers from earth's center would move about 0.94 km/s. But since EML2 is moving at the moon's angular velocity, it is traveling 1.19 km/s. Earth's Hill Sphere is about 1.5 million km in radius. So depending on timing, an EML2 nudge could send a payload out of earth's sphere of influence into a heliocentric orbit.

     Another possibility is the sun's influence could send a payload back towards the earth with a lower perigee:

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     All of these pellets were nudged from EML2. The sun's influence has wrested most of these from earth's influence. But check out pellet number 3 (orange). The sun's influence has dropped this pellet to a perigee deep in earth's gravity well. For a .1 km/s nudge from EML2 we can get a deep perigee that can give a very healthy Oberth benefit. However, such a route takes about 100 days.

Farquhar Route

     Using an lunar gravity assist along with an Oberth enhanced burn deep in the moon's gravity well, EML2 is 9 days and 3.5 km/s from Low Earth Orbit (LEO):

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     This route was found by Robert Farquhar.

Bi-Elliptic Transfers

     If radii of two different orbits differ by a factor of 11.94 or more, a bi-elliptic transfer takes less delta V than Hohmann. EML2 radius / LEO radius is about 67, so LEO to EML2 could definitely be a beneficiary of bi-elliptic.

     From LEO, a 3.1 km/s burn gets us to a hair under a escape. A multitude of elliptical orbits fall under this umbrella!

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     As you can see, it takes almost as much to get as high as EML1 as it does to reach a 1.8 million km apogee. I chose 1.8 million as an apogee since a 450,000 x 1,800,000 km ellipse at perigee has the same altitude and speed as EML2. At perigee a payload can slide right into EML2 with little or no parking burn.

     What's needed is an apogee burn to raise perigee to 450,000 km. A 6738x1,800,000 km ellipses moves very slow at apogee, a mere 0.04 km/s. A 450,000 x 1,800,000 km ellipse doesn't move much faster at apogee, about 0.3 km/s So a .26 km/s apogee burn suffices to raise perigee.

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     So the total budget is 0.26 + 3.1555 km/s. This 3.42 km/s delta V budget is better than a Hohmann but about the same as Farquhar's 9 day route.

     But recall apogee is beyond earth's Hill Sphere. With good timing, the sun can provide the apogee delta v.

Hop's Route from LEO to EML2

     Here's a route I found with my shotgun orbital sim:

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LEO burn is about 3.11 km/s. Payload passes near the moon on the way out, boosting apogee and rotating line of apsides. The sun boosts apogee as well as perigee. Coming back the pellets slide right into EML2 (the circular path alongside the Moon's orbit).

This LEO to EML2 route took 74 days and 3.11 km/s.

EML2 and Reusable Earth Departure Stages.

     Using the Farquhar route, it takes about 0.4 km/s to drop from EML2 to a perigee moving just under escape velocity. At this perigee 0.5 km/s will give Trans Mars Insertion (TMI). After the departure stage separates from the payload it's pushing, it can do a 0.5 km/s braking burn to drop to an ellipse with a near moon apogee. Once at the moon, another 0.4 km/s takes the EDS back to EML2.

     For massive craft moving from between earth's neighborhood and other heliocentric orbits, it makes little sense to climb down to Low Earth Orbit (LEO) and back each trip. It saves time and and delta V to park at EML2 on arrival. If EML2 becomes a stop for interplanetary space craft, a reusable EDS is a good way to depart the earth/moon neighborhood.

     I talk about this in more detail at Reusable Earth Departure Stages.

EML2 and Fast Transits

     Here's a pic of a Non Hohmann Mars transfer:

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     Mars and earth orbits are approximated as circular orbits. A Hohmann orbit will have a 1 A.U. perihelion and a 1.52 A.U. aphelion. The transfer orbit above has perihelion 0.7 A.U. and aphelion 1.53 A.U. Semi-major axis of this orbit is (0.7 + 1.53)/2 A.U. or 1.115 A.U. Orbital period is 1.115^3/2 years which is about 1.18 years.

     The trip to Mars isn't the entire orbital period though, just the turquoise area swept out from departure to destination. The turquoise area is 31.5% of the ellipse's area. 31.5% of 1.18 years is about 135 days or about 4.4 months.

     Departure and arrival Vinf are indicated by the red arrows. These are the vector differences between the transfer orbit's velocity vector and the planet's velocity vector at flyby. A change in direction accounts for most of the Vinf. I'm assuming Mars and Earth are in circular orbits with a zero flight path angle. Therefore the direction difference between vectors can be described with the flight path angle of the transfer orbit's velocity vector.

     In this case the earth departure Vinf is 11.3 km/s. That's a big Vinf! But if falling from EML2, only a 4.9 km/s perigee burn is needed. This is doable.

     At Mars the Vinf is 5.14 km/s. But a periaerion burn of 2.57 km/s brakes the orbit into an (3697x2345 km ellipse. This orbit could be circularized via periaerion drag passes through the upper atmosphere. Since this ellipse has a period less than a day, orbit could be circularized in a few weeks.

     An upper stage can have a 8 km/s delta V budget. Recall it takes about 0.4 km/s to fall from EML2. Therefore let's try to find a route that takes about 7.6 km/s from perigee burn to periaerion burn.

     Trial and error with my Non Hohmann Transfers spreadsheet gives:

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     With chemical rockets departing from EML2, I believe 4 month trips to Mars are doable.

EML2 Proximity to Possible Propellent Sources.

     In terms of delta V, time and distance EML2 is quite close to several possible propellent sources.

     There are thought to be frozen volatiles in the lunar cold traps. Some craters at the lunar poles have floors in permanent shadow. Temperatures can go as low as 30 K. Volatiles that find their way to the cold traps would freeze out and remain. There may be rich deposits of H₂0, CO₂, CH₄, NH₃ and other compounds of hydrogen, carbon, oxygen and nitrogen. These would be valuable for life support as well as propellent.

     The moon's surface is about 2.5 km/s from EML2.

     Also there are proposals to retrieve asteroids and park them in lunar Deep Retrograde Orbits (DROs). DROs are stable lunar orbits that can remain for centuries without station keeping. Planetary Resources would like to retrieve water rich carbonaceous asteroids. Carbonaceous asteroids can contain up to 20% water by mass in the form of hydrated clays. They can also contain compounds of carbon and oxygen.

     LDROs would be about 0.4 km/s from EML2.

Summary

     EML2 would make a great transportation hub. Not only for travel to destinations throughout the solar system but also within our own earth moon neighborhood.

     This entry covers near-term resources on and near Mars and how they might be transported.
     I assume the propellant network described in part 1 has been built, or at least that lunar propellant is available at Earth-Moon Lagrange Point One (EML1).
     The short version of the below is that using hardware similar (or identical) to the part 1 Lunar infrastructure, cargo to and from Mars becomes relatively cheap.
     A set of three NASA-reference mars missions could have their cargo requirements filled for a total of $11.4 billion (including fuel).
     Nitrogen and argon from Mars could be as cheap as $400 per kg at EML1.

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     Again, let's start with the basics.

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     Mars has a surface gravity of 3.711m/s² or 37.6% of Earth gravity, and a surface rotation of about 0.24km/s.
     Mars is blessed with a thin atmosphere, mostly carbon dioxide but with useful amounts of nitrogen and argon.
     Martian soils are rich in perchlorates and relatively rich in water.
     The north and south poles have permanent ice caps, mostly water ice under a layer of frozen CO₂.
     CO₂ slabs violently sublimate in the spring, making the edges of the ice caps fairly dangerous in this season. Otherwise the dust storms are not a serious threat.

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     Takeoff and landing requires 4.1km/s; drag can save as much as 2.4km/s of this on landing.
     Mars has two moons, Phobos and Deimos. Both are believed to be captured asteroids, probably C-type.
     From EML1, Mars transfer costs 0.74km/s. Mars capture costs 0.9km/s and the move to low Mars orbit costs 1.4km/s. Trip total is just over 3km/s.
     From low Mars orbit, reaching Phobos costs 1.4km/s and reaching Deimos costs 1.9km/s.
     Launch opportunities to or from Earth (using a Hohmann transfer orbit) occur every 2.135 years, or about 26 months.

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     Taking all of this together we want a surface location near the equator and a depot in low orbit. If we find significant amounts of water at either Phobos or Deimos then we want a mining outpost there. It would be nice if we could set up a Phobos-anchored tether for orbit changes.

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     Getting cargo to Mars is slow and expensive. The post in part 1 used NASA's Mars Design Reference Mission (DRM) as a baseline to show how in-situ resource utilization (ISRU) could save money on fuel. Let's look at other ways to save.
     For starters, let's use the same cargo tug as described in part 1. This is a 5.5 ton dry vehicle that would deliver 50 tons of payload to EML1 using 77 tons of Lunar fuel and a reusable 15-ton (fueled) drop tank. The tug and payload would refuel with 63 tons of lunar fuel and continue on to Mars with fully propulsive capture to low orbit. These tugs can be used as fuel tankers between Phobos and LMO and can also be used to return cargo to Earth (up to 52 tons to EML1 all propulsive).
     Once cargo is in low Mars orbit, a different vehicle is needed to land it. The cargo tugs are capable, but multiple atmospheric landings in Mars gravity are out of reach. Instead, let's use the Access to Mars SSTO ferry (Strickland, Gopalaswami). This is a 30-ton dry mass, reusable SSTO cargo ferry for moving payloads to and from the Martian surface. The vehicle is 14 meters in diameter, which poses a problem; we would need a larger-diameter booster, a way to assemble the vehicle in orbit or a way to launch such a wide payload on an existing rocket. The design could probably be scaled down, but extrapolating the performance of a smaller version is not trivial. I will assume the problem is solved without specifying how. If reliable fuel is available from Phobos then the lander might be redesigned to refuel both in LMO and on the surface, allowing it to carry more payload each way.
     The cargo lander is intended to fuel up, launch to orbit (with 5 tons of payload or excess fuel), collect a 25-ton payload and land. Repeat as needed. Each 50-ton tug payload would require two lander trips. NASA's reference mission involves 80 tons of cargo to the surface, so two such flights would be required. Note that the ~40-ton DAV would not be required at all; a crew version of the lander can ferry crew to and from the surface. That suggests that a single 50-ton payload could include all of the required surface habitat, science equipment, power, rovers, food, etc. The first mission would require a cargo lander and crew lander as well as an orbital depot, so the first mission would require three 50-ton cargo flights to establish infrastructure along with the surface cargo. Each additional mission would require only one cargo flight.
     Each lander flight requires 95 tons of fuel, and each mission requires two flights (one 25-ton cargo, one 20-ton cargo + crew). This 190 tons requires about 9 tons of ISRU plant and 2 tons of spares. The Phobos ISRU plant should provide 15 tons of fuel for each lander. We might also send a single tug flight back to Earth each period for a cost of 63 tons of propellant from Phobos. That totals about 7.4 tons of ISRU plant and 1.6 tons of spares. A cargo tug can haul fuel from Phobos to the LMO depot, delivering 37.2 tons per trip at a cost of 25.8 tons of fuel.

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     So, let's compare this to the baseline NASA plan over the course of three missions to Mars. The first trip requires three cargo flights to establish the orbital depot, Phobos ISRU base, tanker fleet and lander fleet. A habitat, science payload, ISRU facility, food and other supplies are delivered to the surface. Total cost is two landers (60 tons), three tugs (16.5 tons), 420 tons of Lunar fuel, a ~15-ton depot (200-ton capacity) and 16.4 tons of ISRU plant (with 3.6 tons of spares). That leaves 55 tons of payload available for the surface mission, with the understanding that the surface ISRU system will provide sufficient water and buffer gas.
     As a first-mission flight, the fuel alone will cost $2.24 billion. Using a middle of the road estimate of $30 million per ton for hardware (including surface payloads) that's almost exactly $5 billion. Launch costs would be three Falcon Heavy and a Falcon 9, about $500 million. Total mission cost, $7.75 billion. Note that this does not include the crew launch, transit habitat or crew recovery; these will be discussed in a later post.
     Future flights would require a single cargo flight, 55.5 tons of hardware ($1.67b) and 140 tons of lunar fuel ($135m). All other components are reusable and already present. Total pricetag for the cargo end is $1.8 billion for each additional mission. After about the fifth mission some of the infrastructure hardware will need scheduled replacement, so this cycle of one expensive mission and four cheap missions could continue indefinitely or the replacement costs could be spread across multiple missions. Still, $11.4 billion for three missions to Mars is cheap compared to $2.5 billion for MSL/Curiosity (which was money well spent) or the $156 billion for Apollo to put men on the moon.

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     Looking past the baseline, the Mars surface base is able to supply nitrogen for breathing gas and argon for electric engine fuel. This can be carried back to EML1 in a reusable cargo tug that paid for itself by delivering its first payload to Mars orbit. The same tug can be reused multiple times. Ultimately the cost of shipping along this route is the cost of spares and replacement of the various ISRU bases, depots and tugs. With Earth launch costs in the $3 million per ton range, any advantage is worth pursuing. Cargo from Mars costs about 126kg of spares to Mars per ton delivered to EML1. Cargo to Mars costs about 76kg of spares to the Moon per ton delivered to LMO. Of course each trip requires the use of a cargo tug costing anywhere from $66-$165 million but in terms of ongoing costs we're looking at 6.78 tons from Earth to support a 50-ton cargo from Mars, a better than 7:1 ratio. Another way to say it is for $20 million in launch costs we can get 50 tons of argon and nitrogen (or Mars core samples or Phobos turnings, etc.) at EML1 for use elsewhere once the system is established, a savings of at least $130 million.

For travel throughout cislunar space, I’ve long been an advocate of having depots on both ends of the journey. The LEO depot provides a refueling stop at the first practical point after leaving the ground, and also a spot for bringing vehicles back from lunar space for refueling for their next trip out. The lunar orbit depot plays a similar role for flights to/from the lunar surface, as well potentially, as being a staging location for departures into interplanetary space. By launching from a lunar facility near the top of earth’s gravity well, it’s both possible to use low-thrust trajectories in and out of cislunar space, as well as to do an earth swingby with a departure burn at apogee for high-thrust departures taking maximum advantage of the Oberth effect.

One important question however has been where to place the lunar orbital facility.

Lunar Orbital Facility Orbit Options
A recent FISO telecon presentation by Ryan Whitley and Roland Martinez of NASA JSC describes and discusses several of these staging orbit options. I’ll be reposting snapshots of a few of their slides to introduce the orbits, but here you can find their full presentation:

http://spirit.as.utexas.edu/%7Efiso/telecon/Whitley_4-13-16/Whitley_4-13-16.pdf

They discuss most of the commonly cited options including Low Lunar Orbit, Frozen Orbits, L2 Halo Orbits, Distant Retrograde Orbits, and a more recently discovered option, Near Rectilinear Orbits.

This slide shows some of the smaller lunar orbit options and descriptions (click for larger image):

And this slide shows some of the larger lunar orbit options, with descriptions (click for larger image):

And this slide shows all of the orbits relative to each other to give you a better idea of what they look like (click for larger image):

Comparison of Options

While Whitley and Martinez in their FISO telecon focus on evaluating the various staging orbits from the standpoint of NASA missions using the Orion capsule, they still provide a lot of useful information for evaluating options for the location of a lunar orbital facility/depot. To me, some of the considerations for locating a lunar orbital facility are:

• How frequently do you have opportunities to travel from a LEO facility to the lunar facility, and how frequently you can travel the other direction?¹
• How much delta-V does it take to go between the facility and LEO and the facility and the lunar surface?
• How long is the transit between the location and the lunar surface?
• How useful is the orbit for supporting deep space missions?
• How hard is it to reach various lunar surface destinations from the lunar orbital facility location?
• What is the thermal environment like in the orbit?
• And how much of the lunar surface to destination delta-V can be provided by some sort of propellantless lunar launch scheme²?

Based on these considerations, I’d like to focus the rest of this post on the pros and cons of the two options I consider most interesting–L2 Halo Orbits and Near Rectilinear Orbits.

Pros and Cons of EML-2 Halo Orbits
EML-2 orbits have been my favorite option ever since learning about the low delta-V cost of reaching them via powered lunar swingbys. They have a lot going for them, including:

• One of the lowest delta-V stopping points in the lunar vicinity, requiring only ~3.43km/s of delta-V from LEO.
• Easy access to/from a LEO facility on every LEO-lunar or lunar-LEO window.
• Any-time access to/from anywhere on the lunar surface.
• Low stationkeeping delta-V³
• Benign and cold thermal environment⁴
• Continuous communications with Earth, and most of the farside of the Moon.
• Good staging point for both deep-space and lunar missions.
• Could become a starting location for a lunar space elevator.

But EML-2 does have a few drawbacks:

• Long LEO-EML2 and EML2-LEO transit times⁵ for the low delta-V powered-swingby option.
• Long EML2 to lunar surface (and vice versa) travel times⁶
• It wasn’t clear that a propellantless lunar launch option located at either pole could launch easily to EML2. An elliptical orbit from such a launcher would have its line of apsides pass through the launch location, which would be orthogonal to the Moon-EML2 line. You could launch into a polar LLO, and then do multiple burns from there to EML2, but the propellantless launch option would only provide the first leg of the trip (surface to LLO).

The long trip times mean that the vehicles taking people between LEO and EML2 and between EML2 and the Moon will require much more extensive life support and accommodations than would be needed if the trip were shorter. That will drive up the dry mass of those systems, and by extension the propellant and overall cost of moving people to and from EML2.

Pros, Cons, and Questions Regarding Near Rectilinear Orbits
Starting several months ago, some of my astrogator friends started telling me about NASA’s interest in Near Rectilinear Orbits for exploration missions. After all the talk about Distant Retrograde Orbits, this sounded a bit like the “flavor of the week” syndrome, but the FISO presentation helps explain some of the allure of such orbits:

• Only slightly higher delta-V to/from LEO to NROs compared to LEO to EML2⁷.
• Because the NRO orbit’s perilune is only 2000km from the Moon’s surface, once per 6-8 day orbit, the orbit lines up so that the travel time between NROs and the lunar surface drops to 0.5 days.
• Powered swingby trajectories between LEO and NROs take approximately 5 days each direction, instead of 9-11 for EML2.
• Slight lower delta-V between NROs and the lunar surface compared to EML2.
• The NRO is close enough to an elliptical polar orbit that it might be possible for a polar base to use propellantless launch techniques to fling payloads nearly into NRO, with possibly only minor adjustments and raising the perilune with a burn near apolune half an orbit later⁸

The benefits of shorter transit times are pretty important, but there are still a couple of relative drawbacks and open questions:

• While it’s possible to get from LEO to a given NRO orbit during every lunar injection window, the NRO facility will be at different points in its orbit during each window, which may make a first-orbit rendezvous either infeasible or it might cost additional delta-V. I’d want to get this resolved, because while this isn’t an issue for one-off, ground-launched missions like the NASA folks were thinking of, this would be a real issue for reusable spaceship flights between a LEO and NRO facility.
• Likewise, departures from the NRO may not be in the optimal part of the orbit for the Earth return maneuver when the timing is right to return to the plane of the LEO facility. This isn’t a problem if you’re doing a direct return, but once again is a big pain in the neck for reuse of space hardware. Once again this is something I’d want to analyze more before settling on an NRO orbit.
• Additionally, the NRO facility has LOS with one lunar pole about 86% of the time (while heading out and coming back from apolune), but only sees the other facility for a brief period near perilune. If you’re planning on using propellantless launch methods to send stuff from a polar lunar settlement to the NRO facility, it’s going to be in an NRO with apilune on the opposite side of the moon from your lunar settlement, meaning you’ll only be in contact briefly for maybe 1 day out of a week.
• Because the perilune is only 2000km, the heating environment is going to be warmer than EML2, with slightly higher boiloff, but this is probably only a minor difference–it should still be tons easier to keep cryo boiloff low in an NRO than in LEO.

Conclusions?
While NRO orbits have some really interesting characteristics, I’d really want answers to those first two concerns before I’d pick it for the location of a lunar orbital facility. If you can’t get to it on a regular basis from a given LEO depot without having to do complicated trajectories, or paying big penalties in flight duration or delta-V, then that would likely outweigh the benefits. If on the other hand, it’s not a big deal to adjust the trajectory on the way to and from the NRO facility to enable rendezvous with the facility regardless of where it is within its orbit when the LEO to lunar launch window opens, then it could be a really interesting location for a lunar transportation node. I’ll have to see if I can get some of my astrogator friends to weigh in on those questions. Until then I’m probably still more of a fan of EML-2, in spite of the annoyingly long transit times.

[Update 1: After speaking with an astrogator friend who’s been looking at NROs to support lunar missions, he thinks it might be possible to put an NRO facility in an orbit whose period is synchronized with the average time between launch windows from the LEO facility. If that works, that would mean the NRO facility would be in approximately the same part of its orbit during each trip to/from the Earth. There are questions of if you can make an NRO orbit with a long enough period (~9 days) to make that work, and if the NRO facility could be made to line up both for arrivals and departures from/to Earth, but hopefully he’ll have more opportunity to dig into that further later this year.]

---

1. Due to the orbital motion of the Moon and nodal regression of a LEO facility, you get optimal lunar departure and/or return options about every 7-9 days IIRC. The choice of lunar orbital facility location may constrain this further.
2. Lunar Slings, Mass Drivers, Launch Loops, etc. All the stuff I was supposed to write about in my “The Slings and Arrows of Outrageous Lunar Transportation Schemes” series that I still need to finish
3. <10m/s/yr
4. About 10x lower LOX/LH2 boiloff rate than LEO. In fact with passive insulation you can completely surpress LOX boiloff and even freeze oxygen at EML-2
5. 9-11 days
6. 3-5 days
7. 3.58km/s vs 3.43km/s
8. Which you’d also want to be the burn that brings you to rendezvous with the NRO facility

For this next depot taxonomy post, we’ll finally be talking about what people usually think about when they hear the term orbital propellant depot — larger, cryogenic refueling facilities, focused on enabling large-scale human spaceflight missions, performed by a diverse variety of users, and going to/from a wide variety of destinations. The idea for such propellant depots for enabling interplanetary human spaceflight dates back to at least 1928 with the writings of Guido von Pirquet¹.

This blog post will be focused on what I call “Low-Orbit” human spaceflight depots. These are depots located near the lowest stable orbit around a planetary body. This is the first place at which you can realistically refuel or switch vehicles on your way from a planetary body, and is the last place at which you can refuel or switch vehicles on your way down. As a firm believer in the idea of refueling early and refueling often, low-orbit depots are an important piece of infrastructure for any planetary system humanity wants to travel to/from regularly. A later blog post in the series will talk about “High-Orbit” depots — depots operating in fixed locations² located out near the edge of a planetary system’s gravitational sphere of influence, and will include a discussion of where those types of depots might make sense.

Before we jump into the weeds about low-orbit human spaceflight depots, I did want to address a recent train of thought I’ve seen that suggests that just using tankers and directly refueling a vehicle is superior to having a depot involved. While this could easily be the topic of a series of its own, I wanted to briefly highlight a few of the biggest advantages I can see of having a depot vs just using direct refueling with tankers:

• Flexibility: A depot, properly designed, with published, standardized grappling, refueling, and power-data interfaces, can be agnostic about who it gets propellant from and who it sells propellant to. Depots can quickly take advantage of whatever the cheapest source of propellant is at a given time (RLVs, ISRU, propellantless launch, buying leftover capacity from other missions going to destinations near a depot, atmospheric gathering, etc), and can easily service both smaller missions and bigger ones. Tanker-based approaches tend to be a lot less adaptable, typically being optimized for one or two specific vehicles that needs refueling.
• Robustness: With a fixed installation, that only has to be launched once for a long mission lifetime, you can afford to throw way more resources (dry mass, volume, and power) at making rendezvous, prox-ops, docking (RPOD), and manipulation as safe and reliable as possible. This could include beacons, larger more capable (and/or redundant) relative navigation sensors and comms, longer reach capture robotics that minimize the dry mass requirements on tanker and client vehicle alike, etc.
  • For tankers, on the other hand, you want to minimize parasitic dry mass that has to be launched every time, and for the departing vehicle you want to minimize mass you have to carry through large, high delta-V in-space maneuvers. You could in theory carry a nicer RPOD kit on your departing vehicle that you jettison before leaving LEO, but now you’re amortizing that mass and cost over a much smaller number of missions, or starting to add more complexity than just doing a depot.
  • Another question to ponder is with vehicles that require large numbers of refueling events per mission, is it best to have the client vehicle handle all of those docking maneuvers with its (by definition) less-capable RPOD capabilities? With a depot, each tanker and each client vehicle only has to perform one mission-critical RPOD/refueling operation per mission, whereas with direct refueling via tankers, the client vehicle would typically have to perform a larger number of mission-critical RPOD events. A depot would also have to handle the same larger number of RPOD/refueling events, but as mentioned before, can throw more resources at making these as reliable and safe as possible.
  • Longer refueling cycles from using a direct-tanker refueling approach also increases the MMOD³ risk to the departing vehicle, which by definition can’t afford to throw as many resources at MMOD protection as a depot can.
  • A corollary to this is that depots make the most sense if you use them in a way that offloads as much of the refueling-unique hardware/software as possible from the delivery and client sides of the system to the depot. Ideally a tanker would be a minimally modified upper stage⁴, and client vehicles would also have similarly minimalistic hardware needed to be grappled and receive propellant. If you’re doing your delivery vehicle or your client vehicle in a way that makes you question the utility of a depot, that may be a hint that you’re doing something wrong.
• Non-Integerality: Yes I may have made up that word, but the point is that tankers tend to come in integer quantities. Unless you always design your departing vehicles to use only integer quantities of tankers, you’ll almost always end up having wasted propellant. This is especially true given how launch vehicles, and in-space vehicles tend to increase performance and get upgraded over time, and don’t necessarily upgrade at the same rates. If you only ever had a monopoly/monopsony situation where you only had one tanker provider, and only one vehicle needing tankers, you might be able to keep tanker size locked to an integer fraction of the amount of propellant needed, but in reality that isn’t going to happen. So a tanker-based system is always going to end up wasting propellants, and this is even more the case when you have diverse customer vehicle and delivery vehicle sizes. The more diversity you have in your space transportation ecosystem, the more a depot makes sense.

There are probably other good arguments I’m glossing over, but long-story short, unless you’re interested in a boring monoculture world where only one type of in-space transportation system exists, depots make a lot of sense. So, without further ado, let’s jump into some of the taxonomical considerations of human spaceflight low-orbit depots.

Human Spaceflight Low Orbit Depots

Application: Refueling large transfer stages or in-space transports for ferrying people and cargo between LEO, the Moon, Mars, Venus, and other destinations of interest.

Location: As discussed earlier, this blog post is focused on depots located near the lowest practical stable orbit around a planetary body. Details vary for different planetary bodies, as described below⁵:

• For Earth, the low-orbit depot location is in LEO, ideally in the lowest inclination that still lets you hit required departure asymptotes and maximize the number of economically useful launch sites to send propellant, people, cargo, and materials to/from the depot⁶. You probably only need one or two such low-orbit depots, though if you get to high-enough earth departure throughput there may eventually make sense to spread more depots out at different RAANs⁷. Likely for a first human spaceflight depot, as with the previously discussed smallsat launcher depots, you’ll want to locate it in LEO near other human-occupied facilities like ISS — far enough away to be safe, but close enough to conveniently move between each other, ideally within one work shift⁸.

• For a low-orbit depot around the Moon, this would likely be a polar or near-polar LLO⁹, though due to the very slow rotation and practically zero J2 perturbation¹⁰, if you have a lot of non-polar surface sites, you eventually may want multiple smaller depots in equally RAAN-spaced near-polar LLO planes, and maybe one in an equatorial orbit. If you’re trying to do lunar surface missions, having your depot in LLO makes way more sense than in a higher orbit like NRHO, for reasons I should probably go into in another blog post.

• For Mars this would also likely be a LMO orbit, with an inclination high enough to be able to access any points of interest on the surface, while still being low enough to minimize delta-V penalties¹¹, and keep the nodal precession rate fast enough to minimize phasing orbit time for three-burn departures. You’ll likely also have to put some thought into perturbations from Phobos and Deimos¹².

• For Venus, the extremely low rotation speed and therefor very low J2 pertubation may require you to do multiple smaller depots in similar inclinations but equally RAAN-spaced planes, as you won’t pass over a given point on the surface very frequently, and the very slow nodal precession rate could potentially require very very long phasing orbits for a 3-burn departure. Venus has a deep enough gravity well that you do want to refuel in LVO coming to/from, but it’s not trivial from an orbital dynamics standpoint¹³.

Size: As big as you can practically get away with — ideally you’d want this depot to be at least 2x the propellant capacity as whatever the largest vehicle you’re refueling. So somewhere in the 100-2000mT range, or even bigger¹⁴. Early versions will want to be single-launch if possible, in many cases repurposing at least one of the main propellant tanks from the stage that delivered them to their destination as one of the depot tanks¹⁵. Eventually, it may be possible to do multi-tank depots, but if you can do a single launch depot big enough for refueling two missions, you may be better off making more than one depot instead of trying to make the depot super big.

Propellant Types: For low depots, you’re primarily going to be dealing with large transfer stages (Centaur V, Starship, New Glenn Upper stage), which typically use LOX, and either Methane or LH2 for the fuel. Most of these use autogenous pressurization, and use the main propellants for RCS. So most of the depot will be for LOX, LH2, and/or Methane.

• For Mars or Venus you may eventually also want to store liquid CO for some applications, since it’s an easier ISRU propellant, but that remains TBD.
• A lunar low-orbit depot may also want to stock storable propellants, depending on what lander propellants end up being most popular¹⁶.
• You may eventually also want to store some secondary fluids (Helium or Neon for active cooling loops, life support consumables like water, air, etc), but you may not explicitly need a depot for that function.

Some Other Considerations for Human Spaceflight Low-Orbit Depots

• As mentioned before, human spaceflight depots want to be designed in a way to enable offloading as much of the RPO and docking/or berthing from the vehicles they’re servicing. The less parasitic dry mass that tankers and clients have to lug around on every mission, the better.
• Storing cryogens in low orbits tends to be hard — you have a warm planet blocking half of the sky. So launching the propellant in a subcooled state or even partially frozen (i.e. slushy propellants) can help a lot. Also a lot should be done to minimize heat leaks between the cold part of the depot and any hot sections (habitation, power, etc). If you can’t get to zero boiloff, LH2 is a great thermal sponge, and can be used to chill other propellants, and intercept heat from heat sources before being vented. You may have to vent some hydrogen boiloff, but if you’re smart, you can use that hydrogen boiloff on the way out to eliminate boiloff issues for everything else.
• These depots are also big debris targets¹⁷. Deployable MMOD/MLI¹⁸ solutions could be very helpful to avoid a puncture, which would probably be very hard to patch. Since these depots are fixed, and are typically only performing stationkeeping maneuvers, it may be possible to augment their MLI/MMOD protection over time using in-space assembly/manufacturing techniques¹⁹.
• Especially for low-orbit depots around the Moon/Mars/Venus, there may be a benefit to having some temporary habitation/shelter collocated with the depot, especially if you’re supporting multiple sites, as a search and rescue option during exploration phases, and as a stop-over point to act as a buffer between different sizes of transportation between planets and between the depot and the surface.
• Over time you may want to add in other facilities such as dry docks for assembling, and repairing/maintaining large in-space vehicles/structures, habitation facilities, etc. But they should probably be coorbital, near the depot, not attached (as that will make cryo thermal management all the harder, and the depot is a big hazardous work location, where you should probably minimize the amount of time people spend in close proximity to it). This could be done in two ways — coorbital facilities, spaced where the safe time to travel from one to the other is as short as practical (definitely less than an 8hr work shift if at all possible, and much closer if possible), or by having the two facilities connected by a connecting tether or other structure that includes elevator facilities.
  • If your depot facility starts wanting to have permanent collocated habitation (say for in-space assembly/repair/maintenance of in-space stages), and having the two be coorbital doesn’t work, you’re likely going to want to keep the people separated as far away from propellant tanks as possible, both to minimize heat-leak into the tanks, but also to minimize hazard to the people²⁰.
  • For larger depots, and ones where people will be there more, putting some thought into spatially separating the fuels and oxidizers more could be a good idea. In rockets you often can’t do much to keep the two separated, and many use a common bulkhead, but in a fixed facility it’s more of a possibility. Having fuel and oxidizer that close together for long periods of time is somewhat tempting fate — you kind of have to do it for high efficiency rockets, but there’s something to be said for having your fuel and oxidizer many meters apart for a long-duration facility.
• One area of disagreement I have with other depot advocates is whether propellants should be shipped to a depot as cryogenic propellants (LOX/LH2/LCH4), or if you should ship them as something more storable like water and CO2, and have the depot itself have large-scale electrolyzing, separation, and propellant refrigeration systems. My concern is that while in theory very large solar arrays could be done in space, combining large flexible structures like multi-MW solar arrays and radiators with a facility that sees a lot of docking, propellant slosh, etc seems like a bad idea from a structural dynamics standpoint. Also depots with very large power generation and heat rejection capabilities are likely to come later in the process, since they’ll almost certainly require multiple launches and in-space construction.

Anyhow, I probably could go on, but as with the previous parts of this series, I am only trying to scratch the surface with considerations and operating details, as I introduce each new type of depot. This definitely isn’t the last you’ll hear from me on the topic.

1. Sykora, Fritz, “Guido von Pirquet-Austrian Pioneer of Astronautics,” History of Rocketry and Astronautics, R. Cargill Hall, ed., AAS Publications, San Diego, 1986, p. 151. And yes, I’d love to see someone actually fly a depot before we hit the 100th anniversary of von Pirquet first discussing the idea
2. Or at least orbits/halo orbits where the only maneuvers being performed are for stationkeeping purposes
3. MicroMeteorite and Orbital Debris
4. With some lightweight grapple fixtures, refueling modified T-0 umblicals, and maybe an upgraded comms/controls system
5. Most of these points could easily justify their own blog post, and knowing me, I’ll probably eventually do said blog posts, but I’ll try to keep things in this post brief.
6. I’m currently noodling doing a paper looking specifically at this optimization — higher inclinations for the depot give you the ability to hit higher departure declinations, and are accessible by more launch sites, but at the cost of slower nodal precession driving longer hang-times for a three-burn departure, and you also have decreased payload mass to higher inclinations. My gut says ISS-like orbit is probably not far from the optimal point, but it would be fun to run the numbers
7. As a reminder, RAAN stands for Right Ascension of the Ascending Node, which is a measure of where a given orbit plane crosses the equator heading northward. Here’s a decent wikipedia explainer. When you see a multi-plane constellation, like say OneWeb or Starlink, in a lot of cases the satellites will be put into multiple planes with each plane having the same inclination and altitude, but a different RAAN. The different RAANs help with more frequent passes over the same point in the ground, and for depots by increasing the frequency with which a depot passes through the plane of a specific departure asymptote. If that makes any sense.
8. I could easily see a scenario where the first human spaceflight depot evolves from smallsat launcher depots.
9. Though how low of a LLO is open to debate — lower orbits have higher stationkeeping requirements, and are harder to do sunshields with because the Moon takes up more of the sky surrounding the depot. I haven’t done an optimization analysis myself, but 250-500km above the Moon might reduce both of these effects noticeably compared to the 100-150km parking orbits used for previous lunar missions
10. The J2 parameter is a measure of how oblate or “round about the middle” a planet is. This is typically tied to the rotation speed of the planet. The J2 perturbation is what causes orbital planes to precess slowly. If a planet had no J2 or very low J2, the planes would stay in a fixed orientation relative to the stars, which is problematic if you’re trying to get the plane to line up with a departure asymptote for an interplanetary departure… Earth, Mars, and most of the gas giants have high J2s, while the Moon, Venus, and Mercury all have very low J2s.
11. The equatorial velocity of Mars is ~half that of earth, so it’s not as big of a deal to launch from or into a higher inclination, but it still is some losses
12. I’m not much of a Mars guy, so am handwaving that part a bit
13. If someone is looking for a PhD dissertation topic, this feels like an area that could use a lot more skullsweat investigating.
14. Depending pretty strongly on how successful Starship ends up being, and how open Elon is to using other people’s depots. In a world where Starship either doesn’t pan out, or Elon insists on doing his own thing, the required size for refueling other missions can be a lot more modest initially. Also a world with depots in low-orbits around destination planets, and roving depots and/or high-orbit fixed depots may not need LEO depots to be quite so big.
15. As shown in Part IV of this series.
16. For long duration lunar landers there are definitely differences of opinion about whether storables or cryogens make more sense.
17. Big LEO facilities that don’t want to dodge things all the time is one of those reasons why moving to a leave-no-trace approach to satellite operations is going to be an important part of growing up as a spacefaring civilization.
18. MLI, or multi-layer insulation is a type of very effective insulation for use in vacuum environments. It is typically made of many layers of thin metalized plastic films separated by nets or spacers so that most of the heat has to transfer via radiation instead of conduction or convection. Having more spacing between layers can help. And in some cases, MLI can be combined with MMOD (which also wants to be multiple thin layers with spacing between them), like what our friends at Quest Thermal have worked on.
19. Most hypervelocity impact shielding benefits a lot from extra spacing between bumper layers, so an in-space assembled/manufactured MMOD solution could be particularly useful
20. Though as my wife pointed out, if your depot facility starts wanting anything, it’s probably a sign of an impending robot apocalypse, so the point might be moot.

• 

Space Drydocks

One of the most unique applications of inflatables is being proposed by General Electric: emergency drydocks for and maintenance of orbiting vehicles and spacecraft, some of which by necessity will always be rigid metallic types.

“The drydocks”, reports E. J. Merrick, project engineer in GE's MSVD, “could be as simple as a plastic sausage-skin drawn over and around the entire craft and then inflated.” Once pumped full of breathable air, space repairmen would not need to wear spacesuits as they scrambled over the metallic vehicle to be fixed.

Their job done, the sealed drydock “sausage” is again deflated, folded and packed away in a ferry rocket for use and again. Safety and comfort for the spacemen, Merrick adds, plus reliability and efficiency of the inflatable technique, will make such convenient once shelters ideal for space drydock missions.

For lesser maintenance and jobs that do not require an immense drydock area around the entire vehicle, GE has designed the smaller “space hog.” Based on the earthly sand-hog technique of providing caissons for men to work in, GE's concept is a pressure-tight, non-rigid tube or cylinder with stiffening rings down its length. In different sizes for one spaceman or several, these “space hog” units could easily be inflated outside of a spacecraft, providing a safe temporary environment for specific at any desired spot.

• 

      Engineer-Captain Mikhail Borisovich Andreev sat strapped behind his oversize desk and worked to peel off a few more items from his overflowing day list before his VIP visitor showed up.  Extending beyond his office in every direction were the slate gray bulkheads and oversize machinery of Orbital Shipyard Delta Seven, recently departed from orbit around Halcyon IV, and now in orbit about Eulysta II, known to its former owners as Corlis.  That, at least, was the human transliteration of the unpronounceable Ryall phonemes that made up the true name of the planet.

     At its most basic, an orbital dockyard performed the same functions as its groundside counterparts.  It just did so in micro gravity.  Delta VII had the ability to build anything up to a light cruiser from scratch, and with some monkeying of the cradles with which it enveloped its wounded patients, could perform major surgery even on one of the big blastships.
     At the moment, the big dock’s restorative hangars were empty, and its first customer was to be a mere light cruiser, which didn’t seem to justify the epic journey through the nebula that Andreev and his men had just completed.
     A space dock is not quantitatively different in function from any other spaceship.  To be effective, it had to have compartments conditioned to shirtsleeve environments in which its crew lived.  Unlike a planet-based dock, one of the big spherical dockyards had to be mobile so that it could be moved to where it would be most useful, which meant it required power reactors and both normal space and jump engines.
     In a war that extended across dozens of star systems and hundreds of light-years, it was unreasonable to expect a wounded ship to return to the place of its birth.  It was more efficient for them to jump one or at most two systems back from the front lines and be repaired close to the scene of battle, the better to return as quickly as possible to the fray.

---

     Philip explained the damage that his ship had taken in a few carefully composed sentences.  The repair officer listened, then nodded slowly.
     “Standard Illustrious-class light cruiser, isn’t she?”
     “Yes, sir.  Built last year at Sandar from Terrestrial Space Navy specifications.”
     “Good, then you use all standard modules.  That means that we will be able to work quick and dirty.  Rip out everything that doesn’t work, weld on a new bow section, and then stuff the hull with new equipment still in the packing boxes from the factory.  We won’t even try to repair your old equipment, just ship what seems salvageable back to human space for a depot to handle.  When we get through with Queen Julia, she’ll be better than new.”
     “How long?” Philip asked.
     “A month, six weeks at the max.  That assumes that something with higher priority doesn’t materialize in the foldpoint and have you kicked out of the bay before we have time to finish the job.”
     “Can my crew help?”
     “Sure.  We can always use some trained hands and that way, they will be up to speed on the new stuff when we send you back to space with a shiny new coat of paint inside and out.”
     “Thank you, Captain.  You don’t know how frustrating it has been to turn our backs on the action and limp here when our mates are getting the hell kicked out of them.”
     “Captain Walkirk,” Andreev replied with a wistful tone, “I have been a repair officer for sixteen years and seen ships and men head out into the deep black to fight the enemy, never to return.  I know precisely how frustrating it is...”

---

     Despite Philip’s impatience, repairs on Queen Julia had progressed with surprising speed after the tugs maneuvered the crippled cruiser into the all-encompassing embrace of the big space dock.  As Captain Andreev, the dock commander, had pointed out, repairs were greatly facilitated by the cruiser’s design, which was based on the Terrestrial Space Navy’s Illustrious III class of warships.  Like her Illustrious sisters, Queen Julia used standard modules throughout her hull.

     The damage to Julia was sufficiently extensive that had the cruiser been one of the older ships of the Royal Sandarian Navy, or of the Altan Space Navy for that matter, she would probably have been scrapped.  Repairing those 150-year-old designs would have taken too much time and too many scarce resources.  For modern ships, with their interchangeable parts, repairs were the equivalent of a child’s game of building sticks.

• 

     The space dock technicians had begun the repair by slicing away the cruiser’s smashed-in bow with a laser as powerful as any carried by a blastship.  It had been disconcerting to look at his ship and see it in cross-section, with compartments, passageways, and utility conduits all open to space.  It had been even more disconcerting to watch the minor surgery that had followed the amputation of the bow.  For more than a week, dockyard technicians had swarmed over the ship, cutting out partially melted sections of hull and interior structure, stripping away kilometers of optical cabling that had been clouded by radiation exposure, and emptying equipment racks of components that triggered fault messages when queried by diagnostic routines.  At first, Philip and his crew acted as unskilled helpers in this systematic vandalism, taking direction from the dock’s skilled cadre of ship wreckers.
     As 16- and 20-hour days began to blur together, however, the cruiser’s crew began to take on more of the repair tasks themselves.  Not only were they becoming more skilled, but also the dock’s personnel were increasingly diverted to service other cripples.

     Altogether, Queen Julia spent 22 days surrounded by space dock scaffolding and movable work centers.  At the end of that time, when the ship was once again vacuum tight, Captain Andreev ordered his dock cleared so that he could begin repairs on another victim of the continuing contest over who would control Spica.  Philip had watched from an inter-orbit scooter as the dock’s massive clamshell doors opened and his ship was again exposed to Eulysta’s warming yellow rays.
     Interior work on the ship proceeded apace even while tugs gently shifted the recuperating cruiser to a parking orbit aft of the repair dock.  Repairs continued for four more weeks as Julia’s crew slowly put their ship back together.  The list of things needing fixing seemed endless.  There were networks to synchronize, interface nodes to reconnect, missile launchers to align.  Most of these tasks required the attention of skilled technicians, all were time consuming, and Philip never seemed to have enough labor of the right sort to satisfy even half of the demands for immediately attention.

     Yet, despite workdays that were much too long and infrequent sleep periods, looking back on it, he could not remember a time when he had been happier.

     Orbital Shipyard Delta Seven was more than just an oversize body and fender shop.  In addition to its primary function, it sported a completely equipped orbital hospital where doctors worked to heal injured crewmembers while the dock technicians attempted to repair their ships.

• 

(ed note: the planet Canaan has an asteroid-moon named TerVeen. The latter has been converted into a shirt-sleeve repair dock. "Climbers" are small cloaked starships delivered to their patrol routes by motherships.)

      Westhause continues to explain. “What they did was drill the tunnels parallel to TerVeen’s long axis. They were cutting the third one when the war started. They were supposed to mine outward from the middle when that was finished. The living quarters were tapped in back then, too. For the miners. It was all big news when I was a kid. Eventually they would’ve mined the thing hollow and put some spin on for gravity. They didn’t make it. This tunnel became a wetdock. A Climber returns from patrol, they bring her inside for inspections and repairs. They build the new ones in the other tunnel. Some regular ships too. It has a bigger diameter.”
     In Navy parlance a wetdock is any place where a ship can be taken out of vacuum and surrounded by atmosphere so repair people don’t have to work in suits. A wetdock allows faster, more efficient, and more reliable repairwork. (so this author has Docks and Wetdocks, instead of Wetdocks and Drydocks)
     “Takes a month to run a Climber through the inspections and preventive maintenance. These guys do a right job.”

     The bus surges forward. I try to watch the work going on out in the big tunnel. So many ships! Most of them are not Climbers at all. Half the defense force seems to be in for repairs. A hundred workers on tethers float around every vessel. No lie-in-the-comer refugees up here. Everybody works. And the Pits keep firing away, sending up the supplies.
     Sparks fly in mayfly swarms as people cut and weld and rivet. Machines pound out a thunderous industrial symphony. Several vessels are so far dismantled that they scarcely resemble ships. One has its belly laid open and half its skin gone. A carcass about ready for the retail butcher. What sort of creature feeds on roasts off the flanks of attack destroyers?
     Gnatlike clouds of little gas-jet tugs nudge machinery and hull sections here and there. How the devil do they keep track of what they’re doing? Why don’t they get mixed up and start shoving destroyer parts into Climbers?

---

     Our mother ship is one of several floating in a vast bay. The others have only a few Climbers suckered on. Each is kept stationary by a spiderweb of common rope. The ropes are the only access to the vessel. “They don’t waste much on fancy hardware.” Tractors and pressors would stabilize a vessel in wetdock anywhere else in the Fleet. Vast mechanical brows would provide access.
     “Don’t have the resources,” Westhause says. “‘Task-effective technological focus,’” he says, and I can hear the quotes. “They’d put oars on these damned hulks if they could figure out how to make them work. Make the scows more fuel-effective.”

     I want to hang back and look at the mother, to work out a nice inventory of poetic images. I’ve seen holoportrayals, but there’s never anything like the real thing. I want to catch the flavors of watching hundreds of upright apes hand-over-handing it along with their duffel bags neatly tucked between their legs, as if they were riding very small, limp, limbless ponies. I want to capture the lack of color. Spacers in black uniform. Ships anodized black. The surface of the tunnel itself mostly a dark black-brown, with streaks of rust. The ropes are a sandy tan. Against all mat darkness, in the low-level lighting, without gravity, those lines take on a flat two-dimensionality, so all of them seem equally near or far away.

---

     The bearing and tilt on the camera tell me nothing. Forward. It should be staring at the wall of the wetdock. Instead, the screen shows me an arc of darkness and only a small amount of wall. The lighting seems brilliant by contrast with the darkness.
     High on the wall, at the edge of the black arc, a tiny figure in EVA gear is semaphoring its arms. I wonder what the hell he or she is up to. I’ll probably never know. One of the mysteries of TerVeen.
     Damn! How imperceptive can one man be? We’re moving out. We’re under way already. Must have been for quite a while. That creeping arc of darkness is naked space. The mother is crawling out of TerVeen’s backassward alimentary canal.

     I close my eyes and try to imagine our departure as it would appear to an observer stationed on the wall of the great tunnel. The Climber people come hustling in, hours after the mothercrew has begun its preparations. They swarm. Soon the mother reports all Climbers manned and all hatches sealed and tested. Her people scamper over her body, releasing the holding stays, being careful not to snap them. Winches on the tunnel walls reel them in.
     Small space tugs drift out from pockets in the walls and grapple magnetically to pushing spars extending beyond the mother’s clinging children.
     Behind them, way behind them, a massive set of doors grinds closed. From the observer’s viewpoint they’re coming together like teeth in Brobdingnagian jaws. They meet with a subaudible thud that shakes the asteroid.

     Now another set of doors closes over the first. They snuggle right up tight against the others, but they’re coming in from left and right. Very little tunnel atmosphere will leak past them. Redundancy in all things is an axiom of military technology.
     There are several vessels caught in the bay with the departing mother. They have to cease outside work and button up. Their crews are cursing the departing ship for interrupting their routine. In a few days others will be cursing them.

     Now the great chamber fills with groans and whines. Huge vacuum pumps are sucking the atmosphere from the tunnel. A lot will be lost anyway, but every tonne saved is a tonne that won’t have to be lifted from Canaan.
     The noise of the compressors changes and dwindles as the gas pressure falls. Out in the middle of the tunnel, the tugs slow the evacuation process by using little puffs of compressed gas to move the mother up to final departure position.
     Now a pair of big doors in front of the mother begins sliding away into the rock of the asteroid. These are the inner doors, the redundant doors, and they are much thicker that those that have closed behind her. Great titanium slabs, they’re fifty meters thick. The doors they back up are even thicker. They’re supposed to withstand the worst that can be thrown against them during a surprise attack. If they were breached, the air pressure in the 280 klicks of tunnel would blow ships and people out like pellets out of a scattergun.
     The inner doors are open. The outer jaws follow. The observer can peer down a kilometer of tunnel at a round black disk in which diamonds sparkle. Some seem to be winking and moving around, like fireflies. The tugs puff in earnest. The mother’s motion becomes perceptible.

     A great long beast with donuts stuck to her flanks, moving slowly, slowly, while “Outward Bound” rings in the observer’s ears. Great stuff. Dramatic stuff. The opening shots for a holo-show about the deathless heroes of Climber Fleet One. The mother’s norm-thrusters begin to glow. Just warming up. She won’t light off till there’s no chance her nasty wake will blast back at her tunnelmates.
     The tugs are puffing furiously now. If the observer were to step aboard one, he would hear a constant roar, feel the rumble coming right up through the deckplates into his body. Mother ship’s velocity is up to thirty centimeters per second.
     Thirty cps? Why, that’s hardly a kilometer per hour. This ship can race from star to star in a few hundred thousand blinks of an eye.

     The tugs stop thrusting except when the mother’s main astrogational computers signal that she’s drifting off the cen-terline of the tunnel. A little puff here, a little one there, and she keeps sliding along, very, very slowly. They’ll play “Outward Bound” a dozen times before her nose breaks the final ragged circle and peeps cautiously into her native element. Groundhog coming up for a look around.
     The tugs let go. They have thrusters on both ends. They simply throw it into reverse and scamper back up the tunnel like a pack of fugitive mice. The big doors begin to close.
     The mother slides on into the night, like an infant entering the world. She hasn’t actually put weigh on but has taken it off. She’s coming out the rear end of TerVeen, relative to the asteroid’s orbit around Canaan. The difference in orbital velocities is small, but soon she’ll drift off the line of TerVeen’s orbit.
     Before she does, word will come from Control telling her the great doors are sealed. Her thrusters will come to life, burning against the night, blazing off the dull, knobby surface of TerVeen. She’ll gain velocity. And up along her flanks will gather the lean black shapes of her friends, the attack destroyers.

• 

A momentum exchange tether is a kind of space tether that can be used as a launch system, or to change spacecraft orbits. Momentum exchange tethers create a controlled force on the end-masses of the system due to centrifugal acceleration. While the tether system rotates, the objects on either end of the tether will experience continuous acceleration; the magnitude of the acceleration depends on the length of the tether and the rotation rate. Momentum exchange occurs when an end body is released during the rotation. The transfer of momentum to the released object will cause the rotating tether to lose energy, and thus lose velocity and altitude. However, using electrodynamic tether thrusting, or ion propulsion the system can then re-boost itself with little or no expenditure of consumable reaction mass.

A non-rotating tether is a rotating tether that rotates exactly once per orbit so that it always has a vertical orientation relative to the parent body. A spacecraft arriving at the lower end of this tether, or departing from the upper end, will take momentum from the tether, while a spacecraft departing from the lower end of the tether, or arriving at the upper end, will add momentum to the tether.

A momentum exchange tether is a long thin cable used to couple two objects in space together so that one transfers momentum and energy to the other. A tether is deployed by pushing one object up or down from the other. Once the two objects are separated by enough distance, the difference in the gravitational force at the two locations will cause the objects to be "pulled" apart. This is called the "gravity gradient force". The tether can then be let out at a controlled rate, pulled by the tension caused by the gravity gradient force. Once the tether is deployed, if there are no other forces on the tether it will have an equilibrium orientation that is aligned vertically. There are a number of different concepts for momentum exchange using tethers. Some general categories are:

A bolo is a long rotating cable anywhere in space that is used as a "momentum-energy bank". It could be used to "catch" a payload coming from any given direction (in its plane of rotation) at any given speed (less than its maximum tip speed), and then some time later, "launch" the payload off in some other direction at some other speed. A gravity gradient stabilized bolo orbiting some planet has the property that if the tether is cut, then one-half an orbit later, the separation distance between the two masses is seven times larger than the initial separation. This can be used to deorbit the lower mass, or throw the upper mass to a rendezvous or to escape.

(ed note: momentum-energy bank is just like a financial bank, only using momentum instead of money. You can put momentum in by catching a spacecraft and take momentum out by launching a spacecraft, but the important point is you cannot take out more momentum than the blasted bank currently contains. This means you need a strategy to balance deposits and withdrawals of momentum.

The most common strategy is to have two bolos, for example one at Terra orbit and one at Mars orbit. Each bolo lobs payload modules at the other, and catches modules aimed at them. This keeps the momenum balanced.)

---

Tip Velocity and Material Strength

The maximum tip speed of all these systems is a function of the "launcher to payload mass ratio" of the tether system and the "characteristic velocity" of the material used. The characteristic velocity of the material in a tether is given by the square root of the ratio of the design tensile strength T of the tether to the density D of the tether material. u = (T_d/D)^1/2. In practice, the design tensile strength is usually chosen to be 50% of the measured strength for metals and 25% of the measured short-term individual fiber strength for other materials. Thus, using imperfect materials with reasonable safety margins, the characteristic velocity of most metals and fibers is around 1 km/s, with optimistic predictions for graphite and improved polymers reaching 3 km/s. With the development of a design for a high strength-to-weight tapered Hoytether, the design tensile strength can be safely chosen to be 60% of the measured strength of the individual fibers, allowing commercially available fibers to have characteristic velocities up to 4 km/s.

• 

     "Just the general idea. It starts with the Spider again. Now it's spinning a different kind of web. Rockets are wrong. That's sitting there in your desk as we talk, but I didn't follow it far enough. I should have known you wouldn't stop with the beanstalk, that just gets us up and down from Earth. You wanted a way of moving materials around the whole System without using drives. And the Spider could give you that."...
     ..."Spin another cable," he went on. "Make it like the beanstalk, with superconducting cables and drive train attached to the load cable. This time, put the powersat at the center of the cable, with an equal length on each side of it. Fabricate it in space, but don't ever plan to fly it in and tether it. Leave it out near the orbit of Mars, or in the Belt, or in near Earth—key places in the System. Then start it rotating about its center, like a couple of spokes on a wheel. I assume that you began with just a couple of them, one in the Belt and one near Earth?"
     Regulo nodded calmly. He had finished fiddling with the control panel and now seemed oddly relaxed. "We started with two. That's just the beginning. The more you have, the better the efficiency of the whole operation. I've been thinking we'd build about five thousand of them through the Earth-Belt region."
     "You could handle that many?"
     "With Sycorax? Easily. We can track that number, and more—there are millions of orbits in the data banks already. This is just a few extra ones." Regulo's tone was that of a patient teacher. "I've told you before, Rob," he went on. "Think big. The System's a big place. You have to scale your thinking to match it."...
     ..."Just as you did. You have a rotating cable out in a free orbit—thousands of kilometers of it." He leaned forward, at the same time as Regulo moved his chair farther away from the desk.
     "Now suppose you want to move a space pod from the Belt to the Moon," Rob went on. "You make it rendezvous with the center of the cable, where the powersat sits. The center of mass of the cable would be moving in a free-fall orbit, travelling about the same speed as the pod, so you use hardly any reaction mass to make the rendezvous. You don't need much acceleration from the pod's drives, either, just a fraction of a gee will be enough. Once you have the pod at the middle of the cable, you let it move out along the drive train. As the pod moves from the center it feels a centripetal acceleration. You need to use the drive train on the cable to restrain it. When it reaches the end of the cable, you release it to move in free fall. You've given it a big velocity boost. But the trouble from the point of view of a human on the pod is the acceleration. Out at the end of the cable, it's huge. I looked at a couple of examples. A cable four thousand kilometers long, with end velocity of twenty-four kilometers a second, would give thirty gees at each end. That's what killed the Goblins."...
     ...He put the space pod to a cable rendezvous with a cargo Slingshot—one with high accelerations, never intended for people."
     "Do you have Slingshots for passengers?" Rob moved forward right up to the desk.
     "We built the first two, just a month ago. I could find out which cable your Goblins used easily enough, by checking the angular momentum of all of them. Each time we use a Slingshot we naturally increase or decrease its angular momentum." Regulo stood up, his back to the wall. "We lose angular momentum when we throw a cargo in toward the Sun, and pick it up when we catch something thrown in from Mars or the Belt. Provided we move the same mass of materials in and out, the whole system balances—just like the beanstalk back on Earth.

• 

     My post Orbital Momentum as a Commodity describes how a tether with a healthy anchor mass can catch and throw payloads. I tried to think of ways a tether might restore orbital momentum lost during a catch or throw. Two way traffic is one way to pay back borrowed momentum.

     Well, Mars' moon Phobos masses 1.066×10¹⁶ kg. With this huge momentum bank, catching and throwing payloads would have less effect than a gnat hitching a ride on a Mack truck. A Phobos anchored tether could catch and throw for millennia with little effect on Phobos' orbit.

     The tether illustrated above doesn't suffer the enormous stress of a full blown earth elevator or even a Mars elevator. It could be made from Kevlar with a taper ratio of about 11.

Access to Mars

     The tether foot pictured above moves about 0.6 km/s with regard to Mars surface. This is about 1/10 of the ~6 km/s the typical lander from earth needs to shed. Mars Entry Descent and Landing (EDL) would be vastly less difficult.

     Some have suggested Phobos 1.88 g/cm³ density indicates volatile ices. If so, the moon could also be used as a source of propellent. A Phobos propellent source would make EDL even less of a problem. However Phobos' low density might also be due to voids within a rubble pile.

     On page 2 of the Acceleration of the Human Exploration of The Solar System with Space Elevators Asteroid Initiatives founder Marshall Eubanks takes a look at how the foot of Phobos-Anchored Martian Space Elevator (PAMSE) might interact with Mars' atmosphere:

     The orbital eccentricity of Phobos amounts to 283 km, which is by coincidence comparable to the effective depth of the Martian atmosphere for satellite drag (typically ~ 170 km, but subject to variations due to atmospheric events such as dust storms). The average relative velocity between the lower tip and the surface of Mars is only 534 m/sec, roughly Mach 2 in the cold Martian atmosphere, and slow enough that it should not cause significant heating of the tip. This raises the interesting possibility that the PASME tip could dip down deep into the atmosphere to leave or recover payloads or perform reconnaissance, acting as a supersonic airplane for the period near periapse when it is near the surface.

     Eubanks' 534 m/sec is a little slower than the 0.6 km/s of my tether tip. This might be because I had placed my tether tip 300 km/s above Mars' surface thinking atmospheric friction would destroy a lower tether foot. Eubanks' analysis has changed my view.

     In the Facebook Asteroid Mining Group, Eubanks noted:

     The orbit of Phobos is equatorial, and there is a big mountain in the way, Pavonis Mons, the middle of the Tharsis volcanoes, straddling the equator and by far the highest obstacle in the path of the elevator tip. Maybe a railroad on top of the volcano could match speeds with the elevator tip, once every 3 days or so (when the orbit and volcano aligned). If so, you would have up to 3 minutes to shift cargo on and off. 

     as well as

     …the cool thing is that the tip can be something like a tethered airplane (with wings and flaps, etc.) and you should be able to use that to control oscillations. I was hoping to get money to begin actually "testing" this (i. e. in simulation), but, alas, not so far. 

     Remember, too, with the PAMSE the counterweight has ~ infinite mass, and so any oscillations have to end there. (of course, anchoring a PAMSE in Phobos is left as an exercise for the reader.)

     If Phobos is indeed a loose rubble pile, anchoring the elevator would be difficult. So while Eubanks eased my anxieties on oscillations and atmospheric friction, he calls my attention to a problem I hadn't thought of.

Access to Earth

     6155 km above Phobos the tether is moving faster than escape velocity with a Vinf of 2.65 km/s. This is sufficient to toss a payload down to a 1 A.U. perihelion. This could provide most of the delta V for Trans Earth Insertion.

     A ship coming from Earth would have a Vinf of 2.65 km/s and so rendezvous with this part of the tether might be accomplished with little propellent.

Access to the Main Belt

     7980 km above Phobos the elevator is moving with a Vinf of 3.27 km/s, enough to hurl payloads to a 2.77 A.U. aphelion. This part of the tether might send/receive payloads to/from the Main Belt. There are a lot of asteroids with healthy inclination, though. So there would be substantial plane change expense at times.

Possible Mars exports to the main belt

     One thing about the Main Belt, the pace is much more leisurely. Ceres moves about 1º every 5 days. In contrast earth moves about 1° a day and a satellite in low earth orbit moves about 4° a minute.

     So a month-long, low-thrust ion burn over there looks a lot more like an impulsive burn than it does in our neck of the woods. I believe high ISP ion engines are well suited for travel about the Main Belt.

     The inert gas argon can be used as reaction mass for ion thrusters. Mars' atmosphere is about 2% argon. It is also about 2% nitrogen and 96% carbon dioxide with traces of oxygen and water. Mars also has respectable slabs of water ice at the poles.

     Mars would be a good source of propellent for the entire belt as well as CHON for the volatile poor asteroids in the inner main belt.

     Ion engines don't have the thrust to weight ratio to soft land on the larger asteroids. But asteroids often have high angular velocity (in other words, they spin fast). High angular velocity combined with shallow gravity wells make asteroids amenable to elevators.

     For example the balance point for a Ceres elevator would only be 706 km above Ceres surface, that is the altitude of a Ceres-synchronous orbit. To provide enough tension to remain erect, the elevator would need to extend to an altitude of 2000 km. At 2000 km, the tether tip is moving about 0.46 km/s, a good fraction of the 2.82 km/s needed fro Trans Mars insertion. If this Ceres elevator is Kevlar, taper ratio would be about 1.02.

     If extended to an altitude of 14,500 km, the Ceres elevator top would be moving fast enough for Trans Mars insertion. This would require a taper ratio of around 5 for a Kevlar tether.

Incremental Development

     The tether pictured at the top of this post is ~14,000 km long with a taper ratio of 11 for Kevlar. While much smaller than a full blown Mars elevator, this elevator would still be a massive undertaking. But the whole thing doesn't need to be built overnight. Early stages of the elevator would still be useful.

• 

     Pictured above a Deimos tether drops a payload to a Phobos tether.

     At apoapsis of the large ellipse, payload velocity matches the Deimos tether foot. At periapsis, the velocity matches the speed of the Phobos tether top. Thus payloads can be exchanged between these Martian moons using practically zero reaction mass.

     After descending the Phobos tether, the payload can be dropped to a Mars atmosphere grazing orbit.

     These tethers are a lot shorter than 14,000 km tether we were talking about and taper ratio is close to 1.

No Moons to Dodge

     A full blown Mars elevator capable of throwing payloads to the Main Belt or even earthward would have to dodge Deimos as well as Phobos.

     A Phobos elevator for flinging payloads to Ceres ends well below Deimos' orbit. And of course a Phobos anchored tether doesn't need to dodge Phobos.

Summary

     Tsiolkovsky's rocket equation and big delta V budgets are touted as show stoppers for routine travel to Mars' surface or the Main Belt.

     With judicious use of tethers and orbital momentum, rhinoceros sized delta V budgets are shrunk to hamster sized delta V budgets. No bucky tubes needed, ordinary materials like Kevlar can do the job.

• 

     The landing system was from what Gallagher called their solar tap. They were tapping the electrical potential that exists between a planet and its orbiting proton and electron belts—the belts of ionized particles caught in the planet's magnetic field.
     The landing system was part of a power system that produced, from this one site, enough electricity to power the entire continent on a broadcast basis.
     Broadcast power. It had been known on Earth since the days of Nicola Tesla—the system for putting power on the airwaves the way radio and TV are broadcast. Electric power that you could tune into, the way you tune in a radio.
     With broadcast power, you didn't have to have wires strung around the continent to plug in motors and appliances and furnaces and the like. You didn't have to carry your own fuel in! your ground car. You tuned in your motor to the power frequency, the way you tune in a radio to the frequency of the station you want to hear.
     Earth hadn't had broadcast power, though she'd known how to broadcast it, because the production of power was geared to installations that didn't have sufficient potential that you could waste it on the airwaves. But the power potential in the solar tap was so great you could throw it away on an inverse-square: basis and still be able to tune in at the coast lines, two thousand miles distant, and run anything you wanted to run, from a manufacturing complex to a skimmer. (I'm not sure that is such a good idea. The thrown-away power has to go somewhere, probably turning into waste heat. And you thought greenhouse gases were bad for climate change...)
     The power that exists between the ground potential on any planet and the orbiting proton and electron belts trapped in the magnetic field of any planet, is fed by the solar wind of the sun around which the planet orbits, and it is a practically limitless potential. Electrons from the solar wind make their way in through the magnetic poles of the planet, distribute themselves at its crust, and seep through the insulating atmosphere towards the strong positive potential of the inner proton belt.
     If you make a "short circuit" through the atmosphere by creating an ionized pathway with a laser beam that reaches to the ionosphere, the top of the insulating atmospheric layers, the electrons will jump across the short circuit, changing the groundside potential. When the groundside potential lowers, it makes it possible for more electrons to pour in from the solar wind to equalize the potential. The planet is effectively recharged, and you can short-circuit again.
     It's done in milliseconds, and it's done on a pulse-basis. You turn the laser-beam short circuit off and on in an alternating-current effect, and it's most efficient at a low sonic frequency, although it has radio-frequency overtones.
     There was a group of huge pyramidal structures that were the bases for the solar tap and landing system. A huge, central pyramid was the -tap itself; built of granite with a marble overlay, and of sufficient size to insulate the tremendous bursts of power flow from the ground. The laser installation was on a small platform at the peak of the pyramid, and the control systems were centered well inside where the X rays and other radiation from the flow would not harm the technicians.
     From this, central pyramid, the pulsed power was broadcast across the continent, and even from inside the canteen and at this distance you could hear the deep-throated roar of that power, pulsing through at a frequency within the audible range. Chee-ops, chee-ops, chee-ops, it seemed to say as it shorted in, was cut off, and pulsed in again.
     It was the landing system that used the smaller, satellite pyramids around the big one and that used other factors of the huge central pyramid as well.
     The landing system was a gigantic web of laser beams, angled upward and focused to create a huge electrical discharge spiral that used magnetic induction and repulsion to bring the meteors in. They could bring in any metallic ship as well on that huge spiral, even though the interstellar ships were comparatively fragile; for the gentle cradle of the magnetic induction-propulsion system could raise or lower the gigaton masses as evenly as a freight elevator might bring down a crate of delicate electronic equipment.
     The landing system. From the huge central pyramid, and from each of the smaller ones, two great alternating laser beams angled upward, aimed through tunnels internal to the pyramids and geometrically accurate to a hair. Those beams discharged their alternating spirals into a crisscrossed web of induction-repulsion that caught a ship and either stepped it up from rung to rung of the magnetic induction-propulsion ladder, or cradled it gently downward.

---

     "But where's the meteor we came down in, and what's your plan for it?" I asked. "That's a big plenty of steel."
     "Over there." He pointed to where I could see a glow in the sky at the center of the port complex. "We can use steel. We just use it with know-how."
     I stared at the glow in the sky. "Surely it couldn't have been that hot?"
     "That glow? That's not the meteor. That's the melting tap in operation. But the meteor did come in at a red heat, at least in its surface layers. You see, we land them direct into the furnace, "and as soon as everybody's clear, they change the frequency of the induction current and start melting them down. Saves quite a bit of time, and time's our most precious commodity. The energy we're not worried about—that we've got in plenty. But it takes time to reheat, and if the ejection mechanism doesn't work, it's a couple or three hours' setback to lose the heat that was built up during descent so that we can get the people out. And then we have to reheat the darned thing so we can melt it down."
     "Isn't that a rather expensive way to get steel?" I asked.
     He grinned. "You're just not used to the idea of really planetary power," he said. "Those meteors—asteroids, really—can be brought in, melted down, and ready to use for tool steel at a cost per ton of, say, a hundredth of a solar credit."

---

     The ridiculousness of shooting a planet with the ineffective sting of a fine-focused laser began to creep up on me. The focus on this device was so fine that it would probably make no more than a centimeter-diameter hole in whatever target it hit, and though that's plenty big to play utter havoc with a space vehicle, it would be less than the sting of a mosquito so far as a planet was concerned.
     "Set power pulse to three seconds."
     Gallagher's voice was slightly edged, but Cricket's came back in a singsong that showed no overtones of emotion.
     "Power pulse on three seconds by off point five seconds."
     "Initiate pulse."
     Cricket didn't have to respond to that one because the power machinery did it for her. There was a slow, rhythmic, mmm-pop, mum-pop from the power supply that went on and on and combined with the cheeee to form a now-familiar repetitive pattern; the song of power that I had heard on Betsy Ann: cheee-ops, cheee-ops.
     "I'll be damned," I yelled. "This thing's an upside-down solar tap!"
     There was a choke behind Gallagher's laugh, and his voice had a sweep and flow that spoke of tensions releasing.
     "We're way above the radiation belts," he said, "but the oscillating lens of our zoom focus makes an ionized path from Durango's ionosphere to its ground, and that's all you need for a tap. We didn't have time to build a pyramid down there, so we turned the tap upside-down."

• 

“Getting the power may not cost us that much after aIl,” said Tuna ("Richard Tuna" is actually Charles Pelligrino). “Not when you consider where computer science and robotics are headed during the next fifty to seventy years. A few days ago, at Brookhaven, we came to realize that it may cost us only the expense of developing about thirty small, self-replicating factories—which build factories, which build factories—which, when they reach a certain population density, switch over from building factories to building solar panels. We simply send them like a viral infection tolthe planet Mercury, and carpet one hemisphere with panels. You’d get more than fifty thousand times the U.S. energy budget there.”

“Mercury rotates!” Colby called out.

“What?”

“Sorry, Richie,” Colby said, “but you can’t build your panels on one hemisphere only, because Mercury rotates. Where have you been? People have known that for ten years. I’m sure it rotates. Have I introduced you to the natural resource? Ask him.”

“He’s right,” said Dr. Isaac Asimov. “Mercury rotates with aperiod of about three months, and the energy received per square foot, when Mercury is most distant from the Sun, is seven times what Earth receives. It goes up to ten times during closest approach. The interesting thing is that the one-eighty-degree and zero-degree longitudes do become gravitationally locked, alternately, during apihelion, and are exposed to the Sun almost three times longer than the longitudes ninety degrees away from them.”

“Oh, no,” Tuna groaned, flushing with embarrassment. “I don’t believe it. We have to redesign the whole thing.”

“Please,” Asimov said, raising his hands above his head, as if to show that he bore no weapons. “Don’t blame me. I didn’t make Mercury rotate. Better to find out now than when you get there and look up in the sky and—‘Hey! Why is the Sun moving?’ ”

Tuna groaned. “I don’t believe I missed that. We can redesign it, though. We’ll just have to cover a bit more surface area, and perhaps send three times as many machines to get the project started … make sure at least one cluster of them is in sunshine at any given moment…”

“And you’ll certainly want to test and perfect the machines on the Moon first,” added Asimov. “You know, it’s just one more argument for building a permanent Moon base. By the time you perfect solar panel builders for Mercury, you’ll already be beaming large quantities of clean energy down to Earth from the Moon…” He paused and smiled. “Incidentally, I wrote a science fiction story, way back in 1940, about an interrupted array of panels around the Moon’s equator (I have yet to figure out which story this is. The closest I can find is Reason). So if you ever do build the thing, Richard, you can call it the Asimov Array.”

“We can?”

“I insist on it.”

---

(ed note: The following is from the technical appendix to the novel.

Keep in mind that this was written in 1993 so adjust US dollars and national power consumption for inflation)

The technology for producing antimatter using particle accelerators is presently under development at American and European laboratories. At CERN’s seven-kilometer-circumference synchrotron near Geneva, Switzerland, antiprotons are routinely produced by firing a high-energy beam of protons into a block of tungsten. A trillion (10¹²) antiprotons can be created in this way. A trillion antiprotons may sound like a lot, but they contain the potential annihilation energy of only three hundred joules (roughly equivalent to the “bang” from a cap gun), and the CERN facility gets slapped with a $40,000 electric bill every time the accelerator is turned on (more like $68,000 US in year 2017).

Clearly, simpler and more efficient accelerators are needed. Machines with the proper requirements are presently under development in the United States (and soon to be under intensive development in Japan) for use in fusion reactor research.

“As an example,” explains physicist George Mueller, “in one design being studied, the particle accelerator will produce short bursts of protons with a power beam of 10¹⁴ watts, about one hundred times the present power output of the entire world! (in 2013 the world's average energy consuption was about 12.3×10¹² watts, 10¹⁴ is still one hundred times as much) Of course, since the machine will be operated in 10^-8-second bursts, the average power is very much lower.”

The natural location for antimatter factories, in view of their large power requirements, is in space, where continuous and, from an industrial perspective, limitless solar power is available. “Using the solar flux at the Earth’s distance from the Sun,” adds Mueller, “a light collector about three hundred kilometers on a side could provide the power for a 10¹⁴-watt factory (9×10¹⁰ m². Solar flux at Terra orbit is 1,366 W/m². At 100% efficiency that is 1.2×10¹⁴ watts).

If the efficiency of antiproton production from each high-energy proton in the initial beam could be made as high as 0.1 percent (efficiency of 0.001.) [presently Hiroshi Takahashi is predicting higher efficiencies] (Dr. Robert Forward thinks realistically it will be more like 0.0001)…

…then this machine would produce 10²⁰ antiprotons per second (each antiproton requires 1.5×10^-10 Joules, 1.2×10¹⁴ watts at efficiency of 0.001 can produce 7.98×10²⁰ antiprotons per second),…

…or about one kilogram of antimatter per month.” (antiprotons have a mass of 1.7×10^-27 kilograms. At a rate of 10²⁰ antiprotons/sec 1 kilogram would take 69 days or 2.3 months)

From our earliest brainstorming sessions emerged proposals for a solar panel array, in orbit around Earth, covering an area in excess of ten thousand square kilometers. Even if it should one day become economically feasible to mine, refine, and transport materials from the Moon to Earth orbit, such an array would literally become a gigantic solar sail, requiring a considerable expenditure in thrust (presumably from rockets that would have to be refueled from somewhere) just to keep it from blowing away on the solar wind.

Our attention tumed elsewhere, to a power source more firmly anchored, yet overlooked, perhaps because it is so large that no one noticed it before. If we are correct, the planet Mercury is destined to become the most valuable piece of real estate in the solar system.

Presently, we are eyeing self-replicating, solar-panel-building machines. If humanity plays its cards right, prototypes could be tested on the Moon near 2020, and the rewards that the descendants of these first machines can bring—cheap, clean, and unlimited power for all mankind—are yet another argument for a permanent Moon base. Using the materials available at the lunar surface, they will build solar panel farms, and new solar-panel-building machines. In time, the farms will girdle the Moon’s equator to form the Asimov Array (named, exactly as described in Chapter 4, after an early contributor to the concept). The Asimov Array will provide power for Earth.

However, we dare not use that power for producing and storing large quantities of antimatter on the Moon, or anywhere near Earth, because even a single kilogram of the stuff—a mere handful—carries the explosive potential of forty hydrogen bombs, along with the moral responsibilities that go hand in hand with the possession of such power. (1 kilogram of antimatter + 1 kilogram of matter will make about 1.8×10¹⁷ joules, about 43 megatons. B83 nuclear bomb has highest yield of any in US nuclear arsenal, it has a maximum yield of 1.2 megatons. So 43 MT is about thirty-six hydrogen bombs)

Once perfected, the descendants of the original Asimov Array robots can be sent like a viral infection to the planet Mercury. Assembling replicas of themselves from the substance of their host, their first half decade of habitation will be a latent, incubation phase, during which most of the solar panels manufactured by the machines will be used to power an ever-accelerating chain reaction of machines building more self-replicating machines. As their population approaches a predetermined critical density, more and more of them cease reproduction and join to form solar panel factories, with the result that almost three decades after the arrival of the original twenty or thirty machines, uncountable square kilometers of generator, with an area the size of Rhode lsland being added daily.

When Mercury is farthest from the Sun, each panel will receive 6.7 times as much solar energy as it would receive on the surface of the Moon. This figure rises to fully ten times the lunar surface value as the planet’s eccentric orbit dips twenty-four million kilometers nearer the Sun. In time, self-replicating machines will carpet the Mercuran landscape from pole to pole with solar panels, giving mankind more than 50,000 times the present U.S. electrical energy budget, a power capable of launching at least two interstellar missions per year.(1990 U.S. annual electrical output 2.7×10¹⁶ Joules. ×50,000 = 1.3×10²¹ Joules. Divided by 2 per year = 6.7×10²⁰ Joules = 7.5 kilogram antimatter per interstellar mission.)

Using self-replicators, the world’s future energy problems and even the excess energy required for relativistic flight can be solved for a very small initial investment: the cost of developing as few as a dozen ancestral machines on the Moon. Of course, we should not trouble ourselves to begin immediate development. They would be too expensive to build today, and too inefficient if built from the equipment now at hand—just as a trans-Atlantic airline service and videocassette recorders, though technologically feasible, would have been prohibitively expensive to build in 1925. We must wait, not only for technology to catch up with the idea, but for the idea to become economically viable.

• 

The main dome of Ganymede City covered three square kilometers of the moon's surface. Here on the top level, the observation screens covered the interior of the dome; they could be turned on individually, or en masse to show a panorama of sky and surrounding terrain. A direct view would have left the observation level with too little shielding against the leakage of solar radiation trapped by Jupiter's magnetic field; even though Ganymede City sat in the moon's radiation shadow, the aboveground portion was protected, as an added precaution, by meters of water in the outer shell, piped in from the nearby ice field. Natives called Ganymede City "the big igloo,” because of the liquid that was kept frozen in its insulating space. In addition to the physical shielding against stray radiation and occasional meteors, the Laser-Fed Fusion Reactor powered the super conducting units which cast a magnetic shield over the domes.

Ganymede plowed through a sea of death; but despite this, ships had visited all the Galilean satellites by 2015. Built at the Martian space docks on Phobos and Deimos, the water- and magnetically-shielded tin cans, as the ships came to be called, had penetrated into Jupiter's radiation belts, setting up research bases on Callisto and Ganymede, as well as temporary facilities on a few of the close-in rocks whipping around the edges of the gas giant’s atmosphere.

To make power for building the first underground living quarters, the tin cans had deployed a giant sun mirror. The collector was no more than a few molecules thick, but its huge size and focusing capacity made up for the fact that the sun's intensity was only about four percent of what it was in the vicinity of earth (51 W/m² instead of 1,366 W/m²).

While Ganymede City's first levels were being built, a mass driver track had been constructed beyond what was now the tug port. Using the three-kilometers-per-second escape velocity from the Jovian moon, the track began to toss copper ingots toward Jupiter, whose escape velocity was twenty times greater; this large energy difference was expressed in the form of eddy currents of electricity forming in the copper as it rushed through Jupiter's powerful magnetic field; these were lased back to Ganymede by a small disposable unit, continuing right up to the moment when the ingot hit the atmosphere for a final, dramatic surge of electricity. Current flow was evened out by storage facilities at the receiving station on Ganymede.

As a result of this and other systems, Ganymede became one of the energy-self-sufficient places for science, attracting research and development from earth. The solar mirror was still working; the lofter still threw ingots into Jove's face; and a second fusion LFR had recently been completed.

Ironically, Sam thought, success on Ganymede had slowed the building of facilities on Saturn's moons, as well as delaying development of the larger asteroids such as Ceres. A whole system of worlds waited to come alive out here, offering conditions for industry and research, room for a civilization to grow.

      The objective of this option is to remove the radioactive waste from the Earth, for all time, by ejecting it into outer space. The waste would be packaged so that it would be likely to remain intact under most conceivable accident scenarios. A rocket or space shuttle would be used to launch the packaged waste into space. There are several ultimate destinations for the waste which have been considered, including directing it into the Sun

     The high cost means that such a method of waste disposal could only be appropriate for separated HLW (high level waste) – i.e., long-lived highly radioactive material that is relatively small in volume – rather than spent fuel. The question was investigated in the USA by NASA in the late 1970s and early 1980s. Because of the high cost of this option and the safety aspects associated with the risk of launch failure, it was abandoned.

APPENDIX 3: ANALYSIS OF SPACE DISPOSAL OF TOTAL SOLIDIFIED NUCLEAR WASTE

Disposal of refined waste was described in section 3.9 of the technical report. It was shown, concurring with earlier NASA studies, that refined waste disposal is practical using the space shuttle and a modified full-capability tug for transportation.

Nuclear waste is presently processed to a solidified form consisting of about 25 percent fission product oxides, less than 1 percent actinides, the remainder being inert (nonradio-active) material. The waste is typically canned in "pots" 0.3m in diameter by 2.4m in length (1 x 8 ft). It would be desirable, if economically practical, to dispose of total waste in this form, eliminating completely the need for long-term Earth storage. Accordingly, a brief study of total waste disposal was performed.

3.1 TOTAL WASTE DISPOSAL PAY LOAD CONCEPT

This concept assumes disposal of total so!idified waste, based on current waste solidification technology. The total waste is roughly 1110th as radioactive per unit rass as the partially refined waste discussed above. The total waste package is illustrated in figure 3-1. It appears practical to provide a portable shield for safe handling and for flight crew protection. It is unlikely, however, that such a massive shield could be designed lo survive abort entry and impact. Thc launch system and operational procedures must provide protection from public exposure. The shield is assumed returned to Earth for reuse.

Requirements are stated in table 3-1. Data shown are typical. Waste can be repackaged to some degree in order to tailor the mass per package to capabilities of the transportation system.

3.2 TRANSPORTATION ANALYSES

3.2.1 Transportation Mode Candidates

The total waste requirement is very demanding, both in terms of total mass and in terms of economics, i.e., transportation cost. Consequently, only very low cost Earth launch options were considered. Orbit transfer options included 1-½ stage and common stage (slingshot mode) LO₂/LH₂ OTV's and an electric propulsion option powered by decay heat of the waste itself. The low cost Earth launch options included a low cost heavy lift vehicle (LCHLV) and a second generation single stage-to-orbit (SSTO) shuttle. Where the LCHLV is used as the only Earth launch option, gliders similar to the shuttle orbiter, but without main propulsion systems, delivered to orbit by the LCHLV, are used as waste carriers to provide the needed intact-abort capability. The LCHLV is described in Appendix 2. SSTO concepts have been published in the literature, notably by Salkeld, and have been studied by Boeing on IR&D. The Boeing concept is illustrated in figure 3-2. No effort was spent on SSTO concepts by this study.

3.2.2 Transportation Sequences

Figures 3-3 and 3-4 show the transportation sequences investigation for the SSE destination. The first mode employs a LCHLV and a common-stage LO₂/LH₂ OTV. Intact abort capability during Earth launch is provided by the gliders shown. One shielded waste package is camed in each glider. In orbit, the waste packages are extracted from their shields and installed on the OTV system. The shields are returned to Earth by the gliders. The OTV's operate in slingshot mode with the boost stage recovered and the second stage expended along with the payloads to solar system escape.

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• 

The second mode employs a SSTO to launch the waste packages and small OTV/drop tank systems to orbit. The waste package goes up last; the shield is recovered by the SSTO. The OTV operates in a perigee kick mode; the drop tanks contain enough LO₂/LH₂ to establish a one day elliptic orbit. At first perigee the injection stage fires to SSE with the payload. All OTV elements are expended.

Table 3-2 (missing) provides a summary mission history for the 1-½ stage OTV system.

Thp LCHLV was assumed to have a low orbit payload capability of 200 000 kg as for the power satellite program. The SSTO was assumed to have 30 000 kg low orbit capability, with return payload capability of 24 000 kg. The gliders used with the LCHLV were also assumed to have 24 000 kg return payload capability.

3.2.3 Earth Launch Summary

A summary of Earth launch and OTV requirements for the various options and modes is shown in table 3-3. The ROM busbar surcharge values shown are in cents/kWh, 1975 dollars, and are transportation cost only. They do not include waste processing or packaging costs. Numbers of flights per waste package are indicated with flights per year in parentheses based on 50 and 1,100 waste packages per year, respectively.

3.3 Special Study: Nuclear Waste Disposal in Space Utilization of Waste Decay Heat

It was suggested that the decay heat of nuclear fission waste products might be used to drive a propulsion system to accomplish disposal of the waste to SSE. A typical conceptual system includes a closed-cycle heat engine operating from the decay heat, generating electricity to drive an electric prppulsion system (figure 3-5). Refined and total waste options are examined by the FSTSA study. Only the total waste option appears to be a candidate for this transportation mode because (a) the refined waste as defined by Lewis Research Center has very little thermal power, and (b) it can be handled economically by Shuttle/FCT.

This is an energy-limited problem. The energy available in the waste is finite and must be sufficient to provide the necessary energy change to accomplish the mission. An estimate of the energy available in solidified total waste is presented in figure 3-6. This decay is nearly a straight line on the log/log-plot and therefore may be approximated by q = at^b where q is thermal panel at time t after core shutdown and a and b are curve-fit constants. Decay heat data were obtained from a MIT study and adjusted for representative mass properties of solidified waste. The above expression can be readily integrated to determine !otal thermal energy available over any period t₁ to t₂. Results are shown in figure 3-7.

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The energy required for solar system escape from low Earth orbit at low thrust is roughly equivalent to a delta V of 25 km/sec. This large delta V arises because the low thrust system must first escape Earth at nearly the full 7.73 km/sec required at infinitely low thrust plus a large proportion of the additional 30 km/sec required to escape the solar system at infinitely low thrust. (An impulsive maneuver from low Earth orbit with no gravity losses, can reach solax system escape with a delta V of about 8.8 km/sec).

The energy required to achieve a ΔV of 25 km/sec is a function of jet velocity (Isp) and of the efficiency of converting thermal energy to jet energy. The required energy versus Isp has a minimum.

This function is plotted in figure 3-8 for cycle and thruster system efficiencies of 40% and 70%.

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Comparing this result with figure 3-7 and recognizing the uncertainties in such a brief analysis, the following observations are made:

• There is a question as to whether enough energy for self-propulsion is available in nuclear waste as presently processed. Careful examination of this question and its ramifications should precede any system definition activities.
• A system designed to utilize waste energy for disposal will be sensitive to the "quality," i.e., thermal power, of the waste. It could not dispose of "old" waste and low grade wastes (contaminated shoes, clothing, tools, etc.) except as a payload on high quality wastes.
• The system will have to combine long life with low cost. Propulsive periods on the order of 5-10 years are required.
• A large riumber of vehicles will be under powered flights in various stages of the escape mission at any one time. All would presumably require some degree of monitoring. We have not made an estimate of the number of vehicles (the number clearly depends on the size of each) but a number in the range between 1 00 and 1,000 is likely.

When I fly from Texas to Europe, I pay $3–6 a pound, depending on how well I do buying a ticket. When a satellite or shuttle is launched into space, the customer (or taxpayer) pays over $10,000 a pound. That is the major challenge of space flight: until the cost of going into space drastically decreases, the large-scale exploration and exploitation of space will not occur.

The world currently sends approximately 200 tons of payloads, the equivalent of two 747 freighter flights, into space annually. At $50–500 million a launch, very few cargoes can justify their cost. We have here the classic chicken-and-egg situation. As long as space flight remains very expensive, payloads will be small. As long as payloads remain small, rockets will be expensive.

If annual demand were 5,000 tons instead of 200, the equation would shift. Engineers would have the incentive to design more efficient launch systems. Large, guaranteed payloads could significantly reduce the cost of reaching orbit, ushering in a new, affordable era in space for governments, businesses, universities, and, hopefully, individuals.

Where would this much new cargo come from? Fortunately, there is an answer. Unfortunately, it’s not intuitively attractive, at least at first glance: it’s high-level nuclear waste, the 45,000 tons and 380,000 cubic meters of high-level radioactive spent fuel and process waste and detritus (as opposed to the more abundant but far less dangerous and shorter-lived low-level waste) from six decades of nuclear weapons programs and civilian power plants.

There are three good reasons to send nuclear waste into space. First, it is safe. Second, space disposal is better than the alternative, underground burial. Third, it may finally open the door to widespread utilization of space.

Because of the obvious and real concern about moving such dangerous material anywhere, let alone into space, this proposal justly raises the question of safety. Can nuclear waste be safely launched into earth orbit? The answer is yes. By keeping the launch system on the ground instead of putting it on the vehicle, designing and building unbreakable containers, and arranging multiple layers of safety precautions, we can operate in a judicious and safe manner.

The nuclear waste problem

The problem of nuclear waste disposal is real, especially for future generations. Leaving radioactive wastes on earth creates permanent and tempting targets for terrorism as well as threatening the environment. We have a moral imperative to solve this problem now so we do not burden our children and their children.

For twenty years, the federal government’s preferred solution to the nuclear waste problem is underground disposal, specifically, over 11,000 30–80 ton canisters buried in 160 kilometers of tunnels hundreds of meters underneath Yucca Mountain in northern Nevada. Forty-nine states favor this plan. It’s not hard to guess which state does not.

To be fair to Nevada, any site would draw the same objections from anybody who lost this lottery, yet policymakers remain stuck on the idea of burial. Nevada’s fears are justified: researchers cannot guarantee complete environmental isolation for the thousands of years needed for these wastes to decay harmlessly. A recent report by the Government Accountability Office raised nearly 200 technical and managerial concerns about the site. Even the promise of construction and maintenance jobs has failed to sway a skeptical public.

Historically, garbage has been something to bury or recycle. Consequently, nuclear waste disposal has remained the province of the geologists, who are professionally inclined to look down, not up. That’s shortsighted. The permanent elimination of high-level radioactive waste demands a reconceptualization of the problem. We need to look up, not down. Let’s put high-level radioactive waste where it belongs, far out in space where it will not endanger anyone on earth.

The laser launch solution

Neither the space shuttle nor conventional rockets are up to this task. Not only are they expensive, but they lack the desired reliability and safety as insurance rates demonstrate. Instead, we need to develop a new generation of launch systems where the launcher remains on the ground so the spacecraft is almost all payload, not propellant. As well as being more efficient, ground-launched systems are inherently safer than rockets because the capsules will not carry liquid fuels, eliminating the in-flight danger of an explosion. Nor will the capsules have the pumps and other mechanical equipment of rockets, further reducing the chances of something going wrong.

How would disposal of nuclear wastes in space actually work? In the simplest approach, a ground-based laser system will launch capsules directly out of the solar system. In a more complicated scheme, the laser system will place the capsules into a nuclear-safe orbit, at least 1,100 kilometers above the earth, so that they could not reenter for several hundred years at a minimum. Next, a space tug will attach the capsules to a solar sail for movement to their final destination orbiting around the sun, far, far from earth.

The underlying concept is simple: the launcher accelerates the capsule to escape velocity. Like a gun, only the bullet heads toward the target, not the entire gun. Unlike a shuttle or rocket, ground systems are designed for quick reuse. To continue the analogy, the gun is reloaded and fired again. These systems would send tens or hundreds of kilograms instead of tons into orbit per launch.

Of the three possible technologies—laser, microwave, and electromagnetic railguns—laser propulsion is the most promising for the next decade. In laser propulsion, a laser beam from the ground hits the bottom of the capsule. The resultant heat compresses and explodes the air or solid fuel there, providing lift and guidance. Although sounding like science fiction, the concept is more than just an elegant idea. In October 2000, a 10-kilowatt laser at White Sands Missile Range in New Mexico boosted a two-ounce (50 gram) lightcraft over 60 meters vertically. These numbers seem small, but prove the underlying feasibility of the concept.

American research, currently at Rensselaer Polytechnic Institute in New York with previous work at the Department of Energy’s Lawrence Livermore National Laboratory in California, has been funded at low levels by the United States Air Force, NASA, and FINDS, a space development group. The United States does not have a monopoly in the field. The four International Symposiums on Beamed Energy Propulsion have attracted researchers from Germany, France, Japan, Russia, South Korea, and other countries.

The long-term benefit of a ground-based system will be much greater if it can ultimately handle people as well as plutonium. Dartmouth physics professor Arthur R. Kantrowitz, who first proposed laser propulsion in 1972, considers the concept even more promising today due to more efficient lasers and adaptive optics, the technology used by astronomers to improve their viewing and the Air Force for its airborne anti-ballistic missile laser.

Where should the nuclear waste ultimately go? Sending the capsules out of the solar system is the simplest option because the laser can directly launch the capsule on its way. Both Ivan Bekey, the former director of NASA’s of Advanced Programs in the Office of Spaceflight, and Dr. Jordin T. Kare, the former technical director of the Strategic Defense Initiative Organization’s Laser Propulsion Program, which ran from 1987-90, emphasized solar escape is the most reliable choice because less could go wrong.

A second option, a solar orbit inside Venus, would retain the option of retrieving the capsules. Future generations might actually find our radioactive wastes valuable, just as old mine tailings are a useful source of precious metals today. After all, the spent fuel still contains over three-quarters of the original fuel and could be reprocessed. Terrorists or rogue states might be able to reach these capsules, but if they have that technical capability, stealing nuclear wastes will be among the least of our concerns. This approach is more complex, demanding a temporary earth orbit and a solar sail to move it into a solar orbit, thus increasing the possibility of something going wrong.

Addressing safety

The issue of safety has two components. One is the actual engineering of safe operations. This is demonstrable and testable. The other, equally important, part is the public perception of safety. As University of Missouri nuclear engineering professor William H. Miller, a specialist on nuclear fuel cycle and fuel management, noted, “The obvious problem is public perception. No matter how far you go to show that it is safe, there will always be someone to say ‘what if’.” John W. Poston, a Texas A&M nuclear engineering professor with a forty-six year career in nuclear health physics, agrees, considering convincing people of the safety of space-based disposal as challenging, if not more so, than the actual technical questions.

Safety should appropriately dominate public discussion of this proposal. To succeed, space disposal must demonstrate lower risk and uncertainty than underground disposal. This project must be completely safe technically, but nonetheless will not succeed unless potential supporters and opponents are thoroughly convinced about its safety and efficiency.

Assuring safety is possible. The two major concerns are launching the capsule and ensuring the integrity of the capsule. Laser launching is safer and more reliable than rockets. The absence of rocket propellants and its accompanying propulsion systems eliminates the possibility of an explosion. The major problem would be if the laser failed before the capsule reached escape velocity. Because the capsule will be bullet-shaped, its ballistic characteristics are well known. Thus, if a launch failure occurred, the capsule would land only in known recovery zones. Launch trajectories would be designed to avoid populated areas.

One advantage of a laser launch system is that the safe return from these aborted missions can be demonstrated by testing with inert capsules. Scores of launches could test every conceivable scenario, the equivalent of firing a new rifle to understand all its characteristics. This could not be done with a rocket. If another layer of safety is desired, placing the launch system on an island in the Pacific Ocean will further decrease the chance of an aborted flight landing in a populated area. Such isolation would also improve security.

The capsule itself must protect its radioactive cargo not only from the demands of a normal launch with its severe atmospheric heating and aerodynamic loading, but also from potential accidents ranging from reentry into the atmosphere to a seriously flawed launch that would send the capsule into the high pressures of the ocean’s depths or into land. Summing up the engineering challenges, Bob Carpenter, the program manager for Orbital Sciences’ space nuclear power program, cautioned, “I’m not saying they are insurmountable, but they are major technical issues to be solved.”

Jordin Kare, now an independent aerospace consultant, was more optimistic. The laser can accelerate the capsule slowly in the lower atmosphere, reducing heating. Furthermore, noted NASA nuclear engineer Dr. Robert C. Singleterry, the same aerobraking analyses and technologies that use a planet’s atmosphere to slow down a visiting spacecraft as the Mars Global Surveyor demonstrated in 1997 can ensure the control of a capsule leaving the earth’s atmosphere.

The integrity of a capsule can be demonstrated too. The aerospace industry has accumulated decades of research and experience on how to contain radioactive material in containers that can maintain their integrity despite atmospheric re-entry, accidents, explosions, and other potential catastrophes. They are called nuclear warheads. Designing containers for space disposal is well within the state of the art. Dr. Rowland E. Burns, the engineer who led a NASA study in the mid-1970s on this issue, stated it is feasible to design and construct containers that can safely withstand the demands of even a catastrophic explosion, claiming, “I won’t say you would have to nuke the container to break it, but it would take something like that.”

Materials technology has improved since the 1970s, making even tougher capsules possible. Because launch costs will be relatively inexpensive, engineers can overdesign for safety instead of trying to create the lightest possible container. Fail-proof capsules can be built, though the ratio of waste to shielding will be low.

Ensuring safety must have an inclusionary component. A broadly based panel of stakeholders, including skeptics and opponents, should determine the criteria for tests and scenarios that proponents must pass. Computer simulations and controlled tests, however, will not be enough. Convincing demonstrations such as aborting launches with a mock payload and sending test capsules to reenter the atmosphere will be necessary to calm fears and prove the veracity of safety calculations. Minimum danger must be demonstrated, not assumed. Those opponents who unilaterally reject space-based disposal should be asked to propose an alternative. Nuclear waste will not go away on its own volition.

Expensive and inexpensive

What about the economics? Let’s be honest and upfront in our accounting: Space disposal will ultimately cost tens of billions of dollars, but the federal government has already spent $8 billion researching underground disposal and expects the total cost will be $60 billion. The difference is that future generations will not have to worry about the waste and they will have an infrastructure for reaching space. While technologically impressive, developments in tunnel boring have far less potential. Disposal in any form will be expensive. Space disposal at least offers a major spinoff, inexpensive access to space. Putting a small surcharge—a fraction of a cent per kilowatt-hour of electricity—on power generated by nuclear reactors would handle the operational costs.

How can a system be both expensive and inexpensive? Judging by the costs of other high technology projects such as the Airbus 380 and Boston’s Big Dig, developing a laser launch system will require at least $5–10 billion. This is a lot of money, but historically space technologies are expensive: The Apollo program cost over $150 billion in contemporary dollars. Constructing the actual launch system will require a few billion dollars and operations will consume billions more. And even if the price of a pound to escape velocity is only $100, 5000 tons is $1 billion.

We owe the future as well as ourselves the opportunity to determine whether space-based disposal is the best way to handle nuclear waste. Accordingly, over the next few years, NASA and the Department of Energy should establish three research programs. The first will determine the criteria and acceptance for a demonstration program. The second program will design safe capsules and the third program will test the ground-launched system. For the price of a new hotel in Las Vegas or a day or two of the defense budget, we will have enough information to decide whether to commit large resources to space-based disposal.

Space disposal may not appear the obvious solution to the high-level nuclear waste problem. Nor is disposing of nuclear waste the obvious answer to the question of how to reduce the cost of reaching space. But the immense magnitude of nuclear wastes provides the incentive to develop launch systems that will drastically cut the cost of space exploitation. The result will be lower operating costs, more infrastructure, and more skilled personnel able to develop other areas of space.

The development of the computer may offer a good analogy. Government funding, mostly from the military, intelligence community, and NASA, greatly accelerated research, development, and diffusion of computers since the 1940s. The federal government did this to conduct projects of national significance such as the census, Social Security, weapons research (especially nuclear explosions), cryptoanalysis, and space exploration. Not until the 1970s did the civilian market grow large enough to seize the technological initiative.

Space disposal may prove a similar opportunity. Once a ground launcher is developed and built, constructing additional launchers will be far less costly and risky. The dream of affordable access to space may then come true, opening up the final frontier in ways that we have not dreamed of since the 1960s. As important, we will be acting ethically, providing our children a safer earth and inexpensive access to space for people as well as plutonium.

• 

      Tim started the scanner going as soon as I handed over the controls. He thought I'd picked up our discarded fuel container again — which annoyed me since it showed little faith in my common sense. But he soon saw that it was in a completely different part of the sky and his scepticism vanished.
     'It must be a spaceship,' he said, 'though it doesn't seem a large enough echo for that. We'll soon find out — if it's a ship, it'll be carrying a radio beacon.'
     He tuned our receiver to the beacon frequency, but without result. There were a few ships at great distances in other parts of the sky, but nothing as close as this.
     Norman had now joined us and was looking over Tim's shoulder.
     'If it's a meteor,' he said, let's hope it's a nice lump of platinum or something equally valuable. Then we can retire for life.'
     'Hey!' I exclaimed, 'I found it!'
     'I don't think that counts. You're not on the crew and shouldn't be here anyway.'
     'Don't worry,' said Tim, 'no one's ever found anything except iron in meteors — in any quantity, that is. The most you can expect to run across out here is a chunk of nickel-steel, probably so tough that you won't even be able to saw off a piece as a souvenir.'

     By now we had worked out the course of the object, and discovered that it would pass within twenty miles of us. If we wished to make contact, we'd have to change our velocity by about two hundred miles an hour — not much, but it would waste some of our hard-won fuel and the Commander certainly wouldn't allow it, if it was merely a question of satisfying our curiosity.
     'How big would it have to be,' I asked, 'to produce an echo this bright?'
     'You can't tell,' said Tim. 'It depends what it's made of — and the way it's pointing. A spaceship could produce a signal as small as that, if we were only seeing it end-on.'

     'I think I've found it,' said Norman suddenly. 'And it isn't a meteor. You have a look.'
     He had been searching with the ship's telescope, and I took his place at the eyepiece, getting there just ahead of Tim. Against a background of faint stars a roughly cylindrical object, brilliantly lit by the sun­light, was very slowly revolving in space. Even at first glance I could see it was artificial. When I had watched it turn through a complete revolution, I could see that it was streamlined and had a pointed nose. It looked much more like an old-time artillery shell than a modern rocket. The fact that it was streamlined meant that it couldn't be an empty fuel container from the launcher in Hipparchus: the tanks it shot up were plain, stubby cylinders, since streamlining was no use on the airless Moon.
     Commander Doyle stared through the telescope for a long time when we called him over. Finally, to my joy, he remarked: 'Whatever it is, we'd better have a look at it and make a report. We can spare the fuel and it will only take a few minutes.'

     Our ship spun round in space as we began to make the course-correction. The rockets fired for a few seconds, our new path was rechecked, and the rockets operated again. After several shorter bursts, we had come to within a mile of the mysterious object and began to edge towards it under the gentle impulse of the steering jets alone. Through all these manoeuvres it was impossible to use the telescope, so when I next saw my discovery it was only a hundred yards beyond our port, very gently approaching us.
     It was artificial all right, and a rocket of some kind. What it was doing out here near the Moon we could only guess, and several theories were put forward. Since it was only about ten feet long, it might be one of the automatic reconnaissance missiles sent out in the early days of space-flight. Commander Doyle didn't think this likely: as far as he knew, they'd all been accounted for. Besides, it seemed to have none of the radio and tv equipment such missiles would carry.

     It was painted a very bright red — an odd colour, I thought, for anything in space. There was some lettering on the side — apparently in English, though I couldn't make out the words at this distance. As the projectile slowly revolved, a black pattern on a white background came into view, but went out of sight before I could interpret it. I waited until it came into view again: by this time the little rocket had drifted considerably closer, and was now only fifty feet away.
     'I don't like the look of the thing,' Tim Benton said, half to himself. 'That colour, for instance — red's the sign of danger.'
     'Don't be an old woman,' scoffed Norman. 'If it was a bomb or something like that, it certainly wouldn't advertise the fact.'
     Then the pattern I'd glimpsed before swam back into view. Even on the first sight, there had been something uncomfortably familiar about it. Now there was no longer any doubt.
     Clearly painted on the side of the slowly approaching missile was the symbol of Death — the skull and crossbones.
     Commander Doyle must have seen that ominous warning as quickly as we did, for an instant later our rockets thundered briefly. The crimson missile veered slowly aside and started to recede once more into space. At the moment of closest approach, I was able to read the words painted below the skull and crossbones — and I understood. The notice read:

WARNING !

RADIOACTIVE WASTE !

ATOMIC ENERGY COMMISSION

     'I wish we'd got a Geiger counter on board,' said the Commander thoughtfully. 'Still, by this time it can't be very dangerous and I don't expect we've had much of a dose. But we'll all have to have a blood-count when we get back to base.'

     'How long do you think it's been up here, Sir?' asked Norman.
     'Let's think — I believe they started getting rid of dangerous waste this way back in the 1970s. They didn't do it for long — the Space Corporations soon put a stop to it! Nowadays, of course, we know how to deal with all the by-products of the atomic piles, but back in the early days there were a lot of radioisotopes they couldn't handle. Rather a drastic way of getting rid of them — and a short-sighted solution, too!'
     'I've heard about these waste-containers' said Tim, 'but I thought they'd all been collected and the stuff in them buried somewhere on the Moon.'
     'Not this one, apparently. But it soon will be when we report it. Good work, Malcolm! You've done your bit to make space safer!'
     I was pleased at the compliment, though still a little worried lest we'd received a dangerous dose of radiation from the decaying isotopes in the celestial coffin. Luckily my fears turned out to be groundless — we had left the neighbourhood too quickly to come to any harm.

     We also discovered, a good while later, the history of this stray missile. The Atomic Energy Commission is still a bit ashamed of this episode in its history, and it was some time before it gave the whole story. Finally it admitted the dispatch of a waste-container in 1981 that had been intended to crash on the Moon but had never done so. The astronomers had a lot of fun working out how the thing had got into the orbit where we found it — it was a complicated story involving the gravities of the Earth, Sun and Moon.

The Interplanetary Internet (based on IPN, also called InterPlaNet) is a conceived computer network in space, consisting of a set of network nodes which can communicate with each other. Communication would be greatly delayed by the great interplanetary distances, so the IPN needs a new set of protocols and technology that are tolerant to large delays and errors. While the Internet as it is known today tends to be a busy network of networks with high traffic, negligible delay and errors, and a wired backbone, the Interplanetary Internet is a store-and-forward network of internets that is often disconnected, has a wireless backbone fraught with error-prone links and delays ranging from tens of minutes to even hours, even when there is a connection.

• 

NASA is about to make it a little easier to check your Instagram in zero gravity. Two teams, Science Mission Directorate and Human Exploration and Operations, are working together to finally make interplanetary internet a thing. Previous efforts to bring WiFi throughout the solar system haven’t always been successful, but this time, it could become reality.

It will work using something called Delay/Disruption Tolerant Networking, which is pretty similar to the internet you’re familiar with. But conventional internet doesn’t do well in space, plagued with long delays, noisy channels, and high error rates.

With DTN, even if your connection gets disrupted, it will guarantee data packet delivery once the next communication path opens. Normally, if you lose connection, the data gets dumped. But by removing the need to retransmit during a lag, it saves time and frees up the limited memory used by spacecraft.

Cosmic WiFI

Getting WiFi in space is complex, especially given typical extreme distances and fragile connection links. Even if your internet is traveling at lightspeed, it can take considerable time to send a message from Earth to Mars, for example. NASA previously proposed bringing the internet to the Red Planet in 2009, but due to budget constraints, the Mars Telecommunications Orbiter was scrapped. It would have used high-speed radio signals and laser light beams to send the equivalent of three compact disks of data each day.

DTN will now be deployed with the launch of PACE, or the Plankton, Aerosol, Cloud, ocean Ecosystem Mission, an Earth-monitoring satellite operation that will advance our understanding of climate change. The satellite, slated for launch in 2022, will surveil everything from massive storms to algal blooms to carbon cycles, teaching us more about the health of the planet’s oceans.

But testing DTN goes back to 1998, when the U.S. Defense Advanced Research Projects Agency (DARPA) began its Next Generation Internet initiative, which financed a small team at NASA’s Jet Propulsion Laboratory in Pasadena, California. Their goal was to build a “specialized deep space backbone network of long-haul wireless links” under the guidance of internet pioneer Vinton Cerf. It built upon the Space Communications Protocol Specifications developed by the late Adrian Hooke.

Since then, DTN has been used on Deep Impact, a space probe that in 2005 launched an impactor at a comet called Hartley 2, which caused an explosion equivalent to 4.8 tons of TNT. More recently, NASA tested DTN to remotely drive a Lego car in Germany from the International Space Station and to send a photo at the National Science Foundation’s McMurdo Station in Antarctica, where internet connection is spotty.

“DTN represents a shift in how data will get delivered in the future,” NASA engineer David Israel said in a statement.

DTN is just a part of NASA’s Decade of Light initiative, a growing endeavor to build an internet in the Solar System, including NASA’s Near Earth Network, Space Network and Deep Space Network. As Space Race 2.0 heats up, we’re going to need faster, more reliable ways to connect online — and technologies like DTN will be the path forward to exploring the stars.


/* Increase the timeout each time we retransmit. Note that
* we do not increase the rtt estimate. rto is initialized
* from rtt, but increases here. Jacobson (SIGCOMM 88) suggests
* that doubling rto each time is the least we can get away with.
* In KA9Q, Karn uses this for the first few times, and then
* goes to quadratic. netBSD doubles, but only goes up to *64,
* and clamps at 1 to 64 sec afterwards. Note that 120 sec is
* defined in the protocol as the maximum possible RTT. I guess
* we'll have to use something other than TCP to talk to the
* University of Mars.
*
* PAWS allows us longer timeouts and large windows, so once
* implemented ftp to mars will work nicely. We will have to fix
* the 120 second clamps though!
*/



Thucydides:

Actually, given the great time lags in communications (and probable bandwidth issues), I would not expect interplanetary Tweets, phone calls or even emails.

Point to point communications would most likely resemble texting (including the incomprehensible abbreviations), but you would have anywhere from 1.4 seconds delay to the moon to many hours to Uranus.

This means another possible driver for high performance spacecraft would be mail delivery. (Consider a small car delivering 100 DVDs for Netflixx probably is carrying more information than you can access through your home internet connection in a day.) Contracts, magazines, movies, personal messages and anything else which needs more detail than a text message will all be loaded aboard mail servers on fast packets, and blasted to their destinations via the fastest means possible.

This of course leads to interesting scenarios where protecting and intercepting mail becomes important for intelligence agencies, business and criminals. Mail delivery will involve high levels of security and screening of the mail delivery personnel. I doubt anyone will hijack a mail packet in flight, but having a covert operative on board to hack the mail server and download the interesting information is a very real possibility.

---

Jim Baerg:

This sounds very implausible to me.

The light speed delays mean that all interplanetary communication will resemble emails with attachments more than phone calls.

The immense power required for torchship performance would make interplanetary communication lasers much more economical for sending even terabytes of data, much less an email greeting with a 1 MB photo.

---

Ian_M:

Email in space — Should be no problem with this, even if the distances slow it down to telegraph speeds.

Text messaging in space — Will be restrained to within a few light-seconds, probably no more than 10 or so. Any longer than that and you might as well send a longer message.

Instant messaging — Will probably be restrained to 2-3 light-seconds. Like a phone call, any more delay defeats the whole purpose.

Physical packets — Documents that you don't want easily copied, high-density data storage for non-essential information (Movies, music, etc), items that can't be replicated at the destination (Due to lack of facilities), items that shouldn't be replicated at the destination (Gifts), and people (Diplomats, managers, mediators, etc).

---

Thucydides:

To expand on Citizen Joe's point, would you like to receive "Male enhancement" emails with the olympusmon.mars address? How about "Dear Sir, I am the last surviving member of the Europan Resistance front and need your help to transfer $10 billion solars from the Bank of Callisto...."

On a more serious note, the high bandwidth links would probably be reserved for ship traffic, government and military communications and corporate communications (for companies with the financial clout to get in line for email). Certainly the Uranus Space Navy would not want the high bandwidth links clogged during their showdown with the Imperial Jovian Navy, nor would they want to risk malware or botnet attacks coming through those links; which suggests interplanetary comms would be tightly controlled and subscribers carefully vetted.

---

Ubreputz:

Since power and surface area aren't problems on the ground (and to a lesser extent, in orbit), communications would likely be predominantly TO ships while ship to shore (home base) comms would be limited to a confirmation of receipt. On Earth, we can afford to put up huge arrays to catch the smallest radio signal. Not so much in space. Likewise we can pump a lot of energy into the antenna to send back a longer message with a lot of bandwidth and strength even at stupendous ranges.

Now there are some tricks, like omnidirectional beacons and antenna that act as targets for the directional antennae. But you can only listen to data in the direction of the directional antenna. That might be limited to a single stream. Comm relays would likely have at least 4: Signal in, Signal out, Previous relay, Next Relay. By using multiple relays (at least 4 would get you around the sun) you wouldn't have black outs (except at ship orbit). And then there is the problem with the fragile gimbals needed for all the antennae.

In the end, yes, you can communicate via radio. No, it isn't broadband. No, it won't service a population comparable to the internet. It will likely be biased communication. It will still be expensive. There will be extreme needs that keep every carrier busy. So, although you COULD send a digital copy of Pluto Nash to Pluto, you would never get enough priority to use the carriers and thus it would be simpler to put it on a torch or fling it out the airlock.

---

Jean Remy:

What about relays at Lagrangian points?

Everyone knows where those are, and if someone needs to connect they can just link into the network. I don't even think lasers would even be needed. The relays would have high gain antennas to receive the data on broadband signals, and when a ship or colony needs to link in they can query the closest available platform with a much lower-gain antenna.

If we can be in contact with various probes (like Voyager probes) at interplanetary distances, on 1970s technology, then I hardly think you would need a giant technological leap to create a system-wide comm network.

Granted you won't have the bandwidth of fiberoptic cables, so no browsing for Earth-porn from Callisto, but I don't see communication as much of an impediment. The only real reason to use lasers (that I see) would be for private (read: military) communications that you don't want intercepted, but with a target area of 90 km in radius, it's not really very private anymore.

---

Eric:

Surfing the interplanetary web would be a unique experience. Given the reply times measured in minutes or hours, getting a response for a search query would be more akin to making a request at the Library of Congress and then waiting for the attending librarian to go and bring you a copy of whatever they think you asked for (which may or may not be what you were actually looking for).

Since we survived the dark ages of dial-up modems, I think we could probably put up with the connection speed of the interplanetary net.

---

Ferrell:

I think that interplanetary communications networks would be more like a cluster of 'webs'; one on Earth, one on Mars, one on Luna, ect. The different webs would send updates to each other on a regular schedule via dense-data/high-priority channels, and all other inquries/messages being sent via lower-priority channels. Most of your routine web activities would be with your local internet, but occationally you'd connect with another world's internet via the systen-wide web.

---

Jean Remy:

Interestingly enough, this is the way the "extranet" is postulated to work in the Mass Effect games. Essentially all colonies pack with them a web server full of the most needed and demanded data. If the information you are looking for is not on the server the system opens a communication link (FTL in the game) and hits the closest large colony server, and so on, until it finds the info, and it uploads it into the colony's server for easy retrieval later. In the game data storage is no longer an issue, but I assume if storage space is needed and a specific file hasn't been looked at for a while, it would be deleted.

However, since most early colonies will be science bases, I would assume large amounts of data will be passed back and forth as scientists on Callisto and on Earth look over the data, form theories and send them back and forth. It seems like a very fluid form of data exchange is needed even with local "nets". I also doubt early colonies will be very large, or widespread, and it is far more likely the Callisto colony terminals will be linked on a LAN to a single server rather than a network of servers in the first place.

---

Citizen Joe:

Viral propagation is another option. It doesn't guarantee speed or privacy. The idea would be that a message would be sent to any ship that is heading in the right direction that is within range. This could be a very long route. However, since there would likely be relatively few interplanetary vessels (compared to airplanes), Solar Space Traffic Controllers would have the full list of vessels and thus able to chart a route based on predicted paths. While the InterPlanetary Space ships would carry parcels, they would probably also serve as communication hubs for the viral network.

---

Eric:

That sounds pretty close to how the modern internet functions, with data tracing geographically indirect paths as it goes from place to place. I like the efficiency of such a scheme: it would use the infrastructure that's already there to create an ad-hoc network backbone. On the other hand, you'd be trading out data security unless you've got some very good encryption (or just plain don't care who else reads your messages).

Jean Remy, I was thinking along those same lines. A lot of business/corporate/government entities would end up hosting proxies of their sites and databases on the far-flung colony servers, with periodic updates being beamed back and forth (ultra-secure data being handled differently).

With multiple proxies being hosted for a single database, with data exchange rates measured in minutes, how long do you think the system will last until Murphy's Law takes a server down through version conflicts?

---

Jean Remy:

Viral dissemination works on Earth because of thousands and thousands of privately owned servers.

However, even in the best-case scenario of a very developed interplanetary infrastructure, I don't see a lot of traffic in space. Say two cyclers between every major colonial epicenters (say 2 for Mars, 2 for the Jovian colonies, 2 for Saturn etc...) and a few "moon hopper" shuttles, but those would be so close to their giant primaries getting a Line of Sight on them would be an issue.

However the Comm relay platforms suggested are basically Voyager probes without the scientific instrumentation and a known stable orbit. Rather that throwing your message out omnidirectionally and hope that eventually it will reach your destination (because viral dissemination is kind of like shooting blind) you simply bounce the signal of a set number of known (and if want, secure) predetermined platforms. If your goal is to reach as many people as possible (the entire point of viral dissemination in the first place) then target the Cyclers. Chances are good the passengers en route back and forth have personal computers linked in to the ship's server, which keeps updated by linking in to the platforms.

---

Ian_M:

The general outlines of communications in the plausible mid-future look a bit like the Victorian era. Fast and reliable close to home, a little slower if you want to contact the next 'city' over, slower again but still reliable if you want to contact another 'country', and best of luck to you if you want to contact someone in the backbeyond. In this case city, country, and backbeyond are defined by both linear distance and orbit.

You could even end up with a Pony Express situation. Someone sets up a Planet Express torchship delivery system to carry high information-density packages or mail that requires high-security delivery... And then nine months later someone else sets up a tightbeam relay network that can handle multiple high-bandwidth high-security messages... And suddenly you've got a lot of cheap ponies on the market.

Except in this case the ponies have 45 gW/30 milligee legs.

• 

This is a topic post referring to Purdue University's project Destiny.
Here is my introductory post for the series.
The subject is section 4, Interplanetary Communications Network.

 Headline results: I believe that the cost of this system can be reduced by nearly 50% without altering the underlying performance assumptions.

System

 This study's communications design is a major factor in overall cost. Every effort should be made to reduce this cost; the headline price is $228 billion, which works out to $4,560 per colonist per year for connectivity.

 The basic design using radio links for surface to orbit and optical links for orbit to orbit is sound. The performance numbers seem well-researched, as do the fault rates. The need for a relay is well-established. The amount of data considered is also reasonable: a single HD video stream for each 100 colonists plus two minutes of SD video (or its equivalent in other data) per person per day. My disagreement is that the use of small disposable satellites is an oldspace norm which unnecessarily drives cost.

 What's the alternative? A large multiuser platform with human maintenance missions. ITS has more than enough dV to deliver a maintenance crew and 300 Mg of cargo from Earth to ESL-5 or from Mars to areosynchrous orbit. By delivering discrete components like transponders, processors, batteries, apertures, etc., the mass of the thrust tube and propellant masses for orbital insertion and decommissioning are eliminated. Beyond that, a shared facility offers options for cost recovery. Science missions such as large telescopes may draw funding for shared services like electrical power and maintenance. Organizations like the L5 society may choose to dock habitats to the platforms for similar reasons.

Relay Node

 Consider for example the ESL-5 relay. Baseline costs for the node are $47.5 billion. The baseline plan requires 53 separate Falcon Heavy launches, 12 of which are contingency satellites. Each vehicle has dedicated avionics, GNC, optical apertures, structural bus, thrust tube, propulsion, etc., etc., all of which are discarded once any one system fails.

 For redundancy and to minimize interference, two platforms should be deployed. The physical structures are not significantly affected by exposure to deep space, so their effective service lifetime is greater than the length of the study. By allowing one ITS launch per cycle and alternating the visited platform, each platform is visited roughly every five years. Components can be tested, repaired or replaced if necessary and certified for another 5-year period. System expansion is done by adding more components than are removed.

 The baseline Falcon Heavy launch and operation costs for this location are $6.2 billion in current dollars. 47 fully-refueled ITS launches (at $46 million each, using a higher lifetime flight rate for near-Earth operation and the same $25 million operations cost) would run $2.16 billion. Total savings: $4.04 billion.

 The solar panels are a major portion of that cost, $240 million for ~360 kWe on one satellite, or $12.72 billion for the relay satellite system. The project considers SAFE-series reactors in other contexts, for example a SAFE-800 design producing 240 kWe for about $1.4 million. At steady state, six satellite equivalents will be operational at each platform; 9 SAFE-800 reactors plus one spare would therefore cost $28 million initial plus $4.3 million per refuel for the L5 relay system. Reactors are assumed to have a 60-year operational life, so two complete sets of hardware (40 total reactors) must be built and deployed during the project along with eight refueling cycles. Total cost $90.4 million, assuming launch costs are covered in the per-cycle maintenance visit cost. Total savings: $12.63 billion. Even if the development process for these reactors took several billion dollars it would be a net gain.

 The thrusters, propellant tanks, lines, valves, avionics, guidance and other systems are shared across the entire platform. Because no insertion or deorbit burns are necessary, the amount of stationkeeping propellant required is a small fraction of the baseline amount. I'll assume that mass and cost of these systems is instead spent on a structural truss for mounting components, on a stationkeeping system with ion thrusters or PITs to eliminate toxic fuels, and on a platform pointing system with very long lever arms and robust reaction wheels. No net change.

 The electronic communication components would be unchanged, and remain the primary expense at $268.8 million each ($11.02 billion total without flight spares). However, because a failure of other critical systems does not lead to loss of use of the comm systems, we are able to eliminate the 12 extra satellites and comm sets. Human servicing missions will be able to repair, repurpose or replace as needed, while one extra set on each platform will provide short-term protection against an outage. Total savings, $3.22 billion.

 This node in the communication network can be implemented as a shared facility for a savings of $19.89 billion, a total cost of $27.65 billion. The average colonist population over 100 years is half a million, so this node costs $553 per colonist per year.

 Similar logic can be applied to the Earth-orbit and Mars-orbit nodes.

Earth Node

 The Earth-orbit node would consist of three geosynchronous shared platforms. Due to the high population of GEO satellites, these could be useful places for permanent satellite servicing facilities. A rotating workforce of technicians could keep the platforms operational while also servicing other customers' hardware to offset costs. Over the life of the node, 66 satellite equivalents are required; 15 satellites are required at steady state, or 5 per platform. For simplicity these platforms should be identical to the ESL-5 platforms, though they will use one fewer reactor due to reduced power requirements. To maintain a five-year service interval, ITS flights are required every 20 months which equals 60 total flights. An alternative is to place a permanent manned maintenance facility at EML-1 and use an orbital tug to service platforms as needed; this would provide the added benefit of a one-week response time to problems.

 The baseline cost of this node is $61.1 billion with 14 flight spares. Electrical savings are ($19.2b baseline - $75.6m capital - $46.44m refueling) = $19.08 billion. Launch savings are ($9.36 billion baseline - $390 million ITS/$6.5m) = $8.97 billion. Spares savings are 14 x $189.86 million = $2.66 billion. This node's total savings are $30.71 billion, total cost is $30.39 billion.

Mars Node

 The Mars-orbit node would consist of three areosynchronous shared platforms. Due to the study's power limitations, bandwidth through each satellite is quite a bit lower than through the relay satellites and so more craft are specified. Over the life of the node, 135 satellite equivalents are required; 33 satellites are required at steady state, or 11 per platform. These would ideally be two standard platforms linked together, with 19 reactor modules. The hardware would be delivered via cargo ITS, one per cycle. This vehicle would aerocapture and then rendezvous with a central maintenance facility on Phobos (or possibly Deimos) for offloading, followed by a descent to the surface with remaining cargo and a standard return. Service missions would be dispatched from the Martian surface using one of several Mars-dedicated ITS vehicles, dock with facilities on Phobos, transit to the appropriate platform, perform maintenance and installation tasks, then return first to Phobos and then the surface. The Phobos base may eventually be permanently manned and used as a port for handling electric-propulsion cargo vehicles via docking tether.

 The baseline cost of this node is $119 billion with 21 flight spares. I believe there is a potential savings from increasing the transmission power of the Mars node to match the data rates of the other nodes, but let's stick with the baseline for now. Electrical savings are ($37.44b baseline - $159.6m capital - $98.04m refueling) = $37.18 billion. Launch savings are ($18.25 billion baseline - $1,457 million ITS/$31m) = $16.79 billion. Spares savings are 21 x $189.86 million = $3.99 billion. This node's total savings are $57.96 billion, total cost is $61.04 billion.

Results

 These changes bring the total I-Comm network costs to $119.08 billion, or $2,382 per colonist per (Earth) year / $198.47 per (Earth) month. That's a combined savings of $108.92 billion, or 47.8%. I believe there may be another $30-$60 billion of potential savings in the actual comms components due to standardization and large production runs, and at the Mars end by adding more power through fewer apertures.

 The use of shared facilities encourages the further development of space. Permanent crewed maintenance bases near the three nodes would revolutionize the way satellites are built, launched and operated, and would open the road to asteroid mining and the colonization of free space.